By A. Roy
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Extra resources for A First Course on Aerodynamics
Like transonic flow, hypersonic flow is also highly non linear in nature. These flows are found to occur on reusable space shuttles during atmospheric reentry, long range ballistic missiles which fly through the outer extent of the atmosphere at hypersonic speeds, on rockets at high altitudes, etc. Flow characteristics change significantly as Mach number becomes large which are briefly discussed below. Shock waves are formed closer to the body surface than at supersonic speeds. The thin flow region formed between the shock and the body surface is called as a thin shock layer.
Some interesting facts about hypersonic flows are revealed if we try to impose the limit of M ∞ → ∞ in exact relations for normal and oblique shocks. For example if we apply this limit to the exact normal shock relations we have the following findings: The above relations show that under the limiting condition the downstream Mach number and density ratio attain finite values. However some other ratios can become infinitely large. 4 and q is small. The above relation reiterates the fact that the shock layer is thin at hypersonic speed.
3. Sonic flow (M = 1) Sonic velocity can only be attained in the stream tube, if dA vanishes locally. According to this result supersonic flow can only be generated, if the stream tube has a convergent-divergent distribution of the cross section (de Laval nozzle). The cross section with minimum area, where the local velocity is equal to the speed of sound, is called critical cross section or throat. 37) tells us that Mach number at any section of the nozzle is a function of the area ratio, namely (A/A*), which is the ratio of the local duct area to the sonic throat area.
A First Course on Aerodynamics by A. Roy
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