Electric Propulsion

EXTROVERT Space Propulsion 11

Electric Propulsion


Perspectives on Achievable Performance

Minimum energy expenditure in taking 1 kg of mass to Earth Orbit : 9kWhTo Earth Escape : 18kWh(Is this true? Please check!)

Chemical energy depends of mass of propellant used – upper limit onenergy 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.

EXTROVERT Space Propulsion 11Electrical: 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, JPLwww.islandone.org/APC/lectric/00.html

http://www.islandone.org/APC/lectric/00.html

http://www.islandone.org/APC/lectric/00.html


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 andpressure forces. Most easily used in pulsed operation for short burst ofthrust.)

3. Electrostatic (ion propulsion) Propellant consists of discrete particlesaccelerated by electrostatic forces. Particles (usually atoms) are chargedby electron bombardment.

Here we will concentrate on ion propulsion

(Fig. 9-17 Humble Ion Propulsion)


www.rocket.com/epandse.html

http://www.rocket.com/epandse.html


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. ”

http://www.rocket.com/epandse.html


www.engin.umich.edu/dept/aero/spacelab/images/fmt_small.jpg

EXTROVERT Space Propulsion 11Propellants for Ion Propulsion

Various propellant types have been used. We generally want a cheapeasily ionized, dense propellant with easily accelerated particles.

• Xenon

• Argon

• Krypton

• Cesium

• C60 (Carbon 60)


www.agu.org/sci_soc/articles/ nelson.html

DS1 ion propulsion system.

http://www.agu.org/sci_soc/articles/nelson.html


http://www-ssc.igpp.ucla.edu/dawn/images/CR-1845.gif

.

[email protected]

mailto:[email protected]

EXTROVERT Space Propulsion 11Resistojet

www.islandone.org/ APC/Electric/02.html

Propellants: ammonia, biowastes, hydrazine, hydrogen. Augmented hydrazine thruster: augments catalytic decomposition. Isp ~ 300 lbf-s/lbm Input power: few hundred kilowatts; 60-90% efficiency. 30% better performance than cold gas thrusters

Courtesy Dr. 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 lbf-s/lbm require 250 to 550 Watts of power. Isp 336 lbf-s/lbm and operational lifetimes > 2.6 x 103 Ns demonstrated.

http://www.islandone.org/APC/Electric/02.html

mailto:[email protected]

EXTROVERT Space Propulsion 11Arcjets

: www.projectrho.com/ rocket/rocket3c2.html

http://www.projectrho.com/rocket/rocket3c2.html


http://www.aero.kyushu-u.ac.jp/fml/study/arc-thruster/arcjet-thruster.jpg




Arcjet Thruster Design Considerations for Satellites. NASA preferred reliability practices PD-ED-1253 http://www.hq.nasa.gov/office/codeq/relpract/1253.pdf

http://www.hq.nasa.gov/office/codeq/relpract/1253.pdf







EXTROVERT Space Propulsion 11Hydrazine Resistojets

RCA SATCOM, G-Star, and Spacenet communication satellites utilize hydrazine resistojets manufactured by Olin Rocket Research (now Primex Aerospace Company).

www.islandone.org/ APC/Electric/02.html

http://www.islandone.org/APC/Electric/02.html


http://www.irs.uni-stuttgart.de/RESEARCH/EL_PROP/RES/eak04b12.gif

EXTROVERT Space Propulsion 11Augmented Catalytic Thruster

· Schub (N) : 0,8-0,36 · Betriebsdruck (bar) : 26,5-6,2 · Spez. Impuls (s) : 299· Min. Impuls Bit (mNs) : 88,96· Gesamtimpuls (kNs) : 524,9· Masse (kg) : 0,871· Ventil Leistung (W) : 8,25· Ventil Heizerleistung (W) : 1,54· Katalysebett Heizerleistung (W) : 3,93· Resistojet Heizerleistung (W) : 885-610· Resistojet Spannung (V) : 29,5-24,5 DC· Nomineller Betrieb: 3,0 h im Einzelbetrieb 370 h akkumuliert

Abb. 7 : MR-502A - Widerstandsbeheiztes Triebwerk der Firma Primex Aerospace.

www.irs.uni-stuttgart.de/. ../RES/d_res_usa.html


http://www.irs.uni-stuttgart.de/RESEARCH/EL_PROP/RES/d_res_usa.html



http://www.plasma.inpe.br/LAP_Portal/LAP_Site/Figures/Ion_Thruster.gif

Ion Thruster


http://www.plasma.inpe.br/LAP_Portal/LAP_Site/Figures/Ion_Thruster.gif


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.

http://fluid.ippt.gov.pl/sbarral/ion.html


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.

EXTROVERT Space Propulsion 11System Performance

Components are

• Power Supply

• Power preparation and conditioning

• Thrusters = & 212j p eP m U

Between the output supply and the jet exhaust

η=

& 212 p e

sT

m UP η η η η=T S PP thwhere


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

η =1S

η = −0.1 0.3Sfor a nuclear device that must convert heat energy to electricity with some type of Engine or mechanism (thernoelectric, Brayton engine, Stirling engine)

η =0.92S for electrostatic power preparation.

η ~ 0.98S for steady arcjet systems.

η ~ 0.5 ~ 0.8thdepending on Isp and propellant (Fig. 9.41 from Humble ).


Also, equations are available to estimate the thermal and power preparation efficiencies for various Isp and propellants.

η η = + ln( )pp th A B IspFrom 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%


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 = bsPs where bs is specific mass. For a typical solar array, bs~ 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); bs~ 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 bs~ 0.1 to 0.4 kg/KW of waste heat


Note: Humble also provides a way to estimate mass of the power preparation hardware and the thrustersfor common systems:

bpp = 0.2 kg/KW for arcjet , compared to 20 kg/KW for PPT.

For electrostatic, we can combine the power preparation and thrusters:

b b η+ = ( )Dpp th pp C Isp

From Table 9.11 For Argon, C = 4490; D = -0.781


+ = ( )Dpp thrusters sMass Mass C Isp P

For Argon, with Ps = 10 KW and Isp = 2500,

0.7814490(2500) (10)

99.65pp thrustersMass Mass

kg

−+ =

=

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,

Δ+ +

= = = 0

vg Isys p pc spi

f sys

m m mMMassRatio eM m


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)


Courtesy: Robert.H. Frisbee, JPLhttp://www.islandone.org/APC/Electric/impulse.gif

Optimum Specific Impulse

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Electric Propulsion