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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. F-15B in flight with X-33 Thermal Protection Systems (TPS) on Flight Test Fixture

    NASA Technical Reports Server (NTRS)

    1998-01-01

    In-flight photo of the NASA F-15B used in tests of the X-33 Thermal Protection System (TPS) materials. Flying at subsonic speeds, the F-15B tests measured the air loads on the proposed X-33 protective materials. In contrast, shock loads testing investigated the local impact of the supersonic shock wave itself on the TPS materials. Similar tests had been done in 1985 for the space shuttle tiles, using an F-104 aircraft.

  3. Aerothermal Test of Thermal Protection Systems for X-33 Reusable Launch Vehicle

    NASA Technical Reports Server (NTRS)

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

    1999-01-01

    An array of metallic Thermal Protection System (TPS) panels developed for the windward surface of the X-33 vehicle was tested in the 8-Foot High Temperature Tunnel at the NASA Langley Research Center. These tests were the first aerothermal tests of an X-33 TPS array and the test results will be used to validate the TPS for the X-33 flight program. Specifically, the tests evaluated the structural and thermal performance of the TPS, the effectiveness of the high temperature seals between adjacent panels and the durability of the TPS under realistic aerothermal flight conditions. The effect of varying panel-to-panel step heights, intentional damage to the seals between adjacent panels, and the use of secondary seals were also investigated during the test program. The metallic TPS developed for the windward surface of the X-33, the blanket TPS developed to protect the leeward surfaces of the X-33, and the test program in the 8-Foot High Temperature Tunnel are presented and discussed.

  4. 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.

  5. Closeup of F-15B Flight Test Fixture (FTF) with X-33 Thermal Protection Systems (TPS)

    NASA Technical Reports Server (NTRS)

    1998-01-01

    A close up of the Flight Test Fixture II, mounted on the underside of the F-15B Aerodynamic Flight Facility aircraft. The Thermal Protection System (TPS)samples, which included metallic Inconel tiles, soft Advanced Flexible Reusable Surface Insulation tiles, and sealing materials, were attached to the forward-left side position of the test fixture. In-flight video from the aircraft's on-board video system, as well as chase aircraft photos and video footage, documented the condition of the TPS during flights. Surface pressures over the TPS was measured by thermocouples contained in instrumentation 'islands,' to document shear and shock loads.

  6. 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

  7. Task 4 supporting technology. Part 2: Detailed test plan for thermal seals. Thermal seals evaluation, improvement and test. CAN8-1, Reusable Launch Vehicle (RLV), advanced technology demonstrator: X-33. Leading edge and seals thermal protection system technology demonstration

    NASA Technical Reports Server (NTRS)

    Hogenson, P. A.; Lu, Tina

    1995-01-01

    The objective is to develop the advanced thermal seals to a technology readiness level (TRL) of 6 to support the rapid turnaround time and low maintenance requirements of the X-33 and the future reusable launch vehicle (RLV). This program is divided into three subtasks: (1) orbiter thermal seals operation history review; (2) material, process, and design improvement; and (3) fabrication and evaluation of the advanced thermal seals.

  8. 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.

  9. 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.

  10. 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

  11. 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

  12. 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.

  13. Design, Development, And Testing of Umbilical System Mechanisms for the X-33 Advanced Technology Demonstrator

    NASA Technical Reports Server (NTRS)

    Littlefield, Alan C.; Melton, Gregory S.

    2000-01-01

    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.

  14. 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.

  15. [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).

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. X-33 Turbulent Aeroheating Measurements and Predictions

    NASA Technical Reports Server (NTRS)

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

    2002-01-01

    Measurements and predictions of the X-33 turbulent aeroheating environment have been performed for Mach 6, perfect-gas air conditions. The purpose of this investigation was to compare turbulent aeroheating predictions from two Navier-Stokes codes, LAURA and GASP, with each other and with experimental data in which turbulent flow was produced through either natural transition or forced transition using roughness elements. The wind tunnel testing was conducted at free stream Reynolds numbers of 0.72 x 10(exp 7)/m to 2.4 x 10(exp 7)/m (2.2 x 10(exp 6)/ft to 7.3 x 10(exp 6)/ft) on 0.254 m (10.0-in.) X-33 models at alpha = 40 deg with smooth surfaces, smooth surfaces with discrete trips, and surfaces with simulated bowed thermal protection system panels. Turbulent flow was produced by the discrete trips and bowed panels for all but the lowest Reynolds number, while 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 agreed to within the experimental accuracy (+/= 15%) of the test technique. 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.

  3. 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.

  4. 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.

  5. 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).

  6. The Lifting Body Legacy...X-33

    NASA Technical Reports Server (NTRS)

    Barret, Chris

    1999-01-01

    NASA has a technology program in place to enable the development of a next generation Reusable Launch Vehicle that will carry our future payloads into orbit at a much-reduced cost. The VentureStar, Lifting Body (LB) flight vehicle, is one of the potential reusable launch vehicle configurations being studied. A LB vehicle has no wings and derives its lift solely from the shape of its body, and has the unique advantages of superior volumetric efficiency, better aerodynamic efficiency at high angles-of-attack and hypersonic speeds, and reduced thermal protection system weight. Classically, in a ballistic vehicle, drag has been employed to control the level of deceleration in reentry. In the LB, lift enables the vehicle to decelerate at higher altitudes for the same velocity and defines the reentry corridor which includes a greater cross range. This paper outlines the flight stability and control aspects of our LB heritage which was utilized in the design of the VentureStar LB and its test version, the X-33. NASA and the U.S. Air Force have a rich heritage of LB vehicle design and flight experience. In the initial LB Program, eight LB's were built and over 225 LB test flights were conducted through 1975. Three LB series were most significant in the advancement of today's LB technolocy: the M2-F; the HL-10; and the X-24 series. The M2-F series was designed by NASA Ames Research Center, the HL-10 series by NASA Langley Research Center, and the X-24 series by the U. S. Air Force. LB vehicles are alive again today with the X- 33, X-38, and VentureStar.

  7. Artist concept of X-33 and Reusable Launch Vehicle (RLV)

    NASA Technical Reports Server (NTRS)

    1997-01-01

    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 had 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 tank, and the resulting cost increase and time delay, the X-33 program was cancelled in February 2001.

  8. 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.

  9. 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.

  10. Thermal Protection Materials Development

    NASA Technical Reports Server (NTRS)

    Selvaduray, Guna; Cox, Michael

    1998-01-01

    The main portion of this contract year was spent on the development of materials for high temperature applications. In particular, thermal protection materials were constantly tested and evaluated for thermal shock resistance, high-temperature dimensional stability, and tolerance to hostile environmental effects. The analytical laboratory at the Thermal Protection Materials Branch (TPMB), NASA-Ames played an integral part in the process of materials development of high temperature aerospace applications. The materials development focused mainly on the determination of physical and chemical characteristics of specimens from the various research programs.

  11. 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.

  12. 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

  13. 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.

  14. 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.

  15. Ablative thermal protection systems

    NASA Technical Reports Server (NTRS)

    Vaniman, J.; Fisher, R.; Wojciechowski, C.; Dean, W.

    1983-01-01

    The procedures used to establish the TPS (thermal protection system) design of the SRB (solid rocket booster) element of the Space Shuttle vehicle are discussed. A final evaluation of the adequacy of this design will be made from data obtained from the first five Shuttle flights. Temperature sensors installed at selected locations on the SRB structure covered by the TPS give information as a function of time throughout the flight. Anomalies are to be investigated and computer design thermal models adjusted if required. In addition, the actual TPS ablator material loss is to be measured after each flight and compared with analytically determined losses. The analytical methods of predicting ablator performance are surveyed.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. Thermal protection apparatus

    DOEpatents

    Bennett, Gloria A.; Moore, Troy K.

    1988-01-01

    An apparatus for thermally protecting heat sensitive components of tools. The apparatus comprises a Dewar for holding the heat sensitive components. The Dewar has spaced-apart inside and outside walls, an open top end and a bottom end. An insulating plug is located in the top end. The inside wall has portions defining an inside wall aperture located at the bottom of the Dewar and the outside wall has portions defining an outside wall aperture located at the bottom of the Dewar. A bottom connector has inside and outside components. The inside component sealably engages the inside wall aperture and the outside component sealably engages the outside wall aperture. The inside component is operatively connected to the heat sensitive components and to the outside component. The connections can be made with optical fibers or with electrically conducting wires.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. Orbiter Thermal Protection System Development

    NASA Technical Reports Server (NTRS)

    Greenshields, D. H.

    1977-01-01

    The development of the Space Shuttle Orbiter Thermal Protection System (TPS) is traced from concept definition, through technical development, to final design and qualification for manned flight. A sufficiently detailed description of the TPS design is presented to support an indepth discussion of the key issues encountered in conceptual design, materials development, and structural integration. Emphasis is placed on the unique combination of requirements which resulted in the use not only of revolutionary design concepts and materials, but also of unique design criteria, newly developed analysis, testing and manufacturing methods, and finally of an unconventional approach to system certification for operational flight. The conclusion is drawn that a significant advance in all areas of thermal protection system development has been achieved which results in a highly efficient, flexible, and cost-effective thermal protection system for the Orbiter of the Space Shuttle System.

  12. 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.

  13. 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.

  14. X-33: The operational challenge

    SciTech Connect

    Hickman, R.A.

    1996-03-01

    In a cooperative effort with industry, NASA has embarked upon an effort to develop the next generation of reusable space transportation systems. Operational efficiency is the key attribute that will determine the economic viability of any reusable launch vehicle program. This factor not only controls the recurring cost per flight, but also determines the number of flights over which the program{close_quote}s fixed cost and overhead can be amortized. A single stage to orbit reusable vehicle has the obvious advantage of eliminating expendable hardware and the need for mating stages and interface verification; however, these factors alone are not sufficient to generate the order of magnitude improvement in operational efficiency necessary to justify a new launch system. To achieve these objectives, a clear understanding is needed of the factors that have allowed aircraft to demonstrate operability and reliability levels five orders of magnitude better than launch systems. This understanding must then be reflected in analytical tools which can support detailed design trades. Even with the best of analytical environments, the difficulty in developing an operationally efficient single stage to orbit vehicle should not be underestimated. New technologies and operational concepts must be creatively combined into a design that will initially involve significant development risk. The risk derives not so much from the possibility that the goals cannot be achieved, but from the recognition that any new program with a high investment in new technology is unlikely to succeed without an opportunity to mature these technologies and to correct deficiencies in the basic design. The X-33 Program represents the first comprehensive testbed developed for the investigation and demonstration of operational technologies and design features prior to final system design and deployment. (Abstract Truncated)

  15. Prestressed Thermal-Protection Panels

    NASA Technical Reports Server (NTRS)

    Dunn, T. J.

    1985-01-01

    Panels held securely with minimum of mounting hardware. Each panel held in place by single screw that pulls it into flat shape from its original shallow-dish shape. Shape and prestressing make panel stiff: resists vibration and withstands large mechanical loads. Panel shape and mounting arrangement not limited to thermal-protection systems but also used on aircraft, building walls, or wherever large surfaces must be covered with stiff, flat sheets easily removed for maintenance.

  16. 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.

  17. 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.

  18. 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.}

  19. [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.

  20. The X-33 Extended Flight Test Range

    NASA Technical Reports Server (NTRS)

    Mackall, Dale A.; Sakahara, Robert; Kremer, Steven E.

    1998-01-01

    Development of an extended test range, with range instrumentation providing continuous vehicle communications, is required to flight-test the X-33, a scaled version of a reusable launch vehicle. The extended test range provides vehicle communications coverage from California to landing at Montana or Utah. This paper provides an overview of the approaches used to meet X-33 program requirements, including using multiple ground stations, and methods to reduce problems caused by reentry plasma radio frequency blackout. The advances used to develop the extended test range show other hypersonic and access-to-space programs can benefit from the development of the extended test range.

  1. 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.

  2. [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.

  3. 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.

  4. Hermes thermal protection system overview

    NASA Astrophysics Data System (ADS)

    Chaumette, Daniel; Cretenet, Jean-Claude

    The HERMES thermal protection system for the reentry is a new challenge for the designer. Compared to the system operational to day which is the U.S. Orbiter, the smaller size and higher cross range of HERMES are inducing higher working temperatures and a longer duration for the hot phase of the reentry. Hence the overall weight of the TPS system is comparatively more critical than on the Orbiter. On the other hand since the conception of the Orbiter a lot of new materials, namely ceramic composites, have been developped, and may lead to more efficient concepts of TPS. In the initial studies on HERMES TPS systems a lot of possibilites were considered, including External passive TPS, Hot structures, Active TPS. This selection has been now shortlisted to three basic concepts, with a number of variant or back ups still under consideration: • Ceramic composites hot structures for the nose, leading edges, fins and control surfaces • External insulation : composite ceramic shingles covering a lightweight thermal insulation (or rigid surface insulation (tiles) as a back up solution) for the hot undersurfaces and part of the upper surface. • Flexible surface insulation for the lower temperature upper surfaces. The paper presents details on the concepts being studied, the optimisation methods and the concept selection criteria.

  5. 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.

  6. 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.

  7. 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.

  8. Apollo experience report: Thermal protection subsystem

    NASA Technical Reports Server (NTRS)

    Pavlosky, J. E.; St.leger, L. G.

    1974-01-01

    The Apollo command module was the first manned spacecraft to be designed to enter the atmosphere of the earth at lunar-return velocity, and the design of the thermal protection subsystem for the resulting entry environment presented a major technological challenge. Brief descriptions of the Apollo command module thermal design requirements and thermal protection configuration, and some highlights of the ground and flight testing used for design verification of the system are presented. Some of the significant events that occurred and decisions that were made during the program concerning the thermal protection subsystem are discussed.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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

  19. Reusable thermal protection system development: A prospective

    NASA Technical Reports Server (NTRS)

    Goldstein, Howard

    1992-01-01

    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.

  20. X-33 Tank Failure During Autoclave Fabrication

    NASA Technical Reports Server (NTRS)

    Nettles, Alan T.; Munafo, Paul (Technical Monitor)

    2001-01-01

    The composite liquid hydrogen tank (tank #1 of 2) for the X-33 flight vehicle is made up of four lobes that have a sandwich construction, bonded to a frame of longerons. Lobes 1 and 4 showed local disbonds to the longerons they were bonded to. The 'bad' areas were cut away and patched with new material. The new material was cured by placing the entire tank in a heated autoclave with no pressure. Upon removal from the autoclave, it was noted that lobe 1 had severe skin/core disbonds on the inner and outer skins. The skins on this lobe were cracked as well. The core was disbonded from the inner skin across the entire acreage, except for spots around the lobe perimeter. The outer skin was separated from the core in a region near the center of the lobe. Lobe 1 was removed from the tank on January 13, 1999. Bolts were placed through the lobe to hold it together and the cuts on the inner skin were not continuous, but 'tabs' were left for final cutting and removal. Upon closer inspection of the disbonded basesheet, it was noted that there was a lack of filleting into the honeycomb core. Good fillets are critical to bond strength.

  1. X-33 Tank Failure During Autoclave Fabrication

    NASA Technical Reports Server (NTRS)

    Nettles, Alan T.; Munafo, Paul (Technical Monitor)

    2001-01-01

    During a repair cure cycle on tank #1 of the X-33 liquid hydrogen tanks, a skin to core disbond occurred. Both the inner skin and outer skin of the lobe #1 sandwich panel was noted to have been disbonded and cracked- An investigation was undertaken to determine the cause of this failure. The investigation consisted of reviewing all of the processing data and performing testing on the failed lobe #1, as well as the other lobes, which did not fail during the cure cycle. The tests consisted of residual stress measurements in one of the intact lobes and "plug-pulls" to assess skin to core strength on all of the remaining lobes. Results showed an extremely low bondline strength due to lack of proper filleting of the adhesive, in addition, tests showed a very rapid decrease in strength with increasing temperature, as well as a further decrease in strength with a larger number of cycles. Also, the honeycomb used was not vented so pressure could build up within the cells. All of these factors appeared to be contributors to the failure.

  2. Toughened Thermal Blanket for MMOD Protection

    NASA Technical Reports Server (NTRS)

    Christiansen, Eric L.; Lear, Dana M.

    2014-01-01

    Thermal blankets are used extensively on spacecraft to provide passive thermal control of spacecraft hardware from thermal extremes encountered in space. Toughened thermal blankets have been developed that greatly improve protection from hypervelocity micrometeoroid and orbital debris (MMOD) impacts. These blankets can be outfitted if so desired with a reliable means to determine the location, depth and extent of MMOD impact damage by incorporating an impact sensitive piezoelectric film. Improved MMOD protection of thermal blankets was obtained by adding selective materials at various locations within the thermal blanket. As given in Figure 1, three types of materials were added to the thermal blanket to enhance its MMOD performance: (1) disrupter layers, near the outside of the blanket to improve breakup of the projectile, (2) standoff layers, in the middle of the blanket to provide an area or gap that the broken-up projectile can expand, and (3) stopper layers, near the back of the blanket where the projectile debris is captured and stopped. The best suited materials for these different layers vary. Density and thickness is important for the disrupter layer (higher densities generally result in better projectile breakup), whereas a highstrength to weight ratio is useful for the stopper layer, to improve the slowing and capture of debris particles.

  3. 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.

  4. 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.

  5. 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.

  6. 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

  7. 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.

  8. Aerogel Composites for Aerospace Thermal Protection

    NASA Technical Reports Server (NTRS)

    White, Susan

    2003-01-01

    Aerogel composites formed by infiltrating organic and/or inorganic aerogels into fiber matrix materials enable us to exploit the low thermal conductivity and low density of aerogels while maintaining the strength, structure and other useful properties of a porous fiber matrix. New materials for extreme heating ranges are needed to insulate future spacecraft against the extreme heat of planetary atmospheric entry, but the insulation mass must be minimized in order to maximize the payload. A reusable system passively insulates to survive heating unchanged for relatively low heating. Ablators, which sacrifice mass to control heating, are used to protect vehicles against more extreme heating for a single use thermal protection system (TPS). Aerogel composites were fabricated and tested for spacecraft thermal protection. The high-temperaturey high heat flux tests described in this paper were performed in NASA Ames arc-jet facilities to simulate spacecraft atmospheric entry, and include heating conditions predicted for the forebody and backshell of the Mars Science Lander (MSL) entry probe. The aerogel composites tested showed excellent thermal performance in the arc-jet tests, functioning both as reusuable insulation under lower heat fluxes, and as ablative aerogels under the extreme heating predicted for the MSL forebody.

  9. Thermal protection systems for hypersonic transport vehicles

    NASA Astrophysics Data System (ADS)

    Reich, G.; Hinger, J.; Huchler, M.

    1990-07-01

    Thermal protection systems (TPS) for hypersonic transport vehicles are described and evaluated. During the flight through the atmosphere moderate to high aerodynamic heating rates with corresponding high surface temperatures are generated. Therefore, a reliable light-weight but effective TPS is required, that limits the heat transfer into the central fuselage with the liquid hydrogen tank and that prevents the penetration of the temperature peak during stage separation to the load carrying structure. The heat transfer modes in the insulation are solid conduction, gas convection and radiation. Thermal protection systems based on different phenomena to reduce the heat transfer, like vacuum shingles, inert gas filled shingles, microporous insulations and multiwall structures, are described. It is demonstrated that microporous and multiwall insulations are efficient, light weight and reliable TPSs for future hypersonic transportation systems.

  10. 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.

  11. Lightweight Thermal Protection System for Atmospheric Entry

    NASA Technical Reports Server (NTRS)

    Stewart, David; Leiser, Daniel

    2007-01-01

    TUFROC (Toughened Uni-piece Fibrous Reinforced Oxidation-resistant Composite) has been developed as a new thermal protection system (TPS) material for wing leading edge and nose cap applications. The composite withstands temperatures up to 1,970 K, and consists of a toughened, high-temperature surface cap and a low-thermal-conductivity base, and is applicable to both sharp and blunt leading edge vehicles. This extends the possible application of fibrous insulation to the wing leading edge and/or nose cap on a hypersonic vehicle. The lightweight system comprises a treated carbonaceous cap composed of ROCCI (Refractory Oxidation-resistant Ceramic Carbon Insulation), which provides dimensional stability to the outer mold line, while the fibrous base material provides maximum thermal insulation for the vehicle structure.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. Estimates Of The Orbiter RSI Thermal Protection System Thermal Reliability

    NASA Technical Reports Server (NTRS)

    Kolodziej, P.; Rasky, D. J.

    2002-01-01

    In support of the Space Shuttle Orbiter post-flight inspection, structure temperatures are recorded at selected positions on the windward, leeward, starboard and port surfaces. Statistical analysis of this flight data and a non-dimensional load interference (NDLI) method are used to estimate the thermal reliability at positions were reusable surface insulation (RSI) is installed. In this analysis, structure temperatures that exceed the design limit define the critical failure mode. At thirty-three positions the RSI thermal reliability is greater than 0.999999 for the missions studied. This is not the overall system level reliability of the thermal protection system installed on an Orbiter. The results from two Orbiters, OV-102 and OV-105, are in good agreement. The original RSI designs on the OV-102 Orbital Maneuvering System pods, which had low reliability, were significantly improved on OV-105. The NDLI method was also used to estimate thermal reliability from an assessment of TPS uncertainties that was completed shortly before the first Orbiter flight. Results fiom the flight data analysis and the pre-flight assessment agree at several positions near each other. The NDLI method is also effective for optimizing RSI designs to provide uniform thermal reliability on the acreage surface of reusable launch vehicles.

  1. 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.

  2. 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

  3. 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.

  4. Advanced materials for thermal protection system

    NASA Astrophysics Data System (ADS)

    Heng, Sangvavann; Sherman, Andrew J.

    1996-03-01

    Reticulated open-cell ceramic foams (both vitreous carbon and silicon carbide) and ceramic composites (SiC-based, both monolithic and fiber-reinforced) were evaluated as candidate materials for use in a heat shield sandwich panel design as an advanced thermal protection system (TPS) for unmanned single-use hypersonic reentry vehicles. These materials were fabricated by chemical vapor deposition/infiltration (CVD/CVI) and evaluated extensively for their mechanical, thermal, and erosion/ablation performance. In the TPS, the ceramic foams were used as a structural core providing thermal insulation and mechanical load distribution, while the ceramic composites were used as facesheets providing resistance to aerodynamic, shear, and erosive forces. Tensile, compressive, and shear strength, elastic and shear modulus, fracture toughness, Poisson's ratio, and thermal conductivity were measured for the ceramic foams, while arcjet testing was conducted on the ceramic composites at heat flux levels up to 5.90 MW/m2 (520 Btu/ft2ṡsec). Two prototype test articles were fabricated and subjected to arcjet testing at heat flux levels of 1.70-3.40 MW/m2 (150-300 Btu/ft2ṡsec) under simulated reentry trajectories.

  5. 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.

  6. Overview of the Orion Thermal Protection System

    NASA Technical Reports Server (NTRS)

    Kowal, T. John

    2010-01-01

    The Orion spacecraft is being developed as part of the Constellation Exploration Program and will serve as the United States crewed transportation system to the International Space Station after the retirement of the Space Shuttle in 2010 and as the eventual means to return U.S. astronauts to the Moon. Therefore, Orion is being designed for reentry missions from both low Earth orbit and from Lunar-return trajectories. This presentation will provide an overview of the development of the Orion TPS, a critical component in the development of the spacecraft. The thermal protection system (TPS) that protects the crew module from the extreme environments associated with Earth atmospheric reentry consists of a forward heatshield and an aft backshell. The requirements that drive the design of the TPS will be discussed, including several key requirements that establish a precedent for U.S. human-rated spacecraft. For the first time in U.S. human spaceflight, a vehicle s TPS is being designed with a specific, derived requirement for reliability. Also, due to the increased presence of spacecraft in Earth s orbit in recent decades, requirements for micro-meteoroid/orbital debris damage tolerance are also a driving requirement that has affected the selection of portions of the TPS. The efforts to select materials and to define a preliminary design for both the heatshield and the backshell will be described. This will include a discussion of the design challenges presented by the numerous penetrations on both the backshell and the heatshield. Finally, the verification and validation plan which is currently under development to certify the TPS for human-rated missions will be outlined. To support the execution of this plan, a ground test campaign for both thermal and structural performance is being designed. This test campaign will directly support thermal and thermal/structural analyses that also are fundamental to the certification effort.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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...

  12. 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...

  13. 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...

  14. 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...

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. Thermal protection systems manned spacecraft flight experience

    NASA Technical Reports Server (NTRS)

    Curry, Donald M.

    1992-01-01

    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.

  4. 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... § 179.18 Thermal protection systems. (a) Performance standard. When the regulations in this subchapter..., and upon request, made available for inspection and copying by an authorized representative of...

  5. Ceramic-Fibrous-Insulation Thermal-Protection System

    NASA Technical Reports Server (NTRS)

    Leiser, Daniel; Churchward, Rex; Katvala, Victor; Stewart, David; Balter, Aliza

    1992-01-01

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

  6. 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.

  7. 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.

  8. 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.

  9. The Control System for the X-33 Linear Aerospike Engine

    NASA Technical Reports Server (NTRS)

    Jackson, Jerry E.; Espenschied, Erich; Klop, Jeffrey

    1998-01-01

    The linear aerospike engine is being developed for single-stage -to-orbit (SSTO) applications. The primary advantages of a linear aerospike engine over a conventional bell nozzle engine include altitude compensation, which provides enhanced performance, and lower vehicle weight resulting from the integration of the engine into the vehicle structure. A feature of this integration is the ability to provide thrust vector control (TVC) by differential throttling of the engine combustion elements, rather than the more conventional approach of gimballing the entire engine. An analysis of the X-33 flight trajectories has shown that it is necessary to provide +/- 15% roll, pitch and yaw TVC authority with an optional capability of +/- 30% pitch at select times during the mission. The TVC performance requirements for X-33 engine became a major driver in the design of the engine control system. The thrust level of the X-33 engine as well as the amount of TVC are managed by a control system which consists of electronic, instrumentation, propellant valves, electro-mechanical actuators, spark igniters, and harnesses. The engine control system is responsible for the thrust control, mixture ratio control, thrust vector control, engine health monitoring, and communication to the vehicle during all operational modes of the engine (checkout, pre-start, start, main-stage, shutdown and post shutdown). The methodology for thrust vector control, the health monitoring approach which includes failure detection, isolation, and response, and the basic control system design are the topic of this paper. As an additional point of interest a brief description of the X-33 engine system will be included in this paper.

  10. Optimal Thermal Design of a Multishield Thermal Protection System of Reusable Space Vehicles

    NASA Astrophysics Data System (ADS)

    Maiorova, I. A.; Prosuntsov, P. V.; Zuev, A. V.

    2016-03-01

    We have solved the problem of the optimal thermal design of a multishield thermal protection system of reusable space vehicles due to the choice of the optimal position and materials of radiation shields.

  11. 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.

  12. 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.

  13. 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.

  14. 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

  15. 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

  16. 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

  17. 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

  18. 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

  19. 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.

  20. Active wireless temperature sensors for aerospace thermal protection systems

    NASA Astrophysics Data System (ADS)

    Milos, Frank S.; Karunaratne, K. S. G.

    2003-07-01

    Vehicle system 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 advance inspection and health management technologies for thermal protection systems. This paper summarizes a joint effort by NASA Ames and Korteks to develop active "wireless" sensors that can be embedded in the thermal protection system to monitor subsurface temperature histories. These devices are thermocouples integrated with radio-frequency identification circuits to enable non-contact communication of temperature data through aerospace thermal protection materials. Two generations of prototype sensors are discussed. The advanced prototype collects data from three type-k thermocouples attached to a 25-mm square integrated circuit and can communicate through 7 to 10 cm thickness of thermal protection materials.

  1. 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 ...

  2. 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 ...

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. Design Description of the X-33 Avionics Architecture

    NASA Technical Reports Server (NTRS)

    Reichenfeld, Curtis J.; Jones, Paul G.

    1999-01-01

    In this paper, we provide a design description of the X-33 avionics architecture. The X-33 is an autonomous Single Stage to Orbit (SSTO) launch vehicle currently being developed by Lockheed Martin for NASA as a technology demonstrator for the VentureStar Reusable Launch Vehicle (RLV). The X-33 avionics provides autonomous control of die vehicle throughout takeoff, ascent, descent, approach, landing, rollout, and vehicle safing. During flight the avionics provides communication to the range through uplinked commands and downlinked telemetry. During pre-launch and post-safing activities, the avionics provides interfaces to ground support consoles that perform vehicle flight preparations and maintenance. The X-33 Avionics is a hybrid of centralized and distributed processing elements connected by three dual redundant Mil-Std 1553 data buses. These data buses are controlled by a central processing suite located in the avionics bay and composed of triplex redundant Vehicle Mission Computers (VMCs). The VMCs integrate mission management, guidance, navigation, flight control, subsystem control and redundancy management functions. The vehicle sensors, effectors and subsystems are interfaced directly to the centralized VMCs as remote terminals or through dual redundant Data Interface Units (DIUs). The DIUs are located forward and aft of the avionics bay and provide signal conditioning, health monitoring, low level subsystem control and data interface functions. Each VMC is connected to all three redundant 1553 data buses for monitoring and provides a complete identical data set to the processing algorithms. This enables bus faults to be detected and reconfigured through a voted bus control configuration. Data is also shared between VMCs though a cross channel data link that is implemented in hardware and controlled by AlliedSignal's Fault Tolerant Executive (FTE). The FTE synchronizes processors within the VMC and synchronizes redundant VMCs to each other. The FTE provides

  8. 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.

  9. Thermal Management Coating As Thermal Protection System for Space Transportation System

    NASA Technical Reports Server (NTRS)

    Kaul, Raj; Stuckey, C. Irvin

    2003-01-01

    This paper presents viewgraphs on the development of a non-ablative thermal management coating used as the thermal protection system material for space shuttle rocket boosters and other launch vehicles. The topics include: 1) Coating Study; 2) Aerothermal Testing; 3) Preconditioning Environments; 4) Test Observations; 5) Lightning Strike Test Panel; 6) Test Panel After Impact Testing; 7) Thermal Testing; and 8) Mechanical Testing.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. Active Wireless Temperature Sensors for Aerospace Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

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

    2002-01-01

    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 advance inspection and health management technologies for thermal protection systems. This paper summarizes a joint project between NASA Ames and Korteks to develop active wireless sensors that can be embedded in the thermal protection system to monitor sub-surface temperature histories. These devices are thermocouples integrated with radio-frequency identification circuitry to enable acquisition and non-contact communication of temperature data through aerospace thermal protection materials. Two generations of prototype sensors are discussed. The advanced prototype collects data from three type-k thermocouples attached to a 2.54-cm square integrated circuit.

  16. Thermal protection during percutaneous thermal ablation procedures: interest of carbon dioxide dissection and temperature monitoring.

    PubMed

    Buy, Xavier; Tok, Chung-Hong; Szwarc, Daniel; Bierry, Guillaume; Gangi, Afshin

    2009-05-01

    Percutaneous image-guided thermal ablation of tumor is widely used, and thermal injury to collateral structures is a known complication of this technique. To avoid thermal damage to surrounding structures, several protection techniques have been reported. We report the use of a simple and effective protective technique combining carbon dioxide dissection and thermocouple: CO(2) displaces the nontarget structures, and its low thermal conductivity provides excellent insulation; insertion of a thermocouple in contact with vulnerable structures achieves continuous thermal monitoring. We performed percutaneous thermal ablation of 37 tumors in 35 patients (4 laser, 10 radiofrequency, and 23 cryoablations) with protection of adjacent vulnerable structures by using CO(2) dissection combined with continuous thermal monitoring with thermocouple. Tumor locations were various (19 intra-abdominal tumors including 4 livers and 9 kidneys, 18 musculoskeletal tumors including 11 spinal tumors). CO(2) volume ranged from 10 ml (epidural space) to 1500 ml (abdominal). Repeated insufflations were performed if necessary, depending on the information given by the thermocouple and imaging control. Dissection with optimal thermal protection was achieved in all cases except two patients where adherences (one postoperative, one arachnoiditis) blocked proper gaseous distribution. No complication referred to this technique was noted. This safe, cost-effective, and simple method increases the safety and the success rate of percutaneous thermal ablation procedures. It also offers the potential to increase the number of tumors that can be treated via a percutaneous approach. PMID:19219496

  17. Thermal Protection During Percutaneous Thermal Ablation Procedures: Interest of Carbon Dioxide Dissection and Temperature Monitoring

    SciTech Connect

    Buy, Xavier; Tok, Chung-Hong; Szwarc, Daniel; Bierry, Guillaume; Gangi, Afshin

    2009-05-15

    Percutaneous image-guided thermal ablation of tumor is widely used, and thermal injury to collateral structures is a known complication of this technique. To avoid thermal damage to surrounding structures, several protection techniques have been reported. We report the use of a simple and effective protective technique combining carbon dioxide dissection and thermocouple: CO{sub 2} displaces the nontarget structures, and its low thermal conductivity provides excellent insulation; insertion of a thermocouple in contact with vulnerable structures achieves continuous thermal monitoring. We performed percutaneous thermal ablation of 37 tumors in 35 patients (4 laser, 10 radiofrequency, and 23 cryoablations) with protection of adjacent vulnerable structures by using CO{sub 2} dissection combined with continuous thermal monitoring with thermocouple. Tumor locations were various (19 intra-abdominal tumors including 4 livers and 9 kidneys, 18 musculoskeletal tumors including 11 spinal tumors). CO{sub 2} volume ranged from 10 ml (epidural space) to 1500 ml (abdominal). Repeated insufflations were performed if necessary, depending on the information given by the thermocouple and imaging control. Dissection with optimal thermal protection was achieved in all cases except two patients where adherences (one postoperative, one arachnoiditis) blocked proper gaseous distribution. No complication referred to this technique was noted. This safe, cost-effective, and simple method increases the safety and the success rate of percutaneous thermal ablation procedures. It also offers the potential to increase the number of tumors that can be treated via a percutaneous approach.

  18. Extended Range Communications Support for the X-33

    NASA Technical Reports Server (NTRS)

    Eslinger, Brian; Garza, Reynaldo

    1998-01-01

    Communications support for the X-33 requires addressing several interesting challenges to meet program and range safety requirements. In addition, the heavily cost driven program has provided concessions to the communications support to reduce the cost of networking the extended range. Cost trade-offs showed that by lowering the telemetry data rate from 2 Megabits per second to 1440 Megabits per second that significant cost avoidance could be realized. Also, by adopting standard telecommunications data rate for the uplink data stream, an efficient and integrated solution for the extended range communications could be supported. Meeting the program requirements as well as range safety requirements for this effort are critical to the success of the program. This paper describes some of the important requirements driving the design of the extended range communications support and the design of the system to meet those requirements.

  19. 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.

  20. 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.

  1. Numerical simulation for thermal shock resistance of thermal protection materials considering different operating environments.

    PubMed

    Li, Weiguo; Li, Dingyu; Wang, Ruzhuan; Fang, Daining

    2013-01-01

    Based on the sensitivities of material properties to temperature and the complexity of service environment of thermal protection system on the spacecraft, ultrahigh-temperature ceramics (UHTCs), which are used as thermal protection materials, cannot simply consider thermal shock resistance (TSR) of the material its own but need to take the external constraint conditions and the thermal environment into full account. With the thermal shock numerical simulation on hafnium diboride (HfB2), a detailed study of the effects of the different external constraints and thermal environments on the TSR of UHTCs had been made. The influences of different initial temperatures, constraint strengths, and temperature change rates on the TSR of UHTCs are discussed. This study can provide a more intuitively visual understanding of the evolution of the TSR of UHTCs during actual operation conditions. PMID:23983628

  2. 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.

  3. "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.

  4. 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.

  5. Closed-pore Insulation Thermal Protection System Design Concept Development

    NASA Technical Reports Server (NTRS)

    Varisco, A.; Harris, H. G.

    1973-01-01

    The development of a unique closed-pore ceramic foam insulation (CPI) produced from low cost fly ash cenospheres is reported for space shuttle external thermal protection. Two basic design approaches were developed: bonded and mechanically fastened. A description of the concepts is presented in addition to fabrication and test results.

  6. Thermal Protection Systems for Future NASA Space Vehicles

    NASA Technical Reports Server (NTRS)

    Leiser, Daniel B.; Rasky, Daniel; Arnold, James O. (Technical Monitor)

    2000-01-01

    The proposed first through fourth generation of future NASA Reusable Launch Vehicles (RLV) within NASA will be described, in general, along with their relative goals for improvement in performance (i.e., cost, safety, life, and turnaround time). A brief description of Spaceliner 100 activities representing a means to achieve those goals will be included. Some of the families of thermal protection materials with widely varying characteristics that are being developed for first generation space vehicles at Ames Research Center will be described as well as potential materials and composites for second and third generation applications as systems. These families of materials include functionally gradient material composites that are made from a variety of low-density substrates and moderate to fully dense surface treatments providing the resultant material with both toughness and higher temperature capability opening the envelope of Thermal Protection Systems (TPS) capabilities. Some of the materials truly represent enabling technologies that are required to achieve substantially enhanced thermal protection system performance thereby reducing vehicle risk. Finally the needs for integrated vehicle health monitoring (IVHM) of future vehicles thermal protection systems relative to achieving the goals for third generation reusable launch vehicles and for improving vehicle performance and capabilities reducing risk will be described along with the state of the art in TPS.

  7. Damage Detection/Locating System Providing Thermal Protection

    NASA Technical Reports Server (NTRS)

    Woodard, Stanley E. (Inventor); Jones, Thomas W. (Inventor); Taylor, Bryant D. (Inventor); Qamar, A. Shams (Inventor)

    2010-01-01

    A damage locating system also provides thermal protection. An array of sensors substantially tiles an area of interest. Each sensor is a reflective-surface conductor having operatively coupled inductance and capacitance. A magnetic field response recorder is provided to interrogate each sensor before and after a damage condition. Changes in response are indicative of damage and a corresponding location thereof.

  8. 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.

  9. 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.

  10. Artist concept of X-33 and VentureStar Reusable Launch Vehicle (RLV)

    NASA Technical Reports Server (NTRS)

    1997-01-01

    An artist's conception showing the relative sizes of the X-33 (left) and the proposed operational VentureStar. Although about the same shape, the VentureStar would be twice the size of the half scale X-33. The added size was necessary to accommodate a large payload bay, and the increased fuel supply needed to reach orbital speeds of Mach 25+. The VentureStar was intended to be a low-cost, reusable launch vehicle, while the X-33 fabricated to test the advanced technologies needed to build it. With the cancellation of the X-33 in February 2001, the VentureStar effort also ended.

  11. 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.

  12. 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.

  13. 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.

  14. Intumescent-ablators as improved thermal protection materials

    NASA Technical Reports Server (NTRS)

    Sawko, P. M.; Riccitiello, S. R.

    1977-01-01

    Nitroaromatic amine-based intumescent coatings were improved with regard to their thermal protection ability by adding endothermic decomposing fillers with endotherms at or near the exothermic reaction of the intumescent agent, since the effectiveness of the intumescent coatings without fillers is reduced by the exothermic behavior of the coatings during thermal activation. Fillers were dispersed directly in the base coating. Potassium fluoborate, ammonium fluoborate, zinc borate, and ammonium oxalate function as endothermic ablative materials at specific temperature regions, and also enhance the char formation during the intumescent process.

  15. 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.

  16. 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.

  17. Subsonic and Transonic Dynamic Stability Characteristics of the X-33

    NASA Technical Reports Server (NTRS)

    Tomek, D.; Boyden, R.

    2000-01-01

    Dynamic stability testing was conducted on a 2.5% scale model of the X-33 technology demonstrator sub-orbital flight-test vehicle. This testing was conducted at the NASA Langley Research Center (LaRC) l6-Foot Transonic Wind Tunnel with the LaRC High-speed Dynamic Stability system. Forced oscillation data were acquired for various configurations over a Mach number range of 0.3 to 1.15 measuring pitch, roll and yaw damping, as well as the normal force due to pitch rate and the cross derivatives. The test angle of attack range was from -2 to 24 degrees, except for those cases where load constraints limited the higher angles of attack at the higher Mach numbers. A variety of model configurations with and without control surfaces were employed, including a body alone configuration. Stable pitch damping is exhibited for the baseline configuration throughout the angle of attack range for Mach numbers 0.3, 0.8, and 1.15. Stable pitch damping is present for Mach numbers 0.9 and 0.6 with the exception of angles 2 and 16 degrees, respectively. Constant and stable roll damping were present for the baseline configuration over the range of Mach numbers up to an angle of attack of 16 degrees. The yaw damping for the baseline is somewhat stable and constant for the angle of attack range from -2 to 8 degrees, with the exception of Mach numbers 0.6 and 0.8. Yaw damping becomes highly unstable for all Mach numbers at angles of attack greater than 8 degrees.

  18. Thermal and aerothermal performance of a titanium multiwall thermal protection system

    NASA Technical Reports Server (NTRS)

    Avery, D. E.; Shideler, J. L.; Stuckey, R. N.

    1981-01-01

    A metallic thermal protection system (TPS) concept the multiwall designed for temperature and pressure at Shuttle body point 3140 where the maximum surface temperature is approximately 811 K was tested to evaluate thermal performance and structural integrity. A two tile model of titanium multiwall and a model consisting of a low temperature reusable surface insulation (LRSI) tiles were exposed to 25 simulated thermal and pressure Shuttle entry missions. The two systems performed the same, and neither system deteriorated during the tests. It is indicated that redesign of the multiwall tiles reduces tile thickness and/or weight. A nine tile model of titanium multiwal was tested for radiant heating and aerothermodynamics. Minor design changes that improve structural integrity without having a significant impact on the thermal protection ability of the titanium multiwall TPS are identified. The capability of a titanium multiwall thermal protection system to protect an aluminum surface during a Shuttle type entry trajectory at locations on the vehicle where the maximum surface temperature is below 811 K is demonstrated.

  19. 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.

  20. 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.

  1. 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.

  2. 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).

  3. 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.

  4. 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...

  5. 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...

  6. 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...

  7. 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...

  8. 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...

  9. 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.

  10. 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.

  11. 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.

  12. 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

  13. Fiber optic temperature profiling for thermal protection heat shields

    NASA Astrophysics Data System (ADS)

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

    2014-04-01

    Reliable Thermal Protection System (TPS) sensors are needed to achieve better designs for spacecraft (probe) heatshields for missions requiring atmospheric aero-capture or entry/reentry. In particular, they will allow both reduced risk and heat-shield mass minimization, which will facilitate more missions and allow increased payloads and returns. For thermal measurements, Intelligent Fiber Optic Systems Corporation (IFOS) is providing a temperature monitoring system involving innovative lightweight, EMI-immune, high-temperature resistant Fiber Bragg Grating (FBG) sensors with a thermal mass near that of TPS materials together with fast FBG sensor interrogation. The IFOS fiber optic sensing technology is highly sensitive and accurate. It is also low-cost and lends itself to high-volume production. Multiple sensing FBGs can be fabricated as arrays on a single fiber for simplified design and reduced cost. In this paper, we provide experimental results to demonstrate the temperature monitoring system using multi-sensor FBG arrays embedded in 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. Bioassay of thermal protection afforded by candidate flight suit fabrics.

    PubMed

    Knox, F S; Wachtel, T L; McCahan, G R

    1979-10-01

    The United States Army Aeromedical Research Laboratory (USAARL) porcine cutaneous bioassay technique was used to determine what mitigating effect four thermally protective flight suit fabrics would have on fire-induced skin damage. The fabrics were 4.8-ox twill weave Nomex aramide, 4.5-oz stabilized twill weave polybenzimidazole, 4.8-oz plain weave experimental high-temperature polymer (HT4), and 4.8-oz plain weave Nomex aramide (New Weave Nomex or NWN). Each fabric sample was assayed 20 times in each of four configurations: as a single layer in contact with the skin; as a single layer with a 6.35 mm (0.25 in) air gap between fabric and skin; in conjuction with a cotton T-shirt with no air gaps; and, finally, in conjuction with a T-shirt with a 6.35 mm air gap between T-shirt and fabric. Bare skin was used as a control. A JP-4 fueled furnace was used as a thermal source and was adjested to deliver a mean heat flux of 3.07 cal/cm2/s. The duration of exposure was 5 s. Four hundred burn sites were graded using clinical observation and microscopic techniques. Used as single layers, none of the fabrics demonstrated superiority in providing clinically significant protection. When used with a cotton T-shirt, protection was improved. Protection improved progressively for all fabrics and configuration when an air gap was introduced. The experimental high-temperature polymer consistently demonstrated lower heat flux transmission in all configurations, but did not significantly reduce clinical burns. PMID:518445

  2. The Reusable Launch Vehicle Technology Program and the X-33 Advanced Technology Demonstrator

    NASA Technical Reports Server (NTRS)

    Cook, Stephen A.

    1995-01-01

    The goal of the Reusable Launch Vehicle (RLV) technology program is formulated, and the primary objectives of RLV are listed. RLV technology program implementation phases are outlined. X-33 advanced technology demonstrator is described. Program management is addressed.

  3. Detection of fastener failure in a thermal protection system

    NASA Astrophysics Data System (ADS)

    Derriso, Mark M.; Olson, Steven E.; Braisted, William R.; DeSimio, Martin P.; Rosenstengel, John; Brown, Kevin

    2004-07-01

    This paper presents experimental and analytical studies focused on the development of a structural health monitoring system to assess the condition of mechanical fasteners of a thermal protection system. A realistic thermal protection system component, consisting of a carbon-carbon panel bolted through 15 brackets to a backing structure, is utilized. Mechanical states considered include all bolts fastened to a nominal torque value, or one of the 15 bolts loosened. Four transducers on the backing structure provide actuation and sensing signals. Spectral functions are computed from all single and pair-wise signal combinations. Automated analysis of the spectral functions shows frequency intervals exist over which the function values are indicative of the mechanical state of the test article. These frequency intervals are used to provide features for the structural health monitoring classifier. Finite element analyses provide a physics-based understanding of these features. Statistical pattern recognition methods select a subset of the features. The overall localization accuracy of the structural health monitoring system on test data is 99.1% with 99.7% probability of detecting a damaged condition at a 0.2% probability of a false alarm.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  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. 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.

  11. Nonlinear dynamic phenomena in the space shuttle thermal protection system

    NASA Technical Reports Server (NTRS)

    Housner, J. M.; Edighoffer, H. H.; Park, K. C.

    1981-01-01

    The development of an analysis for examining the nonlinear dynamic phenomena arising in the space shuttle orbiter tile/pad thermal protection system is presented. The tile/pad system consists of ceramic tiles bonded to the aluminum skin of the orbiter through a thin nylon felt pad. The pads are a soft nonlinear material which permits large strains and displays both hysteretic and nonlinear viscous damping. Application of the analysis to a square tile subjected to transverse sinusoidal motion of the orbiter skin is presented and the following nonlinear dynamic phenomena are considered: highly distorted wave forms, amplitude-dependent resonant frequencies which initially decrease and then increase with increasing amplitude of motion, magnification of substrate motion which is higher than would be expected in a similarly highly damped linear system, and classical parametric resonance instability.

  12. 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.

  13. 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.

  14. Room temperature mechanical properties of shuttle thermal protection system materials

    NASA Technical Reports Server (NTRS)

    Sawyer, J. W.; Rummler, D. R.

    1980-01-01

    Tests were conducted at room temperature to determine the mechanical properties and behavior of materials used for the thermal protection system of the space shuttle. The materials investigated include the LI-900 RSI tiles, the RTV-560 adhesive and the .41 cm (.16 thick) strain isolator pad (SIP). Tensile and compression cyclic loading tests were conducted on the SIP material and stress-strain curves obtained for various proof loads and load cyclic conditioning. Ultimate tensile and shear tests were conducted on the RSI, RTV, and SIP materials. The SIP material exhibits highly nonlinear stress-strain behavior, increased tangent modulus and ultimate tensile strength with increased loading rate, and large short time load relaxation and moderate creep behavior. Proof and cyclic load conditioning of the SIP results in permanent deformation of the material, hysteresis effects, and much higher tensile tangent modulus values at large strains.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  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. 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.

  6. 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.

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

    NASA Technical Reports Server (NTRS)

    Addy, Gene; Oleskiw, Myron; Broeren, Andy P.; Orchard, David

    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.

  8. 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

  9. Ocean thermal conversion (OTEC) project bottom cable protection study. Analysis and selection of protection techniques

    SciTech Connect

    Not Available

    1981-10-01

    General guidelines and procedures for cable protection are given for the four proposed Ocean Thermal Energy Conversion (OTEC) plant sites and cable routes, together with seafloor scenarios and protection strategies for each site. Burial of the cable below the seafloor is the recommended and best method of protecting OTEC cables from the hazards existing at all sites, namely, chafe and corrosion, hydrodynamic forces, trawler/dredge, and ship anchor. For landslides and earthquakes the only feasible method of protection, although limited, is to provide slack, in the cable, i.e. lay extra length. Trenches for burying the cable are recommended to be constructed a) by blasting through hard bottom at Hawaii for the first nautical mile (n.m.) and at Puerto Rico for the first 0.9 n.m; b)by a plowing machine at Hawaii for the next 0.5 n.m.; c) by a trenching machine at Guam for the first 0.55 n.m.; d) by a trenching /laying machine at Florida for 110 n.m.; and e) by a conventional floating dredge for 15 n.m. For the outshore segments of the cable routes it is recommenced to lay the cable on th seafloor because bottom sediments are soft enough to permit the cable to bury itself. Except for the Florida route, a normal cable laying vessel is recommended for laying the cable from plant site to landfall and for performing the protection details which are temie concrete cover over the cable at Hawaii for 0.5 n.m. and split pipe and rock anchor at Puerto Rico for 0l2 n.m.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  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. Heat flux instrumentation for Hyflite thermal protection system

    NASA Astrophysics Data System (ADS)

    Diller, T. E.

    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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. Evaluation of protective ensemble thermal characteristics through sweating hot plate, sweating thermal manikin, and human tests.

    PubMed

    Kim, Jung-Hyun; Powell, Jeffery B; Roberge, Raymond J; Shepherd, Angie; Coca, Aitor

    2014-01-01

    The purpose of this study was to evaluate the predictive capability of fabric Total Heat Loss (THL) values on thermal stress that Personal Protective Equipment (PPE) ensemble wearers may encounter while performing work. A series of three tests, consisting of the Sweating Hot Plate (SHP) test on two sample fabrics and the Sweating Thermal Manikin (STM) and human performance tests on two single-layer encapsulating ensembles (fabric/ensemble A = low THL and B = high THL), was conducted to compare THL values between SHP and STM methods along with human thermophysiological responses to wearing the ensembles. In human testing, ten male subjects performed a treadmill exercise at 4.8 km and 3% incline for 60 min in two environmental conditions (mild = 22°C, 50% relative humidity (RH) and hot/humid = 35°C, 65% RH). The thermal and evaporative resistances were significantly higher on a fabric level as measured in the SHP test than on the ensemble level as measured in the STM test. Consequently the THL values were also significantly different for both fabric types (SHP vs. STM: 191.3 vs. 81.5 W/m(2) in fabric/ensemble A, and 909.3 vs. 149.9 W/m(2) in fabric/ensemble B (p < 0.001). Body temperature and heart rate response between ensembles A and B were consistently different in both environmental conditions (p < 0.001), which is attributed to significantly higher sweat evaporation in ensemble B than in A (p < 0.05), despite a greater sweat production in ensemble A (p < 0.001) in both environmental conditions. Further, elevation of microclimate temperature (p < 0.001) and humidity (p < 0.01) was significantly greater in ensemble A than in B. It was concluded that: (1) SHP test determined THL values are significantly different from the actual THL potential of the PPE ensemble tested on STM, (2) physiological benefits from wearing a more breathable PPE ensemble may not be feasible with incremental THL values (SHP test) less than approximately 150-200 W·m(2), and (3) the

  5. 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.

  6. 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

  7. 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

  8. 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.

  9. 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.

  10. 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.

  11. X-33 Metallic TPS Tests in NASA-LaRC High Temperature Tunnel

    NASA Technical Reports Server (NTRS)

    Bouslog, Stanley A.; Moore, Brad; Scanlon, Ron J.; Sawyer, James Wayne

    1998-01-01

    Conclusions The first series of metallic TPS tests in the NASA-LaRC Mach 7 High Temperature Tunnel has been completed. Additional testing is 'in progress and shall provide data for off-design configurations for the metallic TPS. The available data are being analyzed and being used to correlate analytical models to be used for X-33 flight design analysis. The final paper shan present additional data from these tests and comparisons between the data and analytical predictions.

  12. Percutaneous thermal ablation: how to protect the surrounding organs.

    PubMed

    Tsoumakidou, Georgia; Buy, Xavier; Garnon, Julien; Enescu, Julian; Gangi, Afshin

    2011-09-01

    A variety of thermal ablation techniques have been advocated for percutaneous tumor management. Although the above techniques are considered safe, they can be complicated with unintended thermal injury to the surrounding structures, with disastrous results. In the present article we report a number of different insulation techniques (hydrodissection, gas dissection and balloon interposition, warming/cooling systems) that can be applied. Emphasis is given to the procedure-related details, and we present the advantages and drawbacks of the insulation techniques. We also provide tips on avoiding painful skin burns when treating superficial lesions. Finally, we point out the interest of temperature monitoring and how it can be achieved (use of thermocouples, fiberoptic thermosensors, or direct magnetic resonance imaging temperature mapping). The above thermal insulation and temperature monitoring techniques can be applied alone or in combination. Familiarity with these techniques is essential to avoid major complications and to increase the indications of thermal ablation procedures. PMID:21767784

  13. 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.

  14. 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.

  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. X33 Reusable Launch Vehicle Control on Sliding Modes: Concepts for a Control System Development

    NASA Technical Reports Server (NTRS)

    Shtessel, Yuri B.

    1998-01-01

    Control of the X33 reusable launch vehicle is considered. The launch control problem consists of automatic tracking of the launch trajectory which is assumed to be optimally precalculated. It requires development of a reliable, robust control algorithm that can automatically adjust to some changes in mission specifications (mass of payload, target orbit) and the operating environment (atmospheric perturbations, interconnection perturbations from the other subsystems of the vehicle, thrust deficiencies, failure scenarios). One of the effective control strategies successfully applied in nonlinear systems is the Sliding Mode Control. The main advantage of the Sliding Mode Control is that the system's state response in the sliding surface remains insensitive to certain parameter variations, nonlinearities and disturbances. Employing the time scaling concept, a new two (three)-loop structure of the control system for the X33 launch vehicle was developed. Smoothed sliding mode controllers were designed to robustly enforce the given closed-loop dynamics. Simulations of the 3-DOF model of the X33 launch vehicle with the table-look-up models for Euler angle reference profiles and disturbance torque profiles showed a very accurate, robust tracking performance.

  17. 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

  18. 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.

  19. 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.

  20. Design, fabrication, and tests of a metallic shell tile thermal protection system for space transportation

    NASA Technical Reports Server (NTRS)

    Macconochie, Ian O.; Kelly, H. Neale

    1989-01-01

    A thermal protection tile for earth-to-orbit transports is described. The tiles consist of a rigid external shell filled with a flexible insulation. The tiles tend to be thicker than the current Shuttle rigidized silica tiles for the same entry heat load but are projected to be more durable and lighter. The tiles were thermally tested for several simulated entry trajectories.

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

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-02-27

    ... function. II. Further Information DG-1264, was published in the Federal Register on May 02, 2011 (76 FR... 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...

  2. 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.

  3. 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.

  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. 49 CFR 179.18 - Thermal protection systems.

    Code of Federal Regulations, 2010 CFR

    2010-10-01

    ... through the pressure release device, when subjected to: (1) A pool fire for 100 minutes; and (2) A torch fire for 30 minutes. (b) Thermal analysis. (1) Compliance with the requirements of paragraph (a) of this section shall be verified by analyzing the fire effects on the entire surface of the tank car....

  7. 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.

  8. 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.

  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. 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.

  11. 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.

  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. 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.

  14. [The mechanisms of the thermal protective action of atsefen].

    PubMed

    Badyshtov, B A; Losev, A S; Sytnik, S I; Pastushenkov, V A; Makhnycheva, A L; Kolotilinskaia, N V; Mironov, S A; Seredenin, S B

    1995-01-01

    The study of thermoprotective properties of acephen (300 mg, single dosage) showed an increase (compared to that of placebo) in the maximum endurable time under thermal loading in healthy volunteers: optimization of evaporative heat exchange regime; reduction of the final values of reactive anxiety and fatigue and decrease in the rate of skin temperature growth. In combination with the data of biochemical analysis of peripheral blood which demonstrated positive changes in hormone, lipid, protein, and energy metabolism as well as in vegetative regulation, the results testify to a substantial thermoprotective potential of acephen attributed to the wide range of its pharmacological activity. PMID:8704597

  15. 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

  16. 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

  17. 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.

  18. 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.

  19. Status of reusable surface insulation thermal protection system technology programs

    NASA Technical Reports Server (NTRS)

    Greenshields, D. H.; Meyer, A. J.; Tillian, D. J.

    1972-01-01

    The development of three low-density rigidized insulation materials for the shuttle TPS application is reported. These materials consist of one high purity silica system and two systems based on mullite, an aluminum silicate. Both systems consist of fibers joined together with appropriate binders to obtain a rigidized insulation composite. Both material systems require the application of a glassy coating to provide a wear resistant, high emittance surface and to prevent the absorption of water by the fiber matrix. The technology program has addressed the development of water impervious coatings, methods of assembling the materials in design concepts while minimizing the thermal stress in the insulation, achieving compatibility between the RSI material and the structural system, and test evaluations to demonstrate the feasibility of the surface insulation concept.

  20. 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.

  1. 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.

  2. 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.

  3. The employment of a high density plasma jet for the investigation of thermal protection materials

    NASA Astrophysics Data System (ADS)

    Kezelis, R.; Grigaitiene, V.; Levinskas, R.; Brinkiene, K.

    2014-05-01

    This paper describes the results of tests of thermal protection materials (TPM) at conditions that simulate the atmospheric re-entry of space vehicles, tested by means of a high velocity and enthalpy air plasma jet generated with a dc plasma torch. Such a high velocity and enthalpy air plasma jet allows us to investigate TPM by simulating heat flux values varying with time in accordance with real re-entry altitudes and trajectories. The main research interests include the measurements of plasma flow temperature and heat flux for the testing of materials used for thermal protection systems of space vehicles. The test results of investigations of light composite thermal protective system material and graphite are presented.

  4. The effect of various cosmetic pretreatments on protecting hair from thermal damage by hot flat ironing.

    PubMed

    Zhou, Y; Rigoletto, R; Koelmel, D; Zhang, G; Gillece, T W; Foltis, L; Moore, D J; Qu, X; Sun, C

    2011-01-01

    Hot flat irons are used to create straight hair styles. As these devices operate at temperatures over 200 °C they can cause significant damage to hair keratin. In this study, hair thermal damage and the effect of various polymeric pretreatments were investigated using FTIR imaging spectroscopy, DSC, dynamic vapor sorption (DVS), AFM, SEM, and thermal image analysis. FTIR imaging spectroscopy of hair cross sections provides spatially resolved molecular information such as protein distribution and structure. This approach was used to monitor thermally induced modification of hair protein, including the conversion of α-helix to β-sheet and protein degradation. DSC measurements of thermally treated hair also demonstrated degradation of hair keratin. DVS of thermally treated hair shows the reduced water regain and lower water retention, compared to the non-thermally treated hair, which might be attributed to the protein conformation changes due to heat damage. The protection of native protein structure associated with selected polymer pretreatments leads to improved moisture restoration and water retention of hair. This contributes to heat control on repeated hot flat ironing. Thermally stressing hair led to significantly increased hair breakage when subjected to combing. These studies indicate that hair breakage can be reduced significantly when hair is pretreated with selected polymers such as VP/acrylates/lauryl methacrylate copolymer, polyquaternium-55, and a polyelectrolyte complex of PVM/MA copolymer and polyquaternium-28. In addition, polymeric pretreatments provide thermal protection against thermal degradation of keratin in the cortex as well as hair surface damage. The morphological improvement in cuticle integrity and smoothness with the polymer pretreatment plays an important role in their anti-breakage effect. Insights into structure-property relationships necessary to provide thermal protection to hair are presented. PMID:21635854

  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. 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.

  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. 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.

  9. [Expression purification and verification of HBscFv-IFNgamma in Pichia pastoris x33].

    PubMed

    Zhou, Shishui; Wang, Xiaoning

    2008-03-01

    In order to effectively cure hepatitis B virus (HBV), we studied on fusion protein HBscFv-IFNgamma, which was connected with single-chain Fv against HBV surface antigen(HBscFv) and gamma-interferon(IFNgamma) of being used in clinic against HBV. Adopting overlap PCR, the hbscfv and the ifngamma were connected into hbscfv-ifngamma. Then the pPICZalphaA/(hbscfv-ifngamma)(1,2,4) of multi-copy recombinant plasmid were constructed and transformed into Pichia pastoris x33. The engineering strain x4 was screened from transformed x33 and could secretively express HBscFv-IFNgamma. The preliminary verification indicates that HBscFv-IFNgamma has the bioactivity of HBscFv and IFNgamma by SDS-PAGE, Western blotting and ELISA. The supernatant of culturing X4 was purified by 14F7 affinity chromatography to HBscFv-IFNgamma with purity of 95%-98%. The HBscFv-IFNgamma is able to bind 27.9% HBV surface antigen (HBsAg) in the serum of HBV transgenic mice, which shows the antibody of HBscFv-IFNgamma has bioactivity in vivo. Therefore HBscFv-IFNgamma can shed light on the development of a new promising HBV-targeted drug. PMID:18589818

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. Dynamics and protection of tripartite quantum correlations in a thermal bath

    SciTech Connect

    Guo, Jin-Liang Wei, Jin-Long

    2015-03-15

    We study the dynamics and protection of tripartite quantum correlations in terms of genuinely tripartite concurrence, lower bound of concurrence and tripartite geometric quantum discord in a three-qubit system interacting with independent thermal bath. By comparing the dynamics of entanglement with that of quantum discord for initial GHZ state and W state, we find that W state is more robust than GHZ state, and quantum discord performs better than entanglement against the decoherence induced by the thermal bath. When the bath temperature is low, for the initial GHZ state, combining weak measurement and measurement reversal is necessary for a successful protection of quantum correlations. But for the initial W state, the protection depends solely upon the measurement reversal. In addition, the protection cannot usually be realized irrespective of the initial states as the bath temperature increases.

  17. Dynamics and protection of tripartite quantum correlations in a thermal bath

    NASA Astrophysics Data System (ADS)

    Guo, Jin-Liang; Wei, Jin-Long

    2015-03-01

    We study the dynamics and protection of tripartite quantum correlations in terms of genuinely tripartite concurrence, lower bound of concurrence and tripartite geometric quantum discord in a three-qubit system interacting with independent thermal bath. By comparing the dynamics of entanglement with that of quantum discord for initial GHZ state and W state, we find that W state is more robust than GHZ state, and quantum discord performs better than entanglement against the decoherence induced by the thermal bath. When the bath temperature is low, for the initial GHZ state, combining weak measurement and measurement reversal is necessary for a successful protection of quantum correlations. But for the initial W state, the protection depends solely upon the measurement reversal. In addition, the protection cannot usually be realized irrespective of the initial states as the bath temperature increases.

  18. 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.

  19. 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…

  20. Transient aero-thermal mapping of passive Thermal Protection system for nose-cap of Reusable Hypersonic Vehicle

    NASA Astrophysics Data System (ADS)

    Mahulikar, Shripad P.; Khurana, Shashank; Dungarwal, Ritesh; Shevakari, Sushil G.; Subramanian, Jayakumar; Gujarathi, Amit V.

    2008-12-01

    The temperature field history of passive Thermal Protection System (TPS) material at the nose-cap (forward stagnation region) of a Reusable Hypersonic Vehicle (RHV) is generated. The 3-D unsteady heat transfer model couples conduction in the solid with external convection and radiation that are modeled as time-varying boundary conditions on the surface. Results are presented for the following two cases: (1) nose-cap comprised of ablative TPS material only (SIRCA/PICA), and (2) nose-cap comprised of a combination of ablative TPS material with moderate thermal conductivity and insulative TPS material. Comparison of the temperature fields of SIRCA and PICA [Case (1)] indicates lowering of the peak stagnation region temperatures for PICA, due to its higher thermal conductivity. Also, the use of PICA and insulative TPS [Case (2)] for the nose-cap has higher potential for weight reduction than the use of ablative TPS alone.

  1. 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.

  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. Design of a Protection Thermal Energy Storage Using Phase Change Material Coupled to a Solar Receiver

    NASA Astrophysics Data System (ADS)

    Verdier, D.; Falcoz, Q.; Ferrière, A.

    2014-12-01

    Thermal Energy Storage (TES) is the key for a stable electricity production in future Concentrated Solar Power (CSP) plants. This work presents a study on the thermal protection of the central receiver of CSP plant using a tower which is subject to considerable thermal stresses in case of cloudy events. The very high temperatures, 800 °C at design point, impose the use of special materials which are able to resist at high temperature and high mechanical constraints and high level of concentrated solar flux. In this paper we investigate a TES coupling a metallic matrix drilled with tubes of Phase Change Material (PCM) in order to store a large amount of thermal energy and release it in a short time. A numerical model is developed to optimize the arrangement of tubes into the TES. Then a methodology is given, based from the need in terms of thermal capacity, in order to help the choice of the geometry.

  4. 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.

  5. 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

  6. 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.

  7. Atomic level description of the protecting effect of osmolytes against thermal denaturation of proteins

    NASA Astrophysics Data System (ADS)

    Pieraccini, Stefano; Burgi, Luigi; Genoni, Alessandro; Benedusi, Anna; Sironi, Maurizio

    2007-04-01

    The protecting effect of the osmolyte molecule taurine against thermal denaturation of the protein Chimotripsin Inhibitor 2 was modelled using Molecular Dynamics simulations. The protein was simulated in denaturing conditions at different taurine concentrations. Analysis of the molecular details of its behaviour shows that the protective effect of the osmolyte is concentration dependent. Moreover, the influence of taurine on the solvent structure was studied. A concentration dependent ordering effect of taurine on water molecules emerges from solvent structure analysis and is well correlated to the protecting effect observed. Based on these observations an interpretation of the osmoprotective effect is proposed.

  8. 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.

  9. 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'

  10. Adaptable Holders for Arc-Jet Screening Candidate Thermal Protection System Repair Materials

    NASA Technical Reports Server (NTRS)

    Riccio, Joe; Milhoan, Jim D.

    2010-01-01

    Reusable holders have been devised for evaluating high-temperature, plasma-resistant re-entry materials, especially fabrics. Typical material samples tested support thermal-protection-system damage repair requiring evaluation prior to re-entry into terrestrial atmosphere. These tests allow evaluation of each material to withstand the most severe predicted re-entry conditions.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  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. 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...

  19. 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...

  20. 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.

  1. 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.

  2. 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.

  3. Development of thermal runaway preventing ZnO varistor for surge protective device.

    PubMed

    Jeoung, Tae-Hoon; Kim, Young-Sung; Nam, Sung-Pill; Lee, Seung-Hwan; Kang, Jeong-Wook; Kim, Jea-Chul; Lee, Sung-Gap

    2014-12-01

    In this paper, the centre of electrode is suggested for heat conduction. Therefore, the specific reflow soldering process is needed. The comparison of temperature difference among the different areas of ZnO varistors is analyzed. With the nominal surge current, thermal behavior is analyzed. The operation point of temperature for disconnection is proposed. Accordingly, the thermal runaway-preventing ZnO varistors were covered with a fusible alloy, i.e., a thermal fuse, in the process of manufacture, which is expected to ensure there the liability of being resistant to lightning discharge and to ensure stability against thermal runaway in the failure mode. Additionally, it is expected to reduce much more limit voltage than the existing products to which the fuse was separately applied. The thermal runaway-preventing ZnO varistor of the surge protection devices can be widely used as part of the protection provisions of lightning discharge and surge protection demanded in connection with power IT about Green Growth which is nowadays becoming the buzzword in the electric power industry. PMID:25970989

  4. 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.

  5. 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.

  6. NDE of thermal protection system for space shuttle solid rocket booster

    NASA Technical Reports Server (NTRS)

    Myers, R. S.

    1990-01-01

    Potential nondestructive test (NDE) methods were evaluated for detecting debonds and weak bonds in the thermal protection system (TPS) for the space shuttle solid rocket boosters. The primary thermal protection material is a sprayable, thick epoxy coating that is filled with lightweight and thermal insulating materials. Test panels were fabricated with a wide variety of hidden realistic defects, including contact debonds and weak bonds. Nondestructive test results were obtained. Candidate NDE methods evaluated for booster production applications include laser interferometry (e.g., electronic shearography), infrared thermography, radiography (e.g., computed tomography), acousto-ultrasonics, mechanical/acoustic impedance, ultrasonics, acoustic emission, and the tap test. Capabilities, advantages, disadvantages, and relative performances in defect detection of each test method for TPS bonding applications are reported. Electronic shearography NDE was technically the superior method for detecting debonds.

  7. 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.

  8. 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.

  9. 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.

  10. Wireless subsurface microsensors for health monitoring of thermal protection systems on hypersonic vehicles

    NASA Astrophysics Data System (ADS)

    Milos, Frank S.; Watters, David G.; Pallix, Joan B.; Bahr, Alfred J.; Huestis, David L.

    2001-07-01

    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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. Durability of thermal control and environmental protective materials for the SSRMS in simulated LEO environment

    NASA Astrophysics Data System (ADS)

    Chang, S. K.

    1993-06-01

    Nine thermal control and environmental protection materials, selected on the basis of their space pedigree, thermal vacuum stability, and thermo-optical properties, were tested to determine their suitability for the Space Station Remote Manipulator System (SSRMS). The ground based testing was carried out to simulate the effects of atomic oxygen and thermal cycling in the Low Earth Orbit (LEO) environment. These factors are deemed most likely to cause degradation to the selected materials. With the exception of the urethane based coatings, the materials tested demonstrate sufficient resistance to atomic oxygen. The detrimental effect of thermal cycling on the adhesion of the silicate based coatings to aluminum substrate was found to depend on the pigment. A separate experiment on Beta-Cloth showed that its thermo-optical properties remained substantially unchanged as the Teflon coating was progressively removed in a plasma asher.

  17. Thermal protection of re-entry vehicles with the usage of inflatable systems

    NASA Astrophysics Data System (ADS)

    Alifanov, O. M.; Outchvatov, V. I.; Pichkhadze, K. M.

    2003-08-01

    The Report comprises the results of theoretical and experimental research aimed at development of special devices intended to increase the effectiveness of spacecraft (SC) descent-systems onto the surface of planets with atmosphere. The subject of research is the thermal protection of a SC. Inflatable braking devices (IBD) made of elastic materials and substantially reducing the value SC's ballistic parameter are considered as special devices. Parametric evaluations of influence of descent SC's ballistic coefficient on specific convective heat fluxes at the zone critical point [ qc = A1·(σ) 1/2] were the basis of experimental research. A 1 coefficient is a function of a number of parameters (given descent vehicle (DV) configuration, atmospheric parameters, angle of entry, and empirical estimations of experiments). Analysis and experiments have shown that real materials permit to develop inflatable devices, geometrical and mass properties of which allow reducing a value of σ by 2 … 2.5 times in comparison with the existing ones. This leads to a reduction of thermal fluxes by 10 … 15 times. At such an intensity of external convective heat fluxes, the most effective thermal protection of a DV turns out to be a protection composed by means of a so-called "heat shields method". Their emanation provides an equilibrium state of the system. Parametric dependency of shield surface temperature at a critical point is defined as: T K = A2·( q c) 1/8 In accordance with the results of expert estimations, the Report describes the dependency of ballistic coefficients values of DV vs. payload (PL) masses regarding producible modern materials. In the framework of projects fulfilled on the basis of current research, elastic thermal protection was developed in order to protect a critical point zone of DV. The Report gives a comparison of possible mass properties of IBD-based DV with analogous parameters of real DV carrying out ballistic trajectory descent.

  18. Thermal comfort range of a military cold protection glove: database by thermophysiological simulation.

    PubMed

    Zimmermann, Carsten; Uedelhoven, Wolfgang H; Kurz, Bernhard; Glitz, Karl Jochen

    2008-09-01

    The thermal insulation properties of a military wet/cold protection glove of the German Bundeswehr were investigated using the thermophysiological simulation device CYBOR with a heated full-scale hand model. The aim of this study was the physiology related and more reliable estimation of a database for the thermal comfort range of the glove in terms of environmental limit temperatures and maximum safe wearing times (limit times). For that purpose the simulation device CYBOR is equipped with a control feature allowing the simulation of the physiological effect that the blood flow into the hands as the dominant heat source is reduced with decreasing skin temperature (vasoconstriction effect). In the simulation test, the criterion defining the thermal comfort range of the glove was the maintenance of a minimum hand phantom skin temperature of 15 degrees C. For various assumed metabolic rates between 50 and 175 W m(-2) and environmental temperatures down to -22 degrees C, the maximum safe wearing times within the thermal comfort range of the military glove were estimated between only 20 min and almost 1 h. The used simulation scenario for the prediction of environmental limit temperatures, however, tends to deliver too low values in correlation to the estimated limit times and needs further critical consideration. The estimated data concerning the thermal comfort range of the wet/cold protection glove of the German Bundeswehr leads to the recommendation for a use of this model in mild cold climatic regions. The presented thermophysiological simulation procedure for the evaluation of the cold protection properties of gloves in terms of maximum safe wearing times within the thermal comfort range can be a useful tool to establish practical operating instructions for soldiers or civilians acting in cold environments. PMID:18172670

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

    NASA Technical Reports Server (NTRS)

    Dec, John A.; Braun, Robert D.

    2006-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Dec, John A.; Braun, Robert D.

    2005-01-01

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

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

    DOE PAGESBeta

    Marchevsky, M.; Turqueti, M.; Cheng, D. W.; Felice, H.; Sabbi, G.; Salmi, T.; Stenvall, A.; Chlachidze, G.; Ambrosio, G.; Ferracin, P.; et al

    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. 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. PMID:25925745

  5. 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.

  6. 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.

  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. Acoustic Characterization and Impact Sensing for Ceramic Thermal Protection Systems (TPS)

    SciTech Connect

    Kuhr, S. J.; Reibel, R.; Sathish, S.; Jata, K. V.

    2006-03-06

    A study was conducted to understand acoustic wave propagation characteristics in a ceramic matrix composite (CMC) wrapped tile thermal protection system (CMC+ Foam+ RTV+ SIP+ RTV+ Al) and ceramic foam. Sound velocities were measured in three orthogonal directions on the above material. The attenuation coefficients were also determined for a uncoated ceramic foam. Commercially available standard acoustic emission transducers, piezo-wafers and polymer based PVDF (polyvinylidiene fluoride) film were employed in the experiments to acquire the acoustic data. The performance characteristics of these sensors will be discussed in light of impact detection. Variation in the wave propagation characteristics along different directions and the role of processing in causing anisotropic acoustic properties in thermal protection systems will be discussed.

  9. Effects of natural environment on first generation solid rocket booster thermal protection system materials

    NASA Technical Reports Server (NTRS)

    Webb, D. D.

    1988-01-01

    The effort to demonstrate, by real-time exposure, the effects of the natural environment at Kennedy Space Center, Florida, upon the Thermal Protection System (TPS) of the Solid Rocket Booster (SRB) is summarized, and that the overall SRB TPS configuration is verified to meet all requirements for resistance to the conditions associated with outdoor weathering, including: solar radiation; temperature; humidity; precipitation; wind; sand/dust abrasion; static electricity; salt spray; fungus; and atmospheric oxidants. The evaluation criterion for this project was based upon flatwise tensile properties, visual inspection, color change, and thermal performance. Based upon the evaluation of the changes in these properties, it is concluded that properly applied and topcoat-protected TPS can satisfactorily withstand the conditions of the natural environment at KSC for exposures up to six months.

  10. 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.

  11. 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.

  12. STS-28 Columbia, OV-102, thermal protection system (TPS) tile repair

    NASA Technical Reports Server (NTRS)

    1988-01-01

    On Columbia, Orbiter Vehicle (OV) 102, underside, technician prepares surface during thermal protection system (TPS) tile repair / replacement. OV-102 is being refurbished for the STS-28 Department of Defense (DOD) dedicated mission in the Kennedy Space Center (KSC) Orbiter Processing Facility (OPF) high bay 2. Technician stands on scaffolding next to deployed landing gear. View provided by KSC with alternate number KSC-87PC-126.

  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. Design of a built-in health monitoring system for bolted thermal protection panels

    NASA Astrophysics Data System (ADS)

    Yang, Jinkyu; Chang, Fu-Kuo; Derriso, Mark M.

    2003-08-01

    Space vehicles require high performance thermal protection systems (TPS) that provide high temperature insulation capability with lower weight, high strength, and reliable integration with the existing system. Carbon-carbon panels mounted with bracket joints are potential future thermal protection systems with light weight, low creep, and high stiffness at high temperatures. However, the thermal protection system experiences a very harsh high-temperature and aerodynamic environment in addition to foreign object impacts. Damage or failure of panels without being detected can lead to catastrophe. Therefore, knowledge of the integrity of the thermal protection system before each launch and reentry is essential to the success of the mission. The objective of the study is to develop a built-in diagnostic system to assess the integrity of TPS panels as well as to lower inspection and maintenance time and costs. An integrated structural health monitoring system is being developed to monitor the TPS panels. The technology includes investigation of the loosening of bolts which connects TPS panels to the supporting structure, and potentially, identifying the location of damage on the panel caused by external impacts from micrometeorites and other objects. The first generation prototype was manufactured and tested in an acoustic chamber which simulated a re-entry environment to investigate the feasibility of the health monitoring system focusing on its survivability and sensitivity. The preliminary results were very promising. Based on the test results, the second generation design was proposed to improve the performance of the first generation design. To put a reliable and accurate decision on the diagnostics of the TPS panels, an advanced algorithm was developed with the aid of a wavelet transform technique.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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

    2016-08-01

    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. PMID:27420726

  20. 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

  1. 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.

  2. 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.

  3. 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).

  4. 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.

  5. 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.

  6. 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).

  7. 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.

  8. Fiber optic microsensor hydrogen leak detection system on Aerospike X-33

    NASA Astrophysics Data System (ADS)

    Kazemi, Alex A.; Goepp, John W.; Larson, David B.; Wuestling, Mark E.

    2007-09-01

    Commercial and military launch vehicles are designed to use cryogenic hydrogen as the main propellant, which is very volatile, extremely flammable, and highly explosive. Current detection system uses Teflon transfer tubes at small number of vehicle location through which gas samples are drawn and stream analyzed by a mass spectrometer. A concern with this approach is the high cost of the system. Also, the current system does not provide leak location and is not in real time. This system is very complex and cumbersome for production and ground support measurement personnel. This paper describes the successful test of a multipoint fiber optic hydrogen microsensors system on the Linear Aerospike X-33 rocket engine at NASA's Stennis Flight Center. The system consisted of a reversible chemical interaction causing a change in reflective of a thin film of coated Palladium. The sensor using a passive element consisting of chemically reactive microcoatings deposited on the surface of a glass microlens, which is then bonded to an optical fiber. The system uses a multiplexing technique with a fiber optic driver-receiver consisting of a modulated LED source that is launched into the sensor, and photodiode detector that synchronously measures the reflected signal. The system incorporates a microprocessor to perform the data analysis and storage, as well as trending and set alarm function. The paper illustrates the sensor design and performance data under field deployment conditions.

  9. 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.

  10. 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.

  11. 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.

  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. Methods of evaluating protective clothing relative to heat and cold stress: thermal manikin, biomedical modeling, and human testing.

    PubMed

    O'Brien, Catherine; Blanchard, Laurie A; Cadarette, Bruce S; Endrusick, Thomas L; Xu, Xiaojiang; Berglund, Larry G; Sawka, Michael N; Hoyt, Reed W

    2011-10-01

    Personal protective equipment (PPE) refers to clothing and equipment designed to protect individuals from chemical, biological, radiological, nuclear, and explosive hazards. The materials used to provide this protection may exacerbate thermal strain by limiting heat and water vapor transfer. Any new PPE must therefore be evaluated to ensure that it poses no greater thermal strain than the current standard for the same level of hazard protection. This review describes how such evaluations are typically conducted. Comprehensive evaluation of PPE begins with a biophysical assessment of materials using a guarded hot plate to determine the thermal characteristics (thermal resistance and water vapor permeability). These characteristics are then evaluated on a thermal manikin wearing the PPE, since thermal properties may change once the materials have been constructed into a garment. These data may be used in biomedical models to predict thermal strain under a variety of environmental and work conditions. When the biophysical data indicate that the evaporative resistance (ratio of permeability to insulation) is significantly better than the current standard, the PPE is evaluated through human testing in controlled laboratory conditions appropriate for the conditions under which the PPE would be used if fielded. Data from each phase of PPE evaluation are used in predictive models to determine user guidelines, such as maximal work time, work/rest cycles, and fluid intake requirements. By considering thermal stress early in the development process, health hazards related to temperature extremes can be mitigated while maintaining or improving the effectiveness of the PPE for protection from external hazards. PMID:21936698

  14. Construction of new Pichia pastoris X-33 strains for production of lycopene and β-carotene.

    PubMed

    Araya-Garay, J M; Feijoo-Siota, L; Rosa-dos-Santos, F; Veiga-Crespo, P; Villa, T G

    2012-03-01

    In this study, we used the non-carotenogenic yeast Pichia pastoris X33 as a receptor for β-carotene-encoding genes, in order to obtain new recombinant strains capable of producing different carotenoidic compounds. We designed and constructed two plasmids, pGAPZA-EBI* and pGAPZA-EBI*L*, containing the genes encoding lycopene and β-carotene, respectively. Plasmid pGAPZA-EBI*, expresses three genes, crtE, crtB, and crtI*, that encode three carotenogenic enzymes, geranylgeranyl diphosphate synthase, phytoene synthase, and phytoene desaturase, respectively. The other plasmid, pGAPZA-EBI*L*, carried not only the three genes above mentioned, but also the crtL* gene, that encodes lycopene β-cyclase. The genes crtE, crtB, and crtI were obtained from Erwinia uredovora, whereas crtL* was cloned from Ficus carica (JF279547). The plasmids were integrated into P. pastoris genomic DNA, and the resulting clones Pp-EBI and Pp-EBIL were selected for either lycopene or β-carotene production and purification, respectively. Cells of these strains were investigated for their carotenoid contents in YPD media. These carotenoids produced by the recombinant P. pastoris clones were qualitatively and quantitatively analyzed by high-resolution liquid chromatography, coupled to photodiode array detector. These analyses confirmed that the recombinant P. pastoris clones indeed produced either lycopene or β-carotene, according to the integrated vector, and productions of 1.141 μg of lycopene and 339 μg of β-carotene per gram of cells (dry weight) were achieved. To the best of our knowledge, this is the first time that P. pastoris has been genetically manipulated to produce β-carotene, thus providing an alternative source for large-scale biosynthesis of carotenoids. PMID:22159890

  15. 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.

  16. Replacement of Ablators with Phase-Change Material for Thermal Protection of STS Elements

    NASA Technical Reports Server (NTRS)

    Kaul, Raj K.; Stuckey, Irvin; Munafo, Paul M. (Technical Monitor)

    2002-01-01

    As part of the research and development program to develop new Thermal Protection System (TPS) materials for aerospace applications at NASA's Marshall Space Flight Center (MSFC), an experimental study was conducted on a new concept for a non-ablative TPS material. Potential loss of TPS material and ablation by-products from the External Tank (ET) or Solid Rocket Booster (SRB) during Shuttle flight with the related Orbiter tile damage necessitates development of a non-ablative thermal protection system. The new Thermal Management Coating (TMC) consists of phase-change material encapsulated in micro spheres and a two-part resin system to adhere the coating to the structure material. The TMC uses a phase-change material to dissipate the heat produced during supersonic flight rather than an ablative material. This new material absorbs energy as it goes through a phase change during the heating portion of the flight profile and then the energy is slowly released as the phase-change material cools and returns to its solid state inside the micro spheres. The coating was subjected to different test conditions simulating design flight environments at the NASA/MSFC Improved Hot Gas Facility (IHGF) to study its performance.

  17. Role of HSP70 in cytoplasm protection against thermal stress in rohu, Labeo rohita.

    PubMed

    Giri, Sib Sankar; Sen, Shib Sankar; Sukumaran, V

    2014-12-01

    To understand the function of HSP70 of Labeo rohita (LrHSP70) in cellular protection, LrHSP70 ORF cDNA was inserted into the plasmid of pET-32a(+) or pEGFP-L1. Then, the recombinant plasmids were transformed or transfected into Escherichia coli cells, mouse myeloma cells (MPC-11) or fish hepatoma cells (PLHC-1). Western blot results revealed that LrHSP70 was expressed in E. coli cells and molecular weight was estimated to be 70 kDa. In cells, LrHSP70 was over-expressed following thermal or cold stress. Results revealed that LrHSP70 protected prokaryotic cells against thermal or cold extremes as well as played the same role in MPC-11 and PLHC-1 cells. After heat treatment at 42 °C for 1 h, the viability of the cell was declined considerably. PLHC-1 cells with pEGFP-L1/LrHSP70 exhibited a higher survival rate (50%) than wild-type cells (18%) or cells with only pEGFP-L1 (21.2%). When the time lag extended to 2 h, the survival rates were 30%, 3.4% and 5.3% respectively. The present study revealed that LrHSP70 plays an important role in response to thermal and cold stress in fish. PMID:25240978

  18. Development of FIAT-Based Parametric Thermal Protection System Mass Estimating Relationships for NASA's Multi-Mission Earth Entry Concept

    NASA Astrophysics Data System (ADS)

    Sepka, S. A.; Samareh, J. A.

    2014-06-01

    Mass estimating relationships have been formulated to determine a vehicle's Thermal Protection System material and required thickness for safe Earth entry. We focus on developing MERs, the resulting equations, model limitations, and model accuracy.

  19. 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.

  20. 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.

  1. 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

  2. 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. PMID:25012597

  3. 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.

  4. 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.

  5. 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

  6. 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. PMID:25976784

  7. 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.

  8. 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.

  9. 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

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. Molecular outgassing measurements for an element of the Shuttle thermal protection system

    NASA Technical Reports Server (NTRS)

    Visentine, J. T.; Richmond, R. G.; Kelso, R. M.

    1976-01-01

    Molecular outgassing studies were conducted on a thermal-protection material recently developed for the Shuttle Orbiter. Molecular outgassing rates as low as 10 to the minus 11th g/sq cm/sec, condensation coefficients, and molecular desorption rates were measured in four separate experiments using cryogenically and thermoelectrically cooled quartz crystal microbalances. Although the initial outgassing rates are high, they decreased to values in the 10 to the minus 10th g/sq cm/sec range in a reasonable period of time. Outgassing rates do not increase after entry heating although the condensation coefficients at various microbalance collection-surface temperatures become somewhat larger.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  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. 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.

  5. 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:

  6. 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.

  7. 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.

  8. Collisional quenching of atoms and molecules on spacecraft thermal protection surfaces

    NASA Technical Reports Server (NTRS)

    Marinelli, W. J.; Green, B. D.

    1988-01-01

    Preliminary results of a research program to determine energy partitioning in spacecraft thermal protection materials due to atom recombination at the gas-surface interface are presented. The primary focus of the research is to understand the catalytic processes which determine heat loading on Shuttle, Aeroassisted OTV, and NASP thermal protection surfaces in nonequilibrium flight regimes. Highly sensitive laser diagnostics based on laser-induced fluorescence and resonantly-enhanced multiphoton ionization spectroscopy are used to detect atoms and metastable molecules. At low temperatures, a discharge flow reactor is employed to measure deactivation/recombination coefficients for O-atoms, N-atoms, and O2. Detection methods are presented for measuring O-atoms, O2 and N2, and results for deactivation of O2 and O-atoms on reaction-cured glass and Ni surfaces. Both atom recombination and metastable product formation are examined. Radio-frequency discharges are used to produce highly dissociated beams of atomic species at energies characteristic of the surface temperature. Auger electron spectroscopy is employed as a diagnostic of surface composition in order to accurately define and control measurement conditions.

  9. 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.

  10. 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. PMID:24793820

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

    NASA Technical Reports Server (NTRS)

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

    1988-01-01

    The thermal protection systems (TPS) for the Shuttle/Centaur were designed to provide fail-safe thermal protection during prelaunch, launch ascent, and on-orbit operations as well as during potential abort, where the Shuttle and Centaur would return to earth. The TPS selected used a helium-purged polyimide foam beneath three radiation shields for the liquid-hydrogen (LH2) tank and radiation shields only for the liquid-oxygen (LO2) tank. A double-walled vacuum bulkhead separated the two tanks. The LH2 tank had one 1.9 cm-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 gave total prelaunch heating rates of 29.5 and 12.9 kW for the LH2 and LO2 tanks, respectively. Calorimeter tests on a representative sample of the LH2 tank sidewall TPS indicated that the measured unit heating one would rapidly decrease from the prelaunch rate of about 300 W/sq m to a desired rate less than 4 W/sq m once on-orbit.

  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. 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.

  14. 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.

  15. 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.

  16. Performance of thermal control tape in the protection of composite materials

    NASA Technical Reports Server (NTRS)

    Kamenetzky, Rachel R.; Whitaker, Ann F.

    1992-01-01

    The selection of materials for construction of long duration mission spacecraft has presented many challenges to the aerospace design community. After nearly six years in low earth orbit, NASA's Long duration Exposure Facility (LDEF), retrieved in January of 1990, has provided valuable information on both the nature of the space environment as well as the effects of the space environment on potential spacecraft materials. Composites, long a favorite of the design community because of a high strength-to-weight ratio, were flown in various configurations on LDEF in order to evaluate the effects of radiation, atomic oxygen, vacuum, micrometeoroid debris, and thermal variation on their performance. Fiberglass composite samples covered with an aluminum thermal control tape were flown as part of the flight experiment A0171, the Solar Array Materials Passive LDEF Experiment (SAMPLE). Visual observations and test results indicate that the thermal control tape suffered little degradation from the space exposure and proved to be a reliable source of protection from atomic oxygen erosion and UV radiation for the underlying composite material.

  17. New Approach for Thermal Protection System of a Probe During Entry

    NASA Technical Reports Server (NTRS)

    Yendler, Boris; Poffenbarger, Nathan; Patel, Amisha; Bhave, Ninad; Papadopoulos, Periklis

    2005-01-01

    One of the biggest challenges for any thermal protection system (TPS) of a probe is to provide a sufficient barrier for heat generated during descent in order to keep the temperature inside of the probe low enough to support operational temperature of equipment. Typically, such a goal is achieved by having the ceramic tiles and blankets like on the Space Shuttle, silicon based ablators, or metallic systems to cover the probe external surface. This paper discusses the development of an innovative technique for TPS of the probe. It is proposed to use a novel TPS which comprises thermal management of the entry vehicle. It includes: a) absorption of the heat during heat pick load by a Phase Change Material (PCM), b) separation of the compartment which contains PCM from the rest of the space vehicle by a gap with a high thermal resistance, c) maintaining temperature of the internal wall of s/c cabin temperature by transfer heat from the internal wall to the "cold" side of the vehicle and to reject heat into the space during the flight and on a ground, d) utilization of an advanced heat pipe, so called Loop Heat Pipe to transfer heat from the cabin internal wall to the cold side of the s/c and to reject the heat into environment outside of the vehicle. A Loop Heat Pipe is capable of transferring heat against gravity

  18. Development of metallic thermal protection systems for the reusable launch vehicle

    SciTech Connect

    Blosser, M.L.

    1997-01-01

    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. {copyright} {ital 1997 American Institute of Physics.}

  19. 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.

  20. A meningococcal vaccine antigen engineered to increase thermal stability and stabilize protective epitopes

    PubMed Central

    Konar, Monica; Pajon, Rolando; Beernink, Peter T.

    2015-01-01

    Factor H binding protein (FHbp) is part of two vaccines recently licensed for prevention of sepsis and meningitis caused by serogroup B meningococci. FHbp is classified in three phylogenic variant groups that have limited antigenic cross-reactivity, and FHbp variants in one of the groups have low thermal stability. In the present study, we replaced two amino acid residues, R130 and D133, in a stable FHbp variant with their counterparts (L and G) from a less stable variant. The single and double mutants decreased thermal stability of the amino- (N-) terminal domain compared with the wild-type protein as measured by scanning calorimetry. We introduced the converse substitutions, L130R and G133D, in a less stable wild-type FHbp variant, which increased the transition midpoint (Tm) for the N-terminal domain by 8 and 12 °C; together the substitutions increased the Tm by 21 °C. We determined the crystal structure of the double mutant FHbp to 1.6 Å resolution, which showed that R130 and D133 mediated multiple electrostatic interactions. Monoclonal antibodies specific for FHbp epitopes in the N-terminal domain had higher binding affinity for the recombinant double mutant by surface plasmon resonance and to the mutant expressed on meningococci by flow cytometry. The double mutant also had decreased binding of human complement Factor H, which in previous studies increased the protective antibody responses. The stabilized mutant FHbp thus has the potential to stabilize protective epitopes and increase the protective antibody responses to recombinant FHbp vaccines or native outer membrane vesicle vaccines with overexpressed FHbp. PMID:26627237

  1. Space Shuttle Thermal Protection System Repair Flight Experiment Induced Contamination Impacts

    NASA Technical Reports Server (NTRS)

    Smith, Kendall A.; Soares, Carlos E.; Mikatarian, Ron; Schmidl, Danny; Campbell, Colin; Koontz, Steven; Engle, Michael; McCroskey, Doug; Garrett, Jeff

    2006-01-01

    NASA s activities to prepare for Flight LF1 (STS-114) included development of a method to repair the Thermal Protection System (TPS) of the Orbiter s leading edge should it be damaged during ascent by impacts from foam, ice, etc . Reinforced Carbon-Carbon (RCC) is used for the leading edge TPS. The repair material that was developed is named Non- Oxide Adhesive eXperimental (NOAX). NOAX is an uncured adhesive material that acts as an ablative repair material. NOAX completes curing during the Orbiter s descent. The Thermal Protection System (TPS) Detailed Test Objective 848 (DTO 848) performed on Flight LF1 (STS-114) characterized the working life, porosity void size in a micro-gravity environment, and the on-orbit performance of the repairs to pre-damaged samples. DTO 848 is also scheduled for Flight ULF1.1 (STS-121) for further characterization of NOAX on-orbit performance. Due to the high material outgassing rates of the NOAX material and concerns with contamination impacts to optically sensitive surfaces, ASTM E 1559 outgassing tests were performed to determine NOAX condensable outgassing rates as a function of time and temperature. Sensitive surfaces of concern include the Extravehicular Mobility Unit (EMU) visor, cameras, and other sensors in proximity to the experiment during the initial time after application. This paper discusses NOAX outgassing characteristics, how the amount of deposition on optically sensitive surfaces while the NOAX is being manipulated on the pre-damaged RCC samples was determined by analysis, and how flight rules were developed to protect those optically sensitive surfaces from excessive contamination where necessary.

  2. Cardiovascular and thermal consequences of protective clothing: a comparison of clothed and unclothed states.

    PubMed

    Fogarty, Alison; Armstrong, Karen; Gordon, Christopher; Groeller, Herbert; Woods, Brian; Stocks, Jodie; Taylor, Nigel

    2004-08-15

    We have undertaken a laboratory-based examination of the cardiovascular and thermal impact of wearing thermal (heat) protective clothing during fatiguing exercise in the heat. Seven males completed semi-recumbent, intermittent cycling (39.6 degrees C, 45% relative humidity) wearing either protective clothing or shorts (control). Mean core and skin temperatures, cardiac frequency (f(c)), stroke volume (Q), cardiac output (Q), arterial pressure, forearm blood flow (Q(f)), plasma volume change, and sweat rates were measured. In the clothed trials, subjects experienced significantly shorter times to fatigue (52.5 vs. 58.9 min), at lower peak work rates (204.3 vs. 277.4 W), and with higher core (37.9 degrees vs. 37.5 degrees C) and mean skin temperatures (37.3 degrees vs. 36.9 degrees C). There was a significant interaction between time and clothing on f(c), such that, over time, the clothing effect became more powerful. Clothing had a significant main affect on Q, but not Q, indicating the higher Q was chronotropically driven. Despite a greater sweat loss when clothed (923.0 vs. 547.1 g.m(-2) x h(-1); P<0.05), Q(f) and plasma volume change remained equivalent. Protective clothing reduced exercise tolerance, but did not affect overall cardiovascular function, at the point of volitional fatigue. It was concluded that, during moderately heavy, semi-recumbent exercise under hot, dry conditions, the strain on the unclothed body was already high, such that the additional stress imparted by the clothing ensemble represented a negligible, further impact upon cardiovascular stability. PMID:15370864

  3. A meningococcal vaccine antigen engineered to increase thermal stability and stabilize protective epitopes.

    PubMed

    Konar, Monica; Pajon, Rolando; Beernink, Peter T

    2015-12-01

    Factor H binding protein (FHbp) is part of two vaccines recently licensed for prevention of sepsis and meningitis caused by serogroup B meningococci. FHbp is classified in three phylogenic variant groups that have limited antigenic cross-reactivity, and FHbp variants in one of the groups have low thermal stability. In the present study, we replaced two amino acid residues, R130 and D133, in a stable FHbp variant with their counterparts (L and G) from a less stable variant. The single and double mutants decreased thermal stability of the amino- (N-) terminal domain compared with the wild-type protein as measured by scanning calorimetry. We introduced the converse substitutions, L130R and G133D, in a less stable wild-type FHbp variant, which increased the transition midpoint (Tm) for the N-terminal domain by 8 and 12 °C; together the substitutions increased the Tm by 21 °C. We determined the crystal structure of the double mutant FHbp to 1.6 Å resolution, which showed that R130 and D133 mediated multiple electrostatic interactions. Monoclonal antibodies specific for FHbp epitopes in the N-terminal domain had higher binding affinity for the recombinant double mutant by surface plasmon resonance and to the mutant expressed on meningococci by flow cytometry. The double mutant also had decreased binding of human complement Factor H, which in previous studies increased the protective antibody responses. The stabilized mutant FHbp thus has the potential to stabilize protective epitopes and increase the protective antibody responses to recombinant FHbp vaccines or native outer membrane vesicle vaccines with overexpressed FHbp. PMID:26627237

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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

  11. 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.

  12. 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. PMID:25816966

  13. The Evolution of Nondestructive Evaluation Methods for the Space Shuttle External Tank Thermal Protection System

    NASA Technical Reports Server (NTRS)

    Walker, James L.; Richter, Joel D.

    2006-01-01

    Three nondestructive evaluation methods are being developed to identify defects in the foam thermal protection system (TPS) of the Space Shuttle External Tank (ET). Shearography is being developed to identify shallow delaminations, shallow voids and crush damage in the foam while terahertz imaging and backscatter radiography are being developed to identify voids and cracks in thick foam regions. The basic theory of operation along with factors affecting the results of these methods will be described. Also, the evolution of these methods from lab tools to implementation on the ET will be discussed. Results from both test panels and flight tank inspections will be provided to show the range in defect sizes and types that can be readily detected.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. Microstructural characterization of the HRSI thermal protection system for space shuttle

    NASA Technical Reports Server (NTRS)

    Ransone, P. O.; Rummler, D. R.

    1980-01-01

    Components of the space shuttle high temperature reusable surface insulation (HRSI) system were microscopically characterized, both separately and as a system, to obtain information needed for stress analysis models of the thermal protection system. A tension specimen of the HRSI system was loaded in steps and was microscopically observed at each load condition to demonstrate the tension failure mode associated with strain isolation pad (SIP) behavior. A local failure occurred which should be associated with transfer of load through transverse fibers in the SIP. Stress concentrations attributed to the SIP behavior necessitated strengthening of the HRSI by densification of the RSI at the bondline. An HRSI tile was microscopically characterized after the densification process. The densified surface layer blended into the RSI which caused a gradual change in density. The gradation in density does not appear to represent a sharp discontinuity in elastic modulus between the densified layer and the parent material.

  19. Assessment of the Shuttle Orbiter Thermal Protection System undensified tiles under mission dynamic loads

    NASA Technical Reports Server (NTRS)

    Cooper, P. A.; Miserentino, R.; Sawyer, J. W.; Leatherwood, J. D.

    1982-01-01

    A description is provided of the results of a series of base driven random dynamic load tests of the Thermal Protection System (TPS). The results were used to insure the integrity of the undensified TPS prior to the first flight in the wing and mid-fuselage region. The number of specimens and the load ranges investigated were limited. Attention is given to the test specimens, the test procedure, a data analysis, and a discussion of the results. All specimens for both the wing and mid-fuselage regions survived an equivalent of 72 ascent missions and exhibited residual static strength greater than their original proof loads. These results indicate that the undensified tiles had sufficient strength to withstand ascent loads during the first flight.

  20. 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.

  1. An Assessment of Alternate Thermal Protection Systems for the Space Shuttle Orbiter. Volume 1; Executive Summary

    NASA Technical Reports Server (NTRS)

    Hays, D.

    1982-01-01

    Alternate thermal protection system (TPS) concepts to the Space Shuttle Orbiter were assessed. Metallic, ablator, and carbon-carbon concepts which are the result of some previous design, manufacturing and testing effort were considered. Emphasis was placed on improved TPS durability, which could potentially reduce life cycle costs and improve Orbiter operational characteristics. Integrated concept/orbiter point designs were generated and analyzed on the basis of Shuttle design environments and criteria. A merit function evaluation methodology based on mission impact, life cycle costs, and risk was developed to compare the candidate concepts and to identify the best alternate. Voids and deficiencies in the technology were identified, along with recommended activities to overcome them. Finally, programmatic plans, including ROM costs and schedules, were developed for all activities required to bring the selected alternate system up to operational readiness.

  2. 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.

  3. 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.

  4. 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

  5. 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.

  6. Characterization and modeling of an advanced flexible thermal protection material for space applications

    NASA Technical Reports Server (NTRS)

    Clayton, Joseph P.; Tinker, Michael L.

    1991-01-01

    This paper describes experimental and analytical characterization of a new flexible thermal protection material known as Tailorable Advanced Blanket Insulation (TABI). This material utilizes a three-dimensional ceramic fabric core structure and an insulation filler. TABI is the leading candidate for use in deployable aeroassisted vehicle designs. Such designs require extensive structural modeling, and the most significant in-plane material properties necessary for model development are measured and analytically verified in this study. Unique test methods are developed for damping measurements. Mathematical models are developed for verification of the experimental modulus and damping data, and finally, transverse properties are described in terms of the inplane properties through use of a 12-dof finite difference model of a simple TABI configuration.

  7. 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

  8. 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.

  9. 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.

  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. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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

  16. Increasing refiner production by using motor thermal capacity for protection and control

    SciTech Connect

    Grainger, L.G.; McDonald, M.C.

    1997-05-01

    Industrial motors are typically controlled and operated by closely monitoring the stator winding temperatures and limiting the phase currents within the motor manufacturer`s full-load ampacity rating. A different approach to motor operation and control was implemented at the Blue Ridge Lumber medium density fiberboard (MDF) plant at Whitecourt, Alta., Canada. The capacity control of the refiner is based on using the remaining thermal capacity of the motor as the primary control parameter. In this installation, a 4,000-hp totally enclosed water air cooled (TEWAC) squirrel-cage induction motor is continuously operating above the manufacturer`s rated full-load current, but is being controlled by maintaining thermal capacity at 50%. Temporary current loadings well above this are permitted for up to several minutes to accommodate variations in the wood feed stock to the refiner. This was implemented by installing a modern motor protection relay, communication with a programmable logic controller (PLC) system, and the development of operator interface displays to provide plant operators with the necessary information to monitor the motor parameters. Factors which needed to be considered were the electrical power system limitations, the motor cooling effectiveness, and mechanical limitations imposed by the refiner shaft design.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. Evaluation of the thermal performance of fire fighter protective clothing with the addition of phase change material

    NASA Astrophysics Data System (ADS)

    McCarthy, Lee K.

    Fire fighters rely on fire fighter protective clothing (FFPC) to provide adequate protection in the various hazardous environments they may encounter during operations. FFPC has seen significant advancement in technology over the past few decades. The addition of phase change material (PCM) to FFPC is a new technology with potential to enhance the thermal protection provided by the FFPC. To explore this technology, data from bench-scale experiments involving FFPC with PCMs are compared with a theoretical finite difference heat transfer model. The results demonstrate an effective method to mathematically model the heat transfer and provide insight into the effectiveness of improving the thermal protection of FFPC. The experiments confirm that the latent heat absorbed during the phase change reduces temperatures that might be experienced at the fire fighter's skin surface, advancing the high temperature performance of FFPC.

  4. 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.

  5. 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

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. Thermal protection system technology and facility needs for demanding future planetary missions

    NASA Astrophysics Data System (ADS)

    Laub, B.; Venkatapathy, E.

    2004-02-01

    NASA has successfully launched numerous science missions to inner and outer planets in our solar system of which the most challenging were to Venus and Jupiter and the knowledge gained from those missions have been invaluable yet incomplete. Future missions will be built on what we have learned from the past missions but they will be more demanding from both the science as well as the mission design and engineering perspectives. The Solar System Exploration Decadal Survey (SSEDS) produced for NASA by the National Research Council identified a broad range of science objectives many of which can only be satisfied with atmospheric entry probes. The SSEDS recommended new probe/lander missions to both Venus and Jupiter. The Pioneer-Venus probe mission was launched in August 1978 and four probes successfully entered the Venusian atmosphere in December 1978. The Galileo mission was launched in October 1989 and one probe successfully entered the Jovian atmosphere in December 1995. The thermal protection system requirements for these two missions were unlike any other planetary probes and required fully dense carbon phenolic for the forebody heat shield. Developing thermal protection systems to accomplish future missions outlined in the Decadal Survey presents a technology challenge since they will be more demanding than these past missions. Unlike Galileo, carbon phenolic may not be an adequate TPS for a future Jupiter multiprobe mission since non-equatorial probes will enter at significantly higher velocity than the Galileo equatorial probe and the entry heating scales approximately with the cube of the entry velocity. At such heating rates the TPS mass fraction for a carbon phenolic heat shield would be prohibitive. A new, robust and efficient TPS is required for such probes. The Giant Planet Facility (GPF), developed and employed during the development of the TPS for the Galileo probe was dismantled after completion of the program. Furthermore, flight data from the

  15. THERMAL INSTABILITY OF ΔF508 CFTR CHANNEL FUNCTION: PROTECTION BY SINGLE SUPPRESSOR MUTATIONS AND INHIBITING CHANNEL ACTIVITY

    PubMed Central

    Liu, Xuehong; O’Donnell, Nicolette; Landstrom, Allison; Skach, William R.; Dawson, David C.

    2012-01-01

    Deletion of Phe508 from CFTR results in a temperature-sensitive folding defect that impairs protein maturation and chloride channel function. Both of these adverse effects, however, can be mitigated to varying extents by second-site, suppressor mutations. To better understand the impact of second-site mutations on channel function, we compared the thermal sensitivity of CFTR channels in Xenopus oocytes. CFTR-mediated conductance of oocytes expressing wt or ΔF508 CFTR was stable at 22°C and increased at 28°C; a temperature permissive for ΔF508 CFTR expression in mammalian cells. At 37°C, however, CFTR-mediated conductance was further enhanced, whereas that due to ΔF508 CFTR channels decreased rapidly towards background, a phenomenon referred to here as “thermal inactivation.” Thermal inactivation of ΔF508 was mitigated by each of five suppressor mutations, I539T, R553M, G550E, R555K and R1070W; but each exerted unique effects on the severity of, and recovery from, thermal inactivation. Another mutation, K1250A, known to increase open probability (Po) of ΔF508 CFTR channels, exacerbated thermal inactivation. Application of potentiators known to increase Po of ΔF508 CFTR channels at room temperature failed to protect channels from inactivation at 37°C and one, PG-01, actually exacerbated thermal inactivation. Unstimulated ΔF508CFTR channels or those inhibited by CFTRinh-172, were partially protected from thermal inactivation, suggesting a possible inverse relationship between thermal stability and gating transitions. Thermal stability of channel function and temperature-sensitive maturation of the mutant protein appear to reflect related, but distinct facets of the ΔF508 CFTR conformational defect, both of which must be addressed by effective therapeutic modalities. PMID:22680785

  16. 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.

  17. 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.

  18. 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.

  19. Thermal Aggregation of Recombinant Protective Antigen: Aggregate Morphology and Growth Rate

    PubMed Central

    Belton, Daniel J.; Miller, Aline F.

    2013-01-01

    The thermal aggregation of the biopharmaceutical protein recombinant protective antigen (rPA) has been explored, and the associated kinetics and thermodynamic parameters have been extracted using optical and environmental scanning electron microscopies (ESEMs) and ultraviolet light scattering spectroscopy (UV-LSS). Visual observations and turbidity measurements provided an overall picture of the aggregation process, suggesting a two-step mechanism. Microscopy was used to examine the structure of aggregates, revealing an open morphology formed by the clustering of the microscopic aggregate particles. UV-LSS was used and developed to elucidate the growth rate of these particles, which formed in the first stage of the aggregation process. Their growth rate is observed to be high initially, before falling to converge on a final size that correlates with the ESEM data. The results suggest that the particle growth rate is limited by rPA monomer concentration, and by obtaining data over a range of incubation temperatures, an approach was developed to model the aggregation kinetics and extract the rate constants and the temperature dependence of aggregation. In doing so, we quantified the susceptibility of rPA aggregation under different temperature and environmental conditions and moreover demonstrated a novel use of UV spectrometry to monitor the particle aggregation quantitatively, in situ, in a nondestructive and time-resolved manner. PMID:23476645

  20. Silicate-Based Thermal Spray Coatings for Environmental Protection of Silicon Carbide

    NASA Astrophysics Data System (ADS)

    Sagiv, Ari Isaac

    Environmental barrier coatings are a key technology for implementing ceramics in high-temperature, high-moisture environments. One such ceramic, silicon carbide, is a material that can be used in gas turbines. However, silicon carbide oxidizes into silicon dioxide with exposure to oxygen, carbon dioxide, and water vapor and would normally provide protection for the silicon carbide. However, silicon dioxide volatilizes in a gas turbine environment, which leads to the degradation of its mechanical properties, making it unfit for use in a gas turbine. Materials like yttria-monosilicate and barium-strontium doped aluminosilicate (BSAS) both have good environmental coating properties. However, sintered yttria-monosilicate does not bond well to silicon carbide, and thermally sprayed BSAS transforms very slowly from a metastable hexacelsian phase to the desired celsian phase that is necessary for it to bond well to silicon carbide. Coatings of these materials have been produced by plasma spray with some additional work using HVOF. Phase identification has been done by x-ray diffraction, and microstructural analysis has been done using scanning electron microscopy.

  1. Development of Wireless Subsurface Microsensors for Health Monitoring of Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Pallix, Joan; Milos, Frank; Arnold, James O. (Technical Monitor)

    2000-01-01

    Low cost access to space is a primary goal for both NASA and the U.S. aerospace industry. 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, increase safety margins and improve mission reliability. A number of efforts are underway to use existing and emerging technologies to establish new methods for vehicle health monitoring on operational vehicles as well as X-vehicles. This paper summarizes a joint effort between several NASA centers and industry partners to develop rapid wireless diagnostic tools for failure management and long-term TPS performance monitoring of thermal protection systems (TPS) on future RLVS. An embedded wireless microsensor suite is being designed to allow rapid subsurface TPS health monitoring and damage assessment. This sensor suite will consist of both passive overlimit sensors and sensors for continuous parameter monitoring in flight. The on-board diagnostic system can be used to radio in maintenance requirements before landing and the data could also be used to assist in design validation for X-vehicles. For a 3rd generation vehicle, wireless diagnostics should be at a stage of technical development that will allow use for intelligent feedback systems for guidance and navigation control applications and can also serve as feedback for TPS that can intelligently adapt to its environment.

  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. 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.

  4. 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.

  5. Functionalized Biopolymer Particles Enhance Performance of a Tissue-Protective Peptide under Proteolytic and Thermal Stress.

    PubMed

    Dooley, Kevin; Devalliere, Julie; Uygun, Basak E; Yarmush, Martin L

    2016-06-13

    Cutaneous burns are often exacerbated by poor perfusion and subsequent necrosis of the microvasculature surrounding the primary injury. Preservation of these vessels can reduce necrotic tissue expansion and increase success rates of skin graft procedures. Recent work has identified a peptide derived from erythropoietin, ARA290, with the ability to mediate tissue protection in a variety of cell types. Here we demonstrate the advantages of fusing ARA290 to an elastin-like polypeptide (ELP) to salvage microvascular endothelial cells in harsh proteolytic conditions following thermal shock. These fusion proteins were expressed recombinantly in bacterial hosts and rapidly purified by inverse transition cycling. They were shown to spontaneously aggregate into particles at subphysiological temperatures. The bifunctional submicron particles were resistant to digestion in enzymes upregulated after burn injury. Furthermore, the data strongly suggest these ARA290-functionalized particles were superior to treatment with the peptide alone in preventing microvascular cell death in these conditions. The results bring to light an efficient and cost-effective strategy for the delivery therapeutic peptides to proteolytically active wound sites. PMID:27219509

  6. 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.

  7. 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.

  8. Structural tests on a tile/strain isolation pad thermal protection system. [space shuttles

    NASA Technical Reports Server (NTRS)

    Williams, J. G.

    1980-01-01

    The aluminum skin of the space shuttle is covered by a thermal protection system (TPS) consisting of a low density ceramic tile bonded to a matted-felt material called strain insulation pad (SIP). The structural characteristics of the TPS were studied experimentally under selected extreme load conditions. Three basic types of loads were imposed: tension, eccentrically applied tension, and combined in-plane force and transverse pressure. For some tests, transverse pressure was applied rapidly to simulate a transient shock wave passing over the tile. The failure mode for all specimens involved separation of the tile from the SIP at the silicone rubber bond interface. An eccentrically applied tension load caused the tile to separate from the SIP at loads lower than experienced at failure for pure tension loading. Moderate in-plane as well as shock loading did not cause a measurable reduction in the TPS ultimate failure strength. A strong coupling, however, was exhibited between in-plane and transverse loads and displacements.

  9. 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.

  10. 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.

  11. 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.

  12. In-Space Repair of Reinforced Carbon-Carbon (RCC) Thermal Protection System Structures

    NASA Technical Reports Server (NTRS)

    Singh, Mrityunjay

    2005-01-01

    Advanced repair and refurbishment technologies are critically needed for the RCC-based 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 adhesives 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, PLASTER (Patch Laminates and Sealant Technology for Exterior Repair) based systems 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. Aerothermodynamic environment and thermal protection for a Titan aerocapture vehicle. [Saturn satellite atmospheric entry

    NASA Technical Reports Server (NTRS)

    Green, M. J.; Moss, J. N.; Wilson, J. F.

    1984-01-01

    This paper presents thermal protection system (TPS) requirements for a potential Titan aerocapture vehicle. Shock-layer solutions are obtained for a nominal trajectory through the current Titan model atmosphere. Fully laminar and fully turbulent solutions are presented along the blunted fore-cone in the windward symmetry plane of a bent-biconic vehicle. Using these solutions to define the aerothermodynamic environment, transient material-response solutions are obtained for a Galileo-type TPS with a carbon-phenolic ablator heat shield. Shock-layer results indicate that turbulent flow is the more realistic flow condition. They also show that the lengthy aerocapture heating pulse is dominated by convective heating. The TPS results show that the required insulation thickness is uniformly about 4 cm along the fore-cone because of the long heat-soak period. The total heat-shield thickness is 6.4 cm at the stagnation point, and 4.7 cm near the end of the fore-cone. These TPS requirements are greater than those presented in a previous Titan aerocapture study.

  19. 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.

  20. 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