SEARCH · Search NASA

Results for “HEAT SHIELD”

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.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 541 records · Page 30

All sources NASA NTRS DOE OSTI

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 NTRS1973

Development of a closed pore insulation material

A closed pore ceramic foam insulation material (CPI) has been developed that offers possibilities for use as a reusable external heat shield for the NASA manned space shuttle. The outstanding characteristics of CPI are: (1) negligible water absorption due to a noninterconnecting network of cells; (2) high emittance at room and elevated temperature; (3) ability to survive at least 10 simulated reentry cycles to 1500 K using radiant heat lamps to simulate the reentry heat fluxes; (4) ability to survive, with no change in properties or appearance, at least 10 simulated plasma arc jet cycles to 1500 K (with the exception of some stress cracks induced either by the unduly severe nature of the initial arc splash heating pulse or by improper mechanical holding of the specimen in the test fixture); (5) strength (flexure); and (6) a low thermal conductivity throughout the temperature range of interest for the space shuttle.

Tobin, A.↗

NASA NTRS2017

Experimental Investigation of Roughness Effects on Transition Onset and Turbulent Heating Augmentation on a Hemisphere at Mach 6 and Mach 10

An experimental investigation of the effects of distributed surface roughness on boundary-layer transition and turbulent heating has been conducted. Hypersonic wind tunnel testing was performed using hemispherical models with surface roughness patterns simulating those produced by heat shield ablation. Global aeroheating and transition onset data were obtained using phosphor thermography at Mach 6 and Mach 10 over a range of roughness heights and free stream Reynolds numbers sufficient to produce laminar, transitional and turbulent flow. Upstream movement of the transition onset location and increasing heating augmentation over predicted smooth-wall levels were observed with both increasing roughness heights and increasing free stream Reynolds numbers. The experimental heating data are presented herein, as are comparisons to smooth-wall heat transfer distributions from computational flow-field simulations. The transition onset data are also tabulated, and correlations of these data are presented.

Hollis, Brian R.↗

NASA NTRS1981

Early development of ceramic fiber insulation for the Space Shuttle

The evolution of reusable surface insulation is described, with attention to the mullite and other aluminosilicates, zirconia, and silicon carbide-coated carbon compositions experimented with in the late 1960s and early 1970s. Evaluation and development concentrated in this period on such aspects of design and material properties as cold soak performance, attachment, shock impingement, surface coating cracks, and tile gap design and heating. In addressing the central problem of heat-shield thermal conductivity, it was found that for a given density, silica fibers had a lower conductivity than those of mullite. This was due to the one-micron, as opposed to 4.7-micron, diameter of the silica fibers, which resulted in smaller pores and therefore less convective and radiative heat transfer. Attention is also given to tile coating materials and the high-temperature processes by which they were applied.

Buckley, J. D.↗

NASA NTRS1979

Radiation absorption by the C2 band systems for Jupiter entry conditions

Revised values of the absorption cross sections for seven electronic band systems of C2 have been calculated using recently published experimental data for the electronic transition moments. Using these revised C2 cross section values, computations were made for the radiating flow field over a Jupiter entry probe with coupled ablation injection from a carbon-phenolic heat shield. Results are presented which show that radiation absorption within the ablation layer for the spectral range of 4 to 6 eV is less than that predicted using previous C2 absorption cross section values. The effect of the reduced radiation absorption by the C2 molecule is an increase in the radiative heating rates and ablation mass loss rates for the Jupiter entry conditions considered in the study.

Sutton, K.↗

NASA NTRS1978

Opacified silica reusable surface insulation /RSI/ for thermal protection of the Space Shuttle Orbiter

A reusable surface insulation (RSI) material adopted for use on limited areas of the Orbiter's heat shield is described. The RSI is a rigid sintered fibrous silica material containing about 3% by weight of 1200-grit silicon carbide powder. RSI manufacture is described; strength characteristics and thermal conductivity of the final product can be tailored by controlling the isotropy, and the density, coefficient of thermal expansion, and specific heat are reported. Data establishing that opacification of RSI with silicon carbide lowers the thermal conductivity are presented.

Goldstein, H. E.↗

NASA NTRS1980

Preliminary thermal analysis for Saturn entry

A preliminary thermal analysis based on recently defined Saturn entry conditions has been conducted. The study, using viscous-shock-layer and engineering codes employed for Project Galileo, investigated nonequilibrium chemistry effects on the Saturn thermal environment, defined the primary heat-transfer mode for heatshield design, delineated some problem areas for future thermal studies, and validated BIRCHES (Blunt Body Inviscid Radiative and Convective Heating Engineering Solutions) for parametric or design studies. The effect of nonequilibrium chemistry appears to significantly influence only the radiative fluxes with effects localized to the stagnation region. However, the heat-transfer mode pertinent to the overall heat-shield design is convection. The convective results of BIRCHES and a detailed code are in good agreement. The resulting mass-loss rates for the currently prescribed nominal Saturn entry conditions are small when compared with the values for nominal Jupiter entry conditions. With coupled carbon-phenolic ablation injection, the convective heating rates are reduced substantially while the radiative heating rates are increased when compared with the corresponding no-injection results.

Zoby, E. V.↗

NASA NTRS1981

System requirements of Starprobe thermal shield

The mission of the Starprobe Spacecraft is to do in situ exploration of the solar corona to four solar radii. A thermal shield is required to protect the spacecraft equipment from a 3000 sun incident solar energy level caused by this close approach to the sun. The thermal shield requirements discussed include configuration, material requirements, temperature level, penetrations, outgassing and mass loss, and contamination as referenced to spacecraft mission parameters. Further, the size of the spacecraft will be limited by the space transport system orbiter payload bay dimensions, thus defining the upper limit of the relative sizes and locations of the heat shield and spacecraft equipment.

Miyake, R. N.↗

NASA NTRS1991

Chemical-kinetic problems of future NASA missions

Thermochemical nonequilibrium in the shock layer surrounding vehicles entering the atmospheres of earth and Mars at superescape velocities is studied, deriving reaction rate coefficients that reproduce experimental data obtained in shock tubes. Thermodynamic properties and emitted radiation intensities are obtained for shock tube flow and flow in a shock layer over a blunt body. The results indicate that the viscous layer of the ablation product over an ablating heat shield is likely to be in chemical nonequilbrium. For earth entry flight, the thickness of the nonequilbrium region is between and 2 cm at the expected peak radiation point in the aerobraking trajectory, For Martian entry flight it is between 8 and 23 cm. For the earth entry case, nonequilibrium phenomena reduce radiative heating rate, while the opposite occurs for the Martian case. The radiative heat transfer rates are significant for the Mars entry conditions at entry velocities equal to or greater than 7 km/s.

Park, Chul↗

NASA NTRS2005

An Approximate Ablative Thermal Protection System Sizing Tool for Entry System Design

A computer tool to perform entry vehicle ablative thermal protection systems sizing has been developed. Two options for calculating the thermal response are incorporated into the tool. One, an industry-standard, high-fidelity ablation and thermal response program was integrated into the tool, making use of simulated trajectory data to calculate its boundary conditions at the ablating surface. Second, an approximate method that uses heat of ablation data to estimate heat shield recession during entry has been coupled to a one-dimensional finite-difference calculation that calculates the in-depth thermal response. The in-depth solution accounts for material decomposition, but does not account for pyrolysis gas energy absorption through the material. Engineering correlations are used to estimate stagnation point convective and radiative heating as a function of time. The sizing tool calculates recovery enthalpy, wall enthalpy, surface pressure, and heat transfer coefficient. Verification of this tool is performed by comparison to past thermal protection system sizings for the Mars Pathfinder and Stardust entry systems and calculations are performed for an Apollo capsule entering the atmosphere at lunar and Mars return speeds.

Dec, John A.↗

NASA NTRS2006

An Approximate Ablative Thermal Protection System Sizing Tool for Entry System Design

Dec, John A.↗

NASA NTRS2004

GRABER: The Duct Tape of Space and JIMO Heat Conducting Foam

Crack formation in the space shuttle's heat shield during flight poses a major safety concern to everyone on board. Cracking weakens the structure of the shield and lessens the protection it offers against the high temperatures and forces encountered during re-entry. Astronauts need a way to mend these cracks while in space. This is GRABER s function; it can be spackled into the cracks by an astronaut. The material then hardens, or cures, due to being in a vacuum and the heat encountered when it faces the sun. A great deal of work and testing is necessary to create a material that will be workable in a vacuum over a wide range of temperatures, will cure without cracking, will adhere to the sides of the crack, and that can withstand the extreme temperatures of re-entry. A Brookfield PVS Rheometer is being used to characterize GRABER's viscosity at various temperatures and stirring rates. Various compositions of GRABER are being heat treated in a vacuum to determine probably curing times in space. The microstructures of cured samples of each composition are being examined using both optical and electron microscopy. Jupiter s Icy Moon Orbiter (JIMO) will be lifting off sometime around 2013. JIMO will have more power than its predecessor, Galileo, allowing it to change orbits to circle three of Jupiter s moons. Both of the engine types being considered require large heat dissipation systems. These systems will be comprised of heat conductive tubing and plates with a liquid flowing through them. In order to maximize the speed of heat transfer between the tubes and the panels, the in-between areas will be filled with heat conductive silicon carbide foam. Two different foam systems are being considered for this foam. Currently, experimentation is underway with adding Sic, carbon, and carbon fibers to a two part fuel retardant foam. The foam is them pyrolized and its mass and dimensional changes are measured. The structure of the foam will be examined using optical and electron microscopy as well. Work is also planned with a foam system developed by an Italian team.

Gamble, Eleanor A.↗

NASA NTRS1996

Effects of Nonequilibrium at Edge of Boundary Layer on Convective Heat Transfer to a Blunt Body

This investigation is a continuation of a previous study on nonequilibrium convective heat transfer to a blunt body. In the previous study, for relatively high Reynolds number flows, it was found that: nonequilibrium convective heat transfer to a blunt body is not strongly dependent on freestream parameters, provided that the thermochemical equilibrium is reached at the edge of boundary layer; and successful testing of convective heat transfer in an arc-jet environment is possible by duplicating the surface pressure and total enthalpy. The nonequilibrium convective heat transfer computations are validated against the results of Fay and Riddell/Goulard theory. Present work investigates low Reynolds number conditions which are typical in an actual arc-jet flow environment. One expects that there will be departures from the Fay and Riddell/Goulard result since certain assumptions of the classical theory are not satisfied. These departures are of interest because the Fay and Riddell/Goulard formulas are extensively used in arc-jet testing (e.g., to determine the enthalpy of the flow and the catalytic efficiency of heat shield materials). For practical sizes of test materials, density of the test flow (and Reynolds number) in an arc-jet is such that thermochemical equilibrium may not be reached at the edge of boundary layer. For blunt body flows of nitrogen and air, computations will be presented to show the effects of thermochemical nonequilibrium at the boundary layer edge on nonequilibrium heat transfer.

Goekcen, Tahir↗

NASA NTRS2011

Measuring Moisture Levels in Graphite Epoxy Composite Sandwich Structures

Graphite epoxy composite (GEC) materials are used in the construction of rocket fairings, nose cones, interstage adapters, and heat shields due to their high strength and light weight. However, they absorb moisture depending on the environmental conditions they are exposed to prior to launch. Too much moisture absorption can become a problem when temperature and pressure changes experienced during launch cause the water to vaporize. The rapid state change of the water can result in structural failure of the material. In addition, heat and moisture combine to weaken GEC structures. Diffusion models that predict the total accumulated moisture content based on the environmental conditions are one accepted method of determining if the material strength has been reduced to an unacceptable level. However, there currently doesn t exist any field measurement technique to estimate the actual moisture content of a composite structure. A multi-layer diffusion model was constructed with Mathematica to predict moisture absorption and desorption from the GEC sandwich structure. This model is used in conjunction with relative humidity/temperature sensors both on the inside and outside of the material to determine the moisture levels in the structure. Because the core materials have much higher diffusivity than the face sheets, a single relative humidity measurement will accurately reflect the moisture levels in the core. When combined with an external relative humidity measurement, the model can be used to determine the moisture levels in the face sheets. Since diffusion is temperaturedependent, the temperature measurements are used to determine the diffusivity of the face sheets for the model computations.

Nurge, Mark↗

NASA NTRS2016

Final Mission and Navigation Design for the 2016 Mars InSight Mission

Project Overview: NASA’s 12th Discovery Program mission that will place a single geophysical lander on Mars to study its deep interior using seismic (tidal, marsquake, and impact), rotational, and thermal measurements -First mission to unveil the composition, structure and thermal state of martian crust, mantle and core; Rebuild of the Phoenix flight system, upgraded with Juno/GRAIL avionics -Phoenix cruise stage: 3-axis stabilized, 3.2 square meter UTJ (Ultra Triple-Junction) solar array, X-band telecom -Phoenix aeroshell (heat shield, backshell, parachute) -Phoenix lander: radar, pulsed thrusters, legs, payload deck, 5.1 square meter UTJ solar arrays, two redundant Li-ion batteries, UHF (Ultra High Frequency)/X-band telecom -Dual string: Internally or functionally-redundant equipment; Science Payload -SEIS (Seismic Experiment for Interior Structure) / CNES (Centre National d’Etudes Spatiales); Three-axis seismometry, to measure seismic waves traveling through the interior -HP3 (Heat Flow and Physical Properties Probe) / DLR (German Aerospace Center) -Subsurface heat probe, to measure the heat flux from the interior - RISE (Rotation and Interior Structure Experiment) / JPL (Jet Propulsion Laboratory) -Radiometric geodesy, to determine precession & nutation of the planet’s rotation axis. 2016 Disclaimer: The InSight launch in 2016 was suspended due to critical issues with the Seismic Experiment for Interior Structure (SEIS) instrument that could not be fixed prior to the planned launch period. This presentation and its corresponding paper represent the state of the design for the 2016 mission. No attempt has been made to reflect the latest developments. [2018 Note: InSight (Mars Lander) was later launched and arrived on Mars on Nov. 26, 2018.]

Abilleira, F.↗

NASA NTRS2013

Aeroheating Measurement of Apollo Shaped Capsule with Boundary Layer Trip in the Free-piston Shock Tunnel HIEST

An aeroheating measurement test campaign of an Apollo capsule model with laminar and turbulent boundary layer was performed in the free-piston shock tunnel HIEST at JAXA Kakuda Space Center. A 250mm-diameter 6.4%-scaled Apollo CM capsule model made of SUS-304 stainless steel was applied in this study. To measure heat flux distribution, the model was equipped with 88 miniature co-axial Chromel-Constantan thermocouples on the heat shield surface of the model. In order to promote boundary layer transition, a boundary layer trip insert with 13 "pizza-box" isolated roughness elements, which have 1.27mm square, were placed at 17mm below of the model geometric center. Three boundary layer trip inserts with roughness height of k=0.3mm, 0.6mm and 0.8mm were used to identify the appropriate height to induce transition. Heat flux records with or without roughness elements were obtained for model angles of attack 28º under stagnation enthalpy between H(sub 0)=3.5MJ/kg to 21MJ/kg and stagnation pressure between P(sub 0)=14MPa to 60MPa. Under the condition above, Reynolds number based on the model diameter was varied from 0.2 to 1.3 million. With roughness elements, boundary layer became fully turbulent less than H(sub 0)=9MJ/kg condition. However, boundary layer was still laminar over H(sub 0)=13MJ/kg condition even with the highest roughness elements. An additional experiment was also performed to correct unexpected heat flux augmentation observed over H(sub 0)=9MJ/kg condition.

Hideyuki, TANNO↗

NASA NTRS2021

Overview of Dragonfly Entry Aerosciences Measurements (DrEAM)

NASA Ames Research Center (ARC) leads the Dragonfly Entry Aerosciences Measurements (DrEAM) project, which is an aeroshell instrumentation suite on the Dragonfly mission that fulfills the Engineering Science Investigation requirement for the New Frontiers mission. NASA ARC is partnering with NASA Langley Research Center (LaRC) and the German Aerospace Center (DLR) to provide a comprehensive sensor suite, including a DLR-provided Data Acquisition System (DAS). DrEAM will provide key aerothermodynamic data and performance analysis for Dragonfly’s forebody and backshell Thermal Protection System (TPS). Titan’s atmosphere predominantly consists of ni-trogen (~98% by mole) with small amounts of me-thane (~2% by mole) and other trace gases. CN is a strong radiator, and is found in nonequilibrium con-centrations for Titan entry. The accurate modeling of nonequilibrium CN radiation has proven to be a diffi-cult task. Prompted by the Huygens mission, many experimental campaigns and analyses were performed to better understand the aerothermal environments experienced by the probe during Titan entry [1]. How-ever, the Huygens probe carried no heatshield instru-mentation. Therefore, the DrEAM sensor suite will sig-nificantly advance the state-of-the-art not only by documenting the environment and performance of Dragonfly’s entry system but also by making key in situ measurements in Titan’s atmosphere for the first time. Aerothermal environments and TPS response will be measured using sensors whose flight heritage is tak-en from the Mars Entry, Descent, and Landing In-strumentation 2 (MEDLI2) thermocouple plugs and the COMbined Aerothermal and Radiometer Sensor (COMARS) suite [2], with the latter supplied by DLR. The MEDLI2 project used embedded thermocouples to directly measure the in-depth TPS temperature-time history at several locations on the heat shield and backshell of the Mars 2020 entry vehicle. These tem-perature measurements, in turn, can be used to infer surface environments via an inverse analysis proce-dure analogous to that used for MEDLI. For DrEAM, the thermocouple plug subsystem will be known as Dragonfly Sensors for Aero-Thermal Reconstruction (DragSTR). On Schiaparelli, the COMARS suite in-cluded three total surface-mounted heat flux sensors, three pressure sensors, and one radiometer. For DrEAM, the COMARS package will be known as the COmbined Sensor System for Titan Atmosphere (COSSTA). Since the methane concentration in the Titan atmos-phere is directly proportional to the radiative heat flux, the COSSTA measurements will be used to reduce the current uncertainty in the methane volume fraction. Atmospheric density measurements and capsule aero-dynamic data will be obtained through the onboard Inertial Measurement Unit (IMU), supplemented by pressure transducers similar to those used by the MEDLI and MEDLI2 projects. The DrEAM pressure sensors will be known as the Dragonfly Atmospheric Flight Transducers. (DrAFT). The pressure measure-ments, when combined with data from the on-board IMU, will allow for reconstruction of such quantities as vehicle Mach number, freestream density, and atmos-pheric winds. DrAFT measurements will enhance Dragonfly trajectory reconstruction and enable a sepa-ration of the aerodynamics from the atmosphere, as was done for MEDLI [3] and is currently in process for the MEDLI2 flight data set.

EDL↗

NASA NTRS1975

Experimental Evaluation of Silica as a Thermal Protection System for Planetary Entry Probes

The evolution and pursuit of the reflecting heat-shield concept has resulted in a refined and sophisticated high-performance material: ultra-pure, slip-cast, fused silica (SCFS). Silica is "tailored" to the intense planetary entry environments through a combination of its good ablative qualities and outstanding reflective properties in the correct spectral region. Extensive development of SCFS has produced a series of candidate materials which are experimentally evaluated and compared in a simulated entry environment. The detailed response of silica to combined heating is discussed, and the relative merits of SCFS and carbon phenolic are indicated.

David L Peterson↗