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Throttleable hybrid engine for planetary soft landing



The SPARTAN research program aims at developing a throttleable propulsion technology, which is mandatorily needed for any planetary soft and precision landing. It relies on the hybrid engine technology, exploiting its capability of being throttled and its proper performance. This research program is complementary to ESA TRP and Piedmont Regional development programs. It implements and strengthens the technological base in view of the future robotics and manned space exploration missions. The outcomes from the SPARTAN development can be reflected also in many Earth/Space civilian and military applications, exploiting both the throttling capability of the propulsion system and the peculiar characteristics of the hybrid engine technology, like: safety, minimum environmental impact (green propellants), lower life cycle costs, responsiveness, competitive performance, increased reliability, soft ignition, and shutdown The hybrid propulsion system is formed by two major constitutors: the engine itself, housing the fuel, and the oxidizer injection system. The research focuses on three major objectives, needed to achieve the soft and precision landing capabilities: • The engine design, specific for throttling functionality • The oxidizer throttleable device development • The design of the landing case: test bench and testing procedures The development will be supported by establishing an advanced coding, enabling the definition of the fuel and the throttling behavior of the hybrid engine. The design will be supported by development tests: cold injection case, dedicated to the throttling device, and hot firing on subscale model, merging the throttling device and a subscale engine. In parallel it will be developed a landing test and the associated landing model (flying test bed), providing the availability of proven landing model and landing test capabilities. These capabilities will allow demonstrating the soft and precision landing features of a throttleable hybrid propulsion technology
Copyright 2011 by Guido Parissenti and Mario Pessana. Published by the EUCASS association with permission.
Throttleable hybrid engine for planetary soft landing
Guido Parissenti
, Mario Pessana
, Enrico Gaia
, Elio Zaccagnino
, Francesco Santilli
, Daniele Pavarin
, Alberto
, Jan-Erik Ronningen
, Turi Valle
, Patrick Van Put
, Rob Tijsterman
, Robert Popola
, Antonin Pistek
Emanuele Di Sotto
, Didier Grandou
, Valentine Stasse
, Luigi T. De Luca
, Luciano Galfetti
, and Filippo Maggi
Thales Alenia Space Italia S.p.A., Italy
Università degli Studi di Padova, Italy
Nammo Raufoss SA, Norway
Bradford engineering B.V. Netherlands
Technical University of Brno, Czech Republic
GMV, Spain
Politecnico di Milano, Italy
The SPARTAN research program aims at developing a throttleable propulsion technology, which is
mandatorily needed for any planetary soft and precision landing. It relies on the hybrid engine
technology, exploiting its capability of being throttled and its proper performance.
This research program is complementary to ESA TRP and Piedmont Regional development programs.
It implements and strengthens the technological base in view of the future robotics and manned space
exploration missions.
The outcomes from the SPARTAN development can be reflected also in many Earth/Space civilian
and military applications, exploiting both the throttling capability of the propulsion system and the
peculiar characteristics of the hybrid engine technology, like: safety, minimum environmental impact
(green propellants), lower life cycle costs, responsiveness, competitive performance, increased
reliability, soft ignition, and shutdown
The hybrid propulsion system is formed by two major constitutors: the engine itself, housing the fuel,
and the oxidizer injection system.
The research focuses on three major objectives, needed to achieve the soft and precision landing
The engine design, specific for throttling functionality
The oxidizer throttleable device development
The design of the landing case: test bench and testing procedures
The development will be supported by establishing an advanced coding, enabling the definition of the
fuel and the throttling behavior of the hybrid engine. The design will be supported by development
tests: cold injection case, dedicated to the throttling device, and hot firing on subscale model, merging
the throttling device and a subscale engine.
In parallel it will be developed a landing test and the associated landing model (flying test bed),
providing the availability of proven landing model and landing test capabilities. These capabilities will
allow demonstrating the soft and precision landing features of a throttleable hybrid propulsion
1. Introduction
The soft landing on extraterrestrial bodies is a key issue in future space exploration and in long-term vision of
European and Non-European space programs. Soft landing is required for unmanned and necessarily manned
missions, which have to deliver on planet surfaces always more heavy equipments and manned modules that cannot
withstand strong impact loads or that needs to maintain the landing place as unaltered as possible. Such approach is
applicable in the near future to both a return of the man on the Moon to future Mars exploration missions and this is
also envisaged explicitly in European future plans. ESA’s view on The Long-Term International Scenario for Space
Exploration [1] clearly states that the period during Phase 1, through 2016 and perhaps through 2020, will
demonstrate key capabilities such as planetary descent and landing, surface mobility, in-situ resource utilization
(ISRU), and perform valuable in-situ science (see Figure 1).
Figure 1 Long Term Scenario for International Space Exploration
Other statements from ESA and other world wide space agencies strongly show the necessity for efforts in Space
Transportation to define Mars Robotic Exploration [2] and Lunar Cargo Landers [2][3] that would lead to a Mars
Sample Return mission, with precursor program able to demonstrate soft and precision landing capabilities.
Soft and precision landing can be performed only with a chemical throttleable engine, capable to vary its thrust to
follow the control commands to successfully deploy its payload on an unknown surface with non fully predictable
atmospheric or gravitational entry conditions. At worldwide level the state of the art of throttleable engine is
represented by liquid thruster, even if solid thruster throttling is possible with the pintle technology, and different
R&D programs are ongoing almost all outside Europe. Current research activity is focusing on high-thrust cryogenic
liquid bipropellant rocket engines, mainly for Moon descent missions. NASA and P&W ongoing Common
Extensible Cryogenic Engine program is focused in demonstrating the technological feasibility of a LH2/LO2 deep
throttled motor (throttle ratio 13:1 demonstrated on ground tests) for Lunar descent. The study baseline is a modified
RL10 engine. Injector and oxidizer feed lines had been deeply modified to suppress combustion instabilities at low
throttle. Northrop Grumman Space is also developing the TR202 engine for Moon descent missions. Stable
combustion was demonstrated over 10-1 throttle range with a pintle injector with GH2 and LOX propellant.
At European level the research in this field is strongly less respect to US. Qualified throttleable engines basically do
not exist in Europe, being the available items standard liquid thrusters with limited throttling capabilities. In addition,
throttleable liquid thrusters have several problems, being generally designed for fixed thrust with small variations
about the design point for throttling. Extension of this capability to deep throttling is expensive, time-consuming, and
with limited off-design chance due to high system complexity and sensitivity to combustion instabilities/oscillations.
Moreover, existing design is difficult to modify addressing mission requirements changes.
Following the aforementioned needs of the European space community, the design of a fully throttleable engine for
soft planetary landing is the objective of the SPARTAN project, which for this activity is supported by European FP7
program. The engine will provide smooth and wide throttle range coupled with high-thrust and will be based on the
hybrid rocket motors technology. Differently than liquid engines, hybrid rockets are intrinsically simpler, safer, with
a wide throttling capability and have the possibility to use green propellants, which concur to lower the development
cost along with the other aforementioned qualities. Their main drawback is modeling issues on the combustion
process, which will be considered during the research on codes and motor fuels.
2. SPARTAN approach on throttling validation
The aim of the SPARTAN program is to develop a new throttling technology which will be applied to a newly
designed throttleable engine for the soft and precision landing on a planetary surface. This goal will be the result of
an intense research program that will involves 8 partners from 6 different countries from both universities and
The approach followed by the SPARTAN project is to not only design a new motor, but also to test it with full soft
landing test with a new low-cost highly realistic test bed. To perform the test a lander will be developed, in order to
provide the necessary elements to carry out successfully a soft landing on the Earth. The test, optimized for the Earth
environmental conditions, is linked to Mars planet requirements and robotic missions. The test plan and development
will interact with the lander design requirements, leading the lander configuration and engine nominal performances.
During the test, the Lander will be lifted to 100m and dropped by a helicopter. Once the Lander will be lifted up to
the proper height, it will be released for a free fall of about 50 m to achieve 30 m/s of vertical velocity before
activating the propulsion system to its nominal operations.
Figure 2 Full scale Soft Landing Test mission timeline
The test objectives are:
To damp the vertical velocity verifying the engine throttling capabilities;
To maintain the vehicle stability during the velocity damping phase.
It is notably that this approach is strongly challenging but can provide an insight in the technology that is difficult to
achieve in other ways. New propulsion systems concept has been tested till now in Europe exclusively with static
ground tests, with the exception of small motors for sounding rockets or interceptors missiles, fired in safe and low
populated zones like Norway (i.e. Nammo hybrid sounding rocket). Space motors are difficult to test, mainly
regarding their dynamic capabilities and safety issues related to the nature of the propellant, and no drop test has
been performed in Europe to explicitly test a working mock up of a chemically propelled lander. The gap with US
testing method is evident, many drop tests have been performed and also tests with ascending and descending lunar
landers mock-ups. It is possible to cite the Northrop Grumman Lunar Lander X PRIZE Challenge: “a $2,000,000
incentive prize program designed to build an industry of American companies capable of routinely and safely flying
vertical take-off and landing rocket vehicles useful both for lunar exploration and other applications”. The private
companies that are competing for the prize must perform a dynamic test that requires the proposed landers to actually
take off, sustain, move laterally and land.
Free Falling
Altitude: from 100 to 50 m
Speed: from 0 to 30 m/s
No engines thrust
Altitude: from 50 to 0.5 m
Speed: from 30 to 2 m/s
System check
Altitude: from 0.5 to 100 m
Stability check
Helicopter lifting
Altitude: 100 m
Waiting for required
Stability check
The test approach foreseen in the SPARTAN program is totally new for Europe, in the frame of the soft landing
capability verification, with an active and throttling propulsion system. A dynamic test able to verify the performance
of the developed technology on ground allows performing a dynamic end-to-end test in a representative scenario,
providing a new methodology and architecture for testing of landing technology and payloads. Therefore this project
is providing for a real step forward with respect to the state of the art in Europe.
2.1 The Lander
The capability of a mechanical structure to land safely on a planet surface is mandatory for each manned and
unmanned mission. Other than the parachute and the thrusters, which scope is to reduce the initial entry velocity to
few meters per second, the lander structure and most of all the lander legs, have to dissipate the residual energy
reducing at the same time the impact shock to preserve the integrity of the payload.
Also the soft landing test approach, the conclusive validation of the developed throttleable engine, requires the
design of the lander. This will be sized to allocate the engines and all the relevant equipment, and to withstand the
impact loads and the environmental issues that will be identified in the drop test location, such as wind, temperature
and soil composition to avoid dust raising. The Landing Model Structure will be developed by using qualified
material and processes: light weigh high strength carbon fiber structures will be maximized.
Other than the structure the design of the lander includes the Thrust Control Algorithm and GNC and the Storage
system. The Thrust Control Algorithm baseline concept is to have a classical outer loop in which estimation of
kinematics state of the vehicle (position, velocity, attitude) is performed by the Navigation function based on the
available sensor data, a simple guidance logic generates a descent profile (no obstacle avoidance capability is
foreseen at this stage), while the control block will issue the command vector in terms of desired thrust level for the
four hybrid engines. In addition to this, the actual thrust level of individual engines will be evaluated by load cells in
order to establish an inner control loop on the thrusters output.
Figure 3 SPARTAN Landing Model Structure and Landing legs
The Storage System is made of one oxidiser tank and one pressurant vessel. The oxidizer tank will be maintained at
the constant pressure by the pressurant (GHe) contained in a dedicated pressure vessel. In order to avoid gas and
liquid mixing, the oxidizer tank will have an elastic bladder, compatible with the chosen oxidizer (H2O2).
2.2 Drop Test requirements
The mission main objective is that the demonstrator shall be able to perform a “soft landing”, which means that the
lander shall be able to achieve a desired impact velocity ad a desired altitude in order to be able to absorb the residual
energy thanks to the landing legs. The most important requirement is obviously the touchdown velocity, which
impacts the structure of the demonstrator.
A short survey shows that while Deep space 2 was supposed to touchdown at 2.4 m/s (mission failed), Phoenix
landed at 1.6 m/s, and Mars sample return is supposed to land at 2 m/s. A good compromise is hence represented by
the last option, which means that the required landing velocity is assumed to be 2 m/s.
In order to avoid hazards to objects and persons with the motors firing at ground level the lander shall not be posed in
the condition of losing stability during the impact, meaning to avoid stresses to the hinges so to compromise the
integrity of one or more legs. Hence two requirements are derived: that the stability of the lander shall be maximized
even in case of the loss of a leg and that the desired lateral velocity shall be assumed null to avoid stresses that can
damage one or more legs. From the first requirement a four legs configuration is mandatory, being a good
compromise between stability and mass penalty.
Summarizing the final requirements are:
Four legs configuration
Vertical impact velocity: 2 m/s
Lateral impact velocity: about 0 m/s
These requirements shall be fulfilled by the chosen configuration and by the foreseen thrust profile. The thrust
profiles are the desired thrust throttling law that should be followed during the soft landing test starting from the
engines ignitions. They are calculated in order to fulfil the mission timeline and the requirement above on final
vertical velocity, with the objective to simulate a thrust profile able to show the engine throttling capabilities.
Obviously from a GNC and Navigation point of view those possible thrust profiles have to face the external
environmental conditions that would provide variation in attitude and lateral velocity. The objective of the
Navigation algorithm is to follow as close as possible the chosen thrust profile minimizing and compensating
external factors. Figure 4 shows a possible thrust profile for two possible lander configuration of different weight.
Figure 4 Thrust profile option for a 253 kg lander (left) and a 370 kg lander (right)
3. SPARTAN Throttling technology development
As introduced above, hybrid propulsion has been selected by the SPARTAN project being the best candidate for
space exploration applications that requires throttling capabilities because of: (i) its higher ISP compared to both
monopropellant and solid motors; (ii) its intrinsic simplicity, only one feeding line for the fluid oxidizer is required
compared to bipropellant liquid motors; (iii) the thrust chamber is easy and cheap to be build and catastrophic failure
related to throttling operation is very unlikely (differently form solid and liquid engines), and last but not least; (iv)
the inert characteristic of the propellants used. In addition often these non toxic propellants can be considered
“green”, representing a good step forward from the planetary protection point of view. Throttling is achieved just
varying the oxidizer mass flow rate, and that has already been demonstrated. However, full system controllability,
required for soft-landing applications, requires a good knowledge of combustion process especially during unsteady
An overview of the benefits of hybrid technology respect to other propulsion concept is clearly visible in Table 1.
The table summarizes pro and cons of rocket motor technologies, with focus on throttling and soft-landing
Table 1 Competing technologies
Rocket motor type Thrust ISP (s) Throttling
ratio Propellants Advantages Issues
rocket engine low low 10:1 mostly toxic
compact and high
responsivity to
low ISP, limited
maximum thrust,
hazardous propellant
bipropellant rocket
low medium
10:1 toxic and
compact and high
responsivity to
propellant, complex
throttle control
bipropellant rocket
high high 10:1
high ISP, high
thrust, wide
throttling capability
propellants, complex
system, expensive
design and testing
solid rocket motor high medium
susceptible to
compact, storable,
medium ISP
limited throttling
hazardous premixed
hybrid rocket
motor medium
10:1 green
simple, wide
capability, medium
ISP, green
modelling issues on
combustion process,
low TRL of high-
regressing fuels
Hybrid rocket motors appear so to be the perfect candidate for missions where deep throttling coupled with medium
thrust and good ISP are required. Moreover, hybrids offer substantial advantages to development costs due to their
simplicity, safety and green propellants. Research is required on codes and fuels to improve knowledge on the
fundamental physical aspects required to successfully qualify hybrid rocket motors for soft-landing and other mission
with throttling requirements.
Current TRL of hybrid throttleable motors, for in space application, is still low, between 3 and 4. With the proposed
research it would be possible to raise the overall TRL of the technology to 6, demonstrating it in a relevant
environment. To reach this it is required the development of a complete and integrated Lander system, not just a
propulsion system, to execute a complicated task. The concept will be developed in detail and the throttling device is
harmonized with the hybrid propulsion technology to obtain a reliable technology demonstrator.
The benefits from the SPARTAN program are not limited to the field of the space propulsion, with the improvement
of the hybrid engine technology and the widening of its application, but also extended to the expertise which will be
matured in the field of the controlled landing test, which can be exploited for further development steps to validate
further landing features.
3.1 Development logic
In SPARTAN program the development of the throttling concept, as in case of the other R&D activities of the
project, is not a self standing activity but is intimately connected with the different parts of the program. The study
logic concerning this part of the work is represented in Figure 5.
Figure 5 Throttling Device Development Logic
The Throttling Technology Development block is the project pivoting task: it establishes both the throttling
requirements and its concept, which drive the technology development. This result will be achieved be leading an
intensive concurrent design session, harmonising the mission requirements (project reference), the testing
requirements (soft landing on Earth), and each subsystem performance characteristics.
The Hybrid Engine Throttling Capability Development block leads to the development of the throttling technology,
through the advanced coding, the CFD analyses and the development testing (cold and hot on the subscale model).
These main tasks are supported by design activities (the Propulsion System Design and development block) relevant
to the throttling device and the engine itself. A proper concept selection for the throttling device and the engine
preliminary design, coming from the Propulsion System Design and development block supports the start up of the
coding activities and will give input to the final design activities of both the throttling device and the engine
components (injectors and combustion chamber). Development tests will be used to validate the code.
The validated code, and the sub-scale experimental set-up, will be used to design the throttling devices and the
engine and to optimize them, than a full scale model is built and tested. This will be the object of activity clustered
under the Propulsion System Design and development block.
3.2 Throttling Technology for hybrid propulsion
The development of the hybrid engine will rely both to the experience of the partner of the project in this technology
and to the development of new methodologies for studying the critical issue of such thrusters mentioned above. In
particular in order to understand the behavior of the fluids inside the engine a 3D code will be developed. This code
will be a new model to simulate the hybrid rocket motor, from upstream injector to the nozzle.
The 3D unsteady CFD code will be capable of simulating detailed physics of the hybrid combustion phenomena in
case of general hybrid rocket. In this case the greater goal will be to simulate the regression rate local variations and
combustion efficiency.
As mentioned before the code is validated through both a cold test, dedicated to the throttling device only, and a hot
test on a sub scale model, to validate the full chain of devices. The code will run in parallel on super computer.
This model will be divided into three main parts:
1. Simulation of the internal flow of the injector;
2. Atomization/ break-up of the two-phase fluid exiting the injector due to its interaction with air, together
with the combustion simulation. Combustion shall cover internal flow dynamic, species diffusion, reactions
kinetics and heat transfer to the fuel surface in the combustion chamber;
3. Nozzle flow accounting eventually for the post-combustion chamber mixing.
1 2 31 2 3
Figure 6: OpenFoam model subdivision
The subdivisions in the above picture:
1. The first block refers to the injection system;
2. Block number two refers to the combustion code;
3. Block number three refers to the nozzle exhaust.
These blocks are to identify the specific areas in which the rocket physics is simulated, whereas it is important to
highlight that the reservoir and feed-lines will not be modelled or simulated.
The physics behind the throttling technology is strictly related with the hybrid propulsion working model. In
chemical rockets thrust is proportional to exhaust/burned propellant flowing out of the nozzle. At steady-state, in
hybrid motors, this is the sum of injected oxidizer mass and regressed fuel mass. The latter, due to the inherent
burning process of hybrid rocket motors, is proportional to oxidizer mass flow. Thus throttling is simply achieved
varying the oxidizer mass flow rate.
Figure 7 Schematic of the Oxidizer feed system
Let’s consider in an example of schematic hybrid motor oxidizer feed system, which in this case is focused on the
SPARTAN vehicle due to the presence of four thrusters. Downstream the FCV there is the injector, one for each
motor. The baseline is a constant area injector, as simple as a plate with many orifices. The scope is to increase the
oxidizer velocity thus promoting its atomization before reacting inside the combustion chamber with the gasified
fuel. Good atomization of the liquid oxidizer is achieved with a reasonable pressure drop across the injector. If the
pressure difference is reduced below a threshold limit, oxidizer droplets become too coarse, neglecting efficient and
stable combustion. Nevertheless the oxidizer mass flow reduction by means of a FCV has the drawback of reducing
pressure drop at low throttle
The pressure drop across the injector is function of oxidizer mass flow rate, and approximately is expressed as:
Given the oxidizer mass flow rate and fixed injector area, the pressure drop is minimum at minimum
. Thus
atomization is worst at low throttle. The oxidizer mass flow rate
is fixed by the FCV:
PCdAm =
At maximum flow rate P
is at minimum, at minimum oxidizer flow rate, P
is at maximum (see Figure 8).
The combustion chamber pressure (P
) is approximately proportional to the motor thrust.
Figure 8 Pressure drops in injections
Below 40% of nominal thrust flows atomization becomes too inefficient for motors with fixed injectors. To achieve
10:1 thrust modulation, upgrades to this baseline configuration are needed. The objective is thus to limit the
reduction of pressure drop across the injector P
when thrust is reduced, preserving atomization quality. To achieve
this task several concepts can be applied. They have been developed in the past 60 yeas in the framework of LRM
(liquid rocket motor) throttleability studies. Main concepts applicable to liquid oxidizer injection are:
1. High pressure-drop: minimum pressure drop across the injector is fixed at the lowest oxidizer mass flow
rate; however this solution requires increasing the pressure drop at high thrust. This does not neglect good
atomization, but requires loading higher than required pressurant, increasing motor mass. This method is
usually feasible for throttling between 100% and 60%;
2. Dual-manifold injectors: oxidizer mass flow rate is provided to the injector by two separate feed-lines.
This reduces the pressure drop at low thrust but requires to double valves;
3. Gas injection: gas is mixed into the oxidizer flow before injection, reducing the mixture density. The
decrease in density is proportional to the increase of pressure drop and good atomization is provided even at
reduces mass flows. This concept requires adding a gas tank and related feed lines;
4. Moveable injector components: the injector area is changed during throttling, i.e. using a pintle. If area is
reduced pressure drop increases at low oxidizer mass flow, preserving good atomization. This concept add
substantial complexity to the hybrid rocket motor;
The main issues to be addressed in investigating the throttle capability for the development of a performing variable
mass flow injector/system are:
Optimized injector atomization, flow stability and jet shape with variable mass flow/throttle ratio;
Minimal time-lag to throttle command;
Throttle device simplicity, robustness and reliability.
Summarizing, the throttleable device is the sum of two systems: the flow regulation device is required to modulate
the oxidizer mass flow rate, the injection system to atomize the oxidizer and possibly, to guarantee a good
performance and stability of the system over the throttling range. A good injection system shall be able to adapt to
the oxidizer mass flow rate and always provide good oxidizer atomization and high feed-system stiffness
(insensibility to couple with or trigger combustion instabilities). CFD analyses are planned to investigate these
4. SPARTAN Test plan
The full scale soft landing drop test is the final point of a development process which foresees a series of
intermediate tests for both the development of the throttleable thruster and for the validation of its throttling
capabilities. Before reaching the confidence to actually load the lander on the helicopter a series of Full Scale Hot
test with dedicated test bed will be performed. These tests will stress the newly designed engines in restricted and
safe conditions, allowing the validation of the throttling ranges without the worries of environmental condition and
Guidance algorithms application.
Prior to this step a sequence of extremely important cold and hot firing test will be performed on parts of the
throttling device and on fully representative models of the final engine.
4.1 Full Scale Hot test
Nammo is responsible for carrying out the full scale test program of the SPARTAN propulsion system before
commencing the flight test program. All testing will be carried out on Nammo’s own test centre, which contains all
necessary state-of-the-art hardware like: environmental test facilities, 6-DOF (Degree Of Freedom) force cells, high
speed video cameras, calibrated multi channel high-frequency band recording equipment, skilled personnel and
more. Nammo also have seven years of experience in testing hybrid rocket motors with thrust levels up to 30kN.
Two main Full Scale Hot test will be performed:
A static test, devoted to measure SPARTAN vehicle actual thrust and lateral forces. See Figure 9.
A dynamic test, represented by a low altitude drop test (to mitigate the risks for the final full height drop
test from helicopter), in which the SPARTAN vehicle is lifted with a crane or an appositely designed
structure and is dropped locking one or two degrees of freedom.
In the full scale restricted movement drop test (Figure 10 left) the lander is forced to move vertically, hence without
requiring stability control. In the quasi non-restricted movement drop-test (Figure 10 right) the lander is dropped
from low altitude hold by a wire rope for safety purposes and is allowed to start the thrusters and land softly: in case
of failure the wire rope will retain the vehicle to crash on ground.
Figure 9 Lander static test vertical fixture
Figure 10 Full scale restricted movement drop test (left), Full scale quasi non-restricted movement drop test (right)
4.2 Cold and Hot test on single engines or technologies
Prior to the Lander Full Scale Hot tests, during the development phase different models will be manufactured to
verify the validity of the throttling technology or of entire thrusters concepts. In fact the development of the rocket
motor will be conducted at various stages:
1. HW Design phase (HW = Heavy Wall test item)
2. HW test phase
3. FW Design phase (FW = Flight Weight test item)
4. FW test phase
The HW engine will have the same internal geometry and size as the proposed FW configuration; however its
surrounding structure will be oversized. Its construction will also be modular, meaning that it can be re-configured,
assembled and re-assembled in a fast and efficient way. In addition, the HW design will be prepared to house several
sensors for measuring the performance of the rocket motor. HW and FW models and the propellant they contains,
will undergo a series of static Cold and Hot test:
Fuel Characterization: a complete characterization of the fuel will be gives in terms of ballistic and mechanical
behavior and theoretical specific impulse under design conditions and throttling operations. Flame visualizations
with a high speed and high resolution camera will be performed getting detailed information on the quality of the
combustion process. Thanks to a specific implementation of a radial burner, the behavior or formulations under quasi
steady combustion will be evaluated. As a complimentary test, mechanical features will be evaluated through
uniaxial tensile tests performed at one strain rate and ambient temperature. Also dynamic mechanical analysis will be
performed. Elastic, visco-elastic and break-up properties will be reported.
Subscale Cold test: the test is devoted to verify the throttling technology. The setup will consist of a pump to
pressurize the liquid simulant and pump it through the throttle. A closed loop setup is employed so that tests can be
performed at specified conditions for prolonged periods of time. The throttle is controlled by a PC with data
acquisition hardware and control logic. Measured quantities include the mass flow through the throttle and the
pressure before and after the throttle. Therefore the performance of the throttle and the oxidizer feed system,
quantified by mass flow as a function of pressure drop, can be completely characterized.
Figure 11 Cold Test Setup schematic
Full Scale Hot test: these tests will be conducted to validate the throttleable hybrid motor advanced code, a tool
which will be developed and used to design the hybrid motor, and to assess the motor performance at steady-state
and during modulation, before progressing to full-scale testing. The lab model hot fire set-up is made-up of
pressurant and oxidizer tanks, oxidizer feed-line, flow control device, injector and combustion chamber. The
diagnostic includes measure of thrust, pressure, temperature, oxidizer and fuel flows. Majority of diagnostic will be
used both on codes validation and motor performance assessment.
The diagnostic system for assessing the performance of the throttleable motor will be designed and developed to be
used on motor cold and hot-fire testing. The objective of the diagnostics is to assess and quantify the response of the
motor to thrust modulation as function of throttling.
System performance of the throttleable device will be quantified as:
Thrust response characterization (i.e. time, overshoot) during throttling;
Performance and combustion stability of the motor at different throttling settings and during continuous
Variation of functioning parameters of the hybrid motor (i.e. oxidizer and fuel mass flow rates, combustion chamber
pressure, combustion efficiency etc.) will be measured and correlated with the throttling command and resulting
thrust modulation. A schematic of the diagnostic system, to be applied to the motor set-up for hot fire and cold flow
tests, is also shown in Figure 12.
Diagnostic functional schematic
Diagnostic Architecture
Figure 12 Schematic of the diagnostic system
5. Conclusions
The SPARTAN project, which aims are to develop a new throttling concept applied to an appositely designed hybrid
engine in order to be capable of performing a wide thrust range in a dynamic thrust variation regime, and to develop
a soft landing test to demonstrate the technology peculiarities, is presented with the detailed description of the main
present and future activities.
6. Acknowledgement
This work has been approved for support in the frame of the European FP7 program.
[1] ESA’s View on The Long-Term International Scenario for Space Exploration (Road map)
[2] International Space Exploration Coordination Group (ISECG) – Annual Report: 2008, ESA/Europe
[3] International Space Exploration Coordination Group (ISECG) – Annual Report: 2009, ESA/Europe
[4] SPARTAN Proposal, SPARTAN, "SPAce exploration Research for Throatable Adavanced eNgine ", FP7 Grant
agreement no: 262837, 2010-12-21
... Recently, remarkable demonstrations have been performed by Purdue [25], USU [23,24], NAMMO [26,27] and UNIPD [28][29][30][31][32]. Moreover, a program for a Mars softlander demonstrator based on hybrid propulsion has been funded by the European community [33][34][35][36]. ...
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Hybrid rocket motors have several attracting characteristics such as simplicity, low cost, safety, reliability, environmental friendliness. In particular, hybrid rockets can provide complex and flexible thrust profiles not possible with solid rockets in a simpler way than liquid rockets, controlling only a single fluid. Unfortunately, the drawback of this feature is that the mixture ratio cannot be directly controlled but depends on the specific regression rate law. Therefore, in the general case the mixture ratio changes with time and with throttling. Thrust could also change with time for a fixed oxidizer flow. Moreover, propellant residuals are generated by the mixture ratio shift if the throttling profile is not known in advance. The penalties incurred could be more or less significant depending on the mission profile and requirements. In this paper, some proposed ways to mitigate or eliminate these issues are recalled, quantitatively analysed and compared with the standard case. In particular, the addition of energetic additives to influence the regression rate law, the injection of oxidizer in the post-chamber and the altering-intensity swirling-oxidizer-flow injection are discussed. The first option exploits the pressure dependency of the fuel regression to mitigate the shift during throttling. The other two techniques can control both the mixture ratio and thrust, at least in a certain range, at the expense of an increase of the architecture complexity. Moreover, some other options like pulse width modulation or multi-chamber configuration are also presented. Finally, a review of the techniques to achieve high throttling ratios keeping motor stability and efficiency is also discussed.
... The Spartan project was aimed to develop an advanced throttling engine for soft planetary landing[42]. NASA JPL is conducting studies for a Mars Ascent Vehicle (MAV) using hybrid rocket propulsion with paraffin-based fuel and MON30. ...
... Recently few results were presented at conferences and published on journals, the major contributor to the actual knowledge on hybrid rocket motors throttleability is EU FP7 SPARTAN Project (funded by the European Union). 1,2 The program concluded in 2012. Afterward throttleability was studied by Stephen Whitmore of the Utah State University 3,4 and the Beihang University Propulsion Group. ...
Conference Paper
Hybrid rockets technology is growing in interest in recent years. The main reasons of this renewed interest is that hybrid rocket motors are simple, safe to develop and operate, reliable, green and less expensive than traditional propulsive technologies (solids and liquids). Another peculiar feature of hybrid rocket motors is throttleability, i.e. the ability to control the thrust level on demand, and this, in hybrids, is achieved by controlling the oxidizer mass flow to the combustion chamber, hence the need for a flow control valve. At University of Padova, the Hybrid Propulsion Group has been developing lab-scale hybrid rocket motors that are particularly suited to be implemented with a flow control valve, using HTP, a monopropellant, as oxidizer, and a catalytic bed injection which is not affected by the problems of atomization of an injection plate. The flow control valve presented in this paper exploits the cavitation on an actuated pintle in order to choke the mass flow. The advantages of such a device, which is called variable area cavitating venturi, are insensitivity of the mass flow to the downstream pressure variation and linearity of the mass flow with the throat area. The paper presents the design phase of the flow control valve and its static and dynamic characterization achieved during cold flow tests. Moreover throttling fire tests are presented in order to understand the throttling behavior of our lab-scale test motor.
... 11 In recent times when a renewed interest started to gather around hybrid rocket motors, VACV started to be applied to this propulsive technology as well. During the SPARTAN program, focused on the development of a descending vehicle employing throttleable hybrid rockets, 13 MOOG-Bradford realized a VACV based flow control valve to be applied to the rocket engine. The University of Padova was entrusted to characterize this FCV. ...
Conference Paper
Hybrid rocket motors have several potential advantages respect to current used propulsion systems (i.e. solids and liquids) like simplicity, safety, reliability, environmental friendliness, lower cost. A particular positive feature of hybrid rockets is the possibility to control the thrust level operating only on the oxidizer mass flow. Thanks to this it is possible to develop a relatively simple propulsion system that is throttleable on demand without the complex mixture ratio control and related hardware of a liquid system. In the past University of Padua has developed a lab-scale hybrid rocket motor that can be throttled at few different discrete levels with the use of parallel feedlines. To give the possibility of having a continuous throttling capability a new mass flow control has been developed recently. The mass flow control make use of a cavitating pintle. The cavitating pintle acts as a cavitating venturi in order to choke the mass flow and make it independent of downstream pressure. The pintle is used to change the venturi throat area and consequently varying the oxidizer mass flow keeping a constant upstream pressure. The paper presents the design of the cavitating pintle and the experimental campaign composed by cold tests followed by hot fire tests of the lab scale hybrid rocket.
... Advanced paraffin-based fuels are researched at the DLR Lampoldshausen and the HyEnD project at the University of Stuttgart with focus on fuel composition, regression rate and mechanical strength as well as combustion chamber processes and scaling of engines and thrust levels [7,8,9,10,11]. Some interesting examples of applications for HRES are upper stages [12], sounding rockets [13,14,15,16] or lander propulsion systems [17]. To evaluate and analyze the feasibility of a hybrid rocket engine as an alternative for liquid rocket engines (LREs) or solid rocket motors (SRMs), the advantages and disadvantages of hybrid rocket engines need to be highlighted. ...
Conference Paper
Full-text available
Since the turn of the century renewed interest in the Moon has led to several lunar missions by many space faring nations. An economically reasonable approach for a future European mission is the combination of technology demonstration and scientific objectives. A promising emerging technology for exploration applications is hybrid propulsion. In the past toxic hypergolic propellants have been used to carry out lunar landings. Hybrid rocket engines are utilizing green propellants which are advantageous regarding safety and low-cost ground operations. A lunar lander spacecraft based on hybrid propulsion is analyzed. The advantages of hybrid rocket engines like inherent safety, throttle-ability, low system complexity and relatively high specific impulse make it an economical alternative to liquid propellant engines with toxic propellants. New research in advanced propellants for hybrid rocket engines further improves their performance and makes them suitable for in-space applications.
... Hybrid rocket propulsion offers several advantages over traditional liquid or solid rocket propulsion systems such as low cost, high impulse, safety, and throttling, with a broad prospect of development [1][2][3][4] . Hybrid rocket engines are easy to be throttled by changing the mass flow rate of oxidizer, therefore they serve as ideal candidates for variable thrust rocket and suitable for many applications including upper stage motors, sounding rockets, spacecraft, etc [5][6][7] . ...
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The main goal of this paper is to study the characteristics of regression rate of solid grain during thrust regulation process. For this purpose, an unsteady numerical model of regression rate is established. Gas-solid coupling is considered between the solid grain surface and combustion gas. Dynamic mesh is used to simulate the regression process of the solid fuel surface. Based on this model, numerical simulations on a H2O2/HTPB (hydroxyl-terminated polybutadiene) hybrid motor have been performed in the flow control process. The simulation results show that under the step change of the oxidizer mass flow rate condition, the regression rate cannot reach a stable value instantly because the flow field requires a short time period to adjust. The regression rate increases with the linear gain of oxidizer mass flow rate, and has a higher slope than the relative inlet function of oxidizer flow rate. A shorter regulation time can cause a higher regression rate during regulation process. The results also show that transient calculation can better simulate the instantaneous regression rate in the operation process.
In this study, ground tests of a lab-scale hybrid rocket motor were conducted to verify the feasibility of the hybrid propulsion system for lunar lander application. The primary goal is to assess the realizability of hybrid rocket by testing its throttleability and soft landing capability with a scale-down lunar module. A design thrust of 200 N was achieved by clustering four identical 50 N-class gaseous oxygen (GOX)/high-density polyethylene (HDPE) hybrid rocket motors with multi-port solid fuels. Ground tests were carried out via two main experiments: static test and drop test. Static test was focused on the overall performance of the clustering module such as cold injection, uniform oxidizer distribution, throttleability and simultaneous ignition of the four motors, while the drop test was performed to investigate the planned throttle behavior using a 1-D vertical drop test stand. The clustering module was controlled in an open-loop setup with a simple ballistic flight simulation input. The landing velocity of 1.01 m/s was achievable, confirming the possibility of soft landing missions on lunar surface using hybrid rocket motors.
This study demonstrated the performance of flight control systems of a hybrid rocket in a hovering flight test by developing a rocket designated HTTP-3AT powered by High Test Peroxide (HTP, a term used for concentrated hydrogen peroxide, H2O2). Hybrid rocket excels in system simplicity, operational safety, oxidizer storability, cost, and throttling capability compared to current solid and liquid rocket engine systems. Although issues such as the severe oxidizer-to-fuel (O/F) ratio shift during combustion and difficulty in gimbaled thrust vector control (TVC) caused by the lengthy chambers need to be solved, hybrid rocket propulsion is nevertheless a promising propulsion technology for future space exploration. To achieve accurate orbit insertion, thrust magnitude control and TVC of the rocket engines are necessary. However, no organizations have successfully implemented this technology on a practical hybrid rocket, not even using this technology for hovering flight tests. On September 8th, 2020, a hovering flight test of HTTP-3AT was conducted, achieving a steady hover 3 m above ground for 25 s utilizing both attitude and position controls. This test showed that a hybrid rocket could achieve a stable hovering flight with the capability of vertical takeoff and vertical landing (VTVL), demonstrating excellent throttling control and TVC capabilities of hybrid rocket propulsion.
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This paper presents the development of indigenous hybrid rocket technology, using 98% hydrogen peroxide as an oxidizer. Consecutive steps are presented, which started with interest in hydrogen peroxide and the development of technology to obtain High Test Peroxide, finally allowing concentrations of up to 99.99% to be obtained in-house. Hydrogen peroxide of 98% concentration (mass-wise) was selected as the workhorse for further space propulsion and space transportation developments. Over the course nearly 10 years of the technology’s evolution, the Lukasiewicz Research Network—Institute of Aviation completed hundreds of subscale hybrid rocket motor and component tests. In 2017, the Institute presented the first vehicle in the world to have demonstrated in-flight utilization for 98% hydrogen peroxide. This was achieved by the ILR-33 AMBER suborbital rocket, which utilizes a hybrid rocket propulsion as the main stage. Since then, three successful consecutive flights of the vehicle have been performed, and flights to the Von Karman Line are planned. The hybrid rocket technology developments are described. Advances in hybrid fuel technology are shown, including the testing of fuel grains. Theoretical studies and sizing of hybrid propulsion systems for spacecraft, sounding rockets and small launch vehicles have been performed, and planned further developments are discussed.
Conference Paper
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This paper describes the GNC and avionics assembly, integration and validation for hybrid engine demonstrator in the frame of the FP7 funded project – Spartan. A description of the GNC and avionics architecture and composing elements is provided and some validation tests are presented. These tests include a UAV flight test of the navigation system/algorithms, laser altimeter beam interaction with engine plume and HWIL test with robotic arm. The results of the tests performed using the fully integrated demonstrator are also presented.
SPAce exploration Research for Throatable Adavanced eNgine
  • Spartan Proposal
  • Spartan
SPARTAN Proposal, SPARTAN, "SPAce exploration Research for Throatable Adavanced eNgine ", FP7 Grant agreement no: 262837, 2010-12-21