Delivering engine demonstrators for competitive evolutions of the european launchers

The year 2014 was one of achievements for the propulsion projects of the Future Launchers Preparatory Programme (FLPP). Highly relevant results were produced by hot-fire testing of three different engine demonstrators — a solid propulsion demonstrator focusing on pressure oscillations, a 5-kilonewton storable bipropellant engine, and a 25-kilonewton hybrid propulsion demonstrator. These concrete achievements were accompanied by the fast progress of a 115-kilonewton cryogenic expandercycle thrust chamber demonstrator which reached Preliminary Design Review (PDR) by end 2014, with hot-fire tests scheduled in 2017. These achievements give a solid grounding to the continuation and expansion of engine demonstrator projects at ESA Launchers.


INTRODUCTION
The FLPP of ESA Launchers has been since 2004 the technology investigation and maturation programme of the ESA Launchers Directorate. It manages several §agship integrated technology demonstrators as well as transversal generic technology maturation projects. The §agships are compact and §exible projects maturing technologies and system architectures from Technology Readiness Level (TRL) 3 to 6, providing concrete results for the de¦nition of the technical content and the speci¦cations of subsequent development programmes.
The impact of preparatory demonstrators is particularly important in propulsion where the time and investment scales are large. Engine demonstrators are designed as prototypes assembling consistent sets of technologies. They are developed up to hot-¦re testing, maturing the technologies in a system-driven project frame and in a representative engine environment. Thereby the technologies reach TRL6 fully integrated in engine systems, ready to be transferred to turnkey developments and §ight quali¦cation programmes. To a lesser extent, they can also support a running development programme by testing attractive alternative technologies embedded in the demonstrators.
Last year saw the hot-¦re testing of three di¨erent engine demonstrators in excellent conditions, delivering very good results. A solid propulsion demonstrator focusing on pressure oscillations was ¦re tested in March 2014 near Bordeaux, France, followed by a 5-kilonewton storable bipropellant engine demonstrator tested between August and October 2014 in Lampoldshausen, Germany, and ¦nally, a 25-kilonewton hybrid propulsion demonstrator was tested from October 2014 until December 2014 in Raufoss, Norway. In addition to these tests, a 115-kilonewton cryogenic expander-cycle thrust chamber demonstrator, started in 2013, made fast progress, reaching PDR at the end of 2014 and, thereby, keeping to the original planning for hot-¦re tests in 2017. In the frame of this latter project, laser ignitions have been successfully performed at the P8 bench in Lampoldshausen, Germany. Generic and transverse technologies are also matured with this project, such as cryogenic valves for smart engine transients and combustion regime regulation, selective laser melting, and additive manufacturing.
These positive achievements are the result of, among other factors, an optimal tailoring of project management and quality assurance requirements, striking the right balance between the speed and a¨ordability of project execution, on the one hand, and the safety of operations and quality of technical results, on the other. In the frame of these propulsion demonstrator projects, the di¨erent industrial contractors are empowered with the design authority, allowing them the freedom to respond optimally to the technology and functional requirements de¦ned by ESA FLPP, chosen to maximize the bene¦t for future evolutions of the European launchers. The potential applications on launcher evolutions are considered at several levels, from the spin-o¨of an individual technology (e. g., laser ignition and sandwich nozzle) to the validation of a full protoengine as the foundation stone for a subsequent fast new engine development. In this respect, synergies between Ariane and Vega are considered and fostered. This paper presents the di¨erent propulsion demonstrators developed in FLPP, their current status, and future perspectives starting with the hot-¦re tests of 2014 and then describing in more detail the 115-kilonewton cryogenic expander-cycle thrust chamber demonstrator and its associated technology tests.

2014 ¡ A YEAR OF HOT FIRE TESTS
This chapter presents in more detail the three demonstrators that underwent hot-¦re testing last year, the logic behind their design, the test con¦guration, initial results, and the next steps.

POD-X ¡ a Solid Propulsion Demonstrator Focusing on Pressure Oscillations
As a major achievement, a pressure oscillation demonstrator (POD-X) developed in cooperation with CNES, the French Space Agency, culminated in a hot-¦ring test in spring 2014. The ¦rst POD-X ¦ring (con¦guration demonstration model (DM) scale 2/9, Fig. 1  The test validated the POD-X experimental tool in terms of functional, mechanical, thermal, and ablative behavior as well as the interfaces with the test bench (Fig. 2). Presented in October 2014, the Level 1 analysis of the test data has enabled the clear de¦nition of the domains requiring further analysis e¨orts and investments.
The POD-X Critical Design Review was held in May 2015 and concluded on the success of the DM ¦ring test while delineating the future perspectives and options for the next POD-X test campaign. In front of future European launcher developments, re §ections are on-going on the best programmatic options in order to decide on the next demonstrator con¦guration to be tested.

Five-Kilonewton Storable Engine Demonstrator
The goal of this project, started  in mid-2010, is to demonstrate technologies for a pressure-fed engine in the thrust class of 3 to 8 kN (Fig. 3), preparing versatile propulsion solutions for the European launchers. One leading potential application is the propulsion of the fourth stage of the Vega-C launcher. The same engine could also provide an alternative solution for launcher versatility as more diverse payload mission pro¦les appear on the launch market, especially with payloads using all-electric propulsion.
The technologies to be matured in this project were chosen through an analysis of market needs, system studies, and earlier test programmes. The main identi¦ed technologies are: nitrogen tetroxide (NTO) regenerative cooling (all monomethylhydrazine (MMH) as regenerative cooling is not enough below a certain engine size); MMH ¦lm cooling; improved injectors (higher e©ciency and lower manufacturing cost); regenerative nozzle for storable engines (Aestus of Ariane 5 does not have a regenerative nozzle); innovative design of acoustic dampers; seamless uncooled nozzle skirt; and electrovalve demonstrators.

Demonstrator architecture and technology tests
After the completion of the concept studies and preliminary design, several technological component tests were carried out. Meanwhile, the engine long lead time items were released. The technological component tests covered two distinct areas: (i) single injector characterization; and (ii) NTO cooling physical characterization.
Several injector designs have been manufactured, tested in water §ow conditions and, eventually, in real propellant hot-¦ring conditions at the P2 bench in DLR Lampoldshausen, Germany. These tests supported the design justi¦cation of the injectors, especially with re- spect to stability and e©ciency and drove the choice of injectors to be implemented on the engine demonstrator.
Regarding NTO cooling characterization, an electrically heated device was installed around a tube channelling the NTO. The cross section of the tube could be adapted to support the ¦nal design of the cooling channels on the engine. The electrically heated device was equipped with temperature sensors so as to measure accurately the heat §ux toward the tube and in the NTO (Fig. 4). The NTO §ow rate could be modulated. This experiment provided precious results to justify the cooling channel design on the engine and ascertain the operational domain in hot-¦re testing.
For the sake of cost e©ciency, the project ¦rst went into hot-¦re testing under sea-level conditions. Therefore, the ¦rst version of the engine demonstrator (5-kilonewton nominal thrust), tested in 2014, had a short expansion ratio in order to avoid §ow separation (Fig. 5). With these constraints, it was decided to test the NTO regenerative cooling (one of the major new technologies) in the cylindrical part of the combus- tion chamber (Fig. 6).
Thanks to the modularity of the demonstrator concept, a con-¦guration with NTO cooling in a regenerative nozzle part under altitude simulation conditions, with the rest of the engine cooled with fuel, is also possible. This latter con¦guration is closer to the design of a possible future §ight engine.
The results from the ¦rst hot ¦re campaign show, however, that with the current design and certain restrictions, a pure MMH cooling might be possible in a thrust class larger than 5 kN. This would be bene¦cial to reduce the complexity, cost, and weight of a future application

Hot ¦re tests ¡ 2014
The manufacturing of the engine demonstrator and its spare parts was completed as planned in spring 2014. After successful acceptance checks, the engine was delivered at the test bench site, P2, in Lampoldshausen, Germany, and integrated into the test facility (Fig. 7).
The ¦rst successful hot-¦ring test of this engine took place on August 19, 2014, and a total of 60 tests were performed, comprising 54 at sea level and 6 under vacuum at ignition (Fig. 8). The run times ranged from 1 to 110 s, leading to a total run time during the campaign of 422.5 s. The tests covered a wide range of parameters, for example, mixture ratio, combustion pressure, and propellant temperatures. Generally, the hardware performed well and as expected in the complete tested domain. The cooling performance of the regenerative cooling was designed so as to be su©cient in the whole demonstration domain. A ¦lm cooling device was designed as a dedicated technological objective for low-thrust engine applications and tested with two mass §ows in several load points. Vacuum hot-¦rings with short duration into a static vacuum were performed with normal and low propellant temperatures. Those exhibited a similar behavior as the ground level tests, without any signs of instability.
The test campaign lasted from August until October 2014 and provided valuable input for model re¦nement and future developments.

Hot ¦re tests ¡ 2015
The next planned test campaign (CAP-1) will use a capacitively cooled thrust chamber. The preparations for this campaign are progressing, with the ¦rst test planned for summer 2015. This test campaign will focus on the dynamic behavior of the thrust chamber (eigen-frequencies, damping, etc.), using bombs to excite high-frequency modes during the hot runs.

Subprojects
Several subprojects were started to mature peripheral technologies needed for an engine application. Those concern mainly valve and nozzle technologies: two valve demonstrator projects (by Techspace Aero in Belgium and Test Fuchs in Austria) were successfully performed up to a PDR status. One is based on a fully electric actuator, counting on a force-balanced poppet. The other relies on a more classical electropneumatic actuator; and the nozzle (both cooled and uncooled) design and manufacturing will be prepared, for testing of the complete engine in altitude simulation. Already, a seamless uncooled nozzle skirt has been successfully manufactured by Franke in Switzerland out of a single metal sheet (Fig. 9). The process principle is the skewed rolling of a circular blank. The ¦nal wall thickness is around 0.5 mm.

Application to the Vega launcher
After the positive results of the regenerative hot-¦ring test campaign in the third quarter of 2014, a detailed analysis of a derived §ight engine powering the fourth stage of the Vega-C launcher has been running since the early days of 2015. This investigation, preparing for a new engine development and a ¦rst §ight in 2018, is a concrete example of a possible transfer from a technology engine demonstrator to a launcher application. In such a context, the technical results obtained in FLPP are solid and precious references for swiftly de¦ning the engine concept design and development logic.

Hybrid Propulsion ¡ Sounding Rocket Applications
Regarding hybrid propulsion, the FLPP is developing a 25-kilonewton demonstrator engine, with the support of the Norwegian space agency (NSC), aimed at modular sounding rocket or microlauncher applications. A test campaign with an initial battleship design was performed on a Nammo test bench as from October 2014. After two monopropellant tests to tune the catalyst, the demonstrator of the Hybrid Unitary Motor performed 6 additional tests with both propellants (liquid and solid). The run time was progressively increased to 25 s during this campaign. The test results were in line with the predictions considering the thrust and chamber pressure. Based on the results from the ¦rst three ¦rings (Con¦guration 1) where the fuel mass §ow was lower than desirable at the end of the hot run, the design of the motor, especially the injector, was optimized (Fig. 10). Three additional tests with the adapted design (Con¦guration 2) showed an improved fuel regression rate in line with the predictions.
As the tests in 2014 have demonstrated the high potential of this technology, further optimization of the motor design, including a §ight-like con¦guration (weight optimized and minimized residual fuel) is currently performed. This design will be extensively tested and re¦ned in a test campaign starting at the end of 2015. In addition, an in- §ight demonstration using a speci¦cally developed oxidizer system on a sounding rocket is planned. In the long term, the results from this project will be used to develop a modular unitary motor, which can be clustered to achieve tailored characteristics for all stages of di¨erent sounding rockets and other applications.

THE EXPANDER TECHNOLOGY INTEGRATED DEMONSTRATOR
The Expander-cycle Technology Integrated Demonstrator (ETID), started mid-2013, is a major constituent of the FLPP Period-3 programme. It prepares competitive evolutions of upper stage propulsion for Ariane 6 and Vega by assembling technologies that pave the way for the next generation of cryogenic upper stage engines in Europe. The project was able to build on FLPP heritage in cryogenic engines, including: hot-¦re testing demonstration of the Vinci engine in 2007-08 (FLPP 2.1); high-thrust engine PDR, subsystem tests (FLPP 2.1, 2.2); and cryogenic expander-cycle thrust chamber concept studies in 2012 (FLPP 2.2).
The key initial input to this project was a re §ection on the competiveness factors of upper stage engines and how best to respond to them. This process resulted in the high level requirements, as outlined in Table 1.  The mission requirements document (MRD) for this engine was centered on these competiveness factors through high-level speci¦cations; however, few technologies were imposed ¡ the technology push was to come from industry responding to the MRD. Through its inherent versatility, the demonstrator was to optimize potential applications and, therefore, synergies between applications ¡ Ariane 6 upper stage propulsion incremental evolution, Vega-E third stage propulsion (LOx/Methane) but also considering several technologies ben-e¦ting the Vulcain 2 incremental evolution plan.
Considering a ¤test as you §y¥ principle, the representativeness to a potential §ight engine was ensured by ¦rst de¦ning the §ight engine ¡ Flight Engine Image (FEI), and then applying testing constraints to arrive at the engine demonstrator design ¡ ETID (Fig. 11). The core of the project is dedicated to the design, manufacturing, and testing of the thrust chamber technology demonstrator ¡ ID#1. This ¦rst demonstrator is poised to be hot-¦re tested in the ¦rst half of 2017. It should replicate in full scale an expander-cycle thrust chamber designed for an optimized §ight engine.

Design of Flight Engine Image
As outlined above, the ETID is designed to be as close as possible to a future competitive evolution of the expander-cycle engine. To this end, an image engine is considered in this project as the virtual image of the target §ight engine, optimized for the better performance and far lower cost of future evolutions of the European launchers. This image engine is used as a reference frame to identify more con¦dently the technologies to be matured and to de¦ne the most e¨ective demonstrator. The demonstrator maximizes the technological content relevant to the ful¦lment of the image engine missions. The ¦rst step in the project is, therefore, the design of the §ight engine image, which is then adapted to a demonstrator compatible with the P3.2 test bench.
Supporting the mission requirements set out in Table 1, the following elements were also speci¦ed for the FEI: inexistent, or negligible, water condensation on the inner wall of the thrust chamber assembly (TCA); strict cost and lead time requirements on an accepted FEI in series production ¡ consideration of the industrialization, production, and exploitation activities from the start of the design cycle; inspectability and replacability of all critical items ¡ nondestructive inspection (NDI) without disassembly, certain subsystems to be designed as line replaceable units (based on HM7B operations feedback); and technology requirements: • new, low-cost, combustion chamber liner and jacket materials to be used; • use of SLM on injector head, turbopumps; • investigation of aluminum-alloy for the regenerative nozzle; • minimization of helium consumption; • lightweight material for pipes, no bellows; and • improvement (mass, reliability, and accuracy) of the measurement chain.
The technical answer of the prime contractor, supported by their subcontractors, to these high-level requirements is detailed below, outlining the design of a competitive evolution of the expander-cycle engine.

Design methodology and modeling
The project makes best use of the latest propulsion modeling capabilities. These computer models contribute to: establishing the steady-state model; de¦ning the reference operating point; mapping the operating domain; identifying the functional speci¦cations for the subsystems; checking the stability of the engine cycle; characterizing and optimizing the start-up and shutdown transients, de¦ning the valve sequences; optimizing the mechanical design; characterizing the thermal behavior, and spotting the potential issues; and directly supporting the detailed design of the thrust chamber and the regenerative cooling channels.
New design technologies are also investigated and developed, such as RALP ¡ Reliability Analysis and Life Prediction and Probabilistic Methods. Probabilistic methods allow sensitivity assessments of input parameters, identifying those that have the most impact on the life-spread, additionally quantitative margins of safety can be assessed, with the potential to allow mass reductions of up to 10% in certain components. This methodology has limited applicability due to the need for parameterization of all inputs and the need of statistically signi¦cant material sampling, such that only simpler forms can be considered. Proposed applications are pipes, bolted connections, thickness of coatings, etc. Potential applications have been found on the ID#1 hardware on nonsafety critical items. Also, Thrust Chamber Life Prediction Based on Survival Analysis is investigated, the goal being to establish a statistical model for failure prediction of liquid rocket engine combustion devices based on empirical knowledge available from the development and test history of comparable components.
The hot-¦re test results of ID#1 will be compared to the numerical models, improving the validation status of the models and design methodologies used within the project, making them applicable for future development activities.

Thrust chamber assembly
The thrust chamber pro¦le of the demonstrator is speci¦ed to be identical to the one of the optimized §ight engine.

Combustion chamber
The combustion chamber stays on the principle which has proven its mechanical robustness and cooling e©ciency on HM7, Vulcain, and Vinci. This demonstrator project will contribute to enhancing this combustion chamber design, with the use of lighter and cheaper alloys, and optimized cooling channel design. The cooling channel design is optimized for the speci¦ed engine life, maximum heat pickup, minimum pressure drop, and avoidance of massive water condensation. As the performance of the expander cycle is mainly determined by the chamber heat pickup and the fuel turbopump e©ciency, the possibility to increase the heat transfer in the combustion chamber will be tested. New low-cost materials are investigated for the liner and jacket as well as the optimization of production times.

Nozzle
The project is an opportunity to implement a regenerative nozzle part and continue the technological progress on the ¤sandwich¥ design. The regenerative nozzle part reduces the heat pickup assigned to the combustion chamber, therefore optimizing the global mass of the TCA. The most mastered stainless steel option is taken as reference for the realization of the nozzle item for the ¦rst demonstrator in 2017, while the maturation of manufacturing processes is continued for an eventual aluminum alloy regenerative nozzle. On the industrial standpoint, the FLPP contributes to the deployment of an automated manufacturing process and tools in industry, for the bene¦t of a future reliable and high performance production line of such nozzles. In addition, the avoidance of water condensation at the inner wall has led to evolved coolant circulation paths, designed and analyzed by numerical models.

Injector head
The injector head is identi¦ed as a good candidate to demonstrate the additive manufacturing. It is multifunctional, mechanically static, and of complex geometry. Dramatic cost reductions could result of this application. As the injector head is a massive part of the thrust chamber assembly, it is also the object of extended material trade-o¨s in order to reduce its mass. Innovative Advanced Porous Injector (API) designs on the face plate are also investigated.

Igniter system
The ignition system, traditionally installed at the center of the injection head, is proposed to be displaced on a ring mounted under the injector face plate. This layout would bring advantages in terms of accessibility and maintainability of the ignition system and would ease the implementation of multiple or redundant igniters for reliability and in- §ight reignition capability. Innovative laser and spark ignition technologies are to be implemented on ID#1. They are attractive as they are disconnected from the engine functional transients and would result in signi¦cant mass reductions; however, the constraints induced by their electronic systems must be carefully assessed.

Other subsystems Turbopumps
Turbopumps and valves are considered in the FEI design according to a functional architecture similar to the one of the Vinci engine. In particular, valves bypassing the turbines will trim the thrust and the mixture ratio. The ETID is designed such that these components can be added in a modular way in later hot-¦re tests, after ID#1. Regarding the oxygen turbopump, robust dynamic seal package design and mass reduction are seen as the most interesting areas of progress. Moreover, the oxygen turbopump may be a one-to-one synergy between an LOx/LH 2 expander engine and a LOx/Methane engine. The hydrogen turbopump is a key component of the cryogenic expander engine, thanks to the limitation of the engine thrust, it is proposed to aim for a single stage pump impeller, thus contributing signi¦cantly to the general cost and mass reduction objectives of this important subsystem. For both turbopumps, objectives are also assigned in the directions of additive manufacturing, §uid bearing technology, and fast chilldown.

Valves
The cryogenic engine valves bring their own technological axes of progress to the project. The most prominent axis is about the systematic application of electric motor actuators. This choice enables a simpli¦cation of the engine functional layout and much wider possibilities of engine regulation, control, and autotests. Regarding the engine control, the electric motor actuators would be an e©cient tool to explore and optimize the start-up and shutdown engine transients. They would also be involved in the engine closed loop mode. Last but not least, electric motor actuators may provide solutions to enhance the global engine failure tolerance in §ight. The valves are also candidates for additive manufacturing.

Pipes
Regarding the mechanical architecture, a global and systematic process is applied to the choice of material, shapes, and dimensions for the pipes in order to reach a global mass optimum at engine level. This approach includes the search for the best positions of the subsystems. The pipes are speci¦ed to be without bellows.

Engine control system / health monitoring system
The project opens the door to an analysis at engine system level of the advantages of an electronic controller. The controller could incorporate and execute the engine sequences, in particular the valve transients enabled by the electric actuators. The launcher onboard computer would only have to send the engine mode switch orders and receive the signals of proper execution sent back by the engine controller. The engine controller would also enable an automated check of the engine once it is integrated on the upper stage and on the launcher. This would provide very reproducible and low cost leak tests, valve actuation checks, igniter status, etc. all along the launcher production line up to lifto¨. This would serve launcher integration complexity and lifto¨availability, which are the major ¦elds for competitiveness. The engine controller could also contribute to a better engine in- §ight telemetry.
As part of the work on engine HMS, an improved sensor technology envisages the following goals: a static/dynamic pressure sensor, which records the dynamic pressure oscillations while preserving the correct static (dc) level of the signal; and minimizing the temperature sensitivity leading to a drift of the sensor.

Technology Development Testing
Although still in the design phase, the ETID project also contributed to the large number of experimental results obtained by FLPP propulsion projects in 2014. Successful subscale test campaigns, closed with Posttest Reviews in December 2014, for both laser and direct spark igniters have demonstrated the Figure 12 Laser igniter testing on the P8 test bench, Lampoldshausen (Airbus Safran Launchers GmbH) potential of these technologies for multiple ignitions in relevant conditions. As an example, over 30 laser ignitions were demonstrated at the P8 test bench in Lampoldshausen, Germany (Fig. 12). The ignitions took place in a variety of propellant inlet conditions covering both initial and reignition environments. A subscale additive-layer manufactured injector head, coming from DLR national R&D programme, was also successfully demonstrated during this test campaign, con¦rming the choice of this technology for ID#1.
These new igniters technologies, with promising mass and performance ¡ reliably enabling multiple ignitions, have the potential to be applied to other combustion elements ¡ such as the gas generator of Vulcain 2. Further testing on the P8 test bench in 2015 will help to con¦rm the interest of this application.

ID#1 ¡ Hot Fire Tests in 2017
The ID#1 demonstrator will be representative of the TCA of the §ight engine, outlined in paragraph 3.1.2, with additional technology, such as light-weight pipes, as passengers during the hot-¦re tests. The thrust chamber is composed of a combustion chamber body, a regenerative nozzle part, and an uncooled nozzle skirt. As a part of the adaption of the §ight engine design to the demonstrator design, both nozzle parts have had to be shortened to ¦t the constraints of the P3.2 test bench. Di¨erent igniter technologies will be tested on the demonstrator, mounted radially on an igniter ring under the injection head. This TCA successfully passed PDR in September 2014 and is on course for a Mission Readiness Review in summer 2015.
Two sets of test hardware will be produced to have redundancy so that the test campaign is not delayed in case of hardware failure but also to enable testing di¨erent technologies. A nonexhaustive list of technologies included on ID#1 is given below: enhanced heat pickup; cost-e©cient combustion chamber liner and jacket; additive manufacturing for injector head; laser igniter, spark igniter, radial positioning for access, and redundancy; ¤sandwich¥ technology for regenerative nozzle (stainless steel for hot-¦re tests, and development work on aluminum); optimized cooling channel design and cancellation of local overcooling; lightweight uncooled nozzle skirt; To ensure timely delivery of test hardware in this fast-to-demonstration project, the manufacture of certain nozzle components is already undertaken on the forerunner (Fig. 13).
The ID#1 hot-¦re tests are planned for 2017 on the P3.2 test bench in DLR Lampoldshausen, Germany (Fig. 14). This test bench o¨ers high pressure pro-pellant feeding capabilities enabling the ¦ring of thrust chambers without turbopumps. At the P3.2, the project also wishes to make progress in the accurate measurement of thrust and high accuracy in other measurements such as propellant mass §ow. The ID#1 will be heavily instrumented, providing the maximum amount to experimental data with which to qualify technologies and analysis tools, the P3.2 must, therefore, design and implement a highly performant measurement system. The activities with DLR-Lampoldshausen on the preparation of the P3.2 test bench were o©cially started at the end of 2014, with the PDR of the P3.2 test bench modi¦cations planned for summer 2015.

Prospectives
The ETID is an important brick of the propulsion technology roadmap, slated to provide mid-and long-term solutions serving the future competitiveness of Ariane 6, Vega, and beyond. Within the current phase of the project, the ¦rst demonstrator, ID#1, will test innovative TCA technologies under representative conditions on the P3.2 test bench by 2017. In addition, the engine speci¦cations for a §ight image engine, based on the tested technologies, will be §owed down to a complete set of system and subsystem speci¦cations, outlining the design of a potential §ight engine.
Preparation for a possible continuation after the current step is underway, with the potential to maximize the technological returns of the demonstrator. The following are considered to be potential continuations for the project, depending on the occurrence of future opportunities: production and test of a partially or fully integrated ID#02 engine demonstrator, with turbopumps and valves, updated based on the results of ID#01 tests, paving the way for the development of a very competitive LOx/LH 2 expander-cycle engine evolution; and methanization of ID#01, including hot-¦re test campaign, and subsequent testing of a complete LOx/Methane engine demonstrator in the 100kilonewton thrust category.
Additionally, technology development work within this project already has the potential to contribute elements to the improvement of the existing cryogenic engines Vinci and Vulcain 2, for example, the laser ignition, where representative testing of this technology will be demonstrated within FLPP in 2015.
In conclusion, the ETID prepares e¨ectively for low cost, high performance, and high reliability upper stage propulsion, in other words, a highly competitive evolution of the expander-cycle engine for the European launchers as from the mid-2020s. It provides also a real opportunity for synergies with a LOx/Methane engine as a possible very low cost upper stage application.

CONCLUDING REMARKS
The achievements described in this paper set the pace for the continuation and expansion of such engine demonstrator projects at ESA Launchers. The propulsion activities in the ESA FLPP programme cover a large array of categories, from liquid propellants, both cryogenic and storable, to hybrid and solid. The FLPP is thereby preparing concretely and e¨ectively the future competitiveness of the European launchers, with the support of the participating Member States and industrial partners. These propulsion projects are important for the European industrial contractors involved, as they provide to their key competence centers many large and cutting-edge tasks in technology maturation, system design, manufacturing, and testing.
Being at the upstream end of the development chain, the engine demonstrators are the best moment to introduce new structuring requirements with high potential impacts on future competitiveness such as dramatic cost and mass reductions. The technologies are evaluated on their ability to reduce costs, deliver versatile performance, and contribute to lean industrial processes for future evolutions of the European launchers. Integrated engine demonstrators are the most time-and cost-e©cient way to assemble and mature the elected technologies up to real hot-¦ring conditions. Designed as prototypes, these demonstrators prepare well for their transfer into shorter and cost-e¨ective §ight engine developments.
The engine demonstrators anticipate also upcoming regulatory requirements such as the Clean Space policy, REACH (registration, evaluation, autorization, and restriction of chemicals), or material depletion guidelines. For instance, the FLPP is analyzing issues such as the switch from storable to nontoxic propellants or the guaranteed destruction of upper stages in passive reentries. For longer-term launcher evolutions and further steps toward competitiveness, the FLPP is closely involved in demonstration activities in the ¦elds of very low cost LOx/hydrocarbon propulsion and reusability concepts. These activities are focused on lower stage applications.
The propulsion projects in FLPP are coordinated to provide a large, robust, and consistent range of solutions to the European launchers for the 2020s and beyond. They build on a long and successful background experience and help European launchers remain a world-class reference.