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  • 1
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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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    NASA TECHNICAL REPORTS
  • 2
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    Heat-transfer and pressure analysis of rocket exhaust impingement on flat plates and curved panels. (1969)
    In:  Other Sources
    Details
    Publication Date: 2011-08-12
    Description: Pressure and heating rate correlations for rocket exhausts impinging on flat plates and curved panels, generating axisymmetric real gas exhaust plumes
    Keywords: THERMODYNAMICS AND COMBUSTION
    Type: ; 633 (
    Format: text
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    NASA TECHNICAL REPORTS
  • 3
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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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    NASA TECHNICAL REPORTS
  • 4
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    Space shuttle orbiter reentry flow field and heating analysis (1972)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The reentry flow field and thermal environment around the straight wing shuttle orbiter vehicle were determined. Both rarefied and continuum flow fields and associated heating rates on various configurations representative of the orbiter at high angle of attack were calculated. Rarefied flow fields and heating rates were computed by the Monte Carlo direct simulation technique for altitudes above 82.3 km. Continuum inviscid flow fields were calculated by 2-D unsteady and 3-D steady finite difference/artificial viscosity methods and also by a 2-D shock layer analysis technique. Viscous flow fields and heating rates in the continuum regime were computed by a boundary layer integral matrix method for laminar flow and by an aerodynamic surface heating technique for turbulent flow. Shapes considered in the study included flat plates (representing the underside of the orbiter fuselage or the wing MAC), orbiter fuselage cross sections, orbiter wing airfoils, and 3-D orbiter configurations, all at high angle of attack (40 - 60 deg). The theoretical results showed good agreement with measured pressure and heat transfer data.
    Keywords: FLUID MECHANICS
    Type: Space Shuttle Aerothermodyn. Technol. Conf., vol. 1; p 115-156
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Theoretical and experimental investigation of heating from Saturn solid propellant rocket exhausts. (1966)
    In:  Other Sources
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    Publication Date: 2019-05-30
    Description: Heating from Saturn solid propellant S-II and S-IVB ullage motor exhausts and S-IB, S-IVB and Centaur retro motor exhaust theoretically and experimentally analyzed
    Keywords: THERMODYNAMICS AND COMBUSTION
    Type: AIAA PAPER 66-653
    Format: text
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  • 6
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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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    NASA TECHNICAL REPORTS
  • 7
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    Aerothermodynamic environments for Mars entry, Mars return, and lunar return aerobraking missions (1990)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The aeroheating environments to vehicles undergoing Mars aerocapture, earth aerocapture from Mars, and earth aerocapture from the moon are presented. An engineering approach for the analysis of various types of vehicles and trajectories was taken, rather than performing a benchmark computation for a specific point at a selected time point in a trajectory. The radiation into Mars using the Mars Rover Sample Return (MRSR) 2-ft nose radius bionic remains a small contributor of heating for 6 to 10 km/sec; however, at 12 km/sec it becomes comparable with the convection. For earth aerocapture, returning from Mars, peak radiation for the MRSR SRC is only 25 percent of the peak convection for the 12-km/sec trajectory. However, when large vehicles are considered with this trajectory, peak radiation can become 2 to 4 times higher than the peak convection. For both Mars entry and return, a partially ablative Thermal Protection System (TPS) would be required, but for Lunar Transfer Vehicle return an all-reusable TPS can be used.
    Keywords: LUNAR AND PLANETARY EXPLORATION
    Type: AIAA PAPER 90-1701
    Format: text
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  • 8
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    Comparison of viscous shock layer and boundary layer reentry heating techniques for Orbiter nose cap (1986)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A comparison of two viscous shock layer methods and one boundary layer method for predicting the aerodynamic heating around the Orbiter nose cap during STS-5 entry is presented. The object of the study was to compare these methods with one another and with the measured Orbiter flight data for this trajectory. The nonequilibrium, chemically reacting viscous flow fields obtained by these methods are evaluated, and effects on heating rate of wall catalycity variation with time are presented. The effects of shock slip and combined wall/shock slip are considered at high altitudes (above 300,000 ft). Using the variable wall catalycity analysis, it is shown that heating rates can be predicted within a 5.7 percent flight data band for altitudes between 175,000 ft and 265,000 ft in this trajectory.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 86-1350
    Format: text
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  • 9
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    Space Shuttle Orbiter nose cap thermal analysis (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The results of a Space Shuttle Orbiter nose cap entry aeroheating assessment, thermal analysis, and correlation of flight data using multidimensional thermal math models (TMM's) and a chemically reacting boundary-layer program are described in this paper. The object of this study was to verify and revise, if required, the nose cap design heating methods and the TMM's used for flight certification. Flight temperature measurements from two Orbiter vehicles, Columbia and Challenger, have been used in this analysis and provide the basis for verification and correlation of the aerothermodynamic environment. Nose cap thermal response predictions, using TMM's verified from certification tests, show that the aerothermodynamic environment can be satisfactorily predicted using accepted analytical methods.
    Keywords: SPACE TRANSPORTATION
    Type: AIAA PAPER 86-0388
    Format: text
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  • 10
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    Development of AFE aerobrake aerothermodynamic data book (1989)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: 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.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 89-1734
    Format: text
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