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Summary: This paper proposes a reinforcement learning‐based guidance law for Mars powered descent phase, which is an effective online calculation method that handles the nonlinearity caused by the mass variation and avoids collisions. The reinforcement learning method is designed to solve the constrained nonlinear optimization problem by using a critic neural network. Specifically, to cope with the position constraint (i.e., glide‐slope constraint) and the thrust force limit constraint, a modified cost function is proposed, and the associated Hamilton‐Jacobi‐Bellman equation is solved online without using an actor neural network, which significantly reduces the computational burden. The convergence of the critic neural network is proven. Simulation results show the effectiveness of the proposed method. [ABSTRACT FROM AUTHOR]

The planetary powered descent guidance problem for step-shaped hazardous terrains is investigated in this article based on prescribed performance function (PPF) methodology. Initially, the distances between the lander and step-shaped terrains around the landing site have been formulated in a new form boundaryfunction using PPF, in which a new step-shaped boundary PPF is specifically designed to constrain the lateral motion. Furthermore, a fixed-time convergent PPF is chosen to coordinate the vertical motion. Next, to avoid the collision with step-shaped terrains and planetary surface, a feedback guidance algorithm is proposed based on the backstepping method. Considering a large guidance gain is beneficial for the lander to move away from the boundary PPF, but excessive control acceleration will be generated when the landing error is large. To solve this problem, an adaptive guidance gain is designed using Gaussian function. Finally, the feasibility and effectiveness of the proposed algorithm have been verified through typical numerical simulations inspired by realistic Martian terrains. Moreover, this attempt using PPF methodology here can be easily reformulated to adjust the powered descent problem with collision avoidance for a flat surface or glide-slope constraint. [ABSTRACT FROM AUTHOR]

In this paper, a novel multi-stage trajectory transfer and fixed-point landing time optimal guidance method for the lunar surface emergency rescue mission is proposed. Firstly, the whole process motion and dynamics model for the lunar surface emergency rescue with four stages are established. Then, in the initial orbit transfer phase, the Lambert algorithm based on "prediction + correction" is designed for the non spherical gravitational perturbation of the moon. In the powered descent phase, the Hamiltonian function is used to design a time suboptimal explicit guidance law that can be applied in orbit in real time. Finally, aiming at the multi-stage global time optimal guidance, the whole time process guidance law is obtained by establishing the allowable control set for each stage in the whole process. The simulation results show that compared with the piecewise optimal control method, the present method has better optimization effect and shorter whole process time. It is of great significance to the possible emergency rescue mission of manned lunar exploration in the future.

• The main contribution of this study is the development of a new sequential-convex-programming-based technique that can periodically update the real-time guidance solution in consideration of the minimum-fuel Mars powered descent guidance problem with free final time. • To illustrate the validity of the proposed strategy for on-board real-time applications, an on-board test was conducted using the GR740, which is the processor of choice for the next-generation on-board computers of the European Space Agency (ESA) and Korea Aerospace Research Institute (KARI). • To assess the performance of the strategy, numerical simulations were conducted under various conditions. • Furthermore, the results of Monte Carlo simulations revealed that the proposed strategy is adequately robust against environmental and systematic disturbances. An on-board guidance generation approach for Mars pinpoint landing has been developed based on sequential convex programming. To optimize the flight time in the formulation, a discretized form of the landing problem was constructed under the conditions of linearized states, controls, and time increment between consecutive time steps. To implement this strategy in real time, on-board demonstrations were conducted with the GR740, which is the next-generation on-board processor of choice for the European Space Agency. The total number of time steps was determined based on the results of these demonstrations. The numerically simulated results also indicate that the solution obtained via the proposed strategy is close to the optimized GPOPS-II solution under the condition of no disturbance. Even under disturbances such as navigation errors, initial prediction errors, and perturbation forces, the proposed strategy ensures that the spacecraft reaches its target position with near-optimal fuel consumption. [ABSTRACT FROM AUTHOR]

Chandrayaan-2, being the sequel of Indian Lunar Mission, differs with Chandrayaan-1 in targeting a specific lunar orbit. The primary ground trajectory design objective is to achieve the desired lunar orbit, from which the lander attempts soft-landing on the pre-selected landing site. The target lunar orbit is designed such that while landing, the Sun elevation rise angle is 6° on the desired site. This is to ensure the lander touchdown is close to sunrise, which in-turn will maximize the lander mission life to one lunar solar day (approximately 14 Earth days). This makes the Chandrayaan-2 trajectory design unique and challenging in itself. The maneuver strategy has been worked out to acquire the specific target lunar orbit. The objective of the powered descent phase is soft-landing over the desired site satisfying all the sensor constraints. This multi-phase lunar landing optimal control problem has been solved for trajectory design of the powered descent phase. Based on this design, the Chandrayaan-2 spacecraft has been launched on 2 2 nd July, 2019 UT and successfully rendezvoused with the Moon on 2 0 th August, 2019 UT and has attempted a soft-landing on 6 th September, 2019 UT. • Chandrayaan-2 ground trajectory design considering mission constraints • A practical approach proposed for designing the target lunar orbit • A detailed summary of executed Chandrayaan-2 maneuvers • A brief summary of the Chandrayaan-2 spacecraft – hardware as well as payloads [ABSTRACT FROM AUTHOR]

This paper proposes a guidance scheme to achieve an autonomous precision landing on Mars and proposes a practical fixed-time stabilization theorem to analyze the robustness of the guidance. The proposed guidance is mainly based on the fixed-time stabilization method, and it can achieve the precision landing within a pre-defined time. This property enables the proposed guidance to outperform the finite-time stabilization technique which cannot handle uncertainties well and whose convergence time is dependent on initial states. Compared with the existing fixed-time stabilization theorem, the proposed practical fixed-time stabilization theorem can achieve a shorter convergence time and cope with unknown disturbances. When the Mars landing guidance is designed by this proposed theorem, the upper bound of the landing time and the maximum landing error subject to unknown disturbances can be calculated in advance. Theoretical proofs and Monte Carlo simulation results confirm the effectiveness of the proposed theorem and the proposed guidance. Furthermore, the efficacy of the proposed guidance with thrust limitations is also demonstrated by testing of 50 cases with a range of initial positions and velocities. [ABSTRACT FROM AUTHOR]

Autonomous powered-descent guidance algorithm for the pinpoint Mars landing in the presence of various disturbances and uncertainties is necessary for next-generation Mars exploration rover mission. This paper proposes a novel two sliding-surfaces guidance scheme based on fixed-time stabilization technique, which is robust against unknown Martian atmospheric disturbances (with known upper bound). For a fixed-time pinpoint landing mission, the main advantage of the proposed guidance is that the landing mission reliability can be ensured that Martian surface collision would never encounter. The fuel efficiency can be guaranteed in the comparison with the offline fuel optimal solution. The capacity of avoiding collisions is guaranteed by the monotonic convergence design of the proposed sliding modes. Lyapunov stabilization theory is adopted to prove the global stability of the proposed guidance. Monte Carlo numerical simulations are implemented in a realistic scenario and the results confirm the collision avoidance capability, the fuel efficiency and the robustness of the proposed guidance. • Compared with offline fuel-optimal guidance, the guidance is fuel efficient.. • The guidance can avoid the Martian surface collisions with any initial states. • The guidance is robust against unknown disturbances and model. • The whole landing process can be achieved within a pre-defined time. • Monte Carlo simulation verifies the effectiveness and robustness. [ABSTRACT FROM AUTHOR]

A novel zero-effort-miss (ZEM)/zero-effort-velocity (ZEV) optimal feedback guidance is proposed in order to rule out the possibility of Martian surface collision caused by the classical ZEM/ZEV optimal feedback guidance. The main approach is to add a collision avoidance term, which has self-adjustment capacity to ensure the near fuel optimality. Its main improvement is that it can not only successfully avoid collisions with the thruster constraint but also guarantee the near fuel optimality, and both of them are pivotal performances in Mars landing missions. Simulations are made to show the effectiveness of the proposed guidance and the parameters effects are simulated as well to analyze the properties of the proposed guidance. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

This paper proposed a two‐stage robust extended Kalman filter (TREKF) for state estimation of non‐linear uncertain system with unknown inputs. In engineering practice, the extended Kalman filter (EKF) with unknown inputs of the non‐linear uncertain system may be degraded or even diverged. The optimal two‐stage EKF (TEKF) is designed to solve the unknown inputs. The robust EKF (REKF) is considered to solve the non‐linear uncertain system for a long time. However, the information about the non‐linear uncertain system with unknown inputs is always incorrect. To solve this problem, the TREKF is designed by using the advantages of the TEKF and REKF, furthermore, its stability is proved. Finally, the performances of the TREKF, which are compared with the results of the REKF, TEKF and EKF, are verified by illustrating a numerical example of the powered descent phase of Mars EDL (entry, descent and landing). These also verify that the unfavourable effects of the model uncertainties and the unknown inputs are reduced efficiently by using the TREKF for the miniature coherent altimeter and velocimeter and inertial measurement unit integrated navigation during the powered descent phase of Mars EDL. [ABSTRACT FROM AUTHOR]

Pinpoint landing (within 100 m from the target) is essential for future Mars exploration missions. This paper deals with one aspect of the pinpoint landing architecture—the navigation performance improvement during the powered descent phase, and proposes an innovative navigation scheme to obtain the vehicle complete and accurate states. On the basis of dead reckoning relying on the Inertial Measurement Unit, measurements of the Integrated Doppler Radar are adopted to correct the vehicle velocity and altitude. Distance between the vehicle and one Mars Orbiter as well as their line-of-sight relative velocity is measured by a radio sensor, and integrated in the filter to correct the vehicle horizontal position. The innovative navigation system is based on an Extend Kalman Filter. Two observation schemes are developed. One considers measurements of the Integrated Doppler Radar and radio range measurement. Another further considers radio velocity measurement. The performance of the innovative navigation scheme is greatly influenced by the position of the Mars Orbiter with respect to the target. Stochastic analyses are performed to obtain optimal locations of Mars Orbiter. Finally, the innovative navigation scheme performances are assessed through stochastic simulations. Its performance improvements are demonstrated by comparison with the Integrated Doppler Radar only navigation scheme. [ABSTRACT FROM AUTHOR]

As a class of autonomous deep-space exploration robots, Mars lander is simultaneously affected by wind disturbance and actuator fault, which hinders precise landing on Mars. In this article, we propose a composite guidance approach by combining disturbance observer and iterative learning observer together. The Mars wind disturbance is dealt by the disturbance observer providing wind estimation which is rejected through feed-forward channel. Meanwhile, the iterative learning observer is used to estimate the deficiency of actuators. The proposed guidance scheme ensures not only the precision but also the reliability of the Mars guidance system. The stability of the closed-loop system is analyzed. Simulations made in different situations demonstrate the performance of the proposed approach. [ABSTRACT FROM AUTHOR]