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.

Ablation and Thermal Response Program for Spacecraft Heatshield Analysis

Chemical-kinetic parameters governing the e ow in the shock layer over a heat shield of a blunt body entering Earth’ s atmosphere from a hyperbolic orbit are derived. By the use of the assumption that the heat shield is made of carbon phenolic and by allowing for an arbitrary rateof pyrolysis-gasinjection, chemical reactions occurring in the shock layer are postulated, and the collision integrals governing the transport properties, the rate coefe cients of the reactions, and the parameters needed for the bifurcation model and for the e nite-rate kinetic wall boundary conditions are determined using the best available techniques. Sample e owe eld calculations are performed using this set of parameters to show that the heating and surface removal rates are substantially smaller than calculated using theexisting setofsuch parameters and traditionalassumptionsof gas ‐surfaceequilibrium and quasi-steadystate ablation.

Chemical-Kinetic Parameters of Hyperbolic Earth Entry

Science and technology of polymeric ablative materials for thermal protection systems and propulsion devices: A review

The development of a new high- delity methodology for predicting entry  ows with coupled radiation and ablation is described. The prediction methodology consists of an axisymmetric, nonequilibrium, Navier–Stokes  ow solver loosely coupled to a radiation prediction code and a material thermal response code. The methodology is used to simulate the 12.6-km/s Earth atmospheric entry of the Stardust sample return capsule using ablating and nonablating boundary conditions. These  ow simulations are used to size and design the Stardust forebody and afterbody heatshields and develop arcjet test conditions and models. The  ow simulations indicate that the afterbody heating and pressure pro les in time are signi cantly different than the forebody heating and pressure pro les. This result is explained in terms of the pertinent aerothermodynamicsof the  ow eld and the vehicle’s geometry.When applied to the afterbody thermal protection system, these results show that the traditional afterbody heatshield design approach is nonconservative for the Stardust sample return capsule shape and entry conditions.

Aerothermodynamics of the Stardust Sample Return Capsule

▪ Abstract A review of planetary-entry gas dynamics is presented. Evolution of a blunt-body flowfield from a free molecular flow environment to a continuum environment is described. Simulations of near-wake flow phenomena, important for defining aerobrake payload environments, are also discussed. Some topics to be highlighted include aerodynamic coefficient predictions with emphasis on high-temperature gas effects; surface heating and temperature predictions for thermal protection system (TPS) design in a high-temperature, thermochemical nonequilibrium environment; and thermochemical models required for numerical flow simulation. Recent applications involving atmospheric entry into Jupiter (Galileo), Mars (Pathfinder and Global Surveyor), and a planned mission in which dust from the tail of a comet will be returned to Earth (Stardust) will provide context for this discussion.

Planetary-Entry Gas Dynamics

The TITANprogramforpredicting charringmaterial ablationand shapechangeof thermal protectionmaterials is presented. The governing equations include energy conservation and a three-component decomposition model. The surface energy balance condition is solved with a moving grid to calculate the shape change due to surface recession. The governing equations are discretized with a  nite volume approximation with a general body- tted coordinate system. A time-accurate solution is achieved by an implicit time-marching techniquewith Gauss–Seidel line relaxationwith alternating sweeps. Benchmark solutionsare calculated and comparedwith availablesolutions to check code consistency and accuracy. For fully coupled solid– uid simulation, this technique has been directly integrated with both a high- delity Navier–Stokes solver and an aerothermal  ow eld engineering correlation code. Representative computations, including a slender hypersonic reentry vehicle and a  at-faced cylinder model in an arcjet test, are presented and discussed in detail.

Two-Dimensional Implicit Thermal Response and Ablation Program for Charring Materials

The Mars Pathe nder probe contained instrumentation that measured heatshield temperatures during entry. A description of the experiment, the data, and an analysis of the entry environment and material response are presented. Navier ‐Stokes forebody heating calculations show a peak unblown radiative-equilibrium heat e ux of 118W/cm 2 at thestagnation point and120 W /cm 2 on theshoulderforturbulente ow. Theheatload is3.8 kJ /cm 2 on thenose,decreases along thefrustum,then increasesto 2.7 ‐3.1kJ/cm 2 on theshoulder depending on the onset time forturbulence. One-dimensional charringmaterialresponseiscalculated using threedifferentmodels.Stagnationpoint temperature data are consistent with about 85% of fully catalytic laminar heating. Shoulder temperature data are inconclusive, but are consistent with fully catalytic laminar heating or with 85% of fully catalytic heating with early onset of turbulence. Aft temperature data indicate a peak heat e ux and heat load of about 1.3 W /cm2 and 70 J/cm 2 , respectively. The aft heating proe le is about 20 s longer than the forebody heating proe le. Bondline temperaturedata, although not useful forquantitative analysis of aerothermal heating, clearly showtheheatshield had adequate thickness margins for the actual entry.

Mars pathfinder entry temperature data, aerothermal heating, and heatshield material response

A formulation of finite rate ablation surface boundary conditions, including oxidation, nitridation, and sublimation of carbonaceous material with pyrolysis gas injection, based on surface species mass conservation, has been developed. These surface boundary conditions are discretized and integrated with a Navier-Stokes solver. This numerical procedure can predict aerothermal heating, chemical species concentration, and carbonaceous material ablation rates over the heat-shield surface of reentry space vehicles. Two finite rate gas-surface interaction models, based on the work of Park and of Zhluktov and Abe, are considered. Three test cases are studied. The stream conditions of these test cases are typical for Earth reentry from a planetary mission with both oxygen and nitrogen fully or partially dissociated inside the shock layer. Predictions from both gas-surface interaction models are compared with those obtained by using chemical equilibrium ablation tables. Stagnation point convective heat fluxes predicted by using Park's finite rate model are usually below those obtained from chemical equilibrium tables and Zhluktov and Abe's model. Recession predictions from Zhluktov and Abe's model are usually lower than those obtained from Park's model and from chemical equilibrium tables. The effect of species mass diffusion on the predicted ablation rate is also examined.

Y.-K. Chen

Papers

Ablation and Thermal Response Program for Spacecraft Heatshield Analysis

Aerothermodynamics of the Stardust Sample Return Capsule

Two-Dimensional Implicit Thermal Response and Ablation Program for Charring Materials

Mars pathfinder entry temperature data, aerothermal heating, and heatshield material response

Navier-Stokes Solutions with Finite Rate Ablation for Planetary Mission Earth Reentries