Crogenic Tank Pressurant Reqs

8/21/2019 Crogenic Tank Pressurant Reqs

1/104

T ECHN I C A L

NOTE

OF CRYOGENIC PROPELLANT

M .

E .

Nein

und

J

F.

Thompson

eorge

C.

Murshull

Spuce

Flight Center

Ala.

T I O N A L A E R O N A U TI C S A N D SPA CE A D M I N I S T R A T I O N W A S H I N G T O N , D .


2/104

N AS A T N D - 3 1 7 7

E X P E R I M E N T A L A N D A N A L Y T I C A L S T U D IE S OF C R Y O G E N I C

P R O P E L L A N T T A N K P R E SS U R A N T R E Q U I R E M E N T S

By M .

E.

N e i n and J . F. T h o m p s o n

G e o r g e C . M a r s h a l l Space F l ig h t C e n t e r

H u n t s v i l l e , A l a .

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

For sale

by

the Clearingh ouse for Feder al Scienti f i c and Techn ical Information

Springfield, Virginia

22151

- Price

$4.00


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TABLE OF CONTENTS

Page

SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

PRESSURIZATION REQUIREMENTS AND LAUNCH VEHICLE DESIGN

. . . . . . . .

EXPERIMENTALPROGRAM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Test Facilities. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Instrumentation

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Test

Results

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

PRESSURIZATION ANALYSES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Previous Work. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Summary of Analytical Program . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Modifications in the

Program

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Evaluation of P rog ram P ar am ete rs . . . . . . . . . . . . . . . . . . . . . . . . . . .

Comparison with Test Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Conclusions f rom Comparisons with Test Data . . . . . . . . . . . . . . . . . . .

THE EFFECTS

O

SYSTEM PARAMETERS ON PRESSURANT REQUIREMENT . .

CONCLUSIONS AND RECOMMENDATIONS . . . . . . . . . . . . . . . . . . . . . . . . . .

REFERENCES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

BIBLIOGRAPHY

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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4

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87

88

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I I I

U S T OF ILLUSTRATIONS

Figure Title Page

I.

Com par ison of the Weights of the Pro pe lla nt Feed Sys tem s of

Two

Flight Vehicles

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19

2.

Weight of LOX Tank Pr es su ra nt Ver sus Vehicle Th rus t .

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20

3.

Saturn I,

S-I

Stage; Saturn I, S-IV Stage.

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21

4.

Te st Fac ility for Tank Configuration C

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22

5a.

Int eri or of Tank Configuration D. .

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23

5b.

Te st F acility f or Tank Configuration D

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24

6 .

Location of Tem pera ture Prob es.

. . . . . . . . . . . . . . . . . . . . . . .

25

7.

Pr es su ra nt Distribu tor, Tank Configuration D.

. . . . . . . . . .

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26

8.

Temperature Sensors . . . .

.

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27

9.

Comparison of Tem pera ture Sensor Response Time

. .

. . .

.

.

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. . .

28

IO.

Copper Plate Calorimeter . . . . . . . . . . . . . . . . . . . . . . . . . . . .

29

11.

Te mp era tur e Respon se of Copper

Plate

Calorimeter .

.

. . . .

.

. . .

30

12.

Calculated

Free

Convection Heat Tr an sfe r Coefficients, hf, on

Vertical

W a l l .

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31

13.

Comparison Between Experimental and Computed Heat Transfer

Coefficients, Configuration C, Te sts

130-9

and

130-10,

Oxygen as Pressurant . .

. . .

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.

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.

.

.

.

32

14.

Comparison Between Experimental and Computed Heat Transfer

Coefficients, Configuration C, Te st 130-15, Helium

as

Pre ssu rant

. . .

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. . .

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.

33

15.

Comparison Between Ullage P re ss ur e Loss for H e and GN, Pre-

Pr es su ra nt s Under Liquid Slosh and Nonslosh con diti ons in Tank

Configuration C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

34

iv

. .

...... ...

. .. I


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LIST OF ILLUSTRATIONS (Cont'd )

Figure

Title

Page

16.

Measu red and Computed Ullage Gas Concentration Gradie nts

in Tank Configuration C.

. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

35

17.

Experimentally Determined Mass Transfer Mt /Am

(lb/lb) Versus Time

t (sec)

. . . . . . . . . . . . . . . . . . . . . . . . . .

36

18.

Liquid Surface Conditions During Pressurization Test

in

Tank

Configuration C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

37

19. Liquid Sur face and Ullage Conditions During SA-5 Flight. . . . . . . .

38

20. Experimentally Determined Radial Temp erature Gradients . . . . . . 39

21a. Com pari son Between Exp erim enta l and Computed Ullage

Tempe rature Gradient, Tank Configuration C y Test 130-6,

Oxygen as Pressuran t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40

21b.

Com parison Between Expe riment al and Computed Ullage

Tem pera ture Gradient, Tank Configuration C ,

Test

130-6,


21c.

Comparison Between Experimental and Computed Pres sur an t

Flowra te, Tank Configuration C, Test 130-6, Oxygen as

P r e s s u r a n t . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

42

21d.

Ullage

Pressure

and P ress uran t Inlet Temperatur e Histories ,

Tank Configuration C , Te st 130-6, Oxygen

as

Pressuran t . . . . . . .

43

22a. Comp arison Between Experim ental and Computed Ullage

Tem pera ture Grad ient, Tank Configuration C, Te st 130-7,


22b.

Comp arison Between Experimenta l and Computed Ullage T empera -

tu re Gradient, Tank Configuration C

Test

130-7, Oxygen as

P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45

22c.

Comparison Between .Experimental and Computed P res sur an t

Flow rate , Tank Configuration C , Test 130-7, Oxygen as


. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

46

V


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LIST OF ILLUSTRATIONS (Cont'd)

Figure

Title

Page

22d.

Ullage Pressure and Pressu rant Inlet T emperature Histories ,

Tank Configuration

C ,

Test

130-7, Oxygen

as

Pressu ran t .

. . . . . .

47

23a.

Com par iso n Between Expe rimen tal and Computed Ullage

Tem per atu re Gradient, Tank Configuration C, Test 130-9,

Oxygen

as

Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48

23b.

Comparison Between Experimental and Computed Pre ssu ran t

Flow rate, Tank Configuration C y Test 130-9, Oxygen as


. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

49

23c.


Flow rate, Tank Configuration C,

Test

130-9, Oxygen as

Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50

23d.

Ullage Pre ss ur e and Pressuran t Inlet Temperat ure Histories ,

Tank Configuration C y Te st 130-9, Oxygen as Pressurant . . . . . . .

51

24a.

Comparison Between Experimental and Computed Ullage

Tem per atur e Gradient, Tank Configuration C, Test 130- IO,

Oxygen

as

Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52

24b. Com paris on Between Expe rimen tal and Computed Ullage

Tem per atu re Gradient, Tank Configuration C y Test 130-10,

Oxygen as Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53

24c.

Comparison Between Experimental and Computed Pr es su ra nt

Flow rate, Tank Configuration C, Test 130-10, Oxygen

as Pressurant

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

54

24d.

Ullage Pressure and Pre ssu rant Inlet Histories, Tank

Configuration C y

Test

130-10, Oxygen as Pressu ran t .

. . . . . . . . .

55

25a. Com pariso n Between Expe rimen tal and Computed Ullage

Temp eratur e Gradient, Tank Configuration C,

Test

130-15,

Helium as pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56

25b.

Com pariso n Between Expe rimen tal and Computed Ullage

Temp eratur e Gradient, Tank Configuration C y Test 130-15,

Helium as Pressu ran t .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

57

v i


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LIST

OF

ILLUSTRATIONS (Cont'd)

Figure

Title Page

25c. Compariso n Between Experimenta l and Computed Pr es su ra nt

Flow rate, Tank Configuratio n C, Te st 130-15, Helium

as

P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58

25d. Ullage Pressure and Pre ssu ran t Inlet Temperature Histories,

Tank Configuratio n C y Te st 130-15, Helium as P r e s s w a n t .

. . . . . .

59

26a. Comparison

Be

tween experimental and Computed P re ss ur an t

Flow rate, Tank Configuration D, Te st C 003-7a, Oxygen

as

Pressuran t

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

60

26b. Com paris on Between Expe rimental and Computed Tank Wall

Tem per atu res , Tank Configuration

D ,

Te st C007-7aY Oxygen

as P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 61

26c.

Comparison Between Experimental and Computed Tank W a l l

Temperatures, Tank Configuration D, Te st C003-7aY Oxygen

as

P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62

26d. Ullage Pr es su re and Pr es su ra nt Inlet Temperature Hist ories ,

Tank Configuration

D,

Te st C003-7aY Oxygen

as

Pressuran t . . . . . .

63

27a. Compariso n Between Experimen tal and Computed Pr es su ra nt

Flow rate, Tank Configuration

D,

Test COO3-12, Oxygen

as


. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

64

27b.

Comparison Between Experimental and Computed Tank W a l l

Tem per atu res , Tank Configuration

D,

Test COO3-12, Oxygen

as

P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65

27c.

Comparison Between Experimental and Computed Tank Wall

Tem per atu res , Tank Configuration D, Test COO3-12, Oxygen

as

P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66

27d. Ullage Pr es su re and Pre ssu ran t Inlet Tem peratu re Histories

,

Tank Configuration D, Test COO3-12, Oxygen as Pressuran t . . . . . . 67

28a. Compariso n Between Experimenta l and Computed Pr es su ra nt

Flow rate, Tank Configuration D, Test COO3-IO, Helium as

P r e S S U r a I I t . . . . ................................... 68

I

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LIST OF ILLUSTRATIONS (Cont'd)

Figure

Title Page

28b.

Com pari son Between Experimen tal and Computed Tank Wall

Tem pera ture s, Tank Configuration D, Test COO3-10,

Helium

as


. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

69

28c.

Compa rison Between Experimental and Computed Tank Wall

Tem pera ture s, Tank Configuration D,

Test

COO3-10,

Helium as Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70

28d.

Ullage Pre ssu re and Pressu rant Inlet Temperature

His tori es, Tank Configuration D,

Test

COO3-10, Helium

as

Pressurant

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

29a.

Compa rison Between Experimental and Computed Pre ss ur an t

Flowrate,

S-I

Stage LOX Tanks, SA-6 Stati c

Test,

Oxygen as


. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

72

29b.

Ullage Pr es su re and Pre ssur ant Inlet Temperatur e Histories,

S-I Stage LOX Tanks, SA-6 Stati c Test, Oxygen as Pressurant . . . .

73

30.

Comparison Between Experimental and Computed Pr ess ur ant

Flo wr ate , S-I Stage LOX Tanks, SA-5 F ligh t

Test,

Oxygen

as


. . . . . . . . . . . .. . . . . . . . . . . .. . . . . . . . . . . ..

74

3ia.

Compa rison Between Experimental and Computed Pres su ran t

Flowrat e, S-IV Stage LOX Tank, Helium as P r e s s u r a n t .

. . . . . . . .

75

31b.

Ullage Pre ssu re and Press urant Inlet Temperature Histories

S-IV Stage

LOX

Tank, Helium

as

Pressurant . . . . . . . . . . . . . . . .

76

32a.


Flow ra te , S-IV Stage LH, Tank, Hydrogen as Pr,essurant

. . . . . . . .

77

32b.

Ullage

Pr es su re and Press uran t Tempe rature Histories, S-IV

Stage LH,

Tank,

Hydrogen as Pressurant .

. . . . . . . . . . . . . . . . . .

78

33.

Comp arison Between Experimenta l and Computed Ullage

Tem pera ture Histories, Tank Configuration D, Te st 187260,

Nitrogen

as

P r e s s u r a n t .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

79

viii


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LIST OF ILLUSTRATIONS (Concluded)

Figure

Title Page

34.

Comp arison Between Experim ental and Computed Ullage

Tem pera ture Gradient,Tank Configuration C,

Test

130-10,

Oxygen as Pressuran t .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

80

35.

Comp arison of P re ss ur an t Flowrate Pred ictions by Two

Computer Pr ogr am s with Experimental Results. . . . . . . . . . . . . .

81

36.

Comp arison of Ullage Mean Tem pe rat ure Predictio n by Two

Computer Pro gra ms with Experimental Results.

. . . . . . . . . . . . .

82

37.

Comparison Between Experimental and Computed Pres sur an t

Flowra te History, Tank Configuration C y Test 130-6, Oxygen

as


. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

83

38. Schematic of Heat and Ma ss Tr an sf er Conditions in a

Propellant Tank.

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

84

39. Com pari son Between Free

Jet

Velocity Decay and Forced Heat

Transfer Coefficient Decay . . . . . . . . . . . . . . . . . . . . . . . . . . . 85

40. The Effects of Vari ous Design Pa ra m et er s on the Mean

Temperature

at

Cutoff

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

86

ix


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LIST OF TABLES

Table

Title

Pa

ge

I.

Tank Configurations and Test Parameters

. . . . . . . . . . . . . . . . . . . . 16

II

.

Su mm ary of

Test

Conditions

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

17

III.

Parameters

for

Heat and Mass Tra ns fer Calculations. . . . . . . . . . . . . 18

X


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~ DEFINITION OF SYMBOLS

Symbol Definition

A

Tank total surface area

AD

Pres

s u r

ant

distr ibutor area

Constants used in calculation

: i

of g as to w a l l forced

b3

coefficients (Table m )

Constants used in calculation of fr ee

convection heat and ma ss tran sfe r

Coefficients (Table I I I

&

Computer

Program)

Tank wall thickne ss

Cons tants used in calculations of gas to

liquid f orc ed convection heat and m a s s

tran sfer coefficients

D

Tank diameter

(

f t )

-

D

Diffusion coefficien t

( f t2/hr )

EK

Modification fac tor fo r therm al con-

ductivity caus ed by mixing of fluid (Btu/h r f t

OR)

ED

Modification of diffusion coefficient

caused by mixing of gas (f t2/hr)

gC

Constant = 32.17

( lb

m

ft/lbf s e C 2 )

xi


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DEFINITION OF SYMBOLS (Cont‘d)

Symbol

Definition

h

Ullage gas-to-wall. heat tra ns fe r

gw

coefficient

h

Ullage gas-to-wall .

free

convection

C

heat tr an sfe r coefficient

h

Ullage gas-to-wall fo rce d convection

0

heat tr an sfe r coefficient at tank top

h Gas-to-liquid heat transfer coefficient

S

h

Gas-to-liquid f ree convection heat

s c

transfer coefficient

hL

Liquid- to-wall he

at

t ransfer

coefficient

K

Gas

thermal conductivity

L/D

Tank length to di ame ter ra tio

n i

Press uran t f lowrate

M a s s

t ransfer

Mt

Am

Pres suran t ma ss accumulated

P Ullage pressure

r

Tank radiu s

t

Time

T

Temperature

Th

Vehicle thrust

V

Gas velocity

V

Tank volume

(Btu /h r f t2

OR)

(Btu/hr

f t 2

OR)

( B t d h r f t 2 OR)

(Btu/hr f t 2

OR)

( B t d h r f t 2

OR)

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14/104

DEFINITION OF

SYMBOLS (Cont'd)

Symbol

Definition

vd

Volumetric pressurant flowrate at

distributor

V

Reference volumetric pressuran t

Od

flowrate

X

Radial distance from tank wall

Y

Gas-to-liquid m a ss tra ns fe r coefficient

S

Y

Gas-to-liquid fr ee convection ma ss

sc

t ransfer

Y

Gas-to-liquid forced correcti on mas s

so

transfer coefficient

at

tank top

Z

Axial dis tance fr om tank top

Z Axial dis tance of gas-liquid int er

i

face from tank top

Dimensional decay coefficient of

ullage forced heat tr ans fer coefficients

(Table

III)

PLP

Th er ma l expansion coefficient of

liquid

* T

Total time of pressurization

P

Gas viscosity

P

Gas density

@

Molefraction

(ft3/sec)

( f t /hr )

(ft-1)

(set)

( lb / f t

hr

m

(lb

/ f t3 )

m

(-)

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15/104

Symbol

Subscripts

a

C

f

L

m

0

U

W

DEFINITION OF SYMBOLS (Concluded)

Definition

Ambient

Ca lo r imete r

F re e convection

Ullage

gas

Interface

Liquid

Mean

Reference, pressurant inlet

Ullage

W a l l

ACKNOWLEDGEMENT

The contribution of the MSFC Te st La boratory in providing the t es t facilities and

complex instrumen tation and obtaining the experi mental data is gratefully acknowledged.

Invaluable contributions in prog ram definition and analysis of experim ental re su lt s wer e

made by J. Moses, T. Stokes, L. Worlund, and G. Platt of the Fluid Mechanics and

Thermodynamics Branch.

NOTE:

M r .

J.

F. Thompson is currently Assistant Pr ofe sso r, Mississippi State

University, Aeronautical Engineering Department. He was fo rm er ly associat ed with

Propulsion Division, MSFC.

x

iv


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EXPER

IMENTAL AND ANALYTICAL STUD

I ES OF

CRYOGEN I C

PROPELLANT TANK PRESSURANT REQU IREMENTS

SUMMARY

The extensive requ ire me nt for pressu riza tion of cryogenic propellant tanks of

launch and space vehic les ha s dir ect ed attention to the need fo r a ccu rat e methods of

analysis of propellant tank thermodynamics. This paper pre sen ts the resu lts of experi

mental and analytical studies of pres suri zatio n gas requir eme nts for cryogenic liquids.

Experimenta l res ul ts are analyzed for cylindrical and spheroidal tanks ranging

in

s i z e

over four ord ers

of

magnitude. A parame ter study

of

the controlla ble varia bles of a

pressurizat ion system design illustrates their effect on ullage gas te mperatu re.

Pressurizati on data

are

provided for use in the development

o r

chec kou t of analy

tical pressurization models and for design of pres suriz atio n sys tem s

for

future launch

and spac e vehicles. A tank pressurizati on computer program , using recomm ended coef

ficients, can be used to pred ict total and transient pressurant requirements and ullage

tem pera ture gradien ts within 10 percent accuracy.

INTRODUCT ION

Determination of the pre ssu ran t ga s weight for cryogenic propellant tanks

is

com

plex and defies exact analytical trea tme nt because of the interdependent tran sien t

phenomena of heat and mas s tr ans fer that occur simultaneously in

a

propellan t tank.

Mathematical models describ ing the inte rnal thermodynamics of tank pre ssu riz ati on have

been developed by va riou s inv estigator s.

The an alys is by Clark [ i] re pr es en ts an analytical solution of the governing equa

tions that predict the tran sien t tem pera ture , the response of the pres sur ant gas, and

container wall. However, the solution req uir es assumptions, such

as

constant tank pre s

sure and zer o initial ullage, that are not always met with real syst ems . The studies by

.Coxe and Tatum [ 21 are base d on ana lysi s of a system in which the ullage is thermally

mixed a'nd heat

transfer

between the gas and the wall is independent of t ime and sp ace .

Gluck and Kline [

31

used dimensional analysis

to

express gas requirements

as

a

function

of known syst em par am ete rs; they determined, experimentally, quantities of in terfacial

m a s s transfer and gas phase he at transfer.

Epstein

[

41 presented

a

numeric al method for calculation of pre ssu ran t gas

re

quirem ents that contains a numb er of phenomena absen t from previous analytical methods.

However, em piric al data

are

required

to

evaluate many constan ts

and

physical parameters .


17/104

---

To provide

a

reliable method fo r determination of pre ssu ran t gas requirements,

the experimenta l data on pre ssur izat ion obtained by the Marsh all Space Flight Center

during the Sat urn launch vehicle development wer e applied

to

the method of Epstein. Th

physical par am et er s and the previously indeterminate constants were developed. After

modification, this numerical method

is

capable of ac cu ra te prediction of pre ssur izat ion

gas

requirem ents fo r cryogenic propellant tanks.

PRESSURIZATION REQU IREMENTS AND LAUNCH VEHICLE DES IGN

The increa sing s iz e and complexity of sp ace launch vehicle s neces sita tes optimi

zation studies of the many subsys tem s involved in launch vehicle design. The propellant

tank pressurization system is of particular importance because

its

weight

is

large in

compa rison to the weight of other subsy stems. Weight optimization stud ies of propellant

tank pres surizatio n syste*msfor the Saturn V S-IC

stage,

were used to establish the lo

cation of the oxid izer and fuel tanks within the over-al l vehicle configuration

(Fig.

1

and

.Ref. 5).

Even the pressurizatio n system components such

a s

heat exchangers, pres

su ran t lines, and control s, weigh considerably l es s than the pressurizing

gas.

A fu rth er indication of the need for optimization of pre ssu ran t require ments is

illustrated in Figure

2.

The pressurant-mass/tank-pressure ra ti os of typical launch

.vehicles is given

as a

function of vehicle thru st, th ru st being repr esen tati ve of vehicle

size. Although th er e

is a

gr ea t deal of differenc e between the propellant tank configura

tions of tactical mi ssi les and space launch vehicles, a near l inear increase occurs in

pressurant-masdtank-pressure as

vehicle siz e incr ease s. Considering only pres sura nt

gas weight, it app ears advantageous to use helium

as

a pressurant . If, however, the

weight of the pr es su ra nt storage containers

is

included in the weight of the pressurization

system, the use of helium as

a

pressuran t in most instances resu l ts in

a

weight penalty.

F o r vehicl es with high acce leration and low turbo-pump NPSH requirements, it

is

possible to eliminate the pressurization sy ste m, relying only on the

self

pressur iza

tion

of

the saturated propellant (flash boiling)

.

Fl as h boiling pressurizatio n, however,

re su lt s in high pre ss ur an t weight and can only be justified

if it

significantly simplifies ve

hicle design. Because of the infant knowledge of cryoge nic tank press uri zat ion at the

initiation of the Saturn launch vehicle development pr og ra m,

a

long

series

of pressur iza

tion experiments w a s conducted

at

MSFC to obtain sys te m design informa tion and scaling

laws for the la rg e propell ant tanks of the Saturn

I

vehicle. Res ults of this experim ental

program and correlations with analytical studies are pre sent ed in the following sec tions

of this report.

2

. _


18/104

EXPER

IMENTAL PROGRAM

T e s t F a c i l i t i e s

The experi mental w ork was conducted on five tank configurations

at

the Marshall

Space Flight Center:

A .

Saturn

I,

S-I Stage,

Multiple Interconnec ted LOX

Tanks

(Fig. 3a)

B.

Saturn

I ,

S-IV Stage (Fig. 3b) LOX and LH, Tanks

C. A 6. 5 by 39-ft (DxL) cylin dr ica l LOX tank (Fig . 4)

D.

A 13 by 26-ft (DxL) cylindr ical LOX tank (Fig .

5a, Fig. 5b)

E.

A 1 by 3-ft (Dx L) cyl indr ical LOX tank .

The

test

pa ram ete rs for these tank configurations are compared in Table I.

Configurations A and B we re flight vehicles and thus contained the stan dard test instru

mentation of the Satur n propell ant feed sys tem , including continuous liquid level se ns or s,

tank pr es su re , pres sura nt flowrates, and supply tem perature measurements. Configura

tions C, D, and E we re equipped with many thermocouples along the tank axis. Thermo

couples, mounted at severa l rad i i at thre e elevations in these tanks, allowed measur e

men t of

radial

tempera ture gradients.

Wall tem pera ture s were measur ed in Configura

tions C and D by t hermoc ouples on the ins ide and outside

surfaces

of the tank at severa l

locations. The locations of the temper atur e se ns or s in these tanks are shown in Figu res

6a and 6b. Special cal ori me ter pla tes we re mounted in both tanks fo r determination of

gas-to-wall heat tra nsf er coefficients.

Finally, gas sam pling devices were placed

at

seve ral locations to m easu re ullage gas concentration gradients.

Configurations C,

D ,

and E wer e equipped with heat exchangers that provide a

variable pressura nt inlet temperature up to

1000

OR.

The pre ssu ran t gas was introduced

at the top of the conta iner through a distr ibutor (ei ther

a

deflecto r plate-Configuration

C

and E , o r a scr een arrangement-Configuration

D)

to minimize inlet velocities and

disturba nces of the liquid surf ace by impinging gas jets.

Figure 7 shows a typical distri

butor configuration.

Pres sura nt velocit ies

at

the distributo r periphery are given in

Table I fo r the five te st configurations.

The tank Configurations C,

D,

and

E

could

be

sloshed

at

rotational

o r

translatory

oscillation in exce ss of the

first

critical frequen cy of the tank.

Configurations

A ,

C, and

D

we re equipped with cam er as

so

that the conditions inside th e tank could be obs erved.

The resu l ts

of tests

conducted with the five tank configurations

are

presented in Figure s

13 through 38. The conditions of these

tests are

summ arized in Table II.

3


19/104

I

nstru mentation

Analysis of ullage gas temperature h istory re qui red a temperature probe with

fast response charac teristic s and good accuracy.

A

fast

response temperature probe

(Fig.

sa)

consisting of a fork-like suppo rt with a 30 gage CuCo welded thermojunction

was designed

at

MSFC. The length-to-diameter ra ti o of the thermocouple

w i r e

and its

distance from the fork base w ere determined using an analog computer representation

of the heat tran sfe r conditions aroun d the probe assembly . Fig ure 9a shows the res pon s

tim e; 63.2 percen t of the total tem pe rat ur e change w a s attained in eight secon ds when

the probe w as extrac ted from liquid oxygen into

a

gas circulating at a velocity of about

. three feet per second [ 61. The respon se of the prob es during a pressurization test w a s

.al so determined (Fig. 9b) ; the fork-type thermocouple has a good response chara cterist

A thermocouple mounted on a long, rod-like suppo rt (Fig. 8 b ), which was de

signed for liquid meas ure men t in the high vibrat ion environ ment of stat ic and flight

testing, exhibited an extrem ely poor response in the gas phase as indicated in Figure

9a. Respo nse time to 63.2 per cen t of total temp era tu re change was in ex ce ss of

10

minutes.

Commercial temperature probes of the resistan ce thermometer type (Fig. 8c

we re also investigated under these conditions. Although th ei r res pon se was considerably

better than the flight type thermocouple (63 .2 perc ent t empe rature change in approxi

mately 50 seconds) , it was too slow for the press urizati on studies.

Pr es su re measu rements in the ullage space , p res sura nt supply lines, and liquid

discharge lines wer e made with close-coupled pre ss ure transd ucers to a ssu re good

response character is t ics .

The pressu rant flowrate and liquid discharge flowrate me as

ure men ts wer e obtained with turbine type flowmeters.

Liquid level before and durin g

the tests was mea sur ed by capacitance discr ete level probes and continuous delta P

me as ur em en t of the liquid column.

T e s t R e s u t s

~Heat Tr an sfe r Coefficients. Heat tra nsfe r between pre ss ura nt and tank side walls

was m easure d during pressurization tests in Configuration C by two plate calorimeters.

Each calorimete r was

a

12 by i2-inchY 30-gage copp er plate mounted from teflon sp ac er

parallel to and

at

a distan ce of four inches from the tank wall (Fig . I O .

Three thermo

couples, spaced to repr esen t equal calorim eter

areas

and connected

as

a thermo-pile,

provided

a

tempe rature /time history of the copper plate before and during the tests. The

local ullage ga s tem per atu re was measure d in the vicinity of .the calo rim eter (Fig.

11).

The calorimeters wer e located

I1

and 30 feet fr om the top of the test tank.

Fo r determination of he at trans fer coefficients,

it

was ass umed that heat transfer

to the back side of the p late (towa rd tank wall) was by free convection because of the

shielding effect of the plate-to-wall ar ran geme nt . The free convection coefficients for a

one component ga s we re evaluated by the equation of Jackso n and Eck er t [ 71 ; the resul ts

are plotted in Figure 12.

The free convection heat tran sfer coefficient was al so

4


20/104

calcula ted for two component mixtures base d on the tim e and spac e dependent helium-

oxygen concentration in the

tank

The total heat t ransf er

to

the calor im eters was +en

cor rec ted using the calculated free convection effect on the back side. The hea t transfer

coefficients.

to

the front of the cal orim eter pl ates m easu red in Test s 130-9, -10, -15 are

presented

in

Figu res 13 and

14

using gaseous oxygen and helium

as

pressurants. Ullage

gas-to-wall heat

transfer

was also evaluated from wall temper ature measure ments at

-a

location 3.5 fe et from the top

to

the

tank.

Wall

measurements

a t

locations initially below

the liquid surface produced erroneous readings and we re discarded. These coefficients

were corre cted by subtracting the effect of exte rnal heat flux from the meas ured wall

temperature rise. During a flash-boiling

test,

which did not req uir e pre ssu ran t flow,

the wall temperatu re rise indicated an external heat f lux of 13 Bf dm in ft'; this compares

ver y favorably with a calculated f l u x of

15

Btu/min f t2

[

81 and conf irms the method used

for correcti ng wi ll measurements .

Insp ecti on of Figures 13 and 14 shows very good agreement between measu red

and calculated heat tra nsf er coefficients. It is noted tha t the gas-to-wall heat tran sfe r

coefficient is definitely within the forc ed convection regim e fo r 'the oxygen tests, but in

the fre e convection reg ime f or the helium test. AltEiough the heat transfer coefficient by

force d convection diminishes with increasi ng distance from the pre ssu ran t distri buto r

,

the

free

convection contribution (Eq.

I

compensates for this decayto such

a

degree that

a nearly constant heat transfer coefficient is obtained along the tank bulkhead and side

wall.

Sloshing Effects. Pre ssu riz ati on studi es conducted

at

MSFC have shown tha t

there is little benefit derived fr om the use of helium as a main pressuran t for cryogenic

propellants. However ,

t

was determ ined experimentally that prepr essur izat ion with

helium reduce s pr es su re decay during liquid sloshing near the cri tic al frequency. It is

assumed that the helium acts as a buffer zone between the splashin g cryogenic liquid and

th.e condensable press urant , suppressing excessive mas s tran sfer .

Figure

15 shows a typical

tank

pre ssur e h is tory for a stati onar y liquid oxygen

tes t tank

as

compared to

a

pres sure h i s to ry in which the liquid slosh es nea r the first

cr it ic al mode of oscill atio n [ 91. The

tank

was prep ressu rized , with eithe r helium o r

nitrogen, followed by mai n pres sur iza tio n .during liquid expulsion with super-heated

oxygen. The.tank pre ssu re his tory during the s losh tes t (using helium

as

a prepres

surant)

is

nearly identical to the pr es su re hi story of the nonsloshing expulsion test.

In

contrast , prepressurization with gaseous nitrogen resul ted in a marked pr essu re decay

duri ng the sloshing of the liquid, which was not evident during a nonsloshing expulsion

test with gaseous nitrogen prepressurization.

Ullage Gas Concentration Gradients.

Gas flow conditions and the conc entration

of

helium gas in a cryogenic propellant tank during pressurization discharge were s tudied

in test Configurations C and D. Spectrographic analyses

were

made of ga s samp les taken

at

various positions i n the tanks. Samples taken at various elevations in tank Configura

tion C ju st befor e the end of the

tests

yielded the res ul ts shown in Figure 16. In the test

5


21/104

in which helium was use d for prep ressu riza tion and oxygen

as

the main pressurant, the

helium concentration is maximum at 12 fee t above the liquid, and gradually dec rea ses

in both directions.

The concentratio n of oxygen nea r the liquid sur fac e is probably caus ed by accumu

lat ion of

t$e

gaseous oxygen that is initially

in

the ullage before prepressuriz ation. Fo r

comparison, Figure 16 also shows the con cen trat ion of helium above the liquid oxygen

for the case

in

which helium prep ressu riza tion

is

followed by press uri za tio n with helium

during liquid expulsion. The oxygen concen tration at

10 feet

above the liquid interface

was only six perc ent by volume. The total amount of gaseous oxygen in the ullage

was

only slig htly lar ge r than the amount of oxygen in the ullage before prepressuriza tion

(0.77 moles ver sus 0.73 mole s). This indicates that interfacial ma ss tran sfe r, although

sm al l under th ese conditions, was i n the form of evaporation.

Mass Transfer. A comparison of ma ss tra nsfer re su lt s obtained in Configuration

C with result s obtained by Clar k [ I ] is shown in Figure 17. Condensation

in

excess of

30 perce nt of the pr es su ra nt flow was found by Cl ar k during liquid nitrogen expulsion

tests

with

a

I by 3-foot cyli ndrical tank. Simi lar resu lt s were obtained with the MSFC

test

Configuration

E ,

also

shown in Figu re 17. The ma ss

transfer

measured in

test

Configuration C indicates that condensation was 5 to 10 percent. Condensation in the

large r facil i ty

is

less because of the sma lle r wal l-a red vol ume ratio of a larger tank.

Comparing the condensation in the sma ll tank with tha t in the large tank on

t;he

basis of wall -ared volum e rat io, the values a re approximately equal. During tests at

high pressuran t inlet temp erat ure, init ial evaporation noted i n Configuration C diminished

as the

test

proceeded.

However, Cl ark had found inc rea sed condensation at higher pres

surant s e t emperatures in smal l tanks. These conflicting resu lt s point out the incom

plete knowledge of ma ss t ran sfe r.

Condition of Liquid Interface. The condition of the liquid inte rfa ce i n Configura

tion C and d uri ng the launch and flight of

SA-5,

are

shown

in several f rames f rom

a

movie taken inside these tanks (Figs.

18 and 19) . Violent boiling oc cu rr ed during venting

of the

tank

before prepressurization. As the vents wer e closed and prepressurization

proceeded, the liquid surfa ce became nearl y quiescent before discharge.

After

discharge

began,disturbance

of

the liquid surfa ce caused by pres sur ant flow and acceleration of the

liquid

surface

were observed; the disturbance diminished as time and distance between

the surface and the pressuran t inlet increased.

Radial Ullage Temperaiture Gradients.

Radial tempe ratur e gradients obtained

with Configurations C and D ar e shown in Figure

20.

In both cases the radial gradients

were small , and ther e apparently exists little differenc e between the gas flow conditions

in the

two

tanks, even though the gas d istrib utors , baffling, and tank diam eters a r e not

comparable.

6


22/104

The temperature probes at X/ D

-

0.025 in Configuration D, which are located be

tween the a ntislosh baffles

Fig . 5a ,

recorded virtually the same tem perature as probes

at

sma ller radii . It wa s concluded that the gas cir culation

in

the tank is not appreciably

affected by the antislosh baffles, and subdivision of the tank into volume el em ent s perpen

dic ular to the tank axis

is

permissa ble for the pressurization analysis.

Axial ullage Tempe ratu re Gradients. The axial ullage temperature gradients

obtained-in tests 130-6 and 130-7 with Configuration

C

(Fig. 21a, 21b, 22a, and 22b)

became approximately linear

as

the

test

proceeded. Th ese two tests were conducted

with oxygen

as

pressurant

at

about 550"R.

There was a rapid increase in temperat ure

of about 30"R immediately above the liquid interface.

in

these tests, indicating that

ma ss t ransfer was small . In tests 130-9 and

130-10

(F ig s. 23a, 23b, 24a, and 24b)

with the s am e Configuration with oxygen pr es su ra nt

at

a lower temperature, the ullage

temperature gradients are much flatter; the rapid in crea se

in

tempera ture immediately

above the liquid interf ace is

still

in evidence. The ullage tempe ratu re gradient s in this

sa me configuration with helium as pr es su ra nt (T es t 130-15; Fig. 25a, and 25b)

are

con

cave, rat her than linear

as

in the

tests

with oxygen

as

pres sura nt, and the increase in

tem pera ture ju st above the liquid interface is very gradual. The concave shape is to

be

expected in this cas e because the mas s tr ansf er

is

in the form of evapora tion with

an

ullage that

is

predominately helium. The line ar ullage temperature gradients in tests

with oxygen

as

pres sura nt indicate that the mass transfer

is

very sm all with an ullage

that

is

predom inantly oxygen.

Other Test Results. Tests are being perfo rme d with Configuration a, but so far

only

threktests

have been completed. The pre ssur ant distr ibutor in this configuration

was designed to minimize the ga s circulation in the tank, reducin g forced convection

heat tran sfer . While this is the desire d condition for optimum pre ssur izat ion system

operation,

it

is detrime ntal to the response time of the tempera ture pro bes

as

the liquid

interface passes. Pr ec is e ullage temperature gra dients will not be available until this

instrumentation is improved. However, prel imin ary data , with very hot

GOX

used as

pres sura nt, indicate that the te mperatu re gradients are concave rather than linear as

was the c ase in the tests with Configuration C using colder

GOX as

pressurant .

The con

cave temp erat me gradients found in the helium p ress uran t

tests

with Configuration

C

we re al so in evidence with Configuration D.

Pres sura nt flowrates and wall temp erature

gradients from these tests

are

pres ente d in Fi gu res 26a, 26b, 26c, 27a, 27b, 27c,

28a, 28b and 2%.

Pre ssur ant flowrates in the LOX tanks of the Saturn I , S-I stage, during

static

test and flight

are

pres ente d in Fig ure s 29a and 30. Fig ures 31a and 32a show pr es

su ran t flowrates in the

LOX

and

LH,

tanks

of

the Saturn

I , s - IV

stage, during

static test.

Finally, ullage tempera ture hist orie s obtained in

a

very s ma ll tank, Configuration

E ,

containing LN2 pres suri zed with nitrogen

are

given in Fig ure 33.

7


23/104

P R E S S U R I Z A T I O N A N A LY S ES

P rev ious W o

rk

Pres suri zed discharge from cryogenic liquid containers was studied ana-jtically

and experimentally by Clark [I]nder sponsorship of the Army Ballistic Missi le Agency

and later MSFC. The analytic al solution s obtained by Cl ar k w e r e applied to test data

obtained fo r Configuration C. In Figu re

34

the

axial

tempera ture gradient through the

ullage gas

is

shown

as

a fuilction of d ista nce fro m the tank top

or

gasdistributor. Excel

lent agreement with test res ult s

w a s

obtained for an assumed gas-to-wall for ced convec

tion heat

transfer

coefficient of 10 B t d h r f t2"R. Agreement for

a

coefficient of 2 Btu/hr

ft2" R, approximately in the ran ge of

free

convection, w a s poor. This illus trate s one

limitatio n of analytical solutions in which the gas-to-wall heat tr ans fer coefficient en te rs

as

an independent variable.

In spit e of this res tri cti on and the assumption of

initial

ze ro ullage volume, the

method by Clark w a s successfu lly applied in design analyses of the Saturn I pres su riz a

tion system. While Cla rk's analysis as sumed st ratification of the ullage gas and con

sta nt heat tran sfe r coefficient, the analysis by Coxe and Tatum

[ 21

w a s based on the as

sumptio n of a complete thermall y mixed ullage gas and constant heat tran sfe r coefficient.

Figures 35 and 36 com pare tes t r es ul ts obtained with MSFC Configuration C with the

ana lyti cal pre dic tio ns by the method of Coxe and Tatum. Toward the end of the test,

agreement

is

good possibly because the conditions of co nstant heat tr an sfe r coefficients

Gre approached in the larg e ullage n ear the end of the run.

A compa rison of the pre ssu ran t flow requ irem ents with predictions by an analog

computer simulation developed by MSFC, is shown in Figure 37. Rep rese ntat ion of the

pres suri zati on thermodynamics by analog method w a s difficult because of sc alin g problem s

and the ext rem e sensitivit y of the equations to tank pr es su re fluctuation. In Fig ure 35,

36, and 37 pres sura nt flow requirem ents ar e also compared with a digital computer pro

gram developed by Rocketdyne

[ 41

and modified by MSFC [

IO].

This program closely

matches test data. However, the pr og ra m insufficiently describes mass t ran sfer and is

sen sit ive to fluctuations of ullage pre ssu re. The se fluctuations do not app ear in the

ineasured flowrates because they are apparently counteracted by the effects of evaporation

and condensation [ I 1] .

S u m m a r y o f

Analytical

P r o g r a m

Since the Rocketdyne program makes maximum use of the techniques of digital

computer calculations and is not subject to the restr ictiv e assumptions that are made in

othe r programs, this method was chosen by MSFC fo r pressuriza tion syst em analyses.

However, extensive comparisons of the pro gra m with

test

data were required to evaluate'

the physical pa ram ete rs and constants initially contained in the prog ram

as

indeterminate

identities. The equations we re modified when nec essa ry.

8


24/104

This program includes in its calculations

a

pressu rant gas s torage tank, heat

exchan ger, and flow co ntr ol valve. It considers a propellant

tank

with o r witho.ut outs ide

insulation and pressurized with either evaporated propellant or with

a

gas s tor ed under

pres sure in a storage tank in which the gas expands nonadiabatically. The ullage pre s

s u r e

is controlled

by a pressurant

flow

control valve that has finite maximum and mini

mum

ar ea s an d may be ei the r the on-off o r the continuously regu lati ng type. In the pro

pellant

tank

the ullage gas may be

a

two

component

mixture

of e vaporated propellan t and

another gas. The ullage gas temperature, composition, and prop ertie s are considered

functions of time and of axial, but not radial o r circumferential, distance.

Liquid and

wall temperature and properties a re treated in the sam e manner. The heat transfer

modes considered are shown

in

Figure 38.

Mass

transfer within the ullage and at the

gas-liquid interface is considered. The effects on heat and ma ss tr ansfe r caused by

gas circulation, as influenced by pre ssu ran t gas

inlet

velocity,

is

also taken into account.

Modifications in

t he

Program

In

the course of

the

comparison s with test data,

it

was necessary

to

make several

modifications in the program

to

obtain good data corr elat ions . Thes e modifications are

discussed in reference

IO.

The ullage gas-to-wall heat trans fer coefficient, which

decr ease s exponentially from the tank

top, is

written as the sum of a

free

convection

coefficient and a for ced convection coefficient.*

where ho is an input constant.

Thus the forc ed convection coefficient

at

the tank top

is a

lin ear function of the

pressu rant volumetric f lowrate ( ed ) from the distr ibutor.

The free convection coef

ficient (h,) is calculated by the free convection equation,

In the sa me manne r the g.as-to-liquid h eat tra ns fe r coefficient at the gas-liquid

interface is written

*

Schmidt

[

121

also writ es the total heat trans fer coefficient as the sum of the fre e and

forc ed convection coefficients.

9


25/104

Ps

zi

hs = hs c + h so

8)

where h

so

is an input constant.

The

free

convection coef ficie nt h

is

calc ulate d by the equation

s c

T

g

4)

It

was found tha t both fo rce d convection coefficients a t the

tank

top could be

calculated more accurateiy by a for ced convection equation of the st anda rd form expre s

sing the Nusselt number a s

a

function of the Reynolds and Pra ndt l number s:

h r

o

=

dl

e)

d3

dz 2)

k

Thus

,

the ullage gas-to-wall heat tran sfe r coefficient and the gas-to-liquid heat tra ns fe r

coefficient at the gas-liquid interface are better

calculated

to equations (7) and

8 ) .

-PwZ

h

= h + h o e

€F c

-P zi

h = h + h

e

Y

s

sc so

wh ere ho and hso

are

calculated

by

equations

(5) and (6) ,

rather than being input

as

cons tants , and he and hsc

are

calculated by equations

(2 )

and

(4 ) .

It was also found that the liquid-to-wall heat tr ans fer coefficients could be better

calculated according

to

a fr ee convection equation ra th er than being taken

as

constank

10


26/104

As

in the cas e of gas-to-liquid heat tra ns fe r coefficient

at

the gas-liquid inter

face , the ma ss tr ansf er coefficient at the interface was

written

where Y is an input constant.

so

The fre e convection coefficient

Y

) is calculated by the equation

s c

The forced convection ma ss tra nsf er coefficient

at

the tank top can be bette r calc ulated

by a forced convection equation expr essi ng the Sherwood number as a function of the

Reynolds and Schmidt numb ers:

Y r

( d3

.

D

Thus, the mass transfer coefficient at the gas-liquid interface is calcula ted by

where

Yso is

calcula ted according to equation (12) ra th er than being input

as a

con

stant, and

Ysc

is calc ulat ed by equation (11).

Evaluation of Program Parameters

A l l pressuri zation analyse s contain numerous param ete rs that must be known

before pressuriz ation requirem ents can be predicted. These parame ters determine the

heat and mas s tr an sf er coefficients and the distributio n of the se coefficients over the tank

Therefo re, studies were conducted to de termine the relat ive importance of e ach of the

pa ram ete rs involved in the progra m, and extensive comparisons with the re su lts

of the tests wer e made to obtain ,numerical values for these p ara me ter s.

A

summary of the test conditions

is

given in Table

II,

and the values

of

the

important para met ers

are

given in Table

III.

The exponential decay coefficients

Pw

and

ps

in equations

(7) (8 )

and (13) are scale d by the equation:


27/104

p =

0.00117 r2.

The para mete rs not l isted in this table are of s ma ll importan ce and may be taken as

zero.

C o m p a r i s o n w i t h T e st Da ta

The pres sur ant flowrate and ullage

and

wall temperature gradients predicted by

the computer progr am using the calculated constants from Table

111 are

compared with

test data [

13, 14,

15, 16, and 171 in Figures 21 through 30. In

all

comparisons the

ullage pre ssu re, l iquid drain

rate,

ambient heat tran sfer coefficients, and ambient

tem pera ture were input to the computer as functions of time.

Either the pressurant inle

tem pera ture o r the heat exchanger performance cu rve was als o input.

Figures

21

through 25 show comp aris ons with t es t dat a obtained with Configura

tion C described in Table

I

and shown in Figure 4

A s

can be s een from these figures,

the agreem ent between the computer predictions and the

test

data

is

general ly good. The

irre gula ritie s in the computed pres sura nt flowrate, pa rticularly marked in

Tests

130-6

and 130-7 (Fig . 21 and 22) ,

are

caused by the over-sensitivity of the pro gram to change

in the slope of the ullage pre ssu re curve. Both ullage pr es su re cur ves of

Test

130-6

and 130-7 have dep ress ion s in the latter half of the ru ns , while the slopes of the ullage

pre ss ure curv es of the other

tests

were nearly constant. The agre eme nt between the

computed and measured ullage temperature gradients was good throughout the run for

all

the

tests

using oxygen

as

pressur ant. In the

test

with helium

as

pres sura nt (130-15,

Fig.

25)

the pres sura nt flowmeter failed. Storage bottle pre ssu re and tem pera ture

history we re use d fo r calculation of

a n

average flowrate. Therefo re,

it

was not unex

pected that the computed flowrate was somewhat below this value. However, the

agree

ment between computed and measured ullage te mpe ratu re gradients was not as good in

this

test as

in the test with oxygen as pres sura nt. This wa s probably caused by deficien

cies

in the prog ram 's m as s tran sfe r calculations from the assumption that

all

heat trans

fer from the ambient to the propellant is converted to sensible heat rather than latent

heat. In Test s 130-9 and 130-10 the ullage heat t ra ns fe r coefficients we re calculated

from calo rime ter m easurem ents anu wer e compared with those calculated by the com

pute r. Although the assump tion of exponential decay of the ullage heat tr an sf er coeffic

ient with distance fro m the tank top Eq. 7) seems arb i t ra ry , the resul t s

were

in excel

lent agreemen t with the meas ured heat tran sfer coefficients

(Figs.

13 and 14) .

In co mparin g the velocity d ecay of

a

free

jet (Fig.

39,

discusse d in

R e f .

18)

it

was found that the exponential decay of the for ced convection heat transfer coefficient

expressed

as a

velocity decay (v,/vo)

is brack ete d by the velocity decay of a free je t

discharging from

a

circular opening and that of

a

free

jet

discharg ing from an infinite

slit. This

is

analogous to the pre ssu ran t ente ring the tank through the gas distributor.

Comparisons with da ta from the

LOX

tanks of the Saturn I, S-I sta ge during

static

test

and flight

are

presented in Figures 29 and 30.

The ag reem ent between computed and

12


28/104

measured pr essu ran t flowrate and pressu ran t inlet temperature is excellent. Ullage

..

temperature measurements we re not available

in

these tests because instrumentation

on

flight vehicles is limited.

Figure

29 shows

a

comparison of

the

computed and measured

flowra te fro m the flight of SA-5. The ag reem en t was gene rall y good, though not as good

as

in the

static

test of SA-6. Evaluation of SA-5 p res su ra n t req uir em en ts wa s compli

, .

cated by the complex

air

flow pa tte rn around and between the propellan t tanks of the

Saturn I,

S-I

sta ge d uring flight.

The aerodynamic heating was difficult to evaluate; the

only possible approach was to use average values for all propellant tanks. Fig ure s 31 and

32 show com par isons with dat a fro m the LOX and LH, t anks of the Saturn I,

S-IV stage

during, st ati c test.

These

tanks

are not of ordinary cylindrical shape, a s can be seen in

Table

I;

the LOX tank is an oblate spheroid and the LH, tank contains a convex inward

lower bulkhead. By comp uter variation of the ch ar ac te ri st ic tank rad ius used in equa

tions (5),(6),and (12) , it was determined that the prope r c ha rac ter ist ic value should be

about two-thirds of the maximum rad ius fo r the LOX tank. This assumption is theoreti

cally justified because a cylinder having the sam e volume and surface a re a a s an oblate

spheroid has a ra diu s equal to 0.63 ti m es the maxim um radius of the oblate sphero id.

The ag reem ent between computed and meas ure d p re ss ur an t flowrate in the LOX tank

is

excellent, as shown in Figu re 31. Because the pr es su ra nt flowra te in the LH, tank

w a s

a

st ep function,

it could not be matched at all times. However, the gen era l range of flow-

rate, a s computed and measured, is the same, and there

is

excellent agreement between

the computed and mea sur ed total pre ss ur an t weight.

Te st re su lt s with Configuration

D

a r e shown in Figures

26

through

28.

This tank

is an approximate one-third sc al e model of the Saturn

V ,

S-IC stage, LOX tank.

It

is

the la rg est sing le cylindrical LOX tank fro m which test data is cur rent ly available. Com

pari son s of the computer p redictions with data obtained from thre e te st s with this configu

ration is good for pr es su ra nt flowrates and tank wall temperatu res.

The final com parison presented is with data from a very sm al l cylindrical tank

(one foot in dia me te r and thre e fe et long) with flat bulkheads (Configuration E ) .

Although

pre ssu ran t flowrate m easurem ents were not available in this test, the computed and

measured ullage temperature hist orie s are compared in Figure 33 . The agreement is

not as good as obtained in Configuration C, probably because equation

(14)

for the scaling

of the exponential decay coefficients w a s developed for tanks with rounded ra th er than

flat

bulkheads.

Con clus ions from Comparisons wi th Test Data

These comparisons with

test

data cover

a

range of conditions, using oxygen,

helium, and nitrogen as pr es su ra nt s and liquid oxygen, liquid hydrogen, and liquid

nitrogen

as

propel lants in tanks ranging in si ze ove r four or d er s of magnitude. The tank

shape s we re repres entativ e of those commonly used in space vehicles, namely cylinders

with various bulkhead shapes and oblate spheroids.

A s a

re su lt of the evaluation of the

many physical parameters and constants involved in the equations, this program can be

used to predic t total and transie nt pre ss ura nt flow requireme nts, ullage and wall

13


29/104

temperature gradients

,

and gas-to-wall hea t transfer coefficients with an accuracy of

=k5percent.

The numerical values of par am et er s recommended by MSFC for use in the

program are given in Table

III.

There

are

presently no other values available in the

lite ratu re. The ch ar ac te ri st ic dimension used in the calculation of the exponential decay

coefficients was taken

as

the radius of the cylindrical section fo r cylindrical tanks. F or

tanks of other s hapes, some comparison with tes t data was ne ces sar y to determine the

prop er choice for the c hara cter istic radius.

A

value of two-th irds of the maximum rad iu s

appears acceptable for oblate spheroids.

The comp aris on with

test

data indicates a sensitivi ty of the program to sudden

changes in ullage pre ssu re. However, in mos t cas es vehicle design pre ssu res a re

either constant

or

vary

in

a monotonic manner.

It

was fur the r found that considerable

experimental experience with pressurization systems is re qu ir ed before this method of

analysis c an be applied reliably

to

evaluate

a

new system.

THE EFFECTS OF SYSTEM PARAMETERS

ON PRESSURANT REQUIREMENT

Weight optimization of propellant tank pre ssur izat ion sy st em s demands that

a

low

pre ssu ran t density be maintained in the ullage space; t his is analogous to using

a

gas of

low molecular weight and maintaining a high ullage mean tem pe rat ure . Therefo re, 30

pressuriz ation te st s and 120 computer predictions were used to se pa rat e the relative

significance of various controllable pa ra me te rs of press uri zat ion sy ste ms and to de

termine their influence on mean 'ullage tempera ture. Figure

40

presents

a

graphical

illustration of the relative influence of these parameters.

From a central origin, representing

a

ref ere nc e condition (Saturn

V ,

S-IC Stage)

for

all

parameters , the increase (+Y) and decreas e

( -Y) ,

of the ullage mean tempera

ture at cutoff

is

shown

as

a function of variatio n of the pa ra me te rs on the absc issa . The

parameter s were varied over a range expected fo r vehicle design. Thus, pr es su ra nt

inlet temperature ca n incr ease o r dec reas e by a facto r of two from the reference condi

tion, p re ss ur e by

a

factor of three, tank radius by

a

factor of

two,

expulsion time by

a

facto r of three, etc.

It

was indicated that the pre ssu ran t inlet tempera ture exer ts the

greatest

influence on the ullage mean temper atur e. Diminishing re tu rn of this effect did

not exist within the rang e of investigation (530"R to 1200OR). The mean tem perat ure

increased as the ullage pr es su re was increased and also as the tank radius was increased.

Increasing the tank wall thickness, heat capacity,

o r

density caused

a

decrease in the

mean temperature.

The pr es su ra nt distributo r flow a re a (AD) that controls the gas-to

wall forced convection hea t tr an sf er coefficient had a significa nt effect on the mean

tempera ture when the ar ea was reduced, but no effect at

all

when flow ar ea was increased.

This indicates that the p res sur ant inlet velocity fo r the refe renc e sys tems was chosen

at an

optimum point.

Figure 40 also indicates that helium pressurant must be introduced

into a tank at a temperature

I.

tim es higher than oxygen pr es su ra nt to obtain the s am e

14


30/104

ullage mean temperature .

This con firms the

results

of other studies (Fig.

2)

indi

cating that the benefits derived from a helium pressuriz ation system are not based on

weight optimization.

C O N C L U S I O N S A N D R E C O M M E N D A TI O N S

a.

Pressuriz ation data fro m cylindrical and spheroidal tanks ranging in si ze

ove r fou r or de rs of magnitude are available for development or checkout of analyt ical

pres suri zati on models and for design of pres surizatio n sy stem s for future launch and

spa ce vehicles.

b. The Rocketdyne tank pre ssur izat ion prog ram , modified as described herein

and utilizing recomm ended coefficie nts, can be used

to

predict total and transient pr es

sura nt requirements and ullage tempera ture gradients with an accuracy of h5 percent.

c. No significant radi al ullage temperature gradient occu rs, even in tanks with

anti-slosh baffles. This pe rm it s the assumption of one-dimensional stratif icat ion of

the ullage gas for analytical represen tation of p ress uran t require ments.

d.

Heat

tra nsf er between pre ssu ran t and tank walls can differ significantly from

fr e e convection, depending on tank geomet ry and d istributor design.

e. The stro nges t influence on pre ssu ran t weight is exerted by pr essu rant inlet

tem pera ture , for which no diminishing retur n occu rs within

a

tempera ture range com

patible with tank mat eria ls. Other important influencing fac to rs

are

tank radius, distrib

utor flow ar ea , expulsion tim e and aerodynamic heating.

The ef fec t of

wall

heat capacity

is

not

as

significant

as

might be expected.

f . M a s s

t ransfer for large tanks is less than previously repo rted .

g.

Additional work

is

nece ssar y to develop better techniques fo r measu ring

gas concentration gradients and m ass transfer.

George C . Marsha ll Space Flight Center,

National Aer onautic s and Space Administration,

Huntsville, Alabama, July 12,

1965.

15


31/104

C O N F I G U R A T I O N

A C

D E B

HEAT EXCHANGER

9

A R A M E T E R

S A T U R N I

C T L 114 SI C 1/3 1x3 MODEL SIV (LOX)

T E S T

PR EPR ESSU R AN T

PR ESSU R AN T

T A N K P R E S S U R E

( p s i a )

14.7

-

60 20-4 0 14.7 - 6 0 4 6

TIME

OF

DISCHARGE (sec. )

.

150 I 5 0

150- 300

I50

-

400

4 78

PR ESSU R AN T T EMP ( R ) 800

510

T O T A L

D I A M E T E R ( i n . ) I @ l 0 5 4 @

70

L / D ( A PP R OX .)

0.45

T A NK M A T E R I A L A L U M . s s s s s s ALUM.

INSULATION

COMMON BL KH

DISTRIBUTOR FLOW

2.5

A R E A ( F T ~ )

I

V O L U M E F T 3

I

I

8

I

1 5 6

I

3

I

2

I

,

T A B L E

I.

T A N K C O N F IG U R A T IO N S A N D T E S T P A R A M E T E R S


32/104

TABLE 11.

SU MMA RY OF TEST CONDITIONS

TEST; FACILITY I

U L L A G E

P R E S S U R E (ps ia ) I INLET T E M P E R A T U R E PR P R E S S U R A N T PRE-PRESSURANT

P RO P E L L A NT

65

I 450 I GOX H e I

C-003

D

20 750

G O X

I

H e L O 2

7 h

I

c-003

20 900

G

OX H e 1 LO ,

- '21)

c-003

-

101), 40 530

H e

He ' 02 .

1)

Tanks not s loshed

2 ) S l o s h i n g d u r i n g S A -5 f l i g h t u n k n o w n


33/104

IPARAMETER 1

I

b l

I

b 2

b 3

d l

I

I

I d 2 I

I

d 3

I

I

C I

I

I

c 4

I

I C 6

I C 8 I

VALUE

0.06

0.8

0.333

0.06

0.8

0.333

0.

I3

0.333

0.I3

0.333

pw=o.oo

117 r

r

IN

FEET

0

T A B L E 111. P A R A M E T E R S F O R H E A T A N D M ASS

T R A N S F E R C A L C U L A T I O N S

18


34/104

PRESSURIZ AT IO N G A S

H A R D W A R E

FIG U RE I. C O M P AR I SO N O F T H E W E I GH T S O F T H E P R O P E L L A N T

F E E D S Y S T EM S O F T W O F L I G H T V E H I C L E S


35/104

1

- -

WEIGHT OF

PRESSURANT,

W l p ( Ib

G a d p s i a )

-

-

0

l-r

..

..

....

-.

ir

4

1

I

T .

r+ I

I

;r;

.O

.+

-

L

c

.

r

m

L1

-I CENTAUR AC-7

JUPITER, THOR

SATURN 18/

S lV,B STAGE

i

--I

SATURN V /

S

II

STAGE

SATURN I /

S- l

STAGE

-I

,SATURN

V /

S - I C

STAGE

1

F I G U R E 2.

WEICiH'I'

OF

LOX T A N K P R E S S U R A N T V E R SU S V E H I C L E THRUST

2 0


36/104

FIGURE

38.

SATURN

I, S-IV

STAGE

FIGURE 3A.

SATURN

I, S-l

STAGE

FIGURE

3.

SATURN

I,

s-I S T A G E ; S A T U R N I, s-IVST.AGE

21


37/104

..

.

FIGURE 4.

TEST FACILITY FOR

T A N K

CONFIGURATION C

22


38/104

FIGURE 5a. INTERIOR OF TANK CONFIGURATION D

23


39/104

F I G U R E 5b.

T E S T F A C I L I T Y F O R T A N K C O N F I G U R A TI O N

D

24


40/104

a

13

0.0442

+ t d

'7 a

1

l

a

1

R ing

t ,?

a

a

0

9

4

a .a a

13

1.64

f l a

16.65

437

7 a

43 1

CONFIGURATION D

FIGURE 6A

Tank Diameter (ft)

Tank Wal l Thickness (ft)

Number of Baffles

Baff le Weight( lb/ f t2)

Baf f le Spacing a

(ft)

Baff le Length r (it)

Baffle Width (ft)

Perforat ion

%

Cyl indr ical Height (ft)

Top Bulkhead Volume

(ft3)

Bottom Bulkhead Volume ( f t3 )

a

39.3

a

34.3

48.1

a

a

1 Calorimeter

a a .

a

a

CONFIGURATION C

FIGURE 6B

FIGURE 6 .

LOCATION OF TEMPERATURE

PROBES


41/104

FIGURE

7.

PRESSURANT DISTRIBUTOR, TANK CONFIGURATION D

26


42/104

INSULATED WIRE

MSFC HIGH RESPONSE

THERMOCOUPLE

ELEMENT LOCATION

TYPICAL RESISTANCE BULB ( RT B 149 A A)

(CERA MIC COATED ELEMENT ON RTB

144)

e; J- r -

SS

TUBING

6

WALL=

35/1000

TEFLON SEAL PLUG

cu co

ROD THERMOCOUPLE

(EARLY PRESSURIZATION TESTS)

F I G U R E

8.

T E M P E R A T U R E S E N SO R S


43/104

I-

0

f.3

M

W

CONTROLLED LAB ORATORY CONDITIO NS

2

MSFC HIGH RESPONSE

T

C.

I

UYSHIELQED

W

c

40

/ PESISTYNCEBULB R ~ B

44

SHIELDED

0 1

c

E$RLY MODEL TC.

* o

W

a LL

W

IO

20

30

40

50

60

70

a TI ME, f (sec.)

ACTUAL PRESSUR IZATION TEST

MSFC T.C.

350

c

a

-

300

3

r

160 L I

0 20

40

60

80

100

120

B

TIME, t

( se c . )

FIG U RE

9.

C O M P AR I SO N O F T E M P E R A T U R E S EN SO R R E S P O N S E T I M E


44/104

FIGURE IO. COPPER PLATE CALORIIVLFTER


45/104

c

lor im e

ter

i L L

50 IO0

TIME, t ( s e c . )

FIGURE 11.

TEMPERATURE RESPONSE OF COPPER PLATE CALORIMETER


46/104

I

I I I I. I

5

w

c

4

I-

z

w

-

u

LL

3

LL

w

0

0

a

w

LL I

v

z

I '

e

a

I

s

W

I

GAS TO WALL

TEMPERATURE DIFFERENCE AT(OR)

F I G U R E

12.

C A L C U L A T E D F R E E C O N VE C T IO N H E A T T R A N S F E R

C O E F F I C I E N T S ,

hf,

O N V E R T I C A L W A L L


47/104

MEASUREMENT

- COMPUTfR

TEST

To

( oA ) P ( ps i a )

130-9

370

62

130-10

450

65

a

a

W

b-

I-

W

W

E E

a

s

a.

*-


48/104

w

0

0

10

5

0

1

W

w

I

+

NEASUREMENT

- COMPUTER

TEST To('R) P(pria)

130-15 560 30

I O

20

P

AXIAL DISTANCE

FROM

TANK

TOP, Z

(f

t.)

FIG U RE 14.

C O M P A R IS O N B E T W E E N E X P E R I M E N T A L A N D C O M P U T E D

H E A T T R A N S F E R C O E F F I C I E N T S , C O N F IG U R A T IO N

C ,

T E S T

130-15,

H ELIU M A S PRESSU RA N T


49/104

-

70

.

tn

n

2 60

.

W

a

3 50

-I---

cn

w

115

e

40-

I 0

-

70

.-

tn

a

-

s 60

e

w

a

2

50-

\

cn

w

a

He

GN2

SLOSHING LIQUID

50 100 150

TIME,

t

(sec.)

NON-

SLOSH

NG

L IQU ID

-

F I G U R E 15. C O M P A R IS O N B E T W E E N U L L A G E P R E S S U R E

LOSS

F O R H e

A N D G N,

P R E P R E S S U R A N T S U N D E R L I Q U ID S L O S H A N D N O N-

S L O S H C O N D I T I O N S I N T A N K C O N F I G U R A T I O N C


50/104

100

W

e9

Q

J

80

z

02

FOR MAIN PRESSURIZATION

0"

l~

60

0

W

L

d 40

>

I-

=.

W

g

20

AND

MAIN

PRESSURIZATION

W

e

-0

IO

LIQUID

LEVEL

AXIAL DISTANCE

FROM

LIQUID

LEVELpi(ft)

FIGURE 16.

M E A S U R E D AN D C O M P U T E D U L L A G E G AS C O N C E N T R A T IO N

G RA D IEN TS IN TA N K CO N FIG U RA T IO N C


51/104

w

Q,

I I

.4

0.2

z

w

0

0

0

0.2

0.4

0.6

To ( O R )

510

310

51

0

532

530

810

6.5 x

40'

6.5'

x

40'

130-7

3 0

130-8

MSFC

5 5

7260

59-D

4 2

50

I I I I

0

I x3 '

7h 7-

0

50 IO0 I50

G

NIT1ON

TIME,

t

( s e d

5 9 - C

CLARK 5 0

59-J

5 0

F I G U R E

17.

E X P E R I M E N T A L L Y D E T E R M I N E D M A SS T R A N S F E R M t /A,

( L B / L B ) V E R S U S T I M E

t

( S E C )


52/104

Vio len t Bo i l ing Dur ing

Venting

V e nt s C l o se d , S t a r t

of

P r e p r e s s u r i z a t i o n

End of P r e p r e s s u r i z a t i o n

S t a r t

of

Drai ning During Dra ining End

of

Drain ing

FIG U RE

18.

LIQ U ID SU RFA CE CO N D ITIO N S D U RIN G PRESSU RIZA TIO N TEST

IN TA N K CO N FIG U RA TIO N C


53/104

w

I g n i t i o n

During Holddown

Cutoff

Residual Liquid Rising During The Firing of Retro Rockets

SA-5

FIGURE 19.

LIQUID SURFAC E AND ULLAGE CONDITIONS DURING SA-5 FLI GH T


54/104

TEST TANK

0

130-6

C

(6.5'x40')

0 COO3-7A D (13'x 26')

HEGHT ABOVE LIQUID LEVEL

( i n ]

1

I 9 8

400

350

a

O

..

W

a

=>

s

W

e

E

W

+

01

0

0.1 0.2 03 0.4 0.5

W A L L

CENTER

DISTANCE FROM

TANK

WALL,

X / D

FIG U RE 20. E X P E R I M E N T A L L Y D E T E R M I N E D R A DI A L T E M P E R A T U R E G R A DI EN T S

w

c9


55/104

Q.

600

500

n

15:

0

'

400

W

a

3

t

300

W

=

W

c

W

200

2i

.I

.I

3

IO0

l

0

0 IO

20 30

40

TOP

AXIAL DISTANCE

F R O M TANK

TOP,

(ft.)

BOTTOM

F I G U R E

2ia.

C O M PA R IS O N B E T W E E N E X P E R I M E N T A L A N D C O M P U T E D U L L A G E

T E M P E R A T U R E G R A D I E N T , T A N K C O N F IG U R A T IO N C y T E S T 1 3 0 -6 ,

O X Y G E N A S ' P R E S S U R A N T


56/104

600

l

m

a

I

I

I

2 100

F I G U R E 2 i b .

C O M PA R IS O N B E T W E E N E X P E R I M E N T A L A ND C O M P U T E D

U L L A G E T E M P E R A T U R E G R A D I E N T , T A N K C O N F IG U R A T I O N C ,

TEST

130-6 O X Y G E N A S P R E S S U R A N T


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5

4

3

I

z

2

a

e

TEST

3

v

v

-

COMPUTER

I

W

a

I

e

0

I I I

IGNITION AT

t =

TIME,

t ( S e C . 1

F I G U R E 21c.

C O M PA R IS O N B E T W E E N E X P E R I M E N T A L A N D C O M P U T E D

P R E S SU R A N T F L O W R A T E , T A N K C O N F I G UR A T IO N C y T E S T

130- 6 ,

O X Y G E N A S P R E S S U R A N T


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70

60

-

50

.-

cn

c

Q

w

40

--a

400

c

c

I /

W

a

m

30

-a 300

m

E

W

W I

a

Q

W

28

4 2001

3

2

IC

a

-

R

c

01

0

PRESSURE

I

COMPUTER

I I I

I

I I 1 1

50 coo 60

TIME,

t ( s e d

F I G U R E 21d.

U L L A G E P R E S S U R E A N D P R E S S U R A N T I N L E T T E M P E R A T U R E

H ISTO RIES, TA N K CO N FIG U RA TIO N

C y

T E S T

130-6 ,

O X Y G EN

A S PRESSU RA N T


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Ip

Ip

w

\

W

2

F I G U R E

22a.

C O M PA R IS O N B E T W E E N E X P E R I M E N T A L A N D C O M P U T E D

U L L A GE T E M P E R A T U R E G R A D I E N T, T A N K C O N FI G U R A T IO N C y

T E S T 130-7

,

O X Y G E N A S P R E S S U R A N T


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6001

7

F I G U R E 22b.

C O M P A RI SO N B E T W E E N E X P E R I M E N T A L A N D C O M P U T E D

U L L A G E T E M P E R A T U R E G R A D I E N T , T A N K C O N F I G U R A TI O N C ,

T E S T

130-7,

O X Y G EN A S P R E S S U R A N T


61/104

5

4

I

I

I

31

2

I

0

50

IO0

IGNITION

AT

t

=

TIME, t ( s e c . )

FIGURE

22c.

COMPARISON BETWEEN EXPERIMEN TAL AND COMPUTED

PRESSURANT FLOWRATE, TANK CONFIGURATION C, TEST

130-7,

OXYGEN AS PRESSURANT


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60 ;00

t

h

0

I : I

I

50

-j500

Q.

c

W

a

40

3

cn

cn

W

a

e

30

300

t

z

4

Q

a

3 20

3

200

3

cn

cn

w

ac

IO

' = 100

0

0

0

I I I

I I

I

I I

I

I

I

Crogenic Tank Pressurant Reqs

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