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Molecular simulations with a resolved surface microstructure were performed for high-temperature dissociated oxygen reacting with a carbon-carbon composite material. The direct simulation Monte Carlo method was used to simulate the convection and diffusion of reactants toward the microstructure and the transport of surface reaction products away from the microstructure. Simulations were performed with and without gas-phase chemical reactions to determine the relative importance of gas-surface reactions compared to gas-phase reactions next to the material surface. The simulations incorporated reaction probabilities for individual gas-surface collisions based on new reactive scattering data obtained from molecular beam experiments. The experiments provide detailed dynamical data on the scattering of O, O2, CO, and CO2 from a representative material sample, made of vitreous carbon. These experiments indicate that a majority of surface reaction products were produced through thermal mechanisms. For a gas-surface temperature of 800 K, it was found that despite CO being the dominant surface reaction product, a gas-phase reaction forms significant CO2 within the microstructure region. The amount of CO2 production within the microstructure region was shown to be dependent on the local Knudsen number, based on the exposed microstructure height. Finally, preliminary simulations were performed for a real carbon-carbon surface. The surface topography is obtained through x-ray microtomography of an ablated carbon-carbon sample, which is triangulated and used directly within a direct simulation Monte Carlo simulation of the gas-surface interaction. [ABSTRACT FROM AUTHOR]

Two sets of finite-rate gas/surface interaction model between air and the carbon surface are studied. The first set is an engineering model with one-way chemical reactions, and the second set is a more detailed model with two-way chemical reactions. Each of these two proposed models intends to cover the carbon surface ablation conditions including the low-temperature rate-controlled oxidation, the midtemperature diffusion-controlled oxidation, and the high-temperature sublimation. The prediction of carbon surface recession is achieved by coupling a material thermal response code and a Navier-Stokes flow code. The material thermal response code used in this study is the Two-Dimensional Implicit Thermal-Response and Ablation Program, which predicts charring material thermal response and shape change on hypersonic space vehicles. The flow code solves the reacting full Navier-Stokes equations using the data parallel line relaxation method. Recession analyses of stagnation tests conducted in NASA Ames Research Center arc-jet facilities with heat fluxes ranging from 45 to 1100W/cm² are performed and compared with data for model validation. The ablating material used in these arc-jet tests is phenolic impregnated carbon ablator. Computational predictions of surface recession and shape change are in good agreement with measurement for arc-jet conditions. [ABSTRACT FROM AUTHOR]

Thermal protection system design for atmospheric reentry vehicles remains a challenging and complex problem. Recent advances in computational modeling of air-carbon interactions consider competing finite-rate reactions on a limited number of available surface sites. One of the most advanced kinetic models is due to Zhluktov and Abe. However, the Zhluktov and Abe model only describes the oxidation and sublimation of carbon and has no nitridation mechanism. The following study develops several modifications to the Zhluktov and Abe air-carbon model that account for all three reaction mechanisms with the goal of improving cyanogen shock-layer radiation predictions to recent experimental results. First, the study examines two possible paths for carbon nitridation and then assesses the augmented surface reaction model in a representative blunt-body reentry flow. Second, a sensitivity analysis is performed to determine which surface reactions have the most impact on altering cyanogen radiance predictions. It was found that modifications to the oxidation mechanisms significantly improved the comparisons and the nitridation mechanism had a minor effect. Specifically, it was determined that carbon monoxide must be the principal oxidation product at high surface temperatures instead of carbon dioxide, as the Zhluktov and Abe model had originally predicted. More detailed measurements of both the carbon oxidation and nitridation reactions at higher surface temperatures are required to further validate and improve the rate parameters derived in this study. [ABSTRACT FROM AUTHOR]

Vehicles entering planetary atmospheres at high speed require an ablative heat shield to withstand the high thermal energy flux to the body. The interaction between the ablative products and the flowfield is not well characterized. In this study, numerical simulations were conducted to investigate the influence of carbon ablation on shock-layer radiation. Data collected from experiments performed in the X-2 expansion tunnel at the University of Queensland were used to compare to the simulations. The model was a short half-cylinder made of isomolded graphite and was tested in 8.6 km/s8.6km/s Earth entry flow. The graphite model was heated within a temperature range of 1770-2410 K over the course of the experimental campaign. The radiation emitted from the CN violet bands was measured by ultraviolet spectrometry in a spectral range from 353 to 391 nm. The simulations used the Park as well as Zhluktov and Abe finite-rate surface kinetic models for determining the chemical state of the ablating boundary layer. The Park model consistently overpredicted the radiative heat fluxes, but better comparison was achieved when a modified nitridation rate was used. The Zhluktov and Abe predictions displayed good agreement with the measured radiative heat fluxes at 1770 K but remained well below the measured values at the higher wall temperatures. The lack of a nitridation mechanism in the Zhluktov and Abe model is suggested as the reason for these underpredictions. [ABSTRACT FROM AUTHOR]

Results from an in situ analysis of ablative heat shield materials based on stereoscopic photogrammetry during plasma wind tunnel testing are presented. A carbon fiber preform material sample has been investigated in a highenthalpy airflow corresponding to a Hayabusa reentry condition at a 78 km altitude. The applied photogrammetry uses two digital single lens reflex cameras triggered at 3 frames/s to acquire images simultaneously from two different perspectives of the sample exposed to the high-speed flow. The photogrammetric analysis resolves the surface with 25,000 px/cm², approximately 400 dpi, allowing the in situ analysis of the recession phenomena of ablators. Material inhomogeneities are identified during the ablation process, and surface recession is derived with an accuracy of 21 μm. The method is a useful tool for the investigation of temporally resolved volumetric ablation phenomena in low-density carbon phenolic materials. [ABSTRACT FROM AUTHOR]

A new conformal type of phenolic impregnated carbon ablator was manufactured by resin impregnation of carbon felt. Based on property measurements of the conformal material and the existing model for a standard, rigid phenolic impregnated carbon ablator, a midfidelity material response model for the conformal material was developed. The rigid and conformal materials were arcjet tested simultaneously on a sphere-cone geometry in several environments with frustum heat flux up to 380 W/cm². Good agreement between the predictions and data was obtained for recession, surface temperature, and in-depth temperatures. The conformal material has a slightly greater ablation rate but significantly lower thermal diffusivity than the rigid material. [ABSTRACT FROM AUTHOR]

The steps necessary to achieve the strong coupling between a flowfield solver and a material response solver are presented. This type of coupling is required to accurately capture the complex aerothermodynamic physics occurring during hvpersonic atmospheric entries. A blowing boundary condition for the flowfield solver is proposed. This allows the ablating gas calculated by the material response solver to be correctly injected in the boundary layer. A moving mesh algorithm for the flowfield solver that implicitly enforces the geometric conservation law is presented. Using that capability, a mesh movement procedure for surface recession and for accurate shock capturing is proposed. The entire technique is tested using a material response solver with surface ablation and pyrolysis coupled to a hypersonic solver for weakly ionized flows in thermochemieal nonequilibrium. Results using the reentry trajectory of the IRV-2 test vehicle are presented, showing that the surface heat fluxes remain accurate as the vehicle geometry and freestream conditions change. [ABSTRACT FROM AUTHOR]

This study demonstrates that coupling of a material thermal response code and a flow solver with nonequilibrium gas-surface interaction for simulation of charring carbon ablators can be performed using an implicit approach. The material thermal response code used in this study is the three-dimensional version of fully implicit ablation and thermal response program, which predicts charring material thermal response and shape change on hypersonic space vehicles. The flow code solves the reacting Navier-Stokes equations using data-parallel line relaxation method. Coupling between the material response and flow codes is performed by solving the surface mass balance in the flow solver and the surface energy balance in the material response code. Thus, the material surface recession is predicted in the flow code, and the surface temperature and pyrolysis gas injection rate are computed in the material response code. It is demonstrated that the time-lagged explicit approach is sufficient for simulations at low surface heating conditions, in which the surface ablation rate is not a strong function of the surface temperature. At elevated surface heating conditions, the implicit approach has to be taken because the carbon ablation rate becomes a stiff function of the surface temperature, and thus the explicit approach appears to be inappropriate, resulting in severe numerical oscillations of predicted surface temperature. Implicit coupling for simulation of arc-jet models is performed, and the predictions are compared with measured data. Implicit coupling for trajectory-based simulation of Stardust forebody heat shield is also conducted. The predicted stagnation point total recession is compared with that predicted using the chemical equilibrium surface assumption. [ABSTRACT FROM AUTHOR]

Three-dimensional ablation measurements, acquired using solid carbon dioxide for the model material and carried out at Mach 3, are compared to predictions of a computational model that incorporates fluid dynamics and solid-body response. Models were formed in situ by mounting three-dimensional printed molds on a hollow sting, through which liquid carbon dioxide flowed. The blowdown wind tunnel, operating with a vacuum tank connected to its outlet, provided stagnation pressures through a wide range (6.0 to 45 psia) while maintaining a nearly constant stagnation temperature for all tests (~295 K). Three synchronized high-speed cameras captured the model recession from different angles while multiple low-power lasers projected structured arrays of dots onto the otherwise visually uniform model. Commercially available software was applied to quantify the recession data. As expected, recession rates were largest at the nose of each model and increased with stagnation pressure. A quantitative comparison of the recession rate at the stagnation point showed very good agreement between the experiment and the computational model. [ABSTRACT FROM AUTHOR]

A finite element ablation and thermal response program for the simulation of three-dimensional transient thermostructurai analysis has been developed. The three-dimensional governing differential equations and finite element formulation are summarized. A novel probabilistic design methodology for thermal protection systems has been developed. The design methodology is an eight step process beginning with a parameter sensitivity study and is followed by a deterministic analysis whereby an optimum design can determined. The design process concludes with a Monte Carlo simulation where the probabilities of exceeding design specifications are estimated. The design methodology is demonstrated by applying the methodology to the carbon phenolic compression pads of the Crew Exploration Vehicle. The maximum allowed values of bondline temperature and tensile stress are used as the design specifications in this study. Using the design methodology, the probability of exceeding these design specifications was shown to be low and within an acceptable range. [ABSTRACT FROM AUTHOR]

The fully implicit ablation and thermal response program was upgraded by incorporation of key features from the multicomponent ablation thermochemistry code. The upgraded program now has expanded capabilities to compute the muitispecies surface chemistry and the ablation rate as part of the surface energy balance. This new methodology eliminates B' tables, provides blown species as a function of time, and enables calculations that would otherwise be impractical (tables with four or more independent variables) or impossible (such as stacked pyrolyzing ablators). Equations and solution procedures are presented, then representative calculations of equilibrium and non- equilibrium ablation in flight and ground-test environments are discussed. [ABSTRACT FROM AUTHOR]

In previous work, an equilibrium ablation and thermal response model for Phenolic Impregnated Carbon Ablator was developed. In general, over a wide range of test conditions, model predictions compared well with arcjet data for surface recession, surface temperature, in-depth temperature at multiple thermocouples, and char depth. In this work, additional arcjet tests were conducted at stagnation conditions down to 40 W/cm² and 1.6 kPa. The new data suggest that nonequilibrium effects become important for ablation predictions at heat flux or pressure below about 80 W/cm² or 10 kPa, respectively. Modifications to the ablation model to account for nonequilibrium effects are investigated. Predictions of the equilibrium and nonequilibrium models are compared with the arcjet data. [ABSTRACT FROM AUTHOR]

A procedure has been developed to couple a hypersonic reacting flow model, a radiative heat transfer model, and a surface ablation model to study the surface heat transfer and surface ablation rate of atmospheric reentry vehicles. The two-way loose-coupling algorithm is described for each of the models, as is the solution procedure to achieve convergence. Observations on the challenges of the loose-coupling strategy are given. Representative results are presented for two-dimensional benchmark examples and for three-dimensional flow at an angle of attack past a symmetric capsule based on the Crew Exploration Vehicle reentry vehicle. Effects due to the interaction with radiation and ablation are shown for two quantities of interest: the predicted peak surface heat flux and the ablation rate on the vehicle heat shield. Uncertain parameters are identified in each of the submodels, and a preliminary parameter sensitivity study is carried out by varying these values to examine their effects on the heat transfer and ablation rates in the coupled problem. [ABSTRACT FROM AUTHOR]

A multiscale approach is used to model and analyze the ablation of porous materials. Models are developed for the oxidation of a carbon preform and of the char layer of two phenolic impregnated carbon ablators with the same chemical composition but with different structures. Oxygen diffusion through the pores of the materials and in depth oxidation and mass loss are first modeled at the microscopic scale. The microscopic model is then averaged to yield a set of partial differential equations describing the macroscopic behavior of the material. Microscopic and macroscopic approaches are applied with progressive degrees of complexity to gain a comprehensive understanding of the ablation process. Porous medium ablation is found to occur in a zone of the char layer that we call the ablation zone. The thickness of the ablation zone is a decreasing function of the Thiele number. The studied materials are shown to display different ablation behaviors, a fact not captured by current models that are based on chemical composition only. Applied to Stardust's phenolic impregnated carbon ablator, the models explain and reproduce the unexpected drop in density measured in the char layer during Stardust postflight analyses [Stackpoole, M., Sepka, S., Cozmuta, I., and Kontinos, D., "Post-Flight Evaluation of Stardust Sample Return Capsule Forebody Heat-Shield Material," AIAA Paper 2008-1202, Jan. 2008]. [ABSTRACT FROM AUTHOR]

The central focus of this study is to demonstrate that time-accurate solutions for multidimensional ablation and shape change of thermal protection system materials may be obtained by loose coupling of a high-fidelity flow solver with a material thermal response code. In this study, the flow code solves the nonequilibrium Navier-Stokes equations using the data-parallel line-relaxation (DPLR) method. The material response code is the latest version of the Two-dimensional Implicit Thermal Response and Ablation Program (TITAN). In TITAN, the governing equations, which include a three-component decomposition model and a surface energy balance with thermochemical ablation, are solved with a robust moving-grid scheme to predict the shape change caused by surface recession. Coupling between the material response and flow codes is required for many multidimensional ablation simulations, because the magnitude and distribution of the surface heat flux are very sensitive to shape change. This paper demonstrates the application of the TITAN-DPLR system to problems with large-scale recession and shape change. Ablation and thermal response simulations are presented for iso-q and flat-faced arc-jet test models and also for a wedge with a cylindrical leading edge exposed to hypersonic flow at various angles of attack. [ABSTRACT FROM AUTHOR]

The reentry of the Stardust sample return capsule was captured by several optical instruments through an observation campaign aboard the NASA DC-8 airborne observatory. Flow environments obtained from computational fluid dynamics solutions are loosely coupled with material response modeling to predict the surface temperature and the observed continuum emission of Stardust throughout the reentry. The calculated surface temperatures are compared with the data from several spectral instruments onboard the airborne observatory, including the ECHELLE (echelle-based spectrograph for the crisp and high efficient detection of low light emission) camera and conventional spectrometer in Czerny-Turner configuration. The ECHELLE camera recorded spectral intensity at a period in the trajectory before peak heating. The graybody curves corresponding to the average and area-averaged surface temperatures predicted by the computational fluid dynamics and material response coupled simulation have excellent agreement with the recorded data at altitudes lower than 74 km. At these altitudes, the computational fluid dynamics and material response coupling agrees with the surface temperature to within 50 K. The computational fluid dynamics calculation without the material response modeling overestimates surface temperatures because it does not take into account such things as ablation. The overprediction of the computational fluid dynamics and material response simulated surface temperature early in the trajectory coincides with high-emission intensity lines corresponding to thermal paint products. The presence of paint on the heat shield could have contributed to the lower observed surface temperatures and could explain the overprediction by the simulated data, which does not account for the paint. The average surface temperatures resulting from the spectrometer in Czerny-Turner configuration telescope analysis agree to within less than 5% with the average surface temperatures predicted by the material response. This observation period included the point of peak heating. The calculated flux based on the surface temperature agrees well with the observed flux. Surface temperature is one of the critical parameters used in the design of thermal protection systems, because it is an indicator of material performance. The coupled computational fluid dynamics and material response approach employed in the present analysis increases confidence for future missions such as the crew exploration vehicle Orion. [ABSTRACT FROM AUTHOR]

A high-fidelity three-dimensional approach is developed to simulate emission signatures from the shock layer around the Stardust sample return capsule at several points on its best-estimated trajectory. Calculations are performed with various gas chemistry, thermodynamic, and radiation models. Results are compared against calibrated imaging data acquired by a slitless echelle spectrometer. The present analysis is based on flowfields computed without the inclusion of ablation products, and the comparison is focused on radiation from atomic oxygen and nitrogen lines. The purpose is to apply and improve, if necessary, the current models used in nonequilibrium atmospheric entry simulations. [ABSTRACT FROM AUTHOR]

The article presents a study which examines an ablating surface in expansion tunnel flow. Researchers conducted a one-dimensional semi-finite analysis in order to examine whether the surface temperature would be enough to produce vaporization of the insulation. They created an ablating shock layer with an epoxy coating on an axisymmetric model during steady test flow in a high-enthalpy superorbital expansion tunnel. Chemical reactions between the epoxy ablation gases and the air test gas produced high levels of cyanogen.

A study of graphite ablation in reentry flows is carried out by a surface ablation approach integrated with a two-dimensional axisymmetric Navier-Stokes solver. The gas-surface interaction model is based on chemical equilibrium ablation with surface mass and energy balances fully coupled with the numerical solver and can account for both surface oxidation and sublimation. The surface temperature is obtained from the steady-state ablation approximation. This numerical procedure can predict aerothermal heating, chemical species concentrations, and carbon material ablation rate over the heat-shield surface of reentry vehicles. Two-dimensional axisymmetric simulations have been performed to numerically reproduce the ablation of a graphite sphere cone that has been tested in the Interaction Heating Facility at the NASA Ames Research Center. The freestream conditions of the selected test case are typical for Earth reentry from a planetary mission. The predicted ablation rate and surface temperature assuming frozen chemistry in the flow show a good agreement with the available experimental data. The agreement is further improved freezing the nitrogen recombination reaction at the surface to be more consistent with experimental observation, which has shown nitrogen atom recombination not to occur at the graphite surface. [ABSTRACT FROM AUTHOR]

A description of models and boundary conditions required for loose coupling of radiation and ablation physics to a hypersonic flow simulation is provided. Chemical equilibrium routines for varying elemental mass fractions are required in the flow solver to integrate with the equilibrium chemistry assumption employed in the ablation models. The capability also enables an equilibrium catalytic wall boundary condition in the nonablating case. The paper focuses on numerical implementation issues using FIRE II, Mars return, and Apollo 4 applications to provide context for discussion. Variable relaxation factors applied to the Jacobian elements of partial equilibrium relations required for convergence are defined. Challenges of strong radiation coupling in a shock capturing algorithm are addressed. Results are presented to show how the current suite of models responds to a wide variety of conditions involving coupled radiation and ablation. [ABSTRACT FROM AUTHOR]

The article examines the effects of surface ablation and blowing on hypersonic boundary-layer transition based on summarized experimental data. It cites the impact of ablation on transition during hypersonic flight tests. Previous studies on symmetric flowfields without lift, most of which are slender cones at zero angle of attack (AOA), wherein the crossflow instability is not a factor are cited. It concludes that experimental data are more generally available for blowing-induced transition on slender cones near zero AOA.

The purpose of this paper is to describe and demonstrate new capabilities in the two-dimensional implicit thermal response and ablation program. These expanded capabilities include grid options for flight and arcjet geometries, a sizing algorithm for the flight-type geometry, and an orthotropic thermal conductivity model: Applications to analysis of an orthotropic low-density carbon-phenolic material in arcjet and flight environments relevant to the Orion crew module are presented. For the arcjet environment, multidimensional conduction effects strongly influence the in-depth thermal response. For a lunar return flight environment, in the shoulder region of the crew module (where the radius of curvature is smallest), the thermal response is influenced by multidimensional conduction and by the orientation of the orthotropic material. [ABSTRACT FROM AUTHOR]

The Inflatable Reentry and Descent Technology flight performed in 2006 was not nominal. The flight was characterized by a blackout period shorter than expected. Here, we apply two methods that can be used to perform postflight analysis. The first, based on an engineering approach, has already been used for the preflight analysis. The second involves coupled calculations between a computational fluid dynamics and an electromagnetic solver. The main objective of this postflight analysis is to validate this advanced approach using flight data and to assess the validity of the engineering method. Numerical results show that the engineering method overpredicts the blackout duration due to the fact that this method is based on a severe on/off switch. The coupled approach leads to an underestimation of the blackout period. Several uncertainties, such as the influence of ablative material on ionization, could explain this discrepancy. Analysis establishes the validity of the coupled approach, but its accuracy depends strongly on the modeling used for the calculations. [ABSTRACT FROM AUTHOR]

Thermal protection materials are required to protect structural components of space vehicles during the reentry stage, missile launching systems, and solid rocket motors. Novel materials based on nanotechnology creating nontraditional ablators are rapidly changing the technology base for thermal protection systems. In this study, different polymer nanocomposite compositions were created by melt-compounded montmorillonite nanoclays or carbon nanofibers in a neat thermoplastic polyurethane elastomer polymer using twin-screw extrusion. The kinetic and thermophysical properties that are required to analyze the ablation characteristics were measured for selective thermoplastic polyurethane materials. Properties of the nanomodified systems were then compared against those of the current state-of-the-art insulation material, Kevlar®-filled ethylene-propylene-diene monomer and the neat thermoplastic polyurethane elastomer, for the investigation of kinetic parameters using the isoconversion technique. Based on temperatures at peak weight-loss rate, Kevlar-filled ethylene-propylene-diene monomer outranked the proposed formulations at all heating rates. Recognizing that ablation performance is a complex function of kinetic and thermophysical/transport properties, a surrogate test for ablation performance was investigated. Samples of neat and nanomodified thermoplastic elastomer were run in an oxygen-consumption (cone) calorimeter. The peak heat release rates of the nanomodified samples were substantially less than that of the neat thermoplastic elastomer. [ABSTRACT FROM AUTHOR]

Using the coupled ablation and radiation capability recently included in the LAURA flowfield solver, this paper investigates the influence of ablation on the shock-layer radiative heating for Earth entry. The extension of the HARA radiation model, which provides the radiation predictions in LAURA, to treat a gas consisting of the elements carbon, hydrogen, oxygen, and nitrogen is discussed. It is shown that the absorption coefficient of air is increased with the introduction of the carbon and hydrogen elements. A simplified shock-layer model is studied to show the impact of temperature, as well as the abundance of carbon and hydrogen, on the net absorption or emission from an ablation contaminated boundary layer. It is found that the ablation species reduce the radiative flux in the vacuum ultraviolet, through increased absorption, for all temperatures. However, in the infrared region of the spectrum, the ablation species increase the radiative flux, through strong emission, for temperatures above 3000 K. Thus, depending on the temperature and abundance of ablation species, the contaminated boundary layer may either provide a net increase or decrease in the radiative flux reaching the wall. To assess the validity of the coupled ablation and radiation LAURA analysis, a previously analyzed Mars-return case (15.24 km/s), which contains significant ablation and radiation coupling, is studied. Exceptional agreement with previous viscous-shock-layer results is obtained. A 40% decrease in the radiative flux is predicted for ablation rates equal to 20% of the freestream mass flux. The Apollo 4 peak-heating case (10.24 km/s) is also studied. For ablation rates up to 3.4% of the freestream mass flux, the radiative heating is reduced by up to 19%, whereas the convective heating is reduced by up to 87 %. Good agreement with the Apollo 4 radiometer data is obtained by considering absorption in the radiometer cavity. For both the Mars-return and the Apollo 4 cases, coupled radiation alone is found to reduce the radiative heating by 30-60% and the convective heating by less than 5%. [ABSTRACT FROM AUTHOR]

Much has been learned about the physics underlying the transition process at supersonic and hypersonic speeds through years of analysis, experiment, and computation. The application has been principally to simple shapes like plates, cones, and spherical nose tips. But the shapes of the new entry vehicles are not simple. They will invariably be at angle of attack, and so three-dimensional effects will be very important, as will roughness effects due to ablation. This paper will review the physics basis of our present understanding of the transition process. Further, because of the complex geometries, it will address the need for careful experimental work as per the guidelines enunciated years ago by the U.S. Transition Study Group. Following these guidelines is essential to obtaining reliable, usable data for use in refining transition estimation techniques. [ABSTRACT FROM AUTHOR]

Numerical modeling is employed to study the heat transfer modulation between the thermal protection shield and the gas flow that is caused by ejection of underexpanded pyrolysis gases through the cracks in the thermal protection shield. The simulations are performed for an axisymmetric bluff body flying at Mach 7. The influence of the geometry of the thermal protection shield on the heat transfer pattern is studied for two representative shapes. The results are presented for three different flight altitudes (low, ground level; moderate, 20 km; and high, 30 km). At the low altitude, the plume pressure is lower than the pressure behind the detached front shock wave and the plume propagates slowly along the wall surface. At high and moderate altitudes, the plume path (and, consequently, the convective heat transfer between the thermal protection shield and the plume) depends on the plume interaction with the bow shock wave. The effect of viscosity for the plume injection conditions and freestream Mach number considered is found to be negligible at simulated altitudes. The effect of the initial pressure of pyrolysis gas on the plume dynamics is significant. The presence of the blast wave associated with the underexpanded plume alters the heat transfer and increases mixing. Finally, the enhanced heat transfer caused by the emergence of multiple plumes is investigated. [ABSTRACT FROM AUTHOR]

The ablation process of the Stardust thermal protection material is designed to reduce aerodynamic heating during reentry for extreme conditions. The coupling of ablation species with the flowfield is investigated in this work using the direct simulation Monte Carlo method for transitional to near-continuum flows. To model surface thermal and chemical ablation processes, a variable surface temperature wall is obtained assuming a radiative heat flux balanced by convective heat flux. It is found that chemical ablation due to the reaction between thermal protection system carbon materials and gaseous oxygen and nitrogen atoms is dominant compared with thermal ablation. As the altitude decreases, the forebody surface temperature increases, the ablation process becomes more intensive, and the influence of ionization reactions on the flowfield becomes more important due to denser freestream conditions. [ABSTRACT FROM AUTHOR]