Mars Aerocapture Research Articles

Nonsingular Fixed-Time Tracking Guidance for Mars Aerocapture With Neural Compensation

Mars aerocapture is one of the most concerned technologies for future Mars sample-return and manned exploration missions. This article investigates the challenging problem of the reference trajectory tracking guidance for Mars aerocapture under uncertainties. Based on an integral sliding surface, a fixed-time neural adaptive tracking guidance law is synthesized by incorporating the neural network (NN) approximation into the fixed-time integral sliding mode control approach. Benefiting from the integral sliding surface design, the proposed tracking guidance law has no singularity problem inherently existing in terminal sliding mode control. By adopting the NN approximation to compensate for the lumped uncertain term in the feedforward loop, the proposed tracking guidance law is strongly robust against aerodynamic coefficient uncertainties and atmospheric density uncertainty. Stability analysis shows that the radial distance tracking error and its time derivative can stabilize to the small neighborhoods around the origin in fixed time under the proposed tracking guidance law. Finally, the effectiveness and advantages of the proposed tracking guidance law are illustrated through simulations and comparisons.

Correction to: Analytical Guidance for Mars Aerocapture via Drag Modulation

Correction to: Analytical Guidance for Mars Aerocapture via Drag Modulation

Optimal Information Filtering for Robust Aerocapture Trajectory Generation and Guidance

Entry flight is a critical mission phase of planetary aeroassist problems. During atmospheric flight, aleatory-epistemic uncertainty and environmental factors reduce the accuracy of predicted future states for precision targeting. This problem has been approached historically with closed-loop guidance rooted in certainty equivalence. This property separates estimation and control problems, allowing each to be considered independently. In other concept studies, an observer model is neglected altogether in favor of assuming perfect state knowledge. However, a flight system will inevitably have imprecise state information and variability in its underlying dynamics and measurement models. Systemic uncertainty is a fundamental limitation of existing entry guidance approaches. This work seeks to overcome these challenges by posing aerocapture as a robust optimization problem. The cost objective of the maneuver is reformulated to account for uncertainty in atmospheric structure, vehicle performance parameters, and state estimation accuracy using an observer-based consider filter. An expected value performance cost is developed from anticipated measurement conditioning effects. A rapid solution methodology is illustrated using explicit integration strategies with a parameterized control structure. Results for a Mars aerocapture concept study show improvement in the postcapture orbit accuracy with low computational overhead.

Characteristic model and its applications in Mars aerocapture guidance

Mars exploration has been a hot spot around the world in the field of deep space exploration in recent years. Low-cost space transport is one of the key factors in space exploration and development, and aerodynamic assisted orbit change will reduce space vehicle design costs. Aerocapture is a form of aerodynamically assisted orbit shift, referring to the use of the aerodynamic force generated by the planetary atmosphere to minimize the energy of the orbit and the transfer of the hyperbolic orbit to the target orbit. There is only one opportunity to fulfill the aerocapture mission, and the criterion for reliability is very high. Because the aerocapture guidance and accuracy and robustness of the control are very poor, a smart and autonomous guidance system is required. The characteristic model-based guidance approach has a broad application potential and has high adaptability, robustness, and autonomy. The research status and progress of the theory of characteristic model and its application in Mars aerocapture guidance are reviewed, and the problems to be solved are put forward, which provides the basis for further research.

Aerothermodynamic Studies on Low-Ballistic-Coefficient Mars Aerocapture Vehicle with Drag Modulation and Electric Propulsion

A Mars orbiter/lander mission using a micro-satellite (less than 100 kg at Mars arrival) with a deployable membrane aeroshell for the orbit insertion by the aerocapture and the electric propulsion for the trajectory maneuver was considered. The aerodynamic heating environment during the atmospheric flight was investigated solving the thermo-chemical nonequilibrium axi-symmetric Navier-Stokes equations around the spacecraft with the flare-shaped aeroshell. To obtain the appropriate amount of deceleration at the atmospheric pass, the drag modulation technique, in which the aeroshell is timely jettisoned from the spacecraft, was assumed. The hypersonic wind tunnel experiment successfully demonstrated that the backward jettison works well without significant time delay and attitude disturbance. Finally the corridor width for the entry path angle was estimated considering the ability of the orbit insertion, the peak aerodynamic heating, and the capability of the electric propulsion to raise the periapsis altitude. The combination of a low ballistic coefficient flight with the membrane aeroshell and an electric propulsion works well to acquire a finite width of the entry corridor for the aerocapture.

Identifying method of entry and exit conditions for aerocapture with near minimum fuel consumption

Identifying the entry and exit conditions for aerocapture can be an important aid for mission analysis and design. The objectives of this paper are to present methods for computing the exit conditions to transfer a post-aerocapture orbit into a specified target orbit with near minimum fuel consumption and corresponding entry conditions to maximise entry zone and to identify several other uses for the conditions. This paper presents: 1) an approach to the solution of exit conditions according to orbital mechanics and 2) an approach to calculate the entry conditions based on the theory of controllable set. To illustrate their effectiveness, the conditions are employed to characterise the performance of Mars Science Laboratory (MSL)-type vehicle for Mars aerocapture. The entry states dispersion analysis under uncertainties of ballistic coefficient, lift-to-drag ratio and atmospheric density are discussed. Roles for the conditions in choosing nominal entry and exit states are described.

Roadmap to a human Mars mission

We propose a new roadmap for the preparation of the first human mission to Mars. This proposal is based on the work of ISECG and several recent recommendations on human Mars mission architectures. A table is proposed to compare the possible benefits of different preparatory missions. Particular attention is paid to the possibility of qualifying important systems thanks to a heavy Mars sample return mission. It is shown that this mission is mandatory for the qualification of Mars aerocapture at scale-1, EDL systems at scale 1 and Mars ascent. Moreover, it is a good opportunity to test many other systems, such as the heavy launcher and the transportation systems for the trips beyond LEO. These tests were not mentioned in the last ISECG report. This strategy is facilitated in the case of the simplified Mars mission scenarios that have recently been presented because it is suggested that relatively small vehicles with small crew sizes are used in order to optimize the payload mass fraction of the landing vehicles and to avoid the LEO assembly. An important finding of the study is that a human mission to the surface of the Moon is not required for the qualification of the systems of a human mission to Mars. Since affordability is a key criterion, two important missions are proposed in the roadmap. The first is a heavy Mars sample return mission and the second is a manned mission to a high Earth orbit or eventually to the vicinity of the Moon. It is shown that both missions are complementary and sufficient to qualify all the critical systems of the Mars mission.

Development of optimal and robust predictive guidance technique for Mars aerocapture

The paper discusses a predictive guidance technique for the optimal aerocapture maneuver accomplished by a high-mass Mars orbiter to achieve the targeted circular parking orbit with minimal fuel consumption. Unlike the majority of investigated algorithms of the “predictor–corrector” family with a time-invariant control, the guidance algorithm considered in the paper computes a time-dependent control function which, being a reference profile of the angle of attack, provides a minimum to the post-aerocapture propulsive ΔV at the apoapsis of a transfer ellipse. Initially, the two-parametric bounded linear control function has been used, and the algorithm developed found an optimal solution directly by “descending” in a control parameters space. The simulation has shown that an optimum is achieved when the slope of the linear function tends to infinity, i.e., the bounded linear control is being transformed into a one-parametric step function. Such a result proved to be in conformity with the earlier “indirect” optimal solution, namely, derived from Pontryaginʼs maximum principle for a slightly simplified motion model. Thus, the final version of algorithm providing “direct” optimization by solving a nonlinear eigenvalue problem needs to determine only one parameter – switching time of angle of attack. The solution obtained satisfies the terminal condition – a target apoapsis, as well as it is close to optimum. Another issue the paper focuses on is an improvement of robustness of the algorithm under unpredictable model disturbances, especially variations of atmospheric density. The new model adaptation algorithm is developed, which estimates density errors and updates the density profile during the descent phase to be used for control corrections in the ascent phase. Testing the example simulation model with angular motion has shown the runtime effectiveness and high targeting accuracy provided by the proposed predictive technique.

Simultaneous Optimization of Flight Control and Aerodynamic Shape for Aerocapture Vehicle

Aerocapture is expected to reduce the cost of the interplanetary exploration in the future because it utilizes aerodynamic drag to decelerate the vehicle rather than fuel-costly chemical propulsion. However, aerocapture mission involves lots of complicated difficulties such as quite narrow entry path angle, excess aerodynamic heating, aerodynamic load, and various uncertainties. It requires high levels of control technique and aerodynamic characteristics prediction. In this paper, the results of the simultaneous optimization of the flight control and aerodynamic shape for Mars aerocapture vehicle are presented. To maximize the payload mass, fuel mass for orbital transfer and TPS mass must be minimized. For this purpose, objective functions are set as follows: 1) minimize the delta-V for orbit insertion, 2) minimize the total heat load, 3) minimize both the delta-V and total heat load. Firstly, optimization was conducted only for the flight control, assuming the fixed aerodynamic shape and entry path angle. Secondary, simultaneous optimization was conducted. As a result, trade-off between delta-V and total heat load was clarified. Moreover, by executing simultaneous optimization, the guideline of designing aerocapture, which can minimize the objective function in each case 1)-3), was obtained.

Shock Radiation Measurements for Mars Aerocapture Radiative Heating Analysis

NASA's In-Space Propulsion Technology program is supporting the development of shock radiation transport models for aerocapture missions to Mars and Venus. Phenomenological models of nonequilibrium shock radiation will be incorporated into high-fidelity flowfield computations used to predict the aerothermal environments for a Mars or Venus aerocapture entry vehicle. These models are validated with shock radiance measurements obtained at flight-relevant conditions. A comprehensive test series in the NASA Ames Electric Arc Shock Tube facility at a representative freestream condition was recently completed. The facility's optical instrumentation enabled spectral measurements of shocked gas radiation from the vacuum ultraviolet to the near infrared. The instrumentation captured the nonequilibrium postshock excitation and relaxation dynamics of dispersed spectral features. A description of the shock tube facility, optical instrumentation, and examples of the test data are presented.

Trajectory and Attitude Simulation for Mars Aerocapture and Aerobraking

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Seasonal effects on a Mars aerocapture manoeuvre, an ISU project

Seasonal effects on a Mars aerocapture manoeuvre, an ISU project

Thermochemical nonequilibrium viscous shock-layer analysis for a Mars aerocapture vehicle

The effects of thermal and chemical nonequilibrium on the aerothermodynamic environment over a Mars aerocapture vehicle are studied. In order to understand the characteristics of the flow at aerobraking, the classification of the stagnation region flow has been carried out from a viewpoint of the extent of the thermal and chemical nonequilibrium in the shock layer. The numerical analysis of the forebody flowfield over a spacecraft has been made by using the viscous shock-layer equations, including the thermal and chemical nonequilibrium effects of carbon dioxide. The influences of the thermal nonequilibrium on the flow properties in the shock layer become significant at an altitude above 60 km. As for the wall heating rate, however, its influence is not as strong as the influence of the wall catalycity. The sensitivities of the wall heating rate to the uncertainties in the real gas effect models are investigated parametrically. The wall heating rate is strongly affected by the wall catalycity. Consequently, a chemical nonequilibrium flow analysis is necessary for the analysis of the aerothermodynamic environment of a Mars aerocapture vehicle.

Six-degree-of-freedom guidance and control analysis of Mars aerocapture

A six-degree-of-freedom (6DOF) simulation is developed to investigate the control and guidance issues of a Mars aerobraking vehicle. The guidance algorithm used is a predictor-corrector guidance formulation designed to control the exit orbital apoapsis and wedge angle using bank-angle modulation. Major features of this predictor-corrector guidance algorithm include: (1) integration of the 3DOF equations of motion within an inner-loop simulation; (2) load-relief logic; (3) finite roll rates; and (4) an aerodynamic feedback multiplier. The algorithm is capable of successfully guiding the vehicle through combinations of atmospheric density dispersions, aerodynamic mispredictions, and off-nominal atmospheric interface conditions. This study demonstrated that the addition of vehicle dynamics to the Mars aerobraking simulation does not significantly impact mission feasibility. That is, a robust control system design coupled with an adaptive guidance algorithm can assure mission success in the presence of numerous off-nominal conditions.

Aerothermodynamic design feasibility of a Mars aerocapture vehicle

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.