UNCLASSIFIED AD 407 315 - Defense Technical … R-5123 5 FORMA 608 B PLATE REV. 1.58 DA IVISION Of'...

UNCLASSIFIED

AD 407 315AD

DEFENSE DOCUMENTATION CENTERFOR

SCIENTIFIC AND TECHNICAL INFORMATION

CAMERON STATION, ALEXANDRIA, VIRGINIA

ITNCLASSIFIED

NOTICE: When government or other drawings, speci-fications or other data are used for any purposeother than in connection with a definitely relatedgovernment procurement operation, the U. S.Government thereby incurs no responsibility, nor anyobligation whatsoever; and the fact that the Govern-ment may have formilated, furnished, or in any waysupplied the said drawings, specifications, or otherdata is not to be regarded by implication or other-wise as in any manner licensing the holder or anyother person or corporation, or conveying any rightsor permission to manufacture, use or sell anypatented invention that may in any way be relatedthereto.

/--

TISIA 8

ADIVISION OF NORTH AMERICAN AVIATION, INC.CANOGA. PARK. CALIFORNIA

Q

R-5123

EFFECT OF HEAT ENVIRONMENT AND LENGTH

OF ULiAGE FITTING ON OPERATING

DURATION OF TEE ATLAS MA-2/5

VERNIER ENGINE

A DIVISION OF NORTH AMERICAN AVIATION. INC.

6633 CANOGA AVENUECANOGA PARK, CALIFORNIA

( Contract -A04(691)-135

Part I, Item 1as Amendedby Call Number

PREPARED BY

Rocketdyne EngineeringCanoga Park, California

APPROVED BY

SfinAtlas/Th y pite rogram Manager

NO. OF PAGES 73 & vi REVISIONS DATE 30 April 1

DATE REV. BY PAGES AFFECTED REM AR K S

FORM R IS.G (PLATE)

A DIVISION OF NORTH AMERICAN AVIATION. INC

FOREWORD

This report was prepared under G.O. 8333,

in 'compliance with Contract AF04(694)-135.

ABSTRACTCThe results of an analytical study to de-

termine the effect of heat input to the

Atlas MA-2/5 oxidizer start tank on vernier

engine duration are presented. Also pre-

sented are the results of a test program

to determine the saving affected in

residual oxidizer weight through the use

of a special oxidizer start tank ullage

fitting with increased length.

f-

R-5123 iii

FORM 608-9 PLATE REV. 1-58


CONTENTS

Foreword . . . . . . . . . . iii

Abstract . . . . . . . ...

Introduction . . . . . . . . . .

Summary .. . . . . . .... . . 3Conclusions and Recommendations . . . 7

Conclusions . . . . . . . . 7Recommendations . . . . . . . . . 8

Discussion . . . . . . 9Propellant Availability . . . . . . . . 9Start Tank Refill . . . . . . . . . 11

Heating Effects . . . . . . . . . . . 15

Digital. Computer Results . . . . . . . . . . 21.

Special Ullage Fitting Test Program .. . . . . 30

Tank Geometry, Nominal Weights . . . . . . . . 30

Liquid Oxygen Weight Determination,

Test Arrangement and Procedures 32

Test Results . . . . . . . . . 39

References . . 49

Appendix A . . . . . . . . 51

Mathematical Model, Listing of Formulas . . . . 51

List of Formulas, Chronological . . . . 51

Nomenclature . . . . . . 54

Appendix B . . . . . . . 55Digital Computer Program . . . . . . . . 55

iv R-5 123

FORM 608-B PLATE REV. 1-58

A DIVISION OF NORTH AMERICAN AVIATION. IN.

ILLUSTRAT IONS

1. Vernier and Start System Schematic 10

2. Oxidizer Start Tank Pressure vs Time History,

Flight Test Evaluation Report Missile 49D 12

3. Ullage Pressure After Tank Repressurization to

600 psig. ....... . .................. 16

4. Weight of Oxidizer in Start Tank After Booster

Separation, Computed and Test Curves ............... 19

5. Weight of Oxidizer in Start Tank After Booster

Cutoff, Normal Missile Flight.22

6. Density of Oxidizer in Start Tank After Booster

Cutoff, Normal Missile Flight ................ 23

( 7. Weight of Oxidizer in Start Tank vs Time for

Various Heat Fluxes............. . 25

8. Weight-of Oxidizer in Start Tank vs Time for Various

Ullage Fitting Lengths.. .... ........ . 27

9. Weight of Oxidizer in Start Tank vs Time With no

Refill of Start Tank ......................... 28

10. Density of Oxidizer in Start Tank vs Time With no

Refill of Start Tank ......................... 29

11. Tank Geometry Showing Nominal Volumes and Weights . 31

12. Test Schematic............ . . .... 33

13. View of Long Ullage Fitting Test Setup, Showing

Oxidizer Start Tank. ..... . ............. 34

14. View of Long Ullage Fitting Test Setup, Showing

Oxidizer Start Tank .......................... 35

1 5. View of Long Ullage Fitting Test Setup, Showing

Pneumatic Control Package ...................... 36

R-5123 v

FORM 608.0 PLATE REV. 1-58

A DIVISION OF NORTH AMERICAN AVIATION, INCý

16, Weight of Oxidizer to Start Tank vs Time From

Initiation of Test, Vent Configuration B . . . . 40

17. Weight of Oxidizer in Start Tank vs Time From

Initiation of Test, Vent Configuration. . . . . . 41

18. Tank Inlet Pressure, and Weight of Oxidizer in

Start Tank vs Time After Vent, Vent

Configuration C .... . ...... . 47

19. Digital Computer Program Flow Diagram . . . . . . 56

20, Sample of Cathode Ray Tube Plot (CRPL#T) . . . . . . 69

21. Sample of Cathode Ray Tube Plot (CRPL#T) . . . . . 69

22. Sample of Cathode Ray Tube Plot (CRLPT) . . . . . . 70

23. Sample of Cathode Ray Tube Plot (CRPL#T) . . . . . . 70

24. Sample of Cathode Ray Tube Plot (CRPLOT) . . . . 71

25, Sample of Cathode Ray Tube Plot (CRPLT) . . . . . . 71

26. Sample of Cathode Ray Tube Plot (CRPL#T) . ... . 72

27. Sample of Cathode Ray Tube Plot (CRPL0T) . . . 72

28. Sample of Cathode Ray Tube Plot (CRPL#T) 73

TABLES

1. LOX Start Tank Refill Study .l.. .. . . . . 14

2, Ullage Fitting and Vent Configuration Test

History, CTL-l, 1962 ........ 37

3o CTL-1 Tests, 20 November 1962 to 19 December 1962 42

vi R-5123

FORM 608-. PLATE REV. 1-58

X;41104::x<IF r WI&- r I=A DIVISION Or NORTH AMKRICAN AVIATION. INM

INTRODUCTION

This is a presentation of the effort to determine the vernier engine solo

duration capabilities for the MA-2/5 propulsion system during missile

flight. The effort consisted of two basic tasks:

Task 1. Oxidizer System: determine the properties and the quantity

of oxidizer in the engine oxidizer tank at sustainer cutoff

Task 2. Fuel System: determine the properties and the quantity

of fuel in the engine fuel tank at sustainer cutoff

In addition, the weight saved by adoption of an oxidizer tank ullage

fitting with increased length is determined.

R-5123

FOF.M 00 0 PL rT MKV, 1.68


SUMMARY

The effect of start tank refill rate, fuel availability, and heat input

the oxidizer start tank on vernier engine solo duration was studied. Ali

the ability of a special (long) MA-5 oxidizer start tank ullage fitting

affect a saving in residual liquid oxygen weight was determined.

Start Tank Refill

The start tank must refill prior to booster cutoff so that the proper

amount of propellants is available for the proposed vernier engine dura-

tion. No problem has been associated with fuel tank refill but, because

of the physical characteristics of liquid oxygen, some difficulties coul(

be encountered. A series of tests was completed at Rocketdyne's Compo-

nents Test Laboratory (CTL) from May to August 1959 to determine the cap.

bility of the oxidizer tank for refilling within the time from engine

start to booster cutoff. The results of these tests are presented.

Fuel Availability

No serious problem has been anticipated in predicting fuel availability.

The bulk temperature and density of the fuel in the start tank will most

likely be the same as that in the main tank and should remain so through-

out the flight.

R-5123 3

FORM 6088. PLATE REV. 1-58

IVO 4CK RET "IIN EA DIVISION OF NORTH AMERICAN AVIATION. INC.

Heat Effects

After booster cutoff during an Atias vehicle flight, the oxidizer start

tank is pressurized, and remains pressurized throughout the sustainer

operation. Separation of the booster section from the balance of the

vehicle exposes the tank: to the following heat sources-

Io Radiation heating from sustainer exhaust flame

2. Radiation heating from vernier exhaust flame

3, Radiation heating from exhausterator (turbine exhaust)

4. Solar radiation

5. Momentary forced convection due to direct -vernier flame impinge-

ment

6. Radiation from other components

Because of the heat input, a propellant bulk temperature rise occurs and.

a portion of the propellant is vaporized. The result of this interchange

is a loss of propellant for vernier operation and, thus, a duration loss.

A mathematical model has been devised to simulate the reaction of the

system to finite amounts of heat flux, Although, admittedly, the model

does not exactly duplicate the physical phenomena, the results obtained

correlate well with. actual test data,

R-5123

FORM 608 B PLATE REV. I-58


Ui isl~e Vittin:

The Mercury Atlas and other special Atlas vehicles do not require a ver-

nier solo phase. Therefore, it would be desirable to eliminate the oxi-

dizer used during this phase. Elimination of the oxidizer during tanking

would result in a saving in initial vehicle weight, which can be transferi

to the payload. The weight saving was to have been brought about by the

use of a special (long) oxidizer start tank ullage fitting. Testing con-

ducted during the MA-2 R&D program indicated the increased length of the

ullage fitting did not prevent complete filling of the tank. Minimizatior

of oxidizer residual is an alternative to elimination of oxidizer during

tanking. A test program has been completed to determine whether or not

an oxidizer start tank, containing a long ullage fitting and having fille,

completely during tanking, would vent the excess liquid oxygen during

the venting cycle. This report presents the results of that test program.

2

R-5123 5

FORMA 608 B PLATE REV. 1.58

DA IVISION Of' NORTH A4EPRICAN AVIATION. IN.

CONCLUSIONS AND RECOMMENDATIONS

CONCLUSIONS

Analytical Study

The mathematical model may be used to determine, with reasonable accurac,

the density and the amount of propellant remaining in the start tank at

sustainer cutoff.

The nominal standard vernier solo duration for a typical missile is 29

seconds. The nominal derated. vernier (525 lb,SL thrust) solo duration

42 ± 3.7 seconds.(

The amount of fuel available for vernier solo operation is approximately

87 pounds, and its density would be the same as that in the main fuel

tank.

Test Program

It is possible for the oxidizer start tank to fill completely when the

special (long) ullage fitting is used. When the level of liquid oxygen

is above the upper holes of the ullage fitting, the excess oxidizer

will be expelled by expansion of the ullage gas as the start tank is

vented.

R-5123 7

FORM 603.3 PLATE REV. 1-50

A DIVISION, OF NORTH AMERICAN AVIATION. INC.

A weight saving, at booster cutoff, of approximately 70 pounds can be

achieved by the use of the special (long) ullage fitting.

RECOMMENDATIONS

T'he existence of a frost layer on the oxidizer start tank at booster cut-.

uff should be verified during a missile flight.

Additional. component tests should be run to gain more information about

the effect of heat flux to the oxidizer start tank.

s R-5123

FORM 608 1 PLATE REV. 1-58

~Wv< 4C c er "uwerw rA DIVISION OP' NORTH AICMRICAN AVIATION. INC.

DISCUSSION

The vernier and start system supplies propellants to the vernier engines

and fuel to the main thrust chamber igniter fuel valves during the igni-

tion stage. It also supplies propellants to the sustainer and booster

gas generators during thrust buildup. Figure 1 is a schematic of the

vernier and start system with its associated components. The remaining

rocket engine components have been omitted for clarity.

The vert jer and start system requirements are varied according to mission

requirements of the Atlas missile. For some missions no vernier engine

solo operation is required. This poses some problems. In the first case,

it becomes necessary to predict the amount of propellants available. In

the secorid case, an appropriate method for el iminating the propellants

used during solo operation could result in a terminal vehicle weight saving

to be used for increasing the payload.

PROPELLANT AVAILABILITY

The major problem in predicting vernier solo duration is the determination

of the amount of available propellants and their density. However, since

the fuel start tank is located in the main fuel tank and since the fuel

has a low vapor pressure, no serious problems are anticipated in predict-

ing fuel availability or the ability of the fuel tank to be refilled after

engine start. The bulk temperature of the fuel in the start tank will

most likely be the same as that in the main itank. The nomioial weight, of

available fuel will be approximately 87 pounds.

R-5123 9

I -- - -l -i lm

INTEGRATED STARTSYSTEM PNEUMATICCONTROL ASSEMBLY.

I •=" ~~ -- n"w I" X

I FUEL

I .---------------------- STARTTANK

ig I FUEL FILL

AND CHECK

--------- I VALVEI I _I

PRESSURE OPERATED I IOXIDIZER BLEED AND I IBACK PRESSURE VALVE - I I

~~ ... .........' S

--------------... --- -- 'START SYSTEM

• PROPELLANT VALVEL

CUSTOM ERCONNECTPANEL TO SUS'

THRUST

VERNIER FUEL

- - - OXIDIZER----- PNEUMATIC

10

IL

A DIVISION OF NORTH AMERICAN AVIATION, INC.

DISCONNECTS BETWEENBOOSTER AND VERNIER

OXIDIZER FILL AND OXIDIZERCHECK VALVE- 7 PRESSURE

REGULATOR

[---. - .---. -. -•- -. € -

---- -- I"

STRAINE -STRAINER

II . LOW

ST® •PRESSUREI GAS GENERATOR BYPASS-VEN TTO

I f OVERBOARDLOW PRESSURE

RELIEF VALVE S~NO. IBOOSTERVENT AND RELIEF VALVE TURBOPUMP

OXID TURBOPURP) STARTTANK

I FILL AND* CHECK VALVE O

TONO2 OOTBOOSTER

STRAINER--, - FUEL /-FUEL

TRIGNITER O I P SIGNITER

VALVE VALVE

TTOUST CHAMBER OXIDIZER PRESSURE TO NO.SI STRAINERREGULATOR BOOSTER

TO SUSAINERT

THRUSTCHAMBER

•IGNITER FUEL

-ORIFICE- SUS1.Vei r RAINER9 CHECK .. TURBOPUM P

TO SUSTAINER- VALVE--THRUST CHAMBER"

GA

STO SUSTAINE GA GENE ATO

IZER THRUST I- EEAO

JMTC CHAMBER r)"

JMATICORIICE/ '

BOOTSTA-• •Figure i.Vernier and Start .SystemBOOTSTRAP Schemai

MANIFOLD STRAINERSUSTAINER R%-5123


(I

The oxidizer start tank is external to the main oxidizer tank and the

vapor pressure of liquid oxygen is very high. Figure 2 is a pressure

vs time history of a typical tank. The sequence of events are as follow

1. The oxidizer start tank is initially filled from a ground sourc

through the low-pressure bypass line emanating from the No. 1

booster low-pressure duct and also through lines emanating from

the No. 1 booster and sustainer high-pressure ducts.

2, The tank is pressurized for engines start.

3. The tank is vented to allow refill from the sustainer high-

pressure duct.

4. A short time after booster cutoff the tank is repressurized.

( 5, The tank is maintained in a pressurized state for the balance

of sustainer operation.

6. The pressurized liquid oxygen is used for vernier engine solo

operation.

Two critical periods exist: while the tank is vented and refill is in

process, and while the tank is held in a pressurized state throughout th

remainder of the sustainer operation and it is exposed to several heat

sources.

START TANK REFILL

The refill period is critical because the oxidizer start tank must be

filled with the required amount of liquid oxygen for the proposed vernie

duration.

R-5123 11

X cL4: 4-- 1W uw -I RA DIVISION OF NORTH AMERICAN AVIATION. INC

0

__Pr)

0

N 000 ('a

-J

-- U)

UE- ULI-- 0

Ic- N~

U) 0

0 0

zz

-~O I-IJU)V

0 Lu 0 0

0 000r 0L 00

0 (0 U ~ NP'In- -t UflS) c~NV XO1ý

12 11 )0- 4-5i2

A DIVISION OP NORTH AMERICAN AVIATION. INC.

A series Of tests was completed at Rocketdyne's Components Test Laboratory

from May to August 1959 to determine the capability of the oxidizer tank

for refilling within the time from engines start to booster cutoff. The

results of these tests are shown in Table 1.

The tests were run with various vent configurations. Theoretically the

refill race is affected by the back pressure presented to oxidizer flow.

The back iressure or, in this case, the start tank pressure for a vented

tank is a function of the vent configuration and the vapor pressure of

the liquid. The vent configuration used for each test is shown in

Table 1 under Remarks.

The theoretical ullage volume was 50 cubic inches for all tests. AmnbicnL

temperatui-es of 130 F were attained by running the tests :in an environmental

chamber.

In all cases the required liquid oxygen weight was achieved when missile

configuration vents were utilized at test site and simulated altitude

ambient pressures. A high correlation was found between the weighit

of liquid oxygen in the tank and the refill rate. The tests also indicat ed

that refill was accomplished within the indicated time interval, except,

ror Test No. 316.

All of the tests displayed an increase in start tank pressure after repres-

surization, which soon equaled the feedline pressure. The cause was con-

s idered 0o bee either the vaporization of liquid oxygen due to heat input,

or the continuance of some liquid oxygen flow into the tank which compl-essed

the ullage volume.

R-5123 13

LOX START T

LOXLOX Weight, pounds Refill Average Ullage Feed Line

Test Start Time, Rate, Pressure, Pressure, TNo. of Refill Refilled seconds lb/sec psig psig

197 14 144 --- * 0.155 *

198 107 142 * 0.110

210 42 141 -- * 0.158

212 86 143 -- * 0.185 -*

248 140 150 17.5 0.571 620

251 141 150 8.8 1.02 -* 602

231 127 130 37.5 0.08 .* 690

236 121 130 40.0 0.225 -* 670

257 128 158 152.5 0.197 -- * 610

308 144 158 66.0 0.212 50-35 600

309 138 160 57.0 0.386 34-25 590

310 138 158 77.5 0.258 20-15 600

311 135 162 72.5 0.372 39-20 610

313 146 161 77.5 0.193 35 620

314 141 163 55.0 0.400 42 610

316 144 161 137.5 0.124 600-20 620

*Data not available or reading not good.

0L


TABIE 1

LOX START TANK REFILL STUDY

LOXFeed Line LOX AmbientPres,sure, Temperature, Temperature

psig F F Remarks

620 -- * 130 Start tank not vented, vent and relief valve by-passed with 1/4-inch orificed line, orificed gateremoved from oxidizer fill and check valve

602 -- * 130 Start tank not vented, No. 32 orifice in ventbypass line, gate removed from fill and checkvalve

690 * 130 Start tank not vented, No. 80 orifice in ventbypass line

670 * 130 Start tank not vented, No. 47 orifice in vent by-pass line, gate removed from fill and check valve

610 - * 130 Start tank not vented, No. 32 orifice in vent

line

600 -288 130 16 psi APlow pressure relief valve installed,enlarged hole in gate of fill and check valve

590 -290 62 16 psi AP low pressure relief valve installed,enlarged hole in gate of fill and check valve

600 -295 62 No low pressure relief valve installed, ventline open to atmosphere

610 -290 62 10 psiAP low pressure relief valve

620 * 81 10 psi AP low pressure relief valve

610 -* 78 16 psi 6P low pressure relief valve

620 -280 80 No low pressure relief valve installed

R-5123

A DIVIION OF NORTH AMERICAN AVIATION. INC

An analysis was made to identify the cause of the ullage pressure increas

Figure 3 is a graph of calculated ullage pressure vs time and is compared

to a curve taken from an actual missile flight, which was typical of seve:

flights reviewed. The theoretical curve was formulated by assuming the

ullage pressure increase to be caused by additional liquid oxygen flowing

into the tank under fill-line pressure and compressing the ullage gas at

a constant temperature. This was considered a reasonable assumption for

preliminary calculations, because the liquid oxygen bulk would absorb the

greater amount of heat caused by compression of the gas. The difference

between the two curves is attributed to the fact that heat input to the

tank was ignored. It was concluded from the analysis that additional

J liquid oxygen was made available for vernier solo operation and amounted

to approximately 0.79 pound. The exact amount made available is dependent

on the initial conditions.

HEATING EFFECTS

The second critical period is encountered while the tank is held in a

pressurized state after booster separation, during the remaining sustainer

operation. At this time the tank is exposed to the following heat sources

1. Radiation heating from sustainer exhaust flame

2. Radiation heating from vernier exhaust flame

3. Radiation heating from exhausterator (turbine exhaust)

4. Solar radiation

5. Momentary forced convection due to direct vernier flame

impingement

6. Radiation from other components

R-5123 15

A DIVISION OF NORTH AMERICAN, AVIATION. INC

700 IACTUAL FORTYPICAL MISSILEFLIGHT

680 ,_ _ .

Lif

u. 660

CALCULATEDz/_1- 640

I-

•"620x0

6000 2 4 6 8 10 12

TIME FROM REPRESSURIZATION TO600 PSIG, SECONDS

Figure 3. Ullage Pressure After TarnRepressurization to 600 psig

16 R-5123

A DIVISION OF NORTH AMI4R|CAN AVIATION. IN.

(I

Because of the heat input, a bulk temperature rise occurs in the liquid

oxygen, and also a portion is vaporized. The bulk temperature rise cause

a decrease in density and the vaporized liquid oxygen requires more volun

than the original. The result is that a portion of the liquid oxygen in

the tank is forced back into the high pressure line, causing a loss in

liquid oxygen available for vernier solo operation.

A mathematical model was formulated to simulate the reaction of the syste

to finite amounts of heat flux. Appendix A is a step-by-step listing of

the formulas involved. The model is based on the relationship of exposed

tank surface area to heat input. Essentially, it is set up so that the

heat flux to the tank surface area, wetted by liquid oxygen, serves only

to raise its bulk temperature, and the heat flux to the surface area of

the ullage vaporizes liquid oxygen at the liquid-to-gas interface. Also,

the temperature gradient between liquid and gas vaporizes oxidizer. As

liquid oxygen is vaporized and, in turn, forces liquid from the tank,

the wetted surface area and ullage surface area change. This is a func-

tion of time. Therefore, the model evolves into a series of calculations

over specific time increments, using the results of one set of calculatio

as the initial values for the next. The smaller the time increment, the

more accurate the results. Although, admittedly, the model is not an

exact duplication of the physical phenomena involved, as will be shown,

the results obtained correlate well with actual test data.

A previous analytical study at Rocketdyne determined that the net heat

flux incident on the tank from the various sources present during a

missile flight was 6159 Btu/hr-sq ft. This value was used for the first

set of calculations utilizing the mathematical model. The generated

R-5123 17

FORM 608.1 PLATE RXV. 1-58

A DIVISION OF NORT14 AMERICAN AVIATION. INC.

curve of liquid oxygen weight vs time is shown in Fig. 4 and is compared

to a similar curve obtained from Test No. 308 of the test series described

earlier in this report.

Test No. 308 was set up to simulate the calculated heating conditions

around the oxidizer start tank- during a flight by exposing the tank to a

bank of infrared lamps, The infrared bank consisted of seven 2000-watt

lamps and three 1000-watt lamps, arranged. against a semicircular reflec-

tor set on one side of -the tank.

An ambient temperatiure of 1.30 F was obtained by running the test in an

environmental chamber.

A large discrepancy existed between the calculated values for liquid.

oxygen weight vs time and the valves obtained from the test, For the

next attempt with the mathematical model., an average heat flux was de-

termined from the test data, A value of 975 Btu/hr-sq ft -was used.,

The curve generated is shown in Fig. 4. This curve correlated well with

the test curve, the maximum difference between the two being approximately

3 pounds after 200 seconds, The initial conditions for all the curves are

as follows-

Tank volume (corrected for contraction

due to temperature), cu in, 4109

Liquid. oxygen weight, pounds l.60

Liquid oxygen temperature, F -279

Liquid oxygen density, lbi'cu ft 68,65

Tank: pressure, psia 600

18 R-5123

FORM 608 3 PLATE REV. 1.58

IFC31,4K-:MC J N7 W "SFT JA DIVISION OF NORTH AMERICAN AVIATION. INC

50 I0

WLI-.cfl CD )

m z

0 W-

/ o w

-10

/1•1 oi0 03

/c/w/W

0 0 0 00 00

((0 In

SONA~ '1HDJM X-

R-5123 1


It was decided to attempt an explanation of the discrepancy between the

curve using 6159 Btu/hr-sq ft and the test curve to justify the use of

975 Btu/hr-sq ft as the value for heat flux in further calculations. A

thorough check of all formulas and values used for the analytical study

showed that no consideration had been given to the possible presence of

a frost layer on the tank., Also, the efficiency of the lamp bank was

assumed to be higher than the value recommended by the manufacturer.

The heat input used throughout the test was corrected for the proper lamp

efficiency, but no correction was made for the possible consequences of a

frost layer.

An analysis was conducted to evaluate the assumption of a frost layer.

Combined convective and conductive heat transfer effect was calculated

to determine its surface temperature. Using a value of 0.035 Btu/hr-sq

ft-F/ft (corresponding to a frost layer thickness of 0.15 inch) for frost

thermal conductivity, results in a over-all heat transfer coefficient of

2.24 Btu/hr-sq ft-F. The calculated surface temperature for the frost

layer was -230 F, utilizing a heat flux of 975 Btu/hr-sq ft. The radia-

tion effects were also evaluated, Comparing the heat flux obtained from

the test data (975 Btu/hr-sq ft) with 6159 Btu/hr.-sq ft, a value of 0.208

for frost layer surface emissivity, and absorbtivity results. The values

for frost layer surface temperature, emissivity, and absorbtivity compare

favorably with those obtained in Ref. 1.

It was decided that the discrepancy can, in fact, be attributed to neglect

of the surface emissivity and absorbtivity of a frost layer. There is no

reason to doubt that a frost layer, formed on the start tank at or near

sea level, will be transported to higher altitudes.

20 R-5123

FORM 60 B PLATE REV. 1.58

4O4CIMMETIrr t EA OIVIGION OP NORTH AMERICAN AVIATION. INC.

It is desirable to verify the presence of a frost layer during a missile

flight at booster cutoff. A motion picture camera aimed at the tank and

actuated at booster cutoff, or thermocouples located approximately 1/16

inch to 1/8 inch from the oxidizer start tank skin could be used for this

purpose.

If, at some later date, the presence of a frost layer is verified or

disproved, a more exact value for heat flux to the oxidizer start

tank can be used.

DIGITAL COMPUTER RESULTS

The method for calculating the effect of heat input to the oxidizer start

tankc lends itself to the use of a computing machine. A digital computer

program was written and is described in Appendix B. The program is

designed to handle combinations of tank volume, initial propellant weight,

propellant density, heat input, and ullage plug length.

The computer program was used to formulate the results presented below.

Calculations were performed to determine the final weight and density

of the liquid oxygen remaining in the start tank after sustainer cutoff

during a normal missile flight. Figure 5 and 6 plot the weight and den-

sity vs time. Starting time was considered to be approximately 10

secon~ds after booster cutoff, when the tank pressure equals the refill

pressure. The following initial conditions were used:

Tank volume (corrected for contraction due to

temperature), cu in 11109

R-5123 21

1W 0 N•K ¶T DINA DIVISION OF NORTH AMERICAN AVIATION. INC

170

160-

150on0z

0n. 140

• 130

X0-J

120

110

1000 20 40 .60 80 100 120 140

TIME FROM BOOSTER CUTOFF, SECONDS

Figure 5. Weight of Oxidizer in Start Tahk After Booster Cutoff,Normal Missile Flight

22 R-5123

l--LO 4 JC E7 Tj " 1k 4 SEA nlVI-ON OF NORTH AMERICAN AVIATION INC

69.0

68.8

I-.LL

U 68.6

m-J

• 68.4

zw

)< 68.20

68.0

67.80 20 40 60 80 100 120 140

TIME FROM BOOSTER CUTOFF, SECONDS

Figure 6. Density of Oxidizer in Start Tank After 1oosterCutoff, Normal Missile Flight

11-5123 23


Liquid oxygen weight, pounds 160o8

Liquid oxygen temperature, F .-280

Liquid oxygen density, lb/cu ft 68°92

Heat input, :Btu/hr-sq ft 975

Tank pressure, psia 685

Sustainer duration., seconds 135

The final liquid oxygen weight and density available for vernier opera-

tion are 155o74 pounds and. 68°28 lb/cu ft, respectively. When this

information is combined with vernier influence coefficients, the vernier

solo operation can be calculated. This was done and the nominal standard

vernier solo duration was determined to be 29 seconds. The nominal

derated vernier (525 pounds, SL thrust) duration would be 42 seconds,

A set of calculations was performed to show a comparison of liquid

oxygen weight in the start tank vs time for various values of heat

flux, Figure 7 is a. plot of the curves generated. The following initial

conditions were used:

Tank volume, cu in 4109

Liquid oxygen weight, pounds 160

Liquid oxygen temperature, F -279

Liquid oxygen density, lb/cu ft 68.65

Tank pressure, psia 600

24 R-5123

FORM 608.5 PLATE RrV. 1.53

160

155--_ _ _ _ _ _

150

145

140

(nz 135

0a.

S130

xo 12.5

120

110

105

Uo

100 10 20 30 40 50R-5123 TI ME FROM


I I T 560 BTU/HR'-FT 2

1 000 BTU /HR- FT2

42000 BTU/HR-FT

2-5

I-

IE;4 --v 4cul NE JWTIsllWlA DIVISION OF NORTH AMERICAN AVIATION. INC

U) Uf

21 zI gU

C'j2 x 3C D z

-~ C'H

(DW zUDc) - in

00 ~ o 0. 0-0v(DOa. (D O . Dj(z j(1

_~j z~ < <W:5Dppt -J

z E-z

0 - 1

I-N 0

z-CIPWx 0 )

0 000 0 0N~~~ 0 Di

SO~fl~ 'IH9J. XO

R.-5123 2


0

10

. 0 .,

I 0 4

cN C....

'c: cd

N0

-iH

z '

•NH0C

b4)

- - - - - -- - - - - - - - - - -•Z,• .

0

W

~. ~ - ---- -

o (D 0 WO (D 'It N 0 ca (D It NJ 0•" 1I) I0) re IN N NW 0 N - -. . . . .

SONflOd 'IH913M Xo0l

28 R-5123

A DIVISION OF NORTH AMERICAN AVIATION. INC:

C0

0

ui

U-

0 L

0

- z

0

loW

0(0 i) ~ rr) -: 0 o~ (0 r- ( in *~t V)w

(0(0((0(00(0( N N N NN N

Ld flo/91. 'A.LISNJG XO1l

R1-5123 29


It is not recommended. that weight saving be accomplished 'by a reduction in

tank size. The anomalies involved in predicting the -weight of oxidizer

available require the use of an oxidizer pad to ensure the proper tanking,

This could reduce the saving substantially, Also, the fuel weight saving

can be accomplished by incorporating an ullage fitting in the present

tank.

SPECIAL I-IIAGE FITTING TEST PROGRAM

The Mercury Atlas vehicle does not require a vernier solo phase, there-

fore, it is desirable to eliminate the oxidizer normally used during this

phase, A special (long) oxidizer start tank ullage fitting (MCR, MA5-5B)

was adopted as a convenient means to prevent tanking of oxidizer in excess

of that required for starting the engines, Testing conducted during the

MA2 R&D program indicated the increased. ullage fitting length did not

prohibit complete filling of the tank, Minimization of oxidizer residual

is an alternative to the elimination of excess oxidizer during tanking,

A test program was conducted at CTL-1 to determine whether or not an oxi-

dizer start tank, containing a long ullage fitting and having filled com-

pletely during tanking, would vent the excess liquid oxygen during the

venting cycle,

TANK GEOMETRY, NOMINAL WEIGHTS

The long ullage fitting (P/1 N 304411) is identical to the standard. fitting

(P/ 306231), except for the length of the tube, It extends approximately

to the center of the oxidizer start tank (Fig, 11), Holes positioned

30 R-5123

FORM 608 9 PLATE REV. 1-58

A DIVISION OP NORTH AMERICAN AVIATION, INC

ULLAGE FITTING304411 (SHOWN)306231 (STANDARD, NOTSHOWN)

& TANK ASSEMBLY

NOMINAL IL•

ULLAGEDEPTH TUBE, I IN. OD

x 0.083 IN. WALL 12 HOLES ( EACH_, HICKNESS 0.312 INCH DIAMETER)

APPROXIMATE

RADIUS-- BAFFLE

WITH SHORTULLAGE FITTING, WITH SHORT WITH LONG

AND TANK ULLAGE ULLAGECOMPLETELY FITTING FITTING

FILLED

NOMINAL ULLAGE DEPTH, INCHES 0 1.349 9,479

NOMINAL-TANK VOLUME, CUBIC INCHES 4109.5 4109.5 4109.5

NOMINAL ULLAGE VOLUME, CUBIC INCHES 0 54.3 1930.8NOMINAL EFFECTIVE LOX VOLUME, CU1IC INCHES 4109.5 4055.2 2178,7

NOMINAL WEIGHT OF LOX IN TANK, POUNDS 161.97 159.81 85.86

(DENSITY- 68.1 POUNDS / CUBIC FOOT)

Figure 11. Tankc Geometry, Showing Nominal Volmnles and Weights

R-5123 31

x OC ETIjr) qWEA DIVISION OF NORTH AMERICAN AVIATION. INC.

near the lower end of the tube permit venting from the approximate center

of the tank, When the level of liquid oxygen reaches the top of the

uppermost holes, approximately 86 pounds of oxidizer is contained, in the

tank.

UIQUID OXYGEN WEIGHT DETER2EINATION,

FEST ARRANGEMENT' AND PROCEDURES

To determine the effect of the long ullage fitting, tests were conducted

at CTL.-l using the schematic and vent configurations shown: in Fig, 1.2.

Figures 1.3 through 15 show the test setup, The arrangement generally

simulated the missile oxidizer start tank fill and vent configuration,

except that the missile vent line is shorter and less restricted. The

first tests were conducted using vent configuration A, However, liquid

oxygen was sprayed. over an asphalt apron located to the rear of the test

cell, creating a safety hazard, The configuration had to be abandoned,

Configuration B was safe but differed considerably from the missile con-

figuration and delayed tank filiing significantly, Vent configuration

C was adopted as a safe, practical approach, although vent line resistance

was somewhat greater than that of the missile vent, Table 2 shows the

correlation between vent configuration, ullage fitting, and tank test

number,

The weight of oxidizer in the tank was determined by'weighing the tank

with its contents and subtracting the TARE weight, An. effort was made

to avoid erroneous weight readings from artificial restraints on the tank

by the installation of flexible sections in all tank inlet and. outlet

lines,

32 R-5123

FORM 608 B PLATE REV. 1-58

LOAD CELL

PNEUMATICMANI FOLD(551747)

P4

(30623 I, SHORT)( (304411, LONG) TANK

C

(306228)

BLEED VALVE---

FILL AND CHECK VALVE(301040) B VL; $

FITTING CHI(300223) (Ni

R-5123U,

0I,,

A DIVISION OF NORTH AMERICAN AVIATION. I

LOAD CELL

VENT LINE CONFIGURATION

(A) 3/4 INCH DIAMETER, 15 FOOT LENGTH

V VENT LINE (B) 3/4 INCH DIAMETER, 15 FOOT LENGTH;I 1/2 INCH DIAMETER, 100 FOOT LENGTH

(C) 3/4 INCH DIAMETER, 5 FOOT LENGTH

11/2 INCH DIAMETER, 5 FOOT LENGTH

(D) MISSILE CONFIGURATION,

3/4 INCH DIAMETER, 14 FOOT LENGTH;LOW PRESSURE I INCH DIAMETER, 78 INCH LENGTHRELIEF VALVE(9512-45049)

'VENT AND RELIEF VALVE INSTRUMENTATION(304327) PARAMETER RANGE

LOX WEIGHT, POUNDS 0 TO 200

-CHECK VALVE SUPPLY PRESSURE (PI), PSIG 0 TO 200(NA5-26032) TANK INLET PRESSURE (P2 ), PSIG 0 TO 1000

I- HELIUM PRESSURE(P 3),PSIG 0 TOIO00

SHUTOFF VALVE PNEUMATIC REGULATOR OUTLET

FACILITY PRESSURE (P 4),PSIG 0 TO 1000L _LOX SUPPLY TEMPERATURE (T3 ),F -325 TO + 175SUPPLY LOX TANK INLET TEMPERATURE(TI),F -300T0 -250

VENT TEMPERATURE (T2 ),F -325 TO + 175

CHECK VALVE(NA5-26159) Figure 12. Test Schematic

X;.4b -- xU R -V I=A DIVISION OF NORTH AMERICAN AVIATION. INC

rV

1262-12/4/62-SIB

Figure 13. View of Long Ullage Fitting Test Setup,Showing Oxidizer Start Tank

34 R-5123

IFCP - MCIE K,"IsUIW6 REA DIVISION OF NORTH AMERICAN AVIATION. INC

1262-12/4/62-Sic

Figure 14. View of Long IUilage Fitting Test Setup,Showing Oxidizer Start Tank

R1-5123 35

A DIVISION O6 NORTH' AMERICAN AVIATION. IN.

TABLE 2

ULLAGE FITTING AND VENT CONFIGURATION

TEST HISTORY, CTL-1, 1962

Ullage Fitting Vent Configuration Test Numbers

Short A 711-1713

711-1714

711-1715

711-1716

B 711-1751

711-1752

(• 711-1753

711-1754

Long B 711-1756

711-1757

711-1758

C 711-1782

711-1789

711-1790

R-5123 37

FORM 608-3 PLATE REV. t-30


Instrumentation was as indicated in Fig. 11; data was recorded by direct-

inking graphic recorders, Pressure measurements, P1 and P4 2 were useful

during the tests primarily to provide tlhe desired. supply and. start tank

pressures. The temperature measurements *were useful as a means of

determining whether or not oxidizer weight had stabilized at a given

pressure, i,'eo, the three temperatures were relatively stable and in

close agreement when tank weight was stable. In the analysis below,

tank weight and tank inlet pressure are the most important measurements,

In general, the test procedure was as follows:

1. Purge the tank with helium at 600 psig for approximately 1

minute to provide the tank atmosphere anticipated in the

missile.

2. Introduce LOX, under tank. head. pressure (10 to 15 psig) and

chill the lines and tank for 5 minutes.

3. Increase the inlet pressure (P1 in Fig. 11) to 25 to 30 psig

and permit the tank to fill.

4o After weight stabilization, increase the fill pressure to 55 to

60 psig and hold for 90 seconds or until weight stabilization,

thus simulating missile LOX tanki flight pressurization.

5. Pressurize the tank to 600 psig°

6. After 15 to 20 seconds at 600 psig, close the facility LOX

inlet valve and open the bleed valve (to simulate the missile

firing sequence, in which there is a slight LOX loss to the

booster engine after start tank pressurization and before

engine start).

38 R-5123

FORM 808.B PLATE REV. 1.58

A C V I5( N OF I l T 0 AIA R14 A 1 4 VWA"I )•. IN M

7. Permit 15 to ( iou ds f LOX ti itcape from the tank, then close

the bleed val -,

8. Vent the tank

Deviations from this p ic di: e r iadirij ua tests are described in the

test log (Table 3).

TEST RESULTS

Of the 14 tests conduc ; d, i i w re val id -it was evident from the tests

that, at a fill pressu', )f !5 o '0 Isig, the relatively restricted vent

line configuration B d,?.a,7d in iiil iank filling, and the use of the long

ullage tube prevented ,'wrle le ,an1 fillLng with vent configuration B or

C. At a fill pressure Jf O to 60 psig, there was no clear difference be-

tween oxidizer weights -ea(hod uLtt v(wnt configuration B and vent configura-

tion C (Fig. 16 and 17) Theý e,,id nce in these tests does not indicate

that the vent line has any sagn.fi ani effect on tank filling, except

that a relatively restrlctcd vetu, in(, delays filling the tank to the

level of the holes in the ulLago tbe, Therefore, it ie presumed that if

the tank, equipped with a oong iLlage tube, can be filled during component

tests with vent systems somewhai. different from the missile vent line the

tank could be filled in the misnil(ý.

The tank filled completely in testm 711-1782 and 711-1789, with fill pres-

sures of 80 and 75 psig, and hoLd times of 9 and 6-1/2 minutes,

respectively. In each of the six tests with the long ullage fitting, the

tank filled above the ullage tube holes at 50 to 60 psig facility inlet

pressure, presumably as a result of" ullage compression. Therefore,

R-5123 39

xN,10 4- Aw- ]EE NWr " 1r wAA DIVISION OF NORTH AMERICAN AVIATION. INC

IL wd

NZ c,

U) W D0)0

wU !K 00L (0, (D

00

0 §

0z 0 -4

- ___ =.*J- 10- P

NH A 0- CIS

(I) N0)> >OH

0) 0)U) Cw LL ;-,r-

w 0

__- N_ ~ 3

_ ILd

HC

0 _H

(n R-5123d

180

160 TEST 711-1782

140

z 120

00100

s..o

x

-1-

o 60_j

40

20 -.

0 5 10 15( TIME

18 0 ------ --- -----

160 -FACILITY~i7 INLET

VALVECLOSED BLEED VALVE V

S140 -CLOSED 01

10~

3: 1200w

x

80

6056* TIME F

R

I.

CO CK IET NEA DIVISION OF NORTH AMERICAN AVIATION. INC.

711- 1782

_ II/

____ SEE EXPANDED- ___--SCALE ELOW

TANK INLET PRESSURE

25PSIG-~- 50 PSIGf5TO 8O0 PSIG..]

10 15 20 25 30 .35 40 45 50 55 60TIME FROM INITIATION OF TEST, MINUTES i

EXPANDED SCALE

VALVE VENT __OPENED LOX LEVEL REACHES

HOLES IN ULLAGE FITTING

Figure 17. Weight of OxidizerStart Tank vs Time FromInitiation of Test, Ven

S F E Configuration CSSEE FIGURE 20 - -w

57 58

TIME FROM INITIATION OF TEST, MINUTES


TABLE 3

CTL-l TESTS, 20 NOVEMBER 1962 to 19 DECEMBER 1962

Test No. Comments

711-1713 Facility and system checkout, invalid test

711-1714 Bleed valve open at tank pressurization, invalid test

711-1715 LOX weight = 167.5 pounds, regulator out pressure (P 4 ) pickuppoint changed

711-1716 LOX weight = 182 pounds, presumed to be instrumentation error,invalid test

711-1751 Temperature bulb installed at T1, replacing thermocouple,vent line changed, bleed 50 pounds twice to determine the effectof rigid flex lines on recorded weight; no significant effect

711-1752 Inlet pressure varied in steps from 40 to 60 to 80 psig, 50-pound bleed to determine effect on recorded weight, no ex-planation for the extreme weight recorded in test 711-1716

711-1753 No pretest helium purge, no significant effect due toomission of the helium purge

711-1754 No pretest helium purge, no significant effect due toomission of the helium purge

711-1756 First test with long ullage plug, weight increase from 0to 80 pounds in 36 minutes, stable at 48 pounds for anadditional 12 minutes at 25 to 30 psig inlet pressure,weight increase to 105 pounds in 52 seconds at 55 to 60 psig,stable at 85 pounds for 20 minutes at 55 to 60 psig

711-1757 Similar to 711-1756

711-1758 See Fig. 12, fill to approximately 80 pounds in 50 minutes at25 to 30 psig, fill to 121 pounds in an additional 52 secondsat 50 psig, pressurize tank to 626 psig and vent, refill thetank to 131 pounds in 3.5 minutes ("stable" weight = 102pounds after an additional 12.5 minutes), pressurize the tankand vent, final weight after ventihg = 76 to 77 pounds in twotrials

42 R-5123

A OIVISION OV NORTH AMERICAN AVIATION. INC.

TABLE 3

(Continued)

Test No. Comments

711-1782 See Fig. 13, vent line changed, flex line straightened

711-1789 Fill to 93 pounds in 20 minutes at 25 to 30 psig tank inletpressure, fill to 146 pounds in next 9 minutes at 50 psigtarn inlet pressure, fill to 170 pounds in the next 6minutes at 100 psig facility supply pressure (75 psig tankinlet pressure), unexplained weight loss at 609 psig tankpressure, invalid test

711-1790 Fill to 63 to 82 pounds in 26 minutes at 24 to 32 psig,fill to 109 pounds in 66 seconds at 56 psig, bleed 18 pounds,pressurize, vent, weight after venting = 71 pounds

)

R-5123 43

FORM GOII.'PLATE REV. I-3

A DIVISION OF NORTH AMERICAN AVIATION. INc.

it seems unlikely that the missile oxidizer start tank will be filled complete-

ly, 'but it may be filled somewhat above the holes in the end of the

long ullage fitting. The tank inlet pressure in the missile ranges from

55 to 60 psig maximum; this pressure is not expected to be held long

enough to permit tank filling (Fig. 16 and 17).

During test 711-1789, an unexplained oxidizer weight loss of 33 pounds

occurred with the vent, bleed, and facility inlet valves closed. This,

plus the 23 pounds intentionally bled from the tank, left only 11.5 pounds

in the tank just prior to opening of the vent valve. After venting, the

oxidizer weight dropped to 90 pounds in 10 seconds and to 78 pounds in

an additional 40 seconds. It is suspected that the weight loss before

venting was due to incomplete closure of the bleed valve,

in test 711-1782, the weight loss after venting was 53 pounds in the

first 16 seconds (152 pounds to 99 pounds) and an additional 22 pounds in

the next 76 seconds, making a total weight loss of 75 pounds in 92

seconds (Fig. 17)o

The weight-time history of the oxidizer tank during venting 'was described

analytically by considering a polytropic expansion (PV = constant) of

the ullage gas. The exponent, n, was determined empirically by relating

the initial conditions to the point at which the ullage gas begins to

vent through the ullage plug, This final point is discernible by

inspection of the pressure-time history of the tank, A relationship

between tank weight and. ullage pressure at any instant 'was derived as

follows:

1V i Pc o

0 1

44 R-5123

FORM G03.B PLATE REV, 1-58

IE< 4OC: f<IE:-t- -'1-1-IE3

A DIVISION OF NORTH AKE5RICAN AVIATION. INC.

1

n

2, Vi = (V )

1 0 0oP

3P = W =lag pressurev

11

1

n =w = ltrpv[Q expn]n

1

w¢here

W = tank weight

p =oxidizer density

V = ullage volumne

P =ullage pressure

n =polytropic exponent

Subscripts:

o = initial condition (start of venting)

)' i = value at any time

R-5123 45


This relationship is plotted, for test 711-1782,. in Fig. 18 for comparison

with actual test data. The correlation between the two curves is considered

satisfactory.

It can be concluded that if the tank fills, the oxidizer above the ullage

tube holes will be expelled through the vent line until the liquid

level reaches the holes in the ullage tube. Oxidizer will continue to

be expelled at a slower rate until the liquid level stabilizes somewhat

below the ullage tube holes.

The weight-time history of the oxidizer tank will vary as a result of

the level of initial filling and the tank vent line configuration. The

assumption that the tank fills completely and that a weight loss of

50 pounds will occur in 16 seconds with an additional loss of 20 pounds

by 92 seconds will yield conservative weight estimates for missile

trajectory analysis.

46 R-5123

FORM 608-B PLATE REV. 1.58


0 toI.-i

IN

III

_~ _ __P

p

0 U

CD or- )

H ~(D 0

Li Li z

b-0

w H(rLi~ Lii

<~ ~w CIS2cn0

Li- HX Ld 02

> -pLi

-ii i

zcr:

z W 0__XN

0 0

ED Lo 4 N - 8 00 O

SO]NnOd 'iH9I3M XO1l

I.0 0 0 0 0 0 0 00 0 0 0 0 0 0fl- w to q* to Nj -

!DISd '3aflS3W JL3lNI )4NVIR-5123 4


REFERENCES

1. Special Report No. 50, Atmospheric Heat Transfer to Vertical Tanks

Filled With Liquid Oxygen, Arthur D. Little, Inc., I November 1958.

j)

(_

R-5123 49

FORM 608BS PLATE REV, 1.58

A DIVISION Of' NORTH AMERICAN AVIATION. IN.

c

APPENDIX A

MATHEMATICAL MODEL, LISTING OF FORMULAS

LIST OF FORMULAS, CHRONOLOGICAL

The formulas for the mathematical model presented chronologically are:

1 Heat Input to Oxidizer

QL q T ATL t = Btu

2, Increase in Oxidizer Temperature

QLTL2= TL1 + WLL1 CL =F

3. Heat Transferred From Ullage to Vaporize Oxidizer

U AOL (Tul - TLI) tQuv = 3= Btu~uv 3600

4. Heat Input for Vaporizing Oxidizer

Qv = qT ATu t = Btu

5. Specific Heat of Ullage Mixture

C - wul C Pill + W=v Cpvl Btu/lb-Fpu2 W ul + Wv1.

R-5123 51

A DIVISION Or NORTH AMERICAN AVIATION. INC.

6. Ullage Temperature Change

T U T -l Q11V Fu2 ul WTu CPU

7. Weight of Oxidizer Vaporized

Wv2 = = poundCL( '-TLi)+ X+CP(Tu2 Ts)

V

8. Ullage Gas Constant

(Wu l + w 1 ) R 1. + v2 RvRu2 - ul + Wv + V2 f t /

'R

9. Ullage Volume

(. W.Ul + U2 + W2) z 2 Tu2V c cu ft

10, Volume of Oxidizer Vaporized

Wv2VLv - _ cuft

PL2

11. Volume of Oxidizer Remaining in Tank

VL1 DL2

VL - V Lv= cuftPLI

52 R-5123

FORM 608 B PLATE REV. 1-58

A DIVISION OF NORTH AMERICAN AVIATION. INC-

12. Weight of Oxidizer Expelled From Tank

WLE = (vL + v 2 -VT) P 2 = pound

13. Total Weight of Oxidizer Lost for Vernier Solo Operation

WTOT = w + W 2

14, Final Oxidizer Volume

SL2 = VT - Vu2 = cu ft

R-5123 53

A DIVISION Oil NORTH AMERICAN AVIATION. INC.

NOt4ENCIATMR

A = surface area, sq ft

C = heat capacity, Btu/lb-F

C = specific heat (constant pressure), Btu/lb-Fp

q = heat flux, Btu/sec-sq ft

Q = heat input, Btuft - lb

R = specific gas constant, lb llb -R

T = temperature F, R

t = time, second

V = volume, cu ft

W = weight, pound

p = liquid density, lb/cu ft

X = latent heat of vaporization

Subscripts

1 = initial value

2 = final value

L = liquid propellant

s = saturated

T = tank

TOT = total

u = ullage

v = gasified propellant

E = expelled

54 R-5123

FORM 608 S PLATE REV. 1-58

A DIVISfON OF NORTH AMERICAN AVIATION. INC.

APPENDIX B

DIGITAL COMPUTER PROqlRAM

A digital computer program was written to expedite the calculation of

propellant loss and bulk density change using the method of Appendix A.

An important requirement of the program is the approximation of initial

conditions existing in the tank, which are input values. The closer

the approximation to actual conditions, the more accurate will be the

results. From initial conditions the program will compute and plot,

if a CRT plot is available, the following parameters vs time:

1. Propellant Weight

2. Propellant Temperature

3. Propellant Density

4. Propellant Volume

5. Weight of Propellant Vaporized

6. Weight of Propellant Expelled From Tank

7. Total Weight of Propellant Unavailable

8. Ullage Temperature

9. Ullage Volume

A program flow diagram is shown in Fig. 19. A typical input data sheet

is provided on page 62. The case shown uses a standard setup with the

data arranged chronologically. The data location numbers indicated areK assigned automatically, by the program. To run additional cases, it is

R-5123 55

1 [:ý. NC KE3 -T "DY 1744 !E


Set Up Program Subroutine FLIN

Read and Write -p Determine Initial I Determine the

Title Density Given an Density for This

Wi Initial Temperature Time Using the

Read d Write as Input Data I Last Temperature

Input Data Calculated

Call FLIN Subroutine CALC Subroutine FLOSS

Initialize Times - Calculate Propellant Calcul~te

and Subscripts Height, Liquid Tank Temperature

Surface Area, and for This Time

Call CALC Gas to Liquid

Interface Surface Call FLIN

Test to See if Area

Final Time has Call FLOSS Cal

Been Reached 4KContinue w Propellanto Loss for

S esThis Time

Increment

Time and

Subscripts Subroutine PRINT Subrutine CR PL

Call PRINT P Print Out Answers -1Plot

For This Case Parameters

Return to at Selected

Read Data Test to See if Times

for Next CR PIOT is Desired

Case. If no No Yes

More Data, TJob will be Call CR PLOT-

Automatically

Terminated

Figure 19. Digital Computer Program Flow Diagram

56 R-5123


necessary only to indicate the desired changes. Pages 58 through 61 are

a sample of the required data for running an additional case. In this

example it is desired to change the ullage plug length, the total heat

flux, and the time from booster separation to sustainer cutoff. The

case number is automatically changed to the next higher number for each

case run. Therefore, only the number of the first case to be run is

required. A minus sign must be placed in the first column of the last

data card as an end-of-case flag. Each piece of data is limited to

eight digits plus a decimal point and a positive or negative sign. The

decimal point may be located anywhere in a number, but must be included

for the program to work properly.

Pages 62 through 68 are examples of the printout produced by the program.

Page 62 presents input data, and pages 63 through 68 are program results.

Figure 20 through 28 are samples of the CRT plots obtained.

The program is equipped with preset limits which will economize computer

time. The case being computed will be terminated automatically:

1. If the number of iterations to find the propellant level in the

tank for one time interval will exceed 75

2. If the initial propellant volume is inadvertently chosen larger

than the volume of the tank

3. If the propellant density falls without the table values supplied

4. If the number of time intervals exceeds 300

5. If the propellant level falls below the top of the bafflesR

C

R-5123 57

P41~C1Uy

1144

4D)- -fa 0

cn- )~P*4, PC 04

Uq I H 0 1- t-4 4

u 44 4

cq IV- $4 -0 - 0 04;4 64.

0. 0-4 -44 - .44-) 4)-- + 4) 4

L0 p4 mU 04 0 93 -N,.~C 0 ;.4 C 4 3

0~ 0 4- - 4) U1 Q U & (U--4PQ QU r. 0 Cd ;4 m4 Cd WDC "0 4.- C ) ) *4~6.

0.. rd 0 to U) .4 a ~ , C3 *.,4 .,1 * 4 ~ C.. H 0 04 Cd 04 Cd 0 & S4 U) p) 0 0

- W W - r4 0 >. ; 4 0 U) N bD 0 04 (a-U4) m 4)j 4)4 4Puw)0 4) (U P40 4A w Cd 0 44 --4 42 3

4.1 CU $U M4 F4. 0 CH 4) 0 CdC bl) CH. MW(D Q.. )U~~) a4 kH M -. 4 F-4 4) 41U 0 4 )D 0 P-1o H4 04 r-4 H) r-44)4)

CU*4 )C .) w ) -4 M-C U v4-4 w 04 k 04 14 11Cd p cUod H UJ4P 4) .CU M 04 0 a)0 04 -H J-

0 M ~ $4 w)4 =~ .m 14 M $ 9044 9. 040 rq-P 0 H;

644 04- -4 Ud2 144 r4r1 1 -4 ) u-P e- 4 ca4 CU (U r4 4-) (404 cd CU4) )

4-4 r. d Cd 0 14 CU 4+.4 P-4 -4 -, 4 1 CO .0 '14 -P4 b.0- Id y

04 N* 4-,4 -,4 *-A -A. ) *-A -A- v-I O4 Mv r0 4) M. 4)4 Q) C:,I 0 -4 0H3 U

;p) 04 ) 00U d 4q k4 4 0 -4u - 0 bi ;it -H. 04 04 U) U) HvI~- *-a + ) a' * - - )-

U~~ ~ 0 ft -- co. - ' ~ U f ~0 4 NI C to,% .~1f 0 %Z0 N 00 C P-4 0 P-4 P-4 .-I vI r-4 0 -4 P-4 " P- 4 1

o cd CU CU CU cl CU CU eU cd CU dU CU ed cU CU c d CU CU U C U C-4-4)4)P)+) 4)1 +4)4) 414 4 4)P4 4) : 4 4), + 4)434 4P 4

0 0w 0 -4

U. J4-

.i

z .. .... .... . ..

ii: xm :

I'- .............. 0

ctr-

w4m 000 -4 1r :VU n'

Z~- W)O -4, till N P~~-0 a4 'I*-4

0 Z ItcI. Q0 --

-4~ -- Nl

58 R-5123

14

001-4 434 0

4) Cd

41- 0 tS

a) m o@ c0 P.-444)-

0) 9 .@3 CD. @4. @4) w44) c 0 .. 4 2

ci 04 @3 f+ CU 0- 0@3 kC. 0 .0 Cd co

Z4 0 -4 r- C124 H L 4- 4-) 0

4.)C L CU4 0&-4

.0L CdC 4 0 k 0 10-001- 04 T - C)-4 4 0 00 U)

C\ 4-1 '0 r-4 4) @3 Z@q3c ) +ý > 9 ->- oCU0U)" 0 4) 0d -k -

C .) 0 _ +) 0 0 4.) 4 0.-

0 -4-) 41 > CC- ) - -4 0H 4) C

ocU 0@C a*. U0 -4 0 0 4

_ 040 e4 CH4) U) +) 4)-. H 4) ) 4 -

4-) V0) - d - 0 0)

.4 U)U -4 0 4'

4-01 4) 1 0 0

40 43 0 0w a. C4) 0-, 0 4)

CU CU CU 'U CU CU WU C U C U C U C CCU CU CU MU CU r CU CUw 4-)l 4)4P)4 4-) 4) 4)4)4 4) 44 ) 4), 4)4)4

w* C0 coc u- o 0 1W Cd

IL.. ... ..

z .........................*r :: :.: :- :. *:. .:~ : '~

04 0- C-4 Vt~N

D 00 O t^, tEr) -4 (%. - 1- go

.4 ~ ~ C ON 0) -* .0 *o _:v *'\ O C 4 C'-4 Q

o * * -4-I * m If .1C- * 'D0 Lm.'. . .0"0 '0 t'

LS L!ý N1 011LD - M OU L I r L-Ji

R-5123 59fI i

- - ,- -

00

PWI

co 1-4

4) -.- cc .04

0 d l 0 04 - 0M, ., 14.4 0204 0 4

Id 0) 4)0ý4) 4

*rd 4) a

X2 4) 0) 14 *F4 Q0- 0w 4 0) F-4 4) 4)H) 0)ý 44 d) c a m)

0 '1+) ' 0 P4 sd "d =0. D)0. 0 a 4 W)

4)~~ ca- CH 1- 44O - 3 4-4 0D~ 0 s ~ +143 0 bb,*eI 4) @

0~-t 1-4 4) d2W 4q + 14,4 4. - 0 ..0 biD 9 1.y 4)4 2-P - r U 3I*0 P4 bDCdP4P-

q ) 4) 0) 0 0 bfl 4)0 'r0 H -p

0 0 'I 0 94 4) ;4 00)

14o-d &CH 0 0 01-40 , 0 P4 - - H C - -

$4 w2 cdm '.q 0) 4) w0 w 0 a 4)14-4 bO4 -4 s4 f b D) U 10 MCH cd% ca C m 0d0 0 P CZ

14 $4 >

0 W Q 4-) -P.0 r 4 . 0 4) 4) 0 __ w~-~ 0 -'

V 1- CUj FIN -t VIN v 00 a% 0 92 P-4 ol n :r nr 0 C4 ~ ~o3 C) 0 -41 ~ - - ~ . - T .. ult 0 P - P -4 r- F -A - 0 -4 -4 i-4

000 0 0 0 co 0 0 0 0 4)0 d 00 0s 00 Cc 0d co

+ ) )4 ) &0 4) +2 4) 4 4) ) 4 4)1)4 4 4) 4- +) ) 4)Dcco0 co 0a a 0 d 0 a 00 0d 11, 0d 0d 0d co

... ... ... .... . *.. .. ..

... .. . . X0 ,, ..............- ~H - .*..*.*...........

ILL

P--4

Ii. i

imr- Z

z Zl 00 W' c %0 CU ý- 00 cu 0 - -

00 r- r- r- .o I'D -m r4 ,- u- CUwU

- ~ N -m cu Icu 0iC N cm

-Zi~~~ ~ ~ ~ ~ n- MNW-L:JU cjuiIL :

60 R1-5123

4)4

z o4

CL0 Sbb 0

0 ;4 44 P-

0 -P

914)

hl t

.H 44P *t-4

0D 9~ 4- m= 0 0

w a. 0 11 00

od -o 0d C d Cp40 4j 4 4) 4P + ) u-P -4.3 43 -P 4-)

"-4 Q)' Q p) 4 4 mo 90 l. ' 441

.. .. ... .. ... .5 I)

... .. . . . .

L rU 0

0L N

PU-OI-...

D...,.S.......*..*' *. '**' .* . ,Z,** .** , * ,.,

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

Z~~~L LJ-ý . .'J M-.****.***..,*.

R.12 61 . ,......*** .,,*, . *

0 0 00 0 0 0 0 0 0 0

o o - M0 '- ") 0 0% 0 0o 0 0ý n- 0 -t 0 0 0

* 00 0 0 - 0o 0% 0 00, 0% 0 05 co .0 6 0 .0

1n i 0 Un 10 en

o 0 0 0 0 P-. 0 0 0 am) 0 0 to 0 un 0 0 0 0

0'0 0 in 0 .0 0 0 0'- 0 4n m) 0 f-.- Iý 0 0N0 N 0 0 0' w 0 0

* 0 w 0 0 0 0 *o 0on a' .- ý * ' 0 .0 0 0*~o 0 n0 '0 0t 0

a- 0. 0 '0 07 0 N) 0 0t 0 0o~ 0A C; I'# 0 0 0 0 ,0 0

0: In ' ' 0 ' 0 0 0 00! Oj 0 '0N 0 0 0 o 0a 0

0.- 0 0 o 0 0% 0l 0) 0 0 0 0

rn, 9'- 0 - co0 10 0'U) t 0 N 0Nn0 o t

0o U, 0 0 0 10 0 0 .

X, 0 O0N) 0 0D 0 .0Z) I 0 t 0 CV- .0 00 0 0 t0

Liirf U) N 0 4 0 .04 N

en 01 0 0 0 a .0 0 0 ,* 0~- 0 CO 0 0' 0n ' 0 * '(0 '0C) 0 0 o 0 a0 0 ;0- U

000 a 0 0n 0 0~ 0) 0 00 0 0 0 U,0 0 0 0% 0

04 0ý 0: 0A (14 0NN N0o M' 00 00 .0 N 0 10

ýN N4

"'0 '0 '0 '0C, !-N N tN) :Ot4

62 ih5

Wj .0 0 m% '0 in O* pn m % m o in 00 * fp% m- *P M> N - 0 m, P-. P- N (Y4 1%- 10 LA * - n M) tn l0 '0W * N Pel % 0 00 %0 a N M% fn' rn' O% 0 C)% N 0 M'.j N 0% 0 - p- p- p 0 co ?.- Ln N 40 .0 mf w m m% N

in 0 co~ %a CY C3 co &A Mn I) P. A 0% 0' 'I- * NV co CU, co 0: 'A.~ P-' . 0 ' n i i n * i

P- - 6 - V- r- 0- P-: V- v- 0-

4 IwU tA 0 %0 Lm p - Ok- *- N) 0 0% ON P.* -. c0 I- 4oe u' N 0 in~ U'1 0k r-- w 0 * 10% fn CO 0 v.. CO t- 47% UI< f- o% .0 Nl U, m) N %0 ID 01 0 0%. U,0 i 0 0' t.. 0- CIO

'0 CO U) N , 00 n OD U ý 0.0 W) 00.4 0 a0 - -a a- '0 in 0 r.- w~ w 47 0 0 - o, N in tn') a n in o0 '0 t- r.-

x 0 0 0 0 0 1% 0 M 0 'L S. 0 N n fS Sf 0% 0:: IA 0 M 0 0 0 %0 0 -0 %ol 6- 6- 1-6- M - 6- M- N- 6-

0 I m 4-P )

I- 4 Nc 6- 1%- P . - l0 r, M' %0 1%.-; %0 M' C) CV f.- in 1.-wL 0% Il- in ui t.- P- CID to - 0 in Iln NJ 1'0 N 0 N .0 inIA w 0. 0 S a S 0 * * 0

0 1- co 0 N J r,- a, N U; C; N4 I0 C; * G; *0 0U -d OD 0 C4 M( in %o co 0% o'N m) in~ %o P.- 0%0 N -T in

- 6 .-6- w N N ~ N N)Nm N') pn

0% 6- P- 40 in N -0 .0 j% M% * 0 0 1- 1"- on .0 ?%- 0"%w I-- V) 0 t * y l% (0)% 0% 0 in 00 pn in 0 -.1 Nl m)D P. N in OD rQn W ) - N"' 0 0 N N - V 1% n (N MNLM' in 10 10 in M' 0 t0 P- J * 0 r- 0 '- *T C% in 0 N 0

1,- 0% 9 - 0% Ln U, 0 0% .0 M' o' 0 v * n -I 0) N P-

in o' N% r-00'P- %o in in) N - a% op %o'0 in.N N 0 91X6 - 0 0 0 a 0 0 0 tyk 011 0% 0% 0% 0% 0% w

o - 6 - -6 6 - 6

It-5123 6

co Ln %0 0 C3 NY Ok f- 0 m% 10 * %m. C3 --t N N -t 1- N N 1-- -1 * 0 0%.0 0) N 0 in Ln it m 0) N- N~ 10 6. 0%

6 t - 4' mV ul Y m% 10 N m% LA t -

OD I- N) * 0 0 0 0% 0% 0 0 0Va tn. ~ 00 N0 P- -M in - M C)j 0 6%)

-0 !t tn -t 0 % NV - r-- (V LA cV uLA LA"e. '0 0 -t wV 0. MA 0 N 0 in 10 10

ao 0o 07 0% 0 0 S 0

-~ C-4 N N (N N% N N

0 Nj (V) ff) Mi" * N 0 N~ '-0 zt co LA rfrn I~ LA N- LA m~ -D t *- !N 0 N 0 LA

N~ tn, rV) Zo (%0 N o~ N 1- co N- in (3.0 N 1 (% '0 o 0% 10 0 LA Nr- ". 'o 0 N

0 1 0 ' o) ON-. CA r- *1 0 *0 :0 0 ,- 42 . %

-- 0 1 0 (V 0 (% 'o '0 OV rOni 0A tn * * L LA LAn LA LA -0 .0

6% (V CN 10 .0 -t 0 N,- 1 0 Nj tn 10_:r N! 0 CC) 0 N 6% ) LA M t t LA 0f'- J, '0 N,- In N- LA C7% N- ODV N .40,- LAI V ao co M V O 1.- N- N 0 10 rw N- LA10 f- r Nr- .0 0% 00 0 N- r*- 0'% t 07% LA

-- '-~ ; C4 :; VA z0 r- Z LA r" P- - rf)l-' LA), *t N 0 OD .0 -t NV 0 ao '0 tn) p.

(0 OD; (V (V co N- r- N-- N- Nl- .0 '0 '0 '0

ULA o LA o LA 0 LAI 0 LA 0 LA 0 LA 0o - ' N N K) (4) :t * LA LA '0 '0 N-

64 R-5123

10 N Lfl 00 0 w-N 0 0 a0 NY (7 O~. V * N~~ C0 *0 M 0. 0 00 it * ( 0 (O - .0 %0 irO> 40 .0 a- 4-. i0 00 L* Sn l m0 m~ 0 N &

< 0 N " (V C) .CO *ý LM m i N N N N N m~ 4 * A

1 .0 0 07 M PM P. 0 C1 W)0 0 0 01 0 O N t- 0 Cm N N a o N N aN N N N Ny N M\

.0 M~ 04 in V- kA 0 N0 1~( - N~ 0 (V N% 00 CY 0- fV. 0 it v4 q co- -6 th m 64 v-0 0 0 Sn (7 0 0~ N c

0. ( v S in N i it Ln 0o p- co - N _ C0 N. 0' -It N I P

V.: 0-j

3C

x 0% N) .0 m N M, m'. DN'N 01 0. .0 0 - 0 0% 0,D > pf5 %0 0' 0 1- tn, tn ýp c ,' in wV Sn _4 in f- t - *T -

z ( el- (N oo o q% in v- - Fni 0 %o pm, 0 V- * N 0N 5'- S)( mi * vi tn %0 %0 P-' CD 00 Oh 0 0 V- N N -- 4t SA .0

W u 0 0 a 0 0 ol 0 0 0l V_ p i V, r- V_ V_ V-,-

0 0 0 0 0 0 0 0 0 0 0 0 0 0 0

5- 5- *ý Ni -. 0 M- ( Q fl- CX) Mi 0 r__ P.- * N 0 -1 Nco r- - (Nj P- mi u, - LA T Mi N Sa n N 0% 10 z0 tn 0

CA wi co "I 5'- N %o %0 o0 o0 0 n 0 a,' 0

w wA wi co 0 - I%- .0 NO Ni. ?I (V z C- LA LA ' -

-0 %0 10 %a %0 .0 %a %0 %0 -0 10 .0 v0 10 10 '0 '0 .0 '0 .0

wtn. o LA o tn 0 Sn 0 in o, Sn 0 Sn 0 Sn 0 0 tn: 0

R-5123 N iN nL 4 . V651

f -. en CD ~ N ul -- 4 o mO 0 In Nwo wO CO n 0 r- * N 10 t-.- In m% '0

,a 10 0 P'-- 1- 0 ?L- 00 .4 tý- in N n rOD CD -t ' C% mo 10 a In 05 0 Ch m% t-

0'~ N mo a t0 - ds 0 N m~ ii 10 MO

V, 10 m% f- wO * n iA 0 oo N~ wO f- ut .N 4t 0% e- t.- w C 0% t r-- m) (7 I% Nll N 4t '0 00 m- ro 0 0 -t 0', 4 Y 0%OD 0% 0 ~- N 4 In o0 CZ 0% 0 N~ po U)

f4 O o to t t4 14 r4 .?4 4 4 *

C' CO (7, f-~ nO 0ý f- 0 int 0 '0 '0 rt-- . OD CO 1- 0 in co *r 4t (% OD tr 4t

01 N., 0 0%, 0% wO wO w 7. 0 m tO 0 0%r-- .co 0% (7 0 .- N fn *t 0o I' o 1O ' 0

*% -% .- N N N N N% N% N% *N~ to

o; 0' 0; 000 0 0; 0 0 ; 0 0 0

Mo to 0 *t V) co kn N _- ch '0 *t * 0fqO '0 CO) In (IN CO - in I- in a, 0 K) (NN El' 0% CO In e- ' CID co -r *V CO COD0' rf) I- r- n a, N~ tn COD . l In LA I09-4 tO K - 0 CO '0 In mo o 0 CO %0 -T 4N

L4 4 U; S 0 a 0 0 0 0 r0 0

0 o %0 0o o .0 .0 0 10 0 o 10 10 40

Ci4 0 mn 0fn n 01 In 0 in 0 iC t0

66 R1-5123

D O 0 N In (I 'o cO 0 %4 * M0 0% N U40 C0 t- (7 N in -a U N m% in ' N a m0 0 '0 a0 0,

_j No a CV tn r- 0~ fn fl 0 0 w mo N co Mo in ý tO 1 0 0M

m Fn it 6r 0n 6n 0 0 0 0 .0 Ohl-

Co %0 N N0 0% C ' 1 0 o N 0 C 0 *r4J 0) ON m n N o N m w Co ()% 0- ' 0 0% oC7 M- it N 10 0 N W * n i 0 :t 9. 0) C14

W LN o N0 '0M m0 P- 10 'o0 U N N o C) '0 tN N6 L

I- 1- ,- m on w0 m P-0 t N 4 N 10 OD 0 % iN .o'o 0 0% N %0 *Y M 1in Ný rn pn 00 -t* NW in I- m0 t- w% t 0- 0 N y 0 w) m P.- 0% N, 0 0CD Pn m% Co Co w 4 w N N '0 0 0 . '0 .0% iN in n ino i04 0 0 0O m 0 n co o 0 y 0 n 0 ý- 0 % %o a, rn *

C-

w - N4 i - i 0ý in %0 0% %0 %a %0 to o ao %0,c in N4 04 C N N ) 0, *Y 04 N 0Y N Cy Cy N O Cy' 0

OU N On .1 Co N- L0 co CON 0n N Wn (D co tN * 4 N% C%0i OD L 0ý 0 0 N N 0 0 r- 00 0 V O0 0n in I%-0tD o, '0 N a in - 10 N% N %0 N t-- 0o v- - ( in oA * tw > Ný N N N0 r A m (-Y 0. P- 4 0 N0 .0 10

x 0.J~ ~ ~~ No i % 0 C N NY N) 0 N C o n i '

4 0 '0' mt NAT N 0' 0% 0' 4 o n 0 N 10" N 0r Co N% inin% . i 0 N t 0n '0o o on o, o No co, Ny %o ' m. 0D 0 0 0% 00 r- Co g9 M T c 0 N Fn %0 o 00 r 0 0,c

) 4'm ~( N- Nn N: Nn N7 No Nn NP --t 0% NN

C 0 Mf - w% .0 N 0 1 O 4 * ) co tn in vi V% nex 0; a; '00 Co Co P: '0' 1 1 o Co rto J': NA * C4N1' ' o N 0 0 0 1 o 9

..J in, %n LA o in 4A tn it *t 0* *l *m *l m0 1 N N- ~ ~ ~ r p- . __ w - - -

I , 0 9n 0 iN M M 0a i 0 0n 0 n 0 wn 0 o I.' o. 0

R-5123 67

F- K) k, 0 N N CO N ao 10 Co M) C-0 0% m) 00 C N N - m% P- r~- .- -h- -t oo 1o . 0 N 0f m) I-- mO 0% m ON

m 0%.0 M, N) 00 hl t P- 1- 0 Ln

oý LA K) 'A 16 10 'A N 44 00 rl 0% wn iý-. 0 .0 -40 1- cO a N1 *t0% 0 -- K ) t LM 140 t.- co 0 f- N1

q- w- p." C',j N 04

m - K OD N 4- hi LA a p.- 4 -t .m) in CN N) (10 co co CO 0 N co co .- .

K) 4o a0 CO 0 ) 0 4- C% M) co 4 0* ) N - . 0 0% ON Co P ~- 0 .0 LA

LA Lo LA U, Ln Ln 4 + 4 - -I 4* -It4

C; 0' 0 09 0: 0 S B 0: S9 0: 0:

-0 0 0% ul 6% . ) 4t M O r ~ ." t, 4 L N U) r.- '0 N N~ .- Ke) '0 0 .0

ol6 - W) MO - 0 N P.- 0 0% N .- t

) 4t '0 10 10 '0 *-- N wO *I ?1- 0 0 6%00 N -1 10 wO 0 N K) LA 10 mON (716

04 N 0 '- 0 ý 0; (3; CO z LA 4 14 N 1*'0 -0 10 '0 LA tA in u) tA LA A LA LA L

N N N N N N N N N N1 Nl NY N NM

'0 1-ý ,J - N '0 4t 0 :t U) U) 0 co knW- 0O L 0 M) M 0 M) 0 - LA -,0

N~ CO) K) ý .0 '0 0 m) N P.- 0 6% .%0N 1.-! K) CO) M CO m) fl- L COD NY r fl-co I.-I r.- .6 o LA u) 4t 4 Po N N 0

0 M)i 0% 0% '0 0 U, 0 N Nv t-- :Tin '0. I.- IN - 0 ON LA 4N 10 CO (% N~ 10r-- 0. ON N- LA U) ?- (7% LA -- 0- ZP a) IDN~) K)' CO N .- (7% NS - .0 tA 6% '0 '0 0%NI ) 6%j a, a 0 ;* r-- 00 f.- UI) V .0 6%ý 0

0'~ % 0% 6%1 MO CO f- P.- .0

.0 0. 'A 0; A ; a; ; 00 09 0 0

.0 - CJ N% M) Kl 4t * LA tn -0 '0 t.-

68 R-5123

S6.0 ST WEIGHT STUOY

- 90.00

L0X 140.00

WE1 13D. 00GH

T

,20.00 - - - - - - - - - - - - - - - -

110.00

I00.00

90.00*

80.0 On

20.00

40.00

2 .0 40.00 60.00 110.00 100.00 120.00 140.00 16n.00 I£f0. 0I iueTIME (CASE 7.00O

Fiur 20. Sample of Cathode Ray Tube Plot (cRPPft)

-Z R-51230U.

A DIVISION OF NORTH AMERICAN AVIATION. IN

OOi 0LST WEIGHT STUOY2.40 - -

LL ----0

v *02.00

L

*

,,,- - U

E

1.80

1.60

1.40

* 1. 20

1. .00

0.80

0.4

160.00 180.00 20.0n .n O2. 40.0 60.00 80.0 200.00 12f'0 240.00 260. o 8 1 T.1. n

"TIME (CAiSE ..00 1u o

4T) ~Figure 21. Sample of Cathode Ray Tube Plot (CRPI4) 2

LST WEIGHT STUOY

EN

40. D 0O

335.00

50.00

45.00

40.00 __

25.00 __

70*

1 0.00o

"55.o 14.0

TIM (ASEO.O

0 50

X;&40 4 c RE T RwIEA DIVISION OF NORTH AMERICAN AVIATION. INC.

24-44.4?{]06l 000{ 002 -00

LST

WEIGHT STUDY0

- 20.00]

L

0- 40.00

TE

P - 60.0

0E

G- 10. 00

F

-100.00

-120.00

-140.00

-180.00

-220. 90

-240. 00

-260.00 .0.

I 1-300,00 ___•0.o50 20n. .On 0 20.00 40.00 60.00 80.00 100.00 120.00 140.00 160T.L0 10 q - DC, 2. 5.5C

TIME (CASE ?,00. )

Figure 23. Sample of Cathode Ray Tube Plot (CRIPLT)

2U

~0.00 LST WEIGH~T STUD~Y

L

A

E

T50

E

"Go0.000E

F

50.00

.45.00

40. 00

35.00

30. 00

li. 00

20. 00 _____

1 5.00

10.00 ____

tI HE (CASE 7. 00i i . Figure 24. Sample of Cathode R~ay Tube rlot (CRPL#T)

I: R-51230IL.

0 Cr K WE T-IF Y 1k17W 1EA DIVISION OF NORTH AMERICAN AVIATION. INC

0?' 0 .01: LST

WEIGHT STUoY

j U.280.

T

L0.26,x

Ap .2'40 ___

- -- L

.200

.180

*

.1671

111 ]*

.14.0 *.• 2 ¢ •e O O0 ! oJO 1 . 0 1 0 10,_1_._iO l gq 0q -- i L .n

*'H C l' ?O

,T,•) Fg~r•2•. Samp• o c~oa•.120 ____(_L•

iJ*

LST WEIGHT STUOY

4. 80

WT4. 40

L0

E 4.00

XP

E . ...

0 *

3.20 __

2.80 -

2. 40

2-4 - • -- - -- - - - - - - - - -

.20 __

2-0.00 '60. DO 60.00 o0.00r 100.00 120.00 140.00i 160.00 Isn.n0

TIME (CASE 7.00

1 .~ Figure 26. Sample of Cathode Ray Tube Plot (C11PLOT)

i . ..

I

w, 720


004A 000 LST WEIGHT STUDY0~ 0

440

L

4.00*UNAV*

L*A

2.80 __

2. 40

2.0 __

1.600 T*9 n~r

Iftono nn .5 n2000 40.0 6 .00 0.0 10 4 0 tin180n 2

- -TIM-E-- - C--E- - - -

Fi.u0 27. Sampl -of-at--e-Ra--ub--lo-----.#T

R*12

LST WEIGHT STUOY

LL 20O0. 00A -

CC

V

0 1800.00L-UM

1600.00

1 400. 00

1200.00

1000.00

800.00

600.00

*

400. 00

200.00 *

20,00 46.00 60.00

R 1

• 1R-51230

IR OC K__ T W "¥I" EA DIVISION OF NORTH AMERICAN AVIATION. INC.

09 000

-I *: . . __T ~

0.010~ 00.00 1 LI. 10 140.00 260.00 180.00 200. r

TIME (CA*SE 00

Figure 28. Sam~ple of Cathode RayTube Plot (cRPL#T)

73

UNCLASSIFIED AD 407 315 - Defense Technical … R-5123 5 FORMA 608 B PLATE REV. 1.58 DA IVISION Of'...

Documents

Transcript of UNCLASSIFIED AD 407 315 - Defense Technical … R-5123 5 FORMA 608 B PLATE REV. 1.58 DA IVISION Of'...