Appraisal of initiated ESA propulsion developments for Exploration Missions

Transcription

1 8 Session 24 Exploration Appraisal of initiated ESA propulsion developments for Exploration Missions Hartwig Ellerbrock, Munich, Germany Thomas Diedrich Bremen, Germany Martin Riehle Lampolshausen, Germany Perigo, David, ESA/ESTEC Noordwjik, The Netherlands Philip Martin, Munich, Germany 7th ESA Space Propulsion conference, Bordeaux, France, Convention Centre, 7th to 10th May 2012

2 8 Appraisal of initiated ESA propulsion developments for Exploration Missions Copyright 2012 by the authors 3AF Association Aéronautique, Astronautic de France, 6 Rue Galilee, Paris, France1 Hartwig Ellerbrock, Propulsion & Equipment, Munich Germany hartwig.ellerbrock@astrium.eads.net Thomas Diedrich Orbital Systems; Bremen Germany Thomas.diedrich@astrium.eads.net Martin Riehle Propulsion & Equipment, Lampoldshausen, Germany martin.riehle@astrium.eads.net David Perigo, ESA/ESTEC Keplerlaan AG Noordwijk ZH, The Netherlands David.perigo@esa.int Philip Martin Propulsion & Equipment; Munich Germany philipe.martin@astrium.eads.net ABSTRACT This paper gives an overview of the propulsion challenges of the ambitious ESA's Exploration missions to Moon and Mars including the propulsion needs for ascent and decent. These needs will be compared with available propulsion engines /systems as well as with the current initiated technology developments An appraisal will be discussed to consider these technologies for ESA s exploration propulsion systems to Moon and Mars. 1. INTRODUCTION Since the 1960'ies, storable propellant engines and components have been developed and manufactured in Europe. An outstanding product portfolio is today available with in-space flight heritage. A good overview of the available Astrium propulsion portfolio is given in [1]. It shows the Astrium ST propulsion contribution as subsystem provider and as propulsion component provider to commercial and ESA missions programs. For the commercial telecommunication Satellites standard propulsion subsystems were designed and are in production. Most of the propulsion systems for 1 the unique ESA missions are taking benefit of these standard propulsion subsystems. This covers both the Bi-Propellant systems for GEO satellites (Geostationary) and LEO / MEO satellites (low- and medium earth orbit). It has to be noted that in Europe nearly all commercial propulsion subsystems and its components were co-funding development of Industries and ESA by the Telecommunication directorate. Therefore commercial programs giving benefit to ESA missions in two senses first to have a wide product portfolio available of standard propulsion systems and secondly by keeping the competence for these products in the long term run available

3 Telecom / Nav. / EO Spacebus Eurostar Leo/ Family Family Meo Rocsat 2 ALPHABUS Globalstar 1 GSTB V2 Galileo IOV Pleiades Myriade Proteus T errasar-x Cosmo Radarsat Skymed 2 GMES-Sentinel Propulsion systems SRE HME Others Other Nasa Galileo ATV Bi-Prop GEO s Hipparcos hybrid ARD GEOS Giotto Inv. Hybrid OTS ISO Tri-Prop ECS Mars Express IXV MeteoSat (MSG) Cluster MTG Rosetta Olympus Artemis Venus Express Propulsion system & components responsibility at astrium ST AMOS Ulysses 1,2, 3,4, Astrium ST Propulsion Eureca components used Metop XMM In selection / definition process SatComBW Aeolus Small Geo Herschel/Planck ARSAT 1,2 BepiColombo AMOS 6 ExoMars ExoMars Carrier Lander Solar Orbiter last ATV-3 Edoardo Amaldi has docked on the ISS safely on The de-orbiting is planned for the Table 1 show the destination, mission, the planned launch date as well as candidates for the main- and AOCS propulsion. Regarding main propulsion and AOCS propulsion, both are color underlined: with green: available engine / thruster; and with yellow: engine / thruster to be developed /adapted. With red: developments are needed Lunar Lander Fig 1 Astrium s propulsion participation to Space progr Fig 1 gives a good overview of Astrium ST propulsion involvement in the different programs, both commercial and ESA programs. Regarding the commercial spacecrafts between 6 up to 9 spropulsion subsystems will be integrated by Astirum in Lampoldshausen for different customers. Especially for exploration missions the current thruster portfolio available in Europe are not covering all specific mission needs. Further, engines in the thrust range between 1 to 12 kn are today not available in Europe. Based on on-going discussion with the agencies, the needs and requirements for storable propulsion have been identified, and will be discussed in the following. 2. MISSION PROFILES The ESA space exploration missions can be principally split into four groups: 1. ISS access including re-boost capability for ISS, man-rated mission requirements; 2. Missions to the Moon; 3. Mission to the Mars; 4. Human space flight exploration mission 1. Figure 2: Europe's HSF & exploration planned missions. The time table (Fig.2) indicates at wich time Europe is planning to launch the missions. Regarding the ATV-Missions a total of 5 missions are planned. The Destination Mission Year Main Propulsion AOCS (pulsed) LEO MPCV x 500 N EAM (steady) LEO MPCV with Crew capability 2025 ff Moon lunar lander 2018 Mars Mars Mars ExoMars Carrier (Orbiter) Exomars Demo- Lander MSR/ Orbiter/ Lander/ Ascender 24 x 220 N 200N to be adapted Bi-Prop Subsystem as above plus 24 x 220 N kn (steady) 5 x 500 N EAM (steady) 16 x 22 N 6 x 220 N (pulsed) x 400 N (LAE) 20 x 10 N Table 1: Storable propulsion needs. 3 cluster systems with 3 x 400 N CHT400 (pulsed) each (Σ= 9) 1500 N (steady), 12 KN (throttled), 5 kn (steady) 8 x 22 N CHT22 10N, 22 N Up to 2025, the investigations in Europe are focusing on the following: 1. Access to the ISS with the ATV follow on vehicle MPCV. The multi purpose concept shall be extended for Crew transportaion in a second development for a flight in 2025; 2. Robotic Moon missions with lunar lander. 3. Robotic mission to Mars (ExoMars) in 2016 with landing on the surface (without rover mission, because this will be the passenger for a Russian mission. 4. A Mars Sample Return (MSR) mission is in planning status with a sample return to Earth mission starting in The investigated mission profiles have to consider the launch vehicle (mainly Soyuz and Ariane 5, both to be launched from Kourou), the lander mass and the dedicated mission objective. The Mission profiles for the MPCV will be similar to the ATV mission profile exect the fact that the propulsion module will be separated from the MPCV prior to the re-entry to earth. The descent and landing of Lunar Lander is given in Figure 2, Ref. [2]. 1 Remark: The group 3 propulsion needs for human space flight exploration missions are expected to be beyond 2025 and are therefore not discussed within this paper. 2

4 Figure 3: lunar landing phases (adopted from Ref. [2]). The Lunar Lander mission scenario is based on a Soyuz launch from Kourou. After transfer to the Moon and some operation in the Low Lunar Orbit the landing will be performed with the coasting, braking-, and approach-phases prior touchdown of the vehicle on the Moon surface in the south pole region. Due to the absence of any atmosphere, the total landing v on the Moon surface is approximately 2030 m/s. Therefore and due to the higher ISP a bi-propellant propulsion subsystem was selected. The main propulsion braking maneuver requires a 3.5 kn thrust level, and an overall braking time of approx 12 minutes. The thrust level needs to be reduced during descent & landing. This will be realized by a sequential shut of the main thrusters. The additional 200/220 N engines in a pulse mode operational mode will be used to ensure a smooth transition of the sequence. In addition 22 N RCT thrusters are accommodated to be used for attitude control. Just during the contact of the first landing leg with the Moon surface the propulsion system will be completely shut off. The thruster pattern of the Lunar Lander is shown below in Figure 4. The engines needed for e.g. a mission to Mars are quite different, because most of the braking v could be achieved by atmospheric re-entry and the use of a parachute. Due to the low density of the Martian atmosphere, final deceleration has to be performed by rocket engines (NASA's Spirit & Opportunity missions used three solid rocket motors for the braking maneuver). For soft- and precise landing with a liquid propulsion system, the v touch-down requirement is about 40 m/s and therefore different to a Moon landing scenario. The ExoMars landing sequence on Mars is illustrated in Fig. 5 Figure 5: ExoMars descent planned mission. The propulsion subsystem responsibility for the ExoMars lander is under TAS-I responsibility and the design is based on 3 independent clusters each equipped with 3 x 400 N SCA-thrusters developed for the Ariane 5 EPS Version. Due to enhanced requirements feasibility hot firing tests have been successfully concluded for the use of the SCA thruster for the landing subsystem. The Mission Requierements covers both steady state firings and Pulse mode operations. 3. LAUNCH AND TRANSFER The Lunar Lander mission will be launched from Kourou with a Soyuz-Fregat 2.1b launcher not later than The launcher will insert the Lander into a HEO orbit, from which the Lander s propulsion system will be used for entering Lunar Transfer Orbit and Lunar Orbit Insertion. Figure 4: Lunar Lander Thruster Pattern Following transfer to the Moon the Lunar Module will be inserted into a low lunar near circular orbit (at about 100 km altitude) on which preparations for 3

5 landing will be performed. Final descent and landing with the subsequent soft landing will then be initiated aiming for a landing near the Lunar South Pole on the near side of the Moon. Fig. 6 Launch into HEO with Soyuz-Fregat from Kourou The ExoMars mission will be launched from Kourou with a Soyuz launcher no later than The concept is that on top of a carrier the ExoMars lander capsule will be mounted. The capsule will be separated when the spacecraft is on its hyperbolic flight to the Mars. After separation the carrier will be injected into the Mars orbit and is now a Mars- Orbiter. The engines / thruster currently under production are described in Ref [1]. Generally, the Lampoldshausen production plant is flexible in its production rate, therefore additional science and /or exploration mission are not limited by the current production capabilities. The current and envisaged engine / thruster storable development programs of Astrium ST are: The kn engine pre-development pintle injector technology ( ), details will be given in Chapter 5.3 The 500 N EAM development is an Astrium funded development with DLR co-funding; CDR planned for 11/2012 with a qualification in N RCT status is close to CDR awaiting for a customer and the corresponding qualification; principle tests performed based on Lunar Lander requierements. 220 N Adaptation of the 200 N ATV Thruster for Lunar lander (see Chapt. 5.2) In the frame of FLPP Storable engine demonstrator technology program, focus is given on a pressure-fed medium thrust engine (3-8 kn) up to a TRL of 5 to 6, Ref. [8]. In addition, Astrium concluded the study for DLR (system study for VEGA evolution). This study focuses on two concepts in view of a European VEGA launcher for performance improvement, a four-stage version (three solid + liquid kick-stage), and a three stage concept with a third liquid stage based on Aestus II engine. Further to these programs, the need for a low TRL pre-development technology programs has been deemed necessary, with the main focus on storable injection characterization and design tool evolution. Within such a program, a TRL range of 2-4 shall be reached with new technologies. 5. ASTRIUM ST CURRENT DEVELOPMENTS Figure 7: ExoMars carrier & EDS, Ref. [6]. 4. ASTRIUM ST (TP) BI-PROPELLANT THRUSTER / ENGINE PROGRAMS At Astrium ST (TP) the bi-propellant engine / thruster programs are split into two segments: 1. Production programs, and 2. Development programs Lunar Lander propulsion concept The propulsion concept is based on three different engine types which all use MON oxidizer and MMH fuel : - 5 x 500 N EAM engines for continuous thrust operation - 6 x 220 N pulse modulated assist engines, based on ATV heritage - 16 x 22 N ACS thrusters 4

6 Figure 8: Lunar Lander Propulsion Subsystem In addition of the detailed analysis of a failure tolerant design of the propulsion subsystem, the study will give particular focus to pressure drop and cross talk effect analysis of the engine cluster fuel sloshing analysis and mitigation techniques thermal analysis of the clustered engine concept. Within the Phase B1 the following propulsion Breadboard activities were performed. Objectives of the breadboarding activities are the reduction of risks for the development of the key technologies needed or the Lunar Lander mission. One of the mian subsystem is the propulsion subsystem. In the frame of development risk mitigation deticated tests were performed. The following components were in more detail investigated 220N thruster hot firing test The main objectives of these tests were: Check for low frequency instability robustness Performance at 1;2...3 Hz pulse mode frequencies Temperature characterization The main results of the performed test programs shows: smooth and stable thruster operation at reduced trimming domain no indications for hard starts or LF/HF instabilities changed parameters (reduced thruster pressure drop, reduced propellant inlet temperature) had no negative impact on thruster operation. During PMF at pulse frequencies > 1Hz, no signs for thermal instabilities Pulse frequencies of 2.0, 2.5 and 3.0 Hz with duty cycles from MIB (t_on_minimum =28ms) to DC=50% and n=2000 consecutive pulses are possible without violating material constraints. The "Lunar Lander landing and descent profile" performed very smooth and uncritical w.r.t thruster thermal behaviour. In consequence the 200N Thruster demonstrated excellent compliance with the needs. Hydraulic Testing Tests with the hydraulic system of the propulsion subsystem were performed to gather the behaviour of: Priming of the system Static pressure drop Water hammer testing Cross talk assessment Symmetry tests Figure 10 Hydraulic test set-up Figure 9 - hotfiring test of the 200 N Thruster 5

7 Status of Testing: The static and dynamic pressure drop measurements were performed and were in the expected range of Δp but also confirmed excellent symmetries in all thruster operational configurations. The thruster cross talk for the 2 3Hz pulse frequencies were finished. The flight representative sequences ("Lunar Lander landing and descent profile") and the priming tests were completed Failure cases and margin tests performed Validation tests for details of modeling in progress The test results were used for the calibration of the established ECOSIM model. The status of modeling and validation with test results is being presented in [9]. A further investigation was the testing of a pulse mode operation of the Astrium built 22N Thruster for the Lunar Lander application. This test campaign has successfully been performed with an all European configuration. First results are presented in [11]. version to welded titanium version was introduced. Fig. 12: Spacecraft design Figure 13 Connection of thruster to tubing system The complete installation of the propulsion subsystem is shown in Fig. 14 Figure 11: all European configuration of the 22 N thruster 5.2 ExoMars Astrium ST participation Entry Descent and landing System (EDS) As pointed out above the EDS is using 3 independent identical cluster subsystems. Astrium is responsible for the delivery of the 400 N monopropellant thruster designed and in production for the Ariane 5 EPS version. Therefore feasibility testing in accordance with the requierements was performed successfully. In addition changes of the connection from screwed Figure 14 EDS Propulsion subsystem Carrier The carrier concept was originally based on a dual mode-propulsion system using N2H4 as propellant for the AOCS Monopropellant thruster and for the injection maneuver into the Mars orbit an engine operation with N2H4 /NTO. However, such propulsion components are not available in Europe. 6

8 OHB, responsible for propulsion, thermal and structure of the carrier, traded the use of a standard bi-propellant system (e.g. standard Telecom propulsion subsystem). The selected carrier structure with a central tube as shown in Fig. 8 is very similar to the standard structure of Telecom buses. Therefore a standard UPS was selected with wide European heritage and the advantage of lower prices with the use of mainly European components. 5.3 Anthiope Throttleable engine investigation Within the ESA GSTP (General Support Technology Program) a co-funding contract was awarded to Astrium for principle injector design investigation of a throttleable injector.configuration. The investigation is and will be a parallel activity to the Astrium ST own funded development of the HOMER vehicle. The GSTP procet is split into 2 phases, Phase 1A with the manufacturing of the Hardware (injector and metallic chamber and the test program of this battleship hardware. Performance investigations will be performed in two fixed pintle injector design configuration (minimum and maximum). In Phase 1B the pintle injector geometry will be adjustable during the hot firing in order to determine the dynamic thrust vaiation behaviopr. For this the hardware will be redesigned and manufactured including an actuation device for dynamic chaning of the injector geometry. Main objective is the investigation of the functional driving parameters in terms of performances and pressure budget of the pintle-injector design. In Ref [1] the author presented advantages and disadvatages of the pintle injector design. The throttlecapability of a rocket engine with a pintle-injectro design down to a performance level of 30-50% of the nom. Thrust level is linked with a reduced combustion efficiency in the nom operation. This is resulting in a lower ISP than to an engine which is optimized to a nom operation loadpoint. In consequence of a lower ISP the mission has to consider a higher amount of propellant to be carried with. A second important objective is the operation in pulse mode operation. In the modulation mode further thottleling of the thrust could be achieved. However, such a thrust variation requeted from design point of view a minimized dribble volume, meaning the volume in the injector from the valve down to the injection element shall be designed as minimum as possible. This volume is also responsible for the thrust ramp-up. For an engine in the thust range of 3-8 kn a full thrust ramp up within 8 15 msec shall be possible. The max pulsemodulation frequency is limited by the design of valve. Fig. 13 shows a cold-flow test and figure 14 shows the cut-drawing of the engine design. Thrust level is 8KN Figure 15: Injector cold flow test of the GSTP technology demosntrator The Phase 1A is planned to be finalized by the end of 2012 and Phase 1B shall be continued in With the Phase 1A and 1B investigations the main design driving parameter will be known an a potential throttleable engine development program could be started with limited development risks. Figure 16 GSTP Technology demonstator 6. STORABLE PROPULSION NEEDS Launcher Propulsion For launcher propulsion, the need of a high performance engine in the thrust class of 3-8 kn is identified. This engine should be capable to be re-. 7

9 ignitable, with long duration single burn requirement of approx. 20 minutes. Such engine will be able to be used for kick-stages in order to come to more flexible launch vehicle missions as well as for the MPCV in the extend Crew version to separate the MPCV from Launch vehicle when a crew is on board. In addition, the Aestus II engine with high performance has a good potential to be applied for VEGA performance improvement and / or for Ariane 5 dedicated science or robotic missions. Figure 17: ESA's lander & ascent capsule, with different propulsion systems. Around 2025, ESA plans to bring samples back from Mars to Earth. This Mars-Sample-Return mission requires (beside the Mars orbiter and the descent vehicle) also an ascent vehicle to rendezvous with the orbiter and a flight mission back to Earth. For such a mission a throttable 8-12 kn engine is needed for the descent. For ascent the engine shall be operated in steady state with the maximum performance. As discussed before, these diverging requirements could be met by different engines (throttable for descent, high performance for ascent) as illustrated in the artist view of the ascent of the MSR mission, see Fig. 13. For Mars orbit insertion of the orbiter, ESA identified the need of a propulsion module with 1 up to 1.5 kn of thrust. 6. Appraisal of engines for Exploration For the todays planned mission Lunar Lander, Exomars 2016, MPCV and MSR the appraisal regarding the most time and funding consuming component were performed. 1. For the lunar lander mission only qualification of the EAM and the 200 N thrusterare needed. Adaptation effort, no new development. 2. ExoMars carrier (and later on the Orbiter) to use a standard bi-propellant propulsion system (in use for telecom missions) based on a 2 -tank version with the 400 N LAE. Adaptation effort, no new development. 3. For the MSR which calls for an orbit insertion thrust of kn Astrium proposes to use 3 clustered EAM engines. Investigation regarding thermal interferences of the engines to be investigated. Adaptation effort, no new development. However, ESA initiated a development of a kn engine in the MREP program. 4. For the MSR descent & ascent propulsion system it is proposed to use the engine under development of the FLPP programe: thrust level up to 12 kn as throttlabe and steady state could be possible Development of one engine instead of two new ones. For upper stage engines, the continuation the FLPP Storable pre-development program is needed: The technology program will be finalized in 2014 and shall be continued with a development / qualification programme for product availability beyond SUMMARY AND RECOMMENDATIONS The proposed ESA missions up to 2025 were analyzed regarding their engine and thruster needs. The main results and recommendations are: 1. EAM is corner stone product in the 500 N class, to be delta-qualified for commercial application, MPCV and the Lunar Lander. For Lunar Lander the effect of interferences with the 5 clustered Engines to be investigated. From analytical point of view this seems to be not critical. The engine is also suitable to be used in a clustered version to cover kn orbit insertion requirements. It is recommended to identify the requirements for an extension of the planned qualification. 2. The pulse mode operation of the 200 N thruster shows excellent results in accordance to the Lunar Lander mission requierements. Additional Pulse mode operation with the 22 N thruster will mitigate the technical risk for the Lunar Lander program. 3. For MPCV (crew- version) a kn engine could be based on Aestus heritage or could be fulfilled by clustering of 8 kn engiens; the development shall be started eight years in advance to the planned mission. 8

10 4. throttleable engine technology is key for further exploration missions. within the Astrium / ESA pre-development investigation of the Anthiope project major risk for such an engine development will be identified.. The development of a full scale engine shall be considered by ESA with a breadboard demonstration from 2014 onwards. 8. LIST OF ABBREVIATIONS ATV Automated Transfer Vehicle ARV Advanced Return Vehicle AOCS Attitude Orbit Control System CHT Chemical Hydrazine Thruster CDR Critical Design Review EAM European Apogee Motor - Astrium Brand name EDS Entry Decent System ε Ratio between nozzle exit / throat area FLPP Future Launcher Preparation Programme GTO Geostationary Transfer Orbit GEO Geostationary Earth Orbit HSF Human Space Flight ISS International Space Station Isp Specific impulse LAE Liquid Apogee Engine MMH Monomethyl Hydrazine N2H3(CH3) MSR Mars Sample Return (planned ESA Mission) MPCV Multi-Purpose-Cargo (Crew) vehicle NTO Nitrogen Tetroxide N2O4 RCT Reaction Control Thruster SCA System Control Attitude (Launcher Roll- Control System) SRE Science Robotic Exploration TRL Technical readiness level UPS Unified Propulsion System 9. REFERENCES [1] Ellerbrock, Henn, Hagemann: Propulsion Roadmap for Exploration ESA 3AF , 6th Prop. Conf San Sebastian. [2] Berengere, Houdou, "ESA Lunar Lander", ESA/DLR Lunar Lander Event, April 18th 2012 [3] EADS-Astrium space transportation web page, [4] The vision of space exploration; Jet Propulsion Laboratory main/index.html. [5] D. Huzel, and D. Huang, "Modern Engineering for Design of Liquid-Propellant Rocket Engines", 1992 ISBN [6] ESA web page [7] M. Peukert, S. Kraus, M. Riehle, G. Hagemann, and T. Diedrich, "Propulsion Needs for Exploration Lander Application", ESA 3AF , 6th Prop. Conf [8] G. Obermaier, A. Goetz, G. Hagemann, P. Philipp, O. de Bonn, C. Maeding, and H. Burkhardt, "Storable Bi-Propellant Technology Enhancements for Low Thrust Engines - Roadmap and Achievements", ESA-3AF , 6th Prop. Conf [9] Riehle, M.: Propulsion System for the European Lunar Lander - Development Status and Breadboarding Activities"; SP , 7th Space Propulsion Conference 2012, 7-10 May 2012, [10] S.Kraus, M.Riehle,M.Bühner, G.Schulte; "Current state, future applications and further development of the ATV 200N bipropellant thruster"; SP , 7th ESA Space Propulsion conference, Bordeaux, France, Convention Centre, 7th to 10th May [11] S.Kraus, M.Wolf, U.Gotzig, E.Dargies; " Verification Test Results of ASTRIUM's 22N all-european Bipropellant Thruster for Lunar Lander Application"; SP , 7th ESA Space Propulsion conference, Bordeaux, France, Convention Centre, 7th to 10th May

11 10 ESA 3AF