Simulation of the Apollo Command Module aerodynamics at re-entry altitudes.
Apollo Command Module aerodynamics at reentry altitudes simulated by models in hypersonic flow
Boylan, D. E.↗
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Hypersonic laminar boundary layers around slender bodies.
Nonsimilar compressible laminar boundary layer equations for hypersonic flow around slender bodies solved by integration based on steepest descent method
Jeng, D. R.↗
Simulation of the Apollo command module aerodynamics at high altitudes.
Aerodynamic measurements on Apollo CM model at hypersonic flow simulating earth orbital reentry trajectory
Boylan, D. E.↗
Radiative transfer in the low Reynolds number, blunt-body stagnation region at hypersonic speeds.
Radiative transfer effects in low Reynolds number or merged layer regime of hypersonic flow about axisymmetric blunt bodies, including thin shock layer theory
Liu, J. T. C.↗
Downstream cooling effectiveness of massive upstream transpiration from a slender cone.
Radiation equilibrium temperature measured downstream of transpiration cooled gas flow near slender cone vertex in continuous flow hypersonic tunnel
Dunavant, J. C.↗
Aerodynamic interference effects on half-cone bodies with thin wings at Mach 10.03
Aerodynamic interference effects on half-cone bodies with thin wings in hypersonic flow
Townsend, J. C.↗
Heat-transfer measurements at Mach 8 on a flat plate with deflected trailing-edge flap with effects of transition included
Heat transfer measurements on flat plate with trailing edge flap in hypersonic flow
Johnson, C. B.↗
Detailed pressure distribution on a blunted 60-deg half-angle cone at Mach numbers of 6.08 and 9.46
Pressure distribution of spherically-blunted 60 deg half-angle cone in hypersonic flow
Kurtz, D. W.↗
Shock-wave shapes on hypersonic axisymmetric power-law bodies
Shock wave shapes over axisymmetric power law bodies in hypersonic flow, noting asymptotic flow theory implications
Beavers, G. S.↗
Numerical solution of the hypersonic viscous shock-layer equations
Viscous shock layer equations of laminar hypersonic flow past blunt body at moderate to high Reynolds numbers
Davis, R. T.↗
Heat transfer and pressure on a hypersonic blunt cone with mass addition
Surface pressure and heat transfer over blunt conical body in hypersonic flow with uniform mass addition of various gases
Lee, G.↗
Hypersonic gas dynamics of slender bodies
Hypersonic flow past slender bodies, discussing inviscid flows, outer edge singularity of boundary layer and three dimensional interaction on needle- like bodies
Cheng, H. K.↗
Calculation of inviscid surface streamlines and heat transfer on shuttle type configurations. Part 2: Description of computer program
A description of the computer program used for heating rate calculation for blunt bodies in hypersonic flow is given. The main program and each subprogram are described by defining the pertinent symbols involved and presenting a detailed flow diagram and complete computer program listing. Input and output parameters are discussed in detail. Listings are given for the computation of heating rates on (1) a blunted 15 deg half-angle cone at 20 deg incidence and Mach 10.6, (2) a blunted 70 deg slab delta wing at 10 deg incidence and Mach 8, and (3) the HL-10 lifting body at 20 deg incidence and Mach 10. In addition, the computer program output for two streamlines on the blunted 15 deg half-angle cone is listed. For Part 1, see N71-36186.
Dejarnette, F. R.↗
Hypersonic bodies of maximum drag for a given lift-to-drag ratio.
The problem considered in this paper is concerned with the aerodynamic design of the forebody shape of reentry vehicles in the blunt, homothetic, elliptic transversal contour, power-law longitudinal contour, raked-off configurational set. In particular, the forebody shape which maximizes the ratio of the forebody pressure drag to the free-stream dynamic pressure for a given lift-to-drag ratio and given geometric properties is determined. This problem is considered because recent survey articles indicate that its solution will provide useful qualitative design information about manned vehicles entering the earth's atmosphere from any of the foreseeable planetary missions. Single-integral equations relating the lift and drag in Newtonian hypersonic flow to the forebody geometry are derived and used to formulate the optimization problem which is solved by a direct numerical method.
Mcmillan, W., III↗
Cone transitional boundary-layer structure at Me = 14.
Experimental results suggest that the boundary-layer mean profiles in hypersonic flows may be highly transitional in the outer part of the boundary layer before the transition process is detected at the surface. Means of determining the initial location where the transition process begins in the outer part of the boundary layer are considered.
Fischer, M. C.↗
Inviscid surface streamlines and heat transfer on shuttle-type configurations.
A method is developed which calculates laminar, transitional, and turbulent heating rates on arbitrary blunt-nosed three-dimensional bodies at angle of attack in hypersonic flow. The geometry of the body may be specified analytically, or generated from a doubly cubic spline fit to coordinate points. Inviscid surface streamlines are calculated from Euler's equation using a prescribed pressure distribution. Laminar and turbulent heating rates are determined along a streamline by applying the axisymmetric analog to solutions of the axisymmetric boundary-layer equations. The location of the transition region may be specified optionally by geometric location, momentum thickness Reynolds number, or integrated unit Reynolds number along a streamline. Transitional heating rates are then calculated as a weighted average of the local laminar and turbulent values. Either ideal gas or equilibrium air properties may be used.
Dejarnette, F. R.↗
Analysis of thermal environment in the thrustor cavity of space vehicles
Wind tunnel tests to determine the mechanisms involved in producing the high local heating rates in and around the thrustor nozzle in a non-firing condition are reported. The geometry of a typical nozzle installation is described. Procedures for reducing the effects of hypersonic flow interactions in order to prevent excess local heating are explained.
Chiou, J. N.↗
Oblique shock-sound interaction at a freestream Mach number of about 20 in helium.
Detailed experimental measurements of the characteristics and level of pre- and postshock disturbances obtained for typical disturbance modes and shock angles are presented to aid in the evaluation of transition studies conducted in 'noisy' hypersonic wind tunnels. The presented data suggest that the shock is probably generating total temperature fluctuations in a fashion directly related to shock strength. Since in high supersonic and hypersonic flows the total temperature fluctuations are mainly contained in velocity fluctuations, it may be concluded that the predominant disturbances generated are vorticity as long as the postshock flow is at a sufficiently high Mach number.-
Weinstein, L. M.↗