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PRELIMINARY DESIGN OF A 30kN PARAFFINBASED HYBRID ROCKET ENGINE
CONFERENCE PAPER · JULY 2015
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3 AUTHORS:
Raffaele Votta
Marco Di Clemente
CIRA Centro Italiano Ricerche Aerospaziali
CIRA Centro Italiano Ricerche Aerospaziali
35 PUBLICATIONS 54 CITATIONS
45 PUBLICATIONS 63 CITATIONS
SEE PROFILE
SEE PROFILE
Giuliano Ranuzzi
CIRA Centro Italiano Ricerche Aerospaziali
36 PUBLICATIONS 83 CITATIONS
SEE PROFILE
Available from: Marco Di Clemente
Retrieved on: 30 September 2015
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 (N 2O) 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
CMA
HRE
Isp
M
O/F
P
r
RCS
T
=
=
=
=
=
=
=
=
=
=
Aerotherm Chemical Equilibrium
Charring Material Ablation
Hybrid Rocket Engine
Specific Impulse
Mach Number
Oxidizer to fuel ratio
Pressure
Regration Rate
Reaction Control System
Temperature
Subscripts
C
e
T
= Combustion chamber
= 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 LN 2O/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
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
1
American Institute of Aeronautics and Astronautics
2
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 subsystems 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 N 2O 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).
Throat Regression
No 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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Throat Regression
No Throat 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 prechamber 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.
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