# How can I determine specific heats of gases for turbojet engine calculations?

I'm in a combustion lab and we are performing experiments on a turbojet engine. However, I am having some trouble calculating the exit velocity and compressor/turbine powers of the engine.

Consider the nozzle exit, where the exhaust gases go. Assuming 100% adiabatic efficiency of the nozzle and applying the 1st law of thermodynamics gives $$\frac{u_e^2}{2} = h_{0e}-h_e = c_{p,N}(T_{0e}-T_e)=c_{p,N}T_{0e}[1-(p_a/p_{0e})^{(\gamma_n - 1)/\gamma_n}],$$ assuming $p_e = p_a$, the ambient pressure. In the lab, we've set up thermocouples that measure the stagnation temperature and pressure at the exit, as well as ambient conditions. My problem, however, is that I have no idea where to get $c_{p,N}$, the specific heat of the gases in the nozzle from. I've looked up in a reference book that the average specific heat ratio $\gamma_n = 1.36$, but this still doesn't give me the specific heat that I need because I don't know the specific gas constant $R$ for these gases. For reference, it is an SR-30 turbojet with Jet-A fuel.

In addition, I have the same problem when calculating the power outputs. Assuming that all of the power of the turbine is transmitted to the compressor, $$\dot{W_T} = (\dot{m_a} + \dot{m_f})c_{p,T}(T_{04}-T_{05})=\dot{m_a}c_{p,C}(T_{03}-T_{02}) = \dot{W_C}$$ Again, I have no clue what the specific heats are, only their ratio. Is there some way I can calculate these quantities or look them up somewhere? Or even find the specific gas constants somewhere?