This work numerically investigates the importance of controlling the wake of upstream body in the tandem cylinders aiming at aerodynamics noise reduction. The two-dimensional unsteady Reynolds-averaged Navier–Stokes approach with the k– turbulence model and Ffowcs Williams–Hawkings method are employed to simulate the flow field and the aerodynamics noise, respectively. The flow in the porous media is calculated by a volume-averaged model. Analogy to the mass-spring-damper system, one preliminary model is proposed to reveal the key role of stabilizing the upstream body wake. The simulations of different porous materials coating designs are implemented to corroborate the model and provide more details of flow modification by porous materials coating. Results indicate that the porous materials coating designed on the upstream cylinder can decrease the wake impingement on the downstream cylinder via suppressing the vortex shedding. Subsequently, not only the tonal noise but also the broadband noise level of tandem system can be reduced. It is also clarified that the effects of downstream cylinder absorption and downstream cylinder wake control by porous materials coating is not comparable with the upstream cylinder wake control. The present study gives a new idea to flow control and noise reduction of the multibody systems considering efficiency and economy.
Rotor icing is a serious threat to helicopter flight safety and computational fluid dynamics technology is very useful in icing prediction. In this work, a numerical simulation method is presented to calculate three-dimensional rotor icing in hovering flight. The rotor flow fields are obtained using overlapping grids. According to Euler two-phase flow, the droplet trajectories and impingement characteristics are predicted. On the basis of three-dimensional ice accretion model, a new runback water distribution method based on shear force and centrifugal force is proposed to simulate liquid water flow and ice shape. The calculation results are compared with the experimental results under different conditions in order to verify the correctness of the method. Furthermore, the effects of blades rotation on the liquid water content distribution and droplet impingement characteristics are studied, and the effect of centrifugal force on ice shape is analyzed. The results show that the blade will influence others with the blade tip Mach number increasing, and centrifugal force will cause the ice thickness increases on leading edge and decreases at both frozen limitations.
This paper presents a novel guidance law considering the seeker dynamics for manoeuvring targets to achieve short homing time guidance using the Lyapunov method. Based on linear and nonlinear kinematics, a Lyapunov-based guidance law is synthesised to compensate for the seeker’s first-order lag. The closed-form solution of the proposed guidance system is also derived analytically. To implement the proposed guidance law, a Kalman filter algorithm is presented to extract the line-of-sight rate and its higher order derivative. Numerical simulations are carried out to demonstrate the effectiveness of the proposed guidance law under various conditions. Monte Carlo simulations are also performed to test the robustness against measurement noise.
In this article, a new floating plate structure with "drive ring" is presented, which breaks the traditional idea of installing fixed baffle in a tank. In order to analyze the performance of this anti-sloshing structure, the effect of sloshing suppression and the liquid sloshing dynamics of this structure in a Cassni tank with numerical simulation is investigated. Liquid sloshing frequencies of the floating plate with different fill ratio under the circumstances of normal weight and micro-gravity were studied. At last aiming at the liquid sloshing amplitude, the suppression effect of the floating plate and rigid ring baffle was also studied under the circumstances of instantaneous angular and "Bang-Bang" transverse acceleration incentives with different fill ratios. The result of the numerical simulation shows that the movable-type floating plate structures can increase the liquid sloshing frequency and can effectively absorb the kinetic energy of the liquid sloshing and reduce liquid sloshing amplitude.
The surface dielectric barrier discharge plasma actuator driven by nanosecond pulses is recognized as an effective fluid actuator for flow separation control. The operation condition of nanosecond dielectric barrier discharge actuators for separated flow control still requires further study, particularly prioritizing the improvement of the effectiveness and reducing energy consumption in flow separation control implementation. In this study, experiments are conducted using a two-dimensional NASA SC(2)-0712 airfoil in a wind tunnel with a Reynolds number of 0.5 x 106 (25 m/s). The pressure measurement experiments show that the location of actuators affects the efficiency of separation control. Particle image velocimetry results indicate that the most efficient location of the actuator is upstream of the separation point and near the original point of the separated shear layer. Meanwhile, the particle image velocimetry results show the vorticity attaches to the airfoil wall after discharge, which suggests that the reattachment is due to the generation of large-scale vortices. These present structures result in the mixing of the shear layer with the main flow thereby delaying separation and reattaching a separated flow. This study shows the most efficient location related to the separation point. Furthermore, it indicates the reattachment of flow is attributed to the motion of vortexes coherent structure.
Considering three-dimensional formation control for multiple unmanned aerial vehicles, this paper proposes a second-order consensus strategy by utilizing the position and velocity coordinate variables. To maintain the specified geometric configuration, a cooperative guidance algorithm and a cooperative control algorithm are proposed together to manage the position and attitude, respectively. The cooperative guidance law, which is designed as a second-order consensus algorithm, provides the desired pitch rate, heading rate and acceleration. In addition, a synchronization technology is put forward to reduce the influence of the measurement errors for the cooperative guidance law. The cooperative control law, regarding the output of the cooperative guidance law as its input, is designed by deducing the state-space expression of both the longitudinal and lateral motions. The formation stability is analyzed to give a sufficient and necessary condition. Finally, the simulations for the three-dimensional formation control demonstrate the feasibility and effectiveness of the second-order consensus strategy.
The effect of air throttling on supersonic combustion was investigated by experiments in the present paper. Our results indicated that, in the non-reacting flow, a shock train could be generated in the scramjet combustor due to the increased backpressure caused by air throttling, and the wall pressure increased obviously. But when the mass flux rate of air throttling was not large enough, the shock train would oscillate with the flow. In the reacting flow, the flame stabilization was achieved in the combustor without air throttling when the equivalence ratio of kerosene was 0.2 and 0.31, but the flame was blown off when the equivalence ratio of kerosene was 0.45. On the contrary, the kerosene (equivalence ratio: 0.45) was ignited successfully in the combustor with air throttling, and it kept burning all the time in the cases with air throttling –5% (the flux of air throttling was 5% of the inflow flux) and with air throttling –14% (the flux of air throttling was 14% of the inflow flux), but the flame was blown off in the case with air throttling –1.1% after kerosene had burnt 70 ms. The flux of air throttling should be large enough to achieve flame stabilization, and the hydrogen and air throttling should both exist all the time in order to keep the flame burning steadily.
Aerodynamic design is of great importance in the overall design of flight vehicles. In this study, an approach to aerodynamic design optimization is proposed by integrating Bezier curve parameterization and radial basis interpolation to enable large variation of aerodynamic profile during optimization. The Bezier curve uses the shape of a given airfoil and the radial basis function interpolation is applied to smoothly transfer the perturbation to the mesh in the whole flow field. Using design of experiments technique, the prominent design parameters that significantly affect the aerodynamic performance are determined. Aerodynamic optimizations are conducted for a wing airfoil and a blade airfoil to verify the efficiency of the proposed method. Genetic algorithm is employed in both single-objective and multiobjective design cases. Design results show that the present method can significantly improve the aerodynamic performance due to its capability to handle large shape changes of the airfoil. This work provides a useful and powerful tool to aerodynamic design with applications to various flight vehicles.
This paper investigates the robust non-fragile state feedback attitude control problem for uncertain spacecraft, and the problem explored here is subject to H performance constraint, quadratic stability, external disturbances, controller perturbation and control input saturation. The model of spacecraft attitude control system is introduced and transformed into a state space form with parameter uncertainty. Based on Lyapunov theory, sufficient conditions for the existence of robust non-fragile state feedback controller are given based on linear matrix inequalities in terms of additive perturbation and multiplicative perturbation. Then, the design of non-fragile controller subject to required constraints can be regarded as a convex optimization problem based on linear matrix inequalities. By solving linear matrix inequalities to obtain an optimal feasible solution which satisfies all the constraints of the problem and optimizes the objective function, the controller gain matrix can be obtained exactly. Based on the obtained controller, the spacecraft attitude control problem can be solved. The simulation results also demonstrate the effectiveness of the proposed control method.
This paper analyzes the performance of an electric solar wind sail for generating and maintaining a heliocentric circular displaced orbit. Previous research on this subject was based on a simplified mathematical model of the spacecraft thrust. However, recent studies have proposed a more accurate algorithm for evaluating both the modulus and the direction of the propulsive thrust as a function of some important parameters related to the spacecraft attitude. Therefore, a reappraisal of the problem is motivated by the need to revise past results, taking into account new information available on the propulsion system. Within this context, this paper focuses on circular displaced orbits that are characterized in terms of orbital period, heliocentric distance and elevation angle. The attitude configuration and the value of the spacecraft characteristic acceleration required for orbital maintenance are calculated. An in-depth analysis of the linear stability of displaced orbits is given. It is shown that displaced orbits are unstable when the elevation angle exceeds about 20°.
In order to eliminate the influence of outlier in measurement to X-ray pulsar based navigation system, a modified unscented Kalman filter-based filtering algorithm is proposed. First, the outlier detection function is established based on the statistical information of measurement residuals to detect whether the fault occurs in the measurements. Then, the filter gain in regular unscented Kalman filter algorithm is modified so that the dramatic changes of measurement residual are avoided. The modified unscented Kalman filter filtering algorithm is tested on the existing high earth orbit to verify the validity in the presence of the measurement with outliers. Simulation results show that the proposed modified unscented Kalman filter can effectively reduce the impact of the outliers on X-ray pulsar-based navigation system and has higher accuracy than the unscented Kalman filter method.
More-electric aircraft draws considerable attention due to its high efficiency, high reliability, and easy maintenance. Linear electro-hydrostatic actuator is a novel linear actuation system suitable for more-electric aircraft, and offers many advantages over traditional rotary electro-hydrostatic actuator. However, it requires high-frequency reciprocate actuation of linear oscillating motors with both high efficiency and high power factor. In this study, a novel moving-magnet tubular linear oscillating motor with dual-resonance is proposed for linear electro-hydrostatic actuator applications to achieve high efficiency and high power factor simultaneously. Firstly, system impedance model is set up analytically, which helps to analyze the influence of structure parameters on system performance. Based on this model, working efficiency and power factor are compared with the nonresonance design, and validated by experimental results. In order to take real scenario into considerations, the nonlinearity of motor and rectangular-type load is also analytically derived and analyzed with effective resonant frequency tracking method proposed. It shows that dual-resonance design does increase system efficiency and power factor, and can be applied in linear electro-hydrostatic actuator system for more-electric aircraft effectively.
Mixing characteristics of a Mach 2 jet controlled by shifted tabs have been studied at different levels of expansion at the nozzle exit. Two identical rectangular flat tabs of aspect ratios (length/width) 3, 4, 5 and 6, offering 2.5% blockage each, located diametrically opposite, found that the mixing promotion caused by the shifted tab increases with increase of adverse pressure gradient (that is, below NPR 5). On the contrary, the mixing enhancement caused by tab placed at the nozzle exit decreases with increase of adverse pressure gradient. At higher NPRs from 5 to 8 for shifted tab configuration, the amplitude of centerline pitot pressure oscillation is considerably smaller than the uncontrolled jet. At lower NPRs, corresponding to expansion level pe /pa, from 0.383 to 0.511, shifted tab is found to be a better mixing promoter than the tab at the nozzle exit. But for expansion levels from 0.511 to 1.022, mixing promoted by tab at nozzle exit is better than the shifted tabs. Shifted tab at 0.5D results in about 55% reduction in core length, at NPR 3, and the corresponding core length reduction by tabs at 0.25D, 0.5D, and 0D is 25.93%, 22.2%, and 14.81%, respectively.
On-orbit space deployable antennas withstand periodic thermal loads, which will degrade the antenna surface accuracy. Thus, it is necessary to preliminarily adjust the antenna before launch to make it adapt to the space temperature environment as much as possible. Aiming at this problem, a pre-adjustment method based on min-max concept is presented in this study. First, according to the force equilibrium equation of the cable element, the incremental equilibrium equation of the cable mesh antenna is derived, and the incremental expressions of the reflector's surface displacement and cable tension with respect to cable length variation and element temperature variation are also developed. Then, the optimization model for shape pre-adjustment is established using min-max method, and the pre-adjustment process is implemented by sequence planning strategy. Finally, an AstroMesh antenna is employed to demonstrate the validity of the method.
A new unsteady three-dimensional aerodynamic performance prediction approach is established to achieve fast and accurate prediction of the unsteady aerodynamics of cycloidal propellers. This model is developed by the coupling of momentum theory, lifting-line method, free wake model, and the Leishman–Beddoes semi-empirical dynamic stall model. The overall calculation process includes two parts. Firstly, to reduce the computational time and improve the computational convergence, the momentum theory is coupled with the Leishman–Beddoes semi-empirical dynamic stall model to predict a uniform inflow velocity through the cycloidal propeller disc, which is set as the initial induced velocity for iterations in the subsequent process. Then, the blade aerodynamic model, which couples the unsteady lifting-line method with the Leishman–Beddoes dynamic stall model, is used to calculate the unsteady aerodynamic response of blades. The wake of cycloidal propeller is represented by a serious of finite-length shed and trailing vortex elements, and the free wake model is utilized to model the dynamics of cycloidal propeller wake. Predictions from the present model are shown to be agreed reasonably well with the overall experimental data and the computational fluid dynamics results, both in terms of the aerodynamic performance prediction of cycloidal propeller and instantaneous blade force variations.
The increase in wing aspect ratio is gaining interest among aircraft designers in conventional and joined-wing configurations due to the higher lift-to-drag ratios and longer ranges. However, current transport aircraft have relatively small aspect ratios due their increased structural stiffness. The more flexible the wing is more prone to higher deflections under the same operating condition, which may result in a geometrical nonlinear behavior. This nonlinear effect can lead to the occurrence of aeroelastic instabilities such as flutter sooner than in an equivalent stiffer wing. In this work, the effect of important stiffness (inertia ratio and torsional stiffness) and geometric (sweep and dihedral angles) design parameters on aeroelastic performance of a rectangular high aspect ratio wing model is assessed. The torsional stiffness was observed to present a higher influence on the flutter speed than the inertia ratio. Here, the decrease of the inertia ratio and the increase of the torsional stiffness results in higher flutter and divergence speeds. With respect to the geometric parameters, it was observed that neither the sweep angle nor the dihedral angle variations caused a substantial influence on the flutter speed, which is mainly supported by the resulting smaller variations in torsion and bending stiffness due to the geometric changes.
Solving Lambert’s problem is one of the fundamental problems in astrodynamics. In this paper, it is first shown in a systematic fashion how Lambert’s problem can be expressed as a univariate equation of one of the orbital elements or various variables related to the orbital elements. Some of the choices of independent variables reported in the literature in formulating Lambert’s problem are shown to be special cases of such more general treatment, in a clear and logical fashion. This development produces a new formulation of Lambert’s problem in terms of the argument of periapsis. A new efficient algorithm using non-rational Bézier functions is designed to solve Lambert’s problem, which takes full advantage of the monotonicity and boundedness of the argument of periapsis as the independent variable and no initial guess is required for this new algorithm. Numerical comparison results are provided to demonstrate the effectiveness and efficiency of the algorithm.
In this paper, a finite-time attitude control scheme based on the dual-loop framework is proposed for rigid spacecraft in the presence of inertia uncertainties, external disturbances, bounded angular velocity, and control input saturation. The control law was designed based on a dual-loop structure. First, a hyperbolic tangent function was used to design the virtual angular velocity command, which serves as the control input of the outer loop system to stabilize the attitude under angular velocity constraints. Second, a smooth function was used to approximate the saturation and an auxiliary system was constructed to compensate for the effects of actuator saturations. A finite-time control law was designed to track the limited command of the outer loop under the disturbance and control input constraints in the inner loop. The stability of the closed-loop system during saturation is thus guaranteed, and the tracking error converges to zero after controller parameters are selected appropriately. A numerical example illustrates the effectiveness of the proposed attitude controller.
This paper designs a finite-time output tracking controller for air-breathing hypersonic vehicles (AHVs) subjected to disturbances and actuator constraints. After proper derivations, the original model is divided into two independent subsystems undergoing mismatched lumped disturbance. A finite-time disturbance observer (FTDO) is employed to estimate the lumped disturbance, while an auxiliary system combined with a command pre-filter is designed to analyze the effect of input saturation caused by the restrained actuators. Based on the FTDO and the auxiliary system, a novel integral sliding surface is constructed and then a chattering-free nonsingular controller is developed to realize finite-time output tracking in spite of mismatched lumped disturbance and input saturation, which is its major merit compared with other existing AHV controllers. A simulation study is carried out to verify the proposed control scheme.
Methods are reported for less computationally expensive and more accurate implementations of the direct simulation Monte Carlo (DSMC) method for the simulation of high speed gas flows over arbitrarily shaped bodies. A new particle-tracking algorithm with a saving of computational time of up to 10% is reported in which tracking of particles is done with the help of big triangles having vertices lying on the boundary curves. An algorithm has been developed to generate DSMC cells for collision and sampling that contain a prescribed number of molecules. This algorithm is able to generate over 90% cells having the optimum number of seven or eight molecules for simulating collisions. Sampling for macroscopic properties is done on dynamic cells that contain a number of particles varying spatially as a function of the local number density. A criterion for finding the number of particles in sampling cells is presented. This criterion has been found to result in accurate and fast simulation of two-dimensional hypersonic flows of argon over a wedge, and argon and nitrogen over a circular cylinder.
In this paper, an augmented predictive functional control approach is investigated to design a missile autopilot system, which can be expressed as a linear model with state-dependent coefficient matrices. A novel performance index depending on the reference trajectory, the output prediction and the set-point is proposed to improve the closed-loop dynamic performance. An augmented predictive functional control strategy is designed based on the proposed index and the stability is proven by using the Z-transform. In order to demonstrate the performance of the proposed approach, numerical simulations comparing the predictive functional control in the missile autopilot system are performed. Finally, results from comprehensive simulations are presented to evaluate the proposed approach in the presence of input constraints and abrupt disturbances.
In this paper, a system for real-time cooperative monocular visual motion estimation with multiple unmanned aerial vehicles is proposed. Distributing the system across a network of vehicles allows for efficient processing in terms of both computational time and estimation accuracy. The resulting global cooperative motion estimation employs state-of-the-art approaches for optimisation, individual motion estimation and registration. Three-view geometry algorithms are developed within a convex optimisation framework on-board the monocular vision systems of each vehicle. In the presented novel distributed cooperative strategy a visual loop-closure module is deployed to detect any simultaneously overlapping fields of view of two or more of the vehicles. A positive feedback from the latter module triggers the collaborative motion estimation algorithm between any vehicles involved in this loop-closure. This scenario creates a flexible stereo set-up which jointly optimises the motion estimates of all vehicles in the cooperative scheme. Prior to that, vehicle-to-vehicle relative pose estimates are recovered with a novel robust registration solution in a global optimisation framework. Furthermore, as a complementary solution, a robust non-linear Hfilter is designed to fuse measurements from the vehicles’ on-board inertial sensors with the visual estimates. The proposed cooperative navigation solution has been validated on real-world data, using two unmanned aerial vehicles equipped with monocular vision systems.
Wrinkling, a common phenomenon found in space membrane structures, is the main factor affecting the performance, stability, and dynamic characteristics of these membrane structures. This article presents an active control method to improve the surface accuracy of membrane structures. A model of a thin rectangular membrane subjects to uniaxial uniform tensile stress is discussed. Initially, the relationship between the out-of-plane deformation of the wrinkles and the boundary conditions is built with the Föppl–Von Karman plate theory by introducing the slow varying Fourier series. Because vertical tensions perpendicular to the direction of the initial wrinkles are necessary to reduce these wrinkles, reasonable locations and magnitudes of these tensions are the key problems. The finite element method and variational principle method are used to solve this issue. Finally, a manufacturing error is added to the model as an initial defect, and the robustness of the controller is verified. Simulation results show that wrinkles are reduced quickly and effectively with the proposed method.
This paper presents the development of a mathematical approach targeting the modelling and analysis of coupled flap-lag-torsion vibration characteristics of non-uniform continuous rotor blades. The proposed method is based on the deployment of Lagrange’s equation of motion to the three-dimensional kinematics of rotor blades. Modal properties derived from classical-beam and torsion theories are utilized as assumed deformation functions. The formulation, which is valid for hingeless, freely hinged and spring-hinged articulated rotor blades, is reduced to a set of closed-form integral expressions. Numerical predictions for mode shapes and natural frequencies are compared with experimental measurements, non-linear finite element analyses and multi-body dynamics analyses for two small-scale hingeless rotor blades. Excellent agreement is observed. The effect of different geometrical parameters on the elastic and inertial coupling is assessed. Additionally, the effect of the inclusion of gyroscopic damping is evaluated. The proposed method, which is able to estimate the first seven coupled modes of vibration in a fraction of a second, exhibits excellent numerical stability. It constitutes a computationally efficient alternative to multi-body dynamics and finite element analysis for the integration of rotor blade flexible modelling into a wider comprehensive rotorcraft tool.
The cycloidal propellers for micro aerial vehicle scale cyclocopter in hovering status were studied in this paper based on the URANs solver using 2D, 2.5D, 3D half blade and 3D full blade model. The results from all numerical models were validated with the experimental results. It was found that 2.5D model cannot produce more accurate results than 2D model, hence results from 2D model were employed to discuss cycloidal rotor with infinite blade span. It was also indicated that the 3D half blade model produced the same results as 3D full blade model, but was more efficient than 3D model. The numerical simulation results of cycloidal rotor with finite (3D model) and infinite span (2D model) were compared. The results indicated that for the 2D cycloidal rotor with large blade pitching amplitude, there were leading edge and trailing edge vortices due to dynamic stall, which resulted in parallel blade vortex interactions. The parallel blade vortex interactions will also induce the fluctuation of aerodynamic forces. For the 3D blade with small aspect ratio, the flow was dominated by 3D dynamic stall and blade vortex interactions. The 3D flow due to finite blade span resulted in smooth dynamic stall and can weaken the parallel blade vortex interactions induced by dynamic stall vortices, hence no strong aerodynamic force fluctuation was observed. The perpendicular blade vortex interactions caused by blade tip vortices can induce cross flow when the azimuth angle of the rotor is between 270° and 360°, which reduces the strength of downwash in the region where the rotor azimuth angle is between 180° and 360°. This results in much smaller side force. Although sometimes the time-averaged aerodynamic forces obtained by 2D and 3D model were quite close to each other, the physics lying behind is quite different. Hence, it was not correct to use the 2D models to discuss the principles of cycloidal rotors with finite blade span.
Stringers are stiffening members of pressurized aircraft fuselage. They provide support to the fuselage’s skin. A new stringer grid concept is proposed for conventional aircraft fuselage. Optimization is used to find the hexagonal grid that best replaces the original while keeping the same total stringer length. A finite element model is built to analyze the optimal hexagonal grid stiffened structure and compare it with the original orthogonally stiffened structure in terms of eigenfrequencies and static response to external loading. The finite element model is validated through Flugge’s analytical expressions for stiffened shells. Results show that the hexagonal grid stiffened structure yields higher eigenfrequencies with stresses and displacements comparable with that of the original structure.
Random-vibration fatigue evaluation can be of considerable importance in the design phase of aerospace structures due to the severe dynamic loads in service. This paper presents the utilization of modal stress approach to the issue of structural random-vibration fatigue evaluation. Prognosis of random fatigue hotspots by using stress mode shapes is theoretically demonstrated. A two-step procedure is proposed for computational efficiency. Firstly, modal stress analysis is conducted to locate the fatigue hotspots in a dynamic structure. Secondly, the frequency domain-based approach for random fatigue evaluation is performed at these hotspots, as opposed to the computation of the entire structure as before. The capability of stress mode shapes to locate fatigue hotspots is verified by numerical investigations. The finite element model of a single-lap plate structure containing various opening holes was constructed for case study. Six elements were identified as hotspots by using modal stress distributions. Then, random responses and fatigue evaluation of the entire structure were carried out for verification. Good agreement was observed between the fatigue damage contour and the modal stress distributions, which can indicate that the critical positions predicted by stress mode shapes have good accuracy. The calculation time and storage space can be significantly reduced by means of the proposed evaluation procedure. Therefore, the accuracy and efficiency of utilization of modal stress approach in random fatigue evaluation can be ensured.
In order to make the shock train leading edge detection method more possible for operational application, a new detection method based on differential pressure signals is introduced in this paper. Firstly, three previous detection methods, including the pressure ratio method, the pressure increase method, and the standard deviation method, have been examined whether they are also applicable for shock train moving at different speeds. Accordingly, three experimental cases of back-pressure changing at different rates were conducted in this paper. The results show that the pressure ratio and the pressure increase method both have acceptable detection accuracy for shock train moving rapidly and slowly, and the standard deviation method is not applicable for rapid shock train movement due to its running time window. Considering the operational application, the differential pressure method is raised and tested in this paper. This detection method has sufficient temporal resolution for rapidly and slowly shock train moving, and can make a real-time detection. In the end, the improvements brought by the differential pressure method have been discussed.
In this paper, we propose a design philosophy for cooling high-pressure nozzle guide vane endwalls, which exploits the momentum of cooling jets to control vane secondary flows thereby improving endwall cooling uniformity. The impact of coolant-to-mainstream pressure ratio, hole inclination angle, hole diameter, vane potential field, and overall mass flux ratios are considered. Arguments are developed by examining detailed experimental studies conducted in a large-scale low-speed cascade tunnel with engine-realistic combustor geometry and turbulence profiles. Computational fluid dynamics predictions validated by the same are used to extend the parameter space. We show that the global flow field is highly sensitive to the inlet total pressure profile, which in turn can be modified by introducing relatively low mass flow rates of cooling gas at engine realistic coolant-to-mainstream pressure ratios and mass flux ratios. This interaction effect must be understood for successful design of optimised endwall cooling schemes, an effect which is not sufficiently emphasized in much of the literature on this topic. Design guidelines are given that we hope will be of use in industry.
Vapor core pump, unlike traditional centrifugal pumps with bypass throttling, exhibits the superiority in active fuel regulation. The vapor–liquid two-phase flow inside vapor core pump is an interesting process over its full regulation range, which influences performance of the pump. In this paper, an enclosed radial straight blade vapor core pump is selected as the object. Through vapor–liquid two-phase numerical simulation, the configuration of the vapor core in impeller zone and the regulation performance of the pump are obtained under several typical operating conditions. The main innovation in the simulation is that vapor volume fraction inside the pump as a means of actual vapor core description is obtained through numerical method with the Schnerr–Sauer cavitation model based. Vapor-phase volume fraction, velocity, head, and efficiency in relation to the opening or fuel rate are calculated and analyzed with inlet or outlet load changing. Performance testing on a vapor core pump prototype is carried out to verify the numerical results. Comparison between the experiment and the simulation shows the acceptable difference between them and the effectiveness of the numerical method. Study results indicate that a vapor core pump has a wide active regulation range of fuel delivery rate. For pure vapor core conditions, a considerable efficiency loss is unavoidable. Inlet throttling and hydraulic loss in its impeller play a significant role in deviated flow regulation characteristics. The opening size and the inlet or outlet load affect the pump delivery performance, which could be used to choose proper operation conditions or regulation range where neither much efficiency loss nor other performance degeneration on the pump occurs.
The shock stand-off distance for basic spherical models was investigated in the flow regime of 1–2 km/s, a range for which data are unavailable in the open literature. Experiments were conducted on two hypersonic shock tunnels, at five different enthalpy conditions and two types of test gas, and the bow shock was captured using schlieren flow visualization technique. These results were compared with existing empirical relations and a good match was observed. To further corroborate the results, a computational analysis of the same problem was carried out using an in-house computational fluid dynamics code, HiFUN (High Resolution Flow Solver on UNstructured meshes), and the results were again seen to match well with the experiments. The shock stand-off distance is an important parameter for the design of high-speed airplanes, and therefore a solid quantitative understanding of this parameter is vital for all flow regimes. This paper aims to fill in the gap by investigating one of those regimes where available data are currently missing.
A brand-new area of low earth orbit satellite constellation research has been expanded by global navigation satellite system (GNSS) radio occultation (RO) atmosphere sounding technology in the last decade. The possibility of reducing the sounding satellites while keeping the amount of atmosphere soundings increasing to produce low-cost meteorological data product is investigated, and a constellation capable of receiving the RO signals from all the 4-GNSS, such as GPS, GLONASS, Galileo, and Compass, is proposed in this paper. This paper focuses on the mathematical problems on the design of 4-GNSS RO satellite constellation. A forward GNSS RO sounding simulation algorithm based on ideal atmosphere model and two-dimensional radial tracing algorithm is presented for a rapidly and accurately sounding performance prediction of 4-GNSS RO satellite constellations. Then, an improved uniformity evaluation factor named Dual-gate is established for 4-GNSS RO satellite constellation optimization design, which is a combination of the uniformity evaluation factors for the latitudinal distribution of RO soundings and those for the gridding uniformity. On the basis of low earth orbit satellite orbit dynamics and spherical geometry, the impacts of partial constellation parameters on the amount and coverage performance accorded with Dual-gate uniformity evaluation index are derived, and a series of design criteria of 4-GNSS RO satellite constellation are summarized. A simplified 4-GNSS RO satellite constellation model is built on the basis of the design criteria, and an improved ant colony algorithm is used to optimize the parameters of a 4-GNSS RO constellation including as many satellites as COSMIC-2. The simulation result shows that this 4-GNSS RO constellation is capable of obtaining near 3000 atmosphere soundings per 3 h. It obtains 13% more soundings than COSMIC-2 with the same 4-GNSS RO sounding devices, and the uniformity of soundings is increased by 9%.
Transonic centrifugal compressors with high efficiency and wide stable flow range are required in modern gas turbine engines. Blade design with complex three-dimensional features is one of the promising methods to further improve the performance of such cases. Aiming to increase the efficiency while maintaining similar level of the stable flow range, this article investigates aerodynamic potentials of complex three-dimensional features in a transonic centrifugal compressor by multi-point and multi-objective optimizations, in which the camber curves, the sweep feature, and the lean feature have been optimized. During the first round of optimizations, the aforementioned three groups of variables are optimized individually, and their sensitivities to the performance have been analyzed. When optimizing the camber curves, the best result shows an end-bend feature at the front of the hub section, and the efficiency is improved by 1.0% due to the lowered shock strength. When optimizing the sweep feature, the best result presents an S-shape leading edge and a forward sweep feature. The efficiency is increased by 0.5% because of the reduced wake region. The optimized lean feature only improves the efficiency by 0.2%, which shows its relatively low potential. The final round of optimizations couples both the camber curves and the sweep feature, and the best geometry combines both the end-bend and S-shape leading edge patterns. The peak efficiency and the choke mass flow rate have been increased by 2.2% and 8.1%, respectively, which is owing to the combination of the lowered shock strength by optimized camber curves and the reduced wake region by optimized sweep feature. The result indicates significant potential of complex three-dimensional features to improve the performance of transonic centrifugal compressors.
In this paper, a computational fluid dynamics trimming method is proposed and compared with wind tunnel experiment and the blade element method. The NASA’s generic ROBIN helicopter model is adopted for transient simulations to obtain the final main rotor trimming conditions. Totally three steps were applied to the computational fluid dynamics method. The first step is associated with no cyclic pitch motion, the second is regarding pure longitudinal cyclic pitch motion and the last is concerning with pure lateral cyclic pitch motion. At the same time, a simple linear equation system between the roll and pitching moment was established to get the final longitudinal and lateral cyclic pitch angles for the main rotor through the above three steps. An overset grid approach was used where the volume around each blade was modeled in an individual overset grid region. The rotor rotation was resolved with three degrees per time-step. Turbulence was modeled with the well-known SST K-omega model with second-order convection. The helicopter was in straight forward flight with an advance ratio of
This article presents an algorithm for the identification of modal parameters during flutter flight testing when forced excitation is employed and the aircraft possesses several sensors for structural response acquisition. The main novelty of the method, when compared with other classical modal analysis methods, is that the analysis is carried out in intervals of time instead of in the whole duration of the excitation. It means that, even when the response signal is only partially available, some modal parameters may be still identified. Application to analytic signals as well as structural response of modern fighter aircraft using frequency-swept excitation is provided in order to demonstrate the effectiveness, robustness and noise immunity of the proposed method.
Statistical data clearly points out compressor train outage as one of the main reasons for the breakdown of aircraft power units. Outages are caused by various factors, including material conditions as the main one. Therefore, structural ultimate strength analysis of constructional raw material properties used in the production process of such an important part of an engine is not only of tremendous significance during the designing phase but also in a research phase and expertise associated with emergency situations. Considering the above-mentioned factors the article presents the outcomes of chemical composition, morphology, and phase structure of metallic material used to produce first row fan blades of RD-33 turbine jet engine, which is (currently) used by the Polish Air Force. Apart from typical material structure analysis, also the basic mechanical strength properties of materials have been determined including hardness, tensile, and impact tests. These may now constitute the basis for the analysis of substitute materials selection allowing the production of analyzed engine part.
This paper reports on a time-accurate numerical investigation of axial-skewed slot casing treatment in a high-speed axial fan. Twenty-two axial and radial skewed slots are placed over the casing of the rotor and choke to stall unsteady computations are conducted. Results show that endwall fluid is absorbed by the slots from the downstream part of the blade passage and is injected to the upstream part with a swirl contrary to the blade rotation. This gives the injected fluid a great circumferential velocity component in the relative frame of reference. The shock is, hence, pushed toward the trailing-edge plane and the pressure difference between the pressure and suction surfaces is reduced. Consequently, the occurrence of the leading-edge vortex spillage, which is believed to be the key to stall initiation for the current rotor, is postponed to lower mass flow rates.
Space radiators used in aerospace applications are required to have a minimum weight, for the desired heat transfer rate. To ensure that the radiator has an optimal geometry, it is important that the heat transfer rate be calculated accurately. To obtain the heat transfer rate from the radiator accurately, the radiative interaction between the fin surface and fin base and the variation of thermal conductivity with temperature should be included in the analysis. Taking into account these two phenomena, this study was conducted in order to explore the optimal dimensions of a space radiator. The dimensionless nonlinear and nonhomogeneous fin equation is solved using the variation of the parameters method for carrying out the required optimization procedure. The optimization results are presented as convenient correlation equations for suitable ranges of problem parameters.
Stall recovery process for performance enhancement of an axial compressor has been experimentally carried out using air injection at its rotor blade row tip region. Twelve air injectors had been mounted evenly spaced around the compressor casing upstream the rotor blade row. Instantaneous flow velocities at various radial and circumferential positions were measured simultaneously utilizing hot wire anemometry. These unsteady results, separately presented in frequency and time domains, provided to distinguish stall inception process and consequent flow induced fluctuations. Time-dependent responses of the flow field properties within the compressor passage and progressive alleviation of stall cells are demonstrated during the tip injection process. Blade tip air injection worked effectively and enhanced the compressor stall margin about 9%. This attractive result occurred for the case where the total mass flow rate passing through the air injectors was as small as 0.8% of the compressor main flow rate. In addition, this augmentation in the stall margin was accompanied by increase in the compressor delivery total pressure. Air injection at the blade row tip region caused its beneficial effects to extend throughout the blade whole span, especially while working at the near-stall conditions.
The real-time trajectory replanning method which is used for the guidance of Mars entry is investigated in this paper. Comparing with the traditional Mars entry guidance methods, such as the reference-trajectory tracking guidance and predictor–corrector guidance, the real-time trajectory replanning method can increase the reliability of the mission remarkably. When faults occur during the Mars entry phase, a replacement trajectory will be planned quickly. Due to the limited onboard computing capacity, replanning the trajectory onboard is a challenging task. Corresponding to this problem, the neural network is trained to approximate the dynamics of the atmospheric entry. The uncertain factor of the atmospheric density is also included in the neural network. Then, by using the characters of the neural network, the analytical expressions of the Jacobian which are needed in trajectory optimization are derived. Finally, an estimation-replanning guidance procedure is introduced. The numerical simulation shows that the proposed guidance strategy can decrease the error of final states effectively, and the neural network approximation improves the computational speed of the nonlinear programming solver remarkably, which makes the method more suitable for use onboard.
The vortex flow and lift force generated by a 50°-sweep non-slender reverse delta wing were investigated via particle image velocimetry, together with flow visualization and force balance measurement, at Re = 11,000. The non-slender reverse delta wing produced a delayed stall but a lower lift compared to its delta wing counterpart. The stalling mechanism was also found to be triggered by the disruption of the multiple spanwise vortex filaments developed over the upper wing surface. The vortex flowfield was, however, characterized by the co-existence of reverse delta wing vortices and multiple shear-layer vortices. The outboard location of the reverse delta wing vortex further implies that the lift force is mainly generated by the wing lower surface while the upper surface acts as a wake generator. The spatial progression of the flow parameters of the vortex generated by the non-slender reverse delta wing as a function of α was also discussed.
A fluid–structure interaction numerical simulation was performed to investigate the flow field around a flexible flapping wing using an in-house developed computational fluid dynamics/computational structural dynamics solver. The three-dimensional (3D) fluid–structure interaction of the flapping locomotion was predicted by loosely coupling preconditioned Navier–Stokes solutions and non-linear co-rotational structural solutions. The computational structural dynamic solver was specifically developed for highly flexible flapping wings by considering large geometric non-linear characteristics. The high fidelity of the developed methodology was validated by benchmark tests. Then, an analysis of flexible flapping wings was carried out with a specific focus on the unsteady aerodynamic mechanisms and effects of flexion on flexible flapping wings. Results demonstrate that the flexion will introduce different flow fields, and thus vary thrust generation and pressure distribution significantly. In the meanwhile, relationship between flapping frequency and flexion plays an important role on efficiency. Therefore, appropriate combination of frequency and flexion of flexible flapping wings provides higher efficiency. This study may give instruction for further design of flexible flapping wings.
This paper studies the problem of attitude tracking control for spacecraft rendezvous and docking based on a physical ground simulation system. Two finite-time controllers based on quaternion are proposed by using a novel fast nonsingular terminal sliding mode surface associated with the adaptive control, the novel fast nonsingular terminal sliding mode surface not only contains the advantages of the fast nonsingular terminal sliding mode surface, but also can eliminate unwinding caused by the quaternion. The first controller, which is continuous and chattering-free, can compensate unknown constant external disturbances, while the second controller can both compensate parametric uncertainties and varying external disturbances with unknown bounds without chattering. Lyapunov theoretical analysis and simulation results show that the two controllers can make the closed-loop system errors converge to zero in finite time and guarantee the finite-time stability of the system.
Experimental investigations of the effect of inlet blade loading on the rotating stall inception process are carried out on a single-stage low-speed axial compressor. Temporal pressure signals from the six high response pressure transducers are used for the analysis. Pressure variations at the hub are especially recorded during the stall inception process. Inlet blade loading is altered by installing metallic meshed distortion screens at the rotor upstream. Three sets of experiments are performed for the comparison of results, i.e. uniform inlet flow, tip, and hub distortions, respectively. Regardless of the type of distortion introduced to the inflow, the compressor undergoes a performance drop, which is more severe in the hub distortion case. Under the uniform inlet flow condition, stall inception is caused by the modal type disturbance while the stall precursor switched to spike type due to the highly loaded blade tip. In the presence of high blade loading at the hub, spike disappeared and the compressor once again witnessed a modal type disturbance. Hub pressure fluctuations are observed throughout the process when the stall is caused by a modal wave while no disturbance is noticed at the hub in spike type stall inception. It is believed that the hub flow separation contributes to the modal type of stall inception phenomenon. Results are also supported by the recently developed signal processing techniques for the stall inception study.
The effects of the volute’s asymmetry on the performance of a turbocharger centrifugal compressor were studied using steady simulations and theoretical analysis. According to the steady simulation results, it is found that the volute’s asymmetry has significant influence on the performance of the centrifugal compressor. The variation of the stage efficiency due to volute’s asymmetry is up to 4%. Meanwhile, the volute’s asymmetry restricts the compressor stable flow range by imposing a distorted outlet pressure condition and forcing some certain impeller passages to suffer from a worse flow than the others. These certain passages are likely to stall first and trigger the surge, as the stage flow rate further decreases. In other words, the local stall triggers the surge. The relevant flow mechanisms were given to explain the effects based on the three-dimensional flow field, and a new model was developed to demonstrate how the local stall induced by the volute’s asymmetry triggers the system instability.
Boundary layer suction is considered to be an available approach to restraining or even eliminating flow separation and to improve the aerodynamic performance of the compressor. In this paper, a highly loaded axial-flow aspirated compressor based on a low-reaction design concept is investigated in detail to find an appropriate flow control strategy for boundary layer suction to achieve significant performance benefit. The flow control strategy consists mainly of the arrangement of suction hole and the aspirated flow rate. The geometrical models including aspiration cannulas, stator cavity and aspiration channel are novelly applied in this research to approach a real engineering application. Complete compressor maps are predicted by a three-dimensional computational fluid dynamics simulation. The distribution of the typical aerodynamic parameters and the partial flow structures are analysed at design and off-design conditions. Three-dimensional separation near the stall point is effectively suppressed by the aspiration on both hub and shroud, and better performance is achieved by a reasonable increase of aspirated flow rate; a peak efficiency of 0.91 and a total pressure ratio of 1.055 are attained at a total aspirated flow rate of 0.024 kg/s per passage. However, sudden turning of compressor maps at low inlet flow rate occurs without the aspiration on shroud, and a noticeable deterioration in performance occurs with the decreasing of the inlet flow rate. The main reason of performance deterioration is that the flow control efficiency of aspiration is not enough for effectively suppressing three-dimensional separation near the casing corner. The difference in the aspirated flow source is owed to the distinct flow feature at various locations caused by the aspiration efficiency of suction holes. A partial auto-readjustment feature of the suction flow rate with increasing flow separation has been found. The boundary layer suction with an appropriate flow control strategy is necessary for a higher efficiency, highly loaded aspiration compressor.
The present paper numerically conducted full-annulus investigation on the effects of circumferential total pressure inlet distortion on the performance and flow field of the axial transonic counter-rotating compressor. Results reveal that the inlet distortion both deteriorates the performance of the upstream and downstream rotors resulting in reduction of total pressure ratio, efficiency and stall margin of the transonic contra-rotating compressor. Regarding the development of distortion inside compressor, the downstream rotor reinforces the air-flow mixing effects and, thus, attenuates the distortion intensity significantly. Under the distorted inflow conditions, the detached shockwave at the leading edge of downstream rotor interacts with the tip leakage flow and causes the blockage of the blades passage, which is one important reason for the transonic contra-rotating compressor stall.
An optimization model of the optimum area of solar array for a stratospheric solar-powered airship is developed. The objective of the optimization is to reduce the mass of the solar array on an airship by keeping the equilibrium between output power and weight of solar array as a constraint. Based on several parameters of the typical existing airships, the optimization works are carried out to verify the effectiveness of the optimization model. Furthermore, the effects of the wind velocity, airship’s latitude and working date on the optimum area are analyzed in detail. The results of this study demonstrate that the optimization model is a good tool for the preliminary design of solar array on an airship. It can also be found that these influence factors have significant influences on the optimum area, and these factors should be considered together to obtain an optimum and balanced design due to the strong dependence on each other.
Aerial planting is a new and affordable method, which is used for purposes of reforestation and restoring pastures. In this study, flight dynamics of an aerial planting projectile has been simulated using data of the wind tunnel test. By creating the projectile graphical model and using the results of dynamic equations for this model, pitching angle has been simulated graphically by recording the pitch attitude of the graphical model and, by applying spectral filtration the rate of pitch angle at any point of the trajectory is calculated. Primary dynamic model parameters including the coefficients of regression equation of motion are estimated by applying the least squares estimator. Comparison of the estimated coefficients from image processing and coefficients given to the simulation program reveals high accuracy of the model resulting from the image processing data used in this method and approves wind tunnel test data by graphical model and calibration of camera measurement. Also, an experimental test setup has been created and the images of projectile falling down in the presence of fan flow have been captured with high-speed digital camera. By decreasing the ambient light intensity, the images and measurement noise of theta angle increases. By applying recursive least square algorithm, sensitivity of coefficients and robustness of this algorithm has been analyzed. The analysis results indicate that the estimation of coefficients using image processing data has great accuracy.
Hartmann–Sprenger tube is a device in which an underexpanded jet enters a closed end tube that is placed in a specific distance from the nozzle. In specific conditions, a standing shock is built in front of the tube, which oscillates based on the tube resonance frequency and creates oscillatory flows with periodic shock motions along the tube. In these conditions, intensive temperature rise could be observed near the tube end wall. Considering these thermal effects, the device could be used as a combustion starter in the space propulsion systems. The present study focuses on flow analysis in various phases of the oscillatory process in a semi-conical PTFE resonance tube by the aim of numerical simulation results. An experimental test is also performed for validation purposes. The T–S diagram is plotted to describe the thermal effects in detail during the oscillatory processes. Various modes of shock contact with flow front are described. In order to follow up the shock traveling process, the diagrams of changes related to major flow properties inside the tube are used. Generation of small turbulences at the moments of combination of compression waves and beginning of flow entrance is also detected. According to the results, traveling of shock waves through the trapped gas was found to be the major mechanism for heat generation inside the tube. The thermal effects are also compared in the conical and cylindrical tubes. The flow analysis will lead to increase in insight for shock motion and heat generation mechanism in a semi-conical Hartmann–Sprenger tube.
In order to provide a method for evaluating flight control systems with the wind tunnel based virtual flight testing and provide a guide for building virtual flight testing systems, the virtual flight testing evaluation method was researched. The virtual flight testing evaluation method consisted of three parts: virtual flight testing method, virtual flight testing data processing method, and flight control system performance determination method, which were respectively designed for a pitching control system. Then, the hardware-in-the-loop simulation evaluation method was presented, and comparisons between the virtual flight testing and hardware-in-the-loop simulation evaluation method were conducted to highlight the characteristics of virtual flight testing evaluation method. Finally, virtual flight testing simulation models of a sample air vehicle were built and virtual flight testing were simulated to demonstrate the virtual flight testing evaluation method, which is helpful for the understanding of the virtual flight testing evaluation method with more sensibility. The evaluation results show that the virtual flight testing evaluation method designed can be used for flight control system evaluation.
The hinge moment acting on the actuator will cause an out-of-plane moment, which is a destabilizing factor to the angular motion of spinning missiles. A new tuning criterion for the actuator controller is proposed to decrease the out-of-plane moment. It is noted that the integral element does not decrease the out-of-plane moment. A carefully designed proportional–derivative controller with some compromises can ensure the stability of the missile and provide good performance for the actuator.
A cruciform-finned slender body can excite a limit-cycle rolling oscillation at high angles of attack. To suppress the unwanted motions, flow control approaches should be employed if the aerodynamic control surfaces lose control efficiency at high angles of attack. As a promising technology, the ns-dielectric barrier discharge plasma actuator has been successfully used in high-speed and high-Reynolds-number flow control applications. The present work employs a -type ns-dielectric barrier discharge plasma actuator and vortex generators to suppress self-excited rolling oscillation at α = 50°. The free-to-roll tests show that the plasma actuator ignited at F+ ~ 1.5 and that the vortex generators can suppress rolling oscillation. The flow patterns from particle image velocimetry measurement at different cross-sections and rolling angles suggest that the vorticity decrease of the leeward vortices may be the control mechanism for the plasma actuator. For the vortex generators, evident modification of the flow field structure can be observed due to the vortices generated from the vortex generators, which decreases the rolling moment induced by the asymmetry vortices to suppress the self-excited rolling oscillation.
An output feedback observer-based dynamic surface controller is presented for attitude tracking problem of the quadrotor unmanned aerial vehicle, which is subject to measurement noise and external disturbances. The dynamics model of the quadrotor unmanned aerial vehicle is firstly introduced with the quaternion representation. Subsequently, a nonlinear augmented observer is introduced for simultaneously estimating the unavailable states and uncertain disturbances from the measurement of system output. The output feedback controller based on the nonlinear augmented observer is designed with the dynamic surface control technique. The Lyapunov stability analysis shows that the attitude tracking performance is ensured and all signals of the closed-loop system remain bounded. Finally, simulative and experimental results are carried out to illustrate, compared with other observer-based controller, the effectiveness of the proposed method is better.
Damped free vibration of carbon nanotube reinforced composite microplate bounded with piezoelectric sensor and actuator layers are investigated in this study. For the mathematical modeling of sandwich structure, the refined zigzag theory is applied. In addition, to present a realistic model, the material properties of system are supposed as viscoelastic based on Kelvin–Voigt model. Distributions of single-walled carbon nanotubes along the thickness direction of the viscoelastic carbon nanotube reinforced composite microplate are considered as four types of functionally graded distribution patterns. The viscoelastic functionally graded carbon nanotube reinforced composite microplate subjected to electromagnetic field is embedded in an orthotropic visco-Pasternak foundation. Hamilton’s principle is employed to establish the equations of motion. In order to calculate the frequency and damping ratio of sandwich plate, boundary condition of plate is assumed as simply-supported and an exact solution is used. The effects of some significant parameters such as damping coefficient of viscoelastic plates, volume fraction of carbon nanotubes, different types of functionally graded distributions of carbon nanotubes, magnetic field, and external voltage on the damped free vibration of system are investigated. Results clarify that considering viscoelastic property for system to achieve accurate results is essential. Furthermore, the effects of volume fraction and distribution type of carbon nanotubes are remarkable on the vibration of sandwich plate. In addition, electric and magnetic fields are considerable parameters to control the behavior of viscoelastic carbon nanotube reinforced composite microplate. It is hoped that the results of this study could be applied in design of nano/micromechanical sensor and actuator systems.
In this paper, an energy method for flutter analysis of wing using one-way fluid structure coupling was developed. To consider the effect of wing vibration, Reynolds-averaged Navier–Stokes equations based on the arbitrary Lagrangian Eulerian coordinates were employed to model the flow. The flow mesh was updated using a fast dynamic mesh technology proposed by our research group. The pressure was calculated by solving the Reynolds-averaged Navier–Stokes equations through the SIMPLE algorithm with the updated flow mesh. The aerodynamic force for the wing was computed using the pressure on the wing surface. Then the aerodynamic damping of the wing vibration was computed. Finally, the flutter stability for the wing was decided according to whether the aerodynamic damping was positive or not. Considering the first four modes, the aerodynamic damping for wing 445.6 was calculated using the present method. The results show that the aerodynamic damping of the first mode is lower than the aerodynamic damping of higher order modes. The aerodynamic damping increases with the increase of the mode order. The flutter boundary for wing 445.6 was computed using the aerodynamic damping of the first mode in this paper. The calculated flutter boundary is consistent well with the experimental data.
A cubic-spline-based time collocation method is used to solve periodic transonic flows for aeroelastic analysis. The time periodical flow variables are first approximated by cubic splines, and then the time-derivative terms are represented by a source term, which couples the flow solutions at all the sampled time instants. Simulation of two transonic external unsteady flows, including the oscillating NACA 0012 aerofoil and pitching rectangular supercritical wing, were carried out. The cubic-spline-based time collocation method with relatively small number of time instants sampled demonstrate high accuracy, whilst it is several times faster than the conventional time-marching method. The effect of different basis function used in the time collocation methods are also studied by comparing the results to that of the state-of-the-art time spectral method. This method was also validated against aeroelastic experimental data of the vibrating STCF 11 turbine blades. Only single passage computational mesh is involved by adopting the phase-lag periodic conditions. Good accuracy of the proposed time collocation method is also achieved and about one order of speed-up is obtained as compared with time-marching simulation.
The effective means of air fuel mixing and flame holding can be achieved by incorporating cavity in supersonic combustor. Understanding the complex flow field of cavity flow is essential for the design of supersonic combustor. An attempt is made to understand the characteristics of supersonic flow past axisymmetric cavity, and a series of nonreacting experiments are carried out in a blow-down type supersonic flow facility. The facility consists of a supersonic nozzle, issues a flow Mach number of 1.80 into a circular cross sectional supersonic combustor in which axisymmetric cavity is placed. Cavity of two consecutive aft wall angles is the key parameter for the study. The performance of the cavity is investigated based on the static pressure measurement, momentum flux distribution at the exit plane of the combustor, and the stagnation pressure loss of the flow. Wall static pressure distribution revealed that pressure increases with decrease in the secondary aft wall angle below 45° due to stronger recompression of shear layers. Moreover, decreasing primary aft wall angle provides a uniform mixing profile along with decrease in stagnation pressure loss across the combustor.
The paper presents an analysis of an unmanned aerial vehicle gliding in an airliner wake vortex using the dynamic soaring principle. The goal of dynamic soaring is an improvement of flight performance of the unmanned aerial vehicle following an airliner. The paper extends previously published results of an airliner–unmanned aerial vehicle climb regime flight formation analysis. Wake vortex model, the unmanned aerial vehicle basic parameters including drag polar and an airliner climb profile are taken from previous research. Dynamic soaring simulation within wake vortex is performed and evaluated. The presented results provide insight into energy balance and controllability of the unmanned aerial vehicle flying in an airliner wake vortex.
The noteworthy feature of aircraft with distributed propulsion configuration is the integration of a blended-wing-body type airframe and an embedded distributed propulsion system, thus inducing the specific boundary layer ingestion effect. Different boundary layer ingestion effects on the distributed engines may generate asymmetric flow fields on the airframe surface, and then lead to the unique lateral-directional aero-propulsive close coupling. To investigate the lateral-directional aerodynamics influenced by boundary layer ingestion, a new comprehensive computational method based on the differentiated boundary conditions is proposed. This method uses a synthetic three-dimensional computational model including the airframe and multi-engine to analyze the aerodynamic characteristics, and the essential boundary conditions can be extracted from the thermodynamic distributed propulsion system model to represent the different boundary layer ingestion intensities on the left and right engines. Subsequently, detailed model-based analyses of boundary layer ingestion influences on the lateral-directional aerodynamic characteristics are conducted, and the influence regularities under different flight states are revealed. All the results demonstrate that the differentiated boundary layer ingestion intensities on distributed engines can certainly affect the roll and yaw aerodynamic performance of the distributed propulsion configuration aircraft.
Large space structures experience changing thermal environment during orbiting the earth. The resultant temperature gradients induce structural deformations that may downgrade performance of payloads conducting high precision missions and even affect stability of the spacecraft. So, it is extremely important to analyze thermally induced deformation of large space structures for routine operation. In this paper, the ultra-large truss support membrane structure on satellite is characterized and studied. The methodology of thermal quasi-static deformation is formulated and the procedure of thermo-structural analysis is proposed. The thermostructural analysis model with hollow tubes is developed based on finite difference method and finite element method. With heat fluxes from solar radiation, earth radiation, and earth albedo radiation being considered, the temperature distribution filed is obtained from the thermal analysis and then applied to the structural analysis model to calculate quasi-static deformations and root mean square errors with orbital angles. Results show that temperature gradients along circumferential direction of tubes can induce prominent shape error. The proposed method is useful for predicting thermally induced deformation of large space structures and valuable for designing active control systems to compensate for disturbances.
The design of space systems is a complex and multidisciplinary process with multiple conflicting objectives, large number of design variables, and constraints that limits application of the existing multidisciplinary design optimization architectures to this class of design problems. This paper presents an enhanced multidisciplinary design optimization architecture to concurrent holistic design optimization of a satellite system. The proposed multidisciplinary design optimization architecture extends concepts of multidiscipline feasible and bi-level integrated system synthesis into a unified architecture using metamodels. The proposed architecture was evaluated and compared with the existing multidisciplinary design optimization architectures that include all-at-once, bi-level integrated system synthesis, and multidisciplinary design optimization using a remote sensing small satellite in low earth orbit. The satellite design optimization problem deals with the minimization of the total mass of the satellite, involving disciplines of mission analysis, payload, structures, attitude determination and control, communication, command and data handling, power and thermal. The computational performance and accuracy of the proposed architecture were compared with multidisciplinary design optimization benchmark problems. Then the proposed architecture is successfully applied to the satellite system design problem. The results obtained show that metamodel-based bi-level integrated system synthesis-multidisciplinary design optimization architecture presented in this paper provides an effective way of solving large-scale design problems.
A conservative unstructured sliding-mesh technique is developed for the rotor–fuselage aerodynamic interaction simulation. The computational domain is decomposed into a rotational zone and a stationary zone. The rotational zone contains the rotor blades that rotate with the zone, while the stationary zone contains the fuselage which keeps stationary during the simulation. The two zones are connected via a sliding interface, which is designed to be a cylindrical surface consisting of the top, bottom, and side surfaces. The top and bottom surfaces are paved with arbitrary triangles and the side surface is meshed by triangularizing equal-sized and right-angled quadrilaterals. The intersection information between the rotational and stationary sliding interface meshes, such as the number of intersection triangles and the area of each intersection polygon, is the key to the present conservative computation. For the top and bottom surfaces, the point-on-line cases are first identified and the point perturbation operation is carried out to eliminate the potential error due to the presence of a point-on-line case. The neighbor-to-neighbor searching algorithm is applied for efficient determination of the intersection triangles, and the intersection polygon areas are determined by enumerating all the possible intersection cases. For the side surface, the intersection relations and polygon areas can be easily determined based on the enumeration method due to the special triangularization. The present method is validated by simulating the GIT (Georgia Institute of Technology) rotor–fuselage interaction model, and comparing numerical results with experiment measurements. It is demonstrated that the present conservative sliding-mesh method is simple to implement, and is efficient for the prediction of complicated unsteady rotor–fuselage aerodynamic interaction.
Analysis and design of composite helicopter landing gears are challenges. Cross tubes of helicopter landing gears consist of straight tubes at the middle and curved tubes at the sides. In this work, to simulate ground handling, thick laminated composite straight tubes subjected to pure bending moments are studied using a new high-order simple-input analytical method. The accuracy of the proposed method is subsequently verified by comparing the numerical results obtained using the proposed method with finite element method and experimental data. The results show good agreement. High efficiency in terms of computational time is achieved when the proposed method is used as compared with finite element method. In addition, a simple non-dimensional coefficient is proposed to predict interlaminar radial stresses of thick composite straight tubes.
The stator well in a compressor is the space between the rotor and stator inside the mainstream annulus flow. Labyrinth seals are normally used to control the internal flow in the stator well. The upstream and downstream rotating cavities of the labyrinth seal can lead to substantial temperature rise and swirl development in this region. Additionally, due to the centrifugal expansion and thermal expansion, the tip clearance of labyrinth seal changes dramatically at different rotational speeds and temperatures in the stator well. A test rig capable to establish different rotational speeds and pressure ratios was designed according to the simplified model of the labyrinth seal in a compressor stator well (one stage). The leakage flow rate and change in total temperature across the stator well were measured. This paper also contains the experimental results of swirl ratios in the outlet rotating cavity to reveal the swirl development. Special emphasis in this work lies on acquiring the working tip clearance precisely. The set up tip clearance was measured with plug gauges, while the radial displacements of labyrinth ring and stator casing were measured separately with two high precision laser distance sensors. Two-dimensional, axisymmetric swirl flow numerical simulations were performed to get a further understanding about the basic flow characteristics and to evaluate their ability to predict the experimental results. The computational results of discharge coefficient, windage heating, and swirl ratio were compared to those obtained from test rig measurements. Particularly, when calculating, the tip clearance, the inlet parameters, and the outlet parameters of numerical model at a specific rotating speed were set to be the same with the experimental conditions.
Icing and rainfall are the two critical meteorological factors that threaten the aircraft flight safety. A number of previous studies have shown their individual influences on aircraft aerodynamics; however, to date no studies on their coupling effects exist. In this paper, a numerical study is conducted to focus on the aerodynamic performance of a NACA 23012 airfoil exposed to heavy rain in the presence of an ice accretion by supercooled raindrop on the leading edge. An Eulerian–Lagrangian two-phase flow approach developed for rain simulation in our previous work is adopted here with some improvement in the turbulence model. A series of new phenomena about the aerodynamic performance and wake characteristics under the coupling effects of ice accretion and rainfall are found and discussed. These results do not seem to have been published previously and should be of significance to the aircraft industry for improved aircraft design and pilot training.
An optimal pulsed guidance law with a time-varying weighted quadratic cost function that enables imposing a predetermined intercept angle is presented. Due to the characteristic of impulse force, admissible variance of control is redefined. The optimal pulsed guidance law is deduced via extended maximum principle. The optimal pulsed guidance law is eventually transformed to solve the two-point boundary value problem. To decide a shooting point, an efficient algorithm is proposed by combining particle swarm optimization and Kriging surrogate model method. The optimal pulsed guidance law is implemented in several representative engagements. From simulation results, it can be seen that the proposed guidance law can achieve small miss distance with terminal impact angle constraint under different conditions. Moreover, the performance of the proposed guidance law is satisfactory with the comparison of sliding-mode pulsed guidance law.
This paper deals with the closed-loop form of mid-course guidance law design for accelerating missile system, whose acceleration is approximately constant. A midcourse guidance algorithm of feedback form is proposed to satisfy the engagement geometry conditions at the burn-out time for terminal homing performance enhancement. The effect of velocity change due to missile acceleration is explicitly considered in the derivation of the guidance law. The terminal constraint update algorithm is proposed under the assumption that the target trajectory is predicted precisely. Simulation results are provided to show the performance and characteristics of the proposed algorithm.
This work proposes a scheme to select a proper dynamics model for space debris removal which is captured by a Tethered Space Robot. A proper dynamics model is crucial for the parameters estimation and controller design in a Tethered Space Robot mission, in particular, for the retrieval or de-orbiting of uncooperative target. A new dynamics model of the system is derived by treating the base satellite and the space debris as rigid bodies in the presence of offsets, and with the effect of the tether’s flexibility and elasticity. Then the equations of motion are simplified based on the attitude analysis and numerical simulations in different cases. It is concluded that in the Tethered Space Robot’s mission, the strong coupled attitude motions among the base, target satellites, and the tether cannot be ignored during the retrieval, which is totally different from the traditional tethered satellite system. The attitude motions of the system in different conditions are discussed respectively, and a method of the model selection of the system during post-capture and retrieval/de-orbiting phase is proposed, which is a balance of the accuracy and facility. Finally, a control scheme is used to prove this conclusion.
Single hidden layer perceptron neural network controllers combined with dynamic inversion are applied to the tilt-rotor unmanned aerial vehicle and its variant model with the nacelle mounted wing extension. The bandwidths of the inner loop and outer loop of the controller are designed using the timescale separation approach, which uses the combined analysis of the two loops. The bandwidth of each loop is selected to be close to each other using a combination of the pseudo-control-hedging and the pole-placement method. Similar to the previous studies on sigma-pi neural network, the dynamic inversion at hover conditions of the original tilt-rotor model is used as a baseline for both aircraft, and the compatible solution to the Lyapunov equation is suggested. The single hidden layer perceptron neural network minimizes the error of the inversion model through the back-propagation adaptation. The waypoint guidance is applied to the outermost loop of the neural network controller for autonomous flight which includes vertical take-off and landing as well as nacelle conversion. The simulation results under the two wind conditions for the tilt-rotor aircraft and its variant are presented. The south and north-west wind directions are simulated in order to compare with the results from the existing sigma-pi neural network, and the estimation results of the wind are presented.
An aircraft can extract energy from a gradient wind field by dynamic soaring. The paper presents trajectory optimization of an unmanned aerial vehicle for dynamic soaring by numerical analysis and validates the theoretical work through flight test. The collocation approach is used to convert the trajectory optimization problem into parameters optimization. The control and state parameters include lift coefficient, bank angle, positions, flight path angle, heading angle, and airspeed, which are obtained from the parameter optimization software. To validate the results of numerical simulation, the dynamic soaring experiment is also performed and experimental data are analyzed. This research work shows that the unmanned aerial vehicle can gain enough flight energy from the gradient wind field by following an optimal dynamic soaring trajectory. Meanwhile, the variation of flight path angle, heading angle, and airspeed has a significant influence on the energy transform. The solution can provide theoretical guide to unmanned aerial vehicles for extracting maximum energy from gradient wind fields.
A new control scheme for flexible air-breathing hypersonic vehicle is designed in this paper based on non-singular fast terminal sliding mode control and nonlinear disturbance observer. The proposed control scheme is derived from basic back-stepping method, which is capable of handling the higher-order nonlinear system, and a novel terminal sliding mode control method is designed for the last step to promise the finite time convergence and improve the steady-state precision. Meanwhile, a command filter is used to avoid the "explosion of complexity" in traditional back-stepping method. To overcome inevitable uncertainties as well as cross couplings between flexible and rigid modes, NDO is introduced to estimate diverse uncertainties. Thus flexible modes and uncertainties can be suppressed simultaneously. The convergence of overall closed-loop system states is proved via Lyapunov analysis. Numerical simulations show the effectiveness and advantages of the proposed control strategy.
A multi-bump strategy combines automatic optimization and flow control and is employed in airframe/intake–exhaust integration design of flying wing. The numerical simulation method of
This paper presents the basic results of the morphing wing planform optimization of an experimental unmanned air vehicle for minimum drag at steady level flight. The aerodynamic design tool that consists of the three-dimensional panel method, two-dimensional boundary layer solution and generalized reduced gradient method-based optimization is appropriate for fixed wing and morphing wing conceptual and preliminary design. The morphing concept is implemented into the solution with the geometric constraints of the wing planform and the airfoil shape design variables. The drag that is created by other components of the aircraft is calculated according to empirical formulas. Wing drag and aircraft drag comparisons between baseline wing (BASE), optimum fixed wing and morphing wing are discussed with the obtained planform and airfoil shapes.
A numerical survey coupled with six degree-of-freedom flight simulation have been undertaken to study the fuel tank separation trajectory, released from a trainer airplane. Two different spanwise release points for the tank, near and farther from the fuselage under the starboard wing with full and empty fuel were considered. The studies were performed at two free stream Mach numbers of 0.23 and 0.42 at zero angle of attack. Dynamic unstructured tetrahedral mesh approach combined with spring-based smoothing and local remeshing was applied with an implicit, second-order upwind accurate Euler solver. A six degree-of-freedom routine using a fourth-order multi-point time integration scheme was coupled with the flow solver to update the payload trajectory information at each time step. According to the results, the payload installed farther from the fuselage falls down with a higher forward velocity than that located closer, once released from the wing. The spanwise installation point was also found to have a strong impact on the pitch attitude of the released payload. The payload weight has been shown to play a vital role in longitudinal-lateral coupling behavior and the associated moments on the released payload.
The inertial navigation system is one of the most important and common methods of navigation. In this system, accelerometers and gyroscopes are used to measure linear accelerations and angular velocities, respectively. Accelerometers have simpler manufacture techniques, lower cost, and smaller volume and weight in comparison with gyroscopes. Therefore, in some application of navigation systems, non-gyro inertial navigation systems based on accelerometers are used. In this paper, an asymmetric structure of six accelerometers is proposed. Then dynamic relations of this structure are extracted. This structure and its relations can determine linear accelerations and angular velocities, completely. Moreover, the algorithm of inertial navigation in earth centered earth fixed (ECEF) frame is suggested. Error analysis as of the most important issues in inertial navigation is discussed. Thus, bias, misalignment, sensitivity, and noise of accelerometers are modeled appropriately. In addition, a symmetric structure of accelerometers is proposed and its equations are derived. Finally, the designed system, error model of accelerometers, and algorithm of inertial navigation in ECEF frame are simulated. The results of simulation show that the designed system has suitable accuracy and applications for short time navigation. Furthermore, results confirm that the proposed asymmetric structure requires less accelerometer in comparison with symmetric structure.
An active fault-tolerant satellite attitude control scheme based on fault effect classification is presented at the occurrence of faults associated with torques. In this paper, the flexibility and practicability of the fault-tolerant scheme are top priorities. Faults are modeled and divided into additive and multiplicative ones in order to estimate and deal with them specifically and exactly. The additive faults, including additive part of flywheel faults and other uncertain fault torques, are estimated by additive fault estimator and compensated on the basis of nominal controller, whereas the multiplicative faults, denoting torque gain parameter faults of flywheels, are estimated by multiplicative fault estimator and the estimated fault parameters are used for dynamic torque command distribution of flywheels. The final simulation examples and performance comparison of three fault-tolerant schemes show that the proposed scheme based on fault effect classification is an effective, flexible and saving-energy fault-tolerant satellite attitude control scheme. It possesses an engineering value for improving reliability and prolonging on-orbit working lifetime of satellites.
Landing footprint is critical to generate a feasible trajectory onboard for entry vehicles. In this paper, a new landing footprint calculation based on geometry-predicted trajectory is proposed. First, the lateral motion states, turning radius and angle, are analytically solved based on a planned drag acceleration vs. energy (D–E) profile. Through some simple coordinate and geometrical triangle transformation, all the trajectory states are extracted. Second, by calculating the maximum longitudinal range point, the furthest reachable boundary for planned D–E profile is obtained with proposed geometry-predicted trajectory. Finally, the landing footprint can be generated by repeatedly computing the reachable boundaries for all D–E profiles in the entry corridor. Simulation results with Common Aero Vehicle (CAV) model validate the accuracy of the geometry-predicted trajectory. Furthermore, detailed comparison of footprint generation results between the traditional and the proposed are presented at last.
Transient numerical simulations are carried out to study missile motion in a vertical launch system and to estimate the effect of missile exhaust in the adjoining launch structure. Three-dimensional Navier–Stokes equations along with k– turbulence model and species transport equations are solved using commercial computational fluid dynamics software. Dynamic grid movement is adopted and one degree of freedom trajectory equations are integrated with the computational fluid dynamic solver to obtain the instantaneous position of the missile. Multi-zone grid generation approach with sliding interface method through layering technique is adopted to address the changing boundary problem. The computational methodology is applied to study the missile motion in a scale-down test configuration as well as in the flight condition. The computations capture all essential flow features of test and flight conditions in active cell as well as in adjacent cells. Parametric studies are conducted to study the effect geometrical features and measurement uncertainty in the input data. Computed pressures in the adjacent cells in the launch system match better (~12%) with the experimental and flight results compared to distant cells.
Both the artificial potential field method and direct method for the optimal control problem have shortcomings in terms of effectiveness and computational complexity for the trajectory-planning problem. This paper proposes an integrated algorithm combining the artificial potential field method and direct method for planning in a complex obstacle-rich environment. More realistic unmanned aerial vehicle dynamics equations, which are rarely applied in the traditional artificial potential field method, are considered in this paper. Furthermore, an additional control force is introduced to transcribe the artificial potential field model into an optimal control problem, and the equality/inequality constraints on the description of the shape of the obstacles are substituted by the repulsive force originating from all the obstacles. The Legendre pseudospectral method and virtual motion camouflage are both utilized to solve the modified optimal control problem for comparison purposes. The algorithm presented in this paper improves the performance of solving the trajectory-planning problem in an obstacle-rich environment. In particular, the algorithm is suitable for addressing some conditions that cannot be considered by the traditional artificial potential field method or direct method individually, such as local extreme value points and a large numbers of constraints. Two simulation examples, a single cube-shaped obstacle and a different-shaped obstacle-rich environment, are solved to demonstrate the capabilities and performance of the proposed algorithm.
Cable-driven parallel robot is a special kind of robot, which is actuated by cables. It is already applied in the low speed wind tunnel to get aerodynamic measurement of aircraft model, and the aircraft pose could be adjusted by changing the cable length. Whether it can be used in hypersonic wind tunnel still needs further discussion. This paper presents the dynamics and aerodynamics analysis of a large-scale model supported by 6-DOF cable-driven parallel robot to investigate the feasibility of this special kind of suspension system in hypersonic wind tunnel. The description of this setup with a X-51A-like model is given, and then based on the system dynamic equations, aerodynamic force and stiffness matrix are derived. In the simulation, properties of dynamics and aerodynamics are mainly concerned. A typical shock tunnel with flow duration of about 100 milliseconds is taken as an example, and results show that the system is stable enough to meet the fundamental static wind tunnel test. From the cable tension variation under impact load and the sensitivity analysis, it is likely accessible to derive the aerodynamic forces. Compared with the sting suspension method, cable-driven parallel robot has the priority of higher inherent frequency and more flexible degrees. The interference to the flow field induced by cables is also preliminarily proved to be small by the CFD simulation, which can be acceptable and corrected. Researches conducted show the feasibility of cable-driven parallel robot’s application in hypersonic wind tunnel.
In a preparatory study conducted prior to the development of an active space debris removal system, a method for selecting target debris based on information such as the cumulative collision probability, the operational condition of objects, and their sizes and launch dates was developed for use in the protection of four Korea Multi-Purpose Satellite constellation satellites. This method can be used to select candidate removal targets. Two-line element data are used to identify threatening objects with high cumulative collision probability. Using information in the Satellite Catalog database, objects smaller than a certain size or objects that are currently operational were excluded from the selection range. The results of an analysis of the cumulative collision probability, object size, object type, and primary mission information showed that the COSMOS 1328, COSMOS 1862, COSMOS 375, and COSMOS 1606 satellites were suitable targets for an active debris removal mission.
Mesh reflectors are uncertain structures because of the existing errors of dimension and material in the procedure of design and manufacture. These uncertainties have significant impacts on the mechanical and electrical properties, which must be considered during the design phase. Three directly related factors of cable uncertainties in mesh reflectors are considered in this paper, including the initial length, cross-sectional area, and elastic modulus. The analytical relationship between the cable uncertainties and the surface accuracy of mesh reflectors is deduced by interval analysis, and an interval force density method is thus proposed. First, this method is used to analyze the influence of the cable uncertainties on the surface accuracy. Then it is applied into the form-finding optimization of uncertain mesh reflectors to minimize the influence of cable uncertainties on the surface accuracy. Three kinds of cable nets of mesh reflectors are illustrated to analyze the influence of the cable uncertainties on the surface accuracy, and the mesh reflectors with high surface accuracy are obtained by the proposed method. Finally, the influences of both the design values and deviation amplitudes of cable uncertainties on the surface accuracy are revealed.
The paper presents a concept of tandem-wing configuration aircraft that was a Warsaw University of Technology proposal for the personal air transport system. The project was developed at Warsaw University of Technology, within the PPLANE project (FP7 – Personal Plane: Assessment and Validation of Pioneering Concepts for Personal Air Transport Systems). First, analysis of the general concept, advantages and disadvantages of tandem-wing configuration, and possible application as a personal air transport system vehicle are presented. Next, aerodynamic design is analyzed and dynamic stability is tested. All numerical analyses were made by use of the well-tested professional software for aerodynamic (MGAERO) and stability (SDSA) analyses.
This paper investigates the unsteady tip flow characteristics and their effects on the aeroelastic stability of a linear oscillating compressor cascade experimentally. Two test cases with different tip clearance configurations were tested, including a suction side squealer configuration and a control test. The unsteady pressure on the blade surface was measured at three blade vibration frequencies, and then was utilized to establish the aero-damping of the cascade. The results show that the impact of suction side squealer tip clearance on the tip clearance flow mainly exists at the region where it rolls up, and the unsteady flow induced by the vibration blade has no effect on the time-averaged result. In tip area of the blade, the leakage vortex plays a dominant role in determining the distribution of the first harmonic of unsteady static pressure. Compared with the baseline test case, the suction side squealer tip geometry exhibits better aerodynamic stability at high oscillating frequency. The aerodynamic response (phase angle) of the leakage flow lags behind the vibration of blade, and its streamwise variation implies strongly associated with the development of the tip leakage vortex. The spanwise three-dimensional unsteady characteristic of suction side squealer tip geometry is more obvious than that of the baseline test case.
Dynamic combustion characteristics of a rectangular scramjet combustor with single-side expansion were studied experimentally and numerically. Experiments were implemented with an isolator entrance Mach number of 3.46, and an air stagnation temperature of 1430 K. Ethylene was utilized to fuel the combustor over an equivalence ratio range of 0.20 < < 0.63. Results indicated that the combustion modes varied from different equivalence ratios. For an intermediate = 0.375, an intermittent dynamic combustion occurred. During the dynamic process, the flame sometimes stabilized in the jet wake of the top cavity, and at other time it oscillated between dual parallel cavities. The pseudo-shock train traveled periodically along the length of the combustor, and the penetration depths of the two injectors exchanged. Quantitative analysis illustrated that the average frequency of unsteady combustion was approximately 200 Hz. The reason for the occurrence of the self-sustained dynamic process was related to the interactions between the shock-induced separated region and heat release.
This paper presents a new active disturbance rejection controller to solve the altitude and attitude control problem for a quadrotor unmanned aerial vehicle. The proposed method requires only the output information of the system. Using the pitch subsystem as an example, the proposed controller is designed by using dynamic surface control strategy incorporated with tracking differentiator, and extended state observer, which is used to estimate the uncertain disturbance. The estimate states of extended state observer are used to design the dynamic surface control law for altitude and attitude tracking problem of the quadrotor unmanned aerial vehicle. The stability analysis proves that a sufficient condition of the asymptotic stability of the extended state observer is achieved, the asymptotic stability of the closed-loop system can be guaranteed, and the tracking feedback error can made arbitrarily small by adjusting the controller parameters. Several simulation results are presented to corroborate that the proposed control method has better effectiveness and robustness.
This paper presents the results of an investigation of the ballistic limits and failure modes of AA2024-T351 sheets impacted with cubical projectiles. The experiment/test setup was based on EASA CS-25 regulations for fuel tank access covers. The effect of cube orientation on the ballistic limit and failure modes was considered in detail. A 25% variation in ballistic limit was observed with the lowest ballistic limit (202 m/s) observed for the cubical projectile edge impacted on the target. In the cube face impacts, the ballistic limit was higher (223 m/s), and the highest ballistic limit (254 m/s) was observed for the corner impact. The observed differences in the ballistic limit were due to differences in failure mechanism, which resulted in different localised deformations near the projectile impact point, but also led to differences in global dishing deformation.
A pulse detonation engine with the inner diameter of 50 mm and total length of 1500 mm was designed. Gasoline was used as the fuel and air as the oxidizer. The Shchelkin spiral was used as the deflagration to detonation transition accelerator. The direct-connected test of the pulse detonation engine was conducted to find out the detonation initiation and propulsion performance at several different operating frequencies. The experimental results indicated that detonation waves were fully initiated in the pulse detonation engine at the operating frequency range of 1–35 Hz. As the operating frequency increased, the pulse detonation engine average thrust increased nearly linearly, while the optimum equivalence ratio decreased gradually. The volume-specific impulse and mixture-specific impulse had the similar increasing trend at the increased operating frequency. As the operating frequency increased from 0 to 25 Hz, the fuel-specific impulse increased dramatically while the specific fuel consumption decreased quickly. The fuel-specific impulse and specific fuel consumption changed slightly when the operating frequency exceeded 25 Hz. In addition, two groups of computational value of mixture-specific impulse and fuel-specific impulse were obtained by Winternberger model and Yan model considering the two-phase effect. Compared with the experimental results, similar variation trends relative to the operating frequency were achieved by the two computational models. However, the computational values of mixture-specific impulse and fuel-specific impulse by Winternberger model were much higher than that of experimental values. When two-phase effect was considered, the computational values of mixture-specific impulse and fuel-specific impulse were close to that of experimental values.
Predicting boundary layer transition accurately is important to thermal protection and drag reduction of flight vehicles. Up to now, there has been many transition prediction methods. However, most of those methods need boundary layer parameters, which are difficult to obtain in massively parallel execution since some parameters are nonlocal variables, thus greatly limiting the application of those methods. A grid-reorder method is developed to obtain the boundary layer parameters, which is suitable for parallel computing in this paper. With the grid-reorder method the wall normal grid cells can be easily found, and two criteria are used to determine the boundary layer edge in the wall normal direction, then the boundary layer parameters such as boundary layer thickness, boundary layer momentum thickness, boundary layer edge velocity, cross-flow velocity, and so on, can be obtained accurately and efficiently. The method has been coupled to three transition prediction methods, the -Re model, the k-- model, and the transition correlations, to validate its effectiveness. For the -Re model, the cross-flow velocity is obtained with the grid-reorder method, then a cross-flow intermittency factor is developed and introduced into the model, and the inclined prolate spheroid case is used to test the performance of the model. For the k-- model, the grid-reorder method is applied to obtain the boundary layer edge velocity and the inflection point velocity which are of vital importance to form the second-mode timescale for hypersonic transition prediction. For the transition correlations, Re/Me is obtained effectively with the grid-reorder method. The X-51 forebody is selected to test the effectiveness of Re/Me for complex geometries and the results show a good correspondence with the experiment results. The successful application in three transition prediction methods demonstrates that the grid-reorder method has an excellent performance in obtaining the boundary layer parameters and can broaden the application of the existing transition prediction method in engineering.
The existing research on air traffic complexity ignores the effects of air traffic situation structure and, thus, cannot reflect the heterogeneous traffic density distribution in airspace. In this study, the structure of air traffic situation was characterized using the idea of community structure in complex networks. An aircraft cluster model was built, and an aircraft cluster discovery method based on depth-first traversal was proposed. The aircraft cluster division effect was comprehensively represented by cluster performance indices, including cohesion and stability. The routinely recorded radar data in two air traffic control sectors were collected to assess the cluster division results. Through statistics, the threshold intervals with 95% of best performance are 40–60 km and 20–50 km for the two sectors, respectively. The value 40 km was selected to further statistically characterize the aircraft clusters. Compared with K-means clustering, the proposed method does not require the predefined number of clusters and has high stability, which confirms its feasibility into cluster division in dynamic air traffic situation. The structural characteristics of aircraft clusters, including the average intra-cluster horizontal distance, number of clusters, and size and life cycle of clusters, were statistically analyzed. Comparison of cluster structures with the commonly used dynamic density index shows that in air traffic situation with relatively large number big size of clusters, the aircraft trajectory changes more frequently. Structural characterization of aircraft clusters is able to portray the nonuniformity of traffic density distribution, and contributes to comprehensive description of air traffic situation, thus providing a new prospect for analysis of air traffic complexity. Moreover, aircraft cluster division contributes to auto-identification of hot-spots on radar screen, and efficiently eliminates the workload imposed on controllers during judgment of these congestion hot-spots, thereby improving the air traffic operation efficiency.
Mass capture ratio of a hypersonic air intake is one of the most important performance parameters. However, no a priori estimate of its value exists for use in initial design exercise of a hypersonic vehicle. In the present work, an air intake of a non-axisymmetric scramjet engine, designed using stream thrust methodology, is studied using computational fluid dynamic techniques. A large amount of air mass flow rate is observed to spill from the sides, which is not accounted for in the initial design phase. In absence of even an approximate estimate of this spillage, computational fluid dynamic studies become the only available tool to evaluate the mass capture ratio. Simulations are also carried out with a side wall at the intake to stop spillage. Although mass capture ratio and static pressure at combustor entry improve, deterioration in other flow parameters such as static temperature, Mach number and total pressure is observed.
In this article, the flow field around NACA0024 airfoil with step at lower and upper surfaces is experimentally investigated. For this purpose, particle image velocimetry technique based on the instantaneous flow structures is used to investigate the flow field around the airfoil at different times. All the experimental measurements in current study are conducted at very low Reynolds number condition based on the chord of the airfoil (Re=2000) and at angles of attack at 0° and 5° where the flow around airfoils is separated. The differences between vortical structures, mean streamlines, sizes of the wake regions, and vortex shedding of the stepped airfoils compared to unmodified airfoil are observed. The results disclose that using step in airfoil leads to a decrease in the Strouhal number. In addition, the formation of vortices in wake region and their positions at different times are discussed.
This paper investigates the vibration patterns, i.e. rigid motions of shaft and elastic deformation of support structures, of fan rotor system in aero-engine, which differs from traditional flexible rotor systems, and together with its shaft transverse motions due to unbalanced mass. The fan rotor system commonly is composed of one rigid shaft and two flexible support structures (such as squirrel cages), which is effective to decrease the critical speeds avoiding serious shaft vibration due to unbalance. Scaled test rig for realistic fan rotor system is set up according to similarity principles, governing differential equations of which are deduced by means of Lagrangian approach with four degrees of freedom. In contrast to modeling a traditional flexible rotor system, the system stiffness is not determined by the shaft but the two flexible support structures. The rigid shaft only contributes to the inertial items of the governing equations. Parameter values of dynamic model are identified from measurements on the scaled test rig, the modal shapes and the modal energy distributions are calculated. These modal characteristics of the fan rotor system are quite different from those of a traditional flexible rotor system whose stiffness mainly contributed by its elastic shaft even the system values are consistent. The obtained modal characteristics are compared and confirmed by using the simulation results of a corresponding finite element model, in which shaft is built by rotating beam elements and its flexible structures are built by equivalent spring elements. Campbell diagrams of the fan rotor system are used to illustrate the gyroscopic effect with the increasing speeds. And then the unbalance responses are calculated through the deduced analytical formula rapidly and comparisons, including the response spectrum and orbits, the amplitude and phase frequency response curves, and operating deflection shapes, are carried out in the sub- and super-critical range.
This study reports multi-algorithm parallel integrated decision-making for liquid-propellant rocket engine online health condition monitoring to improve reliability and safety, especially for next-generation reusable engines. Fusing multi-algorithm detection information to judge liquid-propellant rocket engine condition is multi-algorithm parallel integrated decision-making main task, and multi-algorithm judgment problem is its central issue; i.e. how to make a global judgment from judgment results of different fault detection methods. Considering opportune fault detection, adequate rocket engine information exploitation, and reliable condition judging, the multi-algorithm parallel integrated decision-making framework for problem definition is presented along with a multi-algorithm parallel integrated decision-making judgment model. For more reliable, efficient global judgment, a method based on the Bayes’ risk function integrating multi-algorithm prior information is adopted. The proposed approach is validated with liquid-propellant rocket engine ground testing data. The results show that the multi-algorithm parallel integrated decision-making judgment model gives very effective and reliable performance relative to the voting method, successfully solving multi-algorithm judgment problems and meeting practical engineering needs.
Like albatross, unmanned aerial vehicles can significantly make use of wind gradient to extract energy by the flight technique named dynamic soaring. The research aims to develop a general optimization method to compute all the possible patterns of dynamic soaring with a small unmanned aerial vehicle. A direct collocation approach based on the Runge-Kutta integrator is proposed to solve the trajectory optimization problem for dynamic soaring. The optimal dynamic soaring trajectories are classified into two patterns: closed trajectory pattern and travelling trajectory pattern by applying terminal constraints of zero horizontal displacement and a certain travelling direction, respectively. Using different terminal constrains for heading angle and initial guesses in the optimization process, the former pattern can be divided into two subtypes: O-shaped and 8-shaped trajectories, while the latter one is divided into C-shaped, α-shaped, S-shaped and -shaped trajectories. The characteristics of these patterns and the correlation among patterns are analyzed and discussed.
The tip leakage vortex breakdown occurs under a few conditions in modern turbines, which leads to extra vortex breakdown losses, but the mechanisms of vortex breakdown and its influencing factors still remain unclear. This paper is a continuation of the previous effort and focuses on the effect of blade rotation on the leakage vortex dynamics in an unshrouded turbine. The analyses on leakage vortex breakdown characteristics are first shown, and then the isolating effects of relative casing motion and Coriolis and centrifugal forces on leakage vortex breakdown and loss are investigated. Based on these, the overall effects of blade rotation on leakage vortex breakdown characteristics are examined. Results indicate that the scraping effects of the casing endwall have a great influence on the blade tip leakage vortex breakdown and loss. However, the effect of Coriolis and centrifugal forces is relatively small. Although the mismatch of both velocity components of the leakage flow and the main flow becomes slightly small with the casing motion, the blade tip mixing loss per unit leakage flow increases due to the fact that the leakage vortex breaks down in advance.
A ballistic error propagation algorithm for glide trajectories of a hypersonic glide vehicle is originally proposed based on the perturbation theory. Perturbation equations that treat perturbations as external inputs and flight state deviations as state variables are established. Based on the reasonable simplification assumptions, the analytic expression of the state transition matrix is deduced and thus the ballistic error propagation model is established. A transposed coordinate frame is introduced to simplify the development of the perturbation equations and the error propagation model. By employing the gravity anomaly as the single perturbation factor, the proposed algorithm is validated and verified by numerical experiments. When compared with the traditional method, the proposed method takes only just a quarter computational costs and restrains the method errors beneath 10% of the total terminal deviations. It is an effort that initially focuses on the error propagation mechanism of the glide trajectory and the proposed model has sufficient precision for the analysis of modeling deviations, thus laying the foundation of correcting the modeling deviations and enhancing the accuracy of vehicle’s flight states.
The history of UAVs is relatively long and many such vehicles are in service for different tasks. They can be used even in environments inhospitable for humans, e.g. because of extreme temperature. Moreover, they can perform a task that is difficult or impossible for a manned aircraft because of its size and usually relatively high airspeed. The photogrammetric tasks belong to this group, especially if we need to take high-resolution pictures during low level flight. The advantages of a small UAV for such mission are more evident if we want to investigate the natural environment, where the wild animals are. The paper presents the small UAV designed for a special task, which is counting of the penguins in Antarctica. Inhabited area, extreme weather conditions, the fearfulness of penguins and the goal of the mission put up certain requirements for the UAV. It had to be a reliable, stable platform, which is able to carry photogrammetric equipment and to perform precise flight to cover all investigated areas. The presented UAV was used on such missions in Antarctica in 2014 and 2015. All mentioned tasks were successfully accomplished.
This paper investigates the distributed finite-time configuration containment control problem for satellite formation with multiple leader satellites under directed communication topology. We consider that only a portion of follower satellites can receive leaders’ information and unknown perturbations and model uncertainties exist in the dynamics models of satellites. By defining the relative configuration error functions and selecting suitable nonsingular terminal sliding mode variables, a fully distributed finite-time configuration containment control scheme is proposed using the matrix properties of graph theory. The Lyapunov method is used to demonstrate the finite-time convergence property of the closed-loop systems. Numerical examples and comparisons with other methods are provided to show the effectiveness and the performance of the proposed control strategy.
In this work, the models and control strategy of the Electric Servo-torque System(ESS) which is used as an experiment rig for conducting dynamic performance and stability tests of aerial vehicle control surface actuation systems are presented. The detailed dynamics of the load motor and loaded flight actuator’s rotating movement in the ESS are analyzed, leading to an integrated load torque synchronization system. The kinematic dynamics of the loaded control surface driving actuator is an important consideration to estimate the trend of torque variation and to improve the performance of the load system. The load control method is expressed in terms of a multi-loop torque control law, which uses feedback and feedforward loops to meet system design requirements. Numerical examples together with experimental results are included to illustrate the effectiveness of the proposed models and control parameters. This brief addressed a specific utilization of the loaded actuator’s dynamics, revealing that it can reduce both the phase lag and the amplitude gain of the load torque in the Electric Servo-torque System.
This study aims at establishing a three-dimensional numerical model, compressor aerodynamic performance analysis model, to simulate the impact of complicated distorted flow on multistage axial flow compressor based on the body force model. The model solves the compressible three-dimensional Euler equations, which are modified to include source terms representing the effect of the blade rows. In this study, the association between blade source terms and entry Mach number together with attack angle could be established with the deviation angle model and loss model. In this paper, compressor aerodynamic performance analysis model is used to evaluate the effect of inlet circumferential total pressure distortion and swirl distortion on a five-stage high-pressure compressor. Calculated operating maps for compressor agree well with the experimental results. Meanwhile, the traveling process of inlet distortions in the multistage compressor is correctly revealed. The wide application prospect of the model can be seen in the area of inlet distortion problems.
In this paper, a new approach based on block-oriented nonlinear models for the modeling and identification of aircraft nonlinear dynamics is proposed. Some of the block-oriented nonlinear models are regarded as flexible structures, which are suitable for the identification of widely applicable dynamic systems. These models are able to approximate a wide range of system dynamics. In general, aircraft flight dynamics is considered as a nonlinear and coupled system whose dynamics—in addition to pilot control inputs—depend on the flight conditions such as Mach number and altitude, which cause the aircraft dynamics to have various operational points. In this study, three types of block-oriented models, namely the Hammerstein, Wiener, and Hammerstein–Wiener models with different nonlinear functions, have been used and compared in order to identify and model the aircraft nonlinear dynamics. These models have been employed in three forms of single-input single-output, multi-input multi-output, and multi-input single-output (MISO); of which, multi-input single-output has been recognized to have fewer errors in aircraft nonlinear dynamics identification. Thus, it has been demonstrated that six separate multi-input single-output models (with three inputs and one output), which have been trained with experimental flight test data, can model the coupled nonlinear six-degree-of-freedom dynamics of a highly maneuverable aircraft.
Aviation industries are vulnerable to the energy crisis and simultaneously posed environmental concerns. Proposed engine technology advancements could reduce the environmental impact and energy consumption. Substituting the source of jet fuel from fossil-based fuel to biomass-based fuel will help reduce emissions and minimize the energy crisis. The present paper addresses the analysis of aircraft engine performance in terms of thrust, fuel flow and specific fuel consumption at different mixing ratio percentages (20%, 40%, 50%, 60% and 80%) of alternative biofuel blends already used in flight test (Algae biofuel, Camelina biofuel and Jatropha biofuel) at different flight conditions. In-house computer software codes, PYTHIA and TURBOMATCH, were used for the analysis and modeling of a three-shaft high-bypass-ratio engine which is similar to RB211-524. The engine model was verified and validated with open literature found in the test program of bio-synthetic paraffinic kerosene in commercial aircraft. The results indicated that lower heating value had a significant influence on thrust, fuel flow and specific fuel consumption at every flight condition and at all mixing ratio percentages. Wide lower heating value differences between two fuels give a large variation on the engine performances. Blended Kerosene–Jatropha biofuel and Kerosene–Camelina biofuel showed an improvement on gross thrust, net thrust, reduction of fuel flow and specific fuel consumption at every mixing ratio percentage and at different flight conditions. Moreover, the pure alternative of Jatropha biofuel and Camelina biofuel gave much better engine performances. This was not the case for the Kerosene–Algae blended biofuel. This study is a crucial step in understanding the influence of different blended alternative biofuels on the performance of aircraft engines.
Numerical simulations were performed to study the influence of the inside diameter of the pulse separation device port on the flow features and the local heat transfer characteristics in a dual pulse solid rocket motor. A lower–upper symmetric Gauss–Seidel implicit dual time-stepping method was applied to address the problem of unsteady flow. A high-resolution upwind scheme (AUSMPW+) and Menter’s shear stress transport turbulence model were employed to solve the Reynolds-averaged Navier–Stokes equations. The conjugate heat transfer strategy was realized by enforcing a common temperature and heat flux at the fluid–solid interface. After validating the accuracy and reliability of the numerical algorithm by comparison with experimental cases, the internal flow of a dual pulse solid rocket motor was simulated. The results show that the magnitude of the velocity, the wall shear stress, and the turbulent kinetic energy downstream of the pulse separation device port decrease with increasing pulse separation device port diameter. The local heat transfer coefficient increases sharply downstream of the pulse separation device port, reaching a maximum within 1–2 diameters downstream of the pulse separation device port, before relaxing back to the fully developed pipe flow value. The peak value of the local heat transfer coefficient reduces as the pulse separation device port diameter increases. Meanwhile with an increasing pulse separation device port diameter, the position of the peak local heat transfer coefficient moves upstream to the head of the first pulse chamber, and appears upstream of the position of the reattachment point by an average of about 28.6%.
Electro-hydrostatic actuator is generally regarded as the preferred solution for more electrical aircraft actuation systems. It is of importance to optimize the weight, efficiency and other key design parameters, during the preliminary design phase. This paper describes a multi-objective optimization preliminary design method of the electro-hydrostatic actuator with the objectives of optimizing the weight and efficiency. Models are developed to predict the weight and efficiency of the electro-hydrostatic actuator from the requirements of the control surface. The models of weight prediction are achieved by using scaling laws with collected data, and the efficiency is calculated by the static energy loss model. The multi-objective optimization approach is used to find the Pareto-front of objectives and relevant design parameters. The proposed approach is able to explore the influence of the level length of linkage, displacement of pump and torque constant of motor on the weight and efficiency of the electro-hydrostatic actuator, find the Pareto-front designs in the defined parameter space and satisfy all relevant constraints. Using an electro-hydrostatic actuator for control surface as a test case, the proposed methodology is demonstrated by comparing three different conditions. It is also envisaged that the proposed prediction models and multi-objective optimization preliminary design method can be applied to other components and systems.
An adaptive fuzzy second-order terminal sliding mode control scheme is developed to solve the problem of rigid spacecraft attitude maneuver in which the external disturbances and uncertain inertia parameters as well as actuator faults are explicitly taken into account. The proposed controller incorporates a second-order terminal sliding mode to obtain finite time convergent performance as well as chattering elimination, meanwhile the fuzzy logical system is implemented to estimate the system uncertainties of the rigid spacecraft. Lyapunov stability analysis shows that the designed control algorithm guarantees the practical finite time stability of the attitude maneuver errors with great robustness to system uncertainties and actuator faults. Numerical simulations are also presented that not only highlights closed-loop performance benefits from the proposed control algorithm, but also illustrates its effectiveness in the presence of disturbances and uncertainties when compared with classic finite time sliding mode control schemes for spacecraft attitude maneuver control.
This paper is concerned with active vibration control of a flexible piezoelectric cantilever plate using a nonlinear radial basis neural network sliding mode control (RBFNN-SMC) algorithm and laser displacement measurement. In order to decouple the low-frequency vibration signals of the bending and torsional modes on measurement, two laser displacement sensors are used. The decoupling method is provided. A hyperbolic tangent function is used instead of the sign function, and the chattering phenomenon is alleviated. Also, the RBFNN is utilized to adjust the switching control gain adaptively to balance the chattering phenomenon and the control effect. The controllers for bending and torsional modes are designed independently. Experimental setup of the flexible piezoelectric cantilever plate with two laser displacement sensors is constructed. Experiments on vibration measurement and control are conducted by using the decoupling method and the designed controller, compared with the classical proportional and derivative (PD) control algorithm. The experimental results demonstrate that the proposed method can decouple the low-frequency bending and torsional vibration signals on measurement. Furthermore, the designed nonlinear RBFNN-SMC can suppress both the bending and torsional vibrations more quickly than the traditional linear PD controller, especially for the small amplitude residual vibration.
Investigations have been made adopting experiments and computations on an ogive-nosed slender body at different angles of attack and Reynolds number of 29,000 based on the model base diameter diameter. The results indicated an increase in the side force at large angles of attack, which is mainly due to the presence of asymmetric vortices in the leeward of the body. The inclusion of a rectangular cross-sectioned ring in the initial portion of the body reduced the side force at higher angles of attack. However, significant side force was experienced at lower angles of attack (30° < α < 40°). Use of a ring of 3% height was found to be suitable for reducing the side force at a higher angle of attack. From the results obtained it was observed that a ring if placed at a different axial location alters the flow field and changes the side force at higher angles of attack. Further studies indicated that placing of rings pair at an axial location of 3.5 and 4.5 times the base diameter reduced the side force to a very low value at all the angles of attack for the present shape of body and flow conditions.
In the active flight phase, the spacecraft suffers the rigorous external vibration environment which comes from the engine at the bottom of rocket. In course of the ground test, the shaker simulates such excitation through the base movement, and the modal parameters identification method which employs the vibration test data is to the more extent in line with the operating state. In this paper, a specified method is applied for parameter estimation of lumped-mass mechanical system in the vibration test. The modal parameters under the fixed and free–free constrained state can be identified at the same time by this technique. In addition, when the analytical model which includes the test article and the shaking table is adopted, the relations between the response transfer function (RTF) and frequency response function (FRF) are clarified, and some misconceptions in the engineering practice are cleared. Because of the uncertain of the connection impedance between the test article and the shaking table, only the partial parameters of the fixed mode can be identified, such as natural frequency, modal damping ratio and modal shape. As the interface force transfer function (IFTF) can eliminate the influence of the connection impedance and the fixed impedance matrix of the test piece is transformed into the free impedance matrix, the identified parameters of the free mode are complete. Furthermore, the influence of the constraint degree of freedom (DOF) on the parameter identification of the IFTF is discussed preliminarily. Finally, the numerical simulation and the vibration test of discrete structure verify the validity of the proposed method.
Artificial neural networks are an established technique for constructing non-linear models of multi-input–multi-output systems based on sets of observations. In terms of aerospace vehicle modeling, however, these are currently restricted to either unmanned applications or simulations, despite the fact that large amounts of flight data are typically recorded and kept for reasons of safety and maintenance. In this paper, a methodology for constructing practical models of aerospace vehicles based on available flight data recordings from the vehicles’ operational use is proposed and applied on the Jetstream G-NFLA aircraft. This includes a data analysis procedure to assess the suitability of the available flight databases and a neural network-based approach for modeling. In this context, a database of recorded landings of the Jetstream G-NFLA, normally kept as part of a routine maintenance procedure, is used to form training datasets for two separate applications. A neural network-based longitudinal dynamic model and gust identification system are constructed and tested against real flight data. Results indicate that in both cases, the resulting models’ predictions achieve a level of accuracy that allows them to be used as a basis for practical real-world applications.
An experimental study was conducted in the present study to investigate the effects of upstream probe disturbance on the compressor cascade performance. The experiments were carried out using a plane cascade test facility where the aerodynamic coupling mechanisms between the inner probe and compressor were particularly evaluated in terms of the complex reversed pressure gradient. Moreover, the influence of probe install position near the cascade leading edge on the downstream flow characteristic was evaluated in detail. Results show that the presence of probe at cascade leading edge can reduce the total pressure recovery coefficient, raise the wake loss at cascade outlet and deteriorate the periodicity of conventional cascade flow field distribution. An optimum circumferential installing position at the upstream of the cascade is found with a minimum disturbance, which is not affected by the varying Mach numbers significantly. The probe installed at the intermediate of flow passage has a smaller disturbance on the downstream cascade performance than that installed in front of cascade leading edge.
Curvilinear stiffener concept has been introduced to aircraft panel structures most recently for possible further weight reduction during optimization design. However, due to enlarged design space and high design complexity, more computational time is needed to optimize curvilinearly stiffened panels. Considering the requirement for both lighter structure and less design time, optimization designs of a unitized panel with stiffeners in three different formats, i.e. curvilinear, oblique, and evenly distributed straight, are conducted under various loading conditions to figure out which stiffener format should be selected in a real design environment. Four single loading cases with different compression-shear ratios and a set of multiple load cases are considered. Evolution strategy with covariance matrix adaption is combined with sequential quadratic programming to seek out the optimum structure with minimum mass taking into account buckling and strength constraints. Comparative analysis of the optimization results indicates that evenly distributed straight format is suitable for compression-dominant loading conditions whereas curvilinear format become superior in case that shear loading is considerable or multiple biaxial compression and shear loads are applied. Besides, curvilinear format adds more design flexibilities to the stiffeners, which enables them to play a more important role in the optimized structures. Furthermore, evolution strategy with covariance matrix adaption is found to be more efficient than particle swarm optimization for this optimization problem. This study can provide a useful guidance for future optimization design of aircraft structures.
To improve the accuracy of the attitude sensor micro electro mechanical system gyroscope in low cost satellite, a nonlinear moving horizon estimation algorithm based on micro-electro mechanical system gyroscope/three-axis magnetometer is proposed in this paper. First, a quaternion micro-electro mechanical system gyroscope/three-axis magnetometer-integrated attitude estimation model is established so as to improve the accuracy of micro-electro mechanical system gyroscope. Thanks to the concealment and autonomy, these two low cost sensors have great potential in the military area. Second, taking advantage of optimal problem in coping with constraints, a real time moving horizon estimation algorithm with equality constraint is designed to deal with the disability of solving quaternion normalization analytically in the frame work of Kalman. In this algorithm, Gauss–Newton iterative method is used to obtain the optimal state estimation in the "window". Meanwhile, strong tracking filter of arrival cost is designed outside of the "window" to enhance system robustness for that three-axis magnetometer is vulnerable to external interference. Third, the proposed MHE is applied in the micro-electro mechanical system gyroscope/three-axis magnetometer attitude estimation system. The simulation results show that the method has higher accuracy and robustness.
Elliptic jet mixing influenced by triangular tabs is demonstrated in this work. Mixing modification of a Mach 2 jet from a convergent-divergent elliptic nozzle of aspect ratio 2, in the presence of two triangular tabs along the major and minor axis at the nozzle exit, at different levels of nozzle expansion has been studied. The results show that the mixing caused by tabs along the minor axis is impressive compared to the uncontrolled jet at all the pressure ratios. But for tabs along the major axis, mixing enhancement is significant only for nozzle pressure ratios above 5. Tabs along the minor axis cause better mixing than tabs along the major axis. The iso-pitot pressure contours reveal that the tabs along the minor axis enhance the mixing by bifurcating the jet. Shadowgraphs show that the tabs render the waves in the jet weaker. The present study demonstrates the superior mixing promotion caused by triangular tab than rectangular tab, studied by Aravindh Kumar and Rathakrishnan (2015).
In impulsive orbital maneuvers, thrust vector misalignment from the center of mass is the serious source of disturbance torque. A high capacity attitude control system is needed to compensate the mentioned large exogenous disturbance. In this paper a new retrofiring control method is proposed and studied which is based on the combination of a 1DoF gimbaled thrust vector control and spin-stabilization method. Spin-axis stabilization and disturbance rejection are considered as two important attitude control objectives. The nonlinear two-body dynamics of a small spacecraft is derived in which dynamical interaction between the nozzle and the body is significant. Reaction control system is not used and the only active control part is a 1DoF gimbal. The spacecraft design efficiency is very important; therefore, the H performance and control gain norm are chosen as two conflicting cost function in the Pareto front multiobjective optimization. Many Pareto fronts are given for some ranges of two favorable parameters: (1) spin rate and (2) spin-axis moment of inertia. Optimization variables are the closed-loop system poles. Moreover, poles region constraint is employed to obtain a well-damped transient response. From the perspective of performance and design efficiency, the optimization results give many attractive outcomes. The resulting system is an efficient design for a small spacecraft. Furthermore, numerical simulations are included to confirm the optimization results and illustrate the superiority of the proposed method compared to the only spin-stabilization.
The acquisition stage in global positioning system receivers provides a coarse estimation of the Doppler shift and the code phase of the incoming signals. The accuracy of the estimation, especially the Doppler shift, greatly influences the subsequent tracking loops. Based on the parameter prediction and the chirp z-transform algorithm, a novel acquisition approach to acquire the Doppler shift accurately is proposed. The code phase and the Doppler shift are predicted first according to the desired trajectory of the vehicle and satellite ephemeris. Then, frequency refinement of the code-stripped signal is conducted within a small interval around the predicted Doppler shift by using the chirp z-transform algorithm. To reduce the computational load, the data sequence is down-sampled with an integrate and dump accumulator without degrading the performance of the proposed algorithm. Results indicate that, with only 1 ms sampled data, the proposed algorithm can achieve a high-frequency accuracy. Besides, the proposed algorithm can acquire signals with the carrier-to-noise ratio down to 31 dB-Hz.
The Republic of Korea plans to launch a lunar orbiter and lander by 2020. There are several ways to enter lunar orbit: direct transfer, phasing loop transfer, weak stability boundary transfer, and spiral transfer trajectory. In this study, trajectory optimization is investigated for a lunar orbiter using a pattern search method that minimizes the required delta-V for direct lunar transfer. This method generates neighborhood points near the initial condition and then determines whether there is a new point that can reduce the value of the objective function. Classical methods require the gradient and acceleration of the objective function, but pattern search does not. Six poll methods and nine search methods are chosen; thus, 54 combinations of poll and search methods are available. The pattern search method can reduce the required delta-V on average by a few meters per second for a time of flight of five days and more than 10 m/s for a time of flight of four or six days, regardless of whether translunar injection is performed at the ascending or descending node.
The aerodynamic performance of a novel trailing edge L-shaped flap design is characterized numerically by means of computational fluid dynamics. The device is primarily thought for aerodynamic load adaptation to flight conditions. The device could be applied on a broad range of flying vehicles, like rotorcraft, for which it has been primarily designed, and also fixed wing aircraft and wind turbines. The operation of this movable surface is twofold. On one side, when the device is deployed downward, it acts as a Gurney flap, allowing the increase of the aerodynamic lift, without severe drawbacks in terms of drag rise. On the other side, when it is deflected upward, it is found capable to significantly alleviate the negative effects of stall. Simulations are carried out on a NACA 0012 airfoil equipped with the present L-shaped device at several angles of attack, both in linear and stall regimes. The reliability of numerical computations is supported by comparisons with pressure measurements and PIV surveys. Moreover, for small angles of attack, a Mach sensitivity analysis is performed, to assess the effects of compressibility on the L-shaped flap. Additionally, this work highlights how such device, when designed appropriately, can even delay the static stall onset.
The observability of the self-calibration and self-alignment system for an inertially stabilized platform is of vital importance, because it determines the solution existence of the system states. This article provides a straightforward and comprehensible method to investigate the observability of the nonlinear inertially stabilized platform’s self-calibration and self-alignment system. The proposed method is based on a principle that a parameter is observable only if it has a unique solution from the system outputs. The effect of the platform coordinates frame on the system observability is discussed in detail. The demonstration results indicate that the system is completely observable if the platform frame is defined based on the input axes of accelerometer triad. Besides, the analysis processes show that a high performance self-calibration and self-alignment can be accomplished if the inertially stabilized platform is kept stationary with the Earth at different positions and alternately rotated around its each axis. The validation of those results is checked by simulations, and the achieved conclusions make outstanding contributions to the development of the optimal torqueing schemes for the inertially stabilized platform’s self-calibration and self-alignment system.
Rotor noise is one of the most important issues for helicopter designer, and high-speed impulsive noise is particularly intense among the various rotor noise sources due to compressibility. Based on Computational Fluid Dynamics/Ffowcs Williams and Hawkings equations with Penetrable Data Surface (CFD/FW-Hpds) methods and hybrid optimization technique, a new optimization design procedure for rotor blade planform with low high-speed impulsive noise characteristics is established. First, in order to accurately capture the unsteady aerodynamic characteristics of rotor, based on the moving-embedded grid methodology, a CFD simulation is developed by solving the compressible Reynolds-average Navier–Stokes equations with Baldwin–Lomax turbulence model. The low dissipation Roe-Monotone Upwind-centered Scheme for Conservation Laws (MUSCL) scheme and highly efficient implicit lower–upper symmetric Gauss–Seidel scheme are used for spatial and temporal discretization, respectively. Second, taking the CFD results as sound pressure information input, the high-speed impulsive noise characteristics generated by transonic rotor are analyzed through a robust numerical method based on FW-Hpds equations. Third, the genetic algorithm and surrogated model based on radial basis function are combined as a hybrid optimization technique; during the optimization process, the blade grids are generated by a highly efficient parameterized method. Aiming at the minimization of the sound pressure level of rotor in forward flight, the parametric effect analyses of blade-tip shapes on transonic noise have been conducted first. Then, optimization analyses based on the rotor blade with double-swept and tapered tip have been accomplished with the aerodynamic performance as constraints. Compared with the baseline blade, it shows that the sound pressure level of rotor with optimized blade-tip shape can be decreased obviously at the present calculating condition due to its weaker transonic delocalization phenomenon in the region of blade tip. In the rotor plane, absolute peak value of sound pressure produced by the optimized blade planform is reduced about 59.1% of that by the baseline one, and the reduced value in sound pressure level is up to 5.6 dB.
The aim of the study presented herein is to numerically predict the behaviour of the airflow around a flying military aircraft with an active intake in which the airflow may enter and travel all the way up to the aerodynamic interface plane (the analytical interface between the inlet and engine). Computational fluid dynamics is used as the basic tool. The geometry created consists of a full-scale military aircraft exposed to different flight conditions. The flow results are mainly focused at the aerodynamic interface plane since the present study is a part of a greater research effort to estimate how the airflow distortion induced to the engine’s face due to the aircraft’s flight attitude, affects the embedded gas turbine’s performance. The obtained results were validated through a direct comparison against similar experimental ones, collected from a wind tunnel environment.
The complex structure and small-batch nature of aircraft landing gear parts lead to complex machining error propagation mechanism and manufacturing complexity. An Extended Machining Error Propagation Network model is presented to quantitatively analyze the complex coupling relationship in the Small-batch Multistage Machining Process of aircraft landing gear parts. Firstly, to depict the coupling relationship quantitatively, the Quality Features are defined to describe the machining precision information of Machining Form Features, and the State Elements are defined to describe the running state information of Machining Elements. Then, Machining Form Features, Machining Elements, Quality Features, and State Elements are identified as different network nodes, and the coupling relationships (such as evolving, locating, machining, and attribute) among these nodes are mapped into network edges. Based on the in-process measuring and sensing data of each machining stage, the topological and physical metrics of the network are explored to analyze the error propagation characteristics. Finally, the machining process of an outer cylinder part from aircraft landing gear is studied to verify the proposed methods.
Engineering design problems can, in general, be discussed under the framework of decision making, namely engineering design decisions. Inherently, accounting for uncertainty factors is an indispensable part in these decision processes. In a sense, the goal of design decisions is to control or reduce the variational effect in decision consequences induced by many uncertainty factors, by optimizing an expected utility objective or other preference functions. In this paper, the value of data in facilitating making engineering design decisions is highlighted, and a data-driven design paradigm for practical engineering problems is proposed. The definition of data in this paradigm is elaborated first. Then the data involvement in a whole stage-based design process is investigated. An overall decision strategy for design problems under the data-driven paradigm is proposed. By a concrete satellite design example, the key ideas of the proposed data-driven design paradigm are demonstrated. Future work is also advised.
The influence of end-wall roughness on the performance of the axial compressor stage was investigated with different values of roughness added to the hub and shroud surfaces of transonic compressor stage, NASA Stage 35. Firstly, the numerical code was validated against the experimental data, which were available from the open literature. Afterwards the model was applied to simulate the effect of end-wall roughness with different amplitudes of dimensionless sand-grain roughness height. Numerical results indicated that the increment of end-wall roughness caused the deterioration of compressor stage performance. To understand the mechanism behind, the distributions of loadings, losses and the detail flow situations near end-wall region were analyzed and discussed. The results show that the overall performance drop is mainly due to the thickened end-wall boundary layer. Near the hub region, the rough hub surface induces larger corner stall and the shock wave moves upstream. Meanwhile, the casing roughness leads to the slight increment of tip leakage mass flow and extension of blockage in the circumferential direction.
With the development of science and technology, the performance of an aero-engine has been given more rigorous requirements. Seal device is an important component part of an aero-engine, and the improvement in its performance may be an efficient way to further improve the performance of an aero-engine. Finger seal is a flexible seal and has higher performance price ratio, therefore it gets more attention and research recently. The phenomenon of noncontact state converting to contact state will occur in every working cycle of finger seal that inevitably lead to the finger seal bearing the impact effect of rotor. But so far, the influence of impact on the finger seal performance has not been discussed and researched. To overcome this shortcoming, the stress–strain curves of C/C composite under different impact velocities are obtained by the Gleeble3500 thermo-simulator system in the paper, and then the elastic modulus of C/C composite in three directions is calculated by experimental data. The effects of impact velocity and impact damping on the impact force are analyzed by means of the impact theory. The new structural stiffness of finger seal and the impact displacement excitation of the rotor are built through impact effect analysis. On this basis, the equivalent dynamic model of C/C composite finger seal with distributed mass is established to evaluate the impact effect. By the model, the difference of calculated results is analyzed under whether considering the impact effect or not. And the effect of impact velocity and coefficient of restitution on the dynamic performance of the finger seal is also analyzed under considering the impact effect, respectively. The above results show that the impact effect has significant influence on the leakage and wear of finger seal, therefore when the performance of the finger seal was analyzed, it is necessary to consider the impact effect.
The main specification in the verification by testing of space hardware vulnerability to shock excitations is the shock response spectrum. Although it compiles the most relevant information needed to describe the overall shock environment characteristics, shock testing still poses various difficulties and uncertainties concerning the suitability and operation of the shock test system used, and the adequate definition of the underlying test parameters. The approach followed from the interpretation of typical shock testing specifications to the development, validation, and characterization of the developed shock test system, including the definition and design of the relevant parameters influencing the attained shock environment, is described in this paper. The shock testing method here presented consists of a pendular in-plane resonant mono-plate shock test apparatus where the structural response of the ringing plate depends upon well-defined controllable parameters (e.g. impact velocity, striker shape, mass, and contact stiffness), which are parametrically determined to achieve the target shock environment specification. The concept and analytical model of two impacting bodies are used in a preliminary analysis to perform a rigid body motion analysis and contact assessment. A detailed finite element model is developed for the definition of the ringing plate dimensions, analysis of the plate dynamics and virtual shock testing. The assembled experimental apparatus is described and a test campaign is undertaken in order to properly characterize and assess the design and test parameters of the system. The developed shock test apparatus and corresponding finite element model are experimentally verified and validated. As a result of this study, a reliable finite element modeling methodology available for future shock test simulation and prediction of the experimental results was created, being an important tool for the adjustment of the shock test input parameters for future works. The developed shock test system was well characterized and is readily available to be used for shock testing of space equipment with varying specifications.
In this paper, a frequency-domain modeling methodology that can be applied to various multi-rotor aerial vehicles is introduced. The primary contribution of this work is a systematic integration of the first-principles modeling and system identification approaches to generate flight dynamics models with good accuracy. The first-principles modeling and model linearization are conducted to obtain an appropriate baseline model for the subsequent system identification. Next, a four-step parameter identification process, which consists of: (1) baseline model determination; (2) data collection and preprocessing; (3) mode-wise parameter identification; and (4) model fidelity validation, is performed in the frequency domain to identify the uncertain parameters. Our method has been applied to two custom-built multi-rotor aircraft (one X-type quadcopter and one QU4D quadcopter) for efficiency demonstration.
A rapid and decoupled three-degree-of-freedom trajectory planning approach is presented for maneuvering entry vehicles. A maneuver coefficient is defined to describe the lateral motion of a three-degree-of-freedom trajectory so that the designs for the longitudinal and lateral trajectories are decoupled. The longitudinal drag profile is planned according to the flight range that considers the maneuver coefficient. The lateral trajectory is controlled by an adjustable heading error corridor to achieve the required maneuver coefficient. The three-degree-of-freedom trajectory is finally generated based on the drag profile and the adjustable corridor. The trajectory planning algorithm is tested on two entry vehicles. Results indicate that this algorithm is capable of planning three-degree-of-freedom trajectories for nominal and maneuvering missions. The feasible region of the maneuver coefficient is investigated for each vehicle. The wide region demonstrates that the proposed planning algorithm is applicable and insensitive to estimation errors of the maneuver coefficient.
Locking mechanism is an important part of landing gear, which is required to lock retractable landing gear during taking off and landing processes of aircraft. This paper introduces an effective numerical simulation forecasting method to investigate the friction and wear of unworn finger lock chuck, where an associated experiment was conducted to verify the correctness of the method. The dynamic explicit procedure was adopted in the simulation process with Abaqus/Explicit solver, the user subroutine VFRIC integrated in the commercial package ABAQUS was coded to study the rate-dependent dynamic friction coefficient during the movement of unworn finger lock chuck. The friction simulation results indicate that large deformation occurs in the finger lock chuck during the unlocking–locking process and the maximum stress lies in the root zone of finger lock chuck, the difference of axial acting force between simulation results and experiment results of unworn finger lock is small. From the perspective of frictional energy dissipation, nodal frictional energy density rate within contact footprint regions was taken as the index to assess the wear severity. Python language commands were programmed to realize the secondary development of ABAQUS post-process, the nodal frictional energy density rate distribution in contact footprint regions of finger lock chuck was graphically displayed. The wear simulation results show that the nodal frictional energy density rate distribution of finger lock chuck concentrates significantly at the regions of two rounded corners in the raised portion of chuck and inner side surface of the raised platform, indicating that wear first occurs in these regions after the unworn finger lock chuck is put into use. The wear simulation results obtained was compared with a worn finger lock chuck after 500 times disengagement-stuck processing wear test, and the results are basically accordant.
Multi-disciplinary shape optimization of a re-entry capsule with aero-thermodynamic, trajectory, stability and the geometry considerations is presented in this research. The method is based on decomposition of the underlying problem into disciplinary routines performing separated analysis for each goal. The objectives of the optimization are maximizing volumetric efficiency, minimizing longitudinal stability derivative and minimizing the ballistic coefficient, subject to constraints on geometry, heating load and deceleration. Utilizing a multi-objective genetic algorithm will result in a collection of Pareto optimal solutions. Then, the multi-disciplinary multi-objective optimization process allows finding a Pareto front of the best shapes. Resulting optimal solutions obviously show the compromises among volumetric efficiency, longitudinal stability and ballistic coefficient. In the end, the results containing dimension's characteristics of the re-entry capsule are presented.
A lunar sampler plays the critical role and is of great importance in lunar explorations. In this paper, an error model was established for a novel flexible lunar sampler which has low weight, small volume, large workspace, and low power consumption. Based on its specific configuration, the forward and inverse kinematic models and the kinetostatic model are developed to formulate the volumetric error model of the mechanism. The error model is built by considering three classes of error sources: flexibility-induced errors, structural parameter-induced errors, and joint clearance errors. The flexible errors are compensated according to the experimental data; the second class of errors are modeled based on complete differential-coefficient theory, while the third class of errors are modeled based on the deterministic method. For the third class in error modeling, contact modes are introduced to build a joint clearance-induced error model. The relationship between different error sources and the output pose error of the sampling head is obtained. Finally, the error distribution in the workspace is evaluated.
A lunar roving vehicle is of great significance in the processing of manned lunar exploration for long-duration space exploration missions. The mechanical model of the lunar roving vehicle’s wheel directly affects the mobility performance. However, the wheel of lunar roving vehicle is a special one with the metal mesh surface, and the soil can penetrate through its surface. In this paper, the mechanical model of a rigid normal wheel is deduced, and four stress correction coefficients are introduced to obtain the mechanical model of the lunar roving vehicle’s wheel. These four stress correction coefficients are identified in several groups of experiment of the lunar roving vehicle’s wheel, and using the functional coefficients can have a high precision for the mechanical model of the lunar roving vehicle’s wheel.
The maximum down range trajectory optimization problem with multiple phases and multiple constraints corresponding to the flight of a boost-glide vehicle is considered. The longitudinal motion model was built as a multiphase optimization problem under constraints. Legendre–Gauss–Radau collocation points were used to transcribe the optimization problem into a finite-dimensional nonlinear programming problem, and the maximization down range trajectory was obtained based on adaptive mesh refinement pseudospectral methods. However, sometimes it is difficult to find interior points without position constraints. A novel optimization strategy based on dynamic programming theory was proposed to search the free interior points more accurately and quickly, which resulted in almost the same optimized trajectory while producing a small mesh. The results of numerical examples showed that the boost-glide vehicle trajectory optimization problem is solved using the adaptive mesh refinement pseudospectral methods.
Balloon-borne astronomy offers an attractive option for experiments that require precise pointing and attitude stabilization, due to a large reduction in the atmospheric interference observed by ground-based systems as well as the low-cost and short development time-scale compared to space-borne systems. The Balloon-borne Imaging Testbed (BIT) is an instrument designed to meet the technological requirements of high-precision astronomical missions, and is a precursor to the development of a facility-class instrument with capabilities similar to the Hubble Space Telescope. The attitude determination and control systems (ADCS) for BIT, the design, implementation, and analysis of which are the focus of this paper, compensate for compound pendulation effects and other sub-orbital disturbances in the stratosphere to within 1–2'' (rms), while back-end optics provide further image stabilization down to 0.05'' (not discussed here). During the inaugural test flight from Timmins, Canada in September 2015, BIT ADCS pointing and stabilization performed exceptionally, with coarse pointing and target acquisition to within <0.1° and fine stabilization to 0.68'' (rms) over long (10–30 min) integrations. This level of performance was maintained during flight for several tracking runs that demonstrated pointing stability on the sky for more than an hour at a time. To refurbish and improve the system for the three-month flight from New Zealand in 2018, certain modifications to the ADCS need to be made to smooth pointing mode transitions and to correct for internal biases observed during the test flight. Furthermore, the level of autonomy must be increased for future missions to improve system reliability and robustness.
This paper investigates the formation control problem of multiple unmanned aerial vehicles (UAVs) with limited communication in a known and realistic obstacle-laden environment. In order to deal with the limited communication constraints, the leader–follower strategy and the virtual leader strategy are integrated into an optimal control framework to formulate this formation control problem. This combination formation framework can be achieved by integrating a redefined directed graph and a proposed information vector. In more practical applications, an obstacle/collision avoidance strategy is achieved by constructing a non-quadratic cost function innovatively using a virtual flow field approach. The proposed optimal control laws, which derive from the local information rather than the global information, are proved to guarantee the stability of the close-loop system by an inverse optimal control approach. The simulation results demonstrate the effectiveness of the formation flight of multiple UAVs with limited communication in an obstacle-laden environment.
The large-scale of unmanned aerial vehicle applications has escalated significantly within the last few years, and the current research is slowly hinting at a move from single vehicle applications to multivehicle systems. As the number of agents operating in the same environment grows, conflict detection and resolution becomes one of the most important factors of the autonomous system to ensure the vehicles’ safety throughout the completion of their missions. The work presented in this paper describes the implementation of the novel distributed reactive collision avoidance algorithm proposed in the literature, improved to fit a swarm of quadrotor helicopters. The original method has been extended to function in dense and crowded environments with relevant spatial obstacle constraints and deconfliction manoeuvres for high number of vehicles. Additionally, the collision avoidance is modified to work in conjunction with a dynamic close formation flight scheme. The solution presented to the conflict detection and Resolution problem is reactive and distributed, making it well suited for real-time applications. The final avoidance algorithm is tested on a series of crowded scenarios to test its performances in close quarters.
This paper presents the large angle attitude manoeuvre control design of a single-axis flexible spacecraft system that consists of a central rigid body and a cantilever beam with bonded piezoelectric sensor/actuator pairs as a flexible appendage. The proposed control strategy combines the attitude controller designed by the adaptive robust control technique with the active vibration controller designed by the positive position feedback control method. The desired angular position of the spacecraft is planned and an adaptive robust attitude control approach based on a projection type adaptation law is proposed to track the planned path and to achieve precise attitude manoeuvre control. Meanwhile, the positive position feedback control method is applied to actively increase the damping of the flexible appendage and to suppress the residual vibration induced by manoeuvre. Improved transient and steady state performance during and after large angle attitude manoeuvre can be both achieved by integration of the technical merits of all these control methods. Analytical and numerical results illustrate the effectiveness of this approach.
Artificial neural networks are an established technique for constructing non-linear models of multi-input-multi-output systems based on sets of observations. In terms of aerospace vehicle modelling, however, these are currently restricted to either unmanned applications or simulations, despite the fact that large amounts of flight data are typically recorded and kept for reasons of safety and maintenance. In this paper, a methodology for constructing practical models of aerospace vehicles based on available flight data recordings from the vehicles’ operational use is proposed and applied on the Jetstream G-NFLA aircraft. This includes a data analysis procedure to assess the suitability of the available flight databases and a neural network based approach for modelling. In this context, a database of recorded landings of the Jetstream G-NFLA, normally kept as part of a routine maintenance procedure, is used to form training datasets for two separate applications. A neural network based longitudinal dynamic model and gust identification system are constructed and tested against real flight data. Results indicate that in both cases, the resulting models’ predictions achieve a level of accuracy that allows them to be used as a basis for practical real-world applications.
In this study, a genetic optimization algorithm is applied to the design of environmentally friendly aircraft departure trajectories. The environmental optimization has been primarily focused on noise abatement and local NOx emissions, whilst taking fuel burn into account as an economical criterion. In support of this study, a novel parameterization approach has been conceived for discretizing the lateral and vertical flight profiles, which reduces the need to include nonlinear side constraints in the multiparameter optimization problem formulation, while still permitting to comply with the complex set of operational requirements pertaining to departure procedures. The resulting formulation avoids infeasible solutions and hence significantly reduces the number of model evaluations required in the genetic optimization process. The efficiency of the developed approach is demonstrated in a case study involving the design of a noise abatement departure procedure at Amsterdam Airport Schiphol in The Netherlands.
An efficient stall compliance prediction method using quick configuration generation, adapted mesh, high fidelity analysis, and wind tunnel test data for trimmed very light aircraft is proposed. The three-dimensional Navier–Stokes equations are used to determine the characteristics of the flow field around the aircraft, and the
Multi-field coupling problems are taken more and more attention mainly because of the higher requirement of load, efficiency, and reliability in aero-engine operation. This research takes an aero-engine compressor as the research object, 3D flow field and structural models are established. For the method of cyclic symmetric, single-sector model is selected as the calculation domain. Considering the influence of former stator wakes, compressor flow field is simulated. The article analyzes the distribution law of unsteady aerodynamic load on rotor blade. Based on Kriging model, load transfer of aerodynamic pressure and temperature is achieved from flow field to blade structure. Then the effects of centrifugal force, aerodynamic pressure and temperature load are discussed on compressor vibration characteristic and structural strength. The results show dominant fluctuation frequencies of aerodynamic load on rotor blade are manly at frequency doubling of stator–rotor interaction, especially at one time frequency (1 x f0). Magnitude and pulsation amplitude on pressure surface are far greater than that on suction surface. Load transfer with Kriging model has a higher precision, it can meet the requirement of multi-field coupling dynamic calculation. In multi-field coupling interaction, temperature load makes the natural vibration frequencies decrease obviously, centrifugal force is the main source of deformation and stress. Bending stress induced by aerodynamic pressure and temperature load can counteract part of bending stress induced by centrifugal force. However, temperature load causes the maximum displacement of blade-disk system to increase.
The fifth Automated Transfer Vehicle was launched on 29 July 2014 with Ariane-5 flight VA 219 into orbit from Kourou, French Guiana. For the first time, the ascent of an Ariane rocket was independently tracked with a Global Navigation Satellite System (GNSS) receiver on this flight. The GNSS receiver experiment OCAM-G was mounted on the upper stage of the rocket. Its receivers tracked the trajectory of the Ariane-5 from lift-off until after the separation of the Automated Transfer Vehicle. This article introduces the design of the experiment and presents an analysis of the data gathered during the flight with respect to the GNSS tracking status, availability of navigation solution, and navigation accuracy.
Cable-driven parallel mechanism is a special kind of parallel robot in which traditional rigid links are replaced by actuated cables. This provides a new suspension method for wind tunnel test, in which an aircraft model is driven by a number of parallel cables to fulfil 6-DOF motion. The workspace of such a cable robot is limited due to the geometrical and unilateral force constraints, the investigation of which is important for applications requiring large flight space. This paper focuses on the workspace analysis and verification of a redundant constraint 6-DOF cable-driven parallel suspension system. Based on the system motion and dynamic equations, the geometrical interference (either intersection between two cables or between a cable and the aircraft) and cable tension restraint conditions are constructed and analyzed. The hyperplane vector projection strategy is used to solve the aircraft’s orientation and position workspace. Moreover, software ADAMS is used to check the workspace, and experiments are done on the prototype, which adopts a camera to monitor the actual motion space. In addition, the system construction is designed by using a built-in six-component balance to measure the aerodynamic force. The results of simulation and tests show a good consistency, which means that the restraint conditions and workspace solution strategy are valid and can be used to provide guidance for the cable-driven parallel suspension system’s application in wind tunnel tests.
The design of guidance and control strategies is a promising study trend of dynamic soaring for small unmanned aerial vehicles, for which the flight modeling and simulation specifically for soaring-capable unmanned aerial vehicles is significant and necessary. The aim of this paper is to propose a flight simulation platform for dynamic soaring. In order to do so, firstly, two different sets of equations of motion of small unmanned aerial vehicles have been derived and their characteristics are compared: one is expressed in body-fixed frame and the other in air-relative flight path frame. Secondly, the latter set is used for energy analysis to maximize energy gain while climbing and diving and minimize energy cost during turning in dynamic soaring. While the former serves to build the dynamic soaring simulation platform, in which a piecewise trajectory-based guidance and control strategy according to the energy analysis is proposed tracking the optimum climb and bank angles and traveling toward desired directions. Simulation results indicate that the unmanned aerial vehicles can perform dynamic soaring toward various directions in different wind fields, follow asymptotically the typical straight-line and circular-orbit paths by repeating soaring cycles.
Based on rigid-body slung-load hypothesis, a nonlinear dynamical model of the helicopter and slung-load system is presented. As for helicopter and slung-load system, the introduction of the rigid-body slung-load results in several extra degrees of freedom and constraints, and this makes the nonlinear equations of helicopter and slung-load motion transfer from 9 orders to 19 orders. The nonlinear equations are linearized by small perturbation hypothesis for stability analysis and they are trimmed by the continuation method. First, the simulation trimmed results are compared with the helicopter flight test data without slung-load and the calculation results of helicopter with mass-point slung-load in the literature. Then, the differences among trimmed states of the helicopter with a rigid-body slung-load, a mass-point slung-load, and without slung-load are studied. The motion modes of helicopter with rigid-body slung-load are calculated in different cases, and the effects of extra slung-load to the helicopter stability are analyzed at the same time. Results showed that the rigid-body slung-load added five new motion modes to the whole helicopter and slung-load system, two of which concerning cable swinging motion are stable, while the others concerning slung-load attitude motion are unstable. At the same time, the introduction of rigid-body slung-load makes the short-period mode and the roll mode of helicopter unstable, which has a strong impact on the flight quality of helicopter.
The true proportional navigation guidance law, the augmented proportional navigation guidance law, or the adaptive sliding-mode guidance law, is designed based on the planar target-to-missile relative motion dynamics. By a proper construction of a nonlinear Lyapunov function for the line-of-sight angular rates in the three-dimensional guidance dynamics, it is shown that the three guidance laws mentioned above are able to ensure the asymptotic convergence of the angular rates as they are directly applied to the three-dimensional guidance environment. Furthermore, considering the missile autopilot dynamics as a first-order lag, we design three-dimensional nonlinear guidance laws by using the backstepping technique for three cases: (1) the target does not maneuver; (2) the information of target acceleration can be acquired; and (3) the target acceleration is not available but its bound is known a priori. In the first step of the backstepping design of the control law, there is no need to cancel the nonlinear coupling terms in the three-dimensional guidance dynamics in such way that the final expressions of the proposed guidance laws are significantly simplified. Thus, the proposed nonlinear Lyapunov function for the line-of-sight angular rates is a generalized function for designing three-dimensional guidance laws. Simulation results of a missile interception mission show that the proposed guidance laws are highly effective.
Seals are used in hydraulic actuators or any other hydraulic devices to prevent passing of hydraulic fluid from one chamber to another, or to prevent external leakage and entry of any foreign contaminants. The primary function of any hydraulic actuator is to efficiently use hydraulic power to drive a load experienced during movement of control surfaces or movable aircraft structure. Efficient sealing helps in achieving this, but with its own friction which should be as minimal as possible. Thus, the estimation of seal friction force has crucial significance in hydraulic actuators, especially in flight control actuators that demand high performance and dynamic behavior characteristics while efficiently driving the load. This paper details the methodology adopted for theoretical estimation of total seal friction force of actuator as well as description of experimental test set-up and test method followed to record the total friction value at different positions of the actuator. The theoretical estimation was done using empirical formulae and graphs for predicting seal friction force by considering the effects of seal squeeze, hydraulic pressure, seal dimensions, seal material and then interpolating the same for the specific type of seals used. An experimental study is also presented in this paper, which can be conducted to validate the theoretically estimated value after building up of development prototypes. The validation is necessary as seal friction force calculation during design phase is an approximation and accurate friction of every seal is difficult to measure as it depends on a number of parameters. Thus, this paper explains the subject issue with the help of a case study which provides the theoretical estimation as well as its validation through an experiment to study this significant aspect of a hydraulic actuator design.
A new dynamic modeling method called the improved thin-layer element method is proposed to apply to the aero-engine bolted joints. The thin-layer elements are partitioned based on the interface contact stress distribution. In addition, the material parameters of the partitioned thin-layer elements are determined by the bolted joints stiffness technique and the fractal contact theory without the experimental results, which allows the engineer to estimate the dynamic characteristics of whole structure before the physical prototype is available. First, the modeling principles of the improved thin-layer element method are studied and the bolted joints stiffness is analyzed. Next, the material parameters of the partitioned thin-layer elements are determined on the basis of the interface contact stress distribution characteristics of the bolted joints. Finally, this method is applied to the simulative casing bolted joints structure and the results are compared with the experimental results in order to verify the proposed method. The results indicate that the improved thin-layer element method is more accurate than the thin-layer elements method, and the material parameters of the partitioned thin-layer elements can be expressed by the structural parameters of the aero-engine bolted joints without updating based on the experiment.
With the development of space exploration, researches on space robot will cause more attentions. However, most existing researches about dynamics and control of space robot concern planar problem, and the effect of flexible panel on dynamics of the system is not considered. In this article, dynamics modeling and active control of a 6-DOF space robot with flexible panels are investigated. Dynamic model of the system is established based on the Jourdain's velocity variation principle and the single direction recursive construction method. The computed torque control method is used to design point-to-point active controller of the space robot. The validity of the dynamic model is verified through the comparison with ADAMS software; the effects of panel flexibility on the system performance and the active controller design are studied in detail. Simulation results indicate that the proposed model is effective to describe the dynamics of space robot; panel flexibility has large influence on the dynamic behavior of space robot; the designed controller can effectively make the robot reach a specified position and the elastic vibration of the panels may be suppressed simultaneously.
Blade–disk–drum (BDD) assemblies are commonly used in rotors of aeroengine. This paper investigates the vibration characteristics of the drum in the BDD assembly. In the analysis, multilevel modeling method based on ANSYS software is proposed. Free vibrations of single drum, disk-drum assembly, and the BDD assembly are calculated and carefully examined to discover the special modal shapes and coupling effects. Then according to the working environment, equivalent aerodynamic loads are imposed on blades and the drum respectively, to explore the vibration behaviors of the drum in the BDD assembly. Results reveal that, vibration of the drum, which may result in rub-impact or/and fatigue failure, can be induced by coupling vibration result from the aerodynamic loads on blades and drum, as well as the drum-dominant mode triggered by the aerodynamic loads on drum.
This paper proposes three-dimensional impact angle control guidance laws based on a sliding mode control technique. Unlike the usual approach of decoupling the engagement dynamics into two mutually orthogonal two-dimensional planes, the guidance laws are derived using coupled engagement dynamics. By using this approach, the control effort required to achieve the objective reduces and the performance of the guidance law is improved. The derivations of guidance laws are done using both conventional as well as nonsingular terminal sliding mode control, which guarantees asymptotic and finite time convergence, respectively, to the desired impact angle. In order to derive the guidance laws, multi-dimensional switching surfaces are used. The stability of the system, with selected switching surfaces, is demonstrated using Lyapunov stability theory. Numerical simulation results are presented to validate the proposed guidance laws for constant speed, as well as a realistic interceptor model with given aerodynamic properties. The simulations show the advantage of using coupled dynamics. The robustness of the proposed guidance laws, with respect to the interceptor’s system lag, is also investigated.
The stall and surge directly impact on the safety and reliability of compressors. The spike-type and modal-type stall inception exist in compressors. At present, few studies pay attention to the stall inception of centrifugal compressor, such as the formation reason for the stall inception and the action by the volute tongue on the stall precursor. This paper investigated the stall characteristics of a high-speed small-flow centrifugal compressor and illustrated the relationship between the volute tongue and the location of stall inception. In addition, the mechanism of stall inception was also clarified. Both the analysis of initial flow structures and the comparison of the frequency spectrum characteristics at different monitoring points show that the spike-type stall occurs at about 115° circumferential position in this centrifugal compressor. The nonaxisymmetric geometry structure of the volute leads to the uneven circumferential pressure distribution. The blockade effect of the volute tongue results in high static pressure area near the volute tongue. The disturbance caused by high static pressure adversely propagates into the diffuser, resulting in the static pressure peak value at different radii. As the pressure peaks adversely migrate to the impeller inlet and induce the leading edge spillover near the corresponding blade, the spike-type stall occurs. Therefore, the volute tongue both induces the stall inception and determines the circumferential position of the stall inception at the centrifugal compressor inlet.
Triple-swirler plays an important role for aero-engine combustors to achieve high temperature rise. In this paper, experimental investigations were carried out to explore the effect of triple-swirler rotational direction on swirling flow field in atmospheric condition. Two-dimensional-planar particle image velocimetry measurements show that the central toroidal recirculation zone (CTRZ) formation is significantly affected by the swirler rotational direction combinations: an obvious CTRZ can be formed by the triple-swirler with co-rotating intermediate swirler and outer swirler, while a much smaller CTRZ was obtained by the triple-swirler with a counter-rotating intermediate and outer swirler. Furthermore, the swirl level of the mixed flow is significantly affected by the rotational direction combination, and the integrated swirl numbers were calculated to help evaluating the swirl level generated by triple-swirlers. The rotational direction combination plays a key role on the tangential velocity distribution. The tangential velocity distribution is not only closely related to rotational direction, but also the swirl number combination and mass proportion of each swirler in a triple-swirler.
In this paper, a modal approach for the fast calculation of flow mesh deformation around a wing is developed based on the elastic solid method of dynamic mesh. The flow mesh domain is assumed to be a pseudo elastic solid. The displacement of the wing and the pseudo elastic solid is continuous at the fluid structure interface. Considering the condition of displacement continuity, the governing equation for the vibration of the wing with the pseudo elastic solid together is derived. The frequencies and mode shapes of the wing and the pseudo elastic solid are computed. Then the nodal displacements for the wing and the flow mesh are computed using modal superposition. The flutter boundary of the AGARD Wing 445.6 is predicted using the present modal approach by considering the first four modes of the wing. The calculated results compare well with the experimental data. The computing time is reduced by 54.8% compared with the pre-existing elastic solid method.
The research reported in this paper investigated the structural design methodology of equipping a jet aircraft modified for maritime patrol with any type of observation system. The study includes installation feasibility, definition of mounting types (fix/retractable), installation location, airworthiness requirements, design considerations, structural design requirements, and static/dynamic analysis considerations. Due to lack of design publication in this field, this study focuses on methodology of electro-optical sensor installation as a flowchart leading to a road map to equip an aircraft with any type of maritime sensor. For validation of above methodology, a case study for installation of an electro-optical system on a particular jet aircraft with new mission is presented. By comparison of system and aircraft specification based on methodology flowchart, a retractable mounting was chosen. To accomplish a safe installation, an optimal mechanism suitable with aircraft fuselage and required backup structure was designed using reinforced cut out. The design was then analyzed for static and dynamic critical load cases using MSC/NASTRAN software.
In the combat environment, wing damage is of particular concern since the wing is the main component to generate lift which affects survivability or safety. Previous research has been mainly concentrated on the aerodynamic behavior of the aircraft with wing damage. However, the military analysts are more concerned about the specific value of survivability (measured by survival probability) so as to correctly plan combat operations. This paper proposes a method for quantitatively describing the relationship of aircraft survivability and damage holes with different sizes and different locations on the wing. Examples show that the survivability decreases with increasing hole size; the magnitude of survivability decrement reduces when moving the damage backwards or moving the damage toward the wing tip; when the effect of damage on the vulnerability is considered for different hole sizes, the survivability is reduced more significantly compared with the cases where the vulnerability is kept constant for all scenarios.
Numerical study on the compressor stage of a KJ-66 micro gas turbine was conducted in this paper through both steady and unsteady Reynolds-averaged Navier–Stokes. The study was conducted for the numerical prediction of micro gas turbine compressor performance at various operation conditions, with special attention given to the transient flow behaviors during compressor operation. The numerical results showed reasonable agreements with experimental data while providing predictions for the charting of compressor performance map at various operation speeds. The simulation results indicated that the increase of operation speed from 80 k r/min to 117 k r/min would leads to an increased peak total pressure ratio from 1.54 to 1.96, while decreasing the peak adiabatic efficiency from 0.73 to 0.55. This paper also provided discussion on details of transient flow field within the compressor stage as well as demonstrated the smooth flow transition through rotor–stator interactions.
Multistage axial compressors are widely used in the gas turbine engines. The strength of rotors is one of the key factors for the reliability of multistage axial compressors. The stresses of rotors at real working conditions can be caused by the centrifugal load, thermal load, and aerodynamic load. It is important to figure out the roles and the mechanism of the three kinds of loads in the stresses generating process. In this paper, the stresses of rotors in a typical five-stage axial compressor are calculated with different kinds of loads by solid–fluid coupling method. The results show that the proportion of the stress caused by centrifugal load is more than 80% of the total stress, which is dominant. The maximum proportion of the stress caused by thermal load is about 20% of the total stress at the front stages. However, the influence of thermal load is quite different from the first stage to the last stage. It is surprising that thermal load can decrease the stresses of the last stage rotor, which is mainly because of the variation of radial temperature gradient at disks for different stages. The proportion of the stress caused by aerodynamic load is usually less than 4%, and it tends to make the stresses at the suction side of the blades lower and enlarge it at the pressure side. According to the above results, centrifugal load is necessary of consideration at the conceptual design phase for the multistage axial compressor rotors. At preliminary three-dimensional design phase, centrifugal load and thermal load should be considered together. At optimized three-dimensional design phase, aerodynamic load cannot be neglected and all the three loads should be considered.
This article derives a nonlinear regression Huber-based divided difference filtering algorithm using a nonlinear regression approach for dynamic state estimation problems with non-Gaussian noises and outliers. In this approach, the nonlinear measurement model is directly used without linear or statistically linear approximation and the Huber-based divided difference filtering problem is solved using a Gauss–Newton approach. This new proposed filter method is then applied to a benchmark problem of estimating the trajectory of an entry body from discrete-time range data measured by a radar tracking station. Simulation results demonstrate the superior performance of the proposed filter as compared to the previous filter algorithms in the presence of non-Gaussian uncertainties.
Awareness of environmental and economic issues associated with fossil fuel has led to the exploration of alternative fuels for aviation. Analysis and measurements of alternative fuel using real aircraft engines are complex and costly. Thus, evaluation only through computation is an option at present. This paper presents an analysis of aircraft engine emissions, particularly NOx and CO, from the blend of bio-synthetic paraffinic kerosene (bio-SPK) fuel with kerosene using a simplified gas emission model. Three different fuels, namely, a conventional aviation fuel Jet-A, Jatropha bio-SPK and Camelina bio-SPK were tested as pure and as blends with Jet-A. Chemical properties of the tested fuels were introduced into HEPHAESTUS, an in-house gas emission software developed in Cranfield University. HEPHAESTUS was developed based on the physics-based approach by incorporating a number of stirred reactors to predict NOx, CO, UHC and soot. Gaseous emissions generated from kerosene were observed to follow the trends provided by the ICAO databank. The capability of HEPHAESTUS in predicting the NOx and CO level from biofuel is yet to be explored. The level of NOx and CO predicted in this study followed the trends shown in the literature, although they quantitatively differed. Compared to Jet-A, NOx decreased and CO increased as the percentage of Jatropha bio-SPK and Camelina bio-SPK in the mixture increased. NOx reduction was consistent with the reduction in flame temperature because NOx generation considered in the model was dominantly based on thermal NOx. In contrast, increases in CO were due to low flame temperature that led to incomplete combustion. The consistency of the results obtained showed that the computational work performed in this study as an initial step toward the prediction of emission level of biofuels was successful. However, further studies on the experimental work or computational fluid dynamic simulation is essential.
Breathing blunt nose technique is one of the promising methods for reducing the drag of blunt-nosed body at hypersonic speeds. The air, traversed by the bow shock positioned ahead of the nose, at the stagnation region is allowed to enter through a hole at the blunt-nose and ejected at the rear part (base region) of the body. This manipulation reduces the positive pressure over the stagnation regions of the nose and increases the pressure at the base, resulting in reduced suction at the base. The simultaneous manifestation of reducing the compression at the nose and suction at the base regions results in reduction of the total drag. The drag reduction caused by the breathing blunt nose technique has been measured in a Mach 7 tunnel. Also, the drag and flow field around the blunt-nosed body, with and without breathing hole, has been computed. The aerodynamic characteristics of the breathing blunt nose model obtained experimentally are compared with the CFD results. It is found that the breathing results in 5% reduction in drag. The lift coefficient also comes down for the model with breathing nose. But the lift-to-drag ratio is found to be the same for both the cases; the blunt-nosed body with and without nose-hole.
Concerning on the problem of low measuring precision of the current micro-inertial sensors, a novel attitude measurement method is proposed to dismiss the drift for remarkable attitude error. According to the output of the onboard three-axis magnetic sensor in the process of projectile flight, a low-cost attitude detection system is designed by using the intersection ratio of the sensor. First, the output model of the onboard three-axis magnetic sensor is established. The mathematical relationship between the characteristic ratio of magnetic sensor output and the pitch angle is then derived. Then, the solution and correction algorithm of the attitude angles are studied. Finally, the effectiveness of the attitude measurement method has been validated by carrying out the semi-physical experiments. The experimental results indicate that the error of attitude angles is within ±1° and the attitude angle error of the combined magnetic sensors is not cumulative. Meanwhile, the geomagnetic field strength is dispensable during the whole calculation process. Compared with the "Zero Crossing Method", the proposed method has shown a nearly two-times higher accuracy and has no limitation of "MAGSONDE window". What is more, this method proves to be more simple and has a doubled update rate in attitude angle calculation.
This paper proposes a new method for selecting an ellipse-shaped geographical area and constructing a routing grid that circumscribes the contour of the designated area. The resulting grid describes the set of points used by the flight trajectory optimization algorithms to determine an aircraft’s optimal flight trajectory as a function of given particular atmospheric conditions. This method was developed with the intent of its employment in the context of Flight Management System trajectory optimization algorithms, but can be used in Air Traffic Management environments as well. The routing grid limits the trajectory’s maximal total ground distance (between the departure and destination airports), maximizes the geographical area (for a better consideration of the wind conditions) and minimizes the number of grid nodes. The novelty of the proposed method resides in the fact that it allows a distinct and independent parameterization and control of the ellipse’s total surface, and the required size of the take-off/landing procedure maneuvering areas at the departure/destination airports. The ellipse contour constructed using this method is, therefore, well adapted to the particular configuration of the trajectory for which the optimization is performed. Each design variables’ influence is presented, as well as a set of routing grids generated for trajectories corresponding to different total flight distances, and were further compared with real flight trajectory data retrieved using the website Flight Aware.
The descent of parachute and re-entry capsule in heavy rain has been rarely researched yet. Study of raindrops distribution on canopy surface in heavy rain environment is a key step in the whole research. In this paper, the discrete phase model of two-phase flow approach is applied to simulate the raindrop trajectories in order to research the problem of raindrops distribution on canopy surface when parachute and re-entry capsule are descending in heavy rain. Numerous cases based on different rainfall rates and vertically descending velocities of a simple hemispherical parachute and re-entry capsule are numerically calculated preliminarily. The simulation results are presented, and it is found that the raindrops trapped by the canopy surface are not even-distributed, and raindrops are concentrated near the bottom edges of canopy surface as a result of high-pressure zone enclosed by the parachute; there is a corresponding critical value of descending velocity of parachute and re-entry capsule which determines whether the raindrops will be trapped by the canopy surface for one particular rainfall rate; only above the critical value of descending velocity of parachute and re-entry capsule the raindrops can be trapped by the canopy surface. The conclusions will be of great significance to the further research of the problem of descent of parachute and re-entry capsule in heavy rain.
A proper orthogonal decomposition (POD) technique has been successfully employed to develop reduced-order models for flow control purposes. For complex flows, higher POD modes also play a significant role in the stability and accuracy of the reduced-order model, thus require a closure, as in turbulent flows. In the presence of nonhomogeneous boundary conditions, developing a closure model becomes a challenging task. This paper discusses nonlinear closure modeling approaches for homogeneous and nonhomogeneous boundary conditions. Burgers’ equations, both one-dimensional and two-dimensional, are considered as the governing equations to develop reduced-order models with different boundary conditions. Homogeneous and nonhomogeneous boundary conditions are considered to demonstrate the effectiveness of the proposed closure modeling technique in boundary control applications. Numerical results show that the proposed closure model improves the accuracy of the reduced-order model.
In this paper, an Aircraft Research Flight Simulator equipped with Flight Dynamics Level D (highest level) was used to collect flight test data and develop new controller methodologies. The changes in the aircraft’s mass and center of gravity position are affected by the fuel burn, leading to uncertainties in the aircraft dynamics. A robust controller was designed and optimized using the H method and two different metaheuristic algorithms; in order to ensure acceptable flying qualities within the specified flight envelope despite the presence of uncertainties. The H weighting functions were optimized by using both the genetic algorithm, and the differential evolution algorithm. The differential evolution algorithm revealed high efficiency and gave excellent results in a short time with respect to the genetic algorithm. Good dynamic characteristics for the longitudinal and lateral stability control augmentation systems with a good level of flying qualities were achieved. The optimal controller was used on the Cessna Citation X aircraft linear model for several flight conditions that covered the whole aircraft’s flight envelope. The novelty of the new objective function used in this research is that it combined both time-domain performance criteria and frequency-domain robustness criterion, which led to good level aircraft flying qualities specifications. The use of this new objective function helps to reduce considerably the calculation time of both algorithms, and avoided the use of other computationally more complicated methods. The same fitness function was used in both evolutionary algorithms (differential evolution and genetic algorithm), then their results for the validation of the linear model in the flight points were compared. Finally, robustness analysis was performed to the nonlinear model by varying mass and gravity center position. New tools were developed to validate the results obtained for both linear and nonlinear aircraft models. It was concluded that very good performance of the business Cessna Citation X aircraft was achieved in this research.
The aim of this study is to find the optimal torsional stiffness and trailing-edge flap locations of the helicopter rotor blade for minimum vibration and flap control power at flap lengths of 6% and 9% of the rotor-blade length. A three level orthogonal array based response surface method using polynomial functions is used to describe both vibration and flap control power. Pareto points minimizing hub vibration and flap control power are found at flap lengths of 6% and 9% of the rotor length. This study also explores the variation in rotor hub vibration and flap control power with flying conditions such as the advance ratio and the thrust-to-solidity ratio at the optimum design points. This gives a useful improved design with about a 60% decrease in hub vibration with a penalization of increased flap power at the normal flying regime of rotor-craft flight.
This paper studies the controller design for a spacecraft approaching a tumbling space target in the presence of model uncertainties and dynamic couplings between relative rotation and relative translation. An alternative relative attitude function is selected to measure the magnitude of the relative attitude motion carefully, and robust adaptive controllers are designed for relative attitude and relative position to guarantee a desirable stabilization performance uniformly for the relative motion. The strict Lyapunov analysis is presented to prove the uniformly ultimately boundedness of relative motion errors. The relative attitude controller is directly developed on the special orthogonal group SO(3) to avoid complexities and ambiguities associated with coordinate-dependent attitude representations. Performance of the controlled overall system is demonstrated via a representative numerical example.
Next-generation aircraft and missile are required to have extreme maneuverability, which should maintain stability at high angle of attack. However, the unsteady flow field surrounding the air vehicle would affect the aerodynamics loads and induce unwanted nonlinear motions, of which the rolling motions are usually generated. In order to investigate the unsteady rolling characteristics of a cruciform-finned slender body, the free-to-roll and force measurement tests including particle image velocimetry measurement have been conducted comprehensively. Different types of rolling motion, trimming at equilibrium positions, self-excited rolling oscillation, and self-excited spinning are observed during the free-to-roll experiment depending on the different angles of attack. The force measurement results show that the rolling motions are related to both the static and dynamic stabilities of the rolling moment at balance points. The stabilities of the rolling moment would change as the angle of attack for the model increases. At last, the flow field results from particle image velocimetry measurement indicate that the unsteady rolling motions may be induced by the interaction between the asymmetry vortices and strake wings, fins.
This paper develops a cooperative controller for multiple Unmanned Aerial Vehicles (UAVs) with application to target tracking. The cooperation between the UAVs is established based on an algebraic graph connection and the target information is provided externally by pinning it into a subset of the network. A backstepping-like technique is employed to design a consensus-based controller for each UAV in order to achieve target tracking in 3-D. The proposed controller computes commanded signals for the speed, flight path angle, and heading angle to track the target. The paper considers both the cases of fixed and dynamically changing communication topologies. It is shown that target tracking is achieved for fixed connection topology, if the graph has a directed spanning tree; and for the dynamically changing topology, if the union of the graphs over finite time intervals has a directed spanning tree. The system’s stability is shown using a Lyapunov function-based approach for these cases. All tracking errors are shown to be bounded as long as the target states and its derivatives up to second order are bounded. Detailed numerical simulations further illustrate the controller performance.
R-bar refers to the local vertical axis pointing radially upward in a satellite-fixed reference frame. Approaching a satellite along the R-bar, especially for rendezvous and docking to geostationary satellites, is advantageous in terms of safety considerations and flight time compared to other options. In this paper, a specialized study on autonomous R-bar proximity operations with respect to a geostationary target from a separation of several kilometers to a few hundreds of meters, commonly referred to as the closing phase, is carried out and a comprehensive solution for both attitude and orbit control in this scenario is proposed. An integrative design of the guidance, navigation, and control for R-bar proximity operations is presented. Impulsive R-bar hopping maneuvers are developed for the trajectory guidance. This method is shown to be passively safe and time efficient. The onboard sensors provide measurements of the line-of-sight, range to the target, attitude and angular velocity in the inertial frame. Due to the sensitivity of the sensor’s pointing in the far-range phase, a sliding mode attitude control law is introduced to align the optical axis with the line-of-sight to the target. Sensor measurements are fused and processed by an extended Kalman filter. Simulation results indicate that the proposed integrative guidance, navigation, and control algorithms are robust to uncertainties and noise, and can be used as a comprehensive solution for R-bar rendezvous and docking mission design during the closing phase.
In the field of structural reliability, the estimation of failure probability often requires large numbers of time-consuming performance function calls. It is a great challenge to keep the number of function calls to a minimum extent. The aim of this paper is to propose an approach to assess the structural reliability in an efficient way. The proposed method could be viewed as a hybrid reliability method which combines the advantages of adaptive importance sampling, low-discrepancy sampling and artificial neural network. In the proposed method, artificial neural network is introduced to alleviate the computational burden of deterministic and boring engineering analysis, and its introduction guarantees the computational efficiency of the proposed method. While the Markov chain process is adopted to generate the experimental samples which are used to construct the artificial neural network, the introduction of Markov chain process guarantees the adaptivity of the proposed method and makes the proposed method applicable for various reliability problems. The proposed method is shown to be very efficient as the estimated failure probability is very accurate and only a small number of calls to the actual performance function are needed. The effectiveness and engineering applicability of the proposed method are demonstrated by several test examples.
Tracking multiple objects with multiple sensors is widely recognized to be much more complex than the single-sensor scenario. This contribution proposes a computationally tractable multi-sensor multi-target tracker. Based on Bayes equation and multi-senor observation model, a new corrector for multi-senor is derived. To lower the complexity of update operation, a parallel track-to-measurement association strategy is applied to the corrector. Hypotheses truncation scheme along with first-moment approximation of multi-target density are also employed to improve the tracking efficiency. The tracker is applied to a couple-sensor scenario. Experiment results validate the advantages of proposed method compared to the standard single-sensor -generalized labeled multi-Bernoulli filter and the iterated-corrector probability hypothesis density filter.
In this paper, two methods are proposed, namely the unified processing method and the distributed processing method, to process the global navigation satellite system observation data in integrated navigation simulation which uses the strapdown inertial navigation system as the reference system and the multi-global navigation satellite systems as the sub-systems. The unified processing method takes all the global navigation satellite systems as a whole as one sub-system while the distributed processing method takes each global navigation satellite system as one sub-system. The centralized filter and federated filter are utilized respectively to process the global navigation satellite system observation data in the unified processing method and the distributed processing method. The mathematical models of the unified processing method and the distributed processing method are given in detail. Through theoretical derivation and mathematical simulations, the performances of the unified processing method and the distributed processing method are investigated and compared, showing that while they have the same position (velocity) accuracy, the distributed processing method offers better efficiency than the unified processing method especially when the number of global navigation satellite systems is large (>3).
Using the global exploration and Kriging-based multi-fidelity analysis methods, this study developed a multi-fidelity aerodynamic database for use in the performance analysis of flight vehicles and for use in flight simulations. Athena vortex lattice, a program based on vortex lattice method, was used as the low-fidelity analysis tool in the multi-fidelity analysis method. The in-house high-fidelity AADL-3D code was based on the Navier–Stokes equations. The AADL-3D code was validated by comparing the data and the analysis results of the Onera M-6 wing and NACA TN 3649. The design of experiment method and the Kriging method were applied to integrate low- and high-fidelity analysis results. General data tendencies were established from the low-fidelity analysis results. The high-fidelity analysis results and the Kriging method were used to generate a surrogate model, from which the low-fidelity analysis results were interpolated. To reduce repeated calculations, three design points were simultaneously added for each calculation. The convergence of three design points was avoided by considering only the peak points as additional design points. The reliability of the final surrogate model was determined by applying the leave-one-out cross-validation method and by obtaining the cross-validation root mean square error. Using the multi-fidelity model developed in this study, a multi-fidelity aerodynamic database was constructed for use in the three degrees of freedom flight simulation of flight vehicles.
A unified ductility criterion for fatigue–creep life prediction is presented based on the static fracture toughness exhaustion and dissipated cyclic strain energy density of high temperature components. It provides a general failure criterion for both low and high cycle fatigue regimes. The effects of mean stress, creep and loading waveform on fatigue life are incorporated into this criterion. Applicability and prediction accuracy of the newly proposed criterion was validated through comparing model predictions to experimental results taken from the literature. The results show that the proposed criterion is robust for different loading conditions and more accurate than other existing strain energy/ductility-based methods.
This paper proposes a fast initial acquisition method for light beacons realized by estimating uncertain regions and rectangle honeycomb spiral scanning. First compute with the global position system (GPS) coordinates of both terminals to obtain the azimuth angle and the pitch angle of an optical antenna, which rotates accordingly and roughly points to the light beacon. Then the uncertain regions are estimated by the measurement errors of the GPS coordinates. The optical antenna scans in the uncertain regions to lead the beacon into the field of view of the charge-coupled device and achieve initial acquisition for the beacon light. Besides, analyze the topology, the speed and the coverage areas of the rectangle honeycomb spiral scanning, conduct fast acquisition experiments with a two-dimensional platform in the estimated uncertain regions. The results verify that fast initial acquisition for beacons can be achieved with the proposed method.
With the goal of reducing dependence on ground tracking systems, satellite autonomous navigation technologies are developed quickly in the recent several decades. However, precise orbit determination at high orbital altitudes is an important and challenging problem. In this paper, the nonlinear real-time orbit determination problem is investigated. Combined with satellite dynamical model, extended Kalman filter is explored to estimate satellite orbit parameters. Further, considering errors occur in linearization processing, two improvements for the extended Kalman filter algorithm, i.e. extended Kalman filter-I and extended Kalman filter-II, are proposed based on Lagrange’s mean value theorem, and respectively focus on choosing better linear expansion point and Jacobian matrix calculation point. Extensive simulations show that extended Kalman filter-I and extended Kalman filter-II significantly enhance orbit accuracy, compared with extended Kalman filter. And the increases in calculation complexity are acceptable. Finally, the robustness of extended Kalman filter-I and extended Kalman filter-II is analyzed by given different initial position errors, and results show that extended Kalman filter-I and extended Kalman filter-II have better robustness than extended Kalman filter.
This study presents a new guidance and control system using a constrained adaptive backstepping method for a space transportation system. In this method, the effects of input saturations by actuator dynamics (e.g. magnitude, rate and bandwidth) are considered to introduce the compensators on the basis of pseudo control hedging. The stability of the proposed entire system is guaranteed by the Lyapunov’ stability theorem. To confirm the realization and robustness of the proposed system, Monte Carlo simulations (MCSs) were performed. In addition, to obtain optimized control performance, a parameter optimization algorithm combined with the MCSs was introduced. Finally, automatic landing simulations using the six degrees-of-freedom nonlinear flight simulation model of the NASA’s Space Shuttle Orbiter were performed to verify the effectiveness of the proposed technique.
This paper deals with the active vibration control of smart truss structure. First, the electro-mechanical coupled dynamic model of the smart structure is constructed. Then, the first-order ordinary differential equation of the control system is presented. After that, an online learning fuzzy control (OLFC) algorithm is proposed to control the structure vibrations. The OLFC algorithm is composed of a reward function, a Q learning algorithm, a rule base generator and a conventional fuzzy controller. The OLFC algorithm learns the rule base by interaction with the plant, and changes rule base generate policy via evaluative reward signal to realize the learning goal. The algorithm only needs little information about the plant to design the reward function. In order to prove the effectiveness of the proposed control algorithm, control responses are presented and compared with conventional fuzzy control method.
Despite the existence of many studies about the structural analysis of a square solar sail, the need for obtaining reliable numerical results still poses a number of practical issues to be solved. The aim of this paper is to propose a new method that improves the existing analysis techniques. In this sense, the solar sail is modeled using distributed sail-boom connections, and its structural behavior in free flight is studied, using the inertia relief method, at different incidence angles of the incoming solar radiation. The proposed approach is able to circumvent the onset of numerical convergence problems by means of suitable strategies. A nonlinear analysis is carried out starting from an initial geometrical configuration in which the whole solar sail is perturbed using a linear combination of the first global buckling modes, obtained with a static eigenvalue analysis. Key points of the procedure are the application of a correct sail pre-stress, a clever choice of the type of elements to be used in the finite element analysis and the use of a suitable mesh refinement. The performance of the new approach have been successfully tested on square solar sails with side length varying from relatively small to medium-to-large sizes, in the range of 10–100 m. A detailed analysis is presented for a reference 20 m x 20 m square solar sail, where the paper shows that the suggested procedure is able to guarantee accurate results without the need of additional stabilization technique. In particular, the vibration global mode shapes and frequencies of the solar sail are correctly described even in presence of unsymmetrical loading conditions. In other terms, the numerical analysis is completed without any convergence problem and any disturbing local modes.
In this work, the Uncertainty and Disturbance Estimation (UDE) approach is employed for the design of a robust flight controller for high performance aircraft. The UDE estimated uncertainty is used to robustify an input–output linearization-based controller designed for nominal system. The UDE theory employing a first-order filter and α-filter is used for the estimation of the composite uncertainty. To address the issue of output derivatives, an observer which too employs the UDE estimated uncertainties for the purpose of robustness is designed. Closed-loop stability for the overall controller-observer system is established. Uncertainties and disturbances in terms of parametric variations, wind gust, unmodeled actuator dynamics and measurement noise are considered to evaluate and compare performances of the two filter-based designs. Further, a performance comparison with some of the existing designs is carried out and the results are presented to demonstrate the efficacy of the proposed work. A notable feature of the proposed design is that the approach neither requires an accurate plant model nor any information about the uncertainties and disturbances.
The nonlinear attitude motion equations of flexible spacecraft described by the Euler angles are expressed in the vector form. Based on dynamic surface control, a new robust dynamic surface sliding mode controller is proposed for the attitude tracking and active vibration suppression of flexible spacecraft in the presence of parameter uncertainty and external disturbances. Then, a novel robust dynamic surface finite time sliding mode controller is proposed with an extended state observer such that the uncertainties can be estimated. Lyapunov stability analyses show that the two controllers can guarantee the asymptotical stability of the attitude control system. The undesirable vibration of flexible spacecraft is also actively suppressed by the modal velocity feedback approach. Finally, simulation results verified the effectiveness of the presented control algorithms.
Manufacturers often develop new products by modifying and extending existing products in order to achieve new market demands while minimizing development time and manufacturing costs. In this research, an efficient derivative design process was developed to efficiently adapt existing aircraft designs according to new requirements. The proposed design process was evaluated using a case study that derives an unmanned aerial vehicle design from a baseline manned 2-seatlight sport aircraft. Multiple unmanned aerial vehicle operational scenarios were analysed to define the requirements of the derivative aircraft. These included patrol, environmental monitoring, and communications relay missions. Each mission has different requirements and therefore each resulting derivative unmanned aerial vehicle design has different geometry, devices, and performance. The derivative design process involved redefining the design requirements and identifying the minimum design variable set that needed to be considered in order to efficiently adapt the baseline design. Uncertainty was considered as well to enhance the reliability of the optimized result when it considered different conditions for each mission. An optimization method based on the possibility based design optimization was proposed to handle uncertainty that arises in the design requirements for the multi-role nature of unmanned aerial vehicles. In this paper, the possibility based design optimization method was implemented with multidisciplinary design optimization technique to derive the derivative unmanned designs based on originally manned aircraft. This approach prevented constraint violation via uncertainty variations in the operating altitude and payload weight for each. The unmanned aerial vehicle derivative designs satisfying the requirements of three different missions were derived from the proposed design process.
In this paper, fault-tolerant attitude tracking control problem is investigated for multiple spacecraft formation flying system with external disturbance, actuator saturation, and faults. A quaternion-based adaptive fault-tolerant control law is proposed based on input normalized neural network. The desired nonlinear smooth function is approximated by using input normalized neural network with an adaptive learning algorithm, and no prior knowledge about spacecraft dynamics is required. Meanwhile, in order to guarantee that the output of input normalized neural network used in the controller is bounded by the corresponding bound of the approximated unknown function, a modified adaptive law is designed to revise the sliding mode manifold. Moreover, the stability of system can be guaranteed by Lyapunov theory. Finally, the validity of the proposed control algorithm is verified through numerical simulations.
Based on L2 optimal control allocation, an autopilot design approach is proposed for the missile with aerodynamic control surfaces and reaction jets. The control system involves a control allocator and a virtual control law. A robust sliding sector with a parameter update law is proposed to deal with unmatched parameter uncertainties and unknown disturbances in the system. Then a control law is designed to produce virtual control effort signals by using the robust adaptive sliding sector. In order to distribute the virtual signals to the aerodynamic control surfaces and reaction jets on the missile, a control allocator is designed by L2 optimal control allocation strategy. Simulation results show that the missile control system tracks the acceleration command fast and smoothly. In the tracking process, aerodynamic control surfaces cooperate with reaction jets, verifying the effectiveness of the proposed approach.
The measurement of spacecraft electrical characteristics and multi-label classification issues are generally including a large amount of unlabeled test data processing, high-dimensional feature redundancy, time-consumed computation, and identification of slow rate. In this paper, a fuzzy c-means offline (FCM) clustering algorithm and the approximate weighted proximal support vector machine (WPSVM) online recognition approach have been proposed to reduce the feature size and improve the speed of classification of electrical characteristics in the spacecraft. In addition, the main component analysis for the complex signals based on the principal component feature extraction is used for the feature selection process. The data capture contribution approach by using thresholds is furthermore applied to resolve the selection problem of the principal component analysis (PCA), which effectively guarantees the validity and consistency of the data. Experimental results indicate that the proposed approach in this paper can obtain better fault diagnosis results of the spacecraft electrical characteristics’ data, improve the accuracy of identification, and shorten the computing time with high efficiency.
A research for designing the optimal lunar vertical landing trajectory to reduce the total energy or mass of propellant is addressed in this paper. Most of these problems can be divided into two phases: breaking and approach phase. The optimal landing trajectory in general does not consider the pitch-up motion so that the landing problem has been only solved in the breaking phase. For this reason, there are some attempts to find the optimal trajectory including the final vertical landing phase by including the equations of angular motion of the vehicle. However, the optimal solution using this approach depends on the scale factor of a cost function because the cost function consists of two different mechanical parameters such as the final mass and total control torque. The final control constraints are augmented for vertical lunar landing instead of the equations of angular motion. The obtained optimal trajectory has an additional positive effect of the image acquisition as well as the final vertical landing.
This paper presents a trajectory parameterization method for calculating emergency flight paths with variable airspeeds under conditions of constant wind. The method is based on the Dubins curve; however, it has been modified to allow for acceleration along the path and finite rate of change in turn rate. The aircraft’s planar trajectory from an initial condition to a terminal condition is parameterized into a small set of path-defining variables. The method uses a number of closed-form solutions and simple iteration schemes to efficiently calculate a path that meets the specified constraints. The parametrized path can then be optimized to minimize a performance objective for real-time emergency path planning. For emergency flight planning, the vertical degree of freedom is treated as a function of the aircraft state and parametric controls, and the optimization is formulated to ensure touchdown at a desired location and aircraft state. The performance of the proposed method is investigated using several test cases, including landing of a commercial jet following total loss of thrust and autorotative recovery of a utility helicopter following total loss of power.
The shock waves are important phenomena in transonic turbines, which cause lots of negative effects on the aerodynamic performance. Much of attention had been paid on reducing the strength of the shock waves via modifying turbine cascade geometry, and it is highly preferred to build experiences on the relationship between the cascade aerodynamic performance and the geometric parameters. The paper presents a numerical study on the aerodynamic optimal transonic turbine cascade and its geometry characteristics. Three typical Russia transonic turbine cascades with different design conditions are selected and optimized using adjoint method at three different back pressures, respectively. Thus, the best geometry parameters for optimum aerodynamic performance can be found. Then the key geometry parameters of optimized cascades are extracted and compared with the original ones. Results show that even the best designs by hands could be less efficient than ones by computer-aided optimizations. Some experiences on how to set the key geometry parameters for a best performance are obtained. The reduced shock profiling is applied to the thermal turbomachinery and machine dynamics transonic turbine by using the adjoint method. The performance of the thermal turbomachinery and machine dynamics transonic turbine was increased significantly.
The theories of circular plates with large deflection and fracture mechanics were employed to investigate the calculating formula of stress intensity factor on prefab gap. The relationship of the opening pressure with physical dimensions of hard pulse separation device (PSD) was analyzed to achieve the feasible configuration of metal diaphragm, which could play a very important role in dual-pulse solid rocket motor. Moreover, six specific single-term tests of the metal diaphragm were conducted to validate the computing precision of designing formula. In addition, three kinds of scale experiments, e.g. the bearing test, opening test, and associated test were performed to explore the working characters of hard pulse separation device. The results indicate that the error between the experimental results and designing results is 4.4%, and the method can be applied to design the physical dimensions of metal diaphragm. The performances of bearing ability, sealability, opening and melt of PSD have been demonstrated. The results show that this kind of PSD can satisfy the requirements of dual-pulse solid rocket motor very well.
Experimental investigations were carried out to study the wake characteristics of a pitching supercritical airfoil at Mach number of 0.6. Flow field inside the wake was measured by a hot-wire anemometry at downstream distances from trailing edge of 0.25 and 0.5 times the chord length. All data were taken at mean incidence angles of 0 and 3°. The amplitudes of oscillation were 1 and 3° while the oscillation frequencies were 3 and 6 Hz. Output signals acquired from sensors were analyzed besides the effects of such parameters as frequency and oscillation amplitude. Moreover, a comprehensive numerical study was carried out for the same airfoil under similar experimental test conditions; then, the results of numerical simulations were analyzed and compared with those of experimental tests. Results of the present research could be summarized as: observation of hysteresis and how it is affected by frequency and amplitude variations, observation of increasing turbulence intensity by root mean square investigation and also increasing signal energy by means of power spectral density diagram for those sensors lied inside the wake, and finally, study of correlation between wake’s interior sensors and exterior ones.
Computational fluid dynamic and order reducing methods have been extensively applied to predict the flutter onset speed of several types of aircrafts. However, the accuracy required by certification standards still ascribes flight testing as the only method available that safely validates the flight envelope of an aircraft. In particular, free-flutter conditions must be demonstrated in the target flight envelope, and several methods have been developed to determine the flutter onset speed in real-time when expanding the envelope during flight testing. Among the methods, the damping versus velocity technique combined with a flutter margin implementation remains the most common technique used for envelope expansion. Even with the popularity and "easy to implement" characteristics of this method, several shortcomings can adversely affect the identification of non-stable conditions during envelope expansion. Notably, the limit cycle oscillations conditions, distinct from flutter, cannot be accurately identified. This study proposes to apply a similar methodology to the flutter margin to anticipate limit cycle oscillations associated with freeplay in the plunge axis of a bi-dimensional airfoil that is aeroelastically representative of the tested aircraft. Analytical considerations are conducted to support this new approach, and a computer model is used to validate the proposed methodology.
When the stagnation pressure of a perfect gas increases, the specific heat and their ratio do not remain constant anymore and start to vary with this pressure. The gas does not stay perfect. Its state equation change and it becomes for a real gas. In this case, the effects of molecular size and intermolecular attraction forces intervene to correct the state equation, the thermodynamic parameters and the value of Prandtl Meyer function. The aim of this work is developing a new form of Prandtl Meyer function based on those assumptions; and determining the effect of stagnation pressure on this function. With the assumptions that Berthelot’s state equation accounts for molecular size and intermolecular force effects, expressions are developed for analysing the supersonic flow for thermally and calorically imperfect gas lower than the dissociation molecules threshold. The supersonic parameters depend directly on the stagnation parameters of the combustion chamber. The application is for air. A computation of error was made in this case to give a limit of the perfect gas and the high temperature models compared to the real gas model.
The current overloaded airspace near large airports leads the stakeholders to develop more fundamental means to improve the use of available air capacity in dense traffic areas as terminal maneuvering area. A smart planning of aircraft trajectories and an appropriate designing of route structure can be a solution. Therefore, to analyze the capability of them, this paper proposes a novel approach to deal with the merging and sequencing problem of arrival traffic in terminal maneuvering area, and examines the performance of diverse conflict resolution strategies and different terminal maneuvering area route structure. An evolutionary algorithm is employed to resolve potential conflicts, altering one or more of the following decision variables: routes, speeds, and entrance time to terminal maneuvering area (slot). Two different fitness functions were developed. One aims to minimize potential conflicts and the other one aims to minimize at the same time potential conflicts and delay induced at runway. The performance of different combinations of decision variables, or conflict resolution strategies, is evaluated through a set of arrival aircraft to Grand Canaria airport. Also, four different route structures are tested by using the same set of incoming aircraft. The results showed that the slot and speed change and mix of all change strategies can successfully resolve all conflicts in an efficient manner. The results also suggested that topologies with one main merging point are more efficient than the ones that have two merging points. Finally, a significant abatement of delay was observed when minimizing conflicts and delay while achieving conflict-free solutions.
Atmospheric icing problem is considered as one of the major hazards to the aviation industry and the existing anti-icing system has numerous disadvantages within it. The recent studies on the super hydrophobic surfaces inspired from lotus leaf are found to be an emerging trend in the ice protection system because of its resistance to the ice formation. But the experimental procedure for creating those surfaces is expensive and it requires specialised equipments. The work of Erbil et al. is identified to be a cost-effective and simple method for creating those surfaces based on the process called phase separation. The experimental investigation of this alternate coating method is scheduled and conducted by creating the polypropylene coating on the aluminium sheet metal. The coated aluminium surface is then tested for its ice repelling property by spraying the super-cooled liquid droplets similar to Rime ice formation. It has shown few excellent results by repelling the water and the ice formation under various atmospheric conditions. Further, the wing model and the coated layer are designed using computer software to analyse the influence of anti-icing coating on the basic stress–strain and flow properties. This novel approach provides significant cost and weight savings through a roadmap for replacing the existing ice protection systems.
Reduced drag, increased lift and, consequently, increased vital ratio and lift-to-drag coefficients are crucial in almost all efficient micro air vehicles. Riblet geometries use a variety of air vehicles. Further investigation on micro air vehicles is, however, necessary for enhanced development. Rectangular riblets on a rectangular micro air vehicle are computationally investigated. In this study, the governing equation of fluid flow is solved numerically; the turbulent model around the NACA S5020 airfoil section is covered by riblets either on both sides or on the upper side of the wings. Results show a difference of behavior in drag reduction due to the angle of attack on the airfoil. When the lift-to-drag coefficient of an angle of attack is at its maximum, an improvement can be observed, where lift-to-drag ratio increases, and drag decreases. Results for the two-side riblets show an increase in the lift-to-drag ratio as well; although the lift-to-drag coefficient and the drag reduction of riblets on both sides were comparatively less than that for riblets on the upside.
The two-dimensional through-flow modeling of turbomachinery is still one of the most powerful tools available to the turbomachinery industry for aerodynamic design, analysis, and post-processing of test data due to its robustness and speed. Although variety of aspects of such a modeling approach are discussed in the publicly available literature for compressors and turbines, not much emphasis is placed on combined modeling of the fan and the downstream splitter of turbofan engines. The current article addresses this void by presenting a streamline curvature through-flow methodology that is suitable for inverse design for such a problem. A new split-flow method for the streamline solver, alternative to the publicly available analysis-oriented method, is implemented and initially compared with two-dimensional axisymmetric computational fluid dynamics on two representative geometries for high and low bypass ratios. The empirical models for incidence, deviation, loss, and end-wall blockage are compiled from the literature and calibrated against two test cases: experimental data of NASA two-stage fan and three-dimensional computational fluid dynamics of a custom-designed transonic fan stage. Finally, experimental validation against GE-NASA bypass fan case is accomplished to validate the complete methodology. The proposed method is a simple extension of streamline curvature method and can be applied to existing compressor methodologies with minimum numerical effort.
This paper studies the 3D flocking control problem for unmanned aerial vehicle swarm when tracking a desired trajectory. In order to allow the unmanned aerial vehicle swarm to form the stable flocking geometry on a same horizontal plane, the altitude consensus algorithm is applied to the unmanned aerial vehicle altitude control channel, using the trajectory altitude as the external input signals. The flocking control algorithm is only performed in the horizontal channel to control the horizontal position of unmanned aerial vehicles. The distributed tracking algorithm, which controls the local averages of position and velocity of each unmanned aerial vehicle, is implemented to achieve the better tracking performance. The improved artificial potential field method is introduced to achieve the smooth trajectory when avoiding obstacles. The practical dynamic and constraints of unmanned aerial vehicles are also taken into account. Numerical simulations are performed to test the performance of the proposed control algorithm.
Global tracking performance analysis in near-real time for the space-based missile warning system is an architecture level demand for the system designers. The traditional method which based on Monte Carlo simulations cannot meet the demand for its time-consuming nature. In this paper, non-simulation rapid performance prediction methods in stereopsis and monocular observation conditions are studied. The latter, though inevitable, has not been taken into account in other literatures. Furthermore, we also find the geometrical symmetry for monocular observation and propose a rapid analysis method called compass of performance based on this symmetry.
In this paper, the relative navigation technique of final approach phase for a tumbling target spacecraft is studied and exploited. It is assumed that the tumbling target is in failure or out of control and there is no good a priori rotation rate information. The Euler’s rotational dynamics is used to propagate the target angular velocity, and the unknown inertia parameter circumstance is also considered. The chaser spacecraft is equipped with three strapdown gyros and accelerometers and a star sensor that determine the absolute motion parameters, and an optical camera that measures relative azimuth and elevation angles to the target spacecraft. On the basis of the rotational and translational motions of both spacecrafts, an angles-only/ strapdown inertial navigation system/celestial navigation system navigation filter is designed. Simulation results indicate that the proposed algorithm can accurately estimate the relative position, velocity, and attitude between two spacecrafts and compensate the biases of the gyros and accelerometers.
This paper considers the situation awareness function associated with an unmanned aerial vehicle arriving at an uncontrolled airfield. Given no air traffic control service available within such a terminal area, the unmanned aerial vehicle needs to establish a good level of situation awareness by using its onboard sensors to detect and track other traffic aircraft. Comparing to the existing works which mainly use sensor observations in the filtering process, this paper exploits the circuit flight rules to provide extra knowledge about the target behaviour. This is achieved by using multiple models to describe the target motions in different flight phases and characterising the phase transition in a stochastic manner. Consequently, an interacting multiple model particle filter with state-dependent transition probabilities is developed to provide the required situation awareness function.
An automated multi-grid overset grid algorithm is presented for accurate simulation of viscous flows around complex configurations. The algorithm is based on the implicit hole cutting technique optimized with an overset grid construction strategy, a grid cutting criterion and a multi-level overset grid cutting method. The enhanced method is more general than the original method, while preserving the high degree of automation of the implicit hole cutting algorithm. Moreover, a mesh sequencing and multi-grid Chimera strategy is proposed to achieve convergence acceleration of the flow calculations. The present Chimera algorithm is demonstrated through the solution of the flows over the NASA Common Research Model configurations. The results show that the present mesh sequencing and multi-grid strategy in the overset grid framework are very effective in increasing convergence of solution in comparison with the standard three-level multi-grid. Wing pressure comparisons indicate the pressure distribution varies with grid solution at the shock, and a higher grid resolution improves shock definition. Abrupt reductions in lift and drag coefficients are correlated with the side-of-body separation bubble as angle of attack is increased. And various modeling and grid resolution have a strong impact on predicting the side-of-body separation. The side-of-body separation decreases with grid refinement. And the bubble size from spalart-allmaras (SA) modeling is larger than that from shear stress transport model.
The layout design problem of a propulsion system is complex and time-consuming process. This is mainly due to geometrical and performance constraints and system requirements. In addition, layout design optimization of a space propulsion system is non-linear, non-convex, and multimodal, which makes it difficult to implement conventional optimization methods to this class of design problems. This paper presents a hybrid optimization algorithm using genetic algorithm and sequential quadratic programming for optimal layout design of a space propulsion system. Previous research works mainly focused on the layout design components with constant parts. However, the approach adopted in this paper involves both variable mass component and hybrid optimization algorithm (GA-SQP) of a space propulsion system. The proposed hybrid optimization algorithm explores globally the design search space to locate the most promising region using genetic algorithm, whereas gradient-based sequential quadratic programming algorithm is used to reduce the computational time with a high degree of accuracy. The results obtained show that the proposed method provides an effective way of solving layout design optimization problem using both variable mass components method and a hybrid optimization for optimal layout design of a space propulsion system.
The emphasis of this paper lies in the development of an efficient approach to reproduce the behaviors of the scramjet-powered hypersonic system with high fidelity. The modeling of the dual-mode scramjet powered hypersonic vehicle dynamics with shock interaction, Ram-to-Scram transition, and finite rate chemistry reaction is firstly introduced. The structure of surrogate model is identified with the implement of iterative fractional factorial design (IFFD). In order to declare the reliability of the surrogate models, -gap metric is applied to distinguish the difference among these surrogate models in terms of closed-loop performance. The results show that the influence of Mach number on the aerodynamics should not be overlooked, and the effect of propulsion system to the aerodynamic pitch moment is dramatic. Further, the partial Kriging model appears to have the closest plants throughout the flight envelope compared with the full Kriging model and polynomials model. Nevertheless, considering the briefness of analytical expression, the polynomials model may be an alternative approach for design-oriented modeling.
Fairings are provided to cover hypersonic air breathing vehicle to protect it from adverse aerodynamic loading and kinetic heating. Separation dynamics of fairings is an important event in the launch of vehicle. Extensive computational fluid dynamics simulations are carried out for the design of fairings and vehicle and selection of time sequences of various separation events. A ground test of fairing separation is conducted in the sled facility to check the structural integrity and functionality of various separation mechanisms and flight hardware. Simulations have been carried out to study the separation dynamics of fairings at test conditions using grid-free Euler solver to get the aerodynamic loads and the loads are integrated to get the trajectory of fairings. The aerodynamic loads are provided to verify the structural integrity of various components and the trajectory of panels is used in the test planning. The pressure distributions on the vehicle are compared with the test results.
The paper deals with the vibration suppression of a bladed disk structure with a piezoelectric network. The piezoelectric network without inductors has a different period (so-called bi-period) from that of the bladed disk. The research focuses on reducing amplified response or localized vibration, a phenomenon existing in bladed disks and induced by mistuning. The study is based on an electromechanical and cyclic-periodic lumped parameter model with two degrees of freedom per sector. Both mechanical mistuning and electrical mistuning have been taken into account. The modified modal assurance criterion is adopted to evaluate the ability of bi-periodic piezoelectric networks for suppressing vibration. The Monte Carlo simulation is used to calculate the modified modal assurance criterion of the system with random mistuning. To validate the numerical results, an experiment research is also carried out. In order to perform a comparison between numerical results and experimental results, the method of equivalent blisk model is introduced to identify the lumped parameters of the experimental model. In the experiment, traveling wave excitations are simulated as dynamical loads to excite the resonant vibration of coupling bladed disks in a rotating state. The results show that with a good design, the bi-periodic piezoelectric network can effectively suppress the amplified-forced-response due to the mistuning of the system, even though there are less piezoelectric patches and electric elements in the system compared to the system with piezoelectric shunt circuits.
This paper deals with the study of the flapwise bending vibration and dynamic stability of rotating functionally graded material plates in thermal environments. A finite element formulation is derived for modal and dynamic stability analyses of rotating functionally graded material plates using first-order shear deformation theory. Temperature-dependent material properties of the plates are considered in the analysis and a simple power law is assumed for composition of constituent materials to vary along the thickness direction. The same power law is also proposed in thermal environments for temperature variation across the thickness of the plate. Some numerical results obtained from the present method are compared with numerical results available in the literature and are found to be in good agreement. Parametric investigation is carried out thoroughly to study the effect of the temperature rise, hub radius, and rotational speed on vibration and the dynamic stability of rotating plate in thermal environment. Bolotin’s method is used to generate the boundaries of stability and instability regions. These instability regions are plotted in the parameter space with the nondimensional dynamic load and excitation frequency. It is observed that the natural frequencies reduce with an increase in temperature rise. Increase in rotational speed and hub radius results in increase of natural frequencies of vibration. The rise in temperature leads to reduction in the dynamic stability of plate. Increase in rotational speed and hub radius enhances the dynamic stability of the rotating plate.
Combustion characteristics in a supersonic combustor with ethylene injection upstream of dual parallel cavities were investigated experimentally in a direct-connected test rig with inflow conditions of Ma = 3.46, Po = 3.60 Mpa and To = 1430 K. The combustor has a two-dimensional rectangular configuration with single-side expansion. Two open cavities with the same size were mounted on the expanded and horizontal walls, respectively. Static pressure distribution in the axial direction was measured along the centerline of the expanded wall. High-speed flame luminosity and schlieren were used to capture the combustion and flow structures at different equivalence ratios. Two clusters of separated and asymmetric flames were found to be stabilized near the dual parallel cavities in all tests. The flame and flow characteristics changed with the combustion modes. For scramjet mode, no obvious flow separation occurred near the walls and the two flames were both stabilized in the cavity shear layer and recirculation zone. For ramjet mode, the high back pressure resulting from intense heat release induced a large-scale recirculation zone upstream of the cavity mounted on the expanded wall, which supplied a favorable combustion condition and the flame was stabilized in the jet-wake. Meanwhile, there was no obvious separation near the horizontal wall, with the local flame stabilized in the cavity shear layer. It is suggested that the combustion near the horizontal wall should be enhanced to improve the combustion performance and avoid a non-uniform flow field at the combustor exit.
Successful navigation of small, unmanned aerial vehicles (UAVs) in cluttered environments is a challenging task, especially in the presence of turbulent winds and state estimation uncertainty. This paper proposes a probabilistic path planner for UAVs operating in cluttered environments. Unlike previous sampling-based approaches which select robust paths from a set of trajectory candidates, the proposed algorithm seeks to modify an initial desired path so that it satisfies obstacle avoidance constraints. Given a desired path, Monte Carlo uncertainty propagation is performed and obstacle collision risks are quantified at discrete intervals along the trajectory. A numerical optimization algorithm is used to modify the flight path around obstacles and reduce probability of collision while maintaining as much of the originally desired path as possible. The proposed path planner is specifically designed to leverage embedded massively parallel computers for near real-time uncertainty propagation. Thus the planner can be run in real-time in a feedback manner, modifying the path appropriately as new measurements are obtained. Example results for a standard quadrotor show the ability of the path planning scheme to successfully generate trajectories in cluttered environments. Trade studies characterize algorithm performance as a function of obstacle density and collision risk acceptability.
Orientation accuracy is a key factor in the design of mechanisms for antenna pointing. Our design uses a redundantly actuated parallel mechanism which may provide an effective way to solve this problem, and even can increase its payload capability and reliability. The presented mechanism can be driven by rotary motors fixed on the base to reduce the inertia of the moving parts and to lower the power consumption. The mechanism is redundantly actuated by three arms, and is used as a two-dimensional antenna tracking and pointing device. Both the forward and inverse kinematics are investigated to find all the possible solutions. Detailed characters of the platform are analyzed to demonstrate the advantages in eliminating singularities and improving pointing accuracy. A method of calculating the overconstrained orientational error is also proposed based on the differential kinematics. All the methods are verified by numerical examples.
The noise radiation characteristics of two-phase pulse detonation combustor and pulse detonation turbine engine were investigated under different operating frequencies utilizing gasoline as fuel and air as oxidizer. The sound pressure data of noise radiation were presented for both single-pulse detonation combustor tube and pulse detonation turbine engine. The experimental results implied that the peak sound pressure level of PDTE exit with inner diameter being 60 mm was about 157 dB under the operating frequencies which varied from 5 Hz to 25 Hz, while the peak sound pressure level of single-pulse detonation combustor tube exit was about 170 dB under the same condition. The far-field jet-noise measurements of the pulse detonation turbine engine showed that radial turbine interacting with the pulse detonation combustor could decrease the peak sound pressure level of pulse detonation combustor with the maximum acoustic attenuation being approximately 14.2 dB for the current test conditions, which could be contributed to the energy extraction by the radial turbine from the pulse detonation combustor exhaust flow. The sound pressure level of both pulse detonation combustor and pulse detonation turbine engine exit was function of directivity angle from the exhaust centerline. In all the experiments, the equivalence ratio of gasoline/air mixture and the fill fraction were 1.2 and 1.0, respectively.
Numerical analysis was conducted of a jaws inlet under different working conditions, including angles of attack of 0° and 3°, varying Mach number, and varying back pressure with a constant-area isolator, to investigate its performance and flow fields of starting and unstarting states. Results reveal that the jaws inlet has an enhanced flow capture capability in starting states, with the mass capture ratio higher than 0.96, but relatively reduced working range of inflow Mach numbers. Its performance at a low Mach number is better than that at a high Mach number. Non-uniform flow fields are observed in unstarting cases at low Mach numbers and high back pressures, while separation structures are confined in the pitching compression section. Further increase in Mach number or decrease in back pressure does not result in significant changes in the separation structures. In the unstarting case under a high back pressure, it is hard to achieve restarting through reductions in the back pressure.
NextGen, SESAR, and other international projects intend to develop the coming air traffic management (ATM) environment to cope with the present problems, including e.g. capacity, efficiency, safety, and environmental impact. These projects also introduce considerable changes in the role of air traffic controllers (ATCOs), and shift some of the present tasks from ground to on-board (pilots) and diversify automation. Due to the consequential effects of these developments, (i) the models describing the present operational circumstances, (ii) the working conditions, and (iii) the working environment (e.g. human–machine interactions, task-work and mental loads, cognitive aspects, situation awareness, decision support) must be reconsidered and adapted to the future role of operators. This paper aims to describe a new approach to model the ATCOs’ working processes, and also introduces a new concept for the ATCOs’ future working environment. The developed model covers the situation awareness and decision making processes, including the subjective aspects of the ATCOs decision making based on their workload, mental condition, knowledge, or practice. The proposed new working environment contains special information displays, the decision support system and the ATCOs workload and mental condition measurement / monitoring system, based on objective, on-line, continuous, and nonintrusive measurement techniques. Some of the core systems developed in collaboration with HungaroControl were exhibited at the World ATM congress in 2014 and 2015.
This paper proposes the adoption of a set of ground support tools to assist the approach controller on issuing and monitoring spacing instructions where interval management operations are carried out. The algorithmic basis of these tools, based on computing feasible maneuvers to achieve the desired spacing and monitoring their safe execution, is described. These tools assume the use of suitable airborne spacing–flight deck interval management systems, and exploit them. These controller support tools comprise two phases: (i) the design of the intended maneuver, making use of simplified geometrical calculations to compute required routes and speeds; and (ii) the monitoring of the safe execution of the intended merging operations. The designed tools have been tested developing a simulator that comprises the complete process. The simulator first generates a set of feasible maneuvers to allow the controllers the selection of the most appropriate one, then simulates the selected maneuver and secondly monitors it to ensure safety and correctness of the operation. The simulation takes into account inaccuracies such as inexact weather (wind) forecast and errors in Automatic Dependent Surveillance – Broadcast measures.
The paper presents a set of visual aids for enhancing remote pilot perception of potential violations of allowed fly areas or conflicts with conventional air traffic when operating remotely piloted aerial vehicles. Assuming a video stream from an on-board camera is available to the remote pilot, visual aids are provided in a head-up display modality by means of reality augmentation techniques. The main visual element consists of a dynamic set of fences allowing for a safe aircraft separation away from no-fly zones and from neighboring vehicles. The shape of the fences is varied according to aircraft current speed and altitude, in order to allow for a sufficient safety margin, also in case of a loss-of-control accident. As a further visual aid, the predicted future position of the aircraft is determined and fence color is changed in order to highlight potential violations of the allowed operational area. The proposed methodology is validated by means of simulations in a reference operational scenario. Results demonstrate the effectiveness of the proposed approach in improving pilot awareness.
In the optimal trajectory design problem for the lunar powered-descent phase, the periapsis of the de-orbit burn phase is usually chosen as a starting point or an initial state. The resultant trajectory in these cases shows that the altitude of the lander increases to gain more sufficient time to reduce large initial horizontal velocity. However, the periapsis of the de-orbit burn phase is not the optimal choice. In this study, the optimal initial phase angle can be found by applying the modified trajectory-optimization problem, where the initial state is considered as a free variable. In this problem, any additional assumption and change in hardware compared with the traditional optimal lunar-landing problem are not imposed except for the initial phase angle. Using the proposed numerical approach, we show that the optimal phase angle is not always equivalent to the periapsis, and fuel consumption can be reduced by changing the starting point of the powered descent phase.
This paper proposes a new traffic alert and collision avoidance system (TCAS) for general aviation (GA) that uses automatic dependent surveillance-broadcast (ADS-B). The proposed GA-TCAS alerts the pilot of upcoming potential conflicts, which are becoming increasingly common as the number of GA movements increases. Conflict detection and avoidance uses the ADS-B and GA maneuvers. The algorithm selects the angle according to the intruder’s heading and position in order to reduce the nuisance alerts. The detection algorithm has three stages: proximate advisory, traffic advisory, and resolution advisory. The relationship between the turn angle and the horizontal miss distance (HMD) are calculated to determine a resolution when an intruder enters the separation bubble. Using the encounter models for visual flight rule (VFR) to evaluate the performance of this TCAS algorithm, the resolution should follow the VFR and prioritize in different situations. In flight simulations, the proposed GA-TCAS is tested using real ultralight flight data. The system performance is demonstrated using flight trajectories from conflict detection and resolution.
Coning motion of spinning carriers is a complex rotating motion with various forms that include single circular motion, double circular motion and multiple circular motion. Due to the fact that it is difficult to describe the real coning motion of double circular motion and multiple circular motion by the common method of attack angle and sideslip angle (A–S), a cone frame and cone angles are defined to describe coning motion. Through analysis of measuring the relationship between coning motion and inertial devices such as gyroscopes and accelerometers, an inertial measuring method is proposed to build the measuring equation and resolving equation. A geometry-solving algorithm of real coning motion is derived in detail, and radiuses of large circle and real cone circle are obtained as well. A flight simulation of a spinning carrier with coning motion is designed and used to verify the measuring method and the geometry-solving algorithm. The result shows that: (1) the inertial measuring method has the same validity as A–S method to describe coning motion, but is superior to A–S method for the reason of providing the rotation information of carriers; (2) due to coupling relationship, the rotating angle is equal to the subtraction of roll angle and precession angle; (3) the real precession angle and the real nutation angle are calculated by the geometry-solving algorithm, and the real coning motion is obtained finally.
A three-dimensional full-annulus computational model for aerodynamic damping prediction based on energy method is demonstrated by the consideration of a transonic rotor which is designed for experimental research of flutter to improve understanding the influence of frequency mistuning and inter-blade phase angle mistuning on aeroelastic stability of transonic compressor. Each individual blade is capable of vibration with its own independent frequency and phase angle, thus modeling a full-annulus rotor with arbitrary mistuned blades. The numerical analyses of alternate, random, linear frequency mistuning, and random inter-blade phase angle mistuning are performed in detail. The studies of tuned rotor show that the aerodynamic damping for the blade first bending mode is most sensitive to inter-blade phase angle. Further investigations on various frequency mistuning and inter-blade phase angle mistuning indicate that inter-blade phase angle mistuning has almost no effect on average aerodynamic damping of rotor, while frequency mistuning significantly changes it. Especially, the aerodynamic damping is about 7 to 11 times at the least stable conditions of the study rotor. Furthermore, the divergence of individual blade aerodynamic damping enlarges due to both frequency and inter-blade phase angle mistuning. The individual blade aerodynamic damping for frequency mistuned blade is affected significantly by the blade local mistuned pattern and mistuning amount. Even more remarkable, the local instability of blade is induced by inter-blade phase angle mistuning at some case. It is beneficial to improve understanding the blade response and onset time differences of flutter in turbomachinery flutter problem.
Study on the motion and thermal states of oil droplet is an important part of research on the oil/air two-phase flow and heat transfer in an aero engine bearing chamber. In this paper, dimensional analysis is applied to the airflow analysis of bearing chamber. That makes the analysis model suitable for a wide range of geometric and operating conditions. Moreover, the temperature solution is added to the oil droplet motion analysis. That could promote the calculation accuracy of the droplet trajectory, velocity, and temperature. Firstly, the similarity criteria of the airflow in a bearing chamber are determined based on the dimensional analysis. The airflow distribution general formulas are proposed based on the numerical results of airflow velocity and temperature. The general formulas include 14 similarity criteria and are suitable for various geometric and operating conditions. The reliability of the general formulas is verified by some available experimental results. Secondly, the difference equations of the oil droplet velocity and temperature are listed by the difference method. The velocity and temperature of the droplet are obtained using a step-by-step method. The influence of droplet diameter, shaft rotational speed, air flow rate, and temperature on the oil droplet trajectory, velocity, and temperature are discussed. Thirdly, a test facility is built in order to investigate into the oil droplet motion and thermal states in a bearing chamber. The trajectory and velocity of the oil droplet are measured by the high-speed photography. Lastly, the proposed theoretical method about the oil droplet motion and thermal states is verified by above measurement results. The work in this paper may have a certain significance for perfecting the research system and improving the research level on the oil/air two-phase flow and heat transfer in an aero engine bearing chamber.
This paper describes current progress in the development of methods to assess aero-engine airframe installation effects. The aerodynamic characteristics of isolated intakes, a typical transonic transport aircraft as well as a combination of a through-flow nacelle and aircraft configuration have been evaluated. The validation task for an isolated engine nacelle is carried out with concern for the accuracy in the assessment of intake performance descriptors such as mass flow capture ratio and drag rise Mach number. The necessary mesh and modelling requirements to simulate the nacelle aerodynamics are determined. Furthermore, the validation of the numerical model for the aircraft is performed as an extension of work that has been carried out under previous drag prediction research programmes. The validation of the aircraft model has been extended to include the geometry with through flow nacelles. Finally, the assessment of the mutual impact of the through flow nacelle and aircraft aerodynamics was performed. The drag and lift coefficient breakdown has been presented in order to identify the component sources of the drag associated with the engine installation. The paper concludes with an assessment of installation drag for through-flow nacelles and the determination of aerodynamic interference between the nacelle and the aircraft.
The stability and control characteristics using a wind tunnel test data process are proposed and developed to investigate the stability and control characteristics of a CS-VLA certified aircraft and to comply with the CS-VLA subpart B at the preliminary design review (PDR) and critical design review (CDR) stage. The aerodynamic characteristics of a 20% scale model are provided and investigated with clean, rudder, aileron, elevator, and winglet effects. The Mach and Reynolds correction methods are proposed to correct the aerodynamics of the scale model for stability and control analysis to obtain more reliable and accurate results of the full-scale model. The aerodynamic inputs and moment of inertia (MOI) comparison between the PDR and CDR stage show good agreement in the trends of stability and control derivatives. The CDR analysis results with the corrected wind tunnel test data and accurate MOI are investigated with respect to the longitudinal and lateral stability, control, and handling qualities to comply with the CS-VLA 173, CS-VLA 177, and CS-VLA 181 for finalizing the configuration in the CDR stage.
A kind of adaptive filter algorithm based on the estimation of the unknown input is proposed for studying the adaptive adjustment of process noise variance of boost phase trajectory. Polynomial model is used as the motion model of the boost trajectory, truncation error is regarded as an equivalent to the process noise and the unknown input and process noise variance matrix is constructed from the estimation value of unknown input according to the quantitative relationship among the unknown input, the state estimation error, and optimal process noise variance. The simulation results show that in the absence of prior information, the unknown input is estimated effectively in terms of magnitude, a positive definite matrix of process noise covariance which is close to the optimal value is constructed real-timely, and the state estimation error approximates the error lower bound of the optimal estimation. The estimation accuracy of the proposed algorithm is similar to that of the current statistical model algorithm using accurate prior information.
The problem of real-time relative orbit estimation of long range and high speed between a noncooperative target and a chaser is addressed. A high precision relative orbit model, in which the Earth J2 perturbation has been taken into consideration, is established for state estimation. As the attitude of the chaser is required for relative orbit estimation, it has also been estimated. The range, elevation, and azimuth between the target and the chaser, which can be measured by a laser rangefinder and a CCD device, are chosen as measuring vector, while a rate gyroscope and one star sensor is installed on the chaser to measure its rate and attitude parameters. On the basis of nonlinear filtering technology, two kinds of filtering systems are designed, namely the separated filtering system and the united filtering system, and the relationship between the two filtering systems are analyzed. Simulation results indicate the united filtering system is more robust than the separated filtering system against realistic initial attitude errors and attitude measurement noises.
Unsteady computations of a counter-rotating axial compressor are performed and analyzed to investigate the unsteady behaviors in the compressor and the role of the tip leakage flow together with the rotating stall inception process. The results show that the oscillation on the pressure side is much stronger than that on the suction surface for both rotors, especially near the tip region where the trajectory of the tip leakage vortex (TLV) interacts with the blades most often. There exists a periodical leading edge spillage of the interface in rotor2 due to the unsteadiness of tip leakage flow (TLF) at near-stall condition. The blockage generated by the TLV increases dramatically due to the increasing strength of the TLV and the backflow phenomenon as the mass flow decreased. The appearance of the frequency components of 0.5 blade passing frequency (BPF) and 1.5BPF from 0.64BPF can be viewed as the rotating stall inception warning. The fluctuation strength of oscillation frequencies of 0.5BPF and 1.5BPF decreases rapidly from leading edge to trailing edge in rotor2, which indicates that the unsteady fluctuation of TLF at the leading edge in rotor2 is responsible for the stall inception of the compressor. Additionally, both the leading edge spillage and trailing edge backflow phenomena are observed for spike initiated rotating stall at stall point.
This experimental investigation presents a new active flow control technique based on plasma actuators applied to a backward facing step whose structure is similar to that formed by the hangar and flight deck of small naval vessels. These experiments were carried out by testing a simple frigate shape model settled at 0° wind over deck in a low-speed wind tunnel. Two different configurations of dielectric barrier discharge plasma actuator have been used to modify the flow downstream of the step. Results obtained investigating the flow by particle image velocimetry prove the capacity of plasma actuators by reducing instabilities and turbulence over the simple frigate shape model.
The effect of the cowl lip angle on the accelerating start process of a hypersonic inlet was investigated. A self-sustaining mechanism of the large-scale separation zone in the start process was studied. The inlet was a simplified, two-dimensional, one-side converging, hypersonic inlet. The movements of the large-scale separation zone of different cowl lip angles were compared and analyzed. The results show that: (1) Cowl lip angle influences the startability of a two-dimensional hypersonic inlet dramatically. With the same contraction ratio, the start Mach number of different cowl lip angles varies. First, it decreases to a minimum value at 4°, and then increases. (2) Large-scale separation zones are the common features of all cowl lip angles in unstart status. Their different development is the main cause of the different start characteristics. (3) The pressure stress is the main factor of the balance of the large-scale separation zone, and the viscous shear stress around it is of little impact. (4) The separation zone sustains itself mainly by the reflection shock of separation shock at the upper wall which is movable and adaptive with the change of Mach number. This is the reason that the separation zone gets a new balance when the incoming flow condition changes.
Dynamics design for complex mechanical systems has become an important research field and development direction at present, capturing attentions of an increasing number of engineers and scientists worldwide. Based on many advantages of the transfer matrix method for multibody system in studying multibody system dynamics, a design problem of a multiple launch rocket system is solved in this paper. Particular attention is addressed to model actions of the exhaust flow on the multiple rocket launcher, which are associated with firing order and firing intervals of rockets. Combined with a genetic algorithm, firing order and firing intervals are optimized to achieve optimum impact point dispersion reduction. The results of numerical simulation and verification tests show good agreement, while the dispersion characteristics of rockets have been improved in a low-cost way.
The problem of formation control of unmanned aerial vehicles is tackled. The control laws were developed based on two methods of formation control. These are the virtual leader formation control method and the non-hierarchical method. A decentralised controller was used, so as to distribute the formation keeping effort among the agents. The virtual leader and the virtual centre were used in specifying the trajectory of the formation. To verify the control laws developed, a simulation study was carried out. In the simulation, a homogenous formation of three agents was used. Even though both formation control methods have their merits and demerits, it was concluded that depending on the situation where the formation was needed, one was better suited than the other. The virtual leader formation method is better suited to situations where the trajectory of the individual agents is more important than keeping the shape of the formation rigidly while the non-hierarchical method is better suited to applications where keeping the shape of the formation is more important.
Considering the transformation in roles of existing air traffic management technologies, future flight operations and flight deck systems will need additional avionics and operational procedures that involve adaptive algorithms and advanced decision support tools. The main purpose of this article is to provide a theoretical framework for tactical 4D-trajectory planning and conflict resolution of an aircraft equipped with novel automation tools. The proposed 4D-trajectory-planning method uses recent algorithmic advances in both probabilistic and deterministic methods to fully benefit from both approaches. We have constructed an aircraft performance model based on Base of Aircraft Data 4 with high-level hybrid flight template automatons and low-level flight maneuver automatons. This multi-modal flight trajectory approach is utilized to generate cost-efficient local trajectory segments instead of solving complex trajectory-generation problems globally. The proposed sampling-based trajectory planning algorithm spatially explores the airspace and provides proper separation through local trajectory segments and guarantees asymptotic optimality under certain conditions. Moreover, we have integrated the cross-entropy method, which transforms the sampling problem into a stochastic optimization problem, rapidly converges on the minimum cost trajectory sequence by utilizing available flight plans, and reduces the amount of sampling. The integration of the proposed strategies lets us solve challenging, real-time in-tactical 4D-trajectory planning problems within the current and the envisioned future realm of air traffic management systems.
A new compressor rotational stall detection algorithm is proposed in this paper, and a well monotonicity exists in the correlation coefficient calculated by the algorithm and compressor surge margin. In order to reconstruct the compressor blade pressure, a trustworthy digital compressor blade pressure model is built. Specially, the model consists of three sub-models, the blade duct pressure model, rotational stall pulsation pressure model, and corrected rotational stall damping model. Simulation results indicate the effectiveness of the stall detection algorithm and compressor stochastic pressure model, which means that the correlation coefficient calculated from the algorithm and stochastic pressure model has a high reliability to take the place of surge margin as the flag of compressor stability.
Owing to the elasticity, the large deformation was brought in the high aspect ratio wing in the flight. The large deformation had a great influence on the flight performance. In this paper, the loosely coupled method was used for the research of high aspect ratio wing aeroelastic problems. The Navier–Stokes equations were solved for fluid domain computation, and the nonlinear finite element method was adopted for solid domain computation. The data exchange program and mesh regeneration progress were adopted for fluid–structure interface problem. Finally, the aerodynamic characteristics of high aspect ratio wing were obtained under different fly conditions. In addition, to validate the proposed method, the flutter analysis of AGARD 445.6 wing is carried out and compared with the experimental data. The numerical result validates the proposed computational fluid dynamics/computational structural mechanics method.
This paper presents a design approach and basic algorithms for a future system that can perform aircraft conflict resolution, arrival scheduling and convective weather avoidance with a high level of autonomy in terminal area airspace. Such a system, located on the ground, is intended to solve autonomously the major problems currently handled manually by human controllers. It has the potential to accomodate higher traffic levels and a mix of conventional and unmanned aerial vehicles with reduced dependency on controllers. The main objective of this paper is to describe the fundamental trajectory and scheduling algorithms that provide the foundation for an autonomous system of the future. These algorithms generate trajectories that are free of conflicts with other traffic, avoid convective weather if present, and provide scheduled times for landing with specified in-trail spacings. The maneuvers the algorithms generate to resolve separation and spacing conflicts include speed, horizontal path, and altitude changes. Furthermore, a method for reassigning arrival aircraft to alternate runways in order to reduce delays is also included. The algorithms generate conflict free trajectories for terminal area traffic, comprised primarily of arrivals and departures to and from multiple airports. Examples of problems solved and performance statistics from a fast-time simulation using simulated traffic of arrivals and departures at the Dallas/Fort Worth International Airport and Dallas Love Field are described.
The coning motion is a basic angular behavior of spinning missiles. Research on the stability of coning motion is always active. In this paper, the integrated nonlinear governing equations of rigid-elastic angular motion for a spinning missile with high fineness ratio are derived firstly following the Lagrangian approach. Secondly, a set of linear equation is obtained under some assumptions considering the first order vibration mode in the form of complex summation for theoretical analysis. Finally, the sufficient and necessary conditions of coning motion dynamic stability for spinning missile with and without an acceleration autopilot are analytically derived and verified by numerical simulations based on the linear equation. It is concluded that the aeroelasticity can shrink the stable region of the design parameters, even lead to a divergent coning motion.
A deceleration guidance law is developed for a reentry warhead with a fixed angle of attack by body geometry and a single moving-mass for attitude control. The guidance model equations, including the dynamic equations of the moving-mass two-body system and the kinematic equations between the warhead and the target, which form the basis of the study, are presented. To analyze the guidance model, the equations are simplified and new variables are introduced. By simplifying the guidance model equations, the transfer functions from the lateral position of the internal mass to the error angle are obtained. Iteration of the error angle is used to control the terminal velocity at a desired value. For the moving-mass reentry warhead studied in this paper, the relationship between the miss distance and the magnitude of the error angle is developed, and precision guidance can be achieved. Nonlinear seven-degree-of-freedom trajectory simulations verify the transfer functions and demonstrate the ability of the guidance law with deceleration control in a typical reentry mission.
This paper presents a robust constraint backstepping control scheme for high performance aircraft attitude tracking in the presence of physical constraints, model uncertainties and unsteady effects. The six-degree-of-freedom aircraft model with unsteady aerodynamics is first established, of which the characteristics are investigated through open-loop analysis. Based on the immersion and invariance approach, a state observer is developed to estimate the unmeasurable unsteady aerodynamic states. The unsteady effects are then compensated in the controller design where the bounds of estimate errors and model uncertainties are used to increase robustness. In addition, a command filter is introduced to impose physical constraints on states and control inputs and overcome the ‘explosion of terms’ problem. An auxiliary system of which the states are applied to the controller design is constructed as well to evaluate the constraint effects. Furthermore, stability of the closed-loop system and convergence of the tracking error are proven by Lyapunov stability theorem. Finally, several numerical simulations are performed to verify the effectiveness and robustness of the proposed control scheme.
As a decomposition-based multidisciplinary design optimization (MDO) method, collaborative optimization (CO) has been widely applied in aerospace design systems. During the execution process of CO method, subsystem optimizations are nested into the system optimization, and numerous subsystem analyses are thus required to achieve the consistency for entire system. For this problem, a non-nested collaborative optimization (NNCO) method is presented in this paper. In this method, system and subsystem optimizations are conducted sequentially and repeatedly, and system consistency is then coordinated with the use of a loop process by updating penalty parameters. The feasibility of this method was firstly tested in solving common analytical problems, and it was further applied in two MDO problems of an observation satellite as well as a maneuver satellite. All of the results have demonstrated the efficacy of this proposed NNCO method, which indicates that it is effective and suitable for dealing with practical multidisciplinary design problems.
According to the problem of autonomous optical navigation, this paper presents an easy and high-precision algorithm to estimate the attitude and position of a lander by using at least three extracted marginal elliptic curves of craters. The geometric and algebraic constraints between the marginal elliptic curves of craters and its 2D images are derived, and then the linear equations about lander’s motion are established by using Kronecker product. Consequently, the laner’s attitude and position relative to targeted features are uniquely acquired from the linear equations. In particular, the algorithm is easy to carry out because all computations involved in this algorithm are linear matrix operations. The extensive experiments over simulated images and parameters demonstrate the robustness, accuracy and effectiveness of our method.
As part of the cross industry efforts to get aircraft flying again during the April/May 2010 eruption of Eyjafjallajokull Rolls-Royce produced a chart that plotted examples of aircraft engine exposure to volcanic ash against the ash concentration the engines had been exposed to. This chart became known as the Rolls-Royce ‘Safe-to-Fly’ chart, and it was used to guide decisions on how the UK Met Office ash concentration charts for the Eyjafjallajokull eruption could be utilised to help aviators plan their flight paths. Over the period 2011–2013, this paper’s authors reassessed the engine data that made up the ‘Safe-to-Fly’ chart, and in particular the data relating to two key exposure events at high ash concentrations, flight BA009 on 24 June 1982 and flight KLM867 on 15 December 1989. Through a combination of reassessment of the original engineering calculations carried out for these events (i.e. calculations based on evidence from engine hardware) and assessments of relevant volcanological and ash cloud visibility data, it has been concluded that these events are unlikely to have occurred at or near the 2000 mg/m3 ash concentration arrived at in 2010; based on current evidence, it is more plausible that these events occurred in ash concentrations of around 200 mg/m3. As a consequence, a revision to the ‘Safe-to-Fly’ chart is recommended. In addition, to more easily present the main considerations associated with flight within ash concentrations where the ash would start to become visible, a new chart is proposed that plots duration of exposure against ash concentration. The points plotted on this chart are the revised understanding of the BA009 and KLM867 events, other relevant engine exposure events and speculative regions where flight would be unsafe and where flight would be safe, but engines would be susceptible to long-term damage.
Ice crystal ingestion at high altitude is a menace to the safe operation of jet engines. Because the core airflow in jet engine has higher temperature, ice crystals may partially melt into droplets when they enter the core airflow. A mixed-phase condition is seen consisting of both water droplets and ice crystals, which will cause ice accretion on both the static surfaces and rotating components in a compressor. This ice accretion may give rise to compressor surge or even mechanical damage of jet engine. In order to analyze this in depth, a numerical method of mixed-phase icing was developed. The Reynolds-averaged Navier–Stokes equations were used for the airflow solution. The Lagrangian method was employed for determining the trajectories of ice crystals and droplets. An ice crystal impingement model was created, in which the breakup and rebounding of ice crystals and splashing of film were considered. A thermodynamic model was proposed for ice crystals and droplets on the basis of the first law of thermodynamics. An icing simulation was developed under mixed-phase conditions with different liquid water contents and ice water contents, and the results were compared with experimental data from the literature. The comparison showed a fairly good correlation, which supports the validity and rationality of the method of mixed-phase icing.
Electromagnetic formation flight leverages electromagnetic force to control the relative position of satellites. This new propulsion technique offers a promising alternative to traditional propellant-based spacecraft formation flying since it does not consume fuel. Due to the restriction of maximum current in coils, the available inter-satellite electromagnetic force is small, and its efficient use is an important issue. In this paper, a modified far-field model is proposed to gain better accuracy of electromagnetic force approximation. Based on this model, an adaptive terminal sliding mode control is proposed to achieve fast trajectory tracking. The given method can guarantee the finite-time convergence of tracking error in the presence of bounded disturbance, input uncertainty, and saturation. Numerical simulation results demonstrate the performance and robustness of the proposed control.
In this article, the combustion chamber of SGT600 gas turbine with 18 Alstom EV burners is numerically simulated to investigate the flow field and combustion properties and analyze the sensitivity of this combustor to diameter of main fuel holes. The three-dimensional simulation is carried out based on the turbulent flame closure model for the two-equation turbulence model, k-, using OpenFOAM code for SGT600 industrial gas turbine burner. An accurate grid combined of structured and unstructured mesh scheme is employed, which is sufficiently fined in areas of high gradients. Grid independency is investigated and validation of the code established comparing the outlet temperature and flame shape with experimental results. Excellent agreement between numerical analyses and experimental data was observed, so that the difference between calculated and measured outlet temperature was 0.1%. On the basis of reasonable matching of the predicted and experimental results for the combustor, the computational fluid dynamics model enables the prediction of the heat shield and combustor’s wall temperature and critical points of operation. In this article, flow-field and operation condition of this combustor and sensitivity analysis with respect to the main fuel holes diameter are numerically investigated.
The performance characteristics of a 2D trapped vortex combustor (TVC) are investigated experimentally in terms of the exhaust gas emission level, combustion efficiency, and the exit temperature uniformity. The present study reveals that the EICO and EIUHC emission levels are sensitive to mainstream Reynolds number (Rems), cavity equivalence ratio (Øc), and primary air velocity (Vp). Besides this, mainstream premixing also has significant impact on the emitted pollutant level. Moreover, the NOx emission level can be reduced to very low levels (<2.6 ppm), which will be difficult to achieve in conventional combustors. Combustion efficiency calculated from these emission results indicates that it is influenced by cavity equivalence ratio and primary air velocity. For particular Rems, Vp, and Øc, combustion efficiency closer to 99% could be achieved for the merged flame cases, especially at higher primary air velocity. On the other hand, for a particular value of Vp, an increase in Rems tends to quench the cavity flame leading to reduction in combustion efficiency. This study also indicates that flame merging helps in efficient utilization of fuel and thus leading to better combustion efficiency. Besides this, the conditions for achieving lower pattern factor (PF) are also brought out from exit temperature measurements. Pattern factor as low as 0.1 could be achieved particularly at higher Vp cases. It can be noted that this value is lower than that of the PF value of swirl combustors (0.2–0.4). Based on the present investigations, it can be concluded that TVC is a viable technology to be considered for future gas turbine application.
This paper introduces the working principle of a high-temperature high-speed wind tunnel and analyzes the necessity of coordinated control of fuel flow-rate for the fuel supply system. In order to achieve a coordinated control of fuel flow-rate, a proportional integral (PI) cross-coupled algorithm is designed. Because PI parameters in PI cross-coupled algorithm are invariant, a better control effect cannot be achieved when there are time-variant parameters in system, so a fuzzy PI adaptive cross-coupled algorithm with model reference is designed. The simulation and experimental results show that the fuzzy PI adaptive cross-coupled algorithm with model reference has better control quality compared with the PI cross-coupling control algorithm.
The capability to "detect and avoid" potential collisions is one of the main technical challenges restricting widespread operations of unmanned aircraft into non-segregated airspaces. In fact, to operate into prescribed environments, an unmanned aircraft needs an onboard technology to replace the capability of the human pilot to "see and avoid" collision hazards. Such a technology is a "sense and avoid" system. This article focuses on the "avoid function" of such a system and proposes a suitable solution. The approach to the problem is to schematize a generic obstacle through a moving ellipsoid that represents the region of space the unmanned aircraft must not violate. The obtained solution enables situations of potential conflict to be detected and avoided through a set of possible actions such as speed changes in magnitude and/or direction. Thousands of test cases have been considered to validate this solution. Simulations show that the proposed algorithm is able to detect and avoid situations of potential conflict in the three-dimensional space and in real-time, even without the assistance of a human operator. As such, it can be considered as a fundamental step for the development of a prototype of "sense and avoid" system for promoting the integration of unmanned aircraft into non-segregated airspaces.
This article uses a control volume method to analyze the circumferential groove casing treatment in a transonic compressor. To analyze the axial momentum transport through the tip gap, the control volume near the casing is divided into two parts: the control volumes inside and outside the tip gap. Besides, the association between the forces acting on the control volume and flow structures is studied by analyzing the distributions of axial momentum flux and axial shear stress. With this method, the flow mechanisms of stall margin improvement due to casing grooves in Rotor 35 are quantitatively analyzed. The analysis is conducted at 100% and 60% design speed with supersonic and subsonic tip speed, respectively. At design speed, the casing grooves decrease the axial shear force and the axial force due to the transport of axial momentum induced by the tip leakage flow. Meanwhile, the bleeding and injecting effect of grooves contribute much to the axial force due to the transport of axial momentum. Based on the axial distribution of the axial forces, the contribution of each groove to the stall margin improvement is assessed. And the grooves that play a major role in stall margin improvement are ascertained. At 60% design speed, because the blade loading is reduced, the axial momentum transport caused by the grooves cannot suppress the boundary layer separation effectively. Consequently, the stall margin of the compressor is not significantly improved by the casing grooves.
Safe planetary landing is considered a key technology for future robotic and manned planetary landing missions. The relay hazard detection and proportion–integration–differentiation avoidance guidance algorithms were used in Chang’e-3 mission, which not only increased the complexity of the guidance system, but also resulted in non-fuel-optimal avoidance guidance from the viewpoint of fuel consumption. To further develop and improve the hazard detection and avoidance scheme of Chang’e-3, novel autonomous hazard avoidance methodologies should be investigated. This paper addresses an innovative hazard detection and avoidance scheme for safe lunar landing from the following three aspects: imaging flash lidar based hazard detection, safe landing site selection strategy, and minimum-fuel hazard avoidance guidance. First, the three-dimensional imaging flash lidar, a developing three-dimensional imaging sensor, is utilized to rapidly and precisely detect three-dimensional terrain of the landing area. Second, the hazard detection and optimum landing site selection strategy inherited from Chang’e-3 are improved and enhanced to estimate the potential obstacles, and select an optimum landing site which is the guidance target of following hazard avoidance. Next, the fuel-optimal hazard avoidance guidance problem is transcribed into as a minimum-fuel consumption optimization problem using the Gauss pseudospectral method, which is easily solved by the open-source software GPOPS. Finally, the validity of the autonomous hazard detection and avoidance guidance scheme proposed in this paper is confirmed by computer simulation.
Estimating the attitude and angular velocity of a malfunctioned satellite is one of the key technologies to achieve robotic on-orbit servicing, which is a challenging space activity. In this paper, a new method of estimating attitude parameters for target satellite by using stereo-vision system is proposed. First, two sets of vectors are constructed according to the feature point coordinates in target frame and stereo-vision system frame respectively, and then the relative attitude between the target and the servicer is determined through QUEST algorithm. Secondly, a new nonlinear estimator is proposed to estimate the attitude and inertial angular momentum of the target and the angular velocity is computed according to the estimated angular momentum and attitude. The stability of the estimator is proved by using the LaSalle’s invariance principle. Thirdly, a switch rule and an adaptive regulating rule are proposed to improve the estimation performance. The former decides when to switch from the rough estimator to the precise one. The latter regulates the gain dynamically in order to reduce the influence of the varying stereo-vision noise in the precise estimating procedure. Finally, numerical simulations and an experiment are carried out to verify the validity of the proposed method.
Ignition enhancement characteristics of partially covered cavity in an ethylene-fueled scramjet combustor were investigated experimentally under the inflow conditions of Ma = 2.1, stagnation pressure P0 = 0.7 MPa, and stagnation temperature T0 = 846 K. It is concluded that under the influence of the baffle, the covered region will become a region with sufficient fuel during the stable combustion process and wall-pressure distribution will be increased due to this restricted region. Comparing with the uncovered cavity under the same equivalence ratio, partially covered cavity is more suitable for the formation and propagation of flame kernel, and ignition in the scramjet combustor could be intensified due to this cavity configuration.
In this study, the effect of perpendicular acoustic excitation on laminar separation bubble, aerodynamic characteristics and stall characteristics over a NACA 2415 aerofoil was investigated experimentally at low Reynolds numbers at varying angles of attack. Extensive experiments such as force and pressure measurements, flow visualization via smoke-wire technique and velocity measurement via hot-wire system were conducted. The experimental results showed that when acoustic excitation of a certain frequency was applied, the length of the laminar separation bubble was shortened owing to the energy added to the flow by acoustic excitation. Because of the shortened laminar separation bubble, coefficient of lift was increased. Furthermore, at the stall angles the separated flow was forced to reattach to the surface of the aerofoil by acoustic forcing, so the stall angle and maximum coefficient of lift were increased, and drag coefficient was decreased.
Although substantial numbers of aerodynamic shape optimization works have been carried out in the past few decades, the effects of boundary layer transition are not considered in the overwhelmingly majority of those previous studies. For more accurate prediction of the flow field, and for exploration of relations between aerodynamic heating and geometrical features, the commonly used local correlation-based transition model
Inerting is an important measure to prevent fire and explosion, and enhance the survivability of an aircraft. In the battle environment, the hits from the projectiles or fragments on the fuel tank may affect the inerting effectiveness. By modifying the traditional scrubbing inerting model, this paper proposes a new inerting analysis method considering the impact of the projectiles. A series of events are introduced including the changes in either of the velocity and temperature of the projectiles, the air increase in the ullage brought by the projectiles, and the heat transfer between the projectiles and the surroundings. Examples show that the fuel load has significant effect on the oxygen concentration and the vapor/air ratio, while the velocity, the mass and the temperature of the projectile have less effect on the two parameters. In addition, the changes in the flight altitude have almost no effect on the vapor/air ratio, while having little effect on the oxygen concentration. Moreover, when the fuel tank suffers multiple hits by the projectiles, the ignition possibility or vulnerability in the ullage will increase significantly.
A genetic algorithm based inverse analysis is done to predict unknown parameters in a trapezoidal extended surface for satisfying a given temperature distribution. An inverse method is adopted to estimate six unknowns involving thermal, surface, and geometric parameters, which helps to identify feasible fin materials, necessary dimensions along with other requirements. Various controlling parameters of genetic algorithm along with random measurement errors have been investigated. Fin efficiencies have been also compared. For satisfying a prescribed temperature distribution, this study shows that many feasible materials exist which may satisfy a given temperature profile, which shall be useful in selecting any material from the available choices depending upon the relevant dimensions, convective and surface requirements. This study also shows that fin dimensions along with the coefficient of thermal conductivity influence the temperature distribution more than other parameters. The maximum variation in the efficiency among the predicted parameters is found to be within 9%.
A numerical study was carried out to investigate the effect of chordwise flexion on the propulsive performance of a two-dimensional flexible flapping wing. The wing undergoes a prescribed sinusoidal heaving motion with a local deflection. A deformable overset grid dynamic mesh method was employed to implement the motion of the grid instantaneously. The effect of flexural pattern, flexural amplitude and flapping frequency in terms of Strouhal number are evaluated. Unsteady flow around the wing is computed using an in-house developed Unsteady Reynolds-Averaged Navier-Stokes (URANS) solver. The results show that the different flexural patterns will create different flow fields, and thus the thrust generation will be significantly varied. The thrust and propulsive efficiency do not increase monotonically with the flexure amplitude while a peak value is revealed. It is found that the wake vortices after the flapping motion assembly behave as a reverse von-Karman vortex street, which can principally create thrust. The thrust is found to increase with increasing Strouhal number. Propulsive efficiency is beneficial from the chordwise flexibility and peaks within the range of 0.2 < St < 0.4, which is evidenced by natural flyers.
The designs of sealing device have prominent influence on the performance of aero-engine. The high temperature environment during the working process of aero-engine also has important influence on the performance of sealing device which is located in the aero-engine. Finger seal has a flexible characteristic and high price performance compared with the other seal devices, thus it gets more attention, and lots of researchers have studied about finger seal’s performance recently. But so far the dynamic performance of finger seal considering temperature effect is not yet analyzed and discussed. Based on this, an equivalent dynamic model based on distributed mass considering temperature effect is proposed in the paper. The effects of environment temperature and heat through friction on the equivalent structural stiffness of finger stick and contact pressure between finger stick and rotor are discussed. Moreover, the data exchange between the dynamic and thermal analysis is confirmed based on the movement relationship between the rotor excitation and finger stick response. Therefore, the dynamic performance analysis of finger seal including thermal-structure coupling is obtained based on an equivalent dynamic method. The effect of temperature on the dynamic performance of finger seal using this model is analyzed, and the effect of C/C composite structural parameters on the finger seal performance is investigated considering the temperature effect. The above results show that the temperature effect has important influence on the performance of finger seal, so it is necessary to consider the temperature effect when the performance of finger seal is analyzed. The current work further improves the theoretical system about finger seal equivalent dynamic research, and has higher academic significance and engineering value.
This paper deals with the problem of passivity and passification for stochastic Markovian jump systems with incomplete transition rates and actuator saturation. By use of the polytopic model, sufficient conditions are built such that the closed-loop constrained system is stochastically passive. The procedure of deriving a state feedback controller is converted into an optimization problem with a set of linear matrix inequalities. Finally, the simulation results of a linearized dynamic model of a vertical take-off and landing aircraft in the vertical plane illustrate the effectiveness of the proposed schemes.
Centrifugal compressors with high pressure ratio have been widely used in small gas turbines, industrial compressors and turbochargers. High mechanical stress and high temperature of the impeller and large compression power required are the key factors that limit the pressure ratio achieved by centrifugal compressors. Cooling is an effective way to reduce the required compression power and the impeller temperature and it is widely used in gas turbines and industrial compressors. In this work, a cooling method named integrated cooling was employed in a high pressure ratio centrifugal compressor. The effects of the integrated cooling on the compressor performance were studied by three-dimensional steady simulation with conjugate heat transfer method using ANSYS CFX commercial solver. It is found that integrated cooling is capable of improving the compressor performance with respect to pressure ratio, efficiency, compression power as well as the impeller maximum temperature. With a cooling temperature of 300 K, the pressure ratio increased by 11.0% and efficiency by about 2.44% at the design condition. Besides, the maximum impeller temperature decreased by about 67 K. This considerable improvement of the compressor performance justifies the integrated cooling, which helps to advance the design of high pressure ratio centrifugal compressors.
The Ornicopter concept is a single-rotor, tailless configuration. By actively flapping its blades, the Ornicopter rotor can propel itself to rotate, and hence does not need a tail rotor. In previous research, the Ornicopter concept has been compared with the Bo-105 conventional helicopter from various aspects, while the Ornicopter has the same design parameters as the Bo-105. Comparisons show that the Ornicopter has one major drawback, namely a small flight envelope. To improve the Ornicopter performance and understand how the Ornicopter should be designed, in this paper, the Ornicopter design is unfrozen and optimized for the flight envelope. The optimization result shows that with a proper design, the Ornicopter performance can be improved dramatically. A similar flight envelope as the Bo-105 can be achieved for the Ornicopter. However, the Ornicopter requires higher power than the Bo-105 due to the inherent characteristics of this concept.
The development pattern of the end wall boundary layer (BL) in whole conditions and its effect on the matching of multistage compressor have been studied in detail in this paper. Moreover, one method of end wall zone blade modification is carried out, using computational fluid dynamics, to improve the stage matching by re-camber. It is found that the pitch-averaged thickness of end wall BL gradually increases along streamline direction and the BL is highly skewed in the pitchwise direction. In addition, the value of BL thickness, mainly depends on the stage pressure rise coefficient
To reduce the influence of nonlinear friction on the dynamic and static performance of gimbal axis for inertial stabilized platform, a novel method based on steady-state error response to identify LuGre friction parameters is proposed in this paper, and a kinetic friction torque testing system using the proposed method is developed. On account of the introduction of structure and electromechanical model for inertial stabilized platform, the relation among the input signals, interfering signals and steady-state error are analyzed. And then the torque ripple compensation function is identified and introduced into the control loop to eliminate the effects of torque ripple on parameter identification of friction model. Based on the friction torque testing system designed, the parameters of LuGre friction model for gimbal axis are got by the method of Two-step identification and dynamic parameter optimization. According to the identified friction model, the tracking process of the framework is compensated. The tracking performance is verified through both simulation and experimental results.
Blade profile plays an extremely important role in the aerodynamic performance of turbomachinary. However, in the process of secondary machining for near-net shape blade and serviced blade, the deviation between nominal profile and measured data can raise serious problems for NC tool path generation. It is even worse that unreasonable reconstruction may be responsible for the deteriorated aerodynamic performance and subsequent high in-service cost. This paper presents a parametric reconstruction strategy for 2D blade profile in reverse engineering. The reconstruction profile approaches to the nominal model and its performance is proved to be more acceptable. This consists in the optimization of parametric geometry (modeling and modification) based on parameter constraints and computational fluid dynamics simulation rather than measured data itself. Some practical examples are used to demonstrate the usefulness and superiority of the strategy.
In this paper, the autocovariance least-squares (ALS) method for a linear time-varying system is proposed for estimating both the process noise covariance and the measurement noise covariance associated with the measurement sensor to mitigate the performance degradation caused by an incorrect information of the sensor errors or by a large change of errors in sensor measurement. To verify the efficiency of the proposed method, simulations were performed for the attitude determination of the lunar lander which combines a gyro and a star tracker measurements assuming that the star tracker errors are increased by fault during the lunar descent and landing. The simulation results show that the attitude error of the proposed method is smaller than the conventional Kalman filter by adaptively tuning the noise covariance of the star tracker measurement errors. Also, the relationship between the ALS estimation accuracy and the innovation sample size and the time lag is discussed.
The paper introduces a dimensionless alternative to the International Civil Aviation Organization (ICAO) target level of safety (TLS) of collision frequency less than
In the wind tunnel, the Mach number in the test section is an important parameter that cannot be calculated accurately. Several mechanism models have been used to estimate it based on the aerodynamics laws in the past decades. However, the accuracy of estimation cannot satisfy measurements. An alternative approach is to design data-driven models. Unfortunately, the high-dimensional input features and the large-scale data sampled from measurements make it difficult to predict the Mach number. To solve the two issues, based on the multivariate fuzzy Taylor theorem, the feature subsets ensemble (FSE) method is proposed in this paper. An FSE is developed on the basis of the set of direct, exhaustive, independent subdivisions of the feature space. Learning on substantially lower dimension feature subsets, the FSE is characterized by low complexity. The FSE models are examined by data of measurements from the wind tunnel of China Aerodynamics Research and Development Center. Experiments show that the FSE speeds up the testing time that would otherwise be infeasible for the individual BP, Bagging and Random Forest. The FSE models meet the requirements of forecasting speed, accuracy and generalization of the Mach number prediction.
A dynamic model was designed and developed using thermodynamic mass conservation equation, energy conservation equation, flow equation of nozzle and motion equation for moving piston and design parameter analysis of free-piston tunnels was proposed. The proposed model was validated by comparing the predictions made with the dynamic model established with the experimental data reported in some of the references and used to run simulation. Simulation results indicated that the target pressure and temperature could be quickly obtained by selecting and adjusting design parameters with the proposed model while the pressure and temperature output of the reservoir were kept long stable with a pneumatic hydraulic driving device. The time domain conditions of free-piston wind tunnel could be established through numerical study at different gas source pressure. The pressure and temperature of reservoir could be obtained by adjusting the valve opening of gas source. The variable specific heat ratio has a remarkable effect on the performance of free-piston wind tunnel. For the variable specific heat ratio, the temperature and pressure of reservoir was lower than the constant specific heat ratio. It was therefore concluded that the proposed model was suitable for performing the design parameter analysis of free-piston wind tunnels, and so, it could be used to facilitate the design of free-piston wind tunnels for long working at high Mach number.
In this paper, a novel inner/outer-loop control framework for the pitch autopilot of a tail-controlled, skid-to-turn missile is proposed. The generalized extended state observer (GESO)-based inner-loop design is constructed to estimate the disturbances and uncertainties and simultaneously compensate them in the input channel. In doing so, the uncertainty of the model can be reduced within an appropriate range for robust control design. The outer-loop
Algorithms for self-separation of aircraft operating in a one-way high density air corridor are discussed. In the air corridor, which is considered as a tube or band-shaped piece of airspace that connects high-demand areas, only aircraft capable of self-separation may operate. It is required that all the aircraft fly in the same direction while maintaining safety without instructions of air traffic controllers. To realize high traffic throughput without losing safety, aircraft should be self-separated in an appropriate manner. Additionally, aircraft must operate inside of the assigned airspace for the air corridor operations. In this paper, an algorithm for self-separation of aircraft in a width-limited band-shaped piece of airspace has been developed. The algorithm determines maneuvers for conflict avoidance using information of surrounding aircraft. It is demonstrated that all the aircraft are able to self-separate without conflict in the width-limited air corridor. It is also demonstrated that self-separation using current state information results in a deadlock in the narrow air corridor. As the corridor width increases, aircraft are able to maintain the separation through the use of heading changes while maintaining their optimum speed. Additionally, it is indicated that the installation of sub-route and utilization of optimum speed information prevent a deadlock even though the corridor width is narrow.
In this paper, the experimental results of an unconventional joined-wing aircraft configuration are presented. The test model uses two different wings, forward and rear, both joined in tandem and forming diamond shapes both in plant and front views. The wings are joined in such a way that it is possible to change the rear wing dihedral angle values and the rear wing sweep angle values in 25 different positions that modify the relative distance and the relative height between the wings. To measure the system aerodynamic coefficients it is necessary to perform wind tunnel tests. The data presented corresponds to the lift, drag and induced drag aerodynamic coefficients, as well as the aerodynamic efficiency and the parameter for minimum required power, from the calculated values of the lift and drag time series measured by a 6-axis force and torque sensor. The results show the influence on the aerodynamic coefficients of the rear wing sweep and dihedral angles parameters. As a main result, it can be concluded that, in general terms, the lift and induced drag aerodynamic coefficients values decrease as both the distance and height between the wings increase, on the other hand, the total drag aerodynamic coefficient decreases if the distance between the wings increases, but nevertheless shows a slight tendency to increase if the height of the rear wing increases, whereas the aerodynamic efficiency and the parameter for minimum required power increase if the distance between the wings increases.
In this paper, a modified chi-square test is proposed to develop an autonomous fault detection and identification system for the navigation system in the lunar lander. The conventional fault detection logics, which is based on state chi-square test have had a limitation on fault identification. The proposed modified chi-square test computes modified chi-square parameter (MCP) by comparing the estimated states which is estimate on local filters to the propagated states. Because the MCP only contains the information of the respective sensor measurement, the MCP from failed measurement is contaminated by the fault. Thus, the MCP from other measurements is not contaminated by the fault, then the MCP from failed sensor can be easily distinguished by finding a diverging MCP signal. Using the proposed method, the fault of the lunar lander can be efficiently detected and isolated.
The increasing volume of traffic in the air transportation is leading to excessive workload on air traffic controllers. Developing automated air traffic management (ATM) tools is a critical technology in reducing the workload of air traffic controllers and hence increasing the airspace capacity. The existing approaches to automated ATM either use overly-simplified air traffic and aircraft dynamics models to reduce computational complexity or end up being computationally intractable for large-scale ATM scenarios. This paper presents a new hybrid system description of modeling the decision process of the air traffic controllers in en-route and approach operations. The model is based on the domain expertise provided by the State Airport Authority and Air Navigation Service Provider (ANSP) of Turkey. The emulation of air traffic controller decision process in the hybrid model provides realistic conflict resolution maneuvers and separation assurance in 3D, while being computationally tractable. The algorithm has polynomial iteration complexity in the number of waypoints of the aircraft, which makes it scalable to large-scale ATM scenarios with more than 100 aircraft. The algorithm is validated on the real air traffic data over the Istanbul region extracted from the ALLFT+ dataset provided by EUROCONTROL, which includes over 7000 flights in a 96-hour period. The developed algorithm is also integrated into a Boeing 737-800 flight deck simulator with a custom radar display to demonstrate the applicability to existing avionics systems.
An approach for improving tracking performance of the predictive functional control (PFC) based on equivalent input disturbance (EID) and generalized extended state observer (GESO) is presented in this paper. Control structure of the proposed method includes two parts, a PFC controller and a state observer, and both of them can be designed separately. The PFC controller is employed to optimize the tracking control of a nonlinear system and its tracking performance heavily depends on the accuracy of the predictive model. Whereas, various internal perturbations and external disturbances always make the predictive model significantly deviate from the nominal model. To address this problem, a state observer is introduced into the proposed approach. All of the above-mentioned uncertainties are regarded as a lumped disturbance, and the lumped disturbance is estimated by EID–GESO as well as the system states in an integrated manner. The estimated disturbance can be eliminated in a negative feedback loop and then, a relatively accurate predictive model for the PFC controller is offered. Closed-loop stability of the composite control is also provided. Compared with previously related work, the notable feature of the proposed design is that the observer-based PFC is extended to nonintegral chain systems subject to mismatched uncertainties and better performance is obtained. Finally, the proposed approach is applied to a missile longitudinal autopilot design, and a comparison with some prominent methods in the presence of significant uncertainties demonstrates the robustness and effectiveness of the proposed design.
An electrostatically charged spacecraft is subject to Lorentz force as it moves through the planetary magnetic field. The induced Lorentz force could be used as propellantless electromagnetic propulsion for orbital maneuvers. Such vehicle is referred to as Lorentz spacecraft. By modeling the Earth’s magnetic field as a tilted dipole that corotates with Earth, a nonlinear dynamical model is developed to characterize the orbital motion of a Lorentz spacecraft relative to an arbitrary Earth orbit. A fast nonsingular terminal sliding mode control law is then proposed, based on which an adaptive controller is designed for Lorentz-augmented spacecraft relative motion to deal with the uncertain parameters and perturbations. Due to the constraint that the direction of Lorentz force is naturally perpendicular to the local magnetic field and the vehicle’s velocity relative to the magnetic field, which does not necessarily coincide with the direction of the required control force, thus, the Lorentz force works as auxiliary propulsion to reduce the fuel consumption in most cases. Then, a fuel-optimal distribution law of Lorentz force and traditional chemical propulsion is proposed. Furthermore, the stability of the closed-loop control system is proved via a Lyapunov-based approach. Numerical simulations substantiate the feasibility and validity of the proposed controller for Lorentz-augmented relative orbital control.
A new dynamic test mechanism has been developed in the unsteady low-speed wind tunnel at Nanjing University of Aeronautics and Astronautics to enhance the abilities of experiments in the unsteady flow cases. The structure of the new test mechanism is described in detail in this paper and the equations used to control the performance of test model’s motions are also listed here. Then a wing-body model was tested in the new test system for the first time, including the forced oscillation tests and free-to-roll test at high angle of attack region. The test results are also discussed.
The problem of relative motion control of spacecraft rendezvous process on elliptical orbit is considered in this paper. Due to the presence of nonlinear dynamics and external disturbances, two robust controllers are developed based on sliding mode control theory. The first one is an optimal sliding mode controller; in which optimal control theory is used to reduce the tracking error and fuel cost, and then integral sliding mode control technique is applied to robustify optimal controller. The other controller is a backstepping sliding mode one that is developed based on nonlinear dynamics of spacecraft rendezvous; having applied the backstepping method to synthesis the tracking errors and Lyapunov functions, a sliding mode controller is developed to guarantee the Lyapunov stability, handling of all nonlinearities, robustness against uncertainties as well as tracking the desired position. It is assumed that the chaser and target spacecraft are in a low Earth orbit and subject to the perturbing effects of J2 and atmospheric drag. In addition, two fault tolerant scenarios, i.e. thruster degradation and short thruster failure are also considered to verify the robustness and efficacy of the control approaches. Simulation results confirm the effectiveness of the proposed controllers in reaching to the desired position in case of actuator fault-free and fault tolerant situations.
This paper investigates free vibration characteristics of a rotating double-tapered functionally graded (FG) beam made of porous material. Material properties of the FG beam vary continuously through thickness direction according to the power-law which modified to approximate material properties for even and uneven distributions of porosities. The governing differential equations of motion are derived based on Euler–Bernoulli beam theory and using Hamilton's principle and then solved utilizing a semi-analytical technique called the differential transform method (DTM). In order to verify the competency and accuracy of the current analysis, a comparative study with previous researches is performed and good agreement is observed. Several important parameters such as power-law exponent, porosity volume fraction, taper ratios, rotational speed and slenderness ratio which have impacts on natural frequencies of such beams are investigated and discussed in detail. It is concluded that these effects play significant role on dynamic behavior of rotating double tapered FG beam. Numerical results are tabulated in several tables and figures that can serve as benchmarks for future analyses of FGM beams with porosity phases.
This paper aims to investigate formation control problem of multi-UAV system with nonuniform time-delays and jointly-connected topologies. No explicit leader exists in the formation team, and therefore a consensus-based distributed formation control protocol which requires only the local neighbor-to-neighbor information between UAVs is proposed for the system. The stability analysis of the proposed formation control protocol is also performed. The research suggests that when the time-delay, communication topology and control protocol satisfy the stability condition, the formation control protocol will guide the multi-UAV system asymptotically converge to the desired velocity and shape the expected formation team, respectively. Numerical simulations verify the effectiveness of the formation control system.
In this study, the effect of varying aspect ratios on aerodynamic performances on the flat plate and NACA 2415 airfoil without/with low frequency pitching movement at low Reynolds (Re) numbers was investigated. The experiments include force measurements for both static and dynamic airfoils and smoke-wire flow visualization for static airfoils at Re numbers of 50,000, 75,000 and 100,000. In the static cases, the stall delayed because of tip vortices at low aspect ratios, as the aspect ratio increased, maximum lift coefficient and the stall angle decreased. On the other hand, the stall mechanism changed and the abrupt stall occurred. As the Re number increased, the mild stall formed. In the dynamic cases, on the flat plates at lower ARs, the oscillation did not change the lift coefficients considerably, and as the Re number increased, the oscillation made the drag force decrease. In the dynamic NACA 2415 wing with AR1, the oscillation improved the aerodynamic performance at down-stroke conditions (especially between 15° and 25°). However, in dynamic NACA 2415 wing with AR2, as the Re number increased, dynamic stall formed over the wings at the angle of attacks between 15° and 25°.
This article deals with aerial manipulators consisting of unmanned helicopters equipped with robotic multi-link arms. This setup has strong potentialities for structure assembly and manipulation tasks, which makes it a valuable resource in outdoor industrial environments. The article presents an elaborate model of the entire system as well as methods for controlling the aerial platform taking into account the motion of the arm. The article presents the design and implementation of a Variable Parameter Integral Backstepping controller which clearly outperforms the results obtained with the standard helicopter controller. Realistic characterization of system sensors has been also accounted for. Simulations experiments confirm the validity of the proposed approach, and the results are also validated with a high-fidelity Modelica simulation model.
For the attitude estimation problem, the quaternion is preferred to describe the rotation of spacecraft, which obeys a unit-norm constraint. However, this constraint is not considered in the most of the traditional estimation algorithms. This work will develop a new estimator named norm-constrained predictive filter to solve this problem based on the predictive filter frame. The optimization theory of Lagrange multiplier is adopted by minimizing a constrained cost function to improve the traditional predictive filter. In addition, the obtained estimator is further analyzed and applied to the spacecraft attitude estimation problem by numerical simulations.
This paper proposes an integrated guidance and control path following and dynamic control allocation method based on constrained optimal quadratic programming for the stratospheric airship with redundant systems. The control system is divided into two levels. A novel planar geometric path following control method is presented as the high-level controller. The method consists of planar position tracking controller, velocity regulator, attitude and altitude stabilizations. The velocity regulator contributes to the directionality of path following and decreases the number of singularity points for the closed-loop system. The control input of the actuators for the low-level controller can be obtained by two principles, one is keeping the control input close to the desired steady-state value, and the other is minimizing the change between the current input and the preceding one as well as the one before. This approach considers both physical constraints and dynamic responses of actuators. When a fault occurs, it can change the weight coefficient of the actuator to ensure the fault is well distributed among the fault-free actuators without reconfiguring the high-level controller. Several simulations illustrate the feasibility of the control law and show advantages for practical application.
The lack of ground-tracking resources has become a primary bottleneck for the Chinese BeiDou navigation satellite system. As crosslinks have been widely recognized as a promising augmentation for the autonomous navigation of the global navigation satellite system, this article studies a new decentralized data fusion method for orbit determination of crosslink-augmented satellite constellations. In the new solution, the system is modeled as a probabilistic graphical model, the dynamical Bayesian network, and a graphical method named junction tree is introduced to analyze the structure of the system. By dividing and arranging the predictions and estimations of junction tree in a proper sequence, a new decentralized solution with centralized equivalent precision is designed. As the solution requires satellites in the constellation cooperating with each other through communications and measurements, it is named junction-tree-based cooperative orbit determination. Simulation results indicate that junction-tree-based cooperative orbit determination has centralized equivalent precision and good robustness after satellite failures, whereas centralized solutions such as extended Kalman filter may suffer from processing center malfunctions. Junction-tree-based cooperative orbit determination could not only serve as an optional choice for global navigation satellite system autonomous navigation but also be used as a general scheme for decentralized data fusion in cooperative systems such as unmanned aerial vehicle formations and so on.
The Flow Corridor is one of the operational concepts in the future air traffic plans. They are planned to be introduced along high demand routes between airport or city pairs. Aircraft capable of self-separation are allowed to fly along it with no instruction by air traffic controllers. All aircraft will be able to fly along each ones’ own optimum trajectory with punctuality in the Flow Corridor by using the self-separation. In this study, the optimum Flow Corridor allocation based on the potential benefit estimation is discussed. The average fuel consumption and flight time of representative aircraft types are analyzed using their actual flight data in Japan. In addition, the fuel optimum flight trajectories of the same type aircraft are analyzed. Through their comparison, it is clarified that the Flow Corridor introduction enables both the flight time and fuel consumption reduction, and that the fuel consumption reduction is remarkably large in both the cruise and descent trajectories. The optimization analyses varying the aircraft weight clarified the height and speed extent among the optimum trajectories of various aircraft types. This determines the cross-sectional shape of the Flow Corridor. The situations of overtaking are also analyzed assuming that all aircraft in the Flow Corridor fly punctually with appropriate landing intervals at the destination airport. It is clarified that the overtaking frequently occurs during both the cruise and descent trajectories. These results provide the requirements for the self-separation during the descent phase.
In this paper we propose a design concept, i.e. constant contraction of discrete streamtubes, to develop a new design method for hypersonic inlets with shape transition, which enables the shape transition inlet to have high compression efficiency while keeping flow uniformity as much as possible. The method first divides the 3D flow field of the inlet into several groups of discrete streamtubes with constant contraction ratio. The axisymmetric flow within each stream tube is then optimized via computational fluid dynamics to have high compression efficiency, and these discrete streamtube groups are finally assembled together by matching their shock wave reflection positions to form a 3D inlet flow field with weak crossflow. This method can be used for the design of shape transition inlets with specified entrance/exit sections. To validate the present methodology, we design a rectangular-to-circular inlet with a design point of Ma 6 and a takeover point of Ma 4, and examine numerically the flow structure and the inlet performances.
Numerical simulations have been carried out to investigate the performance of the axial ventilator equipped with honeycomb structure. The oil /air two-way coupling model based on the Realizable k– turbulence model and the droplet impact model are proposed. Based on verifying the rationality of numerical model, characteristics of flow resistance and oil–gas separation efficiency are calculated and analyzed. The results show that the axial ventilator with honeycomb has favorable separation efficiency, which is estimated as 99.6% for the oil droplet diameter of 5 µm. The honeycomb structure has little effect on flow resistance, but plays a major role in the oil–gas separation of axial ventilator, where the contribution to the oil separation accounts for 80% at least. Besides, the increase of rotation Reynolds number enhances the centrifugal force, resulting in the increase of separation efficiency, while the increase of nondimensional mass flow rate and environmental temperature reduce the residence time of oil droplets in the axial ventilator mainly, resulting in the decreasing of separation efficiency.
In this paper, we analyze an evolutionary multiobjective heuristic algorithm for collision avoidance maneuver optimization when multiple objects in space threaten a satellite. Population-based evolutionary multiobjective heuristic algorithms can stably find optimal or suboptimal solutions on account of their robustness and flexibility. We present a collision avoidance maneuver optimization planning method using the nondominated sorting genetic algorithm II. Our proposed method simultaneously optimizes two goals: the minimization of DeltaV consumption and maximization of the satellite maneuver cycle length. Through an experimental evaluation, we demonstrate the efficient performance of our method in terms of DeltaV consumption and maneuver cycle length under constraints, including allowable DeltaV and burn start time, with consideration of a safety threshold between the user satellite and multiple threatening objects. We perform the simulation once per each burn strategy, i.e. in an in-track direction with a single burn; in three radial, in-track, cross-track direction with a single burn; and three radial, in-track, cross-track directions with two burns. We obtain various solutions, all of which satisfy the safety threshold. For the in-track direction with the single burn strategy, the solution showed merit in the minimization of DeltaV consumption. On the other hand, the results of three radial, in-track, cross-track directions with two burns showed good performance in maximization of the maneuver cycle. We therefore contend that efficient collision avoidance maneuver optimization planning can be achieved by using a multiobjective heuristic algorithm.
Application of composite patches in repair of damaged/aged aircraft structures is one of the most popular repairing methods in aerospace engineering. Since running experiments are difficult, time consuming, expensive, and also require high level of expertise, simulation of the behavior of the patch and also the faulty components after repair can assist designers and engineers in optimization of their designs. In this article, full-scale simulation of a damaged panel that is experimentally repaired with a composite patch will be considered using ABAQUS, a commercial finite element code. The crack growth process is modeled with the extended finite element method and the cohesive zone model (CZM) is used to model the progressive damage in the adhesive of the composite patch repair. Also, sensitivity analysis is performed on the CZM parameters and it is shown that the three parameters i.e. the shear toughness, maximum first shear traction, and penalty parameter for the elastic stiffness are important in the simulation of adhesively bonded composite patch repairs. The calibrated cohesive properties are successfully used to predict the response of the composite patch to the strengthened in damaged structure with considering the linear and nonlinear stage of failure process. The simulation results obtained in different stages have been verified with the existing experimental results.
This paper proposes a new observer-based sliding mode guidance law to intercept maneuvering targets in terminal phase. In order to satisfy the interception, the line of sight angular rate is considered as sliding variable based on parallel navigation idea. For this sliding variable reach to zero in finite time, the guidance law is designed using a new chattering-free observer-based sliding mode control. This algorithm consists of the target acceleration estimation using a new extended state observer. Therefore, with the availability of this estimation, switching function of the sliding mode control is replaced with a smooth function to generate smooth control signal. Also, the stability of closed-loop system is proven by using Lyapunov method. Numerical simulations are presented to illustrate the proposed guidance law potential.
Algorithms have been presented for the low-frequency nonlinear dynamic modeling and stability domain determination of liquid propellant engines. Also, the considerations that can facilitate the modeling process and remove the drawbacks have been presented. The defined algorithms and the stated considerations have been applied for a particular liquid propellant engine. The equations describing the mentioned engine have been classified into 11 subsystems. The simulations have been performed in the MATLAB Simulink software environment. Each simulated subsystem represents one or several physical subsystems whose interactions have been defined in the overall form of engine configuration. The engine modeling results have been validated by the data of cold and warm tests of the subsystems and engine. In the process of determining the stable operation range of the mentioned engine, by linearizing the obtained nonlinear model, which accompanies the correct determination of model input and output, and by applying the Nyquist criterion, the stability analysis of the regulator and engine systems has been performed control system and the stability margin and the cutoff frequencies have been obtained. Finally, the effects of structural and process parameters of the relevant subsystems on the engine’s stable range of performance have been evaluated.
Modern flight management systems (FMS) are able to predict highly precise 4D trajectories for complete flights from take-off to touchdown. The generated trajectories take into account aircraft’s performance parameters, altitude-, speed- and time constraints, given procedures, and the expected weather conditions, as forecasted. Based on these trajectories, a traffic scheduler can generate and guarantee conflict-free traffic—as long as every aircraft sticks to the plan. However, even the most accurate trajectory becomes unrealizable when significant disturbances arise that were not foreseeable at prediction time. Uncertainties like imprecise weather forecasts, inaccurate departure times, and issues with delayed passengers influence a direct implementation of pre-planned 4D schedules in practice. If such disturbing events are rare, the overall plan may be adapted. Too many disturbances might ruin the whole schedule though. In the initial planning stage, larger margins increase overall robustness, but obviously also downgrade the possible optimum of the whole system. This article therefore discusses the trade-off between built-in robustness and efficiency. Resilience is added by increasing the separation between trajectories. Assuming that the mandatory distance for guaranteeing safe operations remains constant, the additional separation provides freedom for aircraft following their trajectories.
This paper describes two systems that can be used to obtain realistic random traffic samples in a terminal area: a real traffic analyser and a synthetic traffic generator. These two systems allow the air traffic management (ATM) engineer to gain insight on the traffic structure of the area under analysis, and allow obtaining realistic traffic samples enabling the evaluation of new operational concepts, the validation or system performance measurement after procedure changes, the analysis of ATM performance under forecasted future traffic changes, etc. Together with the design of the system, the work provides insight of user interfaces and describes the potential uses of such tools in an integrated ATM system.
In future air traffic scenarios, remotely piloted aircraft systems are likely to play an increasingly important role. A major use case and subject of research is loss of command and control data link. Therefore, a novel engineering solution is described, which makes use of knowledge-based information processing methods for the implementation of higher cognitive functions in an artificial agent. This agent shall be integrated aboard the airborne vehicle and shall work in a cooperative relationship with the human pilot on the ground for his or her support. With the provision of functional redundancy in flight guidance and mission management related decision-making the overall system safety shall be improved. The article captures the whole development process of the agent including concept, requirements definition, system design, implementation and evaluation. The interactions between the human and the cognitive agent are discussed and the necessity to augment the conventional human–machine interface is deduced. A scenario-based usability study with a group of pilots in addition to real flight tests to evaluate the developed system is described. In addition, considerations on certifiability of such an artificial cognitive associate system are presented.
In this paper, the effect of aerodynamic damping produced by store rear horizontal fins on the aeroelastic instability of a wing with a large underwing store is investigated in an incompressible flow. The wing is modeled by using the classical Euler–Bernoulli beam theory while the unsteady strip aerodynamic theory based on the Wagner function is used to simulate the aerodynamic field. The underwing store is modeled as a concentrated mass attached to the wing by a rigid pylon and has rear horizontal fins. The numerical results of the developed generic and simple model are compared with available results, and an excellent agreement is observed. It is found that the aerodynamic damping of the store produced by its rear fin can compensate the flutter boundary of the wing-store configuration and therefore for more accurate analysis, in addition to the store weight and location, the effect of store fin must also be considered.
This paper presents a robust observer-based finite-time convergent guidance law to intercept maneuvering targets. To avoid the undesired chattering phenomenon existing in classical sliding mode control, the adaptive super-twisting algorithm is first introduced to design a robust guidance law to drive the line-of-sight (LOS) angular rates to a small region around the origin in finite time. The important feature of the proposed adaptation algorithm is in nonoverestimating the values of the guidance law gains, which is achieved by using a ‘detector’ scheme in the updating law. Moreover, with the proposed adaptive law, the presented guidance law requires no information on target maneuvers. Since the LOS angular rate is difficult for a pursuer to measure accurately for some seekers, a novel adaptive super-twisting observer is proposed to estimate it and a composite guidance law is then synthesized. Detailed finite-time stability analysis and comparison results with other guidance laws demonstrate the superiority of the proposed formulation.
The U-bend which turns flow through 180° is encountered in many applications in mechanical and aerospace engineering systems. One important example occurs in modern turbine blade cooling systems, where the internal cooling passages is threaded radially outwards and inwards to form a multi-pass arrangement where straight passages are connected with U-bends. The main purpose of the present investigation was to focus on finding a 3D U-bend configuration with minimum pressure loss using the 3D CAD-based surface parameterisation method. The design of experiment technique and surrogate design space model were successfully applied by the authors, as opposed to direct numerical optimisation, to reduce the computational cost. A standard Reynolds-averaged Navier–Stokes (RANS) computational fluid dynamics method with the Spalart–Allmaras one equation turbulence model was selected for. Even though the simple RANS with the one equation turbulence model cannot simulate the highly complex U-bend flow physics precisely, the optimisation process was able to identify an optimum U-bend configuration which achieved a 63.3% pressure loss reduction, relative to the datum configuration, yielding the lowest loss U-bend in the literature. The authors also performed careful experiments to confirm their predictions and the performance of the optimum U-bend configuration identified by this work was validated.
Single expansion ramp nozzle (SERN) is widely used in the propulsion system of hypersonic vehicle; it has good effect on weight loss, and also can reduce the nozzle base drag and friction loss effectively. But under the condition of transonic region, the flow is severely over-expanded in the SERN, and the corresponding performance of SERN is sharply declined. In order to improve the SERN performance under over-expansion condition, the passive flow control technique application of passive cavity on SERN is investigated numerically by solving Reynolds average Navier–Stokes equations, and with standard two-equation k- turbulent models adopted. The influences of major geometric parameters of passive cavity on the flow field and performance of passive cavity SERN are deeply explored. Results show passive cavity structure has active influence on the performance of SERN under the over-expansion condition. The shock position moves upstream and separation region increases in the passive cavity SERN. The number of the shock train in the passive cavity SERN decreases. And the second peak of the pressure distribution is obviously higher than that it has for the baseline SERN. The performance of passive cavity SERN is related to the size of separation zone and its starting position, which is determined by the position of the first hole in the passive cavity. Percent of porosity has a dominant influence on the flow field of passive cavity SERN. The position of the first hole in the passive cavity changes with the variation of percent of porosity, and the corresponding starting position of the flow blowing out from the passive cavity is different, resulting in the different position and intensity of the anterior oblique compression shock wave. The axial thrust coefficient of passive cavity SERN decreases with the increase in percent of porosity accordingly. The flow structures of passive cavity SERN change little with variety of aperture and cavity depth when percent of porosity remains constant, the axial thrust coefficient of passive cavity SERN are almost tantamount at this case. Compared to the effect of percent of porosity, the influence of aperture and cavity depth on flow field are much smaller. The influence of aperture and cavity depth on the performance of the passive cavity SERN can be ignored in the design.
Two-dimensional laminar incompressible flow around a National Advisory Committee for Aeronautics 0012 plunging airfoil with different Gurney flap heights of 2% C, 5% C, and 7% C and the oscillation amplitude of h = 0.12 C and frequencies of = 1.3 and 4.67 rad/s at Re = 1850 is carried out in this paper. In addition to the Gurney flap, in order to control the flow over the airfoil, dielectric-barrier discharge plasma actuator is attached to the flap and its impact is investigated on the aerodynamic coefficients. Flow field and aerodynamic characteristics are compared with the clean plunging airfoil. For the Plasma force considered in this study, the 2% C Gurney flap height lead to a more enhancement in the lift curve. The results indicate that at all actuated and non-actuated Gurney flap heights the thrust is reduced. Furthermore, it is observed that by using plasma force at lower frequencies,
Sensors system is considered as one of the most indispensable devices to the mission of capturing a satellite for on-orbit servicing in the space. In this paper, in order to measure the pose of nozzle, six laser range finders are installed in the capturing nozzle device for entering into and capturing nozzle, and a Geometric Axis Approximation (GAA) method is proposed to estimate the pose of satellite nozzle. Six laser range finders are arranged in the configuration of emission radiated at two planes of top and bottom. The pose measurement system is composed of the six laser range finders, applicable for real-time measurement of the ranges from inner surface of nozzle. Based on the properties of nozzle shape and finders arrangement, the GAA method is used to estimate the pose of nozzle by a connecting line between two centers of circles in the top and bottom planes. In the pose estimation of GAA method, solving the nonlinear transcendental equations can be avoided and a simple explicit model is built. Therefore, the GAA method is suitable for real time measurement. In this paper, the theoretical error of GAA method is analyzed, and in the capturing nozzle experiment system, the experiments of pose measurement are carried out. The results of experiments prove the efficiency of GAA method in pose estimation to nozzle based on laser range finders.
Unmanned air vehicles (UAVs), which have been popular in military context, have recently attracted attention of many researchers because of their potential civilian applications. However, before the UAVs can be used for civilian applications, a systematic integration of UAVs in the National Airspace System (NAS) is needed that can allow safe operation of UAVs along with other manned aircrafts. One of the critical capabilities needed for safe operation of the UAVs in the NAS is the ability of UAV to carry out sense and avoid task which would allow the UAV to amend its path to avoid collision with other aircrafts. Despite recent technological advances, such as availability of automatic dependent surveillance broadcast that can transmit position of an aircraft to others in the vicinity, planning a collision-free path autonomously is still challenging in a dynamic environment. The objective of this paper is to develop a methodology for discovering a path for the UAV that meets mission goals, avoids collision, is optimal in terms of path length and, more importantly, is feasible. The paper formulates the problem of path planning using the mathematical paradigm of mixed integer linear programming and provides a solution strategy for solving this problem in the dynamic sense. The tasks of avoidance of obstacles and waypoint navigation are incorporated as constraints in the MILP problem. The paper presents the solution of the MILP problem and verifies the performance of the proposed methodology with regard to optimality of the solution and computational time requirement via several simulated scenarios of different complexities.
A method based on parametric reduced-order models to efficiently predict the transient response of aeroelastic systems with concentrated structural nonlinearities is presented. The approach approximates the nonlinear response in a piecewise-linear manner through time integration of sub-linear reduced-order models; these are parameterized with respect to the nonlinearity and are efficiently obtained by balanced truncation and interpolation. The procedure is applied to the optimization of a wing tip device for passive loads alleviation which features a nonlinear stiffness, showing the effectiveness and efficiency of the methodology.
This paper discusses the outcome of flight experiments relevant to non-cooperative sense and avoid that were carried out in the framework of a collaboration between the Italian Aerospace Research Center and the University of Naples "Federico II". Within the project, aimed at full autonomy for medium/large Unmanned Aircraft Systems, an integrated radar/electro-optical system configuration was adopted, and real-time data fusion/automatic decision-making algorithms were developed to estimate intruders’ motion, and generate and follow proper escape trajectories. All the systems were designed to guarantee collision avoidance without any dependence on the command and control link. The hardware/software sense and avoid system was installed onboard an optionally piloted flying laboratory of Very Light Aircraft category, and a single intruder aircraft of the same class was considered for avoidance experiments. In particular, the paper focuses on the results from radar-based autonomous avoidance flight tests. Performance is analyzed in terms of situational awareness before and during the maneuver, and effectiveness in performing safe avoidance maneuvers while generating smooth commands and minimizing the deviation from nominal trajectory. Statistics relevant to several encounters show that in spite of a small number of valid sensor measurements, the tracking algorithm is able to keep a satisfying level of situational awareness, thus enabling safe avoidance. However, the limits deriving from coarse radar accuracy, low scan rate, and interaction between mechanical scanning and aircraft flight dynamics are also pointed out. While these limits did not impact avoidance safety in the considered tests due to the favorable operating environment in terms of time to collision, radar detection range, and ownship maneuverability, they confirm the need of improved sensing concepts in more challenging avoidance scenarios, such as adoption of radar/electro-optical data fusion and electronic scanning technologies.
The main objective of this paper is to present our vision of the role of automation in future air traffic management (ATM) system. It also includes an analysis and state of the art on ATM and automation. The content is based on HALA! position paper, that looks beyond the framework defined by SESAR and NextGen for the near future, and analyses the feasibility of higher levels of automation in ATM in the far future, taking into account the relationship between humans, organizations, and machines and scoping the allocation of functions, roles, and tasks between them.
Taking into consideration the proposed SESAR&NextGen paradigm shift (trajectory-based operations; proactive, more distributed and autonomous system) and studies developed in ATM automation in the past years, the vision goes toward a more ATM automation system operation questioning even the possibility of fully automated ATM system. The understanding of higher levels of automation in ATM considers the "overall system performance" as main driver for the resulting "optimal" ATM level of automation.
Implementation of above programs will change the human role in the future ATM system. In this regard, the vision considers ATM as sociotechnical system and orchestrated organization, overcoming the former consequences of automation, where some "human errors" caused from automation were found, to this new vision where human motivation for their engagement in the ATM activities, will be promoted.
New roles between humans, organizations, and between them and machines will be derived by considering ATM as a complex sociotechnical multi-agent system. Under this assumption, it will be essential: to maintain a high degree of autonomy among different ATM agents and, simultaneously, an optimized level of orchestration among them.
Three interdependent criteria to support the always controversial decision about where to dynamically allocate the ATM essential functions/tasks are proposed; when is the "best time" to take an operational action, where is the "best place" having the best picture to take it and, finally, who will be the "best player" to implement the associated tasks.
As a result, the paper advocates for a different role of automation devoted to support a temporal, institutional, and physical distributed ATM system.
It is also pointed out that research in ATM automation should always take under consideration the adequate level of automation in the far future ATM system and should always ensure the safety and reliability of a highly automated ATM system.
Finally considers the need to align the research efforts into two different main activities of improvement, devoted to: aircraft trajectory hierarchal, spatial, and temporal cohesion and trajectory management.
The flight control system plays an important role in adjusting the attitude of manned or auto-pilot aircrafts. To reduce the fault diagnosis time and accelerate the maintenance actions, many flight control systems have adopted the design for testability. Testability demonstration for the flight control system is needed to check the indexes of testability such as fault detection rate and fault isolation rate. Currently, the standards and statistical methods for the testability demonstration planning have the problems such as large sample, long test period and it is not optimal for the flight control systems which are of complex structure and high cost. A testability demonstration planning method based on the sequential probability ratio test method is proposed as it can decrease the sample size with almost the same operation characteristic as the classical method. Firstly, the decision factor and rules of the sequential probability ratio test method and truncated decision rules are introduced. Secondly, the establishment of failure mode set based on the failure rate for the sequential probability ratio test method is illustrated. Finally, the demonstration of fault detection rate for a flight control system is implemented with the given method and steps. Software named testability demonstration and evaluation system which can calculate the decision criteria, plot decision chart, select failure mode and make decisions is used to assist the implementation of the test. The result shows that the fault detection rate passes the test with a credible performance and the actual sample size is remarkably decreased while comparing with the classical method.
In this work, a new algorithm is presented with regard to the free initial condition for solving optimal control problems. The reason for presentation of such an algorithm is to develop a suitable method that simplifies difficulties of the optimal control problems that researchers face in common methods of optimal control theory. Also, initial condition as true anomaly is considered to be free in this optimal control problem. To do so, issues such as optimal control theory, orthogonal functions in the Hilbert space, and evolutionary optimizations such as genetic algorithm-particle swarm optimization and imperial competition algorithm are utilized. The algorithm was solved for low-thrust orbital transfer problems which included nonlinear dynamic equations. To validate the algorithm, simplified Edelbaum low-thrust equations are compared with the proposed analytical solutions. Next, the algorithm is investigated for the low-thrust orbital transfers with respect to the equinoctial orbital equations of the minimum-time problem. Results are achieved for two evolutionary optimization methods genetic algorithm-particle swarm optimization and imperial competition algorithm and three orthogonal functions such as Fourier, Chebyshev, and Legendre. Two optimization methods and three orthogonal functions are covered and compared precisely. With respect to the results, this algorithm has the capability to overcome difficulties of the optimal control problems and can be considered as a novelty in this field for the free initial condition problems.
A coupled free-wake/panel method to predict the rotor downwash aerodynamic interaction and helicopter trims is presented. Free-wake method was developed to analyze aerodynamics interference of rotor wake, based on lifting-surface theory and vortex method, which relates the core radius, span station, and circulation of initial tip vortex with the blade-bound circulation distribution, and eliminates the empirical parameters during rotor free wake analysis. Helicopter fuselage with empennage was discretized into source panels. The vortex line mirror method was adopted to account for wake acceleration phenomenon that resulted from the close interaction between rotor wake and fuselage surface. The rotor wake geometry and downwash simulations were investigated including comparisons with the available measured data. Combined with the helicopter flight dynamics model and embedded in the trim procedure, the free-wake/panel coupled method was applied to calculate the rotor wake and its interference on other components of a full-scale UH-60A rotorcraft in level flight. Comparisons among predictions, referenced results, and experimental results are made for rotor wake geometries, blade-bound vortex, blade tip vortex, rotor downwash velocity, and control stick positions, and encouraging results were obtained. To validate the present method, this paper discusses the fundamental formulation, the numerical algorithms, and the simulation results.
At high-cruising speeds, airplane wing structures shall experience air loads of significant magnitudes from different directions. The flexible wing structure with an elevated aspect ratio produces the bend-twist coupling that often exceeds the control limits. The popular aeroelastic control reversal issue occurs because of the unsystematic changes in the aerodynamic quantities. This article presents a detailed investigation on the airplane lateral stability that is subjected to a variety of aerodynamic forces at cruising flight. A novel idea is proposed to find the dynamic stability characteristics of an airplane against the aeroelastic reversal problem. Initially, the pitching moment relating to the change in lift at various angles of attack with longitudinal static stability condition is verified by using computer simulation. Then the slope of wing lift curve, aileron lift curve and the moment coefficient concerning to aileron deflection are computed using computational as well as experimental methods. The changes in pressure distributions about the aileron deflection revealed several facts to enhance the reversal speed. The wind tunnel testing results are perfectly concur with the computational fluid dynamics plots. The outcome of experimental analysis is confirming that aeroelastic control reversal speed enhancement without any major structural optimization is possible. The aeroelastic behaviour of large commercial airplanes and subsonic bombers are the primary interest in the view of application about this analysis.
The importance of the more electric aircraft has been highlighted in many publications, projects and industrial presentations. By definition, the more electric aircraft concept achieves the majority of the required system functionality by using electrically powered sub-systems and components. This manifests itself in much higher electrical power demands on-board aircraft, compared to conventional architectures. This presents many challenges in the design process. To alleviate the risk and choose the optimum architectures for the systems on the aircraft, it is essential to incorporate the characteristics and possible configurations of the electrical network in the conceptual and preliminary design stages. Hence the current practice of performing an electrical load analysis at the detailed design stage is not adequate. To address this gap, this paper presents a viable and robust methodology to define requirements, size components and systems and calculates the electric power requirements at the preliminary design stages. The methodology uses the conventional aircraft, systems and components as the baseline and uses mathematical techniques and logical sequences of component operation, developed through the research, to size electrical load profiles for conventional aircraft. It then adapts this result to the more electric aircraft concept by adding key components that would account for the difference between a conventional system and a more electric system. The methodology presented here makes the design process more robust and aids the choice of the optimum design for the aircraft.
With various modeling technologies applied, the sensor fault detection and isolation scheme based on the decentralized model (also referred to as dedicated observer scheme) becomes a popular approach for sophisticated systems. However, the commonly used modeling approach in many literatures that directly takes measurement values as model inputs may result in residual crosstalks and even false alarms. In this paper, the traditional decentralized model scheme is analyzed and a novel scheme based on the time window interactive prediction structure is proposed. Then, the Elman neural network is applied to model identification due to its nonlinear approximation and online learning properties. Finally, Simulations for comparison using the decoupled longitudinal motion model of some airplane are performed, and the results show that the proposed scheme has higher detection speed, lower false alarm rate and less undetected faults.
The aim of this paper is to account for the effect of the epistemic uncertainty of the input variables’ uncertainty in the nonprobabilistic reliability analysis on the safety of the structure system. Based on the idea of moment-independent sensitivity analysis, a modified sensitivity measure of the nonprobabilistic reliability is constructed to identify the most influential epistemic parameters of interval variables. For calculating the nonprobabilistic reliability sensitivity measures of the epistemic variables, a computational model is established. And a solution method with the advantages of the state-dependent parameter model is employed to improve the computational efficiency and avoid the complex sampling procedure. The numerical examples and engineering examples show that the proposed method of solving the sensitivity measure is reasonable and effective. The sensitivity measure of nonprobabilistic reliability proposed in this paper can give an essential importance sequence of all the epistemic uncertainties and identify key contributing epistemic uncertainties. When the sensitivity measure is larger, the epistemic uncertainty variable will become more important and should collect the data to increase knowledge of parameters. The sensitivity measures can provide the availability guidance to reduce the number of epistemic variables.
Based on the
Electrical load simulator is hardware in the loop simulator used to exert real-time aerodynamics loads on the servo actuation system of a flight vehicle under test according to flight conditions. This article investigates direct torque control of electrical load simulator system using adaptive fuzzy backstepping method. To analyze the effect of extra torque disturbance on electrical load simulator system, detailed mathematical formulations are derived. Considering practical aspects of the proposed method, state vector is estimated using a state predictor, and parameters of the system are estimated using algebraic method. Fuzzy logic system is used to estimate extra torque disturbance acting on electrical load simulator system, but the approximation error may not converge to zero, which may affect control performance. Similarly, the parameters of the system may vary with time; thus the lumped disturbance due to time variation of parameters and fuzzy approximation error is compensated using adaptive control law derived based on estimated error dynamics between actual plant and state predictor. Moreover, to improve transient response, a novel saturation function-based transient performance controller is introduced. The performance of the proposed control is verified using extensive numerical simulations.
To measure the attitude of a satellite, conical earth sensor is usually used on the low- and medium Earth orbit satellite. By detecting the infrared radiation at the horizon using infrared detectors, the conical earth sensor gives a measure of the attitude of a satellite. However, when light from the moon or sun comes into the field of view of the conical earth sensor, it will capture unexpected pulse signals that will induce measurement errors and finally bring about attitude fluctuation of the satellite. In this article, the mechanism of such an interference has been analyzed in depth. By detecting and discriminating the pulse widths of the sun, the moon, and the Earth, a novel method was presented and new software was developed to eliminate the interference. Furthermore, a special ground test platform was set up to verify the proposed method and software. Some real on-orbit flight data were applied as well. Both results showed that the sun and moon’s interference was identified and rejected without corrupting the horizon crossing.
To improve the control precision of nonlinear spacecraft formation flying, the input–output linearization minimum sliding-mode error feedback controller is presented based on the linear-decoupled spacecraft formation model by input–output linearization method incorporating the sliding-mode control. This paper proposes a new strategy to estimate and offset the system-control errors, which include various kinds of uncertainties and disturbances. To facilitate the analysis, the linear-decoupled spacecraft formation model is first given; on which basis, the concept of equivalent control error is introduced to define the entire model error. Based on the minimum sliding-mode covariance constraint, a cost function is formulated to estimate the equivalent control error and fed back to the conventional sliding-mode control. It is shown that the sliding mode after the input–output linearization minimum sliding-mode error feedback controller will approximate to the ideal sliding mode with high-control precision. In addition, the new methodology is applied to spacecraft formation flying. It guarantees global asymptotic convergence of the relative-tracking error in the presence of the large perturbations. More exactly, the two input–output linearization minimum sliding-mode error feedback controller laws (continuous sliding-mode control and nonsingular terminal sliding-mode control) are developed for this spacecraft formation flying system. Several fault-tolerant scenarios are considered to verify that the input–output linearization minimum sliding-mode error feedback controller is still effective in the presence of faults in spacecraft thrusters. Numerical simulations are performed to demonstrate the efficacy of the proposed methodology to maintain and reconfigure the spacecraft formation with existence of initial offsets and large perturbations effects.
A new method, based on singular value decomposition and QR factorization, has been developed and applied to the analysis of F-18 flutter flight test data. The method is capable of identifying the frequency and damping of the critical aircraft modes, those responsible for the flutter phenomenon. The procedure relies on the capability of singular value decomposition for the analysis, modeling, and prediction of data series with periodic features and also on its power to identify matrix rank. The analysis of simulated and real flutter flight test data demonstrates the effectiveness, robustness, noise-immunity, and the capability for automation of the method proposed under specific conditions.
Conventional buffet onset methods for a 2D supercritical airfoil, SC0410, in transonic regime for various Mach number and various angles of attack have been surveyed. The existing methods give good results for high subsonic and transonic regimes, but demand a computational procedure to detect the buffet onset. One of these methods, trailing edge pressure divergence, that have been recognized inappropriate in other studies for supercritical airfoils, shows acceptable result at least for the present supercritical airfoil. A new method has been proposed by the authors for transonic regime that is based on the physical definition of the buffet onset from the surface pressure distribution diagram. This method does not require any special calculations. The data scrutiny shows good agreement by this method in comparison with the conventional schemes.
This paper presents a combinatorial optimization method based on uniform design in combination with response surface methodology and genetic algorithm. Uniform design is used to obtain experimental points and response surface methodology to establish a mathematical regression model. Subsequently, genetic algorithm is employed to acquire optimal solution of the objective function. The optimization method has been applied to a two-dimensional S-shaped transition duct design. The process is performed with two design variables. One defines the drop height ratio which describes wall profile, and the other depicts the length ratio between the axial length of the S-shaped transition duct and the duct inlet height. Total pressure loss coefficient as an aerodynamic performance parameter is selected as the objective function for optimization. The objective function is numerically assessed at design points sampled by uniform design in the experimental domain. The initial transition duct was designed with a radius-change to length ratio 11.6% larger than current engine design limits, and the optimization yields a decrease of 36.9% in total pressure loss and more uniform distributions of parameters at the outlet. The paper shows that the described optimization method can be applied to turbofan engines to increase the radial offset and decrease the axial design space between the fans and cores without jeopardizing performance.
An adaptive model of the human pilot engaged in pursuit tracking tasks that was previously introduced in the literature is modified and applied to the analysis of piloted control of a realistic transport aircraft model. As described, the pilot model requires no guesswork on the part of the analyst as regards initial parameter settings. By means of computer simulation, the adaptive pilot model is shown to exhibit superior performance to its non-adaptive counterpart in a series of configuration changes associated with the vehicle model. The overall validity of the post-adaptive pilot model is assessed by examining the resulting open-loop pilot vehicle dynamics in comparison to that predicted by the crossover model of the human pilot. The pilot modeling approach is proposed as a preliminary analytical tool to be used in the assessment of robust flight control system designs subject to faults or system failures with an eye toward potential loss-of-control.
Different schemes of a propulsion system have a distinguished influence on the overall performance of high altitude airship. There is an optimum power, called optimum power unit, to achieve the lowest propulsion system and energy system weight for a high altitude airship. The paper represents an optimization model of the optimum power unit for a high altitude airship. Firstly, the optimal Latin hypercube design method is applied to obtain the sample points of the distributed low power propulsion system. Secondly, the surrogate model, which is used to establish the optimization model, is obtained by responding surface method based on these sample points. The computational model of the energy system is obtained by the airship’s location and the working time. Finally, the multi-island genetic algorithm is used to find the optimum power unit for a typical high altitude airship. Furthermore, the optimization work under different typical power levels and diameters is carried out to verify the effectiveness of the optimum power unit design method. It has been found that the identical result validates the effectiveness of the optimum power unit design method.
In order to improve the computational efficiency of nonlinear dynamic probabilistic design for aeroengine typical components, a probabilistic design method–extremum response surface method-based support vector machine of regression was proposed. By taking support vector machine of regression as an extremum response surface function, the mathematical model of surface method-based support vector machine of regression was established. The probabilistic design of turbine disk-radial deformation was accomplished based on the surface method-based support vector machine of regression fully considering the influences of the nonlinearity of material property and the dynamic of heat load and mechanical load. The analysis results show that the probabilistic distribution and inverse probabilistic features of input–output parameters and the major factors (rotor speed and gas temperature) are gained legitimately, which provide the useful reference for disk design and blade-tip clearance control more effective of high-pressure turbine). Through the comparison of methods, surface method-based support vector machine of regression is demonstrated to hold high efficiency and high precision in nonlinear dynamic probabilistic design of aeroengine typical components. Moreover, the proposed surface method-based support vector machine of regression is promising to provide a useful insight for disk dynamic optimal design and blade-tip clearance control of aeroengine high-pressure turbine.
The mathematical dynamics model of the tilt-rotor unmanned aerial vehicle (UAV) that has nacelle-fixed auxiliary wings (NFAWs) is presented based on computational fluid dynamics analysis using FLUENT and DATCOM. The advantage of the aerodynamic performance of the NFAW is compared to the performance of the original tilt-rotor UAV in a trim analysis as well as simulation. The inner loop and outer loop of the neural network controller are designed for the tilt-rotor and its NFAW variant. In order to improve the control performance of outer loop, pseudo-control hedging (PCH) is applied to the outer loop as well as the inner loop neural network control. The dynamic inversion on a linear model of the original tilt-rotor at hover conditions is used as a baseline. The sigma-pi neural network (SPNN) adaptation minimizes the error of the inversion model. This error typically occurs due to the use of an approximate tilt-rotor model for helicopter mode instead of the NFAW model throughout the flight envelope from helicopter to airplane mode. The waypoint navigation and the automatic hover guidance are applied to the most outer loop of the neural network controller for the autonomous flight, which consists of nacelle conversion and reconversion as well as automatic take-off and landing. The fast dynamic reference commands generated by the autonomous waypoint guidance are inputted to the outer loop control in order to make the PCH of the outer loop effective. Lastly, the nonlinear simulation results are compared under turbulent wind conditions, in which the NFAW is more negatively affected than the original tilt-rotor model.
This paper investigates the dynamics and failure modes during the impact of an elastic body on a sandwich structure by means of non-linear finite element analysis. The main motivation for the study is the accidental impact of a human body on the interior sandwich structure of a civil aircraft during a crash situation. The considered model is a rectangular simply supported sandwich plate that is loaded dynamically by the centric impact of a spherical body with varying stiffness. In principle, the impactor stiffness has a significant influence on the contact forces between impactor and sandwich structure, and consequently, leads to a change in the impactor deceleration and re-acceleration as well as a change in the contact duration. However, the deformation of a "softer" impactor causes a smoother load introduction. Thus, two questions arise: can the altered stress distribution change the initial failure mode of the sandwich structure? And how are the deformations and deceleration and accelerations of the elastic impactor influenced? As, particularly, the latter question is crucial to human safety in crash situations, the inertia loads exerted on the elastic impactor are evaluated in detail by standard injury criteria.
Roughness modelling at low Reynolds numbers of O(104–105) is of practical importance for micro air vehicles. This paper investigates the roughness modelling behaviour of the low Reynolds number shear stress transport model and the -Re shear stress transport model. Both include modelling flow transition and surface roughness effects. The roughness effects are modelled as sand grain roughness. A series of simulations using the two models have been performed on a NACA0012 aerofoil with comparisons to available experimental data. The results show that both of the models have the capability to reasonably predict the leading edge laminar separation bubble, transition and skin friction and, therefore, lift and drag on smooth surfaces. However, the two models behave very differently for the rough surface aerofoil. While the low Reynolds number shear stress transport model performs well, the -Re model fails to predict the transition on the rough aerofoil surface, resulting in inaccurate lift and drag prediction.
Intercooled turbofan cycles allow higher overall pressure ratios to be reached, which gives rise to improved thermal efficiency. In addition, intercooling allows for the size, weight and exhaust jet velocity of the core to be reduced. For an optimum jet velocity ratio and fixed thrust, the fan pressure ratio and specific thrust are also reduced, which benefits propulsive efficiency. A new intercooled core concept is proposed in this paper, which promises to alleviate limitations identified in previous intercooled turbofan designs. This concept facilitates the installation of the intercooler and reduces core losses at high overall pressure ratios. This engine concept takes advantage of intercooling and the arrangement of the high pressure spool to reach and exceed overall pressure ratios of 80. In addition, given the reduction in core size, bypass ratios beyond 14 have been considered. In order to identify efficiency gains and performance characteristics which are due to the novel arrangement alone, the geared intercooled reversed flow core engine has been compared with a geared intercooled engine with a more conventional core. Finally an optimisation exercise has been carried out to identify the best configuration for both the geared intercooled reversed flow core concept and the conventional core concept. In this paper, it is demonstrated that the geared intercooled reversed flow core concept allows for a 2.3% reduction in block fuel burn. The reductions are due to the improved core efficiency, higher overall pressure ratio as well as efficiency gains from the use of a mixed exhaust. The sensitivity analysis shows that the improvements are highly dependent on pressure losses in the core and bypass stream and that careful design of the mixer chutes and intercooler headers to achieve low losses is essential if the concept gains are to be realised.
In this study, we performed experiments to investigate the effect of sweep angle on the transition location of laminar flow to turbulent flow. Three half wing models were used, each having a different sweep angle but with the same aspect ratio in various angles of attack. Two flat plates were used at the ends of the swept wing models to prevent the flow from rolling up over the wing. By simulating flow over infinity swept wing by eliminating tip vertices, the effect of sweep angle on flow transition phenomenon was investigated. The experiments included the study of transition flow via hot-film sensors, which were glued on the wing surface. We found that the small leading-edge radius and low Reynolds number used in the experiments showed the effect of cross-flow mode is dominant over flow transition, rather than other flow instability modes on the leading edge of the wing. Increasing the swept angle therefore leads to enforcement of cross-flow mode and, in return, causes rapidity of flow transition. The increasing angle of attack makes the location of transition nearer to the leading edge.
This paper compares and analyzes four heuristic algorithms for collision avoidance maneuver optimization (CAMO) when multiple space objects threaten a satellite. Classical gradient-based optimization methods are not appropriate for this kind of problem due to their discontinuities. On the other hand, heuristic algorithms can obtain suboptimal solutions due to their robustness and flexibility. In this paper, we develop CAMO planning methods using four heuristic algorithms. Their performance is compared in terms of the Del-V achieved under constraints on the minimum distance between the user satellite and multiple threatening objects, the maximum burn duration, and the boundary conditions for the maneuver start time. To validate the proposed strategy with the heuristic algorithms, two CAMO problems are analyzed. One is a simple problem using two control parameters (the maneuver start time and Del-V along the in-track direction) when a single threatening object is approaching. The second is a more complex CAMO problem that uses four control parameters (the maneuver start time and Del-V in three directions, i.e. radial, in-track, and cross-track) when four threatening objects are approaching from different angles and at different times. As a result, we minimize Del-V for each CAMO problem while satisfying all constraints. The differential evolution heuristic algorithm is found to exhibit the best performance in terms of minimized Del-V.
The attitude control for reentry vehicle is responsible for the robust operation to avoid the major deterioration from parametric uncertainties and external disturbances. Targeting these practical issues, both the cases with and without a priori knowledge of upper bound on the lumped uncertainty (i.e. the joint effect caused by external disturbance and inertia matrix uncertainty) are addressed, and correspondingly two continuous time-varying sliding mode based attitude controller design strategies are proposed to achieve the robust tracking of the attitude commands while alleviating the control chattering. Firstly, to deal with the case where the upper bound on the second derivative of the lumped uncertainty is known in advance, a nonlinear disturbance observer based continuous time-varying sliding mode control algorithm is developed so that the asymptotic stability of the closed system is guaranteed. Furthermore, in order to address the more practical case that the upper bound on the lumped uncertainty is unavailable, a continuous adaptive time-varying sliding mode control algorithm is derived with the related switching gains adjusted on-line, by which the trajectories of the closed-loop system are guaranteed to be uniformly ultimately bounded. Finally, the proposed strategies are applied to the attitude control of X-33 RLV in the reentry phase to illustrate the effectiveness of the theoretical results.
A spline wavelet collocation method is presented to solve optimal control problem (OCP) of flexible spacecraft, which is often required to reorient and reposition with minimum manoeuvre time or fuel consumption. It is very difficult and computationally expensive to determine the open-loop optimal control inputs for flexible spacecraft, because the optimal control profile is often characterised by discontinuities or switching in the control variables. In our approach, the state and control variables are expanded via cubic spline wavelet decomposition, and then an OCP would be converted into a nonlinear programming problem where the wavelet coefficients are treated as the optimisation variables. As opposed to the usual pseudospectral method based on polynomial approximation, the wavelet advantageous properties of compact representation would inherently make it efficiently and accurately to solve nonlinear programming problem using standard solver. The novel approach is demonstrated by two typical optimal problems. The results show that our approach outperforms Gauss pseudospectral method for discontinuous OCPs arising from the flexible spacecraft.
A numerical study is performed to study the effect of nozzle wall cooling on transition between two different shock structures such as free shock separation and restricted shock separation in an axisymmetric thrust-optimized contour nozzle. In this study, cooling of nozzle wall which is associated to the first half of nozzle length is concerned, and at different cooling rates, the transition between shock structures, hysteresis cycle, and also plateau pressure ratio at which the transition occurs are characterized. To do this, a two-dimensional numerical calculation is accomplished utilizing the commercial CFD software, FLUENT. Validity of current numerical model is confirmed by comparison of nozzle wall pressure, hysteresis cycle, and plateau pressure ratio with experimental and previously published works as well as applying simple energy balance. Numerical results show that the increase in cooling rate causes the transition between shock structures and thus hysteresis cycle to appear at lower values of pressure ratio. It is found that, in the case of nozzle wall cooling, a single point could be realized for transition between shock structures. It is also shown that the effect of nozzle wall cooling is to reduce the plateau pressure ratio at which the transition happens.
Multidisciplinary design optimization is one of the modern design methods. It was developed in several different structures and used to solve some of the theoretical and applied problems. Collaborative optimization is one of the structures of bi-level multidisciplinary design optimization. It comprises system level and discipline level, which is used to solve engineering complex problems. Collaborative optimization structure maximizes options of discrete disciplines and provides a mechanism for coordinating design problem at system level. The present research discusses capability of the collaborative optimization method to solve multidisciplinary problems aiming at reducing the weight of a liquid-propellant system. It is realized by implementing a propellant system design comprising an engine and consumption of fuel and oxidizer. To do this, we calculated engine parameters through response surface methodology. The calculation parameters were optimized by applying a response surface and an engine structure design in the collaborative optimization process at the same time in the form of combustion, geometry, and weight (structure) problems with the evolutionary algorithms. Finally, we compared the obtained results with the reference results and specified the optimization rate achieved for the values of variables. The values included pressure increase of combustion, specific impulse, engine mass reduction, rate of fuel and oxidizer consumption with fixed thrust, and burn time.
Persistent surveillance is a major role envisioned for autonomous unmanned vehicles. The mission of persistent surveillance requires the vehicles to continuously survey a target region. This paper investigates the techniques of persistent surveillance control for a swarm of micro aerial vehicles. We present a flocking algorithm to drive the micro aerial vehicles flying in a coordinate formation with a capability of obstacle avoidance. We propose a new digital pheromone mechanism to control and coordinate the swarms of micro aerial vehicles to search a field of interest and to reduce the uncertainty of every region in the field over time. Simulation results show the effectiveness of our proposed algorithm in generating collision-free persistent surveillance trajectories for a swarm of micro aerial vehicles in a coordinated manner.
The paper presents the automatic control of the aircraft in the longitudinal plane during landing, taking into account the sensor errors and the wind shears. The H-inf control provides robust stability with respect to the uncertainties caused by different disturbances and noise type signals, while the dynamic inversion provides good precision tracking. These techniques are combined and a robust automatic landing system is obtained; by adding an optimal observer and two reference models providing the desired altitude and velocity, one obtained a new automatic landing system which is very suited for landing control in the longitudinal plane. The optimal control law is calculated in two ways, this improving the generality, applicability, and simplicity degree of the automatic landing system. The theoretical results are validated by numerical simulations for a Boeing 747 landing; the simulation results are very good (Federal Aviation Administration accuracy requirements for Category III are met) and show the robustness of the algorithm even in the presence of wind shears and sensor errors. Moreover, the designed control law has the ability to reject the sensor measurement noises, wind gust, and wind shears with low intensity.
The online estimation of the center of mass plays an important role in the attitude and orbit control law design for spacecrafts with significantly time-varying masses. A new method is proposed to estimate the center of mass of a spacecraft by using six accelerometers and three gyros. The six accelerometers are used to measure the accelerations of six different points in three directions, and the three gyros are used to get the angular velocity of the spacecraft. By combining the acceleration and the angular velocity, the angular acceleration can be obtained directly instead of differentiating the angular velocity. In this way, the differential error can be avoided and thus the center of mass estimation precision can be increased. Besides, the requirement on the measurement precisions of gyros and accelerometers can be relaxed. Two configuration modes of the six accelerometers on three directions, 2-2-2 and 3-2-1 are discussed, and based on that the simulation results are generated and evaluated in terms of the root of mean square error of the center of mass estimation. When the measurement precision of accelerometer is higher than
Ground effect is asymmetric when an unmanned aerial vehicle takes off by using the catapult nearest to the edge of the deck from a carrier, because a large part of the wing is out of the deck. Asymmetric ground effect would induce rolling and yawing moments, which are critical factors affecting the safety of takeoff operations. In this research, focus was on asymmetric ground effect, especially on the lateral and directional aerodynamic characteristics. Effects of height, velocity, and wind over deck were studied. Computational fluid dynamics method was used and validated by comparing it with the experimental data presented in early reports. Height is the most important factor that influences lift, rolling moment, and yawing moment. Lateral and directional stabilities are weakened by reducing height. Lateral stability decreased 2.8% and 5.6% as the height was reduced from 1.5 m to 1.2 m and to 1.0 m, respectively. By increasing velocity, lift is increased significantly, while yawing moment is little influenced. Magnitudes of both lift and rolling moment are amplified slightly with the increase of wind over deck. When wind over deck varied from 0 m/s to 15 m/s, lift and rolling moment varied only within 1% and 3.4%, respectively, and thus the effect of wind over deck is secondary.
This paper concentrates on the noise reduction effect of stator lean in rotor–stator interaction. The compressor with a serial stator-blade lean angle has been employed to acoustically test and numerically calculate. The experiment results show that stator-leaned positive has better effect on noise reduction than leaned negative; the tone noise is determinant on total sound pressure level, and the lean angle of the stator should exceed 10°. Based on the results of unsteady calculation about the compressor, the amplitude of the unsteady loading of stator decreases with the increase in the lean angle. Leaned positive stator has lower unsteady force and loading than leaned negative. The phase parameters q and pcle of wakes are almost proportionate to the stator-blade lean angle. The distribution of q and phase close-leading edge (pcle) shows that the compressor with 25° and –20° stator-blade lean angles has the maximum and minimum phase variation, respectively.
A new nonlinear adaptive control scheme based on the immersion and invariance theory is presented to achieve robust velocity and altitude tracking for hypersonic vehicles with parametric uncertainty. The longitudinal dynamics of the hypersonic vehicle are first decomposed into velocity, altitude/flight-path angle, and angle of attack/pitch rate subsystems. Then a non-certainty-equivalent controller based on immersion and invariance, consisting of a control module and a parameter estimator, is designed for each subsystem with all the aerodynamic parameters unknown. The main feature of this method lies in the construction of the estimator, which is a sum of a partial estimate generated from the update law and an additional nonlinear term. The new term is capable of assigning appointed stable dynamics to the parameter estimate error. Stability analysis is presented using Lyapunov theory and shows asymptotical convergence of the tracking error to zero. Representative simulations are performed. Rapid and accurate command tracking is demonstrated in these numerical simulations, which illustrate the effectiveness and robustness of the proposed approach.
An accurate thrust model is extremely important for the navigation and space mission of solar sails. The thrust is deeply affected by the deformation of the highly flexible structure. Thus, in this paper, the exact thrust models for two-point and infinite-point-connected sails are presented by calculating the static deformations for the sail support beam structure with geometrical nonlinearity based on the assumption that the deformation of the sail film coincides with the support beam. And the film is merely regarded as the structure that transfers the solar radiation pressure force to the support beam. The nonlinear finite element model of the support beam with the Von-Karman’s nonlinear strain–displacement relationships is obtained. Then the Newton iteration method is used to calculate the large deformation of the sail structure. The thrust-modification methods are proposed for the two-connected sail. The deformation of the two-point-connected sail is larger than the infinite-point-connected sail, and the thrust loss of the two-point-connected sail is larger than the infinite-point-connected sail by nonlinear static calculations. Some suggestions are given based on the calculation results and relevant analysis. The thrust model should be verified and modified by in-flight data in the future.
Nowadays, the harmonic drive is widely used as the reducer in the spacecraft manipulator, which may influence the dynamical properties of flexible spacecraft manipulator. The alternative thermal environment makes the spacecraft manipulator to experience periodic heating and cooling in the sunlight and shadow region of the Earth. The analysis of dynamic modeling and motion precision of flexible spacecraft manipulator with harmonic drive, considering the alternate thermal field in orbit is of significant importance for spacecraft manipulator designers in the early stage of design. The thermal load influences the motion precision, which reflects whether the mechanism is performed normally or not. In order to evaluate the loss of motion precision, this paper establishes the dynamical model of spacecraft manipulator with harmonic drive considering the alternate thermal field in orbit. A thermal analysis model of flexible spacecraft manipulator with harmonic drive is developed to characterize the thermal response of the whole spacecraft manipulator system subjected to space heat flux. Two different altitudes including low Earth orbit and geosynchronous Earth orbit are considered. Moreover, the transient temperature fields in different orbits of spacecraft manipulator and the effects of the thermal environment factors on the spacecraft manipulator are investigated. Simulation results reveal the evolution process of the transient temperature field of the spacecraft manipulator system. According to the results, the maximum temperature difference for space manipulator can lead to more severe precision loss compared with the minimum temperature difference. In addition, the vibration frequency of angular velocity error is determined by the maximum thermal heat flux. The proposed method is useful for forecasting the temperature distribution of the spacecraft manipulator system, and will provide meaningful information for performance enhancement of the aerospace facilities.
In this paper, a simple and low-cost three-axis gimbal simulator is introduced. This simulator has been constructed in Amirkabir University of Technology and is used for implementation of attitude control algorithms of remote-sensing satellites in a real time condition using three reaction wheels as hardware in the loop test-bed. This simulator is modeled in Solidworks software package to determine its mass properties in order to utilize in obtaining the dynamic model of the simulator. Afterward, an attitude control algorithm is designed. Performance of the designed attitude control algorithm is investigated by implementing it on the simulator.
In this paper, the unsteady aerodynamics of a ducted fan micro air vehicle is investigated using an unstructured overset grid technique and momentum source method. The in-house programmed compressible Navier–Stoke solver is preconditioned for low Mach number flow regime, and a dual time-stepping strategy is employed to guarantee the computing accuracy and efficiency. Momentum source items are added in the Navier–Stoke solver to replace the contra-rotating propellers in numerical simulation which simplify the inherently unsteady flow into a quasi-steady one. The developed method was verified and validated as a reliable tool for predicting the unsteady aerodynamic performance in low Reynolds flow regime. The effects of reduced frequency, flight velocity and propeller speed on the aerodynamic performance of the ducted fan micro air vehicle are evaluated in this paper. Results show that the hysteresis effect of aerodynamic coefficient increases as induced frequency, freestream velocity and propeller speed increases.
The Doppler frequency changes rapidly due to high dynamics of vehicle, which leads to the loose lock and even the abnormal performance of global positioning system (GPS) receiver. To solve this problem, a federated ultra-tight integration algorithm based on pre-filters is proposed to optimal estimate both receiver tracking control commands and inertial navigation system (INS) navigation solutions. Firstly, the INS error model and GPS receiver tracking loop structure are built to present the fundamental architecture of the proposed ultra-tightly coupled system. Meanwhile, in order to reduce the load of the integrated filter, the pre-filters are incorporated to the ultra-tightly coupled system, and the state variables are fed into the integrated Kalman filter. Secondly, the intrinsic relevance between the phase and frequency biases of replica signals and INS states is analyzed to accomplish the deep fusion of INS and tracking loop. Finally, semi-physical simulations are performed by using a GPS signal simulator to generate signals of two high dynamic trajectories. The experimental results indicate that the proposed ultra-tight integration algorithm can achieve a good performance on reliable positioning and robust tracking in high dynamic environments, compared with the conventional approaches such as tightly coupled integration strategy and third-order phase-locked loops.
This paper presents a tracking control problem of flexible air-breathing hypersonic vehicle with input constraint and aerodynamic uncertainty. Without ignoring aero-propulsive and elevator-to-lift couplings, a control-oriented model including aerodynamic uncertainty is firstly established. Then a robust adaptive backstepping control scheme is designed, in which the control-oriented model does not need to be transformed into linear parameterization formulation. Upper bounds of the uncertain terms do not need to be known in advance, which are estimated online by designing robust adaptive laws. To further consider input constraint, a constrained robust adaptive backstepping controller is proposed to simultaneously handle input constraint and aerodynamic uncertainty. Finally, the compared simulation results show the effectiveness of the designed control strategy.
One of the most important components constituting a homing guided missile is the seeker which basically consists of a detector with a servo-tracking loop. The performance of gimbal seeker is evaluated according to the line of sight (LOS) stability. The purpose of this paper is to present, investigate, and analyze the performance of two axes gimbal seeker which must strictly isolate the LOS from the torque disturbances and missile vibrations. The equations of gimbals motion are derived using Lagrange equation considering the missile angular motion and gimbals mass unbalance. The stabilization loop is constructed by identifying its components, then the traditional and cascade loops are defined. The overall control system is built considering the cross coupling unit and simulated in MATLAB for the traditional and cascade control loops. A comparison study is carried out to investigate the gimbal seeker performance under different operational conditions such as missile rates and accelerations. The simulation results prove the efficiency of the proposed cascade control loop which offers better response more than traditional one, and improves further the transient and the steady-state response of two axes gimbal seeker system.
The two-axis gimbaled antenna’s performance can be greatly improved if it is statically balanced. This paper intends to present a novel design of a measurement system for use in statically balancing a two-axis gimbaled antenna mounted on an aircraft. The details of the measurement system and its working principle are explained, including the dynamics of the two-degree-of-freedom flexure-hinge leverage and the control configuration of the measurement system. The measurement principle is proposed after the theoretical measurement uncertainties estimated and the key factors that determine the measurement accuracy are found. By controlling the uncertainty induced from the major factors, the measurement accuracy can be finally controlled. The measurement result is proved sufficiently accurate by means of High-speed centrifuge method.
Broadening the stable combustion range is particularly desirable for future aircraft engines. The triple swirler is considered to be a promising solution. Experiments were conducted to study different triple swirlers on the performance of a triple swirler combustor, which includes several technology innovations at different inlet airflow velocity (40–70 m/s), temperature (296 K, 373 K, and 473 K), and combustor overall fuel–air ratio with fixed atmospheric pressure. The total pressure loss coefficient increases linearly, while the flow drag coefficient decreases nonlinearly as the inlet airflow velocity increases from 40 m/s to 70 m/s. The flow drag of the combustor assembling counter-rotating swirlers for intermediate swirler and outer swirler is less than that of co-rotating swirlers at the same inlet airflow velocity. The ignition overall fuel–air ratio and lean blowout fuel–air ratio decrease along with inlet airflow velocity and temperature increasing on the whole. The triple swirler with swirl number combination labeling "1.5-1-0.8" has better combustion performance than the other one labeling "0.7-1-1.5". At the temperature of 473 K, the lean blowout fuel–air ratio is almost below 0.005 for the triple swirler with swirl number combination labeling "1.5-1-0.8" at different inlet airflow velocity, and from this point, it has proved the feasibility of the design rules of triple swirler combustor in this paper.
A study that examines the use of aircraft as wind sensors in a terminal area for real-time wind estimation in order to improve aircraft trajectory prediction is presented in this paper. We describe not only different sources in the aircraft systems that provide the variables needed to derivate the wind velocity but the capabilities which allow us to present this information for air traffic management applications. Based on wind speed samples from aircraft landing at Madrid-Barajas airport, a real-time wind field will be estimated using a data processing approach through a minimum variance method. Finally, the accuracy of this procedure will be evaluated for this information to be useful to air traffic control.
This paper presents the main results of the test campaign performed on a frigate ship model in a low-speed wind tunnel in order to investigate the ship superstructure airwake by means of particle image velocimetry (PIV). On board wind velocities measurements above the flight deck were carried out by a sonic anemometer and results were compared with these obtained in wind tunnel tests, providing information about the influence of the ship environment on the helicopter safe operational limitations during launch and recovery operations. The first step in the helicopter–ship qualification program is determine the wind limitations in order to build a candidate launch and recovery wind envelope. Thus subsequent steps of the program, additional effects produced by the helicopter rotor and ship motion must be evaluated, and finally flight trial on the ship must be performed to evaluate the pilot workload.
This study investigated the far-distance cooperative rendezvous problem for two spacecrafts. The orbital dynamics equations were represented based on the orbital elements with an improved vernal equinox and were normalized. Pontryagin’s extremum principle was introduced into the dynamics equations and the co-state equations were obtained. A performance evaluation function was created by particle swarm optimization algorithm based on simulated annealing. The convergent co-state initial vector was obtained using an improved particle swarm optimization algorithm. The initial vector was set as the initial value for optimization and a rapid small-population genetic algorithm was applied, before the approximate global optimum was obtained rapidly. The fine adjustment of the search process was performed based on sequential quadratic programming and the results were sufficiently precise. The process of optimization was simulated for problems that involved far-distance coplanar cooperative rendezvous and active-passive rendezvous, which showed that cooperative rendezvous had more advantages than active-passive rendezvous in terms of fuel saving and time.
Wind tunnel and numerical results are presented from a 33% scale model of a Scottish Aviation Bulldog light aircraft. The model was developed using reverse engineering and computer aided design processes from a laser scan of the full scale aircraft. This solid model was subsequently used to provide a basic aerodynamic wind tunnel assessment of the aircraft, specifically in the region behind the canopy. The computer aided design model was also meshed with 3.4 million cells in Ansys ICEM CFD and solved using Ansys Fluent. The CFD solution was verified and validated using comparisons with flight test and type record data. Subsequent comparisons of the CFD pressure data behind the canopy with the wind tunnel data was found to match within a Cp of 0.05 which was within experimental error and scaling effects.
The center of gravity variations have a direct impact on the dynamic and the quality characteristics of the aircraft, which makes the control of the aircraft more difficult after center of gravity shifting. In order to solve this problem, an aircraft model that can simulate both the instantaneous and gradual center of gravity shift has been built as research object. Based on this model, an adaptive nonsingular fast terminal sliding mode controller is proposed to control the research object. Fast nonsingular terminal sliding mode has been combined with adaptive control method in the controller, in which the improved attractor can eliminate the chattering phenomenon and the nonlinear adaptive law can compensate the system disturbance caused by the center of gravity variation. The stability of closed loop is proved by using Lyapunov stability theory. The simulation results show that the proposed controller can realize the fast and precise track of the command.
This paper addresses the positioning problem of an unmanned airship in the presence of parametric uncertainties and external disturbances. A sliding mode controller (SMC) based on fuzzy approximation is proposed that steers an airship to remain fixed over a desired position. First, the dynamic model of an airship is derived and formulated. Second, a SMC is designed to actualize positioning control under the assumption that the airship model is accurately known. However, the airship model is partially or totally unknown in practice. In order to solve this problem, a fuzzy logic system is used to approximate the unknown model of the airship, and an adaptive law is adopted to update the optimal parameters. The stability and convergence of the closed-loop controller is proven by using the Lyapunov stability theorem. Finally, the effectiveness and robustness of the proposed controller are demonstrated via simulation results. Contrasting simulation results indicate that the proposed controller promotes the control precise and has better performance against the SMC.
The staged injection scheme has drawn an increasing attention for the airbreathing hypersonic propulsion system, and the fuel injection angle has a large impact on the mixing improvement between the fuel and the supersonic cross flow. The Reynolds-averaged Navier–Stokes equations associated with the SST k- turbulence model have been employed to investigate the interaction mechanism in the staged sonic injection flow field, and the influences of the injection angle, the injection angle arrangement, and the distance between the injectors on the flow field characteristics have been analyzed comprehensively. At the same time, three grid scales have been used to perform the grid independency analysis, and the predicted results have been compared with the experimental data in the open literature for the single transverse injection scheme. The obtained results show that the penetration height for the cases with the distance between the injectors being 1 mm is the highest in the range considered in the current study, and this may be due to the strongest shock wave/shock wave interaction between the injectors. At the same time, due to the blockage of the fuel injection, the penetration height increases as the supersonic air stream flows downstream, and the influence of the wave system generated by the first and third injectors cannot propagate downstream and upstream, respectively. The multi-port injection scheme can provide better fuel penetration performance than the single one when the flow flux keeps constant, and the multi-port injection scheme with a certain angle can provide a higher total pressure recovery efficiency than the staged transverse injection scheme. Further, the staged transverse injection flow field can provide a better recirculation zone for the mixing between the fuel jet and the boundary layer, and the separation length increases with the increase of the distance between the injectors.
The impingement of hot rocket motor plume inside a canister is simulated numerically by solving three-dimensional Reynolds Averaged Navier–Stokes equations using commercial software. The computed methodology is first validated for cold flow jet impingement in a circular tube for different chamber pressure and the simulations captured all the finer aspects of blow-by flow conditions as reported in the literature. A very good comparison is obtained between experimental and numerical surface pressure distribution. The validated methodology is applied to simulate the hot launch of a missile from a canister. It is observed that for low annular gap between missile body and canister the motor plume interaction became intense and gave rise to a very significant base drag which may constrain the motion of the missile inside the canister.
This paper presents an extension of fuzzy-multi-objective genetic algorithm (MOGA) optimization methodology that could effectively be used to find the overall satisfaction of objective functions (selecting the design variables) in the early stages of design process. The coupling of objective functions due to design variables in an engineering design process will result in difficulties in design optimization problems. The primary application of this methodology is the design of a liquid propellant engine with the maximum specific impulse and the minimum weight. The independent design variables in this model are combustion chamber pressure, exit pressure, oxidizer to fuel mass flow rate. To handle the mentioned problems, a fuzzy-multi-objective genetic algorithm optimization methodology is developed based on Pareto optimal set. Liquid propellant engine, F-1 is modeled to illustrate accuracy and efficiency of proposed methodology.
Cavities are widely used as flameholders in supersonic combustors due to their outstanding potential to stabilize combustion without excessive total pressure loss. A review of cavity-stabilized combustion for scramjet applications is provided in this article. The topics cover the fundamental problems and recent advances regarding cavity-organized combustion in high-speed flows, including combustion stabilization modes and mechanisms, flame stability analyses and correlations, combustion oscillations, and other related issues. Remarkable questions such as cavity-coupled fuel injection, flow and combustion coupling, optimal cavity geometry and scale, auto-ignition and flame propagation interactions, and unsteady effects are discussed. Then, an attempt is made to provide some guidelines for the future research of cavity flameholders.
Tethered satellite with chemical propulsion has broad application prospects in the space debris removal, the orbital transfer of space detector and the orbital rescue of malfunctioning satellite. In orbital maneuvering, tethered satellite with a short constant tether can avoid using certain windlass mechanism of base satellite, which is helpful for the implementation of project. In this article, based on a dumbbell model of tethered satellite, dynamic equations of tethered system in orbital maneuvering are established. Furthermore, taking elastic strain of the tether into account in the dynamic model, as the slackness of tether occurs, the effects on tethered satellite of the degree of slackness, initial states of librational angles and the variation of thrust acceleration are analyzed. To avoid several adverse phenomena aroused by the slackness of tether, such as tether winding or collision between satellites, a thrust control method of tethered satellite with a short constant tether in orbital maneuvering is proposed. In this method, the thrust acceleration imposed on the base satellite can be adjusted to avoid the slackness of tether and damp out the librational angles; meanwhile, it is required that the regulating value of thrust acceleration meets with accuracy requirements of orbital trajectory in practical engineering; therefore, a continuous thrust controller is presented based on the feedback of tether tension; besides, considering in practical engineering that the continuous thrust is always replaced by an impulse thrust, the ranges of impulse thrust parameters, such as impulse width and duty cycle, are studied. Afterwards, an orbital transfer case between two circular orbits is studied to demonstrate the effectiveness of the tethered satellite with a short constant tether in orbital maneuvering. In this case, an orbital transfer strategy for tethered satellite is designed based on a continuous thrust. Numerical simulation results show that the slackness of tether can be eliminated and the librational angles are damped out according to the thrust control scheme in orbital maneuvering; in addition, the stability of tethered system could be guaranteed by the designed thrust controller, which is useful for flight safety.
This article proposes a new strategy that computes the bank-to-turn commands without a singularity problem. To this end, the singularity problem is first analysed, and the main influence factors are found. An extended roll-angle command calculation method is then derived for the missile body coordinate based on the bank-to-turn-90 logic. The auxiliary skid-to-turn manoeuvring and the command increment saturation are induced to eliminate the oscillation of roll-angle command due to the noises in guidance commands. Three control zones are designed to ensure that suitable command calculations are for different conditions. When the strategy is used, the missile tends to maintain a smooth and varied roll-angle command, even if the guidance acceleration commands approach zero at the endgame of guidance. Finally, numerical simulation results are provided, and the validity of the strategy is proven via a comparison between the typical bank-to-turn guidance law and the normal bank-to-turn command calculation method.
Buzz is an important issue for a scramjet engine. A mathematical model of buzz oscillations is necessary for control system design. Control-oriented models of hypersonic vehicle propulsion systems require a reduced-order model that is accurate to some extent but requires less than a few seconds of computational time. To achieve this goal, a reduced-order model of buzz oscillations for a scramjet engine is built by introducing the modeling idea of Moore–Greitzed model for compressors. The introduction of characteristic lines avoids the complex interactions in hypersonic inlet, such as shock–shock interactions and shock–boundary layer interaction. And the inlet characteristics are obtained from the pressure signal of combustor. Based on the established buzz model, we can predict the inlet performance, characterize the stability margin of inlet, reflect the oscillatory characteristics of inlet buzz including the dominant amplitude and frequency and describe the transition process of inlet buzz.
A novel inertial measurement unit scheme with five accelerometers and two gyros (5A2G) is proposed in this paper to achieve high precision measurement in high dynamic environment. The three channels are decoupled during the angular velocity calculation procedure to ensure the precision and efficiency. The yawing and pitching angular velocities are directly measured by gyros, while only the rolling angular velocity is inferred indirectly from the rolling angular information vector composed of rolling angular acceleration and quadratic component of rolling angular velocity. Based on the proposed scheme, the configuration ascertainment problem for acquiring the required installation parameters of accelerometers is transformed into a constraint optimization problem with the objective of minimizing the error of rolling angular information vector. A single channel rolling angular velocity calculation model is then established on the basis of the optimal configuration and the extended Kalman filter algorithm is utilized for state estimation. Simulations are implemented and results indicate that the optimal 5A2G scheme is feasible for high-speed rotating ammunition.
Due to the unconventional nature of the blended wing body (BWB) no off-the-shelf software package exists for its conceptual design. This study details a first step towards the implementation of traditional and BWB-specific design and analysis methods into a software tool to enable preliminary sizing of a BWB. The tool is able to generate and analyze different BWB configurations on a conceptual level. This paper investigates three different BWB configurations. The first configuration is an aft-swept BWB with aft-mounted engines, the second configuration is an aft-swept BWB with wing-mounted engines and the third configuration is a forward-swept BWB with wing-mounted engines. These aircraft comply with the same set of top-level requirements and airworthiness requirements. Each of the designs has been optimized for maximum harmonic range, while keeping its maximum take-off weight constant and identical. Results show that the forward-swept configuration with wing-mounted engines has the highest harmonic range. These findings warrant further investigation in this configuration and other alternative BWB configurations.
The near-field (up to three chord lengths) development of a wing-tip vortex is investigated both numerically and experimentally. The research was conducted in a medium speed wind tunnel on a NACA 0012 square tip half-wing at a Reynolds number of 3.2 x 105. A full Reynolds stress turbulence model with a hybrid unstructured grid was used to compute the wing-tip vortex in the near field while an x-wire anemometer and five-hole probe recorded the experimental results. The mean flow of the computed vortex was in good agreement with experiment as the circulation parameter was within 6% of the experimental value at x/c = 0 for α = 10° and the crossflow velocity magnitude was within 1% of the experimental value at x/c = 1 for α = 5°. The trajectory of the computed vortex was also in good agreement as it had moved inboard by the same amount (10% chord) as the experimental vortex at the last measurement location. The axial velocity excess is under predicted for α = 10°, whereas the velocity deficit is in relatively good agreement for α = 5°. The computed Reynolds shear stress component <u'v'> is in good agreement with experiment at x/c = 0 for α = 5°, but is greatly under predicted further downstream and at all locations for α = 10°. It is thought that a lack of local grid refinement in the vortex core and deficiencies in the Reynolds stress turbulence model may have led to errors in the mean flow and turbulence results respectively.
The design parameters of high-altitude solar-powered aircraft are highly correlative with its flight trajectory. However, it is not an easy work to jointly optimize them in the concept design stage. This paper considers the joint optimization problem of battery mass and flight trajectory for high-altitude solar-powered aircraft. The system model including the aircraft dynamic model, aerodynamic parameters, and thrust model is presented. Then the problem to be optimized is formulated and a new optimization method, which uses the particle swarm optimization and Gauss pseudo-spectral method, is proposed. The Gauss pseudo-spectral method is employed to generate the minimal power consumed by following the flight trajectory in the given configuration of high-altitude solar-powered aircraft, while the particle swarm optimization is used to calculate the optimal battery mass of aircraft. The simulation result shows that the proposed joint optimization method can reduce the battery mass of high-altitude solar-powered aircraft from 16 kg to 13.6 kg, which is equivalent to enhancing its energy density by 19.7%. It can be also seen that the proposed optimization method connects each parameter in a logically clear way and hence provide a perspective for understanding the optimization problem.
The hybrid rocket motor is a kind of chemical propulsion system that has been recently given serious attention by various industries and research centers. The relative simplicity, safety and low cost of this motor, in comparison with other chemical propulsion motors, are the most important reasons for such interest. Moreover, throttle-ability and thrust variability on demand are additional advantages of this type of motor. In this paper, the result of an internal ballistic simulation of hybrid rocket motor in a zero-dimensional form is presented. Further to validate the code, an experimental setup was designed and manufactured. The simulation results are compared with the experimental data and good agreement is achieved. The effect of various parameters on the motor performance and on the combustion products is also investigated. It is found that increasing the oxidizer flow rate, increases the pressure and specific impulse of the motor; however, the slope of the specific impulse for the high flow rate case reduces. In addition, by increasing the combustion chamber pressure, the specific impulse is increased considerably. The initial diameter of the fuel port does not have significant effect on the pressure chamber and on the specific impulse. Addition of a percentage of an oxidizer like ammonium perchlorate to the fuel increases the specific impulse linearly.
A new design method for pulse detonation engines nozzle was developed theoretically. The effects of non-uniform exhaust on the performance of pulse detonation engine were analyzed by constant volume cycle model. The results showed thrust losses induced by the non-uniform exhaust could be decreased by increasing fill pressure ratio. If the fill pressure ratio was larger than 10, the performance losses with a fixed optimal nozzle could be controlled within 3%. The optimal area ratio of the nozzle was obtained when the time-averaged pressure at the nozzle exit equals the ambient pressure. This was also applicable to one-dimensional unsteady frictionless pulse detonation engine model. Thus an optimal area of the nozzle could be calculated by the time-averaged total pressure. Compared with the zero-dimensional results obtained by numerical search technique, the errors of predicted optimal area could be neglected if fill pressure ratio is too large to prevent shock from propagating back to the nozzle. And the errors of predicted optimal area are lower than 5% compared with the results of the one-dimensional unsteady pulse detonation engine model.
This paper addresses the functional verification and performance assessment of an ultra-low shock non-explosive actuator appropriate to space applications of hold-down and release mechanisms. To demonstrate that the design implementation and manufacturing methods have resulted in an engineering model conforming to the set of functional, performance and environmental requirements specified, a space qualification test campaign is typically required. To ensure the readiness of the engineering model and the adequacy of the mechanical and electrical ground support equipment required for the entire qualification test campaign, a set of functional verification procedures and performance characterization tests were systematized and undertaken before the mechanism qualification. A preload monitoring system was developed and calibrated, and the performance of the mechanism was evaluated through the estimation of the release time and the measurement of the self-generated shock. The main results and conclusions taken from these tests are presented and discussed here.
An evaluation model for stratospheric airship energy storage system selection is developed, which provides a new method for quantitative selection of renewable energy storage system. Firstly, some basic properties and their indexes are proposed to evaluate the overall performance of stratospheric airship energy storage system by qualitative analysis of airship’s operation principium and operation environment. Secondly, the weights to be used as subperformance indexes coefficients of the model are obtained with analytic hierarchy process method. The normalization of the subperformance indexes is implemented by physical programming method which takes the decision maker's preferable degree into account and can convert the indexes with different physical meaning and magnitude into nondimensionless satisfaction level evaluation scores with the magnitude in the same quantity level. Thirdly, the weights and evaluation scores of candidate designs’ indexes are combined with the aggregate objective function in the form of specific scores and the optimal energy storage system concept can be found out. Finally, an example of stratospheric airship energy storage system selection is given to illustrate this method. Moreover, the method presented in this paper can be effectively applied to various decision-making scenarios.
A large variety of promising power and propulsion system concepts are being proposed to reduce carbon dioxide and other emissions. However, the best candidate to pursue is difficult to select and it is imperative that major investments are correctly targeted to deliver environmentally friendly, economical and reliable solutions. To conceive and assess gas turbine engines with minimum environmental impact and lowest cost of ownership in a variety of emission legislation scenarios and emissions taxation policies, a tool based on a techno-economic and environmental risk assessment methodology is required. A tool based on this approach has been developed by the authors. The core of the tool is a detailed and rigorous thermodynamic representation of power plants, around which other modules can be coupled (that model different disciplines such as aircraft performance, economics, emissions, noise, weight and cost) resulting in a multidisciplinary framework. This approach can be used for efficient and cost-effective design space exploration in the civil aviation, power generation, marine, and oil and gas fields. In the present work, a conceptual intercooled core aeroengine design was assessed with component technologies consistent with 2020 entry into service via a multidisciplinary optimisation approach. Such an approach is necessary to assess the trade-off between asset life, operating costs and technical specification. This paper examines the influence of fuel consumption, engine weight and direct operating costs with respect to extending the engine life. The principal modes of failure such as creep, fatigue and oxidation, are considered in the engine life estimation. Multidisciplinary optimisation, comprising the main engine design parameters, was carried out with maximum time between overhaul as the objective function. The trade-off between minimum block fuel burn and maximum engine life was examined; the results were compared against the initial engine design and an assessment was made to identify the design changes required for obtaining an improved engine design in terms of direct operating costs. The results obtained from the study demonstrate that an engine optimised for maximum time between overhaul requires a lower overall pressure ratio and specific thrust but this comes at the cost of lower thermal efficiency and higher engine production costs.
A vision-aided terrain referenced navigation (VATRN) approach is addressed for autonomous navigation of unmanned aerial vehicles (UAVs) under GPS-denied conditions. A typical terrain referenced navigation (TRN) algorithm blends inertial navigation data with measured terrain information to estimate vehicle’s position. In this paper, a low-cost inertial navigation system (INS) for UAVs is supplemented with a monocular vision-aided navigation system and terrain height measurements. A point mass filter based on Bayesian estimation is employed as a TRN algorithm. Homograpies are established to estimate the vehicle’s relative translational motion using ground features with simple assumptions. And the error analysis in homography estimation is explored to estimate the error covariance matrix associated with the visual odometry data. The estimated error covariance is delivered to the TRN algorithm for robust estimation. Furthermore, multiple ground features tracked by image observations are utilized as multiple height measurements to improve the performance of the VATRN algorithm.
The attitude control subsystem plays a significant role in the overall performance of the spacecraft. Attitude control subsystem is vitally important to design the control system with rapid response performance, high control precision and insensitive to external perturbations. In this paper a novel fault-tolerant control design technique against faulty thrusters is investigated. This technique uses adaptive sliding mode control with application to spacecraft attitude maneuvering control system. The principle of the proposed fault-tolerant control scheme is to design sliding mode attitude controller using the time variable sliding surface to compensate the effect of partial loss of the actuators effectiveness. This adaptive law calculates the ability of spacecraft maneuvering in following the control input based on kinematic energy of the estimated and real model of the spacecraft. It is shown that the presented controller can accommodate the actuator faults, even while resisting the external disturbances. Moreover, in the control law scheme the effect of actuator saturation/constraint has been considered. An additional advantage of the proposed fault-tolerant control strategy is that the control design does not require a fault detection and isolation mechanism to detect, separate, and identify the actuator faults on-line. The associated stability proof is constructive and accomplished by the development of Lyapunov function candidate, which shows that the attitude orientation and angular velocity will globally asymptotically converge to zero. Moreover, several numerical examples are presented to demonstrate the efficacy of the proposed controller despite the external perturbations, moment of inertia uncertainty and faulty actuators. The numerical results clearly demonstrate the good performance of the adaptive sliding mode control despite the actuator fault comparison with some other controllers.
In this paper, the design of attitude and airspeed controls for a fixed-wing unmanned aerial vehicle by means of an Adaptive Super Twisting Algorithm approach is addressed. In order to implement these controllers and taking into account the difficulties for measuring some of its states, necessary information about inertial attitude and airspeed is estimated using Super-Twisting Observers. This control scheme increases robustness since it is not necessary to know the bound of the perturbations affecting the system or the exact values from the parameters of the system. Furthermore, due to the finite-time converge of the observer, the stability of the closed-loop system is guaranteed. Simulation results illustrate the performance of the proposed scheme under unmodelled dynamics, noisy measurements and external disturbances.
Aerogine noise leads to environment pollution largely when aerogine is tested. In this paper, the power spectrum analysis method of the aeroengine test noise was discussed, and the noise measurement and analysis experiments of a turbojet engine and a turbofan engine tests were carried out. The noise level, main noise resource, and noise characteristics of the two turbojet and turbofan engines were analyzed. Meanwhile, the indoor noise and far-field noise of the turbojet engine were both measured, the noise spread characteristics were analyzed and the noise reduction performance of the test bench was evaluated. The noise generated by the turbojet engine test had the discrete characteristic of high frequency. The higher frequencies when peak values occurred were the blade passage frequencies and the noises with lower frequencies were the broad band noises, especially the jet noise, and the maximum of the peak values occurred at the basic frequencies or harmonic frequencies of the compressor. Meanwhile, the noises generated by the turbofan engine, focused on the high frequencies and the peak values corresponded to the rotation noise of the fan blades. The experimental results were consistent with the theory basically, which indicated that the aeroengine operating status information could be identified by the noise power spectrum analysis. In addition to the aeroengine noise reduction research, the noise power spectrum analysis could also be used to diagnose the fault of the aeroengine structure and performance. On the other hand, the indoor and far-field noise measurement experimental results implied that the noise was suppressed from 136 dB to 85 dB and could provide the reference to the noise reduction design of the aeroengine test bench.
Recently, rapid repair of damaged blade has become the focus of considerable interest for extending its service life. However, due to the defects caused by high temperature and pressure of operations as well as foreign object impact, the turbine blades often undergo the deviations of the actual part profile from its design model, such that this nominal Computer Aided Design (CAD) model cannot be directly used in the process of repair for tool path generation of laser cladding and Numerical Control (NC) machining, thus to nicely repair the damaged or worn blades, it is necessary to reconstruct the surface model of the actual blade. This paper develops a deformable template-based approach to recovering the surface of blade from the cross-sectional profiles. The mathematical model for cross-sectional profile reconstruction is first established and is then solved by an alternate iteration optimization strategy consisting of registration and deformation of the template curve. Since the proposed method can automatically transform and deform the template curve to best fit the cross-sectional points, the compatibility conditions between different sections are automatically satisfied and there is no need for the data preprocessing such as data sorting, parameterization, etc. which are necessary for the traditional surface fitting methods. Undoubtedly, this considerably simplifies the reconstruction problem of the damaged blade and nicely adapts to blade part-to-part variation. Moreover, a method of closest point computation that combines the arithmetic for Bernstein-form polynomials and Bézier curve subdivision is also given based on bintree decomposition to improve the iteration processes of 2D profile reconstruction. Then, according to these reconstructed sectional profiles, the actual blade surface is reconstructed by surface skinning operations. Finally, the proposed method is tested on a sample blade, and the experimental results show that our method can precisely reconstruct the surface of the damaged blade, especially for the blades with area defects.
To improve the performance of aero gas turbine engines, more and more interests have been shown on turbine inter-guide-vane burner based on the ultra-compact combustion concept. To make a universal turbine inter-guide-vane burner, a new concept is proposed using a trapped vortex cavity to replace the high swirling circumferential cavity combustor to address the need to scale the configuration for a larger turbine spool. Three models, including trapped vortex combustor, transition model, and turbine inter-guide-vane burner, are designed. Comparative analysis between combustion performances of three models by using numerical simulation method is carried out. The scale-adaptive simulation turbulence model is used in the simulation process, aiming to reduce the deviation between numerical simulation value and actual value. Finally, the turbine inter-guide-vane burner model is found to be the superior design proposal for turbine inter-guide-vane combustion technology, compared with the other two models.
An L-shaped tab was tested at the trailing edge of an oscillating airfoil to evaluate its effects on blades aerodynamic performance. The tests were conducted on a NACA 23012 pitching airfoil in deep dynamic stall conditions with the L-shaped tab fixed in two different positions. When deployed the tab is attached to the airfoil upper surface so that the end prong protrudes at the airfoil trailing edge. In retracted position the tab features an angle of 9.1° with the airfoil upper surface, since its prong tip touches the airfoil trailing edge. The airloads time histories during a pitching cycle were evaluated by pressure measurements carried out on the airfoil midspan contour. The phase-averaged flow field at the trailing edge region was investigated by means of particle image velocimetry to evaluate the detailed flow physics involved in the use of the device. The experimental results indicate that the use of such a pivoting L-shaped tab can introduce similar effects to those that can be obtained by the use of an active Gurney flap. Thus, the L-shaped tab can be considered an attractive device due to its easier integration on helicopter blades.
In geomagnetic aided navigation, directional matching suitability can be depicted by the directional features extracted from candidate matching areas. First, Gabor filtering and gray-level co-occurrence matrix are used to extract frequency-domain and spatial-domain directional features, respectively. Meanwhile, the parameter settings of the above methods are also discussed in order to make the extracted features correctly reflect the directional matching suitability. Then, adaptive neuro-fuzzy inference system is utilized for modeling the complementary relationship between Gabor filtering and gray-level co-occurrence matrix with the purpose of playing their respective advantages in directional matching suitability analysis. Afterward, a hierarchical decision-making scheme is designed, where the first stage is to use adaptive neuro-fuzzy inference system for selecting an appropriate analysis method (Gabor filtering or gray-level co-occurrence matrix) based on the characteristics of the given candidate matching area, and the second stage is to utilize the selected method for directional matching suitability analysis. Experimental results show that the proposed scheme is effective, and the conclusions can afford credible guidance for geomagnetic matching.
In this work, a real-time vision-based algorithm has been developed and implemented on a flying robot, in order to detect and identify a light beacon in the presence of excessive colored noise and interference. Starting from very basic and simple image analysis techniques including color histograms, filtering techniques, and color space analyses, typical pixel-based characteristics or a model of the light beacon has been progressively established. It has been found that not only are various color space-based characteristics significant, but also the relationships between various channels across different color spaces are of great consequence, in a beacon detection problem, specifically referring to a blue light-emitting diode. A block-based search algorithm comprising of multiple thresholds and linear confidence level calculation has been implemented to search established model characteristics in real-time video image data. During implementation, once excessive noise was encountered during flight tests, a simple and low cost noise and interference filter was developed. This filter very effectively handled all noise encountered in real time. The proposed work was successfully implemented and utilized on GeorgiaTech’s participating aircraft for the International Aerial Robotics Competition by Association for Unmanned Vehicle Systems International for detection of a blue light-emitting diode problem. Major contributions of this work include establishing a multiple threshold search and detection algorithm based on not only various color channels but also their relationships and handling of as much as 40% noisy or interfered video data with successful practical implementation and demonstration of proposed approach.
The singularity is the inherent characteristics of parallel manipulators. At near-singular configurations, the parallel manipulator cannot resist externally applied force/torque along certain directions. The axisymmetric vectoring exhaust nozzle is driven by the 3-SPS + 3-PRS parallel manipulator to change its exit area A9 and make the universal vector of its divergent section. Preventing the 3-SPS + 3-PRS parallel manipulator from falling into the singular configuration is very important for the maneuverability and safety of the jet-thrust aircraft equipping with the axisymmetric vectoring exhaust nozzle. In this paper, a methodology to eliminate the singularity of the 3-SPS + 3-PRS parallel manipulator is presented. At first, with the aid of the configuration homotopy-tracing algorithm, the configuration curves relative to the input parameters are figured out. It is found that the singularity-free zone corresponding to the input parameter exists between the left and the right extreme singular positions. Based on the extended equation algorithm, the curves of singular points going with the input parameters are drawn. By selecting the suitable initial working point and letting the input parameters locate within the singularity-free zones of input parameters determined by these curves, the singularity can be eliminated in the design stage. The method to eliminate the singularity presented in this paper is simple, efficient, and easy to be implemented directly through inspecting the lengths of the input parameters.
Experimental studies conducted during the 70s and 80s of the previous century are numerically simulated. We examine a horizontal duct with a vertical branch having a circular cross section whose diameter is 5 cm. These experiments were conducted by the late Dr Heilig in the Ernst-Mach-Institute (private communication). In both segments of the branched duct pressure transducers were installed. They were used for recording the pressure histories and for deducing the traveling shock wave speed. These results were compared with the present numerical simulation. The numerical simulations were conducted using the commercial code Fluent with the density-based AUSM solver. The solver is second order in both space and time. It is apparent from the results obtained that good agreement exists between the recorded pressure histories and their simulations. Based on the good agreement between recorded and simulated pressures a numerical study was conducted by comparison between two similar branched ducts, one having a circular cross section while the other has a rectangular cross section. Also, the effect that changes in the branched segment orientation have on the resulting flow field were investigated.
Although systems engineering processes and standards are widely used in aircraft development programs, traditional requirements’ engineering practice for commercial aircraft does not explicitly address value perceptions and associated information. In this paper, a value-focused approach is proposed to promote a better understanding of customer-value perceptions and their derivation among different levels for value-based requirements engineering of commercial aircraft. The approach is a four-step process starting from initial customer statements to a customer-value model and leading to a system-value model with associated component-value models. A set of theories and methods are introduced in order to resolve different aspects of the approach regarding the appropriate understanding of customer-value perceptions and the establishment of the value-based requirements’ specification. A case study is used to demonstrate the transformation of airlines’ initial expectation statements into three types of value models. There are two significant benefits of this approach: (a) perceived customer value can be explicitly modeled, simulated, and derived into different levels of the system development and (b) the value model can be subsequently utilized reactively for design evaluations and proactively for design optimization to generate creative design alternatives.
In this paper, an autonomous orbit control of a satellite in Low Earth Orbit is investigated using model predictive control. The absolute orbit control problem is transformed to a relative orbit control problem in which the desired states of the reference orbit are the orbital elements of a virtual satellite which is not affected by undesirable perturbations. The relative motion is modeled by Gauss’s variational equations including J2 and drag perturbations which are the dominant perturbations in Low Earth Orbit. The advantage of using Gauss’s variational equations over the Cartesian formulations is that not only the linearization errors are much smaller, but also each orbital element can be controlled independently. Model predictive control finds the finite horizon optimal firing times of the satellite thrusters. The problem of orbit control has been cast as a linear programming which is a subset of convex optimization problems. As a result, model predictive control can maintain and control orbits of Low Earth Orbit satellites in optimal way, and this modern control technique can be an alternative for traditional analytical-based orbit control methods. Also, a comparison between model predictive control and linear quadratic regulator orbit control showed the superiority of MPC in fuel consumption.
Design of an adaptive dynamic feedback-linearization control law for a quadrotor unmanned aerial vehicle under uncertain parameters is presented. Because the quadrotor carries rotational speed-varying thrusters, it has the advantage of simple mechanism compared to the pitch-varying thrusters. However, it is subjected to slow dynamics in thruster and suffers from uncertainties in efficiency due to power subsystem. Additionally, parametric uncertainties tend to exist such as thruster misalignment, mass, and inertia. The control law is targeted to tracking reference trajectories under such uncertainties. Dynamic feedback-linearization method is employed primarily to produce the small-bandwidth thruster signal. A dynamic observer is used to estimate the states of feedback-linearized system, and Lyapunov-based update laws are derived to compensate for uncertain parameters. The controller and its performance are evaluated using a nonlinear, six-degree-of-freedom dynamic model of a quadrotor unmanned aerial vehicle with a thruster model in the simulation. The results illustrate that the proposed control law enhances tracking performance even with slow and misaligned thruster.
In this paper, an optimal distribution algorithm for a large group of heterogeneous unmanned aerial vehicles is developed. A typical unmanned aerial vehicle cooperative control in a battlefield can be categorized as a hierarchical system that is usually composed of several levels, and the decision making step, or the resource management step, is the main focus of this paper. In the resource management step, the factors to be decided are the proper number and types of unmanned aerial vehicles that will be committed to each operational area to increase the overall performance of the entire group and achieve a successful mission accomplishment. A task assignment algorithm, which is the next level in the cooperative control hierarchy, may begin with a higher chance of success when the number and types of resources are given correctly by the resource management step. This research suggests an optimal resource management algorithm for operations in various combat or civilian missions by solving an integer linear programming problem. A Suppression of Enemy Air Defense (SEAD) mission is considered as the main example in this paper. Finally, the algorithm is supported with number of verifications and numerical simulations in various SEAD mission cases.
A combination of fluid network analysis method with conjugate heat transfer are applied to the improvement design of the integrated cooling structures in a high-performance turbine blade, coupled with the 3D viscous solver for the gas flow field. By comparison with the experimental results of open literatures, the methodology developed is numerically validated. For a high-pressure turbine rotor blade, it is used to rapidly predict and evaluate the aerodynamic and heat transfer performances of its integrated inner cooling structures. According to the computation results, three ways are definitely proposed for the improvement design, including the adjustment of the coolant flow mass entering into the front and rear cavities in a more appropriate flow mass ratio, the improvement of the turning geometries in serpentine channels to minimize the inner coolant flow resistance, and the adjustment of the local cooling structure dimension according to the high temperature region on outer surface of blade. Through the verification of the fully 3D conjugate heat transfer simulation for the fields of gas flow, solid blade and coolant flow, it shows that the maximum temperature on rotor blade surface is reduced obviously, the temperature distribution becomes more uniform after improvement, and the inlet parameters of cooling cavities are matched more reasonably. It is concluded that in this paper the fluid network combined with conjugate heat transfer significantly shortens the aerodynamic and heat transfer design cycle for the turbine blade with integrated cooling structures.
A supersonic combustion organizer that consists of both cavity flameholder and strut injector was applied in a liquid-kerosene-fueled model scramjet. The experimental results indicated that the strut injection can improve the combustion performance. When the strut was mounted near the cavity, transverse injection from the strut gave the best performance. However, the excessive long distance between the upstream strut and the cavity led to upstream spreading of combustion to the isolator and pressure rise at the isolator entrance. Besides that, parallel injection was found difficult to establish effective combustion due to the poor spreading performance, except in the condition that the strut was mounted close to the cavity and wall injection was used simultaneously.
This paper introduces a concept, a baseline design, and a trade study for a new space-based global continuous disaster monitoring system composed of a dual-mode satellite constellation and on-orbit propellant depots. The proposed constellation operates in two different modes: a normal mode and a disaster mode, which are responsible for atmospheric/oceanic imaging and disaster monitoring, respectively. The dual-mode concept enables the system to manage the uncertainties associated with the unknown time and location of a disaster and to enhance its operational efficiency by improving its utilization. The mode-change requires orbit transfers accompanying large amounts of fuel consumption, and this challenge is addressed by an on-orbit refueling system to support the constellation. A reference design for the proposed satellite constellation and the orbiting depot is presented. Orbital parameters and the options for mode-change transfers are explored considering the trade-off relationships among the propellant consumption (to minimize), the response time (to minimize), and the access area in normal mode (to maximize). Options for the number of on-orbit propellant depots and the drift rate for the refueling operation are also explored considering the time to complete the preparation and associated probability to get ready for the next disaster outbreak.
In this study, a general assessment of inverse trigonometric shear deformation theory, recently developed by the authors, is performed and the structural responses (static, buckling, and free vibration) of laminated-composite and sandwich plates are investigated. The in-plane displacement components are expressed in terms of an inverse cotangent function, which yields the nonlinear shear deformation while the constant transverse displacement is assumed over the thickness of the plate. A computationally efficient finite element model for the implementation of above-mentioned theory is proposed. The continuity requirement of the finite element model is maintained as C0 by a suitable choice of nodal field variables. Numerous analysis problems are selected to study the effects of various parameters such as span-to-thickness ratio, lamination sequence, loading conditions, boundary conditions, etc. on the response characteristics of plates. Higher modes are also obtained for the buckling and vibration problems and the ability to investigate higher modes is ensured. The comparison of the present results with the established results in literature indicates the efficiency and range of applicability of the present formulation. Moreover, the formulation is presented in a generalized approach which enables the implementation of all existing seven degree-of-freedom theories in a single computer algorithm thereby making it practically more significant.
Numerical experiments are carried out using commercially available Navier–Stokes solver to investigate the effect of forward-facing parabolic cavity on the heat fluxes over a spherical nosed blunt body. A wide range of parabolic cavities with depths varying between 2 and 10 mm placed at the nose of sphere-cylinder with base diameter 40 mm and overall length 70 mm have been investigated. The ratio of the cavity radius at intersection with y-axis to depth of cavity (r/d) of these cavities varies from 1.5 to 2.5. All computations have been done at a freestream Mach number of 6.2 and sea level atmospheric conditions assuming air to be a thermally perfect gas. The steady-state solutions obtained through time marching solution of axisymmetric Navier–Stokes equations suggest that the total heat transfer rate, area weighted average heat flux and the peak heat fluxes to the body can be favorably reduced for shallow parabolic cavities.
A new analytical tool is proposed to aid in the design and performance evaluation of advanced guided projectile concepts. A projectile linear theory applicable to aerodynamically asymmetric configurations is created, leading to a linear, periodic dynamic system. Utilizing concepts in Floquet theory, stability of asymmetric projectile configurations is explored. While stability of many asymmetric projectile configurations can be accurately predicted using averaged linear, constant coefficient system dynamics, there are some configurations where the use of linear periodic systems theory is required for accurate stability prediction. This fact is shown by comparing stability boundaries of an example projectile configuration using the conventional projectile linear theory model and the new periodic projectile linear theory model.
Most smart projectile control systems generate lateral control forces to guide the round to a target. Experience has shown that under the right combination of body orientation, translational velocity, and angular velocity, relatively low lateral control force inputs can induce instability of the round. To solve this problem, an additional control logic layer is appended to a nominal impact point flight control law to protect it from instability in these infrequent, but consequential situations. To highlight the newly developed control logic, a smart 155 mm spin-stabilized projectile equipped with a rotating paddle control mechanism is considered. For this example configuration, cross range maneuvering occasionally induces instability. Simulation results, using both rigid and multi-body nonlinear flight dynamics models, indicate that the addition of the instability protection layer in the control logic prevents projectile instability while not substantially altering target impact statistics. The nature of this protector design lends itself well to the use of a GPU to perform the calculations, greatly decreasing the computation time needed.
In this paper, modified genetic algorithm has been used as a simultaneous optimizer of recurrent neural network to improve identification and modeling of aircraft nonlinear dynamics. Weighted connections, network architecture, and learning rules are features that play important roles in the quality of neural networks training and their generalizability in order to model nonlinear systems. Therefore, the main focus of this paper is to apply appropriate evolutionary methods in order to simultaneously optimize the parameters of neural networks for the improvement identification and modeling of aircraft nonlinear dynamics. To validate this study, the results have been compared with the recorded data from a fourth generation highly maneuverable fighter aircraft flight test. Furthermore, having been compared to normal genetic algorithm, the results of the present study have showed significant improvement of the neural networks generalization which leads to better identification and modeling of aircraft nonlinear dynamics.
The attitude control of an Earth-pointing spacecraft in a circular orbit, subject to the gravity-gradient torque, is explored. The spacecraft attitude is described using the modified Rodrigues parameters. A series of controllers are designed using the nonlinear H control methodology and are subsequently generated using a Taylor series expansion to approximate solutions of the Hamilton–Jacobi equations. These controllers are applied to the problem of Earth-pointing spacecraft in circular orbits. The controllers are compared using both input–output and initial condition simulations, in an effort to gauge the improvements made possible by nonlinear feedback.
The growth rate of high-speed mixing layer between two dissimilar gases is explored through the model free simulation results. To analyse the cause for the higher mixing layer growth rate in comparison to the existing values reported in literature, the results were compared with the model free simulations of mixing of two high-speed streams of nitrogen (similar gas) at matched temperature and density. The analysis indicates that pressure and density fluctuations no longer remain correlated completely for the mixing layer formed between two dissimilar gases at different temperatures in contrast to the complete pressure density correlation for similar gases. It has been observed that the correlation between temperature and density fluctuations is near –1.0 for dissimilar gases in the mixing layer region and is much higher than for similar gases. It is concluded that mixing layer of similar gases shows a decrease in growth rate due to compressibility effect, while that of dissimilar gases shows a decrease due to dominant temperature effect on density.
The heat-flux measurement is the precondition of the structure and cooling system design for hypersonic aircraft, while no experiment method can be applied to the heat-flux measurement of the sharp leading edge for its small size and high-heat flux. In this article, an experimental method based on the law of inverse conduction is proposed in this paper for the calculation of heat flux from the inner wall temperature of the sharp leading edge. This experimental method proposed includes the method of inner wall transient temperature acquisition, smoothing technique, and the Matlab and Fluent cosimulation postprocessor, where Fluent is used as a 3D unsteady heat conduction solver, and conjugate gradient method is applied for optimization. The effectiveness of the proposed method is verified by its application in the scramjet experiment for the heat-flux measurement on the strut-leading edge, which greatly helps the assessment of thermal environment with a total error of less than 5%.
Electrically controlled rotor (ECR) system has been demonstrated in the primary control of helicopter. Without the restraint of the swashplate, ECR is also an efficient means to enhance the rotor performance by applying higher harmonic flap input. In order to investigate the potential of 2/rev flap input to enhance ECR performance, corresponding aerodynamic model of the ECR rotor and flight dynamic model of the ECR helicopter are established, in which the variation of ECR aerodynamic characteristic due to the deflection of trailing edge flap is emphasized. The analytical study is accomplished by simulating the ECR helicopter in trimmed flight for various combinations of takeoff weight, flight speed, and amplitude and phase angle of the 2/rev flap input. By studying the variation of the profile drag distribution over the rotor disc in detail, the physical essence through which 2/rev flap input affects the ECR power is explored. Subsequently, the parametric study of the blade pre-index angle affecting power reduction is conducted. Finally, the effectiveness of ECR approach for power reduction is compared with that of the conventional individual blade control approach. The simulation results show the potential of the ECR system for power reduction.
To improve operation stability of pulse detonation engine and shorten the distance of deflagration to detonation transition, a series of multi-cycle detonation experiments were investigated with six different air inlet systems. Using air as oxidizer and liquid C8H18 as fuel, the effect of different air inlet systems on pulse detonation engine with frequency of 14 Hz and equivalence ratio of 1.5 was analyzed. Pressure history along detonation tube was recorded by five dynamic piezoelectric pressure transducers. It was approved by a particle image velocimetry that centrifugal forces from rotating airflow had a significant negative impact on the uniformity of fuel distribution in detonation tube. Furthermore, the experimental results indicated that operation stability of pulse detonation engine was increased with the improvement of fuel distribution, and deflagration to detonation transition distance was obviously decreased with the increase of thrust wall sealing. In these different air inlet systems, the pulse detonation engine with air inlet system of reed valve achieved the shortest deflagration to detonation transition distance, the best stability and stable operation of about 1 min at 14 Hz. This study provided references for the development of pulse detonation engine.
For large impact angle control problem (here, the "large impact angle" means the impact angle in the closed interval from –180° to 180°), estimating the time-to-go accurately is the key of impact time and impact angle control guidance (ITIACG). The objectives of this paper are to construct a new impact angle control guidance (IACG) law suitable for large impact angle control and present a time-to-go estimation procedure for the new IACG law suitable for designing ITIACG law. The constructed IACG law is a biased proportional navigation guidance law with large impact angle constraint, the rule of the cosine of the lead angle in the biased term is to guarantee that the lead angle remains in the open interval from –90° to 90°, which is required in the development of time-to-go estimation procedure. To estimate the time-to-go, by introducing a self-convergent angle named as alfa, the closed equations of motion are transformed to a different form, which can be solved conveniently under the assumption of small lead angle. For the case of large lead angle, the time interval of time-to-go is partitioned into n segments, the maximum increment of lead angle is supposed to be a small angle in each segment, the transformed closed equations of motion can be expressed as function of alfa angle and solved analytically. A geometric approach is proposed to determine conservatively a suitable alfa angle to guarantee that the maximum increment of lead angle is a small angle in each segment. The time-to-go estimation procedure for the new IACG law are illustrated. Simulations are performed to verify the effectiveness of the proposed IACG law and the accuracy of the time-to-go estimation procedure.
Aircraft engine sensor fault diagnosis is closely related technology that assists operators in managing the health of gas turbine engine assets. As all gas turbine engines will exhibit performance changes due to usage, the on-board engine model built up initially will no longer track the engine over the course of the engine’s life, and then the model-based method for sensor fault diagnosis tends to be failure. This necessitates the study of the sensor fault diagnosis techniques due to usage over its operating life. Based on our recent results, an integrated approach based on nonlinear on-board model is developed for the gas turbine engine sensor fault diagnostics in this paper. The architecture is mainly composed of dual nonlinear engine models; one is a nonlinear real-time adaptive performance model and the other a nonlinear on-board baseline model. The extended Kalman filter estimator in the nonlinear real-time adaptive performance model is used to obtain the real-time estimates of component performance, and the nonlinear on-board baseline model with performance periodically update to provide the nominal reference in flight. The novel update strategy to sensor fault threshold based on the model errors and noise level is also presented. Important results are obtained on step fault and pulse fault behavior of the engine sensor. The proposed approach is easy to design and tune with long-term engine health degradation. Finally, experiment studies are provided to validate the benefit of the engine sensor fault diagnostics.
Given their hovering ability, static lift airships, such as airships and balloons, are proposed as stratospheric platforms flying at a high altitude of 20 km. The shape of the envelope has a major influence on the lift and drag efficiency of an airship. Furthermore, the efficiency of a conventional actuator, such as an aerodynamic control surface for stratospheric platforms, is decreased by the low-atmospheric density and flight speed. Thus, a new type of effector configuration must be proposed. A new multivectored thrust airship called flat peach is proposed in this paper. The name is attributed to the shape of the airship, which resembles a flat peach that is a cross between a ball and a water droplet. Thus, this airship has a smaller drag coefficient than the spherical airship and higher lift efficiency than a conventional airship. A control allocation strategy among the multivectored thrusters is proposed, and a composite control structure is designed for the airship to realize accurate position control and to decrease energy consumption.
According to the requirements of a high-speed heat-airflow wind tunnel experimental system, a fuel supply system based on variable frequency control technology and proportional throttle valve is designed. The mathematical model of the fuel supply system under the mode of the proportional throttle valve control and the variable frequency pump control is established. Because the fuel supply system has a pure time delay and the change of working conditions can cause the problem of time-variant parameters, a fuzzy proportion integration differentiation control strategy with Smith predictor is proposed. In addition, switching between the pump control mode and the valve control mode will bring disturbances, so two undisturbed switching methods are designed. The simulation and experimental results show that the proposed control strategy can overcome effects of the pure time delay and obtain a satisfactory control performance. The two undisturbed switching control methods designed in this paper can achieve the undisturbed switching of the fuel flow-rate control.
As a complex system, control performance of small rotary-wing unmanned aircraft is easily affected by measurement errors and environment disturbances. This paper proposes a nonlinear disturbance observer-based control to improve control performance. The constant and harmonic disturbance that is generated by the exogenous system with modeling perturbation can be estimated and rejected effectively. The random disturbance with certain bound can be reduced by the feedback control. By solving linear matrix inequality, the parameters for feedback control and nonlinear disturbance observer can be selected simultaneously. Therefore, the system stability can be guaranteed and the control performance can be improved effectively. The effectiveness of the nonlinear disturbance observer-based control is proved by a series of flight tests. Compared with feedback control, the disturbance observer-based control yields a better tracking performance in the presence of disturbances.
This paper addresses trajectory generation problem of a fixed-wing miniature air vehicle, constrained by bounded turn rate, to follow a given sequence of waypoints. An extremal path, named as -trajectory, that transitions between two consecutive waypoint segments (obtained by joining two waypoints in sequence) in a time-optimal fashion is obtained. This algorithm is also used to track the maximum portion of waypoint segments with the desired shortest distance between the trajectory and the associated waypoint. Subsequently, the proposed trajectory is compared with the existing transition trajectory in the literature to show better performance in several aspects. Another optimal path, named as loop trajectory, is developed for the purpose of tracking the waypoints as well as the entire waypoint segments. This paper also proposes algorithms to generate trajectories in the presence of steady wind to meet the same objective as that of no-wind case. Due to low computational burden and simplicity in the design procedure, these trajectory generation approaches are implementable in real time for miniature air vehicles.
An experimental study was conducted to further validate whether the newly proposed flapping rotary wing is suitable for micro air vehicle design. First, the effects of two main kinematical parameters (flapping frequency and initial angle of attack) of flapping rotary wing on lift generation were discussed. It was found that a higher lift can be generated by flapping rotary wing through increasing flapping frequency at a proper initial angle of attack. Second, effect of coupled flapping motion with rotating motion on lift generation was analyzed. It is important that a larger lift was generated by flapping rotary wing than the superposition lifts from purely flapping and purely rotating motions when the initial angle of attack was less than a critical value. Finally, the comparison of the capability of lift generation from the flapping rotary wing and conventional rotary wing was given. It was indicated that the lift from flapping rotary wing was larger than that from conventional rotary wing in the range of Reynolds number from 2600 to 5000 as long as Strouhal number was determined appropriately. The present work suggests that flapping rotary wing may be a feasible and promising wing layout used in the design of micro air vehicle in terms of lift generation.
Carrier-phase ambiguity resolution of Global Navigation Satellite System in applications that require both high accuracy and high integrity is challenging. This paper proposes an efficient partial ambiguity resolution method with integrity risk constrain for high-performance navigation. First, the Global Navigation Satellite System observation model and the integer ambiguity resolution procedure, especially the partial ambiguity resolution using least squares ambiguity decorrelation adjustment, are described. Then an integrity risk constraint method and an improved integrity risk constraint method for ambiguity resolution are presented. Based on these methods, a hybrid strategy is further derived. Last, the simulation and analysis for the integrity risk constraint method versus the hybrid method are performed. Simulation results show that the first method is conservative and thus unnecessarily limits the navigation availability, while the hybrid method presents a tight upper-bound, so that it increases the integrity and accuracy of high-performance navigation significantly.
Rotating blades on helicopters experience reverse flow under high advance ratio conditions. Here, reverse flow is characterized by the flow traveling from the sharp trailing edge to the blunt leading edge. Uncertainty in the blade aerodynamic loads under these conditions has been a limitation during the design of high-speed rotorcraft. In this work, we hypothesize that the reverse flow over a yawed blade includes phenomena similar to the formation of a leading edge vortex on sharp-edged delta wings. Low-speed wind tunnel experiments are reported on a scaled version of a rotor blade in regular and reverse flow over a large range of yaw and moderate ranges of angle of attack. Force measurements indicate a deviation from yawed-wing expectations at high yaw angles. Surface flow visualization via tufts shows the existence of an attached span-wise vortex on the wing.
Electrostatically charged spacecraft accelerates when orbiting a central body with magnetic field due to the induced Lorentz force. This Lorentz force could be used as propellantless propulsion for orbital maneuvers. Such spacecraft is referred to as Lorentz spacecraft. Modeling the Earth’s magnetic field as a tilted magnetic dipole rotating with the Earth, this paper first presents the analytical expressions that characterize the orbital motion of Lorentz spacecraft with respect to inclined low Earth orbit. Using the information from line-of-sight observations and gyro measurements, coupled with the proposed dynamical model, both extended and unscented Kalman filter are designed to perform relative navigation for Lorentz spacecraft. Two scenarios are simulated to illustrate the accuracy of derived analytical solutions and the performance of proposed filters, respectively. Through comparison with previous work, the accuracy of relative motion model has proved to be greatly enhanced. Numerical simulation results also show that unscented Kalman filter presents more accurate relative state estimation for Lorentz spacecraft than extended Kalman filter.
The aim of this study is to present a fully automated computational fluid dynamics-based optimization chain, implementing a radial basis function meta-model combined with an improved Latin hypercube design of experiments strategy. The objective function (aerodynamic performance) is evaluated through computational fluid dynamics calculations by using the commercial code ANSYS-CFX. The optimization strategy is hybridization between a stochastic bi-objective non-dominated sorting genetic algorithm and a gradient-based method known as modified method of feasible direction to get benefit from their combined capabilities. The testing of this optimization chain consisted in finding the optimal operating conditions of an airfoil NACA0012. This methodology may help to a great extent in the better exploration of the design space and to guide numerical and experimental studies to the potentially optimal design parameters.
A dynamic model of a twin ducted-fan vertical takeoff and landing aircraft, the Martin Jetpack, has been developed to study and improve the understanding of the flight mechanics involved with this novel aircraft concept. This article describes the flight mechanics of a twin ducted-fan aircraft and explains in detail the modeling of the forces and moments contributed by the twin ducted-fans, body aerodynamics, control surfaces, gyration, and landing gear interactions. Also, a novel model for the movement of the duct center of pressure has been developed, which allows for the complex duct pitching moment to be predicted. Employing the conventional aircraft modeling methodology, a system of ordinary differential equations that describes the behavior of the aircraft is developed. The equations are solved in real-time using MATLAB–Simulink software to simulate the response to given inputs. A comparison of the flight data with both steady-state (trimmed) and dynamic simulations shows good agreement, which validates the novel duct center of pressure model. The validated model allows the aircraft designer/engineer to efficiently evaluate the sizing of key aerodynamic features and various control methodologies to aid in the design and flying of the Martin Jetpack.
In this study, we develop a coarse-to-fine particle filter algorithm for track-before-detect in order to track a subpixel-sized, low signal-to-noise ratio target in sensor data. The proposed algorithm enhances tracking performance in the presence of target motion uncertainty and it also maintains the computational load without increasing the number of particles. This coarse-to-fine particle filter, which is newly applied to track-before-detect, has two recursive stages: a coarse stage for extensive searches of the target’s state space and a fine stage that narrows down the tracking results. During the coarse stage, particles are propagated with uniformly distributed noise to compensate for highly nonlinear target motion. The fine stage disturbs the particles filtered from the coarse stage using Gaussian distributed noise. Monte Carlo simulation results using artificial image sequences indicate improved performance with the proposed algorithm when uncertain large frame-to-frame pixel differences are caused by nonlinear target motions such as jittering effects. The algorithm is also applied to the real camera image frames to verify its detecting performance.
The focus of this article is on the numerical simulation of compressible flow in a diffusing S-duct inlet; this flow is characterized by secondary flow as well as regions of boundary layer separation. The S-duct geometry produces streamline curvature and an adverse pressure gradient resulting in these flow characteristics. The geometry used in this investigation is based on a NASA Glenn Research Center experimental diffusing S-duct that was studied in the early 1990s. The computational fluid dynamics flow solver ANSYS - FLUENT is employed in the investigation of compressible flow through the S-duct. A second-order accurate, steady, density-based solver is employed in a finite-volume framework. The three-dimensional Reynolds-Averaged Navier-Stokes equations are solved on a structured mesh with a number of turbulence models, namely the Spalart–Allmaras (SA), k-, k- SST, and Transition SST models, and the results are compared with the experimental data. The computed results capture the flow field and pressure recovery with acceptable accuracy when compared with the experimental data. The turbulence model giving the best results is identified.
Electromagnetic formation flying is a novel concept of controlling the relative degrees of freedom of a satellite formation without the expenditure of fuel by using high-temperature superconducting wires to create magnetic dipoles. Micro-electromagnetic formation flying, which is an alternative to electromagnetic formation flying in terms of reduced complexity, uses conventional conductors to replace the high-temperature superconducting coils in electromagnetic formation flying, shortening the separation distances between the electromagnets. This paper investigates the use of micro-electromagnetic formation flying for providing relative position control for unperturbed station-keeping in a multi-satellite array along the cross-track direction that can be used in cross-track interferometric synthetic aperture radar applications. Considering that conventional conductors produce small separation distances between electromagnets, comparatively large baselines can be achieved by positioning multiple satellites consecutively in an array. The existence of equilibrium positions of the satellites is demonstrated. The station-keeping efficiency of the formation satellites is studied. It is found that the electromagnetic dipoles on neighboring satellites should be equal in magnitude and opposite in direction to obtain the maximum station-keeping efficiency of the formation; correspondingly, the equilibrium positions of the satellites along the cross-track direction are symmetrical about the center of mass of the formation. A method for maximizing the station-keeping efficiency of the formation using micro-electromagnetic formation flying is also presented, using feasible designs for small satellite formations as examples.
This article derives an improved robust Huber-based divided difference filter by using the Huber’s technique, in which the nonlinear measurement function is directly used in the nonlinear regression equation instead of the linear or statistical approximation. The presented filtering algorithm exhibits robustness against the deviations from the Gaussian error distribution and has better estimate accuracy compared with the Huber-based divided difference filter. This filter is applied to a benchmark problem of estimating the trajectory of an entry body from discrete-time range data measured by a radar tracking station. Simulation results indicate that the proposed filter algorithm outperforms the previous methods in terms of robustness and accuracy.
Loading an aerospace and automotive seat statically through lap or body blocks is a complex and highly non-linear problem, as the key numerical challenge is to replicate the contact and slipping kinematics between seat, lap block and belt. In addition, severe element distortions and unexpected contact between parts can occur due to the large deformations involved, which result in implicit solvers struggling to find a converged solution. This paper focuses on the use of an explicit Finite Element Analysis (FEA) solver (LS-DYNA3D) for an aircraft seat subject to Certification Specifications CS25.561, although the ideas presented are equally applicable to automotive seat designers. Explicit codes are better able to overcome contact convergence issues and are often used with appropriate damping to achieve a quasi-static solution. This paper reviews the methodology presented in Part I, whereby issues relating to damping, mass and time scaling are outlined in order to overcome the high computational time step costs (Courant-Friedrichs-Lewy (CFL) condition), together with the procedural and error checks required to ensure a quasi-static response. This paper extends the methodology by considering load cases that use lap blocks, such as ‘forward 9g’ and ‘upward 3g’ certification requirements. Alternative modelling approaches to represent the loading mechanism and effect of lap block mass on solution accuracy are discussed. This paper concludes with a verification framework that outlines the quality checks on various model energies and their ratios, where the numerical results are validated against test in terms of displacements and seat kinematics. Thus, ‘Part I’ and ‘Part II’ cover all elements related with the application of an explicit dynamic integration scheme to demonstrate static seat compliance, and together, form a clear framework to assist a Computer Aided Engineering (CAE) analyst involved in applying an explicit integration scheme to solve non-linear quasi-static analyses.
This paper proposes a purely aerodynamic pitch control design scheme for agile missiles that need to reverse their flight directions for tracking a target. The main challenge in performing this kind of manoeuvre is that the missiles may enter the high angle of attack domain in which aerodynamic actuation is ineffective. Assuming that reliable aerodynamic data are available for the flight envelope, a systematic search algorithm is employed to find an aerodynamic control parameter and the motor thrust activation time, so as to enable the missile’s required manoeuvre. The strength of the proposed design scheme is that no third control mechanism such as thrust vectoring is required for the missile control even in the high angle of attack domain. The proposed design scheme is tested on aerodynamic data available in literature, and is shown to be promising via simulations.
In this article a L1 adaptive state feedback controller is presented for the three-dimensional integrated guidance and control of interceptor. The model of three-dimensional integrated guidance and control is first built by using engagement kinematics and interceptor dynamics. The objective is to control the fin deflection angles to make the line-of-sight rates converge to zero. The L1 adaptive control is then applied to design the three-dimensional integrated guidance and control controller. The L1 adaptive controller guarantees uniformly asymptotic and bound transient tracking for the system inputs and outputs, which are important for interceptions with strong robustness and high timeliness. The design issues of L1 adaptive controller are also discussed. In the interception simulation, the L1 adaptive state feedback controller demonstrates the robustness to different kinds of uncertainties, while guaranteeing the transient performance of dynamic interception process.
The force equalization of a hybrid actuation system combining one servo-hydraulic actuator and one electro-mechanical actuator operated in position control and in active/active mode is addressed for safety critical applications such as primary flight controls. In a first step, an accurate virtual test bench is built to facilitate the analysis of force fighting and the assessment of the performance and robustness of the proposed force equalization strategies. It is validated from real experiments performed for the aileron actuator of a single-aisle commercial aircraft. Static force equalization is achieved first by adding equalization offsets in the position control loops as a function of the integral of the force difference between actuators. In order to keep a high level of segregation, the authority for this action is limited to 4% of the total actuator stroke. The dynamic force equalization is performed by forcing the two actuators to follow the same path. Thus, a trajectory generator is introduced to output the required position, velocity and acceleration from the position set point with realistic reproduction of the actuator power limits. Feedforward actions are used to compensate the major and invariant effects such as servo-hydraulic actuators functional flow and electro-mechanical actuator inertial torque. In this way, the pursuit errors are significantly reduced without decreasing robustness. Then, the accurate virtual test bench is used to assess the robustness of the force equalization strategy by analyzing the sensitivity of performance indicators to parameters and operating conditions. It is shown that the proposed force equalization scheme meets all the requirements, including segregation, robustness and simplicity.
This paper presents a vibration control strategy for a flexible manipulator with a collocated piezoelectric sensor/actuator pair. A hybrid vibration controller is proposed by combining the input shaping technique with auto disturbance rejection controller. The parameters of the closed-loop system can be adjusted to the known values by disturbance compensation and linear feedback using the auto disturbance rejection controller. This way, input shaper can be designed without accurate parameters of the flexible manipulator. Both simulation and experiments are conducted to validate the proposed control algorithm. The results verified the effectiveness of the proposed controller in vibration suppression of flexible manipulator.
A finite element model of an aircraft seat subjected to static certification loads (Certification Specifications CS25.561) involves material, geometric and contact non-linearities. Implicit algorithms can model the physics of such problems appropriately but suffer from shortcomings such as significant finite element modelling efforts, high disk space and memory requirements and unconverged solutions. Explicit finite element schemes offer a more robust alternative for convergence for quasi-static loadcases but may come at an even higher computational cost as smaller solution time steps are required, in addition to unwanted inertial effects. A methodology to apply an explicit formulation for simulating static certification loading for an aircraft seat-structure is presented and validated in this article. The first part reviews the design novelties of the triple seat-structure considered, the safety regulations used in aircraft seat certification. The key theoretical aspects of an explicit solver are presented, together with the numerical challenges faced when applied to solving quasi-static problems. Time scaling, mass scaling and damping are common approaches to assist in artificially reducing the computational time but previous articles provide little insight into how to apply these techniques correctly and the level of checking that is required to ensure the quality of the results are unaffected by these modifications. The main focus of this article is to clearly define the procedure to establish appropriate factors for mass scaling, time scaling and damping. Quality checks, such as ratio of kinetic energy to internal energy and their time-histories have been investigated to ensure a quasi-static solution. finite element analysis results are validated against experimental testing for the 8.6 g downward loadcase. Parameters such as kinematic behaviour and deflections at key locations been used for comparison. An acceptable level of correlation between finite element analysis results and physical tests validates the proposed methodology, which will be extended in a future article (Part II) to consider additional contact complexities with the inclusion of body blocks.
The modified quantum-behaved particle swarm optimization algorithm is developed. It has the ability to learn from excellent individuals and precisely update all the particles that are involved in computational fluid dynamics computation. The airfoil parameterization method of the Hicks–Henne form function was also improved. The Reynolds averaged Navier–Stokes equation solver and the multi-objective and nonlinear adaptive value weighting method were used to optimize a transonic and high-aspect-ratio swept-back wing and winglet. The optimization results show that the drag characteristics of the optimized configuration are reduced greatly, the shock-wave amplitude on the wing is reduced, and intense shock wave on the winglet is completely eliminated, thus indicating that this method has strong engineering practicality.
The acceleration loads of projectile during launch process of two-stage light gas gun were studied by the developed computational fluid dynamics program. With the usage of LS-DYNA software, the diaphragm rupture pressure was calculated by finite element method. The influence of different waves rupture diaphragm on the maximum acceleration loads of projectile was analyzed, keeping the configurations of gun unchanged. It is found that the maximum acceleration loads can be reduced and the muzzle velocity objective can be achieved by choosing the ruptured wave appropriately and optimizing other operational parameters. Soft launch capability is provided for launching complex lifting configuration models up to hypervelocity.
Finding the location of feature points in 3D space from 2D vision data in structured environments has been done successfully for years and has been applied effectively on industrial robots. Miniature flying robots flying in unknown environments have stringent weight, space, and security constraints. For such vehicles, it has been attempted here to reduce the number of vision sensors to a single camera. At first, feature points are detected in the image using Harris corner detector, the measurements of which are then statistically corresponded across various images, using knowledge of vehicle’s pose from onboard inertial measurement unit. First approach attempted is that of ego-motion perpendicular to camera axis and acceptable results for 3D feature point locations have been achieved. Next, except for a small region around the focus of expansion, forward translations along the camera axis have also been attempted with acceptable results, which is an improvement to the previous relevant work. The 3D location map of feature points thus obtained is utilizable for trajectory planning while ensuring collision avoidance through 3D space. Reduction of vision sensors to a single camera while utilizing minimum ego-motion space for 3D feature point location is a significant contribution of this work.
In this paper, a guidance scheme is developed for tracking constrained entry trajectory which is updated onboard. From an initial offline trajectory, the guidance system updates trajectories at every step of control command generation. This scheme models state error dynamics as a linear time varying system and updates the trajectory using a pseudospectral method. The solution provides an updated and optimal trajectory from the present position to the terminal state satisfying the path constraints. The guidance system continues updating the online trajectories and generates the control commands. This method is different from tracking methods purely based on linear quadratic regulator theory because it utilizes the pseudospectral method in generating control command; but it is also different from pseudospectral guidance because generation of new reference trajectories is done onboard. The method is validated through a number of test cases for initial state perturbations, aerodynamics and atmosphere modeling errors. In order to demonstrate improved accuracy relative to other methods, the method is also tested against linear quadratic regulator and pseudospectral guidance schemes.
Shape optimization of transonic airfoils requires creating an airfoil that reduces the drag due to transonic shocks by either eliminating them or reducing their strength at a given transonic cruise speed while maintaining the lift. The RAE 2822 and NACA 0012 airfoils are most widely used test cases for validation of computational modeling in transonic flow. This study employs a multi-objective genetic algorithm for shape optimization of RAE 2822 and NACA 0012 airfoils to achieve two objectives, namely eliminating shock and maintaining or increasing the lift at a given transonic Mach number and angle of attack. The commercially available software FLUENT is employed for calculation of the flow field using the Reynolds-averaged Navier–Stokes equations in conjunction with a two-equation turbulence model. It is shown that the multi-objective genetic algorithm can generate superior airfoils compared with the original airfoils by achieving both the objectives.
To get a better understanding on the output uncertainty contributed by an individual variable as well as the correlated variables of models with dependent inputs, a method for decomposing Sobol’s first-order effect indices into uncorrelated variations and correlated variations is investigated. Instead of using Monte Carlo simulation or full tensor product-based numerical integration approaches, a new sparse grid numerical integration method is proposed for estimating Sobol’s main effect indices as well as the two decomposed sensitivity measures. Before conducting the sparse grid numerical integration-based algorithm, an orthogonal transformation is used to transform the dependent input variables and model performance function into independent space as the joint probability density function of the correlated variables cannot be written as the product of univariate density functions. An obvious advantage of the sparse grid numerical integration-based method is that it can decrease the computational cost of the conventional methods significantly while keeping the accuracy level controllable, particularly for high-dimensional problems. The proposed approach is compared with other alternative approaches through theoretical and applied numerical experiments to demonstrate its efficiency, accuracy and high-dimensional adaptivity.
A theoretical framework of nonlinear flight control is exploited and applied to nonlinear longitudinal dynamics of a generic air-breathing hypersonic flight vehicle. A combination of novelty command filtered back-stepping technology and dynamic inversion methodology is adopted for designing a dynamic state-feedback controller that provides stable tracking of the altitude and velocity reference commands. The novel command filtered back-stepping altitude control obviates the need to compute analytic derivatives in the traditional back-stepping design, providing a simple and effective way for controlling non-linear hypersonic flight vehicle. An input-to-state stability-modular approach is presented by combining command filtered back-stepping method with sliding-mode-based integral filters, input-to-state stability analysis, and small-gain theorem. The stability analysis of the closed-loop system including the flexible dynamic, and the convergence of the system outputs are derived. The proposed control scheme is verified in simulations in a climbing maneuver case of separate velocity and altitude reference commands.
The SAFE Structural Analysis procedure is an idealisation error control methodology devised for linear static finite element analysis. This study examines the applicability of this process to non-linear problems. The studied case is the collapse analysis of an aircraft stiffened panel loaded in compression. This article presents the critical investigation of important modelling assumptions, including the joint modelling, boundary conditions, geometrical imperfections and scattering in material parameters. Potential error sources are identified and then analysed using the non-linear finite element solver ABAQUS. The analysis derived an improved finite element model and concrete idealisation error estimates. The finally simulated failure behaviour corresponds well to the data measured in the test.
To analyze the reliability of an airship’s envelope, the methodology of structural reliability analysis is adopted. The basic theory and the detailed steps of the algorithm of the first-order reliability method are discussed. For finding multiple design points, the method of adding bulge to the limit-state function is applied. With regard to the problem of envelope’s reliability, the safety criterion and limit state function of the airship’s envelope are analyzed. The mathematical model of the envelope’s maximum stress is also presented. The reliability simulation of a stratospheric airship’s envelope is taken as an example. Results of sensitivity analyses of the envelope are also obtained.
Powerful actuators are both indispensable and critical components for the flight control system of hypersonic vehicles. More importantly, the performance of actuators has substantial impact on the control ability; therefore, when designing the control system, one needs to fully take into account actuator restraints in order to meet the efficient and precise control demands under complex flight conditions. In this article, the advanced flight control methods concerned with actuator limitations are discussed for hypersonic vehicles. First, the longitudinal model of hypersonic vehicle is established with consideration of the nonlinear coupling dynamics. Second, the actuator constraints with the effect of elastic deformation are introduced to this built model and then the resulting unstable dynamics characteristic and the control limitation conditions are analyzed for hypersonic vehicle. Furthermore, the advanced flight control laws are designed by using the differential geometry principle and the total energy theory. Finally, simulation results verify the feasibility of the proposed methods for hypersonic vehicle.
This paper presents a methodology for the optimal preliminary design of electro-mechanical actuators. The main design drivers, design parameters and degrees of freedom that can be used for preliminary design and optimization of electro mechanical actuator are described. The different types of models used for model-based design (estimation, simulation, evaluation and meta-model), and their associations are presented. The process preferred for its effectiveness in terms of flexibility, and computational time is then described and illustrated with the example of a spoiler electromechanical actuator. The proposed approach, based on meta-models obtained using the surfaces response methods and scaling laws models, is used to explore the influence of anchorage points and transmission ratio on the different design constraints and the overall mass of the actuator.
Wake vortices are an issue affecting both capacity and safety of air traffic and therefore need to be dealt with by appropriate measures and procedures. Today, the only means to prevent wake vortex encounters is procedural separation which however is statical and in many cases conservative. The concept of dynamic separations using wake vortex predictions aims at optimising the separation between consecutive aircraft based on the knowledge of the actual position and strength of the wake vortices. A concept for approach procedures has been developed that involves dynamical calculation of minimum safe distance, adaption of follower aircraft speed and the corresponding approach types. The concept, its implementation and simulation test results will be presented and it will be discussed how it can be applied to contribute to an optimised use of available capacity while maintaining and improving the safety level.
The effects of using porous aluminum particles in solid propellants were studied, with emphasis on the agglomeration phenomena. Burning strands containing either regular (as-received) or porous aluminum were photographed by a high-speed camera, and particulate combustion products were analyzed in a laser particle analyzer. Results obtained from experiments conducted in a pressure-range of 1–34 atmospheres show that porous aluminum particles produce smaller agglomerates than regular aluminum. The median diameter of agglomerates resulting from porous aluminum reached, on average, 70% of the one originating from regular aluminum. This reduction in agglomerate diameter corresponds to a substantial volume (and hence, mass) decrease of approximately 65%. It is assumed that the high-specific area of the porous aluminum particles (10–18 m2/g, similar to that of nano-Al) results in high reactivity, leading to shorter ignition time and hence to the formation of smaller agglomerates.
To eliminate the effect of the uncertain disturbances and improve the control accuracy of spacecraft Attitude Control System, a nonlinear control algorithm named nonsingular terminal sliding-mode feedback controller is proposed in this work, which is mainly made up of nonsingular terminal sliding-mode controller and sliding-mode feedback controller. In the first place, nonsingular terminal sliding-mode controller is designed, which guarantees global asymptotic convergence of the attitude in the presence of the uncertain perturbations from the space. Despite that, it is the influence of the uncertain disturbances that hinder the control accuracy. Then, in order to promote the control accuracy, the sliding-mode feedback controller based on the principle of minimum sliding-mode error is proposed, which is used to compensate the control errors of the nonsingular terminal sliding-mode controller caused by the uncertainties. Hence, the determination principle of the weighting matrix in sliding-mode feedback controller is discussed, and the algorithm structure of the sliding-mode feedback controller is also analyzed, which provides the theoretical basis for the sliding-mode feedback controller. By contrast, an adaptive fuzzy algorithm is designed and introduced into the nonsingular terminal sliding-mode controller to improve the control accuracy, which named the nonsingular terminal fuzzy sliding-mode controller. Last but not the least, several numerical examples are presented to demonstrate the efficacy of the proposed nonsingular terminal sliding-mode feedback controller. Simulation results confirm that the control accuracy of the nonsingular terminal sliding-mode feedback controller is higher than the nonsingular terminal sliding-mode controller and the same as nonsingular terminal fuzzy sliding-mode controller. Not only is the calculation of the nonsingular terminal fuzzy sliding-mode feedback controller smaller than nonsingular terminal fuzzy sliding-mode controller, the adjusted parameters are also fewer than nonsingular terminal fuzzy sliding-mode controller obviously. The numerical results clearly indicate that the proposed nonsingular terminal sliding-mode feedback controller based on the principle of minimum sliding-mode error can compensate control errors accurately and quickly; therefore, it can reduce the effect of the uncertainties from the space indirectly.
There is a greatly persistent wind shear in the upper atmosphere, especially at the altitude of 10–20 km. For the idea of dynamic soaring, the wind shear can be treated as a kind of energy resource for aircraft if the aircraft is flying in a proper manner. Based on the above facts, the influence of wind shear to the performance of high-altitude solar-powered aircraft from a new prospect is systemically studied: The wind shear in the upper atmosphere is treated as a kind of energy resource for aircraft, and to be used to compensate the energy consumed by drag. The results of simulations show that the energy extracted from wind shear can compensate about 30–50% of the energy consumed by drag in climbing and 20–40% in descending for high-altitude aircraft when the strength of wind shear is greater than 0.005 s–1 and smaller than 0.01 s–1. This is a valuable conclusion for the high-altitude aircraft, since the strength of wind shear between 10 km and 20 km has fallen into this interval. By defining the dynamic soaring parameter, it has been found that the dynamic soaring parameter is possibly greater than 1 in the place that great enough strength of wind shear can be found, which implies that it is possible for high-altitude aircraft to perform unpowered flight by dynamic soaring if the wind shear can be unitized properly.
Boundary layer ingestion by fans in propulsion system improves the propulsive efficiency. However, inlet flow distortion will dramatically eliminate these benefits. This paper puts forward a method to deal with inlet flow distortions and examines their impacts on turbofan performance at engine design point. The method models both radial and circumferential distortion and their impacts separately. Firstly, a distorted fan map is calculated by parallel compressor method. Then, the new map is utilised to find the fan exit flow conditions by parallel stream method. Finally, we assume that all the flows mixed well before entering the nozzle without any pressure losses. At all examined fan pressure ratios, boundary layer ingesting improved fuel consumption. However, the benefits reduced by the new method are lower than previous predictions without considering intake distortion. If the fan pressure loss and efficiency drop due to inlet distortion are too high, boundary layer ingestion should not be used with a traditional fan design. Large boundary layer ingestion for future propulsion system should consider new fan blade design.
To test different types of noise abatement approach procedures the Institute of Flight Guidance and the Institute of Aerodynamics and Flow Technology performed flight tests on 6 September 2010 with a Boeing 737-700. In total, 13 approaches to the research airport in Brunswick, Germany (EDVE) were flown while the approach area of the airport was equipped with six noise measurement microphones. Brunswick airport is equipped with an experimental ground based augmentation system which allows the implementation of 49 instrument landing system (ILS) look-alike precision approach procedures with different approach angles simultaneously.
Airplane stability theory was born at the end of the XIX century and matured around 100 years ago, when airplanes were hardly controllable yet. The success and safety of flights in the pioneer years depended upon largely unknown stability and control characteristics. Understanding the modes of airplane motion has been of paramount importance for the development of aviation. The contributions made by a few scientists in the decades preceding and following the first flight by the Wright brothers set the concepts and equations that, with minor notation aspects, have remained almost unchanged till present day.
Approximate solutions for vibrations of flexible beam-type appendages subjected to tip mass are studied while uniform and exponential profiles for arm deployment are simulated. Applying an equivalent dynamical system and following Lagrangian approach, the equations of motion of the system are derived as nonlinear ordinary differential equations (ODEs) (with time-varying coefficients), in which the effect of the tip mass can be considered as some nonlinearity added to the ‘no tip mass’ case dynamics. The approximate closed-form solutions are obtained through a novel methodology using a computer algorithm, in which the solutions of the ‘no tip mass’ case are expanded by imposing quadratic perturbations on the independent variable. The mean square of errors (MSEs) for the obtained approximate analytical solution is computed. Using this method, the amplitude and frequency of the arm response are presented by the algebraic equations, which help the parametric design of such systems. In addition, effects of tip mass as an indicator of nonlinearities added to the system dynamics, on the amplitude and frequency of the beam response, are investigated during arm deployment.
This paper presents robust fault detection based on adaptive thresholds for a three axis satellite. For this purpose, first, the attitude control system (ACS) is described as a quasi linear parameter model that includes both bounded parametric modeling errors and measurement noises. Next, using the interval arithmetic tools, an interval linear parametric varying observer is designed to propagate the effect of satellite parametric uncertainties into the alarm limits. This idea enhances the robustness of fault detection system at the decision making stage. In other words, the adaptive thresholds are generated for evaluating the residuals. Obtained results show that the missing alarm rates are minimized by the developed method; also this approach detects small or incipient faults more effectively than the classical robust fault detection algorithms with constant thresholds. In the next part of paper, an isolation algorithm has been proposed using the fault tree approach. Also, an accommodation system has been designed based on reconfiguration of available actuators. Accordingly, after isolation of faulty reaction wheels using the developed fault tree library, the accommodation system turned them off and replaced the suitable magnetic tourqers instead of faulty reaction wheels. Therefore, despite occurrences of several failures in the ACS, attitude control error is kept limited.
A tight-coupling heat transfer method is proposed in this paper for the analysis of the performance of the hot air ice protection system. The Eulerian method is used for the calculation of the local collection efficiency. The external heat transfer coefficient was computed using the boundary layer integrated method. The thermal conductivity within the wing skin and the internal heat transfer between the hot air flow and the skin were computed using the computational fluid dynamic method. At the same time, the external heat flux boundary condition and the surface temperature were updated automatically by user-defined function to drive the iteration of the tight-coupling heat transfer calculation. The surface temperature results in dry air condition are compared with flight test data and show agreement. The maximum temperature difference between the simulation and the test is 11.5 K. In addition, the method proposed in this paper is applied to the wing hot air ice protection system of a real aircraft in icing conditions. It is found that the surface temperature ranges from 3 to 30 °C under certain flight and icing conditions. Larger droplet diameter or larger liquid water content leads to more runback water which changes the surface temperature not much when the parameters of the bleeding air are the same.
The clearance flow of compressor variable stator vanes has a significant impact on compressor performance. Most clearance flow experimental research work has been carried out in compressor cascades with lower turning angle (lower load). The clearance flow in a higher load compressor cascade is relatively more complicated due to three-dimensional (3D) flow separation. In the present work, the effects of different heights and locations of clearances in a compressor cascade on aerodynamic performance and separation flow were investigated experimentally in a low-speed plane cascade wind tunnel. The objective of this investigation is to study the characteristics of different clearance configurations depending on the clearance height and the clearance axial position, and make an attempt to verify the possibility of a local clearance flow that improves the performance. The cascade outlet section aerodynamic parameters were measured by "L" five-hole probe. The ink-trace flow visualizations on the cascade surface and lower end wall were performed. The results show that with the increase in clearance size, the core area of high loss is moving to the end wall and the morphology of limiting streamlines on the suction surface changes gradually from separation lines to reattachment lines. The separation lines on end wall gradually moved to the middle of the cascade and the separation area increased gradually. The above reasons caused high loss zone gradually extended to the end wall. Thus, the loss at mid-span declined slightly while the loss on the end wall was increasing. Otherwise, cascade load reduced significantly when the clearance height became 3 mm, and the change magnitude of limiting streamline morphology in the flow field which was influenced by clearance flow at the rear of the blade reached to a minimum. So the loss and load came to a minimum, respectively. The impact of the clearance in the rear of a blade on loss and flow is weaker than the other schemes, with a slightly larger loss and separation with respect to the original scheme without a clearance. It may be expected to improve the cascade performance further by an appropriate design of clearance near the trailing.
This paper describes the development of a multi-phase/multi-criteria trajectory optimization framework that has been conceived to support the synthesis of green mission profiles that will allow aircraft to fly optimum flight paths with the lowest possible noise and emissions. The proposed multi-phase/multi-criteria framework is not only suitable to formulate trajectory optimization problems in which noise, emissions, or global warming effects can be simultaneously considered, but also provides the possibility to implement air traffic management constraints that apply to certain flight stages. A case study involving a trip from Amsterdam Airport Schiphol to Munich Franz Josef Strauss International Airport is presented to illustrate the synthesis of green trajectories and to demonstrate the potential for improving the environmental footprint. The optimization results bear out that optimizing with respect to noise can be very rewarding in terms of reducing the local noise impact, without significantly affecting the overall flight-economic performance.
Structural mass optimisation of an aircraft design is important in increasing the likelihood that a high quality airframe is designed of minimal weight whilst providing necessary resistance to load. Analysis of such structures is often performed at a single level of model fidelity, the selection of which can lead to either excessive computational time or reduced accuracy of results. Alternatively, variable-fidelity modelling may be employed to reduce such computational expense whilst maintaining accuracy, traditionally performed using predetermined levels of fidelity for specific periods of the optimisation process. This paper investigates dynamically controlled variable-fidelity modelling during aircraft conceptual design optimisation wherein fidelity is controlled as a dynamic parameter of the optimisation process. Consequently, model fidelity is adapted during optimisation to promote early discovery of promising design characteristics prior to detailed analysis of the best designs available. Models are constructed through the grouping of similar structural members within elements, thus reducing the number of degrees of freedom and subsequent computational effort required for analysis of each design. A case study is performed to verify the results of analysis and obtain benchmark results for optimisation with static model fidelity prior to the investigation of various set-ups of dynamically controlled variable-fidelity modelling. The results of this study indicate improved design quality using dynamically controlled variable-fidelity modelling compared to using static model fidelity whilst reducing the necessary computation time.
In this paper, a fractional order proportional-integral controller is developed for a miniature air vehicle for rectilinear path following and trajectory tracking. The controller is implemented by constructing a vector field surrounding the path to be followed, which is then used to generate course commands for the miniature air vehicle. The fractional order proportional-integral controller is simulated using the fundamentals of fractional calculus, and the results for this controller are compared with those obtained for a proportional controller and a proportional integral controller. In order to analyze the performance of the controllers, four performance metrics, namely (maximum) overshoot, control effort, settling time and integral of the timed absolute error cost, have been selected. A comparison of the nominal as well as the robust performances of these controllers indicates that the fractional order proportional-integral controller exhibits the best performance in terms of ITAE while showing comparable performances in all other aspects.
In this article, a simplified divided difference filter based on the model structure with linear output equations and the assumption of additive Gaussian noise is introduced. By making use of the Huber technique to modify the measurement update equations of the simplified divided difference filter, the new filter exhibits robustness with respect to deviations from the common assumption of Gaussian distributed random measurement errors, for which the simplified divided difference filter exhibits mild degradation in estimation accuracy. In addition, in contrast to standard extended Kalman filter, more accurate estimation and fast convergence are achieved from the poor initial conditions. The proposed Huber-based simplified divided difference filter algorithm has been tested in relative navigation using global position system for spacecraft formation flying in low Earth orbits with real orbit perturbations and non-Gaussian random measurement errors in flight simulations. Simulations results indicate that the proposed filter provides better performance in relative navigation accuracy and robustness when compared to extended Kalman filter and simplified divided difference filter in the presence of non-Gaussian measurement noise.
Composite airframes suffer from complex damage modes during operation. Many investigations tend to look into specific aspects of damage mechanisms but seldom take a systematic view. This paper introduces a new fault tree methodology to synthesize various damage modes of composite structures by identifying possible damage causes. Qualitative analysis is performed incorporating structure importance analysis, probability importance analysis and relative probability importance analysis. Quantitative analysis by Monte Carlo simulation is then conducted as a validation to demonstrate the feasibility of the fault tree for composite damages. A number of options addressing main damage causes are proposed to improve the reliability of composite structures. Engineers from airlines and manufacturers can use this method to prioritize the main damage causes in different situations as a failure preventative tool or damage evaluation. Also, this approach can be extended to provide valuable inputs to other advanced methodologies to perform better diagnosis and prognosis for composites.
In this paper, the adaptive simplified spherical simplex unscented Kalman filter was proposed to calculate angular velocity in gyro-free strapdown inertial navigation system. Firstly, a general angular velocity calculation modeling method with time-varying process noise was proposed, which was not limited to a certain kind of accelerometer configuration. Then aiming at the issues of large amount of calculation of unscented Kalman filter and the time variation of the process noise, and based on the characteristics of additive noise and linear state equation, the adaptive simplified spherical simplex unscented Kalman filter was proposed to estimate the angular velocity. The sampling points were decreased in this method through adopting the spherical simplex sampling strategy and not augmenting the state, thus improving the calculation efficiency. Meanwhile, Sage–Husa suboptimal maximum a posteriori noise estimator was brought in to estimate the process noise in real time in order to settle the problem of filter divergence induced by the time variation. Lastly, the proposed algorithm was simulated and also contrasted with the integration method, the evolution method and the conventional adaptive UKF algorithm. The simulation results indicated that the adaptive simplified spherical simplex unscented Kalman filter algorithm has higher precision than the integration method and evolution method and has higher efficiency than the AUKF, which could effectively improve the calculation precision and meanwhile guarantee the calculation efficiency.
In the presence of plant uncertainties, utilizing an appropriate controller for a smooth output tracking and elimination of high-frequency disturbances, especially in accurate systems is very important. In this paper, a controller is proposed based on the robust and optimal theory to achieve a combination of such characteristics in the face of model parameter variations and unknown disturbances. The proposed controller has been simulated on a three-axis gyro-stabilized MIMO platform and comparison results with a NLPID controller simulation are provided.
The aerodynamic performance of a 2D lumped flexible airfoil in forward flight is studied by solving the incompressible N-S equations coupled with a structural dynamics equation for the motion of the airfoil. A lumped torsional NACA0012 airfoil in forward flight is employed and the flow field, the aerodynamic force, the energy efficiency, and the lumped torsional positions of the airfoil with different flexibilities are investigated. It is found that the flexibility influences the aerodynamic characteristics of the airfoil greatly and if the airfoil has an appropriate flexibility the flexibility can increase the thrust force and the energy efficiency while the mean lift force is almost unchanged. The results also show that a light airfoil possesses a larger propulsive efficiency than a heavy airfoil, and if the lumped torsional position is close enough to the leading edge of the airfoil the shedding of leading edge vortex can be delayed, which results in a greater thrust force.
A dynamic control allocation approach is presented to address the attitude stabilization problem of a rigid spacecraft. The approach is developed by using a least-square support vector machine. Actuator uncertainty including misalignment and magnitude deviation is explicitly addressed. A dynamic inverse control law is firstly designed. A least-square support vector machine-based adaptive compensator is then designed to handle actuators uncertainties, external disturbances and unknown moment of inertia. Lyapunov stability analysis shows that the closed-loop attitude system is asymptotically stable. More specifically, constrained quadratic programming-based robust dynamic control allocation is implemented to manage the redundancy actuators. The goal of minimizing the assumption of total energy is achieved. A numerical example is provided to demonstrate the effectiveness of the proposed scheme.
In this article, we discuss the possibility of integrating image de-blurring techniques in an aerial simultaneous localization and mapping by a single camera (monocular simultaneous localization and mapping (Mono-SLAM)). We use an integrated aerial virtual environment together with a six-degree-of-freedom aircraft flight simulator to show the effectiveness of the approach to generate three-dimensional flight trajectories via integration of image de-blurring in the associated loop of the Mono-SLAM. The objective is to increase the overall performance of a flying mission over an unknown area by means of a vision-only method. The integrated aerial virtual environment produces and collects real-time pictures from a nadir-looking vision sensor mounted on the vehicle. Our MATLAB GUI-based toolbox helps user to investigate an offline Mono-SLAM with a predefined de-blurring method integrated with an estimator which extracts navigational parameters. The system efficient architecture allows effective virtual experiments in a completely unknown environment, without using preloaded maps or predefined features. Simulation outcomes demonstrate the feasibility of navigation of aerial robots in inaccessible environments. Different case studies support the conclusion; nonetheless, we observe a number of nonlinearities from de-blurring filters even for a general aviation aircraft.
For the enhancement of survivability and maneuverability, modern aircraft systems have redundant control effectors. Control allocation is a useful method for distributing control signals among the individual effectors. In order to implement a control allocation scheme, the control system is designed using two-step procedures. In the first step, the control law is designed by adaptive backstepping control. The second step is to design the control allocator. A robust control allocation method is presented in this paper, which is motivated from the concept of the worst-case robust approximation approach. By assuming uncertainties in the control effectiveness matrix, the worst-case robust control allocation problem is investigated. The proposed robust control allocation technique is compared with weighted least squares control allocation. In particular, nonlinear simulations demonstrate that the proposed robust control allocation method has satisfactory performance and robustness for the assumed uncertainties in the control effectiveness matrix.
The primary objective of this work is to investigate lip-thickness effect on the various acoustic emission characteristics of Hartmann whistle. Nozzle-to-cavity distance (stand-off distance), cavity length, nozzle pressure ratio and lip thickness form the pertinent parameters of the present study. Two lip thicknesses considered in the present study are a thick-lipped cavity (5 mm) and a thin-lipped cavity (1 mm). Although lip thickness has negligible effect on the resonance frequency, it has significant influence over the sound pressure levels generated. The results showed that thin-lipped-Hartmann whistles could emit up to 2.4 times the acoustic power as compared with thick-lipped whistles.
A laser-supported air-breathing thruster utilizes the remote laser energy and atmospheric air to boost a vehicle. To calculate the impulse induced by a laser pulse, the operational process was divided into two phases. First, one dimension (1D) laser-supported absorption waves in the air were simulated by an implicit dual-time method, and laser absorption efficiencies were predicted, based on a more accurate absorption model and three temperatures thermal nonequilibrium. Sequentially, impulses for different parabolic thrusters and pulse energies were computed, considering the high-temperature real gas effect. Then experiments were conducted with a ballistic pendulum apparatus. The calculations of 1D absorption waves show that as laser intensity increases, the electron number density would reach the critical value, resulting in a laser reflection and decrease of absorption efficiency. Further calculations for thrusters imply the thrust oscillation due to air-refilling has an evident influence on the total impulse received, and because of a higher thrust peak and longer positive phase time, the flat top and longer configuration would significantly enhance the performance. Experimental results show that the errors of the impulse calculations for two thrusters are 4.2% and 9.4%, respectively, which verifies the calculation model.
The most recent idea of using the exothermic coating to reduce the drag coefficient on blunt noses in hypersonic flows is numerically simulated to study parametrically the effect of material properties. The conjugate heat transfer is considered by simultaneous solution of flow and solid phase governing equations in numerical procedure. The results show that changing some effective parameters in material properties may increase the coating effectiveness significantly. Using such analysis leads to choose the proper materials for specific flight missions.
Linearization of turbofan engines is an effective method for performance control and fault diagnosis. In this study, three different linearization techniques, including the partial derivative method, optimized fitting method and equilibrium manifold, are discussed and compared. First, an optimized fitting method is developed based on the least-square method and an optimization algorithm. To avoid trapping in local optimization solution, the initial values used in the optimization approach are obtained through the partial derivative method. Second, to verify the accuracy and effectiveness of the linear model in the flight envelope, the result of linear modeling method based on equilibrium manifold is analyzed in detail. Finally, an overall assessment of the merits or weaknesses of linearization models is provided based on the obtained results.
This paper defines a method for the optimization of design parameter tolerances. The general architecture of the proposed method is identical to that of the robust design reference method proposed by Taguchi but its content is different as the tolerances are considered as functions to be maximized here, while Taguchi’s method rather considers these tolerances as fixed data. Instead of looking for design parameters that minimize the sensitivity of some performance criteria, the design parameters are calculated so as to obtain maximal tolerance intervals, thus minimizing manufacturing costs. Performance criteria are then considered in terms of optimization constraints: each criterion gives rise to an inequality constraint that specifies the minimum level of performance that the designer wants to achieve. The possibilities offered by this method are illustrated through its use in the preliminary design of a cold-expanded bushing. In this case, tolerance optimization enables the allowable tolerances on the design diameters to be increased and performance levels are defined on the residual radial stress at the bushing/part contact radius and on the residual orthoradial (hoop) stress at the part inner radius.
For autonomous operations, unmanned micro air vehicles depend on novel trajectory generation schemes. Trajectories parallel to the ground are fairly well understood in surveillance and reconnaissance contexts. In addition, trajectories with altitude as one of the navigational parameters are also envisioned for aerial robots, urban terrain coverage, etc. In these missions, trajectory generation in pitch plane becomes an important problem. In this article, a linear algebraic procedure is applied to generate the pitch plane trajectories. That is, when a terrain (latitude, longitude and altitude) is sensed from a global positioning system, a trajectory to navigate the aircraft along the terrain is presented. By linear algebra, the state vector is spanned using a set of known vectors and some of their scalars (coefficients of spanning vectors) are solved in an optimization framework so that the state variables mapping the desired terrain are determined. After the trajectories are generated, normally smoothening is recommended so that the trajectories are continuous for autonomous control development. Thus, an extended Kalman filter is applied to smooth the trajectories. A nonlinear micro air vehicle model is considered to illustrate the trajectory generation and smoothening scheme.
In this research, we focus on an optimal trajectory of a spacecraft rendezvous operation. This optimal trajectory, which is studied in rendezvous problem consists of a controlling distribution, parametric and permitted forms of performance index of minimum fuel-time free; this method will guide the chaser spacecraft toward the target spacecraft in optimal trajectory by using multiple-subarc sequential gradient–restoration algorithm. During trajectory analysis, we will define two problems, P1 and P2. In P1, initial position vector and initial velocity vector are given and fixed and in P2, initial position vector is given and fixed, while initial velocity vector varies and will be free. Previous studies show that the condition of optimal fuel in a solution will be obtained with four subarcs and it shows that for distribution of optimal thrust, the amount of maximum thrust or zero during each subarc will be needed. The main result is that, the new method proposed by this study leads to reduction in performance index and also fuel consumption will be in an appropriate amount. This event takes place in proportion of initial free velocity (problem P2) to initial fixed velocity (problem P1). Another result is that the reduction in fuel consumption and performance index is associated with a remarkable reduction in central processing unit time up to 1650 order. The result of this investigation for rendezvous mission between target spacecraft and chaser spacecraft named OCAT1.1 was extracted and examined on SuperCluster computer.
A kind of robust control of linear electromechanical actuator (LEMA) spoiler system for thrust vector control (TVC) in a spacecraft was investigated. This paper presents an inverse system method (ISM) to achieve precise motion control for the LEMA spoiler system, which is based on a voice coil motor. The dynamic model of the LEMA spoiler system is established, and the inverse system of the LEMA spoiler is obtained. By linearization of the nonlinear LEMA spoiler system, the state feedback control is used to control the spoiler to follow the desired spoiler motion. A state variable observer is designed to estimate the unmeasurable state variables. Simulations and experimental setup are used to demonstrate the effectiveness of the proposed control method. The performance of proportional-integral-differential (PID) control is compared to the ISM control in simulation and ISM control is robust to large parameter variations of 50%. The experimental results show close agreement with the simulation results. The proposed method proved effective by achieving a transient time of 5.5 ms and system bandwidth of 20 Hz, so that ISM control method can more effectively handle the parameters’ perturbation, load disturbance, the variation of work temperature, and the LEMA's nonlinearity.
This paper presents a new guidance method for cruise missiles, which makes use of geomagnetic isolines. The isolines are extracted by using bilinear interpolation in the geomagnetic contour map. Path tracking is implemented in the geomagnetic map and takes into account the dynamics of the missile and its constraints. For this, the dynamic equations for the missile during its cruise phase are established. Using an inverse dynamics guidance law to implement the simulation experiments, the simulation results show that missile flight along a geomagnetic isoline is theoretically feasible. As the path made up of isoline points is determined, the missile can follow the isoline with a path tracking controller.
A new adaptive modeling method for aircraft engine by using equilibrium manifold (EM) and its expansion (EME) model is presented, following research undertaken by the authors at School of Transportation Science and Engineering, Beijing University of Aeronautics and Astronautics, Beijing, China. The property of the expansion model and the effect of mapping design to the form are systematically studied. The model adaptivity analysis is discussed, and this paper also gives the identification procedure of modeling the aircraft engine approximate nonlinear model; the deterioration modification of compressor and the comparison with linear parameter-varying model and Kalman estimator are discussed. Simulations illustrate that modeling accuracy is high and the structure is simpler.
Regeneratively cooled scramjet heat transfer calculation method was developed using three-dimensional calculation for engine solid wall heat conduction. The scramjet thermal environment was determined by engine heat flux measurement and the engine three-dimensional combustion flow calculations. Different heat transfer relationships in the laminar, transition and turbulence fuel flow regions were applied. The cooling fuel flow distribution data inside combustion chamber side panel were obtained by three-dimensional fuel cooling flow calculation. The results of a light weight regeneratively cooled combustion chamber heat transfer tests were adopted to verify the calculation method. The comparisons showed good agreement, indicating that the calculation method is applicable.
As ballistic targets reenter the earth’s atmosphere, they undergo rapid deceleration and may perform spiraling motion that can make accurate tracking very difficult. A variety of filter structures have been proposed for this problem, including variants of the Kalman filter. The present work employs a new target dynamic model combined with an unscented Kalman filter to yield an effective tracking solution. The new target model includes, in addition to position and velocity states, a modified drag coefficient, the target spiral frequency, and two harmonic states describing the periodic rotation of the target body around its velocity vector. The spiral frequency was allowed to have a general variation with time. The unscented Kalman filter with new target model produced very satisfactory tracking for targets with varying frequency. The sensitivity to angle measurement accuracy using a reduced set of measurements was evaluated and a threshold value was determined. Finally, tracking performance was confirmed by Monte Carlo analysis.
Traditional grain designs, which identify the best combination of geometrical parameters to improve the grain performance and meet the flight-mission requirements, are often performed manually. In this article, an integrated framework is presented to perform the design optimization of solid rocket motor propellant grains. In the proposed framework, the level set method is adapted to solid propellant burnback analysis and this technique does not have any restriction on the grain configuration and is capable of handling grains of multi-stage and various burning forms. Along with the level set method, a dedicated algorithm is developed using application programming interfaces of commercial computer-aided design software to transform the initial grain shape into a special data file that can be fed to the level set codes to activate the burnback analysis. Moreover, a hybrid optimization method incorporating genetic algorithm and sequential quadratic programming is exploited to improve the grain design efficiency. Finally, two case studies have been performed to verify the feasibility and general-purpose characteristics of the proposed grain design optimization environment. The results obtained show that the proposed design framework facilitates the grain optimization process and various grain design requirements can be met. The design cycle has been remarkably reduced because of the introduction of hybrid optimization method.
Supersonic single-mode ramjet performance was analyzed using a prescribed two-dimensional conical shock wave in axisymmetric supersonic flow. The ramjet under consideration for the analysis consists of a mixed compression intake, a cylindrical combustion chamber and a supersonic constant convergent–divergent nozzle. A computer program was developed to carry out the analysis based on the formation of multiple conical shock waves at the engine intake at different flight Mach numbers and different altitudes in the range of 1.5–4 Mach and 9000–18,000 m, respectively. Accordingly, a supersonic convergent–divergent nozzle was designed and consequently, the area ratios along the ramjet were calculated to find the correct dimensions for the thrust required. The analysis of the multi-shock system showed that for a given number of conical shocks and Mach numbers, the thrust decreases as the altitude increases. Also, the thrust increases at higher Mach numbers and higher number of conical shocks regardless of the altitude. Furthermore, for M > 2.5, and at number of conical shocks greater than 2, thrust stays constant. The flow rate and the pressure after combustion showed similar trends as the thrust. The multi-shock system of the intake system proposed showed that a limit of a three conical shocks were sufficient for a reasonable pressure recovery for a M > 2, while for a M < 2, a single normal shock wave could be sufficient for different altitudes. Also, pressure recovery is unaffected by the altitude for the same Mach number and increases with lower Mach numbers. Moreover, the increase of number of conical shocks is limited to 3 where no further increase in pressure recovery could be indicated.
Control components of the aircraft environmental control system (AECS), which is fast becoming an increasingly complex system, are of significant importance from the viewpoint of safety. However, few studies have focused on fault diagnosis of the AECS. This study proposes a method based on adaptive threshold and parameter extraction (ATPE) to realize fault detection and isolation for control components in the AECS. To overcome the drawback of a fixed threshold for fault detection, a practical approach is employed by combining a radial basis function (RBF)-based observer with an RBF-based adaptive threshold producer. The RBF neural network observer is used to generate a residual error signal. By comparing the residual error signal with the adaptive threshold, fault occurrence can be detected. To improve the fault isolation accuracy, an RBF fault tracker is used; the parameters of this tracker are extracted for fault isolation along with the residual error, unlike in the case of conventional fault diagnosis methods that are based on a single residual error signal. Finally, an RBF-based fault isolator is adopted to realize fault isolation and classification. Two commonly occurring faults in the control components of the AECS are simulated to verify the performance and effectiveness of the proposed method. The experimental results demonstrate that the proposed method based on ATPE is effective for fault detection and isolation for the control components in the AECS.
This article describes the design, development, build and flight testing of a health monitoring system for the landing gear and the electrical power system on board the Demon prototype unmanned airborne vehicle. Demon is a flying technology demonstrator which successfully flew in September 2010. The Demon can achieve pitch and roll control without the use of hinged control surfaces, by instead using fluidic devices based on the Coanda effect, attaining low-maintenance, high-manoeuvrability operations. A vehicle health monitoring system was added on board between the first and the second flight test campaigns. The integration of the health monitoring system into the vehicle is discussed as a whole. The key health monitoring sub-systems include data logging and real-time measurement of several parameters. This includes systems to measure Voltage and current from the main batteries, landing gear stress, suspension travel, wheel hub acceleration and shock absorber pressure. Wherever possible, the use of commercially available components was maximised to minimise development time and cost. Some example results of system health monitoring during flight trials are presented.
The three-loop autopilot is employed by spinning missiles as well as by many high-performance command or homing guidance missiles currently because it performs well in stabilizing airframe and implementing guidance commands. However, for spinning missiles, the closed-loop system may be dynamically unstable in the form of a divergent coning motion due to the existence of cross-coupling effects. And the stability criteria of the autopilot applicable to the nonspinning missile are no longer valid in the event of the spinning. To address this issue, the structure of a three-loop autopilot of spinning missiles is introduced in this study, for which the sufficient and necessary condition of coning motion stability is analytically derived from the equations in the form of complex summation. The stability criteria are further illustrated by numerical simulation. It is noticed that spinning shrinks the stable region of the design parameters significantly. And the higher the spinning rate, the smaller the stable region becomes.
Non-contacting internal forces among multiple spacecrafts have several advantages such as no propellant consumption and plume contamination and provide a novel approach for spacecraft formation flight control. They differ from traditional thrust in their nature as an internal force, which has potentially complicated the analysis on dynamics and equilibrium of such formations. This article mainly studies the generalized dynamics and relative equilibrium for multi-spacecraft formation with non-contacting internal forces. Treating such a formation as a free multi-rigid-body system connected by force element, Kane method is applied to develop a generalized 6-DOF dynamic model with internal forces accommodated. After verifying the validity of the model for a case of two-spacecraft formation, the relative equilibrium for the generalized model is analyzed and the necessary conditions for circularly restricted static formation with non-contacting internal force are derived, which would provide guidance for formation design and control in the future.
Heavy rainfall greatly affects the aerodynamic performance of the aircraft. Aerodynamic efficiency degradation due to the heavy rain has been the cause of many aircraft accidents. We have studied the effects of heavy rain on the aerodynamic efficiency of NACA 0012 2D airfoil cruise and landing configurations and NACA 0012 3D rectangular wing. Our results show significant increase in drag and decrease in lift in heavy rain environment. For our study we used preprocessing software gridgen for creation of geometry and mesh and fluent as solver. Discrete phase modeling has been used to model the rain particles using two-phase flow approach. The rain particles have been assumed to be inert. In simulated rain environment, both the 2D airfoil and 3D wing showed significant decrease in lift and increase in drag. This study will be quite useful for the designer of the commercial aircrafts and unmanned aerial vehicles and will be helpful for training of the pilots to control the airplanes better in heavy rain environment.
The classical picture of shock evolution in nozzles holds that under over-expanded flow conditions, a single, nominally normal shock exists within the nozzle. Focusing on the highly dynamic flow produced during blow-down of an experimental, high-nozzle pressure ratio, planar nozzle, this article presents visual evidence that shock-trains – here, a pair of parallel, nominally normal shocks – dominate the rapidly evolving flow field. Three principal results are presented in this study. First, high-speed schlieren images of the evolving nozzle flow are reported. Second, a simple qualitative model of shock–boundary layer/recirculation zone interaction is proposed and used to explain observed millisecond-scale shock-train structure. Third, limited wall pressure measurements and schlieren images are combined to propose a second qualitative model of shock-train–boundary layer/recirculation zone evolution on the longer blow-down process time-scale. The results provide insight into millisecond-scale compressible flow dynamics within high-nozzle pressure ratios .
This article presents a fault tolerant attitude determination system for a three-axis satellite including a sun sensor and a magnetometer. The suggested methodology is developed based on all possible rotations between reference and body frames and computation of Euler angles by them. Using the resulted Euler angles, some variance measures have been derived that offer a solution for analytical model-free fault detection. It is demonstrated that by categorizing different computation methods, the contaminated measurement data could be isolated. Also, utilizing the methods in which the contaminated data are not used, we can continue to provide correct Euler angles. The cited features provide a fault tolerant attitude determination system that always generates the correct attitude angles for attitude control purposes. Since these algorithms are model-free, the fault detection and isolation in the attitude determination system is accomplished independent of the health status of actuators in the attitude control system. In this article, through extensive simulation studies, the desired performance and accuracy of the outlined methods are demonstrated.
Flow pattern, aerodynamic performance, vortex-shedding frequency and turbulence intensity were compared between a single wing-blade and two tandem wing-blades for various Reynolds numbers (Re), gap ratios (g*) and angles of attack (α). The wing-blade profile was NACA 0012 which was fabricated from stainless steel. The effects of α comprised the effect of the angle of attack of the front wing-blade (α1) and that of the rear wing-blade (α2). Flow behaviors and flow patterns were visualized using the smoke-wire method. The vortex-shedding frequency and turbulence intensity were measured using a hot-wire anemometer. The smoke-streak flow pattern around the single wing-blade was categorized as five characteristic flow patterns, which are attached surface flow, instability wave in wake, vortical wake, separation from near-leading-edge and bluff-body wake. The smoke-streak flow patterns around the two tandem wing-blades (changing α1 and g*) were classified as attached flow, separation, turbulent flow and vortex street for α2 = 0°. The flow patterns for α1 = 0° (with varying α2 and g*) were attached flow, vortical wake, bluff-body wake and bi-vortex street As α1, α2 and g* changed, the flow patterns were classified as attached flow, vortical wake, bluff-body wake and turbulent flow. The characteristics of St1, St2, I1 and I2 parameters (measured around the tandem wing-blades) were compared with StS and IS (detected around the single wing-blade). The aerodynamic performance was measured using a six-force balancer. For α2 ≥ 30°, the maximum lift coefficient (CL) was reached at g* = 0 because of the equivalent flap effect that was caused from the existence of rear wing-blade.
A global sliding mode controller (GSMC) is proposed for the missile electromechanical actuator (EA) servo system, where exists high uncertainties, such as parameter variations and external disturbances. By the design of an optimal integral switching function based on optimal linear quadratic regulator (LQR) theory, the initial state of system is set on the switching surface, and the optimal sliding mode motion is produced. The proposed GSMC is composed of an optimal linear state feedback controller (OLSFC), and a fuzzy nonlinear robust controller (FNRC), which can be designed respectively. The OLSFC, generated by the designed switching function, intends to minimise a quadratic performance index, and then improves the dynamic performance of system. Meanwhile, the FNRC employs a fuzzy decision maker (FDM), which estimates the upper bound of uncertainties as FNRC’s gain adaptively, and then makes GSMC robust and control input smooth. With the computer simulations on an EA experiment plant, it presents that the proposed scheme possesses good tracking precision, effective suppression against chattering at control input, and strong robustness against system uncertainties.
Thrust vector nozzles are finding place on modern fighter airplanes because of the benefits they provide and also due to diminishing weight penalty of such nozzles. They offer additional benefits in the case of a twin-engine airplane. Different vectoring configurations such as multi-axis vectoring, single-axis pitch vectoring and single-axis vectoring with canted nozzles have been studied with respect to twin-engine airplane configuration. Modeling and integration of thrust vector nozzles with rigid airplane six-degrees-of-freedom equations of motion have been carried out in this article. Using the integrated model, a comparative study is carried out to summarize the capabilities and limitations of various nozzle configurations with respect to performance of an airplane in velocity vector roll and in Herbst maneuvers. The airplane model used in this work is the F-18/HARV and all simulation results have been produced using a nonlinear dynamic inversion controller developed in Matlab/Simulink environment. Results show that a multi-axis thrust vectoring provides additional benefits as compared to single-axis vectoring with canted nozzles in high angle of attack velocity vector roll and in Herbst maneuvers. The single-axis pitch only vectoring has roll control power and lacks in yaw control power, to execute the velocity vector roll maneuver.
An angle-of-attack tracking control system is designed for the hypersonic reentry vehicle, whose aerodynamic parameters vary dramatically during reentry phase. The linear parameter-varying (LPV) theory based on linear fractional transformation (LFT) model (named as LPV–LFT method) is applied to design the controller for hypersonic reentry vehicle. Longitudinal moment trim of the hypersonic reentry vehicle is made along the desired flight trajectory, and a damping feedback loop is firstly designed to improve the system’s damping and static stability. Then, the linear dynamics model with damping feedback loop is established in LFT structure and treated as the controlled plant, and a parameter-varying reference model is utilized to guarantee the transient performance. The effectiveness of the proposed angle-of-attack tracking control system is validated through the frequency domain analysis and step response simulations. Finally, the actual angle-of-attack command tracking simulations using the nonlinear time-varying mathematical dynamics model are carried out to verify the accuracy and robustness of the hypersonic reentry vehicle control system.
The purpose of this paper is to present orientation estimation for small unmanned aerial vehicle (SUAV). An extended Kalman filter with adaptive PR (P denotes the estimation error covariance matrix, and R denotes the measurement noise covariance matrix) is designed to estimate orientation by sensors of gyroscope, accelerometer, and magnetometer integrated in Micro Electronic Mechanic System-based heading reference systems. Since ferromagnetic materials or other magnetic fields near the magnetometer disturb the measurement of local earth magnetic field and the external forces which produce maneuvering acceleration effect the measurement of gravity by the accelerometer, the orientation estimation is disturbed. Accordingly, the error equations of sensors are established using a current statistical model, and then the extended Kalman filter with adaptive PR with 12 state variables is designed. In the filter, the orientation error, gyroscope offset error, magnetic disturbance error, and maneuvering acceleration error are estimated. The swing experiment in hand with the magnetic disturbance and small maneuvering acceleration, and flight experiment for SUAV with the magnetic disturbance and large maneuvering acceleration, are developed. The compensation results show that the orientation is accurately calculated with disturbances. A new methodology for the orientation estimation is proposed, which could also be considered for other special application such as the robot on the ground and the autonomous underwater vehicles. This paper provides a novel realization method for accurate orientation estimation for SUAV. The method can be applied in many applications with a simple hardware.
This study proposes a numerical method for solving transonic moist air flows with non-equilibrium condensation by taking FLUENT as the secondary development platform. The governing equations consist of moist air motion equations with an additional source term in the energy equation, which considers the effects of heat addition, and a set of four ordinary differential equations that are related to the generation of condensate mass. The moist air motion equations are solved using the standard features of the code in FLUENT, whereas the four ordinary differential equations are modelled using the user-defined-scalar transport modelling provided by FLUENT. This numerical method is validated both under internal and external flow conditions in a turbine cascade and over the NACA 0012 airfoil. Further investigations include testing the moist air flows in a compressor cascade channel and over the asymmetric RAE 2822 airfoil. The results show that the non-equilibrium condensation of moist air under transonic flow condition has a significant influence on the flow field structure and the aerodynamic performance of the turbine and compressor cascade.
A modified sequential shifting control algorithm is proposed for changing the helicopter rotor’s speed in a large variation and providing continuous output power to the rotor. Two turbo-shaft engines and two multi-speed gearboxes, coordinating with the rotor, facilitate a wide rotor speed variation and provide continuous torque to the rotor. In the process of shifting, a new control scheme is proposed to design turbo-shaft engines’ power turbine speed controller which is robust linear matrix inequality control, combined with active disturbance rejection control torque feed-forward compensation, so as to weaken engines’ torque disturbance in rotor speed variation process. In the end, some numerical simulations are carried out to verify the feasibility of the shifting algorithm, based on the integrated helicopter/engine system model which can simulate auto-flight tasks of the real system. The simulation results show that the turbo-shaft engine’s torque disturbance has little influence on engine power turbine speed in the process of torque shifting, and the rotor speed can perform large and rapid speed changes smoothly.
The problem of finding a lost target in a noisy environment by a group of flying vehicles is studied in this article. The developed cooperative search algorithm that is decentrally applied on the flying vehicles is a combination of searching guidance and neighborhood laws. The searching guidance law generates an acceleration command to direct each flying vehicle to the position of the lost target. The command is generated based on the information gathered by those flying vehicles that are categorized as neighbors by the neighborhood law. The neighborhood law specifies the sharing network between the flying vehicles for intelligent cooperation. Various neighborhood laws are introduced for tuning the search exploration and exploitation, which influence the performance of the cooperative search algorithm. To evaluate this performance, two approaches are considered. The analytical approach shows that the search process is stable and convergent. In the second approach, numerical simulations demonstrate that properly selecting the neighborhood law significantly enhances the performance of the search.
The conventional model-based diagnosis usually potentially presumes faults are persistent and does not take intermittent faults into account, which is the major cause of the problems of false alarms, cannot duplicate and no fault found in aircraft avionics and present a tremendous challenge to prognostics and health management. Aiming at the problem that the logical automaton proposed by Sampath et al. cannot distinguish between strings or states that are highly probable and those that are less probable, a stochastic automaton approach is given to distinguish the fault types by extending the fault model to include both permanent faults and intermittent faults. The notions of A- and AA-diagnosability of permanent faults and intermittent faults for stochastic automaton are defined. Thereafter, the diagnoser with a probability matrix appended to each transition that can be used to update the probability distribution on the state estimate is constructed. Finally, an example of aeronautic gyroscope is presented to demonstrate the proposed approach, and the analysis results show that this approach is able to discriminate the fault types within bounded delay if the system is A- and AA-diagnosable. In our previous paper, we have extended the logical automaton model, and investigated the stochastic automaton approach in this article.
The computation of heat flux on two current re-entry capsules, European eXPErimental Reentry Testbed (EXPERT) and Orion, has been carried out by a direct simulation Monte Carlo code (DS2V) and by a computational fluid dynamic code (H3NS) in transitional regime, considering both non-reactive and fully catalytic surface. These capsules have been chosen for this analysis because they have been characterized by completely different shapes and re-entry trajectories. DS2V and H3NS use the Gupta and the Park chemical models, respectively. The results showed that the heat flux predicted by DS2V is always higher than that predicted by H3NS. Therefore, a sensitivity analysis of the chemical models on the heat flux has been carried out for both capsules. More specifically, the Park model has been implemented in DS2V as well. The results showed that DS2V and H3NS compute a different chemical composition both in the flow field and on the surface, even when using the same chemical model (Park); therefore, the different results obtained from the two codes can be attributed mostly to the different methodology used in handling all chemical processes.
An aeroelastic analysis is used to investigate the rate dependent hysteresis in piezoceramic actuators and its effect on helicopter vibration control with trailing edge flaps. Hysteresis in piezoceramic materials can cause considerable complications in the use of smart actuators as prime movers in applications such as helicopter active vibration control. Dynamic hysteresis of the piezoelectric stack actuator is investigated for a range of frequencies (5 Hz (1/rev) to 30 Hz (6/rev)) which are of practical importance for helicopter vibration analysis. Bench top tests are conducted on a commercially available piezoelectric stack actuator. Frequency dependent hysteretic behavior is studied experimentally for helicopter operational frequencies. Material hysteresis in the smart actuator is mathematically modeled using the theory of conic sections. Numerical simulations are also performed at an advance ratio of 0.3 for vibration control analysis using a trailing edge flap with an idealized linear and a hysteretic actuator. The results indicate that dynamic hysteresis has a notable effect on the hub vibration levels. It is found that the theory of conic sections offers a straight forward approach for including hysteresis into aeroelastic analysis.
This article proposes an indirect approach for fault diagnosis and fault-tolerant control in the satellite attitude control system with sampled-data measurements. The proposed method is based on a discrete-time approximation model of the continuous attitude dynamics. By considering the fault term as an auxiliary state vector, an augmented plant is constructed. Then an observer is designed to simultaneously estimate the system state vector and the fault term. Specifically, the observer design problem is reformulated as a set of linear matrix inequalities and can be conveniently solved by standard linear matrix inequality tools. The fault-tolerant controller is easily derived using the fault diagnosis result and the H index is adopted to analyze the fault-tolerant control performance. Finally, numerical simulation results are given to demonstrate the effectiveness of the proposed method.
An experimental activity was conducted to investigate the aerodynamic effects of a stream-wise vortex impacting on a NACA 23012 oscillating aerofoil. The experimental setup allowed to study the effects of the blade pitching motion in the interaction with the vortex. The impacting vortex was statistically qualified by means of a three-dimensional hot-wire anemometry, taking into account also the vortex wandering phenomenon. The flow developed on the aerofoil was investigated through particle image velocimetry surveys carried out on different measurement planes along span-wise direction. The experimental study investigated both the light and the deep dynamic stall, representing typical helicopter flight conditions. In particular, in the tested light dynamic stall condition, the phase averaged velocity fields showed that in downstroke, the vortex impact triggers the flow separation on the aerofoil upper surface. Therefore, the vortex interaction can introduce detrimental effects on the blade performance. Moreover, the influence of the target aerofoil oscillating motion on the vortex trajectory was investigated.
This paper considers the formation dynamics with the target satellite in an elliptic orbit. A new state transition matrix for the linearized equations of relative motion is presented by series expansion and some mathematical transformations. The state transition matrix is applicable to any eccentricity elliptic reference orbit. Besides, the state transition matrix is just related to the trigonometric function of the true anomaly of the target satellite, so it is easy to calculate. With the state transition matrix, the contribution of three-order nonlinearity in the differential gravitational acceleration on the relative motion is estimated by a perturbation approach. Numerical simulations are included to evaluate the proposed methods.
The unpredictability and uncertainty of Operationally Responsive Space (ORS) missions have brought new challenges to traditional spacecraft design methods, in which the satellite and the launch vehicle are designed independently. In this article, a novel integrated design method for ORS spacecraft design has been proposed by analyzing the characteristics of the ORS missions according to the system theory. Firstly, the feasibility of the integrated design method is analyzed. Then, qualitative analyses of integrated design method for the subsystems of the traditional satellite and the launch vehicle are carried out. Finally, the advantages of the integrated design method are demonstrated through quantitative research, especially for the uncertain space missions. Simulation results show that the integrated design method is an effective technology to realize unpredictable ORS missions.
This article introduces a tool chain for extended physics-based wing mass estimation. Compared to state-of-the-art tool chains, the physics-based structural modelling is extended beyond the wing primary structure. The structural model also includes the movable trailing edge devices including tracks, the spoilers, the engine pylons and the landing gear. The chain consists of the structural analysis model, models for aerodynamic, fuel, landing gear and engine loads as well as a sizing algorithm. To make the complexity of the model generation process feasible for preliminary aircraft design, a knowledge-based approach is chosen. This means that the analysis models are created partly automatically, which leads to a minimum set of required input parameters for the central model generator. The DLR aircraft parametrisation format Common Parametric Aircraft Configuration Scheme is used as central data model for input and output. Therefore, the chain can be easily included in a wider multidisciplinary aircraft design environment.
Dual combustion ramjet (DCR) is one of the most promising and realistic propulsion systems to realize hypersonic flight. The flow fields and the performance of the full-size DCR in Ma4/17 km and Ma6/25 km flight conditions are investigated through direct-connected experiments and numerical simulations. The pressure distributions from simulations are in agreement with that from experiments under both cold flow and hot flow conditions. Different combustion modes are revealed according to numerical results: purely subsonic flow field is established in the front part of the combustor in Mach 4 condition, and there is a thermal choked throat; only central flow is subsonic in Mach 6 condition, and the peripheral supersonic air results in a lower static temperature (less than 2000 K), which is beneficial to thermal protection of the combustor wall. The thrust increment increases with the increasing of the equivalence ratio (). The thrust increment is 8.1 kN for Mach 4 when = 0.9; however, further increasing the equivalence ratio causes unstart of the supersonic intake. The thrust increment is 3.15 kN for Mach 6 when = 1.0. The equivalence ratio affects the combustion efficiency and the specific impulse with the same trend. The maximum combustion efficiency is 0.91 for Mach 4 and 0.89 for Mach 6. The maximum specific impulse is 13.3 kN·s/kg for Mach 4 and 7.96 kN·s/kg for Mach 6. In general, the performance is good and the DCR is worth further investigating.
This article presents an adaptive filter backstepping control strategy for reusable launch vehicles attitude tracking during reentry phase in the presence of input constraints, model uncertainties and external disturbances. The control-oriented model with uncertainties is constructed, where the uncertainties do not satisfy the linear parameterization assumption. To cope with input constraints, an auxiliary system is introduced, and the states of which are applied to the procedure of control design and stability analysis. Second-order filters are employed to overcome the ‘explosion of terms’ problem inherent in traditional backstepping control. Moreover, the stability of the closed-loop system is proven via Lyapunov technique, and the tracking error can be forced into an arbitrarily small neighborhood around zero (i.e. semi-globally uniformly ultimate bounded tracking). Finally, the 6-degree-of-freedom nonlinear reusable launch vehicle simulation results are presented to verify the effectiveness of the control strategy.
This article deals with the effect of leading edge imperfections on the aerodynamic characteristics of a NACA 632-215 laminar aerofoil at low Reynolds numbers. Wind tunnel tests have been performed at different Reynolds numbers and angles of attack and global aerodynamic loads were measured. To perform these tests, a NACA 632-215 aerofoil was built up in two halves (corresponding to the upper side and to the lower side), the leading edge imperfection here considered being a slight displacement of half aerofoil with respect to the other. From experimental results, a quantitative measure of the influence of the leading edge displacement on the degradation of the aerofoil aerodynamic performances has been obtained. This allows the establishment of a criterion for an acceptance limit for this kind of imperfection.
The characteristics of cavity-based supersonic flow and combustion in a scramjet combustor are investigated both experimentally and numerically. Cavities with depth D = 8 mm, length-to-depth ratio L/D = 4 or 7 and aft angle A = 22.5, 45 or 90° are considered. In non-reacting flows, the cavity shear layer dives deeper into the cavity with decreasing aft angle, resulting in a more intense impingement of the shear layer on the aft wall and a stronger trailing edge shock but weaker oscillations within the cavity as well as a smaller penetration of the upstream-injected hydrogen jet. The cavity with larger aft angle is beneficial to promote the instabilities evolving in the jet-mixing layer and accelerate the breakdown of the counter-rotating vortices, resulting in more rapid fuel–air mixing. In reacting flows, the cavity with larger aft angle also exhibits stronger oscillations and higher combustion efficiency with no greater total pressure loss. The results indicate that cavities with larger aft angle may be more beneficial to enhance the supersonic mixing and combustion as long as the oscillations are not too violent to induce blowout or blowoff of the cavity-stabilized combustion.
This article presents the effectiveness of a tab shedding small-scale vortices of continuously varying size, in enhancing the mixing of axisymmetric Mach 2 jet at different levels of expansion. Corrugations in the form of semicircle, triangle, and square geometries located on the slanting edges of two identical isosceles triangular tabs, placed at diametrically opposite locations of the circular nozzle exit have been studied. The corrugated tabs are found to be the better mixing promoter than uncorrugated tabs. Among the corrugation geometries, the semicircular corrugation with two sharp corners is found to be the best mixing promoter in the presence of marginally overexpanded and almost zero pressure gradients, for the Mach 2 jet. However, for a highly overexpanded state, the performance of semicircular and square corrugated tabs is comparable. A reduction in core length as high as 90% was achieved with semicircular corrugation tabs, compared to uncontrolled jet in the presence of marginally adverse pressure gradient, corresponding to nozzle pressure ratio 7.
The study of detonations and their interactions is vital for the understanding of the high-speed flow physics involved and the ultimate goal of controlling their detrimental effects. However, producing safe and repeatable detonations within the laboratory can be quite challenging, leading to the use of computational studies which ultimately require experimental data for their validation. The objective of this study is to examine the induced flow field from the interaction of a shock front and accompanying products of combustion, produced from the detonation taking place within a non-electrical tube lined with explosive material, with porous plates with varying porosities, 0.7–9.7%. State of the art high-speed schlieren photography alongside high-resolution pressure measurements is used to visualise the induced flow field and examine the attenuation effects which occur at different porosities. The detonation tube is placed at different distances from the plates' surface, 0–30 mm, and the pressure at the rear of the plate is recorded and compared. The results indicate that depending on the level of porosity and the Mach number of the precursor shock front secondary reflected and transmitted shock waves are formed through the coalescence of compression waves. With reduced porosity, the plates act almost as a solid surface, therefore the shock propagates faster along its surface.
The influences of wall functions, reaction model, and wall Prandtl number on the reacting flow field characteristics of a typical cavity-based scramjet combustor have been investigated numerically, and the computed results have been compared with the available experimental data in the open literature. The computed results are in reasonable agreement with the experimental data, but the pressure downstream of the leading edge of the cavity is overpredicted. Meanwhile, the grid discrepancy has been analyzed by employing three different grid scales, namely the coarse, the moderate, and the refined grids. The obtained results show that the moderate grid may be employed with confidence to compute the reacting flow field of the scramjet combustor. The wall functions, reaction model, and wall Prandtl number have a large impact on the pressure distribution along the floor face of the cavity, and this may be due to the influence of the variance of the flow separation point on the top wall for the different cases investigated with different inlet boundary conditions. The static pressure along the floor face of the cavity decreases with an increase in the wall Prandtl number, and then increases when the wall Prandtl number is large enough, namely 1.2 in the current study. This points out that there exists an optimal wall Prandtl number for the prediction of the reacting flow field of the cavity-based scramjet combustor.
A quasi-one-dimensional combustor model is developed using MacCormack’s method to solve simplified Navier–Stokes equations for the intention of simulating the flow parameters in the combustor. The combustor model is capable of predicting the flowfield of scramjet combustor. The Eckert reference technique is adopted in the heat transfer model of the regenerative cooling system. With these two models, the relationships between phenomena are investigated. Simulation results indicate that this coupled model of combustor and regenerative cooling system can be used to investigate the influence of the cooling channel geometry, flight Mach number and fuel equivalence ratio. This coupled model is useful for the pre-design of scramjet engines.
Conjugate heat transfer studies are presented for high speed aerospace vehicle using commercial CFD software. Navier Stokes equations in the fluid domain and transient heat conduction equations in the solid domain are solved simultaneously to obtain the skin temperature history and other flow parameters. The computational methodology is applied to predict the surface temperature of high speed aerospace vehicle after validating the methodology against experimental results. Validation cases include laminar flow past axisymmetric double cone at Mach 4.57 and turbulent flow past circular cylinder at Mach 6.7. Computed flow field including cold wall heat flux, surface temperature distribution, surface temperature history match nicely with experimental as well as other numerical results. Temperature dependent material properties are found to have significant effect on the surface temperature prediction. Computed surface temperature of a high speed aerospace vehicle show good overall match with flight measured values.
The rotor tip clearance in a gas turbine engine varies throughout the engine operating regime. It has considerable influence on the engine performance. Blade to casing rub is imminent at certain operating points of the engine. Mechanical rub at high speeds could damage the total engine hardware. A precise measurement helps to obtain the optimum engine performance with safe engine operation. In this article a typical case study related to fan clearance measurement is discussed, where indications of a proven measurement system is not in agreement with the physical event during engine test. Centrifugal, thermal, assembly and wear effects can affect tip clearance measurement. Centrifugal forces untwist the blade tip, resulting in change in the effective area of the target that is seen by capacitance sensor. Relative component growths due to thermal effect result in the displacement of the sensor from its original position. This could induce error into this measurement. Assembly errors are seen during blade to disc assembly. Wear occurs under the action of centrifugal loading and vibration in compressor blades dovetail roots that are attached to the disc. This leads to wear in involved metal surfaces and it could be a source of error in this measurement. Measurement system also has its own uncertainty. During the current work all sources of errors were evaluated. Probable actual running clearance on the engine and reasons for the mismatch in indication were successfully arrived at through analytical and experimental studies. This work has provided an insight into probable sources of errors and their treatment methodologies using analytical and experimental techniques. This has helped in identifying the changes needed in the calibration procedure, methods to reduce the measurement system uncertainty band and measurement procedure.
Concurrent orbit and attitude determination (COAD) plays a key role in reducing the cost of navigation and control subsystem for small satellites. This article is devoted to the problem of the COAD of satellites. A measurement package consisting of three axis magnetometer (TAM) and a sun sensor is shown to be sufficient to estimate the attitude and orbit information. To this end, an autonomous gyro-less COAD algorithm is proposed and implemented through the centralized data fusion of the TAM and the sun sensor. The set of nonlinear-coupled roto-translation dynamics of the satellite is used with a modified unscented Kalman filter (MUKF) to estimate the full satellite states. The MUKF is specially proposed to substantially cut the run time by minimizing the number of required sigma points. The results indicate that the adopted strategy fulfills the essential requirements of accuracy and the speed of state estimation. Local observability is demonstrated and an extensive Monte Carlo simulation has shown desirable stability characteristics for the proposed algorithm. Additionally, a sensitivity analysis on the orbital elements and sensor characteristics is performed to verify the feasibility and utility of the MUKF over a wider acceptable range of sensory and operating environments.
Large light-weight deployable cable net reflectors are widely used in communication satellites. These are much larger than satellite and need stowing into a relatively small volume to fit into launch vehicle and are deployed in orbit. The cable net reflectors use deployable support structure along with pretensioned cable net. Due to its large dimensions, the deployed natural frequency will be quite low and may not be acceptable from controls point of view. In such cases, the deployed frequency needs to be enhanced to ensure that the same meets the minimum requirements. This article presents a study of the different parameters, which have been varied in order to assess the sensitivity of the same with respect to the deployed frequency. The effect of the flexibility of each sub-assembly has been assumed and this has helped to identify the sub-assembly, which has to be stiffened in order to enhance the overall frequency. Also, the results from this study have provided the design inputs for modifications to be carried out in order to realize a large cable net antenna.
This article addresses the problem of determining the shortest path that connects a given initial configuration (position, heading angle, and flight path angle) to a given rectilinear or a circular path in three-dimensional space for a constant speed and turn-rate constrained aerial vehicle. The final path is assumed to be located relatively far from the starting point. Due to its simplicity and low computational requirements the algorithm can be implemented on a fixed-wing type unmanned air vehicle in real time in missions where the final path may change dynamically. As wind has a very significant effect on the flight of small aerial vehicles, the method of optimal path planning is extended to meet the same objective in the presence of wind comparable to the speed of the aerial vehicles. But, if the path to be followed is closer to the initial point, an off-line method based on multiple shooting, in combination with a direct transcription technique, is used to obtain the optimal solution. Optimal paths are generated for a variety of cases to show the efficiency of the algorithm. Simulations are presented to demonstrate tracking results using a 6-degrees-of-freedom model of an unmanned air vehicle.
Hybrid combustors are of increasing interest for space and civilian propulsion. A test facility has been settled to investigate high-density polyethylene combustion (propellant of length 0.15 m). A parametric study has been conducted on the oxidiser nature (gaseous oxygen diluted in nitrogen, from 31.4 vol.% to 69.2 vol.% of O2), on the oxidiser flow rate (from 28.6 g/s to 53.1 g/s), on the combustor pressure (from 11.4 bar to 25 bar) and on the nozzle diameter (from 6.4 mm to 12.9 mm). The regression rate has been estimated by weight loss (mean value of 0.207 mg/s) and by thermocouples (0.198 mg/s). Its values are compared to existing data through the Marxman law; this enlarges the range of validity of this law. The conduction heat flux in the solid reducer is estimated around 6000–8000 W; which is related to the low regression rate of the solid fuel. The axial thrust has been measured in addition to other parameters (pressures, temperatures and mass flow rates). Solid particles have been gathered at the combustor outlet to conduct additional chemical analyses. These particles were formed at the surface of the reducer and extracted by the oxidiser from the solid surface.
The development of an L1 adaptive control system for the control of satellites in elliptic orbits using solar radiation pressure is the subject of this article. The nonaffine-in-control spacecraft model includes the gravity gradient torque, the control torque produced by two solar flaps, and external time-varying disturbance torque. The objective is to control the pitch angle of the spacecraft using the solar flaps. The design is based on the L1 adaptive control theory for the control of nonlinear nonautonomous uncertain systems. The control system includes an adaptation law based on the state prediction error. Unlike traditional adaptive systems, the control input is obtained by filtering an estimated control signal through a low-pass filter. In the closed-loop system, the designed adaptive law accomplishes large angle maneuver. A special feature of the control system using filtered signal is that it is possible to select large adaptation gains for fast adaptation and to obtain quantifiable performance bounds. Simulation results are presented which show that in the closed-loop system, precise pitch angle control is accomplished, despite parameter uncertainties and external disturbance input in the model.
A review of some human pilot models of the years 1970–1990 – the Kleinman–Baron–Levison optimal control model, the Davidson–Schmidt modified optimal control model, and the Hess optimal control model – has been presented from the perspective of a new model based on the optimal control synthesis of time-delay systems. In the first of the listed models, the ‘central nervous’ reaction of the pilot is naturally defined as a pure time delay in the measurement equation of the system. In the framework of the optimal control theory, the pilot’s behavior is modeled by linear quadratic regulator gain and Kalman–Bucy filter with a linear predictor. Starting from this optimal model of the 1970s, the other two models assumed the Padé approximation of the pure time delay, thus eliminating the linear predictor. In this article, the pure time delay of pilot reaction was reconsidered and divided, for convenience, into two equal parts: for the output measurement equation and for the input control. The pilot model problem has been first defined in the framework of rigorous time-delay synthesis and then solved by making reference to the control separation and duality principles. A closed-form expression of the solution is thereby obtained. The proposed model was then compared by numerical simulations with Kleinman and Hess consacrated models. The analysis of the results shows that this new pilot model is described by a simplified representation, instead denoting similar performance versus previous optimal models – which contains additional insertions as Kleinman–Baron predictor or Padé approximation, respectively. Finally, joint evaluation of the proposed model and Kleinman and Hess models with respect to the well-known Neal–Smith criterion confirms the consistency and viability of the employed strategy as a possible tool for pilot-induced oscillations phenomenon investigation.
In recent times, shock tube flows have been widely employed in many micro-scale devices in the fields of propulsion technology, micro-heat engines, particle delivery systems, and so on. The very small length scales in such micro-shock tubes make the flow physics more complicated compared to the ordinary macro-shock tubes. The major differences in the flow features are the profound influences of wall effects and rarefaction effects. The rarefaction effect alters the boundary layer structure by imparting additional velocity and thermal gradients to the wall-bounded fluid. These phenomena can strongly affect the micro-shock tube flow characteristics such as shock–contact wave speeds, wave propagations, hot and cold zone properties. The main objective of the present work is to produce a detailed understanding on the wave propagation characteristics in a micro-shock tube under rarefied conditions using computational fluid dynamics methods. The shock–contact interface movement under different operating conditions such as Knudsen number and pressure ratio are investigated in detail and compared with the macro-scale shock tube flows. The difference between the numerical and analytical works and their cause is identified and discussed. The results obtained show that the shock strength attenuates rapidly for micro-shock tubes compared to macro-shock tubes. The shock–contact propagation and the distance between them in a micro-shock tube have a strong dependence on rarefaction effects. The more the rarefaction effects are, lesser will be the shock–contact distance. The shock–contact distance decreases as the pressure ratio increases. A strong attenuation in shock strength can also be observed as the rarefaction increases.
Kalman filters are very popular in gas path diagnostics. This algorithm estimates the engine state variables to assess engine health conditions and is accurate in tracking gradual deterioration. However, the performance of the Kalman filter deteriorates when an abrupt fault occurs. There could be a long delay with the Kalman filter in diagnosing the abrupt fault. In addition, the Kalman filter may transfer the abrupt fault on to other components. In this article, an adaptive gas path diagnostic method using strong tracking filter is described that can track gradual deterioration and abrupt fault accurately. The strong tracking filter is an adaptive extended Kalman filter, which introduces suboptimal fading factors into the prediction error covariance of the extended Kalman filter algorithm. The suboptimal fading factors automatically increase when an abrupt fault occurs, therefore, more importance is given to the new measurement in state estimation which allows the filter to quickly track abrupt faults. All of the suboptimal fading factors become one when gradual deterioration occurs, and in this situation, the strong tracking filter becomes the common extended Kalman filter to filter the measurement noise. Therefore, the strong tracking filter can track abrupt faults quickly and accurately, filter measurement noise, and obtain noise-free parameter estimation for gradual deterioration. The strong tracking filter is applied to heavy-duty gas turbine gas path diagnostics for a variety of simulated fault cases to demonstrate the capability of the strong tracking filter in accurately tracking gradual deterioration and abrupt fault.
In this article, solid fuel transient burning behavior under oxidizer gas flow is numerically investigated. It is accomplished by the analysis of regression rate responses to the imposed sudden and oscillatory variations at inflow properties. The conjugate problem is considered by simultaneous solution of flow and solid-phase governing equations to compute the fuel regression rate. The advection upstream splitting method is used as the flow computational scheme in a finite volume method. The ignition phase is completely simulated to obtain the exact initial condition for response analysis. The results show that the transient burning effects that lead to the combustion instabilities and intermittent extinctions could be observed in solid fuels such as solid propellants.
The aerodynamic design of micro air vehicles is challenging since previous studies have shown that the aerodynamic efficiency of airfoils and wings decreases substantially at low Reynolds-numbers. While many MAV approaches investigate biological designs, here a study is conducted on the aerodynamics of paper airplanes, which fly in the same Reynolds-number range as MAV, but have the advantage of simplicity compared to biological counterparts. Flow visualizations and force measurements in a water tunnel as well as large-eddy simulations are presented on one of the simplest paper airplane design: the dart. The results show that the high-sweep delta design of such an airplane provides high lift coefficients at low Reynolds-numbers. Furthermore, the centerfold of the airplane as a mean to improve the aerodynamic performance is identified.
This article focuses on the application of different nonlinear filtering techniques for geomagnetic navigation, including extended Kalman filter, unscented Kalman filter, particle filter and extended Kalman particle filter. The research evaluates the four methods for navigation of missile during its cruise phase. The measurement equations are obtained by using a surface spline method with the real regional geomagnetic data. Simulation results show all the filters have good performance in areas with abundant geomagnetic information. Among the four filters, unscented Kalman filter tops in the convergence rate and precision with a fairly low tuning sensitivity to the flight path. As the performance of unscented Kalman filter effect is influenced by unscented conversion parameters, a method of parameters optimization using genetic algorithm is presented. The method’s feasibility is further demonstrated by robustness analysis of optimal parameters.
The results of a comprehensive experimental campaign are compared to computational fluid dynamics simulations results to assess the modelling capabilities for a NACA 23012 pitching airfoil in deep dynamic stall regime. The experimental campaign involved fast unsteady pressure measurements and particle image velocimetry. Two-dimensional simulations were carried out with EDGE, developed by FOI. The investigated test case consists in a sinusoidal pitching motion with a 10° amplitude and a reduced frequency of 0.1 around a mean angle of attack of 10°. The behaviour of the experimental lift and pitching moment coefficients is in close agreement with the two-dimensional simulations results, also during the downstroke motion where the flow field is characterised by severe unsteadiness conditions. A three-dimensional numerical model was built to evaluate the relevance of three-dimensional effects on the experiments. Three-dimensional simulations were carried out using the commercial code FLUENT. During upstroke motion, three-dimensional simulations results are in better agreement with the experiments, in particular in terms of the lift coefficient curve slope and of the pitching moment coefficient peak. The flow fields evaluated by particle image velocimetry surveys show strong vortical structures moving on the airfoil upper surface during the downstroke motion that are captured only by the three-dimensional model; then, the flow fields comparison demonstrates the importance of three-dimensional effects for a deep dynamic stall condition.
A new direct adaptive failure compensation approach is developed for satellite attitude control systems in the presence of uncertain failures of redundant actuators. The adaptive failure compensation controller is designed via a backstepping design, which can accommodate uncertainties in actuator failure time instants, values, and patterns. The failure uncertainties are estimated directly by adaptive laws and the adaptive satellite attitude control system with actuator failures is analyzed, to show its desired stability and asymptotic tracking properties. Finally, simulation results of a satellite attitude control system with redundant reaction wheels are presented to demonstrate the effectiveness of the proposed adaptive failure compensation scheme.
Fluidlastic isolators are proposed for higher harmonic pitch link loads reduction in helicopter rotors. An aeroelastic simulation of a soft in-plane rotor in high forward flight is conducted to investigate the dynamic characteristics of the coupled rotor and fluidlastic isolators. Using Hamilton’s principle, the system equations of motion are derived based on the generalized force formulation. The results indicate that the application of the fluidlastic isolator can reduce the 4/rev pitch link load by 98.9% in high forward flight with small variations in the other harmonic loads. The isolator has significant influence on the higher harmonic torsional rotation of the blade tip. Increasing the tuning port area ratio can significantly reduce the tuning mass with little variation of the isolation ability. Within 5% variation of the normal rotor speed, the 4/rev isolator can reduce more than 80% of the 4/rev pitch link load. The effects of isolator’s damping, forward speed, and thrust on the performance of the isolator are also studied. The ability to isolate other higher harmonic pitch link loads is also investigated.
Combustion characteristics in a supersonic combustor with normal and angled hydrogen injection upstream of a cavity flameholder are investigated numerically using a hybrid Reynolds-averaged Navier–Stokes/large eddy simulation method acting as a wall-modeled large eddy simulation. A turbulent incoming boundary layer with thickness of inf = 2.5 mm is considered and a recycling/rescaling method is adopted to treat the unsteady inflow. Three injection angles, α = 30°, 60° and 90°, are considered. The results show that combustion efficiency increases with increasing injection angle since the fuel jet with larger injection angle tends to benefit more from the close coupling of flow, mixing and combustion. Moreover, it is found that the heat release distribution in the streamwise direction is more uniformly for larger injection angle than for lower injection angle, which tends to result in higher total pressure recovery in the far downstream regions for larger injection angle as uniformly-distributed heat release seems beneficial to reduce the total pressure loss for the present diffusion combustion.
In this study, due to the innate trans-atmospheric nature of flight of the space transportation system, assessment of the control discipline interaction with aerodynamic, weights and sizing, external fin-stabilizers configuration, and trajectory disciplines in an multidisciplinary design optimization-based platform has been addressed. Parameters considered for the control sub-system optimization are external stabilizing fins geometrical characteristics and attitude control vernier motors thrust value. Specifically, this article addresses optimization of fin–body combinations with geometric constraints for minimizing control moment required by vernier motors as well as total possible control sub-system weight satisfying design constraints. Results show that using external stabilizer fins is not economical from energetic stand point for space transportation system, but is necessary for control subsystems when there are deflection constraints for vernier motors.
Cast in place liquid shim is usually used to fix the gap/mismatch problems, which occur during assembly process of large aircraft structures. The effect of liquid shim on performances of the assembled structures is influenced not only by thickness and mechanical properties of the shim, but also by the location where the shim is used, that is to say the stiffness of the substrates may have impact on the shim’s effect. To study the usage of shim in composite-to-titanium bolted joints, a three-dimensional finite element method is introduced, and the method incorporates the progressive damage of composite materials, elastic-plastic property of the titanium alloy, super-elastic property of the liquid shim, contact relationships between the joint elements, and real assembly conditions of the mechanical joints. After validating through comparing with the experimental results, the modeling method is adopted to simulate the tensile response of the bolted joints with shims. Furthermore, both the influence of liquid shim layer thickness on the mechanical behaviors of composite-to-titanium bolted joints and the influence of the substrate stiffness on the liquid shim effect are studied in detail. Based on the analysis of the results, it can be concluded that the maximum load, initial joint stiffness and design load of the joints decrease with the increase of liquid shim layer’s thickness; and the effect of liquid shim layer relies heavily on the stiffnesses of the substrates and will reduce when the substrates become stiffer.
In this article, a method for satellite constellation build-up is presented based on three-body dynamics. This method enables satellite constellation build-up with single launch and less energy and time by designing round trip trajectories from low Earth orbit to halo orbit and by distributing satellites in orbital planes. Stable and unstable manifolds associated with halo orbit around <inline-formula id="ilm1-0954410013476615"><inline-graphic xlink:href="10.1177_0954410013476615mml-inline1"/>L_1 </inline-formula> and multiple shooting method are used in order to design trajectories. Multiple shooting method is a powerful tool for solving two-point boundary value problems with high rate of convergence. For satellite constellation build-up, a transfer trajectory from low Earth park orbit to the halo orbit and six return trajectories from halo orbit to six orbital planes which form the constellation are needed. Comparison of results of the method presented in this study with other conventional methods shows its relative advantages including reduced energy, time and cost needed for satellite constellation build-up.
The aero-thermodynamic effects of water ingestion on an axial flow compressor performance are presented in this article. Under adverse weather conditions, gas turbine engine performance deteriorates and in extreme cases, this performance deterioration may result in flameout or shutdown of the engine, which means that serious incidents or possibly accidents may occur. When the water droplets enter into the engine they break up into smaller droplets which may bounce, coalesce or splash onto the compressor blades. They also form a liquid film whose motion is influenced by inertia forces, blade friction, aerodynamic drag and pressure gradient. The water liquid film has considerable effects on blade’s geometric characteristics. Apart from the change in its profile due to thickness increase, air shear force and water droplets momentum cause waves in water film’s surface introducing a kind of ‘roughness’ on blade’s surface. The current work focuses on the aero-thermodynamic effects. Its methodology is based on computational fluid dynamics, which is used to solve the flow field of the computational domain. The model consists of an extended inlet, an inlet guide vane, a rotor and a stator blade. Several cases with water ingestion are solved, varying the parameter of water mass and engine rotational speed, simulating adverse weather conditions. On the rotor blade, the water film height and speed are calculated at the equilibrium condition. This condition is achieved when the water mass which flows out of the blade surface equals with this which impacts on it. Taking into account the film thickness at each computational node of the blade surface, the blade’s geometry is changed. Furthermore, an equivalent roughness is introduced and the effects on compressor’s performance are calculated. It is found that deterioration is more pronounced in low rotational speed. For 4% water/air, compressor’s isentropic efficiency deteriorates 8.5% for idle speed and 1.6% for full speed. For the same water mass, mass flow capacity deteriorates 2.4% at idle speed while the change is small for full speed.
A new pipe-routing method for aero-engines is proposed in this article. Careful consideration of the spatial characteristics and the primary engineering constraints of aero-engines yielded a new space representation method as well as a space diving method for the aero-engine surface, which simplify the searching space. A genetic algorithm, along with certain modified strategies, including the ‘initiation’ and ‘direction guideline’, is also developed for pipe-routing in the sub-spaces. Simulation experiments in the context of various scenes have been carried out to explore the applicability and performance of the propose method. Simulation results showed that this novel method can quickly deliver the optimal routes for any aero-engine of large space and with complex mechanical components while avoiding convergence into local optimal value in comparison with some existing published methods.
In the frame of the European Space Agency (ESA) Intermediate eXperimental Vehicle (IXV) project, ESA is coordinating a series of technical assistance activities aimed at verifying and supporting the IXV industrial design and development process. The technical assistance is operated by the Italian Space Agency, by means of the technical support of the Italian Aerospace Research Centre. One of the purposes of the activity is to develop an independent capability for the assessment and verification of the industrial results with respect to the aerothermodynamic characterization of the IXV vehicle. To this aim, CIRA have developed an independent aerothermodynamic database, intended as a tool generating in output the time histories of local quantities for each point of the IXV vehicle surface and for each trajectory, together with an uncertainties model. The whole procedure followed for the definition of the numerical tool and the main results achieved will be presented in this article.
The stealth aircraft, studied in this article, plans a low observability trajectory to evade radars tracking, considering probability of detection and system constraints. An elaborate framework of planning low observability trajectory, which integrated the models of the stealth aircraft and radars, the theory of multi-phase optimal control and the algorithms of adaptive pseudospectral method, is presented in this article. The constraints and temporal features of low observability trajectory are modeled. The optimal objectives of the flight time, the total fuel consumption, and stealth are defined and synthesized. The trajectory planning problem then was formulated as a multi-objective multi-phase optimal control problem, which can availably grasp the radar tracking features to ensure safety in the whole flight process. A hybrid heuristic adaptive pseudospectral method is developed to solve the trajectory planning problem. The novel algorithm integrates prior knowledge, error estimate and solution regularity for adaptive strategy, which improves computational efficiency and convergence speed. The results of experiments show that the proposed method is feasible and the radar tracking features are effectively utilized to optimize the comprehensive efficiency of penetration.
Smoothing thin plate splines, a nonparametric statistical technique for multivariate data fitting, were investigated to predict the aerodynamic performance (output variables) of a generic 3D helicopter fuselage as functions of the pitch angle and of some geometric parameters describing their shape (input variables). In order for the smoothing thin plate splines to be properly applied, a database needed to be constructed containing pairs of input–output variables. To this purpose, a sample helicopter fuselage was chosen and 14 variants were generated modifying the geometric parameters; then, the pertinent lift, drag and pitching moment coefficients were obtained via computational fluid dynamics. The smoothing thin plate splines model was built excluding from the database one fuselage at a time and was then used to determine the aerodynamic performance of the left out configuration: finally, the obtained results were compared with those coming from direct computational fluid dynamics simulations over the same fuselage. The prediction capability of the smoothing thin plate splines models has been confirmed for all the analyzed fuselage geometries.
Condition-based maintenance is currently widely used in the aviation industry with diagnoses obtained from the performance data of the aircraft. Online assessments of the real-time condition and predicted residual life have been of great importance for both mechanics and pilots, especially during flight for the latter. Statistical distribution and feature parameters are believed to be crucial criteria of performance degradation, which facilitate making practical component replacement decisions. Furthermore in terms of observations featuring performance degradation, time-series analysis provides feasible forecasts of residual life from the available working time of aero parameters. The recorded data from constant speed generator drives of aircraft generally demonstrate these characteristics, are non-stationary and have time-varying variance in time-series analysis. The generalized autoregressive conditional heteroskedasticity approach is appropriate to the situation to obtain prediction results. The suitability of the proposed method has been examined through calculating prediction errors with data from an actual life experiment of aviation generator.
This article investigates the effect of asymmetric wing damage on flight dynamic characteristics of a flying-wing single motor unmanned air vehicle. To construct a six degree-of-freedom model of the damaged aircraft, a flying-wing type unmanned aerial vehicle is designed, and the wind tunnel test for damaged configurations is performed to identify the change of aerodynamic coefficients. The changes of mass, center of gravity, and moment of inertia are also calculated for each damage configuration with CATIA. The changed trim states are calculated depending on the severity of damage, and the movements of poles in longitudinal/lateral-directional flight modes are examined to evaluate the change of the dynamic stability and performance. Numerical simulations and eigenvalue analyses are performed to investigate the altered flight dynamics. It is verified that an asymmetrically wing-damaged unmanned air vehicle shows a sluggish roll behavior with longitudinal instability, and the result of this study can be a cornerstone for the future research on reconfigurable flight controller design against aircraft damage.
Particle image velocimetry data are presented from a scaled jet-lift aircraft ground vortex compressible flow. The scaled ground vortex is generated by a vertical compressible jet in cross-flow impinging on a moving ground plan. Particle image velocimetry is used to generate both transient and time-averaged flow statistics from the ground vortex region over a range of nozzle pressure ratios from 2.3 to 3.7, nozzle height-to-diameter ratios(h/dn) from 3 to 10 (where dn = 12.7 mm) and cross-flow velocities (V) from 10 to 20 m/s. These conditions correspond to effective (jet-to-cross flow) velocity ratios of 15Ve-1- 1<60. For each condition, mean and root mean square ground vortex core position was
analysed from sets of 72 instantaneous particle image velocimetry vector maps. Over the range of effective velocity ratios, Ve-1- 1, the particle image velocimetry results showed that the ground vortex mean streamwise position varied from 5dn to 16dn and the root mean square fluctuation in this position varied from 0.7dn to 1.5dn. Further analysis of the ground vortex temporal characteristics did not reveal any dominant non-dimensional frequencies.
An autonomous navigation system integrating both the path following and the autonomous sense & avoid functions is presented in this article. The sense & avoid algorithm was developed to provide an avoidance manoeuvre that ensures a minimum separation between the ownship and all other agents during its execution in a multiple flying threats scenario. The resolution manoeuvre is defined as step variations in the heading angle and altitude autopilots commands. The commands are optimised in order to get the smallest step command necessary to keep a minimum predefined separation between the ownship and the threats. Its computation is based on the estimation of the future trajectory of all the agents and, therefore, on the estimation of aircraft performance during the manoeuvre. The suggested resolution manoeuvre is updated at 1 Hz in order to take into account any unpredictable changes of the threat trajectories. The obtained heading and altitude change commands are displayed on a novel human–machine interface to support the pilot in the planning of the avoidance action. The proposed sense & avoid system is modelled in a Matlab/Simulink® environment for a Piper J3 Cub 40 model aircraft. The threats considered are aircrafts that communicate their states to the system through their Automatic Dependent Surveillance-Broadcast mode S transponders.
It is observed that a diffusion-controlled mechanism applied to the burning of micron-sized particles is not applicable to the combustion of nano-sized particles burning under kinetically controlled conditions. Furthermore, when heat transfer occurs between micron-sized particles and air, Nusselt number can be assumed to be constant and equal to 2, while this number is a function of Knudsen number when heat transfer occurs between nano-sized particles and air. Ignition temperatures of micron- and nano-sized particles are also different. In this article, mass and energy conservation equations for both particle and gas phases are solved. By doing so, flame velocity is obtained. Afterwards, with respect to different combustion characteristics of micron-and nano-sized particles such as ignition temperature, burning time, and Nusselt number, the effect of particle size on the flame velocity of aluminum particles combustion in air is studied and compared with experimental and numerical results. At the equivalence ratio of 0.85, it is shown that flame velocity is proportional to d-0.94 and d-0.56 for micron- and nano-sized aluminum particles, respectively.
Signal scaling is an essential step in spaceflight simulation. Thus far, the third-order polynomial scaling method has been widely used for signal scaling; however, in this method, parameter tuning is complicated and may induce perceptible distortion during large-range monotonic signal scaling. In the simulation of spacecraft return, specifically, that of re-entry, acceleration and angular velocity signals may vary considerably over short time periods. Motion perception is important for training astronauts in this phase. In this study, two strategies are proposed to solve these problems using the ‘scaling scope’ parameter. The first strategy is based on the Hermite interpolation polynomial, and the other is based on third-order polynomial scaling. Two methods were developed which make use of the stable region of third-order polynomial scaling. The first method maximizes the stable region to prevent signal distortion, and the other restricts the scaling scope in the stable region. Based on the dynamic characteristics of spacecraft in the return phase, the signal scaling strategies proposed in this study are simulated for trainees’ perception in a motion-base simulator. Simulations were implemented by utilizing the full curves of spacecraft return phase for the first time, and results show that these methods are more advantageous for parameter tuning and can eliminate signal distortion for all input signals. While these methods have a shortcoming in that the trigger velocity (onset cue) is slowed down, this shortcoming is eliminated by employing the moving cueing algorithm. Both the strategies proposed in this article show good performance and can be applied potentially to the motion simulation of the spacecraft return phase.
This article describes a new methodology allowing a combined exploration of experience feedback databases. Based on a small set of data provided by an airline, the study demonstrates the feasibility and the benefit for safety management of this new approach, which highlights links between human-factor components revealed by crew reports and operational deviations detected through digital flight data. Such a new understanding of the insight of the operations could have a major impact on safety management and should contribute to the proactive safety management culture that many airlines try to promote.
This article presents a new database of speech produced under cognitive load for the purpose of non-invasive psychological stress monitoring. The voices and the heart rates of eight airline pilots were recorded while completing an advanced flight simulation programme in a level D full flight simulator. Focusing on real-world applicability, the experiments were designed to yield the maximum degree of realism possible. Evaluation of physiological reference measures in pilots demonstrates that several heart rate variability parameters correlate with speech features derived from the recorded data. The article discusses the evolution of speech monitoring in aviation and proposes that application-orientated research methods can be useful in designing a system for real-world monitoring.
The problem of matching suitability for geomagnetic aided navigation is investigated from the viewpoint of pattern recognition in this article. In order to improve the classification accuracy of candidate matching areas, an intelligent classification method based on genetic algorithm and support vector machine is proposed. Firstly, the geomagnetic datasets and the factors influencing the classification performance of support vector machine are studied. Then support vector machine is employed as the classifier, and genetic algorithm is utilized for feature selection and support vector machine parameters optimization to improve the classification performance. Afterwards the multi-class support vector machine classifiers based on the one-against-one strategy are constructed for analyzing matching suitability. Experimental results show that the proposed method can greatly improve the classification accuracy of candidate matching areas, and moreover, the conclusions of this article can provide beneficial guidance for geomagnetic matching and route planning.
Two types of rotary valve were designed and fabricated for single-tube pulse detonation rocket engines. According to their driving components, they were named as gear-driven and cam-driven rotary valve, respectively. Preliminary experimental investigations were carried out to test their feasibility for supply of the single-tube pulse detonation rocket engine. Based on the gear-driven rotary valve, the single-tube pulse detonation rocket engine operated at 10 Hz stably. When operating frequency was increased to 15 Hz, inconsecutive detonations happened which was due to signal output block in the encoder. It was found that optimal ignition timing needed to be adjusted to obtain stable operation in the gear-driven rotary-valved pulse detonation rocket engine. Discussion on this was conducted. Operations of the cam-driven rotary-valved pulse detonation rocket engine were also performed. Fully developed and successive detonations at 40 Hz were successfully obtained although an explosion happened in initial tests. Experimental results showed that these two types of rotary valve were able to realize supply control of single-tube pulse detonation rocket engines effectively.
Disturbances induced from platform motion, gimbal imbalance and state couplings, may severely degrade the line-of-sight tracking accuracy of a two-axis gimbaled system and even induce instability. This article intends to present a stable approach to realize optimal disturbance attenuation for a yaw-pitch gimbaled system, in the presence of saturation nonlinearity and multiple disturbances. A dynamic model of the line-of-sight dynamics is formulated and linearized; feedback controllers are synthesized via linear matrix inequalities and convex optimization; state trajectories of the system before and after stabilization are compared to examine the effectiveness of the feedback approach. The simulation results show that the synthesized controllers are effective in stabilizing the system and realizing optimal disturbance attenuation.
A computational fluid dynamics method has been applied to simulate the exhaust gas flow and the additional thrust during a missile launching from concentric canister launcher. The unsteady, axisymmetric Reynolds-averaged Navier–Stokes equations with renormalization group k – turbulence model are numerically solved here. The dynamic mesh method is utilized to simulate the movement of the missile. Computational fluid dynamics results show that the additional thrust is an important thrust and fluctuates with the movement of the missile for launching from concentric canister launcher. The mechanism for producing and influencing the additional thrust is typically relevant to the choking states of exhaust gases at the inlet and outlet of the annular tube of concentric canister launcher, responding to the jet impinging on the bottom of the launcher, the approximate wall jet, and the friction effect in the tube.
The efficient designs of lubrication and heat transfer in an aeroengine bearing chamber require a better understanding of the complex air/oil two-phase flow in the chamber, which contains oil droplet deformation and motion, as well as droplet/wall interactions including wall impingement and deposition behavior. A modified droplet deformation model is proposed to describe the effect of deformation on the motion, and then a splash critical criterion also is established by means of energy conservation to estimate the impingement conditions of droplets. Using the above knowledge, in combination with a secondary droplet characteristic model predicting the outcome of droplet impact with wall, the droplet deformation, motion, and the associated transfer of mass and momentum are calculated in an aeroengine bearing chamber, and the effects of air mass flow rates and shaft speeds are subsequently discussed. This article may contribute to providing initial conditions to study further film flow behavior on the chamber housing.
The article addresses data-driven fault detection in commercial aircraft gas turbine engines in the framework of multi-sensor information fusion and symbolic dynamic filtering. The hierarchical decision and control structure, adopted in this article, involves construction of composite patterns, namely, atomic patterns extracted from single sensors, and relational patterns representing cross-dependence between a pair of sensors. While the underlying theories are presented along with necessary assumptions, the proposed method is validated on the NASA C-MAPSS simulation test bed of aircraft gas turbine engines; both single-fault and multiple-fault scenarios have been investigated. Since aircraft engines undergo natural degradation during the course of their normal operation, the issue of distinguishing between a fault and natural degradation is also addressed.
Assembling aircraft stiffened panels using friction stir welding offers potential to reduce fabrication time in comparison to current mechanical fastener assembly, making it economically feasible to select structurally desirable stiffener pitching and novel panel configurations. With such a departure from the traditional fabrication process, much research has been conducted on producing strong reliable welds, with less examination of the impact of welding process residual effects on panel structural behaviour and the development of appropriate design methods. This article significantly expands the available panel level compressive strength knowledge, demonstrating the strength potential of a welded aircraft panel with multiple lateral and longitudinal stiffener bays. An accompanying computational study has determined the most significant process residual effects that influence panel strength and the potential extent of panel degradation. The experimental results have also been used to validate a previously published design method, suggesting accurate predictions can be made if the conventional aerospace design methods are modified to acknowledge the welding altered panel properties.
A three-dimensional numerical simulation of a supersonic free-stream at Mach 2.5 over a spherical body with a sonic opposing jet from its stagnation point is carried out by solving the three-dimensional Navier–Stokes equations coupled with the standard k– turbulence model. It is aimed to investigate the effects of the jet on the drag reduction on the body and the flow field around the body. The influences of the jet pressure, the nozzle size of the jet, and the angle of attack are systematically studied for the purpose. An unsteady oscillatory motion mode and a nearly steady motion mode are identified depending upon the jet total pressure. There exists a critical jet pressure where the flow mode transition from one to the other happens suddenly and this critical pressure value varies approximately linearly with the jet nozzle exit size inversely. For the zero angle of attack, the results show that there exists a maximum overall drag reduction as the jet pressure changes for each jet nozzle size and the maximum overall drag reduction always happens at the unsteady oscillatory motion mode. The main shock in front of the body is pushed backward by the jet and the displacement of the shock decreases with the increase of the angle of attack, and the drag reduction efficiency also decreases with the angle. Regarding to the mode transition, it is found that the drag rises suddenly when the transition happens for the angle of attack smaller than or equal to 5° but it does not result in the rise for the angle larger than 5°. The results show that the maximum overall drag reduction can be reached as high as 32.6% for the cases studied. The present results provide useful information for drag reduction applications using an opposing jet.
A mixed H/H2 gain-scheduled state-feedback control method is developed for trajectory tracking of spacecraft rendezvous in elliptical orbits. Since the tracking accuracy is vulnerable to exogenous disturbances, the mixed H/H2 control, which takes into consideration both worst-case disturbance-attenuation performance and tracking performance, is particularly attractive for trajectory tracking of spacecraft rendezvous. Owing to the fact that the dynamic model for elliptical-orbit rendezvous is time varying, the feedback gain matrix is formulated as a matrix fraction of parameter-dependent matrix. Parameter-dependent Lyapunov functions are adopted to reduce conservatism caused by fixed quadratic Lyapunov matrices, and slack matrices are introduced to avoid setting a common Lyapunov matrix for different performances. Then, the desired controller can be obtained through a convex optimization with linear matrix inequality constraints. Computer simulations show that the proposed method can (a) handle trajectory tracking of elliptical-orbit rendezvous effectively; (b) provide a balanced performance between disturbance-attenuation performance and tracking performance; and (c) yield results that are less conservative than those obtained through conventional methods.
A methodology to predict the airflow hazard of helicopter flight in icing conditions is developed. By incorporating the existent ice accretion codes into an established basic helicopter flight dynamic model and considering airflow disturbance that mainly covers downdraft, head/tail wind, and left/right wind, the hazardous effects on trims, stability, and controllability of UH-60A single rotor helicopter in icing/ice-free conditions and within/without different types of wind field are investigated. The stability and controllability of helicopter that encounters airflow disturbance from wind velocity of 0 to 2.5–5.0 m/s for forward flight are examined. The indications of all the work are summarized at the end of this article. Furthermore, this method can be used to helicopter inflight safety prediction or airflow hazard avoidance analysis in icing conditions. It can also be laid as the foundation of the further research about the more complex airflow hazard prediction in icing conditions for helicopter flight safety.
The design of the autopilot is one of the most important algorithms of missiles. Performance of the autopilot and its robustness are significant matters to hit a target accurately. The autopilot should satisfy the desired performance under disturbances. In the scope of this study, three autopilots were offered for tracking pitch acceleration command using different control methods: three-loop classic control, pole-placement control and receding horizon predictive control. The aim of the autopilot designed by employing receding horizon predictive control is to minimize the flight control effort, and to make the close-loop system insensitive against modelling uncertainties and stochastic shattering factors. This study comes up with that the missile is able to move in desired performance under disturbances such as control surface misplacement, thrust misalignment, wind and aerodynamic uncertainties with more robustness, less control effort and minimum miss distance and terminal time using an alternative control method instead of classic and pole-placement control methods which are generally referred by the defence industry.
This article describes meta-synthesis information fusion as a novel mode in integrated avionic management systems. With deeper space explorations, avionic health management systems require more perfectly integrated data and information because of the increased spacecraft functionality and the complex software. Meta-synthesis information fusion allows for a more accurate picture of the state of the avionics and therefore allows for better decision making. This study uses a meta-synthesis information fusion application for the hybrid diagnostics, which is a very important part of integrated health management systems. For the meta-synthesis information fusion, specific approaches such as probability theory, neural networks and the Dempster–Shafer evidence theory are used to construct the hybrid diagnostics model. Through this novel meta-synthesis information fusion mode, efficiency as a whole, from input to output is realized and dynamic, real-time diagnostics is achieved. A numerical example is given, which demonstrates the application of the hybrid diagnostics to a radar indicator. By analyzing the feasibility and the pragmatic utility of the hybrid diagnostics meta-synthesis information fusion, the advantages of this mode are shown.
The nonlinear dynamic characteristics of rotating ramjet rotor supported by hybrid gas bearing are studied. The compression inlet flow field at different back pressure levels is analyzed and the normal working back pressure level is determined. The periodic movement phenomenon of normal shock wave in compression inlet is presented. The influence on the compression inlet flow field with the variation of structure dimension is introduced. Then, the nonlinear compression inlet flow force generated from the whirling of the rotor is obtained. The model for the rotating ramjet rotor supported by the hybrid gas bearing is established by the finite element method. The equation of motion for the rotating ramjet rotor is numerically solved and coupled with the gas lubricated Reynolds equation considering the time terms. The vibration characteristics of the rotating ramjet with different supply pressure and unbalanced mass eccentricities are solved by the Newmark method. The orbit trajectory diagram, frequency spectrum diagram, and time response diagram are obtained. Then, the stability of the rotating ramjet rotor system is discussed. The results indicate that the compression inlet is under the condition of high adverse pressure gradient, the shock wave, expansion wave, reflections and crossings of the shock waves, boundary layer–shock wave interference, and separation of the flow, which lead to the unstable flow of the compression inlet. The nonlinear compression inlet flow force can cause sub-synchronous vibration. If the supply pressure and eccentricities are properly designed, the vibration amplitudes can be decreased and the stability will be improved, which will make the foundation for the vibration control of the rotating ramjet system.
A two-dimensional wing aeroelastic model with a modified Leishman–Beddoes model in state-space form is performed to investigate the flutter behavior at low Mach numbers. The two-dimensional wing aeroelastic system employs highly aerodynamic non-linear features. The modifications to the Leishman–Beddoes model are validated by comparing with the experimental results of airfoil airloads at low Mach numbers, and the wing flutter experimental results also verify the model of the two-dimensional wing aeroelastic system. Results demonstrate that a fold bifurcation, which is related to the high-amplitude limit cycle oscillation, stems from the phenomenon of dynamic stall and the emergence of the period-2 limit cycle oscillation is due to the asymmetry of the aeroelastic system.
A dynamic model for the launching system incorporating the influence of the clamp band joint is developed using the finite element method, where both of the launch vehicle and the spacecraft are modeled as Timoshenko beams. The clamp band joint is represented by a massless beam element, of which the element stiffness matrix is developed based on the expressions for the axial and bending stiffnesses of the joint deduced by the authors in the previous works. Dynamic analyses are performed to evaluate the joint influence on the launching system, where the variations of the mass and length of the launching system due to the fuel combustion and stage jettisons during the ascent flight are considered. The dynamic model presented here can be applied to investigate dynamics of launching systems involving the influence of clamp band joints conveniently.
Collision between satellites and debris is a rare event but with high financial consequences. This risk therefore has to be addressed carefully. To support the decision to start a collision avoidance maneuver, a dedicated tool to characterize the risk uncertainty is the probability of collision between the debris and the satellite.<xref ref-type="bibr" rid="bibr1-0954410012467725">1</xref> Crude Monte Carlo could be a way if it could cope with very small probabilities, say 10-6, within the available simulation budget and time. The methodology nowadays in use is a numerical integration made tractable by physical hypothesis and numerical approximation.<xref ref-type="bibr" rid="bibr2-0954410012467725">2</xref> We advocate the adaptive splitting technique, presented in Cérou et al.,<xref ref-type="bibr" rid="bibr3-0954410012467725">3</xref> as it avoids all the hypothesis needed for the numerical integration and clearly outperforms Crude Monte Carlo with respect to rare events. A direct comparison between Crude Monte Carlo and adaptive splitting technique approach is also given on real-life examples.
The main contribution of this article is the proposal of a path-following method for fixed-wing unmanned aerial vehicles. This path-following method employs the multi-loop framework that consists of an outer guidance loop and an inner control loop. The guidance loop relies on the idea of tracking a virtual target. The virtual target is assumed to travel along the defined path and its speed is explicitly specified. This guidance law guarantees the asymptotic convergence to the desired path and can anticipate the transition of the flight path in advance, which reduces the command for the inner control loop. In the inner control loop, the flight control law based on dynamic surface control is derived to overcome the ‘explosion of complexity’ problem in the backstepping design. The numerical simulation result illustrates the effectiveness of the proposed method.
While linear covariance analysis is widely used for navigation system design and analysis, it is often overlooked as a tool for closed-loop guidance navigation and control (GN&C) system design and analysis. This article presents an overview of the techniques and methods required to develop a linear covariance analysis tool for a close-loop GN&C system. Then, using a simple nonlinear closed-loop GN&C problem as a guide, the capabilities of linear covariance analysis for the design and analysis of closed-loop systems are demonstrated. It is shown that linear covariance can be accurately applied to a closed-loop system with time-to-go guidance, dead-reckoning navigation, and a Kalman filter for state estimation. The accuracy and efficiency of linear covariance analysis is shown by direct comparison to Monte Carlo analysis results, and the value of linear covariance analysis is highlighted by presenting several analysis capabilities that are often required in the design and analysis of closed-loop GN&C systems. It is also shown how the efficiency of linear covariance enables new design methodologies, one of which is presented in this article, that would otherwise be prohibitive with Monte Carlo analysis.
The article mainly employed the numerical simulation method to conduct the 2D hypersonic inlet performance research. In this study, the numerical simulation method that includes a two-equation turbulence model was validated through the test case firstly. For the purpose of oppressing unstart phenomena, extending operating range and improving comprehensive performance of the inlet, the bleeding system and isentropic expansion arc surface were applied to the inlet, respectively. Different kinds of configurations were designed through changing the length of the bleeding or changing the radius of the arc surface. The influence of the bleeding and isentropic expansion arc surface on the 2D hypersonic inlet was analyzed systematically based on different objective parameters. The calculative results show that the bleeding system with appropriate value of the length can improve comprehensive performance of hypersonic inlet remarkably either on the design point or the off-design points, though it has a very small negative impact on the inlet in terms of the mass flow ratio coefficient. As for isentropic expansion arc surface, it can improve the inlet’s performance at the design point on the conditions that the radius of isentropic expansion arc surface has an appropriate value. But it has much more disadvantages over the basic configuration reversely when it comes to the off-design conditions or the large size of the radius of isentropic expansion arc surface.
The separation system is crucial for the launch of satellites. Dynamic characteristics of satellite separation are quite complex and difficult to predict. With respect to the helical compression spring mechanism, an approach using transient perturbation analysis is presented. The separation springs and limit switches are mathematically modeled, and disturbing forces and moments are considered. ADAMS and MATLAB software platforms are combined to obtain separation trajectory and attitude parameters. The minimum relative distance is proposed to show whether there is collision between the satellite and launch vehicles. Emphasis is placed on introducing the approach by analyzing a typical separation system. With experimental design and statistical analysis, the influences of perturbation factors are concluded. For example, three angular velocities are approximately linear with center of gravity offsets of the satellite and deviations of spring parameters; however, the effect law of asynchronous time is non-linear. A ground test system of satellite separation is designed and the test results are compared with the analysis, which prove accuracy of the dynamic model and feasibility of the approach.
Line-of-sight stabilization is an important concept for aerospace applications utilizing gimbaled imaging systems. A widely used method for protecting the line-of-sight stabilization system from the disturbing effects of the base vibrations is to mount it on passive vibration isolators. However, these isolators may interact with gimbal controller and drastically limit the stabilization performance. This work deals with line-of-sight stabilization problem in aerospace structures by focusing on the parameters of controller and vibration isolation system. The problem is investigated on an experimental setup for a specific case. Several performance tests are applied on the setup and a relation between isolation parameters and controller bandwidth is obtained. The results are used to generate design constraints.
Experimental study of a small partial admission axial turbine with low aspect ratio blade has been done. Tests were also performed with full admission stator replacing the partial one for the same rotor to assess the losses occurring due to partial admission. Further tests were conducted with stator admission area split into two and three sectors to study the effects of multiple admission sectors. The method of Ainley and Mathieson with suitable correction for aspect ratio in secondary losses, as proposed by Kacker and Okapuu, gives a good estimate of the efficiency. Estimates of partial admission losses are made and compared with experimentally observed values. The Suter and Traupel correlations for partial admission losses yielded reasonably accurate estimates of efficiency even for small turbines though limited to the region of design u/cis. Stenning’s original concept of expansion losses in a single sector is extended to include multiple sectors of opening. The computed efficiency debit due to each additional sector opened is compared with test values. The agreement is observed to be good. This verified Stenning’s original concept of expansion losses. When the expression developed on this extended concept is modified by a correction factor, the prediction of partial admission efficiencies is nearly as good as that of Suter and Traupel. Further, performance benefits accrue if the turbine is configured with increased aspect ratio at the expense of reduced partial admission.
The problem of active feedback control of fluid flows falls into a class of problems in the area of distributed parameter control, typically defined by partial-differential equations. Physical processes modeled by partial-differential equations are infinite-dimensional systems and are often simulated by numerical methods. However, for complex flows, the degrees of freedom may still be of the order of millions and are not practical for direct use in control design and optimization of fluid flow systems. Consequently, ‘reduce-then-control’ strategy is often employed for flow control of many engineering and industrial applications. In this study, we develop a linear quadratic regulator control to suppress fluctuating forces on a circular cylinder using a proper orthogonal decomposition based low-dimensional model. We numerically simulate the flow past a circular cylinder by solving the incompressible Navier–Stokes equations, and record the flow field data over one vortex shedding cycle. Using the data ensemble, we compute the proper orthogonal decomposition basis functions (modes) of the divergence-free velocity and pressure fields. We project the Navier–Stokes equations onto these proper orthogonal decomposition modes to develop a reduced-order model. Later, we modify the model by applying suction on the cylinder surface and adding a control function in the velocity expansion. The nonlinear dynamical system thus developed is linearly unstable due to negative damping in the system. We linearize the system about the mean velocity and apply optimal control. We seek to minimize the fluctuating forces on the cylinder using a reasonable amount of control effort. The novelty in this control strategy lies in feeding back only the dominant mode to suppress the fluctuating forces. On the contrary, feedback of higher modes fails to control and destabilizes the system.
This article aims to present an adaptive and robust cooperative visual localization solution based on stereo vision systems. With the proposed solution, a group of unmanned vehicles, either aerial or ground will be able to construct a large reliable map and localize themselves precisely in this map without any user intervention. For this cooperative localization and mapping problem, a robust nonlinear H filter is adapted to ensure robust pose estimation. In addition, a robust approach for feature extraction and matching based on an adaptive scale invariant feature transform stereo constrained algorithm is implemented to build a large consistent map. Finally, a validation of the solution proposed is presented and discussed using simulation and experimental data.
An experimental study has been conducted on the generation and propagation of compressible vortex rings using helium as a driver gas, with the aim of evaluating the effects of multi-gas operations for real-life applications. The advantage of such system, when compared to a constant gas system based on ambient air, is to effectively increase the Mach number while keeping the pressure ratio constant. Three pressure ratios of ~4, 8 and 12 were set, corresponding to experimental Mach numbers of approximately 1.50, 1.81 and 2.05. The increase in incident Mach number resulted in the variation of the vortex ring and trailing jet structure, and an increase in both the velocity magnitude and vorticity field. Results showed a transition from the regular-reflection shock-cell system at the experimental Mach number approximately 1.50 to the presence of a Mach reflection with a central Mach disc, which grew in size with further increase in incident Mach number. The presence of the Mach disc resulted in the formation of a subsonic jet, internal to the main trailing jet. Its velocity was measured to be in the order of magnitude of 550 m/s, with the speed of sound of helium at 1005 m/s. Results also demonstrated that shear layers formed between the subsonic and main trailing jet have opposing vorticity, with that of the subsonic jet being approximately half in magnitude. Secondary counter-rotating vortex rings were generated ahead of the main vortex and orbited around it. The analysis of the vorticity field has shown that these secondary vortices have a magnitude approximately half of that of the vorticity of the main vortex, and has confirmed that they have an opposite direction of rotation.
In this article, a semi-analytical approach is used to formulate the torsional analysis of beams having a crack. The formulation is based on analytical and accurate numerical analyses. The cross section is decomposed into several segments, including straight and curved segments with or without a crack. A dimensionless formulation is introduced and used to analyze the problem. The cracked segments are analyzed using the finite element method. The other segments are analytically formulated. The crack is supposed to be perpendicular to the boundary and located in the inner or outer surface of the beam. The thickness of the cross section is assumed to be uniform. Some numerical results are provided to show the accuracy of the presented method.
Recent investigations of trajectory options that incorporate solar sails have been motivated by missions to observe planetary poles or to communicate with an outpost at the lunar south pole. Designing reference trajectories and understanding their fundamental dynamics are the necessary first steps toward flying spacecraft in dynamically complicated regimes. However, the existence of a reference orbit alone is insufficient for flight operations. Two variations of a turn-and-hold strategy are examined for flight-path control: an approach that implements multiple turns to achieve a target in an error-free scenario and an approach that incorporates a look-ahead strategy to accommodate representative errors.
A novel, modified model-based fault-tolerant attitude tracking control scheme is derived for the uncertain flexible satellites with four reaction wheels. The stability conditions of this type of satellite are also analyzed. The attitude tracking error dynamics is employed to derive the novel control scheme, which is formulated in the presence of actuator fault uncertainties, moment-of-inertia uncertainties, flexible appendage dynamics uncertainties, space environmental disturbances, and reaction wheel dynamics. The uncertainties are estimated to update the computed torque, which extremely enhances the pointing accuracy and reduces the conservativeness. The large-angle attitude tracking and time-varying attitude trajectory can be stabilized by the novel controller, which solves the constant desired trajectory restriction in most of other control schemes. Numerical results are presented to verify the advantages of the proposed control scheme in comparison with proportional–derivative controller, model-based controller, time delay controller, and modified time delay controller.
A set of exploratory experiments are conducted to test a newly designed strut for fuel injection and flame holding in a liquid-kerosene-fueled dual-mode scramjet combustor. The thickness of the strut is 8 mm and the front blockage is about 8%. To organize stable combustion in a Mach number equal to 2.6 air flow under this thin strut using room-temperature liquid kerosene in a flash wall combustor without any cavity and other flame holders, some oxygen is injected through a set of orifices at the back of the strut, based on which a stable center local flame is generated at the back of the strut and the main flow combustion can be organized around this local flame. Experimental results show that stable combustion can be achieved at the center of the combustor with a wide range of equivalence ratio from 0.19 to 1 based on this center flame strut strategy. Through the analysis of the pressure distribution along the combustor, different combustor modes appear with different equivalence ratio. The article also gives some discussions about different influence of the oxygen to the combustion process under different equivalence ratio.
This article investigates the finite-time attitude stabilization problem by output feedback for a rigid spacecraft without angular velocity measurement. First of all, by employing the finite-time control technique, a finite-time stabilizing controller by state feedback is designed. Then, to address the problem of lack of angular velocity measurement, a semi-global finite-time convergent observer is proposed to recover the unknown angular velocity information in a finite time. Finally, a semi-global finite-time output feedback controller is developed. Rigorous proof shows that the attitude of rigid spacecraft will converge to the equilibrium in a finite time. A simulation example is given to demonstrate the efficiency of the proposed method.
Gravitational gliding during night without electric power is an efficient method to enhance the endurance of high-altitude solar-powered unmanned aerial vehicles. The properties of the maximum endurance path of gravitational gliding are studied in this article. The maximum endurance path problem is formulated as the problem of the maximum endurance can be sustained by unit altitude difference with the constraints of dynamic equations and aerodynamic parameters. The maximum endurance path is generated by Gauss pseudo-spectral method, and a new way to estimate co-state variables in Hamiltonian is proposed. In order to analyze the sensitivity of initial altitude and velocity of solar unmanned aerial vehicles with its endurance performance, the lift coefficient in the interval [0.4, 1.2] and flight envelopes between 0 and 30 km are investigated. The results are as follows: first, the broad range of lift coefficients can improve solar aircrafts’ long-endurance performance; second, the lower the initial altitude, the longer the gliding endurance can be sustained by unit altitude difference; third, it is possible for a solar-powered unmanned aerial vehicle to keep aloft during the whole night just by gravitational potential energy storage, but the issues with turbulence and wind would render this mode of unlimited endurance unfeasible. Thus, gravitational gliding by potential energy storage can only be partly used instead of electric energy storage in application now.
This article considers the attitude tracking problem of a rigid spacecraft involving inertia matrix uncertainty and external disturbance. The adaptive sliding mode control is utilized for the attitude controller design. The major concern is reducing the switching gain generated by current adaptive sliding mode control, thereby alleviating the chattering problem. By eliminating the influence of initial tracking error from the switching gain adaptation, an adaptive integral sliding mode control scheme is first presented. As compared with current adaptive sliding mode control, a much smaller switching gain is produced. Then, a disturbance observer-based adaptive integral sliding mode control design is proposed to further enhance the result. To this end, the joint effect caused by external disturbance and inertia matrix uncertainty, referred as lumped uncertainty, is divided into a slow varying part and a rapid varying part. By compensating the slow varying component via a disturbance observer, the switching gain is only required to be larger than the upper bound on the rapid varying component. The effectiveness of the proposed strategies, especially the switching gain reduction ability, is verified by both theoretical analysis and simulation results.
A nonlinear aircraft tracking filter using a point mass flight dynamics model with three degrees of freedom is presented. While the models used by conventional air traffic control tracking filters are based on simple kinematics, the model for the present filter is based not only on kinematic relations but also on three-dimensional aircraft translational force equations and control variables. This allows for practical and sophisticated implementation of the attitude effects on translational acceleration. The control variables, which consist of the angle of attack, roll angle, and thrust setting, are treated as states with random processes. Tracking with simulation data indicates that the present filter is superior to other single and multiple model-based filters in terms of position and course accuracy, and the model associated with it is insensitive to flight motion types and design parameters. The results of tracking with real flight data also correspond well with those found by tracking with the simulation data.
This article presents the novel results obtained using variance-constrained controllers and maneuvering helicopters also when some helicopter sensors fail. For this purpose, complex, control oriented, and physics-based helicopter models are used. A nonlinear model of the helicopter, which includes blade flexibility, is first linearized around specific maneuvering flight conditions (i.e. level banked turn and helical turn). The resulting linearized models are used for the design of variance-constrained controllers (i.e. output and input variance-constrained controllers). Then, the robustness of the closed-loop systems with respect to modeling uncertainties (i.e. flight conditions and helicopter inertial parameters variations) is studied. Next, variance-constrained controllers are designed for these maneuvering helicopter models when some helicopter sensors fail. Several sensor failure cases are examined and robustness properties of the closed-loop systems with respect to modeling uncertainties are also examined. Limitations of the control design process due to the number and type of failed sensors are investigated as well. Finally, the possibility to adaptively switch between controllers in order to mitigate sensor failure is studied.
An interdisciplinary study on the SpaceLiner orbiter is conducted focusing on the aspect of trimming and its impact on aerodynamical performance and structural design of the wing. Part of this study is to update the previously optimized aerodynamic shape taking into account the aspect of trim capability while maximizing the aerodynamic efficiency. This highly automated process results in a new preliminary definition of the aerodynamic shape of the SpaceLiner orbiter. A dimensioning abort case trajectory is defined following the recent update of the nominal return trajectory. Both serve as a starting point to the investigation of the trim requirement. The study confirms that the updated aerodynamic shape requires only small flap deflection angles for trimming, even under degraded conditions, and thus limits the impact on the lift-to-drag ratio. A structural analysis on the wing attests the importance of limiting the dimensioning flap deflection angles as the forces exercised by the flaps on the hinges can be considerable.
The active cooling with endothermic hydrocarbon fuel is proved to be a very effective approach for scramjet thermal protection. This article has summarized the research status of active cooling of endothermic hydrocarbon fueled scramjet engine from the following five aspects, cooling capacity and heat sink measurement, thermal and catalytic cracking, coking suppression, heat transfer characteristics, and injection, mixing, ignition, and combustion performance, and suggestions on the further study of active cooling of endothermic hydrocarbon fueled scramjets are put forward. The total heat sink of endothermic hydrocarbon fuel is found to be sufficient for scramjet cooling as the additional chemical heat sink is generated from cracking reactions, and catalytic cracking is proved to be better, because of its low cracking starting temperature, high conversion percentage, and good selectivity. As for coking mitigation, approaches are usually made from eliminating the oxygen dissolved in the fuel, reducing the amount of coking-foregoing substances, and doing some special treatments on the metal surface. The heat transfer of endothermic hydrocarbon fuels presents two enhancements. The first one is related to the variation of parameters near the critical point while the second one is a result of cracking reactions. The cracked products are proved to have better performance of injection, mixing, ignition, and combustion, and this is especially attractive for scramjets as effective mixing and stable ignition and combustion has always been difficult.
The mathematical models and a numerical code for numerical simulation of a thermal anti-icing system are presented in this article. Mass conservation is applied to the runback water flow. An energy balance is imposed on the airfoil skin including the water flow. The heat transfer coefficient distributions are obtained using the boundary layer integral method. The external flowfield and the local water collection efficiency data are predicted using an Eulerian method, based on a computation fluid dynamic code and its user-defined functions. Given the input of the electrical power density distribution, the numerical code is able to calculate airfoil equilibrium surface temperature, mass flux of runback water, runback ice mass flux, and range if happens. A user interface is developed to integrate the computation fluid dynamic code to achieve a method for the analysis of a thermal anti-icing system. All the numerical results are compared with both experimental data and other numerical results presented in the literature.
Current rapid growth of the aviation transport sector is deemed to be unsustainable because of the large quantities of greenhouse gas the aircrafts emit at sensitive altitudes. To prevent growth limits on this sector being imposed, the airlines must become cleaner. In the present study, a sustainable engine concept – based on intercooled recuperated turbofan, fuelled by liquid hydrogen, and using water injection for take-off and climb-out – is developed. This concept is intended to provide airlines a clean, efficient alternative to conventional engines. A commercially available computer program is used for modeling the sustainable turbofan concept. The optimization scheme of Guha has been applied for the new engine concept for minimizing fuel consumption and the performance of the sustainable turbofan engine in the possible design space has been determined. Comparisons with an existing conventional engine (of same thrust) revealed the significant improvement in overall efficiency (44%), reduction of emissions surpassing the demanding ACARE 2020 goals, improved longevity of the engine, and simplification in turbo-machinery components trying to offset the increase in weight due to the heat exchangers.
The aim of this article is to use an accurate truncation error estimate in order to perform -extrapolation and mesh adaptation in an unstructured finite volume computational fluid dynamics solver, in the context of a posteriori error estimation. The truncation error is approximated by the so-called -estimation technique, in which a special criterion is defined in order to account for the finite volume discretisation. It is shown that an accurate truncation error evaluation can be obtained on arbitrary geometries as long as restriction of the solution from the fine-to-coarse grid is accurate and the coarse grid possesses the same quality requirements as the fine grid. The accuracy of the truncation error estimation is successfully verified on Euler and Reynolds-averaged Navier–Stokes equations using the method of manufactured solutions. Then, mesh adaptation is performed on aerodynamic configurations where a good improvement of the force coefficients with respect to a classic feature-based indicator is obtained, at a lower cost than performing global refinement.
An N-impulse control scheme for spacecraft formations in elliptical orbits is developed to regulate the differential elements of the deputy spacecraft in the presence of the J2 perturbation. The presented control scheme is an extension of an existing circular-orbit formation control scheme and is shown to perform well at large eccentricities where the circular control scheme fails. For the case of two impulses being applied at arbitrary firing times, a discrete-time approach for ascertaining the stability of the controlled spacecraft formation, under the assumption of small impulsive thrusts, is presented. It is found that stability is guaranteed for the majority of firing time pairs; however, the requisite V can be prohibitive for some firing time pairs. The control scheme and stability predictions for formations in high eccentricity orbits are validated in numerical simulation.
This article investigates the attitude synchronization problem of spacecraft formations. A class of continuous sliding mode control schemes with finite-time convergent property is developed. The control laws are designed by the utilization of behavior-based control approach. An improved version of terminal sliding mode is applied in both the reaching phase and the sliding phase of the control system. In the presence of external disturbances, the proposed control strategies are able to overcome the unexpected phenomenon and can steer the spacecraft formation to a dynamic reference attitude state coordinately subject to arbitrary communication topologies. Numerical simulations are provided to validate the theoretical analysis.
The separation dynamics of a large-scale fairing section in ground test is investigated numerically using a fluid–structure interaction method. The commercial finite element software MSC/Dytran is adopted to establish the dynamic fluid–structure coupling model of the fairing. Two coupling surfaces are constructed for the inner and outer surfaces of the fairing section. The coupling equations are solved using the sequenced-coupling method, in which the fluid and structural problems are examined by the finite volume method and the finite element method, respectively. A comparison between fluid–structure interaction and dynamical response analysis is performed under the conditions with and without atmosphere effect. Results shown that the consideration of atmosphere effect will attenuate the vibration frequency and slow down the center of mass velocity. The effect of aerodynamic interference on the displacement response indicates that a maximum of 13.3% relative displacement can be induced, which may cause collision between the lower trailing portion of fairing section and the core vehicle. Therefore, it can be concluded that the fluid–structure interaction analysis is essential for evaluating and validating the reliability of separation mechanisms in ground tests.
In this article, a novel deployable mechanism that can be depaloyed from a bundle compact configuration onto a large volume double-layer truss structure is proposed. The mechanism is constructed by a set of Myard linkages through specially designed mechanical connections, so that the whole assembled mechanism has single degree of freedom. The model of the multi-objective design for the proposed deployable mechanism is developed. In the optimal design of this mechanism, many design objectives have to be taken into consideration, such as weight, stiffness, packaging/expansion ratio and natural frequency, etc. Many of these design objectives have no explicit analytical expression and may be contradicted with each other. A randomized multi-objective search algorithm is proposed for solving this multi-objective design problem, by using the algorithm, the set of Pareto optimal solutions can be obtained, and the relationship between different objectives is figured out, so that the designers can choose the compromise solutions intuitively. The physical prototype is also fabricated based on the optimized parameters, the stiffness and natural frequency experiments are conducted to evaluate the design. The experimental results demonstrate that the proposed mechanism offers an attractive combination of performance characteristics for both stiffness and natural frequency.
Cost estimation plays an essential role in the development of aerospace systems that are perhaps the most complex, time- and labor-consuming ones. Regarding this matter, it is unavoidable to take a systematic approach to build a realistic model through a deliberative, heuristical and easy-to-do process in the early stages of design. In the current study, complexity index theory is utilized to develop a heuristic complexity-based method to estimate various costs of aerospace systems. This method promises to be logically and practically more reliable and accurate than classical parametric methods. Logically, manipulating a group of parameters, instead of only one or two, reduces the probability of misrepresentation of systems and in the case of incompleteness of input data, reserves the chance for guessing them. Practically, all operations in this method are linear which makes it possible to work with matrices. With its organized and discrete nature, simulated annealing as a heuristic tool is employed to offset undesirable effects of imprecise initial assumptions. This helps to adjust complexity coefficients to more realistic magnitudes, when deriving a specific model from the heuristic complexity-based method. These coefficients may be used to estimate the cost of a new system as well as for sensitivity analysis. As a test scenario, estimation of an acquisition cost of a newly-developed unmanned aerial vehicle is concerned. Sensitivity of the complexity index to a number of complexity inducer parameters is also examined in order to achieve the most affecting parameters. Comparing by previously published results, it is seen that the current model is a remarkably accurate estimator for the acquisition cost of aerospace systems. This model shows a better R2 value, as a statistical measure of regression quality, than an already existing successful model by Technomics Corporation, regarded as a pioneer in this field.
Aiming at increasing the convergence rate and the accuracy simultaneously, an hp-adaptive Radau pseudospectral method is presented to generate a re-entry launch vehicle’s optimal re-entry trajectory. The method determines the number of mesh intervals, the width of the each mesh interval, and the degree of the polynomial in each mesh interval iteratively until a user-specified error tolerance is satisfied. In regions of relatively high curvature, convergence is achieved by dividing a segment into more mesh intervals, while in regions of relatively low curvature, convergence is achieved by increasing the degree of the approximating polynomial within a mesh interval. Simulation results show that the optimized trajectory obtained by the method satisfies the path constraints and the boundary constraints successfully. Moreover, the hp-adaptive Radau pseudospectral method is shown to be more efficient than either a global Radau pseudospectral method or a fixed-low-order Radau pseudospectral method. The results indicate that the hp-adaptive Radau pseudospectral method can be applied for real-time trajectory generation due to its high efficiency and high precision.
In recent years, significant resources have been invested to further improve the efficiency and environmental sustainability of modern aircraft. A possible strategy consists of reducing the induced-drag contribution (40% of total drag) by means of wing tip devices, e.g. winglets. However, these solutions have a negative impact on structural sizing, requiring reinforcements, and aeroelastic stability, requiring mass balancing. The subject of this study is the numerical study of an alternative wing tip device. In particular, two different design concepts are presented, namely discrete and raked options. These solutions improve the aerodynamic efficiency by extending the wing span and feature an integrated aeroelastic passive load alleviation capability. The design of the wing tip devices follows a multi-fidelity approach, closely matching today's best practices in the aerospace industry. In the first part of the study, the design phase is carried out with low-fidelity very efficient tools. In the second part, the most promising solutions are verified with high-fidelity more expensive tools, within the framework of computational aeroelasticity.
Avionics Full DupleX Switched Ethernet is a deterministic communication protocol for distributed embedded real-time applications over asynchronous channels. It is a promising technique that can replace the existing avionics data buses, such as ARINC429 and MIL-STD-1553B. One of the key challenges for deploying and maintaining Avionics Full DupleX Switched Ethernet is to determine the end-to-end transmission delay in such a network. This article aims at handling this challenge effectively applying a completely theoretical analysis. An analytical method based on the network calculus theory is presented in detail to calculate the end-to-end transmission delay in an Avionics Full DupleX Switched Ethernet network, which consists of many interconnect nodes with different scheduling disciplines. Further, this approach is improved by taking into account the effects of source node initial jitter and first in, first out optimization. Additionally, a Matlab/TrueTime simulation platform is constructed to verify the effectiveness of the method. Simulation results show that switches with advanced scheduling algorithms, i.e. static priority scheduling, can significantly improve the delay performance in a multi-hop network.
This article seeks to outline novel simplified model and analytical solution for predicting geometrical shape and film tension of high-altitude balloon. Sphero-conical and ellipsoid-conical models were proposed to depict geometrical configuration of high-altitude balloon subjected to a payload. By considering the effect of atmospheric factors and lifting gas temperature on geometrical shape of balloon, geometrical parameters of equilibrium shape were solved, based on minimum potential energy principle satisfying material constraint. New analytical solution was derived to allow balloon film tensions in meridional and circumferential directions. Finally, new model and its solution were used for predicting geometrical shape and film tension of natural-shape balloon on the ground and on float in stratosphere respectively, demonstrating the practical and effective use of the proposed model.
Launch vehicle design is a complex problem involving a series of disciplines. These disciplines present conflicting objectives and require multidisciplinary design optimization methods in order to handle the couplings and to make the search of compromises easier. Launch vehicle design problem is a specific multidisciplinary design optimization problem because it combines the optimizations of design and trajectory variables. In this article, we present a new multidisciplinary design optimization approach, called the StageWise decomposition for Optimal Rocket Design (SWORD). This method splits up the multidisciplinary design optimization process into different stages and transforms the initial multidisciplinary design optimization problem into the coordination of elementary ones. This method is compared to the standard multidisciplinary design optimization method (multidiscipline feasible method). In this article, we propose a new formulation of the SWORD method and a new dedicated optimization strategy. Results show that using a global search algorithm, the stagewise decomposition for optimal rocket design method allows to find a better optimum than multidiscipline feasible method. Furthermore, with the new proposed optimization strategy, the SWORD method does not require any initialization from the user, allows to quickly find feasible solutions and converge to an optimum in a limited computation time.
This article presents a study on the flow field and aerodynamic characteristics of the propellers and rudders for the design of a small ducted fan aircraft. In the analysis, an equivalent actuation disk was created as a simplified model of the propellers. A momentum source is determined by equivalent actuation disk numerical simulation using ANSYS CFX software based on vortex theory. In addition, the momentum source distribution function for the equivalent actuation disk model is improved by using experimental data. The computational fluid dynamics method in the ANSYS CFX software is then used to simulate the flow field of the propellers slipstream and determine the key aerodynamic and design parameters for the built-up rudders of the aircraft. The improved equivalent actuation disk method is validated by using the mixing plane method and their results show a good comparison. The simulation and analysis results provide the basis for optimal design of the built-up rudder arrangement and layout for a small ducted fan aircraft.
Due to the strong empirical demand of setting the design space variation in optimization design process which strongly affect the design result and efficiency, the adaptive aerodynamic optimization design method based on design variables space reconstruction concept has been established by the spatial statistical analysis of the design variables spatial distribution in the optimization process, which resolves the issue of design variables spatial distribution selection and lead to a better flexibility and convergence capability of the optimization model. As the design variables distribution has been statistically analyzed in the optimization process with the concept of clustering level, on the one hand the design variables spatial variation is reconstructed, on the other hand the design variables of a part of samples is adjusted in the reconstructed design space, both of which result in a better population diversity and a faster convergence with reservation of good genetic information. The NACA 0012 airfoil and NLF(1) 0416 airfoil have been optimized by this method, the results of which have been compared and analyzed with that of optimization of fixed design space variation. The design method approached has been proved of good feasibility with the optimization results, which results in an optimum solution in a larger variables distribution scale and shows better optimization efficiency.