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.
Slide 3
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
Slide 4
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)
Slide 9
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.
Slide 25
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.
Slide 26
EXTROVERTSpace Propulsion 11 System Performance Components are
Power Supply Power preparation and conditioning Thrusters Between
the output supply and the jet exhaust where
Slide 27
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 ).
Slide 28
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%
Slide 29
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
Slide 30
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
Slide 31
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,
Slide 32
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)
