2. Electric PropulsionOverview & Hall Thruster

What is Electric Propulsion?

“The acceleration of gases for propulsion by electrical

heating and/or by electrical and magnetic body forces.”

(Physics of Electric Propulsion, Robert G. Jahn, 1968)

For Ve above 10,000 m/s, a propulsion system needs;

Extremely high temperature of an insulated gas

stream

or

Direct acceleration by applied body forces

EP thrusters are very “fuel efficient” due to high exhaust

velocity Ve.2

Arcjet thruster

(5-10km/s）

Hall thruster

(15-25km/s）

Ion thruster

(30-50km/s）

MPD thruster

(50-100km/s)

How to accelerate?

Electrothermal Electrostatic Electromagnetic

3

How much mass reduction expected?

History of Himawari-5 (Metrological Sat.)

1995 Mar. Launched

2000 Mar. Life-time (5 years)Expired

2003 May Taken over by retired GOES-9

2005 Feb. Himawari-6

2006 Feb. Himawari-7

Momo ⇒ Ion Engine

Propellant for NSSK 207kg(28%) ⇒ 21kg

Propellant for OTV 402kg(54%) ⇒ 40kg

Total mass 747kg ⇒ 199kg

‘Whether satellite Himawari-5’

Hydrazine monopropellant thruster

4

How Large ∆V mission is possible?

Large ∆V = long-range missions (interplanetary flights)

long-time missions (maintenance of satellite)

Ex. Orbit to Mars ∆V = 14 km/s

Typical ion thruster Ve = 30 km/s

5

i f eexp 1.6m m V V

i f eexp 106m m V V

Cf. Typical chemical Ve = 3 km/s

Electric Propulsion vs Chemical Propulsion

Chemical Ve = 4.5 km/s

Electric Ve = 30 km/s

6

Power Supply – The higher the Ve is, The Better?

All EP thrusters must have electric power supply on-board the spacecraft.

The weight of power supply scales monotonically with the power level , which is directly related to Ve.

There exists an optimal Ve for a given ∆V and thrust level.

7

Total Mass vs. Ve (Isp).

Transfer time

Final net mass delivered to GEO as a function of LEO-to-GEO trip time for various propulsion options based on a 1550 kg Atlas IIAS-class payload with 10 kW power. (AIAA-95-2513)

8

2prop e

ths2

m V

P

prop propm m : Transfer Time

Optimum exhaust velocity for OTV by Electric Propulsion (1)

Thrust Efficiency

Propellant consumption rate

Ps: Available Power

Specific power of solar cell panels

α=Ps/mpanel (W/kg)

9

β= mpanel /Ps (kg/W) Typical β=0.05 kg/W

Optimum exhaust velocity for OTV by Electric Propulsion (2)

e,opt 0 thV V Payload ratio and exhaust velocity

pay props

i i i

prop s

i prop

2e

e th

2e

e th e

1

1 1

1 1 exp 12

exp 1 exp2

m mP

m m m

m P

m m

VV

V

VV V

V V

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0.1 1 10P

ay

loa

d r

ati

o

Ve/Vo

ΔV/V0=0.050.1

0.2

0.3

0.5

0.7

10e,optV 10km/s for 30 days, 33km/s for 1 year.

History of Electric Propulsion

Brief History of EP – It’s An Old Idea.

In 1906, R. H. Goddard expressed informally many of the key physical concepts of EP.

In 1911, Konstantin Tsiolkovskiy proposed similar concepts as Goddard.

In 1929, Herman Oberth included a chapter on EP in his classic book “Man into Space.”

In 1950s, Ernst Stuhlinger summarized his studies in his book “Ion Propulsion for Space Flight.”

In 1960s, EP became a major component of space propulsion development. Ion thrusters in US Hall thrusters in the former Soviet Union

12

In the United states in ’90s

Intelsat

Telstar

PAS-5

Intelsat(’93-’96) :40 Resistjets

(Ford Aerospace-SS/Loral)

Telstar (‘93-’95) 24 Arcjets

(GE Aerospace-Lockeed Martin)PanAmSat (‘90-’00)

60 Ion thrusters

(Hughes-Boeing)

Aerojet 550W Arcjet

Boeing NSTAR Ion engine

13

In Russia and Europe

More than 100 Hall thrusters on Cosmos & Meteor series from ’70s.

Russia

SPT100 D-55

Astrium and Alcatel are competing and cooperating for Hall thrusters in 2000s.

Europe

PPS1350 SMART-1

14

In Japan

ETS-series

MELCO Ion thruster ISAS Ion Thruster

On Hayabsa explorer

Ion thrusters No. Operation Satellite

NASA NSTAR Ring Cusp 1 16kh DS1

BoeingXIPS13

XIPS25Ring Cusp

52

24

55kh

14kh

BS601HP

BS702

AstriumUK10

RIT10

Kaufman

RF

2

3

0.7kh

7.7kh

Artemis

Artemis, EURECA

MELCO IES Kaufman 8 0.2kh ETS6, COMETSISAS/NEC μ10 Microwave 4 40 kh HAYABUSA

Type

15

Renewed interests in EP in recent years

More on-board electric power In-space propulsion for manned asteroid/planet

exploration All Electric satellites

16Boeing satellite

Courtesy NASA

Summary of EP overview

So, Do We Really Want an EP System?

In general, large ∆V missions are appropriate for using EP systems.

Factors in determining a suitable propulsion system: Ve, exhaust velocity EP Thrust CP Efficiency EP Power subsystem CP Propellant management subsystem ---- Mission trajectory and time CP/EP Interface with other subsystems CP Reliability (heritage, redundancy) CP Operation of spacecraft (orbit, trajectory) CP/EP Cost CP

(advantage for) EP: electric propulsionCP: chemical propulsion 18

Operation of a GEO Satellite

Propulsion System

NSSKtime/firing

NSSK CycleEWSK

time/firingEWSK Cycle

20-N Bi-prop(x 2 units)

A few minutes1 per two

weeksA few seconds

2 per two weeks

SPT-100/Bi-prop

12 hrs. by SPT-100

2 per dayby SPT-100

A few secondsby a Bi-prop

2 per dayby a Bi-prop

Orbit Control Maneuvering of chemical and Hybrid GEO satellites

19

MBSAT: DBS communications satellite for Japan and Korea

built by SS/L, 2004. The first US-built satellite with SPT-100.

Operation of a GEO Satellite (2)

• For the bi-prop system, 3 maneuvers in 2 weeks.

• For the EP/Bi-prop hybrid system, 56 maneuvers in 2 weeks.

• The hybrid system requires orbit control maneuvering every day due to its low thrust SPT-100.

• One NSSK maneuver takes a few minutes for the bi-prop system and 12 hours for the hybrid system.

• Orbit control ∆V: 95% for NSSK, 5% for EWSK

• Large electric power (~ 2 kW) from both the solar cells and batteries is supplied for the hybrid system.

Complex battery charging management, especially during the eclipse periods in spring and autumn.

20

Cost of EP systems (1)

As a simple comparison, the number of orbital maneuvers for a GEO communications satellite by the hybrid system is 18 times that by the bi-prop system.

An orbital maneuver is proceeded by a ranging, orbit determination, and control planning and followed by another ranging, orbit determination, and assessment.

A rough estimate of cost increase in orbit control operations (with one operator + one orbital engineer) is ~ $8 million in 15 years (orbital lifetime).

21

Cost of EP systems (2)

MBSAT carries the same number of bi-prop thrusters as the satellite that does not carry SPT-100s.

MBSAT: 12 bi-prop thrusters + 4 SPT-100 thrusters

Chemical propulsion satellite: 12 bi-prop thrusters

~ $20 million cost increase due to four SPT-100 thrusters plus gimbals, PPU, and PFS

Trade-off between the bi-propellant mass and the mass of SPT-related hardware and xenon.

The decrease in launch cost is minimal.

Having an EP system, with less propellant mass and longer mission lifetime, would NOT REALLY save the total cost when considering actual satellite procurement and operations!!

22

EP CAN….

can have an advantage with their low thrust level. Extremely small attitude disturbance by their maneuvering. Suitable for satellites with flexible structures. Suitable for missions with fine maneuvering requirement such

as co-location and formation flying. can appeal to missions of extremely high ∆V requirements.

Impractical to carry the necessary propellant mass of a chemical rocket system

Interplanetary missions Deep-space probes

can continue to evolve. Higher reliability Lower cost Greater variety of the combinations of thrust, Isp, and power

levels. 23

Hall Thruster

Hall Thruster Principle: Overview

Electrostatic acceleration

Axial E and radial B

Cathode as electron emitter

Design criteria:

25

BE

Anode

Acceleration

channel

Xe+

Propellant

(Xe)

Hall

current

electron

Cathode

(Hall parameter)

(Larmor radius)

(Ion mean free path)

: Channel length

Hall Thruster Principle: Electric Field

100 ~ 1000 V DC

Plasma is highly

conductive

→ Electric field is shielded

26

Plasma

Channel wall

An

od

e

Cath

od

e s

ideE field

B fieldrz

Radial magnetic field to

confine electron motion

Electron gyroradius must

be much smaller than

channel length

Electron Larmor radius Anode Channel exit Cathode

Φ

Br

Hall Thruster Principle: Hall Current

E x B drift movement of

charged particles

Electron-induced Hall

current

For effective confinement of

electron motion, electron

Hall parameter must be

much larger than 1

(Hall parameter) =

(Gyrofrequency) x (mean free time)

Electron

E

B

E x B direction

electron motion

Hall current

Electron motion with collisions

E

B

Collisions contribute to

cross B-field transports

27

Hall Thruster Principle: Closed Annular Cavity

B

E

Hall current

Operation of Hall thruster

Linear Hall-effect ion source1

1https://www.ulvac.co.jp/ 28

Hall Thruster Principle: Ion Generation

Electrons are very lightweight

Electrons are heated to

20 eV = 232,000 K

Collisional ionization

First ionization energy [eV]:

Argon: 15.8

Krypton: 14.0

Xenon: 12.1

Anode Channel exit Cathode

Te

Φ

Energy gain:

Joule heating

Energy loss:

Ionization,

Wall loss, etc.

Ionization

region

Xe

Electron

Xe+

29

Hall Thruster Principle: Ion Acceleration

Ions are smoothly

accelerated by electric field

Little disturbance by

magnetic field or collisions

Ion Larmor radius Ion mean free path

Exhaust velocity estimation

Discharge voltage:

Acceleration voltage:

Ion mass (Xe):

Channel wall

An

od

e

rz

Ion

If 90% of is ionized,

30

Elemental charge:

Hall Thruster Principle: Recap

Electrostatic acceleration

Axial E and radial B

Cathode as electron emitter

Design criteria:

31

BE

Anode

Acceleration

channel

Xe+

Propellant

(Xe)

Hall

current

electron

Cathode

(Hall parameter)

(Larmor radius)

(Ion mean free path)

Magnetic Layer vs. Anode Layer

32

Magnetic-layer type Channel length > channel width

Dielectric material (BN) to protect ferromagnetic pole pieces

Anode-layer type Very short channel length, so that no need for magnetic pole protection

Conducting metal at the cathode potential

Stationary Plasma Thruster (SPT)

33

Parameter SPT-50 SPT-70 SPT-100 SPT-140

Slot diameter, mm 50 70 100 140

Thrust input power, W 350 700 1350 5000

Average Isp, s 1100 1500 1600 1750

Thrust, mN 20 40 80 300

Thrust efficiency, % 35 45 50 >55

Lifetime, h 1500 3000 9000 >7000

Status Flight Flight Flight Qualified

SPT series specifications

Thruster with anode layer (TAL)

34

Parameter D-38 D-55 D-100 D-110 D-150

Slot diameter, mm 38 55 100 110 150

Thrust input power, W 1500 2500 7500 15000 17500

Average Isp, s 2000 2000 2100 1750 2300

Thrust, mN 100 120 340 240 800

Thrust efficiency, % 70 55 60 60 60

TAL series specifications

Thrust Performances of SPT and TAL

35

Outer diam.: 100mm

Power : 1.35 kW

Isp : 1600 sec

Thrust : 80 mN

Efficiency : 50 %

Magnetic layer type: SPT-100

Outer diam.: 55mm

Power : 2.5 kW

Isp : 2000 sec

Thrust : 120 mN

Efficiency : 55 %

Anode layer type: D-55

Global Exploration Roadmap (GER)

36

GER: Post-ISS international cooperative missions

High-power EP (In-space propulsion) is a key tech.

In-Space Propulsion in US and Europe

37

NASA’s 12.5-kW HERMeS thruster1 Snecma’s 5-kW PPS®50002

1Kamhawi et al., IEPC-2015-07, 2015. 2Duchemin et al., IEPC-2015-127, 2015.

Thrust efficiency : ~ 63%

Specific impulse : ~ 3000 s

RAIJIN: All-Japan Collaboration

38

Plume Interaction

High-Current Neutralizer

Mission & Orbit Analyses

Project RAIJIN

Thruster Head, PPU

Development of UT-58 Anode-Layer Thruster

UT-58 3D Model

Magnetic Field Analysis

39

Thermal Analysis

UT-58: 2 kW-level TAL developed in UT

Performance of UT-58

40

Xenon Argon

Operation of UT-58 with Xenon and Argon

Maximum value Xenon (~3.0 Aeq) Argon (~8.0 Aeq)

Input power, kW 1.92 2.81

Thrust, mN 92.2 51.8

Anode efficiency, % 73.0 23.5

Exhaust velocity, km/s 23.0 18.6

Comparison: UT-58 and D-series

41

Compared with Russia’s

D-series, UT-58 achieved

70% of anode eff. with

1-order small input power

Anode efficiency vs. Input power of UT-58 and D=series

Ion Energy Loss to Wall and Wall Erosion

Ion collision with wall causes

energy losses of ionization

and acceleration

Channel wall erosion by high-

energy ion sputtering limits

thruster lifetime

42

Channel wall

An

od

e

rz

Ion

Channel wall

Sputtering

Magnetic pole pieceSimulation of channel wall recession

Equipotential

lines

B Field Optimization: Plasma Lens Focusing

Plasma

Highly conductive

B field

E field

Metal

43

Electric field lines surrounding

conductorPlasma lens focusing

B Field Optimization: Actual Designs

44

ATON thruster

Magnetic field geometry using

plasma lens focusing

B Field Optimization: Numerical Simulations

45

330 h 600 h

1Cho et al., 48th Joint Propulsion Conference, AIAA 2012-4016, 2012.

Calculated space potential distribution.

Begin of life and end of life.

Ion Attracting Physics: Presheath

Acceleration in presheath

Bohm criterion

46

Convex-shaped magnetic field and

ideal ion acceleration direction

Deflected equipotential lines and

ion-loss region by presheath

Equipotential line

Ion-loss region

Outer channel wall

Inner channel wallInner channel wall

Outer channel wall

Channel wall

Magnetic Shielding: Theory

47

Anode Channel exit Cathode

Te

Z

Magnetic Shielding:

Very concave magnetic field

geometry

Channel wall parallel to

magnetic lines

Φ

High Te

Strong presheath

Low Te

Weak presheath

Magnetic Shielding: Numerical Simulation

48

Unshielded configuration1

Magnetic shielding configuration1 Calculated space potential distribution2

1Hofer et al., Journal of Applied Physics, 2014. 2Mikellides et al., Journal of Applied Physics, 2014.

Magnetic Shielding: JPL Experiments

49

Measured erosion rate of unshielded (US) and magnetic shielding (MS)

configurations1

1Hofer et al., Journal of Applied Physics, 2014.

50

Better 0

2

4

6

8

10

12

14

20 30 40 50 60 70

Gu

ard

-rin

g c

urr

ent

rati

o, %

Anode efficiency, %

gap=14

gap=19

gap=24

Better

Current trade-

off relation

New trade-off

relation

Discharge current

Ion wall currentGuard-ring

current ratio =

UT Experiment: Magnetic Shielding in TAL

Anode Wall

gap = 14 mm

gap = 19 mm

gap = 24 mm

Magnetic and wall geometry

of UT-58

Str

on

g m

ag

ne

tic s

hie

ldin

g

Simulation: Magnetic Shielding in TAL

51

Weak magnetic shielding Strong magnetic shielding

Calculated space potential distribution in UT-58

52

Simulation: Magnetic Shielding in TAL

All ions Ions entering channel wall

Calculated ion production rate distribution of

magnetically shielded UT-58

Discharge Current Oscillation

53

Time history of discharge current

1. Ionization: 10 kHz - 100 kHz

2. Transit-time: 100 kHz - 1 MHz

3. Electron-drift: 1 MHz - 10 MHz

4. Electron-cyclotron: 1 GHz

5. Langmuir: 0.1 GHz - 10 GHz

Inherent oscillation types:

Discharge oscillation causes

Heavy power processing unit

Electromagnetic noise

Ionization Oscillation

Ionization rate

“Predator-Prey” model

54

Principle of ionization oscillationTime

Nu

mb

er

den

sit

y

Frequent

ionization

Neutral

depletion

Infrequent

ionization

Neutral

recovery

Neutral, nnIon, Electron ne

Synchronous Breathing-mode Oscillation

55

Oscillation behavior of plasma1.

1Yamamoto et al., Trans. Japan Soc. Aero. Space Sci. (48), 2005.

Electron Diffusion and Discharge Oscillation

Typical tendency of discharge current and oscillation

vs. magnetic flux density

Classical

diffusion

Transition

region

Anomalous

diffusion

0.4

0.01

Anomalous

diffusion

Transition

regionClassical

diffusion

56

Electron Diffusion and Discharge Oscillation

Xenon

60

10

24

1

Argon

Classical diffusion

region

Transition and

anomalous diffusion

region

57

Summary: Hall Thruster

Principle: Hall current and electrostatic acceleration

High-power thruster for In-space propulsion

B-field optimization and magnetic shielding

Discharge current oscillation

58