I

NASA TECHNICAL NOTE

EXPERIMENTAL AND ANALYTICAL STUDIES OF CRYOGENIC PROPELLANT TANK PRESSURANT REQUIREMENTS

by M . E. Nein undJ. F. Thompson

George C. Murshull Spuce Flight Center Huntsville, Ala.

N A T I O N A L AERONAUTICS AND SPACE A D M I N I S T R A T I O N WASHINGTON, D.

d

NASA TN D-3177

EXPERIMENTAL AND ANALYTICAL STUDIES OF CRYOGENIC

P R O P E L L A N T TANK PRESSURANT REQUIREMENTS

By M. E. Nein and J. F. Thompson

George C. M a r s h a l l Space Flight C e n t e r Huntsvil le, Ala.

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - Price $4.00

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 Program Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . Comparison with Test Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Conclusions from Comparisons with Test Data . . . . . . . . . . . . . . . . . . .

THE EFFECTS OF SYSTEM PARAMETERS ON PRESSURANT REQUIREMENT . . CONCLUSIONS AND RECOMMENDATIONS . . . . . . . . . . . . . . . . . . . . . . . . . .

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

BIBLIOGRAPHY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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I I I 1 1 1

U S T OF ILLUSTRATIONS

Figure Title Page

I. Comparison of the Weights of the Propellant Feed Systems of Two Flight Vehicles .. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

2. Weight of LOX Tank Pressurant Versus Vehicle Thrust . . . . . . . . 20

3. Saturn I, S-I Stage; Saturn I, S-IV Stage. . . . . . . . . . . . . . . . . . . 21

4. Test Facility for Tank Configuration C . . . . . . . . . . . . . . . . . . . 22

5a. Interior of Tank Configuration D. . . . . . . . . . . . . . . . . . . . . . . . 23

5b. Test Facility for Tank Configuration D . . . . . . . . . . . . . . . . . . . 24

6 . Location of Temperature Probes. . . . . . . . . . . . . . . . . . . . . . . . 25

7. Pressurant Distributor, Tank Configuration D. . . . . . . . . . . . . . . 26

8. Temperature Sensors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

9. Comparison of Temperature Sensor Response Time . . . . . . . . . . . 28

IO. Copper Plate Calorimeter . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

11. Temperature Response of Copper Plate Calorimeter . . . . . . . . . . 30

12. Calculated Free Convection Heat Transfer Coefficients, hf, on Vertical Wall . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

13. Comparison Between Experimental and Computed Heat Transfer Coefficients, Configuration C, Tests 130-9 and 130-10, Oxygen as Pressurant . . . . . . . . . . . .. . . . . . . . . . . . . . . . . . . 32

14. Comparison Between Experimental and Computed Heat Transfer Coefficients, Configuration C, Test 130-15, Helium as Pre ssur ant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

15. Comparison Between Ullage Pressure Loss for He and GN, Pre-Pressurants Under Liquid Slosh and Nonslosh conditions in Tank Configuration C . . . . .. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

iv

........ ... . .. I

LIST OF ILLUSTRATIONS (Cont'd)

Figure Title Page

16. Measured and Computed Ullage Gas Concentration Gradients 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 Surface and Ullage Conditions During SA-5 Flight. . . . . . . . 38

20. Experimentally Determined Radial Temperature Gradients . . . . . . 39

21a. Comparison Between Experimental and Computed Ullage Temperature Gradient, Tank Configuration C y Test 130-6, Oxygen as Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40

21b. Comparison Between Experimental and Computed Ullage Temperature Gradient, Tank Configuration C , Test 130-6, Oxygen as Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41

21c. Comparison Between Experimental and Computed Pressurant Flowrate, Tank Configuration C, Test 130-6, Oxygen as P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42

21d. Ullage Pressure and Pressurant Inlet Temperature Histories , Tank Configuration C, Test 130-6, Oxygen as Pressurant. . . . . . . 43

22a. Comparison Between Experimental and Computed Ullage Temperature Gradient, Tank Configuration C, Test 130-7, Oxygen as Pressurant. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44

22b. Comparison Between Experimental and Computed Ullage Tempera-ture 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 Pressurant Flowrate, Tank Configuration C, Test 130-7, Oxygen as P r e s s u r a n t . . .................................... 46

V

LIST OF ILLUSTRATIONS (Cont'd)

Figure Title Page

22d. Ullage Pressure and Pressurant Inlet Temperature Histories, Tank Configuration C , Test 130-7, Oxygen as Pressurant . . . . . . . 47

23a. Comparison Between Experimental and Computed Ullage Temperature Gradient, Tank Configuration C, Test 130-9, Oxygen as Pressurant. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48

23b. Comparison Between Experimental and Computed Pressurant Flowrate, Tank Configuration C y Test 130-9, Oxygen as P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 49

23c. Comparison Between Experimental and Computed Pressurant Flowrate, Tank Configuration C, Test 130-9, Oxygen as Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50

23d. Ullage Pressure and Pressurant Inlet Temperature Histories , Tank Configuration C y Test 130-9, Oxygen as Pressurant . . . . . . . 51

24a. Comparison Between Experimental and Computed Ullage Temperature Gradient, Tank Configuration C, Test 130- IO, Oxygen as Pressurant. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52

24b. Comparison Between Experimental and Computed Ullage Temperature Gradient, Tank Configuration C y Test 130-10, Oxygen as Pressurant. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53

24c. Comparison Between Experimental and Computed Pressurant Flowrate, Tank Configuration C, Test 130-10, Oxygen as Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 54

24d. Ullage Pressure and Pressurant Inlet Histories, Tank Configuration C y Test 130-10, Oxygen as Pressurant . . . . . . . . . . 55

25a. Comparison Between Experimental and Computed Ullage Temperature Gradient, Tank Configuration C, Test 130-15, Helium as pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56

25b. Comparison Between Experimental and Computed Ullage Temperature Gradient, Tank Configuration C y Test 130-15, Helium as Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57

v i

LIST OF ILLUSTRATIONS (Cont'd)

Figure Title Page

25c. Comparison Between Experimental and Computed Pressurant Flowrate, Tank Configuration C, Test 130-15, Helium as P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58

25d. Ullage Pressure and Pressurant Inlet Temperature Histories, Tank Configuration C y Test 130-15, Helium as P r e s s w a n t . . . . . . . 59

26a. Comparison Between experimental and Computed Pressurant Flowrate, Tank Configuration D, Test C 003-7a, Oxygen as Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60

26b. Comparison Between Experimental and Computed Tank Wall Temperatures, Tank Configuration D, Test C007-7aY Oxygen as P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 61

26c. Comparison Between Experimental and Computed Tank Wall Temperatures, Tank Configuration D, Test C003-7aY Oxygen as P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62

26d. Ullage Pressure and Pressurant Inlet Temperature Histories, Tank Configuration D, Test C003-7aY Oxygen as Pressurant . . . . . . 63

27a. Comparison Between Experimental and Computed Pressurant Flowrate, Tank Configuration D, Test COO3-12, Oxygen as P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64

27b. Comparison Between Experimental and Computed Tank Wal l Temperatures, Tank Configuration D, Test COO3-12, Oxygen as Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65

27c. Comparison Between Experimental and Computed Tank Wall Temperatures, Tank Configuration D, Test COO3-12, Oxygen as P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66

27d. Ullage Pressure and Pressurant Inlet Temperature Histories , Tank Configuration D, Test COO3-12, Oxygen as Pressurant . . . . . . 67

28a. Comparison Between Experimental and Computed Pressurant Flowrate, 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. Comparison Between Experimental and Computed Tank Wall Temperatures, Tank Configuration D, Test COO3-10, Helium as Pres su ran t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 69

28c. Comparison Between Experimental and Computed Tank Wall Temperatures, Tank Configuration D, Test COO3-10, Helium as Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70

28d. Ullage Pressure and Pressurant Inlet Temperature Histories, Tank Configuration D, Test COO3-10, Helium as Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

29a. Comparison Between Experimental and Computed Pressurant Flowrate, S-I Stage LOX Tanks, SA-6 Static Test, Oxygen as P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72

29b. Ullage Pressure and Pressurant Inlet Temperature Histories, S-I Stage LOX Tanks, SA-6 Static Test, Oxygen as Pressurant . . . . 73

30. Comparison Between Experimental and Computed Pressurant Flowrate, S-I Stage LOX Tanks, SA-5 Flight Test, Oxygen as P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74

3ia. Comparison Between Experimental and Computed Pressurant Flowrate, S-IV Stage LOX Tank, Helium as P res su ran t . . . . . . . . . 75

31b. Ullage Pressure and Pressurant Inlet Temperature Histories S-IV Stage LOX Tank, Helium as Pressurant . . . . . . . . . . . . . . . . 76

32a. Comparison Between Experimental and Computed Pressurant Flowrate, S-IV Stage LH, Tank, Hydrogen as Pr,essurant . . . . . . . . 77

32b. Ullage Pressure and Pressurant Temperature Histories, S-IV Stage LH, Tank, Hydrogen as Pressurant. . . . . . . . . . . . . . . . . . . 78

33. Comparison Between Experimental and Computed Ullage Temperature Histories, Tank Configuration D, Test 187260, Nitrogen as P res su ran t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79

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LIST OF ILLUSTRATIONS (Concluded)

Figure Title Page

34. Comparison Between Experimental and Computed Ullage Temperature Gradient,Tank Configuration C, Test 130-10, Oxygen as Pressurant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80

35. Comparison of Pressurant Flowrate Predictions by Two Computer Programs with Experimental Results. . . . . . . . . . . . . . 81

36. Comparison of Ullage Mean Temperature Prediction by Two Computer Programs with Experimental Results. . . . . . . . . . . . . . 82

37. Comparison Between Experimental and Computed Pressurant Flowrate History, Tank Configuration C y Test 130-6, Oxygen as P r e s s u r a n t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 83

38. Schematic of Heat and Mass Transfer Conditions in a Propellant Tank. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 84

39. Comparison Between Free Jet Velocity Decay and Forced Heat Transfer Coefficient Decay . . . . . . . . . . . . . . . . . . . . . . . . . . . 85

40. The Effects of Various Design Parameters on the Mean Temperature at Cutoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 86

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LIST OF TABLES

Table Title Page

I. Tank Configurations and Test Parameters . . . . . . . . . . . . . . . . . . . . 16

II. Summary of Test Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

III. Parameters for Heat and Mass Transfer Calculations . . . . . . . . . . . . . 18

X

~ DEFINITION OF SYMBOLS

Symbol Definition

A Tank total surface area

AD Pressurant distributor area

Constants used in calculation:"i of gas to wall forcedb3

coefficients (Table m)

Constants used in calculation of f ree

convection heat and mass transfer

Coefficients (Table III & Computer Program)

Tank wall thickness

Constants used in calculations of gas to

liquid forced convection heat and m a s s

transfer coefficients

D Tank diameter ( ft) -D Diffusion coefficient ( f t2/hr )

EK Modification factor for thermal con-ductivity caused by mixing of fluid (Btu/hr f t OR)

ED Modification of diffusion coefficient caused by mixing of gas ( ft2/hr)

gC Constant = 32.17 ( lbm ft/lbf s e C 2 )

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DEFINITION OF SYMBOLS (Cont‘d)

Symbol Definition

h Ullage gas-to-wall. heat transfer gw coefficient

h Ullage gas-to-wall . free convection C heat transfer coefficient

h Ullage gas-to-wall forced convection 0 heat transfer coefficient at tank top

h Gas-to-liquid heat transfer coefficient S

h Gas-to-liquid free convection heat s c transfer coefficient

hL Liquid- to-wall heat transfer coefficient

K Gas thermal conductivity

L/D Tank length to diameter ratio

n i Pressurant flowrate

M a s s transferMt

Am Pressurant mass accumulated

P Ullage pressure

r Tank radius

t Time

T Temperature

Th Vehicle thrust

V Gas velocity

V Tank volume

(Btu/hrf t2 OR)

(Btu/hr f t2 OR)

( B t d h r f t 2 OR)

(Btu/hr f t2 OR)

( B t d h r f t2 OR)

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DEFINITION OF SYMBOLS (Cont'd)

Symbol Definition

vd Volumetric pressurant flowrate at distributor

V Reference volumetric pressurant Od flowrate

X Radial distance from tank wall

Y Gas-to-liquid mass transfer coefficient S

Y Gas-to-liquid free convection mass sc transfer

Y Gas-to-liquid forced correction mass so transfer coefficient at tank top

Z Axial distance from tank top

Z Axial distance of gas-liquid inter­i face from tank top

Dimensional decay coefficient of ullage forced heat transfer coefficients (Table III)

PLP Thermal expansion coefficient of liquid

*T Total time of pressurization

P Gas viscosity

P Gas density

@ Molefraction

(ft3/sec)

(ft /hr)

(ft-1)

(set)

( lb / f t hr)m

(lb /ft3)m

(-)

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Symbol

Subscripts

a

C

f

L

m

0

U

W

DEFINITION OF SYMBOLS (Concluded)

Definition

Ambient

Calorimeter

Free convection

Ullage gas

Interface

Liquid

Mean

Reference, pressurant inlet

Ullage

Wall

ACKNOWLEDGEMENT

The contribution of the MSFC Test Laboratory in providing the test facilities and complex instrumentation and obtaining the experimental data is gratefully acknowledged. Invaluable contributions in program definition and analysis of experimental results were made by J. Moses, T. Stokes, L. Worlund, and G. Platt of the Fluid Mechanics and Thermodynamics Branch.

NOTE:

Mr. J. F. Thompson is currently Assistant Professor , Mississippi State University, Aeronautical Engineering Department. He was formerly associated with Propulsion Division, MSFC.

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I

EXPER IMENTAL AND ANALYTICAL STUD IES OF CRYOGEN I C PROPELLANT TANK PRESSURANT REQU IREMENTS

SUMMARY

The extensive requirement for pressurization of cryogenic propellant tanks of launch and space vehicles has directed attention to the need for accurate methods of analysis of propellant tank thermodynamics. This paper presents the results of experi­mental and analytical studies of pressurization gas requirements for cryogenic liquids. Experimental resul ts are analyzed for cylindrical and spheroidal tanks ranging in s ize over four orders of magnitude. A parameter study of the controllable variables of a pressurization system design illustrates their effect on ullage gas temperature.

Pressurization data are provided for use in the development o r checkout of analy­tical pressurization models and for design of pressurization systems for future launch and space vehicles. A tank pressurization computer program , using recommended coef­ficients, can be used to predict total and transient pressurant requirements and ullage temperature gradients within 10 percent accuracy.

INTRODUCT ION

Determination of the pressurant gas weight for cryogenic propellant tanks is com­plex and defies exact analytical treatment because of the interdependent transient phenomena of heat and mass transfer that occur simultaneously in a propellant tank. Mathematical models describing the internal thermodynamics of tank pressurization have been developed by various investigators.

The analysis by Clark [ i] represents an analytical solution of the governing equa­tions that predict the transient temperature, the response of the pressurant gas, and container wall. However, the solution requires assumptions, such as constant tank pres­sure and zero initial ullage, that are not always met with real systems. The studies by

.Coxe and Tatum [ 21 are based on analysis of a system in which the ullage is thermally mixed a'nd heat transfer between the gas and the wall is independent of time and space. Gluck and Kline [ 31 used dimensional analysis to express gas requirements as a function of known system parameters; they determined, experimentally, quantities of interfacial mass transfer and gas phase heat transfer.

Epstein [ 41 presented a numerical method for calculation of pressurant gas re­quirements that contains a number of phenomena absent from previous analytical methods. However, empirical data are required to evaluate many constants and physical parameters.

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To provide a reliable method for determination of pressurant gas requirements, the experimental data on pressurization obtained by the Marshall Space Flight Center during the Saturn launch vehicle development were applied to the method of Epstein. The physical parameters and the previously indeterminate constants were developed. After modification, this numerical method is capable of accurate prediction of pressurization gas requirements for cryogenic propellant tanks.

PRESSURIZATION REQU IREMENTS AND LAUNCH VEHICLE DES IGN

The increasing s ize and complexity of space launch vehicles necessitates optimi­zation studies of the many subsystems involved in launch vehicle design. The propellant tank pressurization system is of particular importance because its weight is large in comparison to the weight of other subsystems. Weight optimization studies of propellant tank pressurization syste*msfor the Saturn V, S-IC stage, were used to establish the lo­cation of the oxidizer and fuel tanks within the over-all vehicle configuration (Fig. 1 and .Ref. 5). Even the pressurization system components such as heat exchangers, pres­surant lines, and controls, weigh considerably l e s s than the pressurizing gas.

A further indication of the need for optimization of pressurant requirements is illustrated in Figure 2. The pressurant-mass/tank-pressure ratios of typical launch .vehicles is given as a function of vehicle thrust, thrust being representative of vehicle size. Although there is a great deal of difference between the propellant tank configura­tions of tactical missiles and space launch vehicles, a near linear increase occurs in pressurant-masdtank-pressure as vehicle s ize increases. Considering only pressurant gas weight, it appears advantageous to use helium as a pressurant. If, however, the weight of the pressurant storage containers is included in the weight of the pressurization system, the use of helium as a pressurant in most instances resul ts in a weight penalty.

Fo r vehicles with high acceleration and low turbo-pump NPSH requirements, it is possible to eliminate the pressurization system, relying only on the self pressuriza­tion of the saturated propellant (flash boiling) . Flash boiling pressurization, however, resul ts in high pressurant weight and can only be justified if it significantly simplifies ve­hicle design. Because of the infant knowledge of cryogenic tank pressurization at the initiation of the Saturn launch vehicle development program, a long series of pressuriza­tion experiments was conducted at MSFC to obtain system design information and scaling laws for the large propellant tanks of the Saturn I vehicle. Results of this experimental program and correlations with analytical studies are presented in the following sections of this report.

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EXPER IMENTAL PROGRAM

Test Faci l i t ies The experimental work was conducted on five tank configurations at the Marshall

Space Flight Center:

A. Saturn I, S-I Stage, Multiple Interconnected LOX Tanks (Fig. 3a)

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

C. A 6. 5 by 39-ft (DxL) cylindrical LOX tank (Fig. 4)

D. A 13 by 26-ft (DxL) cylindrical LOX tank (Fig. 5a, Fig. 5b)

E. A 1 by 3-ft (DxL) cylindrical LOX tank.

The test parameters for these tank configurations are compared in Table I. Configurations A and B were flight vehicles and thus contained the standard test instru­mentation of the Saturn propellant feed system, including continuous liquid level sensors, tank pressure , pressurant flowrates, and supply temperature measurements. Configura­tions C, D, and E were equipped with many thermocouples along the tank axis. Thermo­couples, mounted at several radi i at three elevations in these tanks, allowed measure­ment of radial temperature gradients. Wall temperatures were measured in Configura­tions C and D by thermocouples on the inside and outside surfaces of the tank at several locations. The locations of the temperature sensors in these tanks are shown in Figures 6a and 6b. Special calorimeter plates were mounted in both tanks for determination of gas-to-wall heat transfer coefficients. Finally, gas sampling devices were placed at several locations to measure ullage gas concentration gradients.

Configurations C, D , and E were equipped with heat exchangers that provide a variable pressurant inlet temperature up to 1000 OR. The pressurant gas was introduced at the top of the container through a distributor (e i ther a deflector plate-Configuration C and E , o r a screen arrangement-Configuration D) to minimize inlet velocities and disturbances of the liquid surface by impinging gas jets. Figure 7 shows a typical distri­butor configuration. Pressurant velocities at the distributor periphery are given in Table I for the five tes t configurations.

The tank Configurations C, D, and E could be sloshed at rotational o r translatory oscillation in excess of the first critical frequency of the tank. Configurations A, C, and D were equipped with cameras so that the conditions inside the tank could be observed. The resul ts of tests conducted with the five tank configurations are presented in Figures 13 through 38. The conditions of these tests are summarized in Table II.

3

I.

Instru mentation

Analysis of ullage gas temperature history required a temperature probe with fast response characteristics and good accuracy. A fast response temperature probe (Fig. sa) consisting of a fork-like support with a 30 gage CuCo welded thermojunction was designed at MSFC. The length-to-diameter ra t io of the thermocouple w i r e and its distance from the fork base were determined using an analog computer representation of the heat transfer conditions around the probe assembly. Figure 9a shows the response time; 63.2 percent of the total temperature change w a s attained in eight seconds when the probe was extracted from liquid oxygen into a gas circulating a t a velocity of about . three feet per second [ 61. The response of the probes during a pressurization test w a s .also determined (Fig. 9b) ; the fork-type thermocouple has a good response characteristic.

A thermocouple mounted on a long, rod-like support (Fig. 8b ) , which was de­signed for liquid measurement in the high vibration environment of static and flight testing, exhibited an extremely poor response in the gas phase as indicated in Figure 9a. Response time to 63.2 percent of total temperature change was in excess of 10 minutes. Commercial temperature probes of the resistance thermometer type (Fig. 8c) were also investigated under these conditions. Although their response was considerably better than the flight type thermocouple (63.2 percent temperature change in approxi­mately 50 seconds), it was too slow for the pressurization studies.

Pressure measurements in the ullage space, pressurant supply lines, and liquid discharge lines were made with close-coupled pressure transducers to assure good response characteristics. The pressurant flowrate and liquid discharge flowrate meas­urements were obtained with turbine type flowmeters. Liquid level before and during the tests was measured by capacitance discrete level probes and continuous delta P measurement of the liquid column.

Test Resu I t s

~ Heat Transfer Coefficients. Heat transfer between pressurant and tank side walls.-

was measured during pressurization tests in Configuration C by two plate calorimeters. Each calorimeter was a 12 by i2-inchY 30-gage copper plate mounted from teflon spacers parallel to and at a distance of four inches from the tank wall (Fig. I O ) . Three thermo­couples, spaced to represent equal calorimeter areas and connected as a thermo-pile, provided a temperature/time history of the copper plate before and during the tests. The local ullage gas temperature was measured in the vicinity of .the calorimeter (Fig. 11). The calorimeters were located I1 and 30 feet from the top of the test tank.

For determination of heat transfer coefficients, it was assumed that heat transfer to the back side of the plate (toward tank wall) was by free convection because of the shielding effect of the plate-to-wall arrangement. The free convection coefficients for a one component gas were evaluated by the equation of Jackson and Eckert [ 71 ; the resul ts are plotted in Figure 12. The free convection heat transfer coefficient was also

4

calculated for two component mixtures based on the time and space dependent helium-oxygen concentration in the tank The total heat transfer to the calorimeters was +en corrected using the calculated free convection effect on the back side. The heat transfer coefficients. to the front of the calorimeter plates measured in Tests 130-9, -10, -15 are presented in Figures 13 and 14 using gaseous oxygen and helium as pressurants. Ullage gas-to-wall heat transfer was also evaluated from wall temperature measurements at -a location 3.5 feet from the top to the tank. Wall measurements at locations initially below the liquid surface produced erroneous readings and were discarded. These coefficients were corrected by subtracting the effect of external heat flux from the measured wall temperature rise. During a flash-boiling test, which did not require pressurant flow, the wall temperature rise indicated an external heat flux of 13 B f d m i n ft'; this compares very favorably with a calculated f lux of 15 Btu/min f t2 [ 81 and confirms the method used for correcting wi l l measurements.

Inspection of Figures 13 and 14 shows very good agreement between measured and calculated heat transfer coefficients. It is noted that the gas-to-wall heat transfer coefficient is definitely within the forced convection regime for 'the oxygen tests, but in the free convection regime for the helium test. AltEiough the heat transfer coefficient by forced convection diminishes with increasing distance from the pressurant distributor, 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. Pressurization studies conducted at MSFC have shown that there is little benefit derived from the use of helium as a main pressurant for cryogenic propellants. However , it was determined experimentally that prepressurization with helium reduces pressure decay during liquid sloshing near the cri t ical frequency. It is assumed that the helium acts as a buffer zone between the splashing cryogenic liquid and th.e condensable pressurant, suppressing excessive mass transfer.

Figure 15 shows a typical tank pressure history for a stationary liquid oxygen tes t tank as compared to a pressure history in which the liquid sloshes near the first cri t ical mode of oscillation [ 91. The tank was prepressurized, with either helium or nitrogen, followed by main pressurization .during liquid expulsion with super-heated oxygen. The.tank pressure history during the slosh test (using helium as a prepres­surant) is nearly identical to the pressure history of the nonsloshing expulsion test. In contrast, prepressurization with gaseous nitrogen resulted in a marked pressure decay during 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 concentration of helium gas in a cryogenic propellant tank during pressurization discharge were studied in test Configurations C and D. Spectrographic analyses were made of gas samples taken at various positions in the tanks. Samples taken at various elevations in tankConfigura­tion C just before the end of the tests yielded the resul ts shown in Figure 16. In the test

5

in which helium was used for prepressurization and oxygen as the main pressurant, the helium concentration is maximum at 12 feet above the liquid, and gradually decreases in both directions.

The concentration of oxygen near the liquid surface is probably caused by accumu­lation of t$e gaseous oxygen that is initially in the ullage before prepressurization. For comparison, Figure 16 also shows the concentration of helium above the liquid oxygen for the case in which helium prepressurization is followed by pressurization with helium during liquid expulsion. The oxygen concentration at 10 feet above the liquid interface was only six percent by volume. The total amount of gaseous oxygen in the ullage was only slightly larger than the amount of oxygen in the ullage before prepressurization (0.77 moles versus 0.73 moles). This indicates that interfacial mass transfer, although small under these conditions, was in the form of evaporation.

Mass Transfer. A comparison of mass transfer resul ts obtained in Configuration C with results obtained by Clark [ I ] is shown in Figure 17. Condensation in excess of 30 percent of the pressurant flow was found by Clark during liquid nitrogen expulsion tests with a I by 3-foot cylindrical tank. Similar results were obtained with the MSFC test Configuration E, also shown in Figure 17. The mass transfer measured in test Configuration C indicates that condensation was 5 to 10 percent. Condensation in the larger facility is less because of the smaller wall-aredvolume ratio of a larger tank.

Comparing the condensation in the small tank with that in the large tank on t;he basis of wall-aredvolume ratio, the values a r e approximately equal. During tests at high pressurant inlet temperature, initial evaporation noted in Configuration C diminished as the test proceeded. However, Clark had found increased condensation at higher pres­surant s e t temperatures in small tanks. These conflicting resul ts point out the incom­plete knowledge of mass transfer.

Condition of Liquid Interface. The condition of the liquid interface in Configura­tion C and during the launch and flight of SA-5, are shown in several f rames from a movie taken inside these tanks (Figs. 18 and 19) . Violent boiling occurred during venting of the tank before prepressurization. As the vents were closed and prepressurization proceeded, the liquid surface became nearly quiescent before discharge. After discharge began,disturbance of the liquid surface caused by pressurant flow and acceleration of the liquid surface were observed; the disturbance diminished a s time and distance between the surface and the pressurant inlet increased.

Radial Ullage Temperaiture Gradients. Radial temperature gradients obtained with Configurations C and D a re shown in Figure 20. In both cases the radial gradients were small, and there apparently exists little difference between the gas flow conditions in the two tanks, even though the gas distributors, baffling, and tank diameters a r e not comparable.

6

The temperature probes at X/ D - 0.025 in Configuration D, which are located be­tween the antislosh baffles (Fig. 5a),recorded virtually the same temperature as probes at smaller radii. It was concluded that the gas circulation in the tank is not appreciably affected by the antislosh baffles, and subdivision of the tank into volume elements perpen­dicular to the tank axis is permissable for the pressurization analysis.

Axial ullage Temperature 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. These two tests were conducted with oxygen as pressurant at about 550"R. There was a rapid increase in temperature of about 30"R immediately above the liquid interface. in these tests, indicating that mass transfer was small. In tests 130-9 and 130-10 (Figs. 23a, 23b, 24a, and 24b) with the same Configuration with oxygen pressurant at a lower temperature, the ullage temperature gradients are much flatter; the rapid increase in temperature immediately above the liquid interface is still in evidence. The ullage temperature gradients in this same configuration with helium as pressurant (Test 130-15; Fig. 25a, and 25b) are con­cave, ra ther than linear as in the tests with oxygen as pressurant, and the increase in temperature just above the liquid interface is very gradual. The concave shape is to be expected in this case because the mass transfer is in the form of evaporation with an ullage that is predominately helium. The linear ullage temperature gradients in tests with oxygen as pressurant indicate that the mass transfer is very small with an ullage that is predominantly oxygen.

Other Test Results. Tests are being performed with Configuration a, but so far only threktests have been completed. The pressurant distributor in this configuration was designed to minimize the gas circulation in the tank, reducing forced convection heat transfer. While this is the desired condition for optimum pressurization system operation, it is detrimental to the response time of the temperature probes as the liquid interface passes. Precise ullage temperature gradients will not be available until this instrumentation is improved. However, preliminary data, with very hot GOX used as pressurant, indicate that the temperature gradients are concave rather than linear as was the case in the tests with Configuration C using colder GOX as pressurant. The con­cave temperatme gradients found in the helium pressurant tests with Configuration C were also in evidence with Configuration D. Pressurant flowrates and wall temperature gradients from these tests are presented in Figures 26a, 26b, 26c, 27a, 27b, 27c, 28a, 28b and 2%.

Pressurant flowrates in the LOX tanks of the Saturn I, S-I stage, during static test and flight are presented in Figures 29a and 30. Figures 31a and 32a show pres­surant flowrates in the LOX and LH, tanks of the Saturn I, s-IV stage, during static test. Finally, ullage temperature histories obtained in a very small tank, Configuration E , containing LN2 pressurized with nitrogen are given in Figure 33.

7

I

PRESSURIZATION ANALYSES

P revious W o rk

Pressurized discharge from cryogenic liquid containers was studied ana-jtically and experimentally by Clark [I]under sponsorship of the Army Ballistic Missile Agency and later MSFC. The analytical solutions obtained by Clark were applied to test data obtained for Configuration C. In Figure 34 the axial temperature gradient through the ullage gas is shown as a fuilction of distance from the tank top or gasdistributor. Excel­lent agreement with test results was obtained for an assumed gas-to-wall forced convec­tion heat transfer coefficient of 10 B t d h r ft2"R. Agreement for a coefficient of 2 Btu/hr ft2"R, approximately in the range of free convection, was poor. This illustrates one limitation of analytical solutions in which the gas-to-wall heat transfer coefficient enters as an independent variable.

In spite of this restriction and the assumption of initial zero ullage volume, the method by Clark was successfully applied in design analyses of the Saturn I pressuriza­tion system. While Clark's analysis assumed stratification of the ullage gas and con­stant heat transfer coefficient, the analysis by Coxe and Tatum [ 21 was based on the as­sumption of a complete thermally mixed ullage gas and constant heat transfer coefficient. Figures 35 and 36 compare test results obtained with MSFC Configuration C with the analytical predictions by the method of Coxe and Tatum. Toward the end of the test, agreement is good possibly because the conditions of constant heat transfer coefficients Gre approached in the large ullage near the end of the run.

A comparison of the pressurant flow requirements with predictions by an analog computer simulation developed by MSFC, is shown in Figure 37. Representation of the pressurization thermodynamics by analog method was difficult because of scaling problems and the extreme sensitivity of the equations to tank pressure fluctuation. In Figure 35, 36, and 37 pressurant flow requirements a re also compared with a digital computer pro­gram developed by Rocketdyne [ 41 and modified by MSFC [ IO]. This program closely matches test data. However, the program insufficiently describes mass transfer and is sensitive to fluctuations of ullage pressure. These fluctuations do not appear in the ineasured flowrates because they are apparently counteracted by the effects of evaporation and condensation [ I1] .

S u m m a r y o f Analytical Program

Since the Rocketdyne program makes maximum use of the techniques of digital computer calculations and is not subject to the restrictive assumptions that are made in other programs, this method was chosen by MSFC for pressurization system analyses. However, extensive comparisons of the program with test data were required to evaluate' the physical parameters and constants initially contained in the program as indeterminate identities. The equations were modified when necessary.

8

I

This program includes in its calculations a pressurant gas storage tank, heat exchanger, and flow control valve. It considers a propellant tank with or witho.ut outside insulation and pressurized with either evaporated propellant or with a gas stored under pressure in a storage tank in which the gas expands nonadiabatically. The ullage pres­sure is controlled by a pressurant flow control valve that has finite maximum and mini­mum areas and may be either the on-off o r the continuously regulating type. In the pro­pellant tank the ullage gas may be a two component mixture of evaporated propellant and another gas. The ullage gas temperature, composition, and properties a re considered functions of time and of axial, but not radial o r circumferential, distance. Liquid and wall temperature and properties a r e treated in the same 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 mass transfer caused by gas circulation, as influenced by pressurant gas inlet velocity, is also taken into account.

Modifications in the Program In the course of the comparisons with test data, it was necessary to make several

modifications in the program to obtain good data correlations. These modifications are discussed in reference IO. The ullage gas-to-wall heat transfer coefficient, which decreases exponentially from the tank top, is written a s the sum of a free convection coefficient and a forced convection coefficient.*

where ho is an input constant.

Thus the forced convection coefficient at the tank top is a linear function of the pressurant volumetric flowrate ( ed ) from the distributor. The free convection coef­ficient (h,) is calculated by the free convection equation,

In the same manner the g.as-to-liquid heat transfer coefficient at the gas-liquid interface is written

* Schmidt [ 121 also writes the total heat transfer coefficient a s the sum of the free and forced convection coefficients.

9

-Ps zi hs = hs c +hso (8)e

where h so is an input constant.

The free convection coefficient h is calculated by the equations c

T g (4)

It was found that both forced convection coefficients a t the tank top could be calculated more accurateiy by a forced convection equation of the standard form expres­sing the Nusselt number a s a function of the Reynolds and Prandtl numbers:

h r -so = dl (e)d3

dz (2)k

Thus , the ullage gas-to-wall heat transfer coefficient and the gas-to-liquid heat transfer 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 Ys sc so

where ho and hso are calculated by equations (5) and (6) , rather than being input as constants, and he and hsc are calculated by equations (2) and (4).

It was also found that the liquid-to-wall heat transfer coefficients could be better calculated according to a free convection equation rather than being taken as constank

10

A s in the case of gas-to-liquid heat transfer coefficient at the gas-liquid inter­face, the mass transfer coefficient at the interface was written

where Y is an input constant. so

The free convection coefficient (Y ) is calculated by the equations c

The forced convection mass transfer coefficient at the tank top can be better calculated by a forced convection equation expressing the Sherwood number as a function of the Reynolds and Schmidt numbers:

Y r ($d3 . D

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

where Yso is calculated according to equation (12) rather than being input as a con­stant, and Ysc is calculated by equation (11).

Evaluation of Program Parameters

All pressurization analyses contain numerous parameters that must be known before pressurization requirements can be predicted. These parameters determine the heat and mass transfer coefficients and the distribution of these coefficients over the tank. Therefore, studies were conducted to determine the relative importance of each of the parameters involved in the program, and extensive comparisons with the results of the tests were made to obtain ,numerical values for these parameters.

A summary of the test conditions is given in Table II, and the values of the important parameters are given in Table III. The exponential decay coefficients Pw and ps in equations (7) (8) and (13) are scaled by the equation:

I

p = 0.00117 r2. The parameters not listed in this table are of small importance and may be taken as zero.

Comparison w i th Test Data The pressurant flowrate and ullage and wall temperature gradients predicted by

the computer program 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 pressure, liquid drain rate, ambient heat transfer coefficients, and ambient temperature were input to the computer as functions of time. Either the pressurant inlet temperature o r the heat exchanger performance curve was also input.

Figures 21 through 25 show comparisons with tes t data obtained with Configura­tion C described in Table I and shown in Figure 4. A s can be seen from these figures, the agreement between the computer predictions and the test data is generally good. The irregularities in the computed pressurant flowrate, particularly marked in Tests 130-6 and 130-7 (Fig. 21 and 22) , are caused by the over-sensitivity of the program to changes in the slope of the ullage pressure curve. Both ullage pressure curves of Test 130-6 and 130-7 have depressions in the latter half of the runs, while the slopes of the ullage pressure curves of the other tests were nearly constant. The agreement between the computed and measured ullage temperature gradients was good throughout the run for all the tests using oxygen as pressurant. In the test with helium as pressurant (130-15, Fig. 25) the pressurant flowmeter failed. Storage bottle pressure and temperature history were used for calculation of a n average flowrate. Therefore, it was not unex­pected that the computed flowrate was somewhat below this value. However, the agree­ment between computed and measured ullage temperature gradients was not as good in this test as in the test with oxygen as pressurant. This was probably caused by deficien­cies in the program's mass transfer calculations from the assumption that all heat trans­fer from the ambient to the propellant is converted to sensible heat rather than latent heat. In Tests 130-9 and 130-10 the ullage heat transfer coefficients were calculated from calorimeter measurements anu were compared with those calculated by the com­puter. Although the assumption of exponential decay of the ullage heat transfer coeffic­ient with distance from the tank top (Eq. 7) seems arbi t rary, the resul ts were in excel­lent agreement with the measured heat transfer coefficients (Figs. 13 and 14).

In comparing the velocity decay of a free jet (Fig. 39, discussed in Ref . 18) it was found that the exponential decay of the forced convection heat transfer coefficient expressed as a velocity decay (v,/vo) is bracketed by the velocity decay of a free jet discharging from a circular opening and that of a free jet discharging from an infinite slit. This is analogous to the pressurant entering the tank through the gas distributor.

Comparisons with data from the LOX tanks of the Saturn I, S-I stage during static test and flight are presented in Figures 29 and 30. The agreement between computed and

12

measured pressurant flowrate and pressurant inlet temperature is excellent. Ullage .. temperature measurements were not available in these tests because instrumentation on flight vehicles is limited. Figure 29 shows a comparison of the computed and measured flowrate from the flight of SA-5. The agreement was generally good, though not as good as in the static test of SA-6. Evaluation of SA-5 pressurant requirements was compli­, .cated by the complex air flow pattern around and between the propellant tanks of the Saturn I, S-I stage during flight. The aerodynamic heating was difficult to evaluate; the only possible approach was to use average values for all propellant tanks. Figures 31 and 32 show comparisons with data from the LOX and LH, tanks of the Saturn I, S-IV stage during, static 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 computer variation of the characterist ic tank radius used in equa­tions (5 ) ,(6),and (12) , it was determined that the proper characterist ic value should be about two-thirds of the maximum radius for the LOX tank. This assumption is theoreti­cally justified because a cylinder having the same volume and surface a rea as an oblate spheroid has a radius equal to 0.63 t imes the maximum radius of the oblate spheroid. The agreement between computed and measured pressurant flowrate in the LOX tank is excellent, as shown in Figure 31. Because the pressurant flowrate in the LH, tank w a s a step function, it could not be matched a t all times. However, the general range of flow-rate, a s computed and measured, is the same, and there is excellent agreement between the computed and measured total pressurant weight.

Test resul ts with Configuration D a r e shown in Figures 26 through 28. This tank is an approximate one-third scale model of the Saturn V, S-IC stage, LOX tank. It is the largest single cylindrical LOX tank from which test data is currently available. Com­parisons of the computer predictions with data obtained from three tests with this configu­ration is good for pressurant flowrates and tank wall temperatures.

The final comparison presented is with data from a very small cylindrical tank (one foot in diameter and three feet long) with flat bulkheads (Configuration E ) . Although pressurant flowrate measurements were not available in this test, the computed and measured ullage temperature histories are compared in Figure 3 3 . 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 rather than flat bulkheads.

Conclusions from Comparisons w i th Test Data

These comparisons with test data cover a range of conditions, using oxygen, helium, and nitrogen as pressurants and liquid oxygen, liquid hydrogen, and liquid nitrogen as propellants in tanks ranging in s ize over four orders of magnitude. The tank shapes were representative of those commonly used in space vehicles, namely cylinders with various bulkhead shapes and oblate spheroids. A s a resul t of the evaluation of the many physical parameters and constants involved in the equations, this program can be used to predict total and transient pressurant flow requirements, ullage and wall

13

temperature gradients , and gas-to-wall heat transfer coefficients with an accuracy of =k5 percent. The numerical values of parameters recommended by MSFC for use in the program are given in Table III. There are presently no other values available in the literature. The characterist ic dimension used in the calculation of the exponential decay coefficients was taken as the radius of the cylindrical section for cylindrical tanks. Fo r tanks of other shapes, some comparison with tes t data was necessary to determine the proper choice for the characterist ic radius. A value of two-thirds of the maximum radius appears acceptable for oblate spheroids.

The comparison with test data indicates a sensitivity of the program to sudden changes in ullage pressure. However, in most cases vehicle design pressures a r e either constant or vary in a monotonic manner. It was further found that considerable experimental experience with pressurization systems is required before this method of analysis can be applied reliably to evaluate a new system.

THE EFFECTS OF SYSTEM PARAMETERS ON PRESSURANT REQUIREMENT

Weight optimization of propellant tank pressurization systems demands that a low pressurant density be maintained in the ullage space; this is analogous to using a gas of low molecular weight and maintaining a high ullage mean temperature. Therefore, 30 pressurization tes ts and 120 computer predictions were used to separate the relative significance of various controllable parameters of pressurization systems and to de­termine their influence on mean 'ullage temperature. Figure 40 presents a graphical illustration of the relative influence of these parameters.

From a central origin, representing a reference condition (Saturn V, S-IC Stage) for all parameters , the increase (+Y) and decrease (-Y) , of the ullage mean tempera­ture at cutoff is shown as a function of variation of the parameters on the abscissa. The parameters were varied over a range expected for vehicle design. Thus, pressurant inlet temperature can increase o r decrease by a factor of two from the reference condi­tion, pressure by a factor of three, tank radius by a factor of two, expulsion time by a factor of three, etc. It was indicated that the pressurant inlet temperature exerts the greatest influence on the ullage mean temperature. Diminishing return of this effect did not exist within the range of investigation (530"R to 1200OR). The mean temperature increased as the ullage pressure 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 pressurant distributor flow area (AD) that controls the gas-to­wall forced convection heat transfer coefficient had a significant effect on the mean temperature when the a rea was reduced, but no effect a t all when flow area was increased. This indicates that the pressurant inlet velocity f o r the reference systems was chosen at an optimum point. Figure 40 also indicates that helium pressurant must be introduced into a tank at a temperature I.I t imes higher than oxygen pressurant to obtain the same

14

ullage mean temperature. This confirms the results of other studies (Fig. 2) indi­cating that the benefits derived from a helium pressurization system are not based on weight optimization.

CONCLUSIONS AND RECOMMENDATIONS

a. Pressurization data from cylindrical and spheroidal tanks ranging in s ize over four orders of magnitude are available for development or checkout of analytical pressurization models and for design of pressurization systems for future launch and space vehicles.

b. The Rocketdyne tank pressurization program, modified as described herein and utilizing recommended coefficients, can be used to predict total and transient pres­surant requirements and ullage temperature gradients with an accuracy of h5 percent.

c. No significant radial ullage temperature gradient occurs, even in tanks with anti-slosh baffles. This permits the assumption of one-dimensional stratification of the ullage gas for analytical representation of pressurant requirements.

d. Heat transfer between pressurant and tank walls can differ significantly from f ree convection, depending on tank geometry and distributor design.

e. The strongest influence on pressurant weight is exerted by pressurant inlet temperature, for which no diminishing return occurs within a temperature range com­patible with tank materials. Other important influencing factors are tank radius, distrib­utor flow area, expulsion time and aerodynamic heating. The effect of wall heat capacity is not as significant as might be expected.

f . M a s s transfer for large tanks is less than previously reported.

g. Additional work is necessary to develop better techniques for measuring gas concentration gradients and mass transfer.

George C . Marshall Space Flight Center, National Aeronautics and Space Administration,

Huntsville, Alabama, July 12, 1965.

15

CONFIGURATION A C D E B

HEAT EXCHANGER 9 PARAMETER SATURN I CTL 114 SIC 1/3 1x3 MODEL SIV (LOX)

T E S T

PREPRESSURANT PRESSURANT

TANK PRESSURE ( p s i a ) 14.7 - 60 20-40 14.7 - 60 4 6

TIME OF DISCHARGE (sec.) . 150 I50 150- 300 I50 - 400 4 78 PRESSURANT TEMP (" R) 800 510

TOTAL

DIAMETER ( in . ) I@l05 4 @ 70

L / D (APPROX.) 0.45

TANK MATERIAL ALUM. ss ss ss ALUM. INSULATION COMMON BLKH

DISTRIBUTOR FLOW 2.5 AREA ( F T ~ ) I

I

VOLUME FT3 II 8:O

I

1 5 6

I1 3:::

I1 2; II ,I::

TABLE I. TANK CONFIGURATIONS AND TEST PARAMETERS

TABLE 11. SUMMARY OF TEST CONDITIONS

TEST; FACILITY I ULLAGE PRESSURE (psia) I INLET TEMPERATURE PR)PRESSURANT PRE-PRESSURANT PROPELLANT

65 I 450 I GOX He I

C-003 D 20 750 GOX 'I He LO2- 7 h I

c-003 20 900 GOX He 1 LO,- '21) c-003 - 101), 40 530 He He '02 .

1)Tanks not sloshed 2) Sloshing during SA-5 flight unknown

IPARAMETER 1 I b l I

b 2

b3

d l II

I d 2 I I d3 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.oo117 r r IN FEET

0

TABLE 111. PARAMETERS F O R HEAT AND MASS TRANSFER CALCULATIONS

18

PRESSURIZATION GAS HARDWARE

FIGURE I. COMPARISON O F THE WEIGHTS O F THE PROPELLANT F E E D SYSTEMS O F TWO FLIGHT VEHICLES

1

- -

WEIGHT OF PRESSURANT, Wlp ( Ib Gadpsia)-- 0 0

l-r 0

!. ..I1

..

.... -.ir4$ 1

1 II -

T .I r+I

II ;r;

.O .+

-L c).. r

m 0

L1'I -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-IC STAGE

1

FIGURE 2. WEICiH'I' OF LOX TANK PRESSURANT VERSUS VEHICLE THRUST

20

FIGURE 38.

SATURN I, S-IV STAGE

FIGURE 3A.

SATURN I, S-l STAGE

FIGURE 3. SATURN I, s-ISTAGE; SATURN I, s-IVST.AGE

21

. .. . -

FIGURE 4. TEST FACILITY FOR TANK CONFIGURATION C

22

FIGURE 5a. INTERIOR OF TANK CONFIGURATION D

23

FIGURE 5b. T E S T FACILITY FOR TANK CONFIGURATION D

24

a

-13 0.0442

+ t d

'7 a 1 l

a 1 Ring

t,? a a 0

9- 4'a . a a 13 1.64

f l a 16.65 4377 a 431-

CONFIGURATION D FIGURE 6A

Tank Diameter (ft) Tank Wall Thickness (ft) Number of Baffles Baffle Weight(lb/ft2) Baffle Spacing a (ft) Baffle Length r (it) Baffle Width (ft) Perforation % Cylindrical 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 6 B

FIGURE 6. LOCATION OF TEMPERATURE PROBES

FIGURE 7. PRESSURANT DISTRIBUTOR, TANK CONFIGURATION D

26

INSULATED WIRE

MSFC HIGH RESPONSE THERMOCOUPLE

ELEMENT LOCATION

TYPICAL RESISTANCE BULB ( R T B 149AA)(CERAMIC COATED ELEMENT ON RTB 144)

e; J- r-39M -SS TUBING 6 WALL= 35/1000 TEFLON SEAL PLUG

cu co ROD THERMOCOUPLE

(EARLY PRESSURIZATION TESTS)

FIGURE 8. TEMPERATURE SENSORS

I-

0

f.3 M

W CONTROLLED LABORATORY CONDITIONS2I 0 MSFC HIGH RESPONSE T C.

I UYSHIELQEDW c 40 / PESISTYNCEBULB R ~ B144

SHIELDED0 1c

E$RLY MODEL TC.

* oW

a LL W IO 20 30 40 50 60 70

a TI ME, f (sec.)

ACTUAL PRESSURIZATION TEST

MSFC T.C. 350

c. a- 3000

3 r

160 I 1 L I

0 20 40 60 80 100 120

B TIME, t (sec.)

FIGURE 9. COMPARISON O F TEMPERATURE SENSOR RESPONSE TIME

FIGURE IO. COPPER PLATE CALORIIVLFTER

c3 0

lorimeter

i L L50 IO0

TIME, t ( s e c . )

FIGURE 11. TEMPERATURE RESPONSE OF COPPER PLATE CALORIMETER

I I I I I. I

5 w­c

4 I-"z w-u II

LL 3 LL w 00

a w LL I v)z I ' e a I­

s I

WI

GAS TO WALL TEMPERATURE DIFFERENCE AT(OR)

FIGURE 12. CALCULATED FREE CONVECTION HEAT TRANSFER COEFFICIENTS, hf, ON VERTICAL WALL

MEASUREMENT- COMPUTfR TEST To (oA)P(psia) 130-9 370 62

130-10 450 65

a aW Wb- I-W W E Ea as a.

*-<- ,h TOTAL 5 -0 s +

130-10

a w 0 --I n I I 1 - I ­

0 IO 20 30 40 TOP AXIAL DISTANCE FROM TANK TOP,Z (ft.) BOTTOM

FIGURE 13. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED HEAT TRANSFER COEFFICIENTS, CONFIGURATION C , TESTS 130-9 A N I 130-i0,OXYGEN AS PRESSURANT

c

w 0010

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.)

FIGURE 14. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED HEAT TRANSFER COEFFICIENTS, CONFIGURATION C , TEST 130-15, HELIUM AS PRESSURANT

-0 70.­tnn Y

2 60 . -W a3 50 -I---­cn w 115

e 40-I 0

-70 n.-0 tna Y - ­s 60­e w a 2 50- \ cn w a

He GN2

SLOSHING LIQUID +--------I­

50 100 150 TIME, t (sec.)

NON- SLOSH ING LIQUID -

FIGURE 15. COMPARISON BETWEEN ULLAGE PRESSURE LOSS FOR H e AND GN, PREPRESSURANTS UNDER LIQUID SLOSH AND NON-SLOSH CONDITIONS IN TANK CONFIGURATION C

100 We9QJ

80 z 0 02 FOR MAIN PRESSURIZATION

0" l~ 60 0 WL3

d 40> I-=.W g 20 AND MAIN PRESSURIZATION W e

-0 IO LIQUID LEVEL AXIAL DISTANCE FROM LIQUID LEVELpi(ft)

FIGURE 16. MEASURED AND COMPUTED ULLAGE GAS CONCENTRATION GRADIENTS IN TANK CONFIGURATION C

w Q, -I-I

0.4

0.2

z w

0 0

0

0.2

0.4

0.6

To ( O R )

510 310 510 532 530 810

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-7­0 50 IO0 I50

1G NIT1ON TIME, t ( s e d

5 9 - C CLARK 5 0 59-J 5 0

FIGURE 17. EXPERIMENTALLY DETERMINED MASS TRANSFER Mt /A, ( LB/LB) VERSUS TIME t (SEC)

V i o l e n t B o i l i n g During Venting

Vents Closed, 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 Draining During Draining End of Dra in ing

FIGURE 18. LIQUID SURFACE CONDITIONS DURING PRESSURIZATION TEST IN TANK CONFIGURATION C

w

I g n i t i o n During Holddown Cutoff

Res idua l Liquid Ris ing During The F i r i n g of Ret ro Rockets SA-5

FIGURE 19. LIQUID SURFACE AND ULLAGE CONDITIONS DURING SA-5 FLIGHT

TEST TANK 0 130-6 C (6.5'x40')

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

HEGHT ABOVE LIQUID LEVEL (in]1

I 9 8

400

350n a O­

.. W a => s a W eE W+

010 0.1 0.2 03 0.4 0.5

WALL CENTER DISTANCE FROM TANK WALL, X / D

FIGURE 20. EXPERIMENTALLY DETERMINED RADIAL TEMPERATURE GRADIENTS w c9

Q.

600

500 n 15: 0 Y' a

400 W a3 t

300 W

= W c W 200 2i .I .I 3

IO0

lu--.,l-l-* 00 IO 20 30 40

TOP AXIAL DISTANCE FROM TANK TOP, I(ft.) BOTTOM

FIGURE 2ia. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED ULLAGE TEMPERATURE GRADIENT, TANK CONFIGURATION C y T E S T 130-6, OXYGEN AS 'PRESSURANT

600 ---lV

m a I

I I2 100

FIGURE 2ib. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED ULLAGE TEMPERATURE GRADIENT, TANK CONFIGURATION C , TEST 130-6, OXYGEN AS PRESSURANT

5

4

3

I­z 2 ae 0 TEST3 v) v) -COMPUTER I W a I e

0 I I I I

IGNITION AT t = 0 TIME, t (SeC.1

FIGURE 21c. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED PRESSURANT FLOWRATE , TANK CONFIGURATION C y T E S T 130-6, OXYGEN AS PRESSURANT

70

60

- 50 .-0 cn c Q w

40 --a3400 c c I /

W a 3 m 30 -a300m E~W W Ia Q

W 28 4d 2001(3

2 3

IC a -R

c P010

PRESSURE

I

COMPUTER I I I I

I I I I 1 1 50 coo 60

TIME, t (sed

FIGURE 21d. ULLAGE PRESSURE AND PRESSURANT INLET TEMPERATURE HISTORIES, TANK CONFIGURATION C y TEST 130-6, OXYGEN AS PRESSURANT

Ip Ip

w \W

2 3

II ! FIGURE 22a. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED

ULLAGE TEMPERATURE GRADIENT, TANK CONFIGURATION C y T E S T 130-7 , OXYGEN AS PRESSURANT

Im

6001 17

FIGURE 22b. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED ULLAGE TEMPERATURE GRADIENT, TANK CONFIGURATION C, TEST 130-7, OXYGEN AS PRESSURANT

5

4

I I I

31

2

I

I

0 50 IO0 IGNITION AT t = 0 TIME, t ( s e c . )

FIGURE 22c. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED PRESSURANT FLOWRATE, TANK CONFIGURATION C, TEST 130-7, OXYGEN AS PRESSURANT

60 -;600t

h 0 I : I ...I 50 -j500 Y Q. c

W a 403cncn W a e 30 300

tz4Q a3 20 '3200

3 cncn wac: IO '= 100

0 00

I I I

I II I I

I

I

, I

Jog TEST I I1 - COMPUTER I 1 1 1 1 1 1

50 100 150

TIME, t (sec)

FIGURE 22d. ULLAGE PRESSURE AND PRESSURANT INLET TEMPERATURE HISTORIES, TANK CONFIGURATION C , TEST 130-7, OXYGEN AS PRESSURANT

600

500 c a 0 Y

400 L

W a 3

!z a 300 W n =E W I­

200W c3

4 -I 3

IO0

I

0 1L 0 10 20 30 40

TOP AXIAL DISTANCE FROM TANK TOP, 3 (ft.) BOTTOM

FIGURE 23a. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED ULLAGE TEMPERATURE GRADIENT, TANK CONFIGURATION C , T E S T 130-9, OXYGEN AS PRESSURANT

IT­

0 IO 20 30 40

TOP A X I A L DISTANCE FROM TANK TW, 3 (ft.) BOTTOM

FIGURE 23b. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED * PRESSURANT FLOWRATE , TANK CONFIGURATION C ,W

TEST 130-9, OXYGEN AS PRESSURANT

cn 0

6

5

4

3

2

I I

1

I i I

I

0 TEST-COMPUTER

I

I' I

I

FIGURE 23c. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED PRESSURANT FLOWRATE , TANK CONFIGURATION C y TEST 130-9, OXYGEN AS PRESSURANT

./­

70 700

-60 0a 600

h .-0

E 50 Y

e m

G 40 a 3

i e z 200 v)W Re

IO 100

0 0

FIGURE 23d.

I I

TEMPERATURETrI I -COMPUTER I

50 100 150, TIME, t (sec.)

ULLAGE PRESSURE AND PRESSURANT INLET TEMPERATURE HISTORIES, TANK CONFIGURATION C, TEST 130-9, OXYGEN AS PRESSURANT

600

500 n

ae: 4

Y

3 480, w" a 3 c

w e 1E w I- 200 w (3a

00

TOP IO 20 30

AXIAL DISTANCE FROM TANK TOP, t (ft)

40 BOTTOM

FIGURE 24a. COMPARBON BETWEEN EXPERIMENTAL AND COMPUTED ULIAGE TEMPERATURE GRADIENT, TANK CONFIGURATION C y TEST 130-10, OXYGEN AS PRESSUFtANT

0

60(

500 n a 0 o t = 100sec.)Y

a +3 400

0 3 I - COMPUTERw a

3

2 300 Q: w eI

w200 w (3 0-C a 4

100

0 0 10 20 30 40 TO P AXIAL DISTANCE FROM TANK TOP, Z ( f t . ) BOTTOM

FIGURE 24b. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED cn ULLAGE TEMPERATURE GRADIENT, TANK CONFIGURATION C ,w

TEST 130-10, OXYGEN AS PRESSURANT

5

4

3

2

I

0­

0 TESTT

FIGURE 24c. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED PRESSURANT FLOWRATE, TANK CONFIGURATION C , TEST 130-10, OXYGEN AS PRESSURANT

700

n60-0: 600 n

0 Y

a­es 50 e 500

W" I a 3 I

d'40 I- 400s We E30 W 300 I-

I­z20 a 200a 3 # I 00 TEST# WIO a IO0 1-COMPUTER II 1 e I I

0 00 100 zoo 300 TI M E, t (sec.)

FIGURE 24d. ULLAGE PRESSURE AND PRESSURANT INLET HISTORIES, TANK CONFIGURATION C y TEST 130- 10, OXYGEN AS PRESSURANT

600

500 n a 0 r Y

A t = 20 scc. $ 400 t = I20 sec.)TEST1 c

~W a I - COMPUTER I ~=> s 300

e3i W I­w 200 I c e9a4 ,I3

100

0 0 IO 20 30 40

TOP AXIAL DISTANCE FROM TANK TOP, Z (ft.) BOTTOM

FIGURE 25a. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED ULLAGE TEMPERATURE GRADIENT, TANK CONFIGURATION C , TEST 130- 15 , HE LIUM AS PRESSURANT

600 I

500 cs a 0 Y I

400 a I

I-a 300Q:W Q.E W I-200

J3

100

0 0

TOP

c 0 t'70 stc. -a t=170sec.

I -COMPUTER

C

IO 20 30 40 BOTTOM

ANAL DISTANCE FROM TANK TOP, 3 (f t.

FIGURE 25b. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED ULLAGE TEMPERATURE GRADIENT, TANK CONFIGURATION C , TEST 130-15, HELIUM AS PRESSURANT

0.5

0.4

0.3

0.2

0.I

0 0 -I--F

0

I o TEST4- COMPUTER

10050 I50IGNITION AT t = O

FIGURE 25c.

TIME, t ( s e c . ) COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED PRESSURANT FLOWRATE , TANK CONFIGURATION C y TEST 130-15, HELIUM AS PRESSURANT

70 700

60 c

0a 600 Y

c e 0

‘3; 50 500 n. 3 Y +0. a we 40 E 400a e 3 s v, W v) I­2 30 + 300 e z a W a (3 3 e 20 g 200 JA W 3 a e

IO 100 -no TEST - COMPUTER

0 00 1 I

50 100 I50 TIME, t (sed

FIGURE 25d. ULLAGE PRESSURE AND PRESSURANT INLET TEMPERATURE HISTORIES, TANK CONFIGURATION C , TEST 130-15, HELIUM AS PRESSURANT

- COMPUTER

TIME, t(mc 1

FIGLXE, 26%. COMP-4RSSON E;E:T\’I’EEN EXPGRIMENTA L AND COMPUTED PRESSbXANT FLOWR4TE, TANK CONFIGURATION D, TEST C 003-7a, OXYGEN AS PRESSURANT

} TEST 4 t S 5 0 s e c

0 t = I50sec

-COMPUTER

20 25 AXIAL DISTANCE FROM TANK'TOP, E ( f t ) BOTTOM

FIGURE 2Gb. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED TANK WALL TEMPERATURES, TANK CONFIGURATION D , TEST C007-7a, OXYGEN AS PRESSURANT

30

I i l i L I I OO 5 IO 15 20 25 30 TOP BOTTOM

AXIAL DISTANCE FROM TANK TOP, Z ( f t )

FIGURE 26c. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED TANK WALL TEMPERATURES, TANK CONFIGURATION D , T E S T C003-7aY OXYGEN AS PRESSURANT

---- TE MPI ERATU RE -A­

~ 4 0-u)Q l-Y a a

W aI I- ’tW PRESSURE I­z a 400 a I

3 (0 00 TEST 1 v)W I

I I I

200- - COMPUTER 1 L

00 5 0 100 150 TIME, t (sec 1

FIGURE 26d. ULLAGE PRESSURE AND PRESSURANT INLET TEMPERATURE HISTORIES, TANK CONFIGURATION D , TEST C003-7a, OXYGEN AS PRESSURANT

200

5

4

3

2

50

1 I I II

0 TEST

- COMPUTER -

TI

IO0 150 200 TIME, t (sec

FIGURE 27a. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED PRESSURANT FLOWRATE, TANK CONFIGURATION D, TEST COO3-12, OXYGEN AS PRESSURANT

5001 I

I­1 I -I

$ 100

FIGURE 27b. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED TANK WALL TEMPERATURES, TANK CONFIGURATION D , TEST COO3-12, OXYGEN AS PRESSURANT

cis 500 2cis

h a 0 Y

3 300

W a 3

at! 200 2 WI­

1 1

9 100

TOP

FIGURE 27c.

0 t = 100sec

A t = 191 sec } TEST

-COMPUTER

AXIAL DISTANCE FROM TANK TOP, Z ( f t ) BOTTOM

COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED TANK WALL TEMPERATURES, TANK CONFIGURATION D , TEST COO3-12, OXYGEN AS PRESSURANT

50 ;IO00 0 Y

n

2 40 u)n Y n .. a

WE 30 E puuI! I3 v) v)W I-a 2

20 2400 a W 3

v) v)4 W

-I a = IO a 200

0 n 1 I "0 50

TEMPERATURE

1 1 1 1 1 100 150

TIME, t (sec.)

FIGURE 27d. ULLAGE PRESSURE AND PRESSURANT INLET TEMPERATURE HISTORIES, TANK CONFIGURATION D , TEST COO3-12, OXYGEN AS PRESSURANT

200

2.0

n

1

- I

LL I

0 TEST aa 1:1 - COMPUTER g 0.s W a <

a r1 1 II:

I5 I I

0 00 50 100

TIME, t (sec 1

FIGURE 28a. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED PRESSURANT FLOWRATE, TANK CONFIGURATION D, TEST COO3-10, HELIUM AS PRESSURANT

3

500

0t=50sec a

A t=14lsec 400

a C O M P U T E RI J CI a 0

300

2 .. W a 3 I- 200 Q:a~

W nL Wk LLL

A X I A L D I S T A N C E FROM TANK TOP, Z ( f t )

FIGURE 28b. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED TANK WALL TEMPERATURES, TANK CONFIGURATION D , TEST COO3-10, HELIUM AS PRESSURANT

-2 0 500

I I I

3 A t = l O O s e c T E S T

400 A - C O M P U T E R

n

e 0 Y

3 I- - 300 W a 3 I-Q e W 200 e

IO 0

C I

T O P 5 IO 15 20 B O T T O M 3025

A X I A L D I S T A N C E F R O M T A N K T O P , Z (ft)

FIGURE 28c. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED TANK WALL TEMPERATURES, TANK CONFIGURATION D , TEST COO3-10, HELIUM AS PRESSURANT

60 I- 1200

50 -

a n

40 ­

.-0 +2 600' w aI w-L 'J I

. . -I ,.w a z

a a N

4 J2 I0l

0 00

7

PRESSURE

---P-@

1 I TEMPERATURE

=---I

I

I

1 1 1 5 0 100 I50

TIME, t (sec)

FIGURE 28d. ULLAGE PRESSURE AND PRESSURANT INLET TEMPERATURE HISTORIES, TANK CONFIGURATION D , TEST COO3-10, HELIUM AS PRESSUMNT

4 N 50

00 50 100 I50

IGNITION AT t = O TIME, t (sec.1

FIGURE 29a. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED PRESSURANT FLOWRATE, S-I STAGE LOX TANKS, SA-6 STATIC TEST, OXYGEN AS PRESSURANT

00

1400

1200 n

0a c Y

c” n

.O 50- .. 1000 v)n W Y a n 3 ca 2 40 - E 800 a , Q v) I

Wv)W cE 30- 600

I-W 2 a a Q a

32 20- 400 W a Q.

-10 200

0 - 0

FIGURE 29b.

U

I1 TEM~ERATURE

- TEST- COMPUTER

ULLAGE PRESSURE AND PRESSURANT INLET TEMPERATURE HISTORIES, S-I STAGE LOX TANKS, SA-6 STATIC TEST, OXYGEN AS PRESSURANT

3

70

W a 60cn­ w .E

E g I

-COMPUTER I-I

50 100 I50 RANGE TIME, t (sed

FIGURE 30. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED PRESSURANT FLOWRATE , S-I STAGE LOX TANKS, SA-5 FLIGHT TEST, OXYGEN AS PRESSURANT

L

/­0­

;/

100 200 3 TIME, t (sec 1

I I I I I I I

TOTAL PRESSURANT WE1OHT:

TEST 88.0 Ib. -COMPUTER 88.41b.

'0 400 SO0

FIGURE 3ia. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED PRESSURANT FLOWRATE, S-IV STAGE LOX TANK, HELIUM AS PRESSURANT

I

W a !z a 20- 200 W 3 W v) ,-Ia v)J W J a I = IO- a 100 t

II

O L 0 I 0 100 200 300 400 500

TIME, t (sec 1

FIGURE 31b. ULLAGE PRESSURE AND PRESSURANT INLET TEMPERATURE HISTORIES, S-IV STAGE LOX TANK, HELIUM AS PRESSURANT

0.3

\ I I I I I I

-,,,J 1 I I I 1 I TOTAL PRESSURANT WEIGHT:

TEST 85.8 Ib. COMPUTER 85.1 Ib.

100 200 300 400 51 TIME, t ( s e d

FIGURE 32a. COMPAREON BETWEEN EXPERIMENTAL AND COMPUTED PRESSURANT FLOWRATE, S-IV STAGE LH2 TANK, HYDROGEN AS PRESSURANT

--

- 600

- -500a L P

L

- E4003 ­k --.a a b S S U R E-=+l

A

/ e L A -A

- Q -I ik A* ~ 3 0 0 - TEMPERATURE +=+dL W I-C z

- ,4200 3 v) COMPUTER v)W a

- p 100

-OO IO0 200 300 400

-a 500 o n TEST 0 Y - COMPUTER I

400 II1--CI

w a 3 I- 300a a w e z 200w c-

W c3 IO0 a J J =>

0 50 100 TIME, t ( s e d

FIGURE 33. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED ULLAGE TEMPERATURE HISTORIES, TANK CONFIGURATION D , TEST 187260, NITROGEN AS PRESSURANT

M 0

DISTANCE FROM TANK TOP, Z(ft.) FIGURE 34. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED ULLAGE

TEMPERATURE GRADIENT,TANK CONFIGURATION C , TEST 130-10, OXYGEN AS PRESSURANT

i

---

- -.>I--.--&

l // I I I ,

I i 0 TEST-I i -ROCKETDYNE ~

1 PROGRAM ,

LOCKHEED PROGRAM

00 50 IO0 150 IGNITION AT t = 0 TIME, t ( s e c )

FIGURE 35. COMPARISON O F PRESSURANT FLOWRATE PREDICTIONS BY TWO COMPUTER PROGRAMS WITH EXPERIMENTAL RESULTS

----

400

n a 0 Y T

,E 3 5 0 I- I I

C.

w a 3 I­,s300 W e s 0 TEST w I- ROCKETDYNE ;

2 PROGRAM2 2 5 0 LOCKHEED -

E PROGRAM w c9a4

5 0 100 I50TIME, t ( s e d

FIGURE 36. COMPARISON OF ULLAGE MEAN TEMPERATURE PREDICTION BY TWO COMPUTER PROGRAMS WITH EXPERIMENTAL RESULTS

' I o TEST o ROCKETDY NE COMPUTER

-=-ANALOG COMPUTER

I . 100 I50 20050

T I M E , t ( s e c )

FIGURE 37. COMPARISON BETWEEN EXPERIMENTAL AND COMPUTED PRESSURANT FLOWRATE HISTORY, TANK CONFIGURATION C y TEST 130-6, OXYGEN AS PRESSURANT

c

PRESSURANT SUPPLY dT d w

INTERNAL INSULATION EXTERNAL INSULATION CONVECTION CONVECTION dQ dQ

CONDUCTION RADIATION - - I -#Odxdt

FIGURE 38. SCHEMATIC OF HEAT AND MASS TRANSFER CONDITIONS IN A P R O P E L L A N T TANK

84

0.0II 10 100 DISTANCE FROM OUTLET, 310 OR Z/a

FIGURE 39. COMPARISON BETWEEN FREE JET VELOCITY DECAY AND FORCED HEAT TRANSFER COEFFICIENT DECAY

Tm /Tmo

rill la P

r ( I - 2) / po = 60 (psi 4) ADO = 1 7 . 8 ( f t 2 ) eo = 150 (sec) ha = 2 ( b t u / h r f t 2 OR) Ta = 530 (OR) cpg6 = 1.51 (btu/ft20R)

Tmo/Tm

FIGURE 40. THE E F F E C T S OF VARIOUS DESIGN PARAMETERS ON THE MEAN TEMPERATURE A T CUTOFF

--

REFERENCES

I.

2.

3.

4.

5.

6.

7.

8.

9.

I O .

11.

12.

13.

Clark, J. A. et al, Pressurization of Liquid Oxygen Containers, The University of Michigan under contract with the Department of the Army No. DA-20-018- o r D-254, March 1961.

Coxe, E. and Tatum, John, Mai&?ro_pe!la$ Tank -Pregsurization System Study and Test Report, Lockheed-Georgia Company, ContractAF 04(611)6087.

Timmerhaus, K. D., Advances in Cryogenic Engineering, Volume 7 and Volume 8 (1962 and 1963).

Fortran P ropam-fo r the Analysis of a Sinffle Propellant Tank Pressurization System, Rocketdyne, Division of North American Aviation, Inc. , R-3936-1, September 1963.

Platt, G. K . , Nein, M. E . , Vaniman, J. , and Wood, C. C . , Feedsystem Problems Associated with Cryogenic Propellant Engines SAE National-Aeronautical Meeting, Washington, D. C. , 1963.

High Response Gas _Tem_p_er&tureProbe for Pressurization Studies, NASA MSFC Memorandum R -P&VE-PTF- 64-M- 70, I964.

-Eckert, E. R. G. , and Jackson, T. W., Analysis of Turbulent Free Convection Boundary.

Atmospheric Heat Transfer to Vertical Tanks Fil led with Liquid Oxygen, A. D. Little, Inc. ,-Special Report No. 50, Contract AF-04-(647) -130, November 1958.

Moses, J. L. , and Nein, M. E. , Evaluation -of Propellant Sloshing on Pressur­ant Requirement for Large Scale Cryogenic Containers,- -presented at the 1962 Cryogenic Engineering Conference, Los Aiigeles, California.

Modification of Rocketdyne Tank- Ee_ssurization ComputCProgram, NASA MSFC Memorandum R- P&VE- PTF-64-M-83, -1964.

Liu, C. K . , Sloshing and~~-Pressure Decay in Pressurized Cryogenic Tanks of Launch Vehicles, NASA MSFC Internal Note MTP-P&VE-P-G1-23, December 1961.

Schmidt, E., Experiments of Heat Trgsfer by Natural Convection, Chem. Ing. -Tech. 28/1956/No. 3, pages 175-180, OctoberT955.

MSFC Test Laboratory Reports CTL 2 , 36, 34, 37, and CTL Flash Reports dated August 17, November 20, and December 11, 1962.

87

REFERENCES (Concluded)

14. SA-I5 Static Test Report, Preliminary Summary Data, Confidential.

15. SA-5 %light Data.

16. Results of Pressurization Tests on-LN2 Facility, NASA MSFC Memorandum R-P&VE-PE- M-526.

17. S-IV Stage Pressurization Data, NASA MSFC Memorandum R-P&VE-PTF-64-M­38, 1964.

18. Aerospace Applied Thermodynamic Manual, Society of Aeronautical Engineers, Section 2.

BIBLIOGRAPHY

J. E. Myers, C. 0. Bennet: "Momentum, Heat and Mass Transfer," McGraw Hill, New York (1962).

R. B. Scott. "Cryogenic Engineering, D. Van' Nostrand Company, Inc., Princeton (1959).

V. J. Johnson: "A Compendium of the Properties of MaLerials at Low Temperature (Phase I) Part I Properties of Fluids, ' I National Bureau of Standards Cryogenic Laboratory (July 1960).

J. H. Hargis: "Thermophysical Prop-exties of Oxygen and Hydrogen.. .. NASA MSFC Internal Note In R-P&VE-P-63-13 (October 1963).

88 NASA-Langley, 1966 M329

.._,,.-

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