8/17/2019 AIAA Experiments on Heat Transfer in a Cryogenic Engine

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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.

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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

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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.

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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 .

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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

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Propulsion Conf.,

 AIAA

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  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

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L O 2 / L H 2 Rocket Engine,"

 Proceedings

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posium  on

  Space

  Technology  an d  Science,  Tokyo, 1983,  pp .  301-

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Steer , T .

 E., "The

 D esign and

 M anufacture

  of a Liquid Hydrogen

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  Space  Flight,

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  1970,

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 135-142.

8

Gordon,  S., and McBride, B.

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  "Theoretical

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Liquid  ydrogen  with Liquid

  Oxy g e n as

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TM-5-21-59E,

  March

 1959.

R e c o m m e n d e d

  R e a d i n g

  f r o m

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