Proceedings of the Institution of Mechanical Engineers,Part G: Journal of Aerospace Engineering

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 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.

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.

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 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.

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.

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.

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 10 - 5  g, the results have shown that the root of mean square error of the estimated center of mass is less than 10 mm given the location error and the angular misalignment of accelerometers are less than 5 mm and 0 . 5 °, 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.

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.

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.

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.

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.