Earth entry system options for human return missions from the Moon and Mars were analyzed and compared to identify trends among the configurations and trajectory options and to facilitate informed decision making at the exploration architecture level. Entry system options included ballistic, lifting capsule, biconic, and lifting body configurations with direct entry and aerocapture trajectories. For each configuration and trajectory option, the thermal environment, deceleration environment, crossrange and downrange performance, and entry corridor were assessed. In addition, the feasibility of a common vehicle for lunar and Mars return was investigated. The results show that a low lift-to-drag ratio (L/D = 0.3) vehicle provides sufficient performance for both lunar and Mars return missions while providing the following benefits: excellent packaging efficiency, low structural and TPS mass fraction, ease of launch vehicle integration, and system elegance and simplicity. Numerous configuration options exist that achieve this L/D.

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A computer tool to perform entry vehicle ablative thermal protection systems sizing has been developed. Two options for calculating the thermal response are incorporated into the tool. One, an industry-standard, high-fidelity ablation and thermal response program was integrated into the tool, making use of simulated trajectory data to calculate its boundary conditions at the ablating surface. Second, an approximate method that uses heat of ablation data to estimate heat shield recession during entry has been coupled to a one-dimensional finite-difference calculation that calculates the in-depth thermal response. The in-depth solution accounts for material decomposition, but does not account for pyrolysis gas energy absorption through the material.

There are numerous challenges associated with placing a spacecraft in orbit around Mars. Often, trades must be made such as the mass of the payload and the amount of fuel that can be carried. One technique employed to more efficiently place a spacecraft in orbit while maximizing payload mass (minimizing fuel use) is aerobraking. The Mars Odyssey Spacecraft made use of aerobraking to gradually reduce its orbit period from a highly elliptical insertion orbit to its final science orbit. Aerobraking introduces its own unique challenges, in particular, predicting the thermal response of the spacecraft and its components during each aerobraking drag pass. This paper describes the methods used to perform aerobraking thermal analysis using finite element thermal models of the Mars Odyssey Spacecraft's solar array. To accurately model the complex behavior during aerobraking, the thermal analysis must be tightly coupled to the spatially varying, time dependent aerodynamic heating analysis. Also, to properly represent the temperatures prior to the start of the drag pass, the model must include the orbital solar and planetary heat fluxes. It is critical that the thermal behavior be predicted accurately to maintain the solar array below its structural flight allowable temperature limit. The goal of this paper is to describe a thermal modeling method that was developed for this purpose.