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

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.

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.

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.

A precision landing guidance design for the Mars powered descent phase is proposed based on model predictive control (MPC) approach. Dynamics model used for the formulation are convexificated and linearized to adopt the convex optimization technique, which has been suggested by researchers of Jet Propulsion Laboratory. To employ the receding horizon frame and reduce the number of control inputs, the convex optimization problem is augmented with Laguerre functions. To represent the minimum fuel consumption or minimum landing error precisely unlike the optimal control theory, new cost function is designed by combining them with weighting factors. Moreover, the stability of the proposed guidance design for the independent control inputs calculated from each time step is verified by using Lyapunov stability analysis. Finally, numerical simulations are conducted to examine the suggested guidance formulation and to compare the performance with an optimal solution.

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.

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.

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.

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).

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.