This paper is concerned with the mixed H2 and H∞ adaptive control for hypersonic vehicles during reentry phase. The dynamics of the angle of attack in the longitudinal plane is modeled by using a secondorder difference equation known as the characteristic model. In order to improve the transient performance in the presence of large aerodynamic coefficient variations and unknown exogenous disturbances, a more intelligent goldensection robust adaptive control law with the similar structure as the characteristic model with goldensection parameters, is proposed based on our previous work. In addition to easy implementation and guarantee of stability before the convergence of parameter estimation, this new adaptive control law contains an adjustable parameter λ(k) that guarantees H2 and H∞ robustness of the system, and is calculated based on the linear matrix inequality approach. The integration of adaptive control and robust control provides the attitude control system for hypersonic vehicles with the capability to accommodate large flight conditions in a reliable way. Simulations are carried out on a hypersonic vehicle

A novel approach to optimal performance guaranteed control of spacecraft large angle attitude maneuvers is proposed based on sum of squares (SOS) optimization, spacecraft attitude dynamics is represented by a linear parameter varying (LPV) system with modified Rodrigues parameters. Based on the analysis of system performance, a sufficient condition for the existence of state feedback performance guaranteed controller is derived by using Lyapunov approach and SOS optimization. Furthermore, the optimal performance guaranteed controller design problem is converted to a parameter optimization problem with SOS constraints. Simulation results of certain satellite attitude maneuver demonstrate the effectiveness of the proposed design approach.

There are deviations between the Gaussian energy distribution model and the actual distribution of the star spots. In order to solve this problem, the rationality of the Gaussian energy distribution model with the simulation verification is reasoned in this paper. Gaussianbottomcutting energy distribution model is established aiming at the denoising in star tracker preprocessing，and then Gaussiantipbottomcutting is established according to the actual saturation in sampling. And the effect of the two models on the precision of the star centroid extraction are examined and compared accordingly through the simulation analysis. This can provide reference for improving the precision of the star spot centroid extraction.

An autonomous orbit control method using Bplane parameters as targeting parameters is proposed during the phase of approaching the deep space celestial body. First, the navigation system determines the current position and velocity of the spacecraft, which are mapped onto the Bplane to get the deviation from the desired encounter condition by integrating the dynamic model, and then the velocity increment for the orbital correction is derived using linear guidance formula and the error ellipse is utilized to describe the Bplane error. Finally, Monte Carlo simulation results show that the method is capable of guiding the Mars probe to a successful encounter.

Focusing on reverse pneumatic pressurereducing valve (PRV), we develop a dynamic mathematical model in this paper. The PRV dynamic characteristic of working process is simulated and the different parameters that influence the PRV’s response characteristic are analyzed. Then, by adopting linearization technique, the state space model (SSM) of PRV is established. The effects of main parameters on valve stability are obtained. The method that can improve the valve response characteristic and stability performance is also presented.

Considering the requirement of satellite fast maneuver, we propose a strategy of hybrid configuration about momentum wheel in this paper. To ensure the satellite momentum to be zero under fault reconstruction, a nominalangularmomentum allocation method and the corresponding moment distribution matrix are introduced. By the designed algorithm, the biggest moment of singleaxis is obtained under reconfiguration. Numerical results show that the new algorithm is effective in reducing adverse effects by the saturation of small momentum wheel, and hence enhances the control ability of system.

The coupling of satellite attitude motion and antenna flexible vibration has more influence on the antenna slewing, which is examined in this paper. First of all， the flexible boom is simplified as an equivalent cantilever beam structure, and the multibody coupling dynamics is established based on the Lagrange equation. Then the impacts of antenna slewing motion on attitude stabilization and flexible vibration are analyzed according to the coupling dynamic model. Based on the analysis results, the command preprocessing algorithm is introduced to plan the antenna slewing process. Finally, simulation analysis is performed for the antenna slewing motion. The simulation results show that the command preprocessing algorithm can effectively reduce the influence of the slewing on attitude stabilization and flexible vibration.

This paper presents a linear approximation method for the steadystate thermal vacuum test system model, and a kind of algorithm for the online joint estimation of time delay and parameters in a class of time varying systems. An adaptive poles assignment controller and compensator are designed with estimated time delay and plant parameters. The system response performance for a given temperature trajectory in tests is greatly improved.

Reconfiguration technique is developed for advancing the reliability of unified propulsion subsystem (UPS) of navigation satellite in failure mode. The UPS can be recombined through this technique, and the damage to 10N thrusters can be isolated. This report describes the principle of the reconfiguration technique in several commonly used modes. The result can provide a reference for the followup design and fault counterplan for the UPS of navigation satellite.

As to the deficiency of analog rebalance loops in terms of digital output and applications of modern control technology, the paper presents a way to design and implement digital rebalance loop for strapdown gyroscopes. Firstly, this paper introduces three types of realizations of the rebalance loop of strapdown gyroscopes, and proposes a new way to design and implement the rebalance loop. Secondly, we set up the mathematical model for the entire loop. According to the prevalent problem of the quantization error in the digital loop, we adopt the golden section based on characteristic model to design the controller. Finally, we simulate the mathematical model of the loop, and validate the performance of the rebalance loop. Simulation results demonstrate that the way of the loop implementation proposed in this paper can improve effectively the loop performance.

Bearing assembly is one of the most important and key parts in fly wheels. The value and the variation of the friction torque of the bearing assembly have direct influence on the performances of fly wheels. In this paper, the startup friction torque and the lowspeed friction torque are measured by using the developed lowspeed friction torque tester, and the theoretical analysis is also presented. The results from the experiments show that the bearing assembly with a larger startup friction torque will lead to a larger current at the working speed, but the two parameters have no strict proportional relationship. With the increasing preload, the startup friction torque of the bearing assembly increases linearly in the experimental range. With the speed increasing, there are two minimum points of the friction torque which cannot be predicted theoretically. Nevertheless, with the same speed, the friction torque of the bearing assembly under an increasing speed is larger than that under a declining speed.

The deepspaceexploration return vehicles reenter the earth under bad conditions. To assure the safety and accuracy of the return, both the LQR referencetrajectory and the predictorcorrector guidance approaches are researched in this paper. The adaptability to the reentry initial conditions and the landing accuracy of both guidance methods are analyzed. The numerical simulation is done for the return vehicles with the second cosmic velocity. The results show that the predictorcorrector guidance gets smaller range error than the LQR one with a relatively large initial error.