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8/17/2019 AIAA Experiments on Heat Transfer in a Cryogenic Engine
1/5
J O U R N A L O F P R O P U L S I O N A N D P O W E R
Vol. 9, No. 2,
March-April
1993
Experiments
on
Heat Transfer
in a
Cryogenic Engine
Thrust Chamber
N . Sugathan,* K . Srinivasan,t a n d S . Srinivasa M u r th y $
Indian
Institute of Technology, Madras 600036, India
Tests are
conducted
on a cryogenic engine using liquid oxygen as
oxidizer
an d
gaseous hydrogen
as
fuel
with
water as a coolant. Th e coolant flow passage of the thrust
chamber
is of m illed channel
configuration.
M easured
heat
transfer results compare well with
those predicted by a
therma l analysis
using the
standard Bartz
correlation
and the
Hess
and Kunz
correlation
for hot gas
side
a nd
coolant
side heat transfer
coefficients, respectively. This
confirms the conclusions of a recent
theoretical
study by the
authors
in which a comparison of various heat
transfer
correlations was made.
Nomenclature
C = characteristic velocity, m/s
c
p
= specific heat, kJ/kg K
D = diameter of cross section, m
D* =
throat diameter, m
h =
convective heat transfer coefficient, W/m
2
K
k
=
thermal conductivity,
W/m K
M
= Mach number
of flow
m = mass
flow
rate, kg/s
P =
pressure,
bar
Pr
=
Prandtl number
q
=
heat
f lux,
k W /m
2
Re = Reynolds number
r
c
=
radius
of
curvature
of
throat section,
m
T = temperature, K
t
=
wall
thickness measured
from
exposed side,
m
x = axial distance from throat, m
L
=
absolute viscosity,
kg m/s
r = specific heat ratio
Subscripts
a = adiabatic
b =
bulk
c = coolant
ch =
chamber
g = hot
combustion
gas
in j = injection
t = total
w = wall
0 =
stagnation
value
1-4
=
axial
locations as in
Fig.
1.
Superscripts
r
= temperature at depth of 1.5 mm on hot gas side
= temperature at depth of 0.5 mm on hot gas side
Introduction
I
N
the thermal design of cooling systems for the thrust
chambers of rocket engines, reliable estimation of the
cool-
R eceived May
31,1991;
revision received Sept. 25,1991; accepted
for publication Oct. 15, 1992. Copyright © 1992 by the American
Institute
of
Aeronautics
and
Astronautics, Inc.
All rights
reserved.
* R efrigeration and Air Conditioning Laboratory, Department of
Mechanical Engineering; currently Engineer SF, Liquid Propulsion
Systems
Center,
Department of
Space,
Trivandrum 695547,
India.
tRefrigeration and Air
Conditioning
L aboratory, Department of
Mechanical Engineering; currently Associate Professor, Instrumen-
tation and
S ervices U nit, Indian
In stitute of Science, Bangalore
560012,
India.
^Professor
and
Head, Refrigeration
and Air Conditioning
Labo-
ratory,
Department of Mechanical
Engineering.
an t side and the hot gas side heat transfer coefficients is of
great
importance. In an earlier
paper,
1
a critical evaluation
of nine combinations
of
coolant side
and hot gas
side heat
transfer correlations w as carried out by the authors. B y com-
paring
it with limited published experimental data on engine
cooling, it was found that the Hess and Kunz
correlation
2
'
3
for the coolant side
heat
transfer coefficient and the standard
Bartz
correlation
4
for hot gas side heat transfer coefficient,
are
suitable
f or
thermal design
of
regeneratively cooled cryo-
genic
engines. However, it was felt that a detailed experi-
mental verification was essential. In this
article,
such a study
is presented with gaseous hydrogen
as
fuel , liquid oxygen
as
oxidizer,
and
water
as the
coolant.
The
present study also
confirms
the earlier contention that the above-mentioned pair
of
heat transfer coefficients
is the
best suitable
fo r
carrying
out
the thermal design.
Experimenta l
Ground Test Setup
T he
configuration
of a
scaled model
fo r
static testing
of the
cryogenic engine is shown in Fig. 1, and the details are given
in Table 1. T he walls of the thrust chamber of this engine are
made
of a
milled channel copper substrate enclosed
in a
stain-
less steel outer shell. The two materials are bonded together
by vacuum brazing. Coolant entry an d exit manifolds of the
engine
are tig-welded. A
16-element
coaxial injector is used
fo r
admitting t he fuel, gaseous hydrogen (GH2) and the ox-
idizer, liquid oxygen (LOX). A rigimesh porous face plate is
used
to
provide transpiration cooling.
G H
2
© Flow
@
Pressure
Water
©
Temperature
Axial locations
for temperature
measurements
INJECTOR
N O T E : - A l l d i m e n s i o n s a re i n m m
Fig. 1
Engine configuration.
8/17/2019 AIAA Experiments on Heat Transfer in a Cryogenic Engine
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SUGATHAN, SRINIVASAN, AND MURTHY: HEAT TRANSFER IN A THRUST
CHAMBER
241
Ta ble 1 Conf igurat ion of the
subscale
engine
Table 2
Details
of instrumentation
Sea level
chamber
area ratio
Characteristic length
Nozzle
contraction
area
ratio
N ozzle ex tension
Injector
Igniter
No. of coolant
channels
8.46
84 cm
3.09
H2 dump-cooled from area
ratio
8.46-140
Coaxial multielement type
with
18
elements
Centrally
mounted electrical
type
with p r e bu r n e r
.40
Coolant
channel geometry
Chamb er wal l
thickness,
mm
Channel width, mm
Channel height , mm
Cylindrical
portion
1.5
4 .0
1.5
Throat
1.0
1.4
2.0
Nozzle
exit
2.0
7 .4
1.0
HE-LN2cooled heat e x c h a n g e r (only f o r c r y o G H
2
t e s t )
C V - C h e c k valve
R V - R e l i e f v a l v e
V V - V e n t valve
Q C -Quick c o n n e c t o r
SV- Start valve
PR- P r e s s u r e regulator
F -Filter
OR-Or ifice/Flow meter
L
-
Level
indicator
P G - P r e s s u r e g a u g e
PT-
Pressure
transducer
BV- Ball v a l v e
F i g . 2 Schematic of test setup
A schematic
diagram
of the engine test setup is shown in
Fig. 2 . The experimental
facility
includes
test
stand structure,
propellant feed system, auxiliary fluid circuits, and instru-
mentat ion. The test
stand
structure supports the engine in the
horizontal position and the
thrust
is
transmitted
to it through
suitable load cells. T he propellant feed system consists of GH2
a n d L O X subsystems. Each subsystem is composed of facil-
ities fo r pressurization,
filling,
draining, cool d o wn ,
purging,
and monitoring of various
parameters.
Other
main
compo-
nents
of the test facility are the
high-pressure
ru n
tank
fo r
storing GH2 and the superinsulated LOX tank. LOX is trans-
ferred to the engine by an oxygen gas pressurization system.
The liquid feed
lines
ar e
insulated with polyurethane
foam.
Helium gas is
used
for the pilot
pressure source
fo r
regulator
an d also
command
pressure
to cryogenic valves.
Locations of temperature, pressure, an d
flow
transducers
on
the engine hardware are also
shown
in
Fig. 1. Pertinent
details of instrumentation are given in
Table
2. Platinum
resistance sensors ar e used fo r fuel an d oxidizer temperature
measurement . Thermocouples are used for coolant and cham-
ber wall temperature measurement. I n four axial positions of
the chamber
identified
in
Fig.
1,
three
thermocouples each
Pressure transducers
Type
A ccuracy
class
Temperature transducers
Thermocouple types
Gauge
Resistance temperature devices (RTD)
Type
Sensitivity
Response time
A ccuracy
Flowmeters
Type
Accuracy
Digital
panel meter (DPM)
Type
Display
Accuracy
Thin
film/strain gauge
±0.5%
Copper-Constantan and
Chromel-Alumel
24 swg
Platinum
resistance
0.00395 n/°C
100/300 ms
0.75% or 1%
Turbine
0.15%
D u e l
slope
integrating
type
3.5
digit
±0.1%
Table 3 Test condit ions
Param eter
Value
Chamb er
pressure
LOX injection pressure and tempera-
ture
G H 2 injection pressure an d
tempera-
ture
Injection velocity ratio (fuel/oxidizer)
L O X flow rate
GH 2
flow rate
Coolant
w at er flow rate
Coolant temperature
N ominal firing
time
10
± 0.2 bar
10 ± 0.2 bar and 90 K
14.5
± 0.2 bar and 300 K
41
0.58 ± 0.01 kg/s
0.097
±
0.0002 kg/s
6 kg/s
310 K
50
s
(test
1)
20 0 s (test
2)
Hot exposed
Waf
0.5mm
Inner
milled
chamber (Cu)
Fi g .
3 Tem perature probe locations
are installed as shown in
Fig.
3. High
precision
in locating
the thermocouples is ensured by the use of a computerized
numerically controlled
riiilling
machine to drill the holes. T he
thermocouples are spot-welded to the
flat
bottom of the holes.
Experimenta l Procedure
The engine is tested in a pressure-fed
mode
for a 50- and
200-s duration. The test conditions are given in
Table
3. T he
propellant
flow
rates, and
thereby
th e combustion chamber
pressure,
are maintained
constant during both tests
by setting
the injector
upstream
pressures at predetermined values. Also,
the water
flow
rate is maintained at 6 kg/s during tests. This
is
achieved by calibrating the
entire coolant
line
with
th e
engine hardw are prior
to the
test.
Tests
are conducted according to a
predetermined test
se-
quence. Countdown starts 30 min before th e start of engine
ignition,
an d
ends
10 min
after
th e
engine firing. Steady con-
ditions are observed
about
10 s
from
th e
start. Data obtained
8/17/2019 AIAA Experiments on Heat Transfer in a Cryogenic Engine
3/5
242
SUGATHAN, SRINIVASAN,
AND MURTHY:
HEAT
TRANSFER IN A
THRUST CHAMBER
from th e experiments are used in the thermal analysis for
making a comparison
w ith
the predictions made based on the
authors'
earlier
paper.
1
Calculation Procedure
Heat transfer
calculations
using
the measured temperature
data are done
considering
one-dimensional, steady-state heat
conduction in the radial direction as already described by the
authors.
1
Considering th e wall thickness an d rib-effect in heat trans-
fer, th e coolant
side
wall
temperature
T
W tC
an d exposed wall
temperature T
W tg
ar e extrapolated
using
measured tempera-
tures
inside
th e wal l . Here, th e
temperature gradient
in the
ribs between coolant channels is taken as linear.
F or instance, near the inlet point of coolant, the measured
values of temp eratures are
T
W tl
=
440 K
T'
W tl
= 410 K
T
Ctl
=
310 K
an d
the extrapolated
temperatures
are
T
W ig
= 460 K
T
W tC
=
4 0 0 K
Making use of
one-dimensional, steady-state heat conduc-
tion
equation
in the
copper
shell in the direction o f local radius
of
curvature, the local
heat
flux is calculated as
Q
— *w\
• •
u f
• -
w
r)'*' /1
w g wc ( )
= (0.35
x 60)70.0015 14 x
10
3
k W / m
2
Also
which
yields
q
=
h
gtt
=
4.76 k W / m
2
K
(2)
where T
W t0
,
th e
adiabatic wall temperature
of gas
stream
at
th e
location,
is
calculated
using
=
T
g
(3)
wh er e
<
R
is the
boundary- layer recovery
factor which
varies
from 0.9 to
0.99
depending on the Prandtl
n u mber
of the
flowing gas. Equation
(3) is a modification of
standard equation
4
T
*•
w. a
(4)
This
modification
is
done
to account for the influence of static
ga s temperature T
g
,
which is
calculated from
R e f . 4
T
g
=
T
ch
/[l
- 1 ) 1 2 ]
(5)
In th e present
work
T
ch
, th e adiabatic combustion tempera-
ture,
which
is a function of
mixture ratio (O/F = 0.6)
an d
chamber pressure,
is taken from R e f . 8. T
w
^ calculated the
same way is
equal
to
3398
K.
Also,
by
considering heat balance
in
one-dimensional
steady
state
q
=
whereby
wf
T
c
)
= 155
k W / m
2
K
(6 )
(7)
Results and
Discussion
Figures 4-7 present the pressure and temperature data ob-
tained during
the
50-s engine
test, and Figs.
8-11 show similar
data obtained from the 200-s engine test. The combustion
chamber
pressure during
both
th e
tests
was maintained at 10
bar. This w as achieved because pressure drops of both pro-
pellants
in injectors
could
be
predicted
accurately, and the
coefficient
of discharge fo r both propellants were k n o wn pre-
cisely.
The temperatures recorded
from
the inner copper
shell
at different radial depths from
th e
surface exposed
to hot
gas
show
similar
trend. Temperatures are transient up to 10 s
from
th e
start
of
engine
firing before steady state is
attained.
O t h e r
properties
of the combustion product remain con stant;
th e heat transfer coefficient in the hot gas
side
is a function
of chamber pressure only.
15
P
f ini
n . .
0, 1 nj
-5 0 10 20
30
Firing time (s)
50
60
Fig. 4
Pressure
variation as a function of t ime in test 1 .
500
-i <
Tw,2
T
w,1
Tw,4 „
T E S T - 1
300
20 30
40
Firing
time (s)
50 60
Fig. 5 W a l l
temperature variation
as a function of t im e in test 1 a t
0.5-mm depth.
45 0
TEST-1
300
20 30 40
Firing
time ( s)
50
60
Fig. 6
W a l l temperature variation
as a function of t ime in test 1 at
1.5-mm
depth.
8/17/2019 AIAA Experiments on Heat Transfer in a Cryogenic Engine
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SUGATHAN, SRINIVASAN, AND MURTHY: HEAT TRANSFER IN A THRUST CHAMBER
243
350
2 3 2 5 -
300,
T c , 4
T
c , 3
20 30 40
Firing time (s)
T E S T - 1
I
450
50 60
Fig. 7
Coolant
temperature
variat ion
as a function of t ime in
test
1 .
5400 -
310-
300.
0 10
T'w,3
100
Firing
time (s)
150
2 0 0 2 1 0
Fi g .
10
Wall temperature variation
as a function of t ime in test 2 at
1.5-mm
depth.
- 5 0 50 100
150 200
Firing time
( s)
Fi g .
8
Pressure
variat ion as a
function
of
t ime
in test 2.
500 -
5
**°
3
m
p
t
£0
u
350-
300
C
3
______
_̂_
Tw,i
\
n
\
'
TEST-2
I I I . I
0 50 100 150 200
21C
Firing time s)
Fig. 9
W a l l
temperature variat ion as a function of t ime in test 2 at
0.5-mni depth.
The experimentally
determined
hot gas side
heat
transfer
coefficients
ar e
compared with
th e
standard
Bartz
correlation
4
C
'
S
r
r
°
(8)
wh er e 0
is the correction
factor
for
property variation across
the boundary layer, given by
(9)
1 + M
2
(r -
- M
2
(r - l)/2]}
068
P is the chamber static pressure measured at an axial location
of
50 mm
from
the
coolant inlet
as
shown
in
Fig. 1. C
is
calculated at stagnation conditions as given in Re f. 4.
T he coolant side heat transfer
coefficients
derived from th e
experiments are compared with the
Hess
an d Kunz correla-
tion
2
-
3
h
c
=
(k
c
/D)[0.020SRe
0
c
8
Pr°
c
4
(l
+
0.01457^7 )̂] (10)
A s
seen
in Fig. 12 , h
c
derived
from
tests, closely
matches
with
the predictions from E q. (10). T he
S ieder-Tate
correlation
4
E320
300
T
c> 2
0 10
50
100
Firing time (s )
150
2 002 1C
F ig . 11 Coolant temperature variation as a function of t ime in test 2 .
400
300
200
100
E x perimental
a T e s t - 1
y Test -2
— Sieder-Tate
4
— - - NAL°
|
Thrust
chamber axial
distance
(mm)
F ig .
12 Comparison of
coolant
s ide heat transfer
coefficients.
Experimental
n Test-1
y Test-2
Predicted
————— Std.Bartz
4
—— • — Mod.Bartz
7
—
- -
NAL
6
Thrust chamber
axial distance
(mm)
Fi g . 13 Comparison of hot gas si e heat transfer coeff icients .
8/17/2019 AIAA Experiments on Heat Transfer in a Cryogenic Engine
5/5
24 4
SUGATHAN, SRINIVASAN, AND MURTHY: HEAT TRANSFER IN A
THRUST
CHAMBER
700
o 500
S.
400
300
Data derived
from
T e s t - 1 Test-2
§Thrust chamber axial distance
(mm)
Fig. 14 Comparison of
temperatures
at various locations
overpredicts th e heat transfer coefficient, whereas, t h e N a-
tional Aerospace Laboratory (NAL), Japan correlation
6
un-
derpredicts th e heat transfer coefficient.
Similarly,
th e
values
of fi
g
derived from test data match well with
the
predictions
of
Eq. (8) as
shown
in Fig. 13 . Bot h modified Bartz
7
an d
N A L
6
correlations underpredict the gas side heat transfer
coefficient. Coolant a nd wall temperature variations predicted
from
th e
one-dimensional thermal analysis
1
using
th e
pair
of
standard Bartz, and Hess, and
Ku n z
correlations are com-
pared with measured
values
in Fig. 14. In the case of tem-
peratures, the deviations are within
+10.7%
to — 1 . 8 % .
Conclusions
Test
results obtained on a cryogenic
engine
with water as
coolant match well
with
th e predictions
made
from a one-
dimensional thermal analysis using the standard Bartz equa-
tion for the ho t gas side heat
transfer
coefficient, and the Hess
an d
Kunz correlation for the coolant side heat transfer coef-
ficient.
It is
suggested that
this
pair
of
correlations
may be
used for the thermal design of thrust chambers.
References
^ugathan,
N ., Srinivasan, K., and Srinivasa Murthy, S.,
"Com-
parison
of Heat
Transfer
Correlations fo r Cryogenic
E ngine T hrust
Ch ambe r Design," Journal of Propulsion and Power, Vol. 7, No. 6,
1991,
pp. 962-967.
2
ARQUARDF Corp., "Thrust C h a m b e r Cooling Techniques for
Space
Craft Engines: Final Report,"
N A S A C R - 5 0 9 5 9 , July
1963.
3
Hess,
H. L., and Kunz, H. R., "A Study of Forced Convection
Heat
Transfer to Super-Critical
Hydrogen,"
Transactions
of the
American Society
of
Mechanical Engineers, Journal of Heat Transfer,
Vol. 87,
Feb.
1965, pp. 41-48.
4
Huzel,
D .
K .,
an d Hu an g , D .
H., "Design
of
Liquid
P ropellant
R o cke t Engines,"
NASA
SP-125, 1971.
5
Yanagawa, K., Fuji to, T., Katsuda, H., and Miyjima,
H.,
"De-
velopment
of LOX/LH2 Engine
LE-5," AIAA/SAE/ASME
20th
Joint
Propulsion Conf.,
AIAA
Paper
84-1223,
Cincinnati, OH,
J u n e
11-
13,
1984.
6
K u m akawa, A., Niino,
M .,
Yatsuyanagi, N., Gomi, H.,
Saka-
moto, H., an d
Sasaki, M.,
A Study of the
Cooling
of Low
Thrust
L O 2 / L H 2 Rocket Engine,"
Proceedings
o f the 13th International
Sym-
posium on
Space
Technology an d Science, Tokyo, 1983, pp . 301-
306.
7
Steer , T .
E., "The
D esign and
M anufacture
of a Liquid Hydrogen
Thrust
Chamber,"
Space Flight,
Vol.
1, Feb.
1970,
pp.
135-142.
8
Gordon, S., and McBride, B.
.„
"Theoretical
Performance of
Liquid ydrogen with Liquid
Oxy g e n as
Rocket Propellant," NASA
TM-5-21-59E,
March
1959.
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