Our heavyweight helicopter equal in the world does not have
In Rostov started production of the most load-lifting rotary-wing car The Russian holding «Helicopt[...]
Summary of Chapter 4
In this chapter, the effects of compressibility, associated with the changing Mach number in inviscid flow, are reviewed. Thin airfoil and small disturbance assumptions are retained.
At low Mach numbers, 0 < M0 < 0.3, the flow is treated as incompressible flow.
Compressibility effects in the range 0.3 < M0 < 0.7 correspond to subsonic compressible flow and the Prandtl-Glauert rule can be applied, in the absence of shock wave, to evaluate forces and moments by solving the incompressible flow past the thin profile, then the resulting aerodynamic coefficients, Cp, Ci and Cm, o,
are magnified by the multiplying factor 1 /^J1 – M2 The drag coefficient remains Cd = 0.
Above M0 = 1.3, the flow is expected to be essentially supersonic and the linear supersonic theory and Ackeret results, are applicable and yield closed form expressions for the coefficients Cp, Cl and Cm, o.An inviscid wave drag is generally present with Cd > 0. The aerodynamic coefficients depend on Mach number via the multiplication factor 1 /УМ2-!
NearMachone, intherange0.7 < M0 < 1.3, mixed-type flow is experienced with regions of subsonic and supersonic flows coexisting within the flow field. This is the transonic flow regime and the subsonic compressible and linear supersonic theories
are not applicable and yield unphysical results at M0 = 1. The nonlinear transonic small disturbance (TSD) equation bridges the gap between the two linear theories. The derivation of the TSD equation is carried out and characteristic lines and jump conditions associated with a hyperbolic equation are studied. Numerical solution of the TSD equation was needed and the breakthrough came with the Murman-Cole four-operator mixed-type scheme, which is discussed in some detail. The pressure coefficient is obtained from the numerical solution and provide the lift and pitching moment coefficients. The wave drag is shown to reduce to an integral over the shock lines impinging on the profile and hence vanishes in their absence. A rare exact solution of the TSD equation, which represents the near-sonic throat flow in a nozzle, is used to construct numerically subsonic-to-supersonic nozzles producing a uniform supersonic exit flow.
The final part of the chapter is devoted to the discussion of simple wave regions of supersonic flow adjacent to a uniform flow region.
