Characterization of Combustion Performance of Hydrocarbon Fuels using a Bi-Propellant Rocket Engine

Abstract

An experimental investigation of the effect of fuel-to-oxidant ratio and reactant composition on the performance of a lab-scale gaseous rocket engine was performed. The senior design project involved the design, fabrication, and testing of a laboratory scale gaseous reactant rocket engine. A simplified zeroth order combustion model was developed to provide guidance for the design of the rocket system. Conservation equations along with chemical equilibrium results were the fabric of the model development. Peak chamber pressures as a function of reactant delivery rates and nozzle diameters were the primary output of the code. Experiments were conducted as a function of chamber pressure and equivalence ratio. For an equivalence ratio of 1, CH 4 mass flow of 1.26 g/s and oxygen flow rate of 4.48 g/s, the measured chamber pressure was nearly 160 psig having a combustion efficiency of 85%. Theoretical pressure-time profiles were compared to empirically-obtained profiles and the peak pressures agreed well. Combustion efficiencies across a broad range of equivalence ratios were found to be approximately 80%±6. Nomenclaturec = chamber = mass flow rate [ kg/s] C* th = theoretical characteristic velocity [m/s] η = efficiency A t = throat area [m 2 ] st = stoichiometric P = pressure [psi] V = volume [m 3 ] R = gas Constant [kJ/kg-K] T = temperature [K] G ox = oxidizer mass flow rate [g/s] O/F = oxidizer-to-fuel ratio = equivalence ratio D c = chamber diameter [m] L c = chamber length [m] L* = characteristic chamber length [m] MW = molecular weight [kg/kmol] = specific heat ratio function C d = discharge coefficient t = time [s]