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FUTURE IGNITER TECHNOLOGIES
W.H.M. Welland1, B.M.J. Brauers2, E.J. Vermeulen3
Aerospace Propulsion Products B.V., Klundert, The Netherlands
1 Project Manager, Engineering Department, [email protected] 2 Project Manager, Engineering Department, [email protected] 3 Director Engineering and Marketing, [email protected]
ABSTRACT Aerospace Propulsion Products B.V. operates in the space transportation business as a specialist in the development and production of igniters and starters for liquid rocket engines and solid rocket motors. In that role APP is investigating and developing technologies to be able to answer to the future needs of the liquid rocket engine and solid rocket motor producers. APP’s point of view on the future application of igniters and starters, and the technologies that can be considered will be discussed in this paper.
NOMENCLATURE AP Ammonium Perchlorate APP Aerospace Propulsion Products BV HV High Voltage ITAR International Traffic in Arms
Regulations LRE Liquid Rocket Engine MPS SRM (Moteur Propergol Solide) REACH Registration, Evaluation, Authorisation
and Restriction of Chemicals SRM Solid Rocket Motor TNO Netherlands Organisation for Applied
Scientific Research TRL Technology Readiness Level
1. INTRODUCTION APP develops and produces gas generators and their components for aerospace, defence and industrial applications. For aerospace applications these are thrust chamber-, gas generator-, solid rocket motor igniters and turbine pump starters. To be able to cover this range of products different gas and heat generating technologies need to be available. In this paper some technologies that are considered for our future
products are being discussed. The preparation for future technologies is done in communication with possible clients.
2. FUTURE APPLICATIONS IGNITION LRE
A Liquid Rocket Engine (LRE) is a rocket engine that uses propellants in liquid form. LREs have been built as monopropellant rockets using a single type of propellant, bipropellant rockets using two types of propellant, or more exotic tripropellant rockets using three types of propellant. Bipropellant LREs generally use one liquid fuel and one liquid oxidizer, such as liquid hydrogen and liquid oxygen. Most current engines utilize a bipropellant configuration in which fuel and oxidizer are stored in separate tanks. The propellants can be fed into the combustion chamber with high pressure gases or pumps. In contrast to the Solid Rocket Motors (SRMs), liquid systems contain many components. However, the liquid systems usually have the advantage of higher specific impulse. Finally, the LRE can be tested prior to use; the SRM must rely on thorough manufacturing processes to insure high reliability. Liquid Rocket engines can be categorized as follows:
o Gas Generator Cycle (e.g. Vulcain; main stage Ariane 5)
o Staged Combustion Cycle (e.g. High Thrust Engine for Next Generation Launcher)
o Expander Cycle (e.g. VINCI foreseen for Ariane 5)
o Pressure-Fed Cycle (AJ-10 of the Delta II Launcher)
It is essential, for liquid rocket engines that the propellants immediately ignite when they enter the combustion chamber. If the ignition occurs too late after engine start up, too much propellant will be present in the combustion chamber, which
results in too much hot gas when ignition eventually occurs. This can lead to destruction of the combustion chamber. Hence, ignition is a critical step in the operation of a rocket engine. The ignition system has to deliver a certain mass flow rate of a certain temperature into a combustion chamber. Dependent on the local mixture ratios in the combustion chamber and the temperature of the hot igniter gases, ignition occurs. Depending on the category of LRE, a gas generator igniter and a turbine pump starter are sometimes required in addition to a thrust chamber igniter for the complete start-up sequence, like for the Vulcain gas generator cycle engine. The ignition system can be based on different methods: Pyrogen A solid propellant is burnt and delivers the power required to ignite an engine. The principle is very similar to a solid rocket motor only on a much smaller scale. Pyrogen igniters can only function once. Due to the solid propellants the combustion gases can contain very small particles; often this is not allowed in a LRE. Pyrogen devices can be refurbished after cleaning and assembly of new solid propellants. Pyrogen igniters are compact and cost effective when compared to other ignition systems that require one or two feed systems. It is proven technology. Currently APP produces pyrogen igniters and a starter for the Vulcain main engine of Ariane V.
Vulcain turbine pump starter Pyrogen technology development for future applications is addressed in a national programme by APP together with TNO.
Spark Torch Gaseous propellants are fed to an igniter, which ignites the propellant by an electrical spark. The combustion gases flow through a torch into the combustion chamber. Such a system requires high voltage electronics, which generate the spark. A spark torch igniter can burn many times and is thus ideal for re-startable LREs. Gaseous propellants burn clean and will produce limited to no particles in the combustion gases. The spark torch ignition system APP is developing for the VINCI engine shows hardly any wear and tear inside the igniter. Refurbishment of the spark plug after engine tests is necessary once in a while. Spark torch systems are more costly as a system in comparison to a pyrogen igniter. Dependent on the number of re-ignitions the system can become significantly cheaper than a multi-pyrogen device. Catalytic An oxidizer decomposes with the aid of a catalytic bed. The decomposed gases are hot enough to ignite a fuel (which is fed to the igniter). The combustion gases of this igniter have a temperature hot enough to ignite a LRE. The catalytic igniter can be simpler than a spark torch system. The fuels can be liquid or gaseous. Currently, this ignition system is in development by APP for the High Thrust Engine (Next Generation Launcher). In the combustion gases no particles are expected. A catalytic igniter can be refurbished. Most probably the catalytic bed deteriorates after multiple starts. A catalytic ignition system can be cheaper than a spark torch ignition system due to the absence of high voltage (HV) electronics. Hypergolic Two propellant combinations react spontaneously and strongly exothermic when they come into contact. Often the propellants are very toxic and or carcinogenic. In Europe there is a tendency to use greener propellants. As such APP has no experience with these igniters Laser A laser can be used for ignition by focussing the laser beam and therefore increasing the energy density in the focal point. Ignition can be obtained by heating a target, creating a spark or causing dissociation of molecules. These ignition technologies are known for many years and have
frequently been applied in laboratory experiments all over the world. More recently1, laser ignition was announced as a serious candidate to replace the conventional spark ignition system for automotive applications. The low level of maintenance required and the capability to ignite at high gas pressure are given as advantages of the laser technology. Laser technology also brings a large freedom of choice for the ignition location. This freedom is appreciated by engine designers as well as by combustion specialists. Ignition of a rocket engine by means of a laser is applied in the R&D environment for studying ignition phenomena2. For liquid rocket engines, laser ignition has not yet been applied in practice. The non-availability of compact, space hardened and affordable laser systems has prevented the introduction of laser ignition to actual rocket engines. However, application of laser ignition technology for automotive purposes might boost the development of laser technology. With this recent development in mind, APP decided to investigate laser ignition technology for application to liquid rocket engines. Hybrid This ignition system is patented by APP3. Only preliminary studies have been performed. In this ignition system the fuel and oxidizer are physically separated, which improves the safety aspects in comparison to a pyrogen igniter. Such a system could be as follows; an oxidizer is decomposed by a catalytic bed. These hot gases ignite a solid fuel, the total combustion gases are then hot enough to ignite a LRE. Due to the fact that the solid fuel burns up, the igniter has only limited re-start capability. Probably, the combustion gases contain a limited amount of very small particles. A hybrid igniter can be refurbished , after cleaning and installation of a new catalytic bed and solid propellant. A hybrid ignition system can be cheaper than a pyrogen system for a limited amount of re-ignitions.
3. FUTURE APPLICATIONS IGNITION SRM
A Solid Rocket Motor (SRM) is a rocket motor that uses propellant in solid form. In this paper focus is on SRMs that contain composite solid
propellant. SRMs have a long history of being used in both defence and aerospace applications. In aerospace SRMs are used as main stages (like VEGA’s P80, Z23 and Z9 core stages), large boosters (like the two Ariane V MPSs), strap-on boosters (like Atlas V with 1 to 5 solid strap-on boosters). Solid propulsion is a major component for European (and most of the world-wide) launchers and will remain to be so. The VEGA expendable launcher with three solid stages is currently under development in Europe and for the Next Generation Launcher (heavy European launcher after Ariane V) SRM(s) will be part of the expected configuration. The ignition of a SRM is done by a pyrogen igniter. As for the ignition of a LRE it is a critical step in the operation of the rocket motor. It is a single ignition and the power required to ignite the SRM is usually much larger than to ignite a LRE. Furthermore the presence of particles in the combustion gases of the igniter is no problem but beneficial for impingement of the to be ignited solid propellant surface. During the development of the VEGA igniters new technologies were implemented in the pyrogen igniters4: • 2-stage igniters consisting of pyrotechnic and
main igniter (instead of the 3 staged Ariane 5 MPS igniter5) for the P80, Z23 and Z9 stages;
• Use of a consumable carbon composite main case to simplify the configuration and reduce the overall mass and cost.
These new technologies are presently qualified.
Zefiro 9 Igniter: pyrotechnic igniter (left) and main igniter (right)
In line with the additional objective of the P80 program to demonstrate new technologies to be applied in possible Ariane 5 evolutions or follow-ups, it can be concluded that the new technologies applied in the VEGA igniters will be of interest for a New Generation Launcher booster igniter in order to reduce the recurring cost. For SRM igniters focus on future developments is on cost reduction, improvement of safety and use of non-critical ingredients.
4. TECHNOLOGIES The above discussed possible future applications for ignition systems will require updating the presently used technologies or developing new technologies in order to remain competitive, compliant to changed requirements. Therefore APP is investigating different igniter and starter technologies. 4.1 Pyrogen igniters Currently APP and TNO work on pyrogen igniter technology development within a NL national programme in preparation of an ESA programme. First the development drivers for future pyrogen igniters (both for LRE and SRM) have been identified and confirmed by APP and TNO experts in cooperation with the possible European customers (like cost reduction, use of non-critical materials and improvement of safety/transport and storage). Secondly the technologies have been identified that could help to meet one ore more development drivers. Then the technologies that give the best opportunities for APP and TNO to meet the development drivers have been chosen. Presently the demonstration of the feasibility of the selected technologies is ongoing, with a focus on experimental activities up to maximum TRL 3. In 2011 the evaluation of the potential new/improved pyrogen igniter technologies and the definition of steps necessary towards actual application will be finalised. From a large number of technologies the following subjects are chosen for feasibility demonstration: • Develop a tuneable baseline non-chlorine
igniter propellant, focussing on:
o Tuneable thermal performance o Use of non-critical ingredients
(commercially available, no ITAR restrictions, REACH compliant)
• Develop a tuneable baseline AP-based igniter propellant, focussing on:
o Tune baseline AP-based SRM propellant from TNO to igniter propellant
o Use of non-critical ingredients (commercially available, no ITAR restrictions, REACH compliant)
• Production of large size propellant grain Small scale sample tests will be performed to simulate a multi-batch propellant grain in preparation of producing grains larger than P80 igniter grains in APP facility.
• Simplification of ignition train In order to test the use of simpler and non-explosive initiators some initial tests on both Ammonium Nitrate-based and AP-based propellant will be performed.
The new technologies already qualified within the VEGA igniter programme (2-stage igniters and consumable composite case, see also §3) will also be taken into account for other future pyrogen igniters, especially for SRM ignition. 4.2 Spark igniters So far APP has produced several spark torch igniters: • Vinci igniter • P8 igniter • FAST II igniter These igniters are all of the spark torch type and make use of gaseous propellants. H2-O2 for Vinci and P8 while the FAST II igniters uses CH4-O2 as propellants. The sparks are generated by one or more spark plugs, while the high voltage for the spark plugs is delivered by an exciter. The P8 and FAST II igniters are designed for test purposes and are finalized projects. The development of the Vinci ignition system is intended for flight. The development is ongoing and aiming for flight in 2016. APP has been working on the Vinci ignition system since the start of the development of the Vinci engine 6 . The development of Ariane 5 Midlife Evolution has recently been started and thus also the development of the engine and the ignition system.
So far, the ignition system has been designed as a non redundant system. To comply with the latest specifications, the system is updated to include electrical redundancy. This includes actuation of electrical valves and the sparking electronics from the high voltage generation in the exciter to the spark generation at the spark plugs. The redundant igniter therefore features two spark plugs. The first prototype of this igniter has been tested in 2009. During this test campaign a 100% ignition reliability was obtained, while the nominal lifetime was reached and exceeded without problems.
Electrically redundant Vinci igniter This test campaign was conducted at normal ambient conditions of 1 bar. Since the system however has to operate in vacuum, much attention has to be paid to reliable operation in vacuum. The low pressures in particular influence the high voltage electronics behaviour. To prevent undesired discharges, the path with the highest voltages will be made as short as possible; the high voltage generating components, in particular the transformer coil and the spark plug will be built into one assembly, that can be properly insulated and closed hermetically. By this method the application of a high voltage cable, connecting the spark plug and the high voltage electronics and its connectors can be omitted. The failure modes connected with this construction are therefore eliminated by design. The first prototypes of the Coil-on-Plug assemblies will be produced and tested in 2010. After successful testing of the Coil-on-Plug assembly, the assembly will be adapted so that it will fit into the available envelope for the redundant igniter. The entire assembly is then foreseen to be tested on thrust chamber and engine level.
4.3 Catalytic igniters In the framework of ESA’s Future Launcher Preparatory Programme (FLPP), APP and TNO are developing a new generation ignition system based on catalytic decomposition of Rocket Grade Hydrogen Peroxide. The new ignition system features the following benefits over existing systems: - No sensitive high-voltage initiation system
required, as is the case for spark torch igniters - Re-usable - In-flight re-ignition capability - Fuel flexible - ADR legislation not applicable for system
hardware Initially the ignition system will be designed for the High Thrust Engine, a stage combustion engine under development by a consortium of EADS Astrium, Avio and Snecma (Joint Propulsion Team). APP and TNO have successfully completed a catalytic igniter test campaign end of 2006 testing a mono-propellant catalytic igniter using a Silver based catalytic bed. In order to raise the flame temperature for achieving ignition of a rocket engine a bi-propellant igniter was developed using liquid ethanol as bi-propellant fuel. This igniter was tested successfully end of 2008. Currently APP is developing a fuel-flexible bi-propellant catalytic igniter which will accept both gaseous methane and hydrogen as bi-propellant fuel. The igniter is intended for application on LOx-LH2 and LOx-LCH4 engines. The capability of using the same propellants for the igniter as already in use for the engine reduces the number of different fuels to be handled on the platform.
Catalytic bi-propellant igniter tested at TNO.
The principle of the catalytic bi-propellant igniter relies on Rocket Grade Hydrogen Peroxide being decomposed over a catalytic bed producing hot oxygen rich gasses. Igniter fuel is injected downstream of the catalytic bed in the flow of hot decomposition gasses aiming to create a oxygen-fuel mixture favourable for ignition. As the oxygen fuel mixture is heated above the auto-ignition temperature by the incoming hot decomposition gasses, the mixture ignites. Sustained combustion provides the required thermal power output for ignition of the engine or pre-burner. The currently developed Methane/Hydrogen catalytic bi-propellant igniter is designed to be compliant with the following requirements: - Thermal power ≥ 500kW - Flame temperature ≥ 2000K - Operating time ≥ 2.2s - Response time < 60ms
(excl. valve response time) - Lifetime 26 ignitions In order to prevent overheating of the igniter structure and in order to keep the interface temperatures within the required limits the igniter features active “dump combustion” cooling. Part of the fuel is directed around the igniter, cooling the structure. The coolant fuel is injected downstream of the igniter tube into the igniter core flow. Here the fuel is combusted adding to the igniter thermal power and creating a relative wide flame, beneficial for ignition of the engine’s main chamber or pre-burner. As part of the thermal power is generated outside the igniter, the thermal load inside the igniter can be kept to a minimum. Stand-alone testing of the newly designed catalytic bi-propellant igniter is planned June 2010 at the TNO test facility. Testing will also involve testing of an alternative gamma-Al2O3 based catalytic bed. In a later phase a so-called ‘subscale’ bi-propellant catalytic igniter and propellant feed system will be developed for testing on the Astrium P8 test-bench at Lampoldshausen. 4.4 Laser igniters The application of laser ignition for LRE is being discussed. In principle laser ignition could also be investigated for ignition of a solid propellant or
the pyrotechnic charge of an initiator. This is currently not pursued by APP. Laser ignition technology has to compete with other technological means for liquid rocket engine ignition: • Pyrogen igniter (single ignition, high reliability,
proven technology) • Spark torch igniter (multiple ignition, more
complex, high voltage electronics, proven technology)
• Catalytic igniter (multiple ignition, no electronics required, low TRL level)
• Resonance igniter (multiple ignition, complex thermal management)
In potential, laser ignition can have distinct advantages to these technologies: • Improvement of service life (compared to using
a spark plug) • The laser ignition system might be more
compact and lighter than corresponding conventional igniter
• The possibility of multi-point ignition • Possibility of initiating combustion reactions in
a desired spatial region, away from the walls where mixtures are not optimum
• Laser ignition has minimum heat loss and flame quenching
• Laser can ignite mixtures at pressures higher than the limiting ignition by the conventional spark ignition system7
• The application of optical fibres to ignite at multiple locations or chambers with one laser unit (and the associated cost and mass benefit)
It is not yet clear if these theoretical advantages can be realized in practice. If yes, it is also not clear if these advantages are strong enough to compete with the existing systems which are proven and reliable. A first important consideration for evaluating laser ignition technology is the required thermal energy to achieve ignition. The theoretical ignition energy for the ignition of a mixture of a fuel with gaseous oxygen is in the order of 0,1 to 10 mJ. In practice however, high turbulence levels, two phase flow and very low (cryogenic) temperatures cause a high probability of flame quenching. Therefore, in general rocket engine ignition is performed with a hot gas flow from a torch with a thermal power of several hundreds of kW. Maybe locations can be identified in the
combustion chamber where ignition is possible in the range of the theoretical ignition energy, without flame quenching. In such case, laser ignition is might be the only possibility to deliver this energy to the desired location. A simplified ignition system, eliminating the need for a torch, might be the result of this. Another consideration for laser ignition application is ignition at multiple locations. This is relevant for large rocket engines as several combustion chambers are to be ignited (main chamber and gas generator or preburners). It is also relevant for future Reaction Control Systems and Attitude and Orbit Control Systems when thrusters with green propellants are applied, with a need for ignition. A single laser, with fibre optics to transfer the laser energy to the required location, might give cost and mass benefits compared to the other technical options. Today, however, laser optical technology is not yet capable of providing a long lasting reliable transmission at the high energy density required for this application. On the other side, laser
miniaturisation is progressing at a rapid pace driven by the automotive application. To conclude on laser ignition it can be stated that this technology is promising, but for rocket engine application still at a very low TRL. APP will continue in monitoring the development of this technology and is currently involved in creating partnerships to perform investigations into the most relevant aspects of laser ignition technology.
5. APPLICATION – TECHNOLOGY COMBINATIONS
In the table below an overview is given of the igniter technologies that have been discussed in this paper. Per technology the suitability for the different engine / motor types is shown. Furthermore the main advantages and disadvantages per technology are given.
Ignition technology
LRE single ignition
LRE multiple
SRM Pro Con
Pyrogen ++ +* ++ TRL 9Price per ignition relatively low (in comparison to spark torch and catalytic)Predefined performanceCompact (no feed system)
Class 1 devicePredefined performanceSingle ignition per device
Spark torch + ++ - TRL 6Re-ignitableClean combustion productsEasily re-usable for tests
High voltage electronic partsRequires feed system for oxidiser, fuel and possibly purge gas
Catalytic ++ ++ - Re-ignitableNo high voltage partsRe-usable after refurbishmentFuel flexible
TRL ~4Catalytic bed needs refurbishmentRequires feed system (for oxidiser, fuel and possibly purge gas)
Laser ++ ++ -** Compact and lightMulti-point ignition, in desired spatial regionRe-ignitable, re-usable
TRL ~2Theoretical advantages, unknown disadvantages
++ Very suitable+ Suitable- Not suitable
* Multiple devices are required for multiple ignition** On large SRM scale solid propellant ignition by laser is not (yet) deemed feasible
Applicable in Specific characteristics
APP covers a broad range of ignition technologies. With these different technologies APP is able to find technically and commercially attractive solutions for the complete spectrum of required ignition systems foreseen for current and future applications in aerospace.
6. ACKNOWLEDGEMENTS APP would like to thank TNO for their continuing cooperation in the development of the different ignition technologies. Furthermore we would like to acknowledge ESA and the Netherlands Space Office for their support in the field of pyrogen, spark, catalytic and laser ignition
7. REFERENCES 1 University of Liverpool, http://www.liv.ac.uk/researchintelligence/issue36/laserignition.htm 2 Manfletti C., “Theoretical and Experimental Discourse on Laser Ignition in Liquid Rocket Engines”, 2009-a-02, 27th International Symposium on Space Technology and Science, Japan, July 2009. 3 Patent pending (“Igniter for a rocket engine, method for ignition of a rocket engine”) 4 Brauers B., et al., “Qualification and production of the VEGA Solid Rocket Motor Igniters”, AIAA 2009-5322, 45th Joint Propulsion Conference, Denver, Colorado, USA, August 2009. 5 Santili, M., Fontana, A., “Ariane 5 MPS Igniter Subsystem – Design Approach and Development Phase Results” AIAA paper 94-3065, 30th Joint Propulsion Conference, Indianapolis, July 1994. 6 G. Frenken et al., “ Development status of the Ignition system for Vinci”, AIAA-2002-4330, 38th Joint Propulsion Conference, Indianapolis, Indiana, July 2002. 7 B. Bihari, S. Gupta and R. Sekar, “Development of an advanced laser ignition system”, AASRE Conference, Downey, CA, September 2007.
