Thruster Options For Microspacecraft A Review And Evaluation Of Existing Hardware Emerging Technologies UCCS Temperature Control Mueller Microprop Paper

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Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

AIAA

97-3058

Thruster

Options

for

Microspacecraft:

A

Review

and

Evaluation

of

Existing

Hardware

and

Emerging

Technologies

Juergen Mueller*

Jet

Propulsion Laboratory

California

Institute

of

Technology

Pasadena,

CA

91109

State-of-the-art

thruster technologies

are

reviewed

and

evaluated

in

view

of

potential

microspacecraft applications. Microspacecraft

are

defined

in

this study

as

spacecraft with

masses

between

1 and 20 kg.

Based

on

this review

of

existing technologies,

future

development

needs

for

micropropulsion

systems

are

defined

and

advanced

new

micropropulsion

concepts

especially designed with microspacecraft applications

in

mind

are

introduced.

Of the

state-of-

the-art

technologies, hydrazine thrusters

and

small solid rocket motors appear applicable

t o

some

microspacecraft.

Cold

gas

systems

may

provide near-term solutions

to

attitude

control,

at

the

expense

of

leakage concerns

and

large

and

heavy

propellant tankage.

New

thruster concepts,

heavily relying

on

advanced microfabrication technologies have been designed

and

built

at

JPL,

addressing

some

of the

microspacecraft design challenges,

and are

introduced.

I.

INTRODUCTION

Background

and

Significance

Within

the

National Aeronautics

and

Space

Administration

(NASA),

a

research

and

development

initiative

is

currently underway

to

investigate

the

feasibility

of

microspacecraft

in the

1-20

kg

class1.

The

motivation

behind

this development

is the

desire

to

reduce launch

masses

in

order

to

reduce

mission

costs

and

greatly increase

launch

rates. Launch costs

for a

typical interplanetary

mission

may be as

high

as 30% of the

overall mission cost,

and

these costs

may be

reduced significantly

as a

result

of

substantially

reduced spacecraft

masses.

In

addition, microspacecraft mission scenarios

may

be

envisioned where, rather than launching

a

single large

spacecraft,

the

mission

is

accomplished

by a

fleet

of

several

smaller microspacecraft, with

the

scientific payload

distributed among

the

micro-craft

to

reduce

mission

risk.

Loss

of one

microspacecraft would

not

eliminate

the

entire

mission.

A

fleet

of

several small microspacecraft, possibly

in

connection with

a

larger

"mother"-spacecraft,

could also

increase mission flexibility.

For

example,

the

smaller

microspacecraft could

be

placed

on

different

trajectories

around

the

target planet

and

provide

an

almost

instantaneous,

global

survey

of the

target.

A

"mother-craft" could also

Member

of

Technical

Staff,

Technology Group,

Member

AIAA

Advanced

Propulsion

release

smaller

micro-craft

to

perform

riskier

portions

of a

mission.

For

example,

a

close-up investigation

of

Saturn's

ring

objects

may be

envisioned2,

with

a

swarm

of

microspacecraft

decending

into Saturn's rings while

the

"mother-craft", providing high-data rate communication

to

earth

via a

large high-gain antenna,

may

cruise

in a

safe

distance

from

the

rings.

Building

microspacecraft

in the

1-20

kg

class,

however, will necessitate

the

miniaturization

of

every

subsystem

in

order

to

maintain

the

high degree

of

onboard

capability required

to

ensure

an

acceptable scientific return

for

the

mission.

One of the

sub-systems that will

be

included

in

such

a

reduction

in

weight

and

size

is

propulsion. Although

in the

past many very small spacecraft

have lacked propulsion

systems

altogether,

future micro-

spacecraft

will likely require significant propulsion

capability

in

order

to

provide

a

high degree

of

maneuverability

and

capability3.

In

particular, interplanetary

mission scenarios will require propulsion capability

on

microspacecraft

for

course

corrections

as

well

as

attitude

control

to

accurately point

the

spacecraft

for

observation

or

communication3.

Attitude control

in low

Earth orbit

is

possibly achievable

via

other means, such

as

magnetic

torquers,

however propulsive capability

is

required

in

higher

orbits,

interplanetary

space

or

around

some

other

planet,

either

to

directly control

the

spacecraft's attitude

or to

off-

load

momentum

wheels3.

In

addition, very small spacecraft

are

often

launched

in a

"piggy-back" configuration together

with

larger spacecraft

to

save launch

costs.

Propulsive

Copyright

©

1997

by the

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

The

U.S.

Government

has a

royalty-free

license

to

exercise

all

rights

under

the

copyright

claimed herein

for

governmental

purposes.

All

other

rights

are

reserved

by the

copyright

owner

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

capability

may be

required

onboard

the

microspacecraft

to

adjust

its

trajectory

according

to the

desired

mission

objective3.

In

order

to

meet

microspacecraft

propulsion

requirements,

the use of

lightweight,

small

sized,

low-

thrust

and

small

impulse

bit

(I-bit)

systems

will

be

needed.

It is the

purpose

of

this

paper

to

review

and

evaluate

existing

propulsion

hardware

and

emerging

micropropulsion

technologies

with

respect

to

their

applicability

to

microspacecraft.

This

paper

is an

extension

of an

earlier

study

performed

by the

author

at the Jet

Propulsion

Laboratory

(JPL),

distributed

JPL-internally

in

late

19954.

The JPL

study

contained

company-discreet

information

and

was

therefore

not

available

for

public

release.

In the

paper

presented

here,

company-sensitive

information

has

been

eliminated

and new

information

gained

in the

emerging

field

of

micropropulsion

since

the

earlier

study

has

been

added

Definitions

There

exists

a

wide

variety

of

opinions

regarding

the

appropriate

definition

of

what

a

microspacecraft

is.

Table

1

below

gives

a

definition

with

respect

to

mass,

size

and

power

of the

type

of

microspacecraft

that

are

considered

in

this

study.

To

simplify

this

discussion,

three

microspacecraft

classes,

Class

I

through

III,

have

been

defined.

These

microspacecraft

classes

distinguish

themselves

from

each

other

by

their

mass,

power

and

size

ranges.

These

should

be

interpreted

as

approximate

values

.

Class

I

spacecraft,

ranging

in

mass

from

about

5 -

20 kg, are

characterized

by the

fact

that

they

may

still

be

able

to use the

smallest

propulsion

hardware

available

today,

or

currently

under

substantial

development.

This

hardware

will

be

conventionally

integrated

by

interconnecting

it via

conventional

feed

lines.

Development

of

some

new

propulsion

hardware,

taking

miniaturization

to new

extremes

and

possibly

incorporating

advanced

microfabrication

techniques,

such

as

MEMS

(Micro-electromechanical

S_ystems)

technologies,

may

also

be

required.

Figure

1

shows

an

example

of a

Class

I-type

spacecraft,

designed

and

built

at JPL and

referred

to as the

JPL 2nd

Generation

Microspacecraft1'5.

This

craft,

in its

current

design

iteration,

has a

mass

of

around

7-8 kg. The

2nd

Generation

Microspacecraft

is not

designed

for

flight,

but,

rather,

it is an

evolutionary

functional

model

of

such

a

craft,

with

subsystem

hardware

constantly

being

upgraded

to

more

"flight-like"

versions5.

Based

on the 2nd

Generation

Microspacecraft

design

regime,

power

densities

for

microspacecraft

have

been

estimated

at

IW/kg,

resulting

in

10 W and 20W

onboard

power

for the two

Class

I

type

microspacecraft

(10 kg and 20 kg)

listed

in

Table

1.

Low-

and

high-mass

versions

of a

Class

I

microspacecraft

will

be

considered.

Microspacecraft

with

masses

between

1 and 5 kg

have

been

categorized

here

as

Class

II

microspacecraft.

In the

case

of

Class

II

microspacecraft,

development

of

new,

extremely

miniaturized

propulsion

components,

both

for

delta-v

maneuvers

and in

particular

for

attitude

control,

Table

1:

Definition

and

Classifications

of

Microspacecraft

for the

Purpose

of the

Study

Micro

S/C

Class

S/C

Mass

(kg)

S/C

Power

(W)

S/C

Dimension

(m)

Comments

I

20

10

20

10

0.4 Use

conventional

components,

possibly

MEMS.

Conventional

integration

(feed

lines).

0.3

Same

as

above.

II

0.1

MEMS

components,

high

level

of

integration

between

components

of

each

subsystem

and

possibly

between

______subsystems._______________

III

«1

«1

0.03

All

MEMS.

Very

high

level

of

integration

between

all

subsystems

and

within

sub-systems

required.

Strong

feasibility

issues.

Not

considered

in

this

______study._____________

__

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

Fig.

1:

JPL

Second

Generation

Microspacecraft

where

multiple

clusters

of

small

thrusters

are

used,

will

be

required.

These

devices

almost

certainly

will

employ

MEMS

technologies

in

some

fashion.

Also,

because

of the

severe

volume

constraints

on

such

a

spacecraft,

a

high

level

of

integration

will

be

required

between

different

propulsion

components,

and

between

propulsion

and

other

spacecraft

subsystems.

For

example,

in the

case

of

MEMS

based

technologies,

several

propulsion

components,

such

as

thrusters

and

valves,

plus

the

required

control

electronics,

may

be

integrated

onto

a

single

chip,

or a

3D-stack

of

chips.

Integration

approaches

of

this

kind,

not

limited

to

propulsion

however,

are

currently

being

addressed

by

JPL's

Center

for

Integrated

Space

Microsystems

(CISM).

This

requirement

for an

increased

level

of

integration,

in

addition

to an

even

more

pronounced

degree

of

miniaturization

over

the

smallest

availbale

state-of-the-art

propulsion

hardware

makes

this

Class

II

category

of

microspacecraft

different

in

its

design

requirements

from

the

Class

I

microspacecraft.

Even

smaller

microspacecraft

with

total

masses

of

significantly

less

than

1 kg,

categorized

as

Class

m

microspacecraft,

have

recently

been

studied

at

JPL6,

these

however,

will

not be

considered

in

this

study

since

the

design

concepts

discussed

in

Ref.

5 did not

require

propulsion.

If

propulsion

needs

should

arise

for

Class

ffl

spacecraft,

they

will

certainly

require

MEMS-based

technologies6.

These

propulsion

systems

would

likely

be

based

on

significantly

scaled

down

versions

of

MEMS-based

Class

II

systems.

Scope

of

this

Study

The

goal

of

this

study

is to

review

current

propulsion

technology

in

view

of its

applicability

to

Class

I

and

II

microspacecraft,

identify

future

technology

needs

and

to

outline

potential

future

thruster

technology

currently

emerging,

aimed

at

meeting

these

needs.

Microspacecraft

mission

scenarios

may

involve

a

variety

of

propulsive

maneuvers,

such

as

attitude

control,

course

correction,

delta-

v

maneuvers,

orbit

insertion

or

even

landing

and

take-off

from

a

distant

planet.

Depending

on the maneuver and

delta-

v

requirement,

different

propulsion

technologies

will

be

needed

.

This

paper

will

focus

on

relatively

low-thrust

propulsion

systems

that

can be

integrated

with

a

microspacecraft

bus of

either

Class

I or II

type

for the

purposes

of

attitude

control

and

delta-v

maneuvers.

If

take-off

and

landing

operations

are

considered

for

microspacecraft,

a

class

of

propulsion

devices

very

different

from

those

to be

considered

here

will

be

required.

Given

the

large

delta-v

requirements

associated

with

landing

and

take-

off

operations,

even

for

relatively

small

payloads,

fairly

large

chemical

stages

well

exceeding

the

mass

limits

considered

here

may

result.

As a

consequence,

high-thrust,

and

thus

high-flow,

devices

will

be

required.

The

need

to

sustain

high

propellant

flow

rates

will

not

allow

for

miniaturization

of

propulsion

components

used

in

these

applications

significantly

beyond

sizes

already

available

today

(although

significant

research

will

have

to be

devoted

to

such

areas

as

further

component

mass

reduction

and

alternate

propellant

usage).

These

devices

are not

considered

micropropulsion

devices

in the

context

of

this

study

and are

therefore

not

included

in the

following

discussion.

Micropropulsion

subsystems

will

not

only

consist

of

thrusters,

but

will

also

require

miniature

feed

system

components,

such

as

valves,

tanks,

and

pressure

regulators,

etc.

In

some

of

these

areas

considerable

design

challenges

arise

during

miniaturization.

In

particular

the

moving

parts

in

valves

make

miniaturization

difficult.

However,

already

the

thruster

material

to be

reviewed

is so

vast,

that

surveying

miniature

components

in

addition

to

thruster

technologies

could

not be

accommodated

in

this

study.

Evaluation

of

miniature

components

will,

thus,

not be

included

in

this

review.

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

H.

REPRESENTATIVE

MISSION

REQUIREMENTS

Requirements

for

microspacecraft

missions

are

difficult

to

predict

accurately

at

this

early

stage

of

their

development

and

will

vary

greatly

given

the

multitude

of

conceivable

missions.

In

this

section,

an

attempt

is

made

to

present

a

class

of

representative

mission

requirements,

based

on

our

current

understanding

of

these

small

spacecraft

requirements,

to

serve

as a

basis

for

micropropulsion

technology

evaluation.

Both

attitude

control

and

delta-v

requirements

(landing

and

take-off

excluded,

see

Section

I)

have

been

considered.

These

mission

requirements

are

based

on

estimates

provided

in a

recent

workshop

on

micropropulsion,

held

at

NASA's

Jet

Propulsion

Laboratory

(JPL)7.

As

concrete

future

microspacecraft

missions

develop,

this

set of

mission

requirements

will

no

doubt

have

to be

modified.

Delta-v

Requirements

Obviously,

delta-v

requirements

do not

depend

on

the

size

of the

spacecraft

and

therefore

requirements

for

four

representative

small

spacecraft

missions,

currently

under

investigation

at

JPL,

have

been

listed

below

in

Tables

2

through

5 to

serve

as a

reference.

These

missions

include

a

small

body

(asteroid)

rendezvous,

an

outer

planet

(Europa)

orbiter,

a

spacecraft

formation

flight

(DS-3)

and an

earth-

observing

cluster.

Inspecting

Tables

2 and 3, the

large

delta-v

requirements

for

deep

space

missions

becomes

immediately

apparent.

Electric

propulsion

applications

result

in

even

larger

delta-v-requirements

due to

burn

losses8,

which

must

be

offset

by the

higher

specific

impulses

and

more

fuel-

efficient

operation

of

electric

engines.

Use of

electric

propulsion

may

either

lead

to

shorter

trip

times,

or

reduced

spacecraft

masses,

or

both.

The

benefit

of

using

electric

propulsion

in

regard

to

spacecraft

mass

reduction

will

likely

be

even

more

important

for

mass

limited

microspacecraft

missions.

An

additional

requirement

for

chemical

primary

propulsion

is the

need

to

maintain

large

enough

thrust-to-

spacecraft

weight

ratios.

Values

around

0.1

- 0.3 are

typical.

Too

small

a

thrust-to-weight

ratio

will

again

lead

to

burn

losses8

and

increase

the

required

delta-v.

Since

in the

case

of

chemical

engines

this

increased

delta-v

requirement

cannot

be

easily

offset

by a

sufficiently

large

specific

impulse,

too low

a

thrust

for

chemical

primary

propulsion

maneuvers

must

be

avoided.

Too

large

a

thrust

value,

on the

other

hand,

may

generate

accelerations

too

large

to be

tolerated

by the

spacecraft

structure,

in

particular

at

times

well

into

the

mission,

when

portions

of the

spacecraft

structures

may be

deployed.

Attitude

Control

Requirements

In

order

to

estimate

attitude

control

requirements

the

following

assumptions

were

made7:

(1)

fine

pointing

requirements

are

assumed,

defined

by the

desire

to

stay

within

a 0.2 - 2

mrad

deadband

and ACS

firings

occuring

no

more

frequently

than

one

couple

firing

every

20 - 60

sec;

(2)

Slew

rates

of

180°/minute

required

with

one

couple

of

thrusters

firing.

The

spacecraft

was

assumed

to be

cubical

in

shape

with

the

side

of the

cube

being

equal

in

length

to the

dimension

listed

in the

forth

column

of

Table

1. The

resulting

minimum

impulse

bit and

minimum

thrust

requirements

are

listed

in

Table

6 for

microspacecraft

masses

of

1, 10 and 20 kg.

Very

small

impulse

bit

requirements

can

be

noted.

It

should

be

pointed

out in

this

context,

that

the

fine

pointing

requirements

given

above

are not

extreme

for

today's

spacecraft.

HI.

REVIEW

OF

THRUSTER

TECHNOLOGIES

In

this

section,

state-of-the-art

thruster

hardware,

either

available

"off-the-shelf

or

under

significant

development,

will

be

reviewed

and

evaluated

in

view

of

microspacecraft

applications.

This

section

is

structured

into

two

main

parts

,

focusing

on

primary

and

attitude

control

applications,

respectively.

Both

primary

and

attitude

control

sections

have

been

further

sub-divided

into

chemical

and

electric

thruster

sections.

Primary

Thrusters

-

Chemical

Bi-Propellant

Engines

Bi-propellant

engines

are

most

commonly

used

for

primary

propulsion

applications

of

conventional

spacecraft

today

due to its

relatively

high

specific

impulse

performances

and

considerable

flight

heritage.

Advantages

of

bi-propellant

engines

over

other

chemical

systems,

such

as

mono-propellant

thrusters,

are

their

higher

specific

impulse

performance,

leading

to

lower

fuel

weights.

Disadvantages

are

their

relative

complexity,

both

with

respect

to

engine

technology

and the

feed

system.

Since

separate

feed

systems

for

fuel

and

oxidizer,

as

well

as

pressurants

are

required,

component

part

count

is

high,

leading

to

large

propulsion

system

dry

masses.

Therefore,

bi-propellnat

systems

are

usually

used

on

mission

requiring

large

delta-v's

(>

1,000

m/s)

and

large

spacecraft.

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

Table

2:

Delta-v

Requirements

for a

Vesta

Rendezvous

Mission7

Delta-v

Requirements

(m/s)

Primary

_____________ACS'"

Mission

Duration

(yrs)

Chemical

3400*

Electric

7000"

2.5-25

2.5-25

Chemical

4.3

Electrical

2.7

Assumes

Mars

Gravity

Assist

(MGA)

trajectory

Assumes

direct

trajectory

Assumes

250

slews

(180°/min

) and

3-axis

stabilization

in the

range

of 1 - 20 kg S/C

mass,

thrust

0.5 - 10

mN.

Table

3:

Delta-v

Requirements

for a

Europa

Orbiter

Mission7

Delta-v

Requirements

(m/s)

Primary

ACS'"

Chemical

2500'

945

Electric

5500"

5-50

5-50

Mission

Duration

(yrs)

Chemical

4.8

Electrical

5.8

Assumes

Solar

Electric

Venus

Venus

Gravity

Assist

(Se-VVG)

and 345

ms/s

chemical

delta-v

required

for

Jupiter

Orbit

Insertion

(JOI)

and 600

m/s

for

Europa

Orbit

Insertion

(EOI).

Assume

500

180°

/min

slews

and

3-axis

stabilization

(S/C

masses

ranging

between

1 - 20 kg and

thrusts

between

0.5 - 10

mN).

ACS

requirements

may be

higher

for low

thrust

trajectories.

Table

4:

Delta-v

Requirements

for the

Deep

Space

Interferometer

Mission

(DS-3)7

Delta-v

Requirements

(m/s)

Primary________________ACS

Mission

Duration

(yrs)

100-300

0.5

- 1.0

Table

5:

Delta-v

requirements

for an

Earth

Observing

Cluster7

Delta-v

Requirements

(m/s)

Primary'_______________ACS"

Mission

Duration

(yrs)

Chemical

500

Electric

550

5-50

5-50

Chemical

5.0

Electrical

5.0

Primary

delta-v

assumes

200

m/s

for

non-Keplerian

orbit,

250

m/s

for

NSSK,

and 50 - 100

m/s

for

phasing.

ACS

assumes

500

slews

of

180°

/min

and

3-axis

stabilization

(S/C

masses

ranging

between

1 - 20 kg and

thrusts

ranging

between

0.5-10

mN).

Table

6:

Representative

Attitude

Control

Requirements

for

Microspacecraft7

S/C

Mass

(kg)

20

10

1

S/C

Dimension*

(m)

0.4

0.3

0.1

Ibit

(mNs)

0.013

0.005

0.0002

Tmln

(mN)

4.65

1.75

0.06

*

Assume

cubical

spacecraft

shape

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

Consequently,

most

bi-propellant

engines

that

have

been

built

today

provide

fairly

large

thrust

levels.

Some

smaller

engines

in the 5 - 22 N (1 -5

Ibf)

thrust

range

have

been

built

or are

under

development.

Up to

this

point,

applications

for

these

engines

have

been

envisioned

in the

use of

attitude

control

purposes

of

larger

spacecraft

in

order

to

simplify

the

overall

propulsion

sytem,

eliminating

separate

attitude

control

propellant

tanks.

Given

these

goals,

considerable

effort

was

devoted

to

fast

thruster

response

times

and

short

impulse

bits9"13.

Challenges

encountered

when

building

bi-propellant

engines

of

such

a

small

size,

include

combustion

efficiency

losses

due to the

potential

of

reduced

mixing

and

vaporization,

thermal

control

issues

of

chambers,

nozzle

throats

and

injector

heads

and

related

material

issues,

injector

design

issues

and

related

accurate

mixture

ratio

control

issues,

and

possibly,

spacecraft

contamination

issues

due the

potential

of

incomplete

mixing

and

vaporization

inside

the

thrust

chamber.

Mixing

and

vaporization

losses

can

occur

in

small

engines

due to the

small

chamber

size.

In

general,

good

vaporization

is

obtained

in

longer

chambers,

at

higher

chamber

pressures

and for

smaller

injector

orifice

sizes14,

while

better

mixing

is

achieved

in

engines

having

high

chamber-length-to~diameter

ratios

and a

larger

number

of

injector

inlets14.

In

addition,

engine

size

also

plays

a

role

in

the

mixing

of the

propellants14.

Smaller

engines

have

lower

chamber

flow

Reynolds

numbers

and

thus

lead

to

less

turbulent

chambers,

reducing

mixing.

The

limitations

imposed

on

chamber

length

and

diameter

has an

immediate

impact

on the

degree

of

miniaturization

of a

bi-propellant

engine.

Thermal

control

of

small

bi-propellant

engines

is

another

key

design

issue.

Film

cooling,

or

boundary

layer

cooling

(BLC),

is

often

employed

in

bi-propellant

engines

to

keep

the

chamber

wall

temperature

within

its

thermal

and

structural

design

limits.

Here,

a

fuel

is

injected

close

to the

chamber

wall.

Since

the

propellant

mixture

is

fuel

rich,

it

does

not

burn

completely

and

will

shield

the

chamber

walls

from

the

heat

output

of the

combustion

reactions

occurring

closer

to the

center

of the

chamber.

However,

at the

same

time

combustion

efficiencies

are

reduced

due to

incomplete

combustion.

While

for

more

conventionally

sized

engines

about

15 - 30 % of the

fuel

is

commonly

used

for

film

cooling,

these

values

may

reach

up to 30 - 40% for

smaller

engines

in the

22-N

class,

causing

performance

losses12

and

possibly

resulting

in

spacecraft

contamination

concerns

due

to the

possibility

that

liquid

fuel

droplets

may

attach

themselves

to

sensitive

surfaces

(optical

lenses,

solar

cells,

etc.).

Elimination

of

film

cooling

was

done

in the

small

bi-propellant

attitude

control

engines

developed

by

Rockwell

for

the

Kinetic

Energy

Anti-Satellite

(KE

ASAT)

program.10'11

This

increases

combustion

efficiency

and

decreases

injector

head

complexity

since

no

separate

BLC

holes

are

required,

and

should

facilitate

miniaturization.

However,

in

order

to

survive

the

punishing

thermal

environment,

high

temperature

chamber

materials

have

to be

employed.

In the

case

of the KE

ASAT

technology10-11,

a

carbon-silicon

carbide

chamber

was

used.

Despite

use of

this

high

temperature

material,

engine

burn

durations

in a

single

burn

where

limited

to

only

a

little

over

20

seconds.

Another

chamber

material

under

significant

investigation

is

rhenium-

iridium

composite

material.

Rhenium

is

uses

as the

substrate

material

because

of its

high

melting

point

(3453

K)12

and

coated

with

a

iridium

layer

for

chemical

inertness.

Iridium

has a

coefficient

of

thermal

expansion

(CTE)

closely

matching

that

of

rhenium

and a

high

melting

point

of

2727

K12.

Using

this

chamber

material,

specific

impulses

in

excess

of 300 sec

have

been

obtained

in a 22 N

thrust

chamber

over

burn

durations

of a

maximum

of 350

seconds12.

Injector

design

also

requires

careful

attention

in

small

bi-propellant

rocket

engine

development.

Due to the

small

flow

cross

sections

encountered

in

small

engines,

flow

rate

control,

and

thus

mixture

rate

control15,

as

well

as

misalignment

of

impinging

propellant

jets12

could

lead

to

poor

engine

performance

repeatability

or

engine

reliability

problems.

In

addition,

thermal

management

of the

injector

head

is

important

to

ensure

that

heat

diffusion

from

the hot

chamber

material

to the

injector

head

is

minimized

in

order

to

prevent

vaporization

of

propellants

in the

injector

and not

to

exceed

thermal

limits

of the

injector

material,

which

may

be

different

from

the

high-temperature

chamber

material

for

machining

reasons.

Unlike-doublet

injector

types

are

favoured14'

16

because

of

better

mixing

results

and

reduced

heat

load

to the

injector

head

by

displacing

the

flame

front

away

from

the

injector

wall

surfaces16.

As

mentioned

above,

more

injector

elements

will

lead

to

better

mixing,

however,

limited

engine

size

may

limit

the

number

of

injector

elements.

In the

case

of the

Rockwell

engine

discussed

above,

only

a

single

unlike-doublet

injector

element

is

used10'11

(combustion

efficiencies

are

maintained

at

high

levels

due to the

aforementioned

elimination

of the BLC

layer).

Table

7

lists

the

smallest

bi-propellant

engine

technology

available

today.

Caution

has to be

exercised

when

referring

to the

data

presented

in

this

table.

Not all

engines

listed

are

space-qualified

at

this

point

and

some

may

still

be

experiencing

problems

in

their

respective

development

programs.

Schwende

et

al.13

point

out in

their

1993

paper

that

the 4 N

engine

experienced

"anomalies"

due

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

Table

7:

Comparison

of

Small

Bi-Propellant

Engines

Thrust

(N)

4

4.45

10

10

10

22

22

22

22

22

30

156

Manu-

facturer

DASA

Marquardt

DASA

Marquardt

Royal

Ordnance

Marquardt

Atlantic

Research

Aerojet

Aerojet

Royal

Ordnance

Rockwell

Marquardt

Type

Fuel/

Isp

(s)

Weigh

Oxidizer

t

(kg)

MMH/

285

0.27

MON-1

R-2/

MMH/NTO

280

0.43

R-2B

MMH/

290 0.3

MON-1

R-52

MMH/NTO

295

LIT

MMH/NTO

R-6C/

MMH or 289

0.67

R-6D

N2H4/NTO

A0809

MMH/NTO

290

0.55

SSD

MMH/NTO

280

0.59

MMH/NTO

313

Leros

N2H4/MON

285

0.85

20H

MMH/NTO

287 0.1

Divert

N2H4/NTO

-

0.1

Size

Comments

(length*

Max.Dia)

(cm)

Reported

temperature

anomalies

26.1

x <9 No

known

applications

Regenerative

cooled

throat

in

previous

version.

34

flight

units

built.

No

known

applications

-

25x<13

Flight

Applications

for

R-6C,

no

known

application

R-

6D

21.7x5.4

Flight

Applications

18.5x6.9

-

Rh/Ir

chamber.

In

development

20.9x6.6

Under

development

Max.

26 sec in

single

burn.

Max.

accumulative

burn:

77

sec.

1.25

mixture

ratio.

Developed

for

BMDO

20 sec

single

burn

demonstrated

Developed

for

LEAP

Ref

13

17

13

17

18

16

19

20

12

18

10,

11

9

to

too

high

chamber

wall

temperatures.

Also,

although

single

duration

burns

which

is too

short

for

interplanetary

commonly

referred

to as

examples

for the

degree

of

delta-v

maneuvers.

As

mentioned

above,

the

KE

ASAT

miniaturization

achieved

for

bi-propellant

engines,

the KE

developments,

as

well

as

others,

have

focused

on

attitude

ASAT

engines

have

been

tested

only

up to 26

seconds

in

control

applications,

rather

than

primary

propulsion

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

applications.

As a

result,

pulsing

performances

rather

than

long

duration

burns

were

emphasized

and led to the

currently

exhibited

design

performances.

Almost

all

engines

use

nitrogen

tetroxide

(NTO)

and

monomethylhydrazine

(MMH)

as

oxidizer

and

fuels,

respectively,

due to

storage

reasons,

acceptable

performance

values

and

relatively

benign

mixture

ratio

sensitivities.

Using

a

O/F

mixture

ratio

of 1.6

results

in

equal

propellant

volumes

for

fuel

and

oxidizer,

so

that

identical

tanks

can be

used

(reducing

development

cost

and

time)

and the

spacecraft

will

experience

no

center-of-

gravity

(e.g.)

shifts

during

burns.

Given

the

engine

data

presented

in

Table

7, and

using

the

representative

mission

requirements

of

Section

n,

propellant

mass

fractions

of 0.7 and

0.58

can be

computed

for

a

spacecraft

with

delta-v

requirements

of

3400

m/s

and

2500

m/s,

respectively,

at a

specific

impulse

of 290 s

(see

Table

7).

Although

these

values

are not

atypical

for

larger

spacecraft,

for a

microspacecraft

they

may be too

high

given

that

component-mass-to-spacecraft-mass

ratios

are

larger.

For

example,

in the

3400

m/s

case

for a 20 kg

spacecraft,

merely

6 kg of

available

mass

remains.

This

will

have

to

include

the

entire

dry

mass

of the

propulsion

system,

structure

plus

all

other

subsystems.

According

to

Table

7,

one

bi-propellant

engine

alone

may

take

up

about

5% of

that

mass,

even

using

the

lightest

engines

available.

Smaller

delta-v

requirements

around

1000

m/s,

on

the

other

hand,

would

result

in

propellant

mass

fractions

of

0.3 for a 290 s

bi-propellant

system.

However,

a

hydrazine

mono-propellant

system

with

a

specific

impulse

of 220 s

(see

below)

would

result

in a

propellant

mass

fraction

of

0.37.

In the

case

of a 10 or 20 kg

Class

I

spacecraft,

this

difference

would

be a

mere

0.7 or 1.4 kg in

propellant

mass,

respectively.

Given

the

lower

component

part

count

of a

mono-propellant

feed

system,

this

higher

propellant

fraction

can

easily

be

offset

by a

simpler

mono-propellant

system.

It

is,

therefore,

concluded

that

bi-propellant

systems

are not

suitable

for

either high

or low

delta-v

requirements

onboard

a

microspacecraft

due to too

high

dry

weight

of the

system.

Further,

aggressive

miniaturization

may

help,

however,

there

exists

considerable

doubt

that

significantly

smaller,

yet

reliable

and

space-qualifyable

bi-propellant

engine

technology

can be

developed

in

view

of the

design

challenges

for

small

bi-propellent

engines

given

above.

However,

separate

chemical

stages

for

large

delta-v

maneuvers

may

make

use of

bi-propellant

engines.

An

example

of

such

a

stage

is

given

in

Ref.

21,

describing

a

Hydrazine

(N2H4)/

Chlorine

Pentafluoride

(C1F5)

chemical

upper

stage,

developed

for the

Lightweight

Exo-

Atmospheric

Projectile

(LEAP)

program.

Thrust

levels

provided

by the

LEAP

stage

are

somewhat

high

(2056

N) for

microspacecraft

applications

and

there

are

concerns

regarding

the

corrosivity

and

toxicity

of

C1F5.

Stages

like

these,

or

similar

ones

using

more

conventional

propellants,

however,

may

be

required

for

orbit

insertion

maneuvers

around

distant

planetary

bodies,

in

particular

when

these

bodies

are

lacking

an

atmosphere

(no

aerobraking

possible)

or

when

they

are

located

too far

from

the sun

(solar

power

levels

too low to

use

solar

electric

propulsion).

In

addition,

landing

and

take-

off

operations

will

require

bi-propellant

technology.

However,

those

mission

applications

may

require

thrust

levels

well

exceeding

those

obtainable

with

the

engines

listed

in

Table

7 due to

high

stage

masses

and

large

required

vehicle

accelrations

to

overcome

the

gravity

of the

respective

planetary

body.

Mono-Propellant

Hydrazine

Engines

Hydrazine

mono-propellant

thrusters

combine

engine

technology

substantially

simpler

than

that

of bi-

propellant

engines,

relatively

simple

and low

part-count

feed

systems,

and

high

reliability

with

intermediate

performance

(specific

impulses

around

220 s for

state-of-the

art

hydrazine

thruster

technology).

In a

hydrazine

thruster,

the

propellant

is

passed

through

a

catalyst

bed and

decomposed.

The

decomposition

products

are

nitrogen,

hydrogen

and

ammonia.

The

reaction

takes

place

in two

stages:

hydrazine

decomposes

first

through

an

exothermic

reaction

into

ammonia

and

nitrogen.

The

ammonia

then

decomposes

further

through

an

endothermic

reaction

into

hydrogen

and

nitrogen,

however,

leaving

the

overall

reaction

exothermic.

The

degree

of

ammonia

decomposition

depends

on

many

factors,

among

them

feed

pressure,

catalyst

type

and

geometry.

Shell

405 is the

standard

catalyst

used

in the US,

consisting

of 1.5 - 3 mm

dia.

alumina

pellets

coated

with

iridium.

The

catalyst

pellets

are

contained

within

a

mesh

construction

in a so

called

catalyst

bed.

Upon

contact

with

the

iridium

surfaces,

the

hydrazine

decomposition

reaction

is

initiated.

Hydrazine

thrusters

have

been

used

extensively

on

conventional

spacecraft

for

attitude

control

as

well

as

primary

propulsion

sources

for

intermediate

to low

delta-v

requirements

(about

1000

m/s

or

less).

Of

interest

here

are

the

smallest

available

hydrazine

thrusters,

in the 0.9 -

4.45

N

range.

These

engines

are

being

manufactured

in the US by

-

Primex

(formerly

Olin

Aerospace/Rocket

Research),

Kaiser-Marquardt

and TRW

companies,

and

abroad

by

Daimler

Benz

Aerospace

in

Germany.

Typical

engine

characteristics

are

listed

in

Table

8.

The

engine

sizes,

weights

and

thrust

levels

should

allow

for

relatively

easy

integration

on a

Class

I

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

Table

8:

State-of-the-Art

US

Hydrazine

Thrusters

Thrust

Manufacturer

Type

Isp

(N)

(s)

Weight

Size

(kg)

(Lengthx

Max.

Dia)

_________(cm)

Comments

Ref.

0.9

Primex

MR-

210-

0.33

103

220

14.8x3.4

C/E,

D

and

G

models.

Isp and

thrust

feed

pressure

dependent

(340

-

<100

psia).

________Considerable

flight

use.

22

0.9

Marquardt

KMHS

226

0.33

10

14.6x3.2

Isp and

thrust

feed

pressure

dependent.

Flight

use________

17

2.2

Primex

MR-

213-

0.33

11

IE

224

16.9x3.8

Isp and

thrust

feed

pressure

dependent

(370

- 60

psia).

Considerable

________flight

use._______

22

4.45

Primex

MR-

226-

0.33

111C

229

16.9x3.8

Isp and

thrust

feed

pressure

dependent

(400-80

psia).

________Considerable

flight

use.

22

4.45

Marquardt

KMHS

230

0.38

17

20.3x3.2

Isp and

thrust

feed

pressure

dependent.

Flight

use._______

17

TRW

MRE-1

220

0.82

15.2xN/A

Mass

is for

dual

thruster

module.

Isp and

thrust

feed

pressure

dependent.

Considerable

flight

use.

23

18

TRW

MRE-4

230

0.41

20.3x

unknown

Isp and

thrust

feed

pressure

dependent.

Considerable

flight

use.

23

microspacecraft

bus,

mounted

along

the

axis

of the bus for

primary

propulsion

applications.

All

thrusters

listed

have

seen

considerable

flight

use and

potentailly

would

require

only

minimal

re-development

for use as

Class

I

main

engines.

Class

II

microspacecraft

are too

small

to

take

advantage

of

these

existing

technologies.

One

area

of

improvement

in the use of

state-of-the-

art

hydrazine

thrusters

as

Class

I

main

engines

may be

found

in

the

valve

area.

Currently,

a

considerable

weight

fraction

of

a

small

hydrazine

thruster

is

taken

up by the

thruster

valve

(greater

than

50% for the

small

engines

considered

here).

This

fact

may

open

an

opportunity

for

further

weight

reductions.

Since

the

smallest

hydrazine

thrusters

have

been

used

mainly

for

attitude

control

purposes

where

fast

valve

action

is

essential

( on the

order

of 15

ms

on/off),

these

valves

could

possibly

be

replaced

by

slower

valves,

since

primary

propulsion

applications

of

these

thrusters,

as

envisioned

here

for

microspacecraft,

seldom

would

require

engine

pulses

that

short.

Slower

valves,

depending

on

design,

may

require

less

electromagnetic

force

action

to

open

the

valve,

which

might

reduce

electromagnet

masses.

A

disadvantage

of

hydrazine

propellant

is its

toxicity

and

flammability

and

resulting

ground

handling

and

propellant

loading

procedures.

These

procedures

are

obviously

well

established

due to

extensive

hydrazine

thruster

use on

conventional

spacecraft,

but

will

significantly

contribute

to the

cost

of

small

spacecraft.

In

addition,

as was

pointed

out in the

preceding

section,

a

hydrazine

propulsion

system

onboard

a

microspacecraft

is

only

practical

if

small

or

intermediate

delta-v

maneuvers

are

required

(i.e.

<

1000

m/s).

In

these

cases,

mono-propellant

systems

will

have

an

advantage

over

bi-propellant

systems

due

to

reduced

system

complexity,

smaller

component

part

count

and,

thus,

smaller

volume

requirements.

If

higher

delta-v's

are

required

(see

Section

II),

mono-propellant

systems

become

increasingly

heavy

due to

large

propellant

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

requirements.

In

these

cases,

bi-propellant

engines,

likely

to

be

mounted

on a

separate

kick-stage,

or

electric

thruster

options

(see

below)

should

be

preferred.

Thus,

limited

Class

I

microspacecraft

use of

existing

hydrazine

thruster

technology

for

primary

propulsion

applications

appears

reasonable.

HAN-based

Mono-Propellant

Thrusters

Recently,

so

called

HAN/TEAN

thrusters4'24

have

received

attention.

This

thruster

is of the

mono-propellant

type,

using

mixtures

of

HAN

(Hydroxylammonium

Nitrate

-

NH3OH+NO3.),

TEAN

(Triethanolammonium

Nitrate

-

(HOCH2)3HNOH+NO3.)

and

water

as a

propellant.

HAN/DEHAN

mixtures

have

also

been

studied,

consisting

of

HAN,

water

and

DEHAN

(Diethylhydroxylammonium

Nitrate

-

(CH3CH2)HNOH+NO3.)24.

HAN is an

oxygen

rich

component

and

TEAN

or

DEHAN

are

fuel

rich

components.

Due

to the

water

additive,

both

components

can

coexist

in a

mixture

without

detonation,

as

long

as the

water

content

is

maintained

at 10% or

above25.

Exposing

the

mixture

to a

catalyst

causes

a

chemical

reaction

and

exothermic

decomposition

of the

components

into

CO2,

N2

and

H2O4>24.

Mixtures

of

these

propellants

have

been

studied

for

use

as

liquid

gun

propellants

by the US

Army

and

have

been

categorized

according

to

their

composition.

LP1846,

for

example,

consists

of

60.8%

HAN,

19.2%

TEAN

and 20%

water,

while

LP1845

consists

of

63.2%

HAN,

20%

TEAN

and

16.8%

water24,

i.e.

has a

lower

water

content

than

LP1846.

The

amount

of

water

in the

mixture

greatly

influences

the

decomposition

temperature

and,

thus,

the

available

specific

impulse.

Increasing

the

water

content

will

lower

the

flame

temperature.

Jankowski24

quotes

flame

temperatures,

based

on

numerical

calculations,

of

2022

K for

LP1846

and

2125

K for

LP1845,

having

the

lower

water

content.

These

values

result

into

theoretical

specific

impulses,

assuming

a

specific

impulse

efficiency

of

92%,

of

233 and 239

sec24.

Tests

performed

with

HAN-based

propellants

with

different,

not

specified

additives

have

resulted

in

specific

impulses

of 270 sec at a

flame

temperature

of

2500

K24.

The

flame

temperatures

of

HAN/TEAN

combinations

are

very

high

and

approach

values

found

in

small

bi-propellant

engines.

Thus,

many

of the

thermal

design

challenges

found

in the

construction

of

small

bi-

propellant

chambers

would

have

to be

overcome

when

using

high-performing

HAN/TEAN

mixtures

with

low

water

content.

Engine

lifetime

restrictions

may

thus

result

if

high

performance

is

required.

However,

as

mentioned

above,

flame

temperatures

may

be

lowered

at the

expense

of

specific

impulse

performance

if the

water

content

is

raised.

Even

though

not

providing

a

significant

performance

advantage

over

existing

hydrazine

mono-propellant

thrusters

in

those

cases

anymore,

HAN/TEAN

thrusters

still

offer

advantages

due to the low

toxicity

of

both

the

NAN/TEAN

propellant

as

well

as its

reaction

products,

high

storage

densities

(about

40%

higher

densities

than

that

of

hydrazine)

and

lower

environmental

temperature

handling

capabilities. While

hydrazine

freezes

at

about

0 C,

HAN/TEAN

mixtures

may be

used

at

temperatures

as low as

about

-33

C24.

At

this

point,

the

viscosity

of

HAN/TEAN

mixtures

increases

and

propellant

feeding

will

no

longer

be

possible

using

conventional

feed

system

technologies24.

Higher

storage

densities

and

lower

environmental

temperature

handling

capabilities

of

HAN/TEAN

propellants

are

beneficial

for

microspacecraft,

since

they

allow

for

smaller

and

lighter

storage

tanks

and the

elimination

of, or

reduction

in

power

for,

tank

and

line

heaters,

reducing

overall

power

requirements

for the

spacecraft.

Thus,

HAN/TEAN

thrusters

may

find

Class

I

microspacecraft

applications

for

those

reasons.

Reduction

in

engine

size

to

meet

Class

n

requirements

may not be

possible

due to the

high

heat

loads

to be

expected

in a

HAN/TEAN

decomposition

chamber.

Considerable

development

work

will

be

required

to

bring

current

HAN/TEAN

thruster

concepts

to

flight

status,

with

reaction

chamber

thermal

design

issues

being

one of the

most

challenging

steps

in the

development.

Other

Mono-Propellant

Thrusters

Hydrogen

peroxide

(H2O2)

thrusters

are

considered

from

time

to

time

as an

alternative

to

more

conventional

mono-propellant

systems25'26.

Hydrogen

peroxide,

when

subjected

to a

suitable

catalyst,

decomposes

into

water

and

oxygen

in an

exothermic

reaction.

Although

hydrogen

peroxide

has

been

used

in

flight

applications

as a

mono-

propellant

in the

past8'26'29,

it is no

longer

in use due to

propellant

storability

issues8'26"29.

Hydrogen

peroxide

slowly

decomposes

when

heated

or

exposed

to a

catalyst.

Almost

any

organic

substance

can

serve

as

such

a

catalyst29.

If

slow

decomposition

occurs

in

propellant

tanks,

as has

been

observed

in the

past8,

tank

pressure

increases

result

over

time

and

propellant

is

lost

due to its

slow

conversion

into

its

reaction

products

inside

the

propellant

tank.

10

Copyright©

1997, American Institute

of

Aeronautics

and

Astronautics,

Inc.

Solid

Rocket

Motors

Solid

rocket

motors

are

frequently

used

in

kick-

stages

for

orbit

raising

or

orbit

insertion

of

spacecraft,

beginning

with

the

Explorer

1

spacecraft8

and

leading

to the

more

recent

Pioneer-Venus30,

Magellan8'30

and

Galileo

missions,

as

well

as

numerous

commercial

missions

(orbit

raising).

In

solid

motors,

fuel

(typically

aluminum

powder),

oxidizer

(typically

ammonium

perchlorate

-

NH4C1O4)

and

an

organic

binder

(typically

Hydroxyl-terminated

Polybutadiene

-

HTPB)

are

combined

into

a

composite

to

form

the

solid

propellant8'30.

The

advantages

of

solid

rocket

motors

are

their

compact

size

combined

with

a

relatively

high

specific

impulse

performance

-

less

than

that

of bi-

propellant

systems

but

higher

than

that

of

mono-propellant

systems29.

For

obvious

reasons,

solid

motors

also

do not

suffer

from

propellant

leakage

concerns.

Propellant

sublimation

by

exposure

to

space

through

an

open

nozzle

was

a

concern

for the use of

solid

motors

for

deep-space

applications,

but has

been

found

to

have

no

impact

on

motor

performance

after

10 - 15

months

of

in-space

storage8-30.

In the

case

of the

Magellan

mission,

a

Thiokol

STAR

48B

motor

was

fired

for

Venus

orbit

insertion

after

462

days

in

space30.

A

Thiokol

STAR

24

motor

was

fired

after

6.5

months

in

space

for the

Venus

orbit

insertion

of

the

Pioneer

probe.30

Disadvantages

of

solid

motors

are

that

they

are

generally

not

restartable,

and

therefore

do not

allow

for

orbit

trimming.

If

several

delta-v

burns

are

required

it is

necessary

to

stack

multiple

stages

leading

to

system

complexities

and

higher

propulsion

system

dry

masses.

The

issue

of

orbit

trimming

is of

particular

importance

for

solid

motors

since

exact

prediction

of

delivered

total

impulse

is

difficult

to

estimate

due to

uncertainties

in the

expected

grain

temperature,

the

exact

propellant

composition,

and the

amount

of

inert

material

consumed8.

Thus,

a

separate

small

liquid

system

may

have

to

provide

for the

orbit

trimming

maneuvers30.

A

separate

liquid

system

may

also

be

required

for

despin

of the

satellite.

Solid

motors

for

space

applications

are

usually

not

equipped

with

thrust

vectoring

capability.

Although

some

larger

motors

have

been

tested

with

such

a

capability32,

these

nozzle

gimbal

systems

may

be

too

heavy

and

complex

for

very

small

motors,

such

as

those

required

for

microspacecraft

applications.

Thus,

spacecraft

generally

are

spin-stabilized

before

motor

firings,

and

despin

may be

required

after

separation

from

the

stage

depending

on the

mission,.

Table

9

shows

some

of the

smallest

solid

motors

available

today30"33.

In

addition

to the

companies

listed

in

Table

9,

Pacific

Scientific,

Inc.

is

also

building

small

rocket

engines

for

missile

divert

purposes34.

As can be

seen

by

inspecting

Table

9,

envelopes

and

masses

of the

smallest

available

motors

fit

within

the

Class

I

category

of

microspacecraft

and

specific

impulse

performances

are

quite

good.

However,

thrust

levels

are

much

higher

than

desired

and

burn

times

are

relatively

short,

which

would

lead

to

very

large

microspacecraft

accelerations.

For

example,

assuming

a

20 kg

overall

spacecraft

weight

(incl.

motor),

using

a

Thiokol

STAR

6B

motor

results

in

accelerations

at the

beginning

of the

burn

of

about

13 g's and at the end of the

burn

around

18

g's.

The

delta-v

that

can be

achieved

with

this

motor

for a 20 kg

spacecraft

would

be 963

m/s.

A

similar

calculation

for a 10 kg

spacecraft

equipped

with

the

STAR

5A

motor

would

lead

to an

initial

acceleration

of 1.7

g's and an

acceleration

just

prior

to

burn

out of

about

2.2

g's,

and a

delta-v

of 641

m/s.

Achievable

delta-v's

are

limited

by the

reduced

propellant

mass

fractions

found

typical

for

smaller

motors.

The

high

thrust

forces

and

short

burn

times

are a

result

of the

intended

design

applications

for

most

of

these

small

motors,

i.e.

stage

separation

or use as

missile

divert

engines

(missile

attitude

control).

In

both

cases

it is

essential

to

provide

a

relatively

large

thrust

in a

short

amount

of

time.

In the

case

of

microspacecraft

applications,

this

could

lead

to

limitations

of

solid

motor

use due to the

requirement

of

being

able

to

fire

only

in a

stowed

vehicle

configuration

(no

deployments)

and

possibly

costly

re-

qualification

of

spacecraft

components

to

account

for

these

high

accelerations.

An

exception

to the

fast

burning,

high-thrust

small

solid

motors

shown

in

Table

9 is the

STAR

5A.

Even

though

accelerations

in the

example

given

above

are

still

quite

high,

those

values

may be

much

more

tolerable.

The

longer

burn

time

and

smaller

thrust

of

this

motor

were

achieved

by

using

an end

burner

propellant

grain.

This

grain

type

was

extensively

used

in the

past

in the so

called

JATO8

(Jet

Assisted

take-Off)

units

used

in the

1940s

to

assist

in

the

take-off

of

aircraft

from

short

runways

or

assisting

aircraft

requiring

additional

thrust

for

heavy-lift

capability.

This

grain

type

may be the

grain

of

choice

for

microspacecraft

applications.

Even

longer,

lower-thrust

burns

could

be

accomplished

if the

length-to-diameter

ratio

of

the

motor

case

could

be

increased.

Solid

rocket

motors

may

thus

present

an

interesting

alternative

to

more

complex

liquid

systems

where

mission

profiles

are

simple,

require

only

single

burns

and

intermediate

delta-v

values

(<

1000

m/s),

a

small

liquid

system

is

onboard

the

microspacecraft

for

orbit

trimming

or

if the

required

accuracy

of the

actually

delivered

delta-v

is not

too

high.

For

example,

in the

Saturn-ring

explorer

scenario

discussed

in

Section

I2,

quite

sizable

delta-v

changes

may be

achieved

for the

microspacecraft

probes

using

small

solid

motors.

Since

multiple

probes

are

being

used

anyway

to

11

Copyright©

1997,

American Institute

of

Aeronautics

and

Astronautics,

Inc.

Table

9:

State-of-the-Art

Small

Solid

Rocket

Motors

Manufacturer

Thiokol32

Thiokol32

Thiokol"

Thiokol32

Atlantic

Research33

Atlantic

Research33

Atlantic

Research33

Type

STAR

5A*"

STAR

5C

STAR

5CB

STAR6B

-

-

-

Loaded

Weight

(kg)

4.7

4.5

4.5

10.3

0.4

0.5

1.6

Propellant

Weight

(kg)

2.3

2.1

2.1

6.1

-

-

-

Size

(LxD)

(cm)

22.5x13

34x12

34x12

40x18.6

-

-

-

Burn

Time'

32

2.8

2.67

5.9

1

1

-

Isp

(s)

250

266

270

273

-

-

-

Thrust"

(N)

169

1953

2041

2513

222

311

952

Burn

time,

10%

thrust

at

ignition,

90%

thrust

at

shut-down

Burn

time

averaged

End

Burner

account

for the

potential

loss

of

some,

total

impulse

and

delivered

delta-v

uncertainties

for

individual

probes

may be

acceptable.

The

absence

of

leakage

concerns

and the

ability

to

compactly

package

solid

motors

will

be

attractive

for

microspacecraft

applications

Existing

motor

hardware

appears

to fit the

envelope

of

Class

I

spacecraft,

although

longer

burn

times,

lower

thrust

values

and,

thus,

lower

vehicle

accelerations

should

be

aimed

for.

The

benefits

of

using

small

solid

rocket

motors

would

be

even

more

pronounced

for

smaller

microspacecraft,

such

as

types

falling

into

the

Class

II-category.

Here,

compactness

plays

an

even

greater

role

than

for

Class

I

craft.

Class

II

application

of

solid

motors

would

require

further

miniaturization

of

solid

motor

technology

and

dedicated

full

development

programs

to

achieve

the

desired

reductions

in

size,

weight

and

thrust.

Hybrid

Rocket

Motors

In

a

hybrid

rocket

motor

a

solid

fuel

is

combined

with

a

liquid

or

gaseous

oxidizer,

which

is

stored

in a

separate

propellant

tank

and fed

into

the

motor

case27'35'36.

As

a

result

of

this

separation

between

solid

fuel

and

liquid

oxidizer,

hybrid

rockets

exhibit

some

interesting

properties.

Hybrid

rockets

are

restartable,

relatively

safe

when

compared

with

solid

motors,

offer

appreciable

specific

impulse

performances

up to

around

300 s

when

using

storable

propellants,

and

still

offer

a

higher

degree

of

compactness

than

bi-propellant

systems.

Hybrid

rockets,

at

first

glance,

may

thus

seem

to be an

attractive

cross

between

high

performing

bi-propellant

engines

and

compact

solid

motor

technology.

One of the

disadvantages

of

hybrid

motors,

in

particular

when

viewed

in

terms

of

space

applications

(microspacecraft

or

otherwise)

with

long

mission

durations,

is the

limited

choice

of

suitable

storable

propellant

combinations

available

today.

Typically,

Hydroxyl-

terminated

Polybutadiene

(HTPB)

is

being

used

as

fuel

and

liquid

oxygen

(LO2)

or

hydrogen

peroxide

(H2O2)

are

used

as

oxidizers27'28'35"37.

Although

both

oxidizers

may be

suitable

for

launch

applications,

they

are not

storable

over

long

periods

of

time

because

they

are

either

cryogenic

(LO2)

or

may

slowly

decompose

over

time

in the

propellant

tank

before

use

(H2O2

- see

above).

Among

storable

oxidizer

options,

nitrogen

tetroxide

(NTO)

was

used,

but

required

a

separate

ignition

source27.

Chlorine-fluorides,

such

as

C1F3

and

C1F5,

have

also

been

used

as

oxidants.

These

substances,

however,

are

highly

toxic

and

corrosive.

Work

on

hybrid

rocket

motors

started

in the

1930's

to

40's

in

both

Germany

and the US,

with

some

of the

early

work

in

Germany

performed

by

Hermann

Oberth35-36.

Development

has

continued

on an

on-and-off

basis

over

the

years.

Focus

was

placed

mainly

on

launch

applications,

leading

to the

development

of the

H-500

(312,000

N

thrust)

and

the

H-250F

(1,000,000

N

thrust)

engines

developed

by

the

recently

failed

AMROC

company35.

Both

of the

latter

two

motors

used

HTPB

and

LO2

as

propellants.

Extensive

research

on

hybrid

rockets

has

also

been

performed

at

various

university

research

laboratories

around

the

world28.

This

work

was

performed

on

smaller test

devices.

The

smallest

quoted

thrust

value

for a

hybrid

rocket

engine

is

found

in

Sellers

et

al.28

at

ION.

Given

that

attention

in

hybrid

rocket

engine

development

was

focused

mostly

on

large

launch

motors,

12

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

where

oxidizer

storage

issues

play

a

lesser

role

than

for in-

space

applications,

it is

uncertain

whether

hybrid

rockets

may

find

space

applications

for

interplanetary

missions

. In

particular

microspacecraft

applications,

requiring

small

motor technology comparable

in

size

to

solid

motor

technology

given

in

Table

9,

remain

uncertain.

The use of

hybrid

engines

for

these applications

would

require

a

full

development

program, starting

at the

point

of

propellant

selection.

If

successful,

hybrid

engine

technology

could

fill

a

useful

gap

between

high

performing,

yet

complex

bi-

propellant

engines

and

compact

and

simple,

yet

relatively

inflexible

solid

motor

technology.

Primary

Propulsion

-

Electric

Ion

Engines

In

an ion

engine,

the

propellant

(typically

xenon)

is

ionized

in a

plasma

discharge.

Ions

are

extracted

from

the

plasma

via

electrostatic

forces

and

accelerated

across

an

electric

potential

difference

of

about

IkV.

In the

process,

xenon

ions achieve

a

velocity

of

about

30,000

m/s,

corresponding

to a

specific

impulse

of

about

3,000

sec.

An

ion

propulsion subsystem

consists

of

several

components

that

all

will

have

to be

miniaturized

for

microspacecraft

applications.

These

are the

thruster

itself,

the

power

conditioning

unit

providing

the

required

voltages

to

the

engine,

and the

feed

system.

Within

the

thruster

assembly,

critical

components

include

the

cathode

for

certain

engine

types,

the

accelerator

grid

system

and the

neutralizes

(

used

to

neutralize

the ion

beam

to

avoid

charging

the

spacecraft).

Different

types

of ion

engines

are

being

developed

. DC

electron

bombardment

types

use an

electron

current

emitted

from

a

hollow

cathode

inside

the

engine

body

to

ionize

the

propellant

gas by

causing

collisions

between

the

electrons

and the

propellant

gas

atoms.

RF

(radio-frequency)

electron

bombardment

engines

Table

10:

State-of-the-Art

Small

Ion

Engines

use

electrons

accelerated

in an

inductive

coupled

RF

field

to

cause

propellant

ionization.

Advantages

of ion

engines

are

their

large

specific

impulses

which

translate

into

significant

propellant

and

spacecraft

mass

reductions.

This

fact

is of

particular

importance

for

mass

constrained

microspacecraft,

especially

for

interplanetary

missions

which

have

large

delta-v

requirements.

Using

ion

engine

technology

may

lead

to

lighter

overall

spacecraft

masses

and

shorter

mission

trip

times

when

compared

with

chemical

bi-propellant

systems.

In

addition,

xenon

propellant,

when

stored

at

about

2,000

psia

pressure,

takes

on a

supercritical

state

with

a

density

about

twice

that

of

water.

Reduced

propellant

requirements

due

to the

high

specific

impulse

of the

engine

and

high

density

will

allow

for

compact

propellant

storage.

On

the

other

hand,

propellant

mass

reductions

due

to

higher

specific

impulses

will

have

to be

traded

off

with

electric

power

requirements.

Power

requirements

drive

power

conditioning

unit

and

power

supply

masses.

In

order

to

reduce

overall

system

wet

masses,

an

optimum

operating

point

must

be

selected

for the

engine,

allowing

for

both

significant

propellant

savings

and low

power

system

masses.

For

typical

interplanetary

mission

requirements

and

current

power

system

technology,

this

optimum

is

typically

found

around

3,000

s

specific

impulse.

Table

10

lists

the

smallest

available

ion

engine

technology

today38"41.

As can be

seen,

all

current

ion

engine

systems

are too

large

for use on a

microspacecraft,

both

with

respect

to

mass

(incl.

PPU)

and

power

requirements.

These

engines

may,

however,

be

used

on

separate

electrical

stages

just

like

the

bi-propellant

thruster

technology

discussed

above,

and

provide

large

delta-v

changes

for

Class

I -

type

microspacecraft

in

that

configuration.

In the

case

of

electric

stages,

a

dedicated

power

supply

will

have

to be

provided.

Manufacturer

Discharge

Type

Thrust

(mN)

Isp

(s)

Power

(W)

Thruster

Mass

(kg)

PPU

Mass

(kg)

Beam

Diameter

(cm)

Hughes40

DC

17.8

2585

439

5.0

6.8

13

DASA38'39

RF

5-15

3000

240

(5 mN) - 600 (15 mN)

1.6

8.0

(PPU)

1.3 (RF

generator)

10

JPL41

DC

21-31

2500

-

3900

500-900

2.5

-

15

13

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

In

order

to

integrate

ion

engine

technology

onboard

a

microspacecraft

bus

within

the

mass

margins

provided

in

Section

II, new

technologies

will

have

to be

developed.Thruster

sizes

will

have

to be

reduced.

Challenges

to

be

overcome

here

will

be to

maintain

plasmas

in

small

discharge

chambers,

where

electron

wall

losses

may be

high.

Cathodes

and

neutralizers

will

have

to be

miniaturized

and

cold

cathode

technology

may be

explored.

Micromachined

grid

systems,

allowing

for

electrostatic

beam

steering,

eliminating

heavy

engine

gimbals

(see

Section

V) and

miniaturized

power

processing

units

will

be

needed.

Miniature

ion

engine

technology

may

possibly

be

based

on

hollow

cathode

technology,

currently

used

for

conventional

engine

cathodes

and

neutralizers.

Thus,

miniature

ion

engines

have

to be

considered

as

very

advanced

micropropulsion

concepts

that

still

have

to

overcome

many

feasibility

concerns

before

they

can be

seriously

considered

for

microspacecraft

applications.

Hall

Thrusters

Hall

thrusters42'43

are

electrostatic

propulsion

devices

which

use

xenon

propellant.

Plasma

generation

and

ion

beam

accelration

are

different

from

those

found

in

ion

engines

and

lead

to a

more

compact

thruster

technology.

In

a

Hall

thruster42'43,

electrons

emitted

from

a

hollow

cathode

external

to the

thruster

are

accelerated

towards

a

positive

anode

located

upstream

and

inside

an

annular

discharge

chamber.

On

their

way to the

anode,

the

electrons

cross

a

radial

magnetic

field

extending

across

the

annular

chamber.

Due to

Lorentz-force

action,

the

electrons

gyrate

around

the

magnetic

field

lines,

and

drift

azimuthally

through

the

annular

channel,

colliding

with

propellant

gas

atoms

(xenon)

and

ionizing

them.

The

ions

are

accelerated

away

from

the

engine

by the

same

electric

field

that

attracted

the

electrons.

The ion

beam

is

neutralized

by

additional

electrons

streaming

off the

cathode.

Due

to the

high

electron

density

in the

magnetic

field

region,

a

dense

ion

beam

can be

formed,

overcomming

space,

charge

limitation

effects

found

in ion

engines.

Hall

thrusters

are

thus

more

compact

for the

same

delivered

thrust

level

than

ion

engines.

On the

other

hand,

Hall

thruster

typically

deliver

specific

impulses

around

1500

-

2000

s,

making

them

more

suitable

for

near-earth

missions

(orbit

transfer,

repositioning,

etc.),

rather

than

interplanetary

flights.

A

high-Isp

Hall

thruster

would

be an

attractive

alternative

to ion

engines.

Current

Hall

engine

technology

is far too

heavy

and

power

consuming

to be

used

within

the

microspacecraft

design

envelope.

However,

efforts

are

underway

to

miniaturize

the

technology

for

small

satellite

applications43.

Design

goals

are a

power

level

of 50

W

at a

thrust

of 5

mN

and

a

specific

impulse

of

1600

s. The

device

is

estimated

to

be

about

4 mm in

diameter.

Given

the

smaller

channel

dimensions,

larger

magnetic

fields

have

to be

provided

to

achieve

smaller

electron

gyration

radii.

Required

estimated

field

strengths

are 0.5 T. For

even

smaller

devices,

magnetic

field

strengths

would

have

to be

increased

further.

Currently,

samarium-cobalt

permanent

magnets

are

able

to

deliver

about

1 T at

their

surface,

and

thus

further

miniaturization

beyond

the

engine

size

outlined

above

may be

difficult

to

achieve.

Power

levels

even

for the

miniaturized

Hall

thruster

version

currently

being

studied

at

MIT43

are

quite

high

for

the

microspacecraft

considered

in

this

study,

but may be fit

within

the

upper

Class

I

category

if

power

requirements

can

be

relaxed

and a

dedicated

power

supply

is

provided

for the

thruster.

However,

as

mentioned,

a

Hall

thruster

device

appears

unsuitable

for

interplanetary

applications

requiring

large

delta-v's

due to

their

lower

specific

impulse

capability

when

compared

with

ion

engines.

In

addition,

as for ion

engine

applications,

substantial

reductions

in

power

processing

unit

weight

and

size

will

have

to be

made.

Field

Emission

Thrusters

Field

emission

thrusters44"46,

or

Field

Emission

Electric

Propulsion

(FEEP)

devices

as

they

are

commonly

being

referred

to,

have

traditionally

been

envisioned

for

spacecraft

attitude

control,

providing

ultra-fine

control

or

drag

make-up

to

establish

virtual-drag-free

environments

on

scientific

missions.

However,

FEEP

devices,

due to

their

already

small

size

and low

available

thrust

levels,

may

also

be

able

to

provide

primary

propulsion

for

microspacecraft.

In

a

FEEP

device,

thrust

is

generated

through

electrostatic

forces

as in ion

engines

or

Hall

thrusters,

but

the

ionization

mechanism

is

different

yet

again.

In a

FEEP

thruster,

a

liquid

metal

propellant

(Cs)

is fed by

capillary

forces

through

a

small

channel.

The

channel

ends

forming

sharp

edges

that

are

located

opposite

a

negative

electrode,

separated

by a

small

gap

(about

1 mm)

from

the

channel

tip.

The

channel

structure

itself

carries

a

positive

potential.

An

electric

field

develops

between

the two

electrodes

and the

free

surface

of the

liquid

CS

metal

column

near

the tip of the

channel

deforms,

forming

cusps

which

protrude

from

the

surface

of the

liquid.

As the

liquid

forms

ever

sharper

cone

structures

due to the

action

of the

electric

field,

the

local

electric

field

strength

near

these

cusps

increases.

Once

a

local

electric

field

strength

of

about

107

V/cm

is

reached,

electrons

are

ripped

off

the Cs

metal

atoms.

The

electrons

are

collected

14

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

through

the

liquid

metal

column

and the

channel

walls,

and

the

positive

ions

are

accelerated

away

from

the

liquid

columns

through

a gap in the

negative

electrode

by the

same

electric

field

that

created

them.

Accelerating

voltages

between

the two

electrodes

typically

reach

values

of

around

10,000

V,

resulting

in

specific

impulse

values

up to

8,000

s48.

Rectangular,

slit-

shaped

channel

geometries

are

most

frequently

investigated

in

the

laboratory

.

Depending

on

slit

width

and

voltages

applied,

thrust

levels

between

10"6

and

10"3

N

have

been

achieved.

Required

power

levels

for the

thruster

itself

as low

as

60W/mN

have

been

achieved.

In

addition

to

thruster

power,

power

has to be

provided

to a

neutralizer.

The

neutralizer

is

required

to

prevent

spacecraft

charging

while

emitting

the

positive

ion

beam.

Different

neutralizer

concepts

have

been

tested,

among

them

a

hollow

cathode

type

and a

tungsten

filament

configuration47.

In the

hollow

cathode

type,

which

requires

less

power

than

the

tungsten

filament

type,

a Cs

compound

is

heated,

creating

a Cs

vapor

that

is

ionized

in a

hollow

cathode

discharge,

providing

the

electrons

required

for

beam neutralization.

This

neutralizer

type

delivers

0.1

mA per

Watt

of

electric

power,

including

heater

power.

Advantages

of

FEEP

devices

for

microspacecraft

primary

propulsion

applications

are its

high-specific

impulse

and

compact

size

and

propellant

storage

(liquid

metal).

Disadvantages

are

potential

contamination

issues

due

to the use of Cs

propellant,

and

high

voltage

and

power

requirements, requiring

dedicated

power

processing

units

adding

to the

propulsion

system

weight.

FEEP

thrusters

also

require

a

fairly

narrow

operating

temperature

range

to

avoid

solidification

of the Cs

propellant.

Table

11

lists

two

FEEP

thrusters

under

development

at

Centrospazio

in

Italy,

currently

the

only

provider

of

FEEP

thrusters.

As can be

noted,

thrust

specific

power

requirements

are

about

90W/mN

for the

smaller

model

and 75

W/mN

for the

larger

thruster.

Given

that

the

neutralizer

requires

1 W per 0.1 mA, and ,

according

to

Petagna

et

al.49,

an

emitter

current

of

7mA/mN

is

required,

the

power

required

per mN of

thrust

is

(90W/mN

+

10W/rnA*7mA/mN),

i.e.

160W/mN

or 145

W/mN

in

case

of

the

larger

thruster. These power

requirements

will

limit

achievable

thrust

values

with

a

FEEP

thruster

on

board

a 20

W

Class

I

spacecraft

to

about

0.125

-

0.14

mN. At

thrust

values

this

low,

thrust-to-spacecraft

mass

ratios

will

be

around

0.006

mN/kg.

The

Europa

orbiter

mission

currently

being

studied

at JPL

uses solar

elelctric

propulsion

(SEP)

and

has

thrust-to-weight

ratios

of

around

0.23

mN/kg

at the

beginning

of the

mission,

dropping

to

about

0.03

mN/kg

at

the

end of the

mission50.

These

values

are

significantly

Table

11:

State-of-the-Art

FEEP

Thrusters

Manufacturer

Propellant

Thrust

(uN)

Isp

(s)

Power

(W)

Thruster

Mass

(kg)

PPU

Mass

(kg)

Thruster

Size

(cm)

PPU

Size

(cm)

Total

Impulse

(Ns)

Centrospazio

Cs

100

8000

9

0.45

2.9

1.2x1.2x0.8

1.2x1.0x0.6

160

Centrospazio

Cs

800

8000

60

3.5

4.1

2.5x2.5x1.5

2.5x1.3x1.5

60,000

higher

than

those

obtainable

with

a

state-of-the

art

FEEP

system

onboard

a

microspacecraft.

Using

FEEP

thrusters

for

microspacecraft

primary

applications

will

therefore

require

higher

power

levels

than

those

assumed

in

this

study

in

Section

II. In

addition,

the

potential

Cs

contamination

issues

will

have

to be

studied

and

evaluated

for

each

mission.

Total

impulse

capabilities

appear

low for the

smaller

of the two

thrusters

listed

in

Table

11

(corresponding

to

about

450

hrs

of run

time

assuming

steady-state

nominal

thrust

values),

however,

are

quite

high

for the

larger

thruster

(corresponding

to

over

20,000

hrs run

time

assuming

nominal

steady-state

thrust

conditions).

The

latter

would

be

sufficient

for

electric

primary

propulsion

applications.

PPU

masses

appear

compatible

with

the

microspacecraft

mass

allocations.

Future

work

in

field

emission

thruster

technology

is

focusing

on the use of

microfabricated

emitter

arrays45

consisting

of a

series

of

micro-"volcano"

structures

on a

wafer.

The

significance

of

these

arrays

is

that

ions

can be

produced

at

much

lower

voltages,

reducing

power

requirements.

Since

extractable

currents

from

each

micro-

emitter

are

much

lower

than

those

obtained

with

conventionally

machined

emitters,

arrays

of

many

of

these

emitters

will

be

required

to

operate

in

parallel.

Colloid

Thrusters

Colloid

thrusters

were

studied

extensively

during

the

late

1960s

and

early

1970s

for

spacecraft

attitude

control

and

drag-makeup,

but due to

their

failure

to

produce

high

enough

thrusts

at

reasonable

power

levels

fell

out of

favor.

With

the

advent

of

microspacecraft

designs,

a

potential

application

for

microspacecraft

primary

propulsion

may

have

arrived.

In

a

colloid

thruster,

thrust

is

produced

by

electrostatically

accelerating

fine

liquid

droplets

ejected

from

15

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

a

capillary51'52.

A

strong

electric

field

applied

between

the

sharp-edged

exit

of the

capillary

and an

external

electrode

causes

charge

separation

inside

the

liquid

propellant,

which

in

most

cases

is

doped

with

an

additive

to

increase

its

electric

conductivity.

Through

a

combination

of

hydrodynamic

instabilities,

causing

jet

break-up

into

small

liquid

droplets,

and the

action

of the

applied

field

acting

on

the

conductive liquid,

charged

droplets

are

extracted

from

the

capillary

at

high

velocities,

producing

thrust.

Depending

on the

propellant

used,

either

positive

or

negative liquid

droplets

can be

produced.

Most

applications

studied

in the

past

used

glycerol

doped

with

sodium

iodine

to

produce

positive

droplets

and

glycerol

doped

with

2 - 10 %

sulfuric

acid

to

produce

negative

droplets.

The

ability

to

produce both

positive

and

negative

droplets

was

termed

a

"bi-polar"

thruster

by

Perel

et

al.51.

Its

significance

is

that

it can

potentially

be

self-neutralizing,

provided

the

same

amount

of

current

can be

drawn

from

each

set of

capillaries,

eliminating

the

need

for a

separate

neutralizer.

Several

trade-offs

have

to be

made

in the

design

of a

successful

colloid

thruster

to

optimize

performance.

In

order

to

obtain

good

performance,

the

specific

charge,

measured

in

Coulomb

per

droplet

mass,

has to be

high

in

order

to

obtain

high

specific

impulses

at

reasonable

applied

voltages.

The

colloid

thruster

is an

electric

thruster,

and as

such

additional

propulsion

system

masses

associated

with

the

power

supply

or

conditioning

will

have

to be

offset

by

sufficient

propellant

mass

savings

that

can

only

be

obtained

through

a

high

enough specific

impulse.

Specific

impulse

in a

colloid

thruster

is

also

determined

by the so

called

specific

charge

efficiency,

which

measures

the

distribution

of

specific

charge

in

a

droplet

stream.

A

more

"peaked"

specific

charge

distribution will

lead

to

higher

specific

impulses

and

higher

propulsion

system

efficiencies.

Specific

charge

efficiency

in

turn

depends

on

several

parameters,

such

as

electric

field

strengths

(higher

electric

field

strengths

reduce

the

efficiency

since

they

create

higher

charged

particles

in

addition

to

lower

charged

ones),

flow

rate

(lower

mass

flow

rates

result

in

higher

specific

charge

efficiencies),

fluid

conductivity

(a

higher

conductivity

leads

to

lower

specific

charge

efficiencies

since

it

becomes easier

to

produce

more

charges

in a

more

conductive stream)

and

capillary

tip

design

(which

affects

the

specific

charge

efficiency

mostly

through

its

effect

on the

local

electric

field

strength

near

the

tip)51.

Some

of

these

design

considerations

work

against

each

other.

A

large

potential

drop

caused

by a

strong

electric

field,

for

example,

will

accelerate

the

droplets

to

a

higher

exhaust

velocity

(which

would

help

to

raise

the

specific

impulse),

however,

it

will

also

decrease

the

specific

charge

efficiency

(which will

lower

the

specific

impulse

again).

Low

flow

rates,

aiding

in

obtaining

higher

specific

charge

efficiencies

and,

thus,

specific

impulses,

will

also

reduce

thrust.

In

addition

to

these

performance

considerations,

careful

propellant

selection

will

need

to be

made

to

ensure

proper

thruster

function

and

long

lifetime.

High

solvation

capability

(to

take

up

dopants),

low

vapor

pressure

(to

avoid

crystallization

of

dopants

on

capillary

walls

near

tip,

potentially

clogging

the

system),

low

freezing

point

(to

avoid

clogging)

and low

corrosivity

(to

ensure

long

thruster

lifetime)

are key

parameters

in the

selection

of the

propellant51.

The

glycerol

propellants discussed

above

were

found

to

have

a

somehwat

high

vaporization

pressure,

but

were

chosen

for its

superior

ability

to

dissolve

dopants51.

Platinum

capillaries

have

been used

because

of

their

resistance

to

corrosion51.

Based

on the

design

considerations

above,

colloid

thrusters

have been

designed

and

operated with platinum

capillaries

having

ID's

of

about

200

microns,

using

sodium

iodine

and

sulfuric

acid

doped

glycerol

propellants,

producing

thrusts

between

0.2 - 0.5

mN

at

power

levels

of

about

4.4

W/mN,

requiring

voltages

of

+4.4

kV and

-5.8

kV,

depending

on

droplet

polarity51.

Specific

impulses

were

estimated

between

450 - 700

sec51.

In

other

cases,

specific

impulses

of up to

1350

sec at

thrust

levels

of

0.55

mN

were

obtained

(power

levels

were

not

reported

in

that

case)52.

Of

all

micro-electric

primary

propulsion

options

reviewed

so

far, colloid

thrusters

are

quite

possibly

the

most

suited

for

microspacecraft

primary

propulsion

applications

at

this

stage

of

micro-electric

propulsion

development.

Power

requirements

are

much

lower

than

those

of

FEEP

devices

and

fit

well

within

Class

I

constraints.

Thrust

levels

of 0.5 mN

can

easily

be

achieved

using

only

about

2 W of

power

in a

self-neutralizing,

bi-polar

array51.

Thruster

specific

masses

(excluding

power

supply

and

conditioning)

have

been

estimated

at 0.2 - 0.5

kg/W50,

and

would

also

fit

within

Class

I

microspacecraft

envelopes.

However,

specific

impulse

values

are

somehwat

low for

interplanetary

applications.

Thus,

future

work

should

focus

on

obtaining

higher

specific

impulse

values.

Attitude

Control

-

Chemical

Cold

Gas

Thrusters

Cold

gas

thrusters

represent

the

smallest

rocket

engine

technology

available

today.

Cold

gas

systems

are

valued

for

their

low

system

complexity,

small

I-bit

and

thrust

capability

and the

fact

that,

when

using

benign

16

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics, Inc.

propellants

(e.g.

N2),

they

present

no

spacecraft

contamination

problems.

Sometimes

high

reliability

is

also

referred

to as one of the

advantages.

However,

valve

leakage

problems

have

resulted

in

repeated

losses

of

spacecraft

due to

premature

depletion

of

propellant

through

valve

internal

leaks.

This

leakage

problem

is a

result

of a

combination

of

small

amount

of

microscopic

contaminants

(to be

found

in

even

the

cleanest

propulsion

system)

and

high

pressure

propellant

storage.

On the

contamination

side,

the

propellant

tank

is one of the

major

contaminant

sources

(microscopic

metal

flakes

left

over

from

fabrication,

etc.).

These

contaminants

may

locate

themselves

on

valve

seats,

carried

along

with

the

propellant

flow,

and

subsequently

prevent

the

valve

from

sealing

completely.

Even

though

the

remaining

opening

across

the

valve

seat

may be so

small

that

for a

liquid

propellant

application

it

would

pose

no

problem

due

to

the

higher

liquid

viscosity,

in

cold

gas

systems,

where

the

propellant

is

stored

at

very

high

pressures,

propellant

may

escape

even

through

these

microscopic

openings.

Another

disadvantage

of

cold

gas

systems

are

their

low

specific

impulse

performances,

unless

very

light

gases

(H2,

He) are

used.

Neither

hydrogen

or

helium

is

commonly

used,

however,

since

storage

problems

and

large

and

heavy

tankage

would

result

due to low gas

densities,

and

additional

leakage

concerns

would

have

to be

considered

with

these

light

gases.

Table

12

gives

a

list

of

typical

cold

gas

performance

values,

based

on

data

found

in

Refs.

8 and 26.

Of

the

gases

listed,

nitrogen

is by far the gas

most

frequently

used

as a

cold

gas,

due to a

combination

of

reasonable

propellant

storage

density,

performance

and

lack

of

contamination

concerns.

Table

13

lists

some

of the

smallest

cold

gas

thrusters

available

today.

The

Moog

58x125

thruster

is

merely

4.3

cm

in

length

and 1.4 cm

max. dia.,

including

fitting.

The

fitting

accounts

for

roughly

half

of the

size

of the

total

envelope.

Size,

mass

and

power

requirements

fit

well

within

the

Class

I

microspacecraft

envelope.

However,

even

a

cold

gas

thruster

this

size

may

perform

only

marginally

with

respect

to the

impulse

bit

requirements

(compare

data

in

Table

13

with

comments

in

Section

II).

Using

the

data

found

in

Tables

12 and 13,

required

leak

rates

for

microspacecraft

can be

estimated

and

current

cold

gas

thruster

performances

be

evaluated

in

this

regard.

Using

the

information

provided

in

Section

II,

a

total

"delta-

v"

requirement

of 50

m/s

is

assumed.

For a

specific

impulse

of 70 sec

(N2)

and an

assumed

microspacecraft

mass

of

10 kg

(Class

I), the

required

attitude

control

propellant

mass

is 0.7 kg of

nitrogen.

At a

storage

density

of

0.28

g/cm3

for

nitrogen

at

3500

psia

and 0 C, a

tank

volume

of

about

2500

cm3

is

required.

Taking

into

account

the

possibility

of

propellant

leakage,

assume

that

10%

more

propellant

is

loaded

onto

the

spacecraft,

now

requiring

a

tank

volume

of

2750

cm3

at the

same

storage

pressure.

Assuming

a

spherical

tank,

this

translates

into

an

inner

tank

diameter

of

roughly

37 cm.

This

tank

size

is

slightly

larger

than

the

Table

12:

Cold

Gas

Propellant

Performances

Propellant

Hydrogen

Helium

Neon

Nitrogen

Argon

Krypton

Xenon

Freon

12

Freon

14

Methane

Ammonia

Nitrous

Oxide

Carbon

Dioxide

Molecular

Weight

(Kg/Kmol)

2.0

4.0

20.4

28.0

39.9

83.8

131.3

121

88

16

17

44

44

Density

(3500

psia

0 C)

(g/cm3)

0.02

0.04

0.19

0.28

0.44

1.08

2.74'"

-

0.96

0.19

liquid

-

liquid

Isp*

(Theoretical)

(s)

296

179

82

80

57

39

31

46"

55

114

105

67*

67

Isp'

(Measured)

(s)

272

165

75

73

52

37

28

37

45

105

96

61

61

at

25 C.

Assume

exoansion

to

zero

pressure

in

case

of

theoretical

value.

at

38 C

(560

R) and

area

ratio

of

100.

Likely

stored

at

lower

pressure

values

(2000

psia)

to

maximize

propellant

to

tank

weight

ratio.

17

Copyright©

1997,

American Institute

of

Aeronautics

and

Astronautics,

Inc.

Table

13:

Small

Cold

Gas

Thrusters

Manufacturer

Type

Thrust

(N)

Ibit

(Ns)

Isp

(s)

Pressure

(kPa)

Open

Response

(ms)

Power

(pull-in)

(W)

Weight

(g)

Moog55

58x125

0.0045

io-4

65

(N2)

34.5

0.94

2.4

7.34

Moog"

58x115

2.89

-

-

1460

3.5

(spec)

30

13

Marquardt7

-

4.5

-

-

8840

<1.1

-

5.4

envelope

assumed

for a

10-kg

spacecraft,

but is

within

the

right

range.

The

tank,

however,

will

dominate

the

spacecraft

design

layout

completely.

Assuming

further

that

all of the

additional

10% of the

propellant

may be

lost

over

the

course

of

the

mission

(corresponding

to 250

cm3

at

3500

psia

storage

pressure

or

almost

59,000

sec,

assuming

zero

compressibility

of

nitrogen

in

this

rough

estimate),

maximum

allowable

leak

rates

for a 2

year

mission

would

be

9 x

IO'4

scc/s

and for a 3

year

mission

6 x

IO"4

scc/s.

These

small

leak

rate

requirements

are a

consequence

of the

small

spacecraft

size.

Since

smaller

spacecraft

carry

smaller

onboard

propellant

supplies

for the

same

attitude

control

requirements,

less

propellant

can be

afforded

to be

lost

due to

leakage

and

valve

leak

rates

for

microspacecraft

consequently

have

to be

even

lower

than

for

bigger

spacecraft.

Recent

leak

tests

with

cold

gas

thrusters

have

shown

leak

rates

lower

than

the

ones

calculated

above53,

however,

at

thruster

operating

pressures

which

where

a

mere

34.5 kPa,

or

about

one

third

of an

atmosphere.

The low

leak

rates

calculated

above

will

have

to be

maintained

throughout

the

entire

feed

system,

however,

in

particular

across

components

that

are

exposed

to the

full

tank

pressure.

A

cold

gas

thruster

valve

recently

developed

by

Marquardt7

is

subjected

to

pressures

of

about

8900

kPa

(1300

psig)

and

maintains

leak

rates

of

<2.8

x

IO"2

scc/s

GHe7,

corresponding

to

roughly

1

x

IO"2

scc/s

GN2,

assuming

a

ll^jM

(M

being

the

molecular

weight)

dependence

for the

leak

rate.

These

values

are far

higher

than

the

estimated

leak

rate

requirements

given

above.

The

leak

rate

requirements

could

be

relaxed

if

more

propellant

reserves

were

carried.

Performing

the

calculations

above

for the

same

case

of a 10 kg

microspacecraft

requiring

a 50

m/s

attitude

control

budget

over

the

course

of the

mission,

but

raising

the

propellant

margin

to 50%

would

yield

maximum

allowable

leak

rates

of 4 x

IO"3

scc/s

in the

case

of a 2

year

mission

and 2.5 x

IO"3

scc/s

for a 3

year

mission.

While

these

leak

rates

approach

obtainable

values,

it

has to be

noted

that

these

rates

will

have

to be

maintained

over

the

course

of the

entire

mission,

even

after

the

valve

has

been

subjected

to

many

cycles

and

substantial

propellant

flow,

carrying

contaminants

through

the

valve.

In

addition,

raising

the

propellant

reserves

from

10% to 50%

leads

to a

bigger

required

tank

diameter

of 41 cm ID

compared

to 37

cm

before.

These

tank

designs

will

completely

dominate

the

spacecraft

design

layout

and

consume

a

substantial

portion

of

the

overall

spacecraft

mass

and

volume.

An

interesting

alternative

to

conventional

cold

gas

propulsion

using

high

pressure

gas

tanks

is the use of

ammonia

as a

propellant.

As

pointed

out by

Nakazono54,

ammonia

has a

vapor

pressure

of 33

psia

(224

kPa)

at -18

C.

Thus,

even

without

tank

heaters,

sufficient

pressure

could

be

provided

to an

ammonia

cold

gas

thruster

merely

using

the

boil-off

of the

propellant.

As can be

seen

from

Table

12,

specific

impulses

obtainable

with

ammonia

are

higher

than

those

achievable

with

nitrogen.

The

ammonia

system

would

allow

for

liquid

storage,

reducing

tank

size

and

mass

and

propellant

leakage

concerns.

Depending

on

available

vaporization

rates

(dependent

on

tank

temperature),

propellant

flow

rates

may be

limited39.

It may

therefore

be

concluded

that

cold

gas

propulsion

systems

using

high-pressure

gas

supplies

do not

appear

to

be a

viable

option

for

microspacecraft

unless

attitude

control

requirements

can

be

substantially

reduced,

valve

leak

rates

can

be

lowered

by at

least

one

order

of

magnitude,

and

severe

mass

and

volume

constraints

on the

remainder

of the

microspacecraft,

caused

by

large

and

heavy

propellant

tanks,

can

be

tolerated.

Cold

gas

systems

based

on

liquid

storage

of

ammonia,

on the

other

hand,

appear

as a

very

attractive

option

for

microspacecraft

attitude

control.

In

either

case,

obtainable

input

bits

have

to be

reduced

further,

even

for

Class

I

microspacecraft

applications.

This

requires

the

development

of

either

faster

valves

or

smaller

nozzle

throat

areas.

Fabricating

nozzle

throat

diameters

smaller

than

the

ones

obtainable

today

may

require

the

exploration

of new

technologies,

such

as

MEMS

(see

Section

V).

Warm

Gas

Thrusters

Cold

gas

technology

discussed

above

could

be

adapted

to

warm

gas

thruster

concepts.

In

this

case

hydrazine

propellant

is

decomposed

in a

separated

gas

generator,

consisting

in

essence

of a

Shell

405

catalyst

bed,

and

gaseous

hydrazine

decomposition

products

are

then

fed to

a

plenum

and

finally

through

a

cold

gas

thruster54.

Such

a

system

would

not

require

separate

propellant

tanks

if a

conventional

hydrazine

thruster

was

used

as the

microspacecraft

primary

propulsion

device.

In

addition,

several

cold

gas

thrusters

in

existence

today

already

claim

18

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

compatibility

with

hydrazine

decomposition

products.

A

separate

heater

would

be

required

to

heat

the

catalyst

bed in

order

to

allow

for a

sufficient

number

of

starts,

just

as in

conventional

hydrazine

thruster

technology.

By

feeding

the

decomposition

products

into

a

plenum,

from

which

they

can

be

drawn

to the

various

attitude

control

thruster

clusters

upon

demand

would

lower

the

required

number

of

catalyst

starts54.

A

hydrazine

warm

gas

system

is a

very

attractive

option

for

microspacecraft,

in

particular

when

a

hydrazine

propellant

supply

is

already

onboard

for

primary

propulsion

purposes.

Relatively

high

performance,

comparable

to the

ammonia

cold

gas

system

described

above,

can be

combined

with

compact

propellant

storage

and the

relatively

near

term

availability

of the

required

propulsion

components,

drawing

upon

cold

gas

heritage.

As in the

case

of

cold

gas

systems,

either

faster

valves

or

smaller

nozzle

orifices

are

required

to

lower

impulse

bits

to the

requirements

for

both

Class

I and

Class

II

microspacecraft

applications.

Tri-propellant

Thrusters

In

a

tri-propellant

thruster,

a

propellant

mixture

of

hydrogen,

oxygen

and an

inert

gas,

such

as

helium

or,

more

commonly

for

propellant

storage

reasons,

nitrogen

is

used7'57.

The

propellants

are

stored

fully

mixed,

no

separate

tanks

are

required.

The

addition

of the

inert

gas to the

mixture

renders

the

mixture

non-combustible,

until

exposed

to a

suitable

catalyst.

Different

catalysts

are

being

studied,

typically

based

on

nobel

metal

compounds27'58,

although

details

on

catalyst

compositions

are

mostly

treated

as

proprietary

information.

Thruster

performances

range

between

70 s and 140 s of

specific

impulse57.

A

recent

design

by

French7

is

aimed

at 125 sec

Isp

and a

thrust

of

about

2N.

However,

this

design

was not

specifically

intended

for

microspacecraft

use.

Tri-propellant

systems,

even

though

able

to

deliver

higher

performance

than

cold

gas

systems,

suffer

the

same

disadvantage

of

high-pressure

propellant

storage

and the

associated leakage

problems

as

cold

gas

systems.

In

addition,

even

though

required

propellant

masses

can be

reduced

over

conventional

cold

gas

systems

due to the

higher

specific

impulse

performance,

required

high-pressure

propellant

tanks

will

likely

continue

to

dominate

the

spacecraft

design.

Thus,

the

advantages

gained

with

the use of

tri-propellant

systems

over

cold

gas

systems

onboard

microspacecraft

may be

limited.

Hydrazine

Mono-Propellant

Thrusters

Hydrazine

mono-propellant

thrusters

readily

available

today

(see

Table

8) are far too

large

and

heavy

to be

used

in

attitude

control

clusters

around

a

microspacecraft.

The

development

of

new,

miniature

hydrazine

thrusters

is

required.

A

research

initiative

to

study

the

feasibility

of

miniature

hydrazine

thrusters

was

recently

begun

at

JPL,

but

it is too

early

to

present

results

from

this

acitivity.

Bi-propellant

Thrusters

Bi-propellant

thrusters

were

not

considered

for

microspacecraft

attitude

control.

Currently

available

engine

technology

is far too big and

heavy

for use in

microspacecraft

attitude

control

clusters.

Bi-propellant

engines

have

been

considered

for

attitude

control

purposes

on

larger

spacecraft

to

allow

for

easier

integration

into

primary

propulsion

systems,

eliminating

separate

attitude

propellant

supplies.

It was

already

concluded,

however,

that

bi-

propellant

systems

are

probably

not

suitable

for

onboard

microspacecraft

primary

propulsion.

In

addition,

even

though

bi-propellant

thrusters

do

offer

higher

performances,

reducing

propellant

requirements,

required

propellant

masses

for

attitude

control

are

usually

small,

not

providing

an

opportunity

for

large

spacecraft

mass

reductions.

Even

if a

substantial

reduction

of the

typically

small

attitude

control

propellant

mass

could

be

achieved,

it

would

likely

be

offset

by

the

higher

dry

mass

of a

bi-propellant

system

due to

increased

component

parts

count

when

compared

with

much

less

complex

cold

or

warm

gas,

or

even

miniature

hydrazine

systems.

Finally,

in

view

of the

survey

of

state-of-the-art

miniature

bi-propellant

technology

given

above,

there

exists

considerable

doubt

whether

further

substantial

reductions

in

thruster

size

can be

made

while

still

being

able

to

provide

reliable,

space-qualifyable

engine

technology.

Attitude

Control

-

Electric

Pulsed

Plasma

Thrusters

(PPTs)

In

a

pulsed

plasma

thruster,

electrical

power

is

used

to

ablate,

ionize

and

electromagnetically

accelerate

atoms

and

molecules

from

a bar of

solid

Teflon50'59.

The

Teflon

bar is

pushed

against

a

retaining

lid

between

two

electrodes

by

means

of a

negator

spring.

The

electrodes

are

connected

to a

capacitor,

which

is

unable

to

discharge because

the

vacuum

19

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

and

solid

Teflon

bar

between

the

electrodes

do not

provide

a

conductive

path.

A

spark

plug

located

near

the

solid

Teflon

surface

is

fired,

removing

a

portion

of the

Teflon

and

ionizing

it. In the

process,

the

Teflon

bar is

pushed

forward

toward

the lid and

brought

in

position

for the

next

pulse.

The

capacitor

now

discharges

through

the

ionized

Teflon

gas.

The

particles inside

the

Teflon

discharge

consist

of a

variety

of

molecular

fluorocarbons,

which

are

ionized

in the

process

and

give

rise

to a

current

flow

between

the

electrodes.

This

discharge

current

generates

a

strong

magnetic

field

surrounding

it.

Lorentz

forces

acting

on the

discharge

current

as a

result

of

this

magnetic

field

push

the

Teflon

plasma

out of the

thruster

at

high

exhaust

velocities

of

about

10 - 20

km/s.

Non-ionized

particles

are

expanded

from

the

thruster

due to

Joule-heating.

Thrusts

generated

per

pulse

are on the

order

of 10s

to

100s

of

micro-newtons.

However,

because

the

pulse

length

is on the

order

of

milliseconds,

the

capacitor

can be

charged

and

discharged

several

times

per

second, thus

creating

an

accumulative

thrust

in the

micronewton

to

millinewton

range.

Pulsing

frequencies

between

1 Hz and 6

Hz

are

common.

The

ability

of a

PPT

to

produce

very

small

thrust

levels

per

pulse

allows

for the

possibility

to

deliver

very

small

impulse

bits

on the

order

of

less

than

1

mNs.

Future

designs

are

aimed

at

providing

merely

10s of

|J.Ns.

PPT

thrusters

have

been

developed

since

the

1950s

and

have

flown

on

several

US

satellites.

State-of-the

art

thruster

are

able

to

provide

specific

impulses

in the

range

of 800 -

1500

s,

thrusts

of 220 -

1100

\iN,

and

efficiencies

between

5 - 15

%

at a wet

mass

of 5 kg.

Current

PPT

designs

are

therefore

too

large

for

microspacecraft

attitude

control.

Future

designs

are

being

contemplated

and are

predicted

to

have

wet

masses

as low as

0.5 kg.

Providing

12 of

these

thrusters

for

attitude

control

on

a

microspacecraft

will

lead

to a

total

(wet)

system

mass

of

6 kg.

These

mass

values

are

compatible

with

Class

I

mass

guidelines.

With

minimum

impulse

bits

for the

future

,

0.5 kg

thrusters

predicted

at 10 - 100

|iNs,

PPT are

also

able

to

provide

the

minimum

impulse

bit

requirements

in

the

upper

Class

I

microspacecraft

category.

However,

minimum

thrust

requirements

for

attitude control

of

Class

I

spacecraft,

ranging

between

1.75

- 4.5

mN

for the

larger

spacecraft

masses,

can

only

be met at

high

power

levels.

According

to

Ref.

59, a

1.75

mN

thrust requires

a

power

level

of 120 W and

pulsing

frequencies

between

3-6

Hz,

depending

on

capacitor

size.

These

values

far

exceed

the

power

levels

that

will

be

available

for

attitude

control

on

any

of the

microspacecraft

considered

here.

PPTs

can

thus

only

be

used

on

Class

I

craft

if

slew

rate

requirements

can be

reduced

substantially. Current

and

predicted

PPT

technologies

are too

large

and

heavy

for

Class

II

spacecraft.

Field

Emission

Thrusters

Field

emission,

or

FEEP,

thrusters,

surveyed

above,

are

reviewed

here

in

terms

of

their

applicability

as

attitude

control

thrusters

on

microspacecraft.

Large

PPU

masses

and

high

power

requirements

may

prohibit

their

use

in

such

a

function.

Following

the

example

given

in the

cold

gas

thruster

section

above,

for a

typical

Class

I

microspacecraft

mission,

0.7 kg of

nitrogen

gas

would

be

required

to

meet

a 50

m/s

delta-v

budget.

If a

FEEP

system

with

a

specific

impulse

of

8,000

s was

used,

the

required

propellant

mass

would

be

reduced

to

12

g.

While

this

is a

substantial

reduction,

the

required

PPU

mass

is,

according

to

Table

11,

is 2.9 kg for the

smaller

of the two

FEEP

thrusters

considered.

Since

Cs

propellant

can be

stored

in its

liquid

state,

whereas

nitrogen

gas

would

have

to be

stored

in

high-pressure

tanks,

tank

weight

reduction

will

benefit

the

FEEP

system.

According

to

recent

tank

data60,

a

3,244

cm3

tank

(roughly

the

size

required

for

nitrogen

storage

in the

example

above),

capable

of

maintaining

a

maximum

expected

operating

pressure

(MEOP)

of

10,000

psia

(far

more

than

required

for a

storage

pressure

of

3,500

psia

as

assumed),

weighs

around

1.8 kg

(these

data

are

based

on a

cylindrical

tank).

Even

neglecting

tank

masses

for the

FEEP

system,

the

cold

gas

system

will

still

be

lighter

than

the

FEEP

system

by

about

0.4 kg

(although volume

requirements

will

be

higher

due to the

large

tank

volume).

In

addition,

the

FEEP

system

will

not be

able

to

provide

the

minimum

thrust

for the

assumed

slew

rate

requirements

within

the

power constraints

expected

to be

found

on a

microspacecraft.

FEEP

systems

may

find

applications

as

attitude

control

devices

for

microspacecraft

if

slew

rate

requirements

can be

relaxed

significantly

and

very

small

impulse

bits

are

required.

Colloid

Thrusters

Much

the

same

comments

as

made

with

respect

to

the

applicability

of a

field

emission

systems

to

microspacecraft attitude

control

can be

made

for

colloid

systems.

Minimum

thrust

requirements

cannot

be met , and

additional

power

processing

unit

masses

may

offset

propellant

reductions

gained

over

cold

gas

systems

through

the use of the

higher

specific

impulse

colloid

thruster.

Since

current

colloid

thrusters

provide

lower

specific

impulses

and

require

less

power,

the

comparison

between

colloid

and

cold

gas

systems

may be

shifted

somewhat

more

in

favor

of the

colloid

system

than

was the

case

for the

FEEP

system.

20

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

Resistojets

State-of-the

art

resistojet

technology22

is far too

heavy,

and

requires

far too

much

power

(in

excess

of 350

W)

to be

useful

for

microspacecraft

attitude

control.

However,

some

work

on

small

water

resistojets

was

performed61"63.

Work

focused

on

measuring

small

nozzle

performances

using

water

vapor.

No

heater

power

requirements

were

reported.

Water

is not the

most

suitable

propellant

for

resistojet

use to its

high

heat

of

vaporization,

even

though

it

does

simplify

laboratory

testing

due to its

lack

of

safety

and

toxicity

concerns.

Table

14,

based

on

data

from

Ref.

26,

list

several

relevant

properties

of

candidate

propellants

for a

resistojet

system.

Of the

propellants

listed,

ammonia

and

water

immediately

stand

out due to

their

low

molecular

weight, resulting

in

high

specific

impulse

performance.

Of

these

two

propellants,

ammonia

requires

less

heat

to

vaporize

and is

thus

the

propellant

of

choice.

Since

no

small

resistojet

technology

exists,

it is

not

clear

how

well

this

technology

is

suited

for

microspacecraft

use.

Liquid

storage

of

propellants

will

reduce

system

weights

over

high-pressure

cold

gas

storage

systems

.

Comparing

ammonia

resistojets

with

the

ammonia

cold

gas

thruster

discussed

above,

the

resistojet

should

enable

higher

duty

cycles

and

longer

burns,

making

resistojet

technology

more

versatile.

The

disadvantage

is

their

need

for

a

separate

power

supply, adding

to the

system

weight.

Work

is

currently

underway

at JPL and the

Aerospace

Corporation

to

develop

micro-resistojet

technology

using

MEMS

technologies

(see

Section

IV).

IV.

FUTURE

DEVELOPMENT

NEEDS

AND

EMERGING

TECHNOLOGIES

Based

on the

thruster

survey

performed

in

Section

III, technology

needs

for

microspacecraft

propulsion

can be

identified.

Several

emerging

micropropulsion

technologies

are

being

introduced.

These

emerging

technologies

are the

first

propulsion

components

designed

specifically

with

microspacecraft

applications

in

mind.

Identification

of

Technology

Needs

The

results

of the

survey

conducted

in

Section

ffl

are

summarized

in a

matrix,

shown

in

Table

15.

Only

state-

of-the-art

technologies,

or

those

under significant

development,

are

listed

in

this

matrix.

The

technologies

listed

were

evaluated

in

view

of

their

application

to

Class

I

and

II

microspacecraft

attitude

control

and

primary

propulsion

applications

by

grouping

them

into

three

categories:

those

technologies

that

appear

applicable

to the

task

("yes"

category),

those

that

do not

("no"

category),

and

those

that

fall

somewhat

in

between

("maybe").

The

latter

category

serves

to

classify

technologies

that

may

fulfill

some

mission

requirements,

but not

others,

or

those

that

could

be

made

to

fulfill

all

requirements

for the

specific

Table

14:

Properties

of

Candidate

Resistojets

Propellants

Propellant

Ammonia

Propane

Ethyl

Chloride

Butane

Freon

12

Water

Hydrogen

Fluoride

Methanol

Methyl

Chloride

Ethane

Ethyl

Methyl

Ether

Mono

Methyl

Amine

Formula

NH3

C3H8

C2H5C1

C^HIO

CC12F2

H2O

HF

C2H3OH

CH3OH

C2H8

C2H5OCH3

CH3NH2

Molecular

Weight

(kg/Kmol)

17.0

44.1

64.5

58.1

120.9

18.0

20.1

44.0

51.0

30.0

60.0

31.0

Liquid

Density

(g/cm3)

0.6

0.49

0.92

0.57

0.98

1.0

0.99

0.79

0.91

0.56

0.8

0.77

Heat

of

Vaporization

(kj/kg)

1159.7

339.3

388.1

360.2

141.8

2442.5

1505.9

1099.3

376.5

313.7

350.9

873.8

21

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

Table

15:

Matrix

of

Applicability

Status

of

State-of-the-Art Technologies,

or

Technologies

under

Significant

Development,

to

Microspacecraft

Technologies

(State-of-the-Art

Only)

Bi-Propellant

Hydrazine

HAN/TEAN

Peroxide

Solid

Hybrid

Cold

Gas

Ammonia

Cold

Gas

Warm

Gas

Tri-Propellant

Ion

Hall

FEEP

Colloid

PPT

Resistojet

Class

I

Primary

No

Yes

Maybe

No

Yes

Maybe

No

No

No

No

No

No

Maybe

Maybe

No

No

ACS

No

No

No

No

No

No

Maybe

Maybe

Maybe

Maybe

No

No

Maybe

Maybe

Maybe

No

Class

II

Primary

No

No

No

No

Maybe

No

No

No

No

No

No

No

No

No

No

No

ACS

No

No

No

No

No

No

No

No

No

No

No

No

No

No

No

No

microspacecraft

category

and

task

with

only

minor

technology

advances

required.

Inspecting

Table

15, the

lack

of

suitable

thruster

technology

for

microspacecraft

applications

becomes

obvious.

This

is not too

surprising

a

result,

given

that

most

of

the

technologies

reviewed

here were developed

for

spacecraft

much

larger

than

the

spacecraft

considered here.

Some

technologies stand

out,

however,

and may

appear

applicable

to

microspacecraft even

in, or

close

to,

their

current

design

stage.

For

primary

Class

I

applications,

the

smallest available

hydrazine

attitude

control

thrusters

appear

as

a

suitable

and

reliable

thruster

option

if

only

small

to

intermediate

delta-v's

are

required.

Solid

motors

also

may

provide

low to

intermediate

delta-v

capability

within

their

current

design

limits,

although thrust

values

may

need

to be

reduced

further.

However,

no

thruster

options

providing

large

delta-v

and

compatible

with

the

microspacecraft assumptions

are

currently

available.

Both

bi-propellant

chemical

as

well

as ion in

their

current

manifestations

would

require separate

propulsion

stages

due to

high

system

masses

and

volumes

resulting

in

larger

launch

masses.

Possible

exceptions

are

FEEP

and

colloid

thrusters.

While

FEEP

thrusters

still

require

too

much

power,

and

colloid thrusters deliver

fairly

low

specific

impulses,

they

may be

marginally

suitable

at

this

point

of

their

development,

but

will likley

require

further

development.

PPT

thrusters

do not

have

the

total

impulse

capability

and

also

require

too

high

power levels

for

the

minimum

thrust

levels

needed.

For

Class

I

attitude

control

applications,

cold

gas

and

warm

gas

systems

currently

offer

the

greatest

near-term

potential.

In

both

cases,

however,

delivered

impulse

bits

are

too

high. Impulse

bits

maybe

reduced

through

faster

valve

technologies

or

increased

flow

restriction through

the use of

smaller nozzle throats

or

separate

flow

restrictors.

The

last

option

may be the

easiest

to

achieve,

but

fabricating

small

orifices

with

low

tolerances

and

protecting

them

from

clogging

through

contaminants

may

require

the use of new

technologies

(MEMS).

PPTs,

FEEPs

and

colloid

thrusters

may be

applicable

for

Class

I

spacecraft

attitude

control

only

when

slew

rate

requirements

can be

lowered dramatically.

Given

that

current

microspacecraft

mission

requirements

may

still

undergo

considerable

evaluation

and may

vary

from

mission

to

mission, these three technologies have

been

listed

as

possible

candidate

technologies.

If

slew rate requirements

will

not

change

significantly,

large reduction

in

power

requirements

and PPU

specific

masses

will

be

needed.

For the

Class

II

category,

virtually

no

state-of-the-

art

propulsion

technologies

appear

suitable.

Once

again,

the

reason

for

this

is

that current hardware

was

designed

for

22

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

larger

spacecraft

and

currently

available

components

may

even

exceed

the

size

of a

typical

Class

II

spacecraft.

The

only

primary

propulsion

device

that

may be

applicable

to

Class

II

spacecraft

using

state-of-the-art

technology

(yet

still

requiring

a

full

development

program)

may be

solid

motors.

For

attitude

control,

no

current

propulsion

hardware

appears

suitable,

either

because

the

size

of

available

components

do

not

allow

for the

distributed

mounting

of a

dozen

or so

thruster

around

a

Class

II

bus,

or

because

delivered

minimum

impulse

bits

are too

high.

Therefore,

high-specific

impulse

primary

propulsion

for

both

spacecraft

categories,

as

well

as low I-

bit

attitude

control

thruster

technology

for

Class

I, but in

particular

Class

II, are

identified

as

major

propulsion

technology

needs

for the

microspacecraft

considered

here.

Class

II

attitude

control

applications

will

also

require

extreme

miniaturization.

Emerging

Technologies

The

increasing

interest

in

microspacecraft

for

deep-

space

missions

has led to the

exploration

of

technologies

suitable

for

these

spacecraft

at

JPL,

NASA's

lead

center

for

robotic

interplanetary

space

exploration

and one of the

first

potential

users

of

these

technologies.

Several

new

micropropulsion

technologies

were

proposed

in the

course

of

these

activities.

Three

of

these

emerging

micropropulsion

technologies

are

discussed

here.

Among

them

are two

micro-attitude

control

phase-change

thrusters

(vaporizing

liquid

or

resistojet

and

subliming

solid)

and

micro-ion

engines

for

high

delta-v

primary

applications.

All

three

technologies

rely

heavily

on

MEMS-fabrication

techniques,

allowing

for the

potential

of

order-of-magnitude

reductions

in

component

mass

and

size,

possibly

suitable

for

both

Class

I as

well

as

Class

II

applications.

Associated

with

each

technology

considered

here,

however,

are

significant

feasibility

issues

that

will

have

to be

thoroughly

investigated

in

coming

years.

Thus,

the

thruster

technologies

introduced

here

have

to be

viewed

as

very

advanced

micropropulsion

options.

Vaporizing

Liquid

Micro-Thruster

The

vaporizing

liquid

micro-thruster,

or

micro-

resistojet,

is a

concept

that

was

proposed

by

Leifer

and

Mueller

, as

well

as

Janson

at the

Aerospace

Corportaion,

a

couple

of

years

ago7'64.

In

this

thruster

concept,

described

in

detail

in a

companion

paper65,

a

suitable

liquid

(ammonia

or

hydrazine)

is

heated

via a

thin-film

heat

exchanger,

micro-

fabricated

onto

a

silicon

substrate.

A

conceptual

sketch

and

picture

of the

assembled

JPL-device

are

shown

in

Figs.

2

and

3. Two

identical

silicon

wafers,

featuring

thin-film

heaters

and

micro-nozzles,

are

bonded

to a

Pyrex

spacer

that,

in

its

final

assembly,

is

sandwiched

between

the two

silicon

wafers.

The

liquid

propellant,

pressure-fed

through

one of the

openings

machined

into

one of the

silicon

wafers,

enters

the

thruster

and is

vaporized

as it

flows

between

the

heater

elements.

Propellant

vapor

is

then

exhausted

through

the

second

nozzle.

A

recess

machined

into

the

silicon

underneath

the

heater

creates

thermal

chokes

near

the

heater

edges,

reducing

conductive

losses

to the

structure.

Silicon

has a

very

high

thermal

conductivity,

providing

a

bigger

challenge

for

the

thermal

design

of

this

device.

However,

no

other

material

can

currently

be

micromachined

to the

degree

and

flexibility

that

silicon

can,

and

silicon

was

thus

chosen

as

the

substrate

material.

A

key

aspect

of

this

design

is its

simplicity.

It

does

not

contain

any

complex

moving

parts

(such

as

MEMS-pumps,

turbines,

etc.)

which

could

decrease

the

Intel

Thin-Film

Heater

(1of2)

Nozzle

Silicon

Substrate

Fig.

2:

Concept

of the

Vaporizing

Liquid

Thruster

Fig.

3:

Vaporizing

Liquid

Micro-Thruster

Chip

23

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

reliability

of the

device.

Performance

targets

are 0.5 to

several

mN

thrust

(to

meet

minimum

thrust

requirements

for

microspacecraft

attitude

control),

50%

efficiency

and

power

requirements

of a few W or

less.

The

latter condition,

dictated

by

available

assumed

power

levels

onboard

a

microspacecraft,

will

limit

performance.

Specific

impulse

values

around

100 to 150 sec

have

been

estimated

depending

on

propellant

and

heater temperature

obtainable.

However,

given

that

delta-v

budgets

for

attitude control

are

fairly

low

(see

Section

31),

high

performance

is not

necessarily

required.

Impulse bits

will

largely

be

determined

by the

valve

technology

to be

used.

Presently,

no

suitable MEMS

valve

technology

exists

to be

interfaced

with

this

thruster,

but

miniature

conventional

valves,

approaching

in

both

size

and

mass

of

fully

packaged

MEMS

valve

technology,

may

be

used

initially.

With

open-to-close

times

on the

order

of a

few

ms,

impulse

bits

may

range

as low as

10"5

Ns or

less.

Work

on

integrating MEMS technology

with

conventional

valve technology

is

currently

being

explored

at

Marotta

Scientific

Controls,

Inc

under

an

SBIR

program.

Several prototype devices

( see

Fig.

3)

have

been

assembled

and are

awaiting

testing

this

summer.

Focus

of

the

initial

work

is to

optimize heater

designs, search

for

design

options

to

reduce

power

losses

to the

structure,

determine

minimum

heater

lengths,

test

the

reliability

of the

electric

contacting through

thruster

cycling

and

gain

initial

thrust

performance

data.

Work

on

micro-resistojets

is

also

underway

at the

Aerospace

Corporation7.

These

devices

are

similar

in

function,

but

feature

slightly

different

heater

designs7.

Subliming

Solid

Micro-Thruster

Subliming

solid

thruster

concepts

are not new and

substantial

development

work

was

performed

with

these

thrusters

in the

1960's.

Main contributors

to

this

field

were

Rocket

Research66"68

(now

Primex),

Lockheed69'74

(now

Lockheed-Martin),

and

NASA

Goddard75'76.

Some

work

was

also

performed

at

Aerospace

Industries77,

the

Lewis

Research

Center78

and the

Martin-Marietta

company79

(now

Lockheed-Martin

also).

In the

subliming

solid

thruster

concept,

a

solid

propellant

is

chosen

with

a

high

sublimation

pressure,

such

as

ammonium

hydrosulfide

(NH4HS)

or

ammonium

carbamate

(NH4CO2NH2)-

Upon

heating,

gas

pressure builds

up

inside

the

propellant

tank

and

the

vapor

is

vented

through

a

valve

and

nozzle

to

produce

thrust.

The

simplicity

of

this

design,

and the

solid

storability

of the

propellant

appear

to

easily

lend

themselves

to

miniaturization.

Based

on the

1960's

work

in

this

area,

a

subliming

solid

micro-thruster

concept

was

proposed

by the

author

using

MEMS

technology80.

A

conceptual

drawing

of

the

device

is

shown

in

Fig.

4. The

chip

consists

of two

layers,

one

made

from

silicon

the

other

from

Pyrex.

The

silicon

portion

of the

chip

was

machined

at

Sandia

Nat'1

Labs.

Propellant vapor

enters

the

chip

from

the

tank

through

the

circular

Pyrex

hole

and

flows

along

a

recess

machined

into

the

silicon

towards

the

nozzle

orifice.

On its

way,

the

propellant

passes

through

a

micromachined

comb

filter.

This

filter

will

prevent

solid

particles

from

drifting

out of the

tank

and

towards

the

nozzle

in the

zero-g

environment

of

space,

preventing

nozzle

blockage.

The

micro-nozzle

has a

throat area

of

about

50 x 50

ftm

and is

shown

in

Fig.

5. The

square

nozzle

shape

is a

result

of the

anisotropic

etch

used

in the

fabrication

of

this

device, resulting

in

preferential

etching

of

some

crystal

planes

over

others.

The

wall

surfaces

of the

nozzle

are

composed

of

{111}

planes,

etching

slowest

in the

fabrication

process.

Using

this

technique,

simple

converging-diverging

nozzle

shapes

can

easily

be

fabricated.

Nozzle shapes

are not

optimized

in

this design, since

performance

optimization

is

not

a

goal

with

the

current,

first

generation

of

chips,

and

future

design iterations

may

explore

different

contours.

A

mock-up

of the

thruster concept

is

shown

in

Fig.

6,

consisting

of a

tank,

a

valve

and the

thruster chip

assembly.

The

valve

shown

in

Fig.

6 is a

MEMS-type

valve

based

on

thermopneumatic

action,

designed

by

Redwood

Microsystems,

Inc.

This

valve

has

leakage

issues

and

currently

only

serves

as a

place

holder.

It

should

be

pointed

out,

however,

that

due to the

solid

propellant

storage

and

relatively

low

vapor

pressures

inside

the

tank,

leak

rate

requirements

for

valves

may be

relaxed

significantly

over

those

required

for

cold

gas

systems.

Entrance

Hole

'.*>

Propellant

Met

(from

Tank)

Pyrox

Cover

Nozzle

Entrance

{Square

Nozzle

Contour)

PropeBanlOulet

(Through

Nozzle

In

Si-

Substrate

Fig.

4:

Subliming

Solid

Thruster

Concept

24

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

Fig.

5:

Subliming

Solid

Micro-Thruster

Nozzle

and

Filter

Fig.

6:

Mock-up

of a

Subliming

Solid

Micro-thruster

System

Propellant

condensation

may

occur

as

propellant

vapors

are

exposed

to the

cold

surfaces.

In the

event

that

propellant

condensation

leads

to

clogging

of the

filter,

a

heater

was

thin-film

deposited

just

underneath

the

filter

on

the

opposite

side

of the

silicon

substrate

(facing

up in

Fig.

6).

Similar

heaters

may be

required

around

the

valve

to

prevent

recondensation

on the

valve

seat,

preventing

sealing.

Valve

heating

will

necessitate

the

selection

of a

different

valve

design

than

the one

shown

in

Fig.

6

(with

no

thermal

activation

as in the

case

of the

Redwood

valve

shown

in

Fig.

6

allowed).

Advantages

of the

subliming

solid

micro-thruster

concepts

are its

compact

size

and

propellant

storage,

light

weight

tankage

due to low

pressure

requirements,

relative

immunity

to

leak

rate

concerns

and

suitable

projected

performances.

Thrust

values

of

about

0.5

mN

up to

several

mN

are

targeted

with

this design.

If the

chip

can be

interfaced

with

fast

acting

valves

very

low

impulse

bits

would

result.

Disadvantages

of the

system

are its low

performances

(50-

75 s

Isp

estimated)

and

toxicity

issues

associated

with

the

propellants,

requiring

special

handling

during testing

and

propellant

loading.

Micro-Ion

Engines

High

specific

impulse

micropropulsion

devices

will

be

required

in

order

to

achieve

high-delta-v

capability

for

microspacecraft.

MEMS-based

devices

have

been

studied81'82,

however,

were

found

to

have problems associated

with

electron

wall

losses.

In

addition,

for the

thrust

and

power

levels

required,

conventionally

machined

thrusters

may be

used7.

MEMS technology,

however,

does

hold

promise

for

use in

various

ion

engine

components.

MEMS

accelerator

grid

system

technology

is

currently

being

studied

at

JPL.

Using

this

technology,

ion

beams

may be

steered

electrostatically

by

designing

a

special

third,

or

decelerator

grid,

of a

three-grid

system,

allowing

for the

application

different

electric

potentials

to

different

grid

section

(see

Fig

7).

In

order

to

fabricate

MEMS-scale

grids,

electric

breakdown

voltages

of

MEMS

materials,

such

as

silicon

oxide,

will

have

to be

tested.

Insulator

materials,

such

as

Decel

Grid

Accel

Grid

^

Scree

Grid

n

Fig.

7:

Concept

of

Electrostatic

Gimbaling

25

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

silicon

oxide,

will

be

needed

as

grid

insulator

materials

since

free

standing

grids

may not be

practical

at

these

dimensions.

Figure

8

shows

a

test

chip

recently

completed

at JPL to

test

silicon

oxide

breakdown

strengths.

The

oxide

layer

is

deposited

between

two

electrodes

(a

polysilicon

layer

underneath

the

oxide

layer,

accessible

through

a via

etched

through

the

oxide,

and an

aluminum

layer

on top of the

oxide).

Applying

voltage

to the two

contact

pads

shown

in

Fig.

8

will

allow

silicon

oxide

breakdown

voltages

to be

tested.

An

integral

heater

allows

for

breakdown

tests

at

various

temperatures.

Other

areas

in the

development

of

miniature

ion

engine

technology

that

require

attention

in the

future

ate

micro

power

processing

units

and

micro

flow

control

and

feed

systems.

In the

latter

area,

micro-ion

engine

systems

(and

conventional

systems

due to

their

low

flow

rate

requirements)

are

likely

to

benefit

from

the

integration

of

MEMS

technologies.

Another

SBIR

contract

performed

by

Marotta

Scientific

Controls,

Inc

addresses

the

issues

of

microflow

control

on a

chip

in a

device

termed

a

"micro

gas

rheostat."

V.

CONCLUSIONS

Existing

thruster

technologies

were

reviewed

in the

view

of

potential

applications

for

microspacecraft,

defines

as

spacecraft

with

masses

of 1 - 20 kg.

Based

on

this

review,

technology

needs

were

defined

and

several

emerging,

advanced

micropropulsion

concepts

specifically

designed

for

microspacecraft

applications

were

introduced.

Only

a few of the

currently

existing

thruster

technologies

appear

applicable

for

spacecraft

of the

size

considered

here.

For

primary

propulsion

applications,

small

hydrazine

thrusters

and

solid

motors

may

provide

Fig.

8:

Grid

Breakdown

Test

Chip

intermediate

to low

delta-v

capability.

Thrust

values

of

solid

motors

may

have

to be

reduced

further

in

order

to

avoid

excessive

spacecraft

accelerations.

FEEP

and

colloid

thruster

may

possibly

be

used

as

primary

propulsion

devices

if

power

requirements

can be

lowered

in the

case

of

FEEP

devices,

and

specific

impulse

can be

raised

for

colloid

thrusters.

For

attitude

control

functions,

currently

available

cold

gas

systems

approach

in

performance

the

requirements

imposed

by

microspacecraft

designs

with

respect

to

minimum

thrust

and

impulse

bit

values.

Impulse

bits,

however,

will

have

to be

lowered

even

further

beyond

the

values

obtainable

with

todays

smallest

thrusters

for

Class

n

microspacecraft.

In

addition,

significant

leakage

concerns

exist

for

cold

gas

systems

and the

required

high-pressure

storage

tanks

will

completely

dominate

microspacecraft

design

with

respect

to

both

size

and

mass,

even

for

relatively

benign

attitude

control

requirements.

Ammonia

cold

gas

thrusters

or

hydrazine

warm

gas

systems

may

provide

fairly

near-term

solutions

to the

propellant

storage

and

leakage

issues.

For

spacecraft

with

masses

of

less

than

5 kg,

virtually

no

suitable

propulsion

hardware

exists,

neither

for

primary

propulsion,

nor

attitude

control.

In

many

cases,

existing

thrusters

are

larger

than

the

spacecraft

in

question.

Future

technology

needs

for

both

primary

and

attitude

control

propulsion

will

be

required,

in

particular

for

spacecraft

below

5 kg in

mass.

However,

even

for

larger

spacecraft

(5 - 20

kg),

significant

improvements

for

high-

specific

impulse

primary

propulsion

will

need

to be

made

in

order

to

reduce

propellant

masses

,

thus

aiding

in

keeping

microspacecraft

small.

Also,

further

reductions

with

respect

to

I-bit

performance

are

required.

New

advanced

micropropulsion

technologies

introduced

include

both

attitude

control

thrusters

of the

vaporizing

liquid

(resistojet)

and

subliming

solid

thruster

types,

and

micro-ion

engines

for

primary

propulsion

applications.

All

devices

make

heavy

use of

advanced

microfabrication

techniques,

such

as

silicon

micromachining.

Several

devices

were

designed

and

built

over

the

course

of the

past

year

and

represent

the

latest

technology

advances

in the

micropropulsion

area,

specifically

aimed

at

addressing

microspacecraft

technology

needs.

VI.

ACKNOWLEDGMENTS

The

author

would

like

to

thank

Ross

Jones

from

JPL and the

NASA

Advanced

Propulsion

Concept

program

for

supporting

this

work.

He

also

wishes

to

thank

William

Tang,

Wen Li,

Indrani

Chakraborty,

Lilac

Muller,

Thomas

26

Copyright©

1997,

American

Institute

of

Aeronautics

and

Astronautics,

Inc.

George,

Andrew

Wallace,

Russell

Lawton,

Dave

Bame,

Morgan

Parker

and

John

Blandino

for

their

contributions

to

the JPL

micropropulsion

activities,

and

John

Brophy

for

reviewing

this

paper.

The

work

described

in the

paper

was

conducted

at

the Jet

Propulsion

Laboratory

under

contract

with

the

National

Aeronautics

and

Space

Administration.

VH.

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1997,

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