Perform a preliminary design of a 2-stage solid rocket to lift an 4,000 kg payload to LEO. The composite propellants are 17% by mass Al fuel, 67% AP oxidizer and 16% HTPB binder, ignore the other minor components. Use Ae/A*=18 and a combustion chamber pressure of 89 atm. The rocket stages are similar (same and ) with =0.11. Display calculations in excel. List your assumptions and all equations used in your submittal. The propellant density is 1,860 kg/m^3. a) Run CEA to find T02, C*, CF, gamma, Cp and any other needed information. Assume the nozzle mixture is in equilibrium. Input the fuel and binder as fuels, you need to specify the relative percent of each. Use AL(cr) for the fuel and for HTPB binder use C4H6,butadiene. Use NH4CLO4(I) for the oxidizer. You need to calculate the O/F: Oxidizer/Fuel Wt. ratio to enter into the code. the website to use is https://cearun.grc.nasa.gov/ b) Determine the M0, Ms and Mp for both stages and . c) Find the propellant volume for both stages. d) Select A* to give takeoff thrust 2.9 times the takeoff weight. e) Assuming the propellant is a solid cylinder that is the same diameter as the nozzle exit, find the height of the propellant in each stage. f) Find the length of the nose cone with a 9 degree half angle. Assume the payload fits inside the nose cone what is the total length of the rocket neglecting the structure. g) For the r=O/F used find the products for complete combustion, compare to the CEA results.

I mostly need assistance with part e, f & g.