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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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At least 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS2006

Displacements of Metallic Thermal Protection System Panels During Reentry

Bowing of metallic thermal protection systems for reentry of a previously proposed single-stage-to-orbit reusable launch vehicle was studied. The outer layer of current metallic thermal protection system concepts typically consists of a honeycomb panel made of a high temperature nickel alloy. During portions of reentry when the thermal protection system is exposed to rapidly varying heating rates, a significant temperature gradient develops across the honeycomb panel thickness, resulting in bowing of the honeycomb panel. The deformations of the honeycomb panel increase the roughness of the outer mold line of the vehicle, which could possibly result in premature boundary layer transition, resulting in significantly higher downstream heating rates. The aerothermal loads and parameters for three locations on the centerline of the windward side of this vehicle were calculated using an engineering code. The transient temperature distributions through a metallic thermal protection system were obtained using 1-D finite volume thermal analysis, and the resulting displacements of the thermal protection system were calculated. The maximum deflection of the thermal protection system throughout the reentry trajectory was 6.4 mm. The maximum ratio of deflection to boundary layer thickness was 0.032. Based on previously developed distributed roughness correlations, it was concluded that these defections will not result in tripping the hypersonic boundary layer.

Daryabeigi, Kamran↗
NASA NTRS2012

Thermal protection system and related methods

A thermal protection system and a method of manufacturing are disclosed. The thermal protection system may be configured to protect a movable joint, for example, a flexible bearing of a rocket motor nozzle. The thermal protection system includes a series of annular shims separated by a plurality of discrete spacers. Each shim of the series of annular shims may have a larger diameter than the previous shim, and the shims may nest. The shims may comprise a thermally stable material, and the discrete spacers may comprise an elastomer. Optionally, an annular bearing protector may separate the annular shims from the flexible bearing.

Garbe, Duane J.↗
NASA NTRS1997

"TPSX: Thermal Protection System Expert and Material Property Database"

The Thermal Protection Branch at NASA Ames Research Center has developed a computer program for storing, organizing, and accessing information about thermal protection materials. The program, called Thermal Protection Systems Expert and Material Property Database, or TPSX, is available for the Microsoft Windows operating system. An "on-line" version is also accessible on the World Wide Web. TPSX is designed to be a high-quality source for TPS material properties presented in a convenient, easily accessible form for use by engineers and researchers in the field of high-speed vehicle design. Data can be displayed and printed in several formats. An information window displays a brief description of the material with properties at standard pressure and temperature. A spread sheet window displays complete, detailed property information. Properties which are a function of temperature and/or pressure can be displayed as graphs. In any display the data can be converted from English to SI units with the click of a button. Two material databases included with TPSX are: 1) materials used and/or developed by the Thermal Protection Branch at NASA Ames Research Center, and 2) a database compiled by NASA Johnson Space Center 9JSC). The Ames database contains over 60 advanced TPS materials including flexible blankets, rigid ceramic tiles, and ultra-high temperature ceramics. The JSC database contains over 130 insulative and structural materials. The Ames database is periodically updated and expanded as required to include newly developed materials and material property refinements.

Squire, Thomas H.↗
NASA NTRS1974

Effects of selected trajectory parameters on weight trends in the shuttle thermal protection system

An empirical heating model, thermal protection system unit weight correlation, and trajectory analysis were used to develop a computation procedure for studying the sensitivity of thermal protection system weights to variations in pertinent trajectory parameters. The analytical techniques used in developing this computer program are described. Application of the analysis to a Space Shuttle Orbiter configuration was performed to demonstrate gross thermal protection system weight trends with selected entry trajectory parameters.

Curry, D. M.↗
NASA NTRS1982

Multiwall thermal protection system

Multiwall insulating sandwich panels are provided for thermal protection of hypervelocity vehicles and other enclosures. In one embodiment, the multiwall panels are formed of alternate layers of dimpled and flat metal (titanium alloy) foil sheets and beaded scarfed edge seals to provide enclosure thermal protection up to 1000 F. An additional embodiment employs an intermediate fibrous insulation for the sandwich panel to provide thermal protection up to 2000 F. A third embodiment employs a silicide coated columbium waffle as the outer panel skin and fibrous layered intermediate protection for thermal environment protection up to 2500 F. The use of multiple panels on an enclosure facilitate repair and refurbishment of the thermal protection system due to the simple support provided by the tab and clip attachment for the panels.

Jackson, L. R.↗
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 NTRS1982

Transient thermal analysis of a titanium multiwall thermal protection system

The application of the SPAR thermal analyzer to the thermal analysis of a thermal protection system concept is discussed. The titanium multiwall thermal protection system concept consists of alternate flat and dimpled sheets which are joined together at the crests of the dimples and formed into 30 cm by 30 cm (12 in. by 12 in.) tiles. The tiles are mechanically attached to the structure. The complex tile geometry complicates thermal analysis. Three modes of heat transfer were considered: conduction through the gas inside the tile, conduction through the metal, and radiation between the various layers. The voids between the dimpled and flat sheets were designed to be small enough so that natural convection is insignificant (e.g., Grashof number 1000). A two step approach was used in the thermal analysis of the multiwall thermal protection system. First, an effective normal (through-the-thickness) thermal conductivity was obtained from a steady state analysis using a detailed SPAR finite element model of a small symmetric section of the multiwall tile. This effective conductivity was then used in simple one dimensional finite element models for preliminary analysis of several transient heat transfer problems.

Blosser, M. L.↗
NASA NTRSDate not supplied

Multiscale Thermal Conductivity Modeling of 3D Woven Composite Thermal Protection System Materials

Novel thermal protection systems (TPS) for re-entry and hypersonic vehicles utilize 3D woven composite materials with blended carbon filament tows (yarns) and a phenolic resin matrix. The presence of continuous carbon fibers provides superior stiffness and strength compared to legacy TPS, while the 3D weave pattern provides a great deal of design flexibility for both in-plane and through thickness thermomechanical behavior. However, modeling of 3D woven composites is notoriously complex and numerically intensive. The present investigation utilizes a unique ultra-efficient approach, known as Multiscale Recursive Micromechanics (MsRM), wherein recursive semi-analytical micromechanics methods are employed at various length scales within the composite. The MsRM approach has recently been extended to solve for the effective thermal conductivity as well as the local temperature and thermal flux fields throughout the composite. Presented model results focus on the impact of the microstructural geometry representation at each length scale on the material’s effective thermal conductivity, along with the local thermal flux and temperature fields induced in the microstructures, for a novel 3D woven TPS.

Composite↗
NASA NTRSDate not supplied

Development of Advanced Lightweight Polymeric Foam Thermal Protection for Space Launch Vehicles

Lightweight thermal protective materials are essential for temperature regulation during ascent of space launch vehicles, especially for expendable launch vehicles and cryogenic propellant tanks. Among the lightest thermal protective materials commonly in use is polymeric foam, in the form of pour foam or spray-on foam insulation (SOFI). Foam processing is rapid, can accommodate large-scale or small-scale challenging geometries easily, and can be used in a climate controlled or non-climate-controlled setting. Application systems can be fixed or portable. Use-temperatures of foam range from cryogenic to those that will ablate the material.

TPS↗
NASA NTRSDate not supplied

TPS Certification by Analysis: Model-Driven Characterization of Properties and Failure in Woven Thermal Protection Systems

Woven, ablative thermal protection system (TPS) materials provide a robust option for aggressive (re)entries and thus have been baselined for the upcoming Mars Sample Return (MSR) mission's Earth Entry System (EES). The reliability requirements for MSR-EES necessitate understanding of material property variability, which could be significant given the complex structure and anisotropic nature of properties in TPS weaves, as well as the response to potential impact with micrometeoroids or orbital debris during the EES re-entry. The TPS Certification by Analysis effort within the Entry Systems Modeling project seeks to provide computational models and analyses that support the certification against such material-based risks. For the present talk, focus will be given to the characterization of baseline woven TPS material properties and mechanical failure limits, which entails (1) use of computational techniques (e.g., machine learning) to interpret computed tomography images of the weave to generate representative structural models and (2) application of multiscale material modeling approaches to characterize thermomechanical and failure properties.

thermal protection systems↗
NASA NTRS2011

Edgewise Compression Testing of STIPS-0 (Structurally Integrated Thermal Protection System)

The Structurally Integrated Thermal Protection System (SITPS) task was initiated by the NASA Hypersonics Project under the Fundamental Aeronautics Program to develop a structural load-carrying thermal protection system for use in aerospace applications. The initial NASA concept for SITPS consists of high-temperature composite facesheets (outer and inner mold lines) with a light-weight insulated structural core. An edgewise compression test was performed on the SITPS-0 test article at room temperature using conventional instrumentation and methods in order to obtain panel-level mechanical properties and behavior of the panel. Three compression loadings (10, 20 and 37 kips) were applied to the SITPS-0 panel. The panel behavior was monitored using standard techniques and non-destructive evaluation methods such as photogrammetry and acoustic emission. The elastic modulus of the SITPS-0 panel was determined to be 1.146x106 psi with a proportional limit at 1039 psi. Barrel-shaped bending of the panel and partial delamination of the IML occurred under the final loading.

Brewer, Amy R.↗
NASA NTRS1992

Ceramic-Fibrous-Insulation Thermal-Protection System

New composite thermal-protection system developed in which glass-ceramic impregnated into surface of fibrous insulation. Called TUFI for toughened unipiece fibrous insulation developed as replacement for tiles with reaction-cured-glass (RCG) coating. Impregnation of glass-ceramic results in thermal protection system with insulating properties comparable to existing system but with 20 to 100 times more resistance to impact.

Leiser, Daniel↗
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.↗
NASA NTRS2014

Thermal Protection System with Staggered Joints

The thermal protection system disclosed herein is suitable for use with a spacecraft such as a reentry module or vehicle, where the spacecraft has a convex surface to be protected. An embodiment of the thermal protection system includes a plurality of heat resistant panels, each having an outer surface configured for exposure to atmosphere, an inner surface opposite the outer surface and configured for attachment to the convex surface of the spacecraft, and a joint edge defined between the outer surface and the inner surface. The joint edges of adjacent ones of the heat resistant panels are configured to mate with each other to form staggered joints that run between the peak of the convex surface and the base section of the convex surface.

Simon, Xavier D.↗
NASA NTRSDate not supplied

NASA Efforts to Explore Additively Manufactured Thermal Protection Systems

Development of the thermal protection system (TPS) needed to protect external and internal surfaces during launch, ascent, cruise, and reentry, is a very specialized field with expertise developed over seven decades and has enabled successful robotic as well as human exploration. Well known examples are: 1) the ablative heatshield that protected Apollo Astronauts during reentry in 1960’s, 2) reusable TPS used on the Space Shuttle Orbiter, which is now sought after by commercial space industry, and 3) hot structures, such as Carbon-Carbon and other high temperature materials needed to operate control surfaces and sharp leading edges during hypervelocity flight. TPS function demands fail-safe design to ensure mission success. One TPS does not fit all, and each application requires a unique material along with manufacturing and integration approach. TPS mass is a key metric. A robust but inefficient TPS reduces payload mass. Mission- and domain-unique TPS development and flight certification efforts have been expensive with long lead times. As a result, there are limited TPS choices with high cost. For future commercial efforts to be successful, the ability to rapidly develop TPS, based on specific needs and at a lower cost, without compromising either mass efficiency or functionality is necessary. NASA is looking into this. Additive manufacturing (AM) techniques have shown their utility in making complex parts and, through automation, the potential to reduce cost and schedule in some applications. The promise of additive manufacturing to TPS, while exciting, is unexplored. In the past several years, NASA has made exploratory investments in internal development as well as in small business and universities through SBIR and STTR programs. In addition, NASA is bringing together the TPS as well as Additive Manufacturing communities to explore approaches to prioritize focus areas for broad benefits. The very first “Additively Manufactured Thermal Protection System Workshop,” to be held at NASA Johnson Space Center in Houston at the end of March of this year, is the culmination of a year-long effort to engage the research, development, and user communities involved in both TPS as well as additive manufacturing. The workshop participants include representatives from large and small commercial space industry, DoD, Federally Funded Research and Development Centers, University Researchers and NASA personnel. The proposed talk will highlight future TPS needs of both emerging commercial and Government interests, NASA’s development efforts in additively manufactured TPS, and the findings and recommendations from the first AM TPS Workshop.

Ethiraj Venkatapathy↗
NASA NTRSDate not supplied

Thermal Protection Materials and Systems at NASA Ames Research Center

Thermal Protection Systems (TPS) are critical for enabling NASA missions involving high-speed atmospheric flight where the entries usually include descending into the atmosphere followed by a trajectory that aims to burn off energy and result in a controlled landing. NASA Ames focuses on qualifying and certifying TPS for current missions, sustaining TPS for future missions, and developing new TPS for upcoming missions where a heritage solution is not viable. More recently there is also a focus on advancing and transferring technologies that can benefit both commercial and government space needs.Developing mature thermal protection systems is a lengthy process involving advanced tools, extensive research, and testing. Design and analysis tools are used to predict aerothermal environments, aid the design of test and flight hardware, and support the testing for the thermal/mechanical response of thermal protection systems. More recently, advances in computational methods help reduce the time and cost of technological advances, aid in optimized material architecture design, and improve material properties and performance. While high-enthalpy testing that simulates the conditions of space flight remains essential for the evaluation and development of TPS materials, computational tools are already showing promise in reducing the need for widespread testing and can help fast-trackthe design cycle.With the exploration of new destinations EDL instrumentation remains an important element of the heatshield and NASA Ames and partners have developed and delivered instrumentation flight hardware in support of recent Mars missions,including Mars Science Lab (MSL) and Mars 2020 (M2020),as well as Artemis Orion. Sensors installed on the heatshield and backshell of spacecraft provide coveted information about the aerodynamic and aerothermal environment during entry.Overthe years NASA Ames has brought several reusable and ablative TPS materials to a level of readiness to hand off to missions and the Thermal Protection Materials branch continues to serve as a TPS steward for the agency.

Thermal protection materials↗
NASA NTRS2013

Arcjet Testing of Micro-Meteoroid Impacted Thermal Protection Materials

There are several harsh space environments that could affect thermal protection systems and in turn pose risks to the atmospheric entry vehicles. These environments include micrometeoroid impact, extreme cold temperatures, and ionizing radiation during deep space cruise, all followed by atmospheric entry heating. To mitigate these risks, different thermal protection material samples were subjected to multiple tests, including hyper velocity impact, cold soak, irradiation, and arcjet testing, at various NASA facilities that simulated these environments. The materials included a variety of honeycomb packed ablative materials as well as carbon-based non-ablative thermal protection systems. The present paper describes the results of the multiple test campaign with a focus on arcjet testing of thermal protection materials. The tests showed promising results for ablative materials. However, the carbon-based non-ablative system presented some concerns regarding the potential risks to an entry vehicle. This study provides valuable information regarding the capability of various thermal protection materials to withstand harsh space environments, which is critical to sample return and planetary entry missions.

Agrawal, Parul↗