of small rockets. Most high thrust rocket models assume the rocket behaves as an inviscid core with a thin boundary layerl, 2. For relatively large rockets, these.
:0
/,/ NASA
Technical
Memorandum
/-7
106281
! J
AIAA-93-1825
A Laboratory Model of a Hydrogen/Oxygen Engine for Combustion and Nozzle Studies Sybil Lewis
Huang Morren Research Center
Cleveland,
Ohio 41%
Roger
M. Myers
Sverdrup
Technology,
Lewis
Research
Brook
Park,
,--4
I C"
Inc.
Center
Z
0
Group 0
Ohio
Stephen E. Benko Lewis Research Center Cleveland, Lynn
Ohio 0
_
Technology,
_
LLaLLJ
Inc.
Research Center Park, Ohio
Group
..J L3 _i ") _Z Z
"" C?, _4"
and
,4Z3!
Brian
D. Reed
Lewis
Research
Cleveland, Prepared 29th
I---
A. Arrington
Sverdrup Lewis Brook
Ltl t/_ 7..w
Joint
Center
Ohio for the Propulsion
v
Conference
sponsored
by the AIAA,
Monterey,
California,
SAE,
June
and Exhibit ASME,
28-30,
1993
and ASEE
'--_
I
_
c_
I
(_ _=_
_- LL v
A Laboratory
Model
of a Hydrogen/Oxygen Studies
National
Engine
for Combustion
and Nozzle
Sybil H. Morren Aeronautics and Space Administration Lewis Research Center Cleveland,
Ohio
44135
Roger M. Myers Sverdrup Technology, Inc. Lewis Research Center Group Brook Park, Ohio 44142
National
Stephen E. Benko Aeronautics and Space Administration Lewis Research Center Cleveland,
Ohio 44135
Lynn A. Arrington Sverdrup Technology, Inc. Lewis Research Center Group Brook Park, Ohio 44142 and Brian National
D. Reed
Aeronautics and Space Administration Lewis Research Center Cleveland,
Ohio
44135
Abstract A small laboratory diagnostic thruster was developed to augment present low thrust chemical rocket optical and heat flux diagnostics at the NASA Lewis Research Center. The objective of this work was to evaluate approaches for the use of temperature and pressure sensors for the investigation of low thrust rocket fiow fields. The nominal engine thrust was 110 N. Tests were performed at chamber pressures of about 255 kPa, 370 kPa, and 500 kPa with oxidizer to fuel mixture ratios between 4.0 and 8.0. Two gaseous hydrogengaseous oxygen injector designs were tested with 60% and 75% fuel film cooling. The thruster and instrumentation designs were proven to be effective via hot fire testing. The thruster diagnostics provided inner wall temperature and static pressure measurements which were compared to the thruster global performance data. For several operating conditions, the performance data exhibited unexpected trends which were correlated with changes in the axial wall temperature distribution. Azimuthal temperature distributions were found to be a function of operating conditions and hardware configuration. The static pressure profiles showed that no severe pressure gradients were present in the rocket. The results indicated that small differences in injector design can result in dramatically different thruster performance and wall temperature behavior, but that these injector effects may be overshadowed by operating at a high fuel film cooling rate.
Copyright © 1993 by the American Institute of Aeronautics, Inc. No copyright is asserted in the United States under Title 17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for Government purposes. All other rights reserved by the copyright owner.
Introduction
Low
thrust
propulsion
systems
satellite,
and
spacecraft,
docking,
and
separation.
contamination engine
specifically
Current
Changes
small
rockets.
core
with
Most
high thrust
a thin
assumptions
boundary
and result
with
thrust effects
and
gradients
insertion,
space
vehicle,
attitude
control,
lifetime,
development requirements,
the complex models
2. For
reliability,
of low
thrust
improved
4500
and
chemical
predictive
and combustion
the rocket
relatively
large
of rocket
N mixing
tools,
rockets,
performance.
Therefore,
rockets
which
are
However, and
of
as an inviscid these
severe
must
processes
behaves
in the boundary
phenomena.
in small
fluid
assume
predictions
dominant
can exist
and
are essential.
layerl,
are more
launch
performance,
technology
rocket
below
every
as apogee
continued
address
in accurate
levels
combustion species
these
do not fully
duties
in spacecraft
for smaU rockets,
tools
virtually
such
are forcing
To address
designed
design
rockets
and perform
requirements
technology.
are on-board
for small
shear
layers
thermal,
be included
good
and
velocity,
in the
design
process.
Large
engine
Richter
codes
have
been
and Price 3 found
compare
well
with
used
that the predictions
experimental
engine
which
had a relatively
Army,
Navy,
NASA,
results.
thick
Air Force
used in conjunction
with experimental
various
other
codes
for modeling
thin boundary
layer
assumptions
comprehensive
is currently This
a lack
deficiency
gather
codes of local
local
This paper
describes
developed
to augment
chemical measurements evaluating
rockets
using
field
boundary
data
inadequate
of interior
opticall0,11,12 Lewis
wall fields.
temperatures
due
ratio
that the Joint
(TDK)
rockets
are being
code 5
program6, TDK 6 and
and
found
the flows.
the
Although
developed7,8,9,
development
the lack
nozzle
there
and verification.
of a sustained
program
to
rockets.
test results
Center.
works 2
field
for modeling
analyses
and
Previous
did not
prediction
Kinetics
flow
and heat
Research
rockets area
performance
to use for code
and preliminary
results.
and Coats 2 compared
layer
is largely
a high
mixed
and also found
Dimensional
flow field data for small
present
engine
Kehtarnavaz
with
hydrogen/oxygen
flow field
rocket
Navier-Stokes
data
the design
flow
layer
used were
flow
at NASA
rocket
thick
rockets
et al. 4 tested
code, Two
results.
of experimental
and analyze
Smith
boundary
with the reference
small
for small
(JANNAF)
did not agree
more
for modeling
of a small
laboratory
flux 13 diagnostics The
thruster
static
pressures
have
also
used
for low thrust provides
for
thruster
the
simple
simple
purpose diagnostic
of
techniquesto evaluaterocket flow fields. SchoenmanandBlock 14hot fire testedseveral enginesinstrumentedwith axial and azimuthal thermocoupleswhich were brazed or welded into place.Back et al.15usedtranslatingpitot tube and thermocoupleprobesto studythe boundarylayer andheattransfer in a conicalnozzle running with pressurized, heatedair. Richter and Price3 tested a low thrust gaseoushydrogen/gaseousoxygen, regeneratively cooled thruster and measuredchamber exterior wall temperatures. Kacynskiet al.16conducteda heatflux analysisof a high arearatio nozzle andmeasured nozzleexterior(backside)wall temperatureswith spotweldedthermocouples.Rousarand Ewen17had a fairly comprehensivetemperaturemeasurementssystemwith axial and azimuthaltemperaturemeasurementsfor a thin walled chamberwhich wascooled only by the internalfilm cooling. Although simple diagnostic techniques were employed in the past, most engine diagnosticswereweldedor brazedinto place,andazimuthaltemperaturemeasurements wereoften limited to a few axial locations.The water-cooledgaseoushydrogen/oxygen laboratorythrusterdiscussedin this paperemployedinstrumentationdesignedfor easeof maintenanceand repair. The instrumentation and thruster joint sealing techniques requiredto provide maximumflexibility for instrumentationreplacementaredescribedin this paper.The thruster provided axial pressureand axial and azimuthal temperature distributionsin the combustor,throat, andnozzle sections.Testing was conductedwith two injector designswhich employed fuel film cooling (FFC) to protect the chamber wall. Comparisonsof the global performancedata to local temperatureand pressure measurements weremade.The azimuthalperformanceof theinjectors andeffectsdueto differencesin injector designand fuel film cooling rateswerealsoinvestigated.
Apparatus
and Procedure
Chamber
The
thruster
Conductivity nominal
liner
and
(OFHC)
chamber
outer
copper.
pressure
The
chamber,
of 33:1. The nozzle
was conical
of an optimized
were
thruster
fabricated
was
designed
(Pc) of 500 kPa. The overall
a 2.54 cm diameter
instead
housing
a 1.27 cm diameter with a nominal
bell nozzle
contour
throat,
from to deliver
thruster
was
nozzle
fabrication.
High at a
20.3 cm with
expansion
A conical
the thruster
Free
110 N thrust
length
and a nozzle
19 ° half angle.
to simplify
Oxygen
area ratio was used
The laboratorythrusterusedaninstrumentationproceduredevelopedto accommodatethe water cooling design. Instrumentationdetails are shown in Figure 1. The thruster was watercooled usingmilled channelson the back sideof the liner asshown in Figure l(a). The channelswere0.032cm high by 0.032cm wide.The wall thicknessfrom the bottom of the channelsto the hot gassidesurfacewasalso0.032cm. Therewere 11 channelsin the combustion chamber and throat section which bifurcated into 22 channelsin the nozzle.A typical bifurcatedchannelis illustratedin Figure l(a). The channelbifurcation ensureduniform cooling of the large nozzle surface area. A clam shell type outer housing slid
over
shown
the
liner
in Figure
clam
shell
solder. degraded
conductivity
water
manifold
of OFHC
fire
the lower
the outer
soldered cycle
Welding
solder
joints
were
no
and
signs
would
not feasible
of the
have
due to the
by bolts
solder
of water
to
temperature
to braze
reinforced
The
and nozzle
a high
required
limit
manifolds
was split.
throat,
using
were
outlet
housing
was also
temperature
there
inlet
the chamber,
the thermal
The
testing,
where
were
of the copper.
copper.
water
it allowed
joints
because
While
to a hot
at the
seams
because
thruster
integrity
flanges.
prior
completion
The
was not desirable
high
to the liner
the two axial
piece.
the structural
concern
joined
style was chosen
in one
Brazing
was
2, and along
housing
be fabricated
and
joints
leakage
at the were
after
a the
of the tests.
Se_soFs
The
laboratory
and nozzle
thruster
sections.
in Figure
had 30 thermocouples
The
2. Typical
axial and azimuthal
thermocouple
respectively.
As seen in Figure
lands,
were located
which
tightly were
to the interior not
bonded
instrumentation diameter.
The
assessed
the
port
to ensure
outer
The
tolerances.
locations
and the ports
The
cooling
jacket.
channels.
housing,
size
Therefore,
that temperature
spanning
the combustor
of the thermocouples
are identified
are shown
l(a)
ports While
the exterior
in the channel liner
wall mated
(channel
lands)
a water
seal
was
required
in the
thermocouples
were
ports
drilled
were
l(b),
wall
thermocouple
4
were located
and
liner
were
of the instrumentation
anomalies
in Figures
the exterior
Therefore,
thermocouple
small
in 4 rows
the instrumentation
chromel-alumel
mm diameter
the hot gas side surface. verify
l(b),
of the outer
to the
0.85
rows
between
wall
ports.
located
nominally to within
ports
made
measurements
not due to thermocouple
0.80
mm
0.76
mm of
it difficult
were
carefully
placement.
in
to
Figure 1(c) is anexplodedview of a typical thermocoupleinstallation.The graphitefoil washersealedwateron thebacksideof the liner.Graphitefoil materialwasusedbecause of its high temperaturelimit. The fluorinatedpolymer washerwasusedto preventwater from leakingthroughthe backsideof the outerhousing.The fluorinatedpolymer sealfor the liner fitting sealedthe water which flows betweenthe liner fitting and the outer housingfitting. The sealcapandfluorinatedpolymer sealweredrilled through to allow insertionof a liner fitting. The fittings wereinstalledafterthe thrusterouter housingand liner werejoined. The liner fitting and graphitefoil washerwere installedfirst. A pipe sealing compoundmixed with a fluorinated polymer and graphite was applied to the threadsto provide addedsealingcapability. The sealingcap and outer housing fitting with the fluorinated polymer washerslid over the liner fitting. The thermocouplewas theninsertedthroughthe fittings andsecuredin placeby the thermocouplespring. The staticpressureportsusedthe sameinstrumentationportsandinstrumentationfittings as the thermocouples.The pressureports were drilled through the liner wall andwere 0.85 mm in diameter.The liner fitting shown in Figure 1(c) was usedas the pressure sensingline andconnectedto tubing which terminatedat the pressuretransducers.The pressure transducers were located outside of the vacuum tank. The pressure measurements were monitoredcarefully to ensuresteadystatepressureconditions were achievedduringtesting. The instrumentation fabrication and assembly techniques greatly simplified the replacementof faulty sensors. The replacementof a damagedthermocouple only requiredthe unfasteningof the spring andinsertionof a newthermocouple.In caseof a water leak, repair of the sealingsurfacessimply required unscrewingthe fittings and insertingnew washers.All of the instrumentationwaseasily replaced,andrepairswere madein situ without removal of the engine from the test stand.
Injectors Two
platelet
verify
operation
Freedom were
gaseous
designed
designated
modified
designed
by Aerojet
of the laboratory
low-thrust,
originally
injectors, was
injectors
in-house
GenCorp
thruster.
The
hydrogen/gaseous
to be tested
SN. 02 and to improve
with
Propulsion
injectors oxygen
were rocket
a regeneratively
SN. 03, were ignition
nearly
reliability 5
Division
identical
part
18 were
of a Space
technology
cooled
by enlarging
18 The
Injector
the gap
to
Station
program
thruster.
designs.
used
and two
SN. 02
at the base
which preventedthe sparkplug from arcingat the base.However, in a study by Arrington and Reed19using the modified SN. 02 injector, performanceanomalieswere observed. The unmodified injector SN. 03 was testedso that any performance,wall temperature, andpressurechangesdueto injector designcould be identified. The propellant flow pathsfor both injectors areshown in Figure 3. The oxygen entered through a platelet stackandwas injectedradially towardthe sparkplug, upstreamof the spark plug tip. The hydrogen entered a fuel manifold within the injector/combustor cavity. Part of the hydrogen entered the platelet stack and was injected radially just downstreamof the sparkplug tip. The remaininghydrogenflowed down milled channels on the fuel film cooling sleeve.A fuel splitting washer,locatedagainstthe flange of the fuel f'tlm cooling sleeve,determinedthe percentageof hydrogenwhich flowed down the sleeve.At the exit of the fuel film cooling sleeve,the chambercontained an oxidizer rich core
of combustion
were
designed
chamber
for fuel
pressure
also tested compare The
gases
film
(Pc)
core
of 500 kPa,
calculated
ratios
were
as the overall faction.
20 with
60%
FFC,
and
reported
in this paper
total
ratio
(MR)
between
divided
platelet
flow
rate,
of 8.0. The
injectors a nominal
injectors
the fuel splitting
were
washer
test condition,
and
by the quantity
of one minus
the fuel
mixture
of 16 and
ratios
32 with
to the total fuel and oxidizer
to
variation.
for every
flow contained MR
fuel
FFC by changing
stoichiometric
ratio
film. Both
of the
due to fuel film cooling
the core
refer
60%
of 75%
above
mixture
Therefore,
using
cooling
and a mixture
rates
and local changes
mixture
by a hydrogen
cooling
at fuel film cooling global
cooling
surrounded
75%
between FFC.
cab be film
MR of 10 and
All mixture
ratios
flows.
Test Facility The
testing
schematic
was
performed
of the RL-11
gaseous
hydrogen/gaseous
altitude
capsule
ejector
system
the exhaust hot gases based
data
was pumped
gases were
is shown
a 0.9 m
acquisition
NASA
the tank down
system
and
received
rockets
cooled
and displayed
tank.
levels A two
of 1.4 kPa.
diffuser.
RL-11
test cell was
1.8 m long
prior to entering
Center
at thrust
to a soft vacuum
into a water cart
Research
4. The RL-11
chemical
diameter
by a spray
Lewis
in Figure
oxygen
were fired
cooled
in the
Downstream the ejectors.
measured
test
capable
During
N. The
air driven
a hot fire test,
of the diffuser, A personal
and calculated
A
of testing
up to 220 stage,
cell.
the
computer parameters
on display screens.Data wererecordedon strip charts,FM tape, andfloppy disks. An extensivediscussionof thefacility is found in Reference20.
Results
Chamber The
and Sensor
laboratory
cooling After good
Operations
thruster
rates
diagnostics
through
their
the completion water
Minimal
water
effects
were
used
to assess
upon
wall
temperature
of 396 hot fire
sealing
capability
leakage
tests,
and good
Thermocouple
replacements
temperature
measurements
showed
No leaks from
of the high condition results
water
cooling
would
be forced
proved
differences
that in
inaccuracies
and/or
measurements,
and the azimuthal
upon
conditions
anomalies
wall
of the azimuthal
and
could
detect
film
cooling
temperature fuel
film
temperature
rate.
profile
cooling profiles,
was
were
Another
wall
symmetry
varied
widely
below,
were
from
was
that
measurement
in the
results
test
the test
concern
no biases
discussed
every
resulting
large
These
wall
Because
that
changes
of
of the
However,
introduce
rates.
Also,
detected.
a concern
temperature
revealed
in a matter
or distortion
distribution.
would
testing
joints
there
temperature
placement
extensive
housing
to have
and repairs.
made
heating
f'dm
profiles.
proved
maintenance were
fuel
pressure
in a few minutes.
of localized
kg/s,
and
instrumentation
accomplished
of 0.693
fuel
thermocouple
However,
test
rate
injectors static
leak repairs
liner and outer
to the same
uncertainties.
the
were
the thermocouples
injector in
flow
and
for in situ
and water
no evidence
the thruster
different
the thruster
accessibility
was experienced,
a few hours.
thruster.
and Discussion
temperature depending
show
that
the
due to real flow
field effects.
Azimuthal
Symmetry
Azimuthal
symmetry
is a key
Therefore,
the degree
of azimuthal
an accurate
rocket
operating illustrated The
model.
conditions in Figures
only known
assumption
Azimuthal
in the modeling
symmetry
in the flows
temperature
for the two injectors. 1 and 2, and
difference
between
used
Both the same
profiles injectors
must were
was
engine
flow
be determined
fields.
to ensure
measured
as a function
tested
with the chamber
were
fuel sleeve
the two injectors 7
of rocket
and
fuel
the small
splitting modification
of
washers. made
on SN. 02 described
earlier,
symmetry.
The azimuthal
Appendix
A, showed
injectors.
For
temperature displayed
impacts symmetry further
test
field
field models
injector,
the
trends
locations
temperature one
measurements
finding
and
single
only
data
from
which
both
azimuthal both injectors
The
observed
may
must
have
large
of azimuthal be investigated
and the development
in most
were
condition
similar
were
of
thermocouple
very
reproducible results
B are used
for all other
different
In addition,
In order
row
for either
for
A by comparing
locations.
Results
cases
thermocouples.
in Appendix
measurements.
for
in
engines.
for specific
thermocouple
condition
variation
processes
operating
is illustrated
are shown
different
The
distribution
and
which
designs
were not symmetric
for a given
of the wall temperature
from
to simplify
the
to display
the
thermocouple
rows
similar.
Thruster
Measurements
performance
6 show
The
discussed
For 60%
Ispv data
were
comparison velocity
with the local
(C*)
and
ratio for the two thruster nearly
identical,
vacuum
assemblies
therefore
only
data.
Figures
specific
5
impulse
for FFC of 60% and the
Ispv results
will
be
below.
FFC,
thruster
The
assembly
in Figure
linearly
SN. 02 thruster not linear.
to allow
in characteristic
with mixture
C'and
as shown
decreased
was measured
the variations
(Ispv), respectively,
ratios
rows
rockets.
chemical
temperature
the azimuthal
profiles.
in injector
of manufacturing
rows),
This
of the results,
Performance
75%.
axial
upon
75% FFC tests, where
of small
profiles
test
exhibited
temperature
for low thrust
(thermocouple
thermocouple
behavior
and
in the
ratio,
an important
temperature
test to another.
presentation
were
was
effects
with
injectors
differences
characteristics
conditions
the azimuthal
typical
small
rocket
flow
the two
varied
symmetric that
to have
for both injectors,
symmetry
at the high mixture
both the definition
azimuthal
the
condition,
implied
Although
from
that
to support
accurate
the
test
flow
with
distributions
azimuthally
profiles on local
temperature
except
similar,
temperature
was not anticipated
clearly
a given
patterns
which
with
6, but performance increasing
performance performance
SN. 03 performed
curve deviated
mixture decreased
better
differences ratio
than
diminished
for the SN.
monotonically
from that of injector
SN. 02 at lower above
03 thruster
with increasing
mixture
MR=5.0. assembly.
Ispv The
MR, but was
SN. 03 for MR below
6.0.
For FFC=75% the Ispv dataexhibited similar trends for both thruster assemblies.The FFC=75%performancedatadid not decreasemonotonically.BetweenMR of 5.0 and6.0 the Ispvvaluesincreasedlocally nearMR of 6.0.The SN.03 thrusterexhibitedlower Ispv valuesthanthe SN.02 thrusterfor all MRs tested. The performancechangeswith increasingMR for both the SN. 02 and SN. 03 thruster assemblieswith 60% and75% FFC fell into three categories. The first was the linear decreasein Ispvof the SN. 03 thrusterwith 60%FFC. This behaviorwas similar to the trendsobservedin a previousstudy by Reedet al.21 The secondcategorywas the slight deviation from the linear profile exhibited by the SN. 02 thruster assemblywith 60% FFC, andthe third was thehighly non-monotonicbehaviorof the two thrusterassemblies with 75% FFC. This substantial variation in thruster performance behavior with changingMR ledto a detailedexaminationof wall temperaturedistribution behavior.
Temperature The
axial
Measurements temperature
Figures
7 through
located
at 0.0 cm.
thruster
assembly
similar the
axial
for mixture
10. The thruster Figure with
60%
distribution
combustor
converging
distributions
and section
the
FFC.
All the The
sections.
(location
B3)
with MR.
In the nozzle
section
The
temperature
distributions
axial
(Figure
8) exhibited
However, wall
than
8 exhibited
disparity was
for the lower
temperatures
Figure
similar
in wall
sustained
generally FFC results
axial
much
temperature
temperature
(Figure
with decreased
cm and the throat
distributions profiles
for the
in Figure
in Figure
11, the
shown SN.03
obtained
thruster variation
SN.
profiles
sensitivity pronounced
in contrast
11) with increasing monotonically.
of Figure
The
was
in the combustor to the data MR,
a
only in in the
decreased
60%
observed and
FFC
above
resulting
temperature
in Figure
in contrast
with
7 for MR's
changed,
6.0.
to MR than
generally
assembly
distribution
with MR above
SN.03
disappeared.
02 thruster to those
was
temperature
assembly
in
7 followed
As
for the
section
shown
to MR was evident
of the
was
8 are
sensitivity
the temperature
greater
4 and
was at -3.81
temperature
MR tests the temperature those
between
temperature
temperature
in the nozzle
increased
exit plane
7 shows
pattern.
throat
injector
ratios
6.0.
in lower profiles
in Figure
in
7. The
throat
sections
and
7. The
temperature
to the SN. 03, FFC 60%
Figures9 and10 showthe axial temperatureprofiles from the 75% FFC testsfor the SN. 03 and SN. 02 thruster assemblies,respectively.The two thruster assembliesexhibited nearly identical temperatureprofiles. The shapeof the profiles were similar to thoseof Figure7. However,for the75% FFCtestswall temperaturesdecreaseddrasticallyfor MR below 6.0. The sensitivityto MR appearsto havebeensustainedthroughoutthe nozzlein the form of two distinct groupingsof temperatureprofiles. For 75% FFC, both thruster assembliesexhibitedsimilar temperaturevariationswith MR (Figures11).
Static
Pressure
The
static
and
13.
Measurements
pressure
distribution
All the static
boundary
layer
dominated
by mixing
film cooling
Low
Thus
far,
tested
discussed
nominally
500 kPa.
pressures
of nominally
75%
FFC and
tests 370
changes
tests,
only
The
at the
same
Temperatures FFC
behavior
kPa.
These
low chamber
to produce
trends
12
no indication
of shocks
that
temperature
trends
the
particular
to injector
gradients
restricted
with the laboratory
design
or
were and fuel
in the flow.
to comparisons
thruster
SN. 02 underwent
Typical
near
in Figure
a mixture
a slight
conditions near
temperature
increase
which
tests
support
behavior
indicated the
was
global
conditions. 10
The
chamber
pressures
showed
trends
large
for the 60%
in Figure
performance FFC
observed
between
seen
in Figure
15 showed
changes tests.
occurred. For the
MR of 7.0 and 14 near
that the diagnostic
measurements
performance
at low
data
diagnostics
a MR of 8.0 at Pc of 255
shown
MR of 6.0 for 75%
to the non-monotonic pressure
near
profiles at which
FFC.
of
chamber
the thruster
curves
designs
pressure
at off-design
performance
behavior
temperature
chamber
data at the lower
of 6.0 at 75%
non-monotonic
significantly
testing of testing
14. The
ratio
of injector
at the design
additional
Ispv performance
shown
operating
a significant
corresponded
been
corresponding
increased
which
interactions
in Figures
and gave
255 and 370 kPa for the purpose
FFC are
showed
indicates
from pressure
have
Injector
Pc conditions.
non-monotonic
layer
similar,
are shown
Results
two FFC rates
and
shear
for both injectors
were
result
and did not result
the results
at the lower
profiles This
and/or
Pressure
under
for 60%
pressure
separation.
rates,
Chamber
measurements
chamber
60% 8.0
MR of 8.0. continued pressure
Comparison
of Global
The
performance
FFC
exhibited
temperature same
a nearly
linear
trend
in Figure
with increasing
a small
a large
decrease
increase
and a dramatic
assembly
MR.
7. All the profiles
with
chamber,
the throat,
from the SN. 03 thruster
decreasing
distribution,
of the combustion near
6 obtained
are shown
temperature
temperature
Data
data of Figure
profiles
axial
length
and Local
7 displayed
the
temperature
along
section,
a maximum
in the converging
decrease
60%
The corresponding
of Figure
in wall
with
in temperature
along
the
the length
of
the nozzle.
The
performance
measurements
FFC displayed lower
than
monotonic,
of Figure
but non-linear
that of the SN. 03 thruster.
observed
with the SN. 02 thruster
decrease
in performance
temperature similar
to that
represent length
The
of the
for thruster
6.0. The Figures
assembly
linear
temperature 9 and
MR above
trend
10. The
at 75%
in wall temperature
well
with
pronounced
in wall
to a
assembly.
The
6.0 in Figure the
8 was
profiles
temperature in wall
which
along
the
temperature
shapes
were
constant
deviations
assemblies
are shown
to those
of Figure
was a significant
the
section
at MR conditions
below
10 at MR below
from
the
trends
in
7 for
decrease
combustor
9 and
deviations
dramatic
with 60% FFC for MR below
similar
along
in Figures
performance
exhibited
6.0 there
temperatures
seen
FFC
in
did
not
6.0. The
6.0 correlated
of
both
thruster
SN. 03 and SN. 02 at 75% FFC.
A close
examination
behavior
were
distribution.
corresponded
7. However,
for the two thruster
7, but remained
decrease
assemblies
FFC profile
the
large
the
was
6.0 which
assembly
for tests with MR below
in Figure
in wall temperature
to the decrease
with 75%
of the SN. 03 thruster
In addition,
as seen
This contrasted
assemblies
temperature
6.0. However,
was
MR above
an increase
60%
SN. 03.
profiles
temperature.
decrease
chamber.
for
in Figure
MR of 6.0 showed
for both thruster
from the nearly
wall
below
drop
with
the performance
of the SN. 03 thruster
assembly
shown
assembly
In general
8, a steep
to the performance
SN. 03 thruster
of the combustion
Ispv data
In Figure
of the SN. 02 thruster
conditions
observed
characteristics.
at 60% FFC for MR below
relative
behavior
6 for the SN. 02 thruster
of Figures
strongly
correlated
Significant
reduction
of 6.0 for all tests conditions
except
5 through with changes in wall
10 showed
that
changes
in the behavior
temperatures
of the wall
consistently
for the SN. 03 thruster 11
in performance
assembly
temperature
occurred with 60%
below
MR
FFC.
The
performance curves and temperatureprofiles indicated that the SN. 02 and SN. 03 thrusterassembliesoperateddifferently for the sametest conditionswith 60%FFC. The significant differencesin injector performancesuggestthat small differencesin injector configuration can result in significantly different performancecharacteristics.However, these differences can be overshadowedby operating at extreme levels of FFC, as is evidencedby the similarity of the thrusterbehaviorfor casesusing75%FFC.
Summary A laboratorymodelgaseoushydrogen/oxygenthrusterwasdevelopedasa diagnostictool for investigating low thrust rocket flow fields. The thruster fabrication and diagnostic techniqueswere proven effective via hot fire testing. Thruster performanceand wall temperature and static pressuredistributions were measuredfor two injectors as a function of both mixture ratio and fuel film cooling rate. The two injectors were identical in design with the single known exception of a small modification in the upstreamregionof oneof theinjectors. Testoperatingconditionsincluded60%and75% fuel film cooling rates,nominalchamberpressuresof 255kPa,370kPa,and500kPa,and oxidizer to fuel ratiosbetween4.0 and8.0. The thruster and diagnosticsfabrication and assemblyproceduresproved to allow easy and fast repair of all instrumentationas required to ensure accuratetemperatureand pressure measurements. Only minor water leaks and transducer failures were encounteredwhich wererapidly repaired. The instrumentedthrustersustainedhundreds of hot fire testswith only minor maintenance. The thruster diagnostics flow
fields,
changing global small
test
provided
rocket
degree
for both
data that
injector
was highly
and changes
of azimuthal
injectors.
and
to be a practical
data under every
behavior
design,
shown
local
conditions
performance
injector The
and
were
models
azimuthal
performance
upon
highly The
test condition
dependent
symmetry
and requires
designs.
upon
in fuel film cooling
This finding
flow field
was
was found
may further
operating
prove
representative
local
of the
data correlated
and configuration, operating
conditions,
small
rocket
effects
of
with
the
well
and indicated small
that
changes
in
rates. to be dependent
critical
investigation condition. 12
tool for investigating
upon
to the development to quantify
operating
conditions
of accurate
the dependence
rocket
of injector
The global and local data showed that the two injectors had distinctly different performancecharacteristicswith 60%FFC. Becausethe only known difference between the injectors was a small modification to the upstreamregion, the results indicate that smalldifferenceswith injector designsmay leadto largedifferencesin performance. The effects of varying the fuel film cooling rateswere evaluatedby comparingthe data trendswith 60% and75% FFC for both injectors.The two injectors performedsimilarly with 75% FFC, and exhibited the same trends with nearly identical temperature measurements.However, with 60% FFC the data trendswere different than what was observedwith 75% FFC. This indicatesthatthe rocket flow field may be dominatedby the effectsof the fuel film coolingrate at 75%FFC. The thrusterdiagnosticswerecapableof providing local datawhich wererepresentative of rocket flow field processes.The observeddependencyof azimuthal behavior upon operatingcondition,the performancedifferencesdueto smallchangesin injector designs, andthe significanteffectsdueto changesin fuel film cooling may proveimportantto the designand modelingof smallrockets.Although the thrusterdiagnosticsdid not provide quantitativeflow field parameterdata,they identify areaswhich greatlyaffectedthe small rocket operation.Furtherinvestigationof theseareasmay result in the developmentof moreaccuratesmallrocketmodelsandyield improvedlow thrustrocketdesigns.
Acknowledgments The
authors
wish to acknowledge
the
thruster
water
Edward
cooling
J. Pluta,
Stephen
system
support
during
Dorrance,
Waiter
E. Hendricks,
intimately
involved
enhanced
and
H. Culler,
electronic
which
the efforts
the
support William
hot fire
tests.
Patrick
with the thruster
of Roger
D. Scheman
for the
testing
M Furfaro The
Spanos,
fabrication,
authors and and
and wish
Walter
the efforts
for the fabrication of the Pablo
thruster,
of
and
A. Gutierrez
to recognize
Robert
A. Wozniak
who
of Richard
of for S.
were
Czentorycki
this work.
References 1. Smith, T.A., "Boundary Layer NASA Lewis 1030:1 Area Ratio
Development as a Function of Chamber Rocket Nozzle," AIAA Paper 88-3301, 13
Pressure in the July, 1988.
2. Kehtarnavaz, Flow,"
AIAA
H. and Coats, Paper 88-3160,
D. E., "Thick July, 1988.
Boundary
Layer
Assessment
for Nozzle
3. Richter, G.P. and Price, H.G., "Proven, Long Life Hydrogen/Oxygen Thrust Chambers for Space Station Propulsion," 1986 JANNAF Propulsion Meeting, August 1986, NASA TM-88822. 4. Smith, T. A., Pavli, A. J., and Kacynski, K.J., "Comparison of Theoretical and Experimental Thrust Performance of a 1030:1 Area Ratio Rocket Nozzle at a Chamber Pressure of 2413 kN/m 2 (350 psia)", NASA TP-2725, 1987. 5. "JANNAF Rocket Engine Performance Publication 246, April 1975.
Prediction
and Calculation
Manual",
CPIA
6. Nickerson, G.R., Coats, D.E., and Dang, L.D., Engineering and Programming Manual: "Two Dimensional Kinetic Reference Computer Program (TDK)" NASA CR-178628, Software Engineering Associates, Inc., 1985. 7. Kim, S.C. and Van Overbeke, T. J., "Calculations Thrusters," AIAA Paper 90-2490, July, 1990.
of Gaseous
Hydrogen/Oxygen
8. Reed, B.D., Penko, P.F., Comparisons of Flowfields
Schneider, S.J., and Kim, S.C., "Experimental and Analytical in a 110 N (25 lbf) H2/O2 Rocket," AIAA Paper 91-2283.
9. Weiss, J. M.and Merkle, Low Thrust Rocket Engine
C. L., "Numerical Investigation of Reacting Flowfields Combustors," AIAA Paper 91-2080, June 1991.
in
10. Seasholtz, R.G., Zupanc, F.J., and Schneider, S.J., "Spectrally Resolved Rayleigh Scattering Diagnostic for Hydrogen-Oxygen Rocket Plume Studies," Journal of Propulsion and Power, 8, 935-942, 1992. 11. Zupanc, F.J. and Weiss, J.M., "Rocket Plume Flowfield Rayleigh Scattering," AIAA Paper 92-3351, July 1992.
Characterization
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Using
Laser
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From
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Heat
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15. Back, L.H., Cuffel, R.F., and Massier, P.F., "Laminarization of a Turbulent Boundary Layer in Nozzle Flow - Boundary Layer and Heat Transfer Measurements With Wall Cooling," Journal of Heat Transfer, 333-344, August 1970. 16. Kacynski, K.J., Pavli, A.J., and Smith, Transfer on a 1030:1 Area Ratio Rocket,"
T.A.,"Experimental NASA TP-2726,
Evaluation 1987.
of Heat
17. Rousar, D.C. and Ewen, R.L., "Combustion Effects on Film Cooling,"NASA 135052, Aerojet Liquid Rocket Company, February, 1977. 14
CR-
18.Robinson,P.J."SpaceStationAuxiliary ThrustChamberTechnology,"Final Report, NASA CR-185296,AerojetTechsystemsCompany,July 1990. 19.Arrington, L.A. andReed,B.D.,"PerformanceComparisonof Three-Dimensional Rocket," AIAA
Hydrogen Film Coolant Paper 92-3390, July 1992.
20. Arrington, L.A., and Schneider, 90-2503, July, 1990.
Injection
S.J., "Low Thrust
Axisymmetric in a 110 N Hydrogen/Oxygen
Rocket
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21. Reed, B.D., Penko, P.F., Schneider, S.J., and Kim, S.C.,"Experimental Comparision of Flowfields in a 110N(251bf) H2/O2 Rocket," AIAA Paper 1991.
15
and
AIAA
Paper
and Analytical 91-2283, June
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hated
polymerI I !1 seal for liner . ...%;j fitting J 0,
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housing fitting,
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