Y.-K. Chen

Phenolic Impregnated Carbon Ablators (PICA) for Discovery class missions

This paper presents the development of the light weight Phenolic Impregnated Carbon Ablators (PICA) and its thermal performance in a simulated heating environment for planetary entry probes. PICA material was developed as a member of the Light Weight Ceramic Ablators (LCAs) family, and since then, the manufacturing process of this material was significantly unproved. The density of PICA material ranges from 0.224 to 0.321 g/cc having uniform resin distribution within the fibrous substrate. Surface densification was also developed to improve the ablation characteristics of PICA against extremely high stagnation pressures. The thermal performance of PICA was evaluated in the Ames arc jet facility at cold wall heat fluxes from 425 to 3360 W/cm and surface pressures of 0.1 to 0.43 attn. Heat loads used in these tests varied from 6,245 to 33,600 J/cm and are representative of the entry conditions of several proposed Discovery missions. Surface and in-depth temperatures were measured by using optical pyrometers and thermocouples. Surface recession was also measured by using a template and a height gage. The ablation characteristics and efficiency of the PICA is quantified by using the effective heat of ablation, and the thermal penetration response is evaluated by the thermal soak data. In addition, comparison of the thermal performance of standard and surface densified PICA is also discussed.

Forebody TPS sizing with radiation and ablation for the Stardust Sample Return Capsule

The development of a new high-fidelity methodology for predicting entry flows with coupled radiation and ablation is described. The prediction methodology consists of an axisymmetric, nonequilibrium, Navier-Stokes flow 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 at seven trajectory points using ablating and nonablating boundary conditions. These flow simulations are used to size and design the Stardust forebody heatshield and develop arc-jet test conditions and models. The paper describes details of the methodology, results from the flow simulations, and, various heatshield design issues.

Graphite Ablation and Thermal Response Simulation Under Arc -Jet Flow Conditions

The Two -dimensional Implicit Ther mal Response and Ablation program (TITAN) was developed and integrated with a Navier -Stokes solver (GIANTS) for multi - dimensional ablation and shape change simulation of thermal protection systems in hypersonic flow environments. The governing equations in both codes are discretized using the same finite -volume approximation with a general body -fitted coordinate system. Time - dependent solutions are achieved by an implicit time marching technique using Gauss -Siedel line relaxation with alternating sweeps. As the first part of a code validation study, this paper compares TITAN -GIANTS predictions with thermal response and recession data obtained from arc -jet tests recently conducted in the Interaction Heating Facility at NASA Ames Research Center. The test mod els are graphite sphere -cones. Graphite was selected as a test material to minimize the uncertainties from material properties. Recession and thermal response data were obtained from two separate arc -jet test series. The first series was at a heat flux whe re graphite ablation is mainly due to sublimation, and the second series was at a relatively low heat flux where recession is the result of diffusion -controlled oxidation. Ablation and thermal response solutions for both sets of conditions, as calculated by TITAN -GIANTS, are presented and discussed in detail. Predicted shape change and temperature histories generally agree well with the data obtained from the arc -jet tests. Nomenclature

High-Fidelity Charring Ablator Thermal Response Model

Low-density carbon/phenolic is a class of ablative materials that is attractive for space exploration missions that use blunt bodies where weight and performance of the material are of primary importance, but shape preservation is not critical. We consider a relatively simple class, PICA, that consists of carbon fibers impregnated with phenolic as the matrix. A new formulation for models of the response of this class of materials to high-enthalpy environments is summarized. The new formulation consists of conservation equations for species, mass, and energy in porous media. The velocity is obtained using Darcy’s law with the pressure obtained so that mass is conserved. Pyrolysis of the matrix is modeled using a discrete number of progress variables representing the decomposition reaction stages. Each decomposition reaction produces its own set of species. The one-dimensional equations are solved by discretizing in space using a second-order staggered mesh on a moving grid, and an implicit dual time step scheme is used to advance the solution in time. The Charring Ablator Thermal response (CAT) code that implements the formulation is tightly coupled to a chemistry code that enables handling equilibrium as well as finite rate chemistry. It was thoroughly verified against analytical solutions and comparisons of results to results using different numerical methods and techniques. Sample verification cases are summarized showing excellent accuracy. Sample material response cases are presented showing the capability of the code. We have established that models of this type of ablative materials are highly sensitive to the energy and chemistry balance at the gas surface interface and that these balances are significantly related to material performance. Unfortunately, highquality data for the decomposition of phenolic is missing and is needed to enable validation of the formulation.

Ablation and thermal response program for spacecraft heatshield analysis

A fully implicit ablation and thermal response program has been developed for the simulation of one-dimensional transient transport of thermal energy in a multilayer stack of isotropic materials and structure which can ablate from a front surface and decompose in-depth. Equations and numerical procedures for solution are described. Solutions are compared with those of Aerotherm Charring Material Thermal Response and Ablation Program, and with the arcjet data. The code is numerically more stable, and solves much wider range of problems compared with the existing explicit code. Applications of the code for the analysis of aeroshell heatshields of Stardust, Mars 2001, and Mars Microprobe using the advanced Light Weight Ceramic Ablators developed at the NASA Ames Research Center are presented and discussed in detail.

Wake flow calculations with ablation for the Stardust Sample Return Capsule

The aerothermal database generated for the design of the Stardust Sample Return Capsule afterbody heatshield is described. Using high-fidelity Navier-Stokes wake flow simulations with and without forebody ablation products, unique afterbody heating and pressure profiles along Stardusts 12.6 krn/s Earth atmospheric entry trajectory are calculated. The flow simulations indicate that the afterbody heating and pressure profiles in time are significantly different than the forebody pressure and heating profiles; the afterbody heat pulse is longer, producing a higher integrated heat load, and, the peak values in the pressure and the heating occur later in the trajectory than on the forebody. In light of these results, traditional afterbody heatshield design methodology is shown to be nonconservative for the Stardust sample return capsule shape and entry conditions. The afterbody aerothermal environment is explained in terms of the relevant flow physics. The computations are shown to agree qualitatively with Viking base pressure flight data.

Coupled CFD-Ablation Response Model Simulations using the libMesh Framework

The DPLR Navier-Stokes flow solver is coupled to two Ablation Response Model (ARM) codes, CHAR and TITAN, using a modular approach where a central handler code runs the analysis codes iteratively and passes the required boundary condition values back and forth between the codes. The handler code is based on the libMesh software libraries. The libMesh mesh-free interpolation routines allow for the coupled analysis codes to be written in different programming languages and for the boundary point data to be non-point-matched. The boundary data is interpolated using a K-D Tree mesh-free interpolation approach. The basic execution flow for the coupled DPLR-ARM code is presented, and the coupled DPLR-ARM code is applied to arc jet test cases. The libMesh mesh-free interpolation successfully transferred the required boundary condition data between the fluid dynamic and material response codes, and the coupled DPLRARM surface recession rates matched experimental measurements as well as previous computations.

TPS requirements for advanced sample return capsule design

A preliminary trade study of the TPS design for a Mars sample return capsule was performed. The trade study assumed a constant mass of 25 kg and a relative entry velocity of 12.4 km/s. To map out the design space for a Mars sample return mission, four classes of geometries were examined with three entry angles and ballistic coefficients ranging from 18-87 kg/m. A total of 27 point designs and TPS sizing calculations for 9 different geometries were generated. For TPS design, a minimum in TPS mass as a function of ballistic coefficients was identified. It was found that shapes such as a spherical section, a 70° sphere-cone, or a 60° sphere-cone, with a moderate ballistic coefficient near 50 kg/m minimized the required forebody TPS mass. Major topics discussed in the paper are the analysis methodology, the criteria for material selection, design uncertainties, entry environments, and results from the TPS point design calculations.