The control problem of characteristic model with unknown input delay is studied. Using the method of directly solving the roots of the closedloop system, the results show that the minimumvariance controller is not robust to unknown input delay, the prototype goldensection controller can deal with small input delay, and the λ type goldensection controller can deal with larger input delay. Considering a stable, oscillating, unstable twointegrator characteristic model, the stability and robustness performance of the closedloop system are further analyzed respectively for the prototype and the λ type goldensection controller with unknown input delay. For stable controlled plant, the λ type goldensection controller is tolerable with large input delay, and the closedloop system performance is basically unchanged. For unstable plant, the λ type goldensection controller can improve the stability and performance by optimizing λ.

This paper studies the star sensor low frequency error (LFE) inflight calibration problem. The star sensor LFE, which is caused primarily by the periodic thermal distortion, has a great impact on the satellite attitude determination accuracy. To cope with this problem, we propose a novel LFE calibration method. An augmented Kalman filter (AKF) is adopted to estimate the satellite attitude and the LFE parameters simultaneously. It is specified that the attitude determination accuracy can be improved by using the proposed LFE calibration method. The star sensor LFE model for numerical simulation is established based on the telemetry data from real star sensors operating under inorbit conditions. The simulation results illustrate the efficiency of the proposed method.

To solve the problem in recovering the normal attitude from high spin speed for a threeaxis stabilized satellite under actuator failure, underactuated methods of despinning and precession control are proposed. The selection principle of thrusters is introduced, as well as the despinning control via integral multiple spin period duration firing mode and symmetry phase littlepulse firing mode. The precession control based on pulse modulation is also proposed. Moreover, the implementation steps and ways in engineering are presented. The numerical simulation and onorbit results indicate that the methods are efficient and practicable in engineering application. Besides, the methods can be used in both onboard autonomous control and remote control.

The highresolutionimage satellite is required to achieve stringent attitude determination accuracy (2.4″ (3σ) groundbased) in order to provide precise geometric accuracy. Influenced by the star sensor’s low frequency error (LFE), the conventional star sensors/gyros attitude determination system cannot achieve the precision requirement. For this reason, a new method for accurate attitude estimation with gyrobased reference time is proposed, and the satellite attitude estimates at the reference time is obtained on the ground by using the ground control points. The most relative time for achieving the required precision is limited because the effect of gyro’s bias errors and drifts is very significant in long term. Analysis and simulation results are also presented.

A new analytic linear covariance method for powered explicit guidance（PEG）is proposed, and the method is capable of obtaining the explicit expressions of PEG’s three parameters. The method is applied to analyze the final altitude dispersion, velocity magnitude dispersion and flight path angle dispersion with the existence of original state errors and parameter uncertainties in the lunar powered ascent phase. In contrast with Monte Carlo simulation results, the method is proved to be valid.

A mathematical timing modeling and analysis method for control system of satellites are proposed. The importance of mathematical timing modeling is illuminated by the analysis of ESA and NASA’s methods and tools in timing model. To design and analyze the timing, a mathematical timing modeling method is introduced in detail, which includes decompounding model, distributing model, and margin model. The describing tool based on workflow is introduced, by which the parameters can be obtained to analyze the timing and indicate the timing testing.

A finite element simulation modeling and flexible structural dynamics analysis of solar sail spacecraft are investigated in this paper. According to the actual structural characteristics of solar sail spacecraft, a reasonable finite element model is obtained. The reasonable value and direction of preload are given. The buckling modal analysis of stretch arm and the critical buckling load are obtained. The modals without and with preload are analyzed respectively and compared. The modal with preload is demonstrated more reasonable. The modal analysis with preload provides basic data for the design and simulation of solar sail spacecraft control system.

The onboard computer(OBC) that controls a manned spacecraft must be extremely reliable．Such a computer system generally has the ability of Byzantine fault resilience．This paper investigates the reliability of the quadmodular redundancy Byzantine fault tolerant computer. By using a continuoustime discretestate Markov model, the occupancy probabilities of the system are calculated. Several probabilities pertinent to reliability are recorded. The proposed approach shows that which redundancy scheme we adopt should depend on the relationship between Punsafe and Pshutdown, and provides reference for the design of the quadmodular redundancy Byzantine fault tolerant computer.

The onboard computer(OBC) that controls a manned spacecraft must be extremely reliable．Such a computer system generally has the ability of Byzantine fault resilience．This paper investigates the reliability of the quadmodular redundancy Byzantine fault tolerant computer. By using a continuoustime discretestate Markov model, the occupancy probabilities of the system are calculated. Several probabilities pertinent to reliability are recorded. The proposed approach shows that which redundancy scheme we adopt should depend on the relationship between Punsafe and Pshutdown, and provides reference for the design of the quadmodular redundancy Byzantine fault tolerant computer.

Attitude control/momentum management (ACMM) aims to balance the disturbance through active attitude adjustment, which can significantly reduce the fuel consumption for control moment gyro (CMG) momentum desaturation. The adaptive momentum management controller based on parameter identification developed in this paper consists of an online identification loop and a feedback linearization loop. In the feedback linearization loop, the ACMM system is transformed into an equivalent linear system through output transformation and state transformation. Nonlinear control laws for the original system are obtained through the design of the linear controller . The online identification loop identifies the system moment of inertia based on the closed loop control information, which provides feedback linearization with enough robustness to parameter variations of the system. Simulation results based on the space station module transfer operation demonstrate excellent performance even when the deviation between the torque equilibrium attitude (TEA) and the local vertical local horizontal orientation is significant.

Star tracker is a type of optical sensor current used widely in spacecraft attitude control system, which provides the most accurate measurement. We study the direct sun irradiation problem of the APS star tracker on orbit, including its effect on the function and performance of star tracker. The experiment and validation schemes are designed. And a clear conclusion is obtained according to the experiment process and verification results. The analysis and experiment results indicate that the ultraviolet irradiation has no significant effect on optical transmittance, and is in the designed allowance range. The direct sun irradiation has no effect on the function and performance of the APS image sensor and star tracker. The analysis and experiment results can be used as the design basis and reference for star tracker subsequent improvement.  

Considered the characteristic of the diagnosability index allocation, a new method is presented. Firstly the parameters and factors which affect the diagnosability index allocation are analyzed. Then the variables valued in different dimensions and ranges are unitized, and the allocation weight is obtained. Moreover, the method of diagnosability index allocation is presented. The method is showed to be reasonable and simple via an illustration.