Satoshi Hosoda

Development of Multi-Utility Spacecraft Charging Analysis Tool (MUSCAT)

A new numerical software package to analyze spacecraft charging, named ldquomulti-utility spacecraft charging analysis toolrdquo (MUSCAT), has been developed. MUSCAT consists of an integrated graphical user interface tool called ldquoVineyardrdquo and the solver. Vineyard enables satellite engineers to compute spacecraft charging with little knowledge of the numerical calculations. Functions include 3-D satellite modeling, parameter input such as material and orbit environment, data transfer, and visualization of numerical results. Fundamental physical processes of charged-particle-surface interaction are included in the solver. These functions enable MUSCAT to analyze spacecraft charging at geostationary orbit, low Earth orbit, and polar Earth orbit (PEO). The numerical solver code is parallelized for high-speed computation, and the algorithm is optimized to achieve analysis of large-scale PEO satellite in the design phase. Variable time steps are also used to calculate the rapid change of the spacecraft body potential and the gradual change of the differential voltage in a single simulation with a practical number of iterations. In this paper, the functionality, algorithms, and simulation examples of MUSCAT are presented.

Electrostatic Discharge Ground Test of a Polar Orbit Satellite Solar Panel

The Advanced Land Observing Satellite that was launched by the Japan Aerospace Exploration Agency in 2006 carries a large solar-array paddle of 22 times 3 m in polar orbit. The wake side of the solar paddle can be charged to a negative value exceeding -1000 V. Numerical simulations and laboratory experiments are carried out to investigate charging and arcing phenomena on the backside of the solar paddle that has exposed bypass diode boards and silver-Teflon thermal film. Possibility of secondary arc and surge voltage between hot and return ends of power circuit has been investigated. The authors irradiate solar-panel coupons with an electron beam to simulate charging situation near the North Pole. Surface flashover is observed once the insulator potential exceeds -7000 V. The authors have also tested charging situation near the South Pole where carbon fiber-reinforced plastics surface facing dense ionospheric plasma may arc easily once a satellite potential drops to -80 or -90 V. The solar-paddle design has been modified to increase the safety margin against arcing, and tests have been carried out to identify the strength limit under extremely harsh test environment

Charge Neutralization via Arcing on a Large Solar Array in the GEO Plasma Environment

The purpose of this paper is to investigate the arc plasma propagation phenomena in a geostationary-orbit plasma environment. Electrostatic discharge tests were carried out using a large solar-array coupon (400 times 400 mm). The coupon consisted of 50 Si cells. The surface potential of the solar-array coupon immediately after arc inception was measured. The area neutralized due to arc plasma propagation differed in each arc. Some arcs neutralized a large area of solar cells even though the ratio of neutralization was not 100% everywhere. As the distance from the arc site increased, less charge was generally neutralized. The arc plasma propagation velocity was of the order of 104 m/s; but, it differed widely in each arc

Degradation of High-Voltage Solar Array Due to Arcing in Plasma Environment

A degradation test for a solar array coupon against electrostatic discharge was performed under a simulated lowEarth-orbit environment as part of research project to develop the next-generation 400-V high-voltage solar array technology. All tests were performed in a vacuum chamber with a plasma source. An inductance‐capacitance‐ resistance circuit was used to simulate the arc current that would flow by collecting electric charge stored on cover glasses. Arcs were repeated until the solar array coupon showed degradation of electrical output. The locations, current waveform, and voltage waveforms of all the arcs during the tests were recorded. The electrical output of the coupon was measured without opening the vacuum chamber. The arc that damaged a solar cell was identified; the cell was damaged by only one arc, which occurred at the edge of the cell.

Development of 400 V Solar Array Technology for Low Earth Orbit Plasma Environment

To realize a 400 V operation in low Earth orbit (LEO), problems of arcing caused by interaction between spacecraft and surrounding LEO plasma must be overcome. This paper is a summary report of the laboratory tests carried out to develop a 400 V solar array technology. Among various designs tested, a design of covering a solar array surface with transparent film, called film coupon, was the most promising mitigation method to prevent arc inception. The authors carried out various tests on the film coupons considering a realistic situation encountered in orbit. The coupon biased to -400 V in LEO-like plasma had no arc for more than 25 h. Other tests involved UV exposure, atomic-oxygen exposure, thermal cycling, and debris impact. Conductive substrate made of carbon fiber reinforced plastic suffered many arcs at -400 V. Sustained arc between a solar cell and the substrate was also observed upon a simulated debris impact. Therefore, the use of flexible substrate is adequate for 400 V solar array in LEO environment. To avoid the snapover effect near the positive end of the array circuit, only the negative part of the array circuit exceeding the arc-inception threshold should be covered by film, or an electron collector should be deployed

Hayabusa's Way Back to Earth by Microwave Discharge Ion Engines

The cathode-less electron cyclotron resonance ion engines, μ10, propelled the Hayabusa asteroid explorer, launched in May 2003, which is focused on demonstrating the technology needed for a sample return from an asteroid, using electric propulsion, optical navigation, material sampling in a zero gravity field, and direct re-entry from a heliocentric orbit. It rendezvoused with the asteroid Itokawa after a two year deep space flight with a delta-V of 1.4 km/s, 22 kg of xenon propellant consumption and 25800 hours of total accumulated operational time of all the four ion engines added up. Though it succeeded in landing on the asteroid on November 2005, the spacecraft was seriously damaged. This delayed the Earth return in 2010 from the original plan in 2007. Reconstruction on the operational scheme using remaining functions and newly uploaded control logic made Hayabusa leave for Earth in April 2007. After a coasting period of 2008, the ion propulsion was reignited in February 2009. Although most of the neutralizers were degraded and unable to be used by fall of 2009, a combination of an ion source and its neighboring neutralizer has been successfully operated for the last 3230 hours including a series of final trajectory correction maneuvers. Before reentry, the total accumulated operational time reached 39637 hours consuming a total of 47 kg Xenon propellant. Total duration of powered spaceflight is 25590 hours which provided a delta-V of 2.2 km/s and a total impulse of 1 MN·s, approximately. Finally, the spacecraft returned to Earth. Its reentry capsule, which may contain samples from asteroid Itokawa, was retrieved from the Australian outback according to plan .

Status of Microwave Discharge Ion Engines on Hayabusa Spacecraft

[Abstract] The μ10 cathode-less electron cyclotron resonance ion engines made the Hayabusa spacecraft rendezvous with the asteroid Itokawa in 2005. Though the spacecraft was seriously damaged after the successful soft-landing and lift-off, the xenon cold gas jets from the ion engines rescued the Hayabusa. New attitude stabilization method using a single reaction wheel, the ion beam jets, and the solar pressure was established and enabled the homeward journey aiming the Earth return on 2010. The total accumulated operational time of the ion engines reaches 28,000 hours at the end of May 2007.

Sustained Arc Between Primary Power Cables of a Satellite

We investigated the power loss due to the sustained arc between primary satellite power cables. If the multi layer insulator (MLI) film on a satellite is electrically floating, energetic electrons in space will charge this film. We carried out an ESD test on cables with cracks and wrapped with this film. When the negative voltage on the MLI exceeded 600 V, a trigger arc discharge occurred between the MLI and the cables. Subsequently, a secondary arc electric discharge occurred between the cables themselves. After several discharges, this secondary arc caused sustained arc which burned out the cables. The heat caused by arc tracking between the hot and return cables made them burn out. If this phenomenon had occurred in space, the satellite would have suffered great damage.

Verification of Multi-Utility Spacecraft Charging Analysis Tool (MUSCAT) via laboratory test

Multi-utility Spacecraft Charging Analysis Tool (MUSCAT), a spacecraft charging analysis software, has been developed as a joint work of JAXA and KIT. Experiments for the fundamental code validation were carried out at the plasma chamber of LaSEINE in KIT to show accuracy of the solver. We evaluated that the test section in the chamber with respect to the plasma environment by measuring two-dimensional plasma distribution and plasma drift velocity. A cube area of 400mm on a side whose center located at the 550mm downstream from plasma source can be considered as the test section with no plasma flow. The averaged plasma density, temperature and plasma potential within this test section were 3±2x10 12 m -3 , 2±1eV and 10±5V, respectively. The length of test section 400mm corresponds to about 67λ D . Spatial distribution of electric potential and IV characteristic curve were measured with an emissive probe and the Langmuir probe whose electrode were cubic in shape to adjust the rectangular numerical domain of MUSCAT. Comparing those experimental results with the numerical ones, both had good agreements. These results show that the physical functions of MUSCAT simulate charging processes quite well. Also, numerical model of the cell-side of solar array paddle was obtained. Conductor patches whose size is the quarter of total amount of the interconnector exposed area put on the coverglass can simulate the cell-side of a real solar array with respect to current collection.

Recent Progress of Development of Multi -Utility Spacecraft Charging Analysis Tool (MUSCAT)

A new numerical software package to analyze spacecraft charging, named ldquomulti-utility spacecraft charging analysis toolrdquo (MUSCAT), has been developed. MUSCAT consists of an integrated graphical user interface tool called ldquoVineyardrdquo and the solver. Vineyard enables satellite engineers to compute spacecraft charging with little knowledge of the numerical calculations. Functions include 3-D satellite modeling, parameter input such as material and orbit environment, data transfer, and visualization of numerical results. Fundamental physical processes of charged-particle-surface interaction are included in the solver. These functions enable MUSCAT to analyze spacecraft charging at geostationary orbit, low Earth orbit, and polar Earth orbit (PEO). The numerical solver code is parallelized for high-speed computation, and the algorithm is optimized to achieve analysis of large-scale PEO satellite in the design phase. Variable time steps are also used to calculate the rapid change of the spacecraft body potential and the gradual change of the differential voltage in a single simulation with a practical number of iterations. In this paper, the functionality, algorithms, and simulation examples of MUSCAT are presented.