(Georgia Institute of Technology, 2006-01)
Dec, John A.; Braun, Robert D.
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. Engineering correlations are used to estimate stagnationpoint
convective and radiative heating as a function of time. The sizing tool calculates
recovery enthalpy, wall enthalpy, surface pressure, and heat transfer coefficient.
Verification of this tool is performed by comparison to past thermal protection system
sizings for the Mars Pathfinder and Stardust entry systems and calculations are performed
for an Apollo capsule entering the atmosphere at lunar and Mars return speeds.
(Georgia Institute of Technology, 2005-08)
Putnam, Zachary R.; Braun, Robert D.; Rohrschneider, Reuben R.; Dec, John A.
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.
(Georgia Institute of Technology, 2005-06)
Kipp, Devin M.; Dec, John A.; Wells, Grant William; Braun, Robert D.
A Planetary Entry Systems Synthesis Tool, with applications to conceptual design and modeling of entry systems has been developed. This tool is applicable to exploration missions that employ entry, descent and landing or aerocapture. An integrated framework brings together relevant disciplinary analyses and enables rapid design and analysis of the atmospheric entry mission segment. Tool performance has been validated against Mars Pathfinder flight experience and has direct relevance to future NASA robotic and human space exploration systems.