REGULUS is an Iodine-based electric propulsion system. It has been designed and manufactured at the Italian company Technology for Propulsion and Innovation SpA (T4i). REGULUS integrates the Magnetically Enhanced Plasma Thruster (MEPT) and its subsystems, namely electronics, fluidic, and thermo-structural in a volume of 1.5 U. The mass envelope is 2.5 kg, including propellant. REGULUS targets CubeSat platforms larger than 6 U and CubeSat carriers. A thrust T = 0.60 mN and a specific impulse Isp = 600 s are achieved with an input power of P = 50 W; the nominal total impulse is Itot = 3000 Ns. REGULUS has been integrated on-board of the UniSat-7 satellite and its In-orbit Demonstration (IoD) is currently ongoing. The principal topics addressed in this work are: (i) design of REGULUS, (ii) comparison of the propulsive performance obtained operating the MEPT with different propellants, namely Xenon and Iodine, (iii) qualification and acceptance tests, (iv) plume analysis, (v) the IoD.

A study on vortex injection in hybrid rocket engines with nitrous oxide and paraffin has been performed. The investigation followed two paths: first, the flowfield was simulated with a commercial computational fluid dynamics code; then, burn tests were performed on a laboratory-scale rocket. The computational fluid dynamics analysis had the dual purpose to help the design of the laboratory motor and to understand the physics underlying the vortex flow coupled with the combustion process compared with axial injection. Vortex injection produces a more diffuse flame in the combustion chamber and improves the mixing process of the reactants, both aspects concurring to increase the c efficiency. A helical streamline develops downstream of the injection region, and the pitch is highly influenced by combustion, which straightens the flow due to the acceleration in the axial direction imposed by the temperature rise. Experimental tests with similar geometry have been performed. Measured performance shows an increase in regression rate up to 51% and a c efficiency that rises from less than 80% with axial injection up to more than 90% with vortex injection. Moreover, a reduction of the instabilities in the chamber pressure has been measured. Nomenclature A t= nozzle throat area a = multiplication coefficient for regression rate G ox = mass flux in combustion chamber L g = grain length M f = mass of burned fuel M m = molar mass _ m tot = mean total mass flow n = exponential coefficient for regression rate O∕F = oxidizer to fuel mass ratio p, pc = pressure, mean chamber pressure R u = universal gas constant r = radial coordinate T = temperature t = time u r , u z , u θ = radial, tangential, and axial velocity z = axial coordinate μ = dynamic viscosity ρ f = density of fuel ρ = density ϕ i , ϕ e , ϕ m = initial, final, and mean diameter of the grain ω = vortex angular velocity

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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

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