Paraffin-based hybrid rockets offer a great potential towards a green, safer, cheaper and more reliable access to space. As for liquids, the pressurization system has a fundamental impact on hybrid rocket motor performances. In particular, unlike liquid rockets, the oxidizer to fuel ratio cannot be directly controlled in a hybrid motor but it is dependent on the complex coupling between oxidizer mass flow (linked to pressurization) and chamber behavior (fuel regression). Pressure-fed circular port hybrid rockets are attractive for their perceived simplicity. In this paper several solutions for the pressurization system of paraffin-based hybrid rocket motors are investigated. A numerical model has been developed in order to determine the main performance parameters of the hybrid motor with time. For this purposes the prediction of oxidizer and fuel mass flows, tank and chamber pressures, thrust and residual gas in the tank is obtained through the modeling of the principal subsystem's behavior. The lumped parameter code is composed by three sub-model linked together: the combustion chamber, the oxidizer tank and the pressurant tank. In the first part of the paper several solutions are investigated like the blowdown mode, the pressure regulated mode, the use of a cavitating venturi, the use of single and multiple orifices, the use of a digital valve, the heating of the pressurant and finally the eventual combustion of the pressurant. For every technique the main aspects/issues are highlighted. In the second part an equivalent model for self-pressurization is presented. It is shown that if proper designed, self-pressurization is a simple, lightweight and high performing solution. However, because of its temperature sensitivity, for optimal performance a good thermal control is required.
Nomenclatureregression rate law coefficients A = area D = diameter c v,p = specific heats (at constant volume, pressure) e = specific energy E = energy = throat erosion rate g = gravitational acceleration G = mass flux h = specific enthalpy = mass flow m = rocket mass M = mass L = length = regression rate O/F = oxidizer to fuel ratio c* = characteristic velocity 1 Ph.D. student, University of Padua, CISAS G. Colombo, francesco.barato@studenti.unipd.it, Student Member Joint Propulsion Conferences 2 = expansion ratio = density T = thrust, temperature V = volume V a = valve position p = pressure = heat flow = heat transfer coefficient R = gas constant = ratio of specific heats = efficiency = time constant x = vapor mass fraction s = specific entropy S = entropy v = velocity = tank mass factor subscripts a = ambient c = combustion chamber cv = cavitating venturi ev = evaporated i = initial, interface inj = injection f = final, frontal l = liquid n = nozzle prop = propellant ox = oxidizer fuel = fuel vap = vapor p = port pt = pressurant tank press = presurant s = exit t = tank, throat v = vapor, valve or orifice w = wall