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8/19/2019 Effect of Duty Cycle on Catalytic Thruster Degradation_archive
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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 -
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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
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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
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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.
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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-
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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.