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Propulsion options for preliminary formation flying missions studies at CNES

N Pillet*, P Bousquet ♣, E Chesta ♠, R Cledassou γ, M Delpech§, E Hinglais$, CNES, 31400 Toulouse, France

In 2005, 4 formation flying missions have been internally studied in CNES Toulouse, France, in the frame of the PASO (pre-project office structure). The satellite formation comprises either 2 spacecrafts (ASPICS, SIMBOL’X and MAX missions) or 3 spacecrafts (PEGASE) dedicated to astro-physics. We will present here, the main trade off and options concerning the satellite on board propulsion systems. For these missions, the propulsion system has to insure appropriate orbit transfer and control but also formation control (i.e. mutual positioning and attitude of the spacecrafts) which may require fine thrust and dedicated micro propulsion system. Different propulsion technologies (i.e. chemical and electrical propulsion system) such as hydrazine monopropellant, bipropellant, cold gas, FEEP, etc., are proposed and discussed.

Nomenclature

AMPS = ASPICS, MAX, PEGASE and SIMBOL’X CIC = Concurrent Engineering Center FEEP = Field Emission Electric Propulsion HEO = High Elliptical Orbit PASO = Plate-forme d’Architecture des Systèmes Orbitaux PPT = Pulse Plasma Thruster SI = Specific Impulse (s)

I. Introduction Ince beginning of 2005, formation flying missions have been studied at CNES Toulouse center. For each of these missions, a preliminary design of the relevant spacecraft has been proposed. The design of the propulsion

system is presented in this paper and some options are also discussed, depending on the specific requirement of the formation relative position and attitude control. It has to be noted that for the propulsion system, the choice of available technology has always been considered as an important criteria, for these new family of mission. This approach has the benefit to underline the requirements which do require new development.

II. Propulsion and Concurrent Engineering in CNES Toulouse

A. Concurrent engineering facility In the frame of PASO (pre-project office structure) which has been running all phase 0 studies of orbital systems at CNES since 2002, CNES decided in 2004 to create an infrastructure based on concurrent engineering methodology dedicated to preliminary studies of spacecrafts. This entity is called CIC, for “Centre d’ Ingénierie Concourante”. For the build-up of system budgets and the exchange of parameters between disciplines, we make good use of the IDM (« Integrated Design Model ») tool provided by the «Concurrent Design Facility» of ESA and resulting from close cooperation between both agencies. As described in the figure 1, the concurrent engineering work takes place in an open area, where each subsystem of the satellite is represented and can use its computer to exchange data’s with the others. There is also some video

* Propulsion engineer, Propulsion and pyrotechnics service, nicolas.pillet@cnes.fr ♣ Head of mechanical and thermal architecture service, Pierre.bousquet@cnes.fr γ System engineer, science service, rodolphe.cledassou@cnes.fr § GNC engineer, GNC service, michel.delpeche@cnes.fr $ System engineer, pre-project office, emmanuel.hinglais@cnes.fr

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screens, which can either be used for internal use (video projection of a dedicated computer screen) and also for video conference purpose.

Figure 1: Organisation of the concurrent engineering at CNES Toulouse The core of the engineering software used in the frame of concurrent engineering studies is based on an Excel tool, using a common data base, which is regularly updated by the different engineers taking part to the study (ref.1) In addition to this tool, dedicated tools can be used such as CATIA, orbit visualisation tool (including satellite positioning, sun shadow etc ..) like OPAL, etc. Basically, the output of a CIC study is a preliminary design and layout of the spacecraft and the relevant budget such as mass, power budgets, performances and ROM Cost of the mission.

B. Propulsion approach and relevant tools The exchange between propulsion and mission analysis, AOCS, Power system and mechanical architecture is illustrated on the figure 2 and is done via the Excel tool.

Figure 2: Exchanges between propulsion and other sub-system In order to size the appropriate propulsion system, an iterative process between the system and propulsion is needed, mainly to assess the preliminary propellant budget, based on mission delta V and propulsion performance.

For these studies, propulsion needs a quick access to the characteristics of components available worldwide. This would greatly benefit engineers taking part in the CIC session, as basic and necessary information in real time would be readily available. A simple database specification was therefore issued and is now operational. The basic need for

Propellant budget

Propulsion component size / mass

Thruster configuration and ACS need

Mission analysis

AOCS

Propulsion

Mechanical architecture

MIB

Delta V budget

S/C power supply

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each propulsion component is its size, mass, power and performance (thruster). In addition, one needs to know the heritage of a dedicated component.

Database structure:

It was decided to use as far as possible the information available on suppliers’ web sites (product datasheets), if any. If complementary information is necessary, electronic data is asked for via direct contact with the identified supplier and the relevant information is then stored on a dedicated disk.

Three software solutions were compared (Access®, HTML or Excel®). Excel® was finally selected for its simplicity and capacity to use filters for components. The programming is done is visual basic. As shown in the following table, it is possible within each column to select the relevant item in order to refine the research phase. Up to now, 61 propulsion suppliers have been selected in the database, and the heritage of 47 spacecraft is taken into account.

Table 1: Propulsion data base structure

III. Propulsion options for the formation flying missions

A. Mission analysis

For the four studied missions (ASPICS, MAX, PEGASE and SIMBOL-X), a quiet orbit (i.e. the gravity gradient must be as low as possible) is required and an altitude of 40 000 km to fly a formation was considered as a minimum. Therefore, the working orbit of the satellites are either HEO (ASPICS, SIMBOL-X and MAX) or L2 (PEGASE) orbits. The propulsion requirement was divided into:

- orbit transfer from launcher injection to the final orbit - orbit keeping - Formation flying control including anti-collision FDIR. - target change (if any)

The orbit transfer of the satellites could be done in 2 different configurations: attached or separated. In case the satellites are mechanically linked together, the orbit transfer is done by a unique propulsion system attached to one of the satellite. For system reasons (not presented in this paper), this option has not been considered and only the option separate transfer of the satellites has been studied. Therefore, each satellite includes its own propulsion system.

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Satellites

Wet mass allocation

(kg)

Working orbit / launcher

option

Scientific purpose

ASPICS 1 satellite coronagraph 1 satellite occultor

250 kg 100 kg

HEO (500 km X 70 000 km) SOYUZ launch with Pleiade

Solar Corona

MAX 1 satellite Lens 1 satellite detector

900 kg 600 kg

HEO (40 000 X 253 000 km) Dedicated Soyuz Launch

γ ray spectroscopy

PEGASE 2 satellite siderostat 1 satellite combiner

350 kg 700 kg

L2 Dedicated Soyuz Launch

Infra-red interferometry

SIMBOL-X 1 satellite mirror 1 satellite detector

900 kg 400 kg

HEO (40 000 X 253 000 km) Dedicated Soyuz Launch

X-ray telescope

Table 2: general characteristics of the missions

B. ASPICS propulsion options

Figure 3: ASPICS mission configuration The DV and MIB (when necessary) required to both satellites are expressed on the following table:

Coronagraph Spacecraft Occultor spacecraft

Orbit keeping DV = 2 m/s Idem

Formation flying keeping / rupture DV =6 m/s, MIB < 1 mN.s NA

Anti collision FDIR DV = 0.4 m/s, MIB < 1 mN.s NA

Deorbitation DV = 25 m/s Idem

Table 3: DV and MIB requirement for the ASPICS mission

Formation control The formation is kept by the Coronagraph spacecraft.

The 1 year ASPICS mission is dedicated to the study of the solar corona by mean of a telescope mounted on the coronograph spacecraft and working in the visible spectrum. The distance between both satellites is around 100 m, with a requirement of +/- 7 mm in lateral positioning and +/- 100 mm in axial positioning. The formation is kept only on the part of the orbit where the altitude is greater than 20 000 km.

X

Z

Yθ

X

Z

Yθ

Figure 4: Thruster configuration

As illustrated on the figure 4, the thruster configuration is based on 8 thrusters, such as the main DV requirement for the formation flying control (for this mission, this corresponds to the X and Z axes) can be realized with an acceptable geometrical efficiency

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Different thruster’s technologies were considered at the beginning of the studies such as small PPT, hydrazine and cold gas thrusters.

Hydrazine propulsion was discarded due to the low MIB requirement which is not feasible with the appropriate accuracy: 10 mN.s is today the limit of the available European technology, so a factor of 10 was missing. PPT's were considered as good candidates, due to their simplicity and capability of low MIB , such as the APPT 30 currently developed by RIAME (ref. 2). However, this option was discarded for maturity reasons. A cold gas propulsion system is appropriate to control the formation (mutual positioning and attitude of both spacecrafts) and the available 10 mN cold gas thruster from Marotta UK is proposed. For redundancy reasons, it was decided to have one for one redundancy on the cold gas thruster, as for example for the ESA Cryosat satellite (Ref 3)

Thruster MIB and cycling requirement: From GNC studies, the spacecrafts mutual drift (due to perturbing forces / torques) is controlled by a bang bang cycle (on 4 thrusters typically). From the MIB capabability of the thrusters, one can derive the number of thruster valve cycling for a dedicated mission duration. For ASPICS mission, the preliminary estimations for the formation keeping, requires a bang-bang cycle around 200 s on the thrusters. This means approximately 0.3 E06 cycles on the valves for a 2 years mission; which is quite achievable for the available cold gas technology, capable of 1 E06 cycles with 1.5 margin coefficient (Ref 5). Transfer phases, orbit control and deorbiting Coronograph satellite It could be attractive to realize the orbit control and deorbiting maneuvers also with the cold gas system: this will offer the benefit of a unified propulsion system for this spacecraft. However, this will lead, due to the poor Isp of a substantial mass penalty due to both propellant penalty (+ 10 kg) and dry mass penalty (GN2 high pressure tank: approximately + 10 kg). For these reason, a small hydrazine system, based on the Myriade hydrazine propulsion system (ref 4), was proposed in baseline for the coronograph satellite, in addition to the cold gas system. The capacity of the 6 liter Rafael tank is enough for the required 4 kg of hydrazine. However, two 6 liters tanks were incorporated in the design and in the mass budget, for growth potential purpose.

Occultor satellite:

The propulsion system of this satellite is mainly used for deorbiting purpose and directly based on the recurring Myriade hydrazine propulsion system, which have a dry mass of approximately 5 kg and a 4.5 kg of hydrazine.

Figure 5: Myriade Propulsion system, with 1 N from EADS-ST Figure 6: Aspics cold gas propulsion system with 10 mN thrusters from MAROTTA UK.

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2 X 8 X 10 mN cold gas thrusters

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Figure 7: Preliminary propulsion lay out on both satellites

C. MAX and SMBOL-X propulsion options

Figure 8: MAX (left side) and SIMBOL-X (right side) configurations The Max mission is dedicated to the detection of the gamma rays. The gamma rays are focused on the detector plan by the lens which is placed approximately at 90 m from the detector satellite. The main specifications of the formation control are +/- 1 cm in lateral positioning and +/- 10 cm in longitudinal positioning. The SIMBOL’X mission goal is to study the X rays, which are focused by the mirror to the detector satellite. The distance between both satellites is approximately 30m. The requirements in mutual positioning, for the formation control are similar to the MAX mission one. The DV and MIB (when necessary) required to both satellites are expressed on the following table: MAX SIMBOL’X

Detector Lens Detector Mirror

Transfer phase DV = 456 m/s Idem DV =465 m/s Idem

Orbit keeping DV = 10 m/s idem DV = 10 m/s idem

Formation flying DV = 6 m/s / MIB< 1 mN.s NA DV = 2,4 m/s / MIB < 1 mN.s NA

Target change DV = 10 m/s / MIB < 0.2 N.s NA DV = 37 m/s / MIB < 0.2 N.s NA

FDIR anti collision DV = 5 m/s / MIB < 1 mN.s DV = 5m/s DV = 1,0 m/s / MIB < 1 mN.s DV = 3.3 m/s

Table 4: DV and MIB requirement for the ASPICS mission

Detector SC

Lens SC

86 m

Detector SC

Lens SC

86 m

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Formation control For both missions, the formation control is done by the satellite detector which is the lighter one. The formation control and FDIR anti-collision do require a low MIB (1 mN.s) and, for the same reason as for the ASPICS mission, the choice of a cold gas system based on available impulsional thrusters (such as 10 mN ) seems to be the best choice and is therefore the proposed baseline. Transfer phase / orbit keeping and target change. Both bipropellant and monopropellant system were considered for the transfer of the satellites. The mass gain provided by a bi-propellant system (including a 320 s transfer engine) was evaluated and compared to the hydrazine solution. For the heaviest satellites (900 kg) the difference between both solutions was evaluated to be around 40 kg (dry mass + propellant) and this was not judged high enough compare to the extra complexity (2 separate circuits / necessity of a pressurisation system to keep the thrusters in their operational domain…) and cost of a bi-propulsion system w.r.t hydrazine system. Therefore, the hydrazine monopropellant option (Isp = 220 s) was always chosen for these transfer phases.

For the detector satellite of the SIMBOL-X mission, an interesting option for compacity is to use common GN2 high pressure storage for cold gas system and hydrazine tank pressurisation, as described in the following schematic. This allow to use the available Proteus 38 litters membrane tanks from Rafael, with a complementary qualification. For the mirror satellite of SYMBOL-X, the blow-down hydrazine architecture is based on 4 Pleaides 104 litters propellant tanks (EADS-ST).

Figure 9: Propulsion lay-out on detector and lens satellites (MAX)

Figure 10: Proposed schematic for detector (left) and mirror (right) satellites of the SIMBOL-X mission

Transfer of the 900 kg Lens from MAX DV = 476 m/s

Ergol Mass (kg)

Dry Mass (kg)

Total (kg)

MON/MMH 127 30 157 Hydrazine 178 20 198 Table 5: Comparison bi propellant / mono propellant system for the transfer from Soyuz injection to HEO.

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D. Pegase propulsion options

The DV and MIB required to both satellites are expressed on the flowing table:

Combinor Siderostats

Transfer phase DV =165 m/s Idem

Orbit keeping DV = 3m/s (MIB < 0.2 N.s) Idem

Formation control DV = 0.5 m/s (MIB < 50µN.s) DV = 0.3 m/s (MIB < 5µN.s)

Target change NA DV =9.5 m/s (MIB < 500 mN.s) FDIR anti collision DV = 1,0 m/s (MIB < 500 µN.s) Idem

Formation control The Pegase mission supposes a very fine control in the attitude of both siderostat and combinor satellites (required accuracy on the attitude angle is +/- 1 arcsec) and also in mutual positioning of the satellites.

For the attitude control, the GNC studies concluded that available 10 mN cold gas thuster was not acceptable due to a too short control cycle (to keep the satellite in the attitude requirement) and consequently too high propellant consumption and thruster valve cycles.

Two options are proposed for the attitude control of the satellites: - Use of an impulsional cold gas thruster capable of a 50 µN.s MIB (typically 2.5 mN thrust with 20 ms

pulse length). In that case, as for ASPICS, the formation is free between to thruster actuations. This option is submitted to the possibility of using reaction wheels for the attitude control. A first study showed that micro-vibrations induced by the wheels seem acceptable. In case this would not be confirmed, this option will be discarded for the following one.

- Use proportional thruster, either cold gas proportional or FEEP thruster, with a thrust range typically

between 1µN and 100µN. In that case, the mutual positioning of the satellite is done every time keeping the formation in a closer window that in the previous “bang-bang” strategy.

The PEGASE mission is based on Infra Red interferometry in view of collecting data on planets and stars formations. The distance between each siderostat and the central combiner can vary between 25 m and 250 m. Both mutual positioning and attitude requirements are used to size the propulsion system.

Figure 11: PEGASE configuration

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For this last option, the following technologies, which are currently developed on ESA contract, are considered:

Figure 12: Cesium FEEP developed by ALTA (It) Figure 13: Proportional cold gas thruster developed by Marotta UK Proportional cold gas thruster developed either by Laben or Bradford engineering could also be considered for this ambitious mission. Transfer phase and orbit keeping: For the siderostat satellite, the EADS-ST PROTEUS hydrazine propulsion system, based on a 38 litters Rafael tank and 4X1 N EADS-ST thruster, is proposed. For the Combiner, the EADS-ST PLEIADE propulsion system based on a 104 litter EADS-ST capillary tank and 4 X 1 N thrusters is proposed.

IV. Conclusion Some propulsion options for formation flying missions have been discussed and studied in this paper. This is quite interesting to note that for 3 of these missions (i.e. ASPICS, MAX and SIMBOL-X) existing cold gas technology is appropriate for the formation control. Only the PEGASE interferometry mission does require some technology development (either cold gas or FEEP) to achieve the required accuracy of the formation control. For the transfer (from launcher injection to the final orbit) and orbit keeping phases, a simple hydrazine system is considered as the most appropriate for this class of mission. Different hydrazine system, based on available plate-forms, such as Myriade or Proteus, have therefore been proposed.

V. Acknowledgements

The author wishes to acknowledge - MM François Gonzales and Philippe Gillen for their support concerning the concurrent engineering facility

used in CNES Toulouse. - Mr Jean Philippe Sire from the mechanical and thermal service, for its support concerning the CATIA

views of the different configurations.

VI. References Proceedings

1Bousquet, P.. “Concurrent Engineering at CNES”, IAC-05-D.1.03.06, IAF 2005. 2Pillet, N.. “Pulse Plasma Thruster option for myriade deorbiting”, EUCASS, Moscow, July 2005. 3Usbeck , T.. “ A flexible cold gas propulsion system concept for different applications”, 4th European propulsion conference, Sardinia, june 2004.

4Salomé, R.. “The Myriade microsatellite hydrazine propulsion system”, AIAA 02-4153, Joint Propulsion confernce juy 2002.

Thrust range: 0 – 150 µN Thrust resolution: 0.1 µN Response time: < 100 ms Low thrust noise (< 0.1µN / Sqr Hz) SI= 6000 s

Thrust range: 0 – 1000 µN (1 bar inlet). Thrust resolution: < 1µN Response time < 20 ms Actuator voltage: -20/+150 V SI = 70 s

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5Solway, N. “Design and performance of a micro Newton proportional thruster”, 4th European propulsion conference, Sardinia, June 2004.

6Smith, P. “Cryosat cold gas system and component development”, AIAA, Joint Propulsion Conference, July 2004.