A A Small EP Spacecraft for Plasma Physics Experiments

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The 30th_{International Electric Propulsion Conference, Florence, Italy} September 17-20, 2007

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A Small EP Spacecraft for Plasma Physics Experiments

IEPC-2007-349

Presented at the 30th_{International Electric Propulsion Conference, Florence, Italy} September 17-20, 2007

Salvo Marcuccio*_{, Antonio d’Alfonso}†_{and Francesco Pegoraro}‡

University of Pisa, Pisa, 56121, Italy

An active plasma experiment in space is proposed, involving the generation of an artificial magnetosphere (plasma bubble) attached to a small spacecraft. The streaming plasma bubble interacts with the ionosphere exciting several interesting plasma phenomena, difficult or impossible to observe in the laboratory. The spacecraft is equipped with a low-power Alta HT-100 Hall thruster used both as a plasma source and to provide propulsion to change the orbital altitude in a wide range. The proposed experiment makes also an opportunity for flight demonstration and characterization of the Hall thruster, as well as an occasion to characterize the environment effect on the spacecraft subsystems in the LEO-GEO range of orbital altitudes. We present a mission concept and the spacecraft preliminary design for a low-cost mission capable of performing both the active plasma experiments and the thruster flight demonstration.

I. Introduction

small spacecraft equipped with a plasma source, such as an electric engine, is a powerful tool to carry out active plasma experiments in space. While investigation of the natural space plasma environment is as old as the first man-made satellites, active experiments - i.e., research on artificially generated plasma phenomena - are much less common.1,2_{During the last decade, considerable interest emerged for such experiments, aimed at} investigating plasma regimes that cannot be maintained in ground facilities3_{. Among other concepts, the so called} “artificial magnetosphere” is presently considered as a viable way to generate a whole range of artificial plasma phenomena in space. In such concept, a magnetized plasma bubble is generated around the satellite, providing the unprecedented opportunity of studying the nonlinear interaction of two magnetized streaming plasmas in a controlled setting and in the absence of external conductors that, in the laboratory, restrain their dynamics.

The need to make available a plasma source leads quite naturally to considering electric propulsion. Small satellites with electric thrusters can be equipped with last generation, low power electric thrusters, providing orbital maneuvering capabilities previously reserved to large spacecraft. A small satellite with electric propulsion can be guided through complex trajectories at different orbital altitudes, exposing the experimental setup to a wide variety of natural plasma environments. Moreover, in the case of active plasma experiments, it is reasonable to assume that much of the plasma diagnostics for the physics can be also used to monitor the electric thruster performance.

From the above considerations, it was decided to investigate the feasibility of a low cost small spacecraft equipped with a low power Hall thruster acting both as plasma source for the experiments and as propulsion device, and with a multi-purpose set of plasma diagnostic tools. In this paper, we present the preliminary design of a mission

*_{Dept. of Aerospace Engineering; Vice President for Space Systems, Alta S.p.A.; [email protected]} †_{Dept. of Aerospace Engineering}

‡_{Dept. of Physics, [email protected]}

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intended to perform the plasma experiments and the mini-HET flight demonstration. The mission goals, listed in order of decreasing priority, are:

- Perform a set of active plasma physics experiments involving magnetized streaming plasmas; - Demonstrate in-flight performance and operation capability of the HT-100 low power Hall thruster; - Collect data on the radiation belts environment effects on solar cells and electronic components.

In the following sections, we discuss the rationale of the proposed experiment and present a preliminary design of the spacecraft. A maximum simplicity approach was adopted, with the goal of building and launching the spacecraft as a low-cost mission.

II. Active Plasma Experiments in Space

Space plasmas provide the environment for controlled experiments in collisionless plasma regimes under essentially boundary-free conditions that cannot be realized in the laboratory. A small spacecraft capable of generating a sufficiently strong magnetic field and equipped with a plasma source can provide the means to perform an active plasma experiment where a magnetized plasma bubble streams through the ambient (ionospheric) plasma.4 The magnetized plasma inside the bubble is either tied to a dipole magnetic field generated inside the satellite or is inflated by a particle outflow from the satellite. Such an experiment exploits the velocity difference between the orbital motion of the magnetic plasma bubble tied to the satellite and the motion of the ambient ionospheric plasma, co-rotating with the Earth, in order to study the nonlinear dynamics of counterstreaming magnetized plasmas.

In such a system, measurements aimed at investigating basic nonlinear phenomena such as magnetic field-line reconnection, the onset of magnetic wake turbulence, magnetic vortex dynamics and particle acceleration can be performed. In principle, such an experiment could allow us to investigate the interaction of plasma winds, such as the solar wind, with planet magnetospheres under controlled conditions, without the restrictions due to the presence of close boundaries, as happens in laboratory experiments. However, the huge scale reduction from a planet magnetosphere to a satellite-bound plasma bubble leads us to consider plasma regimes that may not allow a physically significant approximation of the magnetosphere/solar wind interaction, or may do so only marginally and within the more complex experimental scheme of an inflated bubble. Nevertheless, important information on the nonlinear dynamics of collisionless plasmas in the frequency range corresponding to whistler waves can be obtained.

The characteristic size of a plasma bubble confined by a dipole magnetic field, defined as the distance at which the strength of the bubble magnetic field equals that of the ambient field, is determined by the strength of the dipole magnetic field that can be generated by the satellite. This, in turn, is limited by the volume, mass and power available to the magnet onboard the spacecraft. Since the dipole field decreases with the third power of the distance from the satellite, it is unpractical to confine plasma bubbles with size more than one order of magnitude larger than the size of the satellite.

In a different scheme,5,6_{the plasma is not confined within the magnetic field imposed by the magnet on the} satellite, but a plasma flow is generated by a large neutral particle outflow from the satellite. If the emitted gas becomes ionized and freezes the magnetic field in its outward flow, it is possible to obtain much larger magnetic bubbles. However for such an inflated bubble a “magnetic connection” problem arises: if the magnetic freezing condition is violated, e.g., by local reconnection processes, part of the inflated bubble will become separated from the satellite - a “tearing” process that is itself important to investigate - thus reducing the effective size of the bubble. Plasma thrusters can provide a natural source of plasma outflow, while the investigation of the interaction between the outflowing plasma and the ambient plasma (in the absence of external boundaries) can be important in predicting the long term performance of the thrusters themselves.

While traditional plasma measurements in space are usually aimed at measuring the undisturbed local properties of the ambient ionospheric plasma, the aim of the proposed active magnetic experiment in space is to excite and investigate the nonlinear collective dynamics of two counterstreaming plasmas. In practice, at least for un-inflated magnetic bubbles, this restricts the spacecraft operational altitude to the range 200 km < h < 2000 km, i.e. to low

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Earth orbits or to the low portions of strongly elongated orbits. For a satellite with size S ~ 1 m, under these conditions the following relations hold:

S > !D! "e, S! "i, S < de (1)

where !_Dis Debye length, !_e, !_i are the electron and ion mean gyroradii and d_e is the collisionless electron skin depth in the ambient plasma. The relative velocity between the satellite and the ambient plasma is around 7 km/s, which is larger than the ion thermal velocity, and smaller than the electron velocity. This ”infrathermal” regime is experimentally interesting, as it can give rise e.g. to Cherenkov emission of ion acoustic waves and, in a magnetized inhomogeneous plasma, to generation of drift-type waves and instabilities.

The bubble transit time, defined as the ratio between the bubble size and the relative satellite ambient-plasma velocity, is of the order of 10-3_{s. This is one order shorter than the ion gyro period. The transit time provides an} effective lower bound to the frequency of the perturbations that can be excited. Thus with a non-inflated bubble it is not possible to excite directly magneto-hydrodynamic phenomena effectively, since the frequency of the disturbances excited by the bubble turns out to be larger than the ion cyclotron frequency in the ambient plasma. Phenomena that can be marginally described within the magneto-hydrodynamic framework can be excited within the bubble in the portion close to the satellite where the ion gyroradii tend to become small but, as the bubble boundaries are approached, a transition must occur towards regimes of the EMHD type where the Hall term in Ohm’s law becomes relevant. This is a very interesting transition to explore experimentally, in particular in view of the role that has been recently attributed to the Hall currents in the evolution of magnetic field line reconnection in collisionless regimes.

III. The HT-100 Thruster

The HT-100 is a 100 W Hall thruster developed by Alta S.p.A. (Fig. 1). The thruster has been widely characterized both in single operation and in a two-thruster configuration.7-9 The HT-100 can run at power levels in the 50 to 200 W range, making it an ideal candidate for small spacecraft missions, usually very limited in power. At design point the HT-100 produces 4.5 mN thrust with 95 W power consumption, at a specific impulse of 950 s and a total thrust efficiency of 22%, with Xe propellant. Maximum thrust as high as 12 mN or specific impulse in excess of 1500 s can be achieved by throttling mass flow rate and adjusting the input power.

Figure 1. Left: a cluster of HT 100 thrusters during tests at Alta; right, a close-up of the two-thruster cluster.

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The thruster is quite simple and rugged in construction, featuring permanent magnets instead of the customary coils, and the auxiliary equipment (propellant flow controller, tank, power electronics) can be kept simple and relatively inexpensive. The main spacecraft-thruster integration issue are the management of the thermal load produced during prolonged operation of the thrusters and the interaction of the plasma beam with the solar panels. Both constraints were taken into account in the spacecraft preliminary design and played a role in the choice of the external shape and solar array arrangement.

IV. Mission profile

In order to limit cost, mission design is based on the assumption of launching the spacecraft as an auxiliary payload on an Ariane 5 flight. The spacecraft dimensions and mass (see next section) are compatible with accommodation on the Ariane 5 platform for auxiliary payloads ASAP 5,10_{in the category “mini-spacecraft” (120} kg to 300 kg). However, the mission profile can be easily adapted to different launch opportunities, even to non-GTO destinations, with little re-configuration of the spacecraft, mainly for thermal control aspects.

Launch to GTO allows the spacecraft to experience a wide range of different ambient plasma and ambient magnetic field conditions. The mission profile (Fig. 2) includes five coasting phases in different elliptical orbits at decreasing apogee and fixed perigee, starting from GTO, five low-thrust transfer phases, and a final quasi-circular LEO. The final orbit is such that de-orbiting is achieved by atmospheric braking and re-entry occurs within 25 years. With this mission profile, the spacecraft travels within the Van Allen radiation belts for a large fraction of the total mission time, allowing for extensive evaluation of the

effects of the environmental plasma on the solar panels. A simple experiment can be hosted to study the environmental effects on the electronics in the Van Allen belts, consisting of two identical memory chips with the same pattern of binary digits. One of the chips is exposed to the incoming radiation, while the other is protected with a metallic shield. Periodic comparison of the content of the two memories will provide indications on the occurrence of single event upsets or permanent damage in the onboard electronics under the prolonged exposure to the trapped radiation.

Plasma experiments are carried out during the coasting orbits. These are geosynchronous elliptical orbits with inclination i = 7 deg and perigee at 560 km. Apogee altitudes are chosen such as to yield orbital

periods that are integer fractions of the Earth rotation period; this may result in simplified scheduling of the communication operations with the ground. Of course, different apogee altitudes could be selected for other reasons, with minimal impact on the overall mission scheme.

The parameters for the final orbit were determined by assuming a spacecraft mass of 200 kg and a value of the ballistic coefficient compatible with the chosen spacecraft configuration. The orbit parameters are shown in Table 1. Velocity at apogee and perigee are also shown, giving an indication of the wide range of relative velocities between the artificial magnetosphere and the ambient plasma along the various phases of the mission.

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GTO Orbit 1 Orbit 2 Orbit 3 Orbit 4 Final Perigee radius, km 6938 6938 6938 6938 6938 6938 Apogee radius, km 42268 33690 26600 18656 9185 7053 Eccentricity 0.71 0.65 0.59 0.46 0.14 0.008 Orbital period, h 10.66 8 6 4 2 1.62 Max eclipse duration, h 1.52 1.35 1.18 0.96 0.53 0.51 Apogee velocity, m/s 1630 2010 2489 3403 6111 7610 Perigee velocity, m/s 9935 9761 9546 9151 8090 7486

Table 1. Initial, intermediate and final orbits.

At GTO and at each of the coasting orbits, i.e. during the non-propelled phases, active plasma experiments are executed by charging the superconducting magnet from the batteries and briefly activating the plasma source. The duration of each single measurement is a few seconds, with a repetition rate of several per minute, so that the thruster firings needed to produce the artificial plasma have a negligible outcome in terms of ∆V, with essentialy no effect on the orbital parameters. Assuming a total of 100 measurements taken at each of the five coasting orbits, and with an energy expenditure of about 10 Wh per experiment, the total energy cost per each phase is about 1 kWh, provided by onboard batteries. With a reasonable choice of the batteries, the charging time can be taken as equal to three times the orbital period at the last coasting orbit, i.e. 4.42 hours. With this choice, the total time needed to perform 100 measurements at each orbit can be calculated. the results are shown in Table 2. The total duration of the experiments is about 116 days.

GTO Orbit 1 Orbit 2 Orbit 3 Orbit 4

Number of orbits 48 66 92 145 300

Total duration, h 515 533 551 582 601

Experiments per orbit 2 1.5 1 0.7 1/3

Table 2. Duration of experiments at each orbit.

Orbital transfers between the coasting orbits are performed firing the thruster in the orbital anti-tangential direction, i.e. opposite to the velocity vector, to lower the apogee, while perigee is kept approximately constant. The spacecraft axis is always perpendicular to the orbital plane. A rough evaluation of the transfer duration was made neglecting the atmospheric braking, assuming that the thruster is operated at 4.5 mN and the spacecraft mass is 200 kg. The thruster is fired for half of the orbital period around perigee for the first three transfers and along all the orbit for the last two transfers (Fig. 3). The results are shown in Table 3. The total duration of the mission is 53 months, 7.5 of which are spent to reach the final de-orbiting altitude. Total propellant mass is 47 kg.

Figure 3. Thruster firing arcs, shown as shaded areas between the orbits.

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Transfer 1 Transfer 2 Transfer 3 Transfer 4 Transfer 5

Total duration, months 5.9 7.1 12.7 16 8.6

Propellant mass, kg 3.7 4.5 8 20 10.7

Table 3. Transfer duration and propellant consumption.

V. Magnet design

The size of the artificial magnetosphere depends on the magnetic field strength in the vicinity of the spacecraft. A reasonable compromise between a large enough artificial magnetosphere and a manageable magnet leads to the requirement of magnetic field intensity at a distance of 10 m from the spacecraft approximately equal to the field intensity at the Earth surface on the equator, B = 3 10-5 T. The magnetic dipole moment needed to produce the required magnetic field is µ = 3.105 A m-2. The magnet design shall be compatible with the specific constraints of a small spacecraft, i.e. compact configuration, low mass and limited power consumption. Conventional copper solenoids were ruled out due to high mass and power consumption. For the sake of simplicity, it was decided to avoid low-temperature superconductor magnets, needing a dedicated active refrigeration system to keep the magnet below the transition temperature. Therefore, the preferred choice is a solenoid made of high-temperature superconductor with a transition temperature high enough to be maintained with a passive thermal control system.

A variety of high-temperature superconductors are available today. Our reference for this study is BSCCO (Bi-Sr-Ca-Cu Oxide), with a transition temperature as high as 125 K and a critical current of 150 A at 77 K, much larger than the current needed to produce the required magnetic dipole. The magnet is made of 60 BSCCO coils, each with external radius 0.5 m and internal radius 0.47 m, manufactured by the Sumitomo Corp. The total length of the BSCCO wire is 5.17 km for a total mass of 43 kg.

The main issue in adopting a superconductor is the need to keep the magnet well below the critical temperature during operation. In our case, active cooling systems were ruled out due to complexity, energy consumption and cost. The solution was to design the spacecraft in order to allow passive thermal conditioning of the magnet. To this end, the spacecraft includes two bays: the upper bay host the magnet, the lower bay accommodates all the other subsystems, including thermally active parts like the thruster and the electronics. The interconnecting structural elements between the bays are designed to minimize thermal flux and the gap between the bays is equipped with multi-layer thermal blankets. Heat produced in the lower bay is radiated to space through a dedicated radiator. The magnet axis is tilted by 7 deg with respect to the spacecraft axis, so that it is normal to the plane of the ecliptic, in order to avoid direct irradiation by sunlight. The magnet is connected to a 1 m2_{radiator with solar absorptivity 0.077} and infrared emissivity 0.79, sitting on top of the spacecraft, whose function is to dispose of heat generated by the non-superconducting jucntions during magent charging and the heat concudcted to the magnet from the lower bay.

Taking in consideration all the heat input sources, including direct sunshine, albedo and infrared emission from the Earth, we can compute the equilibrium temperature of the magnet-radiator assembly as a function of the heat input to the magnet from the lower bay. The results, essentially independent on orbital altitude, are shown in Fig. 4. The magnet temperature stays below 125 K if the thermal power transferred from the lower bay is less than 11 W. Thus, the magnet can be kept at superconducting temperature if sufficient thermal de-coupling is provided between the two bays.

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Power, W

Temperature, K

Fig. 4. Magnet temperature as a function of internal heat input.

VI. Spacecraft configuration

The external dimensions of the spacecraft are constrained by the ASAP 5 interface requirements. For simplicity, body-mounted solar panels are used, with triple junction GA-As solar cells. Sufficient solar panel area (cross-section 1.7 m2_{) and ease of manufacturing can be obtained with a octagonal prism shape, with height 1.5 m and maximum} diameter 1.5 m, fitting almost completely the internal envelope allocated to a mini-spacecraft on ASAP 5. Spinning the spacecraft is not possible, due to the presence of the thruster, so the spacecraft will be 3-axis stabilized using momentum and reaction wheels. Miniature Xe resistojets or other simple thrusters are used for desaturation, sharing the same propellant tank as the main thruster. The cumulated effect on spacecraft attitude of the torque generated by activation of the superconducting magnet is essentially neglibilble, due to the very short duration of the activations.

The spacecraft configuration is shown in Fig. 5 and Fig. 6. The Xe tank position is chosen to minimize the center of mass offset from the geometric center of the spacecraft. Two thrusters are mounted for redundancy. The battery is made of 27 Saft-VES 100 Li-Ion cells, with a total energy storage capacity of about 920 Wh. Due to the limited amount of data collected to be transferred to the ground, there is no need for continuous coverage. Data are stored onboard and released to the ground station periodically, allowing the communication link to be kept to a minumum of bandwidth and to the periods of best visibility from the ground. The total spacecraft mass is 202 kg and the total power installed is about 350 W at end of life. The spacecraft configuration allows for ample mass margins with respect to the allowed mass for an ASAP 5 payload.

The plasma diagnostics suite include a compact three axis magnetometer. A similar instrument, built by Surrey Satellite Technology Limited, has a mass of 295 g and a power consumption of 0.15 W. Six sensors will be mounted at the end of three 6 m telescopic booms, to take measurements into the plasma bubble around the spacecraft. Other instrumentations include a mass spectrometer, a set of Langmuir probes, and a dedicated Hall thruster diagnostic package to measure the plume properties in close proximity to the thruster and to monitor the visual aspect of the plume.

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Fig. 5. Spacecraft external dimensions.

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VII. Conclusion

The preliminary design presented shows that a small spacecraft for active magnetic plasma experiments and electric propulsion flight demonstration can be realized within the constraints of an Ariane 5 auxiliary payload. The orbital transfer strategy was not studied in detail and there is ample room for optimization, possibly leading to further reduction in launch mass. The total mission duration, presently evaluated at about 4.5 years, can be probably reduced by carefully selecting a combination of higher thrust and intermediate orbital altitudes. While cost was not explicitly addressed, the spacecraft is simple enough to justify the assumption that it can be built and flown with a very limited budget.

Acknowledgments

Many useful discussions on plasma experiments in space took place in the frame of the ESA study “Plasma Laboratory in Space”, initiated and monitored by Mr. Alain Hilgers, D/TOS-EES.

References

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