Trajectory, aerothermal conditions, and thermal protection system mass for the MARS 2001 aerocapture mission

Self Cite

352

An implicit ablation and thermal response program is presented for simulation of one-dimensional transient thermal energy transport in a multilayer stack of isotropic materials and structure that can ablate from a front surface and decompose in depth. The governing equations and numerical procedures for solution are summarized. Solutions are compared with those of an existing code, CMA, and also with arcjet data. Numerical experiments show that the new code is numerically more stable and solves a much wider range of problems compared with the older code. To demonstrate its capability, applications for thermal analysis and sizing of aeroshell heatshields for planetary missions of Stardust, Mars Microprobe (Deep Space II), Saturn Entry Probe, and Mars 2001, using advanced lightweight ceramic ablators developed at NASA Ames Research Center, are presented and discussed. Nomenclature a = absorption coef cient, m ¡1 B 0 = dimensionless mass blowing rate, P m=½ e u e C M B a = pre-exponential constant in Eq. (8), s ¡1 C H = Stanton number for heat transfer C M = Stanton number for mass transfer c p = speci c heat, J/kg-K E a = activation temperature in Eq. (8), K F = view factor g = outward pyrolysis mass ux, kg/m 2 -s h = enthalpy, J/kg N h = partial heat of charring, de ned in Eq. (6), J/kg I 0 = radiation source function in Eq. (2), W/m 2 -sr i C = radiant intensity in Cx direction, W/m 2 -sr i ¡ = radiant intensity in ¡x direction, W/m 2 -sr K = extinction coef cient, a C ¾ s , m ¡1 k = thermal conductivity, W/m-K P m = mass ux, kg/m 2 -s P = pressure, N/m 2 q C = conductive heat ux, W/m 2 q R = radiative heat ux, W/m 2 R = universal gas constant, J/kmol-K s = surface recession, m P s = surface recession rate, m/s T = temperature, K u = velocity, m/s x = moving coordinate, y ¡ s, m y = stationary coordinate, m Z ¤ = coef cient in Eq. (9), de ned in Ref. 10 ® = surface absorptance 0 = volume fraction of resin " = surface emissivity µ = time, s · = optical thickness · D = optical thickness for path of length Ḑ = blowing reduction parameter ½ = density, kg/m 3 ¾ = Stefan-Boltzmann constant, W/m 2 -K 4 ¾ s = scattering coef cient, m ¡1 ¿ = mass fraction of virgin material, de ned in Eq. (5) 9 = decomposition reaction order in Eq. (8) Subscripts c = char e = boundary-layeredge g = pyrolysis gas i = density component (A, B, and C ) j = surface species v = virgin w = wall

Section: Mars 2001mentioning

confidence: 99%

Ablation and Thermal Response Program for Spacecraft Heatshield Analysis

Chen

Milos

1999

Self Cite

352

“…Traditionally, aerocapture has been analyzed using a rigid aeroshell similar to those used for entry applications [25,26]. Lift is necessary to control the energy dissipated in the atmosphere in the presence of uncertainties, and so axisymmetric shapes are flown at an angle of attack.…”

Section: Ballute Aerocapturementioning

confidence: 99%

Survey of Ballute Technology for Aerocapture

Rohrschneider

Braun

2007

Ballute aerodynamic decelerators have been studied since early in the space age (1960's), being proposed for aerocapture in the early 1980's. Significant technology advances in fabric and polymer materials as well as analysis capabilities lend credibility to the potential of ballute aerocapture. The concept of the thin-film ballute for aerocapture shows the potential for large mass savings over propulsive orbit insertion or rigid aeroshell aerocapture. The mass savings of this concept enables a number of high value science missions. Current studies of ballute aerocapture at Titan and Earth may lead to flight test of one or more ballute concepts within the next five years. This paper provides a survey of the literature with application to ballute aerocapture. Special attention is paid to advances in trajectory analysis, hypersonic aerothermodynamics, structural analysis, coupled analysis, and flight tests. Advances anticipated over the next 5 years are summarized. Clamped Ballutes Trailing Ballutes Single stage design Asymmetric lifting design Two stage design Circular cross-section torus design Airfoil cross-section torus design Cable Membrane Towed sphere design Clamped torus Clamped Ballutes Trailing Ballutes Single stage design Asymmetric lifting design Two stage design Circular cross-section torus design Airfoil cross-section torus design Cable Membrane Towed sphere design Clamped torus

“…GASP has been used extensively at NASA Ames Research Center in TPS sizing applications for both planetary 12,13 and transatmospheric 14 entry vehicles. The GASP solutions were run with models that are similar, if not identical, to those used in LAURA.…”

Section: Gasp Cfd Codementioning

confidence: 99%

Aeroheating Environments for a Mars Smart Lander

Edquist

Liechty

Hollis

et al. 2006

A proposed Mars Smart Lander is designed to reach the surface via lifting-body atmospheric entry (α = 16 deg) to within 10 km of the target site. CFD predictions of the forebody aeroheating environments are given for a direct entry from a 2005 launch. The solutions were obtained using an 8-species gas in thermal and chemical nonequilibrium with a radiative-equilibrium wall temperature boundary condition. Select wind tunnel data are presented from tests at NASA Langley Research Center. Turbulence effects are included to account for both smooth body transition and turbulence due to heatshield penetrations. Natural transition is based on a momentum-thickness Reynolds number value of 200. The effects of heatshield penetrations on turbulence are estimated from wind tunnel tests of various cavity sizes and locations. Both natural transition and heatshield penetrations are predicted to cause turbulence prior to the nominal trajectory peak heating time. Laminar and turbulent CFD predictions along the trajectory are used to estimate heat rates and loads. The predicted peak turbulent heat rate of 63 W/cm 2 on the heatshield leeward flank is 70% higher than the laminar peak. The maximum integrated heat load for a fully turbulent heat pulse is 38% higher than the laminar load on the heatshield nose. The predicted aeroheating environments with uncertainty factors will be used to design a thermal protection system. NOMENCLATUREmomentum thickness ∞ freestream --------------* Vehicle Analysis Branch, Senior Member, AIAA. † Aerothermodynamics Branch.