Similitude requirements for hypersonic, rarefied, nonequilibrium flow

Similitude requirements for hypersonic, rarefied flow with nonequilibrium chemistry and vibration are presented. The full Navier-Stokes equations with catalytic or noncatalytic walls and with or without slip conditions are nondimensionalized. The heat transfer coefficient is written in terms of fourteen dimensionless parameters and reduced to four by making the binary scaling assumption. Duplication of blunt and sharp nose heat transfer requires the use of air over a geometrically similar model with the same free stream velocity, wall temperature and product of free stream density and characteristic length. Estimates of this heat transfer coefficient are also presented

The addition of algebraic turbulence modeling to program LAURA

The Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) is modified to allow the calculation of turbulent flows. This is accomplished using the Cebeci-Smith and Baldwin-Lomax eddy-viscosity models in conjunction with the thin-layer Navier-Stokes options of the program. Turbulent calculations can be performed for both perfect-gas and equilibrium flows. However, a requirement of the models is that the flow be attached. It is seen that for slender bodies, adequate resolution of the boundary-layer gradients may require more cells in the normal direction than a laminar solution, even when grid stretching is employed. Results for axisymmetric and three-dimensional flows are presented. Comparison with experimental data and other numerical results reveal generally good agreement, except in the regions of detached flow

Unsteady Newton-Busemann flow theory. Part 2: Bodies of revolution

Newtonian flow theory for unsteady flow past oscillating bodies of revolution at very high Mach numbers is completed by adding a centrifugal force correction to the impact pressures. Exact formulas for the unsteady pressure and the stability derivatives are obtained in closed form and are applicable to bodies of revolution that have arbitrary shapes, arbitrary thicknesses, and either sharp or blunt noses. The centrifugal force correction arising from the curved trajectories followed by the fluid particles in unsteady flow cannot be neglected even for the case of a circular cone. With this correction, the present theory is in excellent agreement with experimental results for sharp cones and for cones with small nose bluntness; gives poor agreement with the results of experiments in air for bodies with moderate or large nose bluntness. The pitching motions of slender power-law bodies of revulution are shown to be always dynamically stable according to Newton-Busemann theory

Heat Transfer Analysis without and with Forward Facing Spike Attached to a Blunt Body at High Speed Flow

The present chapter deals with heat transfer analysis around unspiked and spiked bodies at high speeds. A spike attached to a blunt-nosed body drastically alters its flowfield and influences the aerodynamic heating in a high speed flow. The effect of spike length, shape and spike-nose configuration is numerically studied at zero angle of incidence. The numerical analysis describes overall flowfield features over without and with forward facing spike attached to a blunt body at high speed flow. The shock stand-off distance, sonic line, stagnation point velocity gradient and stagnation point heat flux are analyzed and compared with different aerodisk configurations. It is found that the hemispherical aerodisk experiences high wall heat flux as compared to the flat-faced aerodisk. Numerical and experimental studies reveal that the wall heat flux levels are decreased in the presence of the spikes and aerospike as compared to without attached spiked to the blunt-nose basic configuration

Attached and separated boundary layers on highly cooled, ablating and nonablating models at M equals 13.8

Attached and separated boundary layers on highly cooled, ablating and nonablating models at Mach 13.

Effects of nose bluntness and shock-shock interactions on blunt bodies in viscous hypersonic flows

A numerical study was conducted to investigate the effects of blunt leading edges on the viscous flow field around a hypersonic vehicle such as the proposed National Aero-Space Plane. Attention is focused on two specific regions of the flow field. In the first region, effects of nose bluntness on the forebody flow field are investigated. The second region of the flow considered is around the leading edges of the scramjet inlet. In this region, the interaction of the forebody shock with the shock produced by the blunt leading edges of the inlet compression surfaces is analyzed. Analysis of these flow regions is required to accurately predict the overall flow field as well as to get necessary information on localized zones of high pressure and intense heating. The results for the forebody flow field are discussed first, followed by the results for the shock interaction in the inlet leading edge region

Shock-Induced Separation of Transitional Hypersonic Boundary Layers

This thesis presents a joint experimental/CFD investigation of shock-induced boundary layer separations in hypersonic transitional boundary layers with an emphasis on collapse and re-establishment times of the separation bubble. This study also provides high fidelity measurements and excellent characterisation of the flow field in order to provide benchmark data of a challenging flow configuration with which to benchmark next generation CFD solvers. The experiments were conducted in the Imperial College Aeronautics Department Number Two Gun Tunnel, a Mach 8.9 axisymmetric facility with a freestream unit Reynolds number of 47 million An axisymmetric blunt-nosed cylinder fitted with an 8 degree flare forms the primary vehicle for this study, although a 1.3 degree cowl geometry was also used to impinge a shock onto the blunt-nosed cylinder.. The shock boundary layer interaction was designed such that it was separated for a laminar boundary layer and collapsed for a turbulent one. Carefully controlled turbulent spots were generated upstream of the interaction region which passed through the separation causing its collapse and subsequent re-establishment. Two intermittency cases are considered, one where turbulent spot spacing is large and collapse/re-establishment pairs can be considered independent of each other and one where they can not. Experimental surface quantities through the interaction region are measured using either heat-transfer or pressure measurements and schlieren video is used to diagnose the larger shock structure. Further a non-intrusive toluene PLIF method is assessed for use in this facility and shows promise. CFD simulations are done using an in-house operator split Godunov solver with a Baldwin-Lomax turbulence model. CFD simulations show good agreement with experiment and provides information on flow quantities that would be extremely difficult to measure otherwise. Collapse times of the separation bubble were found to be fast in relation to characteristic spot passage times. The collapse process is also fast in relation to the surrounding flows ability to adjust, with collapse associated with significant shock curvature of the immediate outboard shock structures. This leads to unsteadiness, with surface pressure measurements exceeding the range bounded by the laminar separated and turbulent collapsed cases. The severity of the unsteadiness appears to be driven by turbulent spot spacing. Re-establishment is considerably slower, showing asymptotic recovery that is likely driven by viscous diffusion rates, taking many characteristic spot passage times to recover.Open Acces

Hypersonic aerodynamic characteristics of a family of power-law, wing body configurations

The configurations analyzed are half-axisymmetric, power-law bodies surmounted by thin, flat wings. The wing planform matches the body shock-wave shape. Analytic solutions of the hypersonic small disturbance equations form a basis for calculating the longitudinal aerodynamic characteristics. Boundary-layer displacement effects on the body and the wing upper surface are approximated. Skin friction is estimated by using compressible, laminar boundary-layer solutions. Good agreement was obtained with available experimental data for which the basic theoretical assumptions were satisfied. The method is used to estimate the effects of power-law, fineness ratio, and Mach number variations at full-scale conditions. The computer program is included

A numerical study of novel drag reduction techniques for blunt bodies in hypersonic flows

Numerical simulations of a full three-dimensional hemispherical body in hypersonic flow are conducted and innovative techniques involving forward injection of gas from the stagnation point of the sphere are investigated; techniques include annular (ring) and swirled injection both with and without upstream energy deposition. Objectives of the analysis are the assessment of 1) drag reductions achieved on the blunt body (including the detrimental drag effect caused by the forward-facing injection itself) and 2) stability characteristics of the jet. Studies are conducted at free-stream Mach numbers of 10 and 6.5 at standard atmospheric conditions corresponding to 30 km altitude. While centered forward injection without upstream energy deposition is confirmed to be highly unstable either with or without swirl, annular ring injection exhibits a stabilizing influence on the jet. Energy deposition upstream of the body is shown to significantly enhance stability and penetration of the forward injection jet for all techniques --Abstract, page iii