MAE 113
–
Fundamentals of Propulsion
Homework Assignment 4
Winter, 2011
Due: Thursday, February 10
th
in Class
Page 1 of 2
Problem 1
Consider the flow through a rocket engine convergingdiverging nozzle shown in figure 1. Assume that
the gas flow through the nozzle is an isentropic expansion of a calorically perfect gas. In the combustion
chamber, the gas which results from the combustion for the rocket fuel and oxidizer is at a pressure and
temperature of 15 atm (abs) and 2500 K, respectively; the molecular weight and specific heat at constant
pressure of the combustion gas are 12 g/mol and 4157 J/kg K, respectively. The gas expands to
supersonic speed through the nozzle, with a temperature of 1350 K at the nozzle exit.
(a)
Assuming incompressible flow, calculate the pressure at the exit.
(b)
Assuming isentropic flow, calculate the Mach number and velocity at the exit.
P
0
= 15 atm
T
0
= 2500 K
T
e
= 1350 K
Figure 1: Diagram for problem 1
Problem 2
Air flows isentropically in a channel. At section (1) the Mach number is 0.3 and the area is 0.001 m
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 Spring '08
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 Fluid Dynamics, Mach number

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