By Brian J. Cantwell
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Extra resources for Aircraft and Rocket Propulsion
4 , C p = 1005 M2/(sec2-°K). The gas constant is R = 287 M2/(sec2-°K). The ambient temperature and pressure are T 0 = 216K 4 2 P 0 = 2 × 10 N ⁄ M . 28 × 10 J/kg . 0. A normal shock stands in front of the inlet. 5 3 P0 4 e M0 = 3 shock and the stagnation temperature at station 4 is T t4 = 2000K . 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.
44) The pressure in the engine is likely to be very high at this free stream Mach number and so the nozzle is surely choked and we can write P te A e 1 γ m˙ a = ------------------ ------------------------------------- ------------------- γ+1 (1 + f ) -------------------- γ RT te 2 ( γ + 1 γ – 1 ) ----------- 2 . 45) The thrust equation is Ue Ae Pe 2 T ------------- = γ M 0 ( 1 + f ) ------- – 1 + ------ ------ – 1 U0 P0 A0 A0 P0 . 46) Our main concern is to figure out what happens to the velocity ratio and pressure ratio as we control the fuel flow and nozzle exit area.
Thus there is no net mass increase or decrease to the system. Note that there is no assumption that the compression or expansion processes operate isentropically, only that the exhaust is fully expanded. 8 SPECIFIC IMPULSE, SPECIFIC FUEL CONSUMPTION An important measure of engine performance is the amount of thrust produced for a given amount of fuel burned. 41) with units of seconds. The specific fuel consumption is essentially the inverse of the specific impulse. 42) The specific fuel consumption is a relatively easy number to remember of order one.
Aircraft and Rocket Propulsion by Brian J. Cantwell