The separation of Newtonian shock layers

1966 ◽  
Vol 26 (3) ◽  
pp. 563-572 ◽  
Author(s):  
J. R. Ockendon
Keyword(s):  
Hypersonic Flow ◽  
Shock Layer ◽  
Present Approach ◽  
Free Layer ◽  
Newtonian Theory ◽  
Direct Verification ◽  
Continuous Curvature ◽  
Shock Layers ◽  
Layer Solution

The separation points which occur in the Newtonian theory of hypersonic flow are treated by locally modifying the shock-layer equations. This approach leads to direct verification of the free-layer theory for separation on certain bodies with discontinuous curvature. Separation points on bodies with continuous curvature have already been treated by matching the upstream shock-layer solution to the downstream free-layer solution. The agreement that is found between these results and those of the present approach provides further confirmation of the free-layer theory.

MHD Control of the Hypersonic Flow in the Shock Layer Around a Body

2008 ◽  
Author(s):  
Andrea Cristofolini ◽  
Carlo Borghi ◽  
Gabriele Neretti ◽  
Andrea Passaro ◽  
Leonardo Biagioni
Keyword(s):  
Hypersonic Flow ◽  
Shock Layer

Magnetohydrodynamic Interaction in the Shock Layer of a Wedge in a Hypersonic Flow

2006 ◽  
Vol 34 (5) ◽  
pp. 2450-2463 ◽  
Author(s):  
C.A. Borghi ◽  
M.R. Carraro ◽  
A. Cristofolini ◽  
A. Veefkind ◽  
L. Biagioni ◽  
...  
Keyword(s):  
Hypersonic Flow ◽  
Shock Layer ◽  
Magnetohydrodynamic Interaction

scholarly journals Some aspects of hypersonic flow over power law bodies

1969 ◽  
Vol 39 (1) ◽  
pp. 143-162 ◽  
Author(s):  
H. G. Hornung
Keyword(s):  
Mach Number ◽  
Hypersonic Flow ◽  
Blast Wave ◽  
Power Law ◽  
Numerical Solutions ◽  
The Body ◽  
Stagnation Temperature ◽  
Free Layer ◽  
Plane Flow ◽  
Wave Limit

This study concerns the hypersonic flow over blunt bodies in two specific cases. The first is the case when the Mach number is infinite and the ratio of the specific heats approaches one. This is sometimes referred to as the ‘Newtonian limit’. The second is the case of infinite Mach number and very large streamwise distance from the blunt nose with a strong shock wave, or the ‘blast wave limit’. In both cases attention is restricted to power law bodies. Experiments are described of such flows at M∞ = 7·55 in air.The Newtonian flow over bodies of the shape y ∞ xm at zero incidence is shown to be divisible into three regions: the attached layer at small x, the free layer and the blast wave region. As m increases from zero, the free-layer region reduces in extent until it disappears at m = 1/(2+j) (j = 1 and 0 for axisymmetric and plane flow respectively). A difficulty arises in a transition solution of the type given by Freeman (1962b) connecting the free layer with the blast wave result. At m > 2/(3+j) the attached layer merges smoothly into the Lees-Kubota solution which replaces the blast-wave result in this range.In the blast wave limit, solutions were obtained for flow over axisymmetric power law shapes in the range ½γ < m < ½. Second-order results taking account of the body shape are given. These solutions are compared with experimental results obtained in air at a free stream Mach number of 7·55 and stagnation temperature of 630 °K, as well as with numerical solutions at Mach number of 100. The numerical method is tested by comparing solutions corresponding to the experimental conditions with experiment.

Three-dimensional wings in hypersonic flow

1972 ◽  
Vol 54 (2) ◽  
pp. 305-337 ◽  
Author(s):  
R. Hillier
Keyword(s):  
Exact Solution ◽  
Hypersonic Flow ◽  
Shock Layer ◽  
Three Dimensional ◽  
Leading Edge ◽  
The Other ◽  
Calculation Methods ◽  
Conical Flow ◽  
Thin Shock Layer ◽  
Good For

Messiter's thin shock layer approximation for hypersonic wings is applied to several non-conical shapes. Two calculation methods are considered. One gives the exact solution for a particular three-dimensional geometry which possesses a conical planform and also a conical distribution of thickness superimposed upon a surface cambered in the chordwise direction. Agreement with experiment is good for all cases, including that where the wing is yawed. The other method is a more general approach whereby the solution is expressed as a correction to an already known conical flow. Such a technique is applicable to conical planforms with either attached or detached shocks but only to the non-conical planform for the region in the vicinity of the leading edge when the shock is attached.

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