Science.gov

Sample records for aerothermodynamics thermal protection

  1. 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.

  2. Aerothermodynamic performance and thermal protection design for blunt re-entry bodies at L/D = 0.3

    NASA Technical Reports Server (NTRS)

    Caram, Jose M.; Kowal, T. J.

    1993-01-01

    Aerodynamic heating and thermal protection design analyses were performed for three blunt re-entry bodies at an L/D = 0.3 returning from low earth orbit. These configurations consisted of a scaled up Apollo command module, a Viking re-entry vehicle, and an Aeroassist Flight Experiment (AFE) aerobrake, each with a maximum diameter of 4.42 m. The aerothermodynamic analysis determined the equilibrium stagnation point heating rate and heat load for nominal and 3-sigma re-entry trajectories and the distribution of heating along the pitch and yaw planes for each of the vehicles at the time of highest heat flux. Using the predicted heating rates and heating distributions, a Thermal Protection System (TPS) design with flight certified materials was tailored for each of the configurations. Results indicated that the heating to the corner of the Viking aeroshell would exceed current limits of reusable tile material. Also, the maximum heating for the AFE would be 15 percent greater than the maximum heating for the Apollo flying the same trajectory. TPS designs showed no significant advantage in TPS weight between the different vehicles; however, heat-shield areal density comparisons showed the Apollo configuration to be the most efficient in terms of TPS weight.

  3. Integrated Design Engineering Analysis (IDEA) Environment - Aerodynamics, Aerothermodynamics, and Thermal Protection System Integration Module

    NASA Technical Reports Server (NTRS)

    Kamhawi, Hilmi N.

    2011-01-01

    This report documents the work performed during from March 2010 October 2011. The Integrated Design and Engineering Analysis (IDEA) environment is a collaborative environment based on an object-oriented, multidisciplinary, distributed environment using the Adaptive Modeling Language (AML) as the underlying framework. This report will focus on describing the work done in the area of extending the aerodynamics, and aerothermodynamics module using S/HABP, CBAERO, PREMIN and LANMIN. It will also detail the work done integrating EXITS as the TPS sizing tool.

  4. NASA's hypersonic fluid and thermal physics program (Aerothermodynamics)

    NASA Technical Reports Server (NTRS)

    Graves, R. A.; Hunt, J. L.

    1985-01-01

    This survey paper gives an overview of NASA's hypersonic fluid and thermal physics program (recently renamed aerothermodynamics). The purpose is to present the elements of, example results from, and rationale and projection for this program. The program is based on improving the fundamental understanding of aerodynamic and aerothermodynamic flow phenomena over hypersonic vehicles in the continuum, transitional, and rarefied flow regimes. Vehicle design capabilities, computational fluid dynamics, computational chemistry, turbulence modeling, aerothermal loads, orbiter flight data analysis, orbiter experiments, laser photodiagnostics, and facilities are discussed.

  5. Thermal response of integral multicomponent composites to a high-energy aerothermodynamic heating environment with surface temperature to 1800 K

    NASA Technical Reports Server (NTRS)

    Stewart, David A.; Leiser, Daniel B.

    1987-01-01

    Laminated composite insulations developed for potential use on advanced spacecraft operating between GEO and LEO were tested in an aerothermodynamic environment simulating an AOTV aerobraking maneuver (altitude 82.3 km and velocity = 9.0 km/s). Comparisons are discussed between these data and predictions of in-depth temperature response using dynamical thermal conductivity values to 2000 K.

  6. Thermal response of integral multicomponent composites to a high-energy aerothermodynamic heating environment with surface temperature to 1800 K

    SciTech Connect

    Stewart, D.A.; Leiser, D.B.

    1987-08-01

    Laminated composite insulations developed for potential use on advanced spacecraft operating between GEO and LEO were tested in an aerothermodynamic environment simulating an AOTV aerobraking maneuver (altitude 82.3 km and velocity = 9.0 km/s). Comparisons are discussed between these data and predictions of in-depth temperature response using dynamical thermal conductivity values to 2000 K. 8 references.

  7. HEART Aerothermodynamic Analysis

    NASA Technical Reports Server (NTRS)

    Mazaheri, Alireza

    2012-01-01

    This paper presents an assessment of the aerothermodynamic environment around an 8.3 meter High Energy Atmospheric Reentry Test (HEART) vehicle. This study generated twelve nose shape configurations and compared their responses at the peak heating trajectory point against the baseline nose shape. The heat flux sensitivity to the angle of attack variations are also discussed. The possibility of a two-piece Thermal Protection System (TPS) design at the nose is also considered, as are the surface catalytic affects of the aeroheating environment of such configuration. Based on these analyses, an optimum nose shape is proposed to minimize the surface heating. A recommendation is also made for a two-piece TPS design, for which the surface catalytic uncertainty associated with the jump in heating at the nose-IAD juncture is reduced by a minimum of 93%. In this paper, the aeroshell is assumed to be rigid and the inflatable fluid interaction effect is left for future investigations.

  8. Computational Aerothermodynamic Design Issues for Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.; Weilmuenster, K. James; Hamilton, H. Harris, II; Olynick, David R.; Venkatapathy, Ethiraj

    1997-01-01

    A brief review of the evolutionary progress in computational aerothermodynamics is presented. The current status of computational aerothermodynamics is then discussed, with emphasis on its capabilities and limitations for contributions to the design process of hypersonic vehicles. Some topics to be highlighted include: (1) aerodynamic coefficient predictions with emphasis on high temperature gas effects; (2) surface heating and temperature predictions for thermal protection system (TPS) design in a high temperature, thermochemical nonequilibrium environment; (3) methods for extracting and extending computational fluid dynamic (CFD) solutions for efficient utilization by all members of a multidisciplinary design team; (4) physical models; (5) validation process and error estimation; and (6) gridding and solution generation strategies. Recent experiences in the design of X-33 will be featured. Computational aerothermodynamic contributions to Mars Pathfinder, METEOR, and Stardust (Comet Sample return) will also provide context for this discussion. Some of the barriers that currently limit computational aerothermodynamics to a predominantly reactive mode in the design process will also be discussed, with the goal of providing focus for future research.

  9. Computational Aerothermodynamic Design Issues for Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Olynick, David R.; Venkatapathy, Ethiraj

    2004-01-01

    A brief review of the evolutionary progress in computational aerothermodynamics is presented. The current status of computational aerothermodynamics is then discussed, with emphasis on its capabilities and limitations for contributions to the design process of hypersonic vehicles. Some topics to be highlighted include: (1) aerodynamic coefficient predictions with emphasis on high temperature gas effects; (2) surface heating and temperature predictions for thermal protection system (TPS) design in a high temperature, thermochemical nonequilibrium environment; (3) methods for extracting and extending computational fluid dynamic (CFD) solutions for efficient utilization by all members of a multidisciplinary design team; (4) physical models; (5) validation process and error estimation; and (6) gridding and solution generation strategies. Recent experiences in the design of X-33 will be featured. Computational aerothermodynamic contributions to Mars Pathfinder, METEOR, and Stardust (Comet Sample return) will also provide context for this discussion. Some of the barriers that currently limit computational aerothermodynamics to a predominantly reactive mode in the design process will also be discussed, with the goal of providing focus for future research.

  10. Computational Aerothermodynamic Design Issues for Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.; Weilmuenster, K. James; Hamilton, H. Harris, II; Olynick, David R.; Venkatapathy, Ethiraj

    2005-01-01

    A brief review of the evolutionary progress in computational aerothermodynamics is presented. The current status of computational aerothermodynamics is then discussed, with emphasis on its capabilities and limitations for contributions to the design process of hypersonic vehicles. Some topics to be highlighted include: (1) aerodynamic coefficient predictions with emphasis on high temperature gas effects; (2) surface heating and temperature predictions for thermal protection system (TPS) design in a high temperature, thermochemical nonequilibrium environment; (3) methods for extracting and extending computational fluid dynamic (CFD) solutions for efficient utilization by all members of a multidisciplinary design team; (4) physical models; (5) validation process and error estimation; and (6) gridding and solution generation strategies. Recent experiences in the design of X-33 will be featured. Computational aerothermodynamic contributions to Mars Path finder, METEOR, and Stardust (Comet Sample return) will also provide context for this discussion. Some of the barriers that currently limit computational aerothermodynamics to a predominantly reactive mode in the design process will also be discussed, with the goal of providing focus for future research.

  11. X-38 Experimental Aerothermodynamics

    NASA Technical Reports Server (NTRS)

    Horvath, Thomas J.; Berry, Scott A.; Merski, N. Ronald; Fitzgerald, Steve M.

    2000-01-01

    The X-38 program seeks to demonstrate an autonomously returned orbital test flight vehicle to support the development of an operational Crew Return Vehicle for the International Space Station. The test flight, anticipated in 2002, is intended to demonstrate the entire mission profile of returning Space Station crew members safely back to earth in the event of medical or mechanical emergency. Integral to the formulation of the X-38 flight data book and the design of the thermal protection system, the aerothermodynamic environment is being defined through a synergistic combination of ground based testing and computational fluid dynamics. This report provides an overview of the hypersonic aerothermodynamic wind tunnel program conducted at the NASA Langley Research Center in support of the X-38 development. Global and discrete surface heat transfer force and moment, surface streamline patterns, and shock shapes were measured on scaled models of the proposed X-38 configuration in different test gases at Mach 6, 10 and 20. The test parametrics include angle of attack from 0 to 50 degs, unit Reynolds numbers from 0.3 x 10 (exp 6) to 16 x 10 (exp 6)/ ft, rudder deflections of 0, 2, and 5 deg. and body flap deflections from 0 to 30 deg. Results from hypersonic aerodynamic screening studies that were conducted as the configuration evolved to the present shape at, presented. Heavy gas simulation tests have indicated that the primary real gas effects on X-38 aerodynamics at trim conditions are expected to favorably influence flap effectiveness. Comparisons of the experimental heating and force and moment data to prediction and the current aerodynamic data book are highlighted. The effects of discrete roughness elements on boundary layer transition were investigated at Mach 6 and the development of a transition correlation for the X-38 vehicle is described. Extrapolation of ground based heating measurements to flight radiation equilibrium wall temperatures at Mach 6 and 10 were

  12. An Overview of the Space Shuttle Aerothermodynamic Design

    NASA Technical Reports Server (NTRS)

    Martin, Fred

    2011-01-01

    The Space Shuttle Thermal Protection System was one of the three areas that required the development of new technology. The talk discusses the pre-flight development of the aerothermodynamic environment which was based on Mach 8 wind tunnel data. A high level overview of the pre-flight heating rate predictions and comparison to the Orbiter Flight Test (OFT) data is presented, along with a discussion of the dramatic improvement in the state-of-the-art in aerothermodynamic capability that has been used to support the Shuttle Program. A high level review of the Orbiter aerothermodynamic design is discussed, along with improvements in Computational Fluid Dynamics and wind tunnel testing that was required for flight support during the last 30 years. The units have been removed from the plots, and the discussion is kept at a high level.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. Aerothermodynamic Data Base

    NASA Technical Reports Server (NTRS)

    1984-01-01

    Space shuttle aerothermodynamic data, collected from a continuing series of wind tunnel tests, are permanently stored with the Data Management Services (DMS) system. Information pertaining to current baseline configuration definition is also stored. A list of documentation of DMS processed data arranged sequentially and by space shuttle configuration is presented. The listing provides an up to date record of all applicable aerothermodynamic data collected, processed, or summarized during the space shuttle program. Tables are designed to provide survey information to the various space shuttle managerial and technical levels.

  19. Intermediate Experimental Vehicle, ESA Program Aerothermodynamics- Transition And Steps And Gaps Assessment

    NASA Astrophysics Data System (ADS)

    Verand, Jean-Luc; Pelissier, Christian; Sourgen, Frederic; Fontaine, Joelle; Garcon, Francois; Spel, Martin; van Hauwaert, Pierre; Charbonnier, Dominique; Vos, Jan; Vallee, Jean-Jacques; Pibarot, Julien; Tribot, Jean-Pierre; Mareschi, Vincenzo; Ferrarella, Daniella; Rufolo, Giuseppe

    2011-05-01

    The Intermediate eXperimental Vehicle (IXV) project objectives are the design, development, manufacture and ground and flight verification of an autonomous European lifting and aerodynamically controlled re-entry system, which is highly flexible and manoeuvrable. The IXV vehicle is a flying test bed for securing the next step of operational space vehicle development by supporting technology demonstration and system concept through the following objectives: a) Aerothermodynamics b) Advanced In Flight Experiments c) Thermal Protection System d) Guidance Navigation and Control e) System design The assessment of the general aerothermodynamic environment of IXV vehicle is mainly performed considering a smooth simplified geometry. However, the thermal protection system of IXV includes a mono-block ceramic matrix composite nose and an assembly of shingles between which steps and gaps are generated. From an aerothermodynamic point of view, such a distributed roughness layout cannot be ignored in terms of modification of the interaction between the flow and the body. To assess this effect, dedicated Mach number 5.5 wind tunnel tests (ONERA, S3MA facility) and numerical simulations (RTECH and CFS Engineering) have been performed during the phase C2 of the project. The paper presents the general logic of the work, with emphasis on the wind tunnel model design, tests involving infrared thermal measurements as well as the CFD rebuilding of the flow in the wind tunnel and the extrapolation from ground-to-flight.

  20. Aerothermodynamic design feasibility of a Mars aerocapture/aeromaneuver vehicle

    NASA Technical Reports Server (NTRS)

    Florence, D. E.

    1981-01-01

    Lifting aerodynamic configurations have been screened and selected for the Mars aerocapture mission that (1) meet the geometric packaging requirements of the various payloads and the Space Shuttle cargo bay and (2) provide the aerodynamic performance characteristics required to obtain the atmospheric exit steering accuracy and the parachute deployment conditions desired. Hypersonic heat transfer and aerodynamic loads to the vehicle in the CO2 atmosphere are evaluated. Contemporary low density ablative thermal protection materials were selected that meet all the atmospheric entry requirements and provide a minimum mass solution. Results are presented of the aerodynamic configuration and thermal protection materials screening and selection. It is concluded that the aerothermodynamic design of this concept is feasible using state-of-the-art technology.

  1. Orbiter thermal protection system

    NASA Technical Reports Server (NTRS)

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

    1985-01-01

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

  2. Aerothermodynamic Heating Analysis of Aerobraking and Aeromaneuvering Orbital Transfer Vehicles

    NASA Technical Reports Server (NTRS)

    Menees, Gene P.; Davies, Carol B.; Wilson, John F.; Brown, Kevin G.

    1985-01-01

    The thermal-protection requirements of two aeroassisted orbital transfer vehicles (AOTVS) are analyzed for return missions between the geosynchronous and Shuttle orbits. One of the designs is a specialized version of a previously proposed generic aerobraking vehicle that Is capable of only delivery-type operations. The other Is a high-lift aeromaneuvering vehicle that is optimized for low Earth orbit sortie missions involving large, multiple plane-inclination changes. The aerothermal environment of the aerobraking vehicle is analyzed using state-of-the-art methods for nonequilibrium-radiative and convective heating that incorporate refinements unique to the configuration. The heating analysis of the aeromaneuvering vehicle required the development of a flowfield model for rarefied-hypersonic flow over a lifting surface at incidence. The predicted aerothermodynamic heating characteristics for both vehicles are correlated with thermal-control

  3. Aerothermodynamic Design of the Mars Science Laboratory Heatshield

    NASA Technical Reports Server (NTRS)

    Edquist, Karl T.; Dyakonov, Artem A.; Wright, Michael J.; Tang, Chun Y.

    2009-01-01

    Aerothermodynamic design environments are presented for the Mars Science Laboratory entry capsule heatshield. The design conditions are based on Navier-Stokes flowfield simulations on shallow (maximum total heat load) and steep (maximum heat flux, shear stress, and pressure) entry trajectories from a 2009 launch. Boundary layer transition is expected prior to peak heat flux, a first for Mars entry, and the heatshield environments were defined for a fully-turbulent heat pulse. The effects of distributed surface roughness on turbulent heat flux and shear stress peaks are included using empirical correlations. Additional biases and uncertainties are based on computational model comparisons with experimental data and sensitivity studies. The peak design conditions are 197 W/sq cm for heat flux, 471 Pa for shear stress, 0.371 Earth atm for pressure, and 5477 J/sq cm for total heat load. Time-varying conditions at fixed heatshield locations were generated for thermal protection system analysis and flight instrumentation development. Finally, the aerothermodynamic effects of delaying launch until 2011 are previewed.

  4. 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.

  5. 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.

  6. 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.

  7. High-Energy Atmospheric Reentry Test Aerothermodynamic Analysis

    NASA Technical Reports Server (NTRS)

    Mazaheri, Alireza

    2013-01-01

    This paper presents an assessment of the aerothermodynamic environment around an 8.3 meter High Energy Atmospheric Reentry Test (HEART) vehicle. This study generated twelve nose shape configurations and compared their responses at the peak heating trajectory point against the baseline nose shape. The heat flux sensitivity to the angle of attack variations are also discussed. The possibility of a two-piece Thermal Protection System (TPS) design at the nose is also considered, as are the surface catalytic affects of the aeroheating environment of such configuration. Based on these analyses, an optimum nose shape is proposed to minimize the surface heating. A recommendation is also made for a two-piece TPS design, for which the surface catalytic uncertainty associated with the jump in heating at the nose-IAD juncture is reduced by a minimum of 93%. In this paper, the aeroshell is assumed to be rigid and the inflatable fluid interaction effect is left for future investigations

  8. Experimental Aerothermodynamics In Support Of The Columbia Accident Investigation

    NASA Technical Reports Server (NTRS)

    Horvath, Thomas J.

    2004-01-01

    The technical foundation for the most probable damage scenario reported in the Columbia Accident Investigation Board's final report was largely derived from synergistic aerodynamic/aerothermodynamic wind tunnel measurements and inviscid predictions made at NASA Langley Research Center and later corroborated with engineering analysis, high fidelity numerical viscous simulations, and foam impact testing near the close of the investigation. This report provides an overview of the hypersonic aerothermodynamic wind tunnel program conducted at NASA Langley and illustrates how the ground-based heating measurements provided early insight that guided the direction and utilization of agency resources in support of the investigation. Global surface heat transfer mappings, surface streamline patterns, and shock shapes were measured on 0.0075 scale models of the Orbiter configuration with and without postulated damage to the thermal protection system. Test parametrics include angle of attack from 38 to 42 degs, sideslip angles of 38 to 42 degs, sideslip angles of plus or minus 1 deg, Reynolds numbers based upon model length from 0.05 x 10(exp 6) to 6.5 x 10(exp 6), and normal shock density ratios of 5 (Mach 6 Air) and 12 (Mach 6 CF4). The primary objective of the testing was to provide surface heating characteristics on scaled Orbiter models with outer mold line perturbations to simulate various forms of localized surface damage to the thermal protection system. Initial experimental testing conducted within two weeks of the accident simulated a broad spectrum of thermal protection system damage to the Orbiter windward surface and was used to refute several hypothesized forms of thermal protection system damage, which included gouges in the windward thermal protection system tiles, breaches through the wing new the main landing gear door, and protuberances along the wing leading edge that produced asymmetric boundary layer transition. As the forensic phase of the investigation

  9. Ablative Thermal Protection: An Overview

    NASA Technical Reports Server (NTRS)

    Laub, Bernie

    2003-01-01

    Contents include the following: Why ablative thermal protections - TPS. Ablative TPS chronology: strategic reentry systems, solid rocket motor nozzles, space (manned missions and planetary entry probes). Ablation mechanisms. Ablation material testing. Ablative material testing.

  10. 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.

  11. Legacy of the Space Shuttle from an Aerodynamic and Aerothermodynamic Perspective

    NASA Technical Reports Server (NTRS)

    Martin, Fred W.

    2011-01-01

    The development of the Space Shuttle Orbiter thermal protection system heating environment is described from a design stand point that began in the early 1970s. The desire for a light weight, reusable heat shield required the development of new technology, relative to previous manned spacecraft, and a systems approach to the design of the vehicle, entry guidance, and thermal protection system. Several unanticipated issues had to be resolved in both the entry and ascent phases of flight, which are discussed at a high level. During the life of the Program, significant improvements in computing power and numerical methods have been applied to Space Shuttle aerodynamic and aerothermodynamic issues, with the Shuttle Program often being the motivation, and or sponsor of the analysis development.

  12. Aerothermodynamic and stability analyses of a deployable re-entry capsule

    NASA Astrophysics Data System (ADS)

    Carandente, Valerio; Zuppardi, Gennaro; Savino, Raffaele

    2014-01-01

    Recent research projects, in the field of atmospheric re-entry technology, are focused on the design of deployable, umbrella-like Thermal Protection Systems (TPSs). These TPSs are made of flexible high temperature resistant fabrics, folded at launch and deployed in space for de-orbit and re-entry operations. In the present paper two possible sphere-cone configurations for the TPS have been investigated from an aerodynamic point of view. The analyzed configurations are characterized by the same reentry mass and maximum diameter, but have different half-cone angles (45° and 60°). The analyses involve both the evaluation of thermal and aerodynamic loads and the assessment of the capsule longitudinal stability. The aerothermodynamic analysis has been performed for the completely deployed heat shield in transitional and continuum regimes, while the longitudinal stability has been analyzed in free molecular, transitional and continuum regimes, also taking into consideration the heat shield deployment sequence at high altitudes.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. Aerothermodynamic Flight Simulation Capabilities for Aerospace Vehicles

    NASA Technical Reports Server (NTRS)

    Miller, Charles G.

    1998-01-01

    Aerothermodynamics, encompassing aerodynamics, aeroheating, and fluid dynamics and physical processes, is the genesis for the design and development of advanced space transportation vehicles and provides crucial information to other disciplines such as structures, materials, propulsion, avionics, and guidance, navigation and control. Sources of aerothermodynamic information are ground-based facilities, Computational Fluid Dynamic (CFD) and engineering computer codes, and flight experiments. Utilization of this aerothermodynamic triad provides the optimum aerothermodynamic design to safely satisfy mission requirements while reducing design conservatism, risk and cost. The iterative aerothermodynamic process for initial screening/assessment of aerospace vehicle concepts, optimization of aerolines to achieve/exceed mission requirements, and benchmark studies for final design and establishment of the flight data book are reviewed. Aerothermodynamic methodology centered on synergism between ground-based testing and CFD predictions is discussed for various flow regimes encountered by a vehicle entering the Earth s atmosphere from low Earth orbit. An overview of the resources/infrastructure required to provide accurate/creditable aerothermodynamic information in a timely manner is presented. Impacts on Langley s aerothermodynamic capabilities due to recent programmatic changes such as Center reorganization, downsizing, outsourcing, industry (as opposed to NASA) led programs, and so forth are discussed. Sample applications of these capabilities to high Agency priority, fast-paced programs such as Reusable Launch Vehicle (RLV)/X-33 Phases I and 11, X-34, Hyper-X and X-38 are presented and lessons learned discussed. Lastly, enhancements in ground-based testing/CFD capabilities necessary to partially/fully satisfy future requirements are addressed.

  19. Opportunities for research in aerothermodynamics

    NASA Technical Reports Server (NTRS)

    Graham, R. W.

    1983-01-01

    "Aerothermodynamics' involves the disciplines of chemistry, thermodynamics, fluid mechanics and heat transfer which have collaborative importance in propulsion systems. There are growing opportunities for the further application of these disciplines to improve the methodology for the design of advanced gas turbines; particularly, the combustor and turbine. Design procedures follow empirical or cut and try guidelines. The tremendous advances in computational analysis and in instrumentation techniques hold promise for research answers to complex physical processes that are currently not well understood. The transfer of basic research understanding to engineering design should result in shorter, less expensive development commitments for engines. The status and anticipated opportunities in research topics relevant to combustors and turbines is reviewed.

  20. Langley Aerothermodynamic Facilities Complex: Enhancements and Testing Capabilities

    NASA Technical Reports Server (NTRS)

    Micol, J. R.

    1998-01-01

    Description, capabilities, recent upgrades, and utilization of the NASA Langley Research Center (LaRC) Aerothermodynamic Facilities Complex (AFC) are presented. The AFC consists of five hypersonic, blow-down-to-vacuum wind tunnels that collectively provide a range of Mach number from 6 to 20, unit Reynolds number from 0.04 to 22 million per foot and, most importantly for blunt configurations, normal shock density ratio from 4 to 12. These wide ranges of hypersonic simulation parameters are due, in part, to the use of three different test gases (air, helium, and tetrafluoromethane), thereby making several of the facilities unique. The Complex represents nearly three-fourths of the conventional (as opposed to impulse)-type hypersonic wind tunnels operational in this country. AFC facilities are used to assess and optimize the hypersonic aerodynamic performance and aeroheating characteristics of aerospace vehicle concepts and to provide benchmark aerodynamic/aeroheating data fr generating the flight aerodynamic databook and final design of the thermal protection system (TPS) (e.g., establishment of flight limitations not to exceed TPS design limits). Modifications and enhancements of AFC hardware components and instrumentation have been pursued to increase capability, reliability, and productivity in support of programmatic goals. Examples illustrating facility utilization in recent years to generate essentially all of the experimental hypersonic aerodynamic and aeroheating information for high-priority, fast-paced Agency programs are presented. These programs include Phase I of the Reusable Launch Vehicle (RLV) Advanced Technology Demonstrator, X-33 program, PHase II of the X-33 program, X-34 program, the Hyper-X program ( a Mach 5,7, and 10 airbreathing propulsion flight experiment), and the X-38 program (Experimental Crew Return Vehicle, X-CRV). Current upgrades/enchancements and future plans for the AFC are discussed.

  1. Thermal cleavage of the fmoc protection group.

    PubMed

    Höck, Stefan; Marti, Roger; Riedl, Rainer; Simeunovice, Marina

    2010-01-01

    The Fmoc protection group is among the most commonly used protection groups for the amino function. A fast method for the thermal deavage of this protection group under base-free conditions without the need for dibenzofulvene scavengers is presented. The advantages of this method include straightforward testability by means of a simple high-temperature NMR experiment, usually high yields, and good selectivity towards the BOC protection group and t-butyl ethers.

  2. 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.

  3. 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.

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

    NASA Technical Reports Server (NTRS)

    Blosser, M. L.

    1982-01-01

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

  5. 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.

  6. 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.

  7. A Perspective on Computational Aerothermodynamics at NASA

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.

    2007-01-01

    The evolving role of computational aerothermodynamics (CA) within NASA over the past 20 years is reviewed. The presentation highlights contributions to understanding the Space Shuttle pitching moment anomaly observed in the first shuttle flight, prediction of a static instability for Mars Pathfinder, and the use of CA for damage assessment in post-Columbia mission support. In the view forward, several current challenges in computational fluid dynamics and aerothermodynamics for hypersonic vehicle applications are discussed. Example simulations are presented to illustrate capabilities and limitations. Opportunities to advance the state-of-art in algorithms, grid generation and adaptation, and code validation are identified.

  8. Space Shuttle aerothermodynamic data report, phase C

    NASA Technical Reports Server (NTRS)

    1985-01-01

    Space shuttle aerothermodynamic data, collected from a continuing series of wind tunnel tests, are permanently stored with the Data Management Services (DMS) system. Information pertaining to current baseline configuration definition is also stored. Documentation of DMS processed data arranged sequentially and by space shuttle configuration are included. An up-to-date record of all applicable aerothermodynamic data collected, processed, or summarized during the space shuttle program is provided. Tables are designed to provide suvery information to the various space shuttle managerial and technical levels.

  9. Aerothermodynamic Analyses of Towed Ballutes

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.; Buck, Greg; Moss, James N.; Nielsen, Eric; Berger, Karen; Jones, William T.; Rudavsky, Rena

    2006-01-01

    A ballute (balloon-parachute) is an inflatable, aerodynamic drag device for application to planetary entry vehicles. Two challenging aspects of aerothermal simulation of towed ballutes are considered. The first challenge, simulation of a complete system including inflatable tethers and a trailing toroidal ballute, is addressed using the unstructured-grid, Navier-Stokes solver FUN3D. Auxiliary simulations of a semi-infinite cylinder using the rarefied flow, Direct Simulation Monte Carlo solver, DSV2, provide additional insight into limiting behavior of the aerothermal environment around tethers directly exposed to the free stream. Simulations reveal pressures higher than stagnation and corresponding large heating rates on the tether as it emerges from the spacecraft base flow and passes through the spacecraft bow shock. The footprint of the tether shock on the toroidal ballute is also subject to heating amplification. Design options to accommodate or reduce these environments are discussed. The second challenge addresses time-accurate simulation to detect the onset of unsteady flow interactions as a function of geometry and Reynolds number. Video of unsteady interactions measured in the Langley Aerothermodynamic Laboratory 20-Inch Mach 6 Air Tunnel and CFD simulations using the structured grid, Navier-Stokes solver LAURA are compared for flow over a rigid spacecraft-sting-toroid system. The experimental data provides qualitative information on the amplitude and onset of unsteady motion which is captured in the numerical simulations. The presence of severe unsteady fluid - structure interactions is undesirable and numerical simulation must be able to predict the onset of such motion.

  10. DSMC aero-thermo-dynamic analysis of a sample-return capsule

    NASA Astrophysics Data System (ADS)

    Zuppardi, Gennaro; Savino, Raffaele; Boffa, Chiara; Carandente, Valerio

    2012-11-01

    A rarefied aero-thermo-dynamic analysis of a sample Earth Return Capsule during the high energy, high altitude re-entry path from an exploration mission is presented. The altitude interval 70-120 km is considered, where the capsule experiences different flow fields. In fact, the flow regime ranges from continuum low density to near free molecular flow and, even though the free stream velocity is almost constant (13 km/s) in the whole altitude interval, the Mach number changes from 44 to 32 and the Reynolds number, based on the capsule diameter, ranges from 4.92×104 to 9. The computations have been carried out using two direct simulation Monte Carlo codes: DS2V to compute local quantities such as heat flux, thermal and aerodynamic loads at zero angle of attack and DS3V to compute global aerodynamic coefficients in the range of the angle of attack 0-60 deg; The results verified that in this altitude interval the heat flux and the thermal load reasonably satisfy specific requirements for the thermal protection system and that the capsule is longitudinally stable up to an angle of attack of about 40 deg..

  11. Thermal Protection during Percutaneous Thermal Ablation of Renal Cell Carcinoma

    PubMed Central

    Kam, Anthony W.; Littrup, Peter J.; Walther, McClellan M.; Hvizda, Julia; Wood, Bradford J.

    2008-01-01

    Thermal injury to collateral structures is a known complication of thermal ablation of tumors. The authors present the use of CO2 dissection and inserted balloons to protect the bowel during percutaneous radiofrequency (RF) ablation and cryotherapy of primary and locally recurrent renal cell carcinoma. These techniques offer the potential to increase the number of tumors that can be treated with RF ablation or cryotherapy from a percutaneous approach. PMID:15231890

  12. 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.

  13. Aerothermodynamics at NASA-Langley Research Center

    NASA Technical Reports Server (NTRS)

    Weilmuenster, K. James

    2001-01-01

    The Aerothermodynamics Branch at NASA - Langley Research Center is tasked with developing, assessing and applying aerothermodynamic technologies to enable the development of hypersonic aircraft, launch vehicles, and planetary/earth entry systems. To accomplish this mission, the Branch capitalizes on the synergism between the experimental and computational facilities/tools which reside in the branch and a staff that can draw on five decades of experience in aerothermodynamics. The Aerothermodynamics Branch is staffed by 30 scientists/engineers. The staff, of which two-thirds are less than 40 years old, is split evenly between experimentalists and computationalists. Approximately 90 percent of the staff work on space transportation systems while the remainder work on planetary missions. The Branch manages 5 hypersonic wind tunnels which are staffed by 14 technicians, numerous high end work stations and a SGI Origin 2000 system. The Branch also utilizes other test facilities located at Langley as well as other national and international test sites. Large scale computational requirements are met by access to Agency resources.

  14. 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.

  15. 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.

  16. 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

  17. 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.

  18. Aerothermodynamic heating analysis of aerobraking and aeromaneuvering orbital-transfer vehicles

    NASA Technical Reports Server (NTRS)

    Davies, C. B.; Wilson, J. F.; Brown, K. G.; Menees, G. P.

    1984-01-01

    The thermal-protection requirements of two aeroassisted orbital-transfer vehicles (AOTVs) are analyzed for return missions between the geosynchronous and Shuttle orbits. One of the designs is a specialized version of a previously proposed generic aerobraking vehicle that is capable of only delivery-type operations. The other is a high-lift aeromaneuvering vehicle that is optimized for low-earth orbit sortie missions involving large, multiple plane-inclination changes. The aerothermal environment of the aerobraking vehicle is analyzed using state-of-the-art methods for nonequilibrium-radiative and convective heating that incorporate refinements unique to the configuration. The heating analysis of the aeromaneuvering vehicle required the development of a flow-field model for rarefied-hypersonic flow over a lifting surface at incidence. The predicted aerothermodynamic heating characteristics for both vehicles are correlated with thermal-control requirements and flight performance capabilities for the specified mission guidelines. The results help identify technical issues related to the development of future operational systems.

  19. 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.

  20. 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.

  1. Overview of aerothermodynamic loads definition study

    NASA Technical Reports Server (NTRS)

    Gaugler, Raymond E.

    1989-01-01

    Over the years, NASA has been conducting the Advanced Earth-to-Orbit (AETO) Propulsion Technology Program to provide the knowledge, understanding, and design methodology that will allow the development of advanced Earth-to-orbit propulsion systems with high performance, extended service life, automated operations, and diagnostics for in-flight health monitoring. The objective of the Aerothermodynamic Loads Definition Study is to develop methods to more accurately predict the operating environment in AETO propulsion systems, such as the Space Shuttle Main Engine (SSME) powerhead. The approach taken consists of 2 parts: to modify, apply, and disseminate existing computational fluid dynamics tools in response to current needs and to develop new technology that will enable more accurate computation of the time averaged and unsteady aerothermodynamic loads in the SSME powerhead. The software tools are detailed. Significant progress was made in the area of turbomachinery, where there is an overlap between the AETO efforts and research in the aeronautical gas turbine field.

  2. Aerothermodynamic methods for a Mars environmental survey Mars entry

    NASA Technical Reports Server (NTRS)

    Mitcheltree, R. A.

    1994-01-01

    Computational fluid dynamics models for the thermodynamics and transport properties used in an equilibrium version of the Langley aerothermodynamics upwind relaxation algorithm (LAURA) for Mars atmospheric entries are described. In addition, the physical models used in a nonequilibrium version of LAURA for Mars-entry flows are described. Uncertainties exist in defining constants used in the transport properties for the equilibrium model and in many of the physical models for the nonequilibrium version. Solutions from the two codes using the best available constants are examined at the Mars-entry conditions characteristics of the Mars environmental survey mission. While the flowfields are near thermal equilibrium, chemical nonequilibrium effects are present in the entry cases examined. Convective heating at the stagnation point for these flows (assuming fully catalytic wall boundary conditions) is approximately 100 W/cm(exp 2). Radiative heating is negligible.

  3. Aerothermodynamic flow phenomena of the airframe-integrated supersonic combustion ramjet

    NASA Technical Reports Server (NTRS)

    Walton, James T.

    1992-01-01

    The unique component flow phenomena is discussed of the airframe-integrated supersonic combustion ramjet (scramjet) in a format geared towards new players in the arena of hypersonic propulsion. After giving an overview of the scramjet aerothermodynamic cycle, the characteristics are then covered individually of the vehicle forebody, inlet, combustor, and vehicle afterbody/nozzle. Attention is given to phenomena such as inlet speeding, inlet starting, inlet spillage, fuel injection, thermal choking, and combustor-inlet interaction.

  4. 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.

  5. Aerothermodynamic Insight From The HIFIRE Program

    NASA Astrophysics Data System (ADS)

    Kimmel, Roger L.; Adamczak, David; Dolvin, Douglas; Borg, Matthew; Stanfield, Scott

    2011-05-01

    The HIFiRE (Hypersonic International Flight Research and Experimentation) program is a joint venture of the United States Air Force Research Laboratory and Australian Defence Science and Technology Organisation to utilize economical flight research opportunities in the exploration of flight science issues for space access systems. Flights 1 and 5 focus on collecting high-resolution experimental data on critical aerothermodynamic phenomena, including laminar-turbulent transition and shock/boundary layer interactions. Flight 1, successfully flown in March 2010, employed a test article composed of a 7-deg right angle cone, followed by a cylinder and flare. The test article remained attached to the second-stage booster throughout the ballistic trajectory. Flight 5, to be launched in a similar fashion, will feature a 2:1 elliptic cross-section cone as the test article. For both flights significant resources have been invested in pre-flight aerothermodynamic analysis and testing. This manuscript will summarize the overall strategy of the HIFiRE program, review the pre-flight aerothermodynamic analysis for Flights 1 and 5, and present a brief look at preliminary results from the post-flight analysis of Flight 1.

  6. 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.

  7. 49 CFR 179.18 - Thermal protection systems.

    Code of Federal Regulations, 2012 CFR

    2012-10-01

    ... jackets, insulation, and thermal protection. A complete record of each analysis shall be made, retained... 49 Transportation 3 2012-10-01 2012-10-01 false Thermal protection systems. 179.18 Section 179.18... § 179.18 Thermal protection systems. (a) Performance standard. When the regulations in this...

  8. 49 CFR 179.18 - Thermal protection systems.

    Code of Federal Regulations, 2014 CFR

    2014-10-01

    ... jackets, insulation, and thermal protection. A complete record of each analysis shall be made, retained... 49 Transportation 3 2014-10-01 2014-10-01 false Thermal protection systems. 179.18 Section 179.18... § 179.18 Thermal protection systems. (a) Performance standard. When the regulations in this...

  9. 49 CFR 179.18 - Thermal protection systems.

    Code of Federal Regulations, 2013 CFR

    2013-10-01

    ... jackets, insulation, and thermal protection. A complete record of each analysis shall be made, retained... 49 Transportation 3 2013-10-01 2013-10-01 false Thermal protection systems. 179.18 Section 179.18... § 179.18 Thermal protection systems. (a) Performance standard. When the regulations in this...

  10. 49 CFR 179.18 - Thermal protection systems.

    Code of Federal Regulations, 2011 CFR

    2011-10-01

    ... jackets, insulation, and thermal protection. A complete record of each analysis shall be made, retained... 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...

  11. Aerothermodynamic Design of the Mars Science Laboratory Backshell and Parachute Cone

    NASA Technical Reports Server (NTRS)

    Edquist, Karl T.; Dyakonov, Artem A.; Wright, Michael J.; Tang, Chun Y.

    2009-01-01

    Aerothermodynamic design environments are presented for the Mars Science Laboratory entry capsule backshell and parachute cone. The design conditions are based on Navier-Stokes flowfield simulations on shallow (maximum total heat load) and steep (maximum heat flux) design entry trajectories from a 2009 launch. Transient interference effects from reaction control system thruster plumes were included in the design environments when necessary. The limiting backshell design heating conditions of 6.3 W/sq cm for heat flux and 377 J/sq cm for total heat load are not influenced by thruster firings. Similarly, the thrusters do not affect the parachute cover lid design environments (13 W/sq cm and 499 J/sq cm). If thruster jet firings occur near peak dynamic pressure, they will augment the design environments at the interface between the backshell and parachute cone (7 W/sq cm and 174 J/sq cm). Localized heat fluxes are higher near the thruster fairing during jet firings, but these areas did not require additional thermal protection material. Finally, heating bump factors were developed for antenna radomes on the parachute cone

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

    NASA Technical Reports Server (NTRS)

    Culp, J. D.

    1973-01-01

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

  13. Thermal Vacuum Facility for Testing Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

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

    2002-01-01

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

  14. Reusable Metallic Thermal Protection Systems Development

    NASA Technical Reports Server (NTRS)

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

    1998-01-01

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

  15. 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.

  16. Thermal Protection Systems: Past, Present and Future

    NASA Technical Reports Server (NTRS)

    Johnson, Sylvia M.

    2015-01-01

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

  17. Thermal Protection System of the Space Shuttle

    NASA Technical Reports Server (NTRS)

    Cleland, John; Iannetti, Francesco

    1989-01-01

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

  18. 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.

  19. 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.

  20. 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.

  1. Development of AFE aerobrake aerothermodynamic data book

    NASA Technical Reports Server (NTRS)

    Ting, Paul C.; Rochelle, W. C.; Mueller, S. R.; Colovin, J. E.; Scott, C. D.; Curry, D. M.

    1989-01-01

    The computation method developed for the NASA Aeroassist Flight Experiment (AFE) data book generates a design reference for the AFE's aerothermodynamic environment using an optimized technology for a 4100-lb vehicle. This environment is defined by convective, radiative, and total heating rates, radiation equilibrium temperatures, and local surface pressures along the AFE pitch-plane and associated off-pitch planes. The Boundary Layer Integral Matrix Procedure is the major program code used in this analysis; a partially catalytic wall was assumed on the basis of measured recombination rates.

  2. 49 CFR 193.2057 - Thermal radiation protection.

    Code of Federal Regulations, 2010 CFR

    2010-10-01

    ... 49 Transportation 3 2010-10-01 2010-10-01 false Thermal radiation protection. 193.2057 Section 193... GAS FACILITIES: FEDERAL SAFETY STANDARDS Siting Requirements § 193.2057 Thermal radiation protection...) The thermal radiation distances must be calculated using Gas Technology Institute's (GTI) report...

  3. 49 CFR 193.2057 - Thermal radiation protection.

    Code of Federal Regulations, 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...

  4. 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...

  5. 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...

  6. 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...

  7. Computational Aerothermodynamics in Aeroassist Applications

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.

    2001-01-01

    Aeroassisted planetary entry uses atmospheric drag to decelerate spacecraft from super-orbital to orbital or suborbital velocities. Numerical simulation of flow fields surrounding these spacecraft during hypersonic atmospheric entry is required to define aerothermal loads. The severe compression in the shock layer in front of the vehicle and subsequent, rapid expansion into the wake are characterized by high temperature, thermo-chemical nonequilibrium processes. Implicit algorithms required for efficient, stable computation of the governing equations involving disparate time scales of convection, diffusion, chemical reactions, and thermal relaxation are discussed. Robust point-implicit strategies are utilized in the initialization phase; less robust but more efficient line-implicit strategies are applied in the endgame. Applications to ballutes (balloon-like decelerators) in the atmospheres of Venus, Mars, Titan, Saturn, and Neptune and a Mars Sample Return Orbiter (MSRO) are featured. Examples are discussed where time-accurate simulation is required to achieve a steady-state solution.

  8. 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.

  9. Thermal protection using very high temperature ceramics

    NASA Technical Reports Server (NTRS)

    Adamczyk, George R.

    1992-01-01

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

  10. 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.

  11. 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.

  12. Lightweight Nonmetallic Thermal Protection Materials Technology

    NASA Astrophysics Data System (ADS)

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

    2005-02-01

    To fulfill President George W. Bush's "Vision for Space Exploration" (NASA, 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) products 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.

  13. 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.

  14. Aerothermodynamic challenges of the Saenger space-transportation system

    NASA Astrophysics Data System (ADS)

    Hirschel, E. H.

    1991-09-01

    The two-stage-to-orbit Saenger space transportation system is the reference concept of the German hypersonics technology program. The technology development concentrates first on the needs of the lower stage. Its requirements on aerothermodynamics and propulsion integration are sketched. The aerothermodynamic design challenge is discussed and the design tools and the design methodology are reviewed. The calibration of both the computational and the experimental methods, as well as the test of vehicle components like the inlet, control surfaces etc., make the Hypersonic Technology Experimental vehicle (HYTEX) mandatory. Contents and workplan of the technology program 'aerothermodynamics and propulsion integration' are laid out. Selected results from the current work are presented.

  15. 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.

  16. Aerothermodynamic testing requirements for future space transportation systems

    NASA Astrophysics Data System (ADS)

    Paulson, John W., Jr.; Miller, Charles G., III

    1995-03-01

    Aerothermodynamics, encompassing aerodynamics, aeroheating, and fluid dynamic and physical processes, is the genesis for the design and development of advanced space transportation vehicles. It provides crucial information to other disciplines involved in the development process such as structures, materials, propulsion, and avionics. Sources of aerothermodynamic information include ground-based facilities, computational fluid dynamic (CFD) and engineering computer codes, and flight experiments. Utilization of this triad is required to provide the optimum requirements while reducing undue design conservatism, risk, and cost. This paper discusses the role of ground-based facilities in the design of future space transportation system concepts. Testing methodology is addressed, including the iterative approach often required for the assessment and optimization of configurations from an aerothermodynamic perspective. The influence of vehicle shape and the transition from parametric studies for optimization to benchmark studies for final design and establishment of the flight data book is discussed. Future aerothermodynamic testing requirements including the need for new facilities are also presented.

  17. Aerothermodynamic data base. Data file contents report, phase C

    NASA Technical Reports Server (NTRS)

    Lutz, G. R.

    1983-01-01

    Space shuttle aerothermodynamic data, collected from a continuing series of wind tunnel tests, are permanently stored with the Data Management Services (DMS) system. Information pertaining to current baseline configuration definition is also stored. Documentation of DMS processed data arranged sequentially and by space shuttle configuration is listed to provide an up-to-date record of all applicable aerothermodynamic data collected, processed, or summarized during the space shuttle program. Tables provide survey information to the various space shuttle managerial and technical levels.

  18. Shuttle Tethered Aerothermodynamics Research Facilty (STARFAC) instrumentation requirements

    NASA Technical Reports Server (NTRS)

    Wood, G. M.; Siemers, P. M.; Carlomagno, G. M.; Hoffman, J.

    1986-01-01

    The instrumentation requirements for the Shuttle Tethered Aerothermodynamic Research Facility (STARFAC) are presented. The typical physical properties of the terrestrial atmosphere are given along with representative atmospheric daytime ion concentrations and the equilibrium and nonequilibrium gas property comparison from a point away from a wall. STARFAC science and engineering measurements are given as are the TSS free stream gas analysis. The potential nonintrusive measurement techniques for hypersonic boundary layer research are outlined along with the quantitative physical measurement methods for aerothermodynamic studies.

  19. Overview of aerothermodynamic loads definition study

    NASA Technical Reports Server (NTRS)

    Gaugler, Raymond E.

    1991-01-01

    The objective of the Aerothermodynamic Loads Definition Study is to develop methods of accurately predicting the operating environment in advanced Earth-to-Orbit (ETO) propulsion systems, such as the Space Shuttle Main Engine (SSME) powerhead. Development of time averaged and time dependent three dimensional viscous computer codes as well as experimental verification and engine diagnostic testing are considered to be essential in achieving that objective. Time-averaged, nonsteady, and transient operating loads must all be well defined in order to accurately predict powerhead life. Described here is work in unsteady heat flow analysis, improved modeling of preburner flow, turbulence modeling for turbomachinery, computation of three dimensional flow with heat transfer, and unsteady viscous multi-blade row turbine analysis.

  20. Aerothermodynamics of the Mars Global Surveyor Spacecraft

    NASA Technical Reports Server (NTRS)

    Shane, Russell W.; Tolson, Robert H.

    1998-01-01

    The aerothermodynamics characteristics of the Mars Global Surveyor spacecraft are investigated and reported. These results have been used by the Mars Global Surveyor mission planners to design the aerobraking phase of the mission. Analytical and Direct Simulation Monte Carlo computer codes were used with a detailed, three dimensional model of the spacecraft to evaluate spacecraft aerobraking characteristics for flight in free molecular and transitional flow regimes. The spacecraft is found to be aerodynamically stable in aerobraking and planned contingency configurations. Aerodynamic forces, moments, and heating are found to be highly dependent on atmospheric density. Accommodation coefficient. is seen to strongly influence drag coefficient. Transitional flow effects are found to reduce overall solar panel heating. Attitude control thruster plumes are shown to interact with the freestream, diminishing the effectiveness of the attitude control system and even leading to thrust reversal. These plume-freestream interaction effects are found to be highly dependent on freestream density.

  1. 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.

  2. Nonequilibrium effects on the aerothermodynamics of transatmospheric and aerobraking vehicles

    NASA Technical Reports Server (NTRS)

    Hassan, Basil; Candler, Graham V.

    1993-01-01

    A 3D CFD algorithm is used to study the effect of thermal and chemical nonequilibrium on slender and blunt body aerothermodynamics. Both perfect gas and reacting gas air models are used to compute the flow over a generic transatmospheric vehicle and a proposed lunar transfer vehicle. The reacting air is characterized by a translational-rotational temperature and a vibrational-electron-electronic temperature and includes eight chemical species. The effects of chemical reaction, vibrational excitation, and ionization on lift-to-drag ratio and trim angle are investigated. Results for the NASA Ames All-body Configuration show a significant difference in center of gravity location for a reacting gas flight case when compared to a perfect gas wind tunnel case at the same Mach number, Reynolds number, and angle of attack. For the same center of gravity location, the wind tunnel model trims at lower angle of attack than the full-scale flight case. Nonionized and ionized results for a proposed lunar transfer vehicle compare well to computational results obtained from a previously validated reacting gas algorithm. Under the conditions investigated, effects of weak ionization on the heat transfer and aerodynamic coefficients were minimal.

  3. Structural characteristics of the shuttle orbiter ceramic thermal protection system

    NASA Technical Reports Server (NTRS)

    Cooper, P. A.

    1982-01-01

    The thermal protection system (TPS) of the Space Shuttle Orbiter is described as well as the results of dynamic reponse studies conducted in support of the efforts to certify the TPS for flight. The ceramic Thermal Protection System consists of ceramic tiles bonded to felt pads which are in turn bonded to the Orbiter substructure to protect the aluminum substructure from the heat of reentry.

  4. 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.

  5. 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

  6. 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

  7. 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.

  8. Assessment Of The Aerodynamic And Aerothermodynamic Performance Of The USV-3 High-Lift Re-Entry Vehicle

    NASA Astrophysics Data System (ADS)

    Pezzella, Giuseppe; Richiello, Camillo; Russo, Gennaro

    2011-05-01

    This paper deals with the aerodynamic and aerothermodynamic trade-off analysis carried out with the aim to design a hypersonic flying test bed (FTB), namely USV3. Such vehicle will have to be launched with a small expendable launcher and shall re-enter the Earth atmosphere allowing to perform several experiments on critical re-entry phenomena. The demonstrator under study is a re-entry space glider characterized by a relatively simple vehicle architecture able to validate hypersonic aerothermodynamic design database and passenger experiments, including thermal shield and hot structures. Then, a summary review of the aerodynamic characteristics of two FTB concepts, compliant with a phase-A design level, has been provided hereinafter. Indeed, several design results, based both on engineering approach and computational fluid dynamics, are reported and discussed in the paper.

  9. Aero-Thermo-Dynamic Mass Analysis

    NASA Astrophysics Data System (ADS)

    Shiba, Kota; Yoshikawa, Genki

    2016-07-01

    Each gas molecule has its own molecular weight, while such a microscopic characteristic is generally inaccessible, and thus, it is measured indirectly through e.g. ionization in conventional mass analysis. Here, we present a novel approach to the direct measurement of molecular weight through a nanoarchitectonic combination of aerodynamics, thermodynamics, and mechanics, transducing microscopic events into macroscopic phenomena. It is confirmed that this approach can provide molecular weight of virtually any gas or vaporizable liquid sample in real-time without ionization. Demonstrations through analytical calculations, numerical simulations, and experiments verify the validity and versatility of the novel mass analysis realized by a simple setup with a flexible object (e.g. with a bare cantilever and even with a business card) placed in a laminar jet. Owing to its unique and simple working principle, this aero-thermo-dynamic mass analysis (AMA) can be integrated into various analytical devices, production lines, and consumer mobile platforms, opening new chapters in aerodynamics, thermodynamics, mechanics, and mass analysis.

  10. Aero-Thermo-Dynamic Mass Analysis.

    PubMed

    Shiba, Kota; Yoshikawa, Genki

    2016-07-14

    Each gas molecule has its own molecular weight, while such a microscopic characteristic is generally inaccessible, and thus, it is measured indirectly through e.g. ionization in conventional mass analysis. Here, we present a novel approach to the direct measurement of molecular weight through a nanoarchitectonic combination of aerodynamics, thermodynamics, and mechanics, transducing microscopic events into macroscopic phenomena. It is confirmed that this approach can provide molecular weight of virtually any gas or vaporizable liquid sample in real-time without ionization. Demonstrations through analytical calculations, numerical simulations, and experiments verify the validity and versatility of the novel mass analysis realized by a simple setup with a flexible object (e.g. with a bare cantilever and even with a business card) placed in a laminar jet. Owing to its unique and simple working principle, this aero-thermo-dynamic mass analysis (AMA) can be integrated into various analytical devices, production lines, and consumer mobile platforms, opening new chapters in aerodynamics, thermodynamics, mechanics, and mass analysis.

  11. Aero-Thermo-Dynamic Mass Analysis.

    PubMed

    Shiba, Kota; Yoshikawa, Genki

    2016-01-01

    Each gas molecule has its own molecular weight, while such a microscopic characteristic is generally inaccessible, and thus, it is measured indirectly through e.g. ionization in conventional mass analysis. Here, we present a novel approach to the direct measurement of molecular weight through a nanoarchitectonic combination of aerodynamics, thermodynamics, and mechanics, transducing microscopic events into macroscopic phenomena. It is confirmed that this approach can provide molecular weight of virtually any gas or vaporizable liquid sample in real-time without ionization. Demonstrations through analytical calculations, numerical simulations, and experiments verify the validity and versatility of the novel mass analysis realized by a simple setup with a flexible object (e.g. with a bare cantilever and even with a business card) placed in a laminar jet. Owing to its unique and simple working principle, this aero-thermo-dynamic mass analysis (AMA) can be integrated into various analytical devices, production lines, and consumer mobile platforms, opening new chapters in aerodynamics, thermodynamics, mechanics, and mass analysis. PMID:27412335

  12. Aeroassist flight experiment aerodynamics and aerothermodynamics

    NASA Technical Reports Server (NTRS)

    Brewer, Edwin B.

    1989-01-01

    The problem is to determine the transitional flow aerodynamics and aerothermodynamics, including the base flow characteristics, of the Aeroassist Flight Experiment (AFE). The justification for the computational fluid dynamic (CFD) Application stems from MSFC's system integration responsibility for the AFE. To insure that the AFE objectives are met, MSFC must understand the limitations and uncertainties of the design data. Perhaps the only method capable of handling the complex physics of the rarefied high energy AFE trajectory is Bird's Direct Simulation Monte Carlo (DSMC) technique. The 3-D code used in this analysis is applicable only to the AFE geometry. It uses the Variable Hard Sphere (VHS) collision model and five specie chemistry model available from Langley Research Center. The code is benchmarked against the AFE flight data and used as an Aeroassisted Space Transfer Vehicle (ASTV) design tool. The code is being used to understand the AFE flow field and verify or modify existing design data. Continued application to lower altitudes is testing the capability of the Numerical Aerodynamic Simulation Facility (NASF) to handle 3-D DSMC and its practicality as an ASTV/AFE design tool.

  13. Aero-Thermo-Dynamic Mass Analysis

    PubMed Central

    Shiba, Kota; Yoshikawa, Genki

    2016-01-01

    Each gas molecule has its own molecular weight, while such a microscopic characteristic is generally inaccessible, and thus, it is measured indirectly through e.g. ionization in conventional mass analysis. Here, we present a novel approach to the direct measurement of molecular weight through a nanoarchitectonic combination of aerodynamics, thermodynamics, and mechanics, transducing microscopic events into macroscopic phenomena. It is confirmed that this approach can provide molecular weight of virtually any gas or vaporizable liquid sample in real-time without ionization. Demonstrations through analytical calculations, numerical simulations, and experiments verify the validity and versatility of the novel mass analysis realized by a simple setup with a flexible object (e.g. with a bare cantilever and even with a business card) placed in a laminar jet. Owing to its unique and simple working principle, this aero-thermo-dynamic mass analysis (AMA) can be integrated into various analytical devices, production lines, and consumer mobile platforms, opening new chapters in aerodynamics, thermodynamics, mechanics, and mass analysis. PMID:27412335

  14. Uncertainty Assessment of Hypersonic Aerothermodynamics Prediction Capability

    NASA Technical Reports Server (NTRS)

    Bose, Deepak; Brown, James L.; Prabhu, Dinesh K.; Gnoffo, Peter; Johnston, Christopher O.; Hollis, Brian

    2011-01-01

    The present paper provides the background of a focused effort to assess uncertainties in predictions of heat flux and pressure in hypersonic flight (airbreathing or atmospheric entry) using state-of-the-art aerothermodynamics codes. The assessment is performed for four mission relevant problems: (1) shock turbulent boundary layer interaction on a compression corner, (2) shock turbulent boundary layer interaction due a impinging shock, (3) high-mass Mars entry and aerocapture, and (4) high speed return to Earth. A validation based uncertainty assessment approach with reliance on subject matter expertise is used. A code verification exercise with code-to-code comparisons and comparisons against well established correlations is also included in this effort. A thorough review of the literature in search of validation experiments is performed, which identified a scarcity of ground based validation experiments at hypersonic conditions. In particular, a shortage of useable experimental data at flight like enthalpies and Reynolds numbers is found. The uncertainty was quantified using metrics that measured discrepancy between model predictions and experimental data. The discrepancy data is statistically analyzed and investigated for physics based trends in order to define a meaningful quantified uncertainty. The detailed uncertainty assessment of each mission relevant problem is found in the four companion papers.

  15. 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 ...

  16. 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 ...

  17. 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.

  18. 49 CFR 179.18 - Thermal protection systems.

    Code of Federal Regulations, 2010 CFR

    2010-10-01

    ... fire for 30 minutes. (b) Thermal analysis. (1) Compliance with the requirements of paragraph (a) of..., underframes, metal jackets, insulation, and thermal protection. A complete record of each analysis shall be... analyzing the fire effects on the entire surface of the tank car. (2) When the analysis shows the...

  19. Ablation Thermal Protection Systems: Suitability of ablation systems to thermal protection depends on complex physical and chemical processes.

    PubMed

    Ungar, E W

    1967-11-10

    The performance of ablation thermal protection systems is intimately related to the mass transfer, heat transfer, and chemical reactions which occur within the gas boundary layer. Production of a liquid layer and phase change or chemical reaction heat sinks greatly improve materials performance. Materials are available which achieve many goals for thermal protection. However, advanced materials which are now being developed provide hope of further reductions in the weight of heat-shielding structures. PMID:17732614

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

    PubMed

    Mandal, Sumit; Song, Guowen

    2014-10-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1973-01-01

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

  2. Thermal protection of the newborn in resource-limited environments.

    PubMed

    Lunze, K; Hamer, D H

    2012-05-01

    Appropriate thermal protection of the newborn prevents hypothermia and its associated burden of morbidity and mortality. Yet, current global birth practices tend to not adequately address this challenge. Here, we discuss the pathophysiology of hypothermia in the newborn, its prevention and therapeutic options with particular attention to resource-limited environments. Newborns are equipped with sophisticated mechanisms of body temperature regulation. Neonatal thermoregulation is a critical function for newborn survival, regulated in the hypothalamus and mediated by endocrine pathways. Hypothermia activates cellular metabolism through shivering and non-shivering thermogenesis. In newborns, optimal temperature ranges are narrow and thermoregulatory mechanisms easily overwhelmed, particularly in premature and low-birth weight infants. Hyperthermia most commonly is associated with dehydration and potentially sepsis. The lack of thermal protection promptly leads to hypothermia, which is associated with detrimental metabolic and other pathophysiological processes. Simple thermal protection strategies are feasible at community and institutional levels in resource-limited environments. Appropriate interventions include skin-to-skin care, breastfeeding and protective clothing or devices. Due to poor provider training and limited awareness of the problem, appropriate thermal care of the newborn is often neglected in many settings. Education and appropriate devices might foster improved hypothermia management through mothers, birth attendants and health care workers. Integration of relatively simple thermal protection interventions into existing mother and child health programs can effectively prevent newborn hypothermia even in resource-limited environments. PMID:22382859

  3. 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.

  4. 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.

  5. 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.

  6. 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.

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

    NASA Astrophysics Data System (ADS)

    Zhu, Fanglong; Ma, Suqin; Zhang, Weiyuan

    2008-05-01

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

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

    NASA Technical Reports Server (NTRS)

    Wiese, M. R.

    1986-01-01

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

  9. Thermal Protection with 5% Dextrose Solution Blanket During Radiofrequency Ablation

    SciTech Connect

    Chen, Enn Alexandria Neeman, Ziv; Lee, Fred T.; Kam, Anthony; Wood, Brad

    2006-12-15

    A serious complication for any thermal radiofrequency ablation is thermal injury to adjacent structures, particularly the bowel, which can result in additional major surgery or death. Several methods using air, gas, fluid, or thermometry to protect adjacent structures from thermal injury have been reported. In the cases presented in this report, 5% dextrose water (D5W) was instilled to prevent injury to the bowel and diaphragm during radiofrequency ablation. Creating an Insulating envelope or moving organs with D5W might reduce risk for complications such as bowel perforation.

  10. 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.

  11. Thermal protection studies of plastic films and fibrous materials

    NASA Technical Reports Server (NTRS)

    Saad, Michel A.; Altman, Robert L.

    1988-01-01

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

  12. 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.

  13. 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.

  14. 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.

  15. "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.

  16. 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.

  17. The tethered satellite system for low density aerothermodynamics studies

    NASA Technical Reports Server (NTRS)

    Carlomagno, Giovanni M.; De Luca, Luigi; Siemers, P. M., III; Wood, George M., Jr.

    1986-01-01

    The feasibility of the operation of the Tethered Satellite System (TSS) as a continuous open wind tunnel for low-density aerothermodynamic studies (applicable to the design of hypersonic space vehicles including STARFAC, AOTV, and ERV) is considered. The Shuttle Continuous Open Wind Tunnel (SCOWT) program, for the study of the energy and momentum transfer between the tethered satellite and its environmental medium during the TSS/2 mission, is described. Instrumentation and TSS design requirements to meet SCOWT objectives are also considered. SCOWT will provide information on the gasdynamic processes occurring downstream of the bow wave standing in front of the TS, the chemistry and physics of the upper atmosphere related to satellite aerothermodynamics, and TSS's overall experimental envelope of operation.

  18. The definition of the Shuttle Tethered Aerothermodynamic Research Facility

    NASA Technical Reports Server (NTRS)

    Siemers, P. M., III; Wood, G. M., Jr.; Wolf, H.; Flanagan, P. F.; Henry, M. W.

    1985-01-01

    Studies have been conducted to define the feasibility and practical limitations of the Shuttle Orbiter Tethered 'wind-tunnel' concept. This concept, referred to as the Shuttle Tethered Aerothermodynamic Research Facility (STARFAC), is proposed to provide researchers access to altitudes above 90 km to accomplish aerothermodynamic research in the rarefied upper atmosphere. Determining the feasibility and limitations of the concept has required the enhancement and/or development of mission simulation analytical techniques and control laws; the accomplishment of candidate mission simulations; the definition of instrumentation requirements, both for science and engineering; and the establishment of tether and satellite design requirements to meet STARFAC objectives. The results of the study, to date, indicate that such a concept is both feasible and practical. Representative results are presented, as are recommendations for continued studies which would result in program implementation.

  19. Saenger: The reference concept and its technological requirements - aerothermodynamics

    NASA Astrophysics Data System (ADS)

    Hirschel, E. H.

    1991-08-01

    The objectives of the technology program 'aerothermodynamics and propulsion integration' are defined. An overview of the special aerothermodynamic phenomena which must be regarded in the design of the Saenger lower stage which presently stands in the center of the technology program is given. The design tools, which must be provided; the components like the inlet, the afterbody, etc., which must be designed and tested; and the special problems like forebody optimization, heat load determination, upper stage integration, etc., which must be treated, are discussed. The general work plan is presented, showing the major activities up to start of the development of the Saenger space transportation system. It includes the development and manufacturing of the experimental vehicle (HYTEXT) as a means for the validation of the design tools and methods which are achieved in the technology program, and for the creation of a freeflight data base.

  20. 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.

  1. 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.

  2. Downward-deployed tethered platforms for high enthalpy aerothermodynamic research

    NASA Technical Reports Server (NTRS)

    Wood, George M.; Siemers, Paul M.; Squires, R. Kenneth; Wolf, Henry; Carlomagno, Giovanni M.

    1988-01-01

    The data on aerothermodynamic and aerodynamic interactions at altitudes above 50 km is extremely limited because of the relative inaccessibility of the region to research vehicles of any sort. This paper addresses the practicability of using downward deployed satellites tethered to an orbiting host vehicle in order to obtain steady-state data in the upper reaches of the region above 80 or 90 km.

  3. Phase C aerothermodynamic data base. [for space shuttle program

    NASA Technical Reports Server (NTRS)

    Moser, M., Jr.

    1974-01-01

    Summary listings of published documentation of SADSAC processed data arranged chronologically and by shuttle configuration are presented to provide an up-to-date record of all applicable aerothermodynamic data collected, processed, or summarized in the course of the space shuttle program. The various tables or listings are designed to provide survey information to the various space shuttle managerial and technical levels. The various listings of the shuttle test data information, the list contents, and the purpose are described.

  4. 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.

  5. Hypervelocity impact testing of Shuttle Orbiter thermal protection system tiles

    NASA Technical Reports Server (NTRS)

    Christiansen, Eric L.; Ortega, Javier

    1990-01-01

    Results are presented from a series of 22 hypervelocity impact tests carried out on the thermal protection system (TPS) for the Shuttle Orbiter. Both coated and uncoated low-density (0.14 g/cu cm) LI-900 and high-density (0.35 g/cu cm) LI-2200 tiles were tested. The results are used to develop the penetration and damage correlations which can be used in meteoroid and debris hazard analyses for spacecraft with a ceramic tile TPS. It is shown that tile coatings act as a 'bumper' to fragment the impacting projectile, with thicker coating providing increased protection.

  6. Thermal protection performance of survival suits in ice-water.

    PubMed

    Hayward, J S

    1984-03-01

    Five models of dry and insulative survival suits for cold-water immersion were studies for their thermal protection while worn by subjects floating in ice-water at 1 degree C for 6 h. Rectal temperature, extremity skin temperatures, and heart rate were monitored. No significant differences occurred in the thermal or heart rate responses of the subjects wearing the five models of suits. The immersion induced only mild hypothermia; the overall amount of rectal cooling during the 6-h period was 0.80 degree C, or a rate of 0.13 degree C/h. Extremity temperatures fell within the first 2 h to levels that induced strong sensation of cold and moderate shivering. The results show that dry and insulative survival suits can provide excellent protection against fatal levels of hypothermia resulting from ice-water immersion. Predicted survival time, given the very low cooling rate observed, exceeds 1 d if drowning can be avoided. PMID:6721808

  7. Development of processing techniques for advanced thermal protection materials

    NASA Technical Reports Server (NTRS)

    Selvaduray, Guna S.

    1995-01-01

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

  8. The Challenges of Credible Thermal Protection System Reliability Quantification

    NASA Technical Reports Server (NTRS)

    Green, Lawrence L.

    2013-01-01

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

  9. Impact Testing of Orbiter Thermal Protection System Materials

    NASA Technical Reports Server (NTRS)

    Kerr, Justin

    2006-01-01

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

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

    PubMed

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

    1965-07-01

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

  11. Aerothermodynamic Testing of Protuberances and Penetrations on the NASA Crew Exploration Vehicle Heat Shield in the NASA Langley 20-Inch Mach 6 Air Tunnel

    NASA Technical Reports Server (NTRS)

    Liechty, Derek S.

    2008-01-01

    An experimental wind tunnel program is being conducted in support of an Agency wide effort to develop a replacement for the Space Shuttle and to support the NASA s long-term objective of returning to the moon and then on to Mars. This paper documents experimental measurements made on several scaled ceramic heat transfer models of the proposed Crew Exploration Vehicle. Global heat transfer images and heat transfer distributions obtained using phosphor thermography were used to infer interference heating on the Crew Exploration Vehicle Cycle 1 heat shield from local protuberances and penetrations for both laminar and turbulent heating conditions. Test parametrics included free stream Reynolds numbers of 1.0x10(exp 6)/ft to 7.25x10(exp 6)/ft in Mach 6 air at a fixed angle-of-attack. Single arrays of discrete boundary layer trips were used to trip the boundary layer approaching the protuberances/penetrations to a turbulent state. Also, the effects of three compression pad diameters, two radial locations of compression pad/tension tie location, compression pad geometry, and rotational position of compression pad/tension tie were examined. The experimental data highlighted in this paper are to be used to validate CFD tools that will be used to generate the flight aerothermodynamic database. Heat transfer measurements will also assist in the determination of the most appropriate engineering methods that will be used to assess local flight environments associated with protuberances/penetrations of the CEV thermal protection system.

  12. 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).

  13. 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.

  14. 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...

  15. 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...

  16. 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...

  17. 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...

  18. 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...

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. Aerothermodynamics of Blunt Body Entry Vehicles. Chapter 3

    NASA Technical Reports Server (NTRS)

    Hollis, Brian R.; Borrelli, Salvatore

    2011-01-01

    In this chapter, the aerothermodynamic phenomena of blunt body entry vehicles are discussed. Four topics will be considered that present challenges to current computational modeling techniques for blunt body environments: turbulent flow, non-equilibrium flow, rarefied flow, and radiation transport. Examples of comparisons between computational tools to ground and flight-test data will be presented in order to illustrate the challenges existing in the numerical modeling of each of these phenomena and to provide test cases for evaluation of Computational Fluid Dynamics (CFD) code predictions.

  6. User's Manual for the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA)

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.; Cheatwood, F. McNeil

    1996-01-01

    This user's manual provides detailed instructions for the installation and the application of version 4.1 of the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA). Also provides simulation of flow field in thermochemical nonequilibrium around vehicles traveling at hypersonic velocities through the atmosphere. Earlier versions of LAURA were predominantly research codes, and they had minimal (or no) documentation. This manual describes UNIX-based utilities for customizing the code for special applications that also minimize system resource requirements. The algorithm is reviewed, and the various program options are related to specific equations and variables in the theoretical development.

  7. Aerothermodynamics and propulsion integration in the Saenger technology programme

    NASA Astrophysics Data System (ADS)

    Hirschel, E. H.

    1991-12-01

    An overview of the special aerothermodynamic phenomena that must be considered in the design of the German Saenger Space Transportation System lower stage is presented. The design tools required, components that must be designed and tested, and certain problem areas (forebody optimization, heat load determination, and upper-stage separation), are discussed. Attention is given to the development and manufacturing of an experimental vehicle, Hytex, as a means for validating the design tools and methods realized in the technology program and for the creation of a free-flight data base.

  8. 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.

  9. 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.

  10. Design of experiments for thermal protection system process optimization

    NASA Astrophysics Data System (ADS)

    Longani, Hans R.

    2000-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Johnson, Sylvia M.

    2013-01-01

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

  12. 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

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

    PubMed

    Pendergast, David R; Mollendorf, Joseph

    2011-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Ko, William L.; Jenkins, Jerald M.

    1987-01-01

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

  15. Development and application of computational aerothermodynamics flowfield computer codes

    NASA Technical Reports Server (NTRS)

    Venkatapathy, Ethiraj

    1994-01-01

    Research was performed in the area of computational modeling and application of hypersonic, high-enthalpy, thermo-chemical nonequilibrium flow (Aerothermodynamics) problems. A number of computational fluid dynamic (CFD) codes were developed and applied to simulate high altitude rocket-plume, the Aeroassist Flight Experiment (AFE), hypersonic base flow for planetary probes, the single expansion ramp model (SERN) connected with the National Aerospace Plane, hypersonic drag devices, hypersonic ramp flows, ballistic range models, shock tunnel facility nozzles, transient and steady flows in the shock tunnel facility, arc-jet flows, thermochemical nonequilibrium flows around simple and complex bodies, axisymmetric ionized flows of interest to re-entry, unsteady shock induced combustion phenomena, high enthalpy pulsed facility simulations, and unsteady shock boundary layer interactions in shock tunnels. Computational modeling involved developing appropriate numerical schemes for the flows on interest and developing, applying, and validating appropriate thermochemical processes. As part of improving the accuracy of the numerical predictions, adaptive grid algorithms were explored, and a user-friendly, self-adaptive code (SAGE) was developed. Aerothermodynamic flows of interest included energy transfer due to strong radiation, and a significant level of effort was spent in developing computational codes for calculating radiation and radiation modeling. In addition, computational tools were developed and applied to predict the radiative heat flux and spectra that reach the model surface.

  16. Experimental Stage Separation Tool Development in NASA Langley's Aerothermodynamics Laboratory

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Scallion, William I.

    2005-01-01

    As part of the research effort at NASA in support of the stage separation and ascent aerothermodynamics research program, proximity testing of a generic bimese wing-body configuration was conducted in NASA Langley's Aerothermodynamics Laboratory in the 20-Inch Mach 6 Air Tunnel. The objective of this work is the development of experimental tools and testing methodologies to apply to hypersonic stage separation problems for future multi-stage launch vehicle systems. Aerodynamic force and moment proximity data were generated at a nominal Mach number of 6 over a small range of angles of attack. The generic bimese configuration was tested in a belly-to-belly and back-to-belly orientation at 86 relative proximity locations. Over 800 aerodynamic proximity data points were taken to serve as a database for code validation. Longitudinal aerodynamic data generated in this test program show very good agreement with viscous computational predictions. Thus a framework has been established to study separation problems in the hypersonic regime using coordinated experimental and computational tools.

  17. ESA Intermediate Experimental Vehicle. Independent Aerothermodynamic Characterization And Aerodatabase Development

    NASA Astrophysics Data System (ADS)

    Rufolo, Giuseppe C.; Di Benedetto, Sara; Walpot, Louis; Roncioni, Pietro; Marini, Marco

    2011-05-01

    In the frame of the Intermediate eXperimental Vehicle (IXV) project, the European Space Agency (ESA) is coordinating a series of technical assistance activities aimed at verifying and supporting the IXV industrial design and development process. The technical assistance is operated with the support of the Italian Space Agency (ASI), by means of the Italian Aerospace Research Center (CIRA), and the European Space Research and Technology Centre (ESTEC) under the super visioning and coordination of ESA IXV team. One of the purposes of the activity is to develop an independent capability for the assessment and verification of the industrial results with respect to the aerothermodynamic characterization of the IXV vehicle. To this aim CIRA is developing and independent AeroThermodynamics DataBase (ATDB), intended as a tool generating in output the time histories of local quantities (heat flux, pressure, skin friction) for each point of the IXV vehicle and for each trajectory (in a pre-defined envelope), together with an uncertainties model. The reference Computational Fluid Dynamics (CFD) solutions needed for the development of the tool have been provided by ESA-ESTEC (with the CFD code LORE) and CIRA (with the CFD code H3NS).

  18. Aerothermodynamic Environments Definition for the Mars Science Laboratory Entry Capsule

    NASA Technical Reports Server (NTRS)

    Edquist, Karl T.; Dyakonov, Artem A.; Wright, Michael J.; Tang, Chun Y.

    2007-01-01

    An overview of the aerothermodynamic environments definition status is presented for the Mars Science Laboratory entry vehicle. The environments are based on Navier-Stokes flowfield simulations on a candidate aeroshell geometry and worst-case entry heating trajectories. Uncertainties for the flowfield predictions are based primarily on available ground data since Mars flight data are scarce. The forebody aerothermodynamics analysis focuses on boundary layer transition and turbulent heating augmentation. Turbulent transition is expected prior to peak heating, a first for Mars entry, resulting in augmented heat flux and shear stress at the same heatshield location. Afterbody computations are also shown with and without interference effects of reaction control system thruster plumes. Including uncertainties, analysis predicts that the heatshield may experience peaks of 225 W/sq cm for turbulent heat flux, 0.32 atm for stagnation pressure, and 400 Pa for turbulent shear stress. The afterbody heat flux without thruster plume interference is predicted to be 7 W/sq cm on the backshell and 10 W/sq cm on the parachute cover. If the reaction control jets are fired near peak dynamic pressure, the heat flux at localized areas could reach as high as 76 W/sq cm on the backshell and 38 W/sq cm on the parachute cover, including uncertainties. The final flight environments used for hardware design will be updated for any changes in the aeroshell configuration, heating design trajectories, or uncertainties.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. Thermal Response and Ablation Programs for TPS Sizing Computation

    NASA Technical Reports Server (NTRS)

    Chen, Y. K.; Rasky, Daniel J. (Technical Monitor)

    1997-01-01

    The computer programs developed at NASA Ames Research Center for TPS sizing computation have been applied to many NASA's space missions, such as Mars Pathfinder, StarDust, Mars 2001, DS-II, and Saturn Entry Probe. These computer programs include FIAT (Fully Implicit Ablation and Thermal Response Program, MAT (Multi-component Ablation Thermochemistry Program), TPSX (Thermal Protection Systems Expert & Material Properties Database), and TPSGui (Thermal Protection Systems Graphical User Interface). For most planetary missions, the aerothermodynamics and material response are strongly coupled; thus a closed loop iteration technique between the FIAT and CFD (Computational Fluid Dynamics) codes has been developed to obtain the high fidelity bench mark TPS sizing solution. The computer codes and predictive methods are presented and discussed in detail.

  4. 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.

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

    NASA Astrophysics Data System (ADS)

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

    2015-01-01

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

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

    PubMed

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

    2011-12-01

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

  7. Overview of the Aerothermodynamics Analysis Conducted in Support of the STS-107 Accident Investigation

    NASA Technical Reports Server (NTRS)

    Campbell, Charles H.

    2004-01-01

    A graphic presentation of the aerothermodynamics analysis conducted in support of the STS-107 accident investigation. Investigation efforts were conducted as part of an integrated AATS team (Aero, Aerothermal, Thermal, Stress) directed by OVEWG. Graphics presented are: STS-107 Entry trajectory and timeline (1st off-nominal event to Post-LOS); Indications from OI telemetry data; Aero/aerothermo/thermal analysis process; Selected STS-107 side fuselage/OMS pod off-nominal temperatures; Leading edge structural subsystem; Relevant forensics evidence; External aerothermal environments; STS-107 Pre-entry EOM3 heating profile; Surface heating and temperatures; Orbiter wing leading edge damage survey; Internal aerothermal environments; Orbiter wing CAD model; Aerodynamic flight reconstruction; Chronology of aerodynamic/aerothermoydynamic contributions; Acreage TPS tile damage; Larger OML perturbations; Missing RCC panel(s); Localized damage to RCC panel/missing T-seal; RCC breach with flow ingestion; and Aero-aerothermal closure. NAIT served as the interface between the CAIB and NASA investigation teams; and CAIB requests for study were addressed.

  8. 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.

  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. 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.

  11. 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.

  12. Buffet loads on shuttle thermal-protection-system tiles

    NASA Technical Reports Server (NTRS)

    Coe, C. F.

    1982-01-01

    Results of wind-tunnel and acoustic tests to investigate buffet loads on Shuttle Thermal-Protection-System (TPS) tiles are given. Also described is the application of these results to the prediction of tile buffet loads for the first shuttle flight into orbit. The wind-tunnel tests of tiles were conducted at transonic and supersonic Mach numbers simulating flow regions on the Orbiter where shock waves and boundary-layer separations occur. The acoustic tests were conducted in a progressive wave tube at an overall sound pressure level (OASPL) approximately equal to the maximum OASPL measured during the wind-tunnel tests in a region of flow separation. The STS-1 buffet load predictions yielded peak tile stresses due to buffeting that were as much as 20 percent of the total stress for the design-load case when a shock wave was on a tile.

  13. Radiative metallic thermal protection systems - A status report

    NASA Technical Reports Server (NTRS)

    Bohon, H. L.; Shideler, J. L.; Rummler, D. R.

    1977-01-01

    During the early stages of the space shuttle program there were a number of technological uncertainties concerning the applicability of metallic thermal protection systems (TPS) to the multimission environment of the shuttle. To resolve the uncertainties and to advance the state-of-the-art, the NASA-Langley Research Center initiated a broad-based technology program to develop metallic TPS over the temperature range from 810 K to 1590 K. Wind tunnel tests conducted to assess the influence of surface/stream interaction of wavy surfaces on the design of metallic TPS indicate small increases in heat flux and surface drag for flow angles less than 20 deg. Analytical and experimental investigations, recently completed, have significantly improved prediction methods for cyclic creep behavior of TPS components repeatedly exposed to complex mission cycles. Thermal/structural concept optimization studies to minimize mass while maintaining structural integrity have led to advanced designs with unit masses which are competitive with those for shuttle RSI. In addition, the durability and reusability of metallic TPS have been repeatedly demonstrated in tests of full-scale systems. The current state-of-the-art strongly suggests that radiative metallic TPS have come of age.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

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

    NASA Astrophysics Data System (ADS)

    Roslyakov, P. V.

    2016-07-01

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

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

    NASA Technical Reports Server (NTRS)

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

    2003-01-01

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

  20. 75 FR 5495 - Alternate Fracture Toughness Requirements for Protection Against Pressurized Thermal Shock Events...

    Federal Register 2010, 2011, 2012, 2013, 2014

    2010-02-03

    ... Pressurized Thermal Shock Events; Correction AGENCY: Nuclear Regulatory Commission. ACTION: Final rule... (75 FR 13), that amends the NRC's regulations to provide alternate fracture toughness requirements for protection against pressurized thermal shock (PTS) events for pressurized water reactor (PWR)...

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

    NASA Technical Reports Server (NTRS)

    Eidinoff, H. L.; Rose, L.

    1974-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Dec, John A.; Braun, Robert D.

    2011-01-01

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

  3. Study of organic ablative thermal-protection coating for solid rocket motor

    NASA Astrophysics Data System (ADS)

    Hua, Zenggong

    1992-06-01

    A study is conducted to find a new interior thermal-protection material that possesses good thermal-protection performance and simple manufacturing possibilities. Quartz powder and Cr2O3 are investigated using epoxy resin as a binder and Al2O3 as the burning inhibitor. Results indicate that the developed thermal-protection coating is suitable as ablative insulation material for solid rocket motors.

  4. Aerodynamic and Aerothermodynamic Layout of the Hypersonic Flight Experiment Shefex

    NASA Astrophysics Data System (ADS)

    Eggers, Th.

    2005-02-01

    The purpose of the SHarp Edge Flight EXperiment SHEFEX is the investigation of possible new shapes for future launcher or reentry vehicles [1]. The main focus is the improvement of common space vehicle shapes by application of facetted surfaces and sharp edges. The experiment will enable the time accurate investigation of the flow effects and their structural answer during the hypersonic flight from 90 km down to an altitude of 20 km. The project, being performed under responsibility of the German Aerospace Center (DLR) is scheduled to fly on top of a two-stage solid propellant sounding rocket for the first half of 2005. The paper contains a survey of the aerodynamic and aerothermodynamic layout of the experimental vehicle. The results are inputs for the definition of the structural layout, the TPS and the flight instrumentation as well as for the preparation of the flight test performed by the Mobile Rocket Base of DLR.

  5. Numerical methods for aerothermodynamic design of hypersonic space transport vehicles

    NASA Astrophysics Data System (ADS)

    Wanie, K. M.; Brenneis, A.; Eberle, A.; Heiss, S.

    1993-04-01

    The requirement of the design process of hypersonic vehicles to predict flow past entire configurations with wings, fins, flaps, and propulsion system represents one of the major challenges for aerothermodynamics. In this context computational fluid dynamics has come up as a powerful tool to support the experimental work. A couple of numerical methods developed at MBB designed to fulfill the needs of the design process are described. The governing equations and fundamental details of the solution methods are shortly reviewed. Results are given for both geometrically simple test cases and realistic hypersonic configurations. Since there is still a considerable lack of experience for hypersonic flow calculations an extensive testing and verification is essential. This verification is done by comparison of results with experimental data and other numerical methods. The results presented prove that the methods used are robust, flexible, and accurate enough to fulfill the strong needs of the design process.

  6. Aerothermodynamic Facilities And Measurement: Flow Characterization in Shock Tunnels

    NASA Technical Reports Server (NTRS)

    Cavolowsky, John A.; Edwards, Thomas A. (Technical Monitor)

    1995-01-01

    This presentation will examine the key performance aspects of shock tunnels as they relate to their use as aerothermodynamic flow simulation facilities. Assessment of shock tube reservoir conditions and flow contaminants generated in the shock tube will be presented along with their limiting impact on viable test envelopes, Facility nozzle performance as it pertains to test time assessment and nozzle exit flow quality (survey of pressure, temperature, and species) will be addressed. Also included will be a discussion of free stream flow diagnostics, both intrusive and nonintrusive, for measurement of critical flow properties not directly inferred from surface mounted transducers. The use of computational fluid dynamics for purposes of validating experimental measurements as well as predicting performance in regimes where measurements are not feasible or possible will be discussed. The use of CFD for facility research and design will also be presented.

  7. Team Software Development for Aerothermodynamic and Aerodynamic Analysis and Design

    NASA Technical Reports Server (NTRS)

    Alexandrov, N.; Atkins, H. L.; Bibb, K. L.; Biedron, R. T.; Carpenter, M. H.; Gnoffo, P. A.; Hammond, D. P.; Jones, W. T.; Kleb, W. L.; Lee-Rausch, E. M.

    2003-01-01

    A collaborative approach to software development is described. The approach employs the agile development techniques: project retrospectives, Scrum status meetings, and elements of Extreme Programming to efficiently develop a cohesive and extensible software suite. The software product under development is a fluid dynamics simulator for performing aerodynamic and aerothermodynamic analysis and design. The functionality of the software product is achieved both through the merging, with substantial rewrite, of separate legacy codes and the authorship of new routines. Examples of rapid implementation of new functionality demonstrate the benefits obtained with this agile software development process. The appendix contains a discussion of coding issues encountered while porting legacy Fortran 77 code to Fortran 95, software design principles, and a Fortran 95 coding standard.

  8. Applications of the ram accelerator to hypervelocity aerothermodynamic testing

    NASA Technical Reports Server (NTRS)

    Bruckner, A. P.; Knowlen, C.; Hertzberg, A.

    1992-01-01

    A ram accelerator used as a hypervelocity launcher for large-scale aeroballistic range applications in hypersonics and aerodynamics research is presented. It is an in-bore ramjet device in which a projectile shaped like the centerbody of a supersonic ramjet is propelled down a stationary tube filled with a tailored combustible gas mixture. Ram accelerator operation has been demonstrated at 39 mm and 90 mm bores, supporting the proposition that this launcher concept can be scaled up to very large bore diameters of the order of 30-60 cm. It is concluded that high quality data obtained from the tube wall and projectile during the aceleration process itself are very useful for understanding aerothermodynamics of hypersonic flow in general, and for providing important CFD validation benchmarks.

  9. 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

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

    NASA Technical Reports Server (NTRS)

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

    2013-01-01

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

  11. Experimental and Computational Aerothermodynamics of a Mars Entry Vehicle

    NASA Technical Reports Server (NTRS)

    Hollis, Brian R.

    1996-01-01

    An aerothermodynamic database has been generated through both experimental testing and computational fluid dynamics simulations for a 70 deg sphere-cone configuration based on the NASA Mars Pathfinder entry vehicle. The aerothermodynamics of several related parametric configurations were also investigated. Experimental heat-transfer data were obtained at hypersonic test conditions in both a perfect gas air wind tunnel and in a hypervelocity, high-enthalpy expansion tube in which both air and carbon dioxide were employed as test gases. In these facilities, measurements were made with thin-film temperature-resistance gages on both the entry vehicle models and on the support stings of the models. Computational results for freestream conditions equivalent to those of the test facilities were generated using an axisymmetric/2D laminar Navier-Stokes solver with both perfect-gas and nonequilibrium thermochemical models. Forebody computational and experimental heating distributions agreed to within the experimental uncertainty for both the perfect-gas and high-enthalpy test conditions. In the wake, quantitative differences between experimental and computational heating distributions for the perfect-gas conditions indicated transition of the free shear layer near the reattachment point on the sting. For the high enthalpy cases, agreement to within, or slightly greater than, the experimental uncertainty was achieved in the wake except within the recirculation region, where further grid resolution appeared to be required. Comparisons between the perfect-gas and high-enthalpy results indicated that the wake remained laminar at the high-enthalpy test conditions, for which the Reynolds number was significantly lower than that of the perfect-gas conditions.

  12. 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.

  13. 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.

  14. Mechanical Testing of Carbon Based Woven Thermal Protection Materials

    NASA Technical Reports Server (NTRS)

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

    2013-01-01

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

  15. 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.

  16. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1951-01-01

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

  18. 10 CFR 50.61 - Fracture toughness requirements for protection against pressurized thermal shock events.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... pressurized thermal shock events. 50.61 Section 50.61 Energy NUCLEAR REGULATORY COMMISSION DOMESTIC LICENSING... Construction Permits § 50.61 Fracture toughness requirements for protection against pressurized thermal shock... specified in § 50.55a. (2) Pressurized Thermal Shock Event means an event or transient in pressurized...

  19. 10 CFR 50.61 - Fracture toughness requirements for protection against pressurized thermal shock events.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... pressurized thermal shock events. 50.61 Section 50.61 Energy NUCLEAR REGULATORY COMMISSION DOMESTIC LICENSING... Construction Permits § 50.61 Fracture toughness requirements for protection against pressurized thermal shock... specified in § 50.55a. (2) Pressurized Thermal Shock Event means an event or transient in pressurized...

  20. 10 CFR 50.61 - Fracture toughness requirements for protection against pressurized thermal shock events.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... pressurized thermal shock events. 50.61 Section 50.61 Energy NUCLEAR REGULATORY COMMISSION DOMESTIC LICENSING... Construction Permits § 50.61 Fracture toughness requirements for protection against pressurized thermal shock... specified in § 50.55a. (2) Pressurized Thermal Shock Event means an event or transient in pressurized...

  1. Internal Flow Thermal/Fluid Modeling of STS-107 Port Wing in Support of the Columbia Accident Investigation Board

    NASA Technical Reports Server (NTRS)

    Sharp, John R.; Kittredge, Ken; Schunk, Richard G.

    2003-01-01

    As part of the aero-thermodynamics team supporting the Columbia Accident Investigation Board (CAB), the Marshall Space Flight Center was asked to perform engineering analyses of internal flows in the port wing. The aero-thermodynamics team was split into internal flow and external flow teams with the support being divided between shorter timeframe engineering methods and more complex computational fluid dynamics. In order to gain a rough order of magnitude type of knowledge of the internal flow in the port wing for various breach locations and sizes (as theorized by the CAB to have caused the Columbia re-entry failure), a bulk venting model was required to input boundary flow rates and pressures to the computational fluid dynamics (CFD) analyses. This paper summarizes the modeling that was done by MSFC in Thermal Desktop. A venting model of the entire Orbiter was constructed in FloCAD based on Rockwell International s flight substantiation analyses and the STS-107 reentry trajectory. Chemical equilibrium air thermodynamic properties were generated for SINDA/FLUINT s fluid property routines from a code provided by Langley Research Center. In parallel, a simplified thermal mathematical model of the port wing, including the Thermal Protection System (TPS), was based on more detailed Shuttle re-entry modeling previously done by the Dryden Flight Research Center. Once the venting model was coupled with the thermal model of the wing structure with chemical equilibrium air properties, various breach scenarios were assessed in support of the aero-thermodynamics team. The construction of the coupled model and results are presented herein.

  2. 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.

  3. Survey of Aerothermodynamics Facilities Useful for the Design of Hypersonic Vehicles Using Air-Breathing Propulsion

    NASA Technical Reports Server (NTRS)

    Arnold, James O.; Deiwert, G. S.

    1997-01-01

    The dream of producing an air-breathing, hydrogen fueled, hypervelocity aircraft has been before the aerospace community for decades. However, such a craft has not yet been realized, even in an experimental form. Despite the simplicity and beauty of the concept, many formidable problems must be overcome to make this dream a reality. This paper summarizes the aero/aerothermodynamic issues that must be addressed to make the dream a reality and discusses how aerothermodynamics facilities and their modem companion, real-gas computational fluid dynamics (CFD), can help solve the problems blocking the way to realizing the dream. The approach of the paper is first to outline the concept of an air-breathing hypersonic vehicle and then discuss the nose-to-tail aerothermodynamics issues and special aerodynamic problems that arise with such a craft. Then the utility of aerothermodynamic facilities and companion CFD analysis is illustrated by reviewing results from recent United States publications wherein these problems have been addressed. Papers selected for the discussion have k e n chosen such that the review will serve to survey important U.S. aero/aerothermodynamic real gas and conventional wind tunnel facilities that are useful in the study of hypersonic, hydrogen propelled hypervelocity vehicles.

  4. Aerothermodynamic Analysis of the Project FIRE II Afterbody Flow

    NASA Technical Reports Server (NTRS)

    Wright, Micheal J.; Loomis, Mark; Arnold, Jim (Technical Monitor)

    2000-01-01

    35 years later, the Project FIRE II ballistic reentry to Earth at a nominal velocity of 11.4 km/s remains one of the best sources of heating data for the design of sample return capsules. The data from this flight experiment encompass both the thermochemical non-equilibrium and equilibrium flow regimes and include measurements of both radiative and total heating on the forebody and afterbody. Because of this, a number of researchers have performed computational fluid dynamics (CFD) simulations of the forebody of the FIRE II entry vehicle, with generally good results. In particular, Olynick et. al. coupled a Navier-Stokes solver (GIANTS) with a radiation code (NOVAR) and showed excellent agreement in surface heat transfer over the FIRE II trajectory between 1634 and 1651 seconds (77 km to 37 km). However, in most cases the primary motivation of the previous work was to understand and model the coupling between shock layer radiation and aerothermodynamics, and thus the simulations concentrated on the forebody flow only. To our knowledge there have been no prior published attempts to reproduce the afterbody heating data. However, an understanding of this data is critical to our efforts to design the next generation of Earth and planetary entry vehicles and to assess our need for additional flight data.

  5. Aerothermodynamic Environment Definition for the Genesis Sample Return Capsule

    NASA Technical Reports Server (NTRS)

    Cheatwood, F. McNeil; Merski, N. Ronald, Jr.; Riley, Christopher J.; Mitcheltree, Robert A.

    2001-01-01

    NASA's Genesis sample return mission will be the first to return material from beyond the Earth-Moon system. NASA Langley Research Center supported this mission with aerothermodynamic analyses of the sample return capsule. This paper provides an overview of that effort. The capsule is attached through its forebody to the spacecraft bus. When the attachment is severed prior to Earth entry, forebody cavities remain. The presence of these cavities could dramatically increase the heating environment in their vicinity and downstream. A combination of computational fluid dynamics calculations and wind tunnel phosphor thermography tests were employed to address this issue. These results quantify the heating environment in and around the cavities, and were a factor in the decision to switch forebody heat shield materials. A transition map is developed which predicts that the flow aft of the penetrations will still be laminar at the peak heating point of the trajectory. As the vehicle continues along the trajectory to the peak dynamic pressure point, fully turbulent flow aft of the penetrations could occur. The integrated heat load calculations show that a heat shield sized to the stagnation point levels will be adequate for the predicted environment aft of the penetrations.

  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. Survey of aerodynamics and aerothermodynamics efforts carried out in the frame of Mars exploration projects

    NASA Astrophysics Data System (ADS)

    Reynier, Philippe

    2014-10-01

    This contribution is a survey of aerodynamic and aerothermodynamics data related to Mars entry. The survey includes the studies carried out in the frame of projects aiming at preparing exploration missions involving entry probes into Mars atmosphere and the efforts have been concentrated on the aerothermodynamics developments. Russian (including former Soviet Union), European and NASA aerothermodynamics developments for preparing such missions have been accounted for. If a focus has been dedicated to the flight data gathered during Viking and Mars Pathfinder entries, the experimental and numerical activities carried out for the different projects have been also considered. The emphasis has been put on the post-flight analysis of flight experiments. The objective of the activity has been to develop a database of the developments performed for Mars entry that will be of interest for the preparation of future missions and for testing new models related to radiative transfer, and chemical kinetics schemes based on a state-to-state approach.

  8. The use of the tethered satellite system to perform low density aerothermodynamics studies

    NASA Technical Reports Server (NTRS)

    Carlomagno, Giovanni M.; Deluca, Luigi; Siemers, Paul M.; Wood, George M., Jr.

    1988-01-01

    The Tethered Satellite System (TSS) is a cooperative space system development activity of the U.S.A. and Italy. It is comprised of the Tether Satellite (TS) and the deployer. Within TSS, the Shuttle Tethered Aerothermodynamic Research Facility (STARFAC) concept has the potential to provide access to vast portions of the upper atmosphere for atmospheric and aerothermodynamic research. The feasibility and capability of the TSS to operate as a continuous open wind tunnel and to perform low density aerothermodynamic studies are investigated. This is accomplished through a modified version of the TS simulation program (SKYHOOK). The results indicate that STARFAC concept is both feasible and practical. The TS can go below 100 km but, if thrust is used, large velocity variation (delta V) maneuvers and an attitude control are required; if a satellite lift is considered, large tether tension is produced and an attitude control is required.

  9. 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.

  10. 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.

  11. 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

  12. 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.

  13. Development Of Metallic Thermal Protection System For The Expert Re-Entry Vehicle: Design Verification

    NASA Astrophysics Data System (ADS)

    Fatemi, Javad

    2011-05-01

    The thermal protection system of the EXPERT re-entry vehicle is subjected to accelerations, vibrations, acoustic and shock loads during launch and aero-heating loads and aerodynamic forces during re-entry. To fully understand the structural and thermomechanical performances of the TPS, heat transfer analysis, thermal stress analysis, and thermal buckling analysis must be performed. This requires complex three-dimensional thermal and structural models of the entire TPS including the insulation and sensors. Finite element (FE) methods are employed to assess the thermal and structural response of the TPS to the mechanical and aerothermal loads. The FE analyses results are used for the design verification and design improvement of the EXPERT thermal protection system.

  14. [Thermal-spring parks within the ambit of protected areas: general trends in morphology and botany].

    PubMed

    Grossi, F; Battista, A; Ricci, B; Di Lascio, F; Gurgo, A M; Mastroianni, S

    1996-04-01

    It is remembered that, in contemporary era, the thermal parks arose in U.S.A. and in Canada: they represent the first example of national parks in the world. Are described legislation, classifications and various types of protected areas. The species of recommended plants for the formation, increase and/or improvement of italian thermal are mentioned.

  15. 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.

  16. 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...

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. Thermal protection of reentry vehicles by actively cooled nosetips

    NASA Technical Reports Server (NTRS)

    Walker, R. E.; Hidahl, J. W.

    1980-01-01

    Analytical modeling efforts and clear-air ground test results of a transportation-cooled nosetips (TCNT) design are presented. The discrete water injection platelet TCNT described was conceived and created to achieve the performance requirements for severe reentry vehicle trajectories. Thermal performance computer modeling techniques, combing both local heat blockage and boundary layer recovery enthalpy reduction are outlined.

  2. 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.

  3. 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.

  4. 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

  5. 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

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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

  11. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

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

  13. 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.

  14. 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

  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. 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.

  17. 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.

  18. 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.

  19. The artificially blunted leading edge concept for aerothermodynamic performance enhancement

    NASA Astrophysics Data System (ADS)

    Gupta, Anurag

    An innovative aerothermodynamic performance enhancement concept for blunted geometries in hypervelocity flight is described. An Artificially Blunted Leading Edge (ABLE) is sought to be created by the use of a flow-through channel sized to choke at supersonic (in the normal direction) conditions. As a result, a normal shock stands off the channel but the high post-shock pressures have no wall to act on, leading to a reduction in wave drag. The effective blunt body flow structure can be effective at preventing the rise in heat transfer rates at channel entrance lips. In lifting flight, the flow in the channel creates suction at the lip, significantly enhancing lift for non-slender shapes. CFD studies using Reynolds Averaged Navier-Stokes simulations provide proof-of- concept for drag reduction for blunted slender geometries and L/D enhancements for sphere-cones. The ABLE flow mechanism's robustness and its effectiveness at off- design conditions is demonstrated. The computed sphere- cone L/D enhancements are also validated with experimental results from Aeroballistic Range tests. As opposed to straight channels, ABLE variants with curved channels that provide for better volumetric efficiency, reduced viscous drag penalties and better performance were designed and investigated. The channels curve outward and exhaust the flow close to the leading edge. Even while exhausting tangentially, the exhaust-mean flow interactions were shown to enhance or create lift. The force amplification due to such interactions can also be leveraged with the channel flow exhausting nearly normal to the surface. The potential of such thrust vectoring to reduce trim drag and augment directional control in the high-speed regime was demonstrated numerically. To evaluate the concept's effectiveness at improving cd or L/D values without paying any penalties in lift, enclosed volume and peak heating rates, Multidisciplinary Design Optimization techniques are used to characterize the design space

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

    NASA Technical Reports Server (NTRS)

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

    2016-01-01

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

  1. Evaluation of coated columbium for thermal protection systems application. [spacecraft heat shielding material

    NASA Technical Reports Server (NTRS)

    Rummler, D. R.; Black, W. E.

    1975-01-01

    Coated columbium alloys were evaluated for thermal protection system (TPS) application in the maximum operating temperature range from 1370 to 1590 K. Evaluation of materials combinations, subsize panels, and single panel TPS led to the development of a full-size nine-panel TPS array. This array was subjected to simulated shuttle missions of both thermal and acoustic loading. Results are presented which illustrate the structural and thermal adequacy under uniform heating and manufacturability of a full-size coated columbium-alloy TPS which is lightweight, reliable, and reusable.

  2. Mars Science Laboratory Heatshield Aerothermodynamics: Design and Reconstruction

    NASA Technical Reports Server (NTRS)

    Edquist, Karl T.; Hollis, Brian R.; Johnston, Christopher O.; Bose, Deepak; White, Todd R.; Mahzari, Milad

    2013-01-01

    The Mars Science Laboratory heatshield was designed to withstand a fully turbulent heat pulse based on test results and computational analysis on a pre-flight design trajectory. Instrumentation on the flight heatshield measured in-depth temperatures in the thermal protection system. The data indicate that boundary layer transition occurred at 5 of 7 thermocouple locations prior to peak heating. Data oscillations at 3 pressure measurement locations may also indicate transition. This paper presents the heatshield temperature and pressure data, possible explanations for the timing of boundary layer transition, and a qualitative comparison of reconstructed and computational heating on the as-flown trajectory. Boundary layer Reynolds numbers that are typically used to predict transition are compared to observed transition at various heatshield locations. A uniform smooth-wall transition Reynolds number does not explain the timing of boundary layer transition observed during flight. A roughness-based Reynolds number supports the possibility of transition due to discrete or distributed roughness elements on the heatshield. However, the distributed roughness height would have needed to be larger than the pre-flight assumption. The instrumentation confirmed the predicted location of maximum turbulent heat flux near the leeside shoulder. The reconstructed heat flux at that location is bounded by smooth-wall turbulent calculations on the reconstructed trajectory, indicating that augmentation due to surface roughness probably did not occur. Turbulent heating on the downstream side of the heatshield nose exceeded smooth-wall computations, indicating that roughness may have augmented heating. The stagnation region also experienced heating that exceeded computational levels, but shock layer radiation does not fully explain the differences.

  3. 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.

  4. 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.

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

    NASA Technical Reports Server (NTRS)

    Zell, Peter

    2012-01-01

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

  6. The use of the Tethered Satellite System to perform low density aerothermodynamics studies

    NASA Technical Reports Server (NTRS)

    Carlomagno, Giovanni M.; De Luca, Luigi; Siemers, Paul M.; Wood, George M., Jr.

    1988-01-01

    The Tethered Satellite System (TSS) is a cooperative space system development activity being carried out by USA and Italy. Within TSS, the Shuttle Tethered Aerothermodynamic Research Facility (STARFAC) concept has the potential to provide access to vast portions of the upper atmosphere for the purpose of atmospheric and aerothermodynamic research. The implementation of this capability will push Tether System (TS) state of the art to its limits; the primary problems being tether/satellite drag, heating, tension control, deployment/retrieval control. In this paper parametric studies are accomplished to assess some of these problems and to delineate the tradeoffs available to missions design to meet the engineering constraints. The utilization of aerodynamic rather than spherical shapes - (TSS) - as well as elementary satellite thrusting and lift are included in the present study.

  7. Numerical Simulations Of High-Altitude Aerothermodynamics Of A Prospective Spacecraft Model

    NASA Astrophysics Data System (ADS)

    Vashchenkov, P. V.; Kaskovsky, A. V.; Krylov, A. N.; Ivanov, M. S.

    2011-05-01

    The paper describes the computations of aerothermodynamic characteristics of a promising spacecraft (Prospective Piloted Transport System) along its de- scent trajectory at altitudes from 120 to 60 km. The computations are performed by the DSMC method with the use of the SMILE software system and by the engineering technique (local bridging method) with the use of the RuSat software system. The influence of real gas effects (excitation of rotational and vibrational energy modes and chemical reactions) on aerothermodynamic characteristics of the vehicle is studied. A comparison of results obtained by the approximate engineering method and the DSMC method allow the accuracy of prediction of aerodynamic characteristics by the local bridging method to be estimated.

  8. Real-Gas Flow Properties for NASA Langley Research Center Aerothermodynamic Facilities Complex Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Hollis, Brian R.

    1996-01-01

    A computational algorithm has been developed which can be employed to determine the flow properties of an arbitrary real (virial) gas in a wind tunnel. A multiple-coefficient virial gas equation of state and the assumption of isentropic flow are used to model the gas and to compute flow properties throughout the wind tunnel. This algorithm has been used to calculate flow properties for the wind tunnels of the Aerothermodynamics Facilities Complex at the NASA Langley Research Center, in which air, CF4. He, and N2 are employed as test gases. The algorithm is detailed in this paper and sample results are presented for each of the Aerothermodynamic Facilities Complex wind tunnels.

  9. Wind-tunnel based definition of the AFE aerothermodynamic environment. [Aeroassist Flight Experiment

    NASA Technical Reports Server (NTRS)

    Miller, Charles G.; Wells, W. L.

    1992-01-01

    The Aeroassist Flight Experiment (AFE), scheduled to be performed in 1994, will serve as a precursor for aeroassisted space transfer vehicles (ASTV's) and is representative of entry concepts being considered for missions to Mars. Rationale for the AFE is reviewed briefly as are the various experiments carried aboard the vehicle. The approach used to determine hypersonic aerodynamic and aerothermodynamic characteristics over a wide range of simulation parameters in ground-based facilities is presented. Facilities, instrumentation and test procedures employed in the establishment of the data base are discussed. Measurements illustrating the effects of hypersonic simulation parameters, particularly normal-shock density ratio (an important parameter for hypersonic blunt bodies), and attitude on aerodynamic and aerothermodynamic characteristics are presented, and predictions from computational fluid dynamic (CFD) computer codes are compared with measurement.

  10. Overview of X-38 Hypersonic Aerothermodynamic Wind Tunnel Data and Comparison with Numerical Results

    NASA Technical Reports Server (NTRS)

    Campbell, C.; Caram, J.; Berry, S.; Horvath, T.; Merski, N.; Loomis, M.; Venkatapathy, E.

    2004-01-01

    A NASA team of engineers has been organized to design a crew return vehicle for returning International Space Station crew members from orbit. The hypersonic aerothermodynamic characteristics of the X-23/X-24A derived X-38 crew return vehicle are being evaluated in various wind tunnels in support of this effort. Aerothermodynamic data from two NASA hypersonic tunnels at Mach 6 and Mach 10 has been obtained with cast ceramic models and a thermographic phosphorus digital imaging system. General windward surface heating features are described based on experimental surface heating images and surface oil flow patterns for the nominal hypersonic aerodynamic orientation. Body flap reattachment heating levels are examined. Computational Fluid Dynamics tools have been applied at the appropriate wind tunnel conditions to make comparisons with this data.

  11. 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'

  12. 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.

  13. 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.

  14. 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.

  15. Hypersonic research engine/aerothermodynamic integration model, experimental results. Volume 1: Mach 6 component integration

    NASA Technical Reports Server (NTRS)

    Andrews, E. H., Jr.; Mackley, E. A.

    1976-01-01

    The NASA Hypersonic Research Engine (HRE) Project was initiated for the purpose of advancing the technology of airbreathing propulsion for hypersonic flight. A large component (inlet, combustor, and nozzle) and structures development program was encompassed by the project. The tests of a full-scale (18 in. diameter cowl and 87 in. long) HRE concept, designated the Aerothermodynamic Integration Model (AIM), at Mach numbers of 5, 6, and 7. Computer program results for Mach 6 component integration tests are presented.

  16. Hypersonic research engine project. Phase 2: Aerothermodynamic Integration Model (AIM) test report

    NASA Technical Reports Server (NTRS)

    Andersen, W. L.; Kado, L.

    1975-01-01

    The Hypersonic Research Engine-Aerothermodynamic Integration Model (HRE-AIM) was designed, fabricated, and tested in the Hypersonic Tunnel Facility. The HRE-AIM is described along with its installation in the wind tunnel facility. Test conditions to which the HRE-AIM was subjected and observations made during the tests are discussed. The overall engine performance, component interaction, and ignition limits for the design are evaluated.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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,…

  3. 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...

  4. 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...

  5. 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.

  6. 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.

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

    NASA Astrophysics Data System (ADS)

    Ardema, Mark D.

    1995-09-01

    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.

  8. 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.

  9. 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.

  10. 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

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

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

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

  12. 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.

  13. 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.

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

    NASA Astrophysics Data System (ADS)

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

    2012-04-01

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

  15. 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.

  16. 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.

  17. 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.

  18. Analytical characterization of AOTV perigee aerothermodynamic regime. [Aeroassisted Orbital Transfer Vehicle

    NASA Technical Reports Server (NTRS)

    Desautel, D.

    1984-01-01

    In preliminary design studies of AOTV (Aeroassisted Orbital Transfer Vehicle) configurations, it is desirable to have a simple analytical method for characterizing the aerothermodynamic regime at skip trajectory perigee as a function of vehicle L/D and m/CDA (ballistic coefficient). The present study derives an approximate perigee solution from the generalized equations of motion. For a prescribed entry velocity vector, the solution determines a Zeta function (proportional to density divided by m/CDA) at perigee as a parametric function of the exit velocity vector and (constant) L/D. The perigee Zeta function then determines perigee density (or altitude) as a parametric function of m/CDA. The solution allows the following classic aerothermodynamic parameters to be determined at perigee as parametric functions of m/CDA, L/D, and the exit velocity vector: Reynolds Number (viscous effects), Mach Number (compressibility effects), Knudsen Number (rarefaction effects), sphere and disk bow shock standoff distance, Damkohler Number (relaxation effects), viscous correlation parameter (viscid-inviscid interactions), and Stanton Number (convective heat transfer). Results of the analysis are given for low L/D and mid L/D AOTV configurations on return from geosynchronous and L5 orbits. It is concluded the method successfully provides preliminary estimates of the aerothermodynamic parameters through the use of simple algebraic equations and plots.

  19. Impact of ETO propellants on the aerothermodynamic analyses of propulsion components

    NASA Technical Reports Server (NTRS)

    Civinskas, K. C.; Boyle, R. J.; Mcconnaughey, H. V.

    1988-01-01

    The operating conditions and the propellant transport properties used in Earth-to-Orbit (ETO) applications affect the aerothermodynamic design of ETO turbomachinery in a number of ways. Some aerodynamic and heat transfer implications of the low molecular weight fluids and high Reynolds number operating conditions on future ETO turbomachinery are discussed. Using the current SSME high pressure fuel turbine as a baseline, the aerothermodynamic comparisons are made for two alternate fuel turbine geometries. The first is a revised first stage rotor blade designed to reduce peak heat transfer. This alternate design resulted in a 23 percent reduction in peak heat transfer. The second design concept was a single stage rotor to yield the same power output as the baseline two stage rotor. Since the rotor tip speed was held constant, the turbine work factor doubled. In this alternate design, the peak heat transfer remained the same as the baseline. While the efficiency of the single stage design was 3.1 points less than the baseline two stage turbine, the design was aerothermodynamically feasible, and may be structurally desirable.

  20. PRE_X Programme: Aerothermodynamic Objectives and Aeroshape Definition for in Flight Experiments

    NASA Astrophysics Data System (ADS)

    Lambert, O.; Tribot, J.-P.; Saint-Cloud, F.

    2002-01-01

    As the expendable launch vehicles (ELV) are limited in their trend to lower costs, the reusability (Reusable Launch Vehicle, RLV) could be the way to make drastic step. By the year 2001, CNES proposed through the ANGEL phase 1 programme to preprare the required technical maturity before that RLV's become alternatives to ELV's. In such way, system ,propulsion, ground based demonstrations, aero-thermo-dynamics as well as in flight experimentation are planned. This paper is focused on the aero-thermo-dynamics (ATD) and in flight demonstration activities with emphasis on the better understanding of ATD problems emerging from past programmes among them shock wave transitionnal boundary layer interaction on surface control, boundary layer transition, local aerothermodynamic effects, gas- surface interaction, catalycity, base flow prediction,...In order to minimize as small as possible the management risk a first generation of vehicle dubbed Pre_X is designed to validate technological choices and to have as soon as possible re-entry data to calibrate the various tools involved in the future RLV definition. In addition, the main requirement for PRE_X aeroshape definition and the two different design approaches considered by Dassault Aviation and EADS-LV are discussed. Then, the more promising concept for the PRE_X application is presented. Finally, the current status of the ATD activities is given as well as the perspectives.

  1. 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.

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

    DOE PAGES

    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

  3. 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.

  4. 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.

  5. 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.

  6. 71 FR 4958 - Advisory Circular 25.856-2, Installation of Thermal/Acoustic Insulation for Burnthrough Protection

    Federal Register 2010, 2011, 2012, 2013, 2014

    2006-01-30

    ... Federal Aviation Administration Advisory Circular 25.856-2, Installation of Thermal/Acoustic Insulation..., ``Installation of Thermal/Acoustic Insulation for Burnthrough Protection.'' The advisory circular provides... thermal/acoustic insulation. DATES: AC 25.856-2 was issued by the FAA Transport Airplane Directorate...

  7. 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.

  8. Chemically modified thermal-spray zinc anodes for galvanic cathodic protection

    SciTech Connect

    Covino, B.S. Jr.; Bullard, S.J.; Holcomb, G.R.; Russell, J.H.; Cramer, S.D.; Bennett, J.E.; Laylor, H.M.

    1999-12-01

    Humectants, substances that promote the retention of moisture, were applied to new and previously aged thermal-sprayed Zn anodes to improve the performance of galvanic cathodic protection systems. Anodes on steel-reinforced concrete were treated with aqueous solutions of the humectants lithium nitrate (LiNO{sub 3}) and lithium bromide (LiBr). LiBr was the most beneficial humectant, increasing the average galvanic current density of new thermal-sprayed Zn anodes by as much as a factor of six.

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

    NASA Astrophysics Data System (ADS)

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

    2014-03-01

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

  10. 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

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

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

    NASA Technical Reports Server (NTRS)

    Alexander, J. G.

    1973-01-01

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

  20. 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.

  1. 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.

  2. 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

  3. 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.

  4. Electrochemical aging of humectant-treated thermal-sprayed zinc anodes for cathodic protection

    SciTech Connect

    Covino, B.S. Jr.; Holcomb, G.R.; Bullard, S.J.; Russell, J.H.; Cramer, S.D.; Bennett, J.E.; Laylor, H.M.

    1999-07-01

    Humectants, substances that promote the retention of moisture, were studied to determine their effectiveness in improving the performance and extending the service life of both new and previously-aged thermal-sprayed Zn anodes used in impressed current (ICCP) and galvanic cathodic protection (GCP) systems for steel-reinforced concrete structures. Potassium acetate, lithium nitrate, and lithium bromide were applied to a series of thermal-sprayed Zn-coated concrete slabs before starting the ICCP or GCP experiment. All of the humectants altered the behavior of the thermal-sprayed Zn anodes. LiNO{sub 3} was the most beneficial for ICCP anodes and LiBr was the most beneficial for GCP anodes. Circuit resistances for ICCP anodes and galvanic current density for GCP anodes are compared on the basis of electrochemical aging, humidity, and type of humectant.

  5. 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.

  6. 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).

  7. 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.

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

    NASA Technical Reports Server (NTRS)

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

    2010-01-01

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

  9. 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.

  10. 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

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. Performance of a Haynes 188 metallic standoff thermal protection system at Mach 7

    NASA Technical Reports Server (NTRS)

    Avery, D. E.

    1981-01-01

    A flight weight, metallic thermal protection system (TPS) model applicable to reentry and hypersonic vehicles was subjected to multiple cycles of both radiant and aerothermal heating to evaluate its aerothermal performance and structural integrity. The TPS was designed for a maximum operating temperature of 1255 K and featured a shingled, corrugation stiffened corrugated skin heat shield of Haynes 188, a cobalt base alloy. The model was subjected to 3 radiant preheat/aerothermal tests for a total of 67 seconds and to 15 radiant heating tests for a total of 85.9 minutes at 1255 K. The TPS limited the primary structure to temperatures below 430 K in all tests. No catastrophic failures occurred in the heat shields, supports, or insulation system. The TPS continued to function even after exposure to a differential temperature 4 times the design value produced thermal buckles in the outer skin. The shingled thermal expansion joint effectively allowed for thermal expansion of the heat shield without allowing any appreciable hot gas flow into the model cavity, even though the overlap gap between shields increased after several thermal cycles.

  16. Effect of surface catalycity on high-altitude aerothermodynamics of reentry vehicles

    NASA Astrophysics Data System (ADS)

    Molchanova, A. N.; Kashkovsky, A. V.; Bondar, Ye. A.

    2016-10-01

    This work is aimed at the development of surface chemistry models for the Direct Simulation Monte Carlo (DSMC) method applicable to non-equilibrium high-temperature flows about reentry vehicles. Probabilities of the surface processes dependent on individual properties of each particular molecule are determined from the macroscopic reaction rate data. Two different macroscopic finite rate sets are used for construction of DSMC surface recombination models. The models are implemented in the SMILE++ software system for DSMC computations. A comparison with available experimental data is performed. Effects of surface recombination on the aerothermodynamics of a blunt body at high-altitude reentry conditions are numerically studied with the DSMC method.

  17. Distributed-Memory Computing With the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA)

    NASA Technical Reports Server (NTRS)

    Riley, Christopher J.; Cheatwood, F. McNeil

    1997-01-01

    The Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA), a Navier-Stokes solver, has been modified for use in a parallel, distributed-memory environment using the Message-Passing Interface (MPI) standard. A standard domain decomposition strategy is used in which the computational domain is divided into subdomains with each subdomain assigned to a processor. Performance is examined on dedicated parallel machines and a network of desktop workstations. The effect of domain decomposition and frequency of boundary updates on performance and convergence is also examined for several realistic configurations and conditions typical of large-scale computational fluid dynamic analysis.

  18. A Review of Hypersonics Aerodynamics, Aerothermodynamics and Plasmadynamics Activities within NASA's Fundamental Aeronautics Program

    NASA Technical Reports Server (NTRS)

    Salas, Manuel D.

    2007-01-01

    The research program of the aerodynamics, aerothermodynamics and plasmadynamics discipline of NASA's Hypersonic Project is reviewed. Details are provided for each of its three components: 1) development of physics-based models of non-equilibrium chemistry, surface catalytic effects, turbulence, transition and radiation; 2) development of advanced simulation tools to enable increased spatial and time accuracy, increased geometrical complexity, grid adaptation, increased physical-processes complexity, uncertainty quantification and error control; and 3) establishment of experimental databases from ground and flight experiments to develop better understanding of high-speed flows and to provide data to validate and guide the development of simulation tools.

  19. 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.

  20. 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.

  1. 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.

  2. A study on metallic thermal protection system panel for Reusable Launch Vehicle

    NASA Astrophysics Data System (ADS)

    Caogen, Yao; Hongjun, Lü; Zhonghua, Jia; Xinchao, Jia; Yan, Lu; Haigang, Li

    2008-07-01

    A Ni-based superalloy honeycomb thermal protection system (TPS) panel has been fabricated. And a curved Ni-based superalloy honeycomb sandwich has also been fabricated. The preliminary thermal insulation results of a fabricated Ni-based superalloy honeycomb TPS panel (the areal density of this panel is 6.7 kg /m2 and total height is 32 mm) indicate that the maximum temperature of the lower surfaces of the panel is lower than 150∘ C when the temperature of outer surface is held at 650∘ C for 30 min. The flatwise tensile strength and compressive properties of a fabricated Ni-based superalloy honeycomb sandwich coupon was studied at room temperature. A multilayered coating has been developed on the surface of the superalloy honeycomb TPS panel for environmental protection and thermal control. The oxidation weight-change results show that the weight change of the Ni-based superalloy honeycomb sandwich with the oxidation resistant coating is extremely small at 1100∘ C in air for 10 h. The emittance layer of the multilayered coating imparts an emittance in excess of 0.85 during exposure at 850∘ C, which was at least 14% greater than that of the substrate with oxidation resistant alone.

  3. 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.

  4. 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.

  5. 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

  6. 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.

  7. 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

  8. 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.

  9. 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.

  10. Thermal sprayed titanium anode for cathodic protection of reinforced concrete bridges

    NASA Astrophysics Data System (ADS)

    Cramer, S. D.; Covino, B. S.; Holcomb, G. R.; Bullard, S. J.; Collins, W. K.; Govier, R. D.; Wilson, R. D.; Laylor, H. M.

    1999-03-01

    Stable operation of cobalt catalyzed thermal sprayed titanium anodes for cathodic protection (CP) of bridge reinforcing steel was maintained in accelerated tests for a period equivalent to 23 years service at Oregon Department of Transportation (Oregon DOT) bridge CP conditions with no evidence that operation would degrade with further aging. The cobalt catalyst dispersed into the concrete near the anodeconcrete interface with electrochemical aging to produce a more diffuse anode reaction zone. The titanium anode had a porous heterogeneous structure composed of α-titanium containing interstitial oxygen and nitrogen, and a fee phase thought to be Ti(O,N). Splat cooling rates were 10 to 150 K/s, and microstructures were produced by equilibrium processes at the splat solidification front. Nitrogen gas atomization during thermal spraying produced a coating with more uniform composition, less cracking, and lower resistivity than using air atomization.

  11. 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

  12. 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

  13. 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.

  14. Survey of Aerothermodynamics Facilities Useful for the Design of Hypersonic Vehicles Using Air-Breathing Propulsion

    NASA Technical Reports Server (NTRS)

    Arnold, James O.; Deiwert, George S.

    1997-01-01

    This paper surveys the use of aerothermodynamic facilities which have been useful in the study of external flows and propulsion aspects of hypersonic, air-breathing vehicles. While the paper is not a survey of all facilities, it covers the utility of shock tunnels and conventional hypersonic blow-down facilities which have been used for hypersonic air-breather studies. The problems confronting researchers in the field of aerothermodynamics are outlined. Results from the T5 GALCIT tunnel for the shock-on lip problem are outlined. Experiments on combustors and short expansion nozzles using the semi-free jet method have been conducted in large shock tunnels. An example which employed the NASA Ames 16-Inch shock tunnel is outlined, and the philosophy of the test technique is described. Conventional blow-down hypersonic wind tunnels are quite useful in hypersonic air-breathing studies. Results from an expansion ramp experiment, simulating the nozzle on a hypersonic air-breather from the NASA Ames 3.5 Foot Hypersonic wind tunnel are summarized. Similar work on expansion nozzles conducted in the NASA Langley hypersonic wind tunnel complex is cited. Free-jet air-frame propulsion integration and configuration stability experiments conducted at Langley in the hypersonic wind tunnel complex on a small generic model are also summarized.

  15. 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.

  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. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. Personal, closed-cycle cooling and protective apparatus and thermal battery therefor

    SciTech Connect

    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.

  7. 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.

  8. 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:

  9. 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.

  10. Investigation of thermal protection systems effects on viscid and inviscid flow fields for manned entry systems

    NASA Technical Reports Server (NTRS)

    Bartlett, E. P.; Morse, H. L.; Tong, H.

    1971-01-01

    Procedures and methods for predicting aerothermodynamic heating to delta orbiter shuttle vehicles were reviewed. A number of approximate methods were found to be adequate for large scale parameter studies, but are considered inadequate for final design calculations. It is recommended that final design calculations be based on a computer code which accounts for nonequilibrium chemistry, streamline spreading, entropy swallowing, and turbulence. It is further recommended that this code be developed with the intent that it can be directly coupled with an exact inviscid flow field calculation when the latter becomes available. A nonsimilar, equilibrium chemistry computer code (BLIMP) was used to evaluate the effects of entropy swallowing, turbulence, and various three dimensional approximations. These solutions were compared with available wind tunnel data. It was found study that, for wind tunnel conditions, the effect of entropy swallowing and three dimensionality are small for laminar boundary layers but entropy swallowing causes a significant increase in turbulent heat transfer. However, it is noted that even small effects (say, 10-20%) may be important for the shuttle reusability concept.

  11. Aerothermodynamics of Pyrolizing Surfaces in Hypersonic Rarefied Flows

    NASA Technical Reports Server (NTRS)

    Haas, Brian L.; Milos, Frank S.; Arnold, James O. (Technical Monitor)

    1994-01-01

    Direct simulation Monte Carlo (DSMC) calculations of rarefied flows about entry bodies typically employ a fixed surface temperature or a radiative-equilibrium energy balance to compute that temperature. Such boundary conditions neglect any effects of heat capacitance and heat conduction in the spacecraft heat shield and, therefore, provide an upper bound for the surface temperature. Such calculations also neglect pyrolysis from the heat shield which can be significant for a high-energy incident flow at very low densities. Accurate prediction of both heating and aerodynamic forces requires including pyrolysis and surface heat transfer in the models for gas-surface interaction employed in DSMC methods. Although these physical models have long appeared in various continuum flow calculation codes, they have only recently appeared in DSMC codes which are required to simulate rarefied flows during entry at high altitudes. In the current implementation, routines from the widely distributed Charring Material Thermal Response and Ablation (CMA) program are coupled into a DSMC code to calculate the one-dimensional heat transfer into the carbon phenolic heat shield at each point on a vehicle surface. Temperature-dependent material properties, surface re-radiation, and in-depth pyrolysis were included in the calculation, but surface ablation was neglected. Sample calculations for entry of the Galileo probe into the atmosphere of Jupiter demonstrate that including pyrolysis in the model leads to significant differences in predicted aerodynamics. Granted, the drag coefficient does not depend strongly on the surface temperature which can itself be significantly below the radiative equilibrium value during entry. However, the surface mass flux due to pyrolysis of the material is significant once the probe drops to altitudes characterized by transition flow. This leads to a noticeable increase in drag and a decrease in heating compared to a body without pyrolysis.

  12. Quantitative assessment of the relationship between radiant heat exposure and protective performance of multilayer thermal protective clothing during dry and wet conditions.

    PubMed

    Fu, M; Weng, W G; Yuan, H Y

    2014-07-15

    The beneficial effect of clothing on a person is important to the criteria for people exposure to radiant heat flux from fires. The thermal protective performance of multilayer thermal protective clothing exposed to low heat fluxes during dry and wet conditions was studied using two designed bench-scale test apparatus. The protective clothing with four fabric layers (outer shell, moisture barrier, thermal linear and inner layer) was exposed to six levels of thermal radiation (1, 2, 3, 5, 7 and 10kW/m(2)). Two kinds of the moisture barrier (PTFE and GoreTex) with different vapor permeability were compared. The outside and inside surface temperatures of each fabric layer were measured. The fitting analysis was used to quantitatively assess the relationship between the temperature of each layer during thermal exposure and the level of external heat flux. It is indicated that there is a linear correlation between the temperature of each layer and the radiant level. Therefore, a predicted equation is developed to calculate the thermal insulation of the multilayer clothing from the external heat flux. It can also provide some useful information on the beneficial effects of clothing for the exposure criteria of radiant heat flux from fire.

  13. 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.

  14. Development of thermal protection system of the MUSES-C/DASH reentry capsule

    NASA Astrophysics Data System (ADS)

    Yamada, Tetsuya; Inatani, Yoshifumi; Honda, Masahisa; Hirai, Ken'ich

    2002-07-01

    In the final phase of the MUSES-C mission, a small capsule with asteroid sample conducts reentry flight directly from the interplanetary transfer orbit at the velocity over 12 km/s. The severe heat flux, the complicated functional requirements, and small weight budget impose several engineering challenges on the designing of the thermal protection system of the capsule. The heat shield is required to function not only as ablator but also as a structural component. The cloth-layered carbon-phenolic ablator, which has higher allowable stress, is developed in newly-devised fabric method for avoiding delamination due to the high aerodynamic heating. The ablation analysis code, which takes into account of the effect of pyrolysis gas on the surface recession rate, has been developed and verified in the arc-heating tests in the facility environment of broad range of enthalpy level. The capsule was designed to be ventilated during the reentry flight up to about atmospheric pressure by the time of parachute deployment by being sealed with porous flow-restrict material. The designing of the thermal protection system, the hardware specifications, and the ground-based test programs of both MUSES-C and DASH capsule are summarized and discussed here in this paper.

  15. Survey of the supporting research and technology for the thermal protection of the Galileo Probe

    NASA Technical Reports Server (NTRS)

    Howe, J. T.; Pitts, W. C.; Lundell, J. H.

    1981-01-01

    The Galileo Probe, which is scheduled to be launched in 1985 and to enter the hydrogen-helium atmosphere of Jupiter up to 1,475 days later, presents thermal protection problems that are far more difficult than those experienced in previous planetary entry missions. The high entry speed of the Probe will cause forebody heating rates orders of magnitude greater than those encountered in the Apollo and Pioneer Venus missions, severe afterbody heating from base-flow radiation, and thermochemical ablation rates for carbon phenolic that rival the free-stream mass flux. This paper presents a comprehensive survey of the experimental work and computational research that provide technological support for the Probe's heat-shield design effort. The survey includes atmospheric modeling; both approximate and first-principle computations of flow fields and heat-shield material response; base heating; turbulence modelling; new computational techniques; experimental heating and materials studies; code validation efforts; and a set of 'consensus' first-principle flow-field solutions through the entry maneuver, with predictions of the corresponding thermal protection requirements.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.}

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

    NASA Astrophysics Data System (ADS)

    Blosser, Max 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.

  5. 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.

  6. A multi-scale model for geared transmission aero-thermodynamics

    NASA Astrophysics Data System (ADS)

    McIntyre, Sean M.

    A multi-scale, multi-physics computational tool for the simulation of high-per- formance gearbox aero-thermodynamics was developed and applied to equilibrium and pathological loss-of-lubrication performance simulation. The physical processes at play in these systems include multiphase compressible ow of the air and lubricant within the gearbox, meshing kinematics and tribology, as well as heat transfer by conduction, and free and forced convection. These physics are coupled across their representative space and time scales in the computational framework developed in this dissertation. These scales span eight orders of magnitude, from the thermal response of the full gearbox O(100 m; 10 2 s), through effects at the tooth passage time scale O(10-2 m; 10-4 s), down to tribological effects on the meshing gear teeth O(10-6 m; 10-6 s). Direct numerical simulation of these coupled physics and scales is intractable. Accordingly, a scale-segregated simulation strategy was developed by partitioning and treating the contributing physical mechanisms as sub-problems, each with associated space and time scales, and appropriate coupling mechanisms. These are: (1) the long time scale thermal response of the system, (2) the multiphase (air, droplets, and film) aerodynamic flow and convective heat transfer within the gearbox, (3) the high-frequency, time-periodic thermal effects of gear tooth heating while in mesh and its subsequent cooling through the rest of rotation, (4) meshing effects including tribology and contact mechanics. The overarching goal of this dissertation was to develop software and analysis procedures for gearbox loss-of-lubrication performance. To accommodate these four physical effects and their coupling, each is treated in the CFD code as a sub problem. These physics modules are coupled algorithmically. Specifically, the high- frequency conduction analysis derives its local heat transfer coefficient and near-wall air temperature boundary conditions from a quasi

  7. Overexpression of Hsp90 from grass carp (Ctenopharyngodon idella) increases thermal protection against heat stress.

    PubMed

    Wu, Chu-Xin; Zhao, Feng-Yun; Zhang, Yuan; Zhu, Yu-Jiao; Ma, Mei-Sheng; Mao, Hui-Ling; Hu, Cheng-Yu

    2012-07-01

    With homologous DNA probes, we had screened a grass carp heat shock protein 90 gene (CiHsp90). The full sequence of CiHsp90 cDNA was 2793 bp, which could code a 798 amino acids peptide. The phylogenetic analysis demonstrated that CiHsp90 shared the high homology with Zebrafish Grp94. Quantitative RT-PCR analysis showed that CiHsp90 was ubiquitously expressed at lower levels in all detected tissues and up-regulated after heat shock at 34 °C or cold stress at 4 °C. To understand the function of CiHsp90 involving in thermal protection, an expression vector containing coding region cDNA was expressed in E. coli BL21 (DE3) plysS. Upon transfer from 37 °C to 42 °C, these cells that accumulated CiHsp90 peptides displayed greater thermoresistance than the control cells. While incubated at 4°C for different periods, it could also improve the cell viability. After transient transfected recombinant plasmid pcDNA3.1/CiHsp90 into mouse myeloma cell line SP2/0, we found that CiHsp90 could contribute to protecting cells against both thermal and cold extremes. On the contrary, the mutant construct ΔN-CiHsp90 (256-798aa) could abolish the protection activity both in prokaryotic cells and eukaryotic cells. Additionally, both CiHsp90 and ΔN-CiHsp90 peptides could reduce the level of citrate synthase aggregation at the high temperature.

  8. 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.

  9. Consumable and non-consumable thermal spray anodes for impressed current cathodic protection of reinforced concrete structures

    SciTech Connect

    Covino, B.S. Jr.; Cramer, S.D.; Bullard, Sophie J.; Holcomb, Gordon R.; Collins, Wesley K.; McGill, G.E.

    1998-01-01

    A comparison is presented of some of the differences between thermal spray Zn, a consumable anode, and catalyzed thermal spray Ti, a non-consumable anode, used for impressed current cathodic protection of reinforced concrete structures. The thermal spray process for both Ti and Zn is compared using the spray parameters, atomizing gases, spray rate, and cost. The thermal spray Ti and Zn coatings are compared in terms of physical properties, composition, and structure. Results of accelerated laboratory experiments are presented and comparisons between Ti and Zn are made on the effect of electrochemical aging on voltage requirements, bond strength, coating resistivity, water permeability, and anode-concrete interracial composition.

  10. 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.

  11. 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

  12. 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.

  13. 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.

  14. 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.

  15. SSTAC/ARTS review of the draft Integrated Technology Plan (ITP). Volume 8: Aerothermodynamics Automation and Robotics (A/R) systems sensors, high-temperature superconductivity

    NASA Technical Reports Server (NTRS)

    1991-01-01

    Viewgraphs of briefings presented at the SSTAC/ARTS review of the draft Integrated Technology Plan (ITP) on aerothermodynamics, automation and robotics systems, sensors, and high-temperature superconductivity are included. Topics covered include: aerothermodynamics; aerobraking; aeroassist flight experiment; entry technology for probes and penetrators; automation and robotics; artificial intelligence; NASA telerobotics program; planetary rover program; science sensor technology; direct detector; submillimeter sensors; laser sensors; passive microwave sensing; active microwave sensing; sensor electronics; sensor optics; coolers and cryogenics; and high temperature superconductivity.

  16. Flight set 360L007 (STS-33R) field joint protection system, thermal protection system, and systems tunnel components, volume 7

    NASA Technical Reports Server (NTRS)

    1990-01-01

    The performance of the thermal protection system, field joint protection system, and systems tunnel components of flight set 360L007 is presented as evaluated by postflight hardware inspection. The condition of both motors was similar to previous flights. Four aft edge strikes were noted on the ground environment instrumentation thermal protection system. The hits all left a clean substrate, indicating that the damage was caused by nozzle severance debris and/or water impact. No National Space Transportation System debris criteria for missing thermal protection system were violated. Two problem reports were written against the field joint protection system. The first concerned two cracks in the K5NA closeout over the trunnion/vent valve location on the left-hand aft field joint. A similar condition was observed on Flight 5 (360H005). The second problem report referred to a number of small surface cracks between two impact marks on the left-hand forward field joint. Neither area exhibited loose material or any abnormal heat effects, and they have no impact on flight safety.

  17. Aerodynamic and aerothermodynamic trade-off analysis of a small hypersonic flying test bed

    NASA Astrophysics Data System (ADS)

    Pezzella, Giuseppe

    2011-08-01

    This paper deals with the aerodynamic and aerothermodynamic trade-off analysis aiming to design a small hypersonic flying test bed with a relatively simple vehicle architecture. Such vehicle will have to be launched with a sounding rocket and shall re-enter the Earth atmosphere allowing to perform several experiments on critical re-entry technologies such as boundary-layer transition and shock-shock interaction phenomena. The flight shall be conducted at hypersonic Mach number, in the range 6-8 at moderate angles of attack. In the paper some design analyses are shown as, for example, the longitudinal and lateral-directional stability analysis. A preliminary optimization of the configuration has been also done to improve the aerodynamic performance and stability of the vehicle. Several design results, based both on engineering approach and computational fluid dynamics, are reported and discussed in the paper. The aerodynamic model of vehicle is also provided.

  18. Intermediate experimental vehicle, ESA program aerodynamics-aerothermodynamics key technologies for spacecraft design and successful flight

    NASA Astrophysics Data System (ADS)

    Dutheil, Sylvain; Pibarot, Julien; Tran, Dac; Vallee, Jean-Jacques; Tribot, Jean-Pierre

    2016-07-01

    With the aim of placing Europe among the world's space players in the strategic area of atmospheric re-entry, several studies on experimental vehicle concepts and improvements of critical re-entry technologies have paved the way for the flight of an experimental space craft. The successful flight of the Intermediate eXperimental Vehicle (IXV), under ESA's Future Launchers Preparatory Programme (FLPP), is definitively a significant step forward from the Atmospheric Reentry Demonstrator flight (1998), establishing Europe as a key player in this field. The IXV project objectives were the design, development, manufacture and ground and flight verification of an autonomous European lifting and aerodynamically controlled reentry system, which is highly flexible and maneuverable. The paper presents, the role of aerodynamics aerothermodynamics as part of the key technologies for designing an atmospheric re-entry spacecraft and securing a successful flight.

  19. New Hypersonic Shock Tunnel at the Laboratory of Aerothermodynamics and Hypersonics Prof. Henry T. Nagamatsu

    SciTech Connect

    Toro, P. G. P.; Minucci, M. A. S.; Chanes, J. B. Jr; Oliveira, A. C.; Gomes, F. A. A.; Myrabo, L. N.; Nagamatsu, Henry T.

    2008-04-28

    The new 0.60-m. nozzle exit diameter hypersonic shock tunnel was designed to study advanced air-breathing propulsion system such as supersonic combustion and/or laser technologies. In addition, it may be used for hypersonic flow studies and investigations of the electromagnetic (laser) energy addition for flow control. This new hypersonic shock tunnel was designed and installed at the Laboratory for of Aerothermodynamics and Hypersonics Prof. Henry T. Nagamatsu, IEAv-CTA, Brazil. The design of the tunnel enables relatively long test times, 2-10 milliseconds, suitable for the experiments performed at the laboratory. Free stream Mach numbers ranging from 6 to 25 can be produced and stagnation pressures and temperatures up to 360 atm. and up to 9,000 K, respectively, can be generated. Shadowgraph and schlieren optical techniques will be used for flow visualization.

  20. Space Shuttle hypersonic aerodynamic and aerothermodynamic flight research and the comparison to ground test results

    NASA Technical Reports Server (NTRS)

    Iliff, Kenneth W.; Shafer, Mary F.

    1993-01-01

    Aerodynamic and aerothermodynamic comparisons between flight and ground test for the Space Shuttle at hypersonic speeds are discussed. All of the comparisons are taken from papers published by researchers active in the Space Shuttle program. The aerodynamic comparisons include stability and control derivatives, center-of-pressure location, and reaction control jet interaction. Comparisons are also discussed for various forms of heating, including catalytic, boundary layer, top centerline, side fuselage, OMS pod, wing leading edge, and shock interaction. The jet interaction and center-of-pressure location flight values exceeded not only the predictions but also the uncertainties of the predictions. Predictions were significantly exceeded for the heating caused by the vortex impingement on the OMS pods and for heating caused by the wing leading-edge shock interaction.

  1. Massively parallel computational fluid dynamics calculations for aerodynamics and aerothermodynamics applications

    SciTech Connect

    Payne, J.L.; Hassan, B.

    1998-09-01

    Massively parallel computers have enabled the analyst to solve complicated flow fields (turbulent, chemically reacting) that were previously intractable. Calculations are presented using a massively parallel CFD code called SACCARA (Sandia Advanced Code for Compressible Aerothermodynamics Research and Analysis) currently under development at Sandia National Laboratories as part of the Department of Energy (DOE) Accelerated Strategic Computing Initiative (ASCI). Computations were made on a generic reentry vehicle in a hypersonic flowfield utilizing three different distributed parallel computers to assess the parallel efficiency of the code with increasing numbers of processors. The parallel efficiencies for the SACCARA code will be presented for cases using 1, 150, 100 and 500 processors. Computations were also made on a subsonic/transonic vehicle using both 236 and 521 processors on a grid containing approximately 14.7 million grid points. Ongoing and future plans to implement a parallel overset grid capability and couple SACCARA with other mechanics codes in a massively parallel environment are discussed.

  2. Introduction: Assessment of aerothermodynamic flight prediction tools through ground and flight experimentation

    NASA Astrophysics Data System (ADS)

    Schmisseur, John D.; Erbland, Peter

    2012-01-01

    This article provides an introduction and overview to the efforts of NATO Research and Technology Organization Task Group AVT-136, Assessment of Aerothermodynamic Flight Prediction Tools through Ground and Flight Experimentation. During the period of 2006-2010, AVT-136 coordinated international contributions to assess the state-of-the-art and research challenges for the prediction of critical aerothermodynamic flight phenomena based on the extrapolation of ground test and numerical simulation. To achieve this goal, efforts were organized around six scientific topic areas: (1) Noses and leading edges, (2) Shock Interactions and Control Surfaces, (3) Shock Layers and Radiation, (4) Boundary Layer Transition, (5) Gas-Surface Interactions, and (6) Base and Afterbody Flows. A key component of the AVT-136 strategy was comparison of state-of-the-art numerical simulations with data to be acquired from planned flight research programs. Although it was recognized from the onset of AVT-136 activities that reliance on flight research data yet to be collected posed a significant risk, the group concluded the substantial benefit to be derived from comparison of computational simulations with flight data warranted pursuit of such a program of work. Unfortunately, program delays and failures in the flight programs contributing to the AVT-136 effort prevented timely access to flight research data. Despite this setback, most of the scientific topic areas developed by the Task Group made significant progress in the assessment of current capabilities. Additionally, the activities of AVT-136 generated substantial interest within the international scientific research community and the work of the Task Group was prominently featured in a total of six invited sessions in European and American technical conferences. In addition to this overview, reviews of the state-of-the-art and research challenges identified by the six research thrusts of AVT-136 are also included in this special

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. Heat gain from thermal radiation through protective clothing with different insulation, reflectivity and vapour permeability.

    PubMed

    Bröde, Peter; Kuklane, Kalev; Candas, Victor; Den Hartog, Emiel A; Griefahn, Barbara; Holmér, Ingvar; Meinander, Harriet; Nocker, Wolfgang; Richards, Mark; Havenith, George

    2010-01-01

    The heat transferred through protective clothing under long wave radiation compared to a reference condition without radiant stress was determined in thermal manikin experiments. The influence of clothing insulation and reflectivity, and the interaction with wind and wet underclothing were considered. Garments with different outer materials and colours and additionally an aluminised reflective suit were combined with different number and types of dry and pre-wetted underwear layers. Under radiant stress, whole body heat loss decreased, i.e., heat gain occurred compared to the reference. This heat gain increased with radiation intensity, and decreased with air velocity and clothing insulation. Except for the reflective outer layer that showed only minimal heat gain over the whole range of radiation intensities, the influence of the outer garments' material and colour was small with dry clothing. Wetting the underclothing for simulating sweat accumulation, however, caused differing effects with higher heat gain in less permeable garments.

  9. Mission load dynamic tests of two undensified Space shuttle thermal protection system tiles

    NASA Technical Reports Server (NTRS)

    Leatherwood, J. D.; Gowdey, J. C.

    1981-01-01

    Two tests of undensified Space Shuttle thermal protection tiles under combined static and dynamic loads were conducted. The tiles had a density of approximately 144 Kg/cum (LI900 tiles) and were mounted on a strain isolation pad which was 0.41 cm (.160 inch) thick. A combined static and dynamic mission stress histogram representative of the W-3 area of the wing of the orbiter vehicle was applied. The stress histogram was provided by the space shuttle project. Results presented include: tabulation of measured peak and root-mean-square (RMS) accelerations in both compression and tension; peak SIP stress in compression and tension, peak and RMS amplitude response ratios; lateral to vertical response ratios; response time histories; peak stress distributions (histograms), and SIP extension measured both with and without static tension at various mission times.

  10. 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

  11. 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.

  12. 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

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. A thermal physiological comparison of two HAZMAT protective ensembles with and without active convective cooling.

    PubMed

    Williamson, R; Carbo, J; Luna, B; Webbon, B W

    1999-06-01

    Wearing impermeable garments for hazardous materials (HAZMAT) cleanup can often present a health and safety problem for the wearer. Even short duration cleanup activities can produce heat-stress injuries in HAZMAT workers. It was hypothesized that an internal cooling system might increase worker productivity and decrease the likelihood of heat-stress injuries in typical HAZMAT operations. Two HAZMAT protective ensembles were compared during subjects' treadmill exercise. The different ensembles were created using two different suits: a Trelleborg vapor protective suit representative of current HAZMAT suits and a prototype suit developed by engineers at the National Aeronautics and Space Administration (NASA). The two life-support systems used were a current-technology Interspiro Spirolite breathing apparatus and a liquid air breathing system that also provided convective cooling. Twelve local members of a HAZMAT team served as test subjects. They were fully instrumented to allow a complete physiological comparison of their theramal responses to the different ensembles. Results showed that cooling from the liquid air system significantly decreased thermal stress. The results of the subjective evaluations of new design features in the prototype suit were also highly favorable. Incorporation of these new design features could lead to significant operational advantages in the future.

  6. 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.

  7. Ceramic Thermal Protection Materials - How Far Can We Go? (Part II)

    NASA Astrophysics Data System (ADS)

    Hilfer, G.

    2002-01-01

    A space vehicle re-entering the earth's atmosphere is exposed to severe environmental conditions. In particular, certain surface areas of the vehicle such as the stagnation point area or exposed control surfaces have to withstand extraordinary thermal and oxidative loads. These loads that have to be taken by a thermal protection system (TPS) are driven mainly by the geometry of the vehicle, its mass and its re-entry path. As a consequence, small vehicles like the X-38 demonstrator of a re-usable crew return vehicle (CRV) need TPS components capable of withstanding temperatures of 1800°C accompanied by severe aerodynamic and chemical loads. Currently, the only promising materials having the potential of re-usability in such an environment are Si-based ceramics and related derivatives such as C-C/SiC. These materials have an extended oxidation regime leading to the formation of an oxidation-inhibiting SiO2-layer. Nevertheless, a number of parameters may turn this so-called passive oxidation mode into a different oxidation mode which can be characterized by the release of gaseous SiO. This is the so-called active oxidation mode which induces massive degradation of the material. Based on a long-term experimental and theoretical investigation performed on the constituents of SiC and its most important derivatives and oxidation products, a mechanism was proposed describing the relevant parameters which govern the transition from passive to active oxidation of SiC in a re-entry type environment. The crucial reaction process related to this transition was found to be the interaction of SiO and SiO2 with atomic and molecular oxygen, i.e. In a previous publication this mechanism was derived by a thorough study of a large number of related elementary reaction steps and the analysis of experimental findings [1]. In the course of the investigation, however, many other results have been obtained which could not be published within the frame of the above publication. Therefore

  8. 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.

  9. 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

  10. 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.

  11. Hypersonic research engine project. Phase 2: Aerothermodynamic integration model development, data item no. 55-4-21

    NASA Technical Reports Server (NTRS)

    Jilly, L. F. (Editor)

    1975-01-01

    The design and development of the Aerothermodynamic Integration Model (AIM) of the Hypersonic Research Engine (HRE) is described. The feasibility of integrating the various analytical and experimental data available for the design of the hypersonic ramjet engine was verified and the operational characteristic and the overall performance of the selected design was determined. The HRE-AIM was designed for operation at speeds of Mach 3 through Mach 8.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. Coating Layer and Corrosion Protection Characteristics in Sea Water with Various Thermal Spray Coating Materials for STS304

    NASA Astrophysics Data System (ADS)

    Kim, Seong-Jong; Woo, Yong-Bin

    We investigated the optimal method of application and the anticorrosive abilities of Zn, Al, and Zn + 15%Al spray coatings in protecting stainless steel 304 (STS304) in sea water. If a defect such as porosity or an oxide layer, causes STS304 to be exposed to sea water, and the thermal spray coating material will act as the cathode and anode, respectively. The Tafel experiments revealed that Al-coated specimens among applied coating methods had the lowest corrosion current densities. As the corrosion potential decreases with increasing corrosion current density, we estimated the characteristics and lifetime of the protective thermal spray coating layer in the galvanic cell formed by the thermal spray coating layer and STS304.

  18. Shock-induced flow separation and the orbiter thermal protection system

    NASA Astrophysics Data System (ADS)

    Waiter, S.-A.

    The Space Shuttle orbiter's thermal protection system (TPS) is composed of reusable tiles separated by narrow gaps that accommodate the contraction and expansion of the aluminum structure that the tiles protect. When local pressure gradients exist, air flows through the tile gaps and releases heat energy by convection. The gaps represent a heat short to the structure, strain isolator pad (SIP), and filler bar. A typical problem is the pressure gradient created during entry by body flap deflection. After a brief description of how this problem affects the Space Shuttle orbiter, a theoretical and experimental review of the major parameters involved in gap heating are analyzed. Then, a review of well-known classical methods to resolve the gap aeroheating problem in the presence of a pressure gradient is presented, and a few solutions are illustrated to assess the sensitivity of each one. The following section starts with a basic relationship (called "eyeball" because of its simplicity) and follows the results up through the most modern engineering approach available in the literature. It shows that in all cases calculated significant areas of overtemperature were predicted. However, none of these methods could be correlated by experimental data. Lastly, the paper presents the solution obtained by using the most sophisticated method, based upon the Navier-Stokes equations. This approach shows excellent correlation with wind tunnel data. The application to four trajectory time points shows less severe results than the other methods. This can be explained by the introduction of a certain amount of conservatism to account for uncertainties inherent in the previous analyses. No correlation of this "exact solution" with the simple preestablished relationships has been found, indicating that more parameters than expected could be involved. However, an after-the-fact, semi-empirical engineering solution that fits the Navier-Stokes solution with good agreement was established.

  19. Physico-chemical and biological characteristics of thermal springs in Köyceğiz and Dalaman basins in south-western Turkey and recommendations for their protection.

    PubMed

    Kazanci, N; Girgin, S

    2001-01-01

    Many hydrogeological researches exist on the thermal springs of Turkey but limnological researches are very deficient. The results of the limnological research of the thermal springs in the basin of the meromictic lake Köyceğiz and Dalaman Basin are as follows; the thermal springs are of euthermal or chliarothermal types. In the thermal springs thermotolerant diatoms constituted the major plankton flora. Recommendations for the protection of thermal springs are given according to the results of this research.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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

  7. 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.

  8. 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.

  9. 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.

  10. 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

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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

  20. Challenges to Computational Aerothermodynamic Simulation and Validation for Planetary Entry Vehicle Analysis

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.; Johnston, Christopher O.; Kleb, Bil

    2010-01-01

    Challenges to computational aerothermodynamic (CA) simulation and validation of hypersonic flow over planetary entry vehicles are discussed. Entry, descent, and landing (EDL) of high mass to Mars is a significant driver of new simulation requirements. These requirements include simulation of large deployable, flexible structures and interactions with reaction control system (RCS) and retro-thruster jets. Simulation of radiation and ablation coupled to the flow solver continues to be a high priority for planetary entry analyses, especially for return to Earth and outer planet missions. Three research areas addressing these challenges are emphasized. The first addresses the need to obtain accurate heating on unstructured tetrahedral grid systems to take advantage of flexibility in grid generation and grid adaptation. A multi-dimensional inviscid flux reconstruction algorithm is defined that is oriented with local flow topology as opposed to grid. The second addresses coupling of radiation and ablation to the hypersonic flow solver--flight- and ground-based data are used to provide limited validation of these multi-physics simulations. The third addresses the challenges of retro-propulsion simulation and the criticality of grid adaptation in this application. The evolution of CA to become a tool for innovation of EDL systems requires a successful resolution of these challenges.

  1. Aerothermodynamic Measurement and Prediction for Modified Orbiter at Mach 6 and 10

    NASA Technical Reports Server (NTRS)

    Micol, John R.

    1995-01-01

    Detailed heat-transfer rate distributions measured laterally over the windward surface of an orbiter-like configuration using thin-film resistance heat-transfer gauges and globally using the newly developed relative intensity, two-color thermographic phosphor technique are presented for Mach 6 and 10 in air. The angle of attack was varied from 0 to 40 deg, and the freestream Reynolds number based on the model length was varied from 4 x 10(exp 5) to 6 x 10(exp 6) at Mach 6, corresponding to laminar, transitional, and turbulent boundary layers; the Reynolds number at Mach 10 was 4 x 10(exp 5), corresponding to laminar flow. The primary objective of the present study was to provide detailed benchmark heat-transfer data for the calibration of computational fluid-dynamics codes. Predictions from a Navier-Stokes solver referred to as the Langley aerothermodynamic upwind relaxation algorithm and an approximate boundary-layer solving method known as the axisymmetric analog three-dimensional boundary layer code are compared with measurement. In general, predicted laminar heat-transfer rates are in good agreement with measurements.

  2. Challenges to Computational Aerothermodynamic Simulation and Validation for Planetary Entry Vehicle Analysis

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.; Johnston, Christopher O.; Kleb, Bil

    2010-01-01

    Challenges to computational aerothermodynamic (CA) simulation and validation of hypersonic flow over planetary entry vehicles are discussed. Entry, descent, and landing (EDL) of high mass to Mars is a significant driver of new simulation requirements. These requirements include simulation of large deployable, flexible structures and interactions with reaction control system (RCS) and retro-thruster jets. Simulation of radiation and ablation coupled to the flow solver continues to be a high priority for planetary entry analyses, especially for return to Earth and outer planet missions. Three research areas addressing these challenges are emphasized. The first addresses the need to obtain accurate heating on unstructured tetrahedral grid systems to take advantage of flexibility in grid generation and grid adaptation. A multi-dimensional inviscid flux reconstruction algorithm is defined that is oriented with local flow topology as opposed to grid. The second addresses coupling of radiation and ablation to the hypersonic flow solver - flight- and ground-based data are used to provide limited validation of these multi-physics simulations. The third addresses the challenges of retro-propulsion simulation and the criticality of grid adaptation in this application. The evolution of CA to become a tool for innovation of EDL systems requires a successful resolution of these challenges.

  3. Hypersonic Airbreathing Propulsion: An Aerodynamics, Aerothermodynamics, and Acoustics Competency White Paper

    NASA Technical Reports Server (NTRS)

    Drummond, J. Philip; Cockrell, Charles E., Jr.; Pellett, Gerald L.; Diskin, Glenn S.; Auslender, Aaron H.; Exton, Reginald J.; Guy, R. Wayne; Hoppe, John C.; Puster, Richard L.; Rogers, R. Clayton

    2002-01-01

    This White Paper examines the current state of Hypersonic Airbreathing Propulsion at the NASA Langley Research Center and the factors influencing this area of work and its personnel. Using this knowledge, the paper explores beyond the present day and suggests future directions and strategies for the field. Broad views are first taken regarding potential missions and applications of hypersonic propulsion. Then, candidate propulsion systems that may be applicable to these missions are suggested and discussed. Design tools and experimental techniques for developing these propulsion systems are then described, and approaches for applying them in the design process are considered. In each case, current strategies are reviewed and future approaches that may improve the techniques are considered. Finally, the paper concentrates on the needs to be addressed in each of these areas to take advantage of the opportunities that lay ahead for both the NASA Langley Research Center and the Aerodynamic Aerothermodynamic, and Aeroacoustics Competency. Recommendations are then provided so that the goals set forth in the paper may be achieved.

  4. Aerothermodynamic Analysis of Stardust Sample Return Capsule with Coupled Radiation and Ablation. Revised

    NASA Technical Reports Server (NTRS)

    Gupta, Roop N.

    2000-01-01

    An aerothermodynamic analysis of the forebody aeroshell of the Stardust Sample Return Capsule is carried out by using the axisymmetric viscous shock-layer equations with and without fully coupled radiation and ablation. Formulation of the viscous shock-layer equations with shoulder radius as the length scale and implementation of the Vigneron pressure condition allow resolution of the flowfield over the shoulder. With a predominantly supersonic outflow over the shoulder, a globally iterated solution or viscous shock-layer equations can be obtained. The stagnation-point results are obtained along a specified trajectory, whereas detailed calculations along the body are provided at the peak-heating point. The equilibrium calculations with ablation injection are the focus of the present study because of the lack of a general chemical nonequilibrium analysis that accounts for both surface and flowfield effect. The equilibrium calculations also provide a simple way to conserve surface (and flowfield) elemental composition for the current small ablation injection rates, where the surface elemental composition is a mixture of freestream and ablator elements. Therefore, the coupled laminar and turbulent flow solutions with radiation and ablation are obtained by using the equilibrium flow chemistry, whereas a nonequilibrium chemistry model is used for solutions without ablation and turbulence. Various computed results are compared with those obtained by the other researchers.

  5. Adding-Point Strategy for Reduced-Order Hypersonic Aerothermodynamics Modeling Based on Fuzzy Clustering

    NASA Astrophysics Data System (ADS)

    Chen, Xin; Liu, Li; Zhou, Sida; Yue, Zhenjiang

    2016-04-01

    Reduced order models(ROMs) based on the snapshots on the CFD high-fidelity simulations have been paid great attention recently due to their capability of capturing the features of the complex geometries and flow configurations. To improve the efficiency and precision of the ROMs, it is indispensable to add extra sampling points to the initial snapshots, since the number of sampling points to achieve an adequately accurate ROM is generally unknown in prior, but a large number of initial sampling points reduces the parsimony of the ROMs. A fuzzy-clustering-based adding-point strategy is proposed and the fuzzy clustering acts an indicator of the region in which the precision of ROMs is relatively low. The proposed method is applied to construct the ROMs for the benchmark mathematical examples and a numerical example of hypersonic aerothermodynamics prediction for a typical control surface. The proposed method can achieve a 34.5% improvement on the efficiency than the estimated mean squared error prediction algorithm and shows same-level prediction accuracy.

  6. Aerothermodynamics of compressible flow past a flat plate in the slip-flow regime

    NASA Astrophysics Data System (ADS)

    Cheng, Chi-Yang; Dai, Yi; Li, Genong; Hu, Yitao; Lai, Ming-Chia

    2015-11-01

    Compressible flow past a flat plate in the slip-flow regime features a very simple geometry and flow field, but it retains the most relevant and interesting physics in high-speed rarefied gas dynamics. In the slip-flow regime, the aerothermodynamic issues, especially the recovery factors and the convection heat transfer correlation, are the focus of this presentation. We first present the detailed similarity equations, especially the transformed Maxwell's slip and jump boundary conditions, and the equations for the Chapman-Rubesin parameter as well as how we incorporate the variable gas properties and the constitutive scaling model for the Knudsen layer in the similarity equations. The similarity solutions are compared with results published by E. R. van Driest [NACA Technical Note 2597, 1952]. We point out that van Driest's solutions were computed by using no-slip and no-jump boundary conditions. The recovery factor and Nusselt number of the plate are shown as functions of the Reynolds number and the Mach number. Finally, the similarity solutions are also compared with simulations of a two-dimensional computational fluid dynamics model solving the full Navier-Stokes-Fourier equations with slip and jump boundary conditions.

  7. Blunt-Body Entry Vehicle Aerothermodynamics: Transition and Turbulence on the CEV and MSL Configurations

    NASA Technical Reports Server (NTRS)

    Hollis, Brian R.

    2010-01-01

    Recent, current, and planned NASA missions that employ blunt-body entry vehicles pose aerothermodynamic problems that challenge the state-of-the art of experimental and computational methods. The issues of boundary-layer transition and turbulent heating on the heat shield have become important in the designs of both the Mars Science Laboratory and Crew Exploration Vehicle. While considerable experience in these general areas exists, that experience is mainly derived from simple geometries; e.g. sharp-cones and flat-plates, or from lifting bodies such as the Space Shuttle Orbiter. For blunt-body vehicles, application of existing data, correlations, and comparisons is questionable because an all, or mostly, subsonic flow field is produced behind the bow shock, as compared to the supersonic (or even hypersonic) flow of other configurations. Because of the need for design and validation data for projects such as MSL and CEV, many new experimental studies have been conducted in the last decade to obtain detailed boundary-layer transition and turbulent heating data on this class of vehicle. In this paper, details of several of the test programs are reviewed. The laminar and turbulent data from these various test are shown to correlate in terms of edge-based Stanton and Reynolds number functions. Correlations are developed from the data for transition onset and turbulent heating augmentation as functions of momentum thickness Reynolds number. These correlation can be employed as engineering-level design and analysis tools.

  8. Post-Flight Assessment of Avcoat Thermal Protection System for the Exploration Flight Test-1

    NASA Technical Reports Server (NTRS)

    Bose, Deepak; Santos, Jose; Rodriguez, Erika; Mahzari, Milad; Remark, Brian; Muppidi, Suman

    2016-01-01

    On December 5, 2014 NASA conducted the first flight test of its next generation human-class Orion spacecraft. The flight was called the Exploration Flight Test -1 (EFT-1) which lasted for 4 hours and culminated into a re-entry trajectory at 9 km/s. This flight test of the 5-meter Orion Crew Module demonstrated various sub-systems including the Avcoat ablative thermal protection system (TPS) on the heat shield. The Avcoat TPS had been developed from the Apollo-era recipe with a few key modifications. The engineering for thermal sizing was supported by modeling, analysis, and ground tests in arc jet facilities. This paper will describe a postlfight analysis plan and present results from post-recovery inspections, data analysis from embedded sensors, TPS sample extraction and characterization in the laboratory. After the recovery of the vehicle, a full photographic survey and surface scans of the TPS were performed. The recovered vehicle showed physical evidence of flow disturbances, varying degrees of surface roughness, and excessive recession downstream of compression pads. The TPS recession was measured at more than 200 locations of interest on the Avcoat surface. The heat shield was then processed for sample extraction prior to TPS removal using the 7-Axis Milling machine at Marshall Space Flight Center. Around 182 rectangular TPS samples were extracted for subsequent analysis and investigation. The final paper will also present results of sample analysis. The planned investigation includes sidewall imaging, followed by image analysis to characterize TPS response by quantifying different layers in the char and pyrolysis zones. A full postmortem of the instrumentation and sensor ports will also be performed to confirm no adverse effects due to the sensors themselves. A subset of the samples will undergo structural testing and perform detailed characterization of any cracks and integrity of gore seams. Finally, the material will be characterized with layer

  9. Modeling of ultrasonic and terahertz radiations in defective tiles for condition monitoring of thermal protection systems

    NASA Astrophysics Data System (ADS)

    Kabiri Rahani, Ehsan

    Condition based monitoring of Thermal Protection Systems (TPS) is necessary for safe operations of space shuttles when quick turn-around time is desired. In the current research Terahertz radiation (T-ray) has been used to detect mechanical and heat induced damages in TPS tiles. Voids and cracks inside the foam tile are denoted as mechanical damage while property changes due to long and short term exposures of tiles to high heat are denoted as heat induced damage. Ultrasonic waves cannot detect cracks and voids inside the tile because the tile material (silica foam) has high attenuation for ultrasonic energy. Instead, electromagnetic terahertz radiation can easily penetrate into the foam material and detect the internal voids although this electromagnetic radiation finds it difficult to detect delaminations between the foam tile and the substrate plate. Thus these two technologies are complementary to each other for TPS inspection. Ultrasonic and T-ray field modeling in free and mounted tiles with different types of mechanical and thermal damages has been the focus of this research. Shortcomings and limitations of FEM method in modeling 3D problems especially at high-frequencies has been discussed and a newly developed semi-analytical technique called Distributed Point Source Method (DPSM) has been used for this purpose. A FORTRAN code called DPSM3D has been developed to model both ultrasonic and electromagnetic problems using the conventional DPSM method. This code is designed in a general form capable of modeling a variety of geometries. DPSM has been extended from ultrasonic applications to electromagnetic to model THz Gaussian beams, multilayered dielectrics and Gaussian beam-scatterer interaction problems. Since the conventional DPSM has some drawbacks, to overcome it two modification methods called G-DPSM and ESM have been proposed. The conventional DPSM in the past was only capable of solving time harmonic (frequency domain) problems. Time history was

  10. High-Temperature Structures, Adhesives, and Advanced Thermal Protection Materials for Next-Generation Aeroshell Design

    NASA Technical Reports Server (NTRS)

    Collins, Timothy J.; Congdon, William M.; Smeltzer, Stanley S.; Whitley, Karen S.

    2005-01-01

    The next generation of planetary exploration vehicles will rely heavily on robust aero-assist technologies, especially those that include aerocapture. This paper provides an overview of an ongoing development program, led by NASA Langley Research Center (LaRC) and aimed at introducing high-temperature structures, adhesives, and advanced thermal protection system (TPS) materials into the aeroshell design process. The purpose of this work is to demonstrate TPS materials that can withstand the higher heating rates of NASA's next generation planetary missions, and to validate high-temperature structures and adhesives that can reduce required TPS thickness and total aeroshell mass, thus allowing for larger science payloads. The effort described consists of parallel work in several advanced aeroshell technology areas. The areas of work include high-temperature adhesives, high-temperature composite materials, advanced ablator (TPS) materials, sub-scale demonstration test articles, and aeroshell modeling and analysis. The status of screening test results for a broad selection of available higher-temperature adhesives is presented. It appears that at least one (and perhaps a few) adhesives have working temperatures ranging from 315-400 C (600-750 F), and are suitable for TPS-to-structure bondline temperatures that are significantly above the traditional allowable of 250 C (482 F). The status of mechanical testing of advanced high-temperature composite materials is also summarized. To date, these tests indicate the potential for good material performance at temperatures of at least 600 F. Application of these materials and adhesives to aeroshell systems that incorporate advanced TPS materials may reduce aeroshell TPS mass by 15% - 30%. A brief outline is given of work scheduled for completion in 2006 that will include fabrication and testing of large panels and subscale aeroshell test articles at the Solar-Tower Test Facility located at Kirtland AFB and operated by Sandia

  11. Manufacture and engine test of advanced oxide dispersion strengthened alloy turbine vanes. [for space shuttle thermal protection

    NASA Technical Reports Server (NTRS)

    Bailey, P. G.

    1977-01-01

    Oxide-Dispersion-strengthened (ODS) Ni-Cr-Al alloy systems were exploited for turbine engine vanes which would be used for the space shuttle thermal protection system. Available commercial and developmental advanced ODS alloys were evaluated, and three were selected based on established vane property goals and manufacturing criteria. The selected alloys were evaluated in an engine test. Candidate alloys were screened by strength, thermal fatigue resistance, oxidation and sulfidation resistance. The Ni-16Cr (3 to 5)Al-ThO2 system was identified as having attractive high temperature oxidation resistance. Subsequent work also indicated exceptional sulfidation resistance for these alloys.

  12. Conformal Ablative Thermal Protection System for Planetary and Human Exploration Missions

    NASA Technical Reports Server (NTRS)

    Beck, R.; Arnold, J.; Gasch, M.; Stackpole, M.; Wercinski, R.; Venkatapathy, E.; Fan, W.; Thornton, J; Szalai, C.

    2012-01-01

    The Office of Chief Technologist (OCT), NASA has identified the need for research and technology development in part from NASAs Strategic Goal 3.3 of the NASA Strategic Plan to develop and demonstrate the critical technologies that will make NASAs exploration, science, and discovery missions more affordable and more capable. Furthermore, the Game Changing Development Program (GCDP) is a primary avenue to achieve the Agencys 2011 strategic goal to Create the innovative new space technologies for our exploration, science, and economic future. In addition, recently released NASA Space Technology Roadmaps and Priorities, by the National Research Council (NRC) of the National Academy of Sciences stresses the need for NASA to invest in the very near term in specific EDL technologies. The report points out the following challenges (Page 2-38 of the pre-publication copy released on February 1, 2012): Mass to Surface: Develop the ability to deliver more payload to the destination. NASA's future missions will require ever-greater mass delivery capability in order to place scientifically significant instrument packages on distant bodies of interest, to facilitate sample returns from bodies of interest, and to enable human exploration of planets such as Mars. As the maximum mass that can be delivered to an entry interface is fixed for a given launch system and trajectory design, the mass delivered to the surface will require reductions in spacecraft structural mass more efficient, lighter thermal protection systems more efficient lighter propulsion systems and lighter, more efficient deceleration systems. Surface Access: Increase the ability to land at a variety of planetary locales and at a variety of times. Access to specific sites can be achieved via landing at a specific location(s) or transit from a single designated landing location, but it is currently infeasible to transit long distances and through extremely rugged terrain, requiring landing close to the site of

  13. A low noise and high precision linear power supply with thermal foldback protection

    NASA Astrophysics Data System (ADS)

    Carniti, P.; Cassina, L.; Gotti, C.; Maino, M.; Pessina, G.

    2016-05-01

    A low noise and high precision linear power supply was designed for use in rare event search experiments with macrobolometers. The circuit accepts at the input a "noisy" dual supply voltage up to ±15 V and gives at the output precise, low noise, and stable voltages that can be set between ±3.75 V and ±12.5 V in eight 1.25 V steps. Particular care in circuit design, component selection, and proper filtering results in a noise spectral density of 50 nV / √{ Hz } at 1 Hz and 20 nV / √{ Hz } white when the output is set to ±5 V. This corresponds to 125 nV RMS (0.8 μV peak to peak) between 0.1 Hz and 10 Hz, and 240 nV RMS (1.6 μV peak to peak) between 0.1 Hz and 100 Hz. The power supply rejection ratio (PSRR) of the circuit is 100 dB at low frequency, and larger than 40 dB up to high frequency, thanks to a proper compensation design. Calibration allows to reach a precision in the absolute value of the output voltage of ±70 ppm, or ±350 μV at ±5 V, and to reduce thermal drifts below ±1 ppm/∘C in the expected operating range. The maximum peak output current is about 6 A from each output. An original foldback protection scheme was developed that dynamically limits the maximum output current to keep the temperature of the output transistors within their safe operating range. An add-on card based on an ARM Cortex-M3 microcontroller is devoted to the monitoring and control of all circuit functionalities and provides remote communication via CAN bus.

  14. Development of Natural Flaw Samples for Evaluating Nondestructive Testing Methods for Foam Thermal Protection Systems

    NASA Technical Reports Server (NTRS)

    Workman, Gary L.; Davis, Jason; Farrington, Seth; Walker, James

    2007-01-01

    Low density polyurethane foam has been an important insulation material for space launch vehicles for several decades. The potential for damage from foam breaking away from the NASA External Tank was not realized until the foam impacts on the Columbia Orbiter vehicle caused damage to its Leading Edge thermal protection systems (TPS). Development of improved inspection techniques on the foam TPS is necessary to prevent similar occurrences in the future. Foamed panels with drilled holes for volumetric flaws and Teflon inserts to simulate debonded conditions have been used to evaluate and calibrate nondestructive testing (NDT) methods. Unfortunately the symmetric edges and dissimilar materials used in the preparation of these simulated flaws provide an artificially large signal while very little signal is generated from the actual defects themselves. In other words, the same signal are not generated from the artificial defects in the foam test panels as produced when inspecting natural defect in the ET foam TPS. A project to create more realistic voids similar to what actually occurs during manufacturing operations was began in order to improve detection of critical voids during inspections. This presentation describes approaches taken to create more natural voids in foam TPS in order to provide a more realistic evaluation of what the NDT methods can detect. These flaw creation techniques were developed with both sprayed foam and poured foam used for insulation on the External Tank. Test panels with simulated defects have been used to evaluate NDT methods for the inspection of the External Tank. A comparison of images between natural flaws and machined flaws generated from backscatter x-ray radiography, x-ray laminography, terahertz imaging and millimeter wave imaging show significant differences in identifying defect regions.

  15. Acousto-optic signature analysis for inspection of the orbiter thermal protection tile bonds

    NASA Technical Reports Server (NTRS)

    Rodriguez, Julio G.; Tow, D. M.; Barna, B. A.

    1990-01-01

    The goal of this research is to develop a viable NDE technique for the inspection of orbiter thermal protection system (TPS) tile bonds. Phase 2, discussed here, concentrated on developing an empirical understanding of the bonded and unbonded vibration signatures of acreage tiles. Controlled experiments in the laboratory have provided useful information on the dynamic response of TPS tiles. It has been shown that several signatures are common to all the pedigree tiles. This degree of consistency in the tile-SIP (strain isolation pad) dynamic response proves that an unbond can be detected for a known tile and establish the basis for extending the analysis capability to arbitrary tiles for which there are no historical data. The field tests of the noncontacting laser acoustic sensor system, conducted at the Kennedy Space Center (KSC), investigated the vibrational environment of the Orbiter Processing Facility (OPF) and its effect on the measurement and analysis techniques being developed. The data collected showed that for orbiter locations, such as the body flap and elevon, the data analysis scheme, and/or the sensor, will require modification to accommodate the ambient motion. Several methods were identified for accomplishing this, and a solution is seen as readily achievable. It was established that the tile response was similar to that observed in the laboratory. Of most importance, however, is that the field environment will not affect the physics of the dynamic response that is related to bond condition. All of this information is fundamental to any future design and development of a prototype system.

  16. Simultaneously Harvesting Thermal and Mechanical Energies based on Flexible Hybrid Nanogenerator for Self-Powered Cathodic Protection.

    PubMed

    Zhang, Hulin; Zhang, Shangjie; Yao, Guang; Huang, Zhenlong; Xie, Yuhang; Su, Yuanjie; Yang, Weiqing; Zheng, Chunhua; Lin, Yuan

    2015-12-30

    Metal corrosion occurs anytime and anywhere in nature and the corrosion prevention has a great significance everywhere in national economic development and daily life. Here, we demonstrate a flexible hybrid nanogenerator (NG) that is capable of simultaneously or individually harvesting ambient thermal and mechanical energies and used for a self-powered cathodic protection (CP) system without using an external power source. Because of its double peculiarities of both pyroelectric and piezoelectric properties, a polarized poly(vinylidene fluoride) (PVDF) film-based NG was constructed to scavenge both thermal and mechanical energies. As a supplementary, a triboelectric NG was constructed below the pyro/piezoelectric NG to grab ambient mechanical energy. The output power of the fabricated hybrid NG can be directly used to protect the metal surface from the chemical corrosion. Our results not only verify the feasibility of self-powered CP-based NGs, but also expand potential self-powered applications. PMID:26669205

  17. Development of acceptance criteria for batches of silane primer for external tank thermal protection system bonding applications

    NASA Technical Reports Server (NTRS)

    Mikes, F.

    1984-01-01

    Silane primers for use as thermal protection on external tanks were subjected to various analytic techniques to determine the most effective testing method for silane lot evaluation. The analytic methods included high performance liquid chromatography, gas chromatography, thermogravimetry (TGA), and fourier transform infrared spectroscopy (FTIR). It is suggested that FTIR be used as the method for silane lot evaluation. Chromatograms, TGA profiles, bar graphs showing IR absorbances, and FTIR spectra are presented.

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

    NASA Technical Reports Server (NTRS)

    Sepka, Steven Andrew; Zarchi, Kerry Agnes; Maddock, Robert W.; Samareh, Jamshid A.

    2011-01-01

    Mass Estimating Relationships (MERs) have been developed for use in the Program to Optimize Simulated Trajectories II (POST2) as part of NASA's multi-mission Earth Entry Vehicle (MMEEV) concept. MERs have been developed for the thermal protection systems of PICA and of Carbon Phenolic atop Advanced Carbon-Carbon on the forebody and for SIRCA and Acusil II on the backshell. How these MERs were developed, the resulting equations, model limitations, and model accuracy are discussed herein.

  19. Development Of FIAT-Based Thermal Protection System Mass Estimating Relationships For NASA's Multi-Mission Earth Entry Concept

    NASA Technical Reports Server (NTRS)

    Sepka, Steven; Trumble, Kerry A.; Maddock, Robert W.; Samareh, Jamshid

    2012-01-01

    Mass Estimating Relationships (MERs) have been developed for use in the Program to Optimize Simulated Trajectories II (POST2) as part of NASA's multi-mission Earth Entry Vehicle (MMEEV) concept. MERs have been developed for the thermal protection systems of PICA and of Carbon Phenolic atop Advanced Carbon-Carbon on the forebody and for SIRCA and Acusil II on the backshell. How these MERs were developed, the resulting equations, model limitations, and model accuracy are discussed herein.

  20. OAST Space Theme Workshop. Volume 3: Working group summary. 9: Aerothermodynamics (M-3). A: Statement. B: Technology needs (form 1). C. Priority assessment (form 2). D. Additional assessments

    NASA Technical Reports Server (NTRS)

    1976-01-01

    Twelve aerothermodynamic space technology needs were identified to reduce the design uncertainties in aerodynamic heating and forces experienced by heavy lift launch vehicles, orbit transfer vehicles, and advanced single stage to orbit vehicles for the space transportation system, and for probes, planetary surface landers, and sample return vehicles for solar system exploration vehicles. Research and technology needs identified include: (1) increasing the fluid dynamics capability by at least two orders of magnitude by developing an advanced computer processor for the solution of fluid dynamic problems with improved software; (2) predicting multi-engine base flow fields for launch vehicles; and (3) developing methods to conserve energy in aerothermodynamic ground test facilities.

  1. Performance of full size metallic and RSI thermal protection systems in a Mach 7 environment. [Reusable Surface Insulation

    NASA Technical Reports Server (NTRS)

    Bohon, H. L.; Sawyer, J. W.; Hunt, L. R.; Weinstein, I.

    1975-01-01

    The integrity and reusability of three flight-weight metallic and RSI thermal protection systems, designed for the Shuttle entry environment, have been demonstrated. Each model successfully survived over 23 entry thermal cycles without serious degradation. The metallic systems were more tolerant of the hostile environment and provided a higher degree of reusability than did the RSI. Thermal expansion slip joints of the metallic TPS successfully prevented hot gas ingress to the substructure. The RSI demonstrated high damage tolerance and field repairs increased its reusability. Heat-transfer tests to further assess RSI gap heating indicate that stacked tile orientations may impose a penalty on tile thickness. Parameters influencing RSI impingement heating were determined, and the heating data were correlated.

  2. Today`s thermal imaging systems: Background and applications for civilian law enforcement and military force protection

    SciTech Connect

    Bisbee, T.L.; Pritchard, D.A.

    1997-10-01

    Thermal (infrared) imagers can solve many security assessment problems associated with the protection of high-value assets at military bases, secure installations, or commercial facilities. Thermal imagers can provide surveillance video from security areas or perimeters both day and night without expensive security lighting. In the past, thermal imagers required cryogenic cooling to operate. The high cost and maintenance requirements restricted their use. However, recent developments in reliable, linear drive cryogenic coolers and uncooled infrared imagers have dramatically reduced system cost. These technology developments are resulting in greater accessibility and practicality for military as well as civilian security and force protection applications. This paper discusses recent advances in thermal imaging technology including uncooled and cryo-cooled. Applications of Forward Looking InfraRed (FLIR) systems are also discussed, including integration with a high-speed pan/tilt mount and remote control, video frame storage and recall, low-cost vehicle-mounted systems, and hand-held devices. Other facility installation topics will be discussed, such as site layout, assessment ranges, imager positioning, fields-of-view, sensor and alarm reporting systems, and communications links.

  3. Remote sensing of thermal radiation from an aircraft - An analysis and evaluation of crop-freeze protection methods

    NASA Technical Reports Server (NTRS)

    Sutherland, R. A.; Hannah, H. E.; Cook, A. F.; Martsolf, J. D.

    1981-01-01

    Thermal images from an aircraft-mounted scanner are used to evaluate the effectiveness of crop-freeze protection devices. Data from flights made while using fuel oil heaters, a wind machine and an undercanopy irrigation system are compared. Results show that the overall protection provided by irrigation (at approximately 2 C) is comparable to the less energy-efficient heater-wind machine combination. Protection provided by the wind machine alone (at approximately 1 C) was found to decrease linearly with distance from the machine by approximately 1 C/100 m. The flights were made over a 1.5 hectare citrus grove at an altitude of 450 m with an 8-14 micron detector. General meteorological conditions during the experiments, conducted during the nighttime, were cold (at approximately -6 C) and calm with clear skies.

  4. The influence of the free space environment on the superlight-weight thermal protection system: conception, methods, and risk analysis

    NASA Astrophysics Data System (ADS)

    Yatsenko, Vitaliy; Falchenko, Iurii; Fedorchuk, Viktor; Petrushynets, Lidiia

    2016-07-01

    This report focuses on the results of the EU project "Superlight-weight thermal protection system for space application (LIGHT-TPS)". The bottom line is an analysis of influence of the free space environment on the superlight-weight thermal protection system (TPS). This report focuses on new methods that based on the following models: synergetic, physical, and computational. This report concentrates on four approaches. The first concerns the synergetic approach. The synergetic approach to the solution of problems of self-controlled synthesis of structures and creation of self-organizing technologies is considered in connection with the super-problem of creation of materials with new functional properties. Synergetics methods and mathematical design are considered according to actual problems of material science. The second approach describes how the optimization methods can be used to determine material microstructures with optimized or targeted properties. This technique enables one to find unexpected microstructures with exotic behavior (e.g., negative thermal expansion coefficients). The third approach concerns the dynamic probabilistic risk analysis of TPS l elements with complex characterizations for damages using a physical model of TPS system and a predictable level of ionizing radiation and space weather. Focusing is given mainly on the TPS model, mathematical models for dynamic probabilistic risk assessment and software for the modeling and prediction of the influence of the free space environment. The probabilistic risk assessment method for TPS is presented considering some deterministic and stochastic factors. The last approach concerns results of experimental research of the temperature distribution on the surface of the honeycomb sandwich panel size 150 x 150 x 20 mm at the diffusion welding in vacuum are considered. An equipment, which provides alignment of temperature fields in a product for the formation of equal strength of welded joints is

  5. Computational Aerothermodynamic Assessment of Space Shuttle Orbiter Tile Damage: Open Cavities

    NASA Technical Reports Server (NTRS)

    Pulsonetti, Maria; Wood, William

    2005-01-01

    Computational aerothermodynamic simulations of Orbiter windside tile damage in flight were performed in support of the Space Shuttle Return-to-Flight effort. The simulations were performed for both hypervelocity flight and low-enthalpy wind tunnel conditions and contributed to the Return-to-Flight program by providing information to support a variety of damage scenario analyses. Computations at flight conditions were performed at or very near the peak heating trajectory point for multiple damage scenarios involving damage windside acreage reaction cured glass (RCG) coated silica tile(s). The cavities formed by the missing tile examined in this study were relatively short leading to flow features which indicated open cavity behavior. Results of the computations indicated elevated heating bump factor levels predicted for flight over the predictions for wind tunnel conditions. The peak heating bump factors, defined as the local heating to a reference value upstream of the cavity, on the cavity floor for flight simulation were 67% larger than the peak wind tunnel simulation value. On the downstream face of the cavity the flight simulation values were 60% larger than the wind tunnel simulation values. On the outer mold line (OML) downstream of the cavity, the flight values are about 20% larger than the wind tunnel simulation values. The higher heating bump factors observed in the flight simulations were due to the larger driving potential in terms of energy entering the cavity for the flight simulations. This is evidenced by the larger rate of increase in the total enthalpy through the boundary layer prior to the cavity for the flight simulation.

  6. Impact of clocking on the aero-thermodynamics of a second stator tested in a one and a half stage HP turbine

    NASA Astrophysics Data System (ADS)

    Billiard, N.; Paniagua, Guillermo; Dénos, R.

    2008-06-01

    This paper focuses on the experimental investigation of the time-averaged and time-accurate aero-thermodynamics of a second stator tested in a 1.5 stage high-pressure turbine. The effect of clocking on aerodynamic and heat transfer are investigated. Tests are performed under engine representative conditions in the VKI compression tube CT3. The test program includes four different clocking positions, i.e. relative pitch-wise positions between the first and the second stator. Probes located upstream and downstream of the second stator provide the thermodynamic conditions of the flow field. On the second stator airfoil, measurements are taken around the blade profile at 15, 50 and 85% span with pressure sensors and thin-film gauges. Both time-averaged and time-resolved aspects of the flow field are addressed. Regarding the time-averaged results, clocking effects are mainly observed within the leading edge region of the second stator, the largest effects being observed at 15% span. The surface static pressure distribution is changed locally, hence affecting the overall airfoil performance. For one clocking position, the thermal load of the airfoil is noticeably reduced. Pressure fluctuations are attributed to the passage of the upstream transonic rotor and its associated pressure gradients. The pattern of these fluctuations changes noticeably as a function of clocking. The time-resolved variations of heat flux and static pressure are analyzed together showing that the major effect is due to a potential interaction. The time-resolved pressure distribution integrated along the second stator surface yields the unsteady forces on the vane. The magnitude of the unsteady force is very dependent on the clocking position.

  7. Integrated Sensing and Material Damage Identification in Metallic and Ceramic Thermal Protection Systems Using Vibration and Wave Propagation Data

    SciTech Connect

    Sundararaman, S.; White, J.; Jiang, H.; Adams, D.; Jata, K.

    2006-03-06

    Global thermal and impact material damage mechanisms in metallic and ceramic thermal protection systems are detected, located, and quantified using four complementary methods for sensing and data interrogation. First, spatial-temporal beamforming algorithms are used to process active elastic waves measured from remote sensor arrays in two different equilibrium positions of a gamma Ti-Al sheet to localize simulated thermal damage. Damage is located even when it is behind the sensor array and on the edge of the panel; results are shown to be dependent on the equilibrium position considered. Second, an active virtual force method is implemented in a honeycomb Al-Al sandwich panel instrumented with a distributed piezo sensor and actuator array to identify impact and thermal damage using frequency response inversion. Damage is quantified and is similarly diagnosed regardless of the excitation location. Third, passive acoustic transmission measurements through a homogeneous baffled Al panel subject to launch-type sound pressure variations are used to detect and locate material damage. The frequency range with highest transmission is shown to be optimal for damage detection. Fourth, thermal damage in a wrapped ceramic tile with a mock strain isolation pad is identified using active propagating waves. Remote actuation and sensing on the bulkhead and the tile backside are shown to be sufficient for detection even when variability is present in the data.

  8. Primer for identifying cold-water refuges to protect and restore thermal diversity in riverine landscapes

    EPA Science Inventory

    EPA recently released a primer that provides guidance to Region 10 Tribes, States, and local watershed community groups to support the identification, protection, and restoration of critical cold water refuges for the protection of salmonids. This primer will assist these entiti...

  9. 75 FR 13 - Alternate Fracture Toughness Requirements for Protection Against Pressurized Thermal Shock Events

    Federal Register 2010, 2011, 2012, 2013, 2014

    2010-01-04

    ... From the Federal Register Online via the Government Publishing Office NUCLEAR REGULATORY... Pressurized Thermal Shock Events AGENCY: Nuclear Regulatory Commission. ACTION: Final rule. SUMMARY: The Nuclear Regulatory Commission (NRC) is amending its regulations to provide alternate fracture...

  10. SO2 protects the amino nitrogen metabolism of Saccharomyces cerevisiae under thermal stress

    PubMed Central

    Ancín‐Azpilicueta, Carmen; Barriuso‐Esteban, Blanca; Nieto‐Rojo, Rodrigo; Aristizábal‐López, Nerea

    2012-01-01

    Summary Thermal stress conditions during alcoholic fermentation modify yeasts' plasma membrane since they become more hyperfluid, which results in a loss of bilayer integrity. In this study, the influence of elevated temperatures on nitrogen metabolism of a Saccharomyces cerevisiae strain was studied, as well as the effect of different concentrations of SO2 on nitrogen metabolism under thermal stress conditions. The results obtained revealed that amino nitrogen consumption was lower in the fermentation sample subjected to thermal stress than in the control, and differences in amino acid consumption preferences were also detected, especially at the beginning of the fermentation. Under thermal stress conditions, among the three doses of SO2 studied (0, 35, 70 mg l−1 SO2), the highest dose was observed to favour amino acid utilization during the fermentative process, whereas sugar consumption presented higher rates at medium doses. PMID:22452834

  11. Experimental study of unsteady aerothermodynamic phenomena on shock-tube wall using fast-response temperature-sensitive paints

    NASA Astrophysics Data System (ADS)

    Ozawa, Hiroshi

    2016-04-01

    This paper describes an experimental study that used a fast-response temperature-sensitive paint (TSP) to investigate the unsteady aerothermodynamic phenomena occurring on a shock-tube wall. To understand these phenomena in detail, a fast-response TSP with high temperature sensitivity developed for transient temperature measurement was applied to the wall. The shock-tube experiment was carried out under the over-tailored condition, with a pressure ratio of 110 for test gases of air in driver/driven tubes. The following aspects were clarified using the TSP: (a) the TSP could be used to visualize the unsteady aerothermodynamic phenomena and estimate the quantitative heat flux on the shock-tube wall; (b) an x-t diagram based on the TSP response showed shock-tube wall characteristics that included the incident/reflected shocks, laminar-to-turbulent boundary-layer transition, streaks in the turbulent boundary layer, reflected shock/turbulent boundary layer interaction, and waves reflected from a contact surface; (c) the TSP graphically showed that a transition front from the plate's leading edge and turbulent spots moved with 80% of the free-stream velocity behind the incident shock. In addition, the TSP could track the growth of the turbulent spots on the wall.

  12. Aerothermal performance and structural integrity of a Rene 41 thermal protection system at Mach 6.6

    NASA Technical Reports Server (NTRS)

    Deveikis, W. D.; Miserentino, R.; Weinstein, I.; Shideler, J. L.

    1975-01-01

    A flightweight panel based on a metallic thermal-protection-system concept for hypersonic and reentry vehicles was subjected repeatedly to thermal cycling by quartz-lamp radiant heating using a thermal history representative of a reentry heat pulse and to aerodynamic heating at heating rates required to sustain a surface temperature of 1089 K (1960 R). The panel consisted of a corrugated heat shield and support members of 0.05-cm (0.02-in.) thick Rene 41 of riveted construction and 5.08-cm (2-in.) thick silica fibrous insulation packages covered by Rene 41 foil and inconel screening. All tests were conducted in the Langley 8-foot high-temperature structures tunnel with the heat shield corrugations alined in the stream direction. The panel sustained 5.33 hr of intermittent radiant heating and 6.5 min of intermittent aerodynamic heating of up to 1-min duration for differential pressures up to 6.2 kPa (0.9 psi) with no apparent degradation of thermal or structural integrity, as indicated by temperature distributions and results from load deflection tests and vibration surveys of natural frequencies.

  13. Development and Design Application of Rigidized Surface Insulation Thermal Protection Systems, Volume 1. [for the space shuttle

    NASA Technical Reports Server (NTRS)

    1972-01-01

    Materials and design technology of the all-silica LI-900 rigid surface insulation (RSI) thermal protection system (TPS) concept for the shuttle spacecraft is presented. All results of contract development efforts are documented. Engineering design and analysis of RSI strain arrestor plate material selections, sizing, and weight studies are reported. A shuttle prototype test panel was designed, analyzed, fabricated, and delivered. Thermophysical and mechanical properties of LI-900 were experimentally established and reported. Environmental tests, including simulations of shuttle loads represented by thermal response, turbulent duct, convective cycling, and chemical tolerance tests are described and results reported. Descriptions of material test samples and panels fabricated for testing are included. Descriptions of analytical sizing and design procedures are presented in a manner formulated to allow competent engineering organizations to perform rational design studies. Results of parametric studies involving material and system variables are reported. Material performance and design data are also delineated.

  14. Modeling thermal insulation of firefighting protective clothing embedded with phase change material

    NASA Astrophysics Data System (ADS)

    Hu, Yin; Huang, Dongmei; Qi, Zhengkun; He, Song; Yang, Hui; Zhang, Heping

    2013-04-01

    Experiments and research on heat transport through firefighting protective clothing when exposed to high temperature or intensive radiation are significant. Phase change material (PCM) takes energy when changes from solid to liquid thus reducing heat transmission. A numerical simulation of heat protection of the firefighting protective clothing embedded with PCM was studied. We focused on the temperature variation by comparing different thicknesses and position conditions of PCM combined in the clothing, as well as the melting state of PCM and human irreversible burns through a simplified one-dimensional model. The results showed it was superior to place PCM between water and proof layer and inner layer, in addition, greater thickness increased protection time while might adding extra burden to the firefighter.

  15. A design assessment of multiwall, metallic stand-off, and RSI reusable thermal protection systems including space shuttle application

    NASA Technical Reports Server (NTRS)

    Jackson, L. R.; Dixon, S. C.

    1980-01-01

    The design and assessment of reusable surface insulation (RSI), metallic stand off and multiwall thermal protection systems (TPS) is discussed. Multiwall TPS is described in some detail, and analyses useful for design of multiwall are included. Results indicate that multiwall has the potential to satisfy the TPS design goals better than the other systems. The total mass of the stand-off TPS and of the metallic systems require less primary structure mass than the RSI system, since the nonbuckling skin criteria required for RSI may be removed. Continued development of multiwall TPS is required to verify its potential and to provide the necessary data base for design.

  16. Intelligent process development of foam molding for the Thermal Protection System (TPS) of the space shuttle external tank

    NASA Technical Reports Server (NTRS)

    Bharwani, S. S.; Walls, J. T.; Jackson, M. E.

    1987-01-01

    A knowledge based system to assist process engineers in evaluating the processability and moldability of poly-isocyanurate (PIR) formulations for the thermal protection system of the Space Shuttle external tank (ET) is discussed. The Reaction Injection Molding- Process Development Advisor (RIM-PDA) is a coupled system which takes advantage of both symbolic and numeric processing techniques. This system will aid the process engineer in identifying a startup set of mold schedules and in refining the mold schedules to remedy specific process problems diagnosed by the system.

  17. Thermal insulating barrier and neutron shield providing integrated protection for a nuclear reactor vessel

    DOEpatents

    Schreiber, Roger B.; Fero, Arnold H.; Sejvar, James

    1997-01-01

    The reactor vessel of a nuclear reactor installation which is suspended from the cold leg nozzles in a reactor cavity is provided with a lower thermal insulating barrier spaced from the reactor vessel to form a chamber which can be flooded with cooling water through passive valving to directly cool the reactor vessel in the event of a severe accident. The passive valving also includes bistable vents at the upper end of the thermal insulating barrier for releasing steam. A removable, modular neutron shield extending around the upper end of the reactor cavity below the nozzles forms with the upwardly and outwardly tapered transition on the outer surface of the reactor vessel, a labyrinthine channel which reduces neutron streaming while providing a passage for the escape of steam during a severe accident, and for the cooling air which is circulated along the reactor cavity walls outside the thermal insulating barrier during normal operation of the reactor.

  18. Thermal insulating barrier and neutron shield providing integrated protection for a nuclear reactor vessel

    DOEpatents

    Schreiber, R.B.; Fero, A.H.; Sejvar, J.

    1997-12-16

    The reactor vessel of a nuclear reactor installation which is suspended from the cold leg nozzles in a reactor cavity is provided with a lower thermal insulating barrier spaced from the reactor vessel to form a chamber which can be flooded with cooling water through passive valving to directly cool the reactor vessel in the event of a severe accident. The passive valving also includes bistable vents at the upper end of the thermal insulating barrier for releasing steam. A removable, modular neutron shield extending around the upper end of the reactor cavity below the nozzles forms with the upwardly and outwardly tapered transition on the outer surface of the reactor vessel, a labyrinthine channel which reduces neutron streaming while providing a passage for the escape of steam during a severe accident, and for the cooling air which is circulated along the reactor cavity walls outside the thermal insulating barrier during normal operation of the reactor. 8 figs.

  19. Multi-functional materials by powder processing for a thermal protection system with self-cooling capability: Perspirable skin

    NASA Astrophysics Data System (ADS)

    Sun, Li

    Aerodynamic heating generated by the friction between the atmosphere and the space vehicle's surface at reentry can enhance the temperature on the surface as high as 1700°C. A Thermal Protection System (TPS) is needed to inhibit the heat entering into the vehicle. Presently, the completely passive thermal protection is used for TPS. The thermal ablation/erosion and oxidization reaction of the current TPS is the major threat to the safety of the space vehicle. Therefore, a new design for TPS with actively self-cooling capability was proposed by bio-mimicking the perspiration of the human body, henceforth called Perspirable skin. The design of Perspirable Skin consists of core material shrink-fitted into a skin panel such as Reinforced Carbon-Carbon (RCC) Composite. The core material contains a very small Coefficient of Thermal Expansion (CTE) compared to the panel material. As temperature increases, the gap between the core and the skin are produced due to the CTE difference. Compressed gas on board the space vehicle will blow out from the gap once the surface temperature reaches a critical value. The cold gas flows over the surface and mixes with the atmospheric air to compensate for the frictional heat. With Perspirable Skin, the highest temperature on the surface is expected to decrease, and we assumed it to be around half of the present temperature. This dissertation focuses on the selection of the core materials and their manufacturing by powder processing. Based on a series of experiments, several results were obtained: (1) the effect of powder mixing on the compaction capability and sintering capability was determined; (2) a flat 3-layered Al 2O3/ZrO2 Functionally Graded Material (FGM) without cracks was fabricated; (3) the factors contributing to the cracks in the multi-layered materials were investigated; (4) an isotropic negative thermal expansion material, ZrW2O8, as well as its composites with ZrO2 were processed by in-situ reaction of WO3 and ZrO2; (5

  20. Thermal Stress Induced Aggregation of Aquaporin 0 (AQP0) and Protection by α-Crystallin via Its Chaperone Function

    PubMed Central

    Swamy-Mruthinti, Satyanarayana; Srinivas, Volety; Hansen, John E.; Rao, Ch Mohan

    2013-01-01

    Aquaporin 0 (AQP0) formerly known as membrane intrinsic protein (MIP), is expressed exclusively in the lens during terminal differentiation of fiber cells. AQP0 plays an important role not only in the regulation of water content but also in cell-to-cell adhesion of the lens fiber cells. We have investigated the thermal stress-induced structural alterations of detergent (octyl glucoside)-solubilized calf lens AQP0. The results show an increase in the amount of AQP0 that aggregated as the temperature increased from 40°C to 65°C. α-Crystallin, molecular chaperone abundantly present in the eye lens, completely prevented the AQP0 aggregation at a 1∶1 (weight/weight) ratio. Since α-crystallin consists of two gene products namely αA- and αB-crystallins, we have tested the recombinant proteins on their ability to prevent thermal-stress induced AQP0 aggregation. In contrast to the general observation made with other target proteins, αA-crystallin exhibited better chaperone-like activity towards AQP0 compared to αB-crystallin. Neither post-translational modifications (glycation) nor C-terminus truncation of AQP0 have any appreciable effect on its thermal aggregation properties. α-Crystallin offers similar protection against thermal aggregation as in the case of the unmodified AQP0, suggesting that αcrystallin may bind to either intracellular loops or other residues of AQP0 that become exposed during thermal stress. Far-UV circular dichroism studies indicated a loss of αhelical structures when AQP0 was subjected to temperatures above 45°C, and the presence of α-crystallin stabilized these secondary structures. We report here, for the first time, that α-crystallin protects AQP0 from thermal aggregation. Since stress-induced structural perturbations of AQP0 may affect the integrity of the lens, presence of the molecular chaperone, α-crystallin (particularly αA-crystallin) in close proximity to the lens membrane is physiologically relevant. PMID:24312215