EXTROVERT
Space Propulsion 11
Electric Propulsion
EXTROVERT
Space Propulsion 11
Perspectives on Achievable Performance
Hill & Peterson
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
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 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)
l
www.islandone.org/APC/lectric/00.htm
Courtesy: Robert.H. Frisbee, JPL
EXTROVERT
Space 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)
EXTROVERT
Space Propulsion 11
www.rocket.com/epandse.html
EXTROVERT
Space 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. ”
EXTROVERT
www.engin.umich.edu/dept/aero/spacelab/images/fmt_small.jpg
Space Propulsion 11
EXTROVERT
Space 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)
EXTROVERT
Space Propulsion 11
DS1 ion propulsion system.
www.agu.org/sci_soc/articles/ nelson.html
EXTROVERT
Space Propulsion 11
.
http://www-ssc.igpp.ucla.edu/dawn/images/CR-1845.gif
ctrussell@igpp.ucla.edu
EXTROVERT
Resistojet
Courtesy Dr. Robert H. Frisbee.
Space Propulsion 11
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
www.islandone.org/ APC/Electric/02.html
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.
EXTROVERT
Arcjets
Space Propulsion 11
: www.projectrho.com/ rocket/rocket3c2.html
EXTROVERT
Space Propulsion 11
http://www.aero.kyushu-u.ac.jp/fml/study/arc-thruster/arcjet-thru
EXTROVERT
Space Propulsion 11
http://www.aero.kyushu-u.ac.jp/fml/study/arc-thrus
EXTROVERT
Arcjet Thruster Design Considerations for Satellites. NASA preferred
reliability practices PD-ED-1253
http://www.hq.nasa.gov/office/codeq/relpract/1253.pdf
Space Propulsion 11
EXTROVERT
Space Propulsion 11
http://www.aero.kyushu-u.ac.jp/fml/study/arc-thruster/arcjet-thru
EXTROVERT
Space Propulsion 11
http://www.aero.kyushu-u.ac.jp/fml/study/arc-thruster/arcjet-thrus
EXTROVERT
Space Propulsion 11
http://www.aero.kyushu-u.ac.jp/fml/study/arc-thruster/arcjet-thru
EXTROVERT
Space Propulsion 11
Hydrazine 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
EXTROVERT
Space Propulsion 11
EXTROVERT
Space Propulsion 11
Augmented 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
EXTROVERT
Ion Thruster
Space Propulsion 11
http://www.plasma.inpe.br/LAP_Portal/LAP_Site/Figures/Ion_Thruster.
EXTROVERT
Space Propulsion 11
http://www.plasma.inpe.br/LAP_Portal/LAP_Site/Figures/Ion_Th
EXTROVERT
Space Propulsion 11
Ion thrusters used for stationkeeping 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.
fluid.ippt.gov.pl/ sbarral/ion.html
EXTROVERT
Radio-frequency Ion Thruster Assembly (RITA).
Space Propulsion 11
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.
http://cs.space.eads.net/sp/images/RITA_Schematic.jpg
EXTROVERT
Space Propulsion 11
System Performance
Components are
•Power Supply
•Power preparation and conditioning
1
2
Pj  mpUe
2
•Thrusters
Between the output supply and the jet exhaust
1
2
mpUe
Ps  2
T
where
T  SPPth
EXTROVERT
Space Propulsion 11
Efficiencies
S  1
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
S  0.1  0.3
S  0.92
S ~ 0.98
th ~ 0.5 ~ 0.8
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 ).
EXTROVERT
Space 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
ppth  A  B ln(Isp)
At a specific impulse of 2500 sec (Argon), the combined efficiency above is
37.8%
EXTROVERT
Space 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 = 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
EXTROVERT
Space Propulsion 11
Note: Humble also provides a way to estimate mass of the power preparation
hardware and the thrusters
for 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 pp  bthpp  C(Isp)
D
From Table 9.11 For Argon, C = 4490; D = -0.781
EXTROVERT
Space Propulsion 11
Masspp  Massthrusters  C(Isp)D Ps
For Argon, with Ps = 10 KW and Isp = 2500,
Masspp  Massthrusters  4490(2500)0.781(10)
 99.65kg
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,
Mi msys  mp  mpc
MassRatio 

e
Mf
msys
v
g0Isp
EXTROVERT
Space 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)
EXTROVERT
Space Propulsion 11
Optimum Specific Impulse
Courtesy: Robert.H. Fris
http://www.islandone.org/APC/Electric/impulse.gif