By Brian J. Cantwell

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**Extra resources for Aircraft and Rocket Propulsion**

**Sample text**

5) implies that f ( M 4 ) must increase and the Mach number downstream of the burner decreases. There is a limit to the amount of heat that can be added to this flow and the limit occurs when f ( M 4 ) attains its maximum value of one. At this point the flow looks like the following. 6 Step 3 - Introduce sufficient heat to bring the exit Mach number to a value slightly greater than one. 7) before unstart Now suppose the temperature at station 4 is increased very slightly. We have a problem; T t4 is up slightly, P t4 is down slightly but f ( M 4 ) cannot increase.

This will only effect the flow in the inlet and all flow variables in the rest of the engine will remain the same. 7 3/21/11 Ramjet flow field With the flow in the engine subsonic and the shock positioned at the end of the diffuser we have a great deal of margin for further heat addition. 5) is still preserved and the exit Mach number remains one. Let the burner outlet temperature be increased to T t4 = 2100°K . The flow now looks something like this. 8 Step 5 - Increase the heat addition to produce some thrust.

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. 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.

### Aircraft and Rocket Propulsion by Brian J. Cantwell

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