SMART-1 mission description and development status

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Planetary and Space Science 50 (2002) 1323 – 1337 www.elsevier.com/locate/pss

SMART-1 mission description and development status G.D. Raccaa;∗ , A. Marinia , L. Stagnaroa , J. van Doorena , L. di Napolia , B.H. Foingb , R. Lumba , J. Volpa , J. Brinkmanna , R. Gr1unagelc , D. Estublierc , E. Tremolizzoc , M. McKayd , O. Caminod , J. Schoemaekersd , M. Hechlerd , M. Khand , P. Rathsmane , G. Anderssone , K. An8oe , S. Bergee , P. Bodine , A. Edforse , A. Hussaine , J. Kugelberge , N. Larssone , B. Ljunge , L. Meijere , A. M1ortselle , T. Nordeb1acke , S. Perssone , F. Sj1oberge a ESA

Scienti c Projects Department, ESA/ESTEC, Keplerlaan 1, Noordwijk, 2200 AG, The Netherlands b ESA Space Science Department, ESA/ESTEC, Noordwijk, The Netherlands c ESA Directorate of Technical and Operational Support, ESA/ESTEC, Noordwijk, The Netherlands d ESA Directorate of Technical and Operational Support, ESA/ESOC, Darmstadt, Germany e Swedish Space Corporation, Solna, Sweden

Abstract SMART-1 is the ;rst of the Small Missions for Advanced Research in Technology of the ESA Horizons 2000 scienti;c programme. The SMART-1 mission is dedicated to testing of new technologies for future cornerstone missions, using Solar-Electric Primary Propulsion (SEPP) in Deep Space. The chosen mission planetary target is the Moon. The target orbit will be polar with the pericentre close to the South-Pole. The pericentre altitude lies between 300 and 2000 km, while the apocentre will extend to about 10; 000 km. During the cruise phase, before reaching the Moon, the spacecraft thrusting pro;le allows extended periods for cruise science. The SMART-1 spacecraft will be launched in the spring of 2003 as an auxiliary passenger on an Ariane 5 and placed into a Geostationary Transfer Orbit (GTO). The expected launch mass is about 370 kg, including 19 kg of payload. The selected type of SEPP is a Hall-e>ect thruster called PPS-1350. The thruster is used to spiral out of the GTO and for all orbit maneuvers including lunar capture and descent. The trajectory has been optimised by inserting coast arcs and the presence of the Moon’s gravitational ;eld is exploited in multiple weak gravity assists. The Development Phase started in October 1999 and is expected to be concluded by a Flight Acceptance Review in January 2003. The short development time for this high technology spacecraft requires a concerted e>ort by industry, science institutes and ESA centres. This paper describes the mission and the project development status both from a technical and programmatic standpoint. ? 2002 Elsevier Science Ltd. All rights reserved.

1. Mission overview SMART-1 is the ;rst Small Mission for Advanced Research in Technology of the ESA Scienti;c Programme Horizons 2000. The mission is dedicated to the testing of new technologies for preparing future cornerstone missions, using Solar-Electric Propulsion in Deep Space. The ;rst study of this mission has been performed in 1997 Racca (1999), for several con;gurations Racca et al. (1998a), electric propulsion options Racca et al. (1998b), and planetary targets. The mission was ;nally approved by the Science ∗ Corresponding author. Tel.: +31-71-565-4618; +31-71-565-5770. E-mail address: [email protected] (G.D. Racca).

fax:

Programme Committee of the European Space Agency in November 1999, on the basis of the baseline described in this paper. The mission planetary target is the Moon. The cruise phase to the Moon contains periods for cruise science and the science operations phase around the Moon will last 6 months. The spacecraft will be launched in early 2003 as an Ariane 5 auxiliary payload. The SMART-1 spacecraft will be placed in orbit around the Moon, using Solar-Electric Propulsion, after a cruise phase which may last up to 18 months. The overall spacecraft mass shall remain within about 370 kg at launch. This is mainly due to the need to provide the spacecraft with an initial acceleration of about 2 × 10−4 m s−2 . This acceleration is necessary in order to maintain a reasonable time of

0032-0633/02/$ - see front matter ? 2002 Elsevier Science Ltd. All rights reserved. PII: S 0 0 3 2 - 0 6 3 3 ( 0 2 ) 0 0 1 2 3 - X

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8ight and to ensure proper lunar injection capability. In addition an increase of mass is re8ected into a greater launch cost. The payload is composed of technology and scienti;c experiments and its total mass has been maintained just below 19 kg. The seven SMART-1 instruments support nine investigations. The seven instruments are: (1) Electric Propulsion Diagnostic Package (EPDP) (2:4 kg; 18 W): A suite of sensor for thruster diagnostics with ion energy up to 400 eV and S/C contamination monitoring. (2) Spacecraft Potential, Electron and Dust Experiment (SPEDE) (0:8 kg 1:8 W): Langmuir probes on short booms for energy range of a few tens of eV, with plasma density from 1/10 to 1000 particles cm−3 . (3) X/Ka-band Telemetry and Telecommand (TT& C) Experiment (KaTE) (6:2 kg; 26 W): A X-up/X-down and Ka-down Deep-Space Transponder running turbo-codes, allowing up to 500 Kb s−1 data rate from lunar orbit. (4) Demonstration of a Compact Imaging X-ray Spectrometer (D-CIXS) (5:2 kg; 18 W): A 8◦ × 24◦ FOV spectral imager in 0.5 –10 keV range based on Swept Charge Device detectors and micro-collimators. (5) X-ray Solar Monitor (XSM), ancillary to D-CIXS (weight and power included in D-CIXS), allows absolute calibration of its 8uorescence spectra and performs long-term monitoring of solar corona and 8ares with energies in the 0.5 –20 keV range. (6) Advanced Moon Micro-Imager (AMIE) Experiment (2:1 kg; 9 W): A 5:3◦ FOV miniaturised camera with a 4-band ;xed ;lter. The camera is based on high-density 3D multi-chip module electronics. (7) SMART-1 Infrared Spectrometer (SIR) (2:3 kg; 4:1 W): A 1 mrad FOV point spectrometer with 256 channels operating in the 0.9 –2:4 m wavelength range (NIR) for lunar mineralogy.

The three supported guest investigations are: • Laser-link demonstration of a Deep Space optical link acquisition (with AMIE) • On-Board Autonomous Navigation (OBAN) concept veri;cation (with AMIE) • Radio-Science investigations for SMART-1 (RSIS) Electric Propulsion monitoring and demonstration of in-orbit libration measurement method (with KATE and AMIE). Foing et al. (2001) and Racca et al. (2001) describe the science potential and goals of the scienti;c instruments, while Racca and Marini (2000) and Marini et al. (2000) show how the technology experiments prepare for future ESA cornerstone missions. The individual instruments are further described in detailed in this issue of Planetary and Space Science in dedicated papers by the Principal Investigators. Furthermore, a recent paper by Marini et al. (2002) describes the as-built status of the payload instruments and their 8ight operations. 2. Spacecraft bus design To reduce development time and increase reliability, strong emphasis has been placed on design modularity, commonality and the use of standard hardware and software components in the design of the spacecraft. In the same vein, commercial-o>-the-shelf (COTS) hardware and software components are used. The SMART-1 spacecraft platform combines heritage from previous small satellite programs by the Swedish Space Corporation (SSC), the Prime Contractor for the spacecraft platform, and up-to-date technology developments from ESA, commercial space programs and other branches of industry.

Fig. 1. Artist impression of SMART-1 in orbit around the Moon.

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Fig. 2. Close-up view of the SMART-1 Spacecraft -Z panel. The Electric Propulsion thruster PPS-1350 is visible mounted on a Thruster Orientation Mechanism shown with hold down points in their released position. The ring represents the launcher interface. Four of the eight hydrazine thrusters are also visible, canted to provide pure torques in all directions. The solar array extension arm mounted on its drive mechanism is shown on the +Y face.

2.1. System con guration The dimensions and shape of the spacecraft are driven by the mass and geometric envelope constraints of the launcher (I2:4 × 1:2 m cylindrical envelope), the requirement to accommodate the existing electric propulsion system and the power generation requirement. The spacecraft design is roughly cubical with a pair of 3 panel solar arrays attached to the += − Y panels, see Fig. 1. The 49 l Xenon tank, capable of carrying 82 kg of supercritical Xenon, is accommodated in the centre of the structure on top of the electric propulsion thruster. The thruster is mounted on its 2-axis orientation mechanism and acts in the −Z direction. The orientation mechanism is used to compensate for static or slowly varying thrust misalignments, shift of the centre of gravity due to fuel depletion and solar pressure and gravity gradient disturbance. The hydrazine thrusters are located in the four corners of the −Z panel and provide torque for reaction wheel momentum unloading (Fig. 2). The solar arrays, equipped with GaInP/GaAs/Ge solar cells, deliver 1850 W of power. The two wings have an active surface of about 10 m2 , span 14 m between the tips and can be individually rotated to maintain maximum solar input. The solar array power is accurately regulated in the power control and distribution unit to 50 V, and distributed to users via solid-state power controllers (SSPCs). Five high capacity Li-ion battery cells, each providing 130 W h, and associated discharge and charge electronics are provided to

support power transients and eclipse phases. The number of cells and the cell capacity is driven by the requirement to survive a 2:1 h eclipse in a one cell failure condition. The power budget is shown in Table 1. The attitude and orbit control uses primarily as attitude sensor an autonomous star tracker, giving 3-axis quaternion information once a second, and 4 skewed reaction wheels as the main actuator. The necessary momentum in all directions can be provided with only three wheels, hence the system is single failure tolerant. The detumbling and safe attitude modes are supported by omni-directional sun sensors, angular rate sensors and two sets of four 1 N hydrazine thrusters. The += − Y panels, which are maintained at a very shallow angle with the sun direction, are used as radiators and equipped with heat pipes for the high dissipating equipment (i.e. the spacecraft power control and distribution unit and the electric propulsion power processing unit). The radiators are painted white for optimum absorptivity/emissivity ratio. Multi-layer insulation blankets are used on the other external surfaces and high emissivity optical properties are used for the internal structure and units. Critical equipment, such as hydrazine lines and valves, are ;tted with thermistor controlled heaters. Two low gain S-band antennas placed on the += −X panels provide omni-directional ground coverage for telecommand, 2 Kbs−1 , and low rate telemetry, 2 Kb s−1 . Through the medium gain S-band antenna located on the +X panel a telemetry rate of 65 Kb s−1 is provided. The +X panel

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Table 1 Power and energy budget

Subsystem power consumption (W)

EPa mode

EP mode eclipse

Science mode eclipse

26.6 57.1 0.1 15.4 24.2 0.8 49.1 37.0 0.0 210.3

26.6 52.0 0.1 15.4 74.5 20.8 0.0 35.6 0.0 225.0

BDR eOciency (%) Battery current sharing (%) Max eclipse duration (h) Energy storage requirement (W h)

82 1.4 2.1 546

82 1.4 1.6 449

Eclipse entry cell temp (◦ C) Energy storage (W h)—no failures Energy Storage (W h)—one failure Nominal DODd (%) Post-failure DOD (%)

35:0 − 20:0 679:0 − 631:0 543:0 − 505:0 80:4 − 86:0 100:5 − 108:0

35:0 − 20:0 679:0 − 631:0 543:0 − 505:0 66:0 − 71:1 82:7 − 89:0

BOLb (63:5◦ C)

EOLc (100◦ C)

Data handling system Altitude, orbit and control system Release mechanisms TT& C Thermal control system Payload Electric Propulsion Power conditioning and harness Battery charging Total power consumption (W)

26.6 65.0 0.1 21.7 89.2 20.0 1400.0 113.3 29.5 1765.4

26.6 65.0 0.1 21.7 89.2 20.0 1140.0 101.0 29.5 1493.1

Margins Minimum array power capability

82.6 1848.0

20.0 1513.0

a Electric

Propulsion. of life. c End of life. d Depth of discharge. b Beginning

2.2. On-board autonomy and control hierarchy

Table 2 Mass budget

Item AOCS System Unit Mechanisms Power Solar array Hydrazine system TT& C Structure Thermal control N2 H 4

Mass (kg) 19.3 15.0 19.9 45.6 50.7 12.5 7.9 49.1 12.2 4.6

Total-platform EP S/S (PPS-1350) Propellant (Xe) Payload

236.8 29.2 82.0 18.9

Total-mass

366.9

also accommodates the antenna of the X/Ka transponder. During the Earth spiraling phase this panel is roughly facing the Earth. The SMART-1 mass budget is shown in Table 2.

The spacecraft is designed to require the minimum of ground intervention during its operative life. An example of keeping operations cost to a minimum is to use ground stations on an availability basis. During the science phase around the Moon the spacecraft will on average be contacted for 8 h twice per week. During its transfer phase to the Moon, SMART-1 is required to maintain nominal electric propulsion operations with autonomy periods of up to 10 days. A typical timeline of a ground station pass is presented in Fig. 3. Attitude pro;les and electric propulsion timelines with a planning horizon of 10 days are up-linked and stored on-board. The spacecraft is further designed to autonomously restore the nominal electric propulsion operations after any identi;ed single failure occurrence on-board except in case of failure of the spacecraft controller, in which case the switch over to the redundant unit would cause the spacecraft to transit to safe mode. Failures of functions classi;ed as minor, which include the individual science instruments and technology experiments, are not autonomously recovered from, but their failure must not propagate to other functions. Ground is used

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• The ability to re-power the spacecraft controller and initiate software boot-up. This is achieved by implementing a non-volatile con;guration memory in the spacecraft supervisor.

S/C Tracking Activities Telecommanding Opportunity S/C Status Check

Real Time HK HK Playback (If Needed) Science Data Playback AOS

LOS

+8 hrs

0 hrs

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Fig. 3. Typical timeline of a ground station pass.

to recon;gure and restore (partial) availability of these functions. All ground communication with the spacecraft is according to the CCSDS standard and implements the Packet Utilisation Standard for telecommands and telemetry. The mission timeline is executed on-board from the time-tagged telecommand bu>er. All housekeeping telemetry and science instrument data are stored in the on-board mass memory waiting for the ground control to contact the spacecraft. The failure detection, identi;cation and recovery (FDIR) function is implemented on a strictly hierarchical basis, aiming at restoring a function with as local means and as little operational impact as practically possible. Failures related to individual units are handled by local redundancy recon;guration. Failures observable at subsystem and system level are either handled by means of multi-unit recon;gurations or operational mode changes respecting the control hierarchy presented below. Four major authority levels illustrate the adopted control hierarchy Elfving et al. (2000). Level 1 is the highest level of authority and concerns the instantaneous ability to ensure spacecraft safety. Level 2 concerns the ability to restore and maintain durable spacecraft safety. Level 3 concerns the ability to maintain electric propulsion operations according to the mission timeline. Finally, level 4 concerns the ability to conduct the rest of the nominal mission timeline. 2.2.1. Level 1—instantaneous spacecraft safety This level concerns three basic functions that are essential for spacecraft safety at any moment: • The ability to provide power to the essential equipment. This is ensured by implementing multi-redundant solar array shunt regulators and battery management segments together with a priority-based power disconnect network. • The ability to get fundamental telecommands through to the spacecraft. This is ensured by the implementation of a hot redundant telecommand chain, omni-directional antennas, and hardwired decoding and distribution of pulse commands. • The ability to provide essential housekeeping data to ground to allow spacecraft operators to get a concise snapshot of the on-board status. This is ensured by hardwired collection and coding of fundamental status telemetry.

2.2.2. Level 2—durable spacecraft safety This level concerns the functions, in addition to level 1 functions, that are essential for long-lasting spacecraft safety: • The ability to reach and maintain operational thermal conditions. The thermal software task in the spacecraft controller is active and controls the temperature of the safety critical spacecraft equipment within their operational range. • The ability to generate sustained electrical power. The attitude control function is active in its safe mode driving the solar arrays to their home position and pointing them perpendicular to the sun. • The ability to transmit telemetry. The data handling and distribution function is active and transmits essential and software enriched housekeeping data on the nominal down-link channel. 2.2.3. Level 3—electric propulsion operations This level concerns the functions, in addition to level 1 and 2 functions, that are needed to conduct the electric propulsion operation according to the nominal mission timeline: • The ability to conduct the electric propulsion timeline. The time-tagged telecommand storage is accessible and the command dispatch function is active. • The ability to generate adequate electrical power and point the thrust vector in the planned direction. The attitude control function is continuously slewing the spacecraft according to the planned attitude pro;le. The solar arrays are maintained perpendicular to the sun and the spacecraft’s angular momentum is minimised using the 2-axis electric propulsion thruster pointing mechanism. 2.2.4. Level 4—science and technology operations This level concerns the functions, in addition to level 1, 2 and 3 functions, that are needed to conduct the science and technology mission timeline: • The ability to operate the payload and retrieve the generated data. • The ability to Earth point the high data rate antennas, Sand X/Ka-band. 2.3. Avionics implementation The avionics implementation is modular, the main components being the Spacecraft Controller, the power controller, the TM& TC electronics and the remote terminal units

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TC-handler Power Control and Distribution

Battery mngmnt Telemetry

POWER TC-handler

PULSE COMMANDS

MGA LGA2 LGA1

Telemetry MASS MEM

Spacecraft controller

Backup controller

TM-TX-A

MASS MEM

TC&TM Electronics

Payload Heaters

2:1 red

Rate Sensors

2:1 red

AMIE

TC-RX-B

Transponder

Reaction wheels

Thermal control

D-CIXS

Thermistors BAPTA

Antenna switches

TM-TX-B

On-board Computer

Remote Terminals

TC-RX-A

2:1 red

Sun sensors

2:1 red Hydrazine system

Star Tracker

Heaters

SIR Valves

EPDP Electric propulsion

EP gimbal drive

SPEDE

KATE

Fig. 4. SMART-1 avionics architecture.

interfacing with the propulsion, attitude control and thermal hardware (see Fig. 4). We describe the main functions in the following section, see Stagnaro (2001) for more detailed information. All functions are redundant and can be cross recon;gured to communicate with each other through the system data bus. Each spacecraft controller implements a dedicated payload data bus to communicate with the instruments. All of the electronics are implemented in three main boxes: the System Unit (SU), the Remote Terminal Unit (RTU) and the Power Control and Distribution Unit (PCDU). 2.3.1. Spacecraft Controller The Spacecraft Controller is the only on-board computer and is built around a single-chip microprocessor ERC32 developed by ESA. Though built in radiation-resistant silicon technology, it is essentially a copy of the SPARC V7 microprocessor enabling the use of commercial software. The operating system is VxWorks from Wind River System. The application software was developed with Matlab/Simulink/State8ow and autogenerated into a C subroutine with Real-Time Workshop. The microprocessor is clocked at 20 MHz and equipped with 3 Mbytes of SRAM, 2 Mbytes of EEPROM and 512 Mbytes of DRAM used as mass memory. All memory devices, including the mass memory DRAM’s are EDAC protected against Single Event Upsets (SEU). The Spacecraft

Controller also contains two protocol controller devices, for interfacing the system and payload data bus, and the controls of the pyrotechnic devices. The Spacecraft Controller has one cold redundant copy that is switched in its place when a permanent fault arises in the computer. The replacement of the nominal computer with the redundant is controlled by the Supervisor electronics hosted in the PCDU. To determine when the nominal Spacecraft Controller is healthy, the Supervisor monitors the system data bus for a periodic message generated by the on-board computer. In the case this fails to arrive within a speci;c interval the Supervisor will assume that this is the symptom of a fault in the nominal computer, and the redundant Spacecraft Controller will be powered in its place. The Spacecraft Controller also has one internal watchdog to monitor the software behaviour; this watchdog causes the computer to reset if one of the critical tasks fails to perform its duties within the speci;c period. The Supervisor and the internal watchdog expiration time are selected to allow for 2–3 computer resets before the redundancy switch-over takes place. 2.3.2. System and Payload data bus The System and Payload data buses transport information using the same data protocol known as Control Area Network (CAN) standard. The protocol is a well known ;eld-bus standard originally used by the automotive

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2.3.3. Power Control and Distribution Unit (PCDU) The PCDU is responsible for regulating the power bus to 50 V and to distribute the power to users via SSPC’s. The SSPCs act as active fuses by limiting the current to users in case of over consumption or main bus voltage drop. The PCDU also incorporates the spacecraft supervisor which monitors the system data bus and recon;gures the spacecraft controllers in case of anomalies. The SSPC settings are such that non-essential loads are powered-o> prior to more critical functions such as the spacecraft controller, the receiver and the TM& TC electronics, and the PCDU itself. The PCDU also interfaces with the ;ve battery management electronics, each in charge of the battery charge and discharge control of a single lithium-ion cell. 2.3.4. TM& TC electronics The telecommand decoding and telemetry encoding function is handled autonomously by the hardware of the TM& TC electronics, without involvement of the on-board computer. Through this part of the avionics the ground control always has knowledge of the on-board status and can issue recon;guration commands overriding any other on-board commands. One omni-directional communication link in S-band guarantees constant communication with Earth regardless of the spacecraft attitude. The ESA developed Packet Telecommand Decoder device decodes and routes the TC packets coming from ground control and also provides the set of 24 high priority commands distributed as hard-wired pulse commands to di>erent parts of the avionics to force the recon;guration of the spacecraft. Two telemetry stream come from separate sources to the TM& TC electronics and assembled together completely via hardware in telemetry frames. The ;rst of the telemetry streams is marked as Virtual Channel 0 and comes from the PCDU, transporting the essential housekeeping data. The second telemetry stream is marked as Virtual Channel 7 and comes from the Spacecraft Controller, containing both science telemetry, housekeeping and diagnostics. 2.3.5. Remote terminal units The RTU act as interface between the system data bus and the COTS electronic elements used on SMART-1, so

CAN bus A

CAN bus B

Power

DC-converter CAN I/F A CAN I/F B

Dedicated H/W

Power bus 50V

CAN controller

industry, but has gained large acceptance in the industrial application. The data layer is handled by a protocol controller implemented using the COTS VHDL (VHSIC— Very High Speed Integrated Circuits—Hardware Description Language) core from Bosch into one radiation resistant Field Programmable Gate Array (FPGA), while the physical layer uses RS-485 transceivers. Both busses are redundant and the active bus is controlled by the Spacecraft Controller. To reduce cost and smooth the integration between payload and spacecraft, the instrument designers have been provided with the electronic schematic and components to build their interface toward the bus.

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Signal I/F

COTS device or Payload

Fig. 5. Generic Remote Terminal Unit block diagram.

as to provide an homogeneous interface to the Spacecraft Controller. They also act as DC/DC converters for the units they control since the SMART-1 power bus voltage level is incompatible with most commercial parts. There are ;ve RTUs: 1. thermal control RTU serving thermistors and heaters; 2. hydrazine subsystem RTU serving thruster valves, heaters and thermocouples, and pressure transducer; 3. reaction wheel and angular rate sensor RTU; 4. sun sensor, star sensor and solar array drive mechanism RTU; 5. electric propulsion RTU serving the power processing unit and the pressure regulator. The ;rst of the above RTU’s is integrated in the S Unit, while the other are integrated together in a box called RT Unit. A sixth self-standing RTU, is the Thruster Orientation Mechanism (TOM) electronics. Each RTU has a generic part being the DC/DC converter and the CAN bus coupler and controller FPGA, and an application speci;c part dedicated to the equipment it is serving. A conceptual design is shown in Fig. 5. 3. Primary electric propulsion The most important technology to be 8own is the Solar Electric Primary Propulsion (SEPP). Indeed, SMART-1 shall demonstrate the system design, integration, mission analysis, as well as the operation aspects of the SEPP. The requirements for a speci;c engine are strongly dependent upon the size of the spacecraft, its available power and the mission total SV . Since Europe has today a large inventory of electric thrusters, (see e.g., Saccoccia et al., 1998; Racca et al., 1998a) no development for a new electric propulsion engine is required. These thrusters are currently under development or already at quali;cation level for application on telecommunication spacecraft. The use and the capability of electric propulsion as primary propulsion for planetary mission has been described by Racca (2001). All these di>erent types of engine have advantages and disadvantages, but in principle they could be used for primary propulsion in Deep

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Fig. 6. View of the SMART-1 8ight model spacecraft during the Thermal Vacuum test in the Large Solar Simulation facility at ESTEC. The −X panel with the scienti;c instruments SIR, AMIE, D-CIXS and the SPEDE boom is in the foreground. The +Y panel is facing up. The white area is the radiator. The solar array wing will be installed in the centre of the panel. On the same face is also visible the star tracker cut-out besides the solar array attachment point. In the centre of the −Z face, a dummy Electric Propulsion thruster PPS-1350 is mounted together with its high temperature radiator on the thruster orientation mechanism. Furthermore, also one of the two S-band low gain antennas and one coarse sun sensor are visible on −X panel.

Space missions for which SMART-1 is a small technology demonstrator. The selected EP system is based on a Stationary Plasma Thruster (PPS-1350), which constitute a family of electric propulsion engines belonging to the category of Hall-e:ect thrusters. In this type of thruster, electrons from an external cathode enter an annular ceramic discharge chamber, attracted by an anode piece. On their way to the anode, the electrons encounter a radial magnetic ;eld created between inner and outer coils, causing cyclotron motion around the magnetic ;eld lines and dramatically increasing the probability of collision with neutrals. Collisions between drifting electrons and Xenon gas create the plasma. The generated primary ions are accelerated by the negative potential existing in the area near the exit of the chamber due to the Hall-e>ect. The external cathode acts also as a neutraliser, injecting electrons into the beam, in order to maintain zero-charge equilibrium in the thrust beam and on the spacecraft.

The PPS-1350, shown in Fig. 2, has an exit diameter of 100 mm. Even though the thruster can operate at a maximum discharge power of 1500 W, it has also demonstrated its capability to start-up and function at much reduced power (480 W). At beginning of life (BOL), the SMART-1 spacecraft will provide a maximum of 1190 W of nominal discharge power, which will produce a nominal thrust of 68 mN at a speci;c impulse of 1640 s. This family of thruster has a quali;ed lifetime of 7000 h in cycles at maximum power, which corresponds to a total impulse of 2 × 106 N s (Fig. 6). In addition to the thruster, the Electric Propulsion Subsystem (EPS) is composed of: • the Power Processing Unit (PPU) which controls the electrical functions of the thruster • the bang–bang Pressure Regulation Unit (PRU) which provides the regulated pressure level required by the thruster

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Fig. 7. SMART-1 Araine 5 launch con;guration.

• the EP Remote Terminal Unit (EPRTU) which control the PRU and are connected to the S/C system via the CAN bus • the Electrical Filter Unit (FU) which ;lters the conducted emission produced by the thruster discharge current • the redundant Xenon Flow Control unit (XFC) which modulates the mass 8ow rate entering the thruster • the xenon tank for storage • piping and harness between the relevant units including 8exible parts going through the Thruster Orientation Mechanism (TOM). The Xenon gas is stored in a cylindrical composite tank of 49 l with aluminium liner. The tank, from Lincoln Composites USA, has a maximum storage capacity of 82 kg at a maximum e>ective operating pressure of 150 bars and a maximum temperature of 50◦ C. 4. Launcher SMART-1 will be launched as an auxiliary payload by an Ariane 5 vehicle into a Standard Geostationary Transfer Orbit (GTO). The launch con;guration is shown in Fig. 7. SMART-1 will interface with the adapter ACU937V5. The cylinders surrounding the spacecraft are to support the main passenger and to provide a separation device to allow the SMART-1 deployment. As auxiliary passenger SMART-1 cannot negotiate a launch date, but will be launched at the suitable time de;ned by the main passenger between October 2002 and September 2003. Due to this, the mission analysis has to show mission feasibility for a launch any day of a calendar year. 5. Mission analysis 5.1. Trajectory optimisation and design The major objective of the trajectory design is to ;nd a numerically integrated and continuous trajectory starting at

the injection orbit and ending at the operational orbit around the Moon. This has to be performed in a propellant eOcient way and for a given transfer duration which has been selected to be less than 18 months for a departure from a GTO. The optimisation of low-thrust trajectories has been studied by ESA extensively in the early 1980s (Hechler, 1983, 1986). More recently, with the advent of real missions based on Electric Propulsion, the problem has been tackled again in a more operational fashion. Jehn et al. (2000) describes and provides references for the optimisation methods used by ESA/ESOC. The trajectory to be optimised starts from the Earth GTO and ends once in lunar operational orbit. For optimisation purposes the trajectory is divided into four phases: 1. from GTO to an orbit with perigee altitude 20; 000 km, apogee altitude about 68; 000 km and inclination 7◦ ; 2. from the phase 1 ;nal orbit to a 135; 000 km × 338; 000 km, about 30◦ inclination orbit; 3. from the phase 2 ;nal orbit to a complete lunar capture; 4. from the lunar capture to the lunar operational orbit. During the phase 1 trajectory, a continuous tangential thrusting is applied in order to leave the radiation belts zone as soon as possible. The phase 2 trajectory is optimised by applying a method based on the Pontryagin maximum principle (Jehn and Cano, 1999). This method relies on an idea by Ge>roy (1997) to solve the associated boundary value problem. The same method is used for the optimisation of the phase 4 trajectory. The solution to the problem is a trajectory which combines coast and thrust arcs. During the thrust arcs, the engine is ;red in a direction which has an out-of-plane and in-plane component with respect to the velocity vector. The optimum strategy during phase 2 is to raise the apogee at the beginning with thrusting almost parallel to the velocity vector between −100◦ and +120◦ true anomaly. The thrust arcs are then gradually expanded to reach about −130◦ + 145◦ . Towards the end, the thrust arcs are around the apogee to raise the perigee and at the same time to change the inclination. The thrust direction lies between −10◦ and +50◦ out-of-plane and ±30◦ in-plane.

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5.

Earth centred inertial projection on moon orbital plane SPT100: launch by AR5; 2002/09/21 -> 2003/12/29; 350 kg -> 295.764 kg;

4. 3.

Y-coordinate (x 1E5 km)

2. 1. 0. -1. -2. -3. -4. -5. -5.

-4.

-3.

-2.

-1.

0.

1.

2.

3.

4.

5.

X-coordinate (x 1E5 km) (ascending node of moon orbital plane)

Fig. 8. Transfer trajectory from GTO to lunar orbit.

The optimisation during the third phase and the matching of the three phases is calculated using a gradient projection method for a set of parameters de;ning the thrust law directly (Pulido Cobo and Schoenmaekers, (2000). The building blocks of these trajectories are thrust arcs, Moon resonances, Moon swing-bys and the lunar capture (Schoenmaekers et al., 1999). During the ascent spiral, the apogee radius gradually increases. From 200; 000 km onwards, the Moon starts to signi;cantly perturb the orbit once every lunar revolution (i.e. every 27.4 days). These perturbations are called Moon resonances and occur near apogee when the Earth–spacecraft direction is close to the Earth–Moon direction. The Moon perturbation is only signi;cant over a small part of the orbit near the point of closest approach, which is near to apogee. The magnitude of the perturbation increases with decreasing distance to the Moon. The only parameter of a Moon resonance which can be easily controlled is the phasing with the Moon when reaching the apogee. It is controlled by tuning the orbital periods prior to the resonance via the length of the thrust arcs. Moon resonances are encounters with the Moon outside its sphere of in8uence. Once the distance of closest approach to the Moon gets within the sphere of in8uence (i.e. lower than 60; 000 km), we start speaking about Moon swing-bys. The patched conics model provides a convenient way to

qualitatively understand what happens at a Moon swing-by. During the period when the spacecraft is in the sphere of in8uence of the Moon, it is on an hyperbolic trajectory w.r.t. the Moon. The outgoing asymptotic velocity has the same magnitude as the incoming asymptotic velocity. The direction is changed in the plane containing the Moon and the incoming asymptote. The de8ection is towards the Moon and increases with decreasing magnitude of the asymptotic velocity and decreasing distance between the Moon and the incoming asymptote. This allows one to look at a swing-by as an instantaneous velocity change. The spacecraft trajectory prior to the swing-by is propagated ignoring the gravity of the Moon until the point of closest approach is reached. At that point the velocity is changed according to the effect of the Moon gravity. The trajectory, see Fig. 8, is then further propagated ignoring the gravity of the Moon. The lunar capture trajectory and the subsequent transfer to the operational orbit around the Moon is computed by backward integration. The right ascension of the ascending node in the Moon equator as well as the arrival epoch in the operational orbit are free parameters. Starting in the operational orbit, the trajectory is integrated backwards, applying a continuous thrust opposite to the velocity. The thrust spiral ends when the osculating apolune radius reaches 60; 000 km. This provides the spacecraft with a large enough value of the Jacobian integral to leave the sphere of in8uence of the

G.D. Racca et al. / Planetary and Space Science 50 (2002) 1323 – 1337

0.6

1333

Inertial projection on target orbital e plan SPT100: launch by AR5; 2002/09/21 -> 2003/12/29; 350 kg -> 295.764 kg;

0.5

Z-coordinate (x 1E5 km) (moon north pole)

0.4 0.3 0.2 0.1 0.0 -0.1 -0.2 -0.3 -0.4 -0.5 -0.6 -0.6

-0.5 -0.4

-0.3

-0.2 -0.1 0.0 0.1 0.2 X-coordinate (x 1E5 km)

0.3

0.4

0.5

0.6

Fig. 9. Capture and orbit around the Moon. Table 3 Main orbit characteristics of the transfer phase

Orbit

Epoch (days)

Perigee radius (km)

Apogee radius (km)

Orbital period (h)

Thrust arc start/stop (◦)

Thrust duration (h)

GTO End-belts Spiral

0 83 313

6628 20,000 28,000

42,164 68,000 250,000

10.53 25.51 148.07

Cont. −100= + 120 −130= + 145

Cont. 6.53 30

Moon and consequently enter into the sphere of in8uence of the Earth. With the aim to get the spacecraft in an orbit with apogee below the Moon orbit, the transfer to the Earth sphere of in8uence is via the L1 point, which is the cislunar Lagrangian point of the Earth–Moon system. To avoid unstable trajectories, the transfer is direct, so after thruster switch-o>, the spacecraft leaves the sphere of in8uence of the Earth without making an additional revolution around the Moon, see Fig. 9. 5.2. Mission phases

perigee of 250 km has been adopted, instead of the more advantageous 600 km, in order to comply with the regulations to limit the dispersion of space debris at high altitudes. The right ascension of the ascending node depends on the day of launch and on the lift-o> time. As auxiliary passenger, SMART-1 must be prepared to accept all possible values for the Right Ascension of the Ascending Node (RAAN) of a midnight launch during the whole year. SMART-1 will remain for about ten orbits in the GTO, for checkout purposes, before the electric propulsion engine is activated and the transfer manoeuvre is initiated. The characteristics of the orbit are shown in Table 3.

5.2.1. Launch and early orbit phase Ariane 5 injects the spacecraft in GTO with a perigee radius of 6628 km, an apogee radius of 42; 164 km, an inclination of 7◦ and an argument of perigee of 178◦ . The low

5.2.2. Earth spiralling phase After a ten revolutions in GTO, a continuous thrust spiral along the velocity is started up to the point where the perigee radius reaches 20; 000 km. The aim is to limit the

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degradation of the solar arrays in the radiation belts as much as possible. During this phase the apogee radius is increased to about 68; 000 km. This phase lasts about 73 days, depending on the actual performance of the power and electric propulsion systems. The trajectory shown in Fig. 8 has been generated for a launch on 21 September 2002. The right ascension of the ascending node when reaching the perigee radius of 20; 000 km is 147◦ . The apogee is raised with the proper phasing by tangential thrust arcs symmetric around perigee until a close swing-by with the Moon is reached. Typical orbit characteristics of this phase are shown in Table 3. The exact duration and position of the thrust and phases depend on the launch date. However, it is clear that plenty of time will be available for coast arcs, as typically only 20 –25% of the orbital time will be used for thrusting. During the coast arcs, the pointing needs not to be along the thrust direction and therefore cruise science observations can take place. When the apogee reaches a suitable altitude, Moon resonances are used to increase the perigee radius prior to the swing-by, as explained earlier. The time necessary from the GTO to the lunar capture depends on the relative geometry of the GTO and Moon orbits, as do the number and duration of the thrust arcs and hence the propellant used. Depending on these conditions a transfer trajectory duration ranging from 430 to 490 days has been found. Another important aspect of the mission analysis is the avoidance of long eclipses during the transfer trajectory. While short eclipses (6 0:5 h) in the ;rst revolutions will be unavoidable for any launch date, launches around the equinoxes would produce, depending on the various parameters de;ning the trajectory, long eclipses up to about 3:7 h at the orbital apocentre. Since such no-power periods cannot be tolerated by the spacecraft design, they will be avoided by slight changes in the line of apsides previous to these and by delaying the time at which the spacecraft passes through the anti-sun region. In this way the duration of the longest eclipses can be limited to 2 h.

gravitational pull of the Earth. The selection of the argument of perilune turns out to be an important parameter to ensure the orbit’s stability. In addition, (Cano et al., 1999) have shown that the initial RAAN slightly modulates the perilune argument required to have a total lifetime of 180 days, without manoeuvres. The initial RAAN cannot be freely chosen, as it depends on the arrival conditions, hence an appropriate value of the argument of perilune has been calculated for any initial RAAN. As an example, with an initial RAAN of 0◦ , an initial argument of perilune of 305◦ ensures the required stability. The altitude of perilune from 300 km will reach about 1900 km after 90 days while the argument of perilune moves to 270◦ , directly on the South-Pole, and the altitude of the apolune decreases from 10,000 to 8400 km, almost keeping a constant semi-major axis. After another 90 days the altitudes will be restored to the original values, while the argument of perilune evolves to a value of 235◦ . This orbit evolution provides a good coverage of the ±30◦ southern latitudes. The orbital period is practically constant at 14:25 h. The eclipses period and durations are dependent on the initial RAAN. During 180 days of nominal mission, there are eclipse free periods of about 120 days, whose position vary according to the RAAN. The eclipse durations vary from zero to a maximum of 1:4 h (see Table 4). In order to improve the high-resolution coverage of the Moon’s surface, a reduction of the orbital altitude at perilune will be performed if there will be suOcient propellant left. Apocentre altitude of 6000, 4000 and 3000 km have been considered, leaving the pericentre altitude at 300 km. Table 4 shows the characteristics of these orbits together with the propellant consumed and the transfer time from the baseline orbit (B0). By using longer transfer times the propellant usage can be reduced. The decision whether to lower the apolune and to what level will be made according to the propellant usage during the mission.

5.2.3. Lunar capture Lunar capture starts from when the spacecraft is gravitationally bound to the Moon (see Fig. 8), though still very perturbed by the Earth. For this reason is necessary to lower the capture orbit towards the operational orbit by means of thrusting. Fig. 9 shows the capture trajectory from a semi-major axis of about 28; 000 km to an operational orbit of about 7200 km semi-major axis. The manoeuvre duration is in the order of 32 days.

The ground segment can be subdivided into the following three main entities as shown in Fig. 10.

5.2.4. Lunar operational orbit The nominal operational lunar orbit is polar, with perilune near the South-Pole and altitudes of periapsis of 300 and 10; 000 km. However, such a polar orbit has been proved to be rather unstable (Cano et al., 1999). The orbit would decay very rapidly into the surface of the Moon, mainly due to the

6. Ground segment

6.1. Mission Operations Centre (MOC) This will be located in ESOC in Darmstadt. The MOC consists of the Main Control Room (MCR) augmented by the Flight Dynamics Room (FDR) and Dedicated Control Rooms (DCR’s) and Project Support Rooms (PSR’s). Initially during the Launch & Early Orbit Phase (LEOP) and during the Moon Capture phase the MCR will be used for SMART-1 mission control. During the routine operations phase the mission control will be conducted from a Dedicated Control Room (DCR). The control centre is equipped with workstations giving access to the di>erent computer systems used for di>erent tasks of operational data processing.

G.D. Racca et al. / Planetary and Space Science 50 (2002) 1323 – 1337

1335

Table 4 Main orbit characteristics of the operational lunar orbits

Orbit ID

Apocentre altitude (km)

Pericentre altitude (km)

Orbital period (hr)

Eclipse duration (hr)

Transfer time (days)

Propellant required (kg)

B0 B1 B2 B3

10,000 6000 4000 3000

300 300 300 300

14.25 8.52 6.04 4.92

1.43 1.11 1.03 1.01

— 16.5 28.7 36.8

— 4.6 8.4 10.5

IOF 1

TC

IOF 2

TC

IOF …

TC

MOC

TC TM

TC

TM packets

ESMP

TM

STOC

G/S 1

Feedback on TC packets

OPEN NETWORK (Internet)

G/S 2

Remote Archive (ESRIN)

MOC = Mission Operations Centre STOC = Science and Technology Operations Coordination IOF = Instrument Operations Facility ESMP = Experiment Science Master Plan

Fig. 10. Ground segment overview.

All ESA ground stations interface to the MOC via the OPSNET communications network. OPSNET is a closed Wide Area Network for data (telecommand, telemetry, tracking data, station monitoring and control data) and voice. Data traOc is generally provided by a Data Packet Switching System based upon the X.25 international standard. It will support the SMART-1 TM data rate of 65 Kb s−1 transfer to the MOC. As explained earlier, ground stations will be used on an availability basis. However, it is expected that the Perth station in Australia and the Maspalomas or Villafranca stations in Spain will be the main stations. The following main functions of the MOC can be identi;ed: 6.1.1. Spacecraft operations This includes derivation of operational requirements for the spacecraft and de;nition of user requirements for the ground systems to be implemented. The main tasks are

the operations preparation, generation of 8ight operation plans and procedures, execution of spacecraft operations and in-orbit mission evaluations. 6.1.2. Data systems This covers in particular the speci;c mission control hardware and software for spacecraft control, the satellite simulators, and payload data processing. 6.1.3. Flight dynamics This covers mission analysis, orbit and attitude aspects. In particular it includes investigations of launch window, orbit selection and evolution, manoeuvre strategies and optimisations, tracking and navigation and sensor and instrument performances. In addition it encompass the implementation of the speci;c applications software systems required to perform planning and execution of all aspects of 8ight dynamics for the mission.

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6.1.4. Stations and facilities This concerns the implementation of the mission-speci;c ground station facilities and the associated communications systems required to support mission operations. In particular, the design of space–Earth and Earth–space links, including link performance budgets and the planning and execution of the ground station acceptance and compatibility tests. 6.2. Science and Technology Operations Coordination (STOC) The STOC is located in ESTEC in Noordwijk. The STOC interfaces are on one hand towards the Principal Investigators and Technology Investigators, who will refer to the STOC for the coordination of their science and technology operations and for data collection and exploitation. On the other hand, STOC interfaces to the MOC to which it provides the inputs to the FOP (Flight Operation Plan) for the payload commanding at spacecraft level, and the ESMP (Experiment Science Master Plan), containing the speci;c payload telecommand sequences and the operational timelines. The key tasks of the STOC are described in Marini et al. (2002) and can be summarised as follows: • Allocation of observation slots. • Coordination and approval of the use of ground and on-board resources (e.g. TC queues, data volume, mass memory). • Adequate allocation of the combined operations of instruments. • Generation of the Experiments Operations time-line as a function of the mission status. • Supervision on the experiment TC packets generation. • Monitoring of the distribution of the TM packets to PI’s/TI’s. • Raw data storage during the mission and support to science archiving and exploitation. 6.3. Experiment Operations Facilities (EOF) The EOF’s are located at each Principle Investigator site. They are connected to the STOC and to the MOC via the open network and they remotely operate the experiments. Coordination of the scienti;c activities is carried out within the Science and Technology Working Team (STWT) and via the STOC. Each Experiment Team has access to the ESOC Payload control room during commissioning, but no other routine real-time operation is foreseen. The tools and interface to the STOC are de;ned to be of the maximum simplicity. Experiment requests for operations, commands and data delivery are routed via the STOC in an agreed format and procedure.

7. Development status After successful completion of the phase B de;nition study the ESA Science Programme Committee gave its ;nal approval to the mission in November 1999. The project is currently in the ;nal stage of the development phase C/D. The following major milestones at spacecraft level are foreseen: Phase C/D kick-o> Critical Design Review System Integration Ready Environmental Test campaign Flight Acceptance Review Launch readiness Cruise phase (Earth–Moon) Moon operations

October 1999 August 2001 February 2002 October–December 2002 January 2003 mid-March 2003 14 –18 months 6 months

The spacecraft veri;cation activities and the development status of the STOC have been described in Marini et al. (2001). Acknowledgements This article has been drawn from the work performed during the Phase B and C/D by several individuals and organisations. The authors are the key personnel of the main organisations involved; however, they are not the only persons who contributed to the work described. Therefore, the authors thankfully acknowledge the work of their colleagues in Alcatel Espacio (E), Alcatel ETCA (B), APCO (CH), ATERMES (F), Contraves (CH), CRISA (E), DTU (DK), ETCA (B), Dutch Space (NL), General Dynamics (USA), Iberespacio (E), Ithaco (USA), Omnisys (S), Spacebel (B), SAAB Ericsson Space (S), SAFT (F), SEP/SNECMA (F), Terma (DK), TNO (NL), the Principal Investigators of the science and technology experiments: RAL (UK), LABEN (I), Astrium (D), MPIAe (D), CSEM (CH), FIM (SF), Helsinki University (SF) and thank the whole science and technology community for their support to the mission. Special thanks also to Dr. David Heather who refereed an early version of the article and provided many useful comments.

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G.D. Racca et al. / Planetary and Space Science 50 (2002) 1323 – 1337 Ge>roy, S., 1997. GTenTeralisation des Techniques de Moyennation en Contrˆole Optimal—Application aux ProblVeme de Transfert et Rendez-Vous Orbitaux aV PousTee Faible. Doctoral Thesis at CNES, Toulouse, France, 1997. Hechler, M., 1983. AGORA example low thrust asteroid rendezvous trajectories. MAS-WP No. 194, ESOC, Darmstadt, Germany, August. Hechler, M., 1986. Example low thrust comet sample return mission. MAS-WP No. 232, ESOC, Darmstadt, Germany, January. Jehn, R., Cano, J.L., 1999. Optimum low thrust transfer between two orbits. MAS-WP No. 414, ESOC, Darmstadt, Germany, March. Jehn, R., Hechler, M., Rodriguez-Canabal, J., Schoenmaekers, J., Cano, J.L., 2000. Trajectory optimisation for ESA low-thrust interplanetary and lunar missions. CNES Workshop on Low-Thrust Trajectory Optimisation, Toulouse, France, 7–8 March. Marini, A., Racca, G.D., Foing, B., 2000. SMART-1 Technology experiments in preparation to future ESA planetary missions. Fifth International Symposium on Small Satellites, Systems and Services. Proceedings, La Baule, France, June 19 –23. Marini, A.E., Gr1unagel, R., Elfving, A., Racca, G.D., Foing, B.H., 2001. The ESA SMART-1 mission and veri;cation status. 52nd IAF Congress, Paper IAA-01-IAA.11.2.01, Toulouse, France, 1–5 October. Marini, A.E., Lumb, R., Racca, G.D., Foing, B.H., Dias-Almeida, M., 2002. Technology and science from Earth to Moon: SMART-1 experiments and their operations. Proceedings of ESLAB 36: Earth-like Planets and Moons, June 2002, Noordwijk, The Netherlands. Pulido Cobo, J., Schoenmaekers, J., 2000. The gradient method adapted for SMART-1 trajectory optimisation. S1-ESC-RP-5502, April 2000, ESOC, Darmstadt, Germany.

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Racca, G.D., 1999. SMART-1: the ;rst small mission for advanced research in technology. Acta Astronaut. 45 (4 –9), 337–345. Racca, G.D., 2001. Capability of solar electric propulsion for planetary missions. Planet. Space Sci. 49, 1437–1444. Racca, G.D., Marini, A.E., 2000. SMART-1: preparing next generation mission to Mercury. 51st IAF Congress, Paper IAA-00-IAA.11.2.06, Rio de Janeiro, Brazil, 2– 6 October. Racca, G.D., Whitcomb, G.P., Foing, B.H., 1998a. The SMART-1 mission. ESA Bull. 95, 72–81. Racca, G.D., Estublier, D., Marini, A.E., Saccoccia, G., Whitcomb, G.P., 1998b. An overview of the ;rst ESA small mission for advanced research and technology. Fourth International Symposium on Small Satellites, Systems and Services, Proceedings, Antibes, France, September 14 –18. Racca, G.D., Foing, B.H., Coradini, M., 2001. SMART-1: the ;rst time of Europe to the Moon. Earth, Moon Planets 85 –86, 379–390. Saccoccia, G., Estublier, D., Racca, G.D., 1998. SMART-1: A technology demonstration mission for science using electric propulsion, AIAA-98-3330. Fourth AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Cleveland, USA, July 13–15. Schoenmaekers, J., Pulido, J., Cano, J., 1999. SMART-1 Moon mission: trajectory design using the Moon gravity. S1-ESC-RP-5501, ESOC, Darmstadt, Germany, May 1999. Stagnaro, L., 2001. The SMART-1 data handling architecture. Proceedings of Data Systems In Aerospace, ESA SP-483, Noordwijk, The Netherlands.