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Ion Thruster ・Electrospray ・Microwave Engine ・Hollow Cathode Thruster Pulsed Plasma Thruster Pulsed Plasma Thruster

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Micro-satellite propulsion KOIZUMI Hiroyuki (小泉 宏之) Graduate School of Frontier Sciences, Department of Advanced Energy & Department of Aeronautics and Astronautics (基盤科学研究系 先端エネルギー工学専攻，工学系航空宇宙工学専攻兼担) Propulsion and Energy Systems; Dec 21st (2015) Huge cost Compensation Failure is by money ant time never excused No growth of Conservative technology and long-time and human design Small satellite ・Welcome, new service ・Quick technology cycle ・Quick human-resource development Innovation Small satellite Small satellites 100 – 500 kg Microsatellites 10 – 100 kg Nanosatellites 1 – 10 kg Cubesat Unit: 10 x 10 x 10 cm3 Small propulsion; Resource Mass Typically 10 – 20% of the satellite mass Typically If parallel to the mission: < 20% Power If dependent from the mission: < 80% Typical power generation： 1–3 W/kg （e.g 10 kg satellite → 10–30 W） Small propulsion; Resource Don’t lose the merit of small sat. Budget Typical budget for propulsion would be 10-20% Law Strict limitation for High-pressure gas, explosives, toxic materials. ΔV LEO; 100km altitude change 60 m/s LEO; 1 degree orbit inclination 130 m/s LEO 300km, 1-year drag compensation （50kg, 50cm x 50cm, CD1.4） 50 m/s GEO; NSSK (1year) 50 m/s Lunar orbit from GTO 2000 m/s Small Chemical Propulsion ・Cold-gas thruster ・Gas-liquid equilibrium thruster ・mono-propellant propulsion using hydrogen peroxide ・Micro-solid array ・Solid microthruster Cold-gas thruster Cold-gas thruster；Principle Blow-down from a high pressure tank Specific impulse by gas and temperature Heater Resisto-jet thruster M/kg Density Isp/s * mol-1 (24 MPa) Isp CF c / g He 4 0.04 180 N2 28 0.28 80 Xe 131 2.74 31 Propulsion and Energy Systems; Dec 21st (2015) Cold-gas thruster Merit The simplest Not for Demerit Low specific impulse high ΔV Large high-pressure system Heritage Already used in micro-/nano- satellites Propulsion and Energy Systems; Dec 21st (2015) 1) SSTL; Gas Propulsion System Propellant: 500g Xe or 176g N2 Thrust: 20 – 50 mN Isp : 42 s Xe or 100 s N2 Total impulse: 380 Ns 50 kg S/C ΔV = 6.4 m/s Dry mass: 6.7 kg Dimensions: 400 x 254 x 215 mm3 Propulsion and Energy Systems; Dec 21st (2015) Cold-gas thruster For ΔV of ～100 m/s? SSTL; Microsatellite Gas Propulsion System Propellant: 12 kg Xe Thrust: 18 mN Isp : 48 s Total impulse: 5644 Ns 50 kg S/C ΔV = 110 m/s Dry mass: 7.3 kg Propulsion and Energy Systems; Dec 21st (2015) Gas-liquid equilibrium thruster Gas-liquid equilibrium thruster One of the gold-gas thrusters Propellant storage in liquid phase Specific impulse is the same as cold-gas thruster Merit No high pressure system →Reducing system volume Needing a heater for vaporization Demerit Needing a gas/liquid separation device Propulsion and Energy Systems; Dec 21st (2015) 1) Thruster for IKAROS Propellant: 20 kg HFC-134a Dry mass: 約20 kg Thrust：400 mN Isp：40 s Total impulse: 7000 Ns Molecular mass：102 g/mol Vapor pressure: 0.57 MPa Liquid density: 1225 kg/m3 Metal foam for the separation using surface tension and temperature gradient No high-pressure system Propulsion and Energy Systems; Dec 21st (2015) Mono-propellant thruster Hydrogen peroxide thruster Hydrogen peroxide + catalyst →Isp: 80 s Propellant feeding →Bladder + High-pressure gas Propellant tank×２＋Pressure tank ×１ Propulsion and Energy Systems; Dec 21st (2015) Thruster for Hodoyoshi-1/3 Thruster for Hodoyosh-3 Power 3.4W Mass 5.8kg, including 2 kg propellant Size 265×270×85 mm3 Thrust 350 mN Isp >80 s ΔV 30m/s Propulsion and Energy Systems; Dec 21st (2015) Hydrogen peroxide thruster Merit High Isp as CP Demerit Toxity of H2O2 Rejected for H2A secondary payload (UNIFORM1) Needing on-site charge Heritage Hodoyoshi-1/3 Propulsion and Energy Systems; Dec 21st (2015) Solid-propellant thruster Micro-solid array Digital control of micro-solid propellant MEMS fabrication Micro Electro Mechanical Systems using semiconductor device fabrication technologies Propulsion and Energy Systems; Dec 21st (2015) マイクロソリッドアレイ 長所 Ultra miniaturization without gas system Easy fabrication of arrayed structure 短所 Small ΔV e.g. 30 μNs by 1shot 300 mNs by 100x100 array ΔV = 0.3 m/s @ 1kg S/C Propulsion and Energy Systems; Dec 21st (2015) Laser ignition, solid microthruster Laser ignition of multiple pellets Boron & potassium nitrate Propulsion and Energy Systems; Dec 21st (2015) Laser ignition, solid microthruster Merit Cubesat compatible 10 m/s class ΔV Demerit ΔV less than 100 m/s Needing rotating mechanism Isp:150s, 60 shots, 300-cc ΔV : 30 m/s for a 3kg Cube-sat Propulsion and Energy Systems; Dec 21st (2015) Small Electric Propulsion ・Pulsed plasma thruster ・Vacuum arc thruster ・Arc-jet thruster ・Ion thruster ・Electrospray ・Microwave engine ・Hollow cathode thruster Pulsed Plasma Thruster Pulsed plasma thruster Pulsed discharge of capacitor energy Solid propellant (not explosive), passive feed Electromagnetic acceleration Propulsion and Energy Systems; Dec 21st (2015) パルス型プラズマスラスタ；特徴 Merit Simple feeding mechanism High specific impulse（500-1000 s） Demerit Geometric limitation of propellant and ΔV EMI (electro magnetic interference) Component life time of 1 billion shots Heritage Several verifications Propulsion and Energy Systems; Dec 21st (2015) Thruster on EO-1 Spacecraft EO-1 by NASA Launch: 2000 Mass: 570 kg Propulsion Impulse bit: 0.86 mNs Isp：1370 s Power：70 W (1Hz) Total mass: 4.95 kg Propellant: 0.14 kg ←!! Total impulse: 460 Ns →ΔV = 0.81 m/s Propulsion and Energy Systems; Dec 21st (2015) Concept by Univ. Surrey for 3U-cubesat Propulsion module: ¼U Pulsed power module: ¼U Mass 0.34 kg Power 1.5 Volume 480 cm3 Isp 320 s Propellant 1.1 g ΔV for 4.5 kg 2.7 m/s Propulsion and Energy Systems; Dec 21st (2015) PROITERES by OIT (大阪工大) PROITERES Launch by PSLV (Indian rocket) in 2012 10 kg, 30x30x30 cm3, 10 W Coaxial pulse plasma thruster Electrothermal acceleration, no feeding system Wet mass: 2.0 kg (Head:0.3kg, Cap.:0.2kg PPU:1.0kg, Cables:0.5 kg) Total impulse: 5 Ns（ΔV 0.5 m/s） Propulsion and Energy Systems; Dec 21st (2015) Pulsed plasma thruster; Summary • Compatible from cubesat to 100 kg S/C • High technical maturity • Few actual usages, so far • ΔV limitation • Electromagnetic interference Propulsion and Energy Systems; Dec 21st (2015) Ion thruster Ion thruster Electrostatic acceleration of ions Inevitable electron emission Neutral Ion Plasma source Ion acceleration Electron Propellant Power electron emission Propulsion and Energy Systems; Dec 21st (2015) Ion acceleration Example of the ion beam Outside Beam Inside trajectory Ion beam Electrostatic potential Appropriately-designed grids system converges the ion beam Propulsion and Energy Systems; Dec 21st (2015) Types of the plasma generators • Direct current (DC) electron discharge • Radio frequency (RF) discharge • Microwave discharge Propulsion and Energy Systems; Dec 21st (2015) DC-discharge ion thruster Gas injection magnetic field Electron emission 30 V Anode Propulsion and Energy Systems; Dec 21st (2015) RF-discharge ion thruster RF coil Gas injection ICP Induced magnetic field Insulating body Induced current Propulsion and Energy Systems; Dec 21st (2015) Microwave discharge ion thruster Gas injection magnetic field Microwave emission Coaxial cable Antenna or waveguide Electron cyclotron resonance 푓microwave = 푓cyclotoron Propulsion and Energy Systems; Dec 21st (2015) Ion thruster Merit High Isp (1000-3000 s) High ΔV mission Storage in a tank Demerit High-pressure tank and feeding system Complicated = a number of parts Heritage The most heritage in standard-size S/C Small ion thruster: Hodoyoshi-4, PROCYON Propulsion and Energy Systems; Dec 21st (2015) μRIT by Univ. Giessen The smallest thruster in RIT series RF plasma Developed for LISA RIT for (four thruster heads) Dry mass: 14 kg Propellant: 0.4 kg Xe Thrust: 7 – 100 μN Power: 60-86 W (PCU: 26 W) Max. thrust 600 μN Propulsion and Energy Systems; Dec 21st (2015) Research by JPL & UCLA DC-discharge, 30 mm of beam diameter Thrust：3.0 mN Specific impulse：1700 – 3200 s Ion production cost: 400 – 600 V/ion *not included the neutralizer DC-discharge = two cathodes Inside for plasma Outside for neutralizer Propulsion and Energy Systems; Dec 21st (2015) Research by KU (九州大学) Microwave discharge (μ10’s neutralizer based plsma source) Thrust 790 μN Isp 4100 s Microwave power 8 W *not included the neutralizer Beam power 20 W システムには45W必要 (8/0.4 + 20/0.8) Propulsion and Energy Systems; Dec 21st (2015) MIPS, flight model MIPS Flight Model Performance Mass: 8.1 kg (incl. 0.9 kg Xe) summary Volume: 39 x 26 x 15 cm3 Power: 27 W Thrust: 210 μN Isp: 740 s Delta V: 140 m/s (50 kg S/C) Propulsion and Energy Systems; Dec 21st (2015) A small secondary payload of HAYABUSA-2 Earth gravity assist Flyby observation of a small asteroid PROCYON needs 1. Wheel unloading by RCS Multiple thrusters 2. High ΔV for orbit transfer Electric propulsion 3. Trajectory correction maneuver Chemical propulsion ©Google MIPS Miniature Ion Propulsion System Controller Control Xenon-gas Gas system Ion thruster Power supplies Microwave High voltages I-COUPS Ion thruster and COld-gas thruster Unified Propulsion System Controller Cold-gas Control thrusters Xenon-gas Gas system Ion thruster Power supplies Microwave High voltages Cold-gas thruster Thrust (single) 22 mN Specific impulse 24 s Number of thrusters 8 7 W Power consumption (2 thrusters) Cold-gas thruster Thruster valve Sharing gas system is superior to increasing Isp Xenon 2.57 kg Gas system1 (Tank) 2.01 kg Gas system Cold-gas 4.5 kg !! thrusters 0.41 kg Gas system2 (others) Power supplies 2.25 kg 1.31 kg Controller Ion thruster 0.95 kg 0.38 kg I-COUPS; ITU全運転(5分平均プロット ) 論文用 from Dec 28 (2014) to Mar 12 (2015); 運転判定条件:SPS-I (mA)>1.00 250 200 223 h 150 100 50 Total operating time/hours operating Total 0 Accumulated operation time/hours operation Accumulated The miniature propulsion system of PROCYON has operated for more than 6 months, as the first interplanetary micropropulsion.
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  • Cathode Arc Thruster

    Cathode Arc Thruster

    Two-Stage Micro-Propulsion System Based on Micro- Cathode Arc Thruster IEPC-2017-472 Presented at the 35th International Electric Propulsion Conference Georgia Institute of Technology • Atlanta, Georgia • USA October 8 – 12, 2017 Jonathan Kolbeck1 and Michael Keidar2 The George Washington University, Washington, DC,20052, USA The Micro-Cathode Arc Thruster (CAT) developed at The George Washington University is one of the two electric micro-propulsion devices that have been flown on CubeSats to date. We report the development of a two-stage propulsion system based on this technology, which will enable interplanetary CubeSat missions and extended missions in low earth orbit. The device uses a modified version of the CAT system for plasma generation, and an accelerating second stage to accelerate the plasma. I. Introduction mall satellites are a growing market, especially after the introduction of the CubeSat standard in 1999. These S standardized satellites enable small organizations, industry, and educational institutions to launch small, affordable satellites into space. For many countries, such as Slovakia1, Hungary2, Switzerland3, Costa Rica4, and others, CubeSats represent their first means of access to space. Per SpaceWorks’ 2017 forecast (Figure 1), more than 400 satellites in the 1-50 kg class were launched between 2014 and 2016. The market is expected to grow to at least 200 new small satellites per year after 2017. In the last few years, missions have increased in complexity and the mission scopes have broadened. The first CubeSat