Ablation and thermal response program for spacecraft heatshield analysis

Abstract

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 which 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, the Aerotherm Charring Material Thermal Response and Ablation Program, 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, such as Stardust, Mars Microprobe (Deep Space n), Saturn Entry Probe, and Mars 2001, using advanced light-weight ceramic ablators developed at NASA Ames Research Center, are presented and discussed. Nomenclature a B' F g h h I' K k rh = absorption coefficient, m"' = rh / p e u e C M , dimensionless mass blowing rate = pre-exponential constant in Eq.(8), s" 1 = Stanton number for heat transfer = Stanton number for mass transfer = specific heat, J/kg-K = activation energy in Eq.(8), J/kmol = exponential integral function = view factor = outward pyrolysis mass flux, kg/m 2 -s = enthalpy, J/kg = partial heat of charring, defined in Eq.(6), J/kg = radiation source function in Eq.(2), W/m 2 -sr = radiant intensity in +x direction, W/m 2 -sr = radiant intensity in -x direction, W/m 2 -sr = a + o s , extinction coefficient, m" 1 = thermal conductivity, W/m-K = mass flux, kg/m 2 -s * Aerospace Engineer. t Aerospace Engineer. Senior Member AIAA. P = pressure, N/m 2 q c = conductive heat flux, W/m 2 qn = radiative heat flux, W/m 2 R = universal gas constant, J/kmol-K s = surface recession, m s = surface recession rate, m/s T = temperature, K u = velocity, m/s x = y -s, moving coordinate, m y = stationary coordinate, m Z* = coefficient in Eq.(9), defined in Ref. 14 a = surface absorptance e = surface emissivity F = volume fraction of resin K = optical thickness K D = optical thickness for path of length D A, = blowing reduction parameter 9 = time, s p = total density, kg/m 3 a = Stefan-Boltzmann constant, W/m 2 -K 4 a s = scattering coefficient, m" 1 I = mass fraction of virgin material, defined in Eq.(5) T = decomposition reaction order in Eq.(8) subscripts c = char e = boundary-layer edge g = pyrolysis gas i = density component (A, B, and C) j = surface species v = virgin w = wall

Atmospheric entry studies for Uranus

“…Based on previous aerocapture work [22] for Mars, the captured orbit was chosen to be I-Uranus day orbit (�17.24 hours) with periapsis =5000km and apoapsis =109,650km. The total energy associated with this trajectory is -3.50x10 11 m 2 /s 2 .…”

Polymer Nanocomposites as Ablative Materials

“…TPS protect a body or substrate from the severe heating encountered during hightemperature conditions such as in a hypersonic flight through a planetary atmosphere or in a propulsion chamber of solid rocket motor engine [15]. Since TPS is a single point of failure subsystem, it is critical and its performance needs to be validated through modeling and simulation, ground test, and analysis.…”

Projectile Nosetip Thermal Management for Railgun Launch to Space

“…Chen and Milos [5] used a three-component decomposition model with a temperature-dependent rate of gasification for each component; this approach is appropriate for a wide class of ablators and will also be used here. The internal state of the ablator is then linked to the flow field through a surface-energy balance relation [5], [6] that accounts for all energy fluxes to and from the surface (convective heating from the gas, chemical reactions at the surface, ablation species blowing from the surface, surface radiative heating and reradiative cooling, and conduction into the ablator). This link between the fluid and solid models is critical; the approach developed by Keenan and Candler [6], [7], [8] will be used in the proposed work.…”

“…• develop a computational fluid dynamics code to model fully coupled, high-enthalpy reacting flows and material response with ablation and spallation; • validate code with the existing Passive Nosetip Technology (PANT) program data [4] and other available experiments; compare to existing methods, including the fully implicit ablation and thermal (FIAT) method [5]; • optimize the trajectory, vehicle geometry, launch velocity, and rocket assist coupled with EM launcher developments; • evaluate system tradeoffs between active cooling, passive ablation, ultrahigh temperature ceramics, and other emerging high-temperature materials concepts. The aerothermal analysis and design is being performed with validated computational fluid dynamics (CFD) codes for hypervelocity flows that have been developed at the University of Minnesota (UMN) over the past 15 years.…”