electric propulsion for rockets - prof. seitzman...

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Electric Propulsion-1

Copyright © 2003-2004, 2017 by Jerry M. Seitzman. All rights reserved. AE4451 Propulsion

Electric Propulsion for

Rockets


Copyright © 2003-2004, 2017 by Jerry M. Seitzman. All rights reserved.

Electric Propulsion System Elements

Propellant Energy Source

Storage

Feed System

Storage

Conversion

Thruster

Gas (Xe), liquid (N2H4), solid (teflon)

chemical, nuclear, radiation (solar)

Thermodynamic (nozzle), electric, electromagnetic

to electricity at proper V, DC/AC( freq), current,

steady/pulsed

Power source separate

from propellant

Isp not limited by

chemical bond energy

2



EP Power Requirements

• Jet Power

• Increase in Isp (or ue) entails increase in power

– ue2 for constant flowrate

– ue for constant thrust

• For comparison, chemical rocket with thrust of

just 1000 lbf (4.5 kN) and Isp= 350s, Pj=7.7MW

Pj=66 MW for elec. rocket with Isp=3000s

and same thrust

• EP tends to be power-limited to low thrust

(and low acceleration)

2

21

ej umP

ee uum2

1 eu2

1 espgI2

1Minimum possible

power required



Required Supply Power

• Supply Power

• Overall conversion efficiency has 3 main

components

T

j

s

PP

total efficiency of

energy conversion

thppST

energy conversion - of raw

energy source to electricity

power preparation and

conditioning electronics

(losses in electronics)

thruster efficiency -

only part of input

elec. energy

converted to KE

• ~100% for photovoltaics –

direct conversion of photons

(does NOT include fraction

of solar radiation that can be

converted to electricity by

typical solar array, ~18-25%)

• 10-40% for nuclear thermal,

thermodynamic cycle limits = lots of waste heat

• ~92% (electrostatic)

• ~98% (steady arc jets)

• 40-70%

3



Mass Requirements

• What limits available power to EP systems?

– mass of power plant and associated

systems

• If power plant mass is significant fraction of

propellant mass, then some advantages of

higher specific impulse operation are lost

– “payload” may consist mostly of power

plant/electronics for propulsion system



Mass Requirements (con’t)

• Power source mass

sssourcepower PMass nearly linear

relationship

specific mass (e.g., kg/kW)

Note: sometimes used for specific mass, but also used by others for

specific power (kW/kg), i.e., = 1/

• s~7-25 kg/kWelec for solar arrays (depends on cell design,

substrate)

• s~2-4 kg/kWthermal for nuclear reactors (depends on shielding)

– to reject waste heat, also require additional mass for radiators

R~0.1-0.4 kg/kWwaste heat

4



Mass Requirements (con’t)

• Power preparation and conditioning

• Large variation with type of EP device (especially if need high voltage or power, and short pulse forming electronics/switches)

– pp~0.2 kg/kWelec for typical arcjets

– pp~20 kg/kWelec for PPT (pulsed plasma

thrusters)

pppppp PMass



Electric Rocket Propulsion

A. Background

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Electric Propulsion - Accelerators

Electrothermal Electromagnetic Electrostatic

Accel.

Force

Pressure,

Electrically heat

propellant and use

nozzle expansion

Lorentz,

Magnetic and elec.

fields accelerate

charged particles

Electrostatic,

Static electric field

alone accelerates

charged particles

Isp(s) 300-1,500 1,000-10,000 2,000-100,000+

<10-3 <10-4 <10-410-6

eF

mF

p

Weight

Thrust



Electrical Heating

• Current passing through

conductor heats it by

amount proportional to

its resistance

• Wire

• Gas (Plasma)

Discharge

– resistance due

to collisions

RJQ 2

Current

(A=C/s)Resistance

(Ohms)

J

Cathode Anode

e- i+J

VE V E

6



Forces on a Charged Particle

• To examine how to use electrical energy to

accelerate a propellant, consider acceleration

of a particle with mass m and charge q

pBuqEqdt

udm

Electrostatic

Force

Lorentz

ForceCollisional Force

(Momentum Transfer)

E

qe

F

+

Bq mF

+

u

u

B

mF

Elec. Field Mag. Field

(1)



Motion of Charged Particle in E&B Fields

• How does charged particle move in electric and magnetic fields

• Electric field only

– electron lighter, higher accel.

• Magnetic field only

– particle gyrates(centripedal accel.)

– radius of gyration

– frequency of gyration

– No work; B and F perpendicular

Em

q

dt

ud

Bum

q

dt

ud

E

+-

u

B

rg2qB

Bumr

g

m

qBg

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J

B

Induced Magnetic Fields

• In general, current flow induces a B field

– B field induced by

linear current

– B field induced

by coil

B

From Space Propulsion Analysis

and Design, Humble, Henry and

Larson, 1995



Motion of Charged Particle in E&B Fields

• Crossed E and B fields

– E field accelerates

particle upward

– B field causes

acceleration

perpendicular to u

– overall result is drift

velocity normal to E and B

E

B

du

heavy

light

2B

BEu

d

mrg mg 1

8



Plasmas

• Gas composed of equal “amount” of negatively

and positively charged particles

– electrically neutral

– negative particles usually e-

– positive particles are positive ions

– typically most of gas molecules remain

neutral

• Momentum equation for plasma

Bjpuut

u

Current Density

(A/m2) 321

,, uuuu



Induced E Fields – Hall Effect

• Electron acceleration

– lighter e- accelerate more quickly

and accommodate to flow field (most of j)

– collisional coupling (momentum) of electrons

and heavies (ions and neutrals) is weak

plasma

resistivity (m)

en

p

en

BjBujE

e

e

e

Hall Effect

accelerates heavy

particles in charge

neutral plasma

2envmecee

(2)

Induced

E field due

to plasma

motion

electron

pressure

termcollision freq.

electrons with heavies

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Electric Rocket Propulsion

B. Examples



Electrothermal Rockets

• Resistojet (1W-10kW)– walls hotter than gas,

lower Tmax

– can combine withchemical heating, e.g.,hydrazine monopropellant

• Arcjet (1 kW- 10 MW)– high Tarc>Twall

(1-4104 K)– must mix with

rest of gas– very high To,

frozen flow losses in nozzleSpace Propulsion Analysis and Design,

Humble, Henry and Larson, 1995

Insulation

arc unstable

10



Resistojets

• Heating methods

– flow over coils of wire

– flow through hollow tubes

– flow over heated cylinder or plate

• Material limits temperature/performance

– conducting materials < 2700 Krhenium, refractory metals/alloys (tungsten, tantalum, molybdenum), platinum, cermets

– max Isp ~ 300 sec

• Power supply

– AC or DC

– Power TcmRVQp 2



Arcjets

• Gas heated by electrical discharge (arc)

– electrically conducting gas: plasma

• Joule heating

• Nozzle and electrode erosion

– high temperature, current arc

• Isp for H2 ~1200-1500 sec

TcmEjQp

enpjBujjee

2en

p

en

BjBujE

e

e

e

Hall Effect

normal to j

Resistive Heating

Work fromEM force

enpjBjujee

2Work from

pressure grad.

from (2)

no work

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Arcjets for Space Missions• Application designs

– from station keeping of moderate-sized spacecraft (few kW, hydrazine)• 1st satellite station keeping application - December 1994 - Lockheed Martin Astro Space

Series 7000 Telstar 401 satellite (owned by AT&T)

– to piloted Mars mission (100 kW, hydrogen)

• Advantages

– high specific impulse, approximately 600 s (hydrazine), 800 s (ammonia, TRW/AFRL

ESEX project) to 2,000 s (hydrogen)• compared to that used hydrazine decomposition thrusters ( Isp ~220 s) and

resistojets (Isp ~300 s)

• Improves propellant efficiency and allows satellite to carry larger payload instead of the

additional propellant normally required for long duration missions

– similarity between hydrazine arcjets and hydrazine thrusters and resistojets already

integrated into spacecraft for station keeping

• Drawbacks

– excessive heating limited life and/or advanced materials

– higher frozen flow losses in nozzle (high T), equivalently low thrust efficiencies

(~35% of input power is converted to useful thrust)

– possible plume contamination



1.8 kW Hydrazine Arcjet

Space Propulsion Analysis and Design,


• Telstar IV application– stationkeeping

• Hot gas from catalytically decomposed liquidN2H4

• Heated valveprevents N2H4

from freezing

• Isp~570 sec

• ~230 mN(~40mg/s)

12



Hydrazine Arcjet

1999 Test Firing of a Primex (now Aerojet)

MR-510 Hydrazine Arcjet

A shipset of Aerojet MR-510 Hydrazine

Arcjets and a Power Conditioning Unit

• Thrust 222-258 mN

• Input 4340 W (into Power Supply [PCU] for 2

Power arcjets); 2000 W input to each arcjet

from PCU)

• Isp > 570-600 s

. • 16 Lockheed Martin A2100

spacecraft (satellites) with MR-510

systems have been launched

• Each spacecraft has 4 thrusters and

1 PCU; 2 at a time used for N-S

station keeping orbit maneuvers



Arcjet with Applied Magnetic Field

• High current density in arctends to destroy electrodes

• Add soleniod B Field– adds azimuthal drift

velocity to arc– improves azimuthal

symmetry of gas temperature

– reduces local electrodeheating, reduces erosion

– needed for high powers(100 kW)

• Microwave discharge– alternate to arc heating



13



Electrostatic Thrusters• Ion engine example

• Near vacuum required

• Ion source

– usually electronbombardmentplasma

– also ion contact: propellant flowingthrough hot poroustungsten

– field emission: chargedspray droplets/particles

• Electrons added toneutralize exhaust

– thermionic emitters



L

A

Accelerator



Electrostatic Thruster Performance• Maximum exhaust velocity (Isp)

theoretically limited by voltage

difference across accelerator

• High thrust requires high mass

flowrates and therefore current

(and number, n) densities

• Ion current limited

by space-charge

– field from lots of ions

overcomes applied

E field

inletexite

VVm

qu 2

enquj

2

23

max

2

9

4

L

V

m

qj o

s

m

MW

V )volts(800,13

for singly ionized

meter

Farado

1210854.8 permittivity

of free space

Child-Langmuir

Law

accelerator electrode

spacing

22

23

8

maxm

)volts(1044.5

m

Amps

LMW

Vj

for singly ionized

14



Electrostatic Thruster Thrust

• Maximum thrust

• High requires high V and aspect ratio

– space chargeD/L ~1

– use many small ion beams toget more thrust

ee

uqmjAum

eu

q

mAj

maxmax V

m

q

q

mA

L

V

m

qo

22

9

42

23

22

max98 LVA

o

A=cross-sectional area flow

for circular cross-section

of diameter, D 22

max92 VLD

o

NewtonsVLD in1018.6 2212

DL



Electrostatic Thruster Thrust

• For fixed specific impulse

– best propellant for high thrust

• heavy molecules (particles)

• xenon (Xe) good choice (MW=131.3)

espuI

qmA

max

sp

accelo

Iq

mA

L

V

m

q2

23

max

2

9

4

15



Electrostatic Thruster Power

• Electrical energy converted to kinetic

energy with some efficiency

• Must also account for energy used to

create ions – ionization losses

– minimum loss given by ionization

potential for atom times current

VjAVIPower

2

2

1econv

umVI

qm

mIP

IIlossion



Electron Bombardment Xe+ Engine Example

• Operating conditions

– V=700 V, L=2.5 mm, 2200 holes (grids)

each with D=2.0 mm

– MW(Xe)=131.3, I(Xe)=12.08 eV

• Determine

– ,

– ue, Isp

– m

– minimum power required (including

ionization loss)

.

16



Solution

2212

max1018.6 VLD

gridN /1094.17005.2/21018.6 62212

mNgridNgridstotal

3.4/1094.12200 6

,max

smsmMWVue

3.131700800,13800,13

smue

860,31 sIsp

3248

sgume

41034.1

mmqumPPP Ielossionexhaust 2

min 2kg

kgMWm25

27

1019.21067.1

Cq 1910602.1

for singly ionized molec.

WW 18.19.67

WP 1.69min maximum effic. =67.9/69.1

= 98.3%

practically < 60-70%

flown

engines up to

~100-200mN

flown engines up to

few kW



Hall Thruster

• Hall effect most noticeable at low particle densities

• Stationary Plasma Thruster (SPT)

– developed in Russia

– 10’s kW

– axial current flowacross radial Bgenerates azimuthal electron flow

– induced (axial) Hall Eaccelerates ions

– also called “gridless” (neutral) ion thruster

– 1500-2000 s Isp with >50% efficiency using Xe and no oscillations like MPD Space Propulsion Analysis and Design,


17



Electromagnetic Propulsion Systems

• Use applied or induced magnetic fields to

produce acceleration of propellant

– high currents/

powers required

to produce

significant

induced fields

– high power available

only (normally) in

pulsed operationSpace Propulsion Analysis and Design,




Pulsed Plasma Thruster

• Propellant produced by vaporizing solid material with discharge

• B field induced by discharge also acts to accelerate vaporized propellant

• Advantage of simplicity

• Accelerationforce~ j2

(discharge current)



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Magnetoplasmadynamic (MPD) Thrusters

• Resemble arcjets

• Lower flow densities to attain higher exhaust velocity

• Diffuse discharge,low erosion

• Self field requires high J

• Applied B field

– allows higher Vat lower discharge currents

– increase accel.

– larger Hall effect

Self-Field MPD Thruster

Applied-Field MPD ThrusterSpace Propulsion Analysis and Design,




MPD Thrusters (con’t)

• Most efforts focused on applications with

exhaust velocities (Isp) greater than arc jets

• Typically require higher powers than currently

available on in-space vehicles

• Exhaust speed

– limited by erosion

and oscillations at high j

mjue

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electric propulsion for rockets - prof. seitzman...

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