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Aerocapture is a maneuver that utilizes atmospheric drag for orbital adjustments. Drag modulation is an effective aerocapture control method that achieves the desired target orbit by adjusting the ballistic coefficient. Accurate identification of the ballistic coefficient during the atmospheric flight phase is critical for ensuring the control precision and flight safety of the vehicle. This paper proposes a Cubature Kalman Smoothing (CKS) algorithm based on singular value decomposition (SVD) and initial value iteration. First, singular value decomposition is introduced to maintain the stability of the identification process. Second, the sensitivity of the algorithm to the initial values is mitigated by incorporating an iterative initialization method. Finally, the accuracy and robustness of the proposed identification algorithm are validated through Monte Carlo simulations.

Aerocapture, the action of delivering a vehicle from a hyperbolic orbit to a planetary orbit by using the aerodynamic force, could potentially lower fuel consumption. By controlling the direction and size of the aerodynamic force, the vehicle can accurately enter the target orbit. This paper focuses on a preliminary study of the optimal trajectory for aerocapture on the basis of a novel flight control option, which considers lift and drag joint modulation so as to suit variable structure spacecraft. In the preliminary evaluation of such a flight control option, the aerocapture corridors under lift modulation and drag modulation and the influence of the ballistic coefficient on aerocapture were analyzed, demonstrating that joint modulation can achieve complementary advantages compared with pure lift modulation and drag modulation. Based on this flight control option, optimal aerocapture trajectories with different path constraints, target orbital constraints and control variable constraints were found. It bears noting that both the bank angle and the reference area were taken as control variables for lift modulation and drag modulation, respectively, during the atmospheric flight in the process of designing the optimal trajectories. The optimal results indicate that the flight control option with lift and drag joint modulation can greatly broaden the necessary conditions for aerocapture and extend the target orbital range.

Developing spacecraft for efficient aerocapture missions demands managing extreme aerothermal environments, precise controls, and atmospheric uncertainties. Successful designs must integrate vehicle airframe considerations with trajectory planning, adhering to launcher dimension constraints and ensuring robustness against atmospheric and insertion uncertainties. To advance robust multi-objective optimization in this field, a new framework is presented, designed to rapidly analyze and optimize non-thrusting, fixed angle-of-attack aerocapture-capable spacecraft and their trajectories. The framework employs a three-degree-of-freedom atmospheric flight dynamics model incorporating planet-specific characteristics. Aerothermal effects are approximated using established Sutton–Graves, Tauber–Sutton, and Stefan–Boltzmann relations. The framework computes the resulting post-atmospheric pass orbit using an orbital element determination algorithm to estimate fuel requirements for orbital corrective maneuvers. A novel algorithm that consolidates multiple objective functions into a unified cost function is presented and demonstrated to achieve superior optima with computational efficiency compared to traditional multi-objective optimization approaches. Numerical examples demonstrate the methodology’s effectiveness and computational cost at optimizing terrestrial and Martian aerocapture maneuvers for minimum fuel, heat loads, peak heat transfers, and an overall optimal trajectory, including volumetric considerations.

Aerocapture maneuvers refer to a single atmospheric crossing to deplete orbital energy and establish a closed orbit. During the atmospheric flight, adjusting the spacecraft’s vertical lift component in an optimal manner, bang-bang bank control, will minimize the propulsion fuel consumption required to establish the target orbit. However, such methods have been suffering from the performance’s oversensitivity to the control’s instantaneous switching time and poor robustness. To address these problems, we propose a new numerical predictor-corrector guidance algorithm based on the saturation function profile in this paper. The saturation function is used to basically simulate the bang-bang control structure, which enhances the algorithm’s robustness by reducing its dependence on the relevant parameters without losing too much optimality. Monte Carlo simulations in both Earth and Mars scenarios demonstrate the robustness, accuracy, and near-optimal performance of the proposed guidance method.

Small satellite constellations in multiple-inclination, low-circular orbits around Mars and Venus have the potential to perform a range of high-value science investigations within cost-constrained missions. A major challenge for small satellites is that they require large ΔV to enter low-circular orbits, which can drive up both spacecraft mass and cost. Compared to chemical propulsion, which requires large amounts of propellant, and electric propulsion, which requires large solar arrays and comes with long flight times, aerocapture enables direct access to low-circular orbits at Mars and Venus with minimal ΔV. The study shows how drag-modulation aerocapture, when combined with small B-plane targeting maneuvers, allows the delivery of multiple small satellites to various-inclination, low-circular orbits to establish a constellation. Preliminary cost estimates indicate that by reducing the required ΔV for orbit insertion, aerocapture can potentially reduce the cost of a small satellite going to a low-circular Mars orbit compared to propulsive insertion. The ability of low-cost spacecraft to enter planetary orbits will enable a new paradigm of interplanetary missions using small dedicated launch vehicles and planetary constellations at Mars and Venus.

While solar sailing missions have been developed in the past, orbital solar sailing missions are limited in number. Increasing access to the solar system can be achieved by leveraging the potential of solar sailing spacecraft. A feasibility study is performed to analyse a controlled solar sailing mission to Mars using a CubeSat platform. The proposal to send a solar sailing CubeSat to Mars would be revolutionary in delivering small payloads to the red planet by harnessing the solar radiation pressure emitted from the Sun. By utilising solar sailing and aerocapture techniques in the mission design, the need for chemical propulsion for transfer and capture is greatly reduced, minimising launch mass and maximizing payload mass efficiency. In this study, investigations were carried out as mission analyses in the area of trajectory architecture and subsystem overview for systems available today and in the future. The influence and evolution of trajectories under gravitational and environmental conditions on Mars and the interplanetary transfer were scrutinised individually as relevant restricted dynamic systems. For a chosen insertion orbit, several feasible trajectories to Mars based on different solar sail systems were investigated, including optimised ballistic trajectories and the Sundiver transfer for solar sails. For the interplanetary trajectories, the planetary capture manoeuvre of the spacecraft to a serviceable orbit using aerobraking capture is evaluated. The final trajectory design features minimum-time optimisation for the interplanetary transfer of a solar sailing spacecraft from Earth to Mars. As the selected spacecraft platform is limited to low-thrust solar sailing propulsion, severe performance limitations apply to this mission. This work aims to highlight possible interplanetary trajectories for present and future small satellite platforms despite these limitations by using dynamical orbital simulation models.

This paper proposes a novel method to efficiently construct the aerocapture attainability set based on convex optimization. Using dimensionality reduction, constructing the attainability set is equivalent to solving a set of discrete points on its boundary. As solving each of the boundary points is a typical nonlinear optimal control problem, a sequential convex programming method is adopted. The efficiency and accuracy of the proposed algorithm is demonstrated by high-fidelity numerical results. This is the first time that the configuration of the aerocapture attainability set is precisely described by the state variables at atmospheric exit. Since the quantification of the set is significant for assessing the feasibility of performing an aerocapture maneuver, the proposed method can be used as a reliable tool for systematic design for aerocapture mission.

Lunar exploration architecture can be made more flexible and reliable with the support of a low-Earth orbit (LEO) space station. This study therefore evaluated a proposed hybrid optimization scheme to design the entire trajectory of a reusable spacecraft starting from trans-Earth injection (EI) at the perilune and ending at an LEO space station. As such a trajectory has multiple constraints and multiple dynamical models, it is divided into the trans-Earth phase, aerocapture phase, and postatmospheric phase. The optimization scheme is performed at two levels: sublevel and top level. At the sublevel, two novel pseudo rules are proposed to optimize the trans-Earth trajectory so that it satisfies the coplanar constraints of the space station. Then, in the aerocapture phase, the bank angle is optimized to satisfy the mission constraints, and in the atmospheric phase, the one-impulsive maneuver is performed and optimized to insert the spacecraft into the target space station orbit. The multiple phases are connected to each other by boundary conditions where the terminal state of the previous phase is transformed into the initial state of the following phase. At the top level, the vacuum perigee height is selected as the mission design variable based on problem characteristics analysis and a hybrid optimization scheme is conducted to minimize the total velocity increment. The simulation results demonstrate that the proposed hybrid optimization method is effective for the design of an entire trajectory with acceptable velocity cost which is less than that in the previous study. The coplanar constraints of the space station and other mission constraints in each phase are also satisfied. Furthermore, the proposed trajectory design method is shown to be applicable to a reusable spacecraft returning to an LEO space station parked in any arbitrary orbital plane.

Aerocapture can significantly reduce the velocity increment required for a planetary orbital mission and reduce the amount of propellant needed. And it may be one of the key technologies necessary for large-scale space exploration missions in the future. In this paper, the analytical solution of aerocapture based on the piecewise variable ballistic coefficient is studied around the exploration of Mars. An aerocapture analytical predictive guidance algorithm for single ballistic coefficient switching is proposed. The terminal velocity after the ballistic coefficient switching can be obtained by analytical calculation in real time. The adaptive control of the switching time of the ballistic coefficient is realized. The simulation results show that the guidance algorithm is accurate and robust, which can effectively overcome the influence of atmospheric density error, aerodynamic parameter error, and initial state uncertainty.