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8. Electric Propulsion An Overview

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OVERVIEW

 At the beginning of last century, Goddard experimented with electric gas discharge tubes recognizing basic concepts for Electric Propulsion.

 “Wege zur Raumschiffahrt” by Hermann Oberth, 1929.  Since the 1950’s, Electric Propulsion has a well established

research history in government, academia, and industry (US, Soviet Union, Europe, Japan).

 Since the mid 1980’s, the propulsion community has seen a boom in Electric Propulsion research.

 Since the early 1990’s:  all major US communications satellite manufactures (such as

Hughes, SS/LORAL, Lockheed-Martin) have embraced Electric Propulsion.

 Earth and Space Science Missions are increasingly baselined with Electric Propulsion (DS-1, EO-1, ST, CNSR, Mars Sample Return, etc.)

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OVERVIEW

 EP Statistics:  EP Capability:

  East/West and North/South Station Keeping  Orbit Transfer  Orbit Insertion

 US Industry (GEO/LEO):   142 Hydrazine Arcjets ordered or in orbit  Over 100 Hydrazine Resistojets in orbit   10 Xenon Ion Thrusters in orbit

 Russia:  Over 100 Hall Thrusters in orbit

 European and Japanese satellite manufacturers increasingly baseline EP into spacecraft design.

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Flight History of EP Systems

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Ch8 –5

CHARACTERISTICS

 Chemical Propulsion

Payload Propellant Power Thrust

Exhaust (Momentum Flux)

Thruster Powerplant

Payload Propellant Thrust

Exhaust (Momentum Flux)

Engine

 Energy Limited  Chemical bonding energy   Limiting energy release  Restricting specific impulse

 Low payload fraction  Strong but short burns  High thrusts and propellant

mass flows  Moderate exit velocities  Limited final velocity

 Electric Propulsion  Power Limited

 Energy conversion rate  Material restrictions

 Separate power source  Higher energy content supplied

to propellant  High specific impulse

 High exit velocities   Low propellant consumption

 Low thrusts and acceleration  High final velocities  Short travel times

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ELECTRIC PROPULSION BENEFITS

 High Fuel Efficiency Enables Missions (High Isp)  EP provides 2 to 10 times payload increases compared to

chemical  EP reduces trip time up to 3 times for many missions

0 0.1 0.2 0.3 0.4 0.5 0.6

Geo

sync

hron

ous

Nep

tune

Orb

iter

Plut

o O

rbite

r

Jupi

ter G

rand

Tour

Inte

rste

llar

Prec

urso

r

Payl

oad

Mas

s / L

aunc

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ass

Chemical Propulsion Electric Propulsion

Courtesy of NASA GRC

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FUNDAMENTALS

 Definition: Electric propulsion is accomplished by the acceleration of gases by electrical and/or electric and magnetic forces acting on a conducting plasma made up of the propellant gas constituents.

Electrical Energy Used to increase Potential Energy

Potential Energy Converted to Kinetic Energy by Various Methods

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Ch8 –8

CLASSIFICATION

E L E C T R I C S P A C E P R O P U L S I O N

Electrostatic Systems Electromagnetic Systems Electrothermal Systems

Arcjets

Resistojet Thrusters

Microwave-Plasma Thrusters

Radio-Frequency Ion Thruster

Ion Thruster

MagnetoPlasmaDynamic Thrusters

RF Plasma Thrusters

Hall Thruster

Pulsed Inductive Thrusters

VaSIMR

Pulsed Plasma Thrusters

Field Emission Ion Thruster

Contact Ionization Thruster

Kaufman Thruster

Thruster with Anode Layer

Stationary Plasma Thruster

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FUNDAMENTALS

 Thrust: T = mUex

 Specific Impulse: Isp = Thrust/Propellant Flow Rate = Uex/g0

 Efficiency: = Exhaust Power/Input Power = mUex/2Pin

•

• 2

• •

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Ch8 –10

FUNDAMENTALS

Thrust/Power Ratio as Function of Spec. Impulse Th

rust

/Pow

er [N

/kW

]

Specific Impulse [s]

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Ch8 –11

EP/POWER SOURCES FOR SPACE MISSIONS M

issi

on E

nerg

y

Ion Hall MPD, PIT, VASIMR

Piloted

Outer Planets Inner System Sample Returns

Inner Planets,Comets, Asteroids

Earth Orbit

Solar

Nuclear Nuclear

Isotope Solar with AeroCap & Chem

Solar

Solar at Earth with Aerocapture & Chemical at Target

Nuclear

Isotope

ION HALL MPD, PIT, VaSIMR

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GENERAL SPACECRAFT LAYOUT

 Based on Nuclear Power and Electric Propulsion

Spacecraft Bus

Electric Propulsion System

ElectricThruster

PropellantFeed

System

PowerProcessingUnit (PPU)

PropellantTank

Spacecraft SubsystemsC&DH RCSGN&C TCSRF SW

Science Payload

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HIGH POWER EP CANDIDATES

Variable Specific Impulse Magnetoplasma Rocket (VaSIMR)

Hall-Ion Thruster

Pulsed Inductive Thruster (PIT)

200 kWe MAI Li- LFA (MPD)

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ELECTRO-THERMAL SYSTEMS

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ELECTROTHERMAL SYSTEMS

 Resistojet Thruster  Principle of Operation:

  simplest of all EP devices  material limitations   electric power employed to

resistor   energy transfer to propellant   hot propellant expanded in

nozzle

 Performance Characteristics:  Specific Impulse: 200-300 s   Thrust: 2-350 mN  Efficiency: 65-90 %  Power: 10-5000 W   Typical Propellant:

  Hydrogen   Hydrazine

Propellant

Resistors

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EXAMPLES

 Concentric Tubular Resistojet:  Input Power: 3.0 kW  Thrust: 0.176 N  Spec. Impulse: 840 s  Efficiency: 88%  Propellant: H2

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ELECTROTHERMAL SYSTEMS

 Arcjet Thruster  Principle of Operation:

  arc discharge between anode + cathode

  direct propellant heating by arc

  two heating regions   hot propellant expanded in

nozzle   sophisticated PPU and PPC

 Performance Characteristics:  Power: 500-5000 W   Thrust: 2-700 mN  Specific Impulse: 400-1500 s  Efficiency: 40-50 %   Typical Propellant:

  Hydrogen   Hydrazine

Cathode

Propellant

Anode

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Ch8 –18

EXAMPLES

 Regeneratively Cooled Arcjet:  Power Input: 30 kW  Thrust: 3.35 N  Spec. Impulse: 1,010 s  Efficiency: 54%  Propellant: H2

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EXAMPLES

 2-kW Arcjet:  Input Power: 2.2 kW  Spec. Impulse: 600 s  Efficiency: 30%-42%  Propellant: Hydrazine

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ELECTROTHERMAL SYSTEMS

LA

DA

a q

G

Cathode

Anode/Nozzle

Typical dimensions: LA = 0.25 mm; DA = 0.6 mm; G = 0.6 mm

= 30°; = 20°

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ELECTROTHERMAL SYSTEMS

 Microwave Plasma Thruster  Principle of Operation

 microwaves generated in cavity

  electromagnetic coupling induces plasma discharge

  energy transfer to propellant   hot propellant expanded in

nozzle

 Performance Characteristics  Power: 100-600 W   Thrust: 2-20 mN  Specific Impulse: 200-400 s  Efficiency: 30-60 %   Typical Propellant:

  Inert Gases   Nitrogen   Hydrogen

Propellant

ResonantCavity

Microwave-PlasmaDischarge

CoaxialMicrowaveInput

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Co-Axial RF Plasma Thruster

 Power:  RF plasmas are sustained over a very wide power and voltage range.  Direct-drive schemes are feasible without the need of exotic voltage

requirements.  Target operating conditions: 1-100 W (< 1W) at frequencies of 1-1000

MHz, which is tailored to main bus operating voltages of most microspacecraft (< 10 V).

 Mass/Volume:  RF components and power sources/processing have seen tremendous

technological advancements in recent years especially with regard to size and efficiency and a significant reduction in mass and volume.

 The co-axial geometry of this propulsion device lends itself to further miniaturization.

 Lifetime:  Preliminary experiments have shown that electrodes are not subject to

erosion and sputtering in RF plasmas at certain conditions (frequency, electrode separation, gas pressure).

 A micropropulsion system based on this concept could provide lifetimes currently not exhibited by technologies considered for microspacecraft applications.

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Ch8 –23

Co-Axial RF Plasma Thruster

 Attributes:  Low power operation suitable for power limited spacecraft (micro/

nano satellite).  RF operation eliminates erosion enabling very long lifetimes.  Suitable for miniaturization leading to Thruster-on-a-Chip.

 Co-axial design philosophy:  Integration Issues arising from miniaturization:

  Propellant feed,   Power feed,  Diagnostics.

 Operating Principle:  Characterized by low-power, RF glow discharge operation.  RF power ionizes neutral gas and sustains the plasma.  Plasma heats gas and enthalpy of gas increases.  Thrust is generated by thermodynamic expansion of gas through nozzle.

Thruster Body Outer Electrode

Inner Electrode

N-Type Connector

Dielectric Separator

Propellant Feed Port

Dielectric Separator

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Ch8 –24

Co-Axial RF Plasma Thruster

 Proof of concept experiments accomplished:  Plasma operation at < 50 W,  Gas temperature depends on frequency and gap distance,  No erosion observed when compared to DC mode operation at same operating conditions.

 Two different size prototypes exist: RF25-1 RF50-1  Device maintains 50 Ω impedance,  Inner Electrode diameter: ¼” ½”

 Mass: 113 g 229 g  Volume: 55 cm3 85 cm3

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ELECTRO-STATIC SYSTEMS

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Ch8 –26

ION THRUSTER BASICS

 Four basic and separable processes occur in an ion thruster:

1.

1.  Energetic Electron Production,

2.

2.  Ion Production,

3.

3.  Ion Extraction and Acceleration,

4.

4.  Ion Beam Neutralization.

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ELECTROSTATIC SYSTEMS

 Ion Thruster  Principle of Operation

  ion acceleration in electrostatic field

  ion neutralization at exit plane   thrust generation: momentum

change due to electric body forces

 Performance Characteristics  Power: 0.5-5 kW (30 kW)   Thrust: 1-200 mN  Specific Impulse: 1500-5,000 s (15,000s)  Efficiency: 40-80 %   Typical Propellant:

  Inert Gases (Xenon, Krypton)   Mercury

  Ionization Mechanisms:   Bombardment   Field Emission, Radio Frequency, Contact

Ionization

 Research Initiatives:  High-voltage extraction grids  Cathodes  Wear mechanisms

 Participants:  NASA GRC/JPL  Academia   Industry

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OPERATION SCHEMATIC

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HOLLOW CATHODE

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ION OPTICS

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EXAMPLES

 NSTAR 30-cm Ion Thruster:  Power Input: 0.5-2.3 kW  Thrust: 21-92 mN  Spec. Impulse: 2000-3200 s  Efficiency: 42%-62%  Propellant: Xenon

DEEP SPACE 1: First use of Ion Thruster for Primary Propulsion

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ION THRUSTER

 15-cm Ion Engine  Slotted, Carbon-Carbon Grids  Power Input: 1.25 kW  Efficiency: 60% (± 5%)  Specific Impulse: 2,500-4,000 s  Thrust: 20-30 mN

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SAFE-30 POWER TRAIN

Thermal Energy Kinetic Energy

Qth

Power Conversion

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HALL SYSTEMS

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Ch8 –35

HALL THRUSTER SYSTEMS

 Stationary Plasma Thruster (SPT)  Thruster with Anode Layer (TAL)

 Performance Characteristics  Power: 0.5 - 3 kW (50 kW)   Thrust: 20-150 mN  Specific Impulse: 500-2500 s (4000 s)  Efficiency: 40-60 %   Typical Propellant:

  Xenon   Argon

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EXAMPLES

 TAL/D-55:  Power Input: 1.5 kW  Thrust: 38-128 mN  Spec. Impulse: 700-1800 s  Efficiency: 24%-53%  Propellant: Xenon

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EXAMPLES

 SPT-100:  Input Power: 1.35 kW  Thrust: 40-97 mN  Spec. Impulse: 1070-1600 s  Efficiency: 34%-50%  Propellant: Xenon  Extensive Flight History (Russia)

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HALL DEVELOPMENT PATH

 One kilowatt level systems flight proven (TRL 9)  Over 100 thrusters flown on Russian satellites   System flown on DOD experimental S/C (STEX)

 Five-kilowatt level system flight qualified (TRL 8)  NASA/BMDO Express T-160 flight program  Air Force/Loral SPT140 qualification program

 Ten-kilowatt system prototype demo (TRL 6)   1000 hr test of NASA/ Pratt & Whitney T-220

 Fifty-kilowatt breadboard validation (TRL 4)  NASA-457M in-house program  Russian TM-50 pathfinder experiments

Courtesy of NASA GRC

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ELECTRO-MAGNETIC SYSTEMS

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Ch8 –40

ELECTROMAGNETIC SYSTEMS

 Magnetoplasma Dynamic Thruster (MPD)   Principle of Operation

  arc discharge between anode + cathode   self-induced magnetic field   thrust generated by interaction of current and magnetic field   thermal thrust generation

 Performance Characteristics  Power: 1-200 kW   Thrust: 1-2,000 mN  Specific Impulse: 400-8,000 s  Efficiency: 30-50 %   Typical Propellant:

  Inert Gases   Cs, Li, Bi, N2

  Thruster Types:   Self-Induced Magnetic Field   External Magnetic Field

 Participants:  NASA JPL/GRC  Princeton

jj j x B

B

ANODE

CACATHODETHODEPLASMANEUTRAL GAS

INJECTION HOLES

INSULATOR BACKPLATE

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DEVELOPMENT PATH TO MWE PLASMA THRUSTERS

• Anode Texturing • Heat Pipes

• Lithium Propellant • Active Turbulence Suppression

• Plume Shields • Booms

h = 50% Isp = 4,000 s

200 kWe Steady State

100’s of Hrs at 3,000 A

1 - 5 MWe Steady State

h = 60% Isp ≤ 8000 s

10,000 Hrs at 20,000 A

• Multi-Channel Hollow Cathodes • Barium Addition

0.001

0.01

0.1

1

10

100

1000

Cur

rent

Den

sity

(A

/cm

2 )

4000350030002500200015001000Temperature (K)

PBa = 100 Pa

10 Pa

1 Pa

Pure W (110) Pure Ba Ba-W (110)

4

3

2

1

0

-1

Thru

st Co

effic

ient,

C T

25x103 20151050 Current, J (A)

BP (p)BP (b)

AIF (p)AIF (b)

CT (p)

AOF (b)

Total

200 kWe Lithium-fed Thruster

Ships of Space

10-8 g/cm2s at 0.3 m

10-10 g/cm2s at 30 m

Courtesy of NASA JPL

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ELECTROMAGNETIC SYSTEMS

 Pulsed Plasma Thruster  Principle of Operation

  spark plug + capacitor discharge

 main discharge ablates + ionizes Teflon

  self-induced magnetic field   Lorentz force

 Performance Characteristics  Power: 10-200 W   Thrust: 0.05-10 mN  Specific Impulse: 200-2000 s  Efficiency: 5-30 %   Typical Propellant:

  Teflon   Carbon based Polymers

Anode

Cathode

j

d

B

h

V(t) j x B

Insulator

SwitchI(t)

Q(t)

Capacitor

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Ch8 –43

EXAMPLES

 58-J Pulsed Plasma Thruster (PPT):  Power Input: 5-50 W  Thrust: 100-750 N  Spec. Impulse: 1200 s  Efficiency: 10%

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PULSED INDUCTIVE THRUSTER

 Operating Principle  Characterized by µ-second, MW-power pulsed operation.  Nozzle injects propellant covering coil surface, while pulse forming

network triggers discharge of cap bank.  Transient high current in coil generates rapidly changing magnetic

field.  Magnetic field induces strong azimuthal electric field which breaks down propellant.  Cross-product interaction of plasma current and magnetic field in coil accelerates plasma to generate thrust.

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PULSED INDUCTIVE THRUSTER

 Performance Characteristics:  Propellant: Argon, Hydrazine, Ammonia, Carbon Dioxide  Specific Impulse: 2,000 - 8,000 s  Efficiency: 20 - 50 %  Operation Mode: Single Shot  Discharge Voltage: 20 - 30 kV  Bank Capacitance: 9 µF

Impulse as a Function of Propellant Mass Efficiency as a Function of Specific Impulse

Ammonia, 16 kV

0

0.05

0.1

0.15

0.2

0 1 2 3 4 5 6 7 8

Impu

lse

[N-s

]

Mass [mg]

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Ch8 –46

PULSED INDUCTIVE THRUSTER

 Electric Characteristics:  Single Circuit Parameters:

 Nine circuits in parallel   Initial Rise Time: dI/dt = 30 kA/µs  Peak Current: 15 kA  Capacitance: 2 µF  Charge Voltage: 15 - 20 kV

Current Waveform for one Circuit

-10

-5

0

5

10

15

20

0 2 4 6 8 10 12 14

Cur

rent

[kA

]

Time [µs]

(dI/dt)initial≈ 30 kA/µs

Ipeak≈ 15 kA

Imin≈ 50% I

peak

 Technical Challenges:  Switch Technology

 High repetition rate and extreme long lifetime

 High peak currents  High and rapid initial current rise

 High Power Capacitors  Extreme long lifetime  Requires space qualification under

extreme operating conditions

 Powertrain Architecture  Recovery of reflected energy  Pulse shape control for optimum

pulse waveform

 Participants:  NASA MSFC/GRC  TRW

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VaSIMR

 Operating Principle:  Propellant gas is ionized at helicon plasma frequency  Ionized gas enters central chamber and undergoes ion-cyclotron

resonance heating  Hot, ionized plasma is expelled through contoured magnetic nozzle

to provide thrust

gas injection quartz tube

helicon antenna

magnets

ICRH antenna

exhaust region

VX-10 Lab Experiment

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VaSIMR DEVELOPMENT PATH

 Performance Characteristics:  VaSIMR concept scales well with high power  Proposed Flight Experiment (VF-10)

  Power: 10 kW   Specific Impulse: 10,000 s   Thrust: 100 mN   Propellant: H2, D2

 Research Activities:  Helicon Development   ICRH Development

 Participants:  NASA JSC/MSFC  Academia  DOE (LANL, ORNL)

( 1 0 k W

2 0 0 k W

1 0 0 M W

G R O U N D T E S T I N G l e a d s s p a c e t e s t i n g )

3 0 M W M a r s H u m a n

m i s s i o n 2 0 1 6 c a r g o 2 0 1 8 c r e w

3 0 M W M a r s N u c l e a r P a t h f i n d e r

O p e r a t i o n a l 1 M W

d e e p s p a c e n u c l e a r

1 0 0 k W d e m o

R T D M i s s i o n

P o w e r

' 9 6 ' 9 7 ' 9 8 ' 9 9 ' 0 0 ' 0 1 ' 0 3 ' 0 3 ' 0 4 ' 0 5 ' 0 6 ' 0 7 ' 0 8 ' 0 9 ' 1 0 ' 1 1 ' 1 2 ' 1 3 ' 1 4 ' 1 5 ' 1 6 ' 1 7 ' 1 8 ' 1 9 ' 2 0

N u c l e a r

Y e a r

N u c l e a r P o w e r D e v

S o l a r P o w e r D e v S o l a r

S p a c e e x p

2 k W

1 0 M W

1 . 5 M W

3 0 M W

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SUMMARY

 Electric Propulsion Capability:  Pulsed Microthrusters:

  Precision Control   Formation Flying

 kW-Class Systems:  NS/EW-Station Keeping   S/C Insertion  Maintenance and Deorbit  Robotic Planetary

 ENABLING ADVANCED EXPLORATION OF SPACE!  Exploration to expand understanding of space.  Commercial development and utilization of extraterrestrial

resources.  Provide potential human and robotic exploration.

 5-10 kW-Class Systems:  GEO Insertions   Large Platform Deorbit  Earth and Space Science Missions

 500+ kW-Class Systems:  Human/Robotic Exploration  Development of Space

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CONCLUSIVE REMARKS

 Performance is mission enabling  Increase in payload fraction,  Decrease in launch vehicle size (step down),  Reduction in trip time.

 Mission benefits arise from electric propulsion technology  ION - Auxiliary/Primary propulsion for planetary/deep space

missions,  HALL - Auxiliary/Primary propulsion for orbital missions  ELECTROMAGNETIC - Primary propulsion for planetary/deep space

missions.   In many aspects Electric Propulsion is superior over chemical

propulsion for space applications.  New developments in materials, electronics, energy storage and

power processing technology will further advance Electric Propulsion.

Disclaimer: Some pictures are courtesy of colleagues from industry and government.