By Brian J. Cantwell
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Additional resources for Aircraft and Rocket Propulsion
58) The Mach number entering the afterburner is fairly low and so the stagnation pressure ratio of the afterburner is fairly close to (but always less than) one. Station 7 - The entrance to the nozzle. Station 8 - The nozzle throat. Over the vast range of operating conditions of modern engines the nozzle throat is choked or very nearly so. Station e - The nozzle exit. 60) In general the goal of the designer is to minimize heat loss and stagnation pressure loss through the inlet, burner and nozzle.
5 = 8 , A 1 = A 3 = A 4 and A 4 ⁄ A e = 3 . Determine the dimensionless thrust T ⁄ ( P 0 A 1 ) . Do not assume f<<1. Neglect stagnation pressure losses due to wall friction and burner drag. Assume that the static pressure outside the nozzle has recovered to the ambient value. 5 = A 3 . By what proportion would the air mass flow change? Solution - The first point to recognize is that the stagnation pressure at station 4 exceeds the ambient by more than a factor of two - note the pressure outside the nozzle is assumed to have recovered to the ambient value.
7054 . 004 . 2316 . 494 . 72 . 0724 State II - Now increase the inlet throat area to the point where the inlet unchokes. As the inlet throat area is increased the Mach number at station 3 will remain the same since it is determined by the choking at the nozzle exit and the fixed enthalpy rise across the burner. 5 is fixed by the loss across the external shock. 64). 64) is maintained and the inlet shock moves to the left increasing P te . 5 and the nozzle exit is across the burner.
Aircraft and Rocket Propulsion by Brian J. Cantwell