EMC ISSUES ON BEPICOLOMBO SPACECRAFT

Dr. Klaus Kempkens ESA Projects

EADS Astrium GmbH Friedrichshafen, Germany

Abstract— This paper describes critical EMC issues on the BepiColombo Spacecraft, the ESA mission to Mercury. EMC on this spacecraft is driven by a high power consuming electric propulsion system which can generate extremely high transient currents during operation, and the need for steerable high gain antennas for space to ground communication which can illuminate part of the spacecraft under certain link constellations. Specifically on the European orbiter, accommodation of a gradiometer instrument requires measures to achieve DC and low frequency AC magnetic cleanliness.

Keywords — BepiColombo, Electric Propulsion, power transients, antenna electric fields

I. INTRODUCTION BepiColombo is the joint European/ Japanese scientific

mission to Mercury, the innermost planet of our solar system. The mission features a special “3 in 1” spacecraft, consisting of an assembly of 2 orbiters, and a transfer vehicle. The MPO is the European Mercury Planetary Orbiter carrying the ESA payload, while the MMO is a second spacecraft, the Japanese Mercury Magnetospheric Orbiter built under the auspices of JAXA. Both scientific spacecrafts are brought to Mercury by the Mercury Transfer Module (MTM) containing propulsion and power systems for the cruise phase. While the MPO and MMO designs are optimised for operation in Mercury orbit, the MTM provides pre-Mercury functions. The MCS cruise configuration is progressively disassembled approaching the arrival at Mercury by in-flight separations until the 2 orbiters are operational in their respective orbits.

Fig. 1. BepiColombo Spacecraft in cruise configuration

The scientific spacecrafts will operate in different polar orbits around Mercury, with the MPO providing global high resolution coverage focusing on planet surface and interior science, while the MMO is focusing on the planetary environment, in particular the planet’s magnetosphere. During cruise phase, the MTM will provide power to the MPO and MMO, while command & control is performed by the MPO. The transfer phase is accomplished by an electrical propulsion system within the MTM, comprising 4 steerable Xenon-fuelled ion thrusters which operate either singly or in pairs. The power generation and distribution hardware needed during Cruise, generating more than 11 kW, resides in the MTM.

The MPO module provides communication to earth, both in cruise phase and during MPO scientific operation when orbiting Mercury. This is achieved by communication systems featuring steerable antennas. Among the MPO spacecraft payload suite of 11 instruments there is a magnetic gradiometer Mermag whose sensors are mounted on a boom deployed immediately after launch. With regard to EMC, the main critical areas are

• The Power subsystem of the transfer module, powering the Electric propulsion

• Transmitter/ receiver EMC, in both cruise and orbit phase, under consideration of changing antenna orientations

• DC/ AC magnetic performance of the MPO spacecraft to ensure a clean magnetic environment for operation of the Mermag magnetic gradiometer instrument

II. POWER SUBSYSTEM Each of the 3 modules MPO, MMO and MTM features a

power subsystem. The MPO and MMO power subsystems supply standard 28 V regulated power within the module (the MMO is not subject of this paper), with the MMO being supplied by the MPO as long as both modules are connected. During cruise phase the MPO steerable solar array is pointing off the sun, the MPO/ MMO being supplied by the MTM via a 60 V link. In this phase the main power consumer is however the Solar Electrical Propulsion System (SEPS) which drives the size of the MTM power system with its 2 steerable solar array wings. When running at maximum thrust, there are 2 ion thrusters operating in parallel, each providing a force of 145 mN. The ion thrusters feature a complex combination of supply lines (ion beam, anode, acceleration grid, solenoids etc.), running at dedicated voltages in between -265 V and +1850 V. The most important lines are shown in fig. 2.

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These voltages are generated in the SEPS Power Processing Unit (PPU), of which two are implemented on the MTM. Input power at maximum thrust is 5300 W each.

Fig. 2. Propulsion Power Subsystem, main lines

Most critical for EMC are any transients associated with the SEPS, which thus had to be carefully assessed for potential impact on other equipment. To achieve compatibility with the power S/S for maximum load cases during SEPS switch-on and thrust force ramp-up, a limitation of current slopes is necessary, taking into account PCDU regulator performance as well as allowable battery transient currents. In all cases stability of the 100 V PCDU main bus must be ensured, and power quality on derived 60 V MPO supply and internal 28 V bus must be maintained.

The most demanding case is the so-called beam-out. This refers to a temporary short circuit between thruster grids e.g. due to particles from grid erosion. In this event the acceleration grid nominally at -265 V is short-circuited to the beam grid at +1825 V, see figure 3 (green curve). Within 1 µsec the beam current of 2.1 A rises to a peak of more than 45 A, thereby generating strong current transients in a number of thruster connections (figure 3 black and blue curve as examples).

Fig. 3. Beam-out transient at Ion thruster

The effects are two-fold: on one hand, they are present on the PPU secondary side, potentially affecting equipment

located close to the thruster or having harness running in parallel to the thruster – PPU connections; on the other hand they result in transients on the PPU supply lines. The latter have been applied to the MTM PCDU during unit level tests, to verify that the quality of the 100 V power bus is not adversely affected.

Fig. 4. Beam-out transient applied to PCDU (100 V supply)

Impact of the transients on thruster side is minimized by use of shielded high voltage cables throughout, which are, moreover, routed separately from all other S/C harness. A critical area however is at the thruster interface, where the thruster pointing mechanism is located in close vicinity. These mechanisms allow two-axis steering of the ion thrusters, featuring sensitive encoders and sensors. To check electromagnetic compatibility a model of the thruster harness was implemented in the pointing mechanism qualification model, and a dedicated transient test performed injecting worst case transients + appropriate margin into the thruster harness. Impact on the sensor signal integrity was then measured and demonstrates that no mechanism step loss will occur if a beam-out transient happens during mechanism operation.

Fig. 5. Pointing mechanism test set-up

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Fig. 6. Position encoder disturbance due to ion thruster pulse at 70 A/µs

(orange/ blue: encoder signals, green: disturbing transient)

III. TRANSMITTER/ RECEIVER EMC BepiColombo Command & Control is performed by an X-

Band system for Earth to S/C commanding and S/C to ground data transmission. For the science phase this system is supplemented by a Ka-Band downlink for payload data transmission, not used in cruise phase. Consequently the spacecraft emissions must be closely controlled in the sensitive X-Band receiver band around 7167 MHz, while large emissions are generated at the transmit frequencies of 8420 MHz and 32 GHz. Moreover, there is a payload instrument featuring a Ka-Band receiver at 34.4 GHz which is however in general also not used during cruise phase.

The communication links are using a suite of 4 antennas, comprising two Low Gain Antennas (LGA) at opposite spacecraft sides (used for LEOP phase), a steerable Medium Gain Antenna (MGA, main antenna in cruise), and a steerable High Gain Antenna (HGA, nominal antenna for science operation, allowing optimized antenna pointing to maximize the link budget). While LGAs and MGA are transmitting/ receiving in X-band only, the HGA provides communication in both X- and Ka-Band.

The E-field levels from LGA antennas can be calculated with good accuracy following standard far-field formulas.

Calculation of electric fields from MGA and HGA however requires the use of simulation programs which fully consider the near-field behavior taking into account antenna patterns & polarization, and spacecraft geometry. For BepiColombo the commercial software CST Microwave Studio® has been used. The algorithm for the solution of the scattering problem is using the so called “Multilevel Fast Multipole Method” (MLFMM) based on the well-known method of moments (MOM), which is used to calculate the surface current distribution by generating and solving an appropriate linear equation system. A recursive solving scheme is used for the efficient final solution of the numerical problem, making the MLFMM algorithm a very effective numerical tool for the

calculation of electromagnetic field solutions for large objects. The resulting spatial current distribution can then be used to analytically calculate the electromagnetic fields on the surface of the structure as well as the far field [2].

Worst case antenna orientations have been identified from spacecraft CAD models, see figure 7 for an example of HGA antenna pointing onto the spacecraft. These orientations serve as an input to the E-field simulations.

Samples of E-field distributions are shown in figure 8 for HGA illumination of Fine Sun Sensor, and figure 9 for MMO illumination by High Gain Antenna. Local E-fields reach levels up to 240 V/m in X-Band as a worst case.

While the external fields from X-Band and Ka-Band transmitters can reach very high levels depending on location and antenna orientation, the spacecraft-internal equipment is shielded against these large E-fields to a certain extent, mainly by the spacecraft Multilayer Isolation (MLI). The spacecraft design is thermally extremely demanding, the satellite being subjected to tremendous thermal fluxes, from both the sun and the Mercury surface. This can only be survived by a special MLI design which features more than 5 times the number of aluminium layers compared to a standard earth-orbiting spacecraft, e.g. TerraSAR, with the side effect that a shielding efficiency of 23 dB can easily be assumed in X- and Ka-band frequency ranges (backed up by measurements from other projects which show a shielding efficiency of 30 dB already for standard design). This allows limiting the RS E-field requirement for internal equipment and instruments to 32 V/m which is within the standard capabilities of EMC test houses, and can easily be complied with for most equipment.

Fig. 7. Fine Sun Sensor illumination by High Gain Antenna – CAD drawing

Fig. 8. E-field distribution at +X FSS, sample plot at azimuth -120°/

elevation 135°

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Fig. 9. MMO orbiter illumination by Higjh Gain Antenna fields during

cruise (MMO S/C hidden behind sun shield)

External equipment being subjected to extreme E-fields had to undergo dedicated equipment level RS tests to prove compatibility with the antenna emissions.

In a similar manner the impact of S/C equipment in the receiver bands has been assessed, resulting in a specification of E-field levels allowable at RX frequencies depending on external equipment location.

IV. DC/ AC MAGNETICS The MPO module features a differential magnetometer

(gradiometer) instrument, Mermag. Strong limits on DC magnetics have been imposed on the spacecraft to ensure that the instrument measurement accuracy is not adversely affected. Meeting these requirements is traced by equipment level DC magnetic verification which data contributes to an overall magnetic budget assessed analytically. All flight units and instruments undergo a DC magnetic test prior to S/C integration.

The background of maintaining a low DC magnetic moment however is ultimately only an aid to minimize the low frequency AC magnetic emissions of the MPO module in order to ensure proper Gradiometer performance, taking into account the instrument measurement principle. The major task was thus to identify any potentially AC critical unit, and test these at an early stage to judge the potential impact on Mermag performance. The main contributors are

• Strong fluctuating currents (mainly in power units) • Mechanisms with rotating magnetic elements (mainly

motors, e.g. flywheels, Solar Array Drive Mechanism, Antenna Pointing Mechanism)

• Instruments with similar mechanisms (Phebus instrument baffle motor, Mertis instrument pointing unit)

• Parts which change their magnetic performance during flight, depending e.g. on temperature. A number of MLI mounting parts have been identified to show such

behavior, which could be significantly reduced by parts demagnetization prior to MLI manufacturing

AC emission testing of critical units was performed in the frame of EM module EMC testing (see sample figure 11: one axis of Phebus baffle motor). The tests revealed that the Phebus baffle rotating reference magnets generate strong AC fields during baffle operation, requiring the installation of compensation magnets at opposite locations, which was then later performed on the flight instrument.

Ultimately an AC magnetic characterization will be performed on the integrated MPO satellite FM, in the frequency range 0-200 Hz. The satellite will be configured and operated in the most flight representative manner possible, operating all AC magnetic critical equipment in parallel. A differential magnetometer will be accommodated in a location representative for the Mermag instrument. Using a gradiometer allows to perform this test in ambient clean room environment, as long as the external magnetic environment can be kept stable during the test in a limited area of only of 2 – 3 m around the spacecraft.

Fig. 10. Phebus baffle AC magnetic field measurement with gradiometer

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Fig. 11. Phebus baffle magnetic field, X axis, uncompensated

V. CONCLUSION Some EMC aspects of BepiColombo spacecraft have been

presented. Although only a few are desive, due to their gravity close monitoring throughout the project is required.

REFERENCES [1] BepiColombo System Design Report, BC-ASD-RP-00003, Astrium

GmbH [2] CST Microwave Studio, User manual, 2013

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