unclassified ad 407 315 - defense technical … r-5123 5 forma 608 b plate rev. 1.58 da ivision...
Post on 09-Jun-2018
226 Views
Preview:
TRANSCRIPT
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
A DIVISION OF NORTH AMERICAN AVIATION. INC.
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
A DIVISION OF NORTH AMERICAN AVIATION. INC.
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
A DIVISION OF NORTH AMERICAN AVIATION. INC
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
A DIVISION OF NORTH AMERICAN AVIATION. INC.
(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
A DIVISION OF NORTH AMERICAN AVIATION. INC.
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
A DIVISION OF NORTH AMERICAN AVIATION. INC.
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
A DIVISION OF NORTH AMERICAN AVIATION. INC.
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
A DIVISION OF NORTH AMERICAN AVIATION. INC.
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
A DIVISION OF NORTH AMERICAN AVIATION. INC
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
A DIVISION OF NORTH AMERICAN AVIATION. INC.
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
A DIVISION OF NORTH AMERICAN AVIATION. INC.
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
A DIVISION OF NORTH AMERICAN AVIATION. INC.
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
A DIVISION OF NORTH AMERICAN AVIATION. INC.
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
A DIVISION OF NORTH AMERICAN AVIATION. INC
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
A DIVISION OF NORTH AMERICAN AVIATION. INC
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
A DIVISION OF NORTH AMERICAN AVIATION. INC
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
A DIVISION OF NORTH AMERICAN AVIATION. INC.
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
A DIVISION OF NORTH AMERICAN AVIATION. INC.
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
top related