Testing and CFD Simulation of Diaphragm Hybrid Rocket Motors

Bettella, Alberto; Lazzarin, Marta; Bellomo, Nicolas; Barato, Francesco; Pavarin, Daniele; Colombo, G.; Grosse, Matthias

doi:10.2514/6.2011-6023

Cited by 20 publications

(13 citation statements)

References 5 publications

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“…Pressure-fed systems are more interesting in the context of CISAS activities in the past and the near future. CISAS has been testing hybrid rocket motors since 2007 in pressure regulated and self-press mode mainly using N 2 O as oxidizer and paraffin as fuel 19,20,21 . Paraffin-based hybrid rockets have a great potential for small launcher applications.…”

Section: Introductionmentioning

confidence: 99%

Numerical Investigation of Hybrid Rocket Combustion Chamber Transient Behavior

Barato

Faenza

Pavarin

et al. 2013

49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference

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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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Section: Introductionmentioning

confidence: 99%

Numerical Investigation of Hybrid Rocket Combustion Chamber Transient Behavior

Barato

Faenza

Pavarin

et al. 2013

49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference

Self Cite

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“…However, the use of hot gas is hazardous in combination with N 2 O, so that a different expulsion method is necessary. There are methods to enhance the fuel-regression rate, which will help increase gas generator compactness by a reduction of the grain length and complexity [27] [34]. However, if carbon-reinforced resin material and a compact pressurization system is used, an N 2 O-based hybrid gas generator may also become a competitive design in terms of mass and volume.…”

Section: Conclusion/outlookmentioning

confidence: 98%

Use of Hybrid Rocket Technology for Submarine Emergency Deballasting

Grosse¹,

Bettella²

2011

47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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One of the most effective measures for a rapid return of military submarines to the surface in case of life threatening, hazardous or potentially catastrophic failures is the emergency blowing of the forward main ballast tanks. This can be achieved with various systems. Gas generator based modular emergency blow systems show several advantages compared to designs, which use highly compressed air in terms of specific gas work, mass flow and gas storage density. Operational gas generators for emergency blowing use either a liquid hydrazine monopropellant or a solid propellant consisting of GAP and strontium nitrate for gas generation. An alternative hybrid propellant gas generator is proposed to avoid safety issues, which both operational gas generators types possess. Also the relative free selection of the blow gas composition motivates the use of a hybrid propellant combination. A mass model for a hybrid gas generator is developed using design elements of the hydrazine gas generator and applied to a number of different hybrid propellant combinations. The hybrid propellant mass is determined by use of the assumption that the volume reduction due to heat losses, water condensation and gas dissolution in the ballast tank can be compared for a blow gas that is generated from a solid or a hybrid propellant. The effectiveness of the modeled hybrid gas generator is assessed amongst others by blow system mass, system volume and required gas generator quantity. Comparing these preliminary results to the operational solid and liquid propellant gas generator shows that even a nonoptimized hybrid propellant gas generator system is very probably competitive in terms of mass and volume. Nomenclaturea = regression rate constant p t = tank pressure ADN = ammonium dinitramide R = universal gas constant c* = characteristic velocity r ɺ = fuel regression rate GAP = glycidyl azide polymer T c = combustion temperature G ox = oxidizer mass flux T gas = gas temperature HAN = hydroxylammonium nitrate V ba = ballast tank volume HAP = hydroxylammonium perchlorate V gas = gas volume ∆H f = molar enthalpy of formation v gas = specific gas volume HTPB = hydroxyl-term. polybutadiene V i = initial volume IRFNA = inhibited red fuming nitric acid W gas = gas work M = molar mass w gas = specific gas work 2 m c = combusted mass α = mass flux exponent m cool = water coolant mass Φ = average O/F (oxidizer-to-fuel) ratio m gas = gas mass ρ f = final density m i = initial mass ρ fu = fuel density m pro = propellant mass ρ i = initial density p a = ambient pressure η blow = blowing efficiency p ba = pressure in ballast tank η c* = combustion efficiency p c = chamber pressure gas = gas p f = final pressure exp = experimental p i = initial pressure th = theoretical

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“…The data compared between the numerical tools are pressure and C* (characteristic velocity), and the reference values of these parameters are the measurements collected during an experiment performed in summer 2011 at CISAS 15 .…”

Section: B Preliminary Simulations: Comparison With Commercial Cfdmentioning

confidence: 99%

Numerical Investigation of Hybrid Motors for the EU FP7 SPARTAN Program

Lazzarin¹,

Bellomo²,

Faenza³

et al. 2012

48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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The SPARTAN (SPAce exploration Research for Throttleable Advanced eNgine) project is targeted at the development of a soft-lander demonstrator for Mars landing, based on throttleable hybrid rocket engine propulsion technology. Within the project, CISAS "G. Colombo" is in charge of the development of a new advanced CFD code for the hybrid rocket motor unsteady simulation, of other CFD simulations conducted with commercial software to investigate the fluid dynamics of the rockets developed by NAMMO, and of ground fire tests to be conducted at lab-scale. Thus, for the CFD simulations, an Open Source code (OpenFOAM) and a commercial code (Ansys CFX) have been used. The flow field has been investigated in its different aspects, to understand the physics behind the processes of injection and combustion. The main approximations applied to this study when analyzing the flow with commercial codes are steady state analysis and fixed chemistry in the combustion chamber. Both vortex and axial injection have been simulated by means of the CFD: an axial injector has been used for the CISAS lab-scale rocket motor and a vortex injector for the test cases reproducing NAMMO's configurations. Regarding vortex injection, the results highlight that it improves the turbulent mixing between the reactants and pushes the flame near the wall, enhancing the heat flux transmitted and thus the regression rate, thanks to its strong helical flow and centrifugal effects. Further efforts have been oriented to take into consideration regression rate dependency on the wall heat flux: besides injecting a fixed and pre-established fuel mass flow, the latter has been coupled to heat exchange. This study is in a preliminary phase, but the first results show that CFD regression rate as a function of the oxidizer mass flux follows the same trend as predicted by:, with a and n experimentally estimated by NAMMO. Concerning OpenFOAM numerical solver, it provides a transient solution of the flow field, but uses corresponding boundary conditions compared to commercial software, in order to allow a full comparison of the steady state results obtained. The study performed with OpenFOAM is a work in progress. The starting point has been a comparison with the results obtained using commercial CFD, to validate the open source solver and boundary conditions. This comparison has showed that OpenFOAM can predict the experimental results as well as commercial software does. At the moment, a feasibility study is being conducted in order to develop a custom and flexible boundary condition, able to calculate regression rate as a function of the wall heat flux independently on the specific oxidizer/fuel combination. Other studies are also necessary to limit the computational time required by this open source tool.

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Testing and CFD Simulation of Diaphragm Hybrid Rocket Motors

Cited by 20 publications

References 5 publications

Numerical Investigation of Hybrid Rocket Combustion Chamber Transient Behavior

Numerical Investigation of Hybrid Rocket Combustion Chamber Transient Behavior

Use of Hybrid Rocket Technology for Submarine Emergency Deballasting

Numerical Investigation of Hybrid Motors for the EU FP7 SPARTAN Program

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