A LOX/LH2 rocket engine with a chamber pressure of 25 atm is optimized with an isentropic nozzle that is perfectly expanded at ground with a thrat area of 0.065 m^2. The fuel and oxidizer is fed into the non-calorically perfect combustion chamber with a stoichiometric mixture ratio at 250 K. You may assume that O2 and H2 do not dissociate, gamma for the products is 1.2, and that the specific heat curve fits are still valid beyond its range. Hint: the specific gas constant is a function of the degree of reaction found using chemical equilibrium.

a) Find the Isp and thrust produced by the rocket engine.

b) Discus qualitatively how the following values will be affected under the following conditions:

1) The thrust, assuming the nozzle is not calorically perfect

2) The flame temperature, assuming O2 and H2 dissociates

3) The mass flow rate, if operating at a higher altitude

4) The exit velocity, if using a fuel rich mixture.