Presentation on theme: "EXTROVERTSpace Propulsion 11 Electric Propulsion."— Presentation transcript:
EXTROVERTSpace Propulsion 11 Electric Propulsion
EXTROVERTSpace Propulsion 11 Perspectives on Achievable Performance Minimum energy expenditure in taking 1 kg of mass to Earth Orbit : 9kWh To Earth Escape : 18kWh (Is this true? Please check!) Chemical energy depends of mass of propellant used – upper limit on energy per unit mass. H2-O2: 3.7kWh per kg. Upper limit on chemical propulsion specific impulse ~ 500 s Hill & Peterson Nuclear thermal: energy transfer must come across some solid walls: maximum propellant temperature is limited by maximum wall temperature. Max specific impulse may be around 1000s.
EXTROVERTSpace Propulsion 11 Electrical: No upper limit identified on energy transfer per unit mass – no upper limit on specific impulse. Energy source can be solar, or Energy from nuclear fuel, which has extremely high energy density (orders of magnitude >> chemical) Courtesy: Robert.H. Frisbee, JPL www.islandone.org/APC/lectric/00.htm l
EXTROVERTSpace Propulsion 11 Several classes of electric propulsion 1.Electrothermal – resistojets and arcjets (N2H4) 2.Electromagnetic – steady (MPD) and unsteady (pulsed plasma thrusters – PPP) (stream of conducting fluid is accelerated by electromagnetic and pressure forces. Most easily used in pulsed operation for short burst of thrust.) 3.Electrostatic (ion propulsion) Propellant consists of discrete particles accelerated by electrostatic forces. Particles (usually atoms) are charged by electron bombardment. Here we will concentrate on ion propulsion (Fig. 9-17 Humble Ion Propulsion)
EXTROVERTSpace Propulsion 11 www.rocket.com/epandse.html “Functional Model Thruster (FMT) provided by the NASA Glenn Research Center. The FMT is functionally equivalent to the 2.3 kW NSTAR ion thruster that flew on Deep Space 1. NSTAR was the first demonstration of ion thruster technology as primary propulsion on an interplanetary spacecraft. ”
EXTROVERTSpace Propulsion 11 Propellants for Ion Propulsion Various propellant types have been used. We generally want a cheap easily ionized, dense propellant with easily accelerated particles. Xenon Argon Krypton Cesium C60 (Carbon 60)
EXTROVERTSpace Propulsion 11 www.agu.org/sci_soc/articles/ nelson.html DS1 ion propulsion system.
EXTROVERTSpace Propulsion 11 Resistojet www.islandone.org/ APC/Electric/02.html Propellants: ammonia, biowastes, hydrazine, hydrogen. Augmented hydrazine thruster: augments catalytic decomposition. I sp ~ 300 lb f -s/lb m Input power: few hundred kilowatts; 60-90% efficiency. 30% better performance than cold gas thrusters Courtesy Dr. Robert H. Frisbee.Robert H. Frisbee Technology issues: material/propellant compatibility at high temperatures, heat transfer; radiation losses. Heat transfer to gas stream is complicated by the geometries and temperature ranges typical of resistojets. Hydrazine resistojets used on several communication satellites: Four TRW hydrazine thrusters on Ford Aerospace's INTELSAT V satellites for station keeping. Thrust of 0.22 to 0.49 Newtons and Isp 296 lb f -s/lb m require 250 to 550 Watts of power. Isp 336 lb f -s/lb m and operational lifetimes > 2.6 x 10 3 Ns demonstrated.
EXTROVERTSpace Propulsion 11 fluid.ippt.gov.pl/ sbarral/ion.html Ion thrusters used for station- keeping on geostationary satellites since 1997. Demonstrated ability to propel space probes: encounter of NASA Deep Space-1 spacecraft with comet Borrelly in September 2001. Ion thrusters unexpectedly performed the first electric propulsion aided orbit transfer of a satellite, following failed orbital injection of ESA's Artemis mission. 2003: first use of a microwave ion thruster on Japanese Muses-C spacecraft.
EXTROVERTSpace Propulsion 11 http://cs.space.eads.net/sp/images/RITA_Schematic.jpg Radio-frequency Ion Thruster Assembly (RITA). Isp 3000 to 5000 s, adjustable thrust from 15 to 135%, operating life > 20,000 hours 85% less propellant than bipropellant thrusters. A 4100 kg spacecraft in GEO using conventional propellants over its 15 year life would save around 574 kg in propellant mass by using RITA.
EXTROVERTSpace Propulsion 11 System Performance Components are Power Supply Power preparation and conditioning Thrusters Between the output supply and the jet exhaust where
EXTROVERTSpace Propulsion 11 Efficiencies for solar arrays since they produce electricity directly (not this does not account for the 18% to 25% conversion efficiency of a solar array from solar radiation to electricity for a nuclear device that must convert heat energy to electricity with some type of Engine or mechanism (thernoelectric, Brayton engine, Stirling engine) for electrostatic power preparation. for steady arcjet systems. depending on Isp and propellant (Fig. 9.41 from Humble ).
EXTROVERTSpace Propulsion 11 Also, equations are available to estimate the thermal and power preparation efficiencies for various Isp and propellants. From Table 9.11, Humble. For Argon, A = -2.024; B = 0.307 At a specific impulse of 2500 sec (Argon), the combined efficiency above is 37.8%
EXTROVERTSpace Propulsion 11 System Mass It appears that Isp and efficiency get better with more power. When would we not want a system with as much power as we can get? POWER COSTS MASS! Typically, we use a linear relationship: Mass = sPs where bs is specific mass. For a typical solar array, s~ 7 to 25 kg/kW, depending on cell efficiency and substrate type. (see Table 9.10 from Humble) For a typical nuclear reactor, (remember, Ps = thermal power); s~ 2 to 4 kg/kW depending on shielding Note that we require space radiators to reject the heat dissipated by the power systems or reactor. Space Radiator s~ 0.1 to 0.4 kg/KW of waste heat
EXTROVERTSpace Propulsion 11 Note: Humble also provides a way to estimate mass of the power preparation hardware and the thrusters for common systems: pp = 0.2 kg/KW for arcjet, compared to 20 kg/KW for PPT. For electrostatic, we can combine the power preparation and thrusters: From Table 9.11 For Argon, C = 4490; D = -0.781
EXTROVERTSpace Propulsion 11 For Argon, with Ps = 10 KW and Isp = 2500, So, for a given system, we can calculate the power system mass, the radiator mass and the pp + thruster mass. Treating Isp as an independent variable and knowing from the rocket equation,
EXTROVERTSpace Propulsion 11 As Isp increases, Mass ratio decreases, but if Isp increases, Ps increases, system mass decreases, so payload mass decreases. These are competing effects, so there is usually an optimum Isp that results from the compromise. Optimum Isp depends on many systems-level design characteristics (Fig. 9.3 in Humble)