Surface Flow Patterns and Aerodynamic Heating on Space Shuttle Vehicles

Marvin, J. G.; Seegmiller, H. L.; Lockman, W. K.; Mateer, G. G.; Pappas, C. C.; Derose, C. E.

doi:10.2514/3.61748

Cited by 7 publications

(2 citation statements)

References 4 publications

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“…The effect of this spanwise flow is to shrink the separated flow extent 4 and to decrease the pressure in the separated flow region. 5 The results presented by Marvin et al 6 show such three-dimensional flow effects even for control-induced turbulent flow separation.Î kawa 1 acknowledges that spanwise flow tends to reduce the separated flow region. This "shrinking" of the separated flow region due to spanwise flow will by itself, without considering any pressure changes, change the separationinduced loads.…”

mentioning

confidence: 88%

Comment on 'A re-entry control effectiveness methodology for the Space Shuttle Orbiter'

Ericsson¹,

Reding²

1978

Journal of Spacecraft and Rockets

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TECHNICAL COMMENTS 189 equation of velocity reduces to (9) Equation (9) is identical, except for the "approximately equal" notation, to the well-known formula that gives the ideal velocity attained by a conventional rocket in a gravitationless vacuum,t the derivation of which assumes that the exhaust velocity and mass flow rate remain constant. Both of these items, however, are varying in the case of the cold-gas thruster. Accuracy of ApproximationBy comparing Eqs. (1) and (5), it may be seen that the terminal velocity, according to the approximate formula, is slightly overestimated. The maximum percentage error occurs with thrusters using the monatomic gases (k= 3 or j= 1.667), and amounts to about 5% in the extreme case of X= Vi. As indicated in Fig. 1, the difference between the two equations decreases as A: or X grows large, such that the error in any situation does not exceed 1 % for propellant mass fractions ranging up to 10%. tSee, for example, Ref. 5.It may be concluded that for the class of cold-gas thrusters containing small propellant mass, the approximate velocity equation closely represents the integral equation of ideal velocity. The abbreviated formula permits the effect of various design factors to be readily determined.

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mentioning

confidence: 88%

Comment on 'A re-entry control effectiveness methodology for the Space Shuttle Orbiter'

Ericsson¹,

Reding²

1978

Journal of Spacecraft and Rockets

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“…4, are supported by representative wind tunnel data. 5 Based on these pressure distributions and correlations of shock shape from wind tunnel shadowgraphs, streamline pressure distributions were calculated by assuming that the windward pitchplane profile of the shuttle at angle of attack was the generator of an axisymmetric body. Since pressure distributions are relatively insensitive to chemistry, nonequilibrium states along these prescribed-pressure streamlines were then calculated by chemical relaxation from a frozen shock wave.…”

Section: Shuttle Pitch-plane Heatingmentioning

confidence: 99%

Nonequilibrium Viscous-Layer Computations of Space Shuttle TPS Requirements

Tong

Morse

Curry

1975

Journal of Spacecraft and Rockets

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The thermal protection system requirements on the windward pitchplane of a typical shuttle orbiter are evaluated using a recently developed computational procedure. This procedure treats chemically nonequilibrium inviscid and viscous flows for surfaces of arbitrary catalycity. Surface catalycities for typical shuttle materials were evaluated from arc plasma generator data. Computations for the shuttle orbiter demonstrated the significance of entropy swallowing on heat transfer and show that for representative catalycity values, nonequilibrium effects are small though not insignificant. It is shown that the reduced heat transfer rates, caused by surface effects, observed in arc plasma facilities, should be reviewed carefully since these reduced rates will not be as significant for the larger orbiter dimensions. NomenclatureH T = total enthalpy k w (O) = surface catalycity for O recombination k w (N) = surface catalycity for N recombination M e = boundary-layer edge Mach number MOO = freestream or flight Mach number P = pressure Pt 2 = total pressure behind a normal shock wave q = heat flux

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