 Oblique shock

An oblique shock wave, unlike a normal shock, is inclined with respect to the incident upstream flow direction. It will occur when a supersonic flow encounters a corner that effectively turns the flow into itself and compresses. The upstream streamlines are uniformly deflected after the shock wave. The most common way to produce an oblique shock wave is to place a wedge into supersonic, compressible flow. Similar to a normal shock wave, the oblique shock wave consists of a very thin region across which nearly discontinuous changes in the thermodynamic properties of a gas occur. While the upstream and downstream flow directions are unchanged across a normal shock, they are different for flow across an oblique shock wave.
It is always possible to convert an oblique shock into a normal shock by a Galilean transformation.
Contents
Oblique Shock Wave Theory
For a given Mach number, M_{1}, and corner angle, θ, the oblique shock angle, β, and the downstream Mach number, M_{2}, can be calculated. M_{2} is always less than M_{1}. Unlike after a normal shock, M_{2} can still be supersonic (weak shock wave) or subsonic (strong shock wave). Nature tends to focus on weak solution. Discontinuous changes also occur in the pressure, density and temperature, which all rise downstream of the oblique shock wave.
the θβM equation
Using the continuity equation and the fact that the tangential velocity component does not change across the shock, trigonometric relations eventually lead to the θβM equation which shows θ as a function of M_{1} and β.
It is more intuitive to want to solve for β as a function of M_{1} and θ, but this approach is more complicated, the results of which are often contained in tables or calculated through an applet.
Here is a way to calculate β knowing θ and M_{1}.
First define λ:
Then define χ:
Finally β is given by the relation:
where δ=0 corresponds to the strong shock solution and δ=1 to the weak one.
Maximum deflection angle
Within the θβM equation, a maximum corner angle, θ_{MAX}, exists for any upstream Mach number. When θ > θ_{MAX}, the oblique shock wave is no longer attached to the corner and is replaced by a detached bow shock. A θβM diagram, common in most compressible flow textbooks, shows a series of curves that will indicate θ_{MAX} for each Mach number. The θβM relationship will produce two β angles for a given θ and M_{1}, with the larger angle called a strong shock and the smaller called a weak shock. The weak shock is almost always seen experimentally.
The rise in pressure, density, and temperature after an oblique shock can be calculated as follows:
M_{2} is solved for as follows:Oblique Shock Wave Applications
Oblique shock waves are used predominantly in engineering applications when compared with normal shock waves. This can be attributed to the fact that using one or a combination of oblique shock waves results in more favorable postshock conditions (lower postshock temperature, etc.) when compared to utilizing a single normal shock. An example of this technique can be seen in the design of supersonic aircraft engine inlets, which are wedgeshaped to compress air flow into the combustion chamber while minimizing thermodynamic losses. Early supersonic aircraft jet engine inlets were designed using compression from a single normal shock, but this approach caps the maximum achievable Mach number to roughly 1.6. The wedgeshaped inlets are clearly visible on the sides of the F14 Tomcat, which has a maximum speed of Mach 2.34.
Many supersonic aircraft wings are designed around a thin diamond shape. Placing a diamondshaped object at an angle of attack relative to the supersonic flow streamlines will result in two oblique shocks propagating from the front tip over the top and bottom of the wing, with PrandtlMeyer expansion fans created at the two corners of the diamond closest to the front tip. When correctly designed, this generates lift.
Oblique Shock Waves and the Hypersonic Limit
As the Mach number of the upstream flow becomes hypersonic, the equations for the pressure, density, and temperature after the oblique shock wave reach a mathematical limit. The pressure and density ratios can then be expressed as:
For a perfect atmospheric gas approximation using γ = 1.4, the hypersonic limit for the density ratio is 6. However, hypersonic postshock dissociation of O_{2} and N_{2} into O and N lowers γ, allowing for higher density ratios in nature. The hypersonic temperature ratio is:
See also
 Bow shock (aerodynamics)
 Gas dynamics
 Mach reflection
 Moving shock
 Shock wave
 Shock polar
References
 Liepmann, Hans W.; Roshko, A. (2001) [1957]. Elements of Gasdynamics. Dover Publications. ISBN 0486419630.
 Anderson, John D. Jr. (January 2001) [1984]. Fundamentals of Aerodynamics (3rd Edition ed.). McGrawHill Science/Engineering/Math. ISBN 0072373350.
 Shapiro, Ascher H.. The Dynamics and Thermodynamics of Compressible Fluid Flow, Volume 1. Ronald Press. ISBN 9780471066910.
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