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.

Planar laser-induced fluorescence of nitric oxide was used to visualize the interaction of reaction control system jet flows on the afterbody of a hypersonic capsule reentry vehicle at the Calspan-University at Buffalo Research Center's Large Energy National Shock Tunnel I reflected shock tunnel facility. The interaction of pitch and roll jets with the flowfield was investigated. Additionally, thin-film sensors were used to monitor heat transfer on the surface of the model to detect localized heating resulting from the firing of the reaction control system jets. Visualizations of the capsule shear layer using both planar laser-induced fluorescence and schlieren imaging compared favorably. The structure of the roll jet was found to be significantly altered due to interactions with the flowfield. Additionally, the presence of the roll jet appeared to change the nature of the shear layer from steady laminar to unsteady. The pitch jet structure was only disturbed in the far field. Comparison of the planar laser-induced fluorescence jet-fluid visualizations and the surface heat flux distributions indicate that the regions of enhanced aeroheating are not caused by the jet fluid itself impinging on the surface, but rather by the presence of jet-induced horseshoe vortices and shock wave/boundary-layer interactions.

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.

The principal mission of NASA Johnson Space Center is Human Spaceflight. In support of the mission the Applied Aeroscience and CFD Branch has several technical competencies that include aerodynamic characterization, aerothermodynamic heating, rarefied gas dynamics, and decelerator (parachute) systems.

In this paper, we present a novel scheme for modeling the hypersonic atmospheric entry of large vehicles with an ablative thermal protection system. The Favre-averaged thermochemical nonequilibrium Navier–Stokes equations with Spalart–Allmaras turbulence closure, thermodynamic, chemical kinetic, and quasi-steady ablation model are presented. The numerical method is based on a streamline upwind Petrov–Galerkin (SUPG) stabilized finite element formulation. The formulation and implementation of the finite element approximation are discussed in detail. The performance of the scheme is investigated through a series of increasingly complex applications, culminating in the simulation of a three-dimensional ablating heatshield in transitioning flow.

This paper assesses the validity of a stabilized finite element formulation applied to shockwave/turbulent boundary layer interaction (SWTBLI) problems. Our focus is on comparison with two well-defined experimental data sets to evaluate the performance of the Spalart-Allmaras turbulence model. The streamline‐upwind Petrov/Galerkin finite element formulation and the fully implicit numerical treatment of the turbulence model is first presented. We choose for validation two experiments which differ in the mechanism through which the SWTBLI is created. The first case is that of high Mach number flow over a compression ramp, and the second is an externally-generated shockwave of varying strength impinging on a turbulent flat plate boundary layer. A particularly unique aspect of our numerical implementation is that it has been previously verified using the Method of Manufactured Solutions, hence we are confident that in this validation step that the selected mathematical models are implemented correctly. The combination of a verified numerical implementation, along with careful consideration of mesh and iterative convergence, allow us to critically assess validity of the underlying mathematical models.

Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize the interaction of reaction-control-system (RCS) jet flows in the wake of a hypersonic capsule reentry vehicle. The tests were performed at the Calspan University at Buffalo Research Center's (CUBRC) LENS-I reflected shock tunnel facility. This was the first application of PLIF to study RCS jets in a large-scale pulsed hypersonic facility. The LENS-I facility allowed RCS jet flows to be studied while varying the flow enthalpy, Reynolds number, angle of attack and jet configuration. The interaction of pitch and roll jets with the flowfield was investigated. Additionally, thin film sensors were used to monitor heat transfer on the surface of the model to detect any localized heating resulting from the firing of the RCS jets. Tests were conducted with the model held at angles of attack of 18deg and 22deg. The nominal Mach number in all tests was 8, while Reynolds number based on model diameter ranged from 2.2x10(exp 6) - 1.5x10(exp 7). Images were processed using the Virtual Diagnostics Interface (ViDI) system developed at NASA Langley Research Center to provide a three-dimensional display of the experimental data.

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.

We analyze the artificial dissipation introduced by a streamline-upwind Petrov-Galerkin finite element method and consider its effect on the conservation of total enthalpy for the Euler and laminar Navier-Stokes equations. We also consider the chemically reacting case. We demonstrate that in general, total enthalpy is not conserved for the important special case of the steady-state Euler equations. A modification to the artificial dissipation is proposed and shown to significantly improve the conservation of total enthalpy.

Presentation topics include background and motivation; physical modeling including governing equations and thermochemistry; finite element formulation; results of inviscid thermal nonequilibrium chemically reacting flow and viscous thermal equilibrium chemical reacting flow; and near-term effort.

The high-temperature thermodynamic and transport models currentlyimplemented in FIN-S are one of several possible choices, and serveto provide the minimum set required for algorithm developmentIt is expected that these simplified models will be invalidated forcertain problem classes and that more complex models will berequiredSimilar thermochemical models are required by other areas ofPECOS research, e.g. ablation and shock layer radiationThe HTChem library is being developed to consolidate efforts andprovide a common source for requisite high-temperaturethermochemistry and transport property data

A streamline upwind Petrov-Galerkin finite element method is presented for the case of a reacting mixture of thermally-perfect gases, using chemical non-equilibrium. Details of the stabilization scheme and nonlinear solution are presented. The authors have independently implemented the proposed algorithm in two separate codes, for both single temperature and and two temperature models. Example problems invoving a cylinder in Mach 20 crossflow, as well as a three-dimensional blunt nosetip are shown and compared to established codes.