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  • 1
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    A historical perspective of the YF-12A thermal loads and structures program (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Around 1970, the Y-F-12A loads and structures efforts focused on numerous technological issues that needed defining with regard to aircraft that incorporate hot structures in the design. Laboratory structural heating test technology with infrared systems was largely created during this program. The program demonstrated the ability to duplicate the complex flight temperatures of an advanced supersonic airplane in a ground-based laboratory. The ability to heat and load an advanced operational aircraft in a laboratory at high temperatures and return it to flight status without adverse effects was demonstrated. The technology associated with measuring loads with strain gages on a hot structure was demonstrated with a thermal calibration concept. The results demonstrated that the thermal stresses were significant although the airplane was designed to reduce thermal stresses. Considerable modeling detail was required to predict the heat transfer and the corresponding structural characteristics. The overall YF-12A research effort was particularly productive, and a great deal of flight, laboratory, test and computational data were produced and cross-correlated.
    Keywords: Aeronautics (General)
    Type: NASA-TM-104317 , NAS 1.15:104317 , H-2079
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Effect of internal convection and internal radiation on the structural temperatures of Space Shuttle Orbiter (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Structural performance and resizing of the finite-element thermal analysis computer program was used in the reentry heat transfer analysis of the space shuttle orbiter. One midfuselage cross section and one midspan wing segment were selected to study the effects of internal convection and internal radiation on the structural temperatures. The effect of internal convection was found to be more prominent than that of internal radiation in the orbiter thermal analysis. Without these two effects, the calculated structural temperatures at certain stations could be as much as 45 to 90 percent higher than the measured values. By considering internal convection as free convection, the correlation between the predicted and measured structural temperatures could be improved greatly.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TM-100414 , H-1466 , NAS 1.15:100414
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Finite-element reentry heat-transfer analysis of space shuttle Orbiter (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A structural performance and resizing (SPAR) finite-element thermal analysis computer program was used in the heat-transfer analysis of the space shuttle orbiter subjected to reentry aerodynamic heating. Three wing cross sections and one midfuselage cross section were selected for the thermal analysis. The predicted thermal protection system temperatures were found to agree well with flight-measured temperatures. The calculated aluminum structural temperatures also agreed reasonably well with the flight data from reentry to touchdown. The effects of internal radiation and of internal convection were found to be significant. The SPAR finite-element solutions agreed reasonably well with those obtained from the conventional finite-difference method.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TP-2657 , H-1236 , NAS 1.60:2657
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Comparison of flight-measured and calculated temperatures on the space shuttle orbiter (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Structural temperatures and thermal protection system surface temperatures were measured on the space shuttle during the flight of STS 5. The measured data are compared with values calculated at wing stations 134, 240, and 328 and at fuselage station 877. The theoretical temperatures were calculated using the structural performance and resizing finite element thermal analysis program. The comparisons show that the calculated temperatures are, generally, in good agreement with the measured data.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TM-88278 , H-1384 , NAS 1.15:88278
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    NASA TECHNICAL REPORTS
  • 5
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    Effects of forced and free convections on structural temperatures of Space Shuttle Orbiter during reentry flight (1987)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Structural performance and resizing (SPAR) finite-element thermal analysis computer program was used in the heat transfer analysis of the Space Shuttle Orbiter wing subjected to reentry aerodynamic heating. With sufficient external forced convective cooling near the end of the heating cycle, the calculated surface temperatures of the thermal protection system (TPS) agree favorably with the flight data for the entire flight profile. However, the effects of this external forced convective cooling on the structural temperatures were found to be negligible. Both free convection and forced convection elements were introduced to model the internal convection effect of the cool air entering the Shuttle interior. The introduction of the internal free convection effect decreased the calculated wing lower skin temperatures by 20 F at most, 1200 sec after touchdown. If the internal convection is treated as forced convection, the calculated wing lower skin temperatures after touchdown can be reduced to match the flight-measured data very closely. By reducing the TPS thicknesses to certain effective thickness to account for the TPs gap heating, the calculated wing lower skin temperatures prior to touchdown can be raised to agree with the flight data perfectly.
    Keywords: SPACE TRANSPORTATION
    Type: AIAA PAPER 87-1600
    Format: text
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    NASA TECHNICAL REPORTS
  • 6
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    Real-time aerodynamic heating and surface temperature calculations for hypersonic flight simulation (1990)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A real-time heating algorithm was derived and installed on the Ames Research Center Dryden Flight Research Facility real-time flight simulator. This program can calculate two- and three-dimensional stagnation point surface heating rates and surface temperatures. The two-dimensional calculations can be made with or without leading-edge sweep. In addition, upper and lower surface heating rates and surface temperatures for flat plates, wedges, and cones can be calculated. Laminar or turbulent heating can be calculated, with boundary-layer transition made a function of free-stream Reynolds number and free-stream Mach number. Real-time heating rates and surface temperatures calculated for a generic hypersonic vehicle are presented and compared with more exact values computed by a batch aeroheating program. As these comparisons show, the heating algorithm used on the flight simulator calculates surface heating rates and temperatures well within the accuracy required to evaluate flight profiles for acceptable heating trajectories.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TM-4222 , H-1602 , NAS 1.15:4222
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 7
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    Preliminary analysis for a Mach 8 crossflow transition experiment on the Pegasus (R) space booster (1993)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A boundary-layer transition is proposed for a future flight mission of the air-launched Pegasus space booster. The flight experiment requires attaching a glove assembly to the wing of the first-stage booster. The glove design consists of a spring and hook attachment system which allows for thermal growth of a steel 4130 skin. The results from one- and two-dimensional thermal analyses of the initial design are presented. Results obtained from the thermal analysis using turbulent flow conditions showed a maximum temperature of approximately 305 C and a chordwise temperature gradient of less than 8.9 C/cm for the critical areas in the upper glove skin. The temperatures obtained from these thermal analyses are well within the required temperature limits of the glove.
    Keywords: AERONAUTICS (GENERAL)
    Type: NASA-TM-104272 , H-1954 , NAS 1.15:104272 , Society for Experimenal Mechanics, Structural Testing Technology at High Temperature 2 Conference; Nov 08, 1993 - Nov 10, 1993; Ojai, CA; United States
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    NASA TECHNICAL REPORTS
  • 8
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    Reentry Thermal Analysis of a Generic Crew Exploration Vehicle Structure (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Comparative studies were performed on the heat-shielding characteristics of honeycomb-core sandwich panels fabricated with different materials for possible use as wall panels for the proposed crew exploration vehicle. Graphite/epoxy sandwich panel was found to outperform aluminum sandwich panel under the same geometry due to superior heat-shielding qualities and lower material density. Also, representative reentry heat-transfer analysis was performed on the windward wall structures of a generic crew exploration vehicle. The Apollo low Earth orbit reentry trajectory was used to calculate the reentry heating rates. The generic crew exploration vehicle has a graphite/epoxy composite honeycomb sandwich exterior wall and an aluminum honeycomb sandwich interior wall, and is protected with the Apollo thermal protection system ablative material. In the thermal analysis computer program used, the TPS ablation effect was not yet included; however, the results from the nonablation heat-transfer analyses were used to develop a "virtual ablation" method to estimate the ablation heat loads and the thermal protection system recession thicknesses. Depending on the severity of the heating-rate time history, the virtual ablation period was found to last for 87 to 107 seconds and the ablation heat load was estimated to be in the range of 86 to 88 percent of the total heat load for the ablation time period. The thermal protection system recession thickness was estimated to be in the range of 0.08 to 0.11 inches. For the crew exploration vehicle zero-tilt and 18-degree-tilt stagnation points, thermal protection system thicknesses of h = {0.717, 0.733} inches were found to be adequate to keep the substructural composite sandwich temperature below the limit of 300 F.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NASA/TM-2007-214607 , H-2674
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    NASA TECHNICAL REPORTS
  • 9
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    A Method for Calculating Transient Surface Temperatures and Surface Heating Rates for High-Speed Aircraft (2000)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: This report describes a method that can calculate transient aerodynamic heating and transient surface temperatures at supersonic and hypersonic speeds. This method can rapidly calculate temperature and heating rate time-histories for complete flight trajectories. Semi-empirical theories are used to calculate laminar and turbulent heat transfer coefficients and a procedure for estimating boundary-layer transition is included. Results from this method are compared with flight data from the X-15 research vehicle, YF-12 airplane, and the Space Shuttle Orbiter. These comparisons show that the calculated values are in good agreement with the measured flight data.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NASA/TP-2000-209034 , NAS 1.60:209034 , H-2427
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    NASA TECHNICAL REPORTS
  • 10
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    Hypothetical Reentry Thermostructural Performance of Space Shuttle Orbiter With Missing or Eroded Thermal Protection Tiles (2004)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: 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.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NASA/TM-2004-212850 , H-2553
    Format: application/pdf
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    NASA TECHNICAL REPORTS
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