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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS1977

Aerodynamic heating in large cavities in an array of RSI tiles

A large panel of reusable surface insulation (RSI) tiles including lost tile cavities was aerothermally tested in the Langley 8 foot high temperature structures tunnel to determine both the heat load within the cavities and the structural performance of the RSI surrounding the cavities. Tests were conducted with a turbulent boundary layer at a nominal free stream Mach number of 6.6, a total temperature of 1800 K, a Reynolds number per meter of 5 million, and a dynamic pressure of 62 kPa. The maximum aerodynamic heating to the floor of the cavity was two to three times the normal surface heating. The cavity heating rates agreed with data from other facilities and were successfully correlated with an empirical equation. A zippering failure occurred to a tile downstream of a double tile cavity when the separated flow attached to the floor of the cavity and forced the tile from its position.

Hunt, L. R.↗
NASA NTRSDate not supplied

Thermal Protection System P50 Cork Char Response Characterization to Ascent Flight Shear and Heating Environments

NASA’s Space Launch System (SLS) experienced deflagration of the acreage of the Core Stage base heat shield during the Green Run hot-fire test campaigns and Artemis I flight. The burning products from the base heat shield during Green Run hot-fire test operation led to concerns in aerothermal environments and thermal protection system (TPS) performance prior to Artemis I flight. Updated design thermal environments accounting for cork combustion were developed using Green Run measurements and assessed prior to Artemis I, resulting in increased TPS thickness for the base heat shield. This work documents a highly controlled ground test within the NASA Marshall Space Flight Center’s Propulsion Test Branch Hot Gas Facility (HGF) and investigates the P50 cork TPS response due to sensitivity changes in radiative heat rates and surface shear stress. Prior to testing at the HGF, most of the P50 cork panels had a thick layer of char due to the cork combustion process replicated in a thermal vacuum chamber. SLS ascent shear stress and heating rate profile were simulated within the HGF by running the combustion-driven wind tunnel at off-nominal conditions and Mach numbers below 1. This test data provided insight into aerothermal measurements from Artemis I flight as well as informing updated base heating design environments for Artemis II. An in-depth understanding of the TPS P50 cork char response and observations is presented for launch vehicle flight ascent environments.

aerothermodynamics↗
NASA NTRSDate not supplied

Data Fusion of In-Flight Aerothermodynamic Heating Measurements Using Kalman Filtering

On February 18th, 2021, the Mars 2020 entry system successfully delivered the Perseverance rover to the surface of Mars at Jezero Crater. The entry capsule carried instrumentation installed on the heatshield and backshell, named the Mars Entry, Descent, and Landing Instrumentation 2. The instruments included pressure transducers, thermocouples, heat flux gauges, and a radiometer to measure the aerodynamic and aerothermodynamic performance of the entry vehicle. Three of these sensors, a thermocouple plug, heat flux gauge, and a radiometer, were co-located on the backshell. The sensors were exposed to roughly the same aerodynamic heating, but measured these environments in different ways, each with its own set of modeling and measurement error complications. This paper develops a method for blending each of these measurements together in a single algorithm to produce estimates of the aerothermodynamic environments at that backshell location. The approach makes use of the Kalman-Schmidt filter/smoother methodology, where systematic measurement error parameters are modeled as multiplicative states that are estimated by the filter along with the aerothermal states. The results indicate peak convective and radiative heating values of 0.86 and 5.16 W/cm2, respectively, compared to the filter predictive model values of 0.67 and 4.83 W/cm2.

Christopher D Karlgaard↗
NASA NTRSDate not supplied

Data Fusion of In-Flight Aerothermodynamic Heating Measurements Using Kalman Filtering

On February 18th, 2021, the Mars 2020 entry system successfully delivered the Perseverance rover to the surface of Mars at Jezero Crater. The entry capsule carried instrumentation installed on the heatshield and backshell, named the Mars Entry, Descent, and Landing Instrumentation 2. The instruments included pressure transducers, thermocouples, heat flux gauges, and a radiometer to measure the aerodynamic and aerothermodynamic performance of the entry vehicle. Three of these sensors, a thermocouple plug, heat flux gauge, and a radiometer, were co-located on the backshell. The sensors were exposed to roughly the same aerodynamic heating, but measured these environments in different ways, each with its own set of modeling and measurement error complications. This paper develops a method for blending each of these measurements together in a single algorithm to produce estimates of the aerothermodynamic environments at that backshell location. The approach makes use of the Kalman-Schmidt filter/smoother methodology, where systematic measurement error parameters are modeled as multiplicative states that are estimated by the filter along with the aerothermal states. The results indicate peak convective and radiative heating values of 0.86 and 5.16 W/cm$^2$, respectively, compared to the filter predictive model values of 0.67 and 4.83 W/cm$^2$.

Christopher D Karlgaard↗
NASA NTRSDate not supplied

Analysis of MSL/MEDLI Entry Data with Coupled CFD and Material Response

The Mars Science Laboratory (MSL) was protected during its atmospheric entry by an instrumented heatshield using NASA's Phenolic Impregnated Carbon Ablator (PICA) material [1]. PICA is a lightweight carbon fiber/polymeric resin material that offers outstanding performances for protecting probes during planetary entry. The Mars Entry Descent and Landing Instrument (MEDLI) suite on MSL offers unique in-flight validation data for models of material response and atmospheric entry. MEDLI recorded, among other things, time-resolved in-depth temperature data of PICA using thermocouple sensors assembled in the MEDLI Integrated Sensor Plugs (MISP) [2]. The objective of this work is to showcase and analyze the coupling between the material response and the aerothermal environment. As shown in Figure 1, the workflow is divided into the following steps. First, the aerothermal properties are computed in the Data Parallel Line Relaxation (DPLR) code [3] and used with the Nonequilibrium air radiation (NEQAIR) program [8] to compute radiative heating. Second, the thermal response inside the material is computed in the Porous material Analysis Toolbox based on OpenFOAM (PATO) [4,5,6] using a fixed blowing correction parameter. Third, the pyrolysis gases computed in PATO are used as inputs to a blowing boundary condition within DPLR. Fourth, the new environment properties from DPLR are used in NEQAIR to provide an updated solution, then both the updated aerothermal environment and radiative heating are used in PATO without blowing correction. The third and fourth steps are then repeated until convergence in surface temperature is obtained. Convergence in the radiative heating is generally achieved before surface temperature, at which point the radiative heating is no longer updated. Char mass loss rates are forced to zero to produce a non-receding surface condition. For early time points in the trajectory, where flow around the MSL aeroshell is rarefied, the Direct Simulation Monte Carlo (DSMC) code, SPARTA [7], is used to compute the aerothermal environment. Iteration between PATO and SPARTA is not performed due to the computational cost of DSMC simulations. Preliminary results of the coupling between PATO and DPLR for the MSL heatshield atmospheric entry model are presented in Figures 2-4 at 65 seconds after entry interface. Figure 2 shows the surface temperature results from an uncoupled simulation in PATO with the blowing correction parameter applied (left) along with the coupled surface temperature after iteration (right). Figure 3 shows the surface temperature along the centerline from windward to leeward for easier comparison. Figure 4 shows the coupled and uncoupled pyrolysis gas blowing rate. Mars 2020 used a similar heatshield consisting of PICA for thermal protection during entry, descent, and landing. In preparation for Mars 2020 post-flight analysis, the predictive material response capability is benchmarked against flight data from MEDLI. This work represents an important milestone toward the development of validated predictive capabilities for designing thermal protection systems for planetary probes.

Thermal Protection Systems↗
NASA NTRS2018

Significance of DSMC Computed Aerothermal Environments in the Rarefied Regime for Atmospheric Entry Material Response

During Mars atmospheric entry, the Mars Science Laboratory (MSL) was protected by a 4.5 meters diameter ablative heatshield assembled in 113 tiles. The heatshield was made of NASA's flagship ablative material, the Phenolic Impregnated Carbon Ablator (PICA). Prior work compared the traditional one-dimensional and three-dimensional material response models at different locations in the heatshield. It was observed that the flow was basically one-dimensional in the nose and flank regions, but three-dimensional flow effects were observed in the outer flank. The objective of this work is to study the effects of the aerothermal environment on the material response. We extend prior work by computing aerothermal environments using the direct simulation Monte Carlo (DSMC) code SPARTA and the CFD code Data Parallel Line Relaxation (DPLR). SPARTA is used to compute environment in the rarefied regime prior to 48.4s of entry where the Knudsen number is such that the Navier-Stokes equations can be inaccurate. Similarly to previous work, the DPLR software is used to compute the hypersonic environment for laminar then turbulent boundary layer assumptions from 48.4 s up to 100 s after Entry Interface (EI) along the MSL 08-TPS-02/01a trajectory. We observe that extending the aerothermal environments to times prior to 48.4 s modifies the thermal response of the heat shield at the surface and in-depth; however the effects on the recession are minimal. Additionally, using the assumption of a turbulent boundary layer versus a laminar one leads to higher surface and in-depth temperatures, larger recession, and a displacement of the peak heating and peak recession location.

Aerothermal↗
NASA NTRSDate not supplied

ANALYSIS OF THE MSL/MEDLI ENTRY DATA WITH COUPLED CFD AND MATERIAL RESPONSE.

The Mars Science Laboratory (MSL) was protected during its atmospheric entry by an instrumented heat-shield using NASA's Phenolic Impregnated Carbon Ablator (PICA) material. PICA is a lightweight carbon fiber/polymeric resin material that offers out-standing performances for protecting probes during planetary entry. The Mars Entry Descent and Landing Instrument (MEDLI) suite on MSL offers unique in-flight validation data for models of material response and atmospheric entry. MEDLI recorded, among other things, time-resolved in-depth temperature data of PICA using thermocouple sensors assembled in the MEDLI Integrated Sensor Plugs (MISP). The objective of this work is to showcase and analyze the coupling between the material response and the aerothermal environment. As shown in Figure 1, the workflow is divided into the following steps. First, the aerothermal properties are computed in the Data Parallel Line Relaxation (DPLR) code [3] and used with the Nonequilibrium air radiation (NEQAIR) program [8] to compute radiative heating. Second, the thermal response inside the material is computed in the Porous material Analysis Toolbox based on Open-FOAM (PATO) using a fixed blowing correction parameter. Third, the pyrolysis gases computed in PATO are used as inputs to a blowing boundary condition within DPLR. Fourth, the new environment properties from DPLR are used in NEQAIR to provide an updated solution, then both the updated aerothermal environment and radiative heating are used in PATO without blowing correction. The third and fourth steps are then repeated until convergence in surface temperature is obtained. Convergence in the radiative heating is generally achieved before surface temperature, at which point the radiative heating is no longer updated. Char mass loss rates are forced to zero to produce a non-receding surface condition. For early time points in the trajectory, where flow around the MSL aeroshell is rarefied, the Direct Simulation Monte Carlo (DSMC) code, SPARTA, is used to compute the aerothermal environment. Iteration between PATO and SPARTA is not performed due to the computational cost of DSMC simulations. Preliminary results of the coupling between PATO and DPLR for the MSL heatshield atmospheric entry model are presented in Figures 2-4 at 65 seconds after entry interface. Figure 2 shows the surface temperature results from an uncoupled simulation in PATO with the blowing correction parameter applied (left) along with the coupled surface temperature after iteration (right). Figure 3 shows the surface temperature along the centerline from windward to leeward for easier comparison. Figure 4 shows the coupled and uncoupled pyrolysis gas blowing rate. Mars 2020 used a similar heatshield consisting of PICA for thermal protection during entry, descent, and landing. In preparation for Mars 2020 post-flight analysis, the predictive material response capability is benchmarked against flight data from MEDLI. This work represents an important milestone toward the development of validated predictive capabilities for designing thermal protection systems for planetary probes.

Mars Science Laboratory↗
NASA NTRSDate not supplied

Overview of Mars Sample Return – Earth Entry System Woven Roughness Heating Augmentation Test in NASA Langley’s Mach 6 Wind Tunnel

The Mars Sample Return Mission (MSR) is a planned NASA flagship mission in which a sample retrieval lander (SRL) with a rover will be flown to Mars to obtain sample tubes on the surface that were dropped by the Mars 2020 rover [1]. After obtaining the sam-ples, the rover will return and ascend back to Martian orbit onboard the Mars Ascent Vehicle (MAV). Upon return to Earth orbit, the samples will perform Entry, Descent, and Landing (EDL) with the Earth Entry Sys-tem (EES) architecture, and land in Utah. The EES vehicle will utilize a HEEET-variant as its TPS, which will be the first time a woven TPS will be used on a flagship NASA mission [2]. This TPS offers a unique challenge for Computational Fluid Dynamics (CFD) modeling of the aerothermal envi-ronment of the vehicle, as woven roughness heating augmentation has not been extensively investigated experimentally. As a result, in order to validate com-putational models for woven roughness heating aug-mentation, a wind tunnel test campaign at NASA Langley Research Center’s Mach 6 wind tunnel was performed in April of 2023. This test campaign consisted of over a hundred runs with Reynolds numbers spanning from 1-7 mil-lion 1/ft and with six separate wind tunnel models used. A second campaign with a suite of new models will be conducted in Summer 2023 as well as a cam-paign with a flat plate model, both of which are of great interest to the MSR-EES project. The data obtained from this test are extremely vital for the MSR mission, as they will validate CFD roughness heating models which will be directly used to design the TPS of the EES portion of MSR and characterize the heating environment that the entry ve-hicle will experience. Further extensions of the MSR-EES test campaign will continue to provide validation data for developing more effective computational tools.

Jonathan Cheatwood↗
NASA NTRS1998

Engineering Aerothermal Analysis for X-34 Thermal Protection System Design

Design of the thermal protection system for any hypersonic flight vehicle requires determination of both the peak temperatures over the surface and the heating-rate history along the flight profile. In this paper, the process used to generate the aerothermal environments required for the X-34 Testbed Technology Demonstrator thermal protection system design is described as it has evolved from a relatively simplistic approach based on engineering methods applied to critical areas to one of detailed analyses over the entire vehicle. A brief description of the trajectory development leading to the selection of the thermal protection system design trajectory is included. Comparisons of engineering heating predictions with wind-tunnel test data and with results obtained using a Navier- Stokes flowfield code and an inviscid/boundary layer method are shown. Good agreement is demonstrated among all these methods for both the ground-test condition and the peak heating flight condition. Finally, the detailed analysis using engineering methods to interpolate the surface-heating-rate results from the inviscid/boundary layer method to predict the required thermal environments is described and results presented.

Wurster, Kathryn E.↗
NASA NTRS1998

Engineering Aerothermal Analysis for X-34 Thermal Protection System Design

Design of the thermal protection system for any hypersonic flight vehicle requires determination of both the peak temperatures over the surface and the heating-rate history along the flight profile. In this paper, the process used to generate the aerothermal environments required for the X-34 Testbed Technology Demonstrator thermal protection system design is described as it has evolved from a relatively simplistic approach based on engineering methods applied to critical areas to one of detailed analyses over the entire vehicle. A brief description of the trajectory development leading to the selection of the thermal protection system design trajectory is included. Comparisons of engineering heating predictions with wind-tunnel test data and with results obtained using a Navier-Stokes flowfield code and an inviscid/boundary layer method are shown. Good agreement is demonstrated among all these methods for both the ground-test condition and the peak heating flight condition. Finally, the detailed analysis using engineering methods to interpolate the surface-heating-rate results from the inviscid/boundary layer method to predict the required thermal environments is described and results presented.

Wurster, Kathryn E.↗
NASA NTRS2018

Overview of Recent EAST Testing, Modeling & Analysis

Shock Layer Radiation at NASA Ames - Background: Complex aerothermal and thermochemical phenomena of planetary entry define convective and radiative heating. A spacecraft's TPS (Thermal Protective System) mitigates heat transfer to substructure. Successful TPS design relies on verifiable characterization of these phenomena in the anticipated flight environment. - Approach: EAST (Electric Arc Shock Tube) [facility at Ames] simulates high-enthalpy, real-gas phenomena encountered by hypersonic vehicles entering planetary atmospheres by spectrally imaging the flow behind a moving shock wave. - Goal: Validate aerothermal models (DPLR (Data Parallel Line Relaxation Code) and NEQAIR (Nonequilibrium Radiative Transport and Spectra Program)), inform model improvements, reduce uncertainty and quantify design uncertainties. - Recent Relevant Projects: MSL (Mars Science Laboratory) & Mars 2020, InSight, OSIRIS-REx (Origins Spectral Interpretation Resource Identification Security Regolith Explorer), Orion EFT-1 (Exploration Flight-1) and EM-1 (Exploration Mission-1) and New Frontiers.

Brandis, Aaron↗
NASA NTRS2009

Analysis of Aeroheating Augmentation due to Reaction Control System Jets on Orion Crew Exploration Vehicle

The analysis of effects of the reaction control system jet plumes on aftbody heating of Orion entry capsule is presented. The analysis covered hypersonic continuum part of the entry trajectory. Aerothermal environments at flight conditions were evaluated using Langley Aerothermal Upwind Relaxation Algorithm (LAURA) code and Data Parallel Line Relaxation (DPLR) algorithm code. Results show a marked augmentation of aftbody heating due to roll, yaw and aft pitch thrusters. No significant augmentation is expected due to forward pitch thrusters. Of the conditions surveyed the maximum heat rate on the aftshell is expected when firing a pair of roll thrusters at a maximum deceleration condition.

Dyakonov, Artem A.↗
NASA NTRS2002

Heat Flux Sensor Testing

This viewgraph presentation provides information on the following objectives: Developing secondary calibration capabilities for MSFC's (Marshall Space Flight Center) Hot Gas Facility (HGF), a Mach 4 Aerothermal Wind Tunnel; Evaluating ASTM (American Society for Testing and Materials) slug/ thinskin calorimeters against current HGF heat flux sensors; Providing verification of baselined AEDC (Arnold Engineering Development Center) / Medtherm gage calibrations; Addressing future calibration issues involving NIST (National Institute of Standards and Technology) certified radiant gages.

Clark, D. W.↗
NASA NTRS2019

Characterization of Thermal Protection Systems: An Analysis of Optical & Thermal Properties

Thermal Protection System Materials & thermal control coatings were characterized by a variety of methodologies including heat flow meter method (HFM), dynamic scanning calorimetry (DSC), fourier transform infrared spectroscopy (FTIR), and Ultra-Violet & Visible light spectroscopy (UV-VIS). Physical properties were measured to aid in the construction of thermal response models for a number a projects including MEDLI and MEDLI2. Comparing pre-flight predictions with actual flight data can allow for a reduction of margins and improve designs. The models enable design of TPS for flight missions and analysis of flight data to understand aerothermal environments experienced on the mission. This presentation reveals critical information: heat capacity (cp vs T), thermal conductivity ( vs T), and emissivity (E vs T) for Super Lightweight Ablative (SLA), Phenolic Impregnated Carbon Ablative (PICA-D), & Heatshield for Extreme Entry Environment Technology (HEEET). These properties were also measured for various thermal control coatings manufactured by AZ Technology. A portion of this work is incomplete and highlights questions to investigate in the future.

Cole, Brian↗
NASA NTRS1989

The computation of thermo-chemical nonequilibrium hypersonic flows

Several conceptual designs for vehicles that would fly in the atmosphere at hypersonic speeds have been developed recently. For the proposed flight conditions the air in the shock layer that envelops the body is at a sufficiently high temperature to cause chemical reaction, vibrational excitation, and ionization. However, these processes occur at finite rates which, when coupled with large convection speeds, cause the gas to be removed from thermo-chemical equilibrium. This non-ideal behavior affects the aerothermal loading on the vehicle and has ramifications in its design. A numerical method to solve the equations that describe these types of flows in 2-D was developed. The state of the gas is represented with seven chemical species, a separate vibrational temperature for each diatomic species, an electron translational temperature, and a mass-average translational-rotational temperature for the heavy particles. The equations for this gas model are solved numerically in a fully coupled fashion using an implicit finite volume time-marching technique. Gauss-Seidel line-relaxation is used to reduce the cost of the solution and flux-dependent differencing is employed to maintain stability. The numerical method was tested against several experiments. The calculated bow shock wave detachment on a sphere and two cones was compared to those measured in ground testing facilities. The computed peak electron number density on a sphere-cone was compared to that measured in a flight test. In each case the results from the numerical method were in excellent agreement with experiment. The technique was used to predict the aerothermal loads on an Aeroassisted Orbital Transfer Vehicle including radiative heating. These results indicate that the current physical model of high temperature air is appropriate and that the numerical algorithm is capable of treating this class of flows.

Candler, Graham↗
NASA NTRS1996

PICA Forebody Heatshield Qualification for the Stardust Discovery Class Mission

This paper presents the qualification of the light weight Phenolic Impregnated Carbon Ablators (PICA) as the forebody heatshield for the Stardust Discovery Class Mission. The Stardust spacecraft will be launched in early 1999 and fly by Comet Wild-2 to collect cometary and interstellar dust and return them back to earth in the Sample Return Capsule (SRC). This earth re-entry will be the fastest to date, at 12.6 km/s, and therefore requires a heatshield that can withstand very high heating rates and stagnation pressures, as well as provide the necessary insulation to the vehicle structure. The PICA material was developed as part of the Lightweight Ceramic Ablators program at NASA Ames Research Center, and was baselined as the forebody heatshield because of its low density and superior ablation and thermal performance at severe aerothermodynamic conditions. Under a Small Business Innovative Research (SBIR) program with NASA Ames, Fiber Materials, Inc. developed a process to manufacture a single-piece PICA heatshield for the forebody of the SRC, along with witness material for the fabrication of the test models. The test models were fabricated and instrumented by the staff of Lockheed Martin Astronautics in Denver, Colorado. Full body preliminary aerothermal CFD calculations were performed at NASA Ames to determine the heating and stagnation pressure conditions. The Heat shield sizing was also performed at NASA Ames by using a new material response code that accounts for the highly porous characteristics of the PICA material. The ablation and thermal performance of PICA was qualified in the NASA Ames Interaction Heating Arc Jet Facility. A total of 24 models and four test conditions were used to qualify PICA at the predicted peak heat flux, heat load, shear, and stagnation pressure conditions. Surface and in-depth temperatures were measured using optical pyrometers and thermocouples. Surface recession was measured by using a template and a height gage. Several models were tested to evaluate repair procedures, and two models were cold soaked in liquid nitrogen, prior th testing, to investigate the effect of the cold space environment on the performance of the material. In addition, material cored from a demonstration single-piece heatshield was tested to verify that the PICA process can be successfully completed on a large, complex heatshield shape.

Tran, Huy K.↗
NASA NTRS2012

Plume-Free Stream Interaction Heating Effects During Orion Crew Module Reentry

During reentry of the Orion Crew Module (CM), vehicle attitude control will be performed by firing reaction control system (RCS) thrusters. Simulation of RCS plumes and their interaction with the oncoming flow has been difficult for the analysis community due to the large scarf angles of the RCS thrusters and the unsteady nature of the Orion capsule backshell environments. The model for the aerothermal database has thus relied on wind tunnel test data to capture the heating effects of thruster plume interactions with the freestream. These data are only valid for the continuum flow regime of the reentry trajectory. A Direct Simulation Monte Carlo (DSMC) analysis was performed to study the vehicle heating effects that result from the RCS thruster plume interaction with the oncoming freestream flow at high altitudes during Orion CM reentry. The study was performed with the DSMC Analysis Code (DAC). The inflow boundary conditions for the jets were obtained from Data Parallel Line Relaxation (DPLR) computational fluid dynamics (CFD) solutions. Simulations were performed for the roll, yaw, pitch-up and pitch-down jets at altitudes of 105 km, 125 km and 160 km as well as vacuum conditions. For comparison purposes (see Figure 1), the freestream conditions were based on previous DAC simulations performed without active RCS to populate the aerodynamic database for the Orion CM. Other inputs to the analysis included a constant Orbital reentry velocity of 7.5 km/s and angle of attack of 160 degrees. The results of the study showed that the interaction effects decrease quickly with increasing altitude. Also, jets with highly scarfed nozzles cause more severe heating compared to the nozzles with lower scarf angles. The difficulty of performing these simulations was based on the maximum number density and the ratio of number densities between the freestream and the plume for each simulation. The lowest altitude solutions required a substantial amount of computational resources (up to 1800 processors) to simulate approximately 2 billion molecules for the refined (adapted) solutions.

Marichalar, J.↗
NASA NTRS1993

Static and aerothermal tests of a superalloy honeycomb prepackaged thermal protection system

A reusable metallic thermal protection system has been developed for vehicles with maximum surface temperatures of up to 2000 F. An array of two 12- by 12-in. panels was subjected to radiant heating tests that simulated Space Shuttle entry temperature and pressure histories. Results indicate that this thermal protection system, with a mass of 2.201 lbm/ft(exp 2), can successfully prevent typical aluminum primary structure of an entry vehicle like the Space Shuttle from exceeding temperatures greater than 350 F at a location on the vehicle where the maximum surface temperature is 1900 F. A flat array of 20 panels was exposed to aerothermal flow conditions, at a Mach number of 6.75. The panels were installed in a worst-case orientation with the gaps between panels parallel to the flow. Results from the aerothermal tests indicated that convective heating occurred from hot gas flow in the gaps between the panels. Proposed design changes to prevent gap heating occurred from hot gas flow in the gaps between the panels. Proposed design changes to prevent gap heating include orienting panels so that gaps are not parallel to the flow and using a packaged, compressible gap-filler material between panels to block hot gas flow in the gaps.

Gorton, Mark P.↗