Download Electric Propulsion

Propulsion
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Basic propulsion
Propulsion performance
Propulsion types
Launch vehicles
Basic Propulsion
Propulsion Basics
Three primary measures of propulsion system
effectiveness and efficiency are:
1. Thrust - the force available for launch or flight
2. Thrust duration - the time available for acceleration
3. Thrust efficiency - measured by the specific
impulse = Isp
Propulsion Basics
Three primary measures of propulsion system
effectiveness and efficiency are:
1. Thrust - the force available for launch or flight
2. Thrust duration - the time available for acceleration
3. Thrust efficiency - measured by the specific
impulse = Isp
Propulsion Basics
1. Thrust - propulsion thrust used for launch, orbital change,
orbital maintenance, station keeping operations, and
attitude control

Launch - very high thrust required (106 -107 Newtons (1
Newton = 0.2248 lb)
 Examples are the Atlas, Delta, Proton, Ariane 5, Long
March

Apogee and orbit boost - moderate thrust needed (103 -105N)
 Examples are Inertial Upper Stage (IUS), Payload Assist
Module (PAM), Centaur upper stage

Attitude control - low thrust needed (1-100 N or less typical)
- short duration bursts and frequent operational cycles
Propulsion Basics
2. Thrust duration - a measure of the total energy available in
the propulsion system

Short duration (.01 - 10 sec)
Primarily used for attitude control

Intermediate duration 10 – 1,000 sec
Used for launch and boost

Long duration 103 - 107 sec
Used for deep space propulsion (ion, nuclear)
Propulsion Basics
3. Thrust efficiency - measured by specific impulse, Isp (units
are seconds)

Low (1-100 s)
Used in attitude control on smaller spacecraft

Moderate (100-500 s)
Launch & boost; attitude control on larger spacecraft

High (500 – 5,000 s)
Interplanetary propulsion, station keeping
Propulsion Basics
Momentum = p = mass x velocity
 Measured by combustion exhaust mass flow rate
Force = F = dp/dt = d(mv)/dt = vdm/dt + mdv/dt
 F = mdv/dt = ma for constant mass
Thrust = T = Ve dm/dt (mass rate flow times exhaust velocity)
 Thrust comes from exhaust mass reaction force and
the pressure difference between the rocket nozzle
pressure and the ambient (surrounding) pressure
Propulsion Basics
Thrust from internal (nozzle) and external (ambient)
pressure difference
T = PeAe - PaAe = Ae(Pe - Pa)
where Pe = exhaust
pressure, Ae = exhaust chamber area, and Pa= ambient pressure
The greater difference between internal pressure, Pe, and
external pressure, Pa, the greater the force
Total thrust = Ve dm/dt + Ae(Pe - Pa)
Propulsion Basics
Exhaust velocity - Ve
The chemical rocket motor performance is determined
primarily by the exhaust velocity - higher exhaust velocity
improves thrust performance
Ve is often expressed as:
where Ve = exhaust velocity, Ro = universal gas constant
Tc = combustion chamber temp
m = molecular weight of exhaust gas
Pe = exhaust pressure, Pc = chamber pressure
γ = gamma = specific heat ratio (function of fuel
chemistry)
Propulsion Basics
The equation shows that increased exhaust velocity (and
performance) is available by:
 Increasing chamber-to-exhaust pressure ratio (vacuum
of space makes the chemical rocket more efficient in
space than in the atmosphere
 Increasing combustion chamber temperature (limitation
on this temperature is based on the chamber and
nozzle design and materials)
 Decreasing molecular weight - Hydrogen the best fuel
but needs oxidizer (H2 + O2 is one of the best propellant
combinations)
 Lower specific heat ratio - This is a property of the fuel
chemistry and has a limited range available
Propulsion Basics
Specific impulse
 Impulse = I = Thrust (force) x time increment Δt, or I =
TΔt

Specific Impulse = Isp = I/mg
where m = propellant
unit mass, g = gravitational acceleration at the Earth's
surface

Isp = T/mg for time increment Δt of 1 sec, a constant
thrust and a constant mass flow
Propulsion Basics
Specific impulse

Isp is a measure of efficiency either for propellants or
the motor

Defined as the thrust per unit weight flow rate of
propellant

High Isp = high thrust efficiency per unit mass important especially in mass-critical designs
Propulsion Basics

Propulsion Basics
Nozzle design

An important design criteria for the chemical rocket
motor is the exhaust nozzle

Exhaust velocity is maximized and pressure
differences are optimized within the nozzle parameters,
from curvature to internal diameter ratios

This is accomplished by the nozzle first constricting
the exhaust gas flow to increase the exhaust velocity,
then provide a divergent flow to accelerate the gas to
even greater velocity
Propulsion Basics
Nozzle design
Propulsion Basics
Nozzle design

A number of other parameters must be optimized for
the nozzle design, including the shape of the throat
area and the angles of convergence and divergence

Another important element in exhaust nozzle design is
the area ratio and resulting gas expansion, which also
represents the maximum exhaust velocity

This ratio is also important since the nozzle is matched
to the combustion chamber output and the ambient
pressure(s) at the nozzle exit
Propulsion Basics
Nozzle design

The area ratio is defined as the exit area divided by the
throat area, or
e = Aexit/Athroat

Ideally, the nozzle expansion curve would produce the
same pressure at the exit point along the wall as the
ambient pressure at the nozzle exit

While this suggests an extended nozzle, weight
limitations and aerodynamic pressures and loads
restrict the length of the nozzle, as does the cooling
mechanism
Propulsion System
Performance
Propulsion Basics
Propulsion system performance

The performance of a rocket motor or engine can be
measured in several ways, but the most common are
 Isp
 ΔV
 Ve
ΔV is the velocity change, or acceleration times the
thrust duration = a Δt
Propulsion Basics
ΔV

ΔV = acceleration x Δt = F Δt/m = Isp g
 This a dimensional relationship only - the more
accurate representation is:
ΔV ideal performance = Isp g ln[mi/(mi - mp)]
were ln = natural log, mi = initial vehicle mass, mp =
mass of consumed propellant, mi - mp = mf = empty
(final) vehicle weight (less is better)
Propulsion Basics
ΔVideal performance = Isp g ln[mi/(mi - mp)]
is the basic relationship between ΔV (performance) and Isp for a single
stage propulsion system. This is not the same as the ΔV required to attain
proper orbit, or for orbit changes. This equation allows us to calculate the
initial-to-final mass ratio for the launch vehicle or booster stage (plus
payload) by simple algebraic rearrangement.
mi/mf = exp[ ΔVrequired/(g Isp)] (50-200 typical for mi/mf )
To calculate the propellant mass use the same equation, but solve for mp
mp = mi[1- exp[ ΔVrequired/(g Isp)]]
Propulsion Basics
Note that the actual ΔV required needs to include the
gravitational force when departing to or approaching
a planet, and the drag force when passing through an
atmosphere, both reducing the ΔVideal available
The extra ΔV for gravity and drag usually requires
1500 to 2000 m/s for departure from Earth
Propulsion Basics
ΔVactual required = ΔVideal + ΔVg + ΔVD

ΔVg = ΔV loss due to gravity = g x Δt
were, g = acceleration of gravity at height r (9.89 m/s2
at the surface of the Earth), Δt is the time variable
which is to be integrated over the time of flight

ΔVd = ΔV loss due to aerodynamic drag
Orbit/Trajectory
Launch to LEO (400 km) (excluding gravity,
atmospheric drag)
).
Gravity & drag during ascent through atmosphere
(typical)
Flight path correction from vertical to horizontal
Launch to LEO (total)
ΔV (m/s)
Percent Payload
7,750
100% (reference)
1,400
350
9,500
100% (reference)
Launch to GEO
Lunar impact
Lunar landing (soft)
Launch to lunar orbit
Circumlunar mission with LEO return
10,200
12,500
14-15,000
13,500
16,000
10-25%
35-45%
20-30%
20-30%
25-35%
Lunar landing and return
16-18,000
1-4%
Mars
Mars - landing and return
11,390
23-27,000
20-30%
0.1-1%
Venus
Venus - soft landing
11,450
23-25,000
Mercury
Jupiter
Sun
Earth Escape
Escape Solar System
12,500
13,930
30,450
12,700
17,500
Chart of maximum vehicle
velocity versus vehicle
initial-to-final mass ratio
(minitia/lmfinal, or 1/MR) in an
idealized drag -free,
gravitationless flight
environment
Plotted are five Isp values
which show the maximum
ΔV limitations of propulsion
efficiency (Isp)
Single stage rockets are
typically limited to a 1/MR of
approximately 20, while
staged vehicles are capable
of mass ratios of 200 or
more (Sutton)
).
Propulsion Basics
From the above chart, the limitation of staged chemical rocket
motors with a maximum Isp of roughly 450 s is approximately 20,000
m/s with a initial-to-final mass ratio of 150
Included in the final mass fraction, mf, are the vehicle inert mass
(structure, tanks, engine, residual fuel, etc.), as well as the payload
mass. The propellant mass fraction, expressed as 1 - mf/mi, would
ideally be 1, although completely impractical since the entire vehicle
would be propellant. The ideal propellant mass fraction of 1 would
also mean that the final-to-initial mass ratio be zero, or the final mass
is zero. Also, this is an obviously impractical design. Propellant mass
ratios beyond 0.85 require careful design, with a practical limit near
0.95.
Propulsion Basics
Limitations on mass ratio

For deep-space or interplanetary missions, a high
payload capacity is desirable

Ideally a small final-to-initial mass fraction with a
minimal remaining vehicle mass

The propellant mass fraction would be modest to
accommodate a payload, on the order of 0.70 to 0.80
Propulsion Basics
Limitations on mass ratio

Final-to-initial mass ratio (MR) would be 0.1 to 0.2
(1/MR = 10-20)

Another way to look at the ΔV limitation is to use the
final-to-initial mass ratio
mf/mi, = e -ΔV/g Isp = e -(ΔV/Ve) if we substitute the
exhaust velocity Ve for Isp times g
Propulsion Basics
Limitations on mass ratio

For an interplanetary payload, the vehicle's exhaust
velocity needs to be comparable to the mission
velocity requirement (ΔV)

A vehicle with a 400 s Isp would have difficulty
delivering a sizable fraction of the total remaining
vehicle mass as payload beyond Mars since g Isp is of
the order of 4,000 m/s

Needs to be 12,000-14,000 m/s for the ΔV
requirement beyond Mars
• Available payload
mass for chemical
rockets are nearly
insignificant beyond
Jupiter
• Higher Isp electric
boosters are
capable of reaching
our Galaxy's
interstellar
environment
• Missions beyond
Jupiter are possible
for chemical rockets
using gravity
assists
).
Propulsion Basics
Propulsion calculations
Calculate the empty weight of the vehicle from the
known Isp and initial weight
Step 1 Calculate the initial-to-final mass ratio for a lunar
flyby mission with a ΔV of 12,500 m/s and a LOX/LH2
propulsion system with an Isp of 435 sec and a final
mass of 1.811x106 kg (g = 9.80 m/s)
mi/mf = e[ΔVrequired/(g Isp)]
mi/mf = e[12,500 m/s/(9.80 m/s x 435 sec)]
= e2.93 = 18.77
Propulsion Basics
Propulsion calculations
Step 2 Working backwards and knowing to final mass
of the vehicle (empty weight of vehicle plus payload
weight), find the final mass of the vehicle
mf = mi/e[ΔVrequired/(g Isp)] but its easier than this
since mf = 1/(mi/mf) x mi
mf = 1/18.77 x 1.811x16 kg
= 96,480 kg
Propulsion Basics
Propulsion calculations
Step 3 Find the empty mass of the vehicle knowing the
payload mass, which is 8,600 kg
mempty weight = mf - mpayload = 96,480 - 8,600 kg
= 87,880 kg
Propulsion Basics
Propulsion calculations
Step 4 Find the propellant mass required for the
mission using
mp = mi {1-e[-ΔVrequired/(g Isp)]}
= 1.811x106 kg {1 - e[-12,500 m/s/(9.80 m/s x 435 sec]}
= 1.811x106 kg {1 - 0.0533)
= 1.715x106 kg
Propulsion Basics
Propulsion calculations
Step 5 From the previous calculation, the propellant
mass fraction can be found with a simple ratio
mp/mi = 1.715x106 kg/1.811x106 kg
= 0.947 (94.7% fuel)
For comparison, the payload mass fraction is
mpayload/mi = 8,600 kg/1.811x106 kg
= 0.0047 = 0.47% payload with 99.53% vehicle and
propellant
Propulsion Types
Chemical Rockets
1. Chemical (solid, liquid, hybrid)

Basic requirements for a chemical, high
thrust, high performance system
• High temperature chemical reaction
• Moderate to high Isp
• Low exhaust gas molecular weight
Chemical Rockets
Liquid bipropellant - separate fuel and oxidizer
Chemical Rockets
Chemical liquid – Monopropellant

Single fuel with a self-contained oxidizer

Most common fuel is hydrazine N2H4 (goes to NH3, N2,
H2)

Variations are monomethyl hydrazine,
unsymmetrical-dimethyl hydrazine

Hypergolic when passed over a catalyst surface. The
simplicity and moderate performance of this system
make this useful for many spacecraft propulsion
applications
Chemical Rockets
Liquid monopropellant
Oxidizer
Specific
gravity
Liquid oxygen 1.14
(LOX)
Hydrogen
peroxide
(H2O2)
Nitric acid
(HNO3)
Boiling point
(1 atm)
Characteristics
90 K (-183oC, -298oF) •Not hypergolic but can combust spontaneously with many
materials at elevated pressures
•Most commonly used rocket fuel oxidizer
•Non-toxic and non-corrosive
423 K (150oC, 302oF) •Oxygen and heat are released by the decomposition of hydrogen
peroxide into H20 + O2
•Decomposition is spontaneous with exposure to a catalyst such
as platinum or iron oxide
•H2O2 was used to generate gas to drive turbopumps in the V-2, X-1
and X-15
o
o
1.26-1.41 356 K (83 C, 181 F)
•Nitric acid and its variants are highly corrosive
•Red fuming nitric acid is nitric acid + 5-20% nitrogen dioxide;
more stable, less corrosive than pure nitric acid
•Addition of <1% fluorine ion (HF) reduces corrosion (inhibited red
fuming nitric oxide)
•Used as propellant oxidizer with gasoline, amines, and hydrazine
Nitric acid is hypergolic when combined with hydrazine and amines
1.19
•Mildly corrosive unless mixed with water
•Spontaneous combustion occurs when exposed to many materials
•NTO is hypergolic when combined with most fuels
•High vapor pressure requires relatively heavy tank
•Used in numerous Russian rockets, the Titan booster series, and
the Space Shuttle attitude control
•Highly toxic - exposure limit < 5 ppm
Nitrogen
tetroxide
(N2O4, NTO)
1.44
291 K (18oC, 64oF)
Fluorine
1.11
83 K (-190oC, -310oF) •Fluorine and fluorides have been proposed in various fuel
combinations which are highly corrosive, difficult to handle, and
toxic
•No fluorine oxidizers have been used for production rocket
engines
Fuel
RP-1
Specific
gravity
Boiling point
(1 atm)
Characteristics
0.80-0.815 420 K (147oC, 297oF) •Highly-refined kerosene
• Developed as a fuel that could also be used for cooling
high-temperature nozzles and combustion chambers
•
Sulfur, aromatics, and unwanted isomers removed to
permit use at high temps
•
Greater stability, lower toxicity, less residue, higher
performance than other hydrocarbons
•High flash point 336 K
•
Safer, less explosive than many hydrocarbon fuels
including gasoline
•Used in Russian R-7 booster and its derivatives, Soviet N-1, Atlas,
Thor, Delta I-III, Titan I, Saturn I, IB, V (1st stage)
Liquid
hydrogen
(H2, LH2)
0.07
(requires
large fuel
tanks)
20 K (requires
•High specific impulse
extensive insulation
for tank and feed
•Highly flammable when hydrogen gas is mixed with air
lines)
•Increased density possible with supercooled solid or slush
hydrogen (not yet used)
•Non-toxic (breathable gas; can replace nitrogen in an artificial
atmosphere)
•Non toxic exhaust gas when reacted with oxygen
•Hydrogen britalizes most metals, making turbopump design more
challenging than with other fuels
Fuel
Methane
(CH4)
Specific
gravity
0.47
Boiling point
(1 atm)
Characteristics
90 K (-183oC, -298oF) •Hydrocarbon fuel
•Stored cryogenically
•Low cost; freely available from gas wells, biomass decomposition
•Potentially useful for Mars return missions
•Under research but not used in production liquid fuel engines
•Possible fuel for arc-jet or resistojet thrusters
Hydrazine
(N2H4)
0.80
Monomethyl 0.88
hydrazine
(CH3NHNH2,
MMH)
387 K (186oC, 303oF) •Used as both a monopropellant and a bipropellant fuel
•Hypergolic fuel as monopropellant
•Hypergolic when mixed with nitrogen tetroxide and with nitric acid
•Spontaneous ignition also possible if it is spilled as a liquid in air,
and in contact with many materials
•Used in the production of stainless steel, nickel and some
aluminum alloys
•Not used in iron, copper, or some aluminum alloys
•Very long storage life
•Highly toxic
• Exposure limit < 0.1 ppm
• Known carcinogen
•Has been used as a monopropellant for gas generators (e.g. Space
Shuttle hydraulic system), and spacecraft attitude control
360 K (87oC, 189oF) •Better liquid temperature range than hydrazine
•Lower reaction threshold to shock waves than hydrazine
•Slightly lower Isp than for hydrazine
•Highly toxic
•
•
Exposure limit < 0.2 ppm
Suspected carcinogen
Oxidizer
Liquid oxygen (LOX)
Fuel
Isp (theoretical, Isp (theoretical,
vacuum)
1 atm)
Liquid hydrogen (lLH2)
477 s
390 S
LOX
Kerosene (RP-1)
370 s
300 s
LOX
Monomethyl hydrazine
365 s
301 s
LOX
Methane (CH4)
368 s
296 s
Liquid ozone (O3)
Hydrogen
580 s
Nitrogen tetroxide (N2O4)
Hydrazine (N2H4)
334 s
Red fuming nitric acid
RP-1
Hydrogen peroxide (H2O2)
Monopropellant
H2O2
Fluorine (Fl)
Fl
292 s
269 s
154 s (90% H2O2)
RP-1
279 s
Lithium
542 s
Hydrogen
580 s
410 s
Propulsion Basics
Liquid propellant complications
A. Zero gravity fuel feed
Liquid fuels are not confined to any specific region within
the tank in low or micro gravity but there are several ways
to allow for positive fuel feed during low gravity conditions
 Capillary devices - use surface tension to keep gas and
liquid separated in the tank (requires pressurization).
Used on Shuttle and Viking
 Diaphragms and bladders - physically separate gas an
liquid with flexible lining made of elastomer or Teflon
(requires pressurization). Used on Magellan and
Voyager
 Bellows - an expandable metal device to separate gas
from liquid (requires pressurization)
Propulsion Basics
Liquid propellant complications
B. Temperature extremes
 Operating temperature range of liquid propellants is limited
and is a function of pressure
Cryogenic liquids must be kept at low temperatures at
moderate pressures
 Storage system, transfer and pumping system, and the
combustion components must be able to reliably withstand
high temperature extremes (20K [storage] to 6,000K
[combustion])
 Hydrazine and its variations and nitrogen tetroxide do not
require cryogenic systems but may require electric strip
heaters
Propulsion Basics
Liquid propellant complications
C. Oscillations

Liquid propellants can oscillate or slosh while in the tanks,
especially during high thrust during launch

Potential for a serious effect on the tank structure as the
fuel surges up and down

Baffles, reinforcements and overall tank design are used to
reduce this effect
Chemical Rockets
Chemical - Solid
Solid rocket fuel is typically identified by the type of
chemical binder used

HTPB (hydroxy-terminator polybutadiene) is a
stronger binder, more flexible, and faster curing, but
suffers from a slightly lower Isp than PBAN (used on
Delta II, Delta III, Delta IV, Titan IVB and Ariane

PBAN (polybutadiene acrylic acid acrylonitrile) and
uses fast-curing, toxic isocynates with a slightly
higher Isp, is less costly, less toxic, and used in the
larger boosters (Titan III, the Space Shuttle SRBs,
and the new SLS) HTPB is or has been
Electric Propulsion
Electric Propulsion
Electric propulsion entails either
1. Accelerating charged particles; or
2. Heating cold gas
Both processes use electrical power
Electric and/or magnetic fields are used to
accelerate charged (ionized) particles
Electrothermal, microwave solar, nuclear, or arc
current methods are used to heat cold gas
Electric Propulsion
Advantages



High Isp
Reduced propellant mass (or increased payload
mass)
Increased mass savings with increased
duration/distance
Disadvantages




High power requirement
High system mass possible
Ion plume can degrade surfaces and create
charge buildup
High cost
Electric Propulsion
Common electric engines
Electrothermal

Resistojet thruster

Arcjet thruster

Solar thermal thruster

RF-heated (microwave) thruster
Electrostatic

Ion thruster
Electrodynamic

Magnetoplasmadynamic thruster (MPD)

Hall-effect thruster
Pulsed-plasma thruster

Variable specific impulse thruster
Nuclear

Nuclear thermal ion
Electric Propulsion
Electromagnetic field ion engine
Ion production




Electrical (resistive) heating
Electron excitation
Microwave excitation
Nuclear heat source
Ion acceleration



Electrostatic field
Magnetic field
Current flow in electric/magnetic field
Electric Propulsion
Electrostatic ion engine
Electric field accelerates oppositely charged ions
in opposite directions

Force is proportional to field strength and ion
charge
F = qE where E is the electric field strength and q is the ion’s charge
Vexit = 2qE/m where Vexit is the electric charge exit speed and m is

the ion mass

High charge/mass ratio is desired


Xenon common today
Krypton, cesium, mercury also used
Thrust proportional to power/current
T = Pη/(gIsp) where T is thrust, P is thrust power, and η is thrust

efficiency
Electric Propulsion
Electrostatic ion engine requirements


High electrostatic field (voltage)
High charge/mass ratio propellant



Highly-ionized high atomic mass gas
Vacuum conditions
Neutralized ion beam to prevent charge
separation

Electron beam injected into ion beam as it
exits engine
Electric Propulsion
Electrostatic ion engine example
(this example includes
magnetic field confinement)
Electric Propulsion
Electrodynamic ion engine
A. Magnetoplasmadynamic ion engine


Also uses electrostatic field but indirectly
Most applications are the Includes magnetic
field acceleration from a radial current flow


Current flow between cathode and anode produces an
induced magnetic field
Current of ions in self-generated magnetic field
produces axial acceleration
F = qV X B where the accelerating force is equal to
the ion charge times the ion’s velocity as a cross
product with the induced magnetic field strength

Additional magnetic field around device
increases acceleration force
Electric Propulsion
Magnetoplasmadynamic ion engine
Electrodynamic ion
engine example
Electric Propulsion
Electrodynamic ion engine
B. Hall effect (plasma) thruster

Ions accelerated in an axial electrostatic field




Radial magnetic field accelerates ions axially
Lower efficiency than electrostatic engine but
greater thrust range



Electrons forward towards anode
Positive ions rearward towards exhaust
Divergent exhaust plume deceases net exit velocity
Common applications include station-keeping
on geostationary satellites
Highest power requirement and highest thrust
of typical electrical power (EP) thrusters
Electric Propulsion
Hall effect (plasma) thruster
Electric Propulsion
Hall effect (plasma) thruster
Photograph of
experimental Hall-effect
thruster with dual (inner &
outer) magnetic rings
and electron injector
Electric Propulsion
Electrodynamic ion engine
C. Variable specific impulse magnetoplasma
engine





Ions accelerated by electromagnetic field and
microwave (RF) energy
Intended to bridge the gap between high-thrust,
low-specific impulse propulsion systems and
low-thrust, high-specific impulse systems
Can operate in either mode
Developed by astronaut Franklin Chang-Diaz
Also called the variable specific impulse
magnetoplasma rocket (VASIMR)
Electric Propulsion
Electrothermal (heating cold gas)
Two types currently used



Resistojet
Arcjet
Heating mechanisms





Electrical (resistive or arc) heating
Microwave excitation
Nuclear heat source
Solar heating
Modest performance



Isp modest but greater than cold gas
Modest thrust but greater than cold gas
Electric Propulsion
Electrothermal engine - resistojet

Cold gas injected into an expansion chamber is
accelerated by heating the gas

Heating possible with several methods






Non-reactive and reactive gases used
Non-reactive – more efficient


Electrical (resistive) heating
Microwave excitation
Nuclear heat source
Solar heating
Nitrogen, hydrogen examples
Reactive gases – less efficient

Hydrazine, methane are examples
Electric Propulsion
Electrothermal engine resistojet
Electric Propulsion
Electrothermal engine - arcjet
 Gas heating from high current arc
between interior cathode and anode shell
 Higher temperatures than resistojet
possible, hence higher Isp
 Higher power requirement than resistojet
 Non-reactive gases – more efficient
 Reactive gases – less efficient
Electric Propulsion
Electrothermal engine - arcjet
Electric Propulsion
Electrothermal engine - other

Solar heating also possible by directing
focused light on gas chamber

Microwave heating also possible
Electric Propulsion
Nuclear (not electric except for nuclear ion)
 Nuclear heating of cold gas can reach
extremely high Isp
 Hot radioactive core can reach much
higher temps than resistojet and arcjet


Nuclear thermal
Additional acceleration of ionized gas
with electrostatic/magnetic fields also
possible

Nuclear ion
Electric Propulsion
Nuclear thermal (non-electric)

Cold gas heated by
reactor core to high
temperature (5,000 K) and
high Isp
Electric Propulsion
NERVA nuclear thermal
reactor model evolution
Electric Propulsion
NERVA nuclear thermal reactor
Electric Propulsion
Electric propulsion - general
 Power proportional to thrust
 Power supply output proportional to
power supply mass
 Acceleration inversely proportional to
spacecraft mass
 Balance required between thrust power
and the mass required for thrust
Electric Propulsion
Electric propulsion
optimization
Questions?