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Sample records for x-33 thermal protection

  1. X-33 Base Region Thermal Protection System Design Study

    NASA Technical Reports Server (NTRS)

    Lycans, Randal W.

    1998-01-01

    The X-33 is an advanced technology demonstrator for validating critical technologies and systems required for an operational Single-Stage-to-Orbit (SSTO) Reusuable Launch Vehicle (RLV). Currently under development by a unique contractor/government team led by Lockheed- Martin Skunk Works (LMSW), and managed by Marshall Space Flight Center (MSFC), the X-33 will be the prototype of the first new launch system developed by the United States since the advent of the space shuttle. This paper documents a design trade study of the X-33 base region thermal protection system (TPS). Two candidate designs were evaluated for thermal performance and weight. The first candidate was a fully reusable metallic TPS using Inconel honeycomb panels insulated with high temperature fibrous insulation, while the second was an ablator/insulator sprayed on the metallic skin of the vehicle. The TPS configurations and insulation thickness requirements were determined for the predicted main engine plume heating environments and base region entry aerothermal environments. In addition to thermal analysis of the design concepts, sensitivity studies were performed to investigate the effect of variations in key parameters of the base TPS analysis.

  2. Transient Analysis of Thermal Protection System for X-33 Aircraft using MSC/NASTRAN

    NASA Technical Reports Server (NTRS)

    Miura, Hirokazu; Chargin, M. K.; Bowles, J.; Tam, T.; Chu, D.; Chainyk, M.; Green, Michael J. (Technical Monitor)

    1997-01-01

    X-33 is an advanced technology demonstrator vehicle for the Reusable Launch Vehicle (RLV) program. The thermal protection system (TPS) for the X-33 is composed of complex layers of materials to protect internal components, while withstanding severe external temperatures induced by aerodynamic heating during high speed flight. It also serves as the vehicle aeroshell in some regions using a stand-off design. MSC/NASTRAN thermal analysis capability was used to predict transient temperature distribution (within the TPS) throughout a mission, from launch through the cool-off period after landing. In this paper, a typical analysis model, representing a point on the vehicle where the liquid oxygen tank is closest to the outer mold line, is described. The maximum temperature difference between the outer mold line and the internal surface of the liquid oxygen tank can exceed 1500 F. One dimensional thermal models are used to select the materials and determine the thickness of each layer for minimum weight while insuring that all materials remain within the allowable temperature range. The purpose of working with three dimensional (3D) comprehensive models using MSC/NASTRAN is to assess the 3D radiation effects and the thermal conduction heat shorts of the support fixtures.

  3. Thermal Management Design for the X-33 Lifting Body

    NASA Technical Reports Server (NTRS)

    Bouslog, S.; Mammano, J.; Strauss, B.

    1998-01-01

    The X-33 Advantage Technology Demonstrator offers a rare and exciting opportunity in Thermal Protection System development. The experimental program incorporates the latest design innovation in re-useable, low life cycle cost, and highly dependable Thermal Protection materials and constructions into both ground based and flight test vehicle validations. The unique attributes of the X-33 demonstrator for design application validation for the full scale Reusable Launch Vehicle, (RLV), are represented by both the configuration of the stand-off aeroshell, and the extreme exposures of sub-orbital hypersonic re-entry simulation. There are several challenges of producing a sub-orbital prototype demonstrator of Single Stage to Orbit/Reusable Launch Vehicle (SSTO/RLV) operations. An aggressive schedule with budgetary constraints precludes the opportunity for an extensive verification and qualification program of vehicle flight hardware. However, taking advantage of off the shelf components with proven technologies reduces some of the requirements for additional testing. The effects of scale on thermal heating rates must also be taken into account during trajectory design and analysis. Described in this document are the unique Thermal Protection System (TPS) design opportunities that are available with the lifting body configuration of the X-33. The two principal objectives for the TPS are to shield the primary airframe structure from excessive thermal loads and to provide an aerodynamic mold line surface. With the relatively benign aeroheating capability of the lifting body, an integrated stand-off aeroshell design with minimal weight and reduced procurement and operational costs is allowed. This paper summarizes the design objectives of the X-33 TPS, the flight test requirements driven configuration, and design benefits. Comparisons are made of the X-33 flight profiles and Space Shuttle Orbiter, and lifting body Reusable Launch Vehicle aerothermal environments. The X-33

  4. X-33 artist concept - 1999

    NASA Technical Reports Server (NTRS)

    1999-01-01

    An artist's conception of the half scale X-33 demonstrator flying over the southwestern desert. The vehicle was a wedge-shaped lifting body, with two vertical fins and a pair of stub wings. On the fins are the Lockheed-Martin Skunk Works logo, which was the prime contractor. At the rear is the aerospike engine, an experimental design that lacked the nozzles of conventional rockets. The X-33 tested several other new technologies, including composite structures and a metallic thermal protection system. It was hoped that these advances would lead eventually to an operational single-stage-to-orbit reusable launch vehicle called the VentureStar. However, due to technical problems with the composite liquid hydrogen tank, the X-33 program was cancelled in February 2001.

  5. Cyclic Cryogenic Thermal-Mechanical Testing of an X-33/RLV Liquid Oxygen Tank Concept

    NASA Technical Reports Server (NTRS)

    Rivers, H. Kevin

    1999-01-01

    An important step in developing a cost-effective, reusable, launch vehicle is the development of durable, lightweight, insulated, cryogenic propellant tanks. Current cryogenic tanks are expendable so most of the existing technology is not directly applicable to future launch vehicles. As part of the X-33/Reusable Launch Vehicle (RLV) Program, an experimental apparatus developed at the NASA Langley Research Center for evaluating the effects of combined, cyclic, thermal and mechanical loading on cryogenic tank concepts was used to evaluate cryogenic propellant tank concepts for Lockheed-Martin Michoud Space Systems. An aluminum-lithium (Al 2195) liquid oxygen tank concept, insulated with SS-1171 and PDL-1034 cryogenic insulation, is tested under simulated mission conditions, and the results of those tests are reported. The tests consists of twenty-five simulated Launch/Abort missions and twenty-five simulated flight missions with temperatures ranging from -320 F to 350 F and a maximum mechanical load of 71,300 lb. in tension.

  6. X-33 Contractor Design Proposals

    NASA Technical Reports Server (NTRS)

    1996-01-01

    This artist's rendering depicts the three designs submitted for the X-33 proposal for a technology demonstrator of a Single-Stage-To-Orbit (SSTO) Reusable Launch Vehicle (RLV). NASA considered design submissions from Rockwell, Lockheed Martin, and McDonnell Douglas. NASA selected Lockheed Martin's design on 2 July 1996. NASA's Dryden Flight Research Center, Edwards, California, expected to play a key role in the development and flight testing of the X-33. The RLV technology program was a cooperative agreement between NASA and industry. The goal of the RLV technology program was to enable significant reductions in the cost of access to space and to promote the creation and delivery of new space services and other activities that was to improve U.S. economic competitiveness. The X-33 was a wedged-shaped subscale technology demonstrator prototype of a potential future Reusable Launch Vehicle (RLV) that Lockheed Martin had dubbed VentureStar. The company hoped to develop VentureStar early this century. Through demonstration flight and ground research, NASA's X-33 program was to have provided the information needed for industry representatives such as Lockheed Martin to decide whether to proceed with the development of a full-scale, commercial RLV program. A full-scale, single-stage-to-orbit RLV was to have dramatically increased reliability and lowered the costs of putting a pound of payload into space, from the current figure of $10,000 to $1,000. Reducing the cost associated with transporting payloads in Low Earth Orbit (LEO) by using a commercial RLV was to have create new opportunities for space access and significantly improved U.S. economic competitiveness in the world-wide launch marketplace. NASA expected to be a customer, not the operator, of the commercial RLV. The X-33 design was based on a lifting body shape with two revolutionary 'linear aerospike' rocket engines and a rugged metallic thermal protection system. The vehicle also had lightweight components

  7. X-33 Simulation Flown by Steve Ishmael

    NASA Technical Reports Server (NTRS)

    1997-01-01

    Steve Ishmael flies a simulation of the X-33 Advanced Technology Demonstrator at NASA's Dryden Flight Research Center, Edwards, California. This simulation was used to provide flight trajectory data while flight control laws were being designed and developed, as well as to provide aerodynamic design information to X-33 developer Lockheed Martin. The X-33 program was a government/industry effort to design, build and fly a half-scale prototype that was to have demonstrated in flight the new technologies needed for the proposed Lockheed Martin full-scale VentureStar Reusable Launch Vehicle. The X-33 was a wedged-shaped subscale technology demonstrator prototype of a potential future Reusable Launch Vehicle (RLV) that Lockheed Martin had dubbed VentureStar. The company had hoped to develop VentureStar early this century. Through demonstration flight and ground research, NASA's X-33 program was to provide the information needed for industry representatives such as Lockheed Martin to decide whether to proceed with the development of a full-scale, commercial RLV program. A full-scale, single-stage-to-orbit RLV was to have dramatically increased reliability and lowered the costs of putting a pound of payload into space, from the current figure of $10,000 to $1,000. Reducing the cost associated with transporting payloads in Low Earth Orbit (LEO) by using a commercial RLV was to have created new opportunities for space access and significantly improved U.S. economic competitiveness in the world-wide launch marketplace. NASA expected to be a customer, not the operator, of the commercial RLV. The X-33 design was based on a lifting body shape with two revolutionary 'linear aerospike' rocket engines and a rugged metallic thermal protection system. The vehicle also had lightweight components and fuel tanks built to conform to the vehicle's outer shape. Time between X-33 flights was normally to have been seven days, but the program hoped to demonstrate a two-day turnaround between

  8. X-33 Proposal by Rockwell - Computer Graphic

    NASA Technical Reports Server (NTRS)

    1996-01-01

    This artist's rendering depicts the Rockwell International X-33 proposal for technology demonstrator of a Single-Stage-To-Orbit (SSTO) Reusable Launch Vehicle (RLV). NASA considered design submissions from Rockwell, Lockheed Martin, and McDonnell Douglas. NASA selected Lockheed Martin's design on 2 July 1996. NASA's Dryden Flight research Center, Edwards, California, was to have had a key role in the development and flight testing of the X-33. The RLV technology program was a cooperative agreement between NASA and industry. The goal of the RLV technology program was to enable significant reductions in the cost of access to space, and to promote the creation and delivery of new space services and other activities that was to have improved U.S. economic competitiveness. The X-33 design selected for development was a wedged-shaped subscale technology demonstrator prototype of a Reusable Launch Vehicle (RLV) by Lockheed Martin. Through demonstration flight and ground research, NASA's X-33 program was to have provided the information needed for industry representatives such as Lockheed Martin to decide whether to proceed with the development of a full-scale, commercial RLV program. A full-scale, single-stage-to-orbit RLV was to have dramatically increased reliability and lowered the costs of putting a pound of payload into space, from the current figure of $10,000 to $1,000. Reducing the cost associated with transporting payloads in Low Earth Orbit (LEO) by using a commercial RLV was to have created new opportunities for space access and significantly improve U.S. economic competitiveness in the world-wide launch marketplace. NASA expected to be a customer, not the operator, of the commercial RLV. The Lockheed Martin X-33 design was based on a lifting body shape with two revolutionary 'linear aerospike' rocket engines and a rugged metallic thermal protection system. The vehicle also had lightweight components and fuel tanks built to conform to the vehicle's outer shape

  9. The X-33 Flight Test Challenge

    NASA Technical Reports Server (NTRS)

    Borden, David; Ramiscal, Ermin; Howell, John

    1999-01-01

    Low cost access to space has eluded present launch system technologies. Our objective is to reduce the cost of putting a payload into space from $10,000 per pound to $1000 per pound. In July 1996, a cooperative agreement was initiated between the Lockheed Martin Skunk Works and NASA to help accomplish this goal. The X-33 is the first step in the process to make low cost space access a reality. The X-33 is a suborbital, hypersonic lifting body, proof of concept of a reusable launch vehicle. The X-33 flight test program will validate technologies such as a metallic thermal protection system, Linear Aerospike Engines, use of tanks and struts as fundamental structural elements, as well as quick turnaround time. Flight testing will begin in July 2000, with launches originating from Edwards Air Force Base and initial landings at Michael Army Airfield in Utah. Data collected from these flight tests will aid in the decision to build an economically viable single stage to orbit reusable launch vehicle. This paper will explore the technical challenges facing the X-33 Flight Test Team.

  10. [X-33 Systems

    NASA Technical Reports Server (NTRS)

    1999-01-01

    Lockheed Martin Skunk Works has compiled an Annual Performance Report of the X-33/RLV Program. This report consists of individual reports from all industry team members, as well as NASA team centers. This portion of the report is comprised of a status report of Allied-Signal Aerospace's contribution to the program. The following is a summary of the work reviewed under their portion of the agreement: (1) Communication Systems; (2) Environmental Control Systems- Active Thermal Control System (ATCS), Purge and Vent System, Hydrogen Detection System (HDS), Avionics Bay Inerting System (ABIS), and Flush Air Data System (FADS); (2) Landing Systems; (3) Power Management and Generation Systems; (4) Flight Control Actuation System (FCAS)- Electric Power Control & Distribution System (EPCDS), and Battery Power System (BPS); and (5) Vehicle Management Systems (VMS)- VMS Hardware, VMS Software Development Activities, and System Integration Laboratory (SIL).

  11. Design, Development,and Testing of Umbillical System Mechanisms for the X-33 Advanced Technology Demonstrator

    NASA Technical Reports Server (NTRS)

    Littlefield, Alan C.; Melton, Gregory S.

    1999-01-01

    The X-33 Advanced Technology Demonstrator is an un-piloted, vertical take-off, horizontal landing spacecraft. The purpose of the X-33 program is to demonstrate technologies that will dramatically lower the cost of access to space. The rocket-powered X-33 will reach an altitude of up to 100 km and speeds between Mach 13 and 15. Fifteen flight tests are planned, beginning in 2000. Some of the key technologies demonstrated will be the linear aerospike engine, improved thermal protection systems, composite fuel tanks and reduced operational timelines. The X-33 vehicle umbilical connections provide monitoring, power, cooling, purge, and fueling capability during horizontal processing and vertical launch operations. Two "rise-ofF' umbilicals for the X-33 have been developed, tested, and installed. The X-33 umbilical systems mechanisms incorporate several unique design features to simplify horizontal operations and provide reliable disconnect during launch.

  12. X-33 RCS model

    NASA Technical Reports Server (NTRS)

    1998-01-01

    Part of the high pressure nitrogen system used for the 1% scale X-33 reaction control system model. Installed in the Unitary Plan Wind Tunnel for supersonic testing. In building 1251, test section #2.

  13. X-33 RCS model

    NASA Technical Reports Server (NTRS)

    1998-01-01

    Model support system and instumentation cabling of the 1% scale X-33 reaction control system model. Installed in the Unitary Plan Wind Tunnel for supersonic testing. In building 1251, test section #2.

  14. X-33 Development History

    NASA Technical Reports Server (NTRS)

    Butrica, Andrew J.

    1997-01-01

    The problem of dealing with various types of proprietary documents, whether from the Lockheed Martin, the Skunk Works, McDonnell Douglas, Rockwell, and other corporations extant or extinct, remains unresolved. The computerized archive finding aid has over 100 records at present. These records consist of X-33 photographs, press releases, media clippings, and the small number of X-33 project records collected to date.

  15. The X-33 Program Update

    NASA Technical Reports Server (NTRS)

    Dill, Charlie

    2000-01-01

    This viewgraph presentation gives an overview of the X-33 program update, including details on program objectives and plans, the X-33 configuration, technologies used, and X-33 assembly and test status.

  16. Hypersonic Boundary-Layer Transition for X-33 Phase 2 Vehicle

    NASA Technical Reports Server (NTRS)

    Thompson, Richard A.; Hamilton, Harris H., II; Berry, Scott A.; Horvath, Thomas J.; Nowak, Robert J.

    1998-01-01

    A status review of the experimental and computational work performed to support the X-33 program in the area of hypersonic boundary-layer transition is presented. Global transition fronts are visualized using thermographic phosphor measurements. Results are used to derive transition correlations for "smooth body" and discrete roughness data and a computational tool is developed to predict transition onset for X-33 using these results. The X-33 thermal protection system appears to be conservatively designed for transition effects based on these studies. Additional study is needed to address concerns related to surface waviness. A discussion of future test plans is included.

  17. X-33 by Lockheed Martin on Launch Pad - Computer Graphic

    NASA Technical Reports Server (NTRS)

    1996-01-01

    This is an artist's conception of the X-33 technology demonstrator on its launch pad, ready for lift-off into orbit. NASA's Dryden Flight Research Center, Edwards, California, expected to play a key role in the development and flight testing of the X-33, which was a technology demonstrator vehicle for a possible Reusable Launch Vehicle (RLV). The RLV technology program was a cooperative agreement between NASA and industry. The goal of the RLV technology program was to enable significant reductions in the cost of access to space, and to promote the creation and delivery of new space services and other activities that would improve U.S. economic competitiveness. The X-33 was a wedged-shaped subscale technology demonstrator prototype of a potential future Reusable Launch Vehicle (RLV) that Lockheed Martin had dubbed VentureStar. The company hoped to develop VentureStar early this century. Through demonstration flight and ground research, NASA's X-33 program was to have provided the information needed for industry representatives such as Lockheed Martin to decide whether to proceed with the development of a full-scale, commercial RLV program. A full-scale, single-stage-to-orbit RLV was to have dramatically increase reliability and lowered costs of putting a pound of payload into space, from the current figure of $10,000 to $1,000. Reducing the cost associated with transporting payloads in Low Earth Orbit (LEO) by using a commercial RLV was to have created new opportunities for space access and significantly improved U.S. economic competitiveness in the world-wide launch marketplace. NASA expected to be a customer, not the operator, of the commercial RLV. The X-33 design was based on a lifting body shape with two revolutionary 'linear aerospike' rocket engines and a rugged metallic thermal protection system. The vehicle also had lightweight components and fuel tanks built to conform to the vehicle's outer shape. Time between X-33 flights was normally to have been seven

  18. X-33 by Lockheed Martin above Earth - Computer Graphic

    NASA Technical Reports Server (NTRS)

    1996-01-01

    This artist's rendering depicts the NASA/Lockheed Martin X-33 technology demonstrator for a Single-Stage-To-Orbit (SSTO) Reusable Launch Vehicle (RLV) in orbit over the Earth. NASA's Dryden Flight Research Center, Edwards, California., expected to play a key role in the development and flight testing of the X-33. The RLV technology program was a cooperative agreement between NASA and industry. The goal of the RLV technology program was to enable significant reductions in the cost of access to space, and to promote the creation and delivery of new space services and other activities that was to have improved U.S. economic competitiveness. The X-33 was a wedged-shaped subscale technology demonstrator prototype of a potential future Reusable Launch Vehicle (RLV) that Lockheed Martin had dubbed VentureStar. The company had hoped to develop VentureStar early this century. Through demonstration flight and ground research, NASA's X-33 program was to have provided the information needed for industry representatives such as Lockheed Martin to decide whether to proceed with the development of a full-scale, commercial RLV program. A full-scale, single-stage-to-orbit RLV was to have dramatically increased reliability and lowered the costs of putting a pound of payload into space, from the current figure of $10,000 to $1,000. Reducing the cost associated with transporting payloads in Low Earth Orbit (LEO) by using a commercial RLV was to have created new opportunities for space access and significantly improved U.S. economic competitiveness in the world-wide launch marketplace. NASA expected to be a customer, not the operator, of the commercial RLV. The X-33 design was based on a lifting body shape with two revolutionary 'linear aerospike' rocket engines and a rugged metallic thermal protection system. The vehicle also had lightweight components and fuel tanks built to conform to the vehicle's outer shape. Time between X-33 flights was normally to have been seven days, but the

  19. Aerothermal Test of Metallic TPS for X-33 Reusable Launch Vehicle

    NASA Technical Reports Server (NTRS)

    Sawyer, James Wayne; Hodge, Jefferson; Moore, Brad

    1998-01-01

    An array of metallic Thermal Protection System (TPS) panels including the seals developed for the windward surface of the X-33 vehicle is being tested in the Eight Foot High Temperature Tunnel at the NASA Langley Research Center. These tests are the first aerothermal tests of an X-33 TPS array and will be used to validate the TPS for the X-33 flight program. Specifically, the tests will be used to evaluate the structural and thermal performance of the TPS, the effectiveness of the high temperature seals between adjacent tiles and the durability of the TPS under realistic aerothermal flight conditions. The effect of varying step heights, damage to the seals between adjacent panels, and the use of secondary seals will also be investigated during the test program. The metallic TPS developed for the windward surface of the X-33 and the test program in the Eight Foot High Temperature Tunnel is presented and discussed.

  20. X-33. Phase 2

    NASA Technical Reports Server (NTRS)

    1998-01-01

    In response to the Cooperative Agreement, Lockheed Martin Skunk Works has compiled an Annual Performance Report of the X-33/RLV Program. This report consists of individual reports from all industry team members, as well as NASA team centers. The first milestone was hand delivered to NASA MSFC. The second year has been one of significant accomplishment in which team members have demonstrated their ability to meet vital benchmarks while continuing on the technical adventure of the 20th century.

  1. X-33 Phase 2

    NASA Technical Reports Server (NTRS)

    1997-01-01

    In response to Clause 17 of the Cooperative Agreement NCC8-115, Lockheed Martin Skunk Works has compiled an Annual Performance Report of the X-33/RLV Program. This report consists of individual reports from all industry team members, as well as NASA team centers. Contract award was announced on July 2, 1996 and the first milestone was hand delivered to NASA MSFC on July 17, 1996. The first year has been one of growth and progress as all team members staffed up and embarked on the technical adventure of the 20th century... the ultimate goal . . a Single Stage to Orbit (SSTO) Reuseable Launch Vehicle (RLV).

  2. Multiwall thermal protection system

    NASA Technical Reports Server (NTRS)

    Jackson, L. R. (Inventor)

    1982-01-01

    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.

  3. Thermal protection apparatus

    DOEpatents

    Bennett, Gloria A.; Elder, Michael G.; Kemme, Joseph E.

    1985-01-01

    An apparatus which thermally protects sensitive components in tools used in a geothermal borehole. The apparatus comprises a Dewar within a housing. The Dewar contains heat pipes such as brass heat pipes for thermally conducting heat from heat sensitive components to a heat sink such as ice.

  4. Thermal protection apparatus

    DOEpatents

    Bennett, G.A.; Elder, M.G.; Kemme, J.E.

    1984-03-20

    The disclosure is directed to an apparatus for thermally protecting sensitive components in tools used in a geothermal borehole. The apparatus comprises a Dewar within a housing. The Dewar contains heat pipes such as brass heat pipes for thermally conducting heat from heat sensitive components such as electronics to a heat sink such as ice.

  5. X-33 Proposal by Lockheed Martin - Computer Graphic

    NASA Technical Reports Server (NTRS)

    1996-01-01

    This artist's rendering depicts the Lockheed Martin X-33 for a technology demonstrator of a Single-Stage-To-Orbit (SSTO) Reusable Launch Vehicle (RLV), as submitted in the aerospace company's original proposal. NASA selected Lockheed Martin's design on 2 July 1996. NASA's Dryden Flight research Center, Edwards, California, was to have had a key role in the development and flight testing of the X-33. The RLV technology program was a cooperative agreement between NASA and industry. The goal of the RLV technology program was to enable significant reductions in the cost of access to space, and to promote the creation and delivery of new space services and other activities that was to have improved U.S. economic competitiveness. The X-33 was a wedged-shaped subscale technology demonstrator prototype of a potential future Reusable Launch Vehicle (RLV) that Lockheed Martin had dubbed VentureStar. The company had hoped to develop VentureStar early this century. Through demonstration flight and ground research, NASA's X-33 program was to have provided the information needed for industry representatives such as Lockheed Martin to decide whether to proceed with the development of a full-scale, commercial RLV program. A full-scale, single-stage-to-orbit RLV was to have dramatically increased reliability and lowered the costs of putting a pound of payload into space, from the current figure of $10,000 to $1,000. Reducing the cost associated with transporting payloads in Low Earth Orbit (LEO) by using a commercial RLV was to have created new opportunities for space access and significantly improve U.S. economic competitiveness in the world-wide launch marketplace. NASA expected to be a customer, not the operator, of the commercial RLV. The X-33 design was based on a lifting body shape with two revolutionary 'linear aerospike' rocket engines and a rugged metallic thermal protection system. The vehicle also had lightweight components and fuel tanks built to conform to the vehicle

  6. X-33 Proposal by McDonnell Douglas - Computer Graphic

    NASA Technical Reports Server (NTRS)

    1996-01-01

    aerospike' rocket engines and a rugged metallic thermal protection system. The vehicle also had lightweight components and fuel tanks built to conform to the vehicle's outer shape. Time between X-33 flights was normally to have been seven days, but the program hoped to demonstrate a two-day turnaround between flights during the flight-test phase of the program. The X-33 was to have been an unpiloted vehicle that took off vertically like a rocket and landed horizontally like an airplane. It was to have reached altitudes of up to 50 miles and high hypersonic speeds. The X-33 program was managed by the Marshall Space Flight Center and was to have been launched at a special launch site on Edwards Air Force Base. Due to technical problems with the liquid hydrogen fuel tanks, and the resulting cost increase and time delay, the X-33 program was cancelled in February 2001.

  7. Thermal insulation protection means

    NASA Technical Reports Server (NTRS)

    Dotts, R. L.; Smith, J. A.; Strouhal, G. (Inventor)

    1979-01-01

    A system for providing thermal insulation for portions of a spacecraft which do not exceed 900 F during ascent or reentry relative to the earth's atmosphere is described. The thermal insulation is formed of relatively large flexible sheets of needled Nomex felt having a flexible waterproof coating. The thickness of the felt is sized to protect against projected temperatures and is attached to the structure by a resin adhesive. Vent holes in the sheets allow ventilation while maintaining waterproofing. The system is heat treated to provide thermal stability.

  8. X-33 Experimental Aeroheating at Mach 6 Using Phosphor Thermography

    NASA Technical Reports Server (NTRS)

    Horvath, Thomas J.; Berry, Scott A.; Hollis, Brian R.; Liechty, Derek S.; Hamilton, H. Harris, II; Merski, N. Ronald

    1999-01-01

    The goal of the NASA Reusable Launch Vehicle (RLV) technology program is to mature and demonstrate essential, cost effective technologies for next generation launch systems. The X-33 flight vehicle presently being developed by Lockheed Martin is an experimental Single Stage to Orbit (SSTO) demonstrator that seeks to validate critical technologies and insure applicability to a full scale RLV. As with the design of any hypersonic vehicle, the aeroheating environment is an important issue and one of the key technologies being demonstrated on X-33 is an advanced metallic Thermal Protection System (TPS). As part of the development of this TPS system, the X-33 aeroheating environment is being defined through conceptual analysis, ground based testing, and computational fluid dynamics. This report provides an overview of the hypersonic aeroheating wind tunnel program conducted at the NASA Langley Research Center in support of the ground based testing activities. Global surface heat transfer images, surface streamline patterns, and shock shapes were measured on 0.013 scale (10-in.) ceramic models of the proposed X-33 configuration in Mach 6 air. The test parametrics include angles of attack from -5 to 40 degs, unit Reynolds numbers from 1x106 to 8x106/ft, and body flap deflections of 0, 10, and 20 deg. Experimental and computational results indicate the presence of shock/shock interactions that produced localized heating on the deflected flaps and boundary layer transition on the canted fins. Comparisons of the experimental data to laminar and turbulent predictions were performed. Laminar windward heating data from the wind tunnel was extrapolated to flight surface temperatures and generally compared to within 50 deg F of flight prediction along the centerline. When coupled with the phosphor technique, this rapid extrapolation method would serve as an invaluable TPS design tool.

  9. Orbiter thermal protection system

    NASA Technical Reports Server (NTRS)

    Dotts, R. L.; Curry, D. M.; Tillian, D. J.

    1985-01-01

    The major material and design challenges associated with the orbiter thermal protection system (TPS), the various TPS materials that are used, the different design approaches associated with each of the materials, and the performance during the flight test program are described. The first five flights of the Orbiter Columbia and the initial flight of the Orbiter Challenger provided the data necessary to verify the TPS thermal performance, structural integrity, and reusability. The flight performance characteristics of each TPS material are discussed, based on postflight inspections and postflight interpretation of the flight instrumentation data. Flights to date indicate that the thermal and structural design requirements for the orbiter TPS are met and that the overall performance is outstanding.

  10. Generation of an Aerothermal Data Base for the X33 Spacecraft

    NASA Technical Reports Server (NTRS)

    Roberts, Cathy; Huynh, Loc

    1998-01-01

    The X-33 experimental program is a cooperative program between industry and NASA, managed by Lockheed-Martin Skunk Works to develop an experimental vehicle to demonstrate new technologies for a single-stage-to-orbit, fully reusable launch vehicle (RLV). One of the new technologies to be demonstrated is an advanced Thermal Protection System (TPS) being designed by BF Goodrich (formerly Rohr, Inc.) with support from NASA. The calculation of an aerothermal database is crucial to identifying the critical design environment data for the TPS. The NASA Ames X-33 team has generated such a database using Computational Fluid Dynamics (CFD) analyses, engineering analysis methods and various programs to compare and interpolate the results from the CFD and the engineering analyses. This database, along with a program used to query the database, is used extensively by several X-33 team members to help them in designing the X-33. This paper will describe the methods used to generate this database, the program used to query the database, and will show some of the aerothermal analysis results for the X-33 aircraft.

  11. X-33 Linear Aerospike Engine

    NASA Technical Reports Server (NTRS)

    Vinson, John

    1998-01-01

    In July of 1999 two linear aerospike rocket engines will power the first flight of NASA's X-33 advanced technology demonstrator. A successful X-33 flight test program will validate the aerospike nozzle concept, a key technical feature of Lockheed Martin's VentureStar(trademark) reusable launch vehicle. The aerospike received serious consideration for NASA's current space shuttle, but was eventually rejected in 1969 in favor of high chamber pressure bell engines, in part because of perceived technical risk. The aerospike engine (discussed below) has several performance advantages over conventional bell engines. However, these performance advantages are difficult to validate by ground test. The space shuttle, a multibillion dollar program intended to provide all of NASA's future space lift could not afford the gamble of choosing a potentially superior though unproven aerospike engine over a conventional bell engine. The X-33 demonstrator provides an opportunity to prove the aerospike's performance advantage in flight before commiting to an operational vehicle.

  12. Thermal Protection and Control

    NASA Technical Reports Server (NTRS)

    Greene, Effie E.

    2013-01-01

    During all phases of a spacecraft's mission, a Thermal Protection System (TPS) is needed to protect the vehicle and structure from extreme temperatures and heating. When designing TPS, low weight and cost while ensuring the protection of the vehicle is highly desired. There are two main types of TPS, ablative and reusable. The Apollo missions needed ablators due to the high heat loads from lunar reentry. However, when the desire for a reusable space vehicle emerged, the resultant_ Space Shuttle program propelled a push for the development of reusable TPS. With the growth of reqsable TPS, the need for ablators declined, triggering a drop off of the ablator industry. As a result, the expertise was not heavily maintained within NASA or the industry. When the Orion Program initiated a few years back, a need. for an ablator reemerged. Yet, due to of the lack of industry capability, redeveloping the ablator material took several years and came at a high cost. As NASA looks towards the future with both the Orion and Commercial Crew Programs, a need to preserve reusable, ablative, and other TPS technologies is essential. Research of the different TPS materials alongside their properties, capabilities, and manufacturing process was performed, and the benefits of the materials were analyzed alongside the future of TPS. Knowledge of the different technologies has the ability to help us know what expertise to maintain and ensure a lack in the industry does not occur again.

  13. Thermographic testing used on the X-33 space launch vehicle program by BFGoodrich Aerospace

    NASA Astrophysics Data System (ADS)

    Burleigh, Douglas D.

    1999-03-01

    The X-33 program is a team effort sponsored by NASA under Cooperative Agreement NCC8-115, and led by the Lockheed Martin Corporation. Team member BFGoodrich Aerospace Aerostructures Group (formerly Rohr) is responsible for design, manufacture, and integration of the Thermal Protection System (TPS) of the X-33 launch vehicle. The X-33 is a half-scale, experimental prototype of a vehicle called RLV (Reusable Launch Vehicle) or VentureStarTM, an SSTO (single stage to orbit) vehicle, which is a proposed successor to the aging Space Shuttle. Thermographic testing has been employed by BFGoodrich Aerospace Aerostructures Group for a wide variety of uses in the testing of components of the X-33. Thermographic NDT (TNDT) has been used for inspecting large graphite- epoxy/aluminum honeycomb sandwich panels used on the Leeward Aeroshell structure of the X-33. And TNDT is being evaluated for use in inspecting carbon-carbon composite parts such as the nosecap and wing leading edge components. Pulsed Infrared Testing (PIRT), a special form of TNDT, is used for the routine inspection of sandwich panels made of brazed inconel honeycomb and facesheets. In the developmental and qualification testing of sub-elements of the X-33, thermography has been used to monitor (1) Arc Jet tests at NASA Ames Research Center in Mountain view, CA and NASA Johnson Space Center in Houston, TX, (2) High Temperature (wind) Tunnel Tests (HTT) at Nasa Langley Research Center in Langley, VA, and (3) Hot Gas Tests at NASA Marshall Space Flight Center in Huntsville, AL.

  14. Thermographic Testing Using on the X-33 Space Launch Vehicle Program by BFGoodrich Aerospace

    NASA Technical Reports Server (NTRS)

    Burleigh, Douglas

    1999-01-01

    The X-33 program is a team effort sponsored by NASA, under Cooperative Agreement NCC8-115, and led by the Lockheed Martin Corporation. Team member BFGoodrich Aerospace Aerostructures Group (formerly Rohr) is responsible for design, manufacture, and integration of the Thermal Protection System (TPS) of the X-33 launch vehicle. The X-33 is a half-scale, experimental prototype of a vehicle called RLV (Reusable Launch Vehicle) or VentureStar(Trademark), an SSTO (single stage to orbit) vehicle, which is a proposed successor to the aging Space Shuttle. Thermographic testing has been employed by BFGoodrich Aerospace Aerostructures Group for a wide variety of uses in the testing of components of the X-33. Thermographic NDT (TNDT) has been used for inspecting large graphite-epoxy/aluminum honeycomb sandwich panels used on the Leeward Aeroshell structure of the X-33. And TNDT is being evaluated for use in inspecting carbon-carbon composite parts such as the nosecap and wing leading edge components. Pulsed Infrared Testing (PIRT), a special form of TNDT, is used for the routine inspection of sandwich panels made of brazed inconel honeycomb and facesheets. In the developmental and qualification testing of sub-elements of the X-33, thermography has been used to monitor 1) Arc Jet tests at NASA Ames Research Center in Mountainview, CA and NASA Johnson Space Center in Houston, TX, 2) High Temperature (wind) Tunnel Tests (HTT) at NASA Langley Research Center in Langley, VA, and 3) Hot Gas Tests at NASA Marshall Space Flight Center in Huntsville, AL.

  15. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will design, build, and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604BOO02G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate the aerodynamic flight database for the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. Al these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  16. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database i n the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  17. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database in the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  18. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database in the hypersonic regime, The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  19. Thermal Protection Materials

    NASA Technical Reports Server (NTRS)

    Johnson, Sylvia M.

    2011-01-01

    Thermal protection materials and systems (TPS) are required to protect a vehicle returning from space or entering an atmosphere. The selection of the material depends on the heat flux, heat load, pressure, and shear and other mechanical loads imposed on the material, which are in turn determined by the vehicle configuration and size, location on the vehicle, speed, a trajectory, and the atmosphere. In all cases the goal is to use a material that is both reliable and efficient for the application. Reliable materials are well understood and have sufficient test data under the appropriate conditions to provide confidence in their performance. Efficiency relates to the behavior of a material under the specific conditions that it encounters TPS that performs very well at high heat fluxes may not be efficient at lower heat fluxes. Mass of the TPS is a critical element of efficiency. This talk will review the major classes of TPS, reusable or insulating materials and ablators. Ultra high temperature ceramics for sharp leading edges will also be reviewed. The talk will focus on the properties and behavior of these materials.

  20. X-33 Flight Operations Center

    NASA Technical Reports Server (NTRS)

    1999-01-01

    In response to Clause 17 of the Cooperative Agreement NCC8-115, Lockheed Martin Skunk Works has compiled an Annual Performance Report of the X-33/RLV Program. This report consists of individual reports from all industry team members, as well as NASA team centers. Contract award was announced on July 2, 1996 and the first milestone was hand delivered to NASA MSFC on July 17, 1996. With the dedication of the launch site, and continuing excellence in technological achievement, the third year of the Cooperative Agreement has been one of outstanding accomplishment and excitement.

  1. Thermal protection systems for aerobrakes

    NASA Technical Reports Server (NTRS)

    Tompkins, Stephen S.

    1993-01-01

    In summary, advantages of the ablative thermal protection system (TPS) for aerobrakes are: (1) proven reliable TPS systems; (2) well characterized (thermally) with good, existing thermal analysis capability; (3) good candidate materials are available; (4) not sensitive to defects and more difficult to damage then RSI or C-C; (5) design program which demonstrated simple (direct bond) application of large panels; (6) thermal excursions not catastrophic; and (7) no SIP required.

  2. X33 Transient Liftoff Analysis

    NASA Technical Reports Server (NTRS)

    Peck, Jeff; Brunty, Joseph

    2000-01-01

    The successful design of a launch vehicle requires the careful characterization of the various loads the structure will experience over its lifetime. Many of the most demanding load environments occur during the launch/ascent phase of a mission, typically defined as the point of engine start through engine cut off. One of the critical events during the launch phase is the liftoff event. This event imparts high loads on the vehicle due to transient events such as thrust build-up and vehicle release. This paper describes the theory and procedures used to calculate structural loads due to the liftoff event for the Lockheed-Martin X33 technology demonstrator vehicle. These procedures were developed at NASA's Marshall Space Flight Center and verified previously on other advanced launch system concepts and the Space Shuttle system.

  3. Ablative Thermal Protection System Fundamentals

    NASA Technical Reports Server (NTRS)

    Beck, Robin A. S.

    2013-01-01

    This is the presentation for a short course on the fundamentals of ablative thermal protection systems. It covers the definition of ablation, description of ablative materials, how they work, how to analyze them and how to model them.

  4. Lightweight Thermal-Protection System

    NASA Technical Reports Server (NTRS)

    Macconochie, I. O.; Lawson, A. G.; Whiteman, T. C.; Brien, E. P.

    1983-01-01

    Hexagonal honeycomb panels secured by Y-shaped plates form lightweight, easily-maintained thermal-protection system. Honeycomb outer panel and fastener materials are selected to match local heating rates. Typical materials include composites, titanium, superalloys, and refractory metals. Advantages include complete symmetry of components--there are no left- or right-hand parts and no asymmetry in thermal expansion.

  5. Thermal Management and Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Hasnain, Aqib

    2016-01-01

    's rays directly impinging on the system. Heating rate of the lamps were calculated by knowing fraction of emitted energy in a wavelength interval and the filament temperature. This version of the model can be used to predict performance of the system under vacuum with extreme cold or hot conditions. Initial testing of the PTMS showed promise, and the thermal math model predicts even better performance in thermal vacuum testing. ii) Thermal Protection Systems (TPS) are required for vehicles which enter earth's atmosphere to protect from aerodynamic heating caused by the friction between the vehicle and atmospheric gases. Orion's heat shield design has two aspects which needed to be analyzed thermally: i) a small excess of adhesive used to bond the outer AVCOAT layer to the inner composite structure tends to seep from under the AVCOAT and form a small bead in between two bricks of AVCOAT, ii) a silicone rubber with different thermophysical properties than AVCOAT fills the gap between two bricks of AVCOAT. I created a thermal model using TD to determine temperature differences that are caused by these two features. To prevent false results, all TD models must be verified against something known. In this case, the TD model was correlated to CHAR, an ablation modelling software used to analyze TPS. Analyzing a node far from the concerning features, we saw that the TD model data match CHAR data, verifying the TD model. Next, the temperature of the silicone rubber as well as the bead of adhesive were analyzed to determine if they exceeded allowable temperatures. It was determined that these two features do not have a significant effect on the max temperature of the heat shield. This model can be modified to check temperatures at various locations of the heat shield where the composite thickness varies.

  6. X-33 Venture Star - Reusable Launch Vehicle

    NASA Technical Reports Server (NTRS)

    1996-01-01

    In this artist's concept, the X-33 Venture Star, a Reusable Launch Vehicle (RLV), manufactured by Lockheed Martin Skunk Works, is shown in orbit with a deployed payload. The Venture Star was one of the earliest versions of the RLV's developed to replace the aging shuttle fleet. The X-33 program was cancelled in 2001.

  7. X-33 Launch - Computer generated graphic

    NASA Technical Reports Server (NTRS)

    1996-01-01

    This 45-second computer-generated launch sequence begins with a view of the X-33 launch facility located near Haystack Butte on the test range at Edwards AFB, California.The X-33 vehicle is then (hypothetically) raised into position, fueled, and launched, making its roll maneuver and then proceeding on its flightpath.

  8. X-33 Leading the Way to VentureStar(Trademark) in this Decade

    NASA Technical Reports Server (NTRS)

    Austin, Robert E.; Rising, Jerry J.

    2000-01-01

    The X-33, reusable space plane technology demonstrator is on course to begin the flights of the X-33 by the end of 2002 that will serve as a basis for industry and government decisions that could lead to VentureStar(Trademark). Lockheed Martin has placed the VentureStar(Trademark) LLC in it's Space Company and is now competing in an industry wide effort that will permit NASA to select a Second Generation RLV source by 2005. This move provides the focus for firm business planning needed to enable the decision by the time X-33 flies in mid 2002 and possibly with upgraded technologies a year or so later. Since the IAF 50th Congress in Amsterdam, most of the major hardware elements of X-33 have been through their assembly and test. The flight liquid oxygen tank was the first major element to complete final assembly. Aerospike Engine qualification testing has progressed successfully through its test objectives and the two flight engines are in preparation to be delivered to the Assembly Facility in Palmdale. All Thermal Protection System (TPS) metallic panels have completed qualification testing and have been delivered to Palmdale and all remaining TPS elements have been assembled and are ready for delivery. Flight Software and Avionics have been delivered and are in integration testing. In November 1999, the first graphite composite liquid hydrogen tank experienced a debond between the tank inner skin and the honeycomb core in testing. This tank had completed its third successful cryogenic and loads testing at MSFC. Replacement liquid hydrogen tanks have completed design and are in fabrication. The resulting delay from this change of design for the liquid hydrogen tank will be approximately two years.

  9. The Effect of Metallic TPS Panel Bowing on the Surface Heating of the X-33 Vehicle

    NASA Technical Reports Server (NTRS)

    Palmer, Grant; Kontinos, Dean; Langhoff, Stephen R. (Technical Monitor)

    1997-01-01

    The thermal protection system of the windward surface of the X-33 vehicle consists of metallic honeycomb sandwich panels. Thermal gradients experienced during the descent phase of the trajectory result in a different rate of thermal expansion between the inner and outer face sheets of the metallic panels. This causes the panels to bow outward when the temperature of the outer face sheet is larger than that of the inner face sheet and inward when the temperature of the outer face sheet is less than that of he inner face sheet. This results in a quilted-type body surface. Using computational fluid dynamic analysis, this study will determine the effect the metallic TPS panel bowing has on the surface heating.

  10. Risk Reduction on X-33/RLV Engines

    NASA Technical Reports Server (NTRS)

    Crowley, Tim

    1999-01-01

    Risk management has received considerable attention in the X-33 and Reusable Launch Vehicle (RLV) program due to aggressive schedules, limited funding. and planned private investment to develop the commercial VentureStar vehicle. As an X-33 and RLV team member and main propulsion supplier, Boeing Rocketdyn Propulsion and Power has addressed risk through a methodical application of systems engineering in identifying, assessing, and mitigating risks. The methods employed involve rigorous risk mitigation planning early in development, continuous risk monitoring and assessment during the course of development, and the systematic verification of compliance with technical requirements prior to delivery. In addition, an engine system reliability analysis was conducted to reduce risk. In July 1996, NASA selected Lockheed Martin's "Skunk Works" (LMSW) as the lead contractor for the X-33 and RLV program. The X-33 vehicle is a half-scale pathfinder for the full-scale RLV. The LMSW RLV design is a lifting body shaped vehicle employing linear aerospike engine provided propulsion. The initial X-33 flight is planned for the summer of 2000, and the initial VentureStar flight is planned for between 2005 and 2007.

  11. RLV/X-33 operations overview

    SciTech Connect

    Black, S.T.; Eshleman, W.

    1997-01-01

    This paper describes the VentureStar{trademark} SSTO RLV and X-33 operations concepts. Applications of advanced technologies, automated ground support systems, advanced aircraft and launch vehicle lessons learned have been integrated to develop a streamlined vehicle and mission processing concept necessary to meet the goals of a commercial SSTO RLV. These concepts will be validated by the X-33 flight test program where financial and technical risk mitigation are required. The X-33 flight test program totally demonstrates the vehicle performance, technology, and efficient ground operations at the lowest possible cost. The Skunk Work{close_quote}s test program approach and test site proximity to the production plant are keys. The X-33 integrated flight and ground test program incrementally expands the knowledge base of the overall system allowing minimum risk progression to the next flight test program milestone. Subsequent X-33 turnaround processing flows will be performed with an aircraft operations philosophy. The differences will be based on research and development, component reliability and flight test requirements. {copyright} {ital 1997 American Institute of Physics.}

  12. RLV/X-33 operations overview

    NASA Astrophysics Data System (ADS)

    Black, Stephen T.; Eshleman, Wally

    1997-01-01

    This paper describes the VentureStar™ SSTO RLV and X-33 operations concepts. Applications of advanced technologies, automated ground support systems, advanced aircraft and launch vehicle lessons learned have been integrated to develop a streamlined vehicle and mission processing concept necessary to meet the goals of a commercial SSTO RLV. These concepts will be validated by the X-33 flight test program where financial and technical risk mitigation are required. The X-33 flight test program totally demonstrates the vehicle performance, technology, and efficient ground operations at the lowest possible cost. The Skunk Work's test program approach and test site proximity to the production plant are keys. The X-33 integrated flight and ground test program incrementally expands the knowledge base of the overall system allowing minimum risk progression to the next flight test program milestone. Subsequent X-33 turnaround processing flows will be performed with an aircraft operations philosophy. The differences will be based on research and development, component reliability and flight test requirements.

  13. [X-33 Launch and Landing Facilities

    NASA Technical Reports Server (NTRS)

    1999-01-01

    Sverdrup is responsible for the design, construction and activation of the X-33 Flight Operations Center at Edwards Air Force Base and for providing assistance in activating the X-33 Landing Sites. The past year has seen the completion of the construction of the X-33 Flight Operations Center. Construction was completed in December of 1998, with systems checkout and testing continuing into early 1999. Integration of the site with LMCMS and other partner-supplied systems began in December and will continue through rollout of the X-33 vehicle. The construction of the X-33 Launch Complex has been performed within the Edwards AFB and Air Force Research Laboratory (AFRL) systems with no substantial interference to either parties. A high level of cooperation exists between Sverdrup, Edwards AFB, and the Air Force Research Laboratory in the areas of access, training, security, and operations. There have been no conflicts between programs that have not been accommodated. Development of the landing sites is progressing with many of the modifications necessary underway. GSE commitments are in place. The personnel training program developed by Sverdrup for persons entering the launch site construction areas, was modified by Lockheed for use in training and access control to the Center during flight operations to maximize safety and minimize intrusion upon the environment. Close cooperation between Sverdrup, the construction workers, and the environmental biologist permitted construction to proceed in a timely fashion without harm to the wildlife, in particular, the Desert Tortoise. Although the entire X-33 site encompasses approximately 50 acres including a new access road, only the areas directly impacted by the construction were cleared to minimize the impact on the environment. A total of about 30 acres was actually disturbed.

  14. [X-33 Research By NASA Centers

    NASA Technical Reports Server (NTRS)

    1999-01-01

    Lockheed Martin Skunk Works has compiled an Annual Performance Report of the X-33/RLV Program. This report consists of individual reports from all industry team members, as well as NASA team centers. This portion of the report is comprised of overviews of each NASA Center's contribution to the program during the period 1 Apr. 1998 - 31 Mar. 1999.

  15. X-33 Landing - Computer generated graphic

    NASA Technical Reports Server (NTRS)

    1996-01-01

    This 46-second clip has the X-33 aircraft on final approach to Michael AAF in Utah, then with its landing gear down, it flares for touchdown and brakes to a halt. This graphic like the three before it shows an early configuration without vertical stabilizers, which have since been added.

  16. Thermal protection system ablation sensor

    NASA Technical Reports Server (NTRS)

    Gorbunov, Sergey (Inventor); Martinez, Edward R. (Inventor); Scott, James B. (Inventor); Oishi, Tomomi (Inventor); Fu, Johnny (Inventor); Mach, Joseph G. (Inventor); Santos, Jose B. (Inventor)

    2011-01-01

    An isotherm sensor tracks space vehicle temperatures by a thermal protection system (TPS) material during vehicle re-entry as a function of time, and surface recession through calibration, calculation, analysis and exposed surface modeling. Sensor design includes: two resistive conductors, wound around a tube, with a first end of each conductor connected to a constant current source, and second ends electrically insulated from each other by a selected material that becomes an electrically conductive char at higher temperatures to thereby complete an electrical circuit. The sensor conductors become shorter as ablation proceeds and reduced resistance in the completed electrical circuit (proportional to conductor length) is continually monitored, using measured end-to-end voltage change or current in the circuit. Thermocouple and/or piezoelectric measurements provide consistency checks on local temperatures.

  17. Guidance and Control Concepts for the X-33 Technology Demonstrator

    NASA Technical Reports Server (NTRS)

    Dukeman, Gregory A.; Gallaher, Michael W.

    1998-01-01

    The X-33 technology demonstrator is a suborbital precursor to the Reusable Launch Vehicle (RLV) with first flight planned for summer of 1999. The flight test program will include about 15 flights originating from Edwards Air Force Base, California, each with widely varying flight profiles in order to test new thermal protection system (TPS) materials, structures, and linear aerospike engines. The first flights will be relatively short range flights with about a 300 nmi range, maximum Mach number of 7, maximum altitude of 190,000 feet, whereas the latter flights will cover about 800 nmi range, with max altitude of about 260,000 feet and max Mach of about 15. The guidance algorithms must be flexible enough to accommodate these various profiles and to adapt to severe off-nominal dispersions, such as early engine failure (partial or total) where possibly more than half the thrust is lost. An onboard real-time performance monitor will be used to assess the viability of the nominal landing site as well as alternate landing sites that would potentially be used in extreme off-nominal conditions. During ascent, a single entry guidance-related parameter, which is easy to calculate, is used to assess the viability of the nominal landing site as well as alternate landing sites. Real-time adjustment of the stored ascent attitude profile will be performed, as required, to maximize the probability of making it to the nominal landing site. Numerical results are given for various engine-out cases to illustrate the adaptability of the performance monitor.

  18. Thermal protection system and related methods

    NASA Technical Reports Server (NTRS)

    Garbe, Duane J. (Inventor)

    2012-01-01

    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.

  19. Current Technology for Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Scotti, Stephen J. (Compiler)

    1992-01-01

    Interest in thermal protection systems for high-speed vehicles is increasing because of the stringent requirements of such new projects as the Space Exploration Initiative, the National Aero-Space Plane, and the High-Speed Civil Transport, as well as the needs for improved capabilities in existing thermal protection systems in the Space Shuttle and in turbojet engines. This selection of 13 papers from NASA and industry summarizes the history and operational experience of thermal protection systems utilized in the national space program to date, and also covers recent development efforts in thermal insulation, refractory materials and coatings, actively cooled structures, and two-phase thermal control systems.

  20. VentureStar(trademark) Reaping the Benefits of the X-33 Program

    NASA Technical Reports Server (NTRS)

    Sumrall, J.; Lane, C.

    1998-01-01

    Major X-33 flight hardware has been delivered, and assembly of the vehicle is well underway in anticipation of its flight test program commencing in the summer of 1999. Attention has now turned to the operational VentureStar(trademark), the first single-stage-to-orbit (SSTO) reusable launch vehicle. Activities are grouped under two broad categories: (1) vehicle development and (2) market/business planning, each of which is discussed. The mission concept is presented for direct payload delivery to the International Space Station and to low Earth orbit, as well as payload delivery with an upper stage to Geosynchronous Transfer Orbit (GTO) and other high energy orbits. System requirements include flight segment and ground segment. Vehicle system sizing and design status is provided including the application of X-33 traceability and lessons learned. Technology applications to the VentureStar(trademark) are described including the structure, propellant tanks, thermal protection system, aerodynamics, subsystems, payload bay and propulsion. Developing a market driven low cost launch services system for the 21 st Century requires traditional and non-traditional ways of being able to forecast the evolution of the potential market. The challenge is balancing both the technical and financial assumptions of the market. This involves the need to provide a capability to meet market segments that in some cases are very speculative, while at the same time providing the financial community with a credible revenue stream.

  1. Development of Processing Techniques for Advanced Thermal Protection Materials

    NASA Technical Reports Server (NTRS)

    Selvaduray, Guna; Cox, Michael; Srinivasan, Vijayakumar

    1997-01-01

    Thermal Protection Materials Branch (TPMB) has been involved in various research programs to improve the properties and structural integrity of the existing aerospace high temperature materials. Specimens from various research programs were brought into the analytical laboratory for the purpose of obtaining and refining the material characterization. The analytical laboratory in TPMB has many different instruments which were utilized to determine the physical and chemical characteristics of materials. Some of the instruments that were utilized by the SJSU students are: Scanning Electron Microscopy (SEM), Energy Dispersive X-ray analysis (EDX), X-ray Diffraction Spectrometer (XRD), Fourier Transform-Infrared Spectroscopy (FTIR), Ultra Violet Spectroscopy/Visible Spectroscopy (UV/VIS), Particle Size Analyzer (PSA), and Inductively Coupled Plasma Atomic Emission Spectrometer (ICP-AES). The above mentioned analytical instruments were utilized in the material characterization process of the specimens from research programs such as: aerogel ceramics (I) and (II), X-33 Blankets, ARC-Jet specimens, QUICFIX specimens and gas permeability of lightweight ceramic ablators. In addition to analytical instruments in the analytical laboratory at TPMB, there are several on-going experiments. One particular experiment allows the measurement of permeability of ceramic ablators. From these measurements, physical characteristics of the ceramic ablators can be derived.

  2. Computer graphic of Lockheed Martin X-33 Reusable Launch Vehicle (RLV) mounted on NASA 747 ferry air

    NASA Technical Reports Server (NTRS)

    1997-01-01

    This is an artist's conception of the NASA/Lockheed Martin X-33 Advanced Technology Demonstrator being carried on the back of the 747 Shuttle Carrier Aircraft. This was a concept for moving the X-33 from its landing site back to NASA's Dryden Flight Research Center, Edwards, California. The X-33 was a technology demonstrator vehicle for the Reusable Launch Vehicle (RLV). The RLV technology program was a cooperative agreement between NASA and industry. The goal of the RLV technology program was to enable significant reductions in the cost of access to space, and to promote the creation and delivery of new space services and other activities that will improve U.S. economic competitiveness. NASA Headquarter's Office of Space Access and Technology oversaw the RLV program, which was being managed by the RLV Office at NASA's Marshall Space Flight Center, located in Huntsville, Alabama. Responsibilities of other NASA Centers included: Johnson Space Center, Houston, Texas, guidance navigation and control technology, manned space systems, and health technology; Ames Research Center, Mountain View, CA., thermal protection system testing; Langley Research Center, Langley, Virginia, wind tunnel testing and aerodynamic analysis; and Kennedy Space Center, Florida, RLV operations and health management. Lockheed Martin's industry partners in the X-33 program are: Astronautics, Inc., Denver, Colorado, and Huntsville, Alabama; Engineering & Science Services, Houston, Texas; Manned Space Systems, New Orleans, LA; Sanders, Nashua, NH; and Space Operations, Titusville, Florida. Other industry partners are: Rocketdyne, Canoga Park, California; Allied Signal Aerospace, Teterboro, NJ; Rohr, Inc., Chula Vista, California; and Sverdrup Inc., St. Louis, Missouri.

  3. X-33/RLV Program Aerospike Engines

    NASA Technical Reports Server (NTRS)

    1999-01-01

    Substantial progress was made during the past year in support of the X-33/RLV program. X-33 activity was directed towards completing the remaining design work and building hardware to support test activities. RLV work focused on the nozzle ramp and powerpack technology tasks and on supporting vehicle configuration studies. On X-33, the design activity was completed to the detail level and the remainder of the drawings were released. Component fabrication and engine assembly activity was initiated, and the first two powerpacks and the GSE and STE needed to support powerpack testing were completed. Components fabrication is on track to support the first engine assembly schedule. Testing activity included powerpack testing and component development tests consisting of thrust cell single cell testing, CWI system spider testing, and EMA valve flow and vibration testing. Work performed for RLV was divided between engine system and technology development tasks. Engine system activity focused on developing the engine system configuration and supporting vehicle configuration studies. Also, engine requirements were developed, and engine performance analyses were conducted. In addition, processes were developed for implementing reliability, mass properties, and cost controls during design. Technology development efforts were divided between powerpack and nozzle ramp technology tasks. Powerpack technology activities were directed towards the development of a prototype powerpack and a ceramic turbine technology demonstrator (CTTD) test article which will allow testing of ceramic turbines and a close-coupled gas generator design. Nozzle technology efforts were focused on the selection of a composite nozzle supplier and on the fabrication and test of composite nozzle coupons.

  4. Flutter Analysis of the X-33

    NASA Technical Reports Server (NTRS)

    Fowler, Samuel B.

    2000-01-01

    Flutter analysis performed in support of the X33 Advanced Technology Demonstrator is described. Analysis was conducted over a range of flow regimes using several different analysis codes. The finite element and aerodynamic models used in the analysis have undergone several years of development and refinement resulting in a high degree of model detail. The flutter analysis focuses on the area of three critical points within the vehicle's design trajectory at which full sets of external loads have previously been developed. A comparison between several different aerodynamic models is also made for the selected trajectory points.

  5. X-33 Hypersonic Boundary Layer Transition

    NASA Technical Reports Server (NTRS)

    Berry, Scott A.; Horvath, Thomas J.; Hollis, Brian R.; Thompson, Richard A.; Hamilton, H. Harris, II

    1999-01-01

    Boundary layer and aeroheating characteristics of several X-33 configurations have been experimentally examined in the Langley 20-Inch Mach 6 Air Tunnel. Global surface heat transfer distributions, surface streamline patterns, and shock shapes were measured on 0.013-scale models at Mach 6 in air. Parametric variations include angles-of-attack of 20-deg, 30-deg, and 40-deg; Reynolds numbers based on model length of 0.9 to 6.6 million; and body-flap deflections of 0, 10 and 20-deg. The effects of discrete and distributed roughness elements on boundary layer transition, which included trip height, size, location, and distribution, both on and off the windward centerline, were investigated. The discrete roughness results on centerline were used to provide a transition correlation for the X-33 flight vehicle that was applicable across the range of reentry angles of attack. The attachment line discrete roughness results were shown to be consistent with the centerline results, as no increased sensitivity to roughness along the attachment line was identified. The effect of bowed panels was qualitatively shown to be less effective than the discrete trips; however, the distributed nature of the bowed panels affected a larger percent of the aft-body windward surface than a single discrete trip.

  6. Thermal protection in space technology

    NASA Technical Reports Server (NTRS)

    Salakhutdinov, G. M.

    1982-01-01

    The provision of heat protection for various elements of space flight apparata has great significance, particularly in the construction of manned transport vessels and orbital stations. A popular explanation of the methods of heat protection in rocket-space technology at the current stage as well as in perspective is provided.

  7. Current technology for thermal protection systems

    SciTech Connect

    Scotti, S.J.

    1992-10-01

    Interest in thermal protection systems for high-speed vehicles is increasing because of the stringent requirements of such new projects as the Space Exploration Initiative, the National Aero-Space Plane, and the High-Speed Civil Transport, as well as the needs for improved capabilities in existing thermal protection systems in the Space Shuttle and in turbojet engines. This selection of 13 papers from NASA and industry summarizes the history and operational experience of thermal protection systems utilized in the national space program to date, and also covers recent development efforts in thermal insulation, refractory materials and coatings, actively cooled structures, and two-phase thermal control systems. Separate abstracts were prepared for papers of this report.

  8. Transient thermal analysis of a titanium multiwall thermal protection system

    NASA Technical Reports Server (NTRS)

    Blosser, M. L.

    1982-01-01

    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.

  9. Thermal response of integral multicomponent composite thermal protection systems

    NASA Technical Reports Server (NTRS)

    Stewart, D. A.; Leiser, D. B.; Smith, M.; Kolodziej, P.

    1985-01-01

    Integral-multicomponent thermal-protection materials are discussed in terms of their thermal response to an arc-jet airstream. In-depth temperature measurements are compared with predictions from a one-dimensional, finite-difference code using calculated thermal conductivity values derived from an engineering model. The effect of composition, as well as the optical properties of the bonding material between components, on thermal response is discussed. The performance of these integral-multicomponent composite materials is compared with baseline Space Shuttle insulation.

  10. Large thermal protection system panel

    NASA Technical Reports Server (NTRS)

    Myers, Franklin K. (Inventor); Weinberg, David J. (Inventor); Tran, Tu T. (Inventor)

    2003-01-01

    A protective panel for a reusable launch vehicle provides enhanced moisture protection, simplified maintenance, and increased temperature resistance. The protective panel includes an outer ceramic matrix composite (CMC) panel, and an insulative bag assembly coupled to the outer CMC panel for isolating the launch vehicle from elevated temperatures and moisture. A standoff attachment system attaches the outer CMC panel and the bag assembly to the primary structure of the launch vehicle. The insulative bag assembly includes a foil bag having a first opening shrink fitted to the outer CMC panel such that the first opening and the outer CMC panel form a water tight seal at temperatures below a desired temperature threshold. Fibrous insulation is contained within the foil bag for protecting the launch vehicle from elevated temperatures. The insulative bag assembly further includes a back panel coupled to a second opening of the foil bag such that the fibrous insulation is encapsulated by the back panel, the foil bag, and the outer CMC panel. The use of a CMC material for the outer panel in conjunction with the insulative bag assembly eliminates the need for waterproofing processes, and ultimately allows for more efficient reentry profiles.

  11. System Identification of X-33 Neural Network

    NASA Technical Reports Server (NTRS)

    Aggarwal, Shiv

    2003-01-01

    Modern flight control research has improved spacecraft survivability as its goal. To this end we need to have a failure detection system on board. In case the spacecraft is performing imperfectly, reconfiguration of control is needed. For that purpose we need to have parameter identification of spacecraft dynamics. Parameter identification of a system is called system identification. We treat the system as a black box which receives some inputs that lead to some outputs. The question is: what kind of parameters for a particular black box can correlate the observed inputs and outputs? Can these parameters help us to predict the outputs for a new given set of inputs? This is the basic problem of system identification. The X33 was supposed to have the onboard capability of evaluating the current performance and if needed to take the corrective measures to adapt to desired performance. The X33 is comprised of both rocket and aircraft vehicle design characteristics and requires, in general, analytical methods for evaluating its flight performance. Its flight consists of four phases: ascent, transition, entry and TAEM (Terminal Area Energy Management). It spends about 200 seconds in ascent phase, reaching an altitude of about 180,000 feet and a speed of about 10 to 15 Mach. During the transition phase which lasts only about 30 seconds, its altitude may increase to about 190,000 feet but its speed is reduced to about 9 Mach. At the beginning of this phase, the Main Engine is Cut Off (MECO) and the control is reconfigured with the help of aerosurfaces (four elevons, two flaps and two rudders) and reaction control system (RCS). The entry phase brings down the altitude of X33 to about 90,000 feet and its speed to about Mach 3. It spends about 250 seconds in this phase. Main engine is still cut off and the vehicle is controlled by complex maneuvers of aerosurfaces. The last phase TAEM lasts for about 450 seconds and the altitude and speed, both are reduced to zero. The

  12. HOPE and its thermal protection systems

    NASA Astrophysics Data System (ADS)

    Morino, Yoshiki

    HOPE (H-II Orbiting Plane) is an unmanned winged space vehicle which will be launched by the H-II rocket. HOPE will establish technologies for development of the space plane such as rendezvous and docking, deorbiting, reentry, automatic landing technologies, etc. One of the key technologies concerns the development of a lightweight airframe composed of a thermal protection system and heat resistant composite material. Efforts of R&D activities on the HOPE airframe have been concentrated on material development to withstand severe reentry thermal conditions. Polyimide-matrix composite materials and carbon/carbon materials are being studied for application to the main structure. Thermal protection systems being developed to protect the main structure from reentry heating environment include carbon/carbon systems, ceramic tile systems, and titanium multiwall systems. The development status of these candidate systems and materials is reported in this paper.

  13. Reusable thermal protection system development: A prospective

    NASA Astrophysics Data System (ADS)

    Goldstein, Howard

    1992-10-01

    The state of the art in passive reusable thermal protection system materials is described. Development of the Space Shuttle Orbiter, which was the first reusable vehicle, is discussed. The thermal protection materials and given concepts and some of the shuttle development and manufacturing problems are described. Evolution of a family of grid and flexible ceramic external insulation materials from the initial shuttle concept in the early 1970's to the present time is described. The important properties and their evolution are documented. Application of these materials to vehicles currently being developed and plans for research to meet the space programs future needs are summarized.

  14. Improved Thermal-Switch Disks Protect Batteries

    NASA Technical Reports Server (NTRS)

    Darcy, Eric; Bragg, Bobby

    1990-01-01

    Improved thermal-switch disks help protect electrical batteries against high currents like those due to short circuits or high demands for power in circuits supplied by batteries. Protects batteries against excessive temperatures. Centered by insulating fiberglass washer. Contains conductive polymer that undergoes abrupt increase in electrical resistance when excessive current raises its temperature above specific point. After cooling, polymer reverts to low resistance. Disks reusable.

  15. Thermal Protection Materials for Reentry Applications

    NASA Technical Reports Server (NTRS)

    Johnson, Sylvia M.; Stackpoole, Mairead; Gusman, Mike; Loehman, Ron; Kotula, Paul; Ellerby, Donald; Arnold, James; Wercinski, Paul; Reuthers, James; Kontinos, Dean

    2001-01-01

    Thermal protection materials and systems (IRS) are used to protect spacecraft during reentry into Earth's atmosphere or entry into planetary atmospheres. As such, these materials are subject to severe environments with high heat fluxes and rapid heating. Catalytic effects can increase the temperatures substantially. These materials are also subject to impact damage from micrometeorites or other debris during ascent, orbit, and descent, and thus must be able to withstand damage and to function following damage. Thermal protection materials and coatings used in reusable launch vehicles will be reviewed, including the needs and directions for new materials to enable new missions that require faster turnaround and much greater reusability. The role of ablative materials for use in high heat flux environments, especially for non-reusable applications and upcoming planetary missions, will be discussed. New thermal protection system materials may enable the use of sharp nose caps and leading edges on future reusable space transportation vehicles. Vehicles employing this new technology would have significant increases in maneuverability and out-of-orbit cross range compared to current vehicles, leading to increased mission safety in the event of the need to abort during ascent or from orbit. Ultrahigh temperature ceramics, a family of materials based on HfB2 and ZrB2 with SiC, will be discussed. The development, mechanical and thermal properties, and uses of these materials will be reviewed.

  16. Thermal Materials Protect Priceless, Personal Keepsakes

    NASA Technical Reports Server (NTRS)

    2014-01-01

    NASA astronaut Scott Parazynski led the development of materials and techniques for the inspection and repair of the shuttle’s thermal protection system. Parazynski later met Chris Shiver of Houston-based DreamSaver Enterprises LLC and used concepts from his work at Johnson Space Center to develop an enclosure that can withstand 98 percent of residential fires.

  17. Thermal protection system flight repair kit

    NASA Technical Reports Server (NTRS)

    1979-01-01

    A thermal protection system (TPS) flight repair kit required for use on a flight of the Space Transportation System is defined. A means of making TPS repairs in orbit by the crew via extravehicular activity is discussed. A cure in place ablator, a precured ablator (large area application), and packaging design (containers for mixing and dispensing) for the TPS are investigated.

  18. Sprayable Phase Change Coating Thermal Protection Material

    NASA Technical Reports Server (NTRS)

    Richardson, Rod W.; Hayes, Paul W.; Kaul, Raj

    2005-01-01

    NASA has expressed a need for reusable, environmentally friendly, phase change coating that is capable of withstanding the heat loads that have historically required an ablative thermal insulation. The Space Shuttle Program currently relies on ablative materials for thermal protection. The problem with an ablative insulation is that, by design, the material ablates away, in fulfilling its function of cooling the underlying substrate, thus preventing the insulation from being reused from flight to flight. The present generation of environmentally friendly, sprayable, ablative thermal insulation (MCC-l); currently use on the Space Shuttle SRBs, is very close to being a reusable insulation system. In actual flight conditions, as confirmed by the post-flight inspections of the SRBs, very little of the material ablates. Multi-flight thermal insulation use has not been qualified for the Space Shuttle. The gap that would have to be overcome in order to implement a reusable Phase Change Coating (PCC) is not unmanageable. PCC could be applied robotically with a spray process utilizing phase change material as filler to yield material of even higher strength and reliability as compared to MCC-1. The PCC filled coatings have also demonstrated potential as cryogenic thermal coatings. In experimental thermal tests, a thin application of PCC has provided the same thermal protection as a much thicker and heavier application of a traditional ablative thermal insulation. In addition, tests have shown that the structural integrity of the coating has been maintained and phase change performance after several aero-thermal cycles was not affected. Experimental tests have also shown that, unlike traditional ablative thermal insulations, PCC would not require an environmental seal coat, which has historically been required to prevent moisture absorption by the thermal insulation, prevent environmental degradation, and to improve the optical and aerodynamic properties. In order to reduce

  19. Thermal Protection System with Staggered Joints

    NASA Technical Reports Server (NTRS)

    Simon, Xavier D. (Inventor); Robinson, Michael J. (Inventor); Andrews, Thomas L. (Inventor)

    2014-01-01

    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.

  20. Thermal response properties of protective clothing fabrics

    SciTech Connect

    Baitinger, W.F.

    1995-12-31

    In the industrial workplace, it becomes increasingly incumbent upon employers to require employees to use suitable protective equipment and to wear protective apparel. When workers may be subjected to accidental radiant, flame, or electric arc heat sources, work clothing should be used that does not become involved in burning. It is axiomatic that work clothing should not become a primary fuel source, adding to the level of heat exposure, since clothing is usually in intimate contact with the skin. Further, clothing should provide sufficient insulation to protect the skin from severe burn injury. If the worker receives such protection from clothing, action then may be taken to escape the confronted thermal hazard. Published laboratory test methods are used to measure flame resistance and thermal responses of flame resistant fabrics in protective clothing. The purpose of this article is to review these test methods, to discuss certain limitations in application, and to suggest how flame resistant cotton fabrics may be used to enhance worker safety.

  1. Modeling thermal protection outfits for fire exposures

    NASA Astrophysics Data System (ADS)

    Song, Guowen

    2002-01-01

    A numerical model has been developed that successfully predicts heat transfer through thermally protective clothing materials and garments exposed to intense heat. The model considers the effect of fire exposure to the thermophysical properties of materials as well as the air layers between the clothing material and skin surface. These experiments involved characterizing the flash fire surrounding the manikin by measuring the temperature of the flame above each thermal sensor in the manikin surface. An estimation method is used to calculate the heat transfer coefficient for each thermal sensor in a 4 second exposure to an average heat flux of 2.00cal/cm2sec. A parameter estimation method was used to estimate heat induced change in fabric thermophysical properties. The skin-clothe air gap distribution of different garments was determined using three-dimensional body scanning technology. Multi-layer skin model and a burn prediction method were used to predict second and third degree burns. The integrated generalized model developed was validated using the "Pyroman" Thermal Protective Clothing Analysis System with Kevlar/PBIRTM and NomexRTMIIIA coverall garments with different configuration and exposure time. A parametric study conducted using this numerical model indicated the influencing parameters on garment thermal protective performance in terms of skin burn damage subjected to 4 second flash fire exposure. The importance of these parameters is analyzed and distinguished. These parameters includes fabric thermophysical properties, PyromanRTM chamber flash fire characteristics, garment shrinkage and fit factors, as well as garment initial and test ambient temperature. Different skin models and their influence on burn prediction were also investigated using this model.

  2. Testing of the X-33 umbilical system at KSC

    NASA Technical Reports Server (NTRS)

    1999-01-01

    At the Launch Equipment Test Facility, Mike Solomon (left) and Will Reaves (right), both with Lockheed Martin Technical Operations, move in for a close look at part of the X-33 umbilical system. A team of Kennedy Space Center experts developed the umbilical system, comprising panels, valves and hoses that provide the means to load the X-33 with super-cold propellant. The X-33, under construction at Lockheed Martin Skunk Works in Palmdale, Calif., is a half-scale prototype of the planned operational reusable launch vehicle dubbed VentureStar.

  3. Testing of the X-33 umbilical system at KSC

    NASA Technical Reports Server (NTRS)

    1999-01-01

    At the Launch Equipment Test Facility, , Will Reaves and Mike Solomon (kneeling), both with Lockheed Martin Technical Operations, observe parts of the X-33 umbilical system during testing. A team of Kennedy Space Center experts developed the umbilical system, comprising panels, valves and hoses that provide the means to load the X-33 with super-cold propellant. The X-33, under construction at Lockheed Martin Skunk Works in Palmdale, Calif., is a half-scale prototype of the planned operational reusable launch vehicle dubbed VentureStar.

  4. Testing of the X-33 umbilical system at KSC

    NASA Technical Reports Server (NTRS)

    1999-01-01

    At the Launch Equipment Test Facility, Mike Solomon, with Lockheed Martin Technical Operations, studies a part of the X-33 umbilical system during testing. Pointing to the part is Will Reaves, also with Lockheed Martin Technical Operations. A team of Kennedy Space Center experts developed the umbilical system, comprising panels, valves and hoses that provide the means to load the X-33 with super-cold propellant. The X-33, under construction at Lockheed Martin Skunk Works in Palmdale, Calif., is a half-scale prototype of the planned operational reusable launch vehicle dubbed VentureStar.

  5. Testing of the X-33 umbilical system at KSC

    NASA Technical Reports Server (NTRS)

    1999-01-01

    At the Launch Equipment Test Facility, workers check results from testing the X-33 umbilical system. From left are Greg Melton (left), a NASA engineer; Will Reaves, with Lockheed Martin Technical Operations; and Scott Holcomb, also with Lockheed Martin Technical Operations. A team of Kennedy Space Center experts developed the umbilical system, comprising panels, valves and hoses that provide the means to load the X-33 with super-cold propellant. The X-33, under construction at Lockheed Martin Skunk Works in Palmdale, Calif., is a half-scale prototype of the planned operational reusable launch vehicle dubbed VentureStar.

  6. Testing of the X-33 umbilical system at KSC

    NASA Technical Reports Server (NTRS)

    1999-01-01

    At the Launch Equipment Test Facility, Will Reaves (top of stand), with Lockheed Martin Technical Operations, looks over components of the X-33 umbilical system undergoing testing. A team of Kennedy Space Center experts developed the umbilical system, comprising panels, valves and hoses that provide the means to load the X-33 with super-cold propellant. The X-33, under construction at Lockheed Martin Skunk Works in Palmdale, Calif., is a half-scale prototype of the planned operational reusable launch vehicle dubbed VentureStar.

  7. Testing of the X-33 umbilical system at KSC

    NASA Technical Reports Server (NTRS)

    1999-01-01

    At the Launch Equipment Test Facility, Greg Melton (left), a NASA engineer, and Will Reaves (right), with Lockheed Martin Technical Operations, look at components of the X-33 umbilical system that is undergoing testing. A team of Kennedy Space Center experts developed the umbilical system, comprising panels, valves and hoses that provide the means to load the X-33 with super-cold propellant. The X-33, under construction at Lockheed Martin Skunk Works in Palmdale, Calif., is a half-scale prototype of the planned operational reusable launch vehicle dubbed VentureStar.

  8. Testing of the X-33 umbilical system at KSC

    NASA Technical Reports Server (NTRS)

    1999-01-01

    At the Launch Equipment Test Facility, Mike Ynclan, with Dynacs, and Greg Melton, a NASA engineer, look at measurements during testing of the X-33 umbilical system. A team of Kennedy Space Center experts developed the umbilical system, comprising panels, valves and hoses that provide the means to load the X-33 with super-cold propellant. The X-33, under construction at Lockheed Martin Skunk Works in Palmdale, Calif., is a half-scale prototype of the planned operational reusable launch vehicle dubbed VentureStar.

  9. Thermal Protection Materials: Development, Characterization and Evaluation

    NASA Technical Reports Server (NTRS)

    Johnson, Silvia M.

    2012-01-01

    Thermal protection materials and systems (TPS) are used to protect space vehicles from the heat experienced during entry into an atmosphere. The application for these materials is very specialized as are the materials. They must have specific properties to withstand conditions during specific entries. There is no one-size-fits-all TPS as the conditions experienced by a material are very dependent upon the atmosphere, the entry speed, the size and shape of the vehicle, and the location on the vehicle. However, all TPS must be reliable and efficient to ensure mission safety, that is to protect the vehicle while ensuring that payload is maximized. Types of TPS will be reviewed in relation to types of missions and applications. Both reusable and ablative materials will be discussed. Approaches to characterizing and evaluating these materials will be presented. The role of heritage versus new materials will be described.

  10. 49 CFR 179.18 - Thermal protection systems.

    Code of Federal Regulations, 2013 CFR

    2013-10-01

    ... require thermal protection on a tank car, the tank car must have sufficient thermal resistance so that... resistance of the tank car does not conform to paragraph (a) of this section, the thermal resistance of...

  11. Support to X-33/Resusable Launch Vehicle Technology Program

    NASA Technical Reports Server (NTRS)

    2000-01-01

    The X-33 Guidance, Navigation, and Control (GN&C) Peer Review Team (PRT) was formed to assess the integrated X-33 vehicle GN&C system in order to identify any areas of disproportionate risk for initial flight. The eventual scope of the PRT assessment encompasses the GN&C algorithms, software, avionics, control effectors, applicable models, and testing. The initial (phase 1) focus of the PRT was on the GN&C algorithms and the Flight Control Actuation Subsystem (FCAS). The PRT held meetings during its phase 1 assessment at X-33 assembly facilities in Palmdale, California on May 17-18, 2000 and at Honeywell facilities in Tempe, Arizona on June 7, 2000. The purpose of these meetings was for the PRT members to get background briefings on the X-33 vehicle and for the PRT team to be briefed on the design basis and current status of the X-33 GN&C algorithms as well as the FCAS. The following material is covered in this PRT phase 1 final report. Some significant GN&C-related accomplishments by the X-33 development team are noted. Some topics are identified that were found during phase 1 to require fuller consideration when the PRT reconvenes in the future. Some new recommendations by the PRT to the X-33 program will likely result from a thorough assessment of these subjects. An initial list of recommendations from the PRT to the X-33 program is provided. These recommendations stem from topics that received adequate review by the PRT in phase 1. Significant technical observations by the PRT members as a result of the phase 1 meetings are detailed. (These are covered in an appendix.) There were many X-33 development team members who contributed to the technical information used by the PRT during the phase 1 assessment, who supported presentations to the PRT, and who helped to address the many questions posed by the PRT members at and after the phase 1 meetings. In all instances the interaction between the PRT and the X-33 development team members was cordial and very

  12. Outer skin protection of columbium Thermal Protection System (TPS) panels

    NASA Technical Reports Server (NTRS)

    Culp, J. D.

    1973-01-01

    A coated columbium alloy material system 0.04 centimeter thick was developed which provides for increased reliability to the load bearing character of the system in the event of physical damage to and loss of the exterior protective coating. The increased reliability to the load bearing columbium alloy (FS-85) was achieved by interposing an oxidation resistant columbium alloy (B-1) between the FS-85 alloy and a fused slurry silicide coating. The B-1 alloy was applied as a cladding to the FS-85 and the composite was fused slurry silicide coated. Results of material evaluation testing included cyclic oxidation testing of specimens with intentional coating defects, tensile testing of several material combinations exposed to reentry profile conditions, and emittance testing after cycling of up to 100 simulated reentries. The clad material, which was shown to provide greater reliability than unclad materials, holds significant promise for use in the thermal protection system of hypersonic reentry vehicles.

  13. Thermal Vacuum Facility for Testing Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Daryabeigi, Kamran; Knutson, Jeffrey R.; Sikora, Joseph G.

    2002-01-01

    A thermal vacuum facility for testing launch vehicle thermal protection systems by subjecting them to transient thermal conditions simulating re-entry aerodynamic heating is described. Re-entry heating is simulated by controlling the test specimen surface temperature and the environmental pressure in the chamber. Design requirements for simulating re-entry conditions are briefly described. A description of the thermal vacuum facility, the quartz lamp array and the control system is provided. The facility was evaluated by subjecting an 18 by 36 in. Inconel honeycomb panel to a typical re-entry pressure and surface temperature profile. For most of the test duration, the average difference between the measured and desired pressures was 1.6% of reading with a standard deviation of +/- 7.4%, while the average difference between measured and desired temperatures was 7.6% of reading with a standard deviation of +/- 6.5%. The temperature non-uniformity across the panel was 12% during the initial heating phase (t less than 500 sec.), and less than 2% during the remainder of the test.

  14. Reusable Metallic Thermal Protection Systems Development

    NASA Technical Reports Server (NTRS)

    Blosser, Max L.; Martin, Carl J.; Daryabeigi, Kamran; Poteet, Carl C.

    1998-01-01

    Metallic thermal protection systems (TPS) are being developed to help meet the ambitious goals of future reusable launch vehicles. Recent metallic TPS development efforts at NASA Langley Research Center are described. Foil-gage metallic honeycomb coupons, representative of the outer surface of metallic TPS were subjected to low speed impact, hypervelocity impact, rain erosion, and subsequent arcjet exposure. TPS panels were subjected to thermal vacuum, acoustic, and hot gas flow testing. Results of the coupon and panel tests are presented. Experimental and analytical tools are being developed to characterize and improve internal insulations. Masses of metallic TPS and advanced ceramic tile and blanket TPS concepts are compared for a wide range of parameters.

  15. Commercial application of thermal protection system technology

    NASA Technical Reports Server (NTRS)

    Dyer, Gordon L.

    1991-01-01

    The thermal protection system process technology is examined which is used in the manufacture of the External Tank for the Space Shuttle system and how that technology is applied by private business to create new products, new markets, and new American jobs. The term 'technology transfer' means different things to different people and has become one of the buzz words of the 1980s and 1990s. Herein, technology transfer is defined as a means of transferring technology developed by NASA's prime contractors to public and private sector industries.

  16. Thermal Protection Systems: Past, Present and Future

    NASA Technical Reports Server (NTRS)

    Johnson, Sylvia M.

    2015-01-01

    Thermal protection materials and systems (TPS) have been critical to fulfilling humankinds desire to explore space. Composite and ceramic materials have enabled the early missions to orbit, the moon, the space station, Mars with robots, and sample return. Crewed missions to Mars are being considered, and this places even more demands on TPS materials. This talk will give some history on the materials used for earth and planetary entry and the demands placed upon such materials. TPS needs for future missions, especially to Mars, will be identified and potential solutions discussed.

  17. Thermal Protection System of the Space Shuttle

    NASA Technical Reports Server (NTRS)

    Cleland, John; Iannetti, Francesco

    1989-01-01

    The Thermal Protection System (TPS), introduced by NASA, continues to incorporate many of the advances in materials over the past two decades. A comprehensive, single-volume summary of the TPS, including system design rationales, key design features, and broad descriptions of the subsystems of TPS (E.g., reusable surface insulation, leading edge structural, and penetration subsystems) is provided. Details of all elements of TPS development and application are covered (materials properties, manufacturing, modeling, testing, installation, and inspection). Disclosures and inventions are listed and potential commercial application of TPS-related technology is discussed.

  18. Thermal protection materials: Thermophysical property data

    NASA Technical Reports Server (NTRS)

    Williams, S. D.; Curry, Donald M.

    1992-01-01

    This publication presents a thermophysical property survey on materials that could potentially be used for future spacecraft thermal protection systems (TPS). This includes data that was reported in the 1960's as well as more current information reported through the 1980's. An attempt was made to cite the manufacturers as well as the data source in the bibliography. This volume represents an attempt to provide in a single source a complete set of thermophysical data on a large variety of materials used in spacecraft TPS analysis. The property data is divided into two categories: ablative and reusable. The ablative materials have been compiled into twelve categories that are descriptive of the material composition. An attempt was made to define the Arrhenius equation for each material although this data may not be available for some materials. In a similar manner, char data may not be available for some of the ablative materials. The reusable materials have been divided into three basic categories: thermal protection materials (such as insulators), adhesives, and structural materials.

  19. Advanced Metallic Thermal Protection System Development

    NASA Technical Reports Server (NTRS)

    Blosser, M. L.; Chen, R. R.; Schmidt, I. H.; Dorsey, J. T.; Poteet, C. C.; Bird, R. K.

    2002-01-01

    A new Adaptable, Robust, Metallic, Operable, Reusable (ARMOR) thermal protection system (TPS) concept has been designed, analyzed, and fabricated. In addition to the inherent tailorable robustness of metallic TPS, ARMOR TPS offers improved features based on lessons learned from previous metallic TPS development efforts. A specific location on a single-stage-to-orbit reusable launch vehicle was selected to develop loads and requirements needed to design prototype ARMOR TPS panels. The design loads include ascent and entry heating rate histories, pressures, acoustics, and accelerations. Additional TPS design issues were identified and discussed. An iterative sizing procedure was used to size the ARMOR TPS panels for thermal and structural loads as part of an integrated TPS/cryogenic tank structural wall. The TPS panels were sized to maintain acceptable temperatures on the underlying structure and to operate under the design structural loading. Detailed creep analyses were also performed on critical components of the ARMOR TPS panels. A lightweight, thermally compliant TPS support system (TPSS) was designed to connect the TPS to the cryogenic tank structure. Four 18-inch-square ARMOR TPS panels were fabricated. Details of the fabrication process are presented. Details of the TPSS for connecting the ARMOR TPS panels to the externally stiffened cryogenic tank structure are also described. Test plans for the fabricated hardware are presented.

  20. A Further Look at X-33 Entry Guidance and Beyond

    NASA Technical Reports Server (NTRS)

    Lu, Ping

    1999-01-01

    An entry guidance design developed at Iowa State University (ISU) for the X-33 advanced technology demonstrator is outlined and compared with the X-33 entry guidance algorithms developed at NASA Marshall. Both designs are based on the Space Shuttle entry guidance concept, but significant improvements have been made to enhance the performance and reduce the complexity. The ISU design was incorporated into MAVERIC, a high fidelity vehicle simulation software for the X-33, and evaluated in Monte Carlo simulations against random dispersions in propulsion system, wind and atmospheric properties, aerodynamic coefficients, interaction between propulsion and aerodynamics, and navigation data. The simulations clearly demonstrated the capability and precision of the ISU entry guidance design in successfully guiding the X-33 in some rather difficult flight scenarios. As the entry guidance development for the X-33 is completing, this report also offers some review of the strength and limitations of the current Shuttle-based entry guidance framework, and finally some potential candidates for next generation of more capable and cost-effective entry guidance designs are discussed.

  1. 49 CFR 193.2057 - Thermal radiation protection.

    Code of Federal Regulations, 2014 CFR

    2014-10-01

    ... 49 Transportation 3 2014-10-01 2014-10-01 false Thermal radiation protection. 193.2057 Section 193... GAS FACILITIES: FEDERAL SAFETY STANDARDS Siting Requirements § 193.2057 Thermal radiation protection...) The thermal radiation distances must be calculated using Gas Technology Institute's (GTI) report...

  2. 49 CFR 193.2057 - Thermal radiation protection.

    Code of Federal Regulations, 2012 CFR

    2012-10-01

    ... 49 Transportation 3 2012-10-01 2012-10-01 false Thermal radiation protection. 193.2057 Section 193... GAS FACILITIES: FEDERAL SAFETY STANDARDS Siting Requirements § 193.2057 Thermal radiation protection...) The thermal radiation distances must be calculated using Gas Technology Institute's (GTI) report...

  3. 49 CFR 193.2057 - Thermal radiation protection.

    Code of Federal Regulations, 2011 CFR

    2011-10-01

    ... 49 Transportation 3 2011-10-01 2011-10-01 false Thermal radiation protection. 193.2057 Section 193... GAS FACILITIES: FEDERAL SAFETY STANDARDS Siting Requirements § 193.2057 Thermal radiation protection...) The thermal radiation distances must be calculated using Gas Technology Institute's (GTI) report...

  4. 49 CFR 193.2057 - Thermal radiation protection.

    Code of Federal Regulations, 2013 CFR

    2013-10-01

    ... 49 Transportation 3 2013-10-01 2013-10-01 false Thermal radiation protection. 193.2057 Section 193... GAS FACILITIES: FEDERAL SAFETY STANDARDS Siting Requirements § 193.2057 Thermal radiation protection...) The thermal radiation distances must be calculated using Gas Technology Institute's (GTI) report...

  5. 3D Multifunctional Ablative Thermal Protection System

    NASA Technical Reports Server (NTRS)

    Feldman, Jay; Venkatapathy, Ethiraj; Wilkinson, Curt; Mercer, Ken

    2015-01-01

    NASA is developing the Orion spacecraft to carry astronauts farther into the solar system than ever before, with human exploration of Mars as its ultimate goal. One of the technologies required to enable this advanced, Apollo-shaped capsule is a 3-dimensional quartz fiber composite for the vehicle's compression pad. During its mission, the compression pad serves first as a structural component and later as an ablative heat shield, partially consumed on Earth re-entry. This presentation will summarize the development of a new 3D quartz cyanate ester composite material, 3-Dimensional Multifunctional Ablative Thermal Protection System (3D-MAT), designed to meet the mission requirements for the Orion compression pad. Manufacturing development, aerothermal (arc-jet) testing, structural performance, and the overall status of material development for the 2018 EM-1 flight test will be discussed.

  6. Lightweight Nonmetallic Thermal Protection Materials Technology

    NASA Technical Reports Server (NTRS)

    Valentine, Peter G.; Lawrence, Timothy W.; Gubert, Michael K.; Milos, Frank S.; Levine, Stanley R.; Ohlhorst, Craig W.; Koenig, John R.

    2005-01-01

    To fulfill President George W. Bush's "Vision for Space Exploration" (2004) - successful human and robotic missions to and from other solar system bodies in order to explore their atmospheres and surfaces - the National Aeronautics and Space Administration (NASA) must reduce the trip time, cost, and vehicle weight so that the payload and scientific experiments' capabilities can be maximized. The new project described in this paper will generate thermal protection system (TPS) product that will enable greater fidelity in mission/vehicle design trade studies, support risk reduction for material selections, assist in the optimization of vehicle weights, and provide materials and processes templates for use in the development of human-rated TPS qualification and certification plans.

  7. Thermal Protection Test Bed Pathfinder Development Project

    NASA Technical Reports Server (NTRS)

    Snapp, Cooper

    2015-01-01

    In order to increase thermal protection capabilities for future reentry vehicles, a method to obtain relevant test data is required. Although arc jet testing can be used to obtain some data on materials, the best method to obtain these data is to actually expose them to an atmospheric reentry. The overprediction of the Orion EFT-1 flight data is an example of how the ground test to flight traceability is not fully understood. The RED-Data small reentry capsule developed by Terminal Velocity Aerospace is critical to understanding this traceability. In order to begin to utilize this technology, ES3 needs to be ready to build and integrate heat shields onto the RED-Data vehicle. Using a heritage Shuttle tile material for the heat shield will both allow valuable insight into the environment that the RED-Data vehicle can provide and give ES3 the knowledge and capability to build and integrate future heat shields for this vehicle.

  8. Thermal protection using very high temperature ceramics

    NASA Technical Reports Server (NTRS)

    Adamczyk, George R.

    1992-01-01

    The purpose of the paper is to expose the reader to a technology that may solve some of the toughest materials problems facing thermal protection for use in aerospace. Supermaterials has created a system capable of producing unique material properties. Over 10 years and many man-hours have been invested in the development of this technology. Applications range from the food industry to the rigors of outer space. The flexibility of the system allows for customization not found in many other processes and at a reasonable cost. The ranges of materials and alloys that can be created are endless. Many cases with unique characteristics have been identified and we can expect even more with further development.

  9. Design of Transpiration Cooled Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Callens, E. Eugene, Jr.; Vinet, Robert F.

    1999-01-01

    This study explored three approaches for the utilization of transpiration cooling in thermal protection systems. One model uses an impermeable wall with boiling water heat transfer at the backface (Model I). A second model uses a permeable wall with a boiling water backface and additional heat transfer to the water vapor as it flows in channels toward the exposed surface (Model II). The third model also uses a permeable wall, but maintains a boiling condition at the exposed surface of the material (Model III). The governing equations for the models were developed in non-dimensional form and a comprehensive parametric investigation of the effects of the independent variables on the important dependent variables was performed. In addition, detailed analyses were performed for selected materials to evaluate the practical limitations of the results of the parametric study.

  10. High Temperature Aerogels for Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Hurwitz, Frances I.; Mbah, Godfrey C.

    2008-01-01

    High temperature aerogels in the Al2O3-SiO2 system are being investigated as possible constituents for lightweight integrated thermal protection system (TPS) designs for use in supersonic and hypersonic applications. Gels are synthesized from ethoxysilanes and AlCl3.6H2O, using an epoxide catalyst. The influence of Al:Si ratio, solvent, water to metal and water to alcohol ratios on aerogel composition, morphology, surface area, and pore size distribution were examined, and phase transformation on heat treatment characterized. Aerogels have been fabricated which maintain porous, fractal structures after brief exposures to 1000 C. Incorporation of nanofibers, infiltration of aerogels into SiC foams, use of polymers for crosslinking the aerogels, or combinations of these, offer potential for toughening and integration of TPS with composite structure. Woven fabric composites having Al2O3-SiO2 aerogels as a matrix also have been fabricated. Continuing work is focused on reduction in shrinkage and optimization of thermal and physical properties.

  11. Shearographic and thermographic nondestructive evaluation of the space shuttle structure and thermal protection systems (TPS)

    NASA Astrophysics Data System (ADS)

    Davis, Christopher K.

    1996-11-01

    Shearography and thermography have shown promising results on orbiter structure and external tank (ET) and solid rocket booster (SRB) thermal protection systems (TPS). The orbiter uses a variety of composite structure, the two most prevalent materials being aluminum and graphite-epoxy honeycomb. Both techniques have detected delaminations as small at 0.25 inches diameter in the orbiter payload bay doors graphite-epoxy honeycomb structure. Other applications include the robotic manipulator system (RMS) and the rudder speed brake structure. The ET uses spray-on foam insulation (SOFI) as the TPS and the SRB forward section uses marshall sprayable ablative as the TPS. Debonding SOFI damage to the orbiter 'belly' tile and exposes the ET to thermal loading. Voids in SOFI test panels as small as 0.375 inch were detected in 1.75 inch thick foam using a pressure reduction of not more than 10 inches of water or 0.4 pounds per square inch. Preliminary results of the X33 metallic TPS are presented. Ultrasonic testing approved for orbiter bond integrity testing, is time consuming and problematic. No current non-destructive inspection technique is approved for inspection of ET/SRB TPS or the orbiter RMS honeycomb at Kennedy Space Center. Only visual inspections are routinely performed on orbiter structure. The various successes of these two techniques make them good candidates for the aforementioned applications.

  12. Development of the X-33 Aerodynamic Uncertainty Model

    NASA Technical Reports Server (NTRS)

    Cobleigh, Brent R.

    1998-01-01

    An aerodynamic uncertainty model for the X-33 single-stage-to-orbit demonstrator aircraft has been developed at NASA Dryden Flight Research Center. The model is based on comparisons of historical flight test estimates to preflight wind-tunnel and analysis code predictions of vehicle aerodynamics documented during six lifting-body aircraft and the Space Shuttle Orbiter flight programs. The lifting-body and Orbiter data were used to define an appropriate uncertainty magnitude in the subsonic and supersonic flight regions, and the Orbiter data were used to extend the database to hypersonic Mach numbers. The uncertainty data consist of increments or percentage variations in the important aerodynamic coefficients and derivatives as a function of Mach number along a nominal trajectory. The uncertainty models will be used to perform linear analysis of the X-33 flight control system and Monte Carlo mission simulation studies. Because the X-33 aerodynamic uncertainty model was developed exclusively using historical data rather than X-33 specific characteristics, the model may be useful for other lifting-body studies.

  13. X-33 Reusable Launch Vehicle Demonstrator, Spaceport and Range

    NASA Technical Reports Server (NTRS)

    Letchworth, Gary F.

    2011-01-01

    The X-33 was a suborbital reusable spaceplane demonstrator, in development from 1996 to early 2001. The intent of the demonstrator was to lower the risk of building and operating a full-scale reusable vehicle fleet. Reusable spaceplanes offered the potential to lower the cost of access to space by an order of magnitude, compared with conventional expendable launch vehicles. Although a cryogenic tank failure during testing ultimately led to the end of the effort, the X-33 team celebrated many successes during the development. This paper summarizes some of the accomplishments and milestones of this X-vehicle program, from the perspective of an engineer who was a member of the team throughout the development. X-33 Program accomplishments include rapid, flight hardware design, subsystem testing and fabrication, aerospike engine development and testing, Flight Operations Center and Operations Control Center ground systems design and construction, rapid Environmental Impact Statement NEPA process approval, Range development and flight plan approval for test flights, and full-scale system concept design and refinement. Lessons from the X-33 Program may have potential application to new RLV and other aerospace systems being developed a decade later.

  14. Fatigue properties of shuttle thermal protection system

    NASA Technical Reports Server (NTRS)

    Sawyer, J. W.; Cooper, P. A.

    1980-01-01

    Static and cyclic load tests were conducted to determine the static and fatigue strength of the RIS tile/SIP thermal protection system used on the orbiter of the space shuttle. The material systems investigated include the densified and undensified LI-900 tile system on the .40 cm thick SIP and the densified and undensified LI-2200 tile system on the .23 cm (.090 inch) thick SIP. The tests were conducted at room temperature with a fully reversed uniform cyclic loading at 1 Hertz. Cyclic loading causes a relatively large reduction in the stress level that each of the SIP/tile systems can withstand for a small number of cycles. For example, the average static strength of the .40 cm thick SIP/LI-900 tile system is reduced from 86 kPa to 62 kPa for a thousand cycles. Although the .23 cm thick SIP/LI-2200 tile system has a higher static strength, similar reductions in the fatigue strength are noted. Densifying the faying surface of the RSI tile changes the failure mode from the SIP/tile interface to the parent RSI or the SIP and thus greatly increases the static strength of the system. Fatigue failure for the densified tile system, however, occurs due to complete separation or excessive elongation of the SIP and the fatigue strength is only slightly greater than that for the undensified tile system.

  15. Thermal protection systems manned spacecraft flight experience

    NASA Astrophysics Data System (ADS)

    Curry, Donald M.

    1992-10-01

    Since the first U.S. manned entry, Mercury (May 5, 1961), seventy-five manned entries have been made resulting in significant progress in the understanding and development of Thermal Protection Systems (TPS) for manned rated spacecraft. The TPS materials and systems installed on these spacecraft are compared. The first three vehicles (Mercury, Gemini, Apollo) used ablative (single-use) systems while the Space Shuttle Orbiter TPS is a multimission system. A TPS figure of merit, unit weight lb/sq ft, illustrates the advances in TPS material performance from Mercury (10.2 lb/sq ft) to the Space Shuttle (1.7 lb/sq ft). Significant advances have been made in the design, fabrication, and certification of TPS on manned entry vehicles (Mercury through Shuttle Orbiter). Shuttle experience has identified some key design and operational issues. State-of-the-art ceramic insulation materials developed in the 1970's for the Space Shuttle Orbiter have been used in the initial designs of aerobrakes. This TPS material experience has identified the need to develop a technology base from which a new class of higher temperature materials will emerge for advanced space transportation vehicles.

  16. Full Envelope Reconfigurable Control Design for the X-33 Vehicle

    NASA Technical Reports Server (NTRS)

    Cotting, M. Christopher; Burken, John J.; Lee, Seung-Hee (Technical Monitor)

    2001-01-01

    In the event of a control surface failure, the purpose of a reconfigurable control system is to redistribute the control effort among the remaining working surfaces such that satisfactory stability and performance are retained. An Off-line Nonlinear General Constrained Optimization (ONCO) approach was used for the reconfigurable X-33 control design method. Three example failures are shown using a high fidelity 6 DOF simulation (case I ascent with a left body flap jammed at 25 deg.; case 2 entry with a right inboard elevon jam at 25 deg.; and case 3, landing (TAEM) with a left rudder jam at -30 deg.) Failure comparisons between responses with the nominal controller and reconfigurable controllers show the benefits of reconfiguration. Single jam aerosurface failures were considered, and failure detection and identification is considered accomplished in the actuator controller. The X-33 flight control system will incorporate reconfigurable flight control in the baseline system.

  17. Mechanical properties of thermal protection system materials.

    SciTech Connect

    Hardy, Robert Douglas; Bronowski, David R.; Lee, Moo Yul; Hofer, John H.

    2005-06-01

    An experimental study was conducted to measure the mechanical properties of the Thermal Protection System (TPS) materials used for the Space Shuttle. Three types of TPS materials (LI-900, LI-2200, and FRCI-12) were tested in 'in-plane' and 'out-of-plane' orientations. Four types of quasi-static mechanical tests (uniaxial tension, uniaxial compression, uniaxial strain, and shear) were performed under low (10{sup -4} to 10{sup -3}/s) and intermediate (1 to 10/s) strain rate conditions. In addition, split Hopkinson pressure bar tests were conducted to obtain the strength of the materials under a relatively higher strain rate ({approx}10{sup 2} to 10{sup 3}/s) condition. In general, TPS materials have higher strength and higher Young's modulus when tested in 'in-plane' than in 'through-the-thickness' orientation under compressive (unconfined and confined) and tensile stress conditions. In both stress conditions, the strength of the material increases as the strain rate increases. The rate of increase in LI-900 is relatively small compared to those for the other two TPS materials tested in this study. But, the Young's modulus appears to be insensitive to the different strain rates applied. The FRCI-12 material, designed to replace the heavier LI-2200, showed higher strengths under tensile and shear stress conditions. But, under a compressive stress condition, LI-2200 showed higher strength than FRCI-12. As far as the modulus is concerned, LI-2200 has higher Young's modulus both in compression and in tension. The shear modulus of FRCI-12 and LI-2200 fell in the same range.

  18. X-33 Rev-F Turbulent Aeroheating Results From Test 6817 in NASA Langley 20-Inch Mach 6 Air Tunnel and Comparisons With Computations

    NASA Technical Reports Server (NTRS)

    Hollis, Brian R.; Horvath, Thomas J.; Berry, Scott A.

    2003-01-01

    Measurements and predictions of the X-33 turbulent aeroheating environment have been performed at Mach 6, perfect-gas air conditions. The purpose of this investigation was to compare measured turbulent aeroheating levels on smooth models, models with discrete trips, and models with arrays of bowed panels (which simulate bowed thermal protections system tiles) with each other and with predictions from two Navier-Stokes codes, LAURA and GASP. The wind tunnel testing was conducted at free stream Reynolds numbers based on length of 1.8 x 10(exp 6) to 6.1 x 10(exp 6) on 0.0132 scale X-33 models at a = 40-deg. Turbulent flow was produced by the discrete trips and by the bowed panels at ill but the lowest Reynolds number, but turbulent flow on the smooth model was produced only at the highest Reynolds number. Turbulent aeroheating levels on each of the three model types were measured using global phosphor thermography and were found to agree to within .he estimated uncertainty (plus or minus 15%) of the experiment. Computations were performed at the wind tunnel free stream conditions using both codes. Turbulent aeroheating levels predicted using the LAURA code were generally 5%-10% lower than those from GASP, although both sets of predictions fell within the experimental accuracy of the wind tunnel data.

  19. Deterministic Reconfigurable Control Design for the X-33 Vehicle

    NASA Technical Reports Server (NTRS)

    Wagner, Elaine A.; Burken, John J.; Hanson, Curtis E.; Wohletz, Jerry M.

    1998-01-01

    In the event of a control surface failure, the purpose of a reconfigurable control system is to redistribute the control effort among the remaining working surfaces such that satisfactory stability and performance are retained. Four reconfigurable control design methods were investigated for the X-33 vehicle: Redistributed Pseudo-Inverse, General Constrained Optimization, Automated Failure Dependent Gain Schedule, and an Off-line Nonlinear General Constrained Optimization. The Off-line Nonlinear General Constrained Optimization approach was chosen for implementation on the X-33. Two example failures are shown, a right outboard elevon jam at 25 deg. at a Mach 3 entry condition, and a left rudder jam at 30 degrees. Note however, that reconfigurable control laws have been designed for the entire flight envelope. Comparisons between responses with the nominal controller and reconfigurable controllers show the benefits of reconfiguration. Single jam aerosurface failures were considered, and failure detection and identification is considered accomplished in the actuator controller. The X-33 flight control system will incorporate reconfigurable flight control in the baseline system.

  20. Reconfigurable Control Design for the Full X-33 Flight Envelope

    NASA Technical Reports Server (NTRS)

    Cotting, M. Christopher; Burken, John J.

    2001-01-01

    A reconfigurable control law for the full X-33 flight envelope has been designed to accommodate a failed control surface and redistribute the control effort among the remaining working surfaces to retain satisfactory stability and performance. An offline nonlinear constrained optimization approach has been used for the X-33 reconfigurable control design method. Using a nonlinear, six-degree-of-freedom simulation, three example failures are evaluated: ascent with a left body flap jammed at maximum deflection; entry with a right inboard elevon jammed at maximum deflection; and landing with a left rudder jammed at maximum deflection. Failure detection and identification are accomplished in the actuator controller. Failure response comparisons between the nominal control mixer and the reconfigurable control subsystem (mixer) show the benefits of reconfiguration. Single aerosurface jamming failures are considered. The cases evaluated are representative of the study conducted to prove the adequate and safe performance of the reconfigurable control mixer throughout the full flight envelope. The X-33 flight control system incorporates reconfigurable flight control in the existing baseline system.

  1. X-33 Environmental Impact Statement: A Fast Track Approach

    NASA Technical Reports Server (NTRS)

    McCaleb, Rebecca C.; Holland, Donna L.

    1998-01-01

    NASA is required by the National Environmental Policy Act (NEPA) to prepare an appropriate level environmental analysis for its major projects. Development of the X-33 Technology Demonstrator and its associated flight test program required an environmental impact statement (EIS) under the NEPA. The EIS process is consists of four parts: the "Notice of Intent" to prepare an EIS and scoping; the draft EIS which is distributed for review and comment; the final ETS; and the "Record of Decision." Completion of this process normally takes from 2 - 3 years, depending on the complexity of the proposed action. Many of the agency's newest fast track, technology demonstration programs require NEPA documentation, but cannot sustain the lengthy time requirement between program concept development to implementation. Marshall Space Flight Center, in cooperation with Kennedy Space Center, accomplished the NEPA process for the X-33 Program in 13 months from Notice of Intent to Record of Decision. In addition, the environmental team implemented an extensive public involvement process, conducting a total of 23 public meetings for scoping and draft EIS comment along with numerous informal meetings with public officials, civic organizations, and Native American Indians. This paper will discuss the fast track approach used to successfully accomplish the NEPA process for X-33 on time.

  2. Assessment of Thermal Protection Afforded by Hot Water Diving Suits

    DTIC Science & Technology

    1980-07-03

    Assessment of Thermal Protect! " Afforded by Hot Water Diving Suits A AA L. A. Kuehn Diver thermal comfort in cold water is presently only...with proper control oj inlet suit water flow% and temperature, as well as heating of inspired gas, this suit technology suffices for thermal comfort for...technology provides in part to the convective heat loss that it prpsents, sustained long-term thermal comfort in cold water, Webb (W) has defined a

  3. X-33 Metal Model Testing In Low Turbulence Pressure Tunnel

    NASA Technical Reports Server (NTRS)

    1997-01-01

    The countrys next generation of space transportation, a reusable launch vehicle (RLV), continues to undergo wind tunnel testing at NASA Langley Research Center, Hampton, Va. All four photos are a metal model of the X-33 reusable launch vehicle (about 15 inches long by 15 inches wide) being tested for Lockheed Martin Skunk Works in the Low Turbulence Pressure Tunnel (LTPT) at NASA Langley Research Center. Tests are being conducted by members of the Aerothermodynamics Branch. According to Kelly Murphy of Langleys Aerothermodynamics Branch, the aluminum and stainless steel model of the X-33 underwent aerodynamic testing in the tunnel. *The subsonic tests were conducted at the speed of Mach 25,* she said. *Force and moment testing and measurement in this tunnel lasted about one week.* Future testing of the metal model is scheduled for Langleys 16-Foot Transonic Tunnel, from the end of March to mid-April 1997, and the Unitary Wind Tunnel, from mid-April to the beginning of May. Other tunnel testing for X-33 models are scheduled from the present through June in the hypersonic tunnels, and the 14- by 22-Foot Tunnel from about mid-June to mid-July. Since 1991 Marshall Space Flight Center in Huntsville, Ala. has been the lead center for coordinating the Agencys X-33 Reusable Launch Vehicle (RLV) Program, an industry-led effort, which NASA Administrator Daniel S. Goldin has declared the agency's highest priority new program. The RLV Technology Program is a partnership among NASA, the United States Air Force and private industry to develop world leadership in low-cost space transportation. The goal of the program is to develop technologies and new operational concepts that can radically reduce the cost of access to space. The RLV program also hopes to speed the commercialization of space and improve U.S. economic competitiveness by making access to space as routine and reliable as today's airline industry, while reducing costs and enhancing safety and reliability. The RLV

  4. X-33 Metal Model Testing In Low Turbulence Pressure Tunnel

    NASA Technical Reports Server (NTRS)

    1997-01-01

    The countrys next generation of space transportation, a reusable launch vehicle (RLV), continues to undergo wind tunnel testing at NASA Langley Research Center, Hampton, Va. All four photos are a metal model of the X-33 reusable launch vehicle (about 15 inches long by 15 inches wide) being tested for Lockheed Martin Skunk Works in the Low Turbulence Pressure Tunnel (LTPT) at NASA Langley Research Center. Tests are being conducted by members of the Aerothermodynamics Branch. According to Kelly Murphy of Langleys Aerothermodynamics Branch, the aluminum and stainless steel model of the X-33 underwent aerodynamic testing in the tunnel. *The subsonic tests were conducted at the speed of Mach .25,* she said. *Force and moment testing and measurement in this tunnel lasted about one week.* Future testing of the metal model is scheduled for Langleys 16-Foot Transonic Tunnel, from the end of March to mid-April 1997, and the Unitary Wind Tunnel, from mid-April to the beginning of May. Other tunnel testing for X-33 models are scheduled from the present through June in the hypersonic tunnels, and the 14- by 22-Foot Tunnel from about mid-June to mid-July. Since 1991 Marshall Space Flight Center in Huntsville, Ala. has been the lead center for coordinating the Agencys X-33 Reusable Launch Vehicle (RLV) Program, an industry-led effort, which NASA Administrator Daniel S. Goldin has declared the agency's highest priority new program. The RLV Technology Program is a partnership among NASA, the United States Air Force and private industry to develop world leadership in low-cost space transportation. The goal of the program is to develop technologies and new operational concepts that can radically reduce the cost of access to space. The RLV program also hopes to speed the commercialization of space and improve U.S. economic competitiveness by making access to space as routine and reliable as today's airline industry, while reducing costs and enhancing safety and reliability. The RLV

  5. Overview of the X-33 Extended Flight Test Range

    NASA Technical Reports Server (NTRS)

    Mackall, D.; Sakahara, R.; Kremer, S.

    1998-01-01

    On July 1, 1996, the National Aeronautics and Space Administration signed a Cooperative Agreement No. NCC8-115 with Lockheed Martin Skunk Works to develop and flight test the X-33, a scaled version of a reusable launch vehicle. The development of an Extended Test Range, with range instrumentation providing continuous vehicle communications from Edwards Air Force Base Ca. to landing at Malmstrom Air Force Base Montana, was required to flight test the mach 15 vehicle over 950 nautical miles. The cooperative agreement approach makes Lockheed Martin Skunk Works responsible for the X-33 program. When additional Government help was required, Lockheed "subcontracted" to NASA Field Centers for certain work. It was through this mechanism that Dryden Flight Research Center became responsible for the Extended Test Range. The Extended Test Range Requirements come from two main sources: 1) Range Safety and 2) Lockheed Martin Skunk Works. The range safety requirements were the most challenging to define and meet. The X-33 represents a vehicle that launches like a rocket, reenters the atmosphere and lands autonomously like an aircraft. Historically, rockets have been launched over the oceans to allow failed rockets to be destroyed using explosive devices. Such approaches had to be reconsidered for the X-33 flying over land. Numerous range requirements come from Lockheed Martin Skunk Works for interface definitions with the vehicle communication subsystems and the primary ground operations center, defined the Operations Control Center. Another area of considerable interest was the reentry plasma shield that causes "blackout" of the radio frequency signals, such as the range safety commands. Significant work was spent to analyze and model the blackout problem using a cooperative team of experts from across the country. The paper describes the Extended Test Range a, an unique Government/industry team of personnel and range assets was established to resolve design issues and

  6. Thermal Protection System Development, Testing and Qualification

    NASA Astrophysics Data System (ADS)

    Venkatapathy, Ethiraj; Arnold, James; Laub, B.; Hartman, G. J.

    The science community currently has interest in planetary entry probe missions to improve our understanding of the atmospheres of Saturn and Venus [1,2]. As in the case of the Galileo entry probe, such data are critical to the understanding of not only the individual planets but also to further knowledge regarding the formation of the solar system. It is believed that Saturn probes to depths corresponding to 10 bars will be sufficient [1] to provide the desired scientific data. The heating rates for the "shallow" Saturn probes and Venus are in the range of 2 - 5KW/cm2 . It is clear that new, mid-density Thermal Protection System (TPS) materials for such probes can be mission-enabling for mass efficiency [3] and also make the use of smaller vehicles possible from advancements in scientific instrumentation [4]. Past consideration of new Jovian multiprobe missions has been considered problematic without the Giant Planet Arcjet Facility that was used to qualify Carbon Phenolic for the Galileo Probe. This paper describes emerging TPS technology and the proposed use of an affordable, small 5 MW arc jet that can be used for TPS development in test gases appropriate for the aforementioned, new planetary probe applications. Emerging TPS technologies of interest include a mid-density, chopped molded carbon phenolic (CMCP) material around 0.8g/cc and a densified variant of phenolic impregnated carbon ablator (PICA) around 0.5g/cc. The small 5 MW arc jet facility, called the Development Arcjet Facility (DAF) and the methodology of testing TPS, both based on previous work, are discussed. Finally, the applications to Earth entry appropriate to speeds greater than lunar return (11km/s) are discussed as will facility-to-facility validation using air as a test gas. The use of other facilities for development, qualification and certification of TPS for Saturn and Venus is also discussed. [1] Atreya, S. K., et. al. Formation of Giant Planets and Their Atmospheres: Entry Probes for

  7. Deployable Aeroshell Flexible Thermal Protection System Testing

    NASA Technical Reports Server (NTRS)

    Hughes, Stephen J.; Ware, Joanne S.; DelCorso, Joseph A.; Lugo, Rafael A.

    2009-01-01

    Deployable aeroshells offer the promise of achieving larger aeroshell surface areas for entry vehicles than otherwise attainable without deployment. With the larger surface area comes the ability to decelerate high-mass entry vehicles at relatively low ballistic coefficients. However, for an aeroshell to perform even at the low ballistic coefficients attainable with deployable aeroshells, a flexible thermal protection system (TPS) is required that is capable of surviving reasonably high heat flux and durable enough to survive the rigors of construction handling, high density packing, deployment, aerodynamic loading and aerothermal heating. The Program for the Advancement of Inflatable Decelerators for Atmospheric Entry (PAIDAE) is tasked with developing the technologies required to increase the technology readiness level (TRL) of inflatable deployable aeroshells, and one of several of the technologies PAIDAE is developing for use on inflatable aeroshells is flexible TPS. Several flexible TPS layups were designed, based on commercially available materials, and tested in NASA Langley Research Center's 8 Foot High Temperature Tunnel (8ft HTT). The TPS layups were designed for, and tested at three different conditions that are representative of conditions seen in entry simulation analyses of inflatable aeroshell concepts. Two conditions were produced in a single run with a sting-mounted dual wedge test fixture. The dual wedge test fixture had one row of sample mounting locations (forward) at about half the running length of the top surface of the wedge. At about two thirds of the running length of the wedge, a second test surface drafted up at five degrees relative to the first test surface established the remaining running length of the wedge test fixture. A second row of sample mounting locations (aft) was positioned in the middle of the running length of the second test surface. Once the desired flow conditions were established in the test section the dual wedge

  8. Displacements of Metallic Thermal Protection System Panels During Reentry

    NASA Technical Reports Server (NTRS)

    Daryabeigi, Kamran; Blosser, Max L.; Wurster, Kathryn E.

    2006-01-01

    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.

  9. Woven Thermal Protection System (Woven TPS) for Extreme Entry Environments

    NASA Video Gallery

    The Woven Thermal Protection System (WTPS) project explores an innovative way to design, develop and manufacture a family of ablative TPS materials using weaving technology and testing them in the ...

  10. Lightweight Nonmetallic Thermal Protection Materials Technology (LNTPMT) Project

    NASA Technical Reports Server (NTRS)

    Flynn, Kevin; Gubert, Michael

    2005-01-01

    Contents include the following: Exploration systems research and technology program structure. Project objective. Overview of project. Candidate thermal protection system (PS) materials. Definition of reference missions and space environments. Technical performance metrics (TPMs).Testing (types of tests). Conclusion.

  11. NASA Ames Develops Woven Thermal Protection System (TPS)

    NASA Video Gallery

    The Woven Thermal Protection System (WTPS) project explores an innovative way to design, develop and manufacture a family of ablative TPS materials using weaving technology and testing them in the ...

  12. Composite flexible insulation for thermal protection of space vehicles

    NASA Technical Reports Server (NTRS)

    Kourtides, Demetrius A.; Tran, Huy K.; Chiu, S. A.

    1992-01-01

    A composite flexible blanket insulation (CFBI) system considered for use as a thermal protection system for space vehicles is described. This flexible composite insulation system consists of an outer layer of silicon carbide fabric, followed by alumina mat insulation, and alternating layers of aluminized polyimide film and aluminoborosilicate scrim fabric. A potential application of this composite insulation would be as a thermal protection system for the aerobrake of the aeroassist space transfer vehicle (ASTV). It would also apply to other space vehicles subject to high convective and radiative heating during atmospheric entry. The thermal performance of this composite insulation as exposed to a simulated atmospheric entry environment in a plasma arc test facility is described. Other thermophysical properties which affect the thermal response of this composite insulation is included. It shows that this composite insulation is effective as a thermal protection system at total heating rates up to 30.6 W/sq cm.

  13. Composite flexible insulation for thermal protection of space vehicles

    NASA Technical Reports Server (NTRS)

    Kourtides, Demetrius A.; Tran, Huy K.; Chiu, S. Amanda

    1991-01-01

    A composite flexible blanket insulation (CFBI) system considered for use as a thermal protection system for space vehicles is described. This flexible composite insulation system consists of an outer layer of silicon carbide fabric, followed by alumina mat insulation, and alternating layers of aluminized polyimide film and aluminoborosilicate scrim fabric. A potential application of this composite insulation would be as a thermal protection system for the aerobrake of the aeroassist space transfer vehicle (ASTV). It would also apply to other space vehicles subject to high convective and radiative heating during atmospheric entry. The thermal performance of this composite insulation as exposed to a simulated atmospheric entry environment in a plasma arc test facility is described. Other thermophysical properties which affect the thermal response of this composite insulation is included. It shows that this composite insulation is effective as a thermal protection system at total heating rates up to 30.6 W/sq cm.

  14. The X-33 Program, Proving Single Stage to Orbit

    NASA Technical Reports Server (NTRS)

    Austin, Robert E.; Rising, Jerry J.

    1998-01-01

    The X-33, NASA's flagship for reusable space plane technology demonstration, is on course to permit a crucial decision for the nation by the end of this decade. Lockheed Martin Skunk Works, NASA's partner in this effort, has led a dedicated and talented industry and government team that have met and solved numerous challenges within the first 26 months. This program began by accepting the mandate that included two unprecedented and highly challenging goals: 1) demonstrate single stage to orbit technologies in flight and ground demonstration in less than 42 months and 2) demonstrate a new government and industry management relationship working together with industry in the lead.

  15. Space Shuttle Orbiter thermal protection system design and flight experience

    NASA Technical Reports Server (NTRS)

    Curry, Donald M.

    1993-01-01

    The Space Shuttle Orbiter Thermal Protection System materials, design approaches associated with each material, and the operational performance experienced during fifty-five successful flights are described. The flights to date indicate that the thermal and structural design requirements were met and that the overall performance was outstanding.

  16. Space Shuttle Orbiter thermal protection system design and flight experience

    NASA Astrophysics Data System (ADS)

    Curry, Donald M.

    1993-07-01

    The Space Shuttle Orbiter Thermal Protection System materials, design approaches associated with each material, and the operational performance experienced during fifty-five successful flights are described. The flights to date indicate that the thermal and structural design requirements were met and that the overall performance was outstanding.

  17. Results of Two-Stage Light-Gas Gun Development Efforts and Hypervelocity Impact Tests of Advanced Thermal Protection Materials

    NASA Technical Reports Server (NTRS)

    Cornelison, C. J.; Watts, Eric T.

    1998-01-01

    Gun development efforts to increase the launching capabilities of the NASA Ames 0.5-inch two-stage light-gas gun have been investigated. A gun performance simulation code was used to guide initial parametric variations and hardware modifications, in order to increase the projectile impact velocity capability to 8 km/s, while maintaining acceptable levels of gun barrel erosion and gun component stresses. Concurrent with this facility development effort, a hypervelocity impact testing series in support of the X-33/RLV program was performed in collaboration with Rockwell International. Specifically, advanced thermal protection system materials were impacted with aluminum spheres to simulate impacts with on-orbit space debris. Materials tested included AETB-8, AETB-12, AETB-20, and SIRCA-25 tiles, tailorable advanced blanket insulation (TABI), and high temperature AFRSI (HTA). The ballistic limit for several Thermal Protection System (TPS) configurations was investigated to determine particle sizes which cause threshold TPS/structure penetration. Crater depth in tiles was measured as a function of impact particle size. The relationship between coating type and crater morphology was also explored. Data obtained during this test series was used to perform a preliminary analysis of the risks to a typical orbital vehicle from the meteoroid and space debris environment.

  18. Flexible thermal protection materials for entry systems

    NASA Technical Reports Server (NTRS)

    Kourtides, Demetrius A.

    1993-01-01

    Current programs addressed in aeroassist flight experiment are: (1) evaluation of thermal performance of advanced rigid and flexible insulations and reflective coating; (2) investigation of lighter than baseline materials; (3) investigation of rigid insulations which perform well; (4) study of flexible insulations which require ceramic coating; and (5) study of reflective coating effective at greater than 15 percent. In National Aerospace Plane (NASP), the programs addressed are: (1) high and low temperature insulations; and (2) attachment/standoff methodology critical which affects thermal performance.

  19. X-33/RLV System Health Management/ Vehicle Health Management

    NASA Technical Reports Server (NTRS)

    Garbos, Raymond J.; Mouyos, William

    1998-01-01

    To reduce operations cost, the RLV must include the following elements: highly reliable, robust subsystems designed for simple repair access with a simplified servicing infrastructure and incorporating expedited decision making about faults and anomalies. A key component for the Single Stage to Orbit (SSTO) RLV System used to meet these objectives is System Health Management (SHM). SHM deals with the vehicle component- Vehicle Health Management (VHM), the ground processing associated with the fleet (GVHM) and the Ground Infrastructure Health Management (GIHM). The objective is to provide an automated collection and paperless health decision, maintenance and logistics system. Many critical technologies are necessary to make the SHM (and more specifically VHM) practical, reliable and cost effective. Sanders is leading the design, development and integration of the SHM system for RLV and X-33 SHM (a sub-scale, sub-orbit Advanced Technology Demonstrator). This paper will present the X-33 SHM design which forms the baseline for RLV SHM. This paper will also discuss other applications of these technologies.

  20. An empirical analysis of thermal protective performance of fabrics used in protective clothing.

    PubMed

    Mandal, Sumit; Song, Guowen

    2014-10-01

    Fabric-based protective clothing is widely used for occupational safety of firefighters/industrial workers. The aim of this paper is to study thermal protective performance provided by fabric systems and to propose an effective model for predicting the thermal protective performance under various thermal exposures. Different fabric systems that are commonly used to manufacture thermal protective clothing were selected. Laboratory simulations of the various thermal exposures were created to evaluate the protective performance of the selected fabric systems in terms of time required to generate second-degree burns. Through the characterization of selected fabric systems in a particular thermal exposure, various factors affecting the performances were statistically analyzed. The key factors for a particular thermal exposure were recognized based on the t-test analysis. Using these key factors, the performance predictive multiple linear regression and artificial neural network (ANN) models were developed and compared. The identified best-fit ANN models provide a basic tool to study thermal protective performance of a fabric.

  1. Space vehicle integrated thermal protection/structural/meteoroid protection system, volume 1

    NASA Technical Reports Server (NTRS)

    Bartlett, D. H.; Zimmerman, D. K.

    1973-01-01

    A program was conducted to determine the merit of a combined structure/thermal meteoroid protection system for a cryogenic vehicle propulsion module. Structural concepts were evaluated to identify least weight designs. Thermal analyses determined optimum tank arrangements and insulation materials. Meteoroid penetration experiments provided data for design of protection systems. Preliminary designs were made and compared on the basis of payload capability. Thermal performance tests demonstrated heat transfer rates typical for the selected design. Meteoroid impact tests verified the protection characteristics. A mockup was made to demonstrate protection system installation. The best design found combined multilayer insulation with a truss structure vehicle body. The multilayer served as the thermal/meteoroid protection system.

  2. Pre-stressed thermal protection systems

    NASA Technical Reports Server (NTRS)

    Dunn, T. J. (Inventor)

    1984-01-01

    A hexagonal protective and high temperature resistant system for the Space Shuttle Orbiter consists of a multiplicity of pockets formed by hexagonally oriented spacer bars secured on the vehicle substructure. A packing of low density insulating batt material 18 in each pocket, and a thin protective panel of laterally resilient advanced carbon-carbon material surmounting the peripherals bars and packing. Each panel has three stepped or offset lips on contiguous edges. At the center of each pocket is a fully insulated stanchion secured to and connecting the substructure and panel for flexing the panel toward the substructure and thereby prestressing the panel and forcing the panel edges firmly against the spacer bars.

  3. Advanced metallic thermal protection systems for reusable launch vehicles

    NASA Astrophysics Data System (ADS)

    Blosser, Max Leon

    2000-10-01

    Metallic thermal protection systems are a key technology that may help achieve the goal of reducing the cost of space access. A study was performed to develop an understanding of the key factors that govern the performance of metallic thermal protection systems for reusable launch vehicles. Multi-disciplinary background information was assembled and reviewed critically to provide a basis for development of improved metallic thermal protection systems. The fundamentals of aerodynamic heating were reviewed and applied to the development of thermal protection systems. General approaches to thermal protection were categorized and critiqued. The high temperature materials used for thermal protection systems (TPS), including insulations, structural materials, and coatings were reviewed. The history of metallic TPS from early pre-Shuttle concepts to current concepts for a reusable launch vehicle was reviewed for the first time. A current advanced metallic TPS concept was presented and systematically analyzed to discover the most important factors governing the thermal performance of metallic TPS. A large number of relevant factors that influence the thermal analysis and thermal performance of metallic TPS were identified and quantified. Detailed finite element computational models were developed for predicting the thermal performance of variations of the advanced metallic TPS concept mounted on a simple, unstiffened structure. The computational models were also used, in an automated iterative procedure, for sizing the metallic TPS to maintain the structure below a specified temperature limit. A statistical sensitivity analysis method, based on orthogonal matrix techniques used in robust design, was used to quantify and rank the relative importance of the various modeling and design factors considered in this study. Results from this study identify factors that have the most potential to improve metallic TPS performance. The thermal properties of the underlying vehicle

  4. Arcjet Testing of Micro-Meteoroid Impacted Thermal Protection Materials

    NASA Technical Reports Server (NTRS)

    Agrawal, Parul; Munk, Michelle M.; Glaab, Louis J.

    2013-01-01

    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.

  5. Support to X-33/Reusable Launch Vehicle Technology Program

    NASA Technical Reports Server (NTRS)

    2000-01-01

    The Primary activities of Lee & Associates for the referenced Purchase Order has been in direct support of the X-33/Reusable Launch Vehicle Technology Program. An independent review to evaluate the X-33 liquid hydrogen fuel tank failure, which recently occurred after-test of the starboard tank has been provided. The purpose of the Investigation team was to assess the tank design modifications, provide an assessment of the testing approach used by MSFC (Marshall Space Flight Center) in determining the flight worthiness of the tank, assessing the structural integrity, and determining the cause of the failure of the tank. The approach taken to satisfy the objectives has been for Lee & Associates to provide the expertise of Mr. Frank Key and Mr. Wayne Burton who have relevant experience from past programs and a strong background of experience in the fields critical to the success of the program. Mr. Key and Mr. Burton participated in the NASA established Failure Investigation Review Team to review the development and process data and to identify any design, testing or manufacturing weaknesses and potential problem areas. This approach worked well in satisfying the objectives and providing the Review Team with valuable information including the development of a Fault Tree. The detailed inputs were made orally in real time in the Review Team daily meetings. The results of the investigation were presented to the MSFC Center Director by the team on February 15, 2000. Attached are four charts taken from that presentation which includes 1) An executive summary, 2) The most probable cause, 3) Technology assessment, and 4) Technology Recommendations for Cryogenic tanks.

  6. X-33/RLV System Health Management/Vehicle Health Management

    NASA Technical Reports Server (NTRS)

    Mouyos, William; Wangu, Srimal

    1998-01-01

    To reduce operations costs, Reusable Launch Vehicles (RLVS) must include highly reliable robust subsystems which are designed for simple repair access with a simplified servicing infrastructure, and which incorporate expedited decision-making about faults and anomalies. A key component for the Single Stage To Orbit (SSTO) RLV system used to meet these objectives is System Health Management (SHM). SHM incorporates Vehicle Health Management (VHM), ground processing associated with the vehicle fleet (GVHM), and Ground Infrastructure Health Management (GIHM). The primary objective of SHM is to provide an automated and paperless health decision, maintenance, and logistics system. Sanders, a Lockheed Martin Company, is leading the design, development, and integration of the SHM system for RLV and for X-33 (a sub-scale, sub-orbit Advanced Technology Demonstrator). Many critical technologies are necessary to make SHM (and more specifically VHM) practical, reliable, and cost effective. This paper will present the X-33 SHM design which forms the baseline for the RLV SHM, and it will discuss applications of advanced technologies to future RLVs. In addition, this paper will describe a Virtual Design Environment (VDE) which is being developed for RLV. This VDE will allow for system design engineering, as well as program management teams, to accurately and efficiently evaluate system designs, analyze the behavior of current systems, and predict the feasibility of making smooth and cost-efficient transitions from older technologies to newer ones. The RLV SHM design methodology will reduce program costs, decrease total program life-cycle time, and ultimately increase mission success.

  7. Assessment of Thermal Control and Protective Coatings

    NASA Technical Reports Server (NTRS)

    Mell, Richard J.

    2000-01-01

    This final report is concerned with the tasks performed during the contract period which included spacecraft coating development, testing, and applications. Five marker coatings consisting of a bright yellow handrail coating, protective overcoat for ceramic coatings, and specialized primers for composites (or polymer) surfaces were developed and commercialized by AZ Technology during this program. Most of the coatings have passed space environmental stability requirements via ground tests and/or flight verification. Marker coatings and protective overcoats were successfully flown on the Passive Optical Sample Assembly (POSA) and the Optical Properties Monitor (OPM) experiments flown on the Russian space station MIR. To date, most of the coatings developed and/or modified during this program have been utilized on the International Space Station and other spacecraft. For ISS, AZ Technology manufactured the 'UNITY' emblem now being flown on the NASA UNITY node (Node 1) that is docked to the Russian Zarya (FGB) utilizing the colored marker coatings (white, blue, red) developed by AZ Technology. The UNITY emblem included the US American flag, the Unity logo, and NASA logo on a white background, applied to a Beta cloth substrate.

  8. A ceramic matrix composite thermal protection system for hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Riccitiello, Salvatore R.; Love, Wendell L.; Pitts, William C.

    1993-01-01

    The next generation of hypersonic vehicles (NASP, SSTO) that require reusable thermal protection systems will experience acreage surface temperatures in excess of 1100 C. More important, they will experience a more severe physical environment than the Space Shuttle due to non-pristine launching and landing conditions. As a result, maintenance, inspection, and replacement factors must be more thoroughly incorporated into the design of the TPS. To meet these requirements, an advanced thermal protection system was conceived, designated 'TOPHAT'. This system consists of a toughened outer ceramic matrix composite (CMC) attached to a rigid reusable surface insulator (RSI) which is directly bonded to the surface. The objective of this effort was to evaluate this concept in an aeroconvective environment, to determine the effect of impacts to the CMC material, and to compare the results with existing thermal protection systems.

  9. Study of skin model and geometry effects on thermal performance of thermal protective fabrics

    NASA Astrophysics Data System (ADS)

    Zhu, Fanglong; Ma, Suqin; Zhang, Weiyuan

    2008-05-01

    Thermal protective clothing has steadily improved over the years as new materials and improved designs have reached the market. A significant method that has brought these improvements to the fire service is the NFPA 1971 standard on structural fire fighters’ protective clothing. However, this testing often neglects the effects of cylindrical geometry on heat transmission in flame resistant fabrics. This paper deals with methods to develop cylindrical geometry testing apparatus incorporating novel skin bioheat transfer model to test flame resistant fabrics used in firefighting. Results show that fabrics which shrink during the test can have reduced thermal protective performance compared with the qualities measured with a planar geometry tester. Results of temperature differences between skin simulant sensors of planar and cylindrical tester are also compared. This test method provides a new technique to accurately and precisely characterize the thermal performance of thermal protective fabrics.

  10. Alternative High Performance Polymers for Ablative Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Boghozian, Tane; Stackpoole, Mairead; Gonzales, Greg

    2015-01-01

    Ablative thermal protection systems are commonly used as protection from the intense heat during re-entry of a space vehicle and have been used successfully on many missions including Stardust and Mars Science Laboratory both of which used PICA - a phenolic based ablator. Historically, phenolic resin has served as the ablative polymer for many TPS systems. However, it has limitations in both processing and properties such as char yield, glass transition temperature and char stability. Therefore alternative high performance polymers are being considered including cyanate ester resin, polyimide, and polybenzoxazine. Thermal and mechanical properties of these resin systems were characterized and compared with phenolic resin.

  11. Composite flexible insulation for thermal protection of space vehicles

    NASA Technical Reports Server (NTRS)

    Kourtides, Demetrius A.; Tran, Huy K.; Chiu, S. Amanda

    1992-01-01

    A composite flexible blanket insulation (CFBI) system considered for use as a thermal protection system for space vehicles is described. This flexible composite insulation system consists of an outer layer of silicon carbide fabric, followed by alumina mat insulation, and alternating layers of aluminized polyimide film and aluminoborosilicate scrim fabric. A potential application of this composite insulation would be as a thermal protection system for the aerobrake of the Aeroassist Space Transfer Vehicle (ASTV). It would also apply to other space vehicles subject to high convective and radiative heating during atmospheric entry. The thermal performance of this composite insulation as exposed to a simulated atmospheric entry environment in a plasma arc test facility is described. Other thermophysical properties which affect the thermal response of this system are also described. Analytical modeling describing the thermal performance of this composite insulation is included. It shows that this composite insulation is effective as a thermal protection system at total heating rates up to 30.6 W/sq cm.

  12. Thermal modeling of a metallic thermal protection tile for entry vehicles

    NASA Technical Reports Server (NTRS)

    Wiese, M. R.

    1986-01-01

    The thermal Energy Flow Simulation (TEFS) computer program was developed to simulate transient heat transfer through composite solids and predict interfacial temperatures. The program and its usage are described. A simulation of the thermal response of a new thermal protection tile design for the Space Shuttle Orbiter is presented and graphically compared with actual data. An example is also provided which shows the program's usage as a design tool for theoretical models.

  13. Thermal stress and radiation protection principles.

    PubMed

    Kheifets, L; Repacholi, M; Saunders, R

    2003-01-01

    Exposure to radiofrequency (RF) fields can occur in residential, occupational and medical settings. Since many technologies use RF fields, it is important to fully investigate their effects on the human body. Since the demonstrated effect of RF exposure is heating, it is important to critically evaluate studies of elevated temperature effects on the human body, from the cellular and tissue level to the whole body level, including potential effects on the susceptible groups such as the very young and the very old. WHO convened a Workshop in the Spring of 2002 on the subject of Adverse Temperature Levels in the Human. The goal of the workshop was to evaluate most recent data useful for the development of science-based RF exposure limits. This paper outlines radiation protection principles that underline such an evaluation. It discusses the quality of literature needed for sound scientific reviews, provides the hierarchy of scientific evidence used to establish effects, distinguish between biological effects and adverse health consequences and indicates how evidence is evaluated. In addition, criteria for determining the most sensitive effects, the value of an effect that has a dose-response and methods of extrapolation are also described. Finally, the need to account for scientific uncertainty in the formulation of guidance on exposure is discussed.

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

    NASA Technical Reports Server (NTRS)

    Squire, Thomas H.; Milos, Frank S.; Rasky, Daniel J. (Technical Monitor)

    1997-01-01

    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.

  15. Thermal protection studies of plastic films and fibrous materials

    NASA Technical Reports Server (NTRS)

    Saad, Michel A.; Altman, Robert L.

    1988-01-01

    The thermal protection properties of various film and woven materials were studied using an experimental method of radiant heating. The materials studied included aluminized and unaluminized synthetic plastic films and fibrous materials like silicon carbide and phenolic novolac. It is shown that a thin metallized coating with good reflectivity significantly enhances the heat blocking capability of a variety of insulative materials.

  16. European Directions for Hypersonic Thermal Protection Systems and Hot Structures

    NASA Technical Reports Server (NTRS)

    Glass, David E.

    2007-01-01

    This presentation will overview European Thermal Protection Systems (TPS) and Hot Structures activities in Europe. The Europeans have a lot of very good work going on in the area. The presentation will discuss their emphasis on focused technology development for their flight vehicles.

  17. 49 CFR 179.18 - Thermal protection systems.

    Code of Federal Regulations, 2011 CFR

    2011-10-01

    ... 49 Transportation 3 2011-10-01 2011-10-01 false Thermal protection systems. 179.18 Section 179.18 Transportation Other Regulations Relating to Transportation (Continued) PIPELINE AND HAZARDOUS MATERIALS SAFETY ADMINISTRATION, DEPARTMENT OF TRANSPORTATION (CONTINUED) SPECIFICATIONS FOR TANK CARS General Design...

  18. Intelligent, Self-Diagnostic Thermal Protection System for Future Spacecraft

    NASA Technical Reports Server (NTRS)

    Hyers, Robert W.; SanSoucie, Michael P.; Pepyne, David; Hanlon, Alaina B.; Deshmukh, Abhijit

    2005-01-01

    The goal of this project is to provide self-diagnostic capabilities to the thermal protection systems (TPS) of future spacecraft. Self-diagnosis is especially important in thermal protection systems (TPS), where large numbers of parts must survive extreme conditions after weeks or years in space. In-service inspections of these systems are difficult or impossible, yet their reliability must be ensured before atmospheric entry. In fact, TPS represents the greatest risk factor after propulsion for any transatmospheric mission. The concepts and much of the technology would be applicable not only to the Crew Exploration Vehicle (CEV), but also to ablative thermal protection for aerocapture and planetary exploration. Monitoring a thermal protection system on a Shuttle-sized vehicle is a daunting task: there are more than 26,000 components whose integrity must be verified with very low rates of both missed faults and false positives. The large number of monitored components precludes conventional approaches based on centralized data collection over separate wires; a distributed approach is necessary to limit the power, mass, and volume of the health monitoring system. Distributed intelligence with self-diagnosis further improves capability, scalability, robustness, and reliability of the monitoring subsystem. A distributed system of intelligent sensors can provide an assurance of the integrity of the system, diagnosis of faults, and condition-based maintenance, all with provable bounds on errors.

  19. Arc Jet Testing of Thermal Protection Materials: 3 Case Studies

    NASA Technical Reports Server (NTRS)

    Johnson, Sylvia; Conley, Joe

    2015-01-01

    Arc jet testing is used to simulate entry to test thermal protection materials. This paper discusses the usefulness of arc jet testing for 3 cases. Case 1 is MSL and PICA, Case 2 is Advanced TUFROC, and Case 3 is conformable ablators.

  20. Artist concept computer graphic of Lockheed Martin X-33 Advance Technology Demonstrator vehicle in f

    NASA Technical Reports Server (NTRS)

    1998-01-01

    An artist's conception of the X-33 in flight, with the aerospike engine firing. The X-33 demonstrator was designed to test a wide range of new technologies (including the aerospike engine), that would be used in a future single-stage-to-orbit reusable launch vehicle called the VentureStar. Due to technical problems with the liquid hydrogen tank, however, the X-33 program was cancelled in February 2001.

  1. Computer graphic of Lockheed Martin X-33 Reusable Launch Vehicle (RLV)

    NASA Technical Reports Server (NTRS)

    1997-01-01

    Another artist's conception of the X-33, this time after engine shutdown. The vehicle is shown gliding toward its landing site in the southwestern U.S. The X-33 was undertaken to demonstrate the technologies required for a full scale, single-stage-to-orbit launch vehicle. The goal was to substantially reduce the cost of putting payloads into orbit. This proved elusive, and for a variety of reasons, the X-33 was cancelled in February 2001.

  2. Computer graphic of Lockheed Martin X-33 Reusable Launch Vehicle over clouds and water

    NASA Technical Reports Server (NTRS)

    1997-01-01

    Another artist's conception of the X-33, this time after engine shutdown. The vehicle is shown gliding toward its landing site in the southwestern U.S. The X-33 was undertaken to demonstrate the technologies required for a full scale, single-stage-to-orbit launch vehicle. The goal was to substantially reduce the cost of putting payloads into orbit. This proved elusive, and for a variety of reasons, the X-33 was cancelled in February 2001.

  3. Investigation of Fundamental Modeling and Thermal Performance Issues for a Metallic Thermal Protection System Design

    NASA Technical Reports Server (NTRS)

    Blosser, Max L.

    2002-01-01

    A study was performed to develop an understanding of the key factors that govern the performance of metallic thermal protection systems for reusable launch vehicles. A current advanced metallic thermal protection system (TPS) concept was systematically analyzed to discover the most important factors governing the thermal performance of metallic TPS. A large number of relevant factors that influence the thermal analysis and thermal performance of metallic TPS were identified and quantified. Detailed finite element models were developed for predicting the thermal performance of design variations of the advanced metallic TPS concept mounted on a simple, unstiffened structure. The computational models were also used, in an automated iterative procedure, for sizing the metallic TPS to maintain the structure below a specified temperature limit. A statistical sensitivity analysis method, based on orthogonal matrix techniques used in robust design, was used to quantify and rank the relative importance of the various modeling and design factors considered in this study. Results of the study indicate that radiation, even in small gaps between panels, can reduce significantly the thermal performance of metallic TPS, so that gaps should be eliminated by design if possible. Thermal performance was also shown to be sensitive to several analytical assumptions that should be chosen carefully. One of the factors that was found to have the greatest effect on thermal performance is the heat capacity of the underlying structure. Therefore the structure and TPS should be designed concurrently.

  4. Engineering Aerothermal Analysis for X-34 Thermal Protection System Design

    NASA Technical Reports Server (NTRS)

    Wurster, Kathryn E.; Riley, Christopher J.; Zoby, E. Vincent

    1998-01-01

    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.

  5. Engineering Aerothermal Analysis for X-34 Thermal Protection System Design

    NASA Technical Reports Server (NTRS)

    Wurster, Kathryn E.; Riley, Christopher J.; Zoby, E. Vincent

    1998-01-01

    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.

  6. Quantitative thermal diffusivity imaging of disbonds in thermal protective coatings using inductive heating

    NASA Technical Reports Server (NTRS)

    Heath, D. M.; Winfree, William P.

    1990-01-01

    An inductive heating technique for making thermal diffusivity images of disbonds between thermal protective coatings and their substrates is presented. Any flaw in the bonding of the coating and the substrate shows as an area of lowered values in the diffusivity image. The benefits of the inductive heating approach lie in its ability to heat the conductive substrate without directly heating the dielectric coating. Results are provided for a series of samples with fabricated disbonds, for a range of coating thicknesses.

  7. Development of processing techniques for advanced thermal protection materials

    NASA Technical Reports Server (NTRS)

    Selvaduray, Guna S.

    1995-01-01

    The main purpose of this work has been in the development and characterization of materials for high temperature applications. Thermal Protection Systems (TPS) are constantly being tested, and evaluated for increased thermal shock resistance, high temperature dimensional stability, and tolerance to environmental effects. Materials development was carried out through the use of many different instruments and methods, ranging from extensive elemental analysis to physical attributes testing. The six main focus areas include: (1) protective coatings for carbon/carbon composites; (2) TPS material characterization; (3) improved waterproofing for TPS; (4) modified ceramic insulation for bone implants; (5) improved durability ceramic insulation blankets; and (6) ultra-high temperature ceramics. This report describes the progress made in these research areas during this contract period.

  8. The Challenges of Credible Thermal Protection System Reliability Quantification

    NASA Technical Reports Server (NTRS)

    Green, Lawrence L.

    2013-01-01

    The paper discusses several of the challenges associated with developing a credible reliability estimate for a human-rated crew capsule thermal protection system. The process of developing such a credible estimate is subject to the quantification, modeling and propagation of numerous uncertainties within a probabilistic analysis. The development of specific investment recommendations, to improve the reliability prediction, among various potential testing and programmatic options is then accomplished through Bayesian analysis.

  9. Impact Testing of Orbiter Thermal Protection System Materials

    NASA Technical Reports Server (NTRS)

    Kerr, Justin

    2006-01-01

    This viewgraph presentation reviews the impact testing of the materials used in designing the shuttle orbiter thermal protection system (TPS). Pursuant to the Columbia Accident Investigation Board recommendations a testing program of the TPS system was instituted. This involved using various types of impactors in different sizes shot from various sizes and strengths guns to impact the TPS tiles and the Leading Edge Structural Subsystem (LESS). The observed damage is shown, and the resultant lessons learned are reviewed.

  10. Evaluation of Thermal Protection Tile Transmissibility for Ground Vibration Test

    NASA Technical Reports Server (NTRS)

    Chung, Y. T.; Fowler, Samuel B.; Lo, Wenso; Towner, Robert

    2005-01-01

    Transmissibility analyses and tests were conducted on a composite panel with thermal protection system foams to evaluate the quality of the measured frequency response functions. Both the analysis and the test results indicate that the vehicle dynamic responses are fully transmitted to the accelerometers mounted on the thermal protection system in the normal direction below a certain frequency. In addition, the in-plane motions of the accelerometer mounted on the top surface of the thermal protection system behave more actively than those on the composite panel due to the geometric offset of the accelerometer from the panel in the test set-up. The transmissibility tests and analyses show that the frequency response functions measured from the accelerometers mounted on the TPS will provide accurate vehicle responses below 120 Hz for frequency and mode shape identification. By confirming that accurate dynamic responses below a given frequency can be obtained, this study increases the confidence needed for conducting the modal testing, model correlation, and model updating for a vehicle installed with TPS. '

  11. Protection of Measles Virus by Sulfate Ions Against Thermal Inactivation.

    PubMed

    Rapp, F; Butel, J S; Wallis, C

    1965-07-01

    Rapp, Fred (Baylor University College of Medicine, Houston, Tex.), Janet S. Butel, and Craig Wallis. Protection of measles virus by sulfate ions against thermal inactivation. J. Bacteriol. 90:132-135. 1965.-The infectivity of measles virus in water is rapidly destroyed at temperatures of 37 C and above. More than 50% of the infectivity is lost after 1 hr at 25 C, and almost 90% loss of infectivity occurs within 24 hr at 4 C. Magnesium chloride enhances the inactivation of the virus at all temperatures tested. Addition of either magnesium or sodium sulfate protects the virus against thermal inactivation. The stabilizing effect is demonstrable at temperatures ranging from 4 to 56 C, but is especially pronounced through 45 C. Prolonged storage (up to 6 weeks) of the virulent virus at 4 C in 1 m magnesium sulfate permits retention of substantial infectivity, whereas storage at 4 C in either water or 1 m magnesium chloride results in a loss of infectivity approximating 99% after 2 weeks. Magnesium chloride also enhances inactivation of the attenuated vaccine strain of measles virus. The attenuated virus, however, is strongly protected by magnesium sulfate against thermal inactivation, and retention of infectivity for long periods of time at 4 C seems feasible when the virus is kept in 1 m magnesium sulfate.

  12. Interfacial thermal conductance of thiolate-protected gold nanospheres

    NASA Astrophysics Data System (ADS)

    Stocker, Kelsey M.; Neidhart, Suzanne M.; Gezelter, J. Daniel

    2016-01-01

    Molecular dynamics simulations of thiolate-protected and solvated gold nanoparticles were carried out in the presence of a non-equilibrium heat flux between the solvent and the core of the particle. The interfacial thermal conductance (G) was computed for these interfaces, and the behavior of the thermal conductance was studied as a function of particle size, ligand flexibility, and ligand chain length. In all cases, thermal conductance of the ligand-protected particles was higher than the bare metal-solvent interface. A number of mechanisms for the enhanced conductance were investigated, including thiolate-driven corrugation of the metal surface, solvent ordering at the interface, solvent-ligand interpenetration, and ligand ordering relative to the particle surface. Only the smallest particles exhibited significant corrugation. All ligands permitted substantial solvent-ligand interpenetration, and ligand chain length has a significant influence on the orientational ordering of interfacial solvent. Solvent-ligand vibrational overlap, particularly in the low frequency range (<80 cm-1), was significantly altered by ligand rigidity, and had direct influence on the interfacial thermal conductance.

  13. Thermal stability of ceramic coated thermal protection materials in a simulated high-speed earth entry

    NASA Technical Reports Server (NTRS)

    Stewart, David A.; Leiser, Daniel B.

    1988-01-01

    The dimensional stability of ceramic coated thermal protection materials developed for use on advanced entry vehicles is evaluated. Dimensional stability of these ceramic materials were studied as a function of temperature and pressure during exposure to simulated atmospheric entry in an arc-jet facility.

  14. Thermal-Acoustic Analysis of a Metallic Integrated Thermal Protection System Structure

    NASA Technical Reports Server (NTRS)

    Behnke, Marlana N.; Sharma, Anurag; Przekop, Adam; Rizzi, Stephen A.

    2010-01-01

    A study is undertaken to investigate the response of a representative integrated thermal protection system structure under combined thermal, aerodynamic pressure, and acoustic loadings. A two-step procedure is offered and consists of a heat transfer analysis followed by a nonlinear dynamic analysis under a combined loading environment. Both analyses are carried out in physical degrees-of-freedom using implicit and explicit solution techniques available in the Abaqus commercial finite-element code. The initial study is conducted on a reduced-size structure to keep the computational effort contained while validating the procedure and exploring the effects of individual loadings. An analysis of a full size integrated thermal protection system structure, which is of ultimate interest, is subsequently presented. The procedure is demonstrated to be a viable approach for analysis of spacecraft and hypersonic vehicle structures under a typical mission cycle with combined loadings characterized by largely different time-scales.

  15. Thermal Protection Materials Technology for NASA's Exploration Systems Mission Directorate

    NASA Technical Reports Server (NTRS)

    Valentine, Peter G.; Lawerence, Timtohy W.; Gubert, Michael K.; Flynn, Kevin C.; Milos, Frank S.; Kiser, James D.; Ohlhorst, Craig W.; Koenig, John R.

    2005-01-01

    To fulfill the President s Vision for Space Exploration - successful human and robotic missions between the Earth and other solar system bodies in order to explore their atmospheres and surfaces - NASA must reduce trip time, cost, and vehicle weight so that payload and scientific experiment capabilities are maximized. As a collaboration among NASA Centers, this project will generate products that will enable greater fidelity in mission/vehicle design trade studies, support risk reduction for material selections, assist in optimization of vehicle weights, and provide the material and process templates for development of human-rated qualification and certification Thermal Protection System (TPS) plans. Missions performing aerocapture, aerobraking, or direct aeroentry rely on technologies that reduce vehicle weight by minimizing the need for propellant. These missions use the destination planet s atmosphere to slow the spacecraft. Such mission profiles induce heating environments on the spacecraft that demand thermal protection heatshields. This program offers NASA essential advanced thermal management technologies needed to develop new lightweight nonmetallic TPS materials for critical thermal protection heatshields for future spacecraft. Discussion of this new program (a December 2004 new start) will include both initial progress made and a presentation of the work to be preformed over the four-year life of the program. Additionally, the relevant missions and environments expected for Exploration Systems vehicles will be presented, along with discussion of the candidate materials to be considered and of the types of testing to be performed (material property tests, space environmental effects tests, and Earth and Mars gases arc jet tests).

  16. 46 CFR 199.214 - Immersion suits and thermal protective aids.

    Code of Federal Regulations, 2014 CFR

    2014-10-01

    ... 46 Shipping 7 2014-10-01 2014-10-01 false Immersion suits and thermal protective aids. 199.214... Passenger Vessels § 199.214 Immersion suits and thermal protective aids. (a) Each passenger vessel must... an immersion suit. (c) The immersion suits and thermal protective aids required under paragraphs...

  17. 46 CFR 199.214 - Immersion suits and thermal protective aids.

    Code of Federal Regulations, 2013 CFR

    2013-10-01

    ... 46 Shipping 7 2013-10-01 2013-10-01 false Immersion suits and thermal protective aids. 199.214... Passenger Vessels § 199.214 Immersion suits and thermal protective aids. (a) Each passenger vessel must... an immersion suit. (c) The immersion suits and thermal protective aids required under paragraphs...

  18. 46 CFR 199.214 - Immersion suits and thermal protective aids.

    Code of Federal Regulations, 2012 CFR

    2012-10-01

    ... 46 Shipping 7 2012-10-01 2012-10-01 false Immersion suits and thermal protective aids. 199.214... Passenger Vessels § 199.214 Immersion suits and thermal protective aids. (a) Each passenger vessel must... an immersion suit. (c) The immersion suits and thermal protective aids required under paragraphs...

  19. 46 CFR 199.214 - Immersion suits and thermal protective aids.

    Code of Federal Regulations, 2011 CFR

    2011-10-01

    ... 46 Shipping 7 2011-10-01 2011-10-01 false Immersion suits and thermal protective aids. 199.214... Passenger Vessels § 199.214 Immersion suits and thermal protective aids. (a) Each passenger vessel must... an immersion suit. (c) The immersion suits and thermal protective aids required under paragraphs...

  20. 46 CFR 199.214 - Immersion suits and thermal protective aids.

    Code of Federal Regulations, 2010 CFR

    2010-10-01

    ... 46 Shipping 7 2010-10-01 2010-10-01 false Immersion suits and thermal protective aids. 199.214... Passenger Vessels § 199.214 Immersion suits and thermal protective aids. (a) Each passenger vessel must... an immersion suit. (c) The immersion suits and thermal protective aids required under paragraphs...

  1. Ablation Modeling of Ares-I Upper State Thermal Protection System Using Thermal Desktop

    NASA Technical Reports Server (NTRS)

    Sharp, John R.; Page, Arthur T.

    2007-01-01

    The thermal protection system (TPS) for the Ares-I Upper Stage will be based on Space Transportation System External Tank (ET) and Solid Rocket Booster (SRB) heritage materials. These TPS materials were qualified via hot gas testing that simulated ascent and re-entry aerothermodynamic convective heating environments. From this data, the recession rates due to ablation were characterized and used in thermal modeling for sizing the thickness required to maintain structural substrate temperatures. At Marshall Space Flight Center (MSFC), the in-house code ABL is currently used to predict TPS ablation and substrate temperatures as a FORTRAN application integrated within SINDA/G. This paper describes a comparison of the new ablation utility in Thermal Desktop and SINDA/FLUINT with the heritage ABL code and empirical test data which serves as the validation of the Thermal Desktop software for use on the design of the Ares-I Upper Stage project.

  2. Thermal Protective Coating for High Temperature Polymer Composites

    NASA Technical Reports Server (NTRS)

    Barron, Andrew R.

    1999-01-01

    The central theme of this research is the application of carboxylate-alumoxane nanoparticles as precursors to thermally protective coatings for high temperature polymer composites. In addition, we will investigate the application of carboxylate-alumoxane nanoparticle as a component to polymer composites. The objective of this research was the high temperature protection of polymer composites via novel chemistry. The significance of this research is the development of a low cost and highly flexible synthetic methodology, with a compatible processing technique, for the fabrication of high temperature polymer composites. We proposed to accomplish this broad goal through the use of a class of ceramic precursor material, alumoxanes. Alumoxanes are nano-particles with a boehmite-like structure and an organic periphery. The technical goals of this program are to prepare and evaluate water soluble carboxylate-alumoxane for the preparation of ceramic coatings on polymer substrates. Our proposed approach is attractive since proof of concept has been demonstrated under the NRA 96-LeRC-1 Technology for Advanced High Temperature Gas Turbine Engines, HITEMP Program. For example, carbon and Kevlar(tm) fibers and matting have been successfully coated with ceramic thermally protective layers.

  3. MMOD Protection and Degradation Effects for Thermal Control Systems

    NASA Technical Reports Server (NTRS)

    Christiansen, Eric

    2014-01-01

    Micrometeoroid and orbital debris (MMOD) environment overview Hypervelocity impact effects & MMOD shielding MMOD risk assessment process Requirements & protection techniques - ISS - Shuttle - Orion/Commercial Crew Vehicles MMOD effects on spacecraft systems & improving MMOD protection - Radiators Coatings - Thermal protection system (TPS) for atmospheric entry vehicles Coatings - Windows - Solar arrays - Solar array masts - EVA Handrails - Thermal Blankets Orbital Debris provided by JSC & is the predominate threat in low Earth orbit - ORDEM 3.0 is latest model (released December 2013) - http://orbitaldebris.jsc.nasa.gov/ - Man-made objects in orbit about Earth impacting up to 16 km/s average 9-10 km/s for ISS orbit - High-density debris (steel) is major issue Meteoroid model provided by MSFC - MEM-R2 is latest release - http://www.nasa.gov/offices/meo/home/index.html - Natural particles in orbit about sun Mg-silicates, Ni-Fe, others - Meteoroid environment (MEM): 11-72 km/s Average 22-23 km/s.

  4. Thermomechanical analysis of a damaged thermal protection system

    NASA Astrophysics Data System (ADS)

    Ng, Wei Heok

    Research on the effects of damage on the thermomechanical performance and structural integrity of thermal protection systems (TPS) has been limited. The objective of this research is to address this need by conducting experiments and finite element (FE) analysis on damaged TPS. The TPS selected for study is the High-Temperature Reusable Insulation (HRSI) tiles that are used extensively on NASA's Space Shuttle Orbiter. The TPS considered, which consists of a LI-900 tile, the strain isolator pad and the underlying structure, is subjected to the thermal loading and re-entry static pressure of the Access to Space reference vehicle. The damage to the TPS emulates hypervelocity-impact-type damage, which is approximated in the current research by a cylindrical hole ending with a spherical cap. Preliminary FE analysis using several simplifying assumptions, was conducted to determine the accuracy of using an approximate axisymmetric model compared to a complete three-dimensional model for both heat transfer and thermal stress analyses. Temperature results from the two models were found to be reasonable close; however, thermal stress results displayed significant differences. The sensitivity of the FE results to the various simplifying assumptions was also examined and it was concluded that for reliable results, the simplifying assumptions were not acceptable. Subsequently, an exact three-dimensional model was developed and validated by comparison with experimental data. Re-entry static pressures and temperatures were simulated using a high-temperature experimental facility that consists of a quartz radiant heater and a vacuum chamber with appropriate instrumentation. This facility was developed during the course of this dissertation. Temperatures on the top and bottom surfaces of the TPS specimen as well as strains in the underlying structure were recorded for FE model validation. The validated FE model was then combined with improved thermal loads based on the interactions

  5. Development of processing techniques for advanced thermal protection materials

    NASA Technical Reports Server (NTRS)

    Selvaduray, Guna S.

    1994-01-01

    The effort, which was focused on the research and development of advanced materials for use in Thermal Protection Systems (TPS), has involved chemical and physical testing of refractory ceramic tiles, fabrics, threads and fibers. This testing has included determination of the optical properties, thermal shock resistance, high temperature dimensional stability, and tolerance to environmental stresses. Materials have also been tested in the Arc Jet 2 x 9 Turbulent Duct Facility (TDF), the 1 atmosphere Radiant Heat Cycler, and the Mini-Wind Tunnel Facility (MWTF). A significant part of the effort hitherto has gone towards modifying and upgrading the test facilities so that meaningful tests can be carried out. Another important effort during this period has been the creation of a materials database. Computer systems administration and support have also been provided. These are described in greater detail below.

  6. Ballistic Performance of Porous-Ceramic, Thermal-Protection-Systems

    NASA Technical Reports Server (NTRS)

    Christiansen, E. L.; Davis, B. A.; Miller, J. E.; Bohl, W. E.; Foreman, C. D.

    2009-01-01

    Porous-ceramic, thermal protection systems are used heavily in current reentry vehicles like the Space Shuttle and are currently being proposed for the next generation of manned spacecraft, Orion. These materials insulate the structural components of a spacecraft against the intense thermal environments of atmospheric reentry. Furthermore, these materials are also highly exposed to space environmental hazards like meteoroid and orbital debris impacts. This paper discusses recent impact testing up to 9 km/s, and the findings of the influence of material equation-of-state on the simulation of the impact event to characterize the ballistic performance of these materials. These results will be compared with heritage models1 for these materials developed from testing at lower velocities. Assessments of predicted spacecraft risk based upon these tests and simulations will also be discussed.

  7. Thermal Protection System Aerothermal Screening Tests in HYMETS Facility

    NASA Technical Reports Server (NTRS)

    Szalai, Christine E.; Beck, Robin A. S.; Gasch, Matthew J.; Alumni, Antonella I.; Chavez-Garcia, Jose F.; Splinter, Scott C.; Gragg, Jeffrey G.; Brewer, Amy

    2011-01-01

    The Entry, Descent, and Landing (EDL) Technology Development Project has been tasked to develop Thermal Protection System (TPS) materials for insertion into future Mars Entry Systems. A screening arc jet test of seven rigid ablative TPS material candidates was performed in the Hypersonic Materials Environmental Test System (HYMETS) facility at NASA Langley Research Center, in both an air and carbon dioxide test environment. Recession, mass loss, surface temperature, and backface thermal response were measured for each test specimen. All material candidates survived the Mars aerocapture relevant heating condition, and some materials showed a clear increase in recession rate in the carbon dioxide test environment. These test results supported subsequent down-selection of the most promising material candidates for further development.

  8. Development of Processing Techniques for Advanced Thermal Protection Materials

    NASA Technical Reports Server (NTRS)

    Selvaduray, Guna; Lacson, Jamie; Collazo, Julian

    1997-01-01

    During the period June 1, 1996 through May 31, 1997, the main effort has been in the development of materials for high temperature applications. Thermal Protection Systems (TPS) are constantly being tested and evaluated for thermal shock resistance, high temperature dimensional stability, and tolerance to environmental effects. Materials development was carried out by using many different instruments and methods, ranging from intensive elemental analysis to testing the physical attributes of a material. The material development concentrated on two key areas: (1) development of coatings for carbon/carbon composites, and (2) development of ultra-high temperature ceramics (UHTC). This report describes the progress made in these two areas of research during this contract period.

  9. Structural evaluation of candidate space shuttle thermal protection systems

    NASA Technical Reports Server (NTRS)

    Burns, A. B.

    1972-01-01

    The characteristics and development of a lightweight reusable thermal protection system for the space shuttle are discussed. The test articles consisted of metallic substrates with upper surfaces covered with all-silica, reusable, surface insulation material. The material is processed in the form of tiles. The external surfaces of the tiles are provided with a coating system which consists of a borosilicate coating with a silicon carbide emittance agent and impregnation with a hydrophobic agent. The finished tiles are attached to the metal substrate by adhesive bonding. Charts and graphs of the properties of the material are provided.

  10. Coated columbium thermal protection systems: An assessment of technological readiness

    NASA Technical Reports Server (NTRS)

    Levine, S. R.; Grisaffe, S. J.

    1973-01-01

    Evaluation and development to date show that of the coated columbium alloys FS-85 coated with R512E shows significant promise for a reusable thermal protection system (TPS) as judged by environmental resistance and the retention of mechanical properties and structural integrity of panels upon repeated reentry simulation. Production of the alloy, the coating, and full-sized TPS panels is well within current manufacturing technology. Small defects which arise from impact damage or from local coating breakdown do not appear to have serious immediate consequences in the use environment anticipated for the space shuttle orbiter TPS.

  11. MSFC Thermal Protection System Materials on MISSE-6

    NASA Technical Reports Server (NTRS)

    Finckenor, Miria M.; Valentine, Peter G.; Gubert, Michael K.

    2010-01-01

    The Lightweight Nonmetallic Thermal Protection Materials Technology (LNTPMT) program studied a number of ceramic matrix composites, ablator materials, and tile materials for durability in simulated space environment. Materials that indicated low atomic oxygen reactivity and negligible change in thermo-optical properties in ground testing were selected to fly on the Materials on International Space Station Experiment (MISSE)-6. These samples were exposed for 17 months to the low Earth orbit environment on both the ram and wake sides of MISSE-6B. Thermo-optical properties are discussed, along with any changes in mass.

  12. Thermal Protection Materials and Systems: Past, Present, and Future

    NASA Technical Reports Server (NTRS)

    Johnson, Sylvia M.

    2013-01-01

    Thermal protection materials and systems (TPS) protect vehicles from the heat generated when entering a planetary atmosphere. NASA has developed many TPS systems over the years for vehicle ranging from planetary probes to crewed vehicles. The goal for all TPS is efficient and reliable performance. Efficient means using the right material for the environment and minimizing the mass of the heat shield without compromising safety. Efficiency is critical if the payload such as science experiments is to be maximized on a particular vehicle. Reliable means that we understand and can predict performance of the material. Although much characterization and testing of materials is performed to qualify and certify them for flight, it is not possible to completely recreate the reentry conditions in test facilities, and flight-testing

  13. Design of experiments for thermal protection system process optimization

    NASA Astrophysics Data System (ADS)

    Longani, Hans R.

    2000-01-01

    Solid Rocket Booster (SRB) structures were protected from heating due to aeroshear, radiation and plume impingement by a Thermal Protection System (TPS) known as Marshall Sprayable Ablative (MSA-2). MSA-2 contains Volatile Organic Compounds (VOCs) which due to strict environmental legislation was eliminated. MSA-2 was also classified as hazardous waste, which makes the disposal very costly. Marshall Convergent Coating (MCC-1) replaced MSA-2, and eliminated the use of solvents by delivering the dry filler materials and the fluid resin system to a patented spray gun which utilizes Convergent Spray Technologies spray process. The selection of TPS material was based on risk assessment, performance comparisons, processing, application and cost. Design of Experiments technique was used to optimize the spraying parameters. .

  14. Thermal stress analysis of space shuttle orbiter wing skin panel and thermal protection system

    NASA Technical Reports Server (NTRS)

    Ko, William L.; Jenkins, Jerald M.

    1987-01-01

    Preflight thermal stress analysis of the space shuttle orbiter wing skin panel and the thermal protection system (TPS) was performed. The heated skin panel analyzed was rectangular in shape and contained a small square cool region at its center. The wing skin immediately outside the cool region was found to be close to the state of elastic instability in the chordwise direction based on the conservative temperature distribution. The wing skin was found to be quite stable in the spanwise direction. The potential wing skin thermal instability was not severe enough to tear apart the strain isolation pad (SIP) layer. Also, the preflight thermal stress analysis was performed on the TPS tile under the most severe temperature gradient during the simulated reentry heating. The tensile thermal stress induced in the TPS tile was found to be much lower than the tensile strength of the TPS material. The thermal bending of the TPS tile was not severe enough to cause tearing of the SIP layer.

  15. Exercising divers' thermal protection as a function of water temperature.

    PubMed

    Pendergast, David R; Mollendorf, Joseph

    2011-01-01

    Physiological adjustments and passive thermal insulation are not sufficient to protect divers in the cold and warm waters experienced by sport, professional and military divers. In a previous study of resting subjects, divers were protected by actively heated/cooled water that perfused a six-zone (head, torso, arms, hands, legs and feet) tube suit. Subsequently a self-contained diver thermal protection system (DTPS) was developed and used in this study to test male divers (n = 8) wearing a 6-mm foam neoprene wetsuit in water temperatures (T(W)) of 10 degrees C-39 degrees C at 4 feet in depth. The DTPS is a scuba backpack containing five thermoelectric devices that heat/cool water to 30 degrees C, six pumps that circulate the water through a six-zone tube suit via two manifolds, and an electronic controller. Skin temperatures (T(S), n = 17) and core temperature (T(C), capsule) were measured. The DTPS and each zone of the tube suit were also instrumented. Divers were tested with the DTPS operational (protected) and turned off (unprotected) for 90 minutes. In the unprotected condition, T(S) decreased and approached T(W), while T(C) trended to decrease over the exposure time. Mean T(S) as a function of T(W) was T(S) = 0.44 T(W) + 21.23 degrees C while unprotected, but T(S) = 0.19 T(W) + 27.1 degrees C when the diver was protected. The average total heating/cooling power required to protect the diver was 166 +/- 78W, 86 +/- 95W, 9 +/- 75W, 72 +/- 45W, 135 +/- 73W, 279 +/- 87W and 336 +/- 95W at 10, 15, 20, 25, 30, 35 and 39 degrees C water temperatures, respectively. This power requirement was nominally split 4%, 22%, 22%, 14%, 25% and 13% for head, torso, arms, hands, legs and feet, respectively. While unprotected, divers T(S) and T(C) did not remain within acceptable limits in T(W) below 25 degrees C or above 30 degrees C. When using the DTPS, however, they did remain within acceptable limits, and the divers reported they were comfortable.

  16. Ballistic performance of porous-ceramic, thermal protection systems

    NASA Astrophysics Data System (ADS)

    Miller, Joshua E.; Bohl, William E.; Christiansen, Eric C.; Davis, Bruce A.; Foreman, Cory D.

    2012-03-01

    Porous-ceramic, thermal protection systems are used heavily in current reentry vehicles like the Orbiter, and they are currently being proposed for the next generation of US manned spacecraft, Orion. These systems insulate reentry critical components of a spacecraft against the intense thermal environments of atmospheric reentry. Additionally, these materials are highly exposed to space environment hazards like solid particle impacts. This paper discusses impact studies up to 10 km/s on 8 lb/ft3 alumina-fiber-enhanced-thermal-barrier (AETB8) tiles coated with a toughened-unipiece-fibrousinsulation/ reaction-cured-glass layer (TUFI/RCG). A semi-empirical, first principles impact model that describes projectile dispersion is described that provides excellent agreement with observations over a broad range of impact velocities, obliquities and projectile materials. Model extensions to look at the implications of greater than 10 GPa equation of state is also discussed. Predicted penetration probabilities for a vehicle visiting the International Space Station is 60% lower for orbital debris and 95% lower for meteoroids with this model compared to an energy scaled approach.

  17. Ballistic Performance of Porous-Ceramic, Thermal Protection Systems

    NASA Astrophysics Data System (ADS)

    Miller, Joshua; Bohl, William; Christiansen, Eric; Davis, B. Alan; Foreman, Cory

    2011-06-01

    Porous-ceramic, thermal protection systems are used heavily in current reentry vehicles like the Orbiter, and they are currently being proposed for the next generation of US manned spacecraft, Orion. These systems insulate reentry critical components of a spacecraft against the intense thermal environments of atmospheric reentry. Additionally, these materials are also highly exposed to space environment hazards like solid particle impacts. This paper discusses impact testing up to 9.65 km/s on one of these systems. The materials considered are 8 lb/ft3 alumina-fiber-enhanced-thermal-barrier (AETB8) tiles coated with a toughened-unipiece-fibrous-insulation/reaction-cured-glass layer (TUFI/RCG). A semi-empirical, first principals impact model that describes projectile dispersion is described that provides excellent agreement with observations over a broad range of impact velocities, obliquities and projectile materials. A model extension to look at the implications of greater than 10 GPa equation of state measurements is also discussed. Predicted penetration probabilities for a vehicle visiting the International Space Station is 60% lower for orbital debris and 95% lower for meteoroids with this model compared to an energy scaled approach.

  18. Ballistic Performance of Porous-Ceramic, Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Miller, J. E.; Bohl, W. E.; Christiansen, Eric C.; Davis, B. A.; Foreman, C. D.

    2011-01-01

    Porous-ceramic, thermal protection systems are used heavily in current reentry vehicles like the Orbiter, and they are currently being proposed for the next generation of US manned spacecraft, Orion. These systems insulate reentry critical components of a spacecraft against the intense thermal environments of atmospheric reentry. Additionally, these materials are highly exposed to space environment hazards like solid particle impacts. This paper discusses impact studies up to 10 km/s on 8 lb/cu ft alumina-fiber-enhanced-thermal-barrier (AETB8) tiles coated with a toughened-unipiece-fibrous-insulation/ reaction-cured-glass layer (TUFI/RCG). A semi-empirical, first principals impact model that describes projectile dispersion is described that provides excellent agreement with observations over a broad range of impact velocities, obliquities and projectile materials. Model extensions to look at the implications of greater than 10 GPa equation of state is also discussed. Predicted penetration probabilities for a vehicle visiting the International Space Station is 60% lower for orbital debris and 95% lower for meteoroids with this model compared to an energy scaled approach.

  19. Heat flux instrumentation for HYFLITE thermal protection system

    NASA Technical Reports Server (NTRS)

    Diller, T. E.

    1994-01-01

    Tasks performed in this project were defined in a September 9, 1994 meeting of representatives of Vatell, NASA Lewis and Virginia Tech. The overall objective agreed upon in the meeting was 'to demonstrate the viability of thin film techniques for heat flux and temperature sensing in HYSTEP thermal protection systems'. We decided to attempt a combination of NASA's and Vatell's best heat flux sensor technology in a sensor which would be tested in the Vortek facility at Lewis early in 1995. The NASA concept for thermocouple measurement of surface temperature was adopted, and Vatell methods for fabrication of sensors on small diameter substrates of aluminum nitride were used to produce a sensor. This sensor was then encapsulated in a NARloy-Z housing. Various improvements to the Vatell substrate design were explored without success. The basic NASA and Vatell sensor layouts were analyzed by finite element modeling, in an attempt to better understand the effects of material properties, dimensions and thermal differential element location on sensor symmetry, bandwidth and sensitivity. This analysis showed that, as long as the thermal resistivity of the thermal differential element material is much larger (10X) than that of the substrate material, the simplest arrangement of layer is best. During calibration of the sensor produced in this project, undesirable side-effects of combining the heat flux and temperature sensor return leads were observed. The sensor did not cleanly separate the heat flux and temperature signals, as sensors with four leads have consistently done before. Task 7 and 8 discussed in the meeting will be performed with a continuation of funding in 1995. The following is a discussion of each of the tasks performed as outlined in the statement of work dated september 26, 1994. Task 1A was added to cover further investigation into the NASA sensor concept.

  20. Heat flux instrumentation for HYFLITE thermal protection system

    NASA Astrophysics Data System (ADS)

    Diller, T. E.

    1994-12-01

    Tasks performed in this project were defined in a September 9, 1994 meeting of representatives of Vatell, NASA Lewis and Virginia Tech. The overall objective agreed upon in the meeting was 'to demonstrate the viability of thin film techniques for heat flux and temperature sensing in HYSTEP thermal protection systems'. We decided to attempt a combination of NASA's and Vatell's best heat flux sensor technology in a sensor which would be tested in the Vortek facility at Lewis early in 1995. The NASA concept for thermocouple measurement of surface temperature was adopted, and Vatell methods for fabrication of sensors on small diameter substrates of aluminum nitride were used to produce a sensor. This sensor was then encapsulated in a NARloy-Z housing. Various improvements to the Vatell substrate design were explored without success. The basic NASA and Vatell sensor layouts were analyzed by finite element modeling, in an attempt to better understand the effects of material properties, dimensions and thermal differential element location on sensor symmetry, bandwidth and sensitivity. This analysis showed that, as long as the thermal resistivity of the thermal differential element material is much larger (10X) than that of the substrate material, the simplest arrangement of layer is best. During calibration of the sensor produced in this project, undesirable side-effects of combining the heat flux and temperature sensor return leads were observed. The sensor did not cleanly separate the heat flux and temperature signals, as sensors with four leads have consistently done before. Task 7 and 8 discussed in the meeting will be performed with a continuation of funding in 1995. The following is a discussion of each of the tasks performed as outlined in the statement of work dated september 26, 1994. Task 1A was added to cover further investigation into the NASA sensor concept.

  1. 77 FR 11598 - Thermal Overload Protection for Electric Motors on Motor-Operated Valves

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-02-27

    ... COMMISSION Thermal Overload Protection for Electric Motors on Motor-Operated Valves AGENCY: Nuclear... (NRC or Commission) is issuing a revision to Regulatory Guide (RG) 1.106, ``Thermal Overload Protection... NRC's staff for complying with NRC requirements for the application of thermal overload...

  2. Thermal Cycling Assessment of Steel-Based Thermal Barrier Coatings for Al Protection

    NASA Astrophysics Data System (ADS)

    Poirier, Dominique; Lamarre, Jean-Michel; Legoux, Jean-Gabriel

    2015-01-01

    There is a strong interest from the transportation industry to achieve vehicle weight reduction through the replacement of steel components by aluminum parts. For some applications, aluminum requires protective coatings due to its limited wear and lower temperature resistance compared to steel. The objective of this study was to assess the potential of amorphous-type plasma-sprayed steel coatings and conventional arc-sprayed steel coatings as thermal barrier coatings, mainly through the evaluation of their spalling resistance under thermal cycling. The microstructures of the different coatings were first compared via SEM. The amorphicity of the coatings produced via plasma spraying of specialized alloyed steel and the crystalline phases of the conventional arc-sprayed steel coatings were confirmed through x-ray diffraction. The thermal diffusivity of all coatings produced was measured to be about a third of that of bulk stainless steel. Conventional arc-sprayed steel coatings typically offered better spalling resistance under thermal cycling than steel-based amorphous coatings due probably to their higher initial bond strength. However, the presence of vertical cracks in the steel-based amorphous coatings was found to have a beneficial effect on their thermal cycling resistance. The amorphous plasma-sprayed steel coatings presented indications of recrystallization after their exposure to high temperature.

  3. Thermal face protection delays finger cooling and improves thermal comfort during cold air exposure.

    PubMed

    O'Brien, Catherine; Castellani, John W; Sawka, Michael N

    2011-12-01

    When people dress for cold weather, the face often remains exposed. Facial cooling can decrease finger blood flow, reducing finger temperature (T (f)). This study examined whether thermal face protection limits finger cooling and thereby improves thermal comfort and manual dexterity during prolonged cold exposure. T (f) was measured in ten volunteers dressed in cold-weather clothing as they stood for 60 min facing the wind (-15°C, 3 m s(-1)), once while wearing a balaclava and goggles (BAL), and once with the balaclava pulled down and without goggles (CON). Subjects removed mitts, wearing only thin gloves to perform Purdue Pegboard (PP) tests at 15 and 50 min, and Minnesota Rate of Manipulation (MRM) tests at 30 and 55 min. Subjects rated their thermal sensation and comfort just before the dexterity tests. T (f) decreased (p < 0.05 for time × trial interaction) by 15 min of cold exposure during CON (33.6 ± 1.4-28.7 ± 2.0°C), but not during BAL (33.2 ± 1.4-30.6 ± 3.2°C); and after 30 min T (f) remained warmer during BAL (23.3 ± 5.9°C) than CON (19.2 ± 3.5); however, by 50 min, T (f) was no different between trials (14.1 ± 2.7°C). Performance on PP fell (p < 0.05) by 25% after 50 min in both trials; MRM performance was not altered by cold on either trial. Subjects felt colder (p < 0.05) and more uncomfortable (p < 0.05) during CON, compared to BAL. Thermal face protection was effective for maintaining warmer T (f) and thermal comfort during cold exposure; however, local cooling of the hands during manual dexterity tests reduced this physiological advantage, and performance was not improved.

  4. Integrated Thermal Protection Systems and Heat Resistant Structures

    NASA Technical Reports Server (NTRS)

    Pichon, Thierry; Lacoste, Marc; Glass, David E.

    2006-01-01

    In the early stages of NASA's Exploration Initiative, Snecma Propulsion Solide was funded under the Exploration Systems Research & Technology program to develop integrated thermal protection systems and heat resistant structures for reentry vehicles. Due to changes within NASA's Exploration Initiative, this task was cancelled early. This presentation provides an overview of the work that was accomplished prior to cancellation. The Snecma team chose an Apollo-type capsule as the reference vehicle for the work. They began with the design of a ceramic aft heatshield (CAS) utilizing C/SiC panels as the capsule heatshield, a C/SiC deployable decelerator and several ablators. They additionally developed a health monitoring system, high temperature structures testing, and the insulation characterization. Though the task was pre-maturely cancelled, a significant quantity of work was accomplished.

  5. Terahertz Computed Tomography of NASA Thermal Protection System Materials

    NASA Technical Reports Server (NTRS)

    Roth, D. J.; Reyes-Rodriguez, S.; Zimdars, D. A.; Rauser, R. W.; Ussery, W. W.

    2011-01-01

    A terahertz axial computed tomography system has been developed that uses time domain measurements in order to form cross-sectional image slices and three-dimensional volume renderings of terahertz-transparent materials. The system can inspect samples as large as 0.0283 cubic meters (1 cubic foot) with no safety concerns as for x-ray computed tomography. In this study, the system is evaluated for its ability to detect and characterize flat bottom holes, drilled holes, and embedded voids in foam materials utilized as thermal protection on the external fuel tanks for the Space Shuttle. X-ray micro-computed tomography was also performed on the samples to compare against the terahertz computed tomography results and better define embedded voids. Limits of detectability based on depth and size for the samples used in this study are loosely defined. Image sharpness and morphology characterization ability for terahertz computed tomography are qualitatively described.

  6. High temperature electromagnetic characterization of thermal protection system tile materials

    NASA Technical Reports Server (NTRS)

    Heil, Garrett G.

    1993-01-01

    This study investigated the impact of elevated temperatures on the electromagnetic performance of the LI-2200 thermal protection system. A 15-kilowatt CO2 laser was used to heat an LI-2200 specimen to 3000 F while electromagnetic measurements were performed over the frequency range of l9 to 21 GHz. The electromagnetic measurement system consisted of two Dual-Lens Spot-Focusing (DLSF) antennas, a sample support structure, and an HP-8510B vector network analyzer. Calibration of the electromagnetic system was accomplished with a Transmission-Reflection-Line (TRL) procedure and was verified with measurements on a two-layer specimen of known properties. The results of testing indicated that the LI-2200 system's electromagnetic performance is slightly temperature dependent at temperatures up to 3000 F.

  7. Microscopy and microstructure of Shuttle thermal protection system materials

    NASA Technical Reports Server (NTRS)

    Newquist, C. W.; Pfister, A. M.; Miller, A. D.; Scott, W. D.

    1981-01-01

    Examples of the contribution of microstructural analysis to the development of the Space Shuttle tile insulation system are presented, with photographic examples of the scanning electron microscope (SEM) investigations. After the basic thermal protection system materials had been selected, it was neccessary to analyze the mechanical responses of the combined materials; which included: (1) the polymer strain isolation pad (SIP), (2) the room temperature-vulcanizing silicone rubber bond, (RTV), and (3) rigid ceramic fiber reusable surface insulation (RSI). Microstructural analysis was used to provide information on deformation and fracture mechanisms, load transfer mechanisms, and structural alterations occurring before final failure. Both quantitative and qualitative information was obtained in the open, three-dimensional fibrous structures of the ceramic tiles by means of novel techniques of encapsulation and dissolution.

  8. Fracture behavior of the Space Shuttle thermal protection system

    NASA Technical Reports Server (NTRS)

    Komine, A.; Kobayashi, A. S.

    1983-01-01

    Stable crack-growth and fracture-toughness experiments were conducted using precracked specimens machined from LI-900 reusable surface insulation (RSI) tiles of the Space Shuttle thermal protection system (TPS) at room temperature. Similar fracture experiments were conducted on fracture specimens with preexisting cracks at the interface of the tile and the strain isolation pad (SIP). Stable crack growth was not observed in the LI-900 tile fracture specimens which had a fracture toughness of 12.0 kPa sq rt of m. The intermittent subcritical crack growth at the tile-pad interface of the fracture specimens was attributed to successive local pull-outs due to tensile overload in the LI-900 tile and cannot be characterized by linear elastic fracture mechanics. No subcritical interfacial crack growth was observed in the fracture specimens with densified LI-900 tiles where brittle fracture initiated at an interior point away from the densification.

  9. Biologically-Derived Photonic Materials for Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Johnson, Sylvia M.; Squire, Thomas H.; Lawson, John W.; Gusman, Michael; Lau, K.-H.; Sanjurjo, Angel

    2014-01-01

    Space vehicles entering a planetary atmosphere at high velocity can be subject to substantial radiative heating from the shock layer in addition to the convective heating caused by the flow of hot gas past the vehicle surface. The radiative component can be very high but of a short duration. Approaches to combat this effect include investigation of various materials to reflect the radiation. Photonic materials can be used to reflect radiation. The wavelengths reflected depend on the length scale of the ordered microstructure. Fabricating photonic structures, such as layers, can be time consuming and expensive. We have used a biologically-derived material as the template for forming a high temperature photonic material that could be incorporated into a heatshield thermal protection material.

  10. Fracture behavior of the Space Shuttle thermal protection system

    SciTech Connect

    Komine, A.; Kobayashi, A.S.

    1983-09-01

    Stable crack-growth and fracture-toughness experiments were conducted using precracked specimens machined from LI-900 reusable surface insulation (RSI) tiles of the Space Shuttle thermal protection system (TPS) at room temperature. Similar fracture experiments were conducted on fracture specimens with preexisting cracks at the interface of the tile and the strain isolation pad (SIP). Stable crack growth was not observed in the LI-900 tile fracture specimens which had a fracture toughness of 12.0 kPa sq rt of m. The intermittent subcritical crack growth at the tile-pad interface of the fracture specimens was attributed to successive local pull-outs due to tensile overload in the LI-900 tile and cannot be characterized by linear elastic fracture mechanics. No subcritical interfacial crack growth was observed in the fracture specimens with densified LI-900 tiles where brittle fracture initiated at an interior point away from the densification. 11 references.

  11. Aerothermodynamic Assessment of Corrugated Panel Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Brandon, H. J.; Britt, A. H.; Kipp, H. W.; Masek, R. V.

    1978-01-01

    The feasibility of using corrugated panels as a thermal protection system for an advanced space transportation vehicle was investigated. The study consisted of two major tasks: development of improved correlations for wind tunnel heat transfer and pressure data to yield design techniques, and application of the design techniques to determine if corrugated panels have application future aerospace vehicles. A single-stage-to-orbit vehicle was used to assess advantages and aerothermodynamic penalties associated with use of such panels. In the correlation task, experimental turbulent heat transfer and pressure data obtained on corrugation roughened surfaces during wind tunnel testing were analyzed and compared with flat plate data. The correlations and data comparisons included the effects of a large range of geometric, inviscid flow, internal boundary layer, and bulk boundary layer parameters in supersonic and hypersonic flow.

  12. Thermal decomposition of fullerene nanowhiskers protected by amorphous carbon mask

    PubMed Central

    Guo, Hongxuan; Wang, Chengxiang; Miyazawa, Kun’ichi; Wang, Hongxin; Masuda, Hideki; Fujita, Daisuke

    2016-01-01

    Fullerene nanostructures are well known for their unique morphology, physical and mechanical properties. The thermal stability of fullerene nanostructures, such as their sublimation at high temperature is also very important for studying their structures and applications. In this work, We observed fullerene nanowhiskers (FNWs) in situ with scanning helium ion microscopy (HIM) at elevated temperatures. The FNWs exhibited different stabilities with different thermal histories during the observation. The pristine FNWs were decomposed at the temperatures higher than 300 °C in a vacuum environment. Other FNWs were protected from decomposition with an amorphous carbon (aC) film deposited on the surface. Based on high spacial resolution, aC film with periodic structure was deposited by helium ion beam induced deposition (IBID) on the surface of FNWs. Annealed at the high temperature, the fullerene molecules were selectively sublimated from the FNWs. The periodic structure was formed on the surface of FNWs and observed by HIM. Monte Carlo simulation and Raman characterization proved that the morphology of the FNWs was changed by helium IBID at high temperature. This work provides a new method of fabricating artificial structure on the surface of FNWs with periodic aC film as a mask. PMID:27991498

  13. Thermal decomposition of fullerene nanowhiskers protected by amorphous carbon mask

    NASA Astrophysics Data System (ADS)

    Guo, Hongxuan; Wang, Chengxiang; Miyazawa, Kun’Ichi; Wang, Hongxin; Masuda, Hideki; Fujita, Daisuke

    2016-12-01

    Fullerene nanostructures are well known for their unique morphology, physical and mechanical properties. The thermal stability of fullerene nanostructures, such as their sublimation at high temperature is also very important for studying their structures and applications. In this work, We observed fullerene nanowhiskers (FNWs) in situ with scanning helium ion microscopy (HIM) at elevated temperatures. The FNWs exhibited different stabilities with different thermal histories during the observation. The pristine FNWs were decomposed at the temperatures higher than 300 °C in a vacuum environment. Other FNWs were protected from decomposition with an amorphous carbon (aC) film deposited on the surface. Based on high spacial resolution, aC film with periodic structure was deposited by helium ion beam induced deposition (IBID) on the surface of FNWs. Annealed at the high temperature, the fullerene molecules were selectively sublimated from the FNWs. The periodic structure was formed on the surface of FNWs and observed by HIM. Monte Carlo simulation and Raman characterization proved that the morphology of the FNWs was changed by helium IBID at high temperature. This work provides a new method of fabricating artificial structure on the surface of FNWs with periodic aC film as a mask.

  14. Flexible Thermal Protection System Development for Hypersonic Inflatable Aerodynamic Decelerators

    NASA Technical Reports Server (NTRS)

    DelCorso, Joseph A.; Bruce, Walter E., III; Hughes, Stephen J.; Dec, John A.; Rezin, Marc D.; Meador, Mary Ann B.; Guo, Haiquan; Fletcher, Douglas G.; Calomino, Anthony M.; Cheatwood, McNeil

    2012-01-01

    The Hypersonic Inflatable Aerodynamic Decelerators (HIAD) project has invested in development of multiple thermal protection system (TPS) candidates to be used in inflatable, high downmass, technology flight projects. Flexible TPS is one element of the HIAD project which is tasked with the research and development of the technology ranging from direct ground tests, modelling and simulation, characterization of TPS systems, manufacturing and handling, and standards and policy definition. The intent of flexible TPS is to enable large deployable aeroshell technologies, which increase the drag performance while significantly reducing the ballistic coefficient of high-mass entry vehicles. A HIAD requires a flexible TPS capable of surviving aerothermal loads, and durable enough to survive the rigors of construction, handling, high density packing, long duration exposure to extrinsic, in-situ environments, and deployment. This paper provides a comprehensive overview of key work being performed within the Flexible TPS element of the HIAD project. Included in this paper is an overview of, and results from, each Flexible TPS research and development activity, which includes ground testing, physics-based thermal modelling, age testing, margins policy, catalysis, materials characterization, and recent developments with new TPS materials.

  15. Shearographic nondestructive evaluation of Space Shuttle thermal protection systems

    NASA Astrophysics Data System (ADS)

    Davis, Christopher K.; Hooker, Jeffery A.; Simmons, Stephen M.; Tenbusch, Kenneth E.

    1995-07-01

    Preliminary results of shearographic inspections of the shuttle external tank (ET) spray-on foam insulation (SOFI) and solid rocket booster (SRB) Marshall sprayable ablative (MSA-2) epoxy-cork thermal protection systems (TPS) are presented. Debonding SOFI or MSA-2 damage the orbiter `belly' tile and exposes the ET/SRB to thermal loading. Previous work with the ET/SRB showed promising results with shearography. The first area investigated was the jack pad close-out, one of many areas on the ET where foam is applied at KSC. Voids 0.375 inch were detected in 1.75 inch thick foam using a pressure reduction of less than 0.4 psi. Of primary interest are areas of the ET that directly face the orbiter tile TPS. It is estimated that 90% of tile TPS damage on the orbiter `belly' results from debonding SOFI during ascent. Test panels modeling these areas were manufactured with programmed debonds to determine the sensitivity of shearography as a function of debond size, SOFI thickness and vacuum. Results show repeatable detection of debonds with a diameter approximately half the SOFI thickness at less than 0.4 psi pressure reduction. Preliminary results are also presented on inspections of MSA-2 and the remote manipulator system (RMS) honeycomb material.

  16. Nanostructured Thermal Protection Systems for Space Exploration Missions

    NASA Technical Reports Server (NTRS)

    Arnold, J. O.; Chen, Y. K.; Squire, T.; Srivastava, D.; Allen, G., Jr.; Stackpoole, M.; Goldstein, H. E.; Venkatapathy, E.; Loomis, M. P.

    2005-01-01

    Strong research and development programs in nanotechnology and Thermal Protection Systems (TPS) exist at NASA Ames. Conceptual studies have been undertaken to determine if new, nanostructured materials (composites of existing TPS materials and nanostructured composite fibers) could improve the performance of TPS. To this end, we have studied various candidate heatshields, some composed of existing TPS materials (with known material properties), to provide a baseline for comparison with others that are admixtures of such materials and a nanostructured material. In the latter case, some assumptions were made about the thermal conductivity and strength of the admixture, relative to the baseline TPS material. For the purposes of this study, we have made the conservative assumption that only a small fraction of the remarkable properties of carbon nanotubes (for example) will be realized in the material properties of the admixtures employing them. The heatshields studied included those for Sharp leading edges (appropriate to out-of-orbit entry and aero-maneuvering), probes, an out-of-orbit Apollo Command Module (as a surrogate for NASA's new Crew Exploration Vehicle [CEV]), a Mars Sample Return Vehicle and a large heat shield for Mars aerocapture missions. We report on these conceptual studies, which show that in some cases (not all), significant improvements in the TPS can be achieved through the use of nanostructured materials.

  17. Thermal certification tests of Orbiter Thermal Protection System tiles coated with KSC coating slurries

    NASA Technical Reports Server (NTRS)

    Milhoan, James D.; Pham, Vuong T.; Sherborne, William D.

    1993-01-01

    Thermal tests of Orbiter thermal protection system (TPS) tiles, which were coated with borosilicate glass slurries fabricated at Kennedy Space Center (KSC), were performed in the Radiant Heat Test Facility and the Atmospheric Reentry Materials & Structures Evaluation Facility at Johnson Space Center to verify tile coating integrity after exposure to multiple entry simulation cycles in both radiant and convective heating environments. Eight high temperature reusable surface insulation (HRSI) tiles and six low temperature reusable surface insulation (LRSI) tiles were subjected to 25 cycles of radiant heat at peaked surface temperatures of 2300 F and 1200 F, respectively. For the LRSI tiles, an additional cycle at peaked surface temperature of 2100 F was performed. There was no coating crack on any of the HRSI specimens. However, there were eight small coating cracks (less than 2 inches long) on two of the six LRSI tiles on the 26th cycle. There was practically no change on the surface reflectivity, physical dimensions, or weight of any of the test specimens. There was no observable thermal-chemical degradation of the coating either. For the convective heat test, eight HRSI tiles were tested for five cycles at a surface temperature of 2300 F. There was no thermal-induced coating crack on any of the test specimens, almost no change on the surface reflectivity, and no observable thermal-chemical degradation with an exception of minor slumping of the coating under painted TPS identification numbers. The tests demonstrated that KSC's TPS slurries and coating processes meet the Orbiter's thermal specification requirements.

  18. Modern air protection technologies at thermal power plants (review)

    NASA Astrophysics Data System (ADS)

    Roslyakov, P. V.

    2016-07-01

    Realization of the ecologically safe technologies for fuel combustion in the steam boiler furnaces and the effective ways for treatment of flue gases at modern thermal power plants have been analyzed. The administrative and legal measures to stimulate introduction of the technologies for air protection at TPPs have been considered. It has been shown that both the primary intrafurnace measures for nitrogen oxide suppression and the secondary flue gas treatment methods are needed to meet the modern ecological standards. Examples of the environmentally safe methods for flame combustion of gas-oil and solid fuels in the boiler furnaces have been provided. The effective methods and units to treat flue gases from nitrogen and sulfur oxides and flue ash have been considered. It has been demonstrated that realization of the measures for air protection should be accompanied by introduction of the systems for continuous instrumentation control of the composition of combustion products in the gas path of boiler units and for monitoring of atmospheric emissions.

  19. Hypervelocity Impact Test Results for a Metallic Thermal Protection System

    NASA Technical Reports Server (NTRS)

    Karr, Katherine L.; Poteet, Carl C.; Blosser, Max L.

    2003-01-01

    Hypervelocity impact tests have been performed on specimens representing metallic thermal protection systems (TPS) developed at NASA Langley Research Center for use on next-generation reusable launch vehicles (RLV). The majority of the specimens tested consists of a foil gauge exterior honeycomb panel, composed of either Inconel 617 or Ti-6Al-4V, backed with 2.0 in. of fibrous insulation and a final Ti-6Al-4V foil layer. Other tested specimens include titanium multi-wall sandwich coupons as well as TPS using a second honeycomb sandwich in place of the foil backing. Hypervelocity impact tests were performed at the NASA Marshall Space Flight Center Orbital Debris Simulation Facility. An improved test fixture was designed and fabricated to hold specimens firmly in place during impact. Projectile diameter, honeycomb sandwich material, honeycomb sandwich facesheet thickness, and honeycomb core cell size were examined to determine the influence of TPS configuration on the level of protection provided to the substructure (crew, cabin, fuel tank, etc.) against micrometeoroid or orbit debris impacts. Pictures and descriptions of the damage to each specimen are included.

  20. Thermal Protection Studies of Synthetic And Woven Materials

    NASA Technical Reports Server (NTRS)

    Saad, Michel A.; Altman, Robert L.; Ransky, Daniel J. (Technical Monitor)

    1995-01-01

    This paper presents results of experimental study to evaluate the thermal protection properties of synthetic felt and woven materials using an NBS smoke chamber. The chamber was modified to record the weight loss of the samples, which in turn, indicated the effectiveness of the insulation material. The following materials were tested: (a) aluminoborosilicate cloth (NEXTEL); (b) fiber glass cloth; (c) carbonized polyaacrylonitrile and rayon cloth; (d) aromatic nylon felt; (e) SiC (NICALON) CLOTH; and (f) phenolic novolac (KYNOL) cloth. Samples of these were put in front of fiber glass batting containing 18% phenolic resin (Owens Corning PF-204). They were exposed to a radiant heat of 5w cm-2 which resulted in an almost complete resin mass loss within four minutes. Results of this study are shown in various figures, where the mass loss from the fiber glass batting is plotted vs. time. In these figures, solid curves show the percent mass loss of the exposed fiber glass and dashed curves indicate the loss in another fiber glass sample of the same initial mass protected by the material under test.

  1. Thermal-Structural Evaluation of TD Ni-20Cr Thermal Protection System Panels

    NASA Technical Reports Server (NTRS)

    Eidinoff, H. L.; Rose, L.

    1974-01-01

    The results of a thermal-structural test program to verify the performance of a metallic/radiative Thermal Protection System (TPS) under reentry conditions are presented. This TPS panel is suitable for multiple reentry, high L/D space vehicles, such as the NASA space shuttle, having surface temperatures up to 1200 C (2200 F). The TPS panel tested consists of a corrugation-stiffened, beaded-skin TD Ni-20Cr metallic heat shield backed by a flexible fibrous quartz and radiative shield insulative system. Test conditions simulated the critical heating and aerodynamic pressure environments expected during 100 repeated missions of a reentry vehicle. Temperatures were measured during each reentry cycle; heat-shield flatness surveys to measure permanent set of the metallic components were made every 10 cycles. The TPS panel, in spite of localized surface failures, performed its designated function.

  2. Three-Dimensional Finite Element Ablative Thermal Response and Thermostructural Design of Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Dec, John A.; Braun, Robert D.

    2011-01-01

    A finite element ablation and thermal response program is presented for simulation of three-dimensional transient thermostructural analysis. The three-dimensional governing differential equations and finite element formulation are summarized. A novel probabilistic design methodology for thermal protection systems is presented. The design methodology is an eight step process beginning with a parameter sensitivity study and is followed by a deterministic analysis whereby an optimum design can determined. The design process concludes with a Monte Carlo simulation where the probabilities of exceeding design specifications are estimated. The design methodology is demonstrated by applying the methodology to the carbon phenolic compression pads of the Crew Exploration Vehicle. The maximum allowed values of bondline temperature and tensile stress are used as the design specifications in this study.

  3. X-33 Model Tested In Langley's 20-Inch Mach 6 Tunnel

    NASA Technical Reports Server (NTRS)

    1997-01-01

    Thomas Horvath of Langley's Aerothermodynamics Branch examines the surface of a model of the X-33 prior to testing in the 20-Inch Mach 6 Air Wind Tunnel at NASA Langley Research Center. The tests, held during the month of September 1997, were conducted to determine aeroheating characteristics of the X-33. The X-33 vehicle will consist of a lifting body airframe with two cryogenic propellant tanks (liquid hydrogen, LH2, and liquid oxygen, LOX) placed within the aeroshell. The vehicle will have two linear aerospike main engines. The X-33 Design and Flight Demonstration Program key objectives are to reduce business and technical risks to privately finance development and operation of a next-generation space transportation system through ground and flight tests of a spaceplane technology demonstrator, ensure that the X-33 design and major components are usable and scaleable to a full-scale, single-stage-orbit reusable launch vehicle (RLV), demonstrate autonomous capability from takeoff to landing, and verify operability and performance in 'real world' environments.

  4. X-33 Model Tested In Langley's 20-Inch Mach 6 Tunnel

    NASA Technical Reports Server (NTRS)

    1997-01-01

    Thomas Horvath of Langley's Aerothermodynamics Branch uses digital instrumentation to set the angle of attack on a model of the X-33 prior to a wind tunnel test run in the 20-Inch Mach 6 Air Wind Tunnel at NASA Langley Research Center. The tests, held during the month of September 1997, were conducted to determine aeroheating characteristics of the X-33. The X-33 vehicle will consist of a lifting body airframe with two cryogenic propellant tanks (liquid hydrogen, LH2, and liquid oxygen, LOX) placed within the aeroshell. The vehicle will have two linear aerospike main engines. The X-33 Design and Flight Demonstration Program key objectives are to reduce business and technical risks to privately finance development and operation of a next-generation space transportation system through ground and flight tests of a spaceplane technology demonstrator, ensure that the X-33 design and major components are usable and scaleable to a full-scale, single-stage-orbit reusable launch vehicle (RLV), demonstrate autonomous capability from takeoff to landing, and verify operability and performance in 'real world' environments.

  5. A Study of the Effects of Altitude on Thermal Ice Protection System Performance

    NASA Technical Reports Server (NTRS)

    Addy, Harold E., Jr.; Oleskiw, Myron; Broeren, Andy P.; Orchard, Andy P.

    2013-01-01

    Thermal ice protection systems use heat energy to prevent a dangerous buildup of ice on an aircraft. As aircraft become more efficient, less heat energy is available to operate a thermal ice protections system. This requires that thermal ice protection systems be designed to more exacting standards so as to more efficiently prevent a dangerous ice buildup without adversely affecting aircraft safety. While the effects of altitude have always beeing taked into account in the design of thermal ice protection systems, a better understanding of these effects is needed so as to enable more exact design, testing, and evaluation of these systems.

  6. 75 FR 72653 - Alternate Fracture Toughness Requirements for Protection Against Pressurized Thermal Shock Events...

    Federal Register 2010, 2011, 2012, 2013, 2014

    2010-11-26

    ... RIN 3150-AI01 Alternate Fracture Toughness Requirements for Protection Against Pressurized Thermal... Regulations (10 CFR) part 50, section 61a to provide alternate fracture toughness requirements for...

  7. Design of metallic foams as insulation in thermal protection systems

    NASA Astrophysics Data System (ADS)

    Zhu, Huadong

    Metallic foams are novel materials that can be used as thermal insulation in many applications. The low volume fraction of solid, the small cell size and the low conductivity of enclosed gases limit the heat flow in foams. Varying the density, geometry and or material composition from point to point within the foam, one can produce functionally graded foams that may insulate more efficiently. The goal of this research is to investigate the use of functionally graded metal foam in thermal protection systems (TPS) for reusable launch vehicles. First, the effective thermal conductivity of the foam is derived based on a simple cubic unit cell model. Then two problems under steady state of heat transfer have been considered. The first one is the optimization of functionally graded foam insulation for minimum heat transmitted to the structure and the second is minimizing the mass of the functionally graded foam insulation for a given aerodynamic heating. In both cases optimality conditions are derived in closed-form, and numerical methods are used to solve the resulting differential equations to determine the optimal grading of the foam. In order to simplify the analysis the insulation was approximated by finite layers of uniform foams when studying the transient heat transfer case. The maximum structure temperature was minimized by varying the solidity profile for a given total thickness and mass. The principles that govern the design of TPS for transient conditions were identified. To take advantage of the load bearing ability of metallic foams, an integrated sandwich TPS/structure with metallic foam core is proposed. Such an integrated TPS will insulate the vehicle interior from aerodynamic heating as well as carry the primary vehicle loads. Thermal-structural analysis of integrated sandwich TPS panel subjected to transient heat conduction is developed to evaluate their performances. The integrated TPS design is compared with a conventional fibrous Safill TPS design

  8. A Strategy for Integrating a Large Finite Element Model: X-33 Lessons Learned

    NASA Technical Reports Server (NTRS)

    McGhee, David S.

    2000-01-01

    The X-33 vehicle is an advanced technology demonstrator sponsored by NASA. For the past three years the Structural Dynamics & Loads Group of NASA's Marshall Space Flight Center has had the task of integrating the X-33 vehicle structural finite element model. In that time, five versions of the integrated vehicle model have been produced and a strategy has evolved that would benefit anyone given the task of integrating structural finite element models that have been generated by various modelers and companies. The strategy that has been presented here consists of six decisions that need to be made. These six decisions are: purpose of model, units, common material list, model numbering, interface control, and archive format. This strategy has been proved and expanded from experience on the X-33 vehicle.

  9. Development of X-33/X-34 Aerothermodynamic Data Bases: Lessons Learned and Future Enhancements

    NASA Technical Reports Server (NTRS)

    Miller, C. G.

    1999-01-01

    A synoptic of programmatic and technical lessons learned in the development of aerothermodynamic data bases for the X-33 and X-34 programs is presented in general terms and from the perspective of the NASA Langley Research Center Aerothermodynamics Branch. The format used is that of the aerothermodynamic chain, the links of which are personnel, facilities, models/test articles, instrumentation, test techniques, and computational fluid dynamics (CFD). Because the aerodynamic data bases upon which the X-33 and X-34 vehicles will fly are almost exclusively from wind tunnel testing, as opposed to CFD, the primary focus of the lessons learned is on ground-based testing.

  10. Real-Time Trajectory Assessment and Abort Management for the X-33 Vehicle

    NASA Technical Reports Server (NTRS)

    Moise, M. C.; McCarter, J. W.; Mulqueen, J.

    2000-01-01

    The X-33 is a flying testbed to evaluate technologies and designs for a reusable Single Stage To Orbit (SSTO) production vehicle. Although it is sub-orbital, it is trans-atmospheric. This paper will discuss the abort capabilities, both commanded and autonomous, available to the X-33. The cornerstone of the abort capabilities is the Performance Monitor (PM) and it's supporting software. PM is an on-board 3-DOF simulation, which evaluates the vehicle ability to execute the current trajectory. The Abort Manager evaluates the results from PM, and, when indicated, computes and implements an abort trajectory.

  11. Effectiveness of Thermal-Pneumatic Airfoil-Ice-Protection System

    NASA Technical Reports Server (NTRS)

    Gowan, William H., Jr.; Mulholland, Donald R.

    1951-01-01

    Icing and drag investigations were conducted in the NACA Lewis icing research tunnel employing a combination thermal-pneumatic de-icer mounted on a 42-inch-chord NACA 0018 airfoil. The de-icer consisted of a 3-inch-wide electrically heated strip symmetrically located about the leading edge with inflatable tubes on the upper and lower airfoil surfaces aft of the heated area. The entire de-icer extended to approximately 25 percent of chord. A maximum power density of 9.25 watts per square inch was required for marginal ice protection on the airfoil leading edge at an air temperature of 00 F and an airspeed of 300 miles per hour. Drag measurements indicated, that without icing, the de-icer installation increased the section drag to approximately 140 percent of that of the bare airfoil; with the tubes inflated, this value increased to a maximum of approximately 620 percent. A 2-minute tube-inflation cycle prevented excessive ice formation on the inflatable area although small scattered residual Ice formations remained after inflation and were removed intermittently during later cycles. Effects of the time lag of heater temperatures after initial application of power and the insulating effect of ice formations on heater temperatures were also determined.

  12. Thermal Protection System (Heat Shield) Development - Advanced Development Project

    NASA Technical Reports Server (NTRS)

    Kowal, T. John

    2010-01-01

    The Orion Thermal Protection System (TPS) ADP was a 3 1/2 year effort to develop ablative TPS materials for the Orion crew capsule. The ADP was motivated by the lack of available ablative TPS's. The TPS ADP pursued a competitive phased development strategy with succeeding rounds of development, testing and down selections. The Project raised the technology readiness level (TRL) of 8 different TPS materials from 5 different commercial vendors, eventual down selecting to a single material system for the Orion heat shield. In addition to providing a heat shield material and design for Orion on time and on budget, the Project accomplished the following: 1) Re-invigorated TPS industry & re-established a NASA competency to respond to future TPS needs; 2) Identified a potentially catastrophic problem with the planned MSL heat shield, and provided a viable, high TRL alternate heat shield design option; and 3) Transferred mature heat shield material and design options to the commercial space industry, including TPS technology information for the SpaceX Dragon capsule.

  13. Heat flux instrumentation for Hyflite thermal protection system

    NASA Technical Reports Server (NTRS)

    Diller, T. E.

    1994-01-01

    Using Thermal Protection Tile core samples supplied by NASA, the surface characteristics of the FRCI, TUFI, and RCG coatings were evaluated. Based on these results, appropriate methods of surface preparation were determined and tested for the required sputtering processes. Sample sensors were fabricated on the RCG coating and adhesion was acceptable. Based on these encouraging results, complete Heat Flux Microsensors were fabricated on the RCG coating. The issue of lead attachment was addressed with the annnealing and welding methods developed at NASA Lewis. Parallel gap welding appears to be the best method of lead attachment with prior heat treatment of the sputtered pads. Sample Heat Flux Microsensors were submitted for testing in the NASA Ames arc jet facility. Details of the project are contained in two attached reports. One additional item of interest is contained in the attached AIAA paper, which gives details of the transient response of a Heat Flux Microsensors in a shock tube facility at Virginia Tech. The response of the heat flux sensor was measured to be faster than 10 micro-s.

  14. Mechanical Testing of Carbon Based Woven Thermal Protection Materials

    NASA Technical Reports Server (NTRS)

    Pham, John; Agrawal, Parul; Arnold, James O.; Peterson, Keith; Venkatapathy, Ethiraj

    2013-01-01

    Three Dimensional Woven thermal protection system (TPS) materials are one of the enabling technologies for mechanically deployable hypersonic decelerator systems. These materials have been shown capable of serving a dual purpose as TPS and as structural load bearing members during entry and descent operations. In order to ensure successful structural performance, it is important to characterize the mechanical properties of these materials prior to and post exposure to entry-like heating conditions. This research focuses on the changes in load bearing capacity of woven TPS materials after being subjected to arcjet simulations of entry heating. Preliminary testing of arcjet tested materials [1] has shown a mechanical degradation. However, their residual strength is significantly more than the requirements for a mission to Venus [2]. A systematic investigation at the macro and microstructural scales is reported here to explore the potential causes of this degradation. The effects of heating on the sizing (an epoxy resin coating used to reduce friction and wear during fiber handling) are discussed as one of the possible causes for the decrease in mechanical properties. This investigation also provides valuable guidelines for margin policies for future mechanically deployable entry systems.

  15. Mars Science Laboratory Entry Capsule Aerothermodynamics and Thermal Protection System

    NASA Technical Reports Server (NTRS)

    Edquist, Karl T.; Hollis, Brian R.; Dyakonov, Artem A.; Laub, Bernard; Wright, Michael J.; Rivellini, Tomasso P.; Slimko, Eric M.; Willcockson, William H.

    2007-01-01

    The Mars Science Laboratory (MSL) spacecraft is being designed to carry a large rover (greater than 800 kg) to the surface of Mars using a blunt-body entry capsule as the primary decelerator. The spacecraft is being designed for launch in 2009 and arrival at Mars in 2010. The combination of large mass and diameter with non-zero angle-of-attack for MSL will result in unprecedented convective heating environments caused by turbulence prior to peak heating. Navier-Stokes computations predict a large turbulent heating augmentation for which there are no supporting flight data1 and little ground data for validation. Consequently, an extensive experimental program has been established specifically for MSL to understand the level of turbulent augmentation expected in flight. The experimental data support the prediction of turbulent transition and have also uncovered phenomena that cannot be replicated with available computational methods. The result is that the flight aeroheating environments predictions must include larger uncertainties than are typically used for a Mars entry capsule. Finally, the thermal protection system (TPS) being used for MSL has not been flown at the heat flux, pressure, and shear stress combinations expected in flight, so a test program has been established to obtain conditions relevant to flight. This paper summarizes the aerothermodynamic definition analysis and TPS development, focusing on the challenges that are unique to MSL.

  16. Interfacial fracture of Space-Shuttle thermal-protection system

    NASA Technical Reports Server (NTRS)

    Komine, A.; Kobayashi, A. S.

    1982-01-01

    Stable crack growth and fracture at the interface of undensified LI-900 reusable surface insulation (RSI) tile and the Nomex strain isolation pad (SIP) of the Space-Shuttle thermal-protection system (TPS) were modeled by double-edged notch-tension specimens. These specimens were loaded under uniaxial tension or 50-Hz cyclic loading and the resultant stable crack growth leading to eventual fracture was monitored by a videocamera. These tests showed that successive local tear-outs due to local tensile overload in the RSI tile resulted in the interfacial fracture where the crack-tip opening angle, CTOA, of the SIP was related to initiation and intermittent stable crack propagation. Fractures in similar static and dynamic test specimens using densified LI-900 RSI tiles occurred in the undensified regions of the RSI tiles. These failures were consistent with the above failure mechanism based on the local tensile strength of the undensified LI-900 RSI tile. The intermittent stable crack growth of undensified LI-900 RSI tile was reproduced by a deterministic, two-dimensional finite-element model with SIP of variable elastic moduli.

  17. Thermal Protection System design studies for lunar crew module

    NASA Technical Reports Server (NTRS)

    Williams, S. D.; Curry, Donald M.; Bouslog, Stanley A.; Rochelle, William C.

    1993-01-01

    The results of a study to predict aeroheating and Thermal Protection System (TPS) requirements for manned entry vehicles returning to Earth from the moon are presented. The effects of vehicle size and lunar-return strategies on the aerothermodynamic environment and TPS design were assessed. Study guidelines were based on an Apollo Command Module (CM) configuration and lunar return strategies included direct entry and aerocapture followed by Low Earth Orbit entry (LEO). Convective heating was obtained by the boundary layer integral matrix procedure (BLIMP) code, and radiative heating was computed with the QRAD program. The AESOP-STAB code and the AESOP-THERM code were used for TPS analysis for ablating materials and nonablating materials respectively. Results indicated that there was an optimum size for minimum heating and that direct entry had higher heating rates than aerocapture. Aerocapture resulted in higher heat loads and TPS weight. The TPS weight factor was 6-8 percent for all lunar return strategies, with the TPS weight being about 50 percent less than that of the original Apollo CM vehicle.

  18. Development of a test device to characterize thermal protective performance of fabrics against hot steam and thermal radiation

    NASA Astrophysics Data System (ADS)

    Su, Yun; Li, Jun

    2016-12-01

    Steam burns severely threaten the life of firefighters in the course of their fire-ground activities. The aim of this paper was to characterize thermal protective performance of flame-retardant fabrics exposed to hot steam and low-level thermal radiation. An improved testing apparatus based on ASTM F2731-11 was developed in order to simulate the routine fire-ground conditions by controlling steam pressure, flow rate and temperature of steam box. The thermal protective performance of single-layer and multi-layer fabric system with/without an air gap was studied based on the calibrated tester. It was indicated that the new testing apparatus effectively evaluated thermal properties of fabric in hot steam and thermal radiation. Hot steam significantly exacerbated the skin burn injuries while the condensed water on the skin’s surface contributed to cool down the skin tissues during the cooling. Also, the absorbed thermal energy during the exposure and the cooling was mainly determined by the fabric’s configuration, the air gap size, the exposure time and the existence of hot steam. The research provides a effective method to characterize the thermal protection of fabric in complex conditions, which will help in optimization of thermal protection performance of clothing and reduction of steam burn.

  19. Physical and mechanical properties and thermal protection efficiency of intumescent coatings

    NASA Astrophysics Data System (ADS)

    Zverev, V. G.; Zinchenko, V. I.; Tsimbalyuk, A. F.

    2016-04-01

    The new engineering technique for the experimental investigation of physical and mechanical characteristics of thermal protective intumescent coatings is offered. A mathematical model is proposed for predicting the thermal behavior of structures protected by coatings; the model is closed by the studied material characteristics. The heating of a metal plate under standard thermal loading conditions is modeled mathematically. The modeling results are in good agreement with bench test results for metal temperature under the coating. The proposed technique of studying physical and mechanical characteristics can be applied to identify and monitor the state of thermal protective intumescent coatings in the long-term operation.

  20. X-33 Attitude Control System Design for Ascent, Transition, and Entry Flight Regimes

    NASA Technical Reports Server (NTRS)

    Hall, Charles E.; Gallaher, Michael W.; Hendrix, Neal D.

    1998-01-01

    The Vehicle Control Systems Team at Marshall Space Flight Center, Systems Dynamics Laboratory, Guidance and Control Systems Division is designing under a cooperative agreement with Lockheed Martin Skunkworks, the Ascent, Transition, and Entry flight attitude control system for the X-33 experimental vehicle. Ascent flight control begins at liftoff and ends at linear aerospike main engine cutoff (NECO) while Transition and Entry flight control begins at MECO and concludes at the terminal area energy management (TAEM) interface. TAEM occurs at approximately Mach 3.0. This task includes not only the design of the vehicle attitude control systems but also the development of requirements for attitude control system components and subsystems. The X-33 attitude control system design is challenged by a short design cycle, the design environment (Mach 0 to about Mach 15), and the X-33 incremental test philosophy. The X-33 design-to-launch cycle of less than 3 years requires a concurrent design approach while the test philosophy requires design adaptation to vehicle variations that are a function of Mach number and mission profile. The flight attitude control system must deal with the mixing of aerosurfaces, reaction control thrusters, and linear aerospike engine control effectors and handle parasitic effects such as vehicle flexibility and propellant sloshing from the uniquely shaped propellant tanks. The attitude control system design is, as usual, closely linked to many other subsystems and must deal with constraints and requirements from these subsystems.

  1. Optimal Trajectory Reconfiguration and Retargeting for the X-33 Reusable Launch Vehicle

    DTIC Science & Technology

    2004-09-01

    6 Figure 4 X-33 Simulation Coordinate System...between the different components of the optimization routine. B. REDUCED-ORDER DYNAMICAL MODEL 1. Coordinate Systems The simulation uses a three...reentry simulation . where γ = flight path angle, the angle made by the velocity vector and the x-y plane. β = azimuth angle, the angle made by

  2. X-33 Aerodynamic and Aeroheating Computations for Wind Tunnel and Flight Conditions

    NASA Technical Reports Server (NTRS)

    Hollis, Brian R.; Thompson, Richard A.; Murphy, Kelly J.; Nowak, Robert J.; Riley, Christopher J.; Wood, William A.; Alter, Stephen J.; Prabhu, Ramadas K.

    1999-01-01

    This report provides an overview of hypersonic Computational Fluid Dynamics research conducted at the NASA Langley Research Center to support the Phase II development of the X-33 vehicle. The X-33, which is being developed by Lockheed-Martin in partnership with NASA, is an experimental Single-Stage-to-Orbit demonstrator that is intended to validate critical technologies for a full-scale Reusable Launch Vehicle. As part of the development of the X-33, CFD codes have been used to predict the aerodynamic and aeroheating characteristics of the vehicle. Laminar and turbulent predictions were generated for the X 33 vehicle using two finite- volume, Navier-Stokes solvers. Inviscid solutions were also generated with an Euler code. Computations were performed for Mach numbers of 4.0 to 10.0 at angles-of-attack from 10 deg to 48 deg with body flap deflections of 0, 10 and 20 deg. Comparisons between predictions and wind tunnel aerodynamic and aeroheating data are presented in this paper. Aeroheating and aerodynamic predictions for flight conditions are also presented.

  3. Design and Calibration of the X-33 Flush Airdata Sensing (FADS) System

    NASA Technical Reports Server (NTRS)

    Whitmore, Stephen A.; Cobleigh, Brent R.; Haering, Edward A.

    1998-01-01

    This paper presents the design of the X-33 Flush Airdata Sensing (FADS) system. The X-33 FADS uses a matrix of pressure orifices on the vehicle nose to estimate airdata parameters. The system is designed with dual-redundant measurement hardware, which produces two independent measurement paths. Airdata parameters that correspond to the measurement path with the minimum fit error are selected as the output values. This method enables a single sensor failure to occur with minimal degrading of the system performance. The paper shows the X-33 FADS architecture, derives the estimating algorithms, and demonstrates a mathematical analysis of the FADS system stability. Preliminary aerodynamic calibrations are also presented here. The calibration parameters, the position error coefficient (epsilon), and flow correction terms for the angle of attack (delta alpha), and angle of sideslip (delta beta) are derived from wind tunnel data. Statistical accuracy of' the calibration is evaluated by comparing the wind tunnel reference conditions to the airdata parameters estimated. This comparison is accomplished by applying the calibrated FADS algorithm to the sensed wind tunnel pressures. When the resulting accuracy estimates are compared to accuracy requirements for the X-33 airdata, the FADS system meets these requirements.

  4. 49 CFR 179.18 - Thermal protection systems.

    Code of Federal Regulations, 2010 CFR

    2010-10-01

    ... thermal resistance so that there will be no release of any lading within the tank car, except release... resistance of the tank car does not conform to paragraph (a) of this section, the thermal resistance of...

  5. 49 CFR 193.2057 - Thermal radiation protection.

    Code of Federal Regulations, 2010 CFR

    2010-10-01

    .... Each LNG container and LNG transfer system must have a thermal exclusion zone in accordance with... computer model GTI-04/0032 LNGFIRE3: A Thermal Radiation Model for LNG Fires (incorporated by...

  6. NASA Crew Exploration Vehicle, Thermal Protection System, Lessons Learned

    NASA Technical Reports Server (NTRS)

    Venkatapathy, Ethiraj; Reuther, James

    2008-01-01

    The Orion (CEV) thermal protection system (TPS) advanced development project (ADP) was initiated in late 2006 to reduce developmental risk by significant investment in multiple heat shield architectural solutions that can meet the needs both the Low Earth orbit (LEO) and Lunar return missions. At the same time, the CEV TPS ADP was also charged with developing a preliminary design for the heat shield to meet the PDR requirement and at the time of the PDR, transfer the design to Lockheed- Martin, the prime contractor. We reported on the developmental activities of the first 18 months at the IPPW5 in Bordeaux, France, last summer. In June 08, at the time of the IPPW6, the CEV TPS ADP would have nearly completed the preparation for the Orion PDR and would be close to the original three-year mark. We plan to report on the progress at the Atlanta workshop. In the past year, Orion TPS ADP investment in TPS Technology, especially in PICA ablative Heat-shield design, development, testing and engineering (DDTE) has paid off in enabling MSL mission to switch from SLA 561 V heat shield to PICA heat shield. CEV TPS ADP considered SLA 561 V as a candidate for LEO missions and our testing identified failure modes in SLA and as a result, we dropped SLA for further evaluation. This close synergy between two projects is a highly visible example of how investment in technology areas can and does benefit multiple missions. In addition, CEV TPS ADP has been able to revive the Apollo ablative system namely AVCOAT honeycomb architecture as an alternate to the baseline PICA architecture and we plan to report the progress we have made in AVCOAT. CEV TPS ADP has invested considerable resources in developing analytical models for PICA and AVCOAT, material property measurements that is essential to the design of the heat-shield, in arcjet testing, in understanding the differences between different arc jet facilities, namely NASA Ames, NASA JSC and Air Force's AEDC, and in Non

  7. Space Shuttle Orbiter flight heating rate measurement sensitivity to thermal protection system uncertainties

    NASA Technical Reports Server (NTRS)

    Bradley, P. F.; Throckmorton, D. A.

    1981-01-01

    A study was completed to determine the sensitivity of computed convective heating rates to uncertainties in the thermal protection system thermal model. Those parameters considered were: density, thermal conductivity, and specific heat of both the reusable surface insulation and its coating; coating thickness and emittance; and temperature measurement uncertainty. The assessment used a modified version of the computer program to calculate heating rates from temperature time histories. The original version of the program solves the direct one dimensional heating problem and this modified version of The program is set up to solve the inverse problem. The modified program was used in thermocouple data reduction for shuttle flight data. Both nominal thermal models and altered thermal models were used to determine the necessity for accurate knowledge of thermal protection system's material thermal properties. For many thermal properties, the sensitivity (inaccuracies created in the calculation of convective heating rate by an altered property) was very low.

  8. Evaluation of Thermal Control Coatings for Flexible Ceramic Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Kourtides, Demetrius; Carroll, Carol; Smith, Dane; Guzinski, Mike; Marschall, Jochen; Pallix, Joan; Ridge, Jerry; Tran, Duoc

    1997-01-01

    This report summarizes the evaluation and testing of high emissivity protective coatings applied to flexible insulations for the Reusable Launch Vehicle technology program. Ceramic coatings were evaluated for their thermal properties, durability, and potential for reuse. One of the major goals was to determine the mechanism by which these coated blanket surfaces become brittle and try to modify the coatings to reduce or eliminate embrittlement. Coatings were prepared from colloidal silica with a small percentage of either SiC or SiB6 as the emissivity agent. These coatings are referred to as gray C-9 and protective ceramic coating (PCC), respectively. The colloidal solutions were either brushed or sprayed onto advanced flexible reusable surface insulation blankets. The blankets were instrumented with thermocouples and exposed to reentry heating conditions in the Ames Aeroheating Arc Jet Facility. Post-test samples were then characterized through impact testing, emissivity measurements, chemical analysis, and observation of changes in surface morphology. The results show that both coatings performed well in arc jet tests with backface temperatures slightly lower for the PCC coating than with gray C-9. Impact testing showed that the least extensive surface destruction was experienced on blankets with lower areal density coatings.

  9. SAFER Inspection of Space Shuttle Thermal Protection System

    NASA Technical Reports Server (NTRS)

    Scoville, Zebulon C.; Rajula, Sudhakar

    2005-01-01

    In the aftermath of the space shuttle Columbia accident, it quickly became clear that new methods would need to be developed that would provide the capability to inspect and repair the shuttle's thermal protection system (TPS). A boom extension to the Remote Manipulator System (RMS) with a laser topography sensor package was identified as the primary means for measuring the damage depth in acreage tile as well as scanning Reinforced Carbon- Carbon (RCC) surfaces. However, concern over the system's fault tolerance made it prudent to investigate alternate means of acquiring close range photographs and contour depth measurements in the event of a failure. One method that was identified early was to use the Simplified Aid For EVA Rescue (SAFER) propulsion system to allow EVA access to damaged areas of concern. Several issues were identified as potential hazards to SAFER use for this operation. First, the ability of an astronaut to maintain controlled flight depends upon efficient technique and hardware reliability. If either of these is insufficient during flight operations, a safety tether must be used to rescue the crewmember. This operation can jeopardize the integrity of the Extra-vehicular Mobility Unit (EMU) or delicate TPS materials. Controls were developed to prevent the likelihood of requiring a tether rescue, and procedures were written to maximize the chances for success if it cannot be avoided. Crewmember ability to manage tether cable tension during nominal flight also had to be evaluated to ensure it would not negatively affect propellant consumption. Second, although propellant consumption, flight control, orbital dynamics, and flight complexity can all be accurately evaluated in Virtual Reality (VR) Laboratory at Johnson Space Center, there are some shortcomings. As a crewmember's hand is extended to simulate measurement of tile damage, it will pass through the vehicle without resistance. In reality, this force will push the crewmember away from the

  10. Updated Results of Deterministic Reconfigurable Control Design for the X-33 Vehicle

    NASA Technical Reports Server (NTRS)

    Cotting, M. Christopher; Burken, John J.

    1999-01-01

    In the event of a control surface failure, the purpose of a reconfigurable control system is to redistribute the control effort among the remaining working surfaces such that satisfactory stability and performance are retained. An Off-line Nonlinear General Constrained Optimization approach was used for the reconfigurable X-33 control design method. Three examples of failure are shown using a high fidelity 6 DOF simulation (case 1: ascent with a left body flap jammed at 25 deg.; case 2: entry with a right inboard elevon jam at 25 deg. and case 3: landing (TAEM) (Terminal Area Energy Management) with a left rudder jam at -30 deg.) Failure comparisons between responses with the nominal controller and reconfigurable controllers show the benefits of reconfiguration. Single jam aerosurface failures were considered, and failure detection and identification is considered accomplished in the actuator controller. The X-33 flight control system will incorporate reconfigurable flight control in the baseline system.

  11. Bioremediation of Parboiled Rice Effluent Supplemented with Biodiesel-Derived Glycerol Using Pichia pastoris X-33

    PubMed Central

    Gil de los Santos, Diego; Gil Turnes, Carlos; Rochedo Conceição, Fabricio

    2012-01-01

    This paper describes the use of Pichia pastoris X-33 as a bioremediator to reduce the chemical oxygen demand (COD), total Kjeldahl nitrogen (TKN), and phosphorus (P-PO4   3−), after culture in parboiled rice effluent supplemented with p.a. glycerol or a glycerol by-product of the biodiesel industry. The greatest reduction in the COD (55%), TKN (45%), and P-PO4   3− (52%) of the effluent was observed in cultures of P. pastoris X-33 supplemented with 15 g ·L−1 of biodiesel-derived glycerol. Furthermore, the overall biomass yield was 2.1 g ·L−1. These data suggest that biodiesel-derived glycerol is an efficient carbon source for the bioremediation of parboiled rice effluent and biomass production. PMID:22919327

  12. Including Aeroelastic Effects in the Calculation of X-33 Loads and Control Characteristics

    NASA Technical Reports Server (NTRS)

    Zeiler, Thomas A.

    1998-01-01

    Up until now, loads analyses of the X-33 RLV have been done at Marshall Space Flight Center (MSFC) using aerodynamic loads derived from CFD and wind tunnel models of a rigid vehicle. Control forces and moments are determined using a rigid vehicle trajectory analysis and the detailed control load distributions for achieving the desired control forces and moments, again on the rigid vehicle, are determined by Lockheed Martin Skunk Works. However, static aeroelastic effects upon the load distributions are not known. The static aeroelastic effects will generally redistribute external loads thereby affecting both the internal structural loads as well as the forces and moments generated by aerodynamic control surfaces. Therefore, predicted structural sizes as well as maneuvering requirements can be altered by consideration of static aeroelastic effects. The objective of the present work is the development of models and solutions for including static aeroelasticity in the calculation of X-33 loads and in the determination of stability and control derivatives.

  13. An Innovative Structural Mode Selection Methodology: Application for the X-33 Launch Vehicle Finite Element Model

    NASA Technical Reports Server (NTRS)

    Hidalgo, Homero, Jr.

    2000-01-01

    An innovative methodology for determining structural target mode selection and mode selection based on a specific criterion is presented. An effective approach to single out modes which interact with specific locations on a structure has been developed for the X-33 Launch Vehicle Finite Element Model (FEM). We presented Root-Sum-Square (RSS) displacement method computes resultant modal displacement for each mode at selected degrees of freedom (DOF) and sorts to locate modes with highest values. This method was used to determine modes, which most influenced specific locations/points on the X-33 flight vehicle such as avionics control components, aero-surface control actuators, propellant valve and engine points for use in flight control stability analysis and for flight POGO stability analysis. Additionally, the modal RSS method allows for primary or global target vehicle modes to also be identified in an accurate and efficient manner.

  14. Ascent, Transition, Entry, and Abort Guidance Algorithm Design for the X-33 Vehicle

    NASA Technical Reports Server (NTRS)

    Hanson, John M.; Coughlin, Dan J.; Dukeman, Gregory A.; Mulqueen, John A.; McCarter, James W.

    1998-01-01

    One of the primary requirements for X-33 is that it be capable of flying autonomously. That is, onboard computers must be capable of commanding the entire flight from launch to landing, including cases where a single engine failure abort occurs. Guidance algorithms meeting these requirements have been tested in simulation and have been coded into prototype flight software. These algorithms must be sufficiently robust to account for vehicle and environmental dispersions, and must issue commands that result in the vehicle operating, within all constraints. Continual tests of these algorithms (and modifications as necessary) will occur over the next year as the X-33 nears its first flight. This paper describes the algorithms in use for X-33 ascent, transition, and entry flight, as well as for the powered phase of PowerPack-out (PPO) aborts (equivalent in thrust impact to losing an engine). All following discussion refers to these phases of flight when discussing guidance. The paper includes some trajectory results and results of dispersion analysis.

  15. An Inviscid Computational Study of an X-33 Configuration at Hypersonic Speeds

    NASA Technical Reports Server (NTRS)

    Prabhu, Ramadas K.

    1999-01-01

    This report documents the results of a study conducted to compute the inviscid longitudinal aerodynamic characteristics of a simplified X-33 configuration. The major components of the X-33 vehicle, namely the body, the canted fin, the vertical fin, and the body-flap, were simulated in the CFD (Computational Fluid Dynamic) model. The rear-ward facing surfaces at the base including the aerospike engine surfaces were not simulated. The FELISA software package consisting of an unstructured surface and volume grid generator and two inviscid flow solvers was used for this study. Computations were made for Mach 4.96, 6.0, and 10.0 with perfect gas air option, and for Mach 10 with equilibrium air option with flow condition of a typical point on the X-33 flight trajectory. Computations were also made with CF4 gas option at Mach 6.0 to simulate the CF4 tunnel flow condition. An angle of attack range of 12 to 48 deg was covered. The CFD results were compared with available wind tunnel data. Comparison was good at low angles of attack; at higher angles of attack (beyond 25 deg) some differences were found in the pitching moment. These differences progressively increased with increase in angle of attack, and are attributed to the viscous effects. However, the computed results showed the trends exhibited by the wind tunnel data.

  16. Numerical investigation on properties of attack angle for an opposing jet thermal protection system

    NASA Astrophysics Data System (ADS)

    Lu, Hai-Bo; Liu, Wei-Qiang

    2012-08-01

    The three-dimensional Navier—Stokes equation and the k-in viscous model are used to simulate the attack angle characteristics of a hemisphere nose-tip with an opposing jet thermal protection system in supersonic flow conditions. The numerical method is validated by the relevant experiment. The flow field parameters, aerodynamic forces, and surface heat flux distributions for attack angles of 0°, 2°, 5°, 7°, and 10° are obtained. The detailed numerical results show that the cruise attack angle has a great influence on the flow field parameters, aerodynamic force, and surface heat flux distribution of the supersonic vehicle nose-tip with an opposing jet thermal protection system. When the attack angle reaches 10°, the heat flux on the windward generatrix is close to the maximal heat flux on the wall surface of the nose-tip without thermal protection system, thus the thermal protection has failed.

  17. Fabrication of titanium multi-wall Thermal Protection System (TPS) test panel arrays

    NASA Technical Reports Server (NTRS)

    Blair, W.; Meaney, J. E.; Rosenthal, H. A.

    1980-01-01

    Several arrays were designed and tested. Tests included vibrational and acoustical tests, radiant heating tests, and thermal conductivity tests. A feasible manufacturing technique was established for producing the protection system panels.

  18. Thermal degradation study of silicon carbide threads developed for advanced flexible thermal protection systems

    NASA Technical Reports Server (NTRS)

    Tran, Huy Kim; Sawko, Paul M.

    1992-01-01

    Silicon carbide (SiC) fiber is a material that may be used in advanced thermal protection systems (TPS) for future aerospace vehicles. SiC fiber's mechanical properties depend greatly on the presence or absence of sizing and its microstructure. In this research, silicon dioxide is found to be present on the surface of the fiber. Electron Spectroscopy for Chemical Analysis (ESCA) and Scanning Electron Microscopy (SEM) show that a thin oxide layer (SiO2) exists on the as-received fibers, and the oxide thickness increases when the fibers are exposed to high temperature. ESCA also reveals no evidence of Si-C bonding on the fiber surface on both as-received and heat treated fibers. The silicon oxide layer is thought to signal the decomposition of SiC bonds and may be partially responsible for the degradation in the breaking strength observed at temperatures above 400 C. The variation in electrical resistivity of the fibers with increasing temperature indicates a transition to a higher band gap material at 350 to 600 C. This is consistent with a decomposition of SiC involving silicon oxide formation.

  19. Optimization of thermal protection systems for the space vehicle. Volume 2: User's manual

    NASA Technical Reports Server (NTRS)

    1972-01-01

    The development of the computational techniques for the design optimization of thermal protection systems for the space shuttle vehicle are discussed. The resulting computer program was then used to perform initial optimization and sensitivity studies on a typical thermal protection system (TPS) to demonstrate its application to the space shuttle TPS design. The program was developed in FORTRAN IV for CDC 6400 computer, but it was subsequently converted to the FORTRAN V language to be used on the Univac 1108.

  20. Numerical simulations of heat and moisture transport in thermal protective clothing under flash fire conditions.

    PubMed

    Song, Guowen; Chitrphiromsri, Patirop; Ding, Dan

    2008-01-01

    A numerical model of heat and moisture transport in thermal protective clothing during exposure to a flash fire was introduced. The model was developed with the assumption that textiles are treated as porous media. The numerical model predictions were compared with experimental data from different fabric systems and configurations. Additionally, with the introduction of a skin model, the parameters that affect the performance of thermal protective clothing were investigated.

  1. Uses of Advanced Ceramic Composites in the Thermal Protection Systems of Future Space Vehicles

    NASA Technical Reports Server (NTRS)

    Rasky, Daniel J.

    1994-01-01

    Current ceramic composites being developed and characterized for use in the thermal protection systems (TPS) of future space vehicles are reviewed. The composites discussed include new tough, low density ceramic insulation's, both rigid and flexible; ultra-high temperature ceramic composites; nano-ceramics; as well as new hybrid ceramic/metallic and ceramic/organic systems. Application and advantage of these new composites to the thermal protection systems of future reusable access to space vehicles and small spacecraft is reviewed.

  2. Orion Flight Test-1 Thermal Protection System Instrumentation

    NASA Technical Reports Server (NTRS)

    Kowal, T. John

    2011-01-01

    The Orion Crew Exploration Vehicle (CEV) was originally under development to provide crew transport to the International Space Station after the retirement of the Space Shuttle, and to provide a means for the eventual return of astronauts to the Moon. With the current changes in the future direction of the United States human exploration programs, the focus of the Orion project has shifted to the project s first orbital flight test, designated Orion Flight Test 1 (OFT-1). The OFT-1 is currently planned for launch in July 2013 and will demonstrate the Orion vehicle s capability for performing missions in low Earth orbit (LEO), as well as extensibility beyond LEO for select, critical areas. Among the key flight test objectives are those related to validation of the re-entry aerodynamic and aerothermal environments, and the performance of the thermal protection system (TPS) when exposed to these environments. A specific flight test trajectory has been selected to provide a high energy entry beyond that which would be experienced during a typical low Earth orbit return, given the constraints imposed by the possible launch vehicles. This trajectory resulted from a trade study that considered the relative benefit of conflicting objectives from multiple subsystems, and sought to provide the maximum integrated benefit to the re-entry state-of-the-art. In particular, the trajectory was designed to provide: a significant, measureable radiative heat flux to the windward surface; data on boundary transition from laminar to turbulent flow; and data on catalytic heating overshoot on non-ablating TPS. In order to obtain the necessary flight test data during OFT-1, the vehicle will need to have an adequate quantity of instrumentation. A collection of instrumentation is being developed for integration in the OFT-1 TPS. In part, this instrumentation builds upon the work performed for the Mars Science Laboratory Entry, Descent and Landing Instrument (MEDLI) suite to instrument the

  3. Fabrication of titanium thermal protection system panels by the NOR-Ti-bond process

    NASA Technical Reports Server (NTRS)

    Wells, R. R.

    1971-01-01

    A method for fabricating titanium thermal protection system panels is described. The method has the potential for producing wide faying surface bonds to minimize temperature gradients and thermal stresses resulting during service at elevated temperatures. Results of nondestructive tests of the panels are presented. Concepts for improving the panel quality and for improved economy in production are discussed.

  4. Corrosion resistant thermal barrier coating. [protecting gas turbines and other engine parts

    NASA Technical Reports Server (NTRS)

    Levine, S. R.; Miller, R. A.; Hodge, P. E. (Inventor)

    1981-01-01

    A thermal barrier coating system for protecting metal surfaces at high temperature in normally corrosive environments is described. The thermal barrier coating system includes a metal alloy bond coating, the alloy containing nickel, cobalt, iron, or a combination of these metals. The system further includes a corrosion resistant thermal barrier oxide coating containing at least one alkaline earth silicate. The preferred oxides are calcium silicate, barium silicate, magnesium silicate, or combinations of these silicates.

  5. A radiant heating test facility for space shuttle orbiter thermal protection system certification

    NASA Technical Reports Server (NTRS)

    Sherborne, W. D.; Milhoan, J. D.

    1980-01-01

    A large scale radiant heating test facility was constructed so that thermal certification tests can be performed on the new generation of thermal protection systems developed for the space shuttle orbiter. This facility simulates surface thermal gradients, onorbit cold-soak temperatures down to 200 K, entry heating temperatures to 1710 K in an oxidizing environment, and the dynamic entry pressure environment. The capabilities of the facility and the development of new test equipment are presented.

  6. High temperature insulation materials for reradiative thermal protection systems

    NASA Technical Reports Server (NTRS)

    Hughes, T. A.

    1972-01-01

    Results are presented of a two year program to evaluate packaged thermal insulations for use under a metallic radiative TPS of a shuttle orbiter vehicle. Evaluations demonstrated their survival for up to 100 mission reuse cycles under shuttle acoustic and thermal loads with peak temperatures of 1000 F, 1800 F, 2000 F, 2200 F and 2500 F. The specimens were composed of low density refractory fiber felts, packaged in thin gage metal foils. In addition, studies were conducted on the venting requirements of the packages, salt spray resistance of the metal foils, and the thermal conductivity of many of the insulations as a function of temperature and ambient air pressure. Data is also presented on the radiant energy transport through insulations, and back-scattering coefficients were experimentally determined as a function of source temperature.

  7. 10 CFR 50.61 - Fracture toughness requirements for protection against pressurized thermal shock events.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 10 Energy 1 2014-01-01 2014-01-01 false Fracture toughness requirements for protection against... Construction Permits § 50.61 Fracture toughness requirements for protection against pressurized thermal shock... fracture mechanics techniques. This analysis must be submitted at least three years before RTPTS...

  8. Design and research of thermal protective material from short basalt fibres

    NASA Astrophysics Data System (ADS)

    Komkov, MA; Tarasov, VA; Boyarskaya, RA; Filimonov, AS

    2016-10-01

    Design and manufacture issues regarding highly porous thermal protection coatings of products by means of liquid filtration of short basalt fibres and mineral binder are considered. The technological process of manufacture of thermally loaded products from the short basalt fibres of thermal protective material (TPM) in the form of tiles and rings, was developed based on a liquid filtration method. The structural and mechanical properties of the highly porous TPM technological modes were determined. The thermal testing of the pipe model samples was carried out on a thermal bench, which showed the temperature on the coating reaching less than 60°C during a hot air run through the pipe at 400°C.

  9. Multiscale Modeling of Carbon/Phenolic Composite Thermal Protection Materials: Atomistic to Effective Properties

    NASA Technical Reports Server (NTRS)

    Arnold, Steven M.; Murthy, Pappu L.; Bednarcyk, Brett A.; Lawson, John W.; Monk, Joshua D.; Bauschlicher, Charles W., Jr.

    2016-01-01

    Next generation ablative thermal protection systems are expected to consist of 3D woven composite architectures. It is well known that composites can be tailored to achieve desired mechanical and thermal properties in various directions and thus can be made fit-for-purpose if the proper combination of constituent materials and microstructures can be realized. In the present work, the first, multiscale, atomistically-informed, computational analysis of mechanical and thermal properties of a present day - Carbon/Phenolic composite Thermal Protection System (TPS) material is conducted. Model results are compared to measured in-plane and out-of-plane mechanical and thermal properties to validate the computational approach. Results indicate that given sufficient microstructural fidelity, along with lowerscale, constituent properties derived from molecular dynamics simulations, accurate composite level (effective) thermo-elastic properties can be obtained. This suggests that next generation TPS properties can be accurately estimated via atomistically informed multiscale analysis.

  10. Thermal protective visor for entering high temperature areas

    NASA Technical Reports Server (NTRS)

    Burgett, F. A.

    1968-01-01

    Chamber observer suit visor protects the eyes and ears of the wearer while he is performing rescue operations during a fire. The visor is a simple curved sandwich of selected glass plates, gold coated polyester plastic film, and a dead air space, all mounted in an aluminum frame.

  11. Fabrication of prepackaged superalloy honeycomb Thermal Protection System (TPS) panels

    NASA Technical Reports Server (NTRS)

    Blair, W.; Meaney, J. E.; Rosenthal, H. A.

    1985-01-01

    High temperature materials were surveyed, and Inconel 617 and titanium were selected for application to a honeycomb TPS configuration designed to withstand 2000 F. The configuration was analyzed both thermally and structurally. Component and full-sized panels were fabricated and tested to obtain data for comparison with analysis. Results verified the panel design. Twenty five panels were delivered to NASA Langley Research Center for additional evaluation.

  12. X-33 Telemetry Best Source Selection, Processing, Display, and Simulation Model Comparison

    NASA Technical Reports Server (NTRS)

    Burkes, Darryl A.

    1998-01-01

    The X-33 program requires the use of multiple telemetry ground stations to cover the launch, ascent, transition, descent, and approach phases for the flights from Edwards AFB to landings at Dugway Proving Grounds, UT and Malmstrom AFB, MT. This paper will discuss the X-33 telemetry requirements and design, including information on fixed and mobile telemetry systems, best source selection, and support for Range Safety Officers. A best source selection system will be utilized to automatically determine the best source based on the frame synchronization status of the incoming telemetry streams. These systems will be used to select the best source at the landing sites and at NASA Dryden Flight Research Center to determine the overall best source between the launch site, intermediate sites, and landing site sources. The best source at the landing sites will be decommutated to display critical flight safety parameters for the Range Safety Officers. The overall best source will be sent to the Lockheed Martin's Operational Control Center at Edwards AFB for performance monitoring by X-33 program personnel and for monitoring of critical flight safety parameters by the primary Range Safety Officer. The real-time telemetry data (received signal strength, etc.) from each of the primary ground stations will also be compared during each nu'ssion with simulation data generated using the Dynamic Ground Station Analysis software program. An overall assessment of the accuracy of the model will occur after each mission. Acknowledgment: The work described in this paper was NASA supported through cooperative agreement NCC8-115 with Lockheed Martin Skunk Works.

  13. Flight Demonstration of X-33 Vehicle Health Management System Components on the F/A-18 Systems Research Aircraft

    NASA Technical Reports Server (NTRS)

    Schweikhard, Keith A.; Richards, W. Lance; Theisen, John; Mouyos, William; Garbos, Raymond; Schkolnik, Gerald (Technical Monitor)

    1998-01-01

    The X-33 reusable launch vehicle demonstrator has identified the need to implement a vehicle health monitoring system that can acquire data that monitors system health and performance. Sanders, a Lockheed Martin Company, has designed and developed a commercial off-the-shelf (COTS)-based open architecture system that implements a number of technologies that have not been previously used in a flight environment. NASA Dryden Flight Research Center and Sanders teamed to demonstrate that the distributed remote health nodes, fiber optic distributed strain sensor, and fiber distributed data interface communications components of the X-33 vehicle health management (VHM) system could be successfully integrated and flown on a NASA F-18 aircraft. This paper briefly describes components of X-33 VHM architecture flown at Dryden and summarizes the integration and flight demonstration of these X-33 VHM components. Finally, it presents early results from the integration and flight efforts.

  14. Flight Demonstration of X-33 Vehicle Health Management System Components on the F/A-18 Systems Research Aircraft

    NASA Technical Reports Server (NTRS)

    Schweikhard, Keith A.; Richards, W. Lance; Theisen, John; Mouyos, William; Garbos, Raymond

    2001-01-01

    The X-33 reusable launch vehicle demonstrator has identified the need to implement a vehicle health monitoring system that can acquire data that monitors system health and performance. Sanders, a Lockheed Martin Company, has designed and developed a COTS-based open architecture system that implements a number of technologies that have not been previously used in a flight environment. NASA Dryden Flight Research Center and Sanders teamed to demonstrate that the distributed remote health nodes, fiber optic distributed strain sensor, and fiber distributed data interface communications components of the X-33 vehicle health management (VHM) system could be successfully integrated and flown on a NASA F-18 aircraft. This paper briefly describes components of X-33 VHM architecture flown at Dryden and summarizes the integration and flight demonstration of these X-33 VHM components. Finally, it presents early results from the integration and flight efforts.

  15. Photographic copy of early 20” x 33”, black and white ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    Photographic copy of early 20” x 33”, black and white photograph. Located loose in oversized box at the National Museum of American History, Smithsonian Institution, Archives Center, Work and Industry Division, Washington, D.C. Original Photographer unknown. EARLY PHOTOGRAPH OF BRIDGE TAKEN FROM DOWN RIVER NEAR EAST BANK LOOKING SOUTHWEST UP RIVER TOWARD WEST BANK SHOWING STEAM LOCOMOTIVE TRAIN CROSSING BRIDGE. - Huey P. Long Bridge, Spanning Mississippi River approximately midway between nine & twelve mile points upstream from & west of New Orleans, Jefferson, Jefferson Parish, LA

  16. A Monte Carlo Dispersion Analysis of the X-33 Simulation Software

    NASA Technical Reports Server (NTRS)

    Williams, Peggy S.

    2001-01-01

    A Monte Carlo dispersion analysis has been completed on the X-33 software simulation. The simulation is based on a preliminary version of the software and is primarily used in an effort to define and refine how a Monte Carlo dispersion analysis would have been done on the final flight-ready version of the software. This report gives an overview of the processes used in the implementation of the dispersions and describes the methods used to accomplish the Monte Carlo analysis. Selected results from 1000 Monte Carlo runs are presented with suggestions for improvements in future work.

  17. Multidimensional Tests of Thermal Protection Materials in the Arcjet Test Facility

    NASA Technical Reports Server (NTRS)

    Agrawal, Parul; Ellerby, Donald T.; Switzer, Mathew R.; Squire, Thomas H.

    2010-01-01

    Many thermal protection system materials used for spacecraft heatshields have anisotropic thermal properties, causing them to display significantly different thermal characteristics in different directions, when subjected to a heating environment during flight or arcjet tests. This paper investigates the effects of sidewall heating coupled with anisotropic thermal properties of thermal protection materials in the arcjet environment. Phenolic Impregnated Carbon Ablator (PICA) and LI-2200 materials (the insulation material of Shuttle tiles) were used for this study. First, conduction-based thermal response simulations were carried out, using the Marc.Mentat finite element solver, to study the effects of sidewall heating on PICA arcjet coupons. The simulation showed that sidewall heating plays a significant role in thermal response of these models. Arcjet tests at the Aerodynamic Heating Facility (AHF) at NASA Ames Research Center were performed later on instrumented coupons to obtain temperature history at sidewall and various radial locations. The details of instrumentation and experimental technique are the prime focus of this paper. The results obtained from testing confirmed that sidewall heating plays a significant role in thermal response of these models. The test results were later used to verify the two-dimensional ablation, thermal response, and sizing program, TITAN. The test data and model predictions were found to be in excellent agreement

  18. Multidimensional Testing of Thermal Protection Materials in the Arcjet Test Facility

    NASA Technical Reports Server (NTRS)

    Agrawal, Parul; Ellerby, Donald T.; Switzer, Matt R.; Squire, Thomas Howard

    2010-01-01

    Many thermal protection system materials used for spacecraft heatshields have anisotropic thermal properties, causing them to display significantly different thermal characteristics in different directions, when subjected to a heating environment during flight or arcjet tests. The anisotropic effects are enhanced in the presence of sidewall heating. This paper investigates the effects of anisotropic thermal properties of thermal protection materials coupled with sidewall heating in the arcjet environment. Phenolic Impregnated Carbon Ablator (PICA) and LI-2200 materials (the insulation material of Shuttle tiles) were used for this study. First, conduction-based thermal response simulations were carried out, using the Marc.Mentat finite element solver, to study the effects of sidewall heating on PICA arcjet coupons. The simulation showed that sidewall heating plays a significant role in thermal response of these models. Arcjet tests at the Aerodynamic Heating Facility (AHF) at NASA Ames Research Center were performed later on instrumented coupons to obtain temperature history at sidewall and various radial locations. The details of instrumentation and experimental technique are the prime focus of this paper. The results obtained from testing confirmed that sidewall heating plays a significant role in thermal response of these models. The test results were later used to validate the two-dimensional ablation, thermal response, and sizing program, TITAN. The test data and model predictions were found to be in excellent agreement

  19. Apollo experience report: Thermal protection from engine-plume environments

    NASA Technical Reports Server (NTRS)

    Taylor, J. T.

    1972-01-01

    Portions of the combined Apollo spacecraft (the command and service module and the lunar module) are subjected to the impingement of hot exhaust gases from the various propulsion systems of the modules. The configurations of the vehicles and the sources of impinging engine plumes are described. A typical Apollo mission is outlined. Protection and design-verification methods are discussed. Finally, recommendations are made for future spacecraft programs.

  20. Vacuum chamber thermal protection for the APS (Advanced Photon Source)

    SciTech Connect

    Kramer, S.L.; Crosbie, E.A.; Kim, S.; Wehrle, R.; Yoon, M.

    1989-01-01

    The addition of undulators and wigglers into synchrotron storage rings created new problems in terms of protecting the integrity of the ring vacuum chamber. If the photon beam from these devices were missteered into striking an inadequately cooled section of the storage ring vacuum chamber, the structural strength might be reduced sufficiently that the vacuum envelope could be penetrated, resulting in long downtime of the storage ring. The new generation of high-energy synchrotron light sources will produce photon beams of such high power density that cooling of the vacuum chamber will not prevent a potential penetration of the vacuum envelope, and other methods of preventing this occurrence will be required. Since active methods will be used to ensure that the beams are delivered to beam lines for users during normal operation, there is a need for passive protection methods during non-routine operation, such as turning on new beam lines, injection, etc., when the active systems may be disabled. In addition, the passive methods could prevent the problem from arising and provide the rapid time response necessary for the highest power beams, a property that might not be easily and reliably provided by active methods during the early operation of these machines. This paper summarizes the results of a task group that studied the problem and outlines passive methods of protection for the Advanced Photon Source (APS). 2 refs., 3 figs., 1 tab.

  1. Space Shuttle Main Engine nozzle thermal protection system

    NASA Technical Reports Server (NTRS)

    Nordlund, R. M.

    1985-01-01

    Two of the three Space Shuttle Main Engine (SSME) nozzles are exposed to significant reentry aeroheating loads. To ensure reusability of the Nozzle Assembly, the nozzle primary structure must not exceed specific temperature limits. Due to the thermal, pressure, and dynamic flexing of the nozzle during a mission cycle, an appropriate insulating system must have significant flexibility. Recent missions have demonstrated nozzle reentry aeroheating rates and heat loads much higher than predictions, higher than the capability of the original insulating system. A new insulating system has been developed using similar materials in an aerodynamically 'smooth' shape to both reduce the incoming heating and increase radiation cooling.

  2. Thermal performance of an integrated thermal protection system for long-term storage of cryogenic propellants in space

    NASA Technical Reports Server (NTRS)

    Dewitt, R. L.; Boyle, R. J.

    1977-01-01

    It was demonstrated that cryogenic propellants can be stored unvented in space long enough to accomplish a Saturn orbiter mission after 1,200-day coast. The thermal design of a hydrogen-fluorine rocket stage was carried out, and the hydrogen tank, its support structure, and thermal protection system were tested in a vacuum chamber. Heat transfer rates of approximately 23 W were measured in tests to simulate the near-Earth portion of the mission. Tests to simulate the majority of the time the vehicle would be in deep space and sun-oriented resulted in a heat transfer rate of 0.11 W.

  3. Design and Testing of a Diver Thermal Protection Garment

    DTIC Science & Technology

    2008-05-01

    valid OMB control number. PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 1. REPORT DATE (D-MM- YYYY) 2 . REPORT TYPE 13. DATES COVERED (From...supply. The battery modules developed under this grant can provide protection for 8-12 hrs in cold and 2 -4 hrs in warm. The DTPS also acts as a total...Postdoctoral Fellow, I PhD in engineering, 2 Masters degrees in engineering and independent research experience for at least 6 undergraduates and 2

  4. Field repair of coated columbium Thermal Protection System (TPS)

    NASA Technical Reports Server (NTRS)

    Culp, J. D.

    1972-01-01

    The requirements for field repair of coated columbian panels were studied, and the probable cause of damage were identified. The following types of repair methods were developed, and are ready for use on an operational system: replacement of fused slurrey silicide coating by a short processing cycle using a focused radiant spot heater; repair of the coating by a glassy matrix ceramic composition which is painted or sprayed over the defective area; and repair of the protective coating by plasma spraying molybdenum disilicide over the damaged area employing portable equipment.

  5. X-33 Computational Aeroheating/Aerodynamic Predictions and Comparisons With Experimental Data

    NASA Technical Reports Server (NTRS)

    Hollis, Brian R.; Thompson, Richard A.; Berry, Scott A.; Horvath, Thomas J.; Murphy, Kelly J.; Nowak, Robert J.; Alter, Stephen J.

    2003-01-01

    This report details a computational fluid dynamics study conducted in support of the phase II development of the X-33 vehicle. Aerodynamic and aeroheating predictions were generated for the X-33 vehicle at both flight and wind-tunnel test conditions using two finite-volume, Navier-Stokes solvers. Aerodynamic computations were performed at Mach 6 and Mach 10 wind-tunnel conditions for angles of attack from 10 to 50 with body-flap deflections of 0 to 20. Additional aerodynamic computations were performed over a parametric range of free-stream conditions at Mach numbers of 4 to 10 and angles of attack from 10 to 50. Laminar and turbulent wind-tunnel aeroheating computations were performed at Mach 6 for angles of attack of 20 to 40 with body-flap deflections of 0 to 20. Aeroheating computations were performed at four flight conditions with Mach numbers of 6.6 to 8.9 and angles of attack of 10 to 40. Surface heating and pressure distributions, surface streamlines, flow field information, and aerodynamic coefficients from these computations are presented, and comparisons are made with wind-tunnel data.

  6. Computational/Experimental Aeroheating Predictions for X-33. Phase 2; Vehicle

    NASA Technical Reports Server (NTRS)

    Hamilton, H. Harris, II; Weilmuenster, K. James; Horvath, Thomas J.; Berry, Scott A.

    1998-01-01

    Laminar and turbulent heating-rate calculations from an "engineering" code and laminar calculations from a "benchmark" Navier-Stokes code are compared with experimental wind-tunnel data obtained on several candidate configurations for the X-33 Phase 2 flight vehicle. The experimental data were obtained at a Mach number of 6 and a freestream Reynolds number ranging from 1 to 8 x 10(exp 6)/ft. Comparisons are presented along the windward symmetry plane and in a circumferential direction around the body at several axial stations at angles of attack from 20 to 40 deg. The experimental results include both laminar and turbulent flow. For the highest angle of attack some of the measured heating data exhibited a "non-laminar" behavior which caused the heating to increase above the laminar level long before "classical" transition to turbulent flow was observed. This trend was not observed at the lower angles of attack. When the flow was laminar, both codes predicted the heating along the windward symmetry plane reasonably well but under-predicted the heating in the chine region. When the flow was turbulent the LATCH code accurately predicted the measured heating rates. Both codes were used to calculate heating rates over the X-33 vehicle at the peak heating point on the design trajectory and they were found to be in very good agreement over most of the vehicle windward surface.

  7. LH2 Tank Composite Coverplate Development and Flight Qualification for the X-33

    NASA Technical Reports Server (NTRS)

    Wright, Richard J.; Roule, Gerard M.

    2000-01-01

    In this paper, the development history for the first cryogenic pressurized fuel tank coverplates is presented along with a synopsis of the development strategy and technologies which led to success on this program. Coverplates are the large access panels used to access launch vehicle fuel tanks. These structures incorporate all of the requirements for a pressure vessel as well as the added requirement to mount all of the miscellaneous access points required for a fuel management system. The first composite coverplates to meet the requirements for flight qualification were developed on the X-33 program. The X-33 composite coverplates went from an open requirement to successful finished flight hardware with multiple unique configurations, complete with verification testing, in less than eighteen months. Besides the rapid development schedule, these components introduced several new technologies previously unseen in cryogenic composites including solutions to cryogenic shrinkage, self-supporting sealing surfaces, and highly loaded composite bosses with precision sealing interfaces. These components were proven to seal liquid hydrogen at cryogenic temperatures under maximum loading and pressure conditions.

  8. Evaluation of GPS Coverage for the X-33 Michael-6 Trajectory

    NASA Technical Reports Server (NTRS)

    Lundberg, John B.

    1998-01-01

    The onboard navigational system for the X-33 test flights will be based on the use of measurements collected from the Embedded Global Positioning System (GPS)/INS system. Some of the factors which will affect the quality of the GPS contribution to the navigational solution will be the number of pseudorange measurements collected at any instant in time, the distribution of the GPS satellites within the field of view, and the inherent noise level of the GPS receiver. The distribution of GPS satellites within the field of view of the receiver's antenna will depend on the receiver's position, the time of day, pointing direction of the antenna, and the effective cone angle of the antenna. The number of pseudorange measurements collected will depend upon these factors as well as the time required to lock onto a GPS satellite signal once the GPS satellite comes into the field of view of the antenna and the number of available receiver channels. The objective of this study is to evaluate the GPS coverage resulting from the proposed antenna pointing directions, the proposed antenna cone angles, and the effects due to the time of day for the X-33 Michael-6 trajectory from launch at Edwards AFB, California, to the start of the Terminal Area Energy Management (TAEM) phase on approach to Michael AAF, Utah.

  9. Mars transit vehicle thermal protection system: Issues, options, and trades

    NASA Technical Reports Server (NTRS)

    Brown, Norman

    1986-01-01

    A Mars mission is characterized by different mission phases. The thermal control of cryogenic propellant in a propulsive vehicle must withstand the different mission environments. Long term cryogenic storage may be achieved by passive or active systems. Passive cryo boiloff management features will include multilayer insulation, vapor cooled shield, and low conductance structural supports and penetrations. Active boiloff management incorporates the use of a refrigeration system. Key system trade areas include active verses passive system boiloff management (with respect to safety, reliability, and cost) and propellant tank insulation optimizations. Technology requirements include refrigeration technology advancements, insulation performance during long exposure, and cryogenic fluid transfer system for mission vehicle propellant tanking during vehicle buildip in LEO.

  10. Composite multilayer insulations for thermal protection of aerospace vehicles

    NASA Technical Reports Server (NTRS)

    Kourtides, Demetrius A.; Pitts, William C.

    1989-01-01

    Composite flexible multilayer insulation systems (MLI), consisting of alternating layers of metal foil and scrim cloth or insulation quilted together using ceramic thread, were evaluated for thermal performance and compared with a silica fibrous (baseline) insulation system. The systems studied included: (1) alternating layers of aluminoborosilicate (ABS) scrim cloth and stainless steel foil, with silica, ABS, or alumina insulation; (2) alternating layers of scrim cloth and aluminum foil, with silica or ABS insulation; (3) alternating layers of aluminum foil and silica or ABS insulation; and (4) alternating layers of aluminum-coated polyimide placed on the bottom of the silica insulation. The MLIs containing aluminum were the most efficient, measuring as little as half the backface temperature increase of the baseline system.

  11. Dust Erosion Performance of Candidate Motorcase Thermal Protection Materials.

    DTIC Science & Technology

    1980-03-10

    pascal (kPa) 1.013 25 X E +2 bar kilo pasc~l (kPa) 1.000 000 X E +2 barn meter (m) 1.000 000 X E -28 British thermal unit (t~ermochemical) joule (J) 2 2...Gy)* 1.000 000 kilotons§ terajoules 4.183 kip (1009 lbf) newton (N) 4.448 222 X E +3 kip/inch (ksi) kilo pascal (kP) 2 6.894 757 X E +3 ktap newton...inch2 newton/meter (N/m) 1.751 268 X E +2 pound-force/foot2 kilo pascal (kPa) 4.788 026 X E -2 pound-force/inch (psi) kilo pascal (kPa) 6.894 757

  12. Design of a Thermal and Micrometeorite Protection System for an Unmanned Lunar Cargo Lander

    NASA Technical Reports Server (NTRS)

    Hernandez, Carlos A.; Sunder, Sankar; Vestgaard, Baard

    1989-01-01

    The first vehicles to land on the lunar surface during the establishment phase of a lunar base will be unmanned lunar cargo landers. These landers will need to be protected against the hostile lunar environment for six to twelve months until the next manned mission arrives. The lunar environment is characterized by large temperature changes and periodic micrometeorite impacts. An automatically deployable and reconfigurable thermal and micrometeorite protection system was designed for an unmanned lunar cargo lander. The protection system is a lightweight multilayered material consisting of alternating layers of thermal and micrometeorite protection material. The protection system is packaged and stored above the lander common module. After landing, the system is deployed to cover the lander using a system of inflatable struts that are inflated using residual fuel (liquid oxygen) from the fuel tanks. Once the lander is unloaded and the protection system is no longer needed, the protection system is reconfigured as a regolith support blanket for the purpose of burying and protecting the common module, or as a lunar surface garage that can be used to sort and store lunar surface vehicles and equipment. A model showing deployment and reconfiguration of the protection system was also constructed.

  13. Micromechanical Characterization and Testing of Carbon Based Woven Thermal Protection Materials

    NASA Technical Reports Server (NTRS)

    Agrawal, Parul; Pham, John T.; Arnold, James O.; Peterson, Keith; Venkatapathy, Ethiraj

    2013-01-01

    Woven thermal protection system (TPS) materials are one of the enabling technologies for mechanically deployable hypersonic decelerator systems. These materials can be simultaneously used for thermal protection and as structural load bearing members during the entry, descent and landing operations. In order to ensure successful thermal and structural performance during the atmospheric entry, it is important to characterize the properties of these materials, once they have been subjected to entry like conditions. The present paper focuses on mechanical characteristics of pre-and post arc-jet tested woven TPS samples at different scales. It also presents the observations from scanning electron microscope and computed tomography images, and explains the changes in microstructure after being subjected to combined thermal-mechanical loading environments.

  14. Re-design and fabrication of titanium multi-wall Thermal Protection System (TPS) test panels

    NASA Technical Reports Server (NTRS)

    Blair, W.; Meaney, J. E., Jr.; Rosenthal, H. A.

    1984-01-01

    The Titanium Multi-wall Thermal Protection System (TIPS) panel was re-designed to incorporate Ti-6-2-4-2 outer sheets for the hot surface, ninety degree side closures for ease of construction and through panel fastness for ease of panel removal. Thermal and structural tests were performed to verify the design. Twenty-five panels were fabricated and delivered to NASA for evaluation at Langley Research Center and Johnson Space Center.

  15. Ceramic matrix composites - Forerunners of technological breakthrough in space vehicle hot structures and thermal protection system

    SciTech Connect

    Lacombe, A.; Rouges, J.

    1990-01-01

    The current status of carbon-carbon and carbon-silicon carbide composites developed for aerospace applications is reviewed. In particular, attention is given to production facilities and technologies for the manufacture of C-C and C-SiC composites, mechanical and thermal characteristics of carbon-carbon and carbon-silicon carbide materials, applications to thermal structures and protection, and technologies developed to build large C-SiC thermostructural components within the Hermes program. 9 refs.

  16. Liquid Oxygen Propellant Densification Production and Performance Test Results With a Large-Scale Flight-Weight Propellant Tank for the X33 RLV

    NASA Technical Reports Server (NTRS)

    Tomsik, Thomas M.; Meyer, Michael L.

    2010-01-01

    This paper describes in-detail a test program that was initiated at the Glenn Research Center (GRC) involving the cryogenic densification of liquid oxygen (LO2). A large scale LO2 propellant densification system rated for 200 gpm and sized for the X-33 LO2 propellant tank, was designed, fabricated and tested at the GRC. Multiple objectives of the test program included validation of LO2 production unit hardware and characterization of densifier performance at design and transient conditions. First, performance data is presented for an initial series of LO2 densifier screening and check-out tests using densified liquid nitrogen. The second series of tests show performance data collected during LO2 densifier test operations with liquid oxygen as the densified product fluid. An overview of LO2 X-33 tanking operations and load tests with the 20,000 gallon Structural Test Article (STA) are described. Tank loading testing and the thermal stratification that occurs inside of a flight-weight launch vehicle propellant tank were investigated. These operations involved a closed-loop recirculation process of LO2 flow through the densifier and then back into the STA. Finally, in excess of 200,000 gallons of densified LO2 at 120 oR was produced with the propellant densification unit during the demonstration program, an achievement that s never been done before in the realm of large-scale cryogenic tests.

  17. Characterization of an Integral Thermal Protection and Cryogenic Insulation Material for Advanced Space Transportation Vehicles

    NASA Technical Reports Server (NTRS)

    Salerno, L. J.; White, S. M.; Helvensteijn, B. P. M.

    2000-01-01

    NASA's planned advanced space transportation vehicles will benefit from the use of integral/conformal cryogenic propellant tanks which will reduce the launch weight and lower the earth-to-orbit costs considerably. To implement the novel concept of integral/conformal tanks requires developing an equally novel concept in thermal protection materials. Providing insulation against reentry heating and preserving propellant mass can no longer be considered separate problems to be handled by separate materials. A new family of materials, Superthermal Insulation (STI), has been conceiving and investigated by NASA's Ames Research Center to simultaneously provide both thermal protection and cryogenic insulation in a single, integral material.

  18. Simulation of Foam Impact Effects on Components of the Space Shuttle Thermal Protection System. Chapter 7

    NASA Technical Reports Server (NTRS)

    Fahrenthold, Eric P.; Park, Young-Keun

    2004-01-01

    A series of three dimensional simulations has been performed to investigate analytically the effect of insulating foam impacts on ceramic tile and reinforced carbon-carbon components of the Space Shuttle thermal protection system. The simulations employed a hybrid particle-finite element method and a parallel code developed for use in spacecraft design applications. The conclusions suggested by the numerical study are in general consistent with experiment. The results emphasize the need for additional material testing work on the dynamic mechanical response of thermal protection system materials, and additional impact experiments for use in validating computational models of impact effects.

  19. Probabilistic Design of a Mars Sample Return Earth Entry Vehicle Thermal Protection System

    NASA Technical Reports Server (NTRS)

    Dec, John A.; Mitcheltree, Robert A.

    2002-01-01

    The driving requirement for design of a Mars Sample Return mission is to assure containment of the returned samples. Designing to, and demonstrating compliance with, such a requirement requires physics based tools that establish the relationship between engineer's sizing margins and probabilities of failure. The traditional method of determining margins on ablative thermal protection systems, while conservative, provides little insight into the actual probability of an over-temperature during flight. The objective of this paper is to describe a new methodology for establishing margins on sizing the thermal protection system (TPS). Results of this Monte Carlo approach are compared with traditional methods.

  20. Design of thermal protection system for 8 foot HTST combustor

    NASA Technical Reports Server (NTRS)

    Moskowitz, S.

    1973-01-01

    The combustor in the 8-foot high temperature structures tunnel at the NASA-Langley Research Center has encountered cracking over a period of 50-250 tunnel tests within a limited range of the required operating envelope. A program was conducted which analyzed the failed combustor liner hardware and determined that the mechanism of failure was vibratory fatigue. A vibration damper system using wave springs located axially between the liner T-bar and the liner support was designed as an intermediate solution to extend the life of the current two-pass regenerative air-cooled liner. The effects of liner wall thickness, cooling air passage height, stiffener ring geometry, reflective coatings, and liner material selection were investigated for these designs. Preliminary layout design arrangements including the external water-cooling system requirements, weight estimates, installation requirements and preliminary estimates of manufacturing costs were prepared for the most promissing configurations. A state-of-the-art review of thermal barrier coatings and an evaluation of reflective coatings for the gasside surface of air-cooled liners are included.

  1. Earth Entry Requirements for Mars, Europa and Enceladus Sample Return Missions: A Thermal Protection System Perspective

    NASA Technical Reports Server (NTRS)

    Venkatapathy, Ethiraj; Gage, Peter; Ellerby, Don; Mahzari, Milad; Peterson, Keith; Stackpoole, Mairead; Young, Zion

    2016-01-01

    This oral presentation will be given at the 13th International Planetary Probe Workshop on June 14th, 2016 and will cover the drivers for reliability and the challenges faced in selecting and designing the thermal protection system (TPS). In addition, an assessment is made on new emerging TPS related technologies that could help with designs to meet the planetary protection requirements to prevent backward (Earth) contamination by biohazardous samples.

  2. Heat Shielding Characteristics and Thermostructural Performance of a Superalloy Honeycomb Sandwich Thermal Protection System (TPS)

    NASA Technical Reports Server (NTRS)

    Ko, William L.

    2004-01-01

    Heat-transfer, thermal bending, and mechanical buckling analyses have been performed on a superalloy "honeycomb" thermal protection system (TPS) for future hypersonic flight vehicles. The studies focus on the effect of honeycomb cell geometry on the TPS heat-shielding performance, honeycomb cell wall buckling characteristics, and the effect of boundary conditions on the TPS thermal bending behavior. The results of the study show that the heat-shielding performance of a TPS panel is very sensitive to change in honeycomb core depth, but insensitive to change in honeycomb cell cross-sectional shape. The thermal deformations and thermal stresses in the TPS panel are found to be very sensitive to the edge support conditions. Slight corrugation of the honeycomb cell walls can greatly increase their buckling strength.

  3. Development of X-33/X-34 Aerothermodynamic Data Bases: Lessons Learned and Future Enhancements

    NASA Technical Reports Server (NTRS)

    Miller, C. G.

    2000-01-01

    A synoptic of programmatic and technical lessons learned in the development of aerothermodynamic data bases for the X-33 and X-34 programs is presented in general terms and from the perspective of the NASA Langley Research Center Aerothermodynamics Branch. The format used is that of the "aerothermodynamic chain," the links of which are personnel, facilities, models/test articles, instrumentation, test techniques, and computational fluid dynamics (CFD). Because the aerodynamic data bases upon which the X-33 and X-34 vehicles will fly are almost exclusively from wind tunnel testing, as opposed to CFD, the primary focus of the lessons learned is on ground-based testing. The period corresponding to the development of X-33 and X-34 aerothermodynamic data bases was challenging, since a number of other such programs (e.g., X-38, X-43) competed for resources at a time of downsizing of personnel, facilities, etc., outsourcing, and role changes as NASA Centers served as subcontractors to industry. The impact of this changing environment is embedded in the lessons learned. From a technical perspective, the relatively long times to design and fabricate metallic force and moment models, delays in delivery of models, and a lack of quality assurance to determine the fidelity of model outer mold lines (OML) prior to wind tunnel testing had a major negative impact on the programs. On the positive side, the application of phosphor thermography to obtain global, quantitative heating distributions on rapidly fabricated ceramic models revolutionized the aerothermodynamic optimization of vehicle OMLs, control surfaces, etc. Vehicle designers were provided with aeroheating information prior to, or in conjunction with, aerodynamic information early in the program, thereby allowing trades to be made with both sets of input; in the past only aerodynamic data were available as input. Programmatically, failure to include transonic aerodynamic wind tunnel tests early in the assessment phase

  4. Flight Set 360L006 STS-34 field joint protection system, thermal protection system, and systems tunnel components, volume 4

    NASA Technical Reports Server (NTRS)

    Wilkinson, J. P.

    1990-01-01

    The performance of the thermal protection system, field joint protection system, and systems tunnel components of Flight Set 360L006, are documented, as evaluated by postflight hardware inspection. The condition of both motors was similar to previous flights. Sixteen aft edge hits were noted on the ground environment instrumentation thermal protection system. Each hit left a clean substrate, indicating that the damage was caused by nozzle severance debris and/or water impact. No National Space and Transporation System debris criteria for missing thermal protection system were violated. One 5.0 by 1.0 in. unbond was observed on the left hand center field joint K5NA closeout and was elevated to an in-flight anomaly (STS-34-M-4) by the NASA Ice/Debris team. Aft edge damage to the K5NA and an associated black streak indicate that burning debris from the nozzle severance system was the likely cause of the damage. Minor divots caused by debris were seen on previous flights, but this is the first occurrence of a K5NA unbond. Since the unbond occurred after booster separation there is no impact on flight safety and no corrective actions was taken. The right hand center field joint primary heater failed the dielectric withstanding voltage test after joint closeout. The heater was then disabled by opening the circuit breaker, and the redundant heater was used. The redundant heater performed nominally during the launch countdown. A similar condition occurred on Flight 4 when a secondary joint heater failed the dielectric withstanding voltage test.

  5. Heat Shield Employing Cured Thermal Protection Material Blocks Bonded in a Large-Cell Honeycomb Matrix

    NASA Technical Reports Server (NTRS)

    Zell, Peter

    2012-01-01

    A document describes a new way to integrate thermal protection materials on external surfaces of vehicles that experience the severe heating environments of atmospheric entry from space. Cured blocks of thermal protection materials are bonded into a compatible, large-cell honeycomb matrix that can be applied on the external surfaces of the vehicles. The honeycomb matrix cell size, and corresponding thermal protection material block size, is envisioned to be between 1 and 4 in. (.2.5 and 10 cm) on a side, with a depth required to protect the vehicle. The cell wall thickness is thin, between 0.01 and 0.10 in. (.0.025 and 0.25 cm). A key feature is that the honeycomb matrix is attached to the vehicle fs unprotected external surface prior to insertion of the thermal protection material blocks. The attachment integrity of the honeycomb can then be confirmed over the full range of temperature and loads that the vehicle will experience. Another key feature of the innovation is the use of uniform-sized thermal protection material blocks. This feature allows for the mass production of these blocks at a size that is convenient for quality control inspection. The honeycomb that receives the blocks must have cells with a compatible set of internal dimensions. The innovation involves the use of a faceted subsurface under the honeycomb. This provides a predictable surface with perpendicular cell walls for the majority of the blocks. Some cells will have positive tapers to accommodate mitered joints between honeycomb panels on each facet of the subsurface. These tapered cells have dimensions that may fall within the boundaries of the uniform-sized blocks.

  6. NASA Earth-to-Orbit Engineering Design Challenges: Thermal Protection Systems

    ERIC Educational Resources Information Center

    National Aeronautics and Space Administration (NASA), 2010

    2010-01-01

    National Aeronautics and Space Administration (NASA) Engineers at Marshall Space Flight Center, Dryden Flight Research Center, and their partners at other NASA centers and in private industry are currently developing X-33, a prototype to test technologies for the next generation of space transportation. This single-stage-to-orbit reusable launch…

  7. Validation of NASA Thermal Ice Protection Computer Codes. Part 1; Program Overview

    NASA Technical Reports Server (NTRS)

    Miller, Dean; Bond, Thomas; Sheldon, David; Wright, William; Langhals, Tammy; Al-Khalil, Kamel; Broughton, Howard

    1996-01-01

    The Icing Technology Branch at NASA Lewis has been involved in an effort to validate two thermal ice protection codes developed at the NASA Lewis Research Center. LEWICE/Thermal (electrothermal deicing & anti-icing), and ANTICE (hot-gas & electrothermal anti-icing). The Thermal Code Validation effort was designated as a priority during a 1994 'peer review' of the NASA Lewis Icing program, and was implemented as a cooperative effort with industry. During April 1996, the first of a series of experimental validation tests was conducted in the NASA Lewis Icing Research Tunnel(IRT). The purpose of the April 96 test was to validate the electrothermal predictive capabilities of both LEWICE/Thermal, and ANTICE. A heavily instrumented test article was designed and fabricated for this test, with the capability of simulating electrothermal de-icing and anti-icing modes of operation. Thermal measurements were then obtained over a range of test conditions, for comparison with analytical predictions. This paper will present an overview of the test, including a detailed description of: (1) the validation process; (2) test article design; (3) test matrix development; and (4) test procedures. Selected experimental results will be presented for de-icing and anti-icing modes of operation. Finally, the status of the validation effort at this point will be summarized. Detailed comparisons between analytical predictions and experimental results are contained in the following two papers: 'Validation of NASA Thermal Ice Protection Computer Codes: Part 2- The Validation of LEWICE/Thermal' and 'Validation of NASA Thermal Ice Protection Computer Codes: Part 3-The Validation of ANTICE'

  8. Altitude Effects on Thermal Ice Protection System Performance; A Study of an Alternative Simulation Approach

    NASA Technical Reports Server (NTRS)

    Addy, Gene; Wright, Bill; Orchard, David; Oleskiw, Myron

    2015-01-01

    The quest for more energy-efficient green aircraft, dictates that all systems, including the ice protection system (IPS), be closely examined for ways to reduce energy consumption and to increase efficiency. A thermal ice protection systems must protect the aircraft from the hazardous effects of icing, and yet it needs to do so as efficiently as possible. The system can no longer be afforded the degree of over-design in power usage they once were. To achieve these more exacting designs, a better understanding of the heat and mass transport phenomena involved during an icing encounter is needed.

  9. THE INFLUENCE OF THERMAL EVOLUTION IN THE MAGNETIC PROTECTION OF TERRESTRIAL PLANETS

    SciTech Connect

    Zuluaga, Jorge I.; Bustamante, Sebastian; Cuartas, Pablo A.; Hoyos, Jaime H. E-mail: sbustama@pegasus.udea.edu.co E-mail: jhhoyos@udem.edu.co

    2013-06-10

    Magnetic protection of potentially habitable planets plays a central role in determining their actual habitability and/or the chances of detecting atmospheric biosignatures. Here we develop a thermal evolution model of potentially habitable Earth-like planets and super-Earths (SEs). Using up-to-date dynamo-scaling laws, we predict the properties of core dynamo magnetic fields and study the influence of thermal evolution on their properties. The level of magnetic protection of tidally locked and unlocked planets is estimated by combining simplified models of the planetary magnetosphere and a phenomenological description of the stellar wind. Thermal evolution introduces a strong dependence of magnetic protection on planetary mass and rotation rate. Tidally locked terrestrial planets with an Earth-like composition would have early dayside magnetopause distances between 1.5 and 4.0 R{sub p} , larger than previously estimated. Unlocked planets with periods of rotation {approx}1 day are protected by magnetospheres extending between 3 and 8 R{sub p} . Our results are robust in comparison with variations in planetary bulk composition and uncertainties in other critical model parameters. For illustration purposes, the thermal evolution and magnetic protection of the potentially habitable SEs GL 581d, GJ 667Cc, and HD 40307g were also studied. Assuming an Earth-like composition, we found that the dynamos of these planets are already extinct or close to being shut down. While GL 581d is the best protected, the protection of HD 40307g cannot be reliably estimated. GJ 667Cc, even under optimistic conditions, seems to be severely exposed to the stellar wind, and, under the conditions of our model, has probably suffered massive atmospheric losses.

  10. Grid Generation Issues and CFD Simulation Accuracy for the X33 Aerothermal Simulations

    NASA Technical Reports Server (NTRS)

    Polsky, Susan; Papadopoulos, Periklis; Davies, Carol; Loomis, Mark; Prabhu, Dinesh; Langhoff, Stephen R. (Technical Monitor)

    1997-01-01

    Grid generation issues relating to the simulation of the X33 aerothermal environment using the GASP code are explored. Required grid densities and normal grid stretching are discussed with regards to predicting the fluid dynamic and heating environments with the desired accuracy. The generation of volume grids is explored and includes discussions of structured grid generation packages such as GRIDGEN, GRIDPRO and HYPGEN. Volume grid manipulation techniques for obtaining desired outer boundary and grid clustering using the OUTBOUND code are examined. The generation of the surface grid with the required surface grid with the required surface grid topology is also discussed. Utilizing grids without singular axes is explored as a method of avoiding numerical difficulties at the singular line.

  11. Reconfigurable Flight Control Designs With Application to the X-33 Vehicle

    NASA Technical Reports Server (NTRS)

    Burken, John J.; Lu, Ping; Wu, Zhenglu

    1999-01-01

    Two methods for control system reconfiguration have been investigated. The first method is a robust servomechanism control approach (optimal tracking problem) that is a generalization of the classical proportional-plus-integral control to multiple input-multiple output systems. The second method is a control-allocation approach based on a quadratic programming formulation. A globally convergent fixed-point iteration algorithm has been developed to make onboard implementation of this method feasible. These methods have been applied to reconfigurable entry flight control design for the X-33 vehicle. Examples presented demonstrate simultaneous tracking of angle-of-attack and roll angle commands during failures of the right body flap actuator. Although simulations demonstrate success of the first method in most cases, the control-allocation method appears to provide uniformly better performance in all cases.

  12. Surface Catalytic Efficiency of Advanced Carbon Carbon Candidate Thermal Protection Materials for SSTO Vehicles

    NASA Technical Reports Server (NTRS)

    Stewart, David A.

    1996-01-01

    The catalytic efficiency (atom recombination coefficients) for advanced ceramic thermal protection systems was calculated using arc-jet data. Coefficients for both oxygen and nitrogen atom recombination on the surfaces of these systems were obtained to temperatures of 1650 K. Optical and chemical stability of the candidate systems to the high energy hypersonic flow was also demonstrated during these tests.

  13. Advances in hypersonic vehicle synthesis with application to studies of advanced thermal protection system

    NASA Technical Reports Server (NTRS)

    Ardema, Mark D.

    1995-01-01

    This report summarizes the work entitled 'Advances in Hypersonic Vehicle Synthesis with Application to Studies of Advanced Thermal Protection Systems.' The effort was in two areas: (1) development of advanced methods of trajectory and propulsion system optimization; and (2) development of advanced methods of structural weight estimation. The majority of the effort was spent in the trajectory area.

  14. Design, development and test of shuttle/Centaur G-prime cryogenic tankage thermal protection systems

    NASA Technical Reports Server (NTRS)

    Knoll, Richard H.; Macneil, Peter N.; England, James E.

    1987-01-01

    The thermal protection systems for the shuttle/Centaur would have had to provide fail-safe thermal protection during prelaunch, launch ascent, and on-orbit operations as well as during potential abort. The thermal protection systems selected used a helium-purged polyimide foam beneath three rediation shields for the liquid-hydrogen tank and radiation shields only for the liquid-oxygen tank (three shields on the tank sidewall and four on the aft bulkhead). A double-walled vacuum bulkhead separated the two tanks. The liquid-hydrogen tank had one 0.75-in-thick layer of foam on the forward bulkhead and two layers on the larger area sidewall. Full scale tests of the flight vehicle in a simulated shuttle cargo bay that was purged with gaseous nitrogen gave total prelaunch heating rates of 88,500 Btu/hr and 44,000 Btu/hr for the liquid-hydrogen and -oxygen tanks, respectively. Calorimeter tests on a representative sample of the liquid-hydrogen tank sidewall thermal protection system indicated that the measured unit heating rate would rapidly decrease from the prelaunch rate of approx 100 Btu/hr/sq ft to a desired rate of less than 1.3 Btu/hr/sq ft once on orbit.

  15. Aerothermodynamic heating environment and thermal protection materials comparison for manned Mars-earth return vehicles

    NASA Technical Reports Server (NTRS)

    Henline, William D.

    1991-01-01

    The aerothermodynamic environment during the earth return portion of a specific manned Mars mission is studied. Particular attention is given to the earlier smaller crew return capsule and its thermal protection system requirements. Data are presented on the stagnation region of a generic Mars return capsule. Insulation material thicknesses required to maintain allowable structural temperatures throughout a prolonged heat soak period are estimated.

  16. Development of a Sheathed Miniature Aerothermal Reentry Thermocouple for Thermal Protection System Materials

    NASA Technical Reports Server (NTRS)

    Martinez, Edward R.; Weber, Carissa Tudryn; Oishi, Tomo; Santos, Jose; Mach, Joseph

    2011-01-01

    The Sheathed Miniature Aerothermal Reentry Thermocouple is a micro-miniature thermocouple for high temperature measurement in extreme environments. It is available for use in Thermal Protection System materials for ground testing and flight. This paper discusses the heritage, and design of the instrument. Experimental and analytical methods used to verify its performance and limitations are described.

  17. The Relationship between Physical Activity and Thermal Protective Clothing on Functional Balance in Firefighters

    ERIC Educational Resources Information Center

    Kong, Pui W.; Suyama, Joe; Cham, Rakie; Hostler, David

    2012-01-01

    We investigated the relationship between baseline physical training and the use of firefighting thermal protective clothing (TPC) with breathing apparatus on functional balance. Twenty-three male firefighters performed a functional balance test under four gear/clothing conditions. Participants were divided into groups by physical training status,…

  18. GCD TechPort Data Sheets Thermal Protection System Materials (TPSM) Project

    NASA Technical Reports Server (NTRS)

    Chinnapongse, Ronald L.

    2014-01-01

    The Thermal Protection System Materials (TPSM) Project consists of three distinct project elements: the 3-Dimensional Multifunctional Ablative Thermal Protection System (3D MAT) project element; the Conformal Ablative Thermal Protection System (CA-TPS) project element; and the Heatshield for Extreme Entry Environment Technology (HEEET) project element. 3D MAT seeks to design, develop and deliver a game changing material solution based on 3-dimensional weaving and resin infusion approach for manufacturing a material that can function as a robust structure as well as a thermal protection system. CA-TPS seeks to develop and deliver a conformal ablative material designed to be efficient and capable of withstanding peak heat flux up to 500 W/ sq cm, peak pressure up to 0.4 atm, and shear up to 500 Pa. HEEET is developing a new ablative TPS that takes advantage of state-of-the-art 3D weaving technologies and traditional manufacturing processes to infuse woven preforms with a resin, machine them to shape, and assemble them as a tiled solution on the entry vehicle substructure or heatshield.

  19. Self-Protection of Electrochemical Storage Devices via a Thermal Reversible Sol-Gel Transition.

    PubMed

    Yang, Hui; Liu, Zhiyuan; Chandran, Bevita K; Deng, Jiyang; Yu, Jiancan; Qi, Dianpeng; Li, Wenlong; Tang, Yuxin; Zhang, Chenguang; Chen, Xiaodong

    2015-10-07

    Thermal self-protected intelligent electrochemical storage devices are fabricated using a reversible sol-gel transition of the electrolyte, which can decrease the specific capacitance and increase and enable temperature-dependent charging and discharging rates in the device. This work represents proof of a simple and useful concept, which shows tremendous promise for the safe and controlled power delivery in electrochemical devices.

  20. Performance of thermal control tape in the protection of composite materials to space environmental exposure

    NASA Technical Reports Server (NTRS)

    Kamenetzky, R. R.; Whitaker, A. F.

    1992-01-01

    Thermal control tape flown on the Long Duration Exposure Facility (LDEF) experiment A0171 has shown to be effective in protecting epoxy fiberglass composites from atomic oxygen and ultraviolet degradation. The tape adhesive performed well. The aluminum, however, appeared to have become embrittled by the 5.8 years of space radiation exposure.

  1. Identifying, protecting and restoring fine-scale thermal heterogeneity in Oregon coastal streams

    EPA Science Inventory

    The functional role of thermal heterogeneity to fish in warm streams has been well recognized in the scientific literature, and is increasingly invoked as an important aspect of biodiversity conservation. Water temperature standards designed to protect cold-water taxa are also be...

  2. Identifying, Protecting, and Restoring (?) Fine-Scale Thermal Heterogeneity in Streams

    EPA Science Inventory

    The functional role of thermal heterogeneity to fish in warm streams has been well recognized in the scientific literature, and is increasingly invoked as an important aspect of biodiversity conservation. Water temperature standards designed to protect cold-water taxa are also be...

  3. Robotic system for the servicing of the orbiter thermal protection system

    NASA Technical Reports Server (NTRS)

    Graham, Todd; Bennett, Richard; Dowling, Kevin; Manouchehri, Davoud; Cooper, Eric; Cowan, Cregg

    1994-01-01

    This paper describes the design and development of a mobile robotic system to process orbiter thermal protection system (TPS) tiles. This work was justified by a TPS automation study which identified tile rewaterproofing and visual inspection as excellent applications for robotic automation.

  4. Flight Performance of an Advanced Thermal Protection Material: Toughened Uni-Piece Fibrous Insulation

    NASA Technical Reports Server (NTRS)

    Leiser, Daniel B.; Gordon, Michael P.; Rasky, Daniel J. (Technical Monitor)

    1995-01-01

    The flight performance of a new class of low density, high temperature, thermal protection materials (TPM), is described and compared to "standard" Space Shuttle TPM. This new functionally gradient material designated as Toughened Uni-Piece Fibrous Insulation (TUFI), was bonded on a removable panel attached to the base heatshield of Orbiter 105, Endeavor.

  5. Flight Performance of an Advanced Thermal Protection Material: Toughened Uni-Piece Fibrous Insulation

    NASA Technical Reports Server (NTRS)

    Leiser, Daniel B.; Gordon, Michael P.; Rasky, Daniel J. (Technical Monitor)

    1995-01-01

    The flight performance of a new class of low density, high temperature thermal protection materials (TPM) is described and compared to "standard" Space Shuttle TPM. This new functionally gradient material designated as Toughened Uni-Piece Fibrous Insulation (TUFI), was bonded on a removable panel attached to the base heat shield of Orbiter 105, Endeavour.

  6. Solid rocket booster thermal protection system materials development. [space shuttle boosters

    NASA Technical Reports Server (NTRS)

    Dean, W. G.

    1978-01-01

    A complete run log of all tests conducted in the NASA-MSFC hot gas test facility during the development of materials for the space shuttle solid rocket booster thermal protection system are presented. Lists of technical reports and drawings generated under the contract are included.

  7. Aerodynamic heating environment definition/thermal protection system selection for the HL-20

    NASA Astrophysics Data System (ADS)

    Wurster, K. E.; Stone, H. W.

    1993-09-01

    Definition of the aerothermal environment is critical to any vehicle such as the HL-20 Personnel Launch System that operates within the hypersonic flight regime. Selection of an appropriate thermal protection system design is highly dependent on the accuracy of the heating-environment prediction. It is demonstrated that the entry environment determines the thermal protection system design for this vehicle. The methods used to predict the thermal environment for the HL-20 Personnel Launch System vehicle are described. Comparisons of the engineering solutions with computational fluid dynamic predictions, as well as wind-tunnel test results, show good agreement. The aeroheating predictions over several critical regions of the vehicle, including the stagnation areas of the nose and leading edges, windward centerline and wing surfaces, and leeward surfaces, are discussed. Results of predictions based on the engineering methods found within the MINIVER aerodynamic heating code are used in conjunction with the results of the extensive wind-tunnel tests on this configuration to define a flight thermal environment. Finally, the selection of the thermal protection system based on these predictions and current technology is described.

  8. Aerodynamic heating environment definition/thermal protection system selection for the HL-20

    NASA Technical Reports Server (NTRS)

    Wurster, K. E.; Stone, H. W.

    1993-01-01

    Definition of the aerothermal environment is critical to any vehicle such as the HL-20 Personnel Launch System that operates within the hypersonic flight regime. Selection of an appropriate thermal protection system design is highly dependent on the accuracy of the heating-environment prediction. It is demonstrated that the entry environment determines the thermal protection system design for this vehicle. The methods used to predict the thermal environment for the HL-20 Personnel Launch System vehicle are described. Comparisons of the engineering solutions with computational fluid dynamic predictions, as well as wind-tunnel test results, show good agreement. The aeroheating predictions over several critical regions of the vehicle, including the stagnation areas of the nose and leading edges, windward centerline and wing surfaces, and leeward surfaces, are discussed. Results of predictions based on the engineering methods found within the MINIVER aerodynamic heating code are used in conjunction with the results of the extensive wind-tunnel tests on this configuration to define a flight thermal environment. Finally, the selection of the thermal protection system based on these predictions and current technology is described.

  9. Filler bar heating due to stepped tiles in the shuttle orbiter thermal protection system

    NASA Technical Reports Server (NTRS)

    Petley, D. H.; Smith, D. M.; Edwards, C. L. W.; Patten, A. B.; Hamilton, H. H., II

    1983-01-01

    An analytical study was performed to investigate the excessive heating in the tile to tile gaps of the Shuttle Orbiter Thermal Protection System due to stepped tiles. The excessive heating was evidence by visible discoloration and charring of the filler bar and strain isolation pad that is used in the attachment of tiles to the aluminum substrate. Two tile locations on the Shuttle orbiter were considered, one on the lower surface of the fuselage and one on the lower surface of the wing. The gap heating analysis involved the calculation of external and internal gas pressures and temperatures, internal mass flow rates, and the transient thermal response of the thermal protection system. The results of the analysis are presented for the fuselage and wing location for several step heights. The results of a study to determine the effectiveness of a half height ceramic fiber gap filler in preventing hot gas flow in the tile gaps are also presented.

  10. Lightweight Ablative and Ceramic Thermal Protection System Materials for NASA Exploration Systems Vehicles

    NASA Technical Reports Server (NTRS)

    Valentine, Peter G.; Lawrence, Timothy W.; Gubert, Michael K.; Milos, Frank S.; Kiser, James D.; Ohlhorst, Craig W.; Koenig, John R.

    2006-01-01

    As a collaborative effort among NASA Centers, the "Lightweight Nonmetallic Thermal Protection Materials Technology" Project was set up to assist mission/vehicle design trade studies, to support risk reduction in thermal protection system (TPS) material selections, to facilitate vehicle mass optimization, and to aid development of human-rated TPS qualification and certification plans. Missions performing aerocapture, aerobraking, or direct aeroentry rely on advanced heatshields that allow reductions in spacecraft mass by minimizing propellant requirements. Information will be presented on candidate materials for such reentry approaches and on screening tests conducted (material property and space environmental effects tests) to evaluate viable candidates. Seventeen materials, in three classes (ablatives, tiles, and ceramic matrix composites), were studied. In additional to physical, mechanical, and thermal property tests, high heat flux laser tests and simulated-reentry oxidation tests were performed. Space environmental effects testing, which included exposures to electrons, atomic oxygen, and hypervelocity impacts, was also conducted.

  11. A new approach to characterize the effect of fabric deformation on thermal protective performance

    NASA Astrophysics Data System (ADS)

    Li, Jun; Li, Xiaohui; Lu, Yehu; Wang, Yunyi

    2012-04-01

    It is very important to evaluate thermal protective performance (TPP) in laboratory-simulated fire scenes as accurately as possible. For this paper, to thoroughly understand the effect of fabric deformation on basic physical properties and TPP of flame-retardant fabrics exposed to flash fire, a new modified TPP testing apparatus was developed. Different extensions were employed to simulate the various extensions displayed during different body motions. The tests were also carried out with different air gaps. The results showed a significant decrease in air permeability after deformation. However, the change of thickness was slight. The fabric deformation had a complicated effect on thermal protection with different air gaps. The change of TPP depended on the balance between the surface contact area and the thermal insulation. The newly developed testing apparatus could be well employed to evaluate the effect of deformation on TPP of flame-resistant fabrics.

  12. Woven Thermal Protection System (WTPS) a Novel Approach to Meet NASA's Most Demanding Reentry Missions

    NASA Technical Reports Server (NTRS)

    Stackpoole, Mairead

    2014-01-01

    NASA's future robotic missions to Venus and outer planets, namely, Saturn, Uranus, Neptune, result in extremely high entry conditions that exceed the capabilities of current mid-density ablators (PICA or Avcoat). Therefore mission planners assume the use of a fully dense carbon phenolic heat shield similar to what was flown on Pioneer Venus and Galileo. Carbon phenolic (CP) is a robust Thermal Protection System (TPS) however its high density and thermal conductivity constrain mission planners to steep entries, high heat fluxes, pressures and short entry durations, in order for CP to be feasible from a mass perspective. The high entry conditions pose certification challenges in existing ground based test facilities. In 2012 the Game Changing Development Program in NASA's Space Technology Mission Directorate funded NASA ARC to investigate the feasibility of a Woven Thermal Protection System (WTPS) to meet the needs of NASA's most challenging entry missions. This presentation will summarize maturation of the WTPS project.

  13. Wireless Subsurface Microsensors for Health Monitoring of Thermal Protection Systems on Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Milos, Frank S.; Watters, David G.; Pallix, Joan B.; Bahr, Alfred J.; Huestis, David L.; Arnold, Jim (Technical Monitor)

    2001-01-01

    Health diagnostics is an area where major improvements have been identified for potential implementation into the design of new reusable launch vehicles in order to reduce life cycle costs, to increase safety margins, and to improve mission reliability. NASA Ames is leading the effort to develop inspection and health management technologies for thermal protection systems. This paper summarizes a joint project between NASA Ames and SRI International to develop 'SensorTags,' radio frequency identification devices coupled with event-recording sensors, that can be embedded in the thermal protection system to monitor temperature or other quantities of interest. Two prototype SensorTag designs containing thermal fuses to indicate a temperature overlimit are presented and discussed.

  14. Thermal Growth and Performance of Manganese Cobaltite Spinel Protection Layers on Ferritic Stainless Steel SOFC Interconnects

    SciTech Connect

    Yang, Zhenguo; Xia, Guanguang; Simner, Steven P.; Stevenson, Jeffry W.

    2005-08-01

    To protect solid oxide fuel cells (SOFCs) from chromium poisoning and improve metallic interconnect stability, manganese cobaltite spinel protection layers with a nominal composition of Mn1.5Co1.5O4 were thermally grown on Crofer22 APU, a ferritic stainless steel. Thermal, electrical and electrochemical investigations indicated that the spinel protection layers not only significantly decreased the contact area specific resistance (ASR) between a LSF cathode and the stainless steel interconnect, but also inhibited the sub-scale growth on the stainless steel by acting as a barrier to the inward diffusion of oxygen. A long-term thermal cycling test demonstrated excellent structural and thermomechanical stability of these spinel protection layers, which also acted as a barrier to outward chromium cation diffusion to the interconnect surface. The reduction in the contact ASR and prevention of Cr migration achieved by application of the spinel protection layers on ferritic stainless steel resulted in improved stability and electrochemical performance of SOFCs.

  15. Ablation, Thermal Response, and Chemistry Program for Analysis of Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Milos, Frank S.; Chen, Yih-Kanq

    2010-01-01

    In previous work, the authors documented the Multicomponent Ablation Thermochemistry (MAT) and Fully Implicit Ablation and Thermal response (FIAT) programs. In this work, key features from MAT and FIAT were combined to create the new Fully Implicit Ablation, Thermal response, and Chemistry (FIATC) program. FIATC is fully compatible with FIAT (version 2.5) but has expanded capabilities to compute the multispecies surface chemistry and ablation rate as part of the surface energy balance. This new methodology eliminates B' tables, provides blown species fractions as a function of time, and enables calculations that would otherwise be impractical (e.g. 4+ dimensional tables) such as pyrolysis and ablation with kinetic rates or unequal diffusion coefficients. Equations and solution procedures are presented, then representative calculations of equilibrium and finite-rate ablation in flight and ground-test environments are discussed.

  16. A Strategy for Integrating a Large Finite Element Model Using MSC NASTRAN/PATRAN: X-33 Lessons Learned

    NASA Technical Reports Server (NTRS)

    McGhee, D. S.

    1999-01-01

    The X-33 vehicle is an advanced technology demonstrator sponsored by NASA. For the past 3 years the Structural Dynamics and Loads Branch of NASA's Marshall Space Flight Center has had the task of integrating the X-33 vehicle structural finite element model. In that time, five versions of the integrated vehicle model have been produced and a strategy has evolved that would benefit anyone given the task of integrating structural finite element models that have been generated by various modelers and companies. The strategy that has been presented here consists of six decisions that need to be made: purpose of models, units, common materials list, model numbering, interface control, and archive format. This strategy has been proven and expanded from experience on the X-33 vehicle.

  17. Performance of thermal-sprayed zinc anodes treated with humectants in cathodic protection systems

    SciTech Connect

    Bullard, Sophie J.; Covino, Bernard S., Jr.; Cramer, Stephen D.; Holcomb, Gordon R.; Russell, James H.; Bennett, John E.; Milius, John K.; Cryer, Curtis B.; Soltesz, Steven M.

    2001-01-01

    Thermal-sprayed Zn anodes are used for impressed current cathodic protection (ICCP) systems in Oregon's reinforced concrete coastal bridges to minimize corrosion damage. Thermal-sprayed Zn performs well as an ICCP anode but the voltage requirement can increase with increasing electrochemical age. It also performs well as a galvanic (GCP) anode but current output can decrease with increasing electrochemical age. Past research has shown that increasing moisture at the Zn anode-concrete interface improves the operation of the thermal-sprayed Zn anode. Humectants, hygroscopic materials that are applied to the surface of the Zn-anode, can increase the moisture at the zinc-concrete interface, thereby improving the performance and extending the anode service life. Results are given for humectant-treated (LiBr and LiNO3) thermal-sprayed Zn anodes used in the laboratory electrochemical aging studies and in field studies on the Yaquina Bay Bridge, Oregon, USA.

  18. Thermal, Radiation and Impact Protective Shields (TRIPS) for Robotic and Human Space Exploration Missions

    NASA Technical Reports Server (NTRS)

    Loomis, M. P.; Arnold, J. L.

    2005-01-01

    New concepts for protective shields for NASA s Crew Exploration Vehicles (CEVs) and planetary probes offer improved mission safety and affordability. Hazards include radiation from cosmic rays and solar particle events, hypervelocity impacts from orbital debris/ micrometeorites, and the extreme heating environment experienced during entry into planetary atmospheres. The traditional approach for the design of protection systems for these hazards has been to create single-function shields, i.e. ablative and blanket-based heat shields for thermal protection systems (TPS), polymer or other low-molecular-weight materials for radiation shields, and multilayer, Whipple-type shields for protection from hypervelocity impacts. This paper introduces an approach for the development of a single, multifunctional protective shield, employing nanotechnology- based materials, to serve simultaneously as a TPS, an impact shield and as the first line of defense against radiation. The approach is first to choose low molecular weight ablative TPS materials, (existing and planned for development) and add functionalized carbon nanotubes. Together they provide both thermal and radiation (TR) shielding. Next, impact protection (IP) is furnished through a tough skin, consisting of hard, ceramic outer layers (to fracture the impactor) and sublayers of tough, nanostructured fabrics to contain the debris cloud from the impactor before it can penetrate the spacecraft s interior.

  19. Aerothermodynamics and thermal control analysis for the heat sink thermal protection system of a flyback booster

    NASA Astrophysics Data System (ADS)

    Jones, Gregory R.; Diamant, Kevin D.

    1989-01-01

    An analysis of the heat sink thermal control of a flyback booster intended for the Space Shuttle Transportation system is presented. Aeroheatng code predictions have been validated by comparison with NASA wind tunnel test data for a similarly-shaped vehicle configuration. Results for high angle-of-attack heating parameters are presented for a Reynolds number of 10 to the 6th. Skin-thickness requirements for a monocoque wall construction have been determined using titanium and aluminum as the skin material options.

  20. Protection heater design validation for the LARP magnets using thermal imaging

    SciTech Connect

    Marchevsky, M.; Turqueti, M.; Cheng, D. W.; Felice, H.; Sabbi, G.; Salmi, T.; Stenvall, A.; Chlachidze, G.; Ambrosio, G.; Ferracin, P.; Izquierdo Bermudez, S.; Perez, J. C.; Todesco, E.

    2016-03-16

    Protection heaters are essential elements of a quench protection scheme for high-field accelerator magnets. Various heater designs fabricated by LARP and CERN have been already tested in the LARP high-field quadrupole HQ and presently being built into the coils of the high-field quadrupole MQXF. In order to compare the heat flow characteristics and thermal diffusion timescales of different heater designs, we powered heaters of two different geometries in ambient conditions and imaged the resulting thermal distributions using a high-sensitivity thermal video camera. We observed a peculiar spatial periodicity in the temperature distribution maps potentially linked to the structure of the underlying cable. Two-dimensional numerical simulation of heat diffusion and spatial heat distribution have been conducted, and the results of simulation and experiment have been compared. Imaging revealed hot spots due to a current concentration around high curvature points of heater strip of varying cross sections and visualized thermal effects of various interlayer structural defects. Furthermore, thermal imaging can become a future quality control tool for the MQXF coil heaters.

  1. Protection heater design validation for the LARP magnets using thermal imaging

    DOE PAGES

    Marchevsky, M.; Turqueti, M.; Cheng, D. W.; ...

    2016-03-16

    Protection heaters are essential elements of a quench protection scheme for high-field accelerator magnets. Various heater designs fabricated by LARP and CERN have been already tested in the LARP high-field quadrupole HQ and presently being built into the coils of the high-field quadrupole MQXF. In order to compare the heat flow characteristics and thermal diffusion timescales of different heater designs, we powered heaters of two different geometries in ambient conditions and imaged the resulting thermal distributions using a high-sensitivity thermal video camera. We observed a peculiar spatial periodicity in the temperature distribution maps potentially linked to the structure of themore » underlying cable. Two-dimensional numerical simulation of heat diffusion and spatial heat distribution have been conducted, and the results of simulation and experiment have been compared. Imaging revealed hot spots due to a current concentration around high curvature points of heater strip of varying cross sections and visualized thermal effects of various interlayer structural defects. Furthermore, thermal imaging can become a future quality control tool for the MQXF coil heaters.« less

  2. Ceramic-ceramic shell tile thermal protection system and method thereof

    NASA Technical Reports Server (NTRS)

    Riccitiello, Salvatore R. (Inventor); Smith, Marnell (Inventor); Goldstein, Howard E. (Inventor); Zimmerman, Norman B. (Inventor)

    1986-01-01

    A ceramic reusable, externally applied composite thermal protection system (TPS) is proposed. The system functions by utilizing a ceramic/ceramic upper shell structure which effectively separates its primary functions as a thermal insulator and as a load carrier to transmit loads to the cold structure. The composite tile system also prevents impact damage to the atmospheric entry vehicle thermal protection system. The composite tile comprises a structurally strong upper ceramic/ceramic shell manufactured from ceramic fibers and ceramic matrix meeting the thermal and structural requirements of a tile used on a re-entry aerospace vehicle. In addition, a lightweight high temperature ceramic lower temperature base tile is used. The upper shell and lower tile are attached by means effective to withstand the extreme temperatures (3000 to 3200F) and stress conditions. The composite tile may include one or more layers of variable density rigid or flexible thermal insulation. The assembly of the overall tile is facilitated by two or more locking mechanisms on opposing sides of the overall tile assembly. The assembly may occur subsequent to the installation of the lower shell tile on the spacecraft structural skin.

  3. Atomic Oxygen Durability Evaluation of Protected Polymers Using Thermal Energy Plasma Systems

    NASA Technical Reports Server (NTRS)

    Banks, Bruce A.; Rutledge, Sharon K.; Degroh, Kim K.; Stidham, Curtis R.; Gebauer, Linda; Lamoreaux, Cynthia M.

    1995-01-01

    The durability evaluation of protected polymers intended for use in low Earth orbit (LEO) has necessitated the use of large-area, high-fluence, atomic oxygen exposure systems. Two thermal energy atomic oxygen exposure systems which are frequently used for such evaluations are radio frequency (RF) plasma ashers and electron cyclotron resonance plasma sources. Plasma source testing practices such as ample preparation, effective fluence prediction, atomic oxygen flux determination, erosion measurement, operational considerations, and erosion yield measurements are presented. Issues which influence the prediction of in-space durability based on ground laboratory thermal energy plasma system testing are also addressed.

  4. An atmosphere protection subsystem in the thermal power station automated process control system

    NASA Astrophysics Data System (ADS)

    Parchevskii, V. M.; Kislov, E. A.

    2014-03-01

    Matters concerned with development of methodical and mathematical support for an atmosphere protection subsystem in the thermal power station automated process control system are considered taking as an example the problem of controlling nitrogen oxide emissions at a gas-and-oil-fired thermal power station. The combined environmental-and-economic characteristics of boilers, which correlate the costs for suppressing emissions with the boiler steam load and mass discharge of nitrogen oxides in analytic form, are used as the main tool for optimal control. A procedure for constructing and applying environmental-and-economic characteristics on the basis of technical facilities available in modern instrumentation and control systems is presented.

  5. Characterization of Textiles Used in Chefs' Uniforms for Protection Against Thermal Hazards Encountered in the Kitchen Environment.

    PubMed

    Zhang, Han; McQueen, Rachel H; Batcheller, Jane C; Ehnes, Briana L; Paskaluk, Stephen A

    2015-10-01

    Within the kitchen the potential for burn injuries arising from contact with hot surfaces, flames, hot liquid, and steam hazards is high. The chef's uniform can potentially offer some protection against such burns by providing a protective barrier between the skin and the thermal hazard, although the extent to which can provide some protection is unknown. The purpose of this study was to examine whether fabrics used in chefs' uniforms were able to provide some protection against thermal hazards encountered in the kitchen. Fabrics from chefs' jackets and aprons were selected. Flammability of single- and multiple-layered fabrics was measured. Effect of jacket type, apron and number of layers on hot surface, hot water, and steam exposure was also measured. Findings showed that all of the jacket and apron fabrics rapidly ignited when exposed to a flame. Thermal protection against hot surfaces increased as layers increased due to more insulation. Protection against steam and hot water improved with an impermeable apron in the system. For wet thermal hazards increasing the number of permeable layers can decrease the level of protection due to stored thermal energy. As the hands and arms are most at risk of burn injury increased insulation and water-impermeable barrier in the sleeves would improve thermal protection with minimal compromise to overall thermal comfort.

  6. Recession Curve Generation for the Space Shuttle Solid Rocket Booster Thermal Protection System Coatings

    NASA Technical Reports Server (NTRS)

    Kanner, Howard S.; Stuckey, C. Irvin; Davis, Darrell W.; Davis, Darrell (Technical Monitor)

    2002-01-01

    Ablatable Thermal Protection System (TPS) coatings are used on the Space Shuttle Vehicle Solid Rocket Boosters in order to protect the aluminum structure from experiencing excessive temperatures. The methodology used to characterize the recession of such materials is outlined. Details of the tests, including the facility, test articles and test article processing are also presented. The recession rates are collapsed into an empirical power-law relation. A design curve is defined using a 95-percentile student-t distribution. based on the nominal results. Actual test results are presented for the current acreage TPS material used.

  7. Development and Verification of the Charring, Ablating Thermal Protection Implicit System Simulator

    NASA Technical Reports Server (NTRS)

    Amar, Adam J.; Calvert, Nathan; Kirk, Benjamin S.

    2011-01-01

    The development and verification of the Charring Ablating Thermal Protection Implicit System Solver (CATPISS) 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 (FEM) with first and second order fully implicit time integrators. The governing equations are fully coupled and are solved in parallel via Newton s method, while the linear system is solved via the Generalized Minimum Residual method (GMRES). Verification results from exact solutions and Method of Manufactured Solutions (MMS) are presented to show spatial and temporal orders of accuracy as well as nonlinear convergence rates.

  8. Arc-heating tests of thermal protection system materials for the H-2 rocket fairing

    NASA Astrophysics Data System (ADS)

    Hirabayashi, Noriaki; Matsuzaki, Takashi; Fukushima, Yukio; Nakamura, Tomihisa; Fujita, Takeshi

    1992-10-01

    Materials suitable for the Thermal Protection System (TPS) of the rocket fairing for the Japanese next generation launch vehicle named H-2 were tested using an arc-heated wind tunnel. The main constituents of the candidate materials were epoxy resin and silicon rubber. The heating rate was from 60 to 100 kW/(sq m) in accordance with the maximum heating rate of the nose region of an ascending H-2 rocket. Test results showed that polycondensed silicon rubber was the best candidate as the main component of the TPS, and also that a SiC coating was indispensable. In addition, based on results and a preliminary estimation of the skin temperatures of the fairing nose region, a SiC coating of about 2 mm in thickness is considered sufficient to achieve thermal protection capability.

  9. Multi-tube thermal fuse for nozzle protection from a flame holding or flashback event

    DOEpatents

    Lacy, Benjamin Paul; Davis, Jr., Lewis Berkley; Johnson, Thomas Edward; York, William David

    2012-07-03

    A protection system for a pre-mixing apparatus for a turbine engine, includes: a main body having an inlet portion, an outlet portion and an exterior wall that collectively establish a fuel delivery plenum; and a plurality of fuel mixing tubes that extend through at least a portion of the fuel delivery plenum, each of the plurality of fuel mixing tubes including at least one fuel feed opening fluidly connected to the fuel delivery plenum; at least one thermal fuse disposed on an exterior surface of at least one tube, the at least one thermal fuse including a material that will melt upon ignition of fuel within the at least one tube and cause a diversion of fuel from the fuel feed opening to at least one bypass opening. A method and a turbine engine in accordance with the protection system are also provided.

  10. Development and Verification of the Charring Ablating Thermal Protection Implicit System Solver

    NASA Technical Reports Server (NTRS)

    Amar, Adam J.; Calvert, Nathan D.; Kirk, Benjamin S.

    2010-01-01

    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.

  11. Analysis of Linear Aerospike Plume Induced X-33 Base Heating Environment

    NASA Technical Reports Server (NTRS)

    Wang, Ten-See

    1998-01-01

    Computational analysis is conducted to study the effect of an linear aerospike engine plume on the X-33 base-heating environment during ascent flight. To properly account for the freestream-body interaction and to allow for potential plume-induced flow-separation, the thermo-flowfield of the entire vehicle at several trajectory points is computed. A sequential grid-refinement technique is used in conjunction with solution-adaptive, patched, and embedded grid methods to limit the model to a manageable size. The computational methodology is based on a three-dimensional, finite-difference, viscous flow, chemically reacting, pressure-based computational fluid dynamics formulation, and a three-dimensional, finite-volume, spectral-line based weighted-sum-of-gray-gases absorption, computational radiation heat transfer formulation. The computed forebody and afterbody surface pressure coefficients and base pressure characteristic curves are compared with those of a cold-flow test. The predicted convective and radiative base-heat fluxes, the effect of base-bleed, and the potential of plume-induced flow separation are presented.

  12. Performance Tests of a Liquid Hydrogen Propellant Densification Ground System for the X33/RLV

    NASA Technical Reports Server (NTRS)

    Tomsik, Thomas M.

    1997-01-01

    A concept for improving the performance of propulsion systems in expendable and single-stage-to-orbit (SSTO) launch vehicles much like the X33/RLV has been identified. The approach is to utilize densified cryogenic liquid hydrogen (LH2) and liquid oxygen (LOX) propellants to fuel the propulsion stage. The primary benefit for using this relatively high specific impulse densified propellant mixture is the subsequent reduction of the launch vehicle gross lift-off weight. Production of densified propellants however requires specialized equipment to actively subcool both the liquid oxygen and liquid hydrogen to temperatures below their normal boiling point. A propellant densification unit based on an external thermodynamic vent principle which operates at subatmospheric pressure and supercold temperatures provides a means for the LH2 and LOX densification process to occur. To demonstrate the production concept for the densification of the liquid hydrogen propellant, a system comprised of a multistage gaseous hydrogen compressor, LH2 recirculation pumps and a cryogenic LH2 heat exchanger was designed, built and tested at the NASA Lewis Research Center (LeRC). This paper presents the design configuration of the LH2 propellant densification production hardware, analytical details and results of performance testing conducted with the hydrogen densifier Ground Support Equipment (GSE).

  13. Design and fabrication of metallic thermal protection systems for aerospace vehicles

    NASA Technical Reports Server (NTRS)

    Varisco, A.; Bell, P.; Wolter, W.

    1978-01-01

    A program was conducted to develop a lightweight, efficient metallic thermal protection system (TPS) for application to future shuttle-type reentry vehicles, advanced space transports, and hypersonic cruise vehicles. Technical requirements were generally derived from the space shuttle. A corrugation-stiffened beaded-skin TPS design was used as a baseline. The system was updated and modified to incorporate the latest technology developments and design criteria. The primary objective was to minimize mass for the total system.

  14. A Non Rigid Reusable Surface Insulation Concept for the Space Shuttle Thermal Protection System

    NASA Technical Reports Server (NTRS)

    Alexander, J. G.

    1973-01-01

    A reusable thermal protection system concept was developed for the space shuttle that utilizes a flexible, woven ceramic mat insulation beneath an aerodynamic skin and moisture barrier consisting of either a dense ceramic coating or a super alloy metallic foil. The resulting heat shield material has unique structural characteristics. The shear modulus of the woven mat is very low such that bending and membrane loads introduced into the underlying structural panel remain isolated from the surface skin.

  15. Heath Monitoring of Thermal Protection Systems - Preliminary Measurements and Design Specifications

    NASA Technical Reports Server (NTRS)

    Scott, D. A.; Price, D. C.

    2007-01-01

    The work reported here is the first stage of a project that aims to develop a health monitoring system for Thermal Protection Systems (TPS) that enables a vehicle to safely re-enter the Earth's atmosphere. The TPS health monitoring system is to be integrated into an existing acoustic emissions-based Concept Demonstrator, developed by CSIRO, which has been previously demonstrated for evaluating impact damage of aerospace systems.

  16. Analysis of gap heating due to stepped tiles in the shuttle thermal protection system

    NASA Technical Reports Server (NTRS)

    Petley, D. H.; Smith, D. M.; Edwards, C. L. W.; Carlson, A. B.

    1983-01-01

    Analytical methods used to investigate entry gap heating in the Shuttle orbiter thermal protection system are described. Analytical results are given for a fuselage lower-surface location and a wing lower-surface location. These are locations where excessive gap heating occurred on the first flight of the Shuttle. The results of a study to determine the effectiveness of a half-height ceramic fiber gap filler in preventing hot-gas flow in the tile gaps are also given.

  17. Study on the primary structural members and thermal protection system for HOPE

    NASA Astrophysics Data System (ADS)

    Kobayashi, Tomoyuki; Matsushita, T.; Yamamoto, M.; Ohnishi, T.; Atsumi, M.

    The research objective is to develop high performance main structural materials and Thermal Protection System (TPS) for the H-II Orbiting Plane (HOPE). Preliminary evaluation tests and fundamental tests of heat-resistant fiber-reinforced plastics for the HOPE body structure have been evaluated. For TPS, heat shield concept tests and preliminary evaluation tests have been conducted of high performance TPS based on Ti-alloy, Ni-alloy and advanced carbon/carbon.

  18. The Development of HfO2-Rare Earth Based Oxide Materials and Barrier Coatings for Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Zhu, Dongming; Harder, Bryan James

    2014-01-01

    Advanced hafnia-rare earth oxides, rare earth aluminates and silicates have been developed for thermal environmental barrier systems for aerospace propulsion engine and thermal protection applications. The high temperature stability, low thermal conductivity, excellent oxidation resistance and mechanical properties of these oxide material systems make them attractive and potentially viable for thermal protection systems. This paper will focus on the development of the high performance and high temperature capable ZrO2HfO2-rare earth based alloy and compound oxide materials, processed as protective coating systems using state-or-the-art processing techniques. The emphasis has been in particular placed on assessing their temperature capability, stability and suitability for advanced space vehicle entry thermal protection systems. Fundamental thermophysical and thermomechanical properties of the material systems have been investigated at high temperatures. Laser high-heat-flux testing has also been developed to validate the material systems, and demonstrating durability under space entry high heat flux conditions.

  19. Ballistic Performance of Porous Ceramic Thermal Protection Systems at 9 km/s

    NASA Technical Reports Server (NTRS)

    Miller, Joshua E.; Bohl, W. E.; Foreman, C. D.; Christiansen, Eric L.; Davis, B. A.

    2009-01-01

    Porous-ceramic, thermal-protection-systems are used heavily in current reentry vehicles like the Orbiter, and they are currently being proposed for the next generation of manned spacecraft, Orion. These materials insulate the structural components and sensitive electronic components of a spacecraft against the intense thermal environments of atmospheric reentry. Furthermore, these materials are also highly exposed to space environmental hazards like meteoroid and orbital debris impacts. This paper discusses recent impact testing up to 9 km/s on ceramic tiles similar to those used on the Orbiter. These tiles have a porous-batting of nominally 8 lb/cubic ft alumina-fiber-enhanced-thermal-barrier (AETB8) insulating material coated with a damage-resistant, toughened-unipiece-fibrous-insulation (TUFI) layer.

  20. Ballistic Performance of Porous-Ceramic, Thermal Protection Systems to 9 km/s

    NASA Technical Reports Server (NTRS)

    Miller, Joshua E.; Bohl, William E.; Foreman, Cory D.; Christiansen, Eric C.; Davis, Bruce A.

    2010-01-01

    Porous-ceramic, thermal protection systems are used heavily in current reentry vehicles like the Orbiter, and they are currently being proposed for the next generation of US manned spacecraft, Orion. These materials insulate the structural components and sensitive components of a spacecraft against the intense thermal environments of atmospheric reentry. These materials are also highly exposed to solid particle space environment hazards. This paper discusses recent impact testing up to 9.65 km/s on ceramic tiles similar to those used on the Orbiter. These tiles are a porous-ceramic insulator of nominally 8 lb/ft(exp 3) alumina-fiber-enhanced-thermal-barrier (AETB8) coated with a damage-resistant, toughened-unipiece-fibrous-insulation/reaction-cured-glass layer (TUFI/RCG).

  1. Impact of cabin environment on thermal protection system of crew hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Zhu, Xiao Wei; Zhao, Jing Quan; Zhu, Lei; Yu, Xi Kui

    2016-05-01

    Hypersonic crew vehicles need reliable thermal protection systems (TPS) to ensure their safety. Since there exists relative large temperature difference between cabin airflow and TPS structure, the TPS shield that covers the cabin is always subjected to a non-adiabatic inner boundary condition, which may influence the heat transfer characteristic of the TPS. However, previous literatures always neglected the influence of the inner boundary by assuming that it was perfectly adiabatic. The present work focuses on studying the impact of cabin environment on the thermal performance. A modified TPS model is created with a mixed thermal boundary condition to connect the cabin environment with the TPS. This helps make the simulation closer to the real situation. The results stress that cabin environment greatly influences the temperature profile inside the TPS, which should not be neglected in practice. Moreover, the TPS size can be optimized during the design procedure if taking the effect of cabin environment into account.

  2. Protection of alodine coatings from thermal aging by removable polymer coatings.

    SciTech Connect

    Wagstaff, Brett R.; Bradshaw, Robert W.; Whinnery, LeRoy L., Jr.

    2006-12-01

    Removable polymer coatings were evaluated as a means to suppress dehydration of Alodine chromate conversion coatings during thermal aging and thereby retain the corrosion protection afforded by Alodine. Two types of polymer coatings were applied to Alodine-treated panels of aluminum alloys 7075-T73 and 6061-T6 that were subsequently aged for 15 to 50 hours at temperatures between 135 F to 200 F. The corrosion resistance of the thermally aged panels was evaluated, after stripping the polymer coatings, by exposure to a standard salt-fog corrosion test and the extent of pitting of the polymer-coated and untreated panels compared. Removable polymer coatings mitigated the loss of corrosion resistance due to thermal aging experienced by the untreated alloys. An epoxide coating was more effective than a fluorosilicone coating as a dehydration barrier.

  3. The relative effects of entry parameters on thermal protection system weight. [space shuttle orbiters

    NASA Technical Reports Server (NTRS)

    Hirasaki, P. N.

    1971-01-01

    Shielding a spacecraft from the severe thermal environment of an atmospheric entry requires a sophisticated thermal protection system (TPS). Thermal computer program models were developed for two such TPS designs proposed for the space shuttle orbiter. The multilayer systems, a reusable surface insulation TPS, and a re-radiative metallic skin TPS, were sized for a cross-section of trajectories in the entry corridor. This analysis indicates the relative influence of the entry parameters on the weight of each TPS concept. The results are summarized graphically. The trajectory variables considered were down-range, cross-range, orbit inclination, entry interface velocity and flight path angle, maximum heating rate level, angle of attack, and ballistic coefficient. Variations in cross-range and flight path angle over the ranges considered had virtually no effect on the required entry TPS weight. The TPS weight was significantly more sensitive to variations in angle of attack than to dispersions in the other trajectory considered.

  4. Thermal Properties of Microstrain Gauges Used for Protection of Lithium-Ion Cells of Different Designs

    NASA Technical Reports Server (NTRS)

    Jeevarajan, Judith

    2011-01-01

    The purpose of this innovation is to use microstrain gauges to monitor minute changes in temperature along with material properties of the metal cans and pouches used in the construction of lithium-ion cells. The sensitivity of the microstrain gauges to extremely small changes in temperatures internal to the cells makes them a valuable asset in controlling the hazards in lithium-ion cells. The test program on lithium-ion cells included various cell configurations, including the pouch type configurations. The thermal properties of microstrain gauges have been found to contribute significantly as safety monitors in lithium-ion cells that are designed even with hard metal cases. Although the metal cans do not undergo changes in material property, even under worst-case unsafe conditions, the small changes in thermal properties observed during charge and discharge of the cell provide an observable change in resistance of the strain gauge. Under abusive or unsafe conditions, the change in the resistance is large. This large change is observed as a significant change in slope, and this can be used to prevent cells from going into a thermal runaway condition. For flexible metal cans or pouch-type lithium-ion cells, combinations of changes in material properties along with thermal changes can be used as an indication for the initiation of an unsafe condition. Lithium-ion cells have a very high energy density, no memory effect, and almost 100-percent efficiency of charge and discharge. However, due to the presence of a flammable electrolyte, along with the very high energy density and the capability of releasing oxygen from the cathode, these cells can go into a hazardous condition of venting, fire, and thermal runaway. Commercial lithium-ion cells have current and voltage monitoring devices that are used to control the charge and discharge of the batteries. Some lithium-ion cells have internal protective devices, but when used in multi-cell configurations, these protective

  5. Altitude Effects on Thermal Ice Protection System Performance; a Study of an Alternative Approach

    NASA Technical Reports Server (NTRS)

    Addy, Harold E., Jr.; Orchard, David; Wright, William B.; Oleskiw, Myron

    2016-01-01

    Research has been conducted to better understand the phenomena involved during operation of an aircraft's thermal ice protection system under running wet icing conditions. In such situations, supercooled water striking a thermally ice-protected surface does not fully evaporate but runs aft to a location where it freezes. The effects of altitude, in terms of air pressure and density, on the processes involved were of particular interest. Initial study results showed that the altitude effects on heat energy transfer were accurately modeled using existing methods, but water mass transport was not. Based upon those results, a new method to account for altitude effects on thermal ice protection system operation was proposed. The method employs a two-step process where heat energy and mass transport are sequentially matched, linked by matched surface temperatures. While not providing exact matching of heat and mass transport to reference conditions, the method produces a better simulation than other methods. Moreover, it does not rely on the application of empirical correction factors, but instead relies on the straightforward application of the primary physics involved. This report describes the method, shows results of testing the method, and discusses its limitations.

  6. Wireless Subsurface Sensors for Health Monitoring of Thermal Protection Systems on Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Milos, Frank S.; Arnold, Jim (Technical Monitor)

    2001-01-01

    Health diagnostics is an area where major improvements have been identified for potential implementation into the design of new reusable launch vehicles (RLVs) in order to reduce life cycle costs, to increase safety margins, and to improve mission reliability. NASA Ames is leading the effort to develop inspection and health management technologies for thermal protection systems. This paper summarizes a joint project between NASA Ames and industry partners to develop "wireless" devices that can be embedded in the thermal protection system to monitor temperature or other quantities of interest. These devices are sensors integrated with radio-frequency identification (RFID) microchips to enable non-contact communication of sensor data to an external reader that may be a hand-held scanner or a large portal. Both passive and active prototype devices have been developed. The passive device uses a thermal fuse to indicate the occurrence of excessive temperature. This device has a diameter under 0.13 cm. (suitable for placement in gaps between ceramic TPS tiles on an RLV) and can withstand 370 C for 15 minutes. The active device contains a small battery to provide power to a thermocouple for recording a temperature history during flight. The bulk of the device must be placed beneath the TPS for protection from high temperature, but the thermocouple can be placed in a hot location such as near the external surface.

  7. Thermal Aggregation of Calcium-Fortified Skim Milk Enhances Probiotic Protection during Convective Droplet Drying.

    PubMed

    Wang, Juan; Huang, Song; Fu, Nan; Jeantet, Romain; Chen, Xiao Dong

    2016-08-03

    Probiotic bacteria have been reported to confer benefits on hosts when delivered in an adequate dose. Spray-drying is expected to produce dried and microencapsulated probiotic products due to its low production cost and high energy efficiency. The bottleneck in probiotic application addresses the thermal and dehydration-related inactivation of bacteria during process. A protective drying matrix was designed by modifying skim milk with the principle of calcium-induced protein thermal aggregation. The well-defined single-droplet drying technique was used to monitor the droplet-particle conversion and the protective effect of this modified Ca-aggregated milk on Lactobacillus rhamnosus GG. The Ca-aggregated milk exhibited a higher drying efficiency and superior protection on L. rhamnosus GG during thermal convective drying. The mechanism was explained by the aggregation in milk, causing the lower binding of water in the serum phase and, conversely, local concentrated milk aggregates involved in bacteria entrapment in the course of drying. This work may open new avenues for the development of probiotic products with high bacterial viability and calcium enrichment.

  8. Development of Thermal Protection Materials for Future Mars Entry, Descent and Landing Systems

    NASA Technical Reports Server (NTRS)

    Cassell, Alan M.; Beck, Robin A. S.; Arnold, James O.; Hwang, Helen; Wright, Michael J.; Szalai, Christine E.; Blosser, Max; Poteet, Carl C.

    2010-01-01

    Entry Systems will play a crucial role as NASA develops the technologies required for Human Mars Exploration. The Exploration Technology Development Program Office established the Entry, Descent and Landing (EDL) Technology Development Project to develop Thermal Protection System (TPS) materials for insertion into future Mars Entry Systems. An assessment of current entry system technologies identified significant opportunity to improve the current state of the art in thermal protection materials in order to enable landing of heavy mass (40 mT) payloads. To accomplish this goal, the EDL Project has outlined a framework to define, develop and model the thermal protection system material concepts required to allow for the human exploration of Mars via aerocapture followed by entry. Two primary classes of ablative materials are being developed: rigid and flexible. The rigid ablatives will be applied to the acreage of a 10x30 m rigid mid L/D Aeroshell to endure the dual pulse heating (peak approx.500 W/sq cm). Likewise, flexible ablative materials are being developed for 20-30 m diameter deployable aerodynamic decelerator entry systems that could endure dual pulse heating (peak aprrox.120 W/sq cm). A technology Roadmap is presented that will be used for facilitating the maturation of both the rigid and flexible ablative materials through application of decision metrics (requirements, key performance parameters, TRL definitions, and evaluation criteria) used to assess and advance the various candidate TPS material technologies.

  9. A modernized high-pressure heater protection system for nuclear and thermal power stations

    NASA Astrophysics Data System (ADS)

    Svyatkin, F. A.; Trifonov, N. N.; Ukhanova, M. G.; Tren'kin, V. B.; Koltunov, V. A.; Borovkov, A. I.; Klyavin, O. I.

    2013-09-01

    Experience gained from operation of high-pressure heaters and their protection systems serving to exclude ingress of water into the turbine is analyzed. A formula for determining the time for which the high-pressure heater shell steam space is filled when a rupture of tubes in it occurs is analyzed, and conclusions regarding the high-pressure heater design most advisable from this point of view are drawn. A typical structure of protection from increase of water level in the shell of high-pressure heaters used in domestically produced turbines for thermal and nuclear power stations is described, and examples illustrating this structure are given. Shortcomings of components used in the existing protection systems that may lead to an accident at the power station are considered. A modernized protection system intended to exclude the above-mentioned shortcomings was developed at the NPO Central Boiler-Turbine Institute and ZioMAR Engineering Company, and the design solutions used in this system are described. A mathematical model of the protection system's main elements (the admission and check valves) has been developed with participation of specialists from the St. Petersburg Polytechnic University, and a numerical investigation of these elements is carried out. The design version of surge tanks developed by specialists of the Central Boiler-Turbine Institute for excluding false operation of the high-pressure heater protection system is proposed.

  10. Ultrahigh Temperature Ceramics for Thermal Protection of Next Generation Space Vehicles

    NASA Technical Reports Server (NTRS)

    Loehman, R. E.; Ellerby, D. T.; Gusman, M. I.; Stackpoole, M.; Johnson, S. M.; Arnold, James (Technical Monitor)

    2001-01-01

    Materials with improved properties are needed for thermal protection of next generation space vehicles. Sharp leading edges on these vehicles will have to withstand exposure to high temperatures (> 2200 C or 4000 F) and severe thermal cycling in both neutral and oxidizing environments. These extreme conditions will require materials that possess superior oxidation resistance, low creep, and excellent thermal shock properties. This presentation will first discuss the system requirements for thermal protection of advanced space vehicles and then show how they are driving development of new materials systems. The presentation will focus on ultrahigh temperature ceramics (UHTCs) that are promising candidates for such applications. ZrB2 and HfB2 and composites of those ceramics with SiC are two particular families of UHTCs that are currently under development for sharp leading edges. These ceramics are appealing because their melting temperatures are 3245 C (5873 F) for ZrB2 and 3380 C (6116 F) for HfB2 and because they may form protective, oxidation resistant coatings in use. The mechanical properties of the UHTCs are strongly dependent on phase purity and the processing route used to make them, both of which factors are being actively investigated. For example, oxide impurities could form glassy grain boundary phases that soften at high temperatures and make the ceramic susceptible to creep deformation. Results from scanning and transmission electron microscopic studies of the UHTCs have shown how their processing can be improved to give better properties. This presentation will discuss the UHTC characterization results in some detail, focusing particularly on the structure and composition of the ceramic grain boundaries. The presentation will conclude with some remarks on how the properties of these promising UHTCs can be further improved and how they might be made more economically.

  11. Ground Vibration Test Planning and Pre-Test Analysis for the X-33 Vehicle

    NASA Technical Reports Server (NTRS)

    Bedrossian, Herand; Tinker, Michael L.; Hidalgo, Homero

    2000-01-01

    This paper describes the results of the modal test planning and the pre-test analysis for the X-33 vehicle. The pre-test analysis included the selection of the target modes, selection of the sensor and shaker locations and the development of an accurate Test Analysis Model (TAM). For target mode selection, four techniques were considered, one based on the Modal Cost technique, one based on Balanced Singular Value technique, a technique known as the Root Sum Squared (RSS) method, and a Modal Kinetic Energy (MKE) approach. For selecting sensor locations, four techniques were also considered; one based on the Weighted Average Kinetic Energy (WAKE), one based on Guyan Reduction (GR), one emphasizing engineering judgment, and one based on an optimum sensor selection technique using Genetic Algorithm (GA) search technique combined with a criteria based on Hankel Singular Values (HSV's). For selecting shaker locations, four techniques were also considered; one based on the Weighted Average Driving Point Residue (WADPR), one based on engineering judgment and accessibility considerations, a frequency response method, and an optimum shaker location selection based on a GA search technique combined with a criteria based on HSV's. To evaluate the effectiveness of the proposed sensor and shaker locations for exciting the target modes, extensive numerical simulations were performed. Multivariate Mode Indicator Function (MMIF) was used to evaluate the effectiveness of each sensor & shaker set with respect to modal parameter identification. Several TAM reduction techniques were considered including, Guyan, IRS, Modal, and Hybrid. Based on a pre-test cross-orthogonality checks using various reduction techniques, a Hybrid TAM reduction technique was selected and was used for all three vehicle fuel level configurations.

  12. Development and Qualification of Alternate Blowing Agents for Space Shuttle External Tank Thermal Protection System

    NASA Technical Reports Server (NTRS)

    Williams, Charles W.; Cavalaris, James G.

    1994-01-01

    The Aerospace industry has a long history of using low density polyurethane and polyurethane-modified isocyanurate foam systems as lightweight, low cost, easily processed cryogenic Thermal Protection Systems (TPS) for ascent vehicles. The Thermal Protection System of the Space Shuttle External Tank (ET) is required so that quality liquid cryogenic propellant can be supplied to the Orbiter main engines and to protect the metal structure of the tanks from becoming too hot from aerodynamic heating, hence preventing premature break-up of the tank. These foams are all blown with CFC-1 I blowing agent which has been identified by the Environmental Protection Agency (EPA) as an ozone depleting substance. CFCs will not be manufactured after 1995, Consequently, alternate blowing agent substances must be identified and implemented to assure continued ET manufacture and delivery. This paper describes the various testing performed to select and qualify HCFC-1 41 b as a near term drop-in replacement for CFC-11. Although originally intended to be a one for one substitution in the formulation, several technical issues were identified regarding material performance and processability which required both formulation changes and special processing considerations to overcome. In order to evaluate these material changes, each material was subjected to various tests to qualify them to meet the various loads imposed on them during long term storage, pre-launch operations, launch, separation and re-entry. Each material was tested for structural, thermal, aeroshear, and stress/strain loads for the various flight environments each encounters. Details of the development and qualification program and the resolution of specific problems are discussed in this paper.

  13. Computational techniques for design optimization of thermal protective systems for the space shuttle vehicle. Volume 2: User's manual

    NASA Technical Reports Server (NTRS)

    1971-01-01

    A modular program for design optimization of thermal protection systems is discussed. Its capabilities and limitations are reviewed. Instructions for the operation of the program, output, and the program itself are given.

  14. Characterization of thermally sprayed coatings for high-temperature wear-protection applications

    SciTech Connect

    Li, C.C.

    1980-03-01

    Under normal high-temperature gas-cooled reactor (HTGR) operating conditions, faying surfaces of metallic components under high contact pressure are prone to friction, wear, and self-welding damage. Component design calls for coatings for the protection of the mating surfaces. Anticipated operating temperatures up to 850 to 950/sup 0/C (1562 to 1742/sup 0/F) and a 40-y design life require coatings with excellent thermal stability and adequate wear and spallation resistance, and they must be compatible with the HTGR coolant helium environment. Plasma and detonation-gun (D-gun) deposited chromium carbide-base and stabilized zirconia coatings are under consideration for wear protection of reactor components such as the thermal barrier, heat exchangers, control rods, and turbomachinery. Programs are under way to address the structural integrity, helium compatibility, and tribological behavior of relevant sprayed coatings. In this paper, the need for protection of critical metallic components and the criteria for selection of coatings are discussed. The technical background to coating development and the experience with the steam cycle HTGR (HTGR-SC) are commented upon. Coating characterization techniques employed at General Atomic Company (GA) are presented, and the progress of the experimental programs is briefly reviewed. In characterizing the coatings for HTGR applications, it is concluded that a systems approach to establish correlation between coating process parameters and coating microstructural and tribological properties for design consideration is required.

  15. Thermal tolerance affects mutualist attendance in an ant-plant protection mutualism.

    PubMed

    Fitzpatrick, Ginny; Lanan, Michele C; Bronstein, Judith L

    2014-09-01

    Mutualism is an often complex interaction among multiple species, each of which may respond differently to abiotic conditions. The effects of temperature on the formation, dissolution, and success of these and other species interactions remain poorly understood. We studied the thermal ecology of the mutualism between the cactus Ferocactus wislizeni and its ant defenders (Forelius pruinosus, Crematogaster opuntiae, Solenopsis aurea, and Solenopsis xyloni) in the Sonoran Desert, USA. The ants are attracted to extrafloral nectar produced by the plants and, in exchange, protect the plants from herbivores; there is a hierarchy of mutualist effectiveness based on aggression toward herbivores. We determined the relationship between temperature and ant activity on plants, the thermal tolerance of each ant species, and ant activity in relation to the thermal environment of plants. Temperature played a role in determining which species interact as mutualists. Three of the four ant species abandoned the plants during the hottest part of the day (up to 40 °C), returning when surface temperature began to decrease in the afternoon. The least effective ant mutualist, F. pruinosus, had a significantly higher critical thermal maximum than the other three species, was active across the entire range of plant surface temperatures observed (13.8-57.0 °C), and visited plants that reached the highest temperatures. F. pruinosus occupied some plants full-time and invaded plants occupied by more dominant species when those species were thermally excluded. Combining data on thermal tolerance and mutualist effectiveness provides a potentially powerful tool for predicting the effects of temperature on mutualisms and mutualistic species.

  16. Thermal tolerance affects mutualist attendance in an ant-plant protection mutualism

    PubMed Central

    Fitzpatrick, Ginny; Lanan, Michele C.; Bronstein, Judith L.

    2014-01-01

    Mutualism is an often-complex interaction among multiple species, each of which may respond differently to abiotic conditions. The effects of temperature on the formation, dissolution, and success of these and other species interactions remain poorly understood. We studied the thermal ecology of the mutualism between the cactus Ferocactus wislizeni and its ant defenders (Forelius pruinosus, Crematogaster opuntiae, Solenopsis aurea, and Solenopsis xyloni) in the Sonoran Desert, USA. The ants are attracted to extrafloral nectar produced by the plants and in exchange protect the plants from herbivores; there is a hierarchy of mutualist effectiveness based on aggression toward herbivores. We determined the relationship between temperature and ant activity on plants, the thermal tolerance of each ant species, and ant activity in relation to the thermal environment of plants. Temperature played a role in determining which species interact as mutualists. Three of the four ant species abandoned the plants during the hottest part of the day (up to 40°C), returning when surface temperature began to decrease in the afternoon. The least effective ant mutualist, F. pruinosus, had a significantly higher critical thermal maximum than the other three species, was active across the entire range of plant surface temperatures observed (13.8-57.0°C), and visited plants that reached the highest temperatures. F. pruinosus occupied some plants full-time and invaded plants occupied by more dominant species when those species were thermally excluded. Combining data on thermal tolerance and mutualist effectiveness provides a potentially powerful tool for predicting the effects of temperature on mutualisms and mutualistic species. PMID:25012597

  17. Fiber-optic temperature profiling for thermal protection system heat shields

    NASA Astrophysics Data System (ADS)

    Black, Richard J.; Costa, Joannes M.; Zarnescu, Livia; Hackney, Drew A.; Moslehi, Behzad; Peters, Kara J.

    2016-11-01

    To achieve better designs for spacecraft heat shields for missions requiring atmospheric aero-capture or entry/reentry, reliable thermal protection system (TPS) sensors are needed. Such sensors will provide both risk reduction and heat-shield mass minimization, which will facilitate more missions and enable increased payloads and returns. This paper discusses TPS thermal measurements provided by a temperature monitoring system involving lightweight, electromagnetic interference-immune, high-temperature resistant fiber Bragg grating (FBG) sensors with a thermal mass near that of TPS materials together with fast FBG sensor interrogation. Such fiber-optic sensing technology is highly sensitive and accurate, as well as suitable for high-volume production. Multiple sensing FBGs can be fabricated as arrays on a single fiber for simplified design and reduced cost. Experimental results are provided to demonstrate the temperature monitoring system using multisensor FBG arrays embedded in a small-size super-light ablator (SLA) coupon which was thermally loaded to temperatures in the vicinity of the SLA charring temperature. In addition, a high-temperature FBG array was fabricated and tested for 1000°C operation, and the temperature dependence considered over the full range (cryogenic to high temperature) for which silica fiber FBGs have been subjected.

  18. Thermal Protection of the Huygens Probe During Titan Entry: Last Questions

    NASA Technical Reports Server (NTRS)

    Bouilly, Jean-Marc

    2005-01-01

    CASSINI-HUYGENS mission is a cooperation between NASA and ESA, dedicated to the exploration of the Saturnian system. In the framework of this mission, the entry of the HUYGENS probe in the atmosphere of TITAN will be of major scientific interest. One of the essential points of the HUYGENS mission is therefore the good behavior of the thermal shield designed to maintain the aerodynamic shape and to protect the probe from excessive heating during the atmospheric entry on TITAN. The design and the qualification of this thermal shield were carried out between 1992 and 1995 (development phase). Currently, the final definition of mission parameters is being completed. As the performance of the thermal shield is one of all the parameters considered at system level, it is therefore necessary to reassess the thermal response of the TPS, taking into account some updated information that was not yet available during the development phase. After some recall of the results of 1992 to 1995, the paper will present a status of the current work on TPS.

  19. X-33 (Rev-F) Aeroheating Results of Test 6770 in NASA Langley 20-Inch Mach 6 Air Tunnel

    NASA Technical Reports Server (NTRS)

    Berry, Scott A.; Horvath, Thomas J.; Kowalkowski, Matthew K.; Liechty, Derek S.

    1999-01-01

    Aeroheating characteristics of the X-33 Rev-F configuration have been experimentally examined in the Langley 20-Inch Mach 6 Air Tunnel (Test 6770). Global surface heat transfer distributions, surface streamline patterns, and shock shapes were measured on a 0.013-scale model at Mach 6 in air. Parametric variations include angles-of-attack of 20-deg, 30-deg, and 40-deg; Reynolds numbers based on model length of 0.9 to 4.9 million; and body-flap deflections of 10-deg and 20-deg. The effects of discrete roughness elements on boundary layer transition, which included trip height, size, and location, both on and off the windward centerline, were investigated. This document is intended to serve as a quick release of preliminary data to the X-33 program; analysis is limited to observations of the experimental trends in order to expedite dissemination.

  20. Sea buckthorn seed oil protects against the oxidative stress produced by thermally oxidized lipids.

    PubMed

    Zeb, Alam; Ullah, Sana

    2015-11-01

    Thermally oxidized vegetable ghee was fed to the rabbits for 14 days with specific doses of sea buckthorn seed oil (SO). The ghee and SO were characterized for quality parameters and fatty acid composition using GC-MS. Rabbits serum lipid profile, hematology and histology were investigated. Major fatty acids were palmitic acid (44%) and oleic acid (46%) in ghee, while SO contains oleic acid (56.4%) and linoleic acid (18.7%). Results showed that oxidized vegetable ghee increases the serum total cholesterol, LDL-cholesterols, triglycerides and decrease the serum glucose. Oxidized ghee produced toxic effects in the liver and hematological parameters. Sea buckthorn oil supplementation significantly lowered the serum LDL-cholesterols, triglycerides and increased serum glucose and body weight of the animals. Sea buckthorn oil was found to reduce the toxic effects and degenerative changes in the liver and thus provides protection against the thermally oxidized lipids induced oxidative stress.

  1. Study of heat sink thermal protection systems for hypersonic research aircraft

    NASA Technical Reports Server (NTRS)

    Vahl, W. A.; Edwards, C. L. W.

    1978-01-01

    The feasibility of using a single metallic heat sink thermal protection system (TPS) over a projected flight test program for a hypersonic research vehicle was studied using transient thermal analyses and mission performance calculations. Four materials, aluminum, titanium, Lockalloy, and beryllium, as well as several combinations, were evaluated. Influence of trajectory parameters were considered on TPS and mission performance for both the clean vehicle configuration as well as with an experimental scramjet mounted. From this study it was concluded that a metallic heat sink TPS can be effectively employed for a hypersonic research airplane flight envelope which includes dash missions in excess of Mach 8 and 60 seconds of cruise at Mach numbers greater than 6. For best heat sink TPS match over the flight envelope, Lockalloy and titanium appear to be the most promising candidates

  2. Woven Thermal Protection System (WTPS) a Novel Approach to Meet Nasa's Most Demanding Reentry Missions

    NASA Technical Reports Server (NTRS)

    Stackpoole, Margaret M.; Ellerby, Donald T.; Gasch, Matt; Ventkatapathy, Ethiraj; Beerman, Adam; Boghozian, Tane; Gonzales, Gregory; Feldman, Jay; Peterson, Keith; Prabhu, Dinesh

    2014-01-01

    NASA's future robotic missions to Venus and other planets, namely, Saturn, Uranus, Neptune, result in extremely high entry conditions that exceed the capabilities of current mid density ablators (PICA or Avcoat). Therefore mission planners assume the use of a fully dense carbon phenolic heatshield similar to what was flown on Pioneer Venus and Galileo. Carbon phenolic is a robust TPS, however, its high density and thermal conductivity constrain mission planners to steep entries, high fluxes, pressures and short entry durations, in order for CP to be feasible from a mass perspective. The high entry conditions pose certification challenges in existing ground based test facilities. In 2012 the Game Changing Development Program in NASA's Space Technology Mission Directorate funded NASA ARC to investigate the feasibility of a Woven Thermal Protection System to meet the needs of NASA's most challenging entry missions. This presentation will summarize the maturation of the WTPS project.

  3. Evaluation of nondestructive testing techniques for the space shuttle nonmetallic thermal protection system

    NASA Technical Reports Server (NTRS)

    Tiede, D. A.

    1972-01-01

    A program was conducted to evaluate nondestructive analysis techniques for the detection of defects in rigidized surface insulation (a candidate material for the Space Shuttle thermal protection system). Uncoated, coated, and coated and bonded samples with internal defects (voids, cracks, delaminations, density variations, and moisture content), coating defects (holes, cracks, thickness variations, and loss of adhesion), and bondline defects (voids and unbonds) were inspected by X-ray radiography, acoustic, microwave, high-frequency ultrasonic, beta backscatter, thermal, holographic, and visual techniques. The detectability of each type of defect was determined for each technique (when applicable). A possible relationship between microwave reflection measurements (or X-ray-radiography density measurements) and the tensile strength was established. A possible approach for in-process inspection using a combination of X-ray radiography, acoustic, microwave, and holographic techniques was recommended.

  4. Monitoring of Thermal Protection Systems Using Robust Self-Organizing Optical Fiber Sensing Networks

    NASA Technical Reports Server (NTRS)

    Richards, Lance

    2013-01-01

    The general aim of this work is to develop and demonstrate a prototype structural health monitoring system for thermal protection systems that incorporates piezoelectric acoustic emission (AE) sensors to detect the occurrence and location of damaging impacts, and an optical fiber Bragg grating (FBG) sensor network to evaluate the effect of detected damage on the thermal conductivity of the TPS material. Following detection of an impact, the TPS would be exposed to a heat source, possibly the sun, and the temperature distribution on the inner surface in the vicinity of the impact measured by the FBG network. A similar procedure could also be carried out as a screening test immediately prior to re-entry. The implications of any detected anomalies in the measured temperature distribution will be evaluated for their significance in relation to the performance of the TPS during re-entry. Such a robust TPS health monitoring system would ensure overall crew safety throughout the mission, especially during reentry

  5. Evaluation of Advanced Thermal Protection Techniques for Future Reusable Launch Vehicles

    NASA Technical Reports Server (NTRS)

    Olds, John R.; Cowart, Kris

    2001-01-01

    A method for integrating Aeroheating analysis into conceptual reusable launch vehicle RLV design is presented in this thesis. This process allows for faster turn-around time to converge a RLV design through the advent of designing an optimized thermal protection system (TPS). It consists of the coupling and automation of four computer software packages: MINIVER, TPSX, TCAT and ADS. MINIVER is an Aeroheating code that produces centerline radiation equilibrium temperatures, convective heating rates, and heat loads over simplified vehicle geometries. These include flat plates and swept cylinders that model wings and leading edges, respectively. TPSX is a NASA Ames material properties database that is available on the World Wide Web. The newly developed Thermal Calculation Analysis Tool (TCAT) uses finite difference methods to carry out a transient in-depth I-D conduction analysis over the center mold line of the vehicle. This is used along with the Automated Design Synthesis (ADS) code to correctly size the vehicle's thermal protection system JPS). The numerical optimizer ADS uses algorithms that solve constrained and unconstrained design problems. The resulting outputs for this process are TPS material types, unit thicknesses, and acreage percentages. TCAT was developed for several purposes. First, it provides a means to calculate the transient in-depth conduction seen by the surface of the TPS material that protects a vehicle during ascent and reentry. Along with the in-depth conduction, radiation from the surface of the material is calculated along with the temperatures at the backface and interior parts of the TPS material. Secondly, TCAT contributes added speed and automation to the overall design process. Another motivation in the development of TCAT is optimization.

  6. CFD Analysis of Flexible Thermal Protection System Shear Configuration Testing in the LCAT Facility

    NASA Technical Reports Server (NTRS)

    Ferlemann, Paul G.

    2014-01-01

    This paper documents results of computational analysis performed after flexible thermal protection system shear configuration testing in the LCAT facility. The primary objectives were to predict the shear force on the sample and the sensitivity of all surface properties to the shape of the sample. Bumps of 0.05, 0.10,and 0.15 inches were created to approximate the shape of some fabric samples during testing. A large amount of information was extracted from the CFD solutions for comparison between runs and also current or future flight simulations.

  7. Substructure procedure for including tile flexibility in stress analysis of shuttle thermal protection system

    NASA Technical Reports Server (NTRS)

    Giles, G. L.

    1980-01-01

    A substructure procedure to include the flexibility of the tile in the stress analysis of the shuttle thermal protection system (TPS) is described. In this procedure, the TPS is divided into substructures of (1) the tile which is modeled by linear finite elements and (2) the SIP which is modeled as a nonlinear continuum. This procedure was applied for loading cases of uniform pressure, uniform moment, and an aerodynamic shock on various tile thicknesses. The ratios of through-the-thickness stresses in the SIP which were calculated using a flexible tile compared to using a rigid tile were found to be less than 1.05 for the cases considered.

  8. Position Paper External Tank Thermal Protection System (TPS) Manually Sprayed fly-as-is Foam Certification

    NASA Technical Reports Server (NTRS)

    Stadler, John H.

    2009-01-01

    During manufacture of the existing External Tanks (ETs), the Thermal Protection System (TPS) foam manual spray application processes lacked the enhanced controls/procedures to ensure that defects produced were less than the critical size. Therefore the only remaining option to certify the "fly-as-is" foam is to verify ET120 tank hardware meets the new foam debris requirements. The ET project has undertaken a significant effort studying the existing "fly-as-is" TPS foam. This paper contains the findings of the study.

  9. Dynamic and static modeling of the shuttle orbiter's thermal protection system

    NASA Technical Reports Server (NTRS)

    Housner, J. M.; Giles, G. L.; Vallas, M.

    1982-01-01

    The dynamic and static analysis methods used to model the nonlinear structural behavior of the Shuttle Orbiter's tile/pad thermal protection system are discussed. The structural evaluation of the tile/pad system is complicated by the nonlinear stiffening, hysteresis and viscosity exhibited by the pad material. Application of the analysis to square tiles subject to sinusoidal and random excitation is presented along with appropriate test data. Correlation is considered good. In order to treat the stress analysis of thousands of individual tiles, a nonlinear static analysis was developed which utilizes equivalent static loads derived from the dynamic environment. Tensile stress at the bondline is examined in thousands of unique tiles.

  10. Manned Evaluation of the NCSC Diver Thermal Protection (DTP) Passive System Prototype.

    DTIC Science & Technology

    1979-08-01

    AD-AOB 820 NAVY EXPERIMENTAL DIVING UNIT PANAMA CITY FL FIG 6/17 MANNED EVALUATION OF THE NCSC DIVER THERMAL PROTECTION (DTP) PA--ETC(U) AUG 79 C A...PIANTADOSI, D J SALL, M L NUCKOLS UNCLASSIFIED NEDU-13-79 NLrrnuuuu/ulbin IInunulnllnl. mmII NAVY EXPERIMENTAL DIVING UrNT~ s0 3 .10 098 DEPARTMENT OF...THE NAVY NAVY EXPERIMENTAL DIVING UNIT Panama City, Florida 32407 NAVY EXPERIMENTAL DIVING UNIT REPORT NO. 13-79 - ANNED EVALUATION OF THE NCSC DIVER

  11. Moisture absorption characteristics of the Orbiter thermal protection system and methods used to prevent water ingestion

    NASA Technical Reports Server (NTRS)

    Schomburg, C.; Dotts, R. L.; Tillian, D. J.

    1983-01-01

    The Space Shuttle Orbiter's silica tile Thermal Protection System (TPS) is beset by the moisture absorption problems inherently associated with low density, highly porous insulation systems. Attention is presently given to the comparative success of methods for the minimization and/or prevention of water ingestion by the TPS tiles, covering the development of water-repellent agents and their tile application techniques, flight test program results, and materials improvements. The use of external films for rewaterproofing of the TPS tiles after each mission have demonstrated marginal to unacceptable performance. By contrast, a tile interior waterproofing agent has shown promise.

  12. Effect of load eccentricity and substructure deformation on ultimate strength of shuttle orbiter thermal protection system

    NASA Technical Reports Server (NTRS)

    Sawyer, J. W.

    1981-01-01

    The effect of load eccentricity and substructure deformation on the ultimate strength and stress displacement properties of the shuttle orbiter thermal protection system (TPS) was determined. The LI-900 Reusable Surface Insulation (RSI) tiles mounted on the .41 cm thick Strain Isolator Pad (SIP) were investigated. Substructure deformations reduce the ultimate strength of the SIP/tile TPS and increase the scatter in the ultimate strength data. Substructure deformations that occur unsymmetric to the tile can cause the tile to rotate when subjected to a uniform applied load. Load eccentricity reduces SIP/tile TPS ultimate strength and causes tile rotation.

  13. Effect of surface catalysis on heating to ceramic coated thermal protection systems for transatmospheric vehicles

    NASA Technical Reports Server (NTRS)

    Stewart, David A.; Kolodziej, Paul; Henline, William D.; Pincha, Elizabeth M. W.

    1988-01-01

    This paper describes the effect of surface catalysis on the heat transfer rate to the heat shield of a typical Transatmospheric Vehicle (TAV) during ascent and atmospheric entry. Surface kinetics and optical properties obtained from arc-jet tests on candidate thermal protection systems (coated metals) were used in a reacting boundary layer code to estimate the heating distribution along the surface of a TAV. Thermochemical stability of the coatings is described in terms of reduction in emittance and loss of opacifiers from the coatings during the arc-jet tests.

  14. Techniques for aerothermal tests of large, flightweight thermal protection panels in a Mach 7 wind tunnel

    NASA Technical Reports Server (NTRS)

    Deveikis, W. D.; Bruce, W. E., Jr.; Karns, J. R.

    1974-01-01

    Thermal performance and structural integrity are experimentally evaluated in the Langley 8-ft high temperature structures tunnel, which uses a combustion products test medium to provide realistic combinations of aerodynamic heating and loading. Recently developed techniques provide independent control of rate and magnitude of surface heating and differential pressure, protection against adverse acoustics buffeting during facility starting and stopping, programed radiant heating before exposing test panels to the high energy stream, and infrared radiometry for detailed surface temperatures. These techniques were verified repeatedly by return of useful data on metallic and nonmetallic panel concepts of reusable surface insulation.

  15. Development of 3D Woven Ablative Thermal Protection Systems (TPS) for NASA Spacecraft

    NASA Technical Reports Server (NTRS)

    Feldman, Jay D.; Ellerby, Don; Stackpoole, Mairead; Peterson, Keith; Venkatapathy, Ethiraj

    2015-01-01

    The development of a new class of thermal protection system (TPS) materials known as 3D Woven TPS led by the Entry Systems and Technology Division of NASA Ames Research Center (ARC) will be discussed. This effort utilizes 3D weaving and resin infusion technologies to produce heat shield materials that are engineered and optimized for specific missions and requirements. A wide range of architectures and compositions have been produced and preliminarily tested to prove the viability and tailorability of the 3D weaving approach to TPS.

  16. Refurbishment cost study of the thermal protection system of a space shuttle vehicle, phase 2

    NASA Technical Reports Server (NTRS)

    Haas, D. W.

    1972-01-01

    The labor costs and techniques associated with the refurbishment and maintenance of representative thermal protection system (TPS) components and their attachment concepts suitable for space shuttle application are defined, characterized, and evaluated from the results of an experimental test program. This program consisted of designing selected TPS concepts, fabricating and assembling test hardware, and performing a time and motion study of specific maintenance functions of the test hardware on a full-scale- mockup. Labor requirements and refurbishment techniques, as they relate to the maintenance functions of inspection, repair, removal, and replacement were identified.

  17. Measured catalycities of various candidate space shuttle thermal protection system coatings at low temperatures

    NASA Technical Reports Server (NTRS)

    Scott, C. D.

    1973-01-01

    Atom recombination catalytic rates for surface coatings of various candidate thermal protection system materials for the space shuttle vehicle were obtained from measurements in arc jet, air flow. The coatings, chrome oxides, siliconized carbon/carbon, hafnium/tantalum carbide on carbon/carbon, and niobium silicide, were bonded to the sensitive surface of transient slug calorimeters that measured the heat transfer rates to the coatings. The catalytic rates were inferred from these heat transfer rates Surface temperatures of the calorimeters varied from approximately 300 to 410 K.

  18. Conversion of hydrocarbon fuel in thermal protection reactors of hypersonic aircraft

    NASA Astrophysics Data System (ADS)

    Kuranov, A. L.; Mikhaylov, A. M.; Korabelnikov, A. V.

    2016-07-01

    Thermal protection of heat-stressed surfaces of a high-speed vehicle flying in dense layers of atmosphere is one of the topical issues. Not of a less importance is also the problem of hydrocarbon fuel combustion in a supersonic air flow. In the concept under development, it is supposed that in the most high-stressed parts of airframe and engine, catalytic thermochemical reactors will be installed, wherein highly endothermic processes of steam conversion of hydrocarbon fuel take place. Simultaneously with heat absorption, hydrogen generation will occur in the reactors. This paper presents the results of a study of conversion of hydrocarbon fuel in a slit reactor.

  19. Thermal Protection Materials and Systems: Where Have We Been, Where are We Going?

    NASA Technical Reports Server (NTRS)

    Johnson, Sylvia M.

    2016-01-01

    Thermal protection materials and systems (TPS) have been critical to fulfilling humankind's desire to explore space. Composite and ceramic materials have enable the early missions to orbit, the moon, the space station, Mars with robots, and sample return. Crewed missions to Mars are being considered, and this places even more demands on TPS materials. This talk will give some history on the materials used for earth and planetary entry and the demands placed upon such materials. TPs needs for future missions, especially to Mars, will be identified and potential solutions discussed.

  20. Development and flight qualification of the C-SiC thermal protection systems for the IXV

    NASA Astrophysics Data System (ADS)

    Buffenoir, François; Zeppa, Céline; Pichon, Thierry; Girard, Florent

    2016-07-01

    The Intermediate experimental Vehicle (IXV) atmospheric re-entry demonstrator, developed within the FLPP (Future Launcher Preparatory Programme) and funded by ESA, aimed at developing a demonstration vehicle that gave Europe a unique opportunity to increase its knowledge in the field of advanced atmospheric re-entry technologies. A key technology that has been demonstrated in real conditions through the flight of this ambitious vehicle is the thermal protection system (TPS) of the Vehicle. Within this programme, HERAKLES, Safran Group, has been in charge of the TPS of the windward and nose assemblies of the vehicle, and has developed and manufactured SepcarbInox® ceramic matrix composite (CMC) protection systems that provided a high temperature resistant non ablative outer mould line (OML) for enhanced aerodynamic control. The design and flight justification of these TPS has been achieved through extensive analysis and testing:

  1. Validation of NASA Thermal Ice Protection Computer Codes. Part 3; The Validation of Antice

    NASA Technical Reports Server (NTRS)

    Al-Khalil, Kamel M.; Horvath, Charles; Miller, Dean R.; Wright, William B.

    2001-01-01

    An experimental program was generated by the Icing Technology Branch at NASA Glenn Research Center to validate two ice protection simulation codes: (1) LEWICE/Thermal for transient electrothermal de-icing and anti-icing simulations, and (2) ANTICE for steady state hot gas and electrothermal anti-icing simulations. An electrothermal ice protection system was designed and constructed integral to a 36 inch chord NACA0012 airfoil. The model was fully instrumented with thermo-couples, RTD'S, and heat flux gages. Tests were conducted at several icing environmental conditions during a two week period at the NASA Glenn Icing Research Tunnel. Experimental results of running-wet and evaporative cases were compared to the ANTICE computer code predictions and are presented in this paper.

  2. Personal, closed-cycle cooling and protective apparatus and thermal battery therefor

    DOEpatents

    Klett, James W.; Klett, Lynn B.

    2004-07-20

    A closed-cycle apparatus for cooling a living body includes a heat pickup body or garment which permits evaporation of an evaporating fluid, transmission of the vapor to a condenser, and return of the condensate to the heat pickup body. A thermal battery cooling source is provided for removing heat from the condenser. The apparatus requires no external power and provides a cooling system for soldiers, race car drivers, police officers, firefighters, bomb squad technicians, and other personnel who may utilize protective clothing to work in hostile environments. An additional shield layer may simultaneously provide protection from discomfort, illness or injury due to harmful atmospheres, projectiles, edged weapons, impacts, explosions, heat, poisons, microbes, corrosive agents, or radiation, while simultaneously removing body heat from the wearer.

  3. Measuring the spectral emissivity of thermal protection materials during atmospheric reentry simulation

    NASA Technical Reports Server (NTRS)

    Marble, Elizabeth

    1996-01-01

    Hypersonic spacecraft reentering the earth's atmosphere encounter extreme heat due to atmospheric friction. Thermal Protection System (TPS) materials shield the craft from this searing heat, which can reach temperatures of 2900 F. Various thermophysical and optical properties of TPS materials are tested at the Johnson Space Center Atmospheric Reentry Materials and Structures Evaluation Facility, which has the capability to simulate critical environmental conditions associated with entry into the earth's atmosphere. Emissivity is an optical property that determines how well a material will reradiate incident heat back into the atmosphere upon reentry, thus protecting the spacecraft from the intense frictional heat. This report describes a method of measuring TPS emissivities using the SR5000 Scanning Spectroradiometer, and includes system characteristics, sample data, and operational procedures developed for arc-jet applications.

  4. Revitalizing the Space Shuttle's Thermal Protection System with Reverse Engineering and 3D Vision Technology

    NASA Technical Reports Server (NTRS)

    Wilson, Brad; Galatzer, Yishai

    2008-01-01

    The Space Shuttle is protected by a Thermal Protection System (TPS) made of tens of thousands of individually shaped heat protection tile. With every flight, tiles are damaged on take-off and return to earth. After each mission, the heat tiles must be fixed or replaced depending on the level of damage. As part of the return to flight mission, the TPS requirements are more stringent, leading to a significant increase in heat tile replacements. The replacement operation requires scanning tile cavities, and in some cases the actual tiles. The 3D scan data is used to reverse engineer each tile into a precise CAD model, which in turn, is exported to a CAM system for the manufacture of the heat protection tile. Scanning is performed while other activities are going on in the shuttle processing facility. Many technicians work simultaneously on the space shuttle structure, which results in structural movements and vibrations. This paper will cover a portable, ultra-fast data acquisition approach used to scan surfaces in this unstable environment.

  5. Quantitative assessment of the relationship between radiant heat exposure and protective performance of multilayer thermal protective clothing during dry and wet conditions.

    PubMed

    Fu, M; Weng, W G; Yuan, H Y

    2014-07-15

    The beneficial effect of clothing on a person is important to the criteria for people exposure to radiant heat flux from fires. The thermal protective performance of multilayer thermal protective clothing exposed to low heat fluxes during dry and wet conditions was studied using two designed bench-scale test apparatus. The protective clothing with four fabric layers (outer shell, moisture barrier, thermal linear and inner layer) was exposed to six levels of thermal radiation (1, 2, 3, 5, 7 and 10kW/m(2)). Two kinds of the moisture barrier (PTFE and GoreTex) with different vapor permeability were compared. The outside and inside surface temperatures of each fabric layer were measured. The fitting analysis was used to quantitatively assess the relationship between the temperature of each layer during thermal exposure and the level of external heat flux. It is indicated that there is a linear correlation between the temperature of each layer and the radiant level. Therefore, a predicted equation is developed to calculate the thermal insulation of the multilayer clothing from the external heat flux. It can also provide some useful information on the beneficial effects of clothing for the exposure criteria of radiant heat flux from fire.

  6. A novel approach for fit analysis of thermal protective clothing using three-dimensional body scanning.

    PubMed

    Lu, Yehu; Song, Guowen; Li, Jun

    2014-11-01

    The garment fit played an important role in protective performance, comfort and mobility. The purpose of this study is to quantify the air gap to quantitatively characterize a three-dimensional (3-D) garment fit using a 3-D body scanning technique. A method for processing of scanned data was developed to investigate the air gap size and distribution between the clothing and human body. The mesh model formed from nude and clothed body was aligned, superimposed and sectioned using Rapidform software. The air gap size and distribution over the body surface were analyzed. The total air volume was also calculated. The effects of fabric properties and garment size on air gap distribution were explored. The results indicated that average air gap of the fit clothing was around 25-30 mm and the overall air gap distribution was similar. The air gap was unevenly distributed over the body and it was strongly associated with the body parts, fabric properties and garment size. The research will help understand the overall clothing fit and its association with protection, thermal and movement comfort, and provide guidelines for clothing engineers to improve thermal performance and reduce physiological burden.

  7. Design of Inorganic Water Repellent Coatings for Thermal Protection Insulation on an Aerospace Vehicle

    NASA Technical Reports Server (NTRS)

    Fuerstenau, D. W.; Ravikumar, R.

    1997-01-01

    In this report, thin film deposition of one of the model candidate materials for use as water repellent coating on the thermal protection systems (TPS) of an aerospace vehicle was investigated. The material tested was boron nitride (BN), the water-repellent properties of which was detailed in our other investigation. Two different methods, chemical vapor deposition (CVD) and pulsed laser deposition (PLD), were used to prepare the BN films on a fused quartz substrate (one of the components of thermal protection systems on aerospace vehicles). The deposited films were characterized by a variety of techniques including X-ray diffraction, X-ray photoelectron spectroscopy, and scanning electron microscopy. The BN films were observed to be amorphous in nature, and a CVD-deposited film yielded a contact angle of 60 degrees with water, similar to the pellet BN samples investigated previously. This demonstrates that it is possible to use the bulk sample wetting properties as a guideline to determine the candidate waterproofing material for the TPS.

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

    NASA Technical Reports Server (NTRS)

    Brewer, Amy R.

    2011-01-01

    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.

  9. The Langley thermal protection system test facility: A description including design operating boundaries

    NASA Technical Reports Server (NTRS)

    Klich, G. F.

    1976-01-01

    A description of the Langley thermal protection system test facility is presented. This facility was designed to provide realistic environments and times for testing thermal protection systems proposed for use on high speed vehicles such as the space shuttle. Products from the combustion of methane-air-oxygen mixtures, having a maximum total enthalpy of 10.3 MJ/kg, are used as a test medium. Test panels with maximum dimensions of 61 cm x 91.4 cm are mounted in the side wall of the test region. Static pressures in the test region can range from .005 to .1 atm and calculated equilibrium temperatures of test panels range from 700 K to 1700 K. Test times can be as long as 1800 sec. Some experimental data obtained while using combustion products of methane-air mixtures are compared with theory, and calibration of the facility is being continued to verify calculated values of parameters which are within the design operating boundaries.

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

    NASA Technical Reports Server (NTRS)

    Gorton, Mark P.; Shideler, John L.; Webb, Granville L.

    1993-01-01

    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.

  11. Ballistic Performance Model of Crater Formation in Monolithic, Porous Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Miller, J. E.; Christiansen, E. L.; Deighton, K. D.

    2014-01-01

    Porous monolithic ablative systems insulate atmospheric reentry vehicles from reentry plasmas generated by atmospheric braking from orbital and exo-orbital velocities. Due to the necessity that these materials create a temperature gradient up to several thousand Kelvin over their thickness, it is important that these materials are near their pristine state prior to reentry. These materials may also be on exposed surfaces to space environment threats like orbital debris and meteoroids leaving a probability that these exposed surfaces will be below their prescribed values. Owing to the typical small size of impact craters in these materials, the local flow fields over these craters and the ablative process afford some margin in thermal protection designs for these locally reduced performance values. In this work, tests to develop ballistic performance models for thermal protection materials typical of those being used on Orion are discussed. A density profile as a function of depth of a typical monolithic ablator and substructure system is shown in Figure 1a.

  12. A Collaborative Analysis Tool for Thermal Protection Systems for Single Stage to Orbit Launch Vehicles

    NASA Technical Reports Server (NTRS)

    Alexander, Reginald; Stanley, Thomas Troy

    2001-01-01

    Presented is a design tool and process that connects several disciplines which are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to all other systems, as is the case with SSTO vehicles with air breathing propulsion, which is currently being studied by the National Aeronautics and Space Administration (NASA). In particular, the thermal protection system (TPS) is linked directly to almost every major system. The propulsion system pushes the vehicle to velocities on the order of 15 times the speed of sound in the atmosphere before pulling up to go to orbit which results in high temperatures on the external surfaces of the vehicle. Thermal protection systems to maintain the structural integrity of the vehicle must be able to mitigate the heat transfer to the structure and be lightweight. Herein lies the interdependency, in that as the vehicle's speed increases, the TPS requirements are increased. And as TPS masses increase the effect on the propulsion system and all other systems is compounded. To adequately calculate the TPS mass of this type of vehicle several engineering disciplines and analytical tools must be used preferably in an environment that data is easily transferred and multiple iterations are easily facilitated.

  13. Development of thermal protection system of the MUSES-C/DASH reentry capsule

    NASA Astrophysics Data System (ADS)

    Yamada, Tetsuya; Inatani, Yoshifumi; Honda, Masahisa; Hirai, Ken'ich

    2002-07-01

    In the final phase of the MUSES-C mission, a small capsule with asteroid sample conducts reentry flight directly from the interplanetary transfer orbit at the velocity over 12 km/s. The severe heat flux, the complicated functional requirements, and small weight budget impose several engineering challenges on the designing of the thermal protection system of the capsule. The heat shield is required to function not only as ablator but also as a structural component. The cloth-layered carbon-phenolic ablator, which has higher allowable stress, is developed in newly-devised fabric method for avoiding delamination due to the high aerodynamic heating. The ablation analysis code, which takes into account of the effect of pyrolysis gas on the surface recession rate, has been developed and verified in the arc-heating tests in the facility environment of broad range of enthalpy level. The capsule was designed to be ventilated during the reentry flight up to about atmospheric pressure by the time of parachute deployment by being sealed with porous flow-restrict material. The designing of the thermal protection system, the hardware specifications, and the ground-based test programs of both MUSES-C and DASH capsule are summarized and discussed here in this paper.

  14. Development of Metallic Thermal Protection Systems for the Reusable Launch Vehicle

    NASA Technical Reports Server (NTRS)

    Blosser, Max L.

    1996-01-01

    A reusable Thermal Protection System (TPS) that is not only lightweight, but durable, operable and cost effective is one of the technologies required by the Reusable Launch Vehicle (RLV) to achieve the goal of drastically reducing the cost of delivering payload to orbit. Metallic TPS is one of the systems being developed to meet this challenge. Current efforts involve improving the superalloy honeycomb TPS concept, which consists of a foil-gage metallic box encapsulating a low density fibrous insulation, and evaluating it for RLV requirements. The superalloy honeycomb TPS concept is mechanically attached to the vehicle structure. Improvements include more efficient internal insulation, a simpler, lighter weight configuration, and a quick-release fastener system for easier installation and removal. Evaluation includes thermal and structural analysis, fabrication and testing of both coupons and TPS panels under conditions simulating RLV environments. Coupons of metallic honeycomb sandwich, representative of the outer TPS surface, were subjected to low speed impact, hypervelocity impact, and rain erosion testing as well as subsequent arcjet exposure. Arrays of TPS panels have been subjected to radiant heating in a thermal/vacuum facility, aerodynamic heating in an arcjet facility and acoustic loading.

  15. Consumable and non-consumable thermal spray anodes for impressed current cathodic protection of reinforced concrete structures

    SciTech Connect

    Covino, B.S. Jr.; Cramer, S.D.; Bullard, Sophie J.; Holcomb, Gordon R.; Collins, Wesley K.; McGill, G.E.

    1998-01-01

    A comparison is presented of some of the differences between thermal spray Zn, a consumable anode, and catalyzed thermal spray Ti, a non-consumable anode, used for impressed current cathodic protection of reinforced concrete structures. The thermal spray process for both Ti and Zn is compared using the spray parameters, atomizing gases, spray rate, and cost. The thermal spray Ti and Zn coatings are compared in terms of physical properties, composition, and structure. Results of accelerated laboratory experiments are presented and comparisons between Ti and Zn are made on the effect of electrochemical aging on voltage requirements, bond strength, coating resistivity, water permeability, and anode-concrete interracial composition.

  16. On-Line Thermal Barrier Coating Monitoring for Real-Time Failure Protection and Life Maximization

    SciTech Connect

    Dennis H. LeMieux

    2004-10-01

    Under the sponsorship of the U. S. Department of Energy's National Energy Laboratory, Siemens Westinghouse Power Corporation proposes a four year program titled, ''On-Line Thermal Barrier Coating (TBC) Monitor for Real-Time Failure Protection and Life Maximization'', to develop, build and install the first generation of an on-line TBC monitoring system for use on land -based advanced gas turbines (AGT). Federal deregulation in electric power generation has accelerated power plant owner's demand for improved reliability availability maintainability (RAM) of the land-based advanced gas turbines. As a result, firing temperatures have been increased substantially in the advanced turbine engines, and the TBCs have been developed for maximum protection and life of all critical engine components operating at these higher temperatures. Losing TBC protection can therefore accelerate the degradation of substrate components materials and eventually lead to a premature failure of critical component and costly unscheduled power outages. This program seeks to substantially improve the operating life of high cost gas turbine components using TBC; thereby, lowering the cost of maintenance leading to lower cost of electricity. Siemens Westinghouse Power Corporation has teamed with Indigo Systems; a supplier of state-of-the-art infrared camera systems, and Wayne State University, a leading research organization.

  17. ON-LINE THERMAL BARRIER COATING MONITORING FOR REAL-TIME FAILURE PROTECTION AND LIFE MAXIMIZATION

    SciTech Connect

    Dennis H. LeMieux

    2002-04-01

    Under the sponsorship of the U. S. Department of Energy's National Energy Laboratory, Siemens Westinghouse Power Corporation proposes a four year program titled, ''On-Line Thermal Barrier Coating (TBC) Monitor for Real-Time Failure Protection and Life Maximization,'' to develop, build and install the first generation of an on-line TBC monitoring system for use on land-based advanced gas turbines (AGT). Federal deregulation in electric power generation has accelerated power plant owner's demand for improved reliability availability maintainability (RAM) of the land-based advanced gas turbines. As a result, firing temperatures have been increased substantially in the advanced turbine engines, and the TBCs have been developed for maximum protection and life of all critical engine components operating at these higher temperatures. Losing TBC protection can therefore accelerate the degradation of substrate components materials and eventually lead to a premature failure of critical component and costly unscheduled power outages. This program seeks to substantially improve the operating life of high cost gas turbine components using TBC; thereby, lowering the cost of maintenance leading to lower cost of electricity. Siemens Westinghouse Power Corporation has teamed with Indigo Systems, a supplier of state-of-the-art infrared camera systems, and Wayne State University, a leading research organization in the field of infrared non-destructive examination (NDE), to complete the program.

  18. On-Line Thermal Barrier Coating Monitoring for Real-Time Failure Protection and Life Maximization

    SciTech Connect

    Dennis H. LeMieux

    2005-04-01

    Under the sponsorship of the U. S. Department of Energy's National Energy Laboratory, Siemens Westinghouse Power Corporation proposes a four year program titled, ''On-Line Thermal Barrier Coating (TBC) Monitor for Real-Time Failure Protection and Life Maximization'', to develop, build and install the first generation of an on-line TBC monitoring system for use on land-based advanced gas turbines (AGT). Federal deregulation in electric power generation has accelerated power plant owner's demand for improved reliability availability maintainability (RAM) of the land-based advanced gas turbines. As a result, firing temperatures have been increased substantially in the advanced turbine engines, and the TBCs have been developed for maximum protection and life of all critical engine components operating at these higher temperatures. Losing TBC protection can therefore accelerate the degradation of substrate components materials and eventually lead to a premature failure of critical component and costly unscheduled power outages. This program seeks to substantially improve the operating life of high cost gas turbine components using TBC; thereby, lowering the cost of maintenance leading to lower cost of electricity. Siemens Westinghouse Power Corporation has teamed with Indigo Systems, a supplier of state-of-the-art infrared camera systems, and Wayne State University, a leading research organization in the field of infrared non-destructive examination (NDE), to complete the program.

  19. ON-LINE THERMAL BARRIER COATING MONITORING FOR REAL-TIME FAILURE PROTECTION AND LIFE MAXIMIZATION

    SciTech Connect

    Dennis H. LeMieux

    2003-07-01

    Under the sponsorship of the U. S. Department of Energy's National Energy Laboratory, Siemens Westinghouse Power Corporation proposes a four year program titled, ''On-Line Thermal Barrier Coating (TBC) Monitor for Real-Time Failure Protection and Life Maximization,'' to develop, build and install the first generation of an on-line TBC monitoring system for use on land-based advanced gas turbines (AGT). Federal deregulation in electric power generation has accelerated power plant owner's demand for improved reliability, availability, and maintainability (RAM) of the land-based advanced gas turbines. As a result, firing temperatures have been increased substantially in the advanced turbine engines, and the TBCs have been developed for maximum protection and life of all critical engine components operating at these higher temperatures. Losing TBC protection can, therefore, accelerate the degradation of substrate component materials and eventually lead to a premature failure of critical components and costly unscheduled power outages. This program seeks to substantially improve the operating life of high cost gas turbine components using TBC; thereby, lowering the cost of maintenance leading to lower cost of electricity. Siemens Westinghouse Power Corporation has teamed with Indigo Systems, a supplier of state-of-the-art infrared camera systems, and Wayne State University, a leading research organization in the field of infrared non-destructive examination (NDE), to complete the program.

  20. ON-LINE THERMAL BARRIER COATING MONITORING FOR REAL-TIME FAILURE PROTECTION AND LIFE MAXIMIZATION

    SciTech Connect

    Dennis H. LeMieux

    2003-10-01

    Under the sponsorship of the U. S. Department of Energy's National Energy Laboratory, Siemens Westinghouse Power Corporation proposes a four year program titled, ''On-Line Thermal Barrier Coating (TBC) Monitor for Real-Time Failure Protection and Life Maximization,'' to develop, build and install the first generation of an on-line TBC monitoring system for use on land-based advanced gas turbines (AGT). Federal deregulation in electric power generation has accelerated power plant owner's demand for improved reliability, availability, and maintainability (RAM) of the land-based advanced gas turbines. As a result, firing temperatures have been increased substantially in the advanced turbine engines, and the TBCs have been developed for maximum protection and life of all critical engine components operating at these higher temperatures. Losing TBC protection can, therefore, accelerate the degradation of substrate component materials and eventually lead to a premature failure of critical components and costly unscheduled power outages. This program seeks to substantially improve the operating life of high cost gas turbine components using TBC; thereby, lowering the cost of maintenance leading to lower cost of electricity. Siemens Westinghouse Power Corporation has teamed with Indigo Systems, a supplier of state-of-the-art infrared camera systems, and Wayne State University, a leading research organization in the field of infrared non-destructive examination (NDE), to complete the program.

  1. International Space Station (ISS) Soyuz Vehicle Descent Module Evaluation of Thermal Protection System (TPS) Penetration Characteristics

    NASA Technical Reports Server (NTRS)

    Davis, Bruce A.; Christiansen, Eric L.; Lear, Dana M.; Prior, Tom

    2013-01-01

    The descent module (DM) of the ISS Soyuz vehicle is covered by thermal protection system (TPS) materials that provide protection from heating conditions experienced during reentry. Damage and penetration of these materials by micrometeoroid and orbital debris (MMOD) impacts could result in loss of vehicle during return phases of the mission. The descent module heat shield has relatively thick TPS and is protected by the instrument-service module. The TPS materials on the conical sides of the descent module (referred to as backshell in this test plan) are exposed to more MMOD impacts and are relatively thin compared to the heat shield. This test program provides hypervelocity impact (HVI) data on materials similar in composition and density to the Soyuz TPS on the backshell of the vehicle. Data from this test program was used to update ballistic limit equations used in Soyuz TPS penetration risk assessments. The impact testing was coordinated by the NASA Johnson Space Center (JSC) Hypervelocity Impact Technology (HVIT) Group [1] in Houston, Texas. The HVI testing was conducted at the NASA-JSC White Sands Hypervelocity Impact Test Facility (WSTF) at Las Cruces, New Mexico. Figure

  2. On-Line Thermal Barrier Coating Monitoring for Real-Time Failure Protection and Life Maximization

    SciTech Connect

    Dennis H. LeMieux

    2005-10-01

    Under the sponsorship of the U. S. Department of Energy's National Energy Laboratory, Siemens Power Generation, Inc proposed a four year program titled, ''On-Line Thermal Barrier Coating (TBC) Monitor for Real-Time Failure Protection and Life Maximization'', to develop, build and install the first generation of an on-line TBC monitoring system for use on land-based advanced gas turbines (AGT). Federal deregulation in electric power generation has accelerated power plant owner's demand for improved reliability availability maintainability (RAM) of the land-based advanced gas turbines. As a result, firing temperatures have been increased substantially in the advanced turbine engines, and the TBCs have been developed for maximum protection and life of all critical engine components operating at these higher temperatures. Losing TBC protection can therefore accelerate the degradation of substrate components materials and eventually lead to a premature failure of critical component and costly unscheduled power outages. This program seeks to substantially improve the operating life of high cost gas turbine components using TBC; thereby, lowering the cost of maintenance leading to lower cost of electricity. Siemens Power Generation, Inc. has teamed with Indigo Systems, a supplier of state-of-the-art infrared camera systems, and Wayne State University, a leading research organization in the field of infrared non-destructive examination (NDE), to complete the program.

  3. Fracture Toughness Evaluation of Space Shuttle External Tank Thermal Protection System Polyurethane Foam Insulation Materials

    NASA Technical Reports Server (NTRS)

    McGill, Preston; Wells, Doug; Morgan, Kristin

    2006-01-01

    Experimental evaluation of the basic fracture properties of Thermal Protection System (TPS) polyurethane foam insulation materials was conducted to validate the methodology used in estimating critical defect sizes in TPS applications on the Space Shuttle External Fuel Tank. The polyurethane foam found on the External Tank (ET) is manufactured by mixing liquid constituents and allowing them to react and expand upwards - a process which creates component cells that are generally elongated in the foam rise direction and gives rise to mechanical anisotropy. Similarly, the application of successive foam layers to the ET produces cohesive foam interfaces (knitlines) which may lead to local variations in mechanical properties. This study reports the fracture toughness of BX-265, NCFI 24-124, and PDL-1034 closed-cell polyurethane foam as a function of ambient and cryogenic temperatures and knitline/cellular orientation at ambient pressure.

  4. Prediction methods of skin burn for performance evaluation of thermal protective clothing.

    PubMed

    Zhai, Li-Na; Li, Jun

    2015-11-01

    Most test methods use skin burn prediction to evaluate the thermal protective performance of clothing. In this paper, we reviewed different burn prediction methods used in clothing evaluation. The empirical criterion and the mathematical model were analyzed in detail as well as their relationship and limitations. Using an empirical criterion, the onset of skin burn is determined by the accumulated skin surface energy in certain periods. On the other hand, the mathematical model, which indicates denatured collagen, is more complex, which involves a heat transfer model and a burn model. Further studies should be conducted to examine the situations where the prediction methods are derived. New technologies may be used in the future to explore precise or suitable prediction methods for both flash fire tests and increasingly lower-intensity tests.

  5. Heat gain from thermal radiation through protective clothing with different insulation, reflectivity and vapour permeability.

    PubMed

    Bröde, Peter; Kuklane, Kalev; Candas, Victor; Den Hartog, Emiel A; Griefahn, Barbara; Holmér, Ingvar; Meinander, Harriet; Nocker, Wolfgang; Richards, Mark; Havenith, George

    2010-01-01

    The heat transferred through protective clothing under long wave radiation compared to a reference condition without radiant stress was determined in thermal manikin experiments. The influence of clothing insulation and reflectivity, and the interaction with wind and wet underclothing were considered. Garments with different outer materials and colours and additionally an aluminised reflective suit were combined with different number and types of dry and pre-wetted underwear layers. Under radiant stress, whole body heat loss decreased, i.e., heat gain occurred compared to the reference. This heat gain increased with radiation intensity, and decreased with air velocity and clothing insulation. Except for the reflective outer layer that showed only minimal heat gain over the whole range of radiation intensities, the influence of the outer garments' material and colour was small with dry clothing. Wetting the underclothing for simulating sweat accumulation, however, caused differing effects with higher heat gain in less permeable garments.

  6. High Temperature Damping Behavior of Plasma-Sprayed Thermal Barrier and Protective Coatings

    NASA Technical Reports Server (NTRS)

    Zhu, Dongming; Miller, Robert A.; Duffy, Kirsten P.; Ghosn, Louis J.

    2010-01-01

    A high temperature damping test apparatus has been developed using a high heat flux CO 2 laser rig in conjunction with a TIRA S540 25 kHz Shaker and Polytec OFV 5000 Vibrometer system. The test rig has been successfully used to determine the damping performance of metallic and ceramic protective coating systems at high temperature for turbine engine applications. The initial work has been primarily focused on the microstructure and processing effects on the coating temperature-dependence damping behavior. Advanced ceramic coatings, including multicomponent tetragonal and cubic phase thermal barrier coatings, along with composite bond coats, have also been investigated. The coating high temperature damping mechanisms will also be discussed.

  7. Micromechanical analysis and design of an integrated thermal protection system for future space vehicles

    NASA Astrophysics Data System (ADS)

    Martinez, Oscar

    Thermal protection systems (TPS) are the key features incorporated into a spacecraft's design to protect it from severe aerodynamic heating during high-speed travel through planetary atmospheres. The thermal protection system is the key technology that enables a spacecraft to be lightweight, fully reusable, and easily maintainable. Add-on TPS concepts have been used since the beginning of the space race. The Apollo space capsule used ablative TPS and the Space Shuttle Orbiter TPS technology consisted of ceramic tiles and blankets. Many problems arose from the add-on concept such as incompatibility, high maintenance costs, non-load bearing, and not being robust and operable. To make the spacecraft's TPS more reliable, robust, and efficient, we investigated Integral Thermal Protection System (ITPS) concept in which the load-bearing structure and the TPS are combined into one single component. The design of an ITPS was a challenging task, because the requirement of a load-bearing structure and a TPS are often conflicting. Finite element (FE) analysis is often the preferred method of choice for a structural analysis problem. However, as the structure becomes complex, the computational time and effort for an FE analysis increases. New structural analytical tools were developed, or available ones were modified, to perform a full structural analysis of the ITPS. With analytical tools, the designer is capable of obtaining quick and accurate results and has a good idea of the response of the structure without having to go to an FE analysis. A MATLABRTM code was developed to analytically determine performance metrics of the ITPS such as stresses, buckling, deflection, and other failure modes. The analytical models provide fast and accurate results that were within 5% difference from the FEM results. The optimization procedure usually performs 100 function evaluations for every design variable. Using the analytical models in the optimization procedure was a time saver

  8. A method for determining structural properties of RCC thermal protection material. [Reinforced Carbon-Carbon

    NASA Technical Reports Server (NTRS)

    Wakefield, R. M.; Fowler, K. R.

    1978-01-01

    A method was developed for evaluation and prediction of effects of oxidation of the graphitic substrate on structural properties of Reinforced Carbon-Carbon (RCC) thermal protection material. Test specimens of RCC material were exposed to successive periods of convective heating in a plasma-jet facility to simulate the chemical reactions of Shuttle atmospheric entry. After each period of testing, the test specimen mass loss and performance in a nondestructive flexure test were determined. A computational model of the RCC specimen was developed for the NASA Structural Analysis (NASTRAN) program and validated by comparison of calculated and experimental results of flexure tests. The elastic moduli and ultimate loads in tension and compression were then computed for various levels of substrate oxidation.

  9. Mission load dynamic tests of two undensified Space shuttle thermal protection system tiles

    NASA Technical Reports Server (NTRS)

    Leatherwood, J. D.; Gowdey, J. C.

    1981-01-01

    Two tests of undensified Space Shuttle thermal protection tiles under combined static and dynamic loads were conducted. The tiles had a density of approximately 144 Kg/cum (LI900 tiles) and were mounted on a strain isolation pad which was 0.41 cm (.160 inch) thick. A combined static and dynamic mission stress histogram representative of the W-3 area of the wing of the orbiter vehicle was applied. The stress histogram was provided by the space shuttle project. Results presented include: tabulation of measured peak and root-mean-square (RMS) accelerations in both compression and tension; peak SIP stress in compression and tension, peak and RMS amplitude response ratios; lateral to vertical response ratios; response time histories; peak stress distributions (histograms), and SIP extension measured both with and without static tension at various mission times.

  10. Mechanical properties of the Shuttle Orbiter thermal protection system strain isolator pad

    NASA Technical Reports Server (NTRS)

    Sawyer, J. W.

    1982-01-01

    An experimental investigation conducted to determine the static and fatigue properties of the Strain Isolator Pad (SIP) used on the Shuttle Orbiter Thermal protection system is described. Static tension-compression and shear test results show that the SIP is highly nonlinear and that it possesses a large hysteresis, a large low modulus region for low stress levels, and stress-strain properties that are highly sensitive to strain rate and previous load history. In addition, the shear properties are also found to be sensitive to forces applied normal to the plane of the pad and to the orientation of the material. For the undensified tile/SIP system, static and fatigue failure takes place at the SIP/tile interface at low stress levels and for a small number of cycles.

  11. Edge softening of the Shuttle TPS strain isolation pad. [Thermal Protection System

    NASA Technical Reports Server (NTRS)

    Ransone, P. O.; Rummler, D. R.

    1982-01-01

    Tensile tests and an analytical investigation were performed to characterize the edge softening behavior of the strain isolation pad (SIP) between the Orbiter skin and thermal protection system. The tensile tests were carried out with varying sizes of disk-shaped specimens bonded between aluminum disks. The specimens strength and stiffness were determined on the basis of specimen size, and an analytical model of the microstructural stress-strain characteristics was developed. Strength and stiffness were found to decrease near the free edges because through-the-thickness fibers located there were not anchored. No size dependence at maximum load was observed in specimens between 0.75-4.0 in. thick. In-plane and out-of-plane coupling in deformation was detected. The model gave accurate predictions of the tensile behavior of the SIP as a function of distance to a free edge.

  12. Photoelastic tests on models of thermal protection system for space shuttle orbiter

    NASA Technical Reports Server (NTRS)

    Prabhakaran, R.; Cooper, P. A.

    1980-01-01

    The thermal protection system (TPS) of the space shuttle orbiter vehicle, consisting of ceramic tile/adhesive/strain isolator pad/adhesive/aluminum substructure, was modeled photoelasticity. A highly sensitive photoelastic material was used in the models to show the nature of the stress-transfer between the strain isolation pad (SIP) and the ceramic tile through the RTV-adhesive layer. Isochromatic fringe patterns were obtained for models subjected to tension and combined tension and bending. Tests indicated that the load-transfer between the SIP and the photoelastic material occurred at discrete locations causing stress concentrations in the photoelastic material. Stress concentration factors of the order of 1.9 were measured, but as the observed photoelastic response was an integrated effect through the model thickness, the local stress concentration factors at the SIP/tile interface could be even higher.

  13. Computer program for nonlinear static stress analysis of shuttle thermal protection system: User's manual

    NASA Technical Reports Server (NTRS)

    Giles, G. L.; Wallas, M.

    1981-01-01

    User documentation is presented for a computer program which considers the nonlinear properties of the strain isolator pad (SIP) in the static stress analysis of the shuttle thermal protection system. This program is generalized to handle an arbitrary SIP footprint including cutouts for instrumentation and filler bar. Multiple SIP surfaces are defined to model tiles in unique locations such as leading edges, intersections, and penetrations. The nonlinearity of the SIP is characterized by experimental stress displacement data for both normal and shear behavior. Stresses in the SIP are calculated using a Newton iteration procedure to determine the six rigid body displacements of the tile which develop reaction forces in the SIP to equilibrate the externally applied loads. This user documentation gives an overview of the analysis capabilities, a detailed description of required input data and an example to illustrate use of the program.

  14. An assessment of the impact of transition on advanced winged entry vehicle thermal protection system mass

    NASA Technical Reports Server (NTRS)

    Wurster, K. E.

    1981-01-01

    This study examines the impact of turbulent heating on thermal protection system (TPS) mass for advanced winged entry vehicles. Four basic systems are considered: insulative, metallic hot structures, metallic standoff, and hybrid systems. TPS sizings are performed using entry trajectories tailored specifically to the characteristics of each TPS concept under consideration. Comparisons are made between systems previously sized under the assumption of all laminar heating and those sized using a baseline estimate of transition and turbulent heating. The relative effect of different transition criteria on TPS mass requirements is also examined. Also investigated are entry trajectories tailored to alleviate turbulent heating. Results indicate the significant impact of turbulent heating on TPS mass and demonstrate the importance of both accurate transition criteria and entry trajectory tailoring.

  15. Determination of Acreage Thermal Protection Foam Loss From Ice and Foam Impacts

    NASA Technical Reports Server (NTRS)

    Carney, Kelly S.; Lawrence, Charles

    2015-01-01

    A parametric study was conducted to establish Thermal Protection System (TPS) loss from foam and ice impact conditions similar to what might occur on the Space Launch System. This study was based upon the large amount of testing and analysis that was conducted with both ice and foam debris impacts on TPS acreage foam for the Space Shuttle Project External Tank. Test verified material models and modeling techniques that resulted from Space Shuttle related testing were utilized for this parametric study. Parameters varied include projectile mass, impact velocity and impact angle (5 degree and 10 degree impacts). The amount of TPS acreage foam loss as a result of the various impact conditions is presented.

  16. The Relationship Between Physical Activity and Thermal Protective Clothing on Functional Balance in Firefighters

    PubMed Central

    Kong, Pui W.; Suyama, Joe; Cham, Rakié; Hostler, David

    2016-01-01

    We investigated the relationship between baseline physical training and the use of firefighting thermal protective clothing (TPC) with breathing apparatus on functional balance. Twenty-three male firefighters performed a functional balance test under four gear/clothing conditions. Participants were divided into groups by physical training status, and task performance was analyzed. There was an effect of equipment and training status on performance with the group reporting both aerobic and resistance training performing better than the group reporting no physical training. In conclusion, firefighters walk more slowly as a strategy to maintain balance when wearing TPC, which may be suboptimal given the emergent nature of fire suppression. This result was most prominent in the group reporting no physical training. PMID:23367817

  17. Compton imaging tomography for nondestructive evaluation of spacecraft thermal protection systems

    NASA Astrophysics Data System (ADS)

    Romanov, Volodymyr; Burke, Eric; Grubsky, Victor

    2017-02-01

    Here we present new results of in situ nondestructive evaluation (NDE) of spacecraft thermal protection system materials obtained with POC-developed NDE tool based on a novel Compton Imaging Tomography (CIT) technique recently pioneered and patented by Physical Optics Corporation (POC). In general, CIT provides high-resolution three-dimensional Compton scattered X-ray imaging of the internal structure of evaluated objects, using a set of acquired two-dimensional Compton scattered X-ray images of consecutive cross sections of these objects. Unlike conventional computed tomography, CIT requires only one-sided access to objects, has no limitation on the dimensions and geometry of the objects, and can be applied to large multilayer non-uniform objects with complicated geometries. Also, CIT does not require any contact with the objects being imaged during its application.

  18. Evaluation of protective coatings under thermal insulation at high temperatures by the use of an innovative design

    SciTech Connect

    Lasarte, C.; Rincon, O.T. de; Montiel, A.

    1994-12-31

    In order to disseminate the existing information on protective systems that have given good performance results, NACE published Document 6H-189 through its technical groups working on coatings for carbon and stainless steels under insulation and corrosion under thermal insulation. This report is unique in its kind and, in the opinion of the authors of this paper, the next step should be the characterization of each of these systems in combination with different insulating materials. Based on NACE Document No. 6H-189, the design of a probe was developed to evaluate, in a salt chamber, the protective coatings which were supposed to work under thermal insulation at high temperatures (30--1,500 C) . This paper describes the results obtained with different combinations of protective coatings (Silicone-Aluminum, Zinc-Rich and Aluminum Metallizing), and thermal insulators (mineral wool, fiber glass and calcium silicate).

  19. Collaborative Analysis Tool for Thermal Protection Systems for Single Stage to Orbit Launch Vehicles

    NASA Technical Reports Server (NTRS)

    Alexander, Reginald Andrew; Stanley, Thomas Troy

    1999-01-01

    Presented is a design tool and process that connects several disciplines which are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system and in the case of SSTO vehicles with air breathing propulsion, which is currently being studied by the National Aeronautics and Space Administration (NASA); the thermal protection system (TPS) is linked directly to almost every major system. The propulsion system pushes the vehicle to velocities on the order of 15 times the speed of sound in the atmosphere before pulling up to go to orbit which results high temperatures on the external surfaces of the vehicle. Thermal protection systems to maintain the structural integrity of the vehicle must be able to mitigate the heat transfer to the structure and be lightweight. Herein lies the interdependency, in that as the vehicle's speed increases, the TPS requirements are increased. And as TPS masses increase the effect on the propulsion system and all other systems is compounded. To adequately determine insulation masses for a vehicle such as the one described above, the aeroheating loads must be calculated and the TPS thicknesses must be calculated for the entire vehicle. To accomplish this an ascent or reentry trajectory is obtained using the computer code Program to Optimize Simulated Trajectories (POST). The trajectory is then used to calculate the convective heat rates on several locations on the vehicles using the Miniature Version of the JA70 Aerodynamic Heating Computer Program (MINIVER). Once the heat rates are defined for each body point on the vehicle, then insulation thickness that are required to maintain the vehicle within structural limits are calculated using Systems Improved Numerical Differencing Analyzer (SINDA) models. If the TPS masses are too heavy for the performance of the vehicle the process may be repeated altering the trajectory or some other input to reduce

  20. Cross-protection between controlled acid-adaptation and thermal inactivation for 48 Escherichia coli strains.

    PubMed

    Haberbeck, Leticia Ungaretti; Wang, Xiang; Michiels, Chris; Devlieghere, Frank; Uyttendaele, Mieke; Geeraerd, Annemie H

    2017-01-16

    Given the importance of pH reduction and thermal treatment in food processing and food preservation strategies, the cross-protection between acid adaptation and subsequent thermal inactivation for 48 Escherichia coli strains was investigated. Those strains were selected among 188 E. coli strains according to their odds of growth under low pH conditions as determined by Haberbeck et al. (2015) [Haberbeck, L.U., Oliveira, R.C., Vivijs, B., Wenseleers, T., Aertsen, A., Michiels, C., Geeraerd, A.H., 2015. Variability in growth/no growth boundaries of 188 different Escherichia coli strains reveals that approximately 75% have a higher growth probability under low pH conditions than E. coli O157:H7 strain ATCC 43888. Food Microbiol. 45, 222-230]. E. coli cells were acid and nonacid-adapted during overnight growth in controlled acidic pH (5.5) and neutral pH (7.0), respectively, in buffered Lysogenic Broth (LB). Then, they were heat inactivated at 58°C in non-buffered LB adjusted to pH6.2 and 7.0. Thus, four conditions were tested in total by combining the different pH values during growth/thermal inactivation: 5.5/6.2, 5.5/7.0, 7.0/6.2 and 7.0/7.0. Acid adaptation in buffered LB at pH5.5 increased the heat resistance of E. coli strains in comparison with nonacid-adaptation at pH7.0. For instance, the median D58-value of strains inactivated at pH7.0 was approximately 6 and 4min for acid-adapted and nonacid-adapted strains, respectively. For the nonacid-adapted strains, the thermal inactivation at pH6.2 and 7.0 was not significantly (p=0.06) different, while for the acid-adapted strains, the thermal treatment at pH6.2 showed a higher heat resistance than at pH7.0. The correlation between the odds of growth under low pH previously determined and the heat resistance was significant (p<0.05). Remarkably, a great variability in heat resistance among the strains was observed for all pH combinations, with D58-values varying between 1.0 and 69.0min. In addition, highly heat

  1. Improvements in Thermal Protection Sizing Capabilities for TCAT: Conceptual Design for Advanced Space Transportation Systems

    NASA Technical Reports Server (NTRS)

    Olds, John R.; Izon, Stephen James

    2002-01-01

    The Thermal Calculation Analysis Tool (TCAT), originally developed for the Space Systems Design Lab at the Georgia Institute of Technology, is a conceptual design tool capable of integrating aeroheating analysis into conceptual reusable launch vehicle design. It provides Thermal Protection System (TPS) unit thicknesses and acreage percentages based on the geometry of the vehicle and a reference trajectory to be used in calculation of the total cost and weight of the vehicle design. TCAT has proven to be reasonably accurate at calculating the TPS unit weights for in-flight trajectories; however, it does not have the capability of sizing TPS materials above cryogenic fuel tanks for ground hold operations. During ground hold operations, the vehicle is held for a brief period (generally about two hours) during which heat transfer from the TPS materials to the cryogenic fuel occurs. If too much heat is extracted from the TPS material, the surface temperature may fall below the freezing point of water, thereby freezing any condensation that may be present at the surface of the TPS. Condensation or ice on the surface of the vehicle is potentially hazardous to the mission and can also damage the TPS. It is questionable whether or not the TPS thicknesses provided by the aeroheating analysis would be sufficiently thick to insulate the surface of the TPS from the heat transfer to the fuel. Therefore, a design tool has been developed that is capable of sizing TPS materials at these cryogenic fuel tank locations to augment TCAT's TPS sizing capabilities.

  2. Monitoring of Thermal Protection Systems and MMOD using Robust Self-Organizing Optical Fiber Sensing Networks

    NASA Technical Reports Server (NTRS)

    Richards, Lance

    2014-01-01

    The general aim of this work is to develop and demonstrate a prototype structural health monitoring system for thermal protection systems that incorporates piezoelectric acoustic emission (AE) sensors to detect the occurrence and location of damaging impacts, such as those from Micrometeoroid Orbital Debris (MMOD). The approach uses an optical fiber Bragg grating (FBG) sensor network to evaluate the effect of detected damage on the thermal conductivity of the TPS material. Following detection of an impact, the TPS would be exposed to a heat source, possibly the sun, and the temperature distribution on the inner surface in the vicinity of the impact measured by the FBG network. A similar procedure could also be carried out as a screening test immediately prior to re-entry. The implications of any detected anomalies in the measured temperature distribution will be evaluated for their significance in relation to the performance of the TPS during reentry. Such a robust TPS health monitoring system would ensure overall crew safety throughout the mission, especially during reentry.

  3. Thermal Protection System Mass Estimating Relationships for Blunt-Body, Earth Entry Spacecraft

    NASA Technical Reports Server (NTRS)

    Sepka, Steven A.; Samareh, Jamshid A.

    2015-01-01

    System analysis and design of any entry system must balance the level fidelity for each discipline against the project timeline. One way to inject high fidelity analysis earlier in the design effort is to develop surrogate models for the high-fidelity disciplines. Surrogate models for the Thermal Protection System (TPS) are formulated as Mass Estimating Relationships (MERs). The TPS MERs are presented that predict the amount of TPS necessary for safe Earth entry for blunt-body spacecraft using simple correlations that closely match estimates from NASA's high-fidelity ablation modeling tool, the Fully Implicit Ablation and Thermal Analysis Program (FIAT). These MERs provide a first order estimate for rapid feasibility studies. There are 840 different trajectories considered in this study, and each TPS MER has a peak heating limit. MERs for the vehicle forebody include the ablators Phenolic Impregnated Carbon Ablator (PICA) and Carbon Phenolic atop Advanced Carbon-Carbon. For the aftbody, the materials are Silicone Impregnated Reusable Ceramic Ablator (SIRCA), Acusil II, SLA-561V, and LI-900. The MERs are accurate to within 14% (at one standard deviation) of FIAT prediction, and the most any MER under predicts FIAT TPS thickness is 18.7%. This work focuses on the development of these MERs, the resulting equations, model limitations, and model accuracy.

  4. Shearographic Non-destructive Evaluation of Space Shuttle Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Davis, Christopher K.; Hooker, Jeffery A.; Simmons, Stephen A.; Tenbusch, Kenneth E.

    1995-01-01

    Preliminary results of shearographic inspections of the shuttle external tank (ET) spray-on foam insulation (SOFI) and solid rocket booster (SRB) Marshall sprayable ablative (MSA-2) epoxy-cork thermal protection systems (TPS) are presented. Debonding SOFI or MSA-2 damage the orbiter 'belly' tile and exposes the ET/SRB to thermal loading. Previous work with the ET/SRB showed promising results with shearography. The first area investigated was the jack pad close-out, one of many areas on the ET where foam is applied at KSC. Voids 0.375 inch were detected in 1.75 inch thick foam using a pressure reduction of less than 0.4 psi. Of primary interest are areas of the ET that directly face the orbiter tile TPS. It is estimated that 90% of tile TPS damage on the orbiter 'belly' results from debonding SOFI during ascent. Test panels modeling these areas were manufactured with programmed debonds to determine the sensitivity of shearography as a function of debond size, SOFI thickness and vacuum. Results show repeatable detection of debonds with a diameter approximately half the SOFI thickness at less than 0.4 psi pressure reduction. Preliminary results are also presented on inspections of MSA-2 and the remote manipulator system (RMS) honeycomb material.

  5. Shearographic non-destructive evaluation of space shuttle thermal protection systems

    NASA Technical Reports Server (NTRS)

    Hooker, Jeffrey A.; Simmons, Stephen M.; Davis, Christopher K.; Tenbusch, Kenneth E.

    1995-01-01

    Preliminary results of shearographic inspections of the shuttle external tank (ET) spray-on foam insulation (SOFI) and solid rocket booster (SRB) Marshall sprayable ablative (MSA-2) epoxy-cork thermal protection systems (TPS) are presented. Debonding SOFI or MSA-2 damage the orbiter 'belly' tile and exposes the ET/SRB to thermal loading. Previous work with the ET/SRB showed promising results with shearography. The first area investigated was the jack pad close-out, one of many areas on the ET where foam is applied at KSC. Voids 0.375 inch were detected in 1.75 inch thick foam using a pressure reduction of less than 0.4 psi. Of primary interest are areas of the ET that directly face the orbiter tile TPS. It is estimated that 90% of tile TPS damage on the orbiter 'belly' results from debonding SOFI during ascent. Test panels modeling these areas were manufactured with programmed debonds to determine the sensitivity of shearography as a function of debond size, SOFI thickness and vacuum. Results show repeatable detection of debonds with a diameter approximately half the SOFI thickness at less than 0.4 psi pressure reduction. Preliminary results are also presented on inspections of MSA-2 and the remote manipulator system (RMS) honeycomb material

  6. CHAP III- CHARRING ABLATOR PROGRAM FOR ADVANCED INVESTIGATION OF THERMAL PROTECTION SYSTEMS FOR ENTRY

    NASA Technical Reports Server (NTRS)

    Stroud, C. W.

    1994-01-01

    The transient response of a thermal protection material to heat applied to the surface can be calculated using the CHAP III computer program. CHAP III can be used to analyze pyrolysis gas chemical kinetics in detail and examine pyrolysis reactions-indepth. The analysis includes the deposition of solid products produced by chemical reactions in the gas phase. CHAP III uses a modelling technique which can approximate a wide range of ablation problems. The energy equation used in CHAP III incorporates pyrolysis (both solid and gas reactions), convection, conduction, storage, work, kinetic energy, and viscous dissipation. The chemically reacting components of the solid are allowed to vary as a function of position and time. CHAP III employs a finite difference method to approximate the energy equations. Input values include specific heat, thermal conductivity, thermocouple locations, enthalpy, heating rates, and a description of the chemical reactions expected. The output tabulates the temperature at locations throughout the ablator, gas flow within the solid, density of the solid, weight of pyrolysis gases, and rate of carbon deposition. A sample case is included, which analyzes an ablator material containing several pyrolysis reactions subjected to an environment typical of entry at lunar return velocity. CHAP III is written in FORTRAN IV for batch execution and has been implemented on a CDC CYBER 170 series computer operating under NOS with a central memory requirement of approximately 102K (octal) of 60 bit words. This program was developed in 1985.

  7. Final analysis and design of a thermal protection system for 8-foot HTST combustor

    NASA Technical Reports Server (NTRS)

    Moskowitz, S.

    1973-01-01

    The cylindrical shell combustor with T-bar supports in the 8-foot HTST at the NASA-Langley Research Center encountered vibratory fatigue cracking over a period of 50-250 tunnel tests within a limited range of the required operating envelope. A preliminary design study provided several suitable thermal protection system designs for the combustor, one of which was a two-pass regenerative type air-cooled omega-shaped segment liner. A final design layout of the omega segment liner was prepared and analyzed for steady-state and transient conditions. The design of a support system for the fuel spray bar assembly was also included. Detail drawings suitable for fabrication purposes were also prepared. Liner design problems defined during the preliminary study included (1) the ingress of gas into the attachment bulb section of the omega segment, (2) the large thermal gradient along the leg of the omega bulb attachment section and, (3) the local peak metal temperature at the radius between the liner ID and the leg of the bulb attachment. These were resolved during the final design task. Analyses of the final design of the omega segment liner indicated that all design goals were met and the design provided the capability of operating over the required test envelope with a life expectancy substantially above the goal of 1500 cycles.

  8. Thermal Protection System Mass Estimating Relationships For Blunt-Body, Earth Entry Spacecraft

    NASA Technical Reports Server (NTRS)

    Sepka, Steven A.; Samareh, Jamshid A.

    2015-01-01

    Mass estimating relationships (MERs) are developed to predict the amount of thermal protection system (TPS) necessary for safe Earth entry for blunt-body spacecraft using simple correlations that are non-ITAR and closely match estimates from NASA's highfidelity ablation modeling tool, the Fully Implicit Ablation and Thermal Analysis Program (FIAT). These MERs provide a first order estimate for rapid feasibility studies. There are 840 different trajectories considered in this study, and each TPS MER has a peak heating limit. MERs for the vehicle forebody include the ablators Phenolic Impregnated Carbon Ablator (PICA) and Carbon Phenolic atop Advanced Carbon-Carbon. For the aftbody, the materials are Silicone Impregnated Reusable Ceramic Ablator (SIRCA), Acusil II, SLA- 561V, and LI-900. The MERs are accurate to within 14% (at one standard deviation) of FIAT prediction, and the most any MER can under predict FIAT TPS thickness is 18.7%. This work focuses on the development of these MERs, the resulting equations, model limitations, and model accuracy.

  9. Parametric Weight Comparison of Current and Proposed Thermal Protection System (TPS) Concepts

    NASA Technical Reports Server (NTRS)

    Myers, David E.; Martin, Carl J.; Blosser, Max L.

    1999-01-01

    A parametric weight assessment of advanced metallic panel, ceramic blanket, and ceramic tile thermal protection systems (TPS) was conducted using an implicit, one-dimensional (1 -D) thermal finite element sizing code. This sizing code contained models to ac- count for coatings, fasteners, adhesives, and strain isolation pads. Atmospheric entry heating profiles for two vehicles, the Access to Space (ATS) rocket-powered single-stage-to-orbit (SSTO) vehicle and a proposed Reusable Launch Vehicle (RLV), were used to ensure that the trends were not unique to a particular trajectory. Eight TPS concepts were compared for a range of applied heat loads and substructural heat capacities to identify general trends. This study found the blanket TPS concepts have the lightest weights over the majority of their applicable ranges, and current technology ceramic tiles and metallic TPS concepts have similar weights. A proposed, state-of-the-art metallic system which uses a higher temperature alloy and efficient multilayer insulation was predicted to be significantly lighter than the ceramic tile systems and approaches blanket TPS weights for higher integrated heat loads.

  10. Thermal Protection System Evaluation Using Arc-jet Flows: Flight Simulation or Research Tool?

    NASA Technical Reports Server (NTRS)

    Stewart, David A.; Venkatapathy, Ethiras (Technical Monitor)

    2002-01-01

    The arc-jet has been used to evaluate thermal protection systems (TPS) and materials for the past forty years. Systems that have been studied in this environmerd include ablators, active, and passive TPS concepts designed for vehicles entering planetary and Earth atmospheres. The question of whether arc-jet flow can simulate a flight environment or is it a research tool that provides an aero-thermodynamic heating environment to obtain critical material properties will be addressed. Stagnation point tests in arc-jets are commonly used to obtain material properties such as mass loss rates, thermal chemical stability data, optical properties, and surface catalytic efficiency. These properties are required in computational fluid dynamic codes to accurately predict the performance of a TPS during flight. Special facilities have been developed at NASA Ames Research Center to approximate the flow environment over the mid-fuselage and body flap regions of proposed space-planes type vehicles. This paper compares flow environments generated in flight over a vehicle with those created over an arc-jet test articles in terms of scale, chemistry, and fluid dynamic properties. Flight experiments are essential in order to validate the material properties obtained from arc-jet tests and used to predict flight performance of any TPS being considered for use on a vehicle entering the Earth atmosphere at hypersonic speed.

  11. Ceramic Matrix Composite (CMC) Thermal Protection Systems (TPS) and Hot Structures for Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Glass, David E.

    2008-01-01

    Thermal protection systems (TPS) and hot structures are required for a range of hypersonic vehicles ranging from ballistic reentry to hypersonic cruise vehicles, both within Earth's atmosphere and non-Earth atmospheres. The focus of this paper is on air breathing hypersonic vehicles in the Earth's atmosphere. This includes single-stage to orbit (SSTO), two-stage to orbit (TSTO) accelerators, access to space vehicles, and hypersonic cruise vehicles. This paper will start out with a brief discussion of aerodynamic heating and thermal management techniques to address the high heating, followed by an overview of TPS for rocket-launched and air-breathing vehicles. The argument is presented that as we move from rocket-based vehicles to air-breathing vehicles, we need to move away from the insulated airplane approach used on the Space Shuttle Orbiter to a wide range of TPS and hot structure approaches. The primary portion of the paper will discuss issues and design options for CMC TPS and hot structure components, including leading edges, acreage TPS, and control surfaces. The current state-of-the-art will be briefly discussed for some of the components. The two primary technical challenges impacting the use of CMC TPS and hot structures for hypersonic vehicles are environmental durability and fabrication, and will be discussed briefly.

  12. Thermal-sprayed zinc anodes for cathodic protection of steel-reinforced concrete bridges

    SciTech Connect

    Bullard, Sophie J.; Covino, Bernard S., Jr.; Cramer, Stephen D.; McGill, Galen E.

    1996-01-01

    Thermal-sprayed zinc anodes are being used in Oregon in impressed current cathodic protection (ICCP) systems for reinforced concrete bridges. The U.S. Department of Energy, Albany Research Center, is collaborating with the Oregon Department of Transportation (ODOT) to evaluate the long-term performance and service life of these anodes. Laboratory studies were conducted on concrete slabs coated with 0.5 mm (20 mil) thick, thermal-sprayed zinc anodes. The slabs were electrochemically aged at an accelerated rate using an anode current density of 0.032 A/m2 (3mA/ft2). Half the slabs were preheated before thermal-spraying with zinc; the other half were unheated. Electrochemical aging resulted in the formation at the zinc-concrete interface of a thin, low pH zone (relative to cement paste) consisting primarily of ZnO and Zn(OH)2, and in a second zone of calcium and zinc aluminates and silicates formed by secondary mineralization. Both zones contained elevated concentrations of sulfate and chloride ions. The original bond strength of the zinc coating decreased due to the loss of mechanical bond to the concrete with the initial passage of electrical charge (aging). Additional charge led to an increase in bond strength to a maximum as the result of secondary mineralization of zinc dissolution products with the cement paste. Further charge led to a decrease in bond strength and ultimately coating disbondment as the interfacial reaction zones continued to thicken. This occurred at an effective service life of 27 years at the 0.0022 A/m2 (0.2 mA/ft2) current density typically used by ODOT in ICCP systems for coastal bridges. Zinc coating failure under tensile stress was primarily cohesive within the thickening reaction zones at the zinc-concrete interface. There was no difference between the bond strength of zinc coatings on preheated and unheated concrete surfaces after long service times.

  13. Prediction of In-Space Durability of Protected Polymers Based on Ground Laboratory Thermal Energy Atomic Oxygen

    NASA Technical Reports Server (NTRS)

    Banks, Bruce A.; deGroh, Kim K.; Rutledge, Sharon; DiFilippo, Frank J.

    1996-01-01

    The probability of atomic oxygen reacting with polymeric materials is orders of magnitude lower at thermal energies (greater than O.1 eV) than at orbital impact energies (4.5 eV). As a result, absolute atomic oxygen fluxes at thermal energies must be orders of magnitude higher than orbital energy fluxes, to produce the same effective fluxes (or same oxidation rates) for polymers. These differences can cause highly pessimistic durability predictions for protected polymers and polymers which develop protective metal oxide surfaces as a result of oxidation if one does not make suitable calibrations. A comparison was conducted of undercut cavities below defect sites in protected polyimide Kapton samples flown on the Long Duration Exposure Facility (LDEF) with similar samples exposed in thermal energy oxygen plasma. The results of this comparison were used to quantify predicted material loss in space based on material loss in ground laboratory thermal energy plasma testing. A microindent hardness comparison of surface oxidation of a silicone flown on the Environmental Oxygen Interaction with Materials-III (EOIM-III) experiment with samples exposed in thermal energy plasmas was similarly used to calibrate the rate of oxidation of silicone in space relative to samples in thermal energy plasmas exposed to polyimide Kapton effective fluences.

  14. Testing Lunar Return Thermal Protection Systems using Sub-Scale Flight Test Vehicles

    NASA Technical Reports Server (NTRS)

    Chen, George; De Jong, Christian; Ivanov, Mark; Ong, Chester; Seybold, Calina; Hash, David

    2007-01-01

    A key objective of NASA's Vision for Space Exploration is to revisit the lunar surface. Such an ambitious goal requires the development of a new human-rated spacecraft, the Orion Crew Exploration Vehicle (CEV), to ferry crews to low earth orbit and to the moon. The successful conclusion of both types of missions will require a thermal protection system (TPS) capable of protecting the vehicle and crew from the extreme heat of atmospheric reentry. As a part of the TPS development, various materials are being tested in arcjet tunnels; however, the combined lunar return aerothermal environment of high heat flux, shear stress, and surface pressure cannot be duplicated using only existing ground test facilities. To ensure full TPS qualification, a flight test program using sub-scale Orion capsules has been proposed to test TPS materials and heat shield construction techniques under the most stressing combination of lunar return aerothermal environments. Originally called Testing Of Reentry Capsule Heat Shield, or TORCH, but later renamed LEX, for Lunar Reentry Experiment, the proposed flight test program is presented along with the driving requirements and descriptions of the vehicle and the TPS instrumentation suite slated to conduct in-flight measurements.

  15. Improving Metallic Thermal Protection System Hypervelocity Impact Resistance Through Design of Experiments Approach

    NASA Technical Reports Server (NTRS)

    Poteet, Carl C.; Blosser, Max L.

    2001-01-01

    A design of experiments approach has been implemented using computational hypervelocity impact simulations to determine the most effective place to add mass to an existing metallic Thermal Protection System (TPS) to improve hypervelocity impact protection. Simulations were performed using axisymmetric models in CTH, a shock-physics code developed by Sandia National Laboratories, and validated by comparison with existing test data. The axisymmetric models were then used in a statistical sensitivity analysis to determine the influence of five design parameters on degree of hypervelocity particle dispersion. Several damage metrics were identified and evaluated. Damage metrics related to the extent of substructure damage were seen to produce misleading results, however damage metrics related to the degree of dispersion of the hypervelocity particle produced results that corresponded to physical intuition. Based on analysis of variance results it was concluded that the most effective way to increase hypervelocity impact resistance is to increase the thickness of the outer foil layer. Increasing the spacing between the outer surface and the substructure is also very effective at increasing dispersion.

  16. Application experience of gas-thermal aluminum coatings to protect the pipes for underground construction and repair of heat networks

    NASA Astrophysics Data System (ADS)

    Kolpakov, A. S.

    2013-11-01

    Questions of sacrificial protection for pipes of underground heat networks with aluminum against the external corrosion are considered. The description of pilot production of pipes with a plasma aluminum coating and the deposition of a sacrificial gas-plasma aluminum coating on weld joints of pipelines and the zone of their thermal influence during assemblage is presented. Examples of repairing the segments of distribution heat networks by the pipes with the tread protection are presented.

  17. Design of an integral thermal protection system for future space vehicles

    NASA Astrophysics Data System (ADS)

    Bapanapalli, Satish Kumar

    Thermal protection systems (TPS) are the features incorporated into a spacecraft's design to protect it from severe aerodynamic heating during high-speed travel through planetary atmospheres. The ablative TPS on the space capsule Apollo and ceramic tiles and blankets on the Space Shuttle Orbiter were designed as add-ons to the main load-bearing structure of the vehicles. They are usually incompatible with the structure due to mismatch in coefficient of thermal expansion and as a result the robustness of the external surface of the spacecraft is compromised. This could potentially lead to catastrophic consequences because the TPS forms the external surface of the vehicle and is subjected to numerous other loads like aerodynamic pressure loads, small object high-speed impacts and handling damages during maintenance. In order to make the spacecraft external surface robust, an Integral Thermal Protection System (ITPS) concept has been proposed in this research in which the load-bearing structure and the TPS are combined into one single structure. The design of an ITPS is a formidable task because the requirement of a load-bearing structure and a TPS are often contradictory to one another. The design process has been formulated as an optimization problem with mass per unit area of the ITPS as the objective function and the various functions of the ITPS were formulated as constraints. This is a multidisciplinary design optimization problem involving heat transfer and structural analysis fields. The constraints were expressed as response surface approximations obtained from a large number of finite element analyses, which were carried out with combinations of design variables obtained from an optimized Latin-Hypercube sampling scheme. A MATLABRTM code has been developed to carry out these FE analyses automatically in conjunction with ABAQUSRTM . Corrugated-core structures were designed for ITPS applications with loads and boundary conditions similar to that of a Space

  18. A Base Drag Reduction Experiment on the X-33 Linear Aerospike SR-71 Experiment (LASRE) Flight Program

    NASA Technical Reports Server (NTRS)

    Whitmore, Stephen A.; Moes, Timothy R.

    1999-01-01

    Drag reduction tests were conducted on the LASRE/X-33 flight experiment. The LASRE experiment is a flight test of a roughly 20% scale model of an X-33 forebody with a single aerospike engine at the rear. The experiment apparatus is mounted on top of an SR-71 aircraft. This paper suggests a method for reducing base drag by adding surface roughness along the forebody. Calculations show a potential for base drag reductions of 8-14%. Flight results corroborate the base drag reduction, with actual reductions of 15% in the high-subsonic flight regime. An unexpected result of this experiment is that drag benefits were shown to persist well into the supersonic flight regime. Flight results show no overall net drag reduction. Applied surface roughness causes forebody pressures to rise and offset base drag reductions. Apparently the grit displaced streamlines outward, causing forebody compression. Results of the LASRE drag experiments are inconclusive and more work is needed. Clearly, however, the forebody grit application works as a viable drag reduction tool.

  19. Aerothermal and structural performance of a cobalt-base superalloy thermal protection system at Mach 6.6

    NASA Technical Reports Server (NTRS)

    Sawyer, J. W.

    1977-01-01

    A flightweight, metallic thermal protection system (TPS) applicable to reentry and hypersonic vehicles was subjected to multiple cycles of both radiant and aerothermal heating in order to evaluate its aerothermal performance and structural integrity. Good structural integrity and thermal performance were demonstrated by the TPS under both a radiant and aerothermal heating environment typical of a shuttle entry. The shingle-slip joints effectively allowed for thermal expansion of the panel without allowing any appreciable hot gas flow into the TPS cavity. The TPS also demonstrated good structural ruggedness.

  20. Thermomechanical response of metal foam sandwich panels for structural thermal protection systems in hypersonic vehicles

    NASA Astrophysics Data System (ADS)

    Rakow, Joseph F.

    Sandwich panels with metal foam cores are proposed for load-bearing structural components in actively cooled thermal protection systems for aerospace vehicles. Prototype acreage metal foam sandwich panels (MFSP's) are constructed and analyzed with the central goal of characterizing the thermomechanical response of the system. MFSP's are subjected to uniform temperature fields and equibiaxial loading in a novel experimental load frame. The load frame exploits the mismatch of coefficients of thermal expansion and allows for thermostructural experimentation without the endemic conflict of thermal and mechanical boundary conditions. Back-to-back strain gages and distributed thermocouples capture the in-plane response of the panels, including buckling and elastic-plastic post-buckling. The out-of-plane response is captured via moire interferometry, which provides a visualization of evolving mode shapes throughout the post-buckling regime. The experimental results agree with an analytical prediction for critical temperatures in sandwich panels based on a Rayleigh-Ritz minimization of the energy functional for a Reissner-Mindlin plate. In addition, a three-dimensional finite element model of the non-linear thermomechanical response of the panel-frame experimental system is developed and the results are shown to agree well with the experimentally identified response of MFSP's. Central to analytical and numerical characterization of MFSP's is an understanding of the response of metal foam under shear loading. The shear response of metal foam is captured experimentally, providing density-dependent relationships for material stiffness, strength, and energy absorption. Speckle photography is employed to identify microstructural size effects in the distribution of strain throughout metal foam under shear loading. In addition, a micromechanical model is established for the density-dependent shear modulus of metal foam, which allows for the coupling of cell-level imperfections

  1. Thermal Protection System Cavity Heating for Simplified and Actual Geometries Using Computational Fluid Dynamics Simulations with Unstructured Grids

    NASA Technical Reports Server (NTRS)

    McCloud, Peter L.

    2010-01-01

    Thermal Protection System (TPS) Cavity Heating is predicted using Computational Fluid Dynamics (CFD) on unstructured grids for both simplified cavities and actual cavity geometries. Validation was performed using comparisons to wind tunnel experimental results and CFD predictions using structured grids. Full-scale predictions were made for simplified and actual geometry configurations on the Space Shuttle Orbiter in a mission support timeframe.

  2. Extravehicular Activity Probabilistic Risk Assessment Overview for Thermal Protection System Repair on the Hubble Space Telescope Servicing Mission

    NASA Technical Reports Server (NTRS)

    Bigler, Mark; Canga, Michael A.; Duncan, Gary

    2010-01-01

    The Shuttle Program initiated an Extravehicular Activity (EVA) Probabilistic Risk Assessment (PRA) to assess the risks associated with performing a Shuttle Thermal Protection System (TPS) repair during the Space Transportation System (STS)-125 Hubble repair mission as part of risk trades between TPS repair and crew rescue.

  3. Thermal overload protection for electric motors on safety-related motor-operated valves: Generic Issue II. E. 6. 1

    SciTech Connect

    Rothberg, O.

    1988-06-01

    NRC regulatory positions, as stated in Regulatory Guide 1.106, Revision 1, have been identified by the Office for Analysis and Evaluation of Operational Data (AEOD) as potential contributors to valve motor burnout. AEOD is particularly concerned about the allowed policy of bypassing thermal overload devices during normal or accident conditions. Regulatory Guide 1.106 favors compromising the function of thermal overload devices in favor of completing the safety-related action of valves. The purpose of this study was to determine if the guidance contained in Regulatory Guide 1.106 is appropriate and, if not, to recommend the necessary changes. This report describes thermal overload devices commonly used to protect safety-related valve operator motors. The regulatory guidelines stated in Regulatory Guide 1.106 along with the limitations of thermal overload protection are discussed. Supplements and alternatives to thermal overload protection are also described. Findings and conclusions of several AEOD reports are discussed. Information obtained from the standard review plan, standard technical specifications, technical specifications from representative plants, and several papers are cited.

  4. Woven Thermal Protection System Based Heat-shield for Extreme Entry Environments Technology (HEEET)

    NASA Technical Reports Server (NTRS)

    Chinnapongse, Ronald; Ellerbe, Donald; Stackpoole, Maragaret; Venkatapathy, Ethiraj; Beerman, Adam; Feldman, Jay; Peterson Keith; Prabhu, Dinesh; Dillman, Robert; Munk, Michelle

    2013-01-01

    NASA's future robotic missions utilizing an entry system into Venus and the outer planets, namely, Saturn, Uranus, Neptune, result in extremely severe entry conditions that exceed the capabilities of state of the art low to mid density ablators such as PICA or Avcoat. Therefore mission planners typically assume the use of a fully dense carbon phenolic heat shield similar to what was flown on Pioneer Venus and Galileo. Carbon phenolic (CP) is a robust TPS material however its high density and relatively high thermal conductivity constrain mission planners to steep entries, with high heat fluxes and pressures and short entry durations, in order for CP to be feasible from a mass perspective. The high entry conditions pose challenges for certification in existing ground based test facilities and the longer-­-term sustainability of CP will continue to pose challenges. In 2012 the Game Changing Development Program (GCDP) in NASA's Space Technology Mission Directorate funded NASA ARC to investigate the feasibility of a Woven Thermal Protection System (WTPS) to meet the needs of NASA's most challenging entry missions. This project was highly successful demonstrating that a Woven TPS solution compares favorably to CP in performance in simulated reentry environments and provides the opportunity to manufacture graded materials that should result in overall reduced mass solutions and enable a much broader set of missions than does CP. Building off the success of the WTPS project GCDP has funded a follow on project to further mature and scale up the WTPS concept for insertion into future NASA robotic missions. The matured WTPS will address the CP concerns associated with ground based test limitations and sustainability. This presentation will briefly discuss results from the WTPS Project and the plans for WTPS maturation into a heat-­-shield for extreme entry environment.

  5. Woven Thermal Protection System Based Heat-shield for Extreme Entry Environments Technology (HEEET)

    NASA Technical Reports Server (NTRS)

    Ellerby, Donald; Venkatapathy, Ethiraj; Stackpoole, Margaret; Chinnapongse, Ronald; Munk, Michelle; Dillman, Robert; Feldman, Jay; Prabhu, Dinesh; Beerman, Adam

    2013-01-01

    NASA's future robotic missions utilizing an entry system into Venus and the outer planets, namely, Saturn, Uranus, Neptune, result in extremely high entry conditions that exceed the capabilities of state of the art low to mid density ablators such as PICA or Avcoat. Therefore mission planners typically assume the use of a fully dense carbon phenolic heat shield similar to what was flown on Pioneer Venus and Galileo. Carbon phenolic is a robust TPS material however its high density and relatively high thermal conductivity constrain mission planners to steep entries, with high heat fluxes and pressures and short entry durations, in order for CP to be feasible from a mass perspective. The high entry conditions pose challenges for certification in existing ground based test facilities and the longer-term sustainability of CP will continue to pose challenges. In 2012 the Game Changing Development Program (GCDP) in NASA's Space Technology Mission Directorate funded NASA ARC to investigate the feasibility of a Woven Thermal Protection System (WTPS) to meet the needs of NASA's most challenging entry missions. This project was highly successful demonstrating that a Woven TPS solution compares favorably to CP in performance in simulated reentry environments and provides the opportunity to manufacture graded materials that should result in overall reduced mass solutions and enable a much broader set of missions than does CP. Building off the success of the WTPS project GCDP has funded a follow on project to further mature and scale up the WTPS concept for insertion into future NASA robotic missions. The matured WTPS will address the CP concerns associated with ground based test limitations and sustainability. This presentation will briefly discuss results from the WTPS Project and the plans for WTPS maturation into a heat-shield for extreme entry environment.

  6. CMC thermal protection system for future reusable launch vehicles: Generic shingle technological maturation and tests

    NASA Astrophysics Data System (ADS)

    Pichon, T.; Barreteau, R.; Soyris, P.; Foucault, A.; Parenteau, J. M.; Prel, Y.; Guedron, S.

    2009-07-01

    Experimental re-entry demonstrators are currently being developed in Europe, with the objective of increasing the technology readiness level (TRL) of technologies applicable to future reusable launch vehicles. Among these are the Pre-X programme, currently funded by CNES, the French Space Agency, and which is about to enter into development phase B, and the IXV, within the future launcher preparatory programme (FLPP) funded by ESA. One of the major technologies necessary for such vehicles is the thermal protection system (TPS), and in particular the ceramic matrix composites (CMC) based windward TPS. In support of this goal, technology maturation activities named "generic shingle" were initiated beginning of 2003 by SPS, under a CNES contract, with the objective of performing a test campaign of a complete shingle of generic design, in preparation of the development of a re-entry experimental vehicle decided in Europe. The activities performed to date include: the design, manufacturing of two C/SiC panels, finite element model (FEM) calculation of the design, testing of technological samples extracted from a dedicated panel, mechanical pressure testing of a panel, and a complete study of the attachment system. Additional testing is currently under preparation on the panel equipped with its insulation, seal, attachment device, and representative portion of cold structure, to further assess its behaviour in environments relevant to its application The paper will present the activities that will have been performed in 2006 on the prediction and preparation of these modal characterization, dynamic, acoustic as well as thermal and thermo-mechanical tests. Results of these tests will be presented and the lessons learned will be discussed.

  7. Results of tests of Insta-Foam Thermal Protection System (TPS) material for protection of equipment inside the SRB aft skirt

    NASA Technical Reports Server (NTRS)

    Dean, W. G.

    1982-01-01

    The objective of these tests was to determine whether Insta-Foam can be used successfully to protect items inside the solid rocket booster aft skirt during reentry. On some of the early Space Shuttle flights the aft skirt heat shield curtain failed during reentry. This allowed the hot gases to damage some of the equipment, etc., inside the skirt. For example, some of the propellant lines were overheated and ruptured and some of the NSI (nozzle severance) cables were damaged. It was suggested that the Insta-Foam thermal protection system be sprayed over these lines, etc., to protect them during future flights in case of a curtain failure. The tests presented were devised and run to check out the feasibility of this idea.

  8. Investigation of Post-Flight Solid Rocket Booster Thermal Protection System

    NASA Technical Reports Server (NTRS)

    Nelson, Linda A.

    2006-01-01

    After every Shuttle mission, the Solid Rocket Boosters (SRBs) are recovered and observed for missing material. Most of the SRB is covered with a cork-based thermal protection material (MCC-l). After the most recent shuttle mission, STS-114, the forward section of the booster appeared to have been impacted during flight. The darkened fracture surfaces indicated that this might have occurred early in flight. The scope of the analysis included microscopic observations to assess the degree of heat effects and locate evidence of the impact source as well as chemical analysis of the fracture surfaces and recovered foreign material using Fourier Transform Infrared Spectroscopy and Scanning Electron Microscopy/Energy Dispersive Spectroscopy. The amount of heat effects and presence of soot products on the fracture surface indicated that the material was impacted prior to SRB re-entry into the atmosphere. Fragments of graphite fibers found on these fracture surfaces were traced to slag inside the Solid Rocket Motor (SRM) that forms during flight as the propellant is spent and is ejected throughout the descent of the SRB after separation. The direction of the impact mark matches with the likely trajectory of SRBs tumbling prior to re-entry.

  9. Laser Shearography Inspection of TPS (Thermal Protection System) Cork on RSRM (Reusable Solid Rocket Motors)

    NASA Technical Reports Server (NTRS)

    Lingbloom, Mike; Plaia, Jim; Newman, John

    2006-01-01

    Laser Shearography is a viable inspection method for detection of de-bonds and voids within the external TPS (thermal protection system) on to the Space Shuttle RSRM (reusable solid rocket motors). Cork samples with thicknesses up to 1 inch were tested at the LTI (Laser Technology Incorporated) laboratory using vacuum-applied stress in a vacuum chamber. The testing proved that the technology could detect cork to steel un-bonds using vacuum stress techniques in the laboratory environment. The next logical step was to inspect the TPS on a RSRM. Although detailed post flight inspection has confirmed that ATK Thiokol's cork bonding technique provides a reliable cork to case bond, due to the Space Shuttle Columbia incident there is a great interest in verifying bond-lines on the external TPS. This interest provided and opportunity to inspect a RSRM motor with Laser Shearography. This paper will describe the laboratory testing and RSRM testing that has been performed to date. Descriptions of the test equipment setup and techniques for data collection and detailed results will be given. The data from the test show that Laser Shearography is an effective technology and readily adaptable to inspect a RSRM.

  10. Metallic Thermal Protection System Technology Development: Concepts, Requirements and Assessment Overview

    NASA Technical Reports Server (NTRS)

    Dorsey, John T.; Poteet, Carl C.; Chen, Roger R.; Wurster, Kathryn E.

    2002-01-01

    A technology development program was conducted to evolve an earlier metallic thermal protection system (TPS) panel design, with the goals of: improving operations features, increasing adaptability (ease of attaching to a variety of tank shapes and structural concepts), and reducing weight. The resulting Adaptable Robust Metallic Operable Reusable (ARMOR) TPS system incorporates a high degree of design flexibility (allowing weight and operability to be traded and balanced) and can also be easily integrated with a large variety of tank shapes, airframe structural arrangements and airframe structure/material concepts. An initial attempt has been made to establish a set of performance based TPS design requirements. A set of general (FARtype) requirements have been proposed, focusing on defining categories that must be included for a comprehensive design. Load cases required for TPS design must reflect the full flight envelope, including a comprehensive set of limit loads, However, including additional loads. such as ascent abort trajectories, as ultimate load cases, and on-orbit debris/micro-meteoroid hypervelocity impact, as one of the discrete -source -damage load cases, will have a significant impact on system design and resulting performance, reliability and operability. Although these load cases have not been established, they are of paramount importance for reusable vehicles, and until properly included, all sizing results and assessments of reliability and operability must be considered optimistic at a minimum.

  11. In-Space Repair and Refurbishment of Thermal Protection System Structures for Reusable Launch Vehicles

    NASA Technical Reports Server (NTRS)

    Singh, M.

    2007-01-01

    Advanced repair and refurbishment technologies are critically needed for the thermal protection system of current space transportation systems as well as for future launch and crew return vehicles. There is a history of damage to these systems from impact during ground handling or ice during launch. In addition, there exists the potential for in-orbit damage from micrometeoroid and orbital debris impact as well as different factors (weather, launch acoustics, shearing, etc.) during launch and re-entry. The GRC developed GRABER (Glenn Refractory Adhesive for Bonding and Exterior Repair) material has shown multiuse capability for repair of small cracks and damage in reinforced carbon-carbon (RCC) material. The concept consists of preparing an adhesive paste of desired ceramic with appropriate additives and then applying the paste to the damaged/cracked area of the RCC composites with an adhesive delivery system. The adhesive paste cures at 100-120 C and transforms into a high temperature ceramic during reentry conditions. A number of plasma torch and ArcJet tests were carried out to evaluate the crack repair capability of GRABER materials for Reinforced Carbon-Carbon (RCC) composites. For the large area repair applications, Integrated Systems for Tile and Leading Edge Repair (InSTALER) have been developed and evaluated under various ArcJet testing conditions. In this presentation, performance of the repair materials as applied to RCC is discussed. Additionally, critical in-space repair needs and technical challenges are reviewed.

  12. Flutter Analysis of the Thermal Protection Layer on the NASA HIAD

    NASA Technical Reports Server (NTRS)

    Goldman, Benjamin D.; Dowell, Earl H.; Scott, Robert C.

    2013-01-01

    A combination of classical plate theory and a supersonic aerodynamic model is used to study the aeroelastic flutter behavior of a proposed thermal protection system (TPS) for the NASA HIAD. The analysis pertains to the rectangular configurations currently being tested in a NASA wind-tunnel facility, and may explain why oscillations of the articles could be observed. An analysis using a linear flat plate model indicated that flutter was possible well within the supersonic flow regime of the wind tunnel tests. A more complex nonlinear analysis of the TPS, taking into account any material curvature present due to the restraint system or substructure, indicated that significantly greater aerodynamic forcing is required for the onset of flutter. Chaotic and periodic limit cycle oscillations (LCOs) of the TPS are possible depending on how the curvature is imposed. When the pressure from the base substructure on the bottom of the TPS is used as the source of curvature, the flutter boundary increases rapidly and chaotic behavior is eliminated.

  13. Detection of bolt loosening in C C composite thermal protection panels: II. Experimental verification

    NASA Astrophysics Data System (ADS)

    Yang, Jinkyu; Chang, Fu-Kuo

    2006-04-01

    The research presented in this paper is motivated by the need for reliable inspection technology for the detection of bolt loosening in carbon-carbon (C-C) thermal protection system (TPS) panels using minimal human intervention. Based on the diagnostic scheme proposed in part I of the study, a new PZT (lead zirconate titanate)-embedded sensor washer was developed to constitute the sensor network. The sensor suite was included in the C-C TPS prototype without jeopardizing the integrity of the original fastening components. The sensor-embedded washer enhances the remote sensing capability and achieves sufficient sensitivity by guiding the diagnostic waves to propagate primarily through the inspection areas. After evolution of the sensor design and appropriate algorithm development, the verification tests were conducted using a shaker which simulated the acoustic environments during the re-entry process. The test results revealed that the proposed system successfully identified the loss of the preload for the bolted joints under loosening and gave the correct diagnostic results. The sensors were found to be durable under cyclic mechanical loads without major failures, and the diagnostic scheme was capable of locating a loosened bracket as well as discriminating major failure modes 1 and 2: panel and bracket loosening in the bracket.

  14. Detection of bolt loosening in C C composite thermal protection panels: I. Diagnostic principle

    NASA Astrophysics Data System (ADS)

    Yang, Jinkyu; Chang, Fu-Kuo

    2006-04-01

    A concept demonstrator of the structural health monitoring (SHM) system was developed to autonomously detect the degradation of the mechanical integrity of the standoff carbon-carbon (C-C) thermal protection system (TPS) panels. This system enables us to identify the location of the loosened bolts, as well as to predict the torque levels of those bolts accordingly. In the process of building the proposed SHM prototype, efforts have been focused primarily on developing a trustworthy diagnostic scheme and a responsive sensor suite. In part I of the study, an attenuation-based diagnostic method was proposed to assess the fastener integrity by observing the attenuation patterns of the resultant sensor signals. The attenuation-based method is based on the damping phenomena of ultrasonic waves across the bolted joints. The major advantage of the attenuation-based method over the conventional diagnostic methods is its local sensing capability of loosened brackets. The method can further discriminate the two major failure modes within a bracket: panel-joint loosening and bracket-joint loosening. The theoretical explanation of the attenuation-based method is performed using micro-contact theory and structural/internal damping principles, followed by parametric model studies and appropriate hypothesis testing.

  15. Room temperature shear properties of the strain isolator pad for the shuttle thermal protection system

    NASA Technical Reports Server (NTRS)

    Sawyer, J. W.; Waters, W. A., Jr.

    1981-01-01

    Tests were conducted at room temperature to determine the shear properties of the strain isolator pad (SIP) material used in the thermal protection system of the space shuttle. Tests were conducted on both the .23 cm and .41 cm thick SIP material in the virgin state and after fifty fully reversed shear cycles. The shear stress displacement relationships are highly nonlinear, exhibit large hysteresis effects, are dependent on material orientation, and have a large low modulus region near the zero stress level where small changes in stress can result in large displacements. The values at the higher stress levels generally increase with normal and shear force load conditioning. Normal forces applied during the shear tests reduces the low modulus region for the material. Shear test techniques which restrict the normal movement of the material give erroneous stress displacement results. However, small normal forces do not significantly effect the shear modulus for a given shear stress. Poisson's ratio values for the material are within the range of values for many common materials. The values are not constant but vary as a function of the stress level and the previous stress history of the material. Ultimate shear strengths of the .23 cm thick SIP are significantly higher than those obtained for the .41 cm thick SIP.

  16. Effect of strain isolator pad modulus on inplane strain in Shuttle Orbiter thermal protection system tiles

    NASA Technical Reports Server (NTRS)

    Sawyer, J. W.

    1983-01-01

    The thermal protection system used on the Space Shuttle orbiter to determine strains in the reusable surface insulation tiles under simulated flight loads was investigated. The effects of changes in the strain isolator pad (SIP) moduli on the strains in the tile were evaluated. To analyze the SIP/tile system, it was necessary to conduct tests to determine inplane tension and compression modulus and inplane failure strain for the densified layer of the tiles. It is shown that densification of the LI-900 tile material increases the modulus by a factor of 6 to 10 and reduces the failure strain by about 50%. It is indicated that the inplane strain levels in the Shuttle tiles in the highly loaded regions are approximately 2 orders of magnitude lower than the failure strain of the material. It is concluded that most of the LI-900 tiles on the Shuttle could be mounted on a SIP with tensile and shear stiffnesses 10 times those of the present SIP without inplane strain failure in the tile.

  17. Backscatter x-ray development for space vehicle thermal protection systems

    SciTech Connect

    Bartha, Bence B.; Hope, Dale; Vona, Paul; Born, Martin; Corak, Tony

    2011-06-23

    The Backscatter X-Ray (BSX) imaging technique is used for various single sided inspection purposes. Previously developed BSX techniques for spray-on-foam insulation (SOFI) have been used for detecting defects in Space Shuttle External Tank foam insulation. The developed BSX hardware and techniques are currently being enhanced to advance Non-Destructive Evaluation (NDE) methods for future space vehicle applications. Various Thermal Protection System (TPS) materials were inspected using the enhanced BSX imaging techniques, investigating the capability of the method to detect voids and other discontinuities at various locations within each material. Calibration standards were developed for the TPS materials in order to characterize and develop enhanced BSX inspection capabilities. The ability of the BSX technique to detect both manufactured and natural defects was also studied and compared to through-transmission x-ray techniques. The energy of the x-ray, source to object distance, angle of x-ray, focal spot size and x-ray detector configurations were parameters playing a significant role in the sensitivity of the BSX technique to image various materials and defects. The image processing of the results also showed significant increase in the sensitivity of the technique. The experimental results showed BSX to be a viable inspection technique for space vehicle TPS systems.

  18. Methodology for Flight Relevant Arc-Jet Testing of Flexible Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Mazaheri, Alireza; Bruce, Walter E., III; Mesick, Nathaniel J.; Sutton, Kenneth

    2013-01-01

    A methodology to correlate flight aeroheating environments to the arc-jet environment is presented. For a desired hot-wall flight heating rate, the methodology provides the arcjet bulk enthalpy for the corresponding cold-wall heating rate. A series of analyses were conducted to examine the effects of the test sample model holder geometry to the overall performance of the test sample. The analyses were compared with arc-jet test samples and challenges and issues are presented. The transient flight environment was calculated for the Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Earth Atmospheric Reentry Test (HEART) vehicle, which is a planned demonstration vehicle using a large inflatable, flexible thermal protection system to reenter the Earth's atmosphere from the International Space Station. A series of correlations were developed to define the relevant arc-jet test environment to properly approximate the HEART flight environment. The computed arcjet environments were compared with the measured arc-jet values to define the uncertainty of the correlated environment. The results show that for a given flight surface heat flux and a fully-catalytic TPS, the flight relevant arc-jet heat flux increases with the arc-jet bulk enthalpy while for a non-catalytic TPS the arc-jet heat flux decreases with the bulk enthalpy.

  19. In-flight load testing of advanced shuttle thermal protection systems

    NASA Technical Reports Server (NTRS)

    Trujillo, B. M.; Meyer, R., Jr.; Sawko, P. M.

    1983-01-01

    NASA Ames Research Center has conducted in-flight airload testing of some advanced thermal protection systems (TPS) at the Dryden Flight Research Center. The two flexible TPS materials tested, felt reusable surface insulation (FRSI) and advanced flexible reusable surface insulation (AFRSI), are currently certified for use on the Shuttle orbiter. The objectives of the flight tests were to evaluate the performance of FRSI and AFRSI at simulated launch airloads and to provide a data base for future advanced TPS flight tests. Five TPS configurations were evaluated in a flow field which was representative of relatively flat areas without secondary flows. The TPS materials were placed on a fin, the Flight Test fixture (FTF), that is attached to the underside of the fuselage of an F-104 aircraft. This paper describes the test approach and techniques used and presents the results of the advanced TPS flight test. There were no failures noted during post-flight inspections of the TPS materials which were exposed to airloads 40 percent higher than the design launch airloads.

  20. Health Monitoring Technology for Thermal Protection Systems on Reusable Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Milos, Frank S.; Watters, D. G.; Heinemann, J. M.; Karunaratne, K. S.; Arnold, Jim (Technical Monitor)

    2001-01-01

    Integrated subsystem health diagnostics is an area where major improvements have been identified for potential implementation into the design of new reusable launch vehicles (RLVs) in order to reduce life cycle costs, to increase safety margins, and to improve mission reliability. This talk summarizes a joint effort between NASA Ames and industry partners to develop rapid non-contact diagnostic tools for health and performance monitoring of thermal protection systems (TPS) on future RLVs. The specific goals for TPS health monitoring are to increase the speed and reliability of TPS inspections for improved operability at lower cost. The technology being developed includes a 3-D laser scanner for examining the exterior surface of the TPS, and a subsurface microsensor suite for monitoring the health and performance of the TPS. The sensor suite consists of passive overlimit sensors and sensors for continuous parameter monitoring in flight. The sensors are integrated with radio-frequency identification (RFID) microchips to enable wireless communication of-the sensor data to an external reader that may be a hand-held scanner or a large portal. Prototypes of the laser system and both types of subsurface sensors have been developed. The laser scanner was tested on Shuttle Orbiter Columbia and was able to dimension surface chips and holes on a variety of TPS materials. The temperature-overlimit microsensor has a diameter under 0.05 inch (suitable for placement in gaps between ceramic TPS tiles) and can withstand 700 F for 15 minutes.

  1. In-Space Repair of Reinforced Carbon-Carbon Thermal Protection System Structures

    NASA Technical Reports Server (NTRS)

    Singh, Mrityunjay

    2006-01-01

    Advanced repair and refurbishment technologies are critically needed for the thermal protection system of current space transportation system as well as for future Crew Exploration Vehicles (CEV). The damage to these components could be caused by impact during ground handling or due to falling of ice or other objects during launch. In addition, in-orbit damage includes micrometeoroid and orbital debris impact as well as different factors (weather, launch acoustics, shearing, etc.) during launch and re-entry. The GRC developed GRABER (Glenn Refractory Adhesive for Bonding and Exterior Repair) material has shown multiuse capability for repair of small cracks and damage in reinforced carbon-carbon (RCC) material. The concept consists of preparing an adhesive paste of desired ceramic with appropriate additives and then applying the paste to the damaged/cracked area of the RCC composites with adhesive delivery system. The adhesive paste cures at 100-120 C and transforms into a high temperature ceramic during simulated entry conditions. A number of plasma torch and ArcJet tests were carried out to evaluate the crack repair capability of GRABER materials for Reinforced Carbon-Carbon (RCC) composites. For the large area repair applications, integrated system for tile and leading edge repair (InSTALER) have been developed. In this presentation, critical in-space repair needs and technical challenges as well as various issues and complexities will be discussed along with the plasma performance and post test characterization of repaired RCC materials.

  2. Parametric Weight Comparison of Advanced Metallic, Ceramic Tile, and Ceramic Blanket Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Myers, David E.; Martin, Carl J.; Blosser, Max L.

    2000-01-01

    A parametric weight assessment of advanced metallic panel, ceramic blanket, and ceramic tile thermal protection systems (TPS) was conducted using an implicit, one-dimensional (I-D) finite element sizing code. This sizing code contained models to account for coatings fasteners, adhesives, and strain isolation pads. Atmospheric entry heating profiles for two vehicles, the Access to Space (ATS) vehicle and a proposed Reusable Launch Vehicle (RLV), were used to ensure that the trends were not unique to a certain trajectory. Ten TPS concepts were compared for a range of applied heat loads and substructural heat capacities to identify general trends. This study found the blanket TPS concepts have the lightest weights over the majority of their applicable ranges, and current technology ceramic tiles and metallic TPS concepts have similar weights. A proposed, state-of-the-art metallic system which uses a higher temperature alloy and efficient multilayer insulation was predicted to be significantly lighter than the ceramic tile stems and approaches blanket TPS weights for higher integrated heat loads.

  3. Hypothetical Reentry Thermostructural Performance of Space Shuttle Orbiter With Missing or Eroded Thermal Protection Tiles

    NASA Technical Reports Server (NTRS)

    Ko, William L.; Gong, Leslie; Quinn, Robert D.

    2004-01-01

    This report deals with hypothetical reentry thermostructural performance of the Space Shuttle orbiter with missing or eroded thermal protection system (TPS) tiles. The original STS-5 heating (normal transition at 1100 sec) and the modified STS-5 heating (premature transition at 800 sec) were used as reentry heat inputs. The TPS missing or eroded site is assumed to be located at the center or corner (spar-rib juncture) of the lower surface of wing midspan bay 3. For cases of missing TPS tiles, under the original STS-5 heating, the orbiter can afford to lose only one TPS tile at the center or two TPS tiles at the corner (spar-rib juncture) of the lower surface of wing midspan bay 3. Under modified STS-5 heating, the orbiter cannot afford to lose even one TPS tile at the center or at the corner of the lower surface of wing midspan bay 3. For cases of eroded TPS tiles, the aluminum skin temperature rises relatively slowly with the decreasing thickness of the eroded central or corner TPS tile until most of the TPS tile is eroded away, and then increases exponentially toward the missing tile case.

  4. Estimation of surface heat flux for ablation and charring of thermal protection material

    NASA Astrophysics Data System (ADS)

    Qian, Wei-qi; He, Kai-feng; Zhou, Yu

    2016-07-01

    Ablation of the thermal protection material of the reentry hypersonic flight vehicle is a complex physical and chemical process. To estimate the surface heat flux from internal temperature measurement is much more complex than the conventional inverse heat conduction problem case. In the paper, by utilizing a two-layer pyrogeneration-plane ablation model to model the ablation and charring of the material, modifying the finite control volume method to suit for the numerical simulation of the heat conduction equation with variable-geometry, the CGM along with the associated adjoint problem is developed to estimate the surface heat flux. This estimation method is verified with a numerical example at first, the results show that the estimation method is feasible and robust. The larger is the measurement noise, the greater is the deviation of the estimated result from the exact value, and the measurement noise of ablated surface position has a significant and more direct influence on the estimated result of surface heat flux. Furthermore, the estimation method is used to analyze the experimental data of ablation of blunt Carbon-phenolic material Narmco4028 in an arc-heater. It is shown that the estimated surface heat flux agrees with the heating power value of the arc-heater, and the estimation method is basically effective and potential to treat the engineering heat conduction problem with ablation.

  5. Backscatter X-Ray Development for Space Vehicle Thermal Protection Systems

    NASA Astrophysics Data System (ADS)

    Bartha, Bence B.; Hope, Dale; Vona, Paul; Born, Martin; Corak, Tony

    2011-06-01

    The Backscatter X-Ray (BSX) imaging technique is used for various single sided inspection purposes. Previously developed BSX techniques for spray-on-foam insulation (SOFI) have been used for detecting defects in Space Shuttle External Tank foam insulation. The developed BSX hardware and techniques are currently being enhanced to advance Non-Destructive Evaluation (NDE) methods for future space vehicle applications. Various Thermal Protection System (TPS) materials were inspected using the enhanced BSX imaging techniques, investigating the capability of the method to detect voids and other discontinuities at various locations within each material. Calibration standards were developed for the TPS materials in order to characterize and develop enhanced BSX inspection capabilities. The ability of the BSX technique to detect both manufactured and natural defects was also studied and compared to through-transmission x-ray techniques. The energy of the x-ray, source to object distance, angle of x-ray, focal spot size and x-ray detector configurations were parameters playing a significant role in the sensitivity of the BSX technique to image various materials and defects. The image processing of the results also showed significant increase in the sensitivity of the technique. The experimental results showed BSX to be a viable inspection technique for space vehicle TPS systems.

  6. Optimization of thermal protection systems for the space shuttle vehicle. Volume 1: Final report

    NASA Technical Reports Server (NTRS)

    1972-01-01

    A study performed to continue development of computational techniques for the Space Shuttle Thermal Protection System is reported. The resulting computer code was used to perform some additional optimization studies on several TPS configurations. The program was developed in Fortran 4 for the CDC 6400, and it was converted to Fortran 5 to be used for the Univac 1108. The computational methodology is developed in modular fashion to facilitate changes and updating of the techniques and to allow overlaying the computer code to fit into approximately 131,000 octal words of core storage. The program logic involves subroutines which handle input and output of information between computer and user, thermodynamic stress, dynamic, and weight/estimate analyses of a variety of panel configurations. These include metallic, ablative, RSI (with and without an underlying phase change material), and a thermodynamic analysis only of carbon-carbon systems applied to the leading edge and flat cover panels. Two different thermodynamic analyses are used. The first is a two-dimensional, explicit precedure with variable time steps which is used to describe the behavior of metallic and carbon-carbon leading edges. The second is a one-dimensional implicity technique used to predict temperature in the charring ablator and the noncharring RSI. The latter analysis is performed simply by suppressing the chemical reactions and pyrolysis of the TPS material.

  7. A Collaborative Analysis Tool for Thermal Protection Systems for Single Stage to Orbit Launch Vehicles

    NASA Technical Reports Server (NTRS)

    Alexander, Reginald A.; Stanley, Thomas Troy

    1999-01-01

    Presented is a design tool and process that connects several disciplines which are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system and in the case of SSTO vehicles with air breathing propulsion, which is currently being studied by the National Aeronautics and Space Administration (NASA); the thermal protection system (TPS) is linked directly to almost every major system. The propulsion system pushes the vehicle to velocities on the order of 15 times the speed of sound in the atmosphere before pulling up to go to orbit which results high temperatures on the external surfaces of the vehicle. Thermal protection systems to maintain the structural integrity of the vehicle must be able to mitigate the heat transfer to the structure and be lightweight. Herein lies the interdependency, in that as the vehicle's speed increases, the TPS requirements are increased. And as TPS masses increase the effect on the propulsion system and all other systems is compounded. To adequately determine insulation masses for a vehicle such as the one described above, the aeroheating loads must be calculated and the TPS thicknesses must be calculated for the entire vehicle. To accomplish this an ascent or reentry trajectory is obtained using the computer code Program to Optimize Simulated Trajectories (POST). The trajectory is then used to calculate the convective heat rates on several locations on the vehicles using the Miniature Version of the JA70 Aerodynamic Heating Computer Program (MINIVER). Once the heat rates are defined for each body point on the vehicle, then insulation thicknesses that are required to maintain the vehicle within structural limits are calculated using Systems Improved Numerical Differencing Analyzer (SINDA) models. If the TPS masses are too heavy for the performance of the vehicle the process may be repeated altering the trajectory or some other input to

  8. Automated sample-changing robot for solution scattering experiments at the EMBL Hamburg SAXS station X33.

    PubMed

    Round, A R; Franke, D; Moritz, S; Huchler, R; Fritsche, M; Malthan, D; Klaering, R; Svergun, D I; Roessle, M

    2008-10-01

    There is a rapidly increasing interest in the use of synchrotron small-angle X-ray scattering (SAXS) for large-scale studies of biological macromolecules in solution, and this requires an adequate means of automating the experiment. A prototype has been developed of an automated sample changer for solution SAXS, where the solutions are kept in thermostatically controlled well plates allowing for operation with up to 192 samples. The measuring protocol involves controlled loading of protein solutions and matching buffers, followed by cleaning and drying of the cell between measurements. The system was installed and tested at the X33 beamline of the EMBL, at the storage ring DORIS-III (DESY, Hamburg), where it was used by over 50 external groups during 2007. At X33, a throughput of approximately 12 samples per hour, with a failure rate of sample loading of less than 0.5%, was observed. The feedback from users indicates that the ease of use and reliability of the user operation at the beamline were greatly improved compared with the manual filling mode. The changer is controlled by a client-server-based network protocol, locally and remotely. During the testing phase, the changer was operated in an attended mode to assess its reliability and convenience. Full integration with the beamline control software, allowing for automated data collection of all samples loaded into the machine with remote control from the user, is presently being implemented. The approach reported is not limited to synchrotron-based SAXS but can also be used on laboratory and neutron sources.

  9. Modeling of ultrasonic and terahertz radiations in defective tiles for condition monitoring of thermal protection systems

    NASA Astrophysics Data System (ADS)

    Kabiri Rahani, Ehsan

    Condition based monitoring of Thermal Protection Systems (TPS) is necessary for safe operations of space shuttles when quick turn-around time is desired. In the current research Terahertz radiation (T-ray) has been used to detect mechanical and heat induced damages in TPS tiles. Voids and cracks inside the foam tile are denoted as mechanical damage while property changes due to long and short term exposures of tiles to high heat are denoted as heat induced damage. Ultrasonic waves cannot detect cracks and voids inside the tile because the tile material (silica foam) has high attenuation for ultrasonic energy. Instead, electromagnetic terahertz radiation can easily penetrate into the foam material and detect the internal voids although this electromagnetic radiation finds it difficult to detect delaminations between the foam tile and the substrate plate. Thus these two technologies are complementary to each other for TPS inspection. Ultrasonic and T-ray field modeling in free and mounted tiles with different types of mechanical and thermal damages has been the focus of this research. Shortcomings and limitations of FEM method in modeling 3D problems especially at high-frequencies has been discussed and a newly developed semi-analytical technique called Distributed Point Source Method (DPSM) has been used for this purpose. A FORTRAN code called DPSM3D has been developed to model both ultrasonic and electromagnetic problems using the conventional DPSM method. This code is designed in a general form capable of modeling a variety of geometries. DPSM has been extended from ultrasonic applications to electromagnetic to model THz Gaussian beams, multilayered dielectrics and Gaussian beam-scatterer interaction problems. Since the conventional DPSM has some drawbacks, to overcome it two modification methods called G-DPSM and ESM have been proposed. The conventional DPSM in the past was only capable of solving time harmonic (frequency domain) problems. Time history was

  10. Post-Flight Assessment of Avcoat Thermal Protection System for the Exploration Flight Test-1

    NASA Technical Reports Server (NTRS)

    Bose, Deepak; Santos, Jose; Rodriguez, Erika; Mahzari, Milad; Remark, Brian; Muppidi, Suman

    2016-01-01

    On December 5, 2014 NASA conducted the first flight test of its next generation human-class Orion spacecraft. The flight was called the Exploration Flight Test -1 (EFT-1) which lasted for 4 hours and culminated into a re-entry trajectory at 9 km/s. This flight test of the 5-meter Orion Crew Module demonstrated various sub-systems including the Avcoat ablative thermal protection system (TPS) on the heat shield. The Avcoat TPS had been developed from the Apollo-era recipe with a few key modifications. The engineering for thermal sizing was supported by modeling, analysis, and ground tests in arc jet facilities. This paper will describe a postlfight analysis plan and present results from post-recovery inspections, data analysis from embedded sensors, TPS sample extraction and characterization in the laboratory. After the recovery of the vehicle, a full photographic survey and surface scans of the TPS were performed. The recovered vehicle showed physical evidence of flow disturbances, varying degrees of surface roughness, and excessive recession downstream of compression pads. The TPS recession was measured at more than 200 locations of interest on the Avcoat surface. The heat shield was then processed for sample extraction prior to TPS removal using the 7-Axis Milling machine at Marshall Space Flight Center. Around 182 rectangular TPS samples were extracted for subsequent analysis and investigation. The final paper will also present results of sample analysis. The planned investigation includes sidewall imaging, followed by image analysis to characterize TPS response by quantifying different layers in the char and pyrolysis zones. A full postmortem of the instrumentation and sensor ports will also be performed to confirm no adverse effects due to the sensors themselves. A subset of the samples will undergo structural testing and perform detailed characterization of any cracks and integrity of gore seams. Finally, the material will be characterized with layer

  11. High-Temperature Structures, Adhesives, and Advanced Thermal Protection Materials for Next-Generation Aeroshell Design

    NASA Technical Reports Server (NTRS)

    Collins, Timothy J.; Congdon, William M.; Smeltzer, Stanley S.; Whitley, Karen S.

    2005-01-01

    The next generation of planetary exploration vehicles will rely heavily on robust aero-assist technologies, especially those that include aerocapture. This paper provides an overview of an ongoing development program, led by NASA Langley Research Center (LaRC) and aimed at introducing high-temperature structures, adhesives, and advanced thermal protection system (TPS) materials into the aeroshell design process. The purpose of this work is to demonstrate TPS materials that can withstand the higher heating rates of NASA's next generation planetary missions, and to validate high-temperature structures and adhesives that can reduce required TPS thickness and total aeroshell mass, thus allowing for larger science payloads. The effort described consists of parallel work in several advanced aeroshell technology areas. The areas of work include high-temperature adhesives, high-temperature composite materials, advanced ablator (TPS) materials, sub-scale demonstration test articles, and aeroshell modeling and analysis. The status of screening test results for a broad selection of available higher-temperature adhesives is presented. It appears that at least one (and perhaps a few) adhesives have working temperatures ranging from 315-400 C (600-750 F), and are suitable for TPS-to-structure bondline temperatures that are significantly above the traditional allowable of 250 C (482 F). The status of mechanical testing of advanced high-temperature composite materials is also summarized. To date, these tests indicate the potential for good material performance at temperatures of at least 600 F. Application of these materials and adhesives to aeroshell systems that incorporate advanced TPS materials may reduce aeroshell TPS mass by 15% - 30%. A brief outline is given of work scheduled for completion in 2006 that will include fabrication and testing of large panels and subscale aeroshell test articles at the Solar-Tower Test Facility located at Kirtland AFB and operated by Sandia

  12. Manufacture and engine test of advanced oxide dispersion strengthened alloy turbine vanes. [for space shuttle thermal protection

    NASA Technical Reports Server (NTRS)

    Bailey, P. G.

    1977-01-01

    Oxide-Dispersion-strengthened (ODS) Ni-Cr-Al alloy systems were exploited for turbine engine vanes which would be used for the space shuttle thermal protection system. Available commercial and developmental advanced ODS alloys were evaluated, and three were selected based on established vane property goals and manufacturing criteria. The selected alloys were evaluated in an engine test. Candidate alloys were screened by strength, thermal fatigue resistance, oxidation and sulfidation resistance. The Ni-16Cr (3 to 5)Al-ThO2 system was identified as having attractive high temperature oxidation resistance. Subsequent work also indicated exceptional sulfidation resistance for these alloys.

  13. The capacity for thermal protection of photosynthetic electron transport varies for different monoterpenes in Quercus ilex.

    PubMed

    Copolovici, Lucian O; Filella, Iolanda; Llusià, Joan; Niinemets, Ulo; Peñuelas, Josep

    2005-09-01

    Heat stress resistance of foliar photosynthetic apparatus was investigated in the Mediterranean monoterpene-emitting evergreen sclerophyll species Quercus ilex. Leaf feeding with fosmidomycin, which is a specific inhibitor of the chloroplastic isoprenoid synthesis pathway, essentially stopped monoterpene emission and resulted in the decrease of the optimum temperature of photosynthetic electron transport from approximately 38 degrees C to approximately 30 degrees C. The heat stress resistance was partly restored by fumigation with 4 to 5 nmol mol(-1) air concentrations of monoterpene alpha-pinene but not with fumigations with monoterpene alcohol alpha-terpineol. Analyses of monoterpene physicochemical characteristics demonstrated that alpha-pinene was primarily distributed to leaf gas and lipid phases, while alpha-terpineol was primarily distributed to leaf aqueous phase. Thus, for a common monoterpene uptake rate, alpha-terpineol is less efficient in stabilizing membrane liquid-crystalline structure and as an antioxidant in plant membranes. Furthermore, alpha-terpineol uptake rate (U) strongly decreased with increasing temperature, while the uptake rates of alpha-pinene increased with increasing temperature, providing a further explanation of the lower efficiency of thermal protection by alpha-terpineol. The temperature-dependent decrease of alpha-terpineol uptake was both due to decreases in stomatal conductance, g(w), and increased volatility of alpha-terpineol at higher temperature that decreased the monoterpene diffusion gradient between the ambient air (F(A)) and leaf (F(I); U = g(w)[F(A) - F(I)]). Model analyses suggested that alpha-pinene reacted within the leaf at higher temperatures, possibly within the lipid phase, thereby avoiding the decrease in diffusion gradient, F(A) - F(I). Thus, these data contribute to the hypothesis of the antioxidative protection of leaf membranes during heat stress by monoterpenes. These data further suggest that fumigation

  14. Conformal Ablative Thermal Protection System for Planetary and Human Exploration Missions

    NASA Technical Reports Server (NTRS)

    Beck, R.; Arnold, J.; Gasch, M.; Stackpole, M.; Wercinski, R.; Venkatapathy, E.; Fan, W.; Thornton, J; Szalai, C.

    2012-01-01

    The Office of Chief Technologist (OCT), NASA has identified the need for research and technology development in part from NASAs Strategic Goal 3.3 of the NASA Strategic Plan to develop and demonstrate the critical technologies that will make NASAs exploration, science, and discovery missions more affordable and more capable. Furthermore, the Game Changing Development Program (GCDP) is a primary avenue to achieve the Agencys 2011 strategic goal to Create the innovative new space technologies for our exploration, science, and economic future. In addition, recently released NASA Space Technology Roadmaps and Priorities, by the National Research Council (NRC) of the National Academy of Sciences stresses the need for NASA to invest in the very near term in specific EDL technologies. The report points out the following challenges (Page 2-38 of the pre-publication copy released on February 1, 2012): Mass to Surface: Develop the ability to deliver more payload to the destination. NASA's future missions will require ever-greater mass delivery capability in order to place scientifically significant instrument packages on distant bodies of interest, to facilitate sample returns from bodies of interest, and to enable human exploration of planets such as Mars. As the maximum mass that can be delivered to an entry interface is fixed for a given launch system and trajectory design, the mass delivered to the surface will require reductions in spacecraft structural mass more efficient, lighter thermal protection systems more efficient lighter propulsion systems and lighter, more efficient deceleration systems. Surface Access: Increase the ability to land at a variety of planetary locales and at a variety of times. Access to specific sites can be achieved via landing at a specific location(s) or transit from a single designated landing location, but it is currently infeasible to transit long distances and through extremely rugged terrain, requiring landing close to the site of

  15. Development of Natural Flaw Samples for Evaluating Nondestructive Testing Methods for Foam Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Workman, Gary L.; Davis, Jason; Farrington, Seth; Walker, James

    2007-01-01

    Low density polyurethane foam has been an important insulation material for space launch vehicles for several decades. The potential for damage from foam breaking away from the NASA External Tank was not realized until the foam impacts on the Columbia Orbiter vehicle caused damage to its Leading Edge thermal protection systems (TPS). Development of improved inspection techniques on the foam TPS is necessary to prevent similar occurrences in the future. Foamed panels with drilled holes for volumetric flaws and Teflon inserts to simulate debonded conditions have been used to evaluate and calibrate nondestructive testing (NDT) methods. Unfortunately the symmetric edges and dissimilar materials used in the preparation of these simulated flaws provide an artificially large signal while very little signal is generated from the actual defects themselves. In other words, the same signal are not generated from the artificial defects in the foam test panels as produced when inspecting natural defect in the ET foam TPS. A project to create more realistic voids similar to what actually occurs during manufacturing operations was began in order to improve detection of critical voids during inspections. This presentation describes approaches taken to create more natural voids in foam TPS in order to provide a more realistic evaluation of what the NDT methods can detect. These flaw creation techniques were developed with both sprayed foam and poured foam used for insulation on the External Tank. Test panels with simulated defects have been used to evaluate NDT methods for the inspection of the External Tank. A comparison of images between natural flaws and machined flaws generated from backscatter x-ray radiography, x-ray laminography, terahertz imaging and millimeter wave imaging show significant differences in identifying defect regions.

  16. Acousto-optic signature analysis for inspection of the orbiter thermal protection tile bonds

    NASA Technical Reports Server (NTRS)

    Rodriguez, Julio G.; Tow, D. M.; Barna, B. A.

    1990-01-01

    The goal of this research is to develop a viable NDE technique for the inspection of orbiter thermal protection system (TPS) tile bonds. Phase 2, discussed here, concentrated on developing an empirical understanding of the bonded and unbonded vibration signatures of acreage tiles. Controlled experiments in the laboratory have provided useful information on the dynamic response of TPS tiles. It has been shown that several signatures are common to all the pedigree tiles. This degree of consistency in the tile-SIP (strain isolation pad) dynamic response proves that an unbond can be detected for a known tile and establish the basis for extending the analysis capability to arbitrary tiles for which there are no historical data. The field tests of the noncontacting laser acoustic sensor system, conducted at the Kennedy Space Center (KSC), investigated the vibrational environment of the Orbiter Processing Facility (OPF) and its effect on the measurement and analysis techniques being developed. The data collected showed that for orbiter locations, such as the body flap and elevon, the data analysis scheme, and/or the sensor, will require modification to accommodate the ambient motion. Several methods were identified for accomplishing this, and a solution is seen as readily achievable. It was established that the tile response was similar to that observed in the laboratory. Of most importance, however, is that the field environment will not affect the physics of the dynamic response that is related to bond condition. All of this information is fundamental to any future design and development of a prototype system.

  17. A low noise and high precision linear power supply with thermal foldback protection

    NASA Astrophysics Data System (ADS)

    Carniti, P.; Cassina, L.; Gotti, C.; Maino, M.; Pessina, G.

    2016-05-01

    A low noise and high precision linear power supply was designed for use in rare event search experiments with macrobolometers. The circuit accepts at the input a "noisy" dual supply voltage up to ±15 V and gives at the output precise, low noise, and stable voltages that can be set between ±3.75 V and ±12.5 V in eight 1.25 V steps. Particular care in circuit design, component selection, and proper filtering results in a noise spectral density of 50 nV / √{ Hz } at 1 Hz and 20 nV / √{ Hz } white when the output is set to ±5 V. This corresponds to 125 nV RMS (0.8 μV peak to peak) between 0.1 Hz and 10 Hz, and 240 nV RMS (1.6 μV peak to peak) between 0.1 Hz and 100 Hz. The power supply rejection ratio (PSRR) of the circuit is 100 dB at low frequency, and larger than 40 dB up to high frequency, thanks to a proper compensation design. Calibration allows to reach a precision in the absolute value of the output voltage of ±70 ppm, or ±350 μV at ±5 V, and to reduce thermal drifts below ±1 ppm/∘C in the expected operating range. The maximum peak output current is about 6 A from each output. An original foldback protection scheme was developed that dynamically limits the maximum output current to keep the temperature of the output transistors within their safe operating range. An add-on card based on an ARM Cortex-M3 microcontroller is devoted to the monitoring and control of all circuit functionalities and provides remote communication via CAN bus.

  18. Simultaneously Harvesting Thermal and Mechanical Energies based on Flexible Hybrid Nanogenerator for Self-Powered Cathodic Protection.

    PubMed

    Zhang, Hulin; Zhang, Shangjie; Yao, Guang; Huang, Zhenlong; Xie, Yuhang; Su, Yuanjie; Yang, Weiqing; Zheng, Chunhua; Lin, Yuan

    2015-12-30

    Metal corrosion occurs anytime and anywhere in nature and the corrosion prevention has a great significance everywhere in national economic development and daily life. Here, we demonstrate a flexible hybrid nanogenerator (NG) that is capable of simultaneously or individually harvesting ambient thermal and mechanical energies and used for a self-powered cathodic protection (CP) system without using an external power source. Because of its double peculiarities of both pyroelectric and piezoelectric properties, a polarized poly(vinylidene fluoride) (PVDF) film-based NG was constructed to scavenge both thermal and mechanical energies. As a supplementary, a triboelectric NG was constructed below the pyro/piezoelectric NG to grab ambient mechanical energy. The output power of the fabricated hybrid NG can be directly used to protect the metal surface from the chemical corrosion. Our results not only verify the feasibility of self-powered CP-based NGs, but also expand potential self-powered applications.

  19. Methods of Evaluating Protective Clothing Relative to Heat and Cold Stress: Thermal Manikin, Biomedical Modeling, and Human Testing

    DTIC Science & Technology

    2011-01-01

    code) 2011 Journal Article-Journal of Occup and Env Hygiene Methods of Evaluating Protective Clothing Relative to Heat and cold Stress: Thermal...of Occupational and Environmental Hygiene , 8: 588–599 ISSN: 1545-9624 print / 1545-9632 online DOI: 10.1080/15459624.2011.613291 Methods of...increases insulation and reduces water vapor permeability, both of which 588 Journal of Occupational and Environmental Hygiene October 2011 D ow nl oa de

  20. Development of acceptance criteria for batches of silane primer for external tank thermal protection system bonding applications

    NASA Technical Reports Server (NTRS)

    Mikes, F.

    1984-01-01

    Silane primers for use as thermal protection on external tanks were subjected to various analytic techniques to determine the most effective testing method for silane lot evaluation. The analytic methods included high performance liquid chromatography, gas chromatography, thermogravimetry (TGA), and fourier transform infrared spectroscopy (FTIR). It is suggested that FTIR be used as the method for silane lot evaluation. Chromatograms, TGA profiles, bar graphs showing IR absorbances, and FTIR spectra are presented.