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  • 1
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    Mars environmental survey probe aerobrake preliminary design study (1993)
    In:  Other Sources
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    Publication Date: 2011-08-24
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4650); 30; 4; p. 431-437.
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  • 2
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    Navier-Stokes solutions with surface catalysis for Martian atmospheric entry (1993)
    In:  Other Sources
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    Publication Date: 2011-08-24
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4650); 30; 1; p. 32-42.
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  • 3
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    CFD Analysis of Hypersonic Flowfields With Surface Thermochemistry and Ablation (1997)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: In the past forty years much progress has been made in computational methods applied to the solution of problems in spacecraft hypervelocity flow and heat transfer. Although the basic thermochemical and physical modeling techniques have changed little in this time, several orders of magnitude increase in the speed of numerically solving the Navier-Stokes and associated energy equations have been achieved. The extent to which this computational power can be applied to the design of spacecraft heat shields is dependent on the proper coupling of the external flow equations to the boundary conditions and governing equations representing the thermal protection system in-depth conduction, pyrolysis and surface ablation phenomena. A discussion of the techniques used to do this in past problems as well as the current state-of-art is provided. Specific examples, including past missions such as Galileo, together with the more recent case studies of ESA/Rosetta Sample Comet Return, Mars Pathfinder and X-33 will be discussed. Modeling assumptions, design approach and computational methods and results are presented.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: Proceedings of the Eighth Annual Thermal and Fluids Analysis Workshop: Spacecraft Analysis and Design; S2.1-S2.19; NASA-CP-3359
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  • 4
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    Navier-Stokes solutions with surface catalysis for Martian atmospheric entry (1992)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: In this study numerical solutions have been obtained for two-dimensional axisymmetric hypersonic nonequilibrium CO2 flow over a high angle blunt cone with appropriate surface boundary conditions to account for energy and mass conservation at the body surface. The flowfield is described by the Navier-Stokes equations and multicomponent conservation laws which account for both translational and internal vibrational nonequilibrium effects. Complete forebody solutions have been obtained for the peak heating point of the Mars entry trajectory specified in the proposed NASA MESUR (Mars Environmental Survey) project. In these solutions, radiative equilibrium wall temperature and surface heating distributions are determined over the MESUR aeroshell forebody for entry velocity equal to 7 km/sec with varying degrees of surface catalysis. The effects of gas kinetics, surface catalysis, transport properties, and vibrational relaxation times on the surface heating are examined. The results identify some important issues in the prediction of surface heating for flows in thermochemical nonequilibrium and show that the Navier-Stokes code used herein is effective for thermal protection system design and materials selection.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: AIAA PAPER 92-2946
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  • 5
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    MESUR probe aerobrake preliminary design study (1992)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Aerobrake design for the Mars Environmental Survey (MESUR) vehicles is considered which is intended for both a nominal entry velocity of 7 km/sec and a high-speed case of 9 km/sec. Topics discussed include the entry environment, the thermal protection requirements for several types of heat shield materials, the structural design of the aeroshell, and the total aerobrake masses and mass fractions. For the nominal 7 km/sec entry, a silicone elastometric charring ablator, SLA-561, was found to be the lightest heat shield material. For the 7 km/sec entry, the mass fraction of the aerobrake was 13.2 percent. For the 9 km/sec entry, the heat shield consisted of the medium-density ablator AVCOAT-5026; SLA-561 was used on part of the conical skirt. The aerobrake mass fraction in this case was 18 percent. It is recommended that separate aerobrakes be designed for probes entering at 7 and 9 km/sec.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: AIAA PAPER 92-2952
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  • 6
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    Aerobrake design studies for manned Mars missions (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The dimensions of aerobrakes and associated heat shields are calculated as a fraction of the vehicle mass required for a high-velocity manned Mars entry. The entry speed and deceleration limit are assumed to be 8.6 km/sec and 5 earth g, respectively, to consider vehicles with low lift-drag ratio (L/D) and ballistic coefficients of 100 and 200 kg/sq m, as well as a vehicle with a medium L/D and a ballistic coefficient of 375 kg/sq m. The aerobrake mass plus the heat shield divided by an optimized, blunt-shaped vehicle's total mass is 15 and 13 percent for ballistic coefficients of 100 and 200 kg/sq m, respectively. For a winged vehicle the mass fraction is 17 percent because the higher ballistic coefficient requires more thermal protection to account for the greater temperatures generated. It is concluded that aerobraking is more efficient than propulsive braking because the mass fraction for a propulsive system would be 4 or 5 times greater than those calculated for aerobraking.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: AIAA PAPER 91-1344
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  • 7
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    Trajectory Based Heating and Ablation Calculations for MESUR Pathfinder Aeroshell (1994)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: Based on the geometry of Mars Environment Survey (MESUR) Pathfinder aeroshell and an estimated Mars entry trajectory, two-dimensional axisymmetric time dependent calculations have been obtained using GIANTS (Gauss-Siedel Implicit Aerothermodynamic Navier-Stokes code with Thermochemical Surface Conditions) code and CMA (Charring Material Thermal Response and Ablation) Program for heating analysis and heat shield material sizing. These two codes are interfaced using a loosely coupled technique. The flowfield and convective heat transfer coefficients are computed by the GIANTS code with a species balance condition for an ablating surface, and the time dependent in-depth conduction with surface blowing is simulated by the CMA code with a complete surface energy balance condition. In this study, SLA-561V has been selected as heat shield material. The solutions, including the minimum heat shield thicknesses over aeroshell forebody, pyrolysis gas blowing rates, surface heat fluxes and temperature distributions, flowfield, and in-depth temperature history of SLA-561V, are presented and discussed in detail.
    Keywords: Fluid Mechanics and Thermodynamics
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  • 8
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    Comparison of Coupled Radiative Flow Solutions with Project Fire 2 Flight Data (1995)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A nonequilibrium, axisymmetric, Navier-Stokes flow solver with coupled radiation has been developed for use in the design or thermal protection systems for vehicles where radiation effects are important. The present method has been compared with an existing now and radiation solver and with the Project Fire 2 experimental data. Good agreement has been obtained over the entire Fire 2 trajectory with the experimentally determined values of the stagnation radiation intensity in the 0.2-6.2 eV range and with the total stagnation heating. The effects of a number of flow models are examined to determine which combination of physical models produces the best agreement with the experimental data. These models include radiation coupling, multitemperature thermal models, and finite rate chemistry. Finally, the computational efficiency of the present model is evaluated. The radiation properties model developed for this study is shown to offer significant computational savings compared to existing codes.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA/TM-95-207283 , NAS 1.15:207283 , AIAA Paper 94-1955 , Joint Thermophysics and Heat Transfer Conference; Jun 20, 1994 - Jun 23, 1994; Colorado Springs, CO; United States|Journal of Thermophysics and Heat Transfer; 9; 4; 586-594
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  • 9
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    Chemical nonequilibrium Navier-Stokes solutions for hypersonic flow over an ablating graphite nosetip (1993)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The general boundary conditions including mass and energy balances of chemically equilibrated or nonequilibrated gas adjacent to ablating surfaces have been derived. A computer procedure based on these conditions was developed and interfaced with the Navier-Stokes solver for predictions of the flow field, surface temperature, and surface ablation rates over re-entry space vehicles with ablating Thermal Protection Systems (TPS). The Navier-Stokes solver with general surface thermochemistry boundary conditions can predict more realistic solutions and provide useful information for the design of TPS. A test case with a proposed hypersonic test vehicle configuration and associated free stream conditions was developed. Solutions with various surface boundary conditions were obtained, and the effect of nonequilibrium gas as well as surface chemistry on surface heating and ablation rate were examined. The solutions of the GASP code with complete ablating surface conditions were compared with those of the ASC code. The direction of future work is also discussed.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: AIAA PAPER 93-2836 , ; 12 p.|AIAA, Thermophysics Conference; Jul 06, 1993 - Jul 09, 1993; Orlando, FL; United States
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  • 10
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    Trajectory based, 3-dimensional heating and ablation calculations for the Apollo Lunar/Earth return capsule (1993)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: Three dimensional (3-D) surface heating and ablation rate distributions have been obtained based on the geometry of the project Apollo Lunar/Earth return capsule and high Reynolds number inviscid/boundary layer methods. This application is based on the principles of the pseudo-three dimensional axisymmetric streamline analogy. Inviscid surface streamlines are determined for the forebody portion of the Apollo capsule using a CFD solver. Streamlines are obtained for the 20 deg angle-of-attack flight case, and the associated streamline metrics are determined. Boundary layer heating and ablation calculations are performed along selected Apollo forebody streamlines using the non-similar boundary layer code. Surface radiative heating estimates were obtained and the validity of the streamline tracing analogy is examined. These comparisons and the Apollo trajectory heating results are discussed.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: AIAA PAPER 93-2788 , ; 17 p.|AIAA, Thermophysics Conference; Jul 06, 1993 - Jul 09, 1993; Orlando, FL; United States
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