A method for determining boundary node displacements for three-dimensional ablation solvers is developed. The approach is applicable to a wide range of mesh motion algorithms used to determine internal node displacements. The method leverages radial basis functions to analytically define the geometry to which the boundary nodes should be constrained. Constrained optimization, elasticity, and spring analogy solvers are utilized to redistribute the nodes on the evolved geometry surface while maintaining mesh quality. The method is implemented in the CHarring Ablator Response (CHAR) code, and a realistic demonstration case is presented on the Boundary Layer Transition (BOLT) flight experiment configuration. The method is shown to be robust for highly-curved geometries with large domain deformations.

An overview of the contact boundary conditions implemented in the CHarring Ablator Response (CHAR) code is presented. CHAR is a one-, two-, and three-dimensional unstructured continuous Galerkin finite element heat conduction and ablation solver with both direct and inverse modes. Special focus is put on the thermal and gas transport contact boundary conditions available in CHAR including finite contact conductance, thermal, and gas flow boundary conditions at contact interfaces. The derivation of gas flow contact terms to account for a multi-species pyrolysis gas mixture, including mass diffusion terms, is presented. Finally, verification of each contact boundary condition is carried out by comparing to analytical solutions or by employing the Method of Manufactured Solutions (MMS).

Surface heating measurements in shock tunnels operating at high enthalpy have previously showed anomalously high heating, the mechanisms for which have been the subject of much speculation. However, recent tests in JAXA’s High Enthalpy Shock Tunnel (HIEST) have presented compelling evidence that the heating anomaly is due to radiation from the shock layer, produced by impurities within the shock. The present paper describes test series performed in HIEST with a flat-faced cylinder model and an Apollo capsule model at angle of attack. The flat-faced cylinder is instrumented with three types of thermocouple heat flux gauges to separate contributions from radiative and convective heating. Most of the augmentation is shown to be due to radiation in this configuration. The Apollo capsule model is instrumented with a trip element and thermocouple gauges to study transition at high enthalpy, but transition data are obscured by the presence of radiation. Both test articles are embedded with fiber optics to measure the magnitude and characteristics of the radiative heating. Analyses of the radiation spectra, estimates of radiative heating corrections and comparison to computational fluid dynamic simulation of the shock tunnel are presented. The data show the radiative correction to be meaningful at high pressure or high enthalpy conditions; however, additional discrepancies remain.

An overview of the capabilities of the CHarring Ablator Response (CHAR) code is presented. CHAR is a one-, two-, and three-dimensional unstructured continuous Galerkin finite-element heat conduction and ablation solver with both direct and inverse modes. Additionally, CHAR includes a coupled linear thermoelastic solver for determination of internal stresses induced from the temperature field and surface loading. Background on the development process, governing equations, material models, discretization techniques, and numerical methods is provided. Special focus is put on the available boundary conditions including thermochemical ablation and contact interfaces, and example simulations are included. Finally, a discussion of ongoing development efforts is presented.

The EFT-1 Avcoat heatshield was instrumented with 34 plugs containing multiple thermocouples in-depth with an objective being to measure the flight aerothermal environment. This paper presents a discussion of the instrumentation and the techniques used to reconstruct the heating environment from the measured in-depth temperatures. The inverse heat transfer problem algorithms, models and assumptions will be outlined, and available results will be presented.

With the recent development of multi-dimensional thermal protection system (TPS) material response codes, the capability to account for surface-to-surface radiation exchange in complex geometries is critical. This paper presents recent efforts to implement such capabilities in the CHarring Ablator Response (CHAR) code developed at NASA's Johnson Space Center. This work also describes the different numerical methods implemented in the code to compute geometric view factors for radiation problems involving multiple surfaces. Verification of the code's radiation capabilities and results of a code-to-code comparison are presented. Finally, a demonstration case of a two-dimensional ablating cavity with enclosure radiation accounting for a changing geometry is shown.

This paper describes the design and results of a surface catalysis flight experiment flown on the Orion Multipurpose Crew Vehicle during Exploration Flight Test 1 (EFT1). Similar to previous Space Shuttle catalytic tile experiments, the present test consisted of a highly catalytic coating applied to an instrumented TPS tile. However, the present catalytic tile experiment contained significantly more instrumentation in order to better resolve the heating caused by the change in surface catalytic efficiency at the interface between two distinct materials. In addition to collecting data with unprecedented spatial resolution of the overshoot phenomenon, the experiment was also designed to prove if such a catalytic would be seen in turbulent flow in high enthalpy regimes. A detailed discussion of the results obtained during EFT1 is presented, as well as the challenges associated with data interpretation of this experiment. Results of material testing carried out in support of this flight experiment are also shown. Finally, an inverse heat conduction technique is employed to reconstruct the flight environments at locations upstream and along the catalytic coating. The data and analysis presented in this work will greatly contribute to our understanding of the catalytic overshoot phenomenon, and have a significant impact on the design of future spacecraft.

Reconstruction of flight aerothermal environments often requires the solution of an inverse heat transfer problem, which is an ill-posed problem of specifying boundary conditions from discrete measurements in the interior of the domain. This paper will present the algorithms implemented in the CHAR code for use in reconstruction of EFT-1 flight data and future testing activities. Implementation nuances will be discussed, and alternative hybrid-methods that are permitted by the implementation will be described. Results will be presented for a number of one-dimensional and multi-dimensional problems

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.

Laminar stagnation region heating augmentation is investigated in the AEDC Tunnel 9 at Mach 10 by performing high frequency surface pressure and heat transfer measurements on the Orion CEV capsule at zero degree angle-of-attack for unit Reynolds numbers between 0.5 and 15 million per foot. Heating augmentation increases with Reynolds number, but is also model size dependent as it is absent on a 1.25-inch diameter model at Reynolds numbers where it reaches up to 15% on a 7-inch model. Heat transfer space-time correlations on the 7-inch model show that disturbances convect at the boundary layer edge velocity and that the streamwise integral scale increases with distance. Therefore, vorticity amplification due to stretching and piling-up in the stagnation region appears to be responsible for the stagnation point heating augmentation on the larger model. This assumption is reinforced by the f(exp -11/3) dependence of the surface pressure spectrum compared to the f(exp -1) dependence in the free stream. Vorticity amplification does not occur on the 1.25- inch model because the disturbances are too large. Improved free stream fluctuation measurements will be required to determine if significant vorticity is present upstream or mostly generated behind the bow shock.

ven with great advances in computational techniques and computing power during recent decades, the modeling of unsteady separated flows, such as those encountered in the wake of a re-entry vehicle, continues to be one of the most challenging problems in CFD. Of most interest to the aerothermodynamics community is accurately predicting transient heating loads on the base of a blunt body, which would result in reduced uncertainties and safety margins when designing a re-entry vehicle. However, the prediction of heat transfer can vary widely depending on the turbulence model employed. Therefore, selecting a turbulence model which realistically captures as much of the flow physics as possible will result in improved results. Reynolds Averaged Navier Stokes (RANS) models have become increasingly popular due to their good performance with attached flows, and the relatively quick turnaround time to obtain results. However, RANS methods cannot accurately simulate unsteady separated wake flows, and running direct numerical simulation (DNS) on such complex flows is currently too computationally expensive. Large Eddy Simulation (LES) techniques allow for the computation of the large eddies, which contain most of the Reynolds stress, while modeling the smaller (subgrid) eddies. This results in models which are more computationally expensive than RANS methods, but not as prohibitive as DNS. By complimenting an LES approach with a RANS model, a hybrid LES/RANS method resolves the larger turbulent scales away from surfaces with LES, and switches to a RANS model inside boundary layers. As pointed out by Bertin et al., this type of hybrid approach has shown a lot of promise for predicting turbulent flows, but work is needed to verify that these models work well in hypersonic flows. The very limited amounts of flight and experimental data available presents an additional challenge for researchers. Recently, a joint study by NASA and CUBRC has focused on collecting heat transfer data on the backshell of a scaled model of the Orion Multi-Purpose Crew Vehicle (MPCV). Heat augmentation effects due to the presence of cavities and RCS jet firings were also investigated. The high quality data produced by this effort presents a new set of data which can be used to assess the performance of CFD methods. In this work, a hybrid LES/RANS model developed at North Carolina State University (NCSU) is used to simulate several runs from these experiments, and evaluate the performance of high fidelity methods as compared to more typical RANS models. .

This paper describes the addition of an overset grid capability to the DPLR flow solver for hypersonic flow in thermochemical nonequilibrium. Modifications to the preexisting flow solver were simplified through the use of DiRTlib, a “solver neutral” library of overset utilities. The new capability is demonstrated on a series of examples, including the Orion Crew Module and other reentry vehicles. For the overset grids used in these examples, the hole cutting and interpolation stencils were determined using SUGGAR, a generalized grid assembly code that can naturally accommodate both the three-dimensional and true two-dimensional cell-centered discretization schemes in DPLR. First a series of building-block examples are presented which highlight aspects of the new capability and assess the technique with comparisons to baseline, block-structured discretizations. The new capability is then exercised for the specific case of a tension tie geometry protruding from the Orion heatshield at both wind tunnel and flight conditions. The addition of overset capability to the DPLR flow solver is seen to be an essential feature for analyzing increasingly complex geometries in thermochemical nonequilibrium.

The development and verification of the Charring Ablating Thermal Protection Implicit System Solver is presented. This work concentrates on the derivation and verification of the stationary grid terms in the equations that govern three-dimensional heat and mass transfer for charring thermal protection systems including pyrolysis gas flow through the porous char layer. The governing equations are discretized according to the Galerkin finite element method with first and second order implicit time integrators. The governing equations are fully coupled and are solved in parallel via Newton's method, while the fully implicit linear system is solved with the Generalized Minimal Residual method. Verification results from exact solutions and the Method of Manufactured Solutions are presented to show spatial and temporal orders of accuracy as well as nonlinear convergence rates.

The development and verification of a one-dimensional material thermal response code with ablation is presented. The implicit time integrator, control volume finite element spatial discretization, and Newton's method for nonlinear iteration on the entire system of residual equations have been implemented and verified for the thermochemical ablation of internally decomposing materials. This study is a continuation of the work presented in "One-Dimensional Ablation with Pyrolysis Gas Flow Using a Full Newton's Method and Finite Control Volume Procedure" (AIAA-2006-2910), which described the derivation, implementation, and verification of the constant density solid energy equation terms and boundary conditions. The present study extends the model to decomposing materials including decomposition kinetics, pyrolysis gas flow through the porous char layer, and a mixture (solid and gas) energy equation. Verification results are presented for the thermochemical ablation of a carbon-phenolic ablator which involves the solution of the entire system of governing equations.

The Exploration Flight Test 1 vehicle included roughly 100 near surface thermocouples on the after body of the vehicle. The temperature traces at each of these instruments have been used to perform inverse environment reconstruction to determine the aerothermal environment experienced during re-entry of the vehicle. This paper provides an overview of the reconstructed environments and identifies critical aspects of the environment. These critical aspects include transition and reaction control system jet influence. A blind test of the process and reconstruction tool was also performed to build confidence in the reconstructed environments. Finally, an uncertainty quantification analysis was also performed to identify the impact of each of the uncertainties on the reconstructed environments.