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216

J. SPACECRAFT

VOL. 18, NO. 3

A I A A 8 1 -4 1 5 1

Effect

of Duty Cycle on Catalytic Thruster Degradation

H. C. Hearn*

Lockheed Missiles & Space Company, Inc.

Sunnyvale,

Calif.

The thermal cycling and low-temperature operation inherent in three-axis stabilization attitude control can

lead

to

several forms

of

hyd razine catalytic thruster degradation including catalyst attrition

and

voiding

catalyst bed comp action chamber pressure spiking and injector tube thermal effects. Each of these mechanism s

is

influenced

b y the

thruster operating characteristics

or

i r i n g

duty cycle

and the

duty cycle plays

a

m ajor role

in

determining the

type

and

rate

of

thruster degradation. Test exp erience show s that p articular thruster designs

m ay

be greatly affected b y a change in duty cycle and that testing over a

variety

of

duty

cycles is required to

demonstrate insensitivity to m ission ap plication.

Introduction

I

T is

general ly known within

th e

industry that

th e

life

of

hydrazine catalyt ic thrusters

is

dependent

on the way

they

ar e operated.

Total

impulse, number

of

pulses, pulse width

an d frequency, operat ing temperature, and the operating

environment all contr ibute to thruster wear an d degradation.

This duty cycle dependenc e

ha s

become increasingly evident

in

th e

application

of

these thrusters

to

three-axis stabilization

of

spacecraf t ; differences have been observed in the rate of

per fo rmance degradat ion

and in the

type

of

anomalies

ex -

perienced. It has become clear, for example, that demon-

stration of great capability in one application does not

necessarily imply

a

similar capability under

a di f ferent

type

o f

use. These

findings

have significance with regard

to the

goal

of

producing s tandardized thrusters that will satisfy th e

requirements of a variety of missions and thus minimize costs

fo r thruster development and quali f icat ion.

This paper addresses th e

effects

of duty cycle on per-

formance of

5-lbf (22-N) thrusters used

for

three-axis

thermal effects. Catalyst bed poisoning due to aniline or

decomposition gases

is not

included since data relative

to

5-lbf

(22-N) thrusters

are not

conclusive. Poisoning

has

been

identified

in

primarily

0.1-0.2-lbf (0.44-0.89-N)

thrusters

operating

at

temperatures below 400

°F

(478

K), and

this

phenomenon

has

been discussed

i n

other papers .

3> 4

Following

th e treatment of these individual effects, th e overall in -

teraction between duty cycle

an d

thruster design

will be

illustrated

in the

section

on

du ty-cycle sensitivity.

Catalyst Attrition and V oiding

Although

a

duty cycle

can be

characterized

in

terms

of

cold

starts, pulse width, pulsing frequency, thermal cycling, and

other variables,

the key

parameter

appears to be

catalyst

bed

temperature. Analyses

an d

experiments dealing with catalyst

breakup

5

show that thruster temperature

is a

significant

factor

in

several mechanisms, including catalyst particle

wetting with liquid hydrazine, impingement of liquid

hydrazine on a hot particle, an d differential thermal ex -



M A Y - JU N E 1981

C A T A L Y T I C T H R U S T E R D E G R A D A T I O N

21 7

u

£

100

-

H

f

u

Q

LU

Q Q

at

50

-

PULSES/DAY

D

2 5 0 0

x

O

3 0 0 0

A

4000

+ 5000

_ 0 6000

X

s~~ ~~

J

[) 10 20 30

A

6000 1-—

• • ^ ^ c T

i i i

i

i

40 50 60 70 80

TOTAL IMPULSE Ib

f

-sec, THOUSANDS)

Fig. 1 Effect of duty cycle on catalyst loss.

D

z:>u

D

u

22

°

S

I

F

r

o

o

O 1.0

A

T

H

R

U

S

R

o

0

1° °

U

AA

A

A

A

A

/ A

A

^

A

>0(

b 1

o o ^

0

00

0

0

o

°co

i

o

0

INJ

TUBE

CLOGCIN

CATALYST

^COMPACTION

INJECTOR

r

/THERMAL

~

CHOKING

tr

I 1

3 100 200 300 400 500 600

PULSES THOUSANDS)

the engine. Figure 2 shows that performance did steadily

degrade unti l the thruster was shut down after 536,000 pulses .

The 73%

reduction

in steady-state thrust was

accompanied

b y

a comparable drop in

pulsing impulse

bit, and

there

w as

also

a

gradua l decrease in steady-state 7

sp

.

The relatively

detailed

analysis devoted to this thruster

during the

test

revealed

that

bed

com pact ion

was the primary

cause

of thrust

degradation even

before disassembly and

injector

flow were conducted.

Since the thru ster w as receiving

a progressively lower than

normal

flow dur ing the test, th e

problem was to

distinguish between

catalyst bed effects an d

injector tube clogging. It was

known

that

some clogging was

occurring

based

on previous experience, a review of chamber-

pressure traces, and the effect of thruster

temperature

on

impulse bit.

However ,

for injector-passage

clogging

to explain

th e

entire

thrust decay would have required a decrease of

about 0.012

in .

(0.305

mm) in the

original 0.017-in. (0.432

mm) diameter injector

passages.

Since this would have been

abou t

three times the reduction experienced in earlier testing,

it w as

considered

unlikely. Also,

7

sp

variations

during the test

were consistent

with

catalyst effects.

Fo r

example, following

a

repressurization

from

120 to 200

psia (827-1379 kPa) ,

the

next

health

check

at 90 psia

(621

k Pa )

revealed a 25 f low-rate

increase

accompanied by a 6-s

increase

in

7

sp

. Subsequent

to

this event, th e

flow

rate

an d

7sp

resumed a steady decrease.

Although the 7

sp

is

affected

by flow

rate

to a

small

degree, th e

magnitude of the changes observed was more consistent

with

th e effect of catalyst fines on bed impedance an d ammonia

dissociation.

In fact , th e shifts in ammonia dissociation

required

(approximately

15-20%) are within th e

range

which

has

been observed

i n

other

thruster testing.

The

repressur izat ions tha t

were a

part

of the

mission duty

cycle provided

a way in

which

to

separate

th e

effects

of bed

compaction from thermal choking in the injector passages.

When the

feed

pressure was increased

significantly,

the in-

creased

flow

and injector cooling

eliminated

the hydrazine

boiling in the injector which

caused part

of the

thrust

reduction

at low

feed pressures.

A

catalyst

bed

flow

coef-

ficient (describing the

physical

state of the bed) can be



218

H . C . H E A R N

J . S P A C E C R A F T

200

160

a:

a.

C Q

0 .2 0.3 0 .4

TIME SEC)

Fig. 3

Chamber

pressure spik e.

1000

ID

I-

<

5

X

u

3 PSIA

FEED PRESSURE

1

PSIA

H

HIGH ST

SPIKE

AMPLITUDES

SEVERE I

SPIKING I

DECREASING AMPLITUDE

I

I

I

I I

10

15

2 2 5

Fig. 4

PULSE

NUMBER

Effect of thruster temp erature on spik ing.

3000

2000

CQ

1

1

20

30

70

NUMBER OF MANEUVER SEQUENCES

Fig.

5

Effect

of

duty

cycle on spiking.

Z

o

200

100

I

1 2 3 4 5 6 7 8 9 1



M A Y - JU N E

1981

C A T A L Y T IC T H R U S T E R D E G R A D A T I O N

219

Pressure spiking has been

shown

to be

more than

just a

curiosity; it can have a significant im pact on thruster life and

performance. Experiments

have shown

5

that spiking causes

pressure and temperature transients which

subject catalyst

particles to

severe

impact

loads,

pressure crushing, an d

abrasion.

These effects result in particle fracture and

generation of fines at a rate dependent on magnitude and

number of spikes. The

role that

spiking can play in catalyst

attrition and performance has been clearly demonstrated in

thruster

testing

(discussed in a

later

section) and by

flight

experience. In the

case

of one

flight

thruster ,

severe spiking

resulted in the rupture of a chamber pressure transducer line.

In another case,

intermittent

thruster valve leakage was at-

tributed to

spiking

magnitudes greater than the

feed

pressure.

Equat ions

were derived

to model the

injector

tube

fluid

dynamics resulting from pressure spiking in the

catalyst

bed

and predictions were

made

regarding the

transport

of

particle

contaminat ion back to the valve seat area. The analysis

verified that the duration and

magnitude

of the spikes ob-

served

was sufficient to

reverse

th e

injector

flow an d cause

valve

contaminat ion.

Injector Tube Thermal Effects

A general review of

thruster

flow anomalies due to

con-

tamin ants in the

propellant

an d

feed

system was reported in a

recent

paper .

9

That review

covered

instances

of tube

plugging

an d flow

reduction

as well as

thruster

valve leakage fo r

primarily 0.1-Ibf

(0.44-N) level thrusters. This paper

will

examine

thermal

anomalies that

have

been

observed on

5-lbf

(22-N)

thrusters

an d that ar e associated with contaminant

deposition in injector

tubes.

M o st of the deposition i s believed

to occur after a pulse is

completed,

as

propellant

evaporation

in

the small

injector passages

leads to an accumulation of

soluble contam inants on the tube walls . Hydrazine

impu ri t ies

that

have

been implicated are the nonvolatiles,

with

some

experience

indicating a predominance of silicon

compounds .

The

roughening

of the tube surface can lead to an increase in

heat t ransfer to the propellant, resulting in boiling and two-

phase f low. This

thermal

choking

can cause sharp

reductions

u.uz

g

l/>

V4-

.0

~ 0.01

Q

O

U

~ 0.0

03

I

P -0.01

0

- -0.02

0%

< b

00°

0 0 00

0

0

00

o °

0

I

I I I I

100

500

00 300 400

PULSES (THOUSANDS)

Fig.

7

Impulse

b it

degradation.

that

involved

0.025-,0.050-, an d

0.100-s

pulsing

at a rate of 1

per second. This pulsing w as

performed

subsequent to steady-

state firings which resulted in high injector

temperatures

of

approximately 775°F (686 K) prior to the pulsing. On two

occasions

when th e initial

temperature

was higher

than

usual

[approximately 865°F

(736 K )] and the feed

pressure

was low

[approximately 110

psia

(758 kPa)], the thruster chamber

pressure and impulse

bits

were sharply reduced during the

pulsing;

a

significantly

lower

nozzle

temperature at the end of

each

sequence was

indicative

of

reduced flow

through the

engine. On subsequent pulsing health checks where the initial

injector temperature was lower, the

chamber

pressures and

impulse bits returned

to normal . It appears that when th e

initial

injector

temperature is

above

a certain level, the pulsing

flow

rate is reduced by boiling in the passages and thu s is not

sufficient to

cool

th e injector an d

allow impulse recovery

to

norm al values. The

occurrence

of these phenom ena illustrates

th e importance of evaluating each mission

application

fo r

dut y cycles that might cause adverse thruster

performance.



220

H.C. H E A R N

J. SPACECRAFT

Q

01

CO

20

10

THRUSTER

A

10

20

30

50

TOTAL

IMPULSE

(lb-sec, THOUSANDS)

Fig. 8 Duty-cycle sensitivity—thruster A .

30

20

10

THRUSTER B

III

10 20 30 40 50

T O T A L I M P U L S E

(l b

f

-sec, T H O U S A N D S )

Fig.

9

Duty-cycle

sensitivity—thruster

B .

thermal

choking.

Although the possibility of some change in

u

Q

50

40

30

20

10

THRUSTER C

10

20

30 40

50

TOTAL

IMPULSE

(lb

f

-sec,

THOUSANDS)

Fig. 10

Duty-cycle sensitivity—thruster

C .

formance

degradation and

early

test termination. Although

thruster

C was not tested to these

specific duty cycles,

th e

A F R P L

test

at

2500

pulses/day

showed

that it was probably

the best of the three

thrusters

for the Type I

duty cycle.

Type II is a 90-min

duty cycle involving

a

mixture

of

minimum

impulse bits

an d

longer

pulse widths

while

the

level

of activity

an d

thruster

temperature is much

lower

than Type

I. Type Ila

involved

an average

rate

of about 65 0

pulses/day

and a

temperature range

of

275-650°F (408-617

K).

Type lib

had a

rate

of 4 20 pulses/day and a temperature range of 225-

350°F

(381-450

K). It is first seen from th e figures tha t th e

thrusters exhibit a higher catalyst loss rate on the Type-II duty

cycle

than

on Type I; this is consistent

with

the lower tem-



M A Y - JU N E 1981

CATALYTIC THRUSTER

DEGRADATION

221

occurred after valve closure,

and the

total delivered impulse

w as 4 0

greater

than normal.

Analysis

showed

that flow into

the thruster was

higher

due to the

lower

pressure in the

chamber

and the

liquid

hydrazine

tha t

collected in the thruste r

subsequently

decomposed and

resulted

in a late and ab-

normally

high impulse bit.

The test results described here

reveal

a situation in which a

dif ferent

thruster

design was more suitable for

each

of the

three types of duty cycle.

Conversely,

it was not

possible

to

select a

single

design

that

would

perform

satisfactorily

for all

duty

cycles.

For example, the change in

duty cycle

showed

that

one

thruster design

w as

mor e

prone to

pressure spiking;

this led to a sharp increase in the catalyst attrition rate,

creating an increase in

flow

impedance as

well

as hazardous

operat ion. Although the

results

obtained here are not

necessarily

transferrable

to

other

thruster

designs, they serve

to illustrate the role

that

duty cycle plays in thruster per-

formance and

life. They

also

help

to explain why

claims

of

duty cycle insensitivity

have been greeted with skepticism

in

cases where

a particular thruster

design

has not

actually

been

tested to a

variety

o f duty cycles.

Conclusions

The degradation and types of anomalies

experienced

in

hydrazine catalytic thrusters are determined by the mission

duty cycle and length of

operat ion.

Experience with 5-lbf

thrusters

shows

that the thermal

cycling

and low-temperature

operation inherent in three-axis stabilization

duty cycles

can

be a source of concern with regard to catalyst attrit ion, bed

compact ion,

pressure spiking, and injector tube effects.

These

degradation modes

ar e

influenced

by operating

temperature,

feed

pressure,

pulse width, and pulse frequency and are

themselves

interdependent. It is significant that a particular

thruster

design

may be greatly affected by a change in

duty

cycle,

an d that a design

shown

to be superior in one ap-

plication may not be optimum should

requirements

change.

Therefore,

testing over a

variety

of du ty

cycles

is required to

demonstrate relative insensitivity to mission app lication.

A c k n o w l e d g m e n t

The p reparat ion of

this

paper was aided by W. Zeber, who

provided information and

helpful

comments and who has

made

significant

contributions

in the areas of thruster

spiking,

injector-tube

effects, and thruster life prediction.

References

^ackheim,

R.L., Carlson, R.A., M ackl in ,

H., and Lee, D.H.,

Evolution of a S tandardized M onoprope l lan t

Rocket Engine Design

Concept,

AIAA/SAE 10th

Propulsion Conference,

Paper

74-1133,

Oct.

1974.

2

Pylan t , A.J., Evanson, B.E. , and Hood, J.F., Product Im -

provement of a 5-Pound

Thrust

R C S

Reaction

Engine Module,

AIAA/SAE 10th Propulsion Conference, Paper 74-1137, Oct. 1974.

3

H o l c o mb , L ., M a t ts o n , L. , and Oshico, R. ,

Ef fects

of Aniline

Impurities on M onopropel lant Hydrazine Thruster Performance,

Journal of

Spacecraft

and

Rockets Vol. 14,

M arc h

1977, pp.

141-148.

4

Ri c h a r d s ,

R.E. and Grabbi , R .,

0.2Ib

M onoprope l lan t

Hydrazine Thruster

Test

Results with M IL-Grade and Viking-Grade

Propellant,

AIAA/SAE 12th

Propuls ion

C onference ,

Paper 76-657,

July

1976.

5

Tavlor, W.F.

et

al., External

Catalyst

Breakup

Phenomena,

AFRPL-TR-76-47, June 1976.

6

M o s el ey , V . A . ,

Fricke,

H.D.,

Hin te rman ,

T.H.,

Long

Life

Monopropellant Design Criteria, AFRPL-TR-75-45 , July 1975.

7

Saenz, A., Life Eva lua t ion of Long-Life

Hydrazine

Engines ,

JANNA F Propuls ion

M ee ting,

Vol. 3,

CPIA

Pub. 300, July 1979.

8

Greer,

H., Vacuum

Startup of Reactors for Catalytic

Decom-

position

of

Hydrazine, Journal of

Spacecraft

a nd

Rockets Vol. 7,

M ay 1970,

pp. 522-528.

9

Schre ib , R.,

Flow Anomalies in

Small Hydrazine Thrusters,

AIAA/SAE/ASME 15th

Joint Propulsion

Conference ,

Paper 79 -

1303, June 1979.

10

Technical Interchange

M eeting

a t LM S C ,

Rocket Research

C o.

presentation,

Docum ent No. 79-H-245,

Feb. 16, 1979.

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