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PRELIMINARYDESIGNOFA30kNPARAFFIN-BASEDHYBRIDROCKETENGINE

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PRELIMINARY DESIGN OF A 30kN PARAFFIN-BASED HYBRID ROCKET ENGINE

Raffaele Votta1, Marco Di Clemente2,Giuliano Ranuzzi3, Lorenzo Pellone4 Italian Aerospace Research Centre, Capua, 81043, Italy

One of the objectives of the HYPROB program is the development of tools, design and manufacturing methodologies applicable to hybrid systems and the verification of these technological developments through the realization and test of a 30 kN rocket engine. The technological demonstrator is based on nitrous oxide (N2O) as oxidizer and paraffin as propellant and will have most attractive capabilities of hybrid systems compared to solid or liquid engines, as throttability and re-ignition. The project targets to achieve a TRL of enabling technologies for hybrid propulsion equal to 5. The paper reports the study logic defined to achieve the goals of the project and the status of the technical activities related to the preliminary design of the demonstrator, based on a design tool developed ad-hoc.

Nomenclature

ACE = Aerotherm Chemical Equilibrium CMA = Charring Material Ablation HRE = Hybrid Rocket Engine Isp = Specific Impulse M = Mach Number O/F = Oxidizer to fuel ratio P = Pressure r = Regration Rate RCS = Reaction Control System T = Temperature Subscripts C = Combustion chamber e = Exit

I. Introduction he development of propulsion systems based on hydrocarbons, as methane, or hybrid is becoming a technology asset for launchers and new generation space transportation systems, through the development of demonstrators

and several research programs focused on the main critical aspects [7]. Methane is one of the most interesting choice regarding liquid rocket engines since it is able to guarantee good

performance of the propulsion system, it can be easily stocked, it is characterized by a low level of risk and no toxicity; moreover production costs are rather low.

On the other hand, hybrid rocket engines, based on LN2O/Wax or LO2/Wax, combine, in principle, some of the advantages of a solid rocket (above all, easiness in storage) and liquid (as performances) even though the use of hybrid engines is not yet optimized to be used in operational flight conditions.

1 PhD, Researcher, Aerothermodynamics Group, Via Maiorise, r.votta@cira.it 2 PhD, Researcher, Fluid Mechanics Group, Via Maiorise, m.diclemente@cira.it 3 PhD, Researcher, Aerothermodynamics Group, Via Maiorise, g.ranuzzi@cira.it 4 Researcher, Aeronautics Group, Via Maiorise, l.pellone@cira.it

T

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Hybrid technology can be applied, in principle, in all rocket propulsion systems of launchers and space vehicles, from large size engines for first stage of launchers of 200 t thrust to medium size engines for second and third stages of 10 t thrust, or small size engines for RCS.

Hybrid propulsion is receiving more and more attention from the space community because it is able to conjugate operative simplicity, and therefore guarantee reliability, of solid propulsion systems and the versatility and performances of liquid propulsion systems, for the aspects related to thrust throttability and re-ignition capabilities. In the last decade, this technology has been developed due to the necessity of cost reduction (related to storage, production and safety) and the tendency to substitute toxic propellants as MMH-NTO or hydrazine (also due to the restrictive legislation recently approved in Europe), and to maintain, at the same time, a propulsion system able to guarantee thrust modularity and multiple ignitions. All those aspects have definitely renewed the interest of the space community in the hybrid rocket engines.

The main advantages of hybrid propulsion, at least from a potential point of view, can be summarized as follows: 1) thrust throttability; 2) re-ignition or multiple ignitions to guarantee adaptive capabilities to different mission conditions and orbiting

manoeuvres; 3) robustness of design and its higher reliability respect liquid propulsion systems due to the limited use of

turbo-pumps or other moveable devices; 4) possible higher performances respect solid propulsion systems; 5) reduced environmental impact. Despite these potential advantages, and despite the base knowledge have been widely developed in the past

years, the development of hybrid propulsion systems is not yet fully exploited due to the technical difficulty related to the low maturity of some enabling technologies, but also due to a certain reticence to change developed design methodologies mostly based on liquid or solid propulsion systems. Moreover, hybrid technology has also some disadvantages that need to be addressed before considering this possibility as real alternative to liquid or solid propulsion. The main drawbacks are low fuel regression rate which requires complex grain geometries and low combustion efficiency related to the different physical state of propellants, to the physics of combustion and to oxidant injection system.

Strategic objectives of the HYPROB program have been defined in [1] in agreement with national stakeholders, such as Italian Space Agency and aerospace industries. In order to pursuit those objectives, different projects have been defined and in particular, regarding the development of the hybrid rocket engines, the objective is to develop design capabilities of a hybrid engine, based on N2O/Wax, as technology demonstrator of national research and developments activities [2]. Paraffin wax overcomes one of the disadvantages of HRE concerning their low burning rates, since it forms a thin, hydro-dynamically unstable liquid layer on the melting surface of the fuel, enhancing the regression rate with entrainment of droplets from the liquid-gas interface [3,4]. Among the paraffin wax, the microcrystalline formulation, being composed of a predominance of non-linear hydrocarbons, has the advantage to form microcrystals, which guarantee a higher tenacity and then a better workability in comparison to the linear paraffin waxes.

Starting from these objectives and constrains, the resulting technical requirements for the Hybrid project are: - design, manufacturing and testing of a technological demonstrator, to achieve a TRL equal to 5; - technology development on alternative oxidizers, as O2 and H2O2, in order to achieve a TRL of relevant sub-

systems equal to 4; - development of a numerical platform, to be used in the demonstrator test rebuilding, to increase design tools

capabilities [5]. Hybrid rocket engine technological demonstrator, whose preliminary design will be presented in the present

work, will have an initial thrust of 30 kN, and it will be based on nitrous oxide and paraffin; the most attractive capabilities of hybrid systems compared to solid or liquid engines, namely throttability, with a ratio 3:1 within two limiting thrust values, and multiple ignition, will be also implemented.

At the same time, high performance solid propellant, based on paraffin, is being developed within the project in order to define a proper formulation able to achieve the requested value of regression rate, to match the performance requirements of the demonstrator, and to achieve suitable resistance and mechanical properties to manufacture a large scale grain [6].

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II. HRE DEMONSTRATOR SYSTEM DESIGN

A. GRAIN DIMENSIONING The overall configuration of the hybrid rocket engine has been carried out by considering the following main

requirements: initial thrust equal to 30 kN, thrust throttability with a ratio 3:1, multiple ignitions, ignition time equal to 30s. In this section the status of the design of the demonstrator will be reported. The trade-off on the grain dimension has been made by considering the high level requirements in terms of performances of the engine, afterword the design of the main sub-systems has been carried out based on engineering correlations and numerical tools.

A typical architecture for the thruster has been considered, with liquid oxidizer and solid fuel presented in Figure 1 and a cylindrical port has been chosen. The rocket is constituted by a solid grain enveloped in an external metallic case, the injection head providing the injection of the oxidizer into the combustion chamber at suitable conditions, and the expansion nozzle. Along the feeding line, connecting the thrust chamber to the tank, oxidizer regulation valve for mass flow rate control are located.

Figure 1 - General arrangement of the hybrid rocket

The 30 kN demonstrator has been preliminary sized using an in-house developed design tool composed by two

modules: trade-off module, which carries out the preliminary sizing of the solid grain by varying geometriacal and operational parameters of the engine, and performance module which evaluates the performances of the selected configuration. The procedure that has been definedis schematically shown in Figure 2: the first step of the numerical procedure of the trade off module relies with the interface Rocket Propulsion Analysis [8] used to evaluate theoretical 0-D rocket performance, optimal mixture ratio to maximize the specific impulse. Trade off module is able to analyze several configurations by varying grain length, port diameter, oxidizer mass flow rate, regression rate model and chamber pressure and to evaluate for each of them the main performance parameters (Thrust, Isp, Tc, Pc, O/F, etc.). As the mass flow rate is updated with the grain consumption (taken into account by implementing different regression rate models) the tool updates the main parameters with time by reading a proper external database.

Figure 2 - Design and performance tool procedure

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Once a geometric configuration has been selected, the performance module is used to analyse the performance of the engine with time with proper graphical outputs. Engine geometries file output, including the expansion nozzle, are also provided by the tool. Moreover, the module is able to take into account nozzle ablation for performances prediction.

Moreover, the developed tool is able also to design nozzle profile with the following features: Convergent c ranging from 18° to 45°; parametric variation of convergent length; spline contour with continuity on first derivatives criterions with throat and chamber regions (second

derivative continuity taken into account in Ansys post-processing. Divergent d ranging from 12° to 18°; throat region; contour with continuity of first derivative criterions with divergent and convergent regions (second derivative

continuity taken into account in Ansys post-processing). Once selected the propellants, from RPA8 computations the ideal performance parameters are available. From

thrust requirement (i.e., 30 kN), total and, of course, fuel and oxidizer mass flow rates are known, considering an optimal O/F ratio. For what concerns the regression of the fuel grain a firing time 30 s + 5s as margin has been considered and Grosse model [10,16] has been taken into account.

More than 600 configurations have been analyzed by varying chamber pressure, grain length and port diameter. Figure 3 and Figure 4 show thrust and Isp, respectively, as function of grain length for different initial port diameter. Initial chamber pressure has been assumed equal to 40 bar, lower than the N2O vaporization pressure (about 50 bar at T=293 K) but adequately high to reach the designed thrust.

The configuration has been chosen to maximize Isp, therefore the following parameters have been selected: grain length equal to 0.8 m; internal port diameter equal to 0.1 m; external diameter equal to 0.35 m.

The performed choice of length is, moreover, in the region diagram (see Figure 3) where the variation of thrust with port dimeter is negligible (about 1%), therefore as the firing time increases the thrust is constant with the grain consumption and, as the nominal thrust is not effected by re-ignition, at least from a theoretical point of view.

Figure 3 - Thrust vs grain length.

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Figure 4 - Isp vs grain length

As clear from the previous diagrams, there is a certain region in terms of grain length where performances are

quite insensitive to port diameter, that means constant along the burning time. This is is part due to the selected regression model due to Grosse, experimentally determined for the same paraffin considered in the present work, but also to the fact that for these values some different effects on engine performances vary and partially compensate each other. It is know that in hybrid engines, O/F ratio decreases with time due to the increasing of grain port diameter and the decreasing of regression rate, and therefore fuel mass flow, while oxidizer mass flow is often kept as constant. This means that combustion paramaters, as chamber temperature, thrust coefficient and characteristic velocity, also vary during the burning time and follow the theoretical values shown in Figure 5; the maximum of the three curves reported in this figure takes place at different values of O/F ratio therefore there is a range of O/F variation where the variation of chamber temperature, thrust coefficient and characteristic velocity partially compensates to provide an almost constant thrust and specific impulse.

Figure 5 - Combustion parameters vs. O/F ratio

The present configuration has been analyzed in terms of performance parameters as function of burning time. In

particular, as already underlined, a negligible decrease (about 1%) of thrust and chamber pressure is predicted during the burning time (see Figure 6 and Figure 7) and also specific impulse is almost constant as shown in Figure 8). It is worth to underline that present results do not consider throat regression due to ablation that will cause a stronger decrease of chamber pressure with time and a consequently decay of all performance parameters.

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Figure 6 - Thrust vs time

Figure 7 - Chamber pressure vs time

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Figure 8 - Isp vs time

B. HRE DEMONSTRATOR SUB-SYSTEMS 1. Injection system The design of injection system has been carried out considering different possible design solutions as pintle

injector or rotating plates [11]. The final configuration considers a fixed injector geometry and un upstream flow control valve and it has been selected through the Analytical Hierarchy Method to consider the aspects related to functionality, technology innovation, reliability, realization costs. Traditional rocket propellants are usually used as sub-cooled due to their low vapor pressure. Assuming that they behave as incompressible liquids in the injection system, the steady state Bernoulli equation can be safely used to properly design the injector. In particular, when the discharge coefficient is well characterized and the mass flow rate is imposed, the evaluation of the injection area is trivial. Unfortunately, this method is not applicable to injectors of hybrid engines that use high vapor pressure oxidizers such as nitrous oxide which vapor pressure, at a temperature of 293 K is about 50 bar. Indeed, this simple relation cannot take into account compressibility effects and the development of two-phase flow within the orifice. In the operating conditions considered for the present analysis, and for typical injection geometries, as the bulk downstream pressure drops below the vapor pressure, a significant vapor formation occurs within the injector. Simultaneously, a maximum flow rate is achieved and the mass flow rate chokes. This regime is known as critical flow since the mass flow rate is independent of the downstream pressure and it allows changing the mass flow rate simply by regulating the upstream pressure through the flow control valve and independently from the combustion chamber pressure.

Initial dimensioning of the single injector has been carried out through a modified version of Bernoulli’s equation, considering the mass flow rate that is necessary to insert into the combustion chamber:

(1)

In detail, the global injection area has been calculated considering the low thrust operating condition, this is due

to the possibility of increasing the mass flow rate, and thus the thrust, simply acting on the regulation valve. Once area of the total injection plate has been estimated through eq.1, number and dimension of the single injector can be determined. The initial dimensioning foresees 60 straight injectors with a diameter of 2.4 mm each, and a predicted mass flow rate equal to about 180 g/s for the maximum thrust condition and 80 g/s for the mimimum thrust condition. Furthermore, in order to have a fully developed flow inside the injector, an aspect ratio of 7.5 has been

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chosen, resulting in a total injector length of 18 mm. It’s of utmost importance to highlight that those geometrical characteristics are not problematic in terms of manufacturability.

This preliminary design has been also verified through detailed numerical analyses carried out in cavitating conditions with Open Foam solver [12]; numerical results are close to the prediction of the engineering method, with an error on the estimated mass flow rate through the single injector less than 5%. Numerical computations highlighted that the cavitation starts the injector inlet, with a steep phase change at the injector exit as shown in Figure 9 where density contours for one of the analysed configuration are reported.

Figure 9 Density contour through single injector

2. Nozzle Nozzle configuration has been defined through a numerical campaign with the aim to verify the influence of

convergent length, convergent angle, radius of curvature at the throat and divergent angle on the overall performaces and wall heat flux. More than 30 configurations have been analysed; results for three meaningful configurations are shown in Figure 10.

Figure 10 Nozzles configurations

Wall heat flux determined on these configurations is shown in Figure 11: the analysis shows that the highest

peak of heat flux has been obtained for test case 1, where the convergent is characterized by the smallest length. The lowest peaks are associated to the longest convergent, but in this case the highest average and integral values is

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determined. The selected configuration (id. 25), characterized by a divergent angle of the conical part equal to 18 deg, represents a compromise between two opposite effects: higher throat heat fluxes with lower total thermal loads and lower throat heat fluxes with higher total thermal loads.

Figure 11 Heat flux distribution along the nozzle wall

It is well known that in hybryd and solid propulsion systems, nozzle erosion will cause a decay of performances.

The performance tool is able predict this decay considering the enlargement of the throat are with respect the initial design value. A conservative value throat erosion for a graphite made nozzle has been used (r= 0.025 m/s), from [3].

Figure 12 shows nozzle profile without erosion (buning time t = 0 s) and at the end of the firing time (i.e., t = 35 seconds).

No Throat Regression

Throat Regression

Figure 12 Nozzle throat erosion

A strong decrease of chamber pressure (see Figure 13) is predicted with respect the no throat regression case

(about 35%), in both cases (with and without nozzle erotion) the throat has been verified to be chocked (Me= 2.73 and 2.95 respectively).

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No Throat RegressionThroat Regression

Figure 13 Nozzle throat erosion; Chamber Pressure

A sensible variation of thrust is predicted (about 8%) but still within the thrust requirement (see Figure 14)and a

low variation Isp is predicted ( see Figure 15) with respect the no throat regression case (about 5%). Anyway, Isp and thrust vary in percentage less than Pc in 0-D approach.

Figure 14 Nozzle throat erosion; Thrust

Figure 15 Nozzle throat erosion; Specific Impulse

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Detailed analyses considering the effect of ablation on the graphite nozzle are currently on-going. In the refined approach, a coupled solution of the flowfield with the thermal protection system considering the ablation of the material and the reduction of the heat flux entering the solid part.

C. HRE Final Configuration After the sizing of the main component of the HRE, the final configuration at PDR step of the actual project is

highlighted in Figure 16. In particular, the injector plate is equipped with an igniter based on GOx/GCH4, developed in past research project [14]. Pre-combustion chamber and post-combustion chamber have been equipped with thermal insulation. As already told, nozzle is made of graphite, surrounded by metallic case connetted with the metallic grain case with a proper flange.

Figure 16 HRE final configuration

III. NUMERICAL VERIFICATION Overall design has been verified to evaluate rocket performances (chamber pressure, efficiency) through

numerical computations on the entire configuration, by using the in-house developed code NEXT [5]. A pre-chamber length equal to 10% of grain length has been considered. Three different geometries corresponding to initial configuration and grain recession after 7s and 15s have been analysed. Transient numerical computations have been carried out with 2nd order spatial accuracy assuming an ideal gas. Turbulence has been modelled through the k- ω SST turbulence model whereas the EDM mono-step model has been consider to take into account turbulence/combustion coupling. Being the EDM essentially based on the hypothesis of fast chemistry, the influence of the chemical system used by the model is particularly related to the number of chemical species used: the higher the species number, the closer to reality the simulation. The same approach used in [13] has been considered in the present application to model a one-phase combustion process, where all the reactants are in a gaseous state.

In Figure 17, temperature contours in the combustion chamber and nozzle are shown; the flame is closed on the axis at the throat location. Predicted chamber pressure is equal to 36.6 bar, slightly below the design value (equal to 40 bar).

Figure 17 Temperature contours in the combustion chamber

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IV. CONCLUSIONS In the frame of the national research program HYPROB, HYBRID project is focused on the development of

enabling technologies for hybrid rocket propulsion. In particular, the developed technologies will be integrated into a ground demonstrator whose preliminary design has been presented. The project is approaching the Preliminary Design Review and subsequently small and medium scale tests will be carried out to validate the overall design before the test campaign on the full scale demonstrator.

Acknowledgments This work has been carried out within the Hyprob program, funded by the Italian Ministry of Education,

University and Research whose financial support is highly appreciated. The authors greatly thank all the HYBRID project team, and in particular Daniele Cardillo, Gianpaolo Elia, Mario Panelli and Marco Invigorito.

References 1 Borrelli S., De Matteis P., Schettino A. and Ferrigno F., “The Hyprob Program Mastering key technologies, design and testing capabilities for space transportation rocket propulsion evolution”, 63rd International Astronautical Conference, Naples, Italy, 2012. 2 Di Clemente M., Votta R., Ranuzzi G., Ferrigno F., “Hybrid Rocket Technology in the frame of the Italian Hyprob Program”, 8th European Symposium on Aerothermodynamics for Space Vehicles, Lisboa, Portugal, 2015 3 Karabeyouglu, M. A., D. Altman, B. J. Cantwell. 2002. Combustion of liquefying hybrid propellants: Part 1, general theory. J. Propul. Power.I. 18 (3), pp. 610-620. 4 Karabeyouglu, M. A., D. Altman, B. J. Cantwell. 2002. Combustion of liquefying hybrid propellants: Part 2, Stability of liquid films. J. Propul. Power.I. 18 (3), pp. 621-630. 5 Ranuzzi G., Cardillo D., Invigorito M., Numerical Investigation of a N2O-Paraffin Hybrid Rocket Engine Combusting Flowfield”, 6th European Conference for Aeronautics and Space Sciences (EUCASS), Krakow, Poland, 2015. 6 Piscitelli F., Saccone G., Gianvito A., Cosentino G., Mazzola L., “Microcrystalline paraffin wax as fuel for Hybrid Rocket Engine”, 6th European Conference for Aeronautics and Space Sciences (EUCASS), Krakow, Poland, 2015. 7 Karabeyoglu M.A., “Hybrid Rocket Propulsion for Future Space Launch”, Aero/Astro 50th Year Anniversary, 2008 8 http://www.propulsion-analysis.com/ 9 Sutton, G., P., History of Liquid Propellant Engines in United States, Journal of Propulsion and Power, Vol. 19, No. 6, 2003 10 Grosse, M. “Effect of a Diaphragm on Performance and Fuel Regression of a Laboratory Scale Hybrid Rocket Motor Using Nitrous Oxide and Paraffin”, AIAA 2009-5113, 45th Joint Propulsion Conference and Exhibit, Denver, USA, 2009 11 Elia G., Invigorito M., Di Clemente M., Votta R., Ranuzzi G., “Injection System Design of an N2O-Paraffin Hybrid Rocket Demonstrator”, Young International Conference, Aachen, Germany, 2015 12 Invigorito M., Elia G., Panelli M., “Hybrid Rocket Injection Hybrid Rocket Injection Hybrid Rocket Injection Hybrid Rocket Injection Hybrid Rocket Injection Hybrid Rocket Injection”, 6th European Conference for Aeronautics and Space Sciences (EUCASS), Krakow, Poland, 2015 13 M. Lazzarin et al. “CFD Simulation of a Hybrid Rocket Motor with Liquid Injection”, AIAA 2011-5537, 47th Joint Propulsion Conference and Exhibit, San Diego, USA, 2011 14 Battista F., Ferraiuolo M., Martucci A., Fragiacomo M., Natale P., Ricci D., Roncioni P., French A., Salvatore V., “Modelling, Testing and Design Considerations of a GOx/GCh4 Igniter for Hyprob SSBB Single Injector Thrust Chamber, Journal of British Interplanetary Society, vol. 67, pp.65-73, February 2014 15 - Lazzarin, M., Barato, F., Bettella, A., Pavarin, D. “Computational Fluid Dynamics Simulation of Regression Rate in Hybrid Rockets,” Journal of Propulsion and Power, Vol. 29, No. 6, November–December 2013 16 Bellomo, N., Lazzarin, M., Barato, F., Bettella, A., Pavarin, D. “Investigation of Effect of Diaphragms on the Efficiency of Hybrid Rockets,” Journal of Propulsion and Power, Vol. 30, No. 1, January–February 2014