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  • 1
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    Space Shuttle Orbiter carbon-carbon oxidation performance (1991)
    In:  Other Sources
    Details
    Publication Date: 2011-08-24
    Description: Reusable, oxidation protected reinforced carbon carbon (RCC) has been successfully flown on forty Shuttle Orbiter flights. Thermal testing of the silicon carbide coated, reinforced carbon-carbon to determine its oxidation characteristics has been performed in both radiant and convective (plasma arc jet) heating test facilities. Subsurface oxidation of the RCC substrate as a result of oxygen penetrating micro cracks (fizzures) in the coating was characterized as a function of temperature and pressure for both convective and radiant environments. High temperature testing was performed to establish coating recession for over-temperature flight conditions experienced on abort trajectories. Suggested methods for using these test data to establish multi-mission reuse (i.e., mission life) and single mission limits are presented.
    Keywords: COMPOSITE MATERIALS
    Type: In: Damage and oxidation protection in high temperature composites. Vol. 1; Proceedings of the Symposium, 112th ASME Winter Annual Meeting, Atlanta, GA, Dec. 1-6, 1991 (A93-53937 23-24); p. 47-64.
    Format: text
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  • 2
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    Thermal protection system trade studies on ablators and reuseable surface insulation (1972)
    In:  CASI
    Details
    Publication Date: 2005-11-30
    Description: Analytical trade studies are presented that consider passive TPS configurations using the following material categories: (1) reuseable surface insulation - surface-coated rigidized ceramic fiber; (2) low density charring ablators; and (3) carbon-carbon and high density ablators for leading edge areas. Emphasized are effects on TPS weight by variations in entry trajectories and material thermal characteristics.
    Keywords: THERMODYNAMICS AND COMBUSTION
    Type: NASA Space Shuttle Technol. Conf.; p 303-334
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  • 3
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    Aeroassist flight experiment heating-rate sensitivity study (1991)
    In:  Other Sources
    Details
    Publication Date: 2011-08-24
    Keywords: AERODYNAMICS
    Type: Journal of Thermophysics and Heat Transfer (ISSN 0887-8722); 5; 456-462
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  • 4
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    Assessing the orbiter thermal environment using flight data (1984)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Keywords: SPACE TRANSPORTATION
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4650); 21; 534-541
    Format: text
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  • 5
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    An analytical and experimental study for surface heat flux determination (1977)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Description: A numerical method by which data from a single embedded thermocouple can be used to predict the transient thermal environment for both high- and low-conductivity materials is described. The results of an investigation performed to verify the method clearly demonstrate that accurate transient surface heating conditions can be obtained from a thermocouple 1.016 cm from the surface in a low-conductivity material. Space Shuttle Orbiter thermal protection system materials having temperature- and pressure-dependent properties and typical Orbiter entry heating conditions were used to verify the accuracy of the analytical procedure. Analytically generated, as well as experimental, data were used to compare predicted and measured surface temperatures.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Journal of Spacecraft and Rockets; 14; Oct. 197
    Format: text
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  • 6
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    Nonlinear least squares - An aid to thermal property determination. (1973)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Nonlinear least squares techniques can be used to determine effective thermal conductivity values from experimental data. Comparisons between measured and predicted conductivity values indicate that the analytically determined values can be used with confidence in performing thermal protection system analyses. A study was performed to compare the relative efficiencies of different minimizing techniques; the method of Peckham was the most efficient.
    Keywords: THERMODYNAMICS AND COMBUSTION
    Type: AIAA Journal; 11; May 1973
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  • 7
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    An implicit-iterative solution of the heat conduction equation with a radiation boundary condition (1977)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Description: For the problem of predicting one-dimensional heat transfer between conducting and radiating mediums by an implicit finite difference method, four different formulations were used to approximate the surface radiation boundary condition while retaining an implicit formulation for the interior temperature nodes. These formulations are an explicit boundary condition, a linearized boundary condition, an iterative boundary condition, and a semi-iterative boundary method. The results of these methods in predicting surface temperature on the space shuttle orbiter thermal protection system model under a variety of heating rates were compared. The iterative technique caused the surface temperature to be bounded at each step. While the linearized and explicit methods were generally more efficient, the iterative and semi-iterative techniques provided a realistic surface temperature response without requiring step size control techniques.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: International Journal for Numerical Methods in Engineering; 11; 10, 1; 1977
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  • 8
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    Orbiter thermal protection system (1985)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: 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.
    Keywords: SPACE TRANSPORTATION
    Type: Space Shuttle Tech. Conf., Pt. 2; p 1062-1081
    Format: application/pdf
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  • 9
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    Aerodynamic heating on AFE due to nonequilibrium flow with variable entropy at boundary layer edge (1991)
    In:  CASI
    Details
    Publication Date: 2013-08-29
    Description: A method of predicting the aerobrake aerothermodynamic environment on the NASA Aeroassist Flight Experiment (AFE) vehicle is described. Results of a three dimensional inviscid nonequilibrium solution are used as input to an axisymmetric nonequilibrium boundary layer program to predict AFE convective heating rates. Inviscid flow field properties are obtained from the Euler option of the Viscous Reacting Flow (VRFLO) code at the boundary layer edge. Heating rates on the AFE surface are generated with the Boundary Layer Integral Matrix Procedure (BLIMP) code for a partially catalytic surface composed of Reusable Surface Insulation (RSI) times. The 1864 kg AFE will fly an aerobraking trajectory, simulating return from geosynchronous Earth orbit, with a 75 km perigee and a 10 km/sec entry velocity. Results of this analysis will provide principal investigators and thermal analysts with aeroheating environments to perform experiment and thermal protection system design.
    Keywords: AERODYNAMICS
    Type: ESA, Aerothermodynamics for Space Vehicles; p 371-376
    Format: text
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  • 10
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    TPS design for aerobraking at Earth and Mars (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was made to determine the feasibility of using an aerobrake system for manned and unmanned missions to Mars, and to Earth from Mars and lunar orbits. A preliminary thermal protection system (TPS) was examined for five unmanned small nose radius, straight bi-conic vehicles and a scaled up Aeroassist Flight Experiment (AFE) vehicle aerocapturing at Mars. Analyses were also conducted for the scaled up AFE and an unmanned Sample Return Cannister (SRC) returning from Mars and aerocapturing into Earth orbit. Also analyzed were three different classes of lunar transfer vehicles (LTV's): an expendable scaled up modified Apollo Command Module (CM), a raked cone (modified AFT), and three large nose radius domed cylinders. The LTV's would be used to transport personnel and supplies between Earth and the moon in order to establish a manned base on the lunar surface. The TPS for all vehicles analyzed is shown to have an advantage over an all-propulsive velocity reduction for orbit insertion. Results indicate that TPS weight penalties of less than 28 percent can be achieved using current material technology, and slightly less than the most favorable LTV using advanced material technology.
    Keywords: THERMODYNAMICS AND STATISTICAL PHYSICS
    Type: NASA-TM-104739 , S-645 , NAS 1.15:104739
    Format: application/pdf
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