Spacecraft Electrical Propulsion Technology Overview - PowerPoint Presentation

1. Spacecraft Electrical Propulsion Technology Overview

2. Electric Propulsion DefinitionsElectric propulsion (EP) : any propulsion system which uses electrical power to provide propellant acceleration, and hence thrust.Excludes systems which only use power for thermal control, valve actuation, etc.Implicitly includes “hybrid” systems, which use electrical power to accelerate combustion products of chemical thrustersSolar electric propulsion (SEP) uses solar energy (from solar arrays) as power sourceAlternatives are radio-isotope electric propulsion(REP) and nuclear electric propulsion (NEP) Comparatively low level of development, and not covered in this course in any further detail

3. EP System ArchitectureFor all EP, the following elements are required:Propellant storage and feed systemElectrical power supply and power conditioningPower conditioning is usually considered to be part of the EP system; the power supply is part of the overall spacecraft power system architectureThe thruster, which converts the electrical power into propellant kinetic energy, providing thrustMany EP systems also use thrust pointing mechanisms (gimbals), to ensure thrust vector passes through spacecraft centre of mass

4. Chemical vs. Electric PropulsionThrust levels CPS - 0.5N to >400NShort impulsive firingsμN / mN possible with cold gas EPS - Typically in the μN / mN range Longer firing times to provide same ΔV Impulsive firings (high thrust, short duration) not feasible Specific impulseCPS - up to ~320sEPS - current state-of-the-art ~1500 to 4500s for medium to large thrustersPropellant mass saving to be offset against increased hardware massMay require additional solar array, batteries, etc.Thrust variabilityCPS - thrust level set by thruster inlet pressuresCan only change average thrust level by pulsed operationEPS - can vary thrust level (and sometimes Isp) in orbit, to meet variable mission needs

5. Chemical vs. Electric PropulsionFine thrust controlEPS - can provide very low thrust levels (μN level) with fine thrust resolution and low noiseCPS - cannot achieve less than around 50 to 100mN with reasonable specific impulse μN level with proportional cold gas, which has specific impulse of around 60 to 70s maximumPowerCPS - low power for valves and heaters onlyEPS - Power increases (roughly) linearly with thrust and specific impulseExample 1: Eurostar 3000 station-keeping requires around 80mN and 1500s Isp corresponding power requirement is around 1.4kWExample 2: BepiColombo requires up to 290mN and 4200s Isp; corresponding power requirement is around 10kWCostEPS typically more expensive than CPS

6. Comparison of Propulsion Systems

7. GTO to GEO Transfer

8. EP EfficiencyEP translates electrical energy into kinetic energy in the propellantSimple power balance equation lead to definition of overall EP system efficiency η,which defines relationship between power, thrust and specific impulse: (P is the input power)Current Efficiency:Thruster efficiencies ~ 45 to 65%EP system efficiencies (including power conditioner losses) ~ 40 to 60%Some technologies have efficiencies above or below this rangeEach technology/product has optimum operating range, with maximum efficiency

9. EP Efficiency

10. Categories of EP Systems Electrothermal:• Power heats propellant• Expansion through a nozzle (same as for a chemical thrusterElectrostatic:• Propellant is ionized• Accelerated usingelectrostatic fieldsElectromagnetic:• Propellant is ionized• Accelerated byinteraction of dischargecurrent with magneticfield (JxB Lorentz force)

11. Electrothermal EP Systems

12. Electrothermal Thruster - Operations and PerformanceThrust generated by heating and acceleration of propellant through a nozzle Reminder:Hence specific impulse maximized for propellants with hig specific heat and high temperatureHigh cp => low molecular weightLow molecular weight gaseous propellants are more difficult to store (tank mass and volume scales approximately with inverse of molecular weight)Trade of propellant mass gains against tank mass penalty

13. Electrothermal - ResistojetsPropellant heated by electrical coils before expansionWide range of units, with various propellantsSSTL low power multi-propellant unit(rated up to 100mN)Temperatures typically ~800 to 1500KPerformance range:Specific impulse 45s (Xenon) to >800s (hydrogen)10mN to >600mNCorresponding power range 30W to 3kWThruster efficiencies 35% to 85%

14. Electrothermal – EHT ResistojetsEHT - Electrothermal hydrazine thrustersHeats the decomposition products of a monopropellant hydrazine thruster to boost the specific impulsePerformance summary:Specific impulse 290 to 300s150 to 800mNSpecific power typically 1 to 2 W/mNEfficiencies apparently >100%; due to contribution from chemical energy released during decomposition

15. Electrothermal – ArcjetsElectrical discharge in the propellant flow adds energy to the propellantUse either an inert propellant, or are combined with monopropellant hydrazine thrusterHydrazine arcjets used on Lockheed Martin telecommunications platformsPerformance summary:Aerojet (USA) thrusters: 213 to 258mN, 500 to 615s Isp, up to 2kW (7 to 8 W/mN)Corresponding thruster developments efficiency 35% to 40%European aimed at lower thrust and power, but higher specific impulse levels up to 800sLife limited by erosion of electrodes

16. Electrostatic EP Systems

17. Electrostatic Thruster - Operations and PerformanceThrust generated by acceleration of positive ions through an electrostatic fieldSpecific impulse primarily driven by the accelerating voltage Vq/m is the ion charge/mass ratioSpecific impulse maximized for small ions for a given voltage, BUT …..Ionization losses are greatest for small ionsConsequently thruster efficiencies are lowest with low atomic mass propellants this outweighs the possible gains in maximizing the specific impulseHigh atomic mass propellants are preferredMore complex molecules avoided as they can dissociate during ionizationMultiple charged ions are also avoided (as far as is possible), as their production and acceleration is less efficient than that for singly charged ions

18. Electrostatic Thruster - Operations and PerformanceXenon is the “industry standard” propellant for electrostatic systems High atomic mass (~131) Low 1st ionization potentialThrust is given by the ion current Ibeam for a given specific impulse Ion beam emitted by thruster requires neutralization, to avoid spacecraft chargingPower and flow rate to neutralizer have to be taken into account for overall thruster specific impulse and efficiency calculations

19. Electrostatic - Hall EffectIons are produced within an annular discharge chamber by electron bombardmentElectrons provided by external cathodePropellant flowed into discharge chamber through anode at rear of discharge chamberMagnets trap the electrons, increasing ionization rateIons accelerated by electric field between anode and spaceCathodes also provide electrons for plume neutralizationHETs typically operate at 300 to 350VHigher voltages are used for higher specific impulse optionsBeam divergence relatively high (45° half cone angle)

20. Electrostatic - Hall EffectOriginally developed in RussiaSPT-100 thruster adopted by Airbus, TAS and SS Loral for their telecommunications platforms; current “industry standard”Operational on many Russian satellites; also operating on 8 Airbus spacecraft (including Intelsat 10, Inmarsat 4, Kasat and Yahsat) Nominal operating point 83mN, 1500s Isp, 1.35kW thrusterSPT-140D now baselined for electric orbit raising (EOR) missionsNominal operating point 280mN, 1800s Isp, 4.5kW thrusterCeramic discharge chamber (BN/SiO2 mix)Subject to wear from ion bombardment; limits life to ~ 8500 to 10000 hoursSlight loss in thrust and specific impulse as ceramic erodesWide range of thrusters in development worldwide: 0.5 mN to 1NModerate throttling range from a single thruster (up to 3:1)895 to 3000s Isp10W to 20kW thruster input powerSpecific power levels 15 to 25 W/mNCorresponding thruster efficiencies ~20% for small HETs, up to >60% for large thrusters

21. Electrostatic - Hall EffectEuropean thrusters:Safran: PPS1350Almost identical to SPT-100 (slightly increased Isp, up to 1700s)Flight Heritage: Flown on Smart-1Adopted by Thales for Alphasat (first Alphabus spacecraft)Safran:PPS5000:170 to 200mN, 2000 to 2200s Isp, 5kW input powerLimited endurance tests performedSitael:HT-100:Miniature HET (6 to 18mN range)TRL6 (Technology Readiness Level) ; in qualification, should be flown in 2018 HT-400 (20 to 50 mN) and HT-5K (150 to 350mN)Both at ~TRL5HT-20K (1N) at lower development level

22. Electrostatic - Ion2 main types of ion thruster, depending on the ionisation technique used:Radio-Frequency (RF) Discharge:Xenon flowed into discharge chamberRF-current applied to coil; results in a primary axial magnetic field being induced inside the ioniser, which generates a secondary circular electric field (E) Free electrons (from neutraliser) ionise the propellantOnce ionisation process is initially triggered the process is self-sustaining, with all electrons required for steady state

23. Xenon ions accelerated by potential difference between screen and accelerator grids, producing thrustNeutralizer provides electrons to prevent spacecraft chargingTypically operate at beam voltages of 1 to 2 kV or higher, depending on the specific impulse requiredAcceleration grid is normally negatively biased (a few hundred volts); this prevents external electrons from being attracted into the discharge chamberAcceleration grid erodes due to ion impingementIon optics design avoids direct impingement; erosion caused by charge exchange between fast ions and slow neutrals in vicinity of gridsGrid life normally ~ >25000 hoursElectrostatic - Ion

24. High throttling range (~ 2:1 to 5:1)Operation of a single thruster over a range of beam voltages (i.e. variable Isp) is possible but difficultGrid ion optics are optimized for a given voltage (and hence Isp)Ion thruster efficiencies ~35% for small thrusters up to ~80% for large thrusters Low beam divergence (typically <15° half cone angle)Electrostatic - Ion

25. European thrusters: RIT series developed by ASL RF ionisationRIT-10; flown on Eureca (experiment) and Artemis - 15 mN, 3325s, 460WRIT-22; under development; 75 to 150mN, 4620s, 4720W at 150mNFurther development with RIT-2X (up to 200mN, 4300s, 5.8kW) RIT-μX under development - up to 500μN, 3000s, 50W T5/6 series developed by QinetiQKauffman typeT5 flown on Artemis - 18mN, 3200s, 500WT5 also flow on GOCE with fine thrust control down to ~1mNT6 under qualification; 75 to 150mN, 4200s, 4800W at 150mNAlso some developments in ItalyElectrostatic - Ion

26. Ions extracted from liquid metal surface by evaporation, and accelerated by electrostatic fieldWith strong electric field, liquid metal surface deforms into Taylor conesShape is determined by balance between electrostatic and surface tension forcesAbove ~ 109 V/m, ions ripped from cone; these ions are then accelerated by the electric fieldNeedle effectively forms anode, with neutralizer as cathodePropellant reservoir embedded within overall thruster design; propellant feed is by capillary effectsElectrostatic – FEEP (Field Emission Electric Propulsion)

27. Single needle can provide thrust levels up to around 10μNAbove this level ion current is so large that emission geometry is disrupted, and droplets are emitted at lower efficiency2 main developments aimed at increasing thrust level up to >100μN: Clustering of needles, developed ARC-Seibersdorf (Austria)Slit thruster, developed by Alta (Italy)For the slit thruster, a series of Taylor cones is generated along the length of the slitUse of metallic propellants raises unique concerns regarding contamination of spacecraftElectrostatic – FEEP (Field Emission Electric Propulsion)

28. Overall voltage levels of around 7 to 9 kVElectrostatic – FEEP (Field Emission Electric Propulsion)

29. Balance of surface tension and electrostatic forces create microscopic droplets of size typically 10 to 100 nm diameterSemi-conductive liquid usedDroplets are extracted and accelerated by either:The same applied voltage, orSeparate extractor and accelerator voltagesDC operation - needle effectively forms anode, with neutralizer as cathodeAlternatively, needles can separately biased +ve and –ve within an array, so no neutraliser neededMain development is in USAFlown as part of US share on LISA pathfinder (referred to in USA as ST7-DRS)Within Europe, Queen Mary (University of London) also developing this technologyElectrostatic – Colloid

30. Electrostatic – ColloidSingle needle can provide only a few μNHowever srray of needles used to provide higher thrust levelsThrust controlled by acceleration voltage and/or propellant flow ratePropellant feed is typically at constant pressureBeam voltages up to 10kVTypical performance summary, based on Busek CMNT (9 emitters per thruster head)

31. Electro Magnetic EP Systems

32. Discharge is created in the propellant, ionizing it and forming a plasmaDischarge current interacts with a magnetic field; this produces a Lorentz force (JxB), which accelerates the plasma producing thrustMagnetic field either externally applied, or self induced by the discharge currentElectromagnetic Thruster - Operations and Performance

33. Electromagnetic Thruster – PPT (Pulsed Plasma Thrusters)Propellant typically a solid fuel bar (e.g. PTFE)Energy storage unit (ESU) or capacitor placed in parallel with electrodes is charged to a high voltageIgnitor (mounted close to the propellant) produces a spark when the ESU is dischargedAblates and ionizes the surface of the propellantCurrent flowing through the vaporized propellant interacts with the magnetic field created by this current to accelerate the propellant out of the engineAs the propellant bar is eroded, the spring pushes it forward for the next pulseCapacitor is then charged up again from a power supply and the pulse cycle repeated

34. Electromagnetic Thruster – PPT (Pulsed Plasma Thrusters)Inherently pulsed operation; can be operated up to several Hz Flown on several Russian and US spacecraft, from the 1960s through to 1980s Within Europe, there are some development activities at University of Stuttgart and Surrey UniversitySurrey University PPT flying on SSTL TechDemoSatPerformance summary:0.8 to 1.4 mNSpecific impulse around 1300 to 1500s Input power 70 to 100W

35. EP System Performance Summary

36. Main EP Design Considerations

37. EP Interface to Other S/C SystemsPower ManagementEP SystemThermal DesignS/C AccommodationAOCSContaminationCorrosionEMCChargingCorona

38. Main EP Design Considerations Mission Requirements for EP Design:• Low thrust levels correspond to relatively long periods of thrusters operation• If thrust level is made too low, this will result in: ΔV losses, due to firing away from optimum points on the orbit Full thrust arcs (for deep space missions); but some allowance has to be made forperformance uncertainties, contingences, SC reconfiguration activities, etc.EP raises the additional issues to be addressed for the overall spacecraft,related to the main interfaces

39. Main EP Design Considerations Power: design of the spacecraft power system has to take into account:Power range required by EP systemManaging power transients during EP system start-up, shut down, and contingencies (e.g. thruster failure) Thermal design: design of thermal system has to manage:Dissipation from electronics (including power system dissipation); this can be up to around 10% of the EP system powerDissipation from thrusters; although most heat dissipated in the thruster is radiated locally, some thrusters require heat transfer back through their mountings Thruster plume may induce heating of spacecraft surfaces on which it impingesAccommodation:Electric propulsion thrusters typically have much lower thrust densities(thrust/volume) than chemical thrusters, and so take up more space on thespacecraft

40. Main EP Design ConsiderationsAOCS: The low thrust operations require different control schemes Plume impingement can induce additional torques Thrusters are frequently mounted on pointing mechanisms to compensate for required thrust vector changes as CG position changes through spacecraft life, and other additional torques Plumes may also refract or absorb light to AOCS sensorsErosion and contamination:High energy ions ejected in the plume can erode spacecraft surfaces; similarly, non volatile propellants and thruster erosion products may condense on and contaminate the spacecraftEMC:Electric propulsion thrusters give radiated emissions; however, these are normally very low at 1GHz and above, and so do not normally interfere with communications bands attenuationPlumes can also refract or absorb electromagnetic radiation, giving distortion or attenuation

41. Main EP Design ConsiderationsCharging:• “Neutralized” plume is actually quasi-neutral mix of charged particles• Different mobility between electrons and heavy ions can result in differential coupling of plume to spacecraft surfaces and different potentialsFor example, exposed SA interconnects at positive potentials (up to 50 or 100V,depending on power system design) can attract highly mobile electrons in plume;this results in the plume plasma being at positive potential with respect to spacecraft structure, and can result in some additional parasitic currents

42. Main EP Design Considerations Corona discharge:• Some EP technologies involve the use of high voltages• Can result in corona discharge under certain low pressure conditionsThe following is an example of the Paschen curve for air:• Critical considerations are pressure levels, voltages, separation and the gasNot normally a problem for ambient pressures or high vacuumGeometry can also be an issue (e.g. anything which can cause a electric field strength concentration, such as spikes)

43. Other EP Thechnologies

44. Electrostatic - HEMPTHEMPT - High Efficiency Multistage Plasma Thruster• Developed by Thales Electron Devices, Germany• Similar in principle to a HET; uses a magnetic multistage cusp/mirror confinement of plasma Higher plasma densities, with the plasma being held away from the thruster walls• Performance summary:18 to 330 mN; throttling range of around 8:1 from a single thrusterSpecific impulse 1300 to 3750sInput power 300W to 12.5 kWCorresponding efficiencies ~25% to 45%

45. Electromagnetic Thruster - MPDMPD - Magnetoplasmadynamic thruster• Also referred to as Lorentz Force Accelerator (LFA) or MPD arcjet• Central cathode surrounded by a concentric cylindrical anode• High-current arc between the anode and cathode• Propellant is accelerated by the Lorentz force• Two main types, applied-field and self-field• Self-field - magnetic field is generated by the discharge current• Applied-field - magnetic field coil used Applied fields arenecessary at lower power levels; self-field configurations are too weak• Performance of up to 200N thrust, 11000s Isp (or more), with power levels of the order of MW• Low power for this technology is considered to be for power levels below 100kW• Main development activity is in the USA (NASA, USAF)• Some activity at Centrospazio and Stuttgart University in Europe• Small experimental MPD (1kW, 20 mN) flown by IHI (Japan)• High erosion of electrodes due to high current loads

46. Electromagnetic Thruster - VASIMRVASIMR - Variable Specific Impulse Magetodynamic RocketTrades Isp for thrust at constant power• 3 magnetic stages perform specific interrelated functions:• First stage handles the main injection of propellant gas and its ionisation• Second (“RF booster”) acts as an amplifier to further energise the plasma• Third stage is a magnetic nozzle, which converts the energy of the fluid into directed flow• Performance at 1 MW power:• >20N at low specific impulse (5000s); ~4N at high specific impulse (30000s)• Overall efficiency around 50% (reduces at low specific impulse)

47. EP Performance Summary

48. EP Performance Summary OverviewSpecific impulse and specific power:Includes indicative TRL for European developments

49. Summary of Current EP ApplicationsElectro thermal:• Resistojets used on various small satellites• Arcjets used for NSSK by Lockheed Martin Electrostatic:• HETs used for NSSK (including by Airbus), and for exploration (Smart-1)• More recently, also for GTO to GEO transfer (Electric Orbit Raising)• Gridded ion used for NSSK (mainly US), exploration (DS1, DAWN, Hayabusa, BepiColombo in work), LEO drag compensation (GOCE)• Colloid used for micropropulsion fine control (LPF) Electromagnetic:• PPTs used on various small satellites

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