LMS qualification and flight acceptance T/V test summary ... · Acceptance T /V Test Summary and Thermal Design Final Report 4. 2. 2 Determine thermal performance in a lunar environment,

LMS Qualification and Flight Acceptance T /V Test Summary and Thermal Design Final Report

NO.

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PAQI

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Of .... lU!IIU-. ~Dhtllan DATE 10/6/72

This ATM summarizes the analytical and test results of the Lunar Mass Spectrometer thermal control system. The report is divided into three sections. The first section introduces the hardware and develops the mathematical thermal model of the experiment. The second section summarizes the Qualification and Flight Acceptance thermal vacuum testing program, and the third section presents the lunar surface predictions of the hardware.

Prepared By~D. Toel e

Checked By .J/: "£_l1' ttr.:;:) G. Psaros

Approved By (. ~ E. Granholm

NO.

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.. 'llii?ID Jp IIIIDe yablina Dhllalan

LMS Qualification and Flight Acceptance T/V Test Summary and Thermal Design Final Report

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SUMMARY

The LMS maintains thermal control. within the -10 to + 125°F operating temperature goals, by means of insulated surfaces and spectrally selective coatings during lunar day, and insulated sur-

, J

~.~"'·' faces and electronics and heater power dissipation during lunar night.

A 107 node thermal model was developed utilizing the Bendix thermal analyzer computer program which accounts for all modes of heat transfer including radiation and conduction. This model was used to optimize the thermal during the initial design phase, by considering such things as, spectrally selective coatings and radiator surfaces, radiator areas, and minimiz~tion of support structure heat leaks. ·

The Qualification and Flight Acceptance thermal vacuum tests are surmnarized including test setups and pertinent results. Detailed cor.:. relations studies are summarized and compared with the analytical model demonstrating the validity of the model in predicting the thermal response.

L'Ullar predictions are presented which are based on the test correlated thermal math model. These predictions include both a clean and dust covered model deployed equatorially and at a 200 latitude. A clean non-dust degraded experiment with 7. 7 watts of power dissipation will operate within a -4/+110°F range for lunar night and noon, respectively. Degradation of all unprotected surfaces lOOo/o results in an increase of lunar n~~n temperature to 11 7°F. An accumulation of 3% dust on the mirror surface· 'can be tolerated before the + 125°F operating goal is reached.

Deployment at a 20° latitude has the effect of decreasing the lunar noon temperature by SOF.

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Paragraph

1.0 Scope

~-Qualification and Flight Acceptance T /V Test Summary and Thermal Design Final Report

TABLE OF CONTENTS

Contents

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Page

1

SECTION I - EXPERIMENT OBJECTIVES AND THERMAL ANALYSES

2.0

3.0

4.0

5.0

Background and Purpose

Design Specification Requirements

3.1 Temperature Requirements

3.2 Power Requirement

Thermal Design and Analysis Goals

4.1 Design Goals

4. 2 . Thermal Analyses Goals

Analytical Thermal Model

5.1 Nodal Description

5.2 Thermal Resistances

5.3 Solar Heating

5.4 Lunar Surface T e:r::Q.perature i ,,.

s. 5 Electronics Power I

5.6 Analyzer Support Structure

Multilayer Bag and Masking . ' '

3 .::·~

4..:::'

4<

4:..!

5 '

5

5..: :;-.

7 I 7

7

7 .

11

7 i

'12

--~-121 •··

• ! 5. 7

5.8 LMS Radiator Plate Thermal --"---~·-~_._:.:'"~.c..· _....o:._._..._. ....:-"---~ ..... '1 2 i 7

Coatings Optical Properties 13

s. 9 Electronics to Analyzer Cable_ . --·-···-··- -----·-· -· ---.1.3

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TABLE OF CONTENTS (CONT. )

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Paragraph Contents Page

------- 13 ""!l'

6.0

7.0

8.0

5.10 Heat Leak Summary

S .11 Design Parameter Study ---------- _____ · 13'

SECTION II- QUAL AND FLIGHT THERMAL VACUUM TESTS

Background

Test Objectives

---------- --·---------------

----------····--· ------

Description of Test Conditions and Test Installation

R.1

8. z 8. 3

8.4

8. 5

8. 6

8. 7

Thermal Test Conditions

Vacuum Chambers

·Lunar Surface Simulator

Solar Simulation

Experiment Location

Thermocouple Location

Experiment Models

8. 7. 1

8. 7. z

Qual Model

Flight Model

.. -·----··-

18 I"

zv zo

zo zo zo 30

30

30

30

30

9. 0 Results and Discussion 33

9. 1 Graphical Summary of Results

9. I. 1

9. 1. z Qualification Test

Flight Test

33

33

38

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TABLE OF CONTENTS (CONT. )

Paragraph

9.2

Contents

Fact~rs Affecting Correlation of the Analytical Model with Test Data---

Page

10.0

33

9.2.1 9.2.2 9.2..3

View Factors . --------------- 38 Effects of Infrared Solar Simulation --- -- 38 Effects of Flight Test Analyzer Simulator -----·- --------------- 38

Comparison of Data and Correlated Temperatures ----------···-. ··- ·---------·--- 43

SECTION Ill- LUNAR PREDICTIONS

Lunar Predictions 47

10.1 Nominal Operational of LMS with Clean Thermal Surfaces -·-··--- 47

10.2 Degraded or Dust Covered Radiator and Cover Surfaces 47

10.3 Lunar Noon Non-Operating Condition 47

10.3.1 10.3.2

Dust Cover On Radiator Exposed - No Power Dissipation

47

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Table

3.1

3.Z

5.1

8.1

9.1

9.Z

9.3

TABLES

Title

Radiator Plate T·emperature Goals

Power Requirement ---- ·--- ------- __ 4

LMS Heat Balance Sununary ----·----------16

Qualification and Flight Test Conditions -----~.Zl

Configuration Factors - Qual and Flight T /V Test 39

Solar Simulation on Qual and Flight Thermal Surfaces _______________ 40

Lunar Mass Spectrometer Qualification Thermal Vacuum Test Data and Correlation ---~44

Lunar Mass Spectrometer Flight Acceptance Thermal Vacuum Test Data and Correlation -----":1:45

iv

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Figure

5.1

5.2

5.3

5.4

5.5

5.6

8.1

ILLUSTRATIONS

Title

LMS Experiment Nodal Model

LMS Electronics Nodal Model

LMS Radiator Plate and Analyzer Plate Assembly Nodal Model

Lunar Surface Temperature vs. Sun Angle and Latitude

LMS Flight Model Predictions Lunar Noon

LMS Flight Model Predictions Lunar Night

LMS Lunar Surface Thermocouple Locations

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PAGE ,y !I,

DATE

8

10

11

14

IS

8.2 LMS Qual Test Lunar Surface Temperatures ----23

8.3

8.4

8.5

8.6

8.7

8.8

8.1()

LMS Flight Acceptance Test Lunar Surface Temperatures

LMS Qual Test Radiometer Response

LMS Flight Acceptance Test Radiometer Response

LMS Qual Test Thermocouple Locations



Qual Model T /V Test Setup

Flight Modtll T /V Tfut S~tup

24

25

26

27 !

28 '

29

31

32

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LMS Qualification and Flight Acceptance T /V Test Summary and Thermal Design Final Report r~-~pace

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Figure

9.1

9.2

9.3

9.4

9.5

10.1

10.2

ILLUSTRATIONS (CONT.)

Title

LMS Qual T /V Test Thermocouple Temperatures (Events Summary) ------

LMS Qual T /V Test Thermocouple Temperatures

LMS Qual T /V Test Thermocouple Temperatures

LMS Flight Acceptance Test Thermistor HK-41 Temperature Plot and Events Summary

LMS Flight Acceptance Test Analyzer Simulator Temperatures

Lunar Mass Spectrometer Radiator Temperature as a Function of Solar Angle

Effects of Dust on the Solar Absorbtance of the LMS Radiator Coatings

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Page

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35

37 I

:.../ 41 '

42 /

48

49' .!

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·- 'tMs 'dua1l~icaH~rl1and Flight Accepta.ace T /V Test Summary and Thermal Design Final Report

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I

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The purpose of this report is to document the development of the Lunar Mass Spectrometer Thermal Analysis, and the results and conclusions resulting from the thermal vacuum testing of the LMS Qualification and Flight models in the BxA thermal vacuum test chamber. This report is in three parts: Experiment Objectives and Thermal Analysis, Qualification and Flight Thermal Vacuum Testing, and Lunar Predictions.

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SECTION I

EXPERIMENT OBJECTIVES AND THERMAL ANALYSES

PAOI

DATI

This section summarizes the objectives of the LMS experiment and the development of the thermal model and analysis.

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2. 0 BACKGROUND AND PURPOSE

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The Lunar Mass Spectrometer is basically a magnetic - deflection mass spectrometer designed to identify the constituents of the lunar surface atmosphere. It consists of two major subsystems, the electronics section

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and the gas analyzer section. The subsystems are joined to a common baseplate with each subsystem having its own fiberglass cover. The electronics subsystem is contained within a 40 layer thermal bag.

Thermal control is maintained by means of a second- surf ace mirrored radiator plate and internal power heat dissipation. This control will maintain the electronics heat sink between -10°F (lunar night) and +125°F (lunar noon). The gas analyzer subsystem contains no thermal control and is subject to

0 0 the lunar temperature range of -300 F to +250 F.

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3. 0 DESIGN SPECIFICATION REQUIREMENTS

3. 1 Temperature Requirements

DATI

Design temperature goals per the LMS project group are as follows:

TABLE 3.1

RADIATOR PLATE TEMPERATURE

Operational Goals Day +125°F

Night -10°F

Survival Night -30°F

3. 2 Power Requirement

Reference than the following:

states that the LMS power requirement shall be less

Operational

. 1/ Surv1va

TABLE 3. 2

POWER

Day

Night

Night

13. 8 Watts

12 Watts

8 Watts

The actual power dissipation used for the purpose of thermal analysis of the Flight Model was, 7. 7 watts within the electronics bag, and 1. 69 watts on the analyzer baseplate.

/.,"WOspaae stems Dhrlalcn


4.0 THERMAL DESIGN AND ANALYSIS GOALS

4.1 De sign Goals

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DATI

4. 1. 1 Maintain the electronics within the specified limits of temperature and power.

4. 1. 2

4. 1. 3

4. 1. 4

Provide means of dissipating electronics heat during lunar day

Minimize the effects of solar heat

Provide adequate temperature control during lunar night

Minimize heat leaks

Provide electronic components with good heat paths to the thermal plate

Provide thermal control to..± 25° off equator deployment.

Minimize dust effect resulting from:

Astronaut activity

LM ascent

Minimize effects of time dependent degradation of thermal surfaces.

4. 1. 5 Provide astronaut activity and deployment constraints commensurate with good thermal control.

4.2 Thermal Analyses Goals

4.2.1 Determine heat leak or gain through:

Cables

Structural supports

Experiment cover

Masking

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4. 2. 2 Determine thermal performance in a lunar environment,

4. 2. 3 Recommend thermal coatings.

4. 2. 4 Develop a thermal math model.

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DATI

4. 2. 5 Utilize data obtained from the performance of DVT to design Qual. and Flight models.

4. 2. 6 Utilize data from the DVT and ALSEP system qualification and flight thermal - vacuum tests to verify the performance of lunar hardware.

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LMS QualUication and Flight Acceptance T /V Test Summary and Thermal Design Final Report

PAGI OP 51

DATI

5. 0 ANALYTICAL THERMAL MODEL

5. 1 Node Description

The thermal model of LMS (Figures I, 2, and 3) was developed utilizing the Bendix Thermal Analyzer Computer Program. The model consists of 108 nodes representing various components of the LMS in addition to the lunar surface and space. These nodes and their physical significance are listed in Appendix A.

5. 2 Thermal Resistances

There are 590 resistances in the LMS thermal model. These resistances represent all heat exchange, both through conduction and radiation, between all nodal surfac£_S of the experiment and the environment. These resistances, their interconnecting nodes, configuration factors, and surface areas are listed in Appendix A.

5. 3 v

Solar Heating r-··f r l· • ' j

j; \ .,c,,i . • ~·

The absorbed solar flux is simulated in model by apply equivalent heat to the surface node of the irradiated surface. Surfaces irradiated are mirror surfaces, masking, analyzer baseplate, and analyzer cover surfaces.

5. 4 Lunar Surface Temeerature

Variation of the lunar surface temperature as a function of deployment angle and solar angle was considered in the analysis. Temperatures of the surface as a function of these variables are indicated in Figure 4, During lunar night the lunar surface was taken as - 300°F.

5. 5 Electronics Power

The power input to the LMS consists of operational electronics power, analyzer filament power. back-up resistor heater power, and standby survival power. These four inputs are handled as heat inputs to the appropriate electronics nodes in the thermal model.

The input power data of each model was obtained during LMS functional testing of Qual. and flight hardware. These functional tests were per Bendix TP 2347420 and 2368966.

A(':. (_ f;' /;'(.!to. ( ·, f'-Jc'.' ''' ;t)•• LMfExPERIMENTAL NODAL MODEL - THERMAL AND Al'fALYZEl\ COVER ASSI:MBLIES

. 1

' I

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I I

1 I

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Figure S.'-2 L

MULTILAYER BAG (NODES &5, 66, 67, 68 AND 69)

THERMAL COVER

,/

PREAMP/DISCRIMINATOR {NODES 107, 108 AND 109)

MOTHER BOARD {PART OF CABLE ASSY) ~ (NODE 100)

EMISSION CONTROL (NODES 111 AND liZ)

..

CONNECTOR SUPPORT (NODES 120, 130 AND 140)

LUNAR MASS SPECTROMETER ELECTRONICS NODAL MODEL

Pa~e 9 ~f l1

COUNTER AND DATA COMPRESSOR (NODES 1 0 1 AND 1 02)

RADIATOR PLATE (NODES 11, 21, 31, 41, 51 AND 61) '

\-- .___

...

DECODER AND SIGNAL CONDITIONING (NODES 103 AND 104)

HOUSEKEEPING MULTIPLEXER (NODES 105 AND 106)

POWJ!:l\ SUPPLY MODULE (NODES 161 THRU 168)

- --~ -.......,.,..,_---~~-..."'---

ANALYZER CONTROL ELECTRONICS (NODES Ill AND 114)

TRANSITION BOARD (PART OF CABLE ASSY)

(NODE 98)

CONTINGENCY HEATER (NODE 75)

;... CONTROL AND MONITOR ,;_,-/ (NODE 110)

H. V. RELAYS

•

l-- ___ -t..__ r. . '' -l.Jtx~tt'omd'

I -

l r ----------------.

·Suppo-rt Structure: {Node•.- 4_2, 62, 71 and nl.

Page l,tO oi' .61·

:_ _.)! ·- { \ ~

Figur,e $. 3

L'tiNAR MASS ~PtCTROMli:tti. ft.ADIATOR~ Pl,;AT.E ANa ANALYZER.J\SSENBIJ! -~ .,-; ,(.. i\._._'"· ·I :.•. r ·

s~pp.;rt Structure Isolator ; · ·, ~T)r. '(Nodes 12 4'-"-d 32): ~, ) · \,

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i_ ~idastd~: { 1(No~ea. 1 O,,_zo, 3.P,

1 40,. SO IUI4 60)

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'- -- ~-- - t

Analyzer Baseplate {Nodes·<Jo,- 73; 760 't? and np·, )

·Analyzer Cable· (Nodes 204,' ·205,· 206, 208 and·210} 1 ,)

___.,-..__

Second Sur(aee Mirron · ·; .,. ··' · (Nodes J througb.o) <.:::- \

I .I

.. .

r

;_, Cable Ree~ Support • Cable Reel Support

rmal Isolators

-Titanium - S\&pport Structure Isolator·; - '· -

Radiator·Plate

j 'l•·

ton Solll'ee · ·

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5. 6 Analyzer Support Structure

The analyzer baseplate is structurally mounted to the electronics radiator plate via two titanium support structures. These structures provide a direct path of heat leak through the multilayer bag to the baseplate, which is not thermally controlled. To minimize this as a problem, a thermal isolator (Figure 3) was designed employing a titanium clevis and fiber spacers. In addition, heat leak through the analyzer cover via radiation was minimized by aluminizing the analyzer cover.

5. 7 Multilayer Bag and Masking

A multilayer bag was produced from 40 layers of double- sided aluminized 1/4 mil mylar and spacers. This bag completely encloses the electronics and is mounted to the underside of the radiator plate. The bag's thermal resistance was obtained fron1 the Lockheed Acc~ptance Test Report TXA 2464. Heat leak through the bag was simulated in 'the model by means of an effective emittance. E = 0. 004 7.

The radiator plate masking is used to eliminate extraneous heat leak from the radiator plate edge and mounting hardware etc. In addition, the masking allows for changes in exposed radiator area to accommodate change in power dissipation without major design changes. This masking is made from 20 layers of double- sided aluminized I/4 mil. mylar with spacers. It has an outer layer of aluminized teflon with teflon side out, resulting in an a/E = 0. 20/0.69. The effective internal emissivity is E = 0. 060.

5. 8 LMS Radiator Plate Thermal Coating Optical Properties

The radiator plate consists of a 0. 060 inch thick plate of 2024- T6 aluminum coated with 60, 1 inch square second- surface mirrors. These mirrors are 1 inch square and 0. 006 inch thick. The mirrors have optical properties per acceptance specification OCLI Spec. SI-100 stating, solar absorbance a= 0. 050 + 0. 005, and hemispherical emissivity E = 0. 80 + 0. 02. The actual values used in the thermal model were a/E = 0. 08/0. 8. The a= 0. 08 was used to accommodate the effect of the cracks between each mirror.

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5. 9 Electronics to Analyzer Cable

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The LMS cable consists of two sections. There exists a section between the electronics and the analyzer comprised of copper coax high voltage cables leading to high voltage filaments, and manganin wires leading to the interconnect board mounted on the analyzer baseplate. The ribbon cable which mates with the ALSEP central station also interfaces with the interconnect board. The thermal conductivity per foot used for the section between the electronics and analyzer was 0. 0214 BTU- Ft.. The

0

I -1/2 inch length between the radiator plate clip and the atil!y.zfr baseplate clip has a conductivity of 0. I 72 BTU

HR°F

5. I 0 Heat Leak Summary

Table 1 summarizes total heat leaks for nominal noon, night, and degraded noon with a mirror a= 0. 10 and 100% dust coverage on unprotected surfaces.

5.11 Design Parameter Study

The final LMS thermal design is influenced by variations in both internal power dissipation and final adjustment to the radiator masking size and associated radiator area. Figures 5 and 6 present the radiator plate temperature variation as a function of both radiator area and internal power dissipation for lunar noon and lunar night.

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LMS FLIGHT MODEL PREDICTIONS LUNAR NIGHT

I

Page .1. of 91'

-80 ~----~~---------+--------~;~--~--~--------~--1

i j

8.

Internal ~w~Jr 'Dissi·pattion-Watts !

9 10

15'0

14.0.

130

<_ 1W . . ?

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80

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LMS FLIGHT MODEL PREDICTIONS LUNAR NOON

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Jj ~-------~-------~Q_--------~---

5 '7 l

8

IS!t~er:lfid FoweT .Dissipation- Watts

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Page 15 of 51

9 10

Page 16 of 51

TABLE 5.1

LMS HE.A T B.AL.ANCE SUMMARY

! Heat Leak (Watts)

Path Night Noon Degraded Noon

40 in2 radiator -4.430 -8.066 -8.615

Masking -0.436 -0.2.83 -0.340 I

Isolator to Analyzer -1.2.40 +0.470 +0.802.

Thermal Bag -0.0930 +0. 160 +0. 163

Electronic Cable -1.075 +0.318 +0. 540

Thermal Plate -0.095 +0.088 +0. 188

TOTAL -7.36 -7.32. -7.36

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SECTION II

QUAL AND FLIGHT THERMAL VACUUM TESTS

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This section summarizes the Qual and Flight thermal vacuum test results and correlations.

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6. 0 BACKGROUND

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DATE

The Lunar Mass Spectrometer Qualification Thermal Vacuum Tests were performed during the period of 20 May 1972 thru 7 June 1972. The Flight Acceptance tests were performed during the period 01 July 1972 thru 11 July 1972, both at the Bendix Aerospace Systems Division in Ann Arbor, Michigan. This report presents the thermal data obtained during those tests together with a correlation of the thermal mathematical model with experimental data. Also included is a brief description of the test installation.

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7. 0 TEST OBJECTIVES

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DATI

The objectives of the LMS Qualification and Flight Acceptance Tests were to subject the LMS and integrated ALSEP to the lunar thermal/ vacuum environment demonstrating the ability of the LMS design to withstand these extremes without loss of function.

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8. 0 DESCRIPTION OF TEST CONDITIONS AND TEST INSTALLATION

8. I Thermal Test Conditions Qual Test and Flight Acceptance Test

The LMS Qualification test was correlated for three thermal test conditions. These conditions simulate the lunar noon and night phases of operating in addition to the Design Limit Condition which simulates a +280°F lunar surface and I. 25 solar load. (See Table 2)

The Flight Acceptance Test was correlated for the Acceptance Lunar Noon and Acceptance Night conditions. (See Table 3)

8. 2 Thermal Vacuum Chamber

The test was conducted i~ the 20' x 27' thermal vacuum chamber in a vacuum greater than 1 X 10- Torr.

8. 3 Lunar Surface Simulator

The lunar surface was simulated by a metal plate (5' x 5 1) with 8 inch vertical lips. The surface was horizontally mounted in the northwest section of the 20' x Zf.,,'"' cA,amber. The surface was coated with black paint to simulate the emittance of the lunar surface. Surface temperature was controlled by means of resistance heaters beneath the surface and vertical, lips, and cooling coils beneath the surface.

8. 4 Solar Simulation

Solar simulation was achieved by utilizing an array of quartz infrared electric heat lamps. These lamps are arranged in pairs and were directed through the center line of the experiment at right angles to the horizontal approximately 18 inches above the experiment surface. The intensity of radiation impingement was determined by measuring the intensity of radiation absorbed in a radiometer coated with a second surface mirror. The plot of this intensity for both Qual and Flight is indicated in Figures 5 and 6.

LMS Qualification and Flight Acceptance T /V Test Summary and Thermal Design Final Report //-~era 111 aoe

, }Stems Dfvlelon

TABLES.!

QUALIFICATION TEST CONDITIONS

Solar Lunar Surface Intensity Temperature

Condition Description (Suns) (o F)

1 Qual Lunar Noon 1.0 +250

2 Qual Lunar Night -- -300

3 Design Limit Noon 1. 25 +280

FLIGHT ACCEPTANCE TEST CONDITIONS

Solar Lunar Surface Intensity Temgerature

Condition Description (Suns) ( F)

1 Acceptance Noon 1.0 +250

2 Acceptance Night -- -300

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DATI

Power Dis sipatior

(Watts)

7.36

7.36

7.36

Power Dissipation

(Watts)

7.70

7.70

~

LS 17

~

$ LS 6

/

I

...... / TFta~ &1 \(

L~UO'

tsi t9

US4 s

LS7 s

LS15

LMS LUNAR SURFACE THERMOCOUPLE LOCATIONS ~ ..-· - -· -·

8

.JS s

LS5 9

(L§ s

s LS14

'

Page 22 of 51

$ LSll / LS3 s

Q LS12

LS6 s

s LS13

1,8'9

t9

s ""

350

. 250

I 150

1.1..

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ILl It: ::;, ta: It: ...... -so Q...

::t: ILl t-

-15

-25

-35

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Figure 8. 2

;

BENDIX AEROSPACE SYSTEMS DIVISION ALSEP ARRAY 'E' THERMAL VACUUH QUALIFICATION TEST

ZERO TIHE • 000001 OF 05/20/72

Page 23 of:

~ 180 LMS LUNAR SURFACE TEMPERATURE AT LOCATION 01 (!) 183 LrtS LUNAfl SURFACE TEI'IPEI'IATUR~ RT LOCATION 0\1 A 188 LHS LUNAR SURFACE TEMPERATURE AT LOCATION 09 ~ 192 LHS LUNAR SURFACE TEHPERRTUftE AT LOCATION 13 x 196 LMS LUNAR SURFACE TEHPERATURE AT LOCATION 17

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-··· :- 'l': .....,-

[F ............

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till ~ ~

~ .. -~ '

~ :

1 30 60 90 120 150 180 210 2ij0 270 300 330 360 . 390

ELAPSED TIHE - HOURS·

~ .... --

ij20 4!

350

250 I

150

.....

I 50 1.&.1 a: ::::t 1-a: a: w -so a.. :r 1.&.1 I-

-150

-25

-35

I

0

Figure 8. 3

BENDIX AER~SPACE SYSTEMS OIVISI~N ALSEP ARRAY 'E' THERMAL VACUUM ACCEPTANCE TEST

ZERO TIME = 000001 ~F 07/01/72

Page 24 of 51

m 160 LMS LUNAR SURFACE TEMPERATURE AT LOCATION 01 ~ 163 LHS LUNAR SURfACE TEMPERATURE AT LOCATION oq & 188 LMS LUNAR SURFACE TEMPERATURE AT LOCATION 09 ~ 192 LMS LUNAR SURFACE TEMPERATURE AT LOCATION 13 X 196 LMS LUNAR SURFACE TEMPERATURE AT LOCATION 17

·-

- ""

~~ l!lr- ......

~

\ Jl

~ I j

' L. I

20 40 60 80 100 120 140 160 180 200 220 2110 ELAPSED TIME - H~URS

~

... I I I

260 280 3 0

600

500

1-

~ 400 0 .... u :;z

1- 300 0 0 lL.

w a:

.a: 5 200 r.n

tr w ~

~ 10 ta: :X

0

-10

)

Figure 8.4

~ 016 LMS RROIOHETER RESPONSE

~~ ill~ v

I

-....

W'

I 0 30 60 90

~

BENDIX AEROSPACE SYSTEMS DIVISION ALSEP ARRAY 'E' THERMAL VACUUM QUALIFICATION TEST

ZERO TIME : 000001 OF 05/20/72

~ I.._

- u .....,., ;..,. ~

~ I-

'""-

120 150 180 210 2ijQ 270 300 ELAPSED TIME - HDURS

· Page 25 oi 51

~ ,. l I

I

-

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330 360 390 "20 1150

2o

600

500

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- 300 ~

0 .... w ;r; a: ~ 200 rn

c::: w IL

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0

-1~

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Figure 8. 5

~016 LNS RAOICHETER RESPONSE

q ~

.........,_

'"u I

'

20 qo 60 80

BENDIX AEROSPACE SYSTEMS DIVISION ALSEP ARRAY 'E' THERMAL VACUUM ACCEPTANCE TEST

ZERO TIME = DOOOOl OF 07/01/72

.

-~ ... -

- -

100 120 140 160 180 200 ELAPSED TIHE - HOURS

220

_ Page 26 of 51

' '!$

/ ~

' I I

v I 1 J

~ \

. 2110 260 280 300

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r

-I rJ)

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f

r -. -"·-·- ;"~ ;1

0 0 0 0

r------ s.;s ±:zs--1

r---------+---------1

I

1 -rk--S-1

-~ I

o -k{i±.>s l LMS- ~

@

COVER ASSY- ELECTRG~!CS s:=C.Tl0N 'REF DVJG 2?.47~75

3.00 !.. 25'

l c=d'

,Page 28 of 51

1.1o L ~.2sl

I 3.E.\ -: . .2-s

0

~LM<i-s

Figure 8. 8

0

LMS-10

i--3.f.~--1 I ±~~ '

tl 2.so±.2s

---+-l

Page 2~ of S'l

.. ___ ] -LMS-5

(MOUNT iloi'C.RMOC.OUf>LE. ON FAR SI"OE.)

-,')

l . c-·-:=J .:J

CO\IER:. ANAL '\'ZER REF DWG. :?.~751\

----I .--- .. -~-·- -. --- -··-- -· ·r-

1 ! 3.>:.3 ~ . .::.=:

-----===F=,.----4j l

!----S-b\0 !.25 ·----~

. ·---;:.,__,.,

THERMI\L CO\JF.R ASSY REF DW6 2.347510

ATM 1111 I LMS Qualification and Flight Acceptance T /V Test Summary and Thermal Design Final Report

8. 5 Experiment Location

, ... 30

DATI

The LMS was centered on the 5' x 5' surface and placed on a 5 inch aluminum standoff. This standoff provided a more accurate view factor of the lunar surface. Figures a. 6 and a. 7 are photos of the deployed Qual and .Flight models.

8. 6 Thermocouple Locations

Ten thermocouples were designated to measure the Qualification model experiment temperatures. These are indicated in Figure 8. 8 thru 8.10. There were no thermocouples attached to the Flight model. The temperature data obtained was housekeeping (HK-41) data via experiment thermistors. Thermocouple monitors located on the lunar surface simulator are indicated in Figure 10.

8. 7 Experiment Models

8. 7. 1 Qual Model

Qual model consisted of the complete LMS experiment including dust cover. It had a radiator plate masked to 34 in2 and dissipated 7.4 watts internal to the multilayer bag.

8.7.2 Flight Model

Flight model consisted of the complete LMS experiment including

or

dust cover.. Although the analyzer section was attached, analyzer power dissipation was removed to an analyzer simulator located on the lunar surface. This simulator temperature was maintained between 50 and 70° F. The model had a radiator plate masked to 40 in2 and dissipated 7. 7 watts internal to the multilayer bag.

51

Pot p E-i ~

U)

E-i U

)

~

E-i

> -E-i ~

0'

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. 0

00

0 ~

Q)

~

f.! ~

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•.-1

~

a

(

l/"'flf'GIIJ8-. , • ..._Dhlllan

LMS Qualification and Flight Acceptance T /V Test Surmnary and Thermal Design Final Report

9.0 RESULTS AND DISCUSSION

9. 1 Graphical Sununary of Results

9. 1. 1 Qualification Test

9. 1. 1. 1 Radiator Plate - Analyzer Plate Events Plot (Figure 11)

MO.

iATM 1111

, ... 33

DATI

This plot presents a chronological picture of the radiator plate and analyzer base plate temperature with the major events superimposed. Critical conditions from a thermal standpoint occur at lunar noon (189 hours) and lunar night (226 hours). The radiator plate/ analyzer plate was 111/130 oF and 0/ -108°F for lunar noon and night respectively. The design limit noon condition, which occurred at 348 hours resulted in a radiator plate/ analyzer baseplate temperature of +120/140°F.

9. 1. 1. 2 LMS 01, 03, 04 and 05 (Figure 12 ).

This curve summarizes the radiato.r plate, analyzer baseplate and analyzer cover together with the cable reel temperatures. The analyzer cover ran +158/136/ -200°F for design limit/lunar noon/night respectively. The cable reel lying on the lunar surface beneath a multilayer blanket attained +251/224/-289°F for the three conditions.

9. 1. 1. 3 LMS 06,07, 08, 09, 10 (Figure 13).

The plots in figure 13 include all the thermal cover and electronics cover thermocouples. Surmnarizing these data; the thermal cover end temperature was 145/124/-101 oF for design limit/lunar noon/night respectively, the electronics cover ends were 117/96/-200 "F and 165/143/-194 oF, the thermal cover bottom was 203/184/ -127°F, and the electronics cover bottom was 224/195/ -230°F.

lilY, MO.

OF 51

..._ "'-

I

w a: => ,_ cr: a: !.L! Q_

z: w ,_

. Page 34 of 5'

Figure 9. 1


ZERO TIME = 000001 OF 05/20/72

~070 LMS 01 RADIATOR PLATE !CABLE SIOEl TEMPERATURE ~ 072 LMS 03 ANALYZER BASEPLATE TEMPERATURE

Radiator Area - 34 in'

I! 00 Acceptance. Noon - 1. 0 Sun -Operate Power 7.4 wattsAnalyzer Power 1. 69 watts

I I

300 1 et Acceptance Night - Operate Power 7. 4 watts - Analyzer Power 1. 69 watts

2oo I

Design Limit Noon0

- 1. 2.5 Suns -Lunar Surface 280 F - Operate

'Power 7.4 watts- .Analyzer \Power 1. 69 watts

Operate Power Plus Backup Heater 7. 89 watte - No Analyzer Standby - Power 6. 72 watte

Pow\r .

1 , J ' \

100 A~ ~ urvrvv

0

I -100 I

.....

.... - ~

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)(

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.· _3cc I I I I I I I I I I I I I I I I I

0 30 60 90 120 150 180 210 2110 270 300 330 360 390 1120 ItS ELAPSED TIHE - HOURS

11

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200

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-20

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.Lgure 9.2

l!l 070 L11S 01 .t. .073 L11S OQ


ZERO TIME = DOOOOl OF 05/20/72

. Page 35 of 51

RADIATOR PLATE !CABLE SlOE! TE11PERATURE (!) 072 LI1S 03 ANALYZER BASEPLATE TE11PEAATURE CABLE REEL TEMPERATURE ~ 0714 LHS OS ANALYZER CASE TEHPERRTURE

f /

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--------~ ---- ------

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1.!20 1.!50

•

• Figure 9. 2 a

150

125

100

i::

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25

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-25

m KK-~1 L"S ELECTRONICS TEMP.

-

11 (1 }\ 1/ v

BENDIX AEROSPACE SYSTEHS DIVISION ALSEP ARRAY E THERMAL VACUUH QUALIFICATION TEST

ZERO TIME : 0000 HOURS ON 05/20/72

(v -""

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...____ ~--L ______ -- -- -- -~--- -- - ----

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Page 36 of 51

~

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v 330 360 390 ij2Q ijSO

I 400

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l 200

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} I 100 UJ c::: ::::> III

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igure 9. 3

1!J 075 LMS 06 A 077 LMS 08 X 079 LMS 10

a ,:I. ..... ,..

0 "'n -v

. Pagf:/37 of 51·


ZERO TIME = 000001 OF 05/20/72

ELECTRONICS COVER ENO TEMPERATURE (!) 076 LMS 07 ELECTRONICS COVER BOTTOM TEMPERATURE ELECTRONICS COVER ENO TEMPERATURE ~ 078 L11S 09 THER11AL COVER 80TTCI1 TEMPERATURE THERMAL COVER ENO TEMPERATURE

p -~

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j"·'fi"GII •ae . ..._Dht.IDn

9. 1. 2 Flight Test


9. 1. 2. 1 Radiator Plate Internal Thermistor HK-41 (Figure 14)

NO.

ATM 1111

PAGI

DATI

This plot presents a chronological picture of the radiator plate thermistor with the major events superimposed. Critical conditions from a thermal standpoint occur at lunar noon (154 hours) and lunar night (210 hours). The radiator plate and analyzer baseplate represented by housekeeping data read during the tests were, lll/140°F during lunar noon, and -11/ -120°F during lunar night.

9. 1. 2. 2 Analyzer Simulator Temperature (Figure 15)

This plot indicates the temperature maintained in the analyzer simulator box which was located on the lunar surface but removed from the LMS experiment.

9.2 Factors Affecting Correlation of Test Data and Analysis

9. 2. 1 View Factors

Table 4 tabulates the experiment to chamber view factors of the experiment critical surfaces. Locating the experiment on a 5 inch aluminum standoff provides an accurate view factor of all surfaces except the edge of the masking which has negligible effect.

9.2.2 Effects of Infrared Solar Simulation

Because the percentage of heat absorbed by the mirrored surface is greater under infrared irradiation then solar irradiation, the heat load on the mirrored surfaces is simulated by measuring one sun equivalent heat load on a mirrored ~surface radiometer. In simulating the mirror heat load, the heat loads on the S-13-G white coating, and the aluminized teflon are under simulated. The extent of the undersimulation is tabulated in Table 5. These loads have the effect of decreasing the equilibrium temperature 3°F from the lunar noon predicted temperature.

9.2.3 Effect of Flight Test Analy,zer Simulator

Qual Model dissipated 1. 69 watts of power in the analyzer baseplate which forms one side of the electronics cover. Dissipation of this power

ltiV. NO.

51 OF

Item

Mirrors

Masking Top

Masking Edge

-Electronic's

Cover

Analyzer Cover

LMS Qualification and Flight Acceptance T/V Test Summary and Thermal Design Final Report

TABLE 9.1

CONFIGURATION FACTORS

QUAL AND FLIGHT T/V TEST

Test

Lunar Cold Lunar

Actual

Surface Wall Surface

0.0 1.0 0.0

0. 0 1.0 0.0

0.452 0.548 0. 50

0. 50 0.50 0.50

0.50 0.50 0.50

MO. RIV. MO.

ATM 1111

OF 51

DATE

OZC! Error

Cold Wall

1.0 0.0

1.0 0.0

0. 50 9.5

0. 50 0.0

o. 50 o.o

~ ' ~' f "'

~- ~ , a

Condition

Design Limit

Noon

Acceptance Noon


--·/ ' TABLE 9. 2

SOLAR SIMULATION ON QUAL AND

FLIGHT THERMAL SURF ACES

Surface Solar Load (Suns)

Second S-13-G Surface

Mirror

1. 25 0.472

1.0 0.378

lATM 1111 I ......

PAGI '40 0, 51

OATI

Aluminized Teflon

Masking

0.431

0.345

150

125

100

1L

I 75 LLJ a: ::> 1-a: a: LLJ 50 Q..

::1: w 1-

25

0

-25 0

Figure 9. 4

[!) Jil(-111 LHS ELECTRONICS TEHP.

I A

) ~u --

. 20 110 60 80

BENDIX AEROSPACE SYSTEMS DIVISION ALSEP ARRAY E THERMAL VACUUM ACCEPTANCE TEST

ZERO TIME : 0000 HOURS ON 07/01/72

Acceptance Noon- 1. 0 Sun /

Page 4i of 51

. I

. Operate Power 7. 7 watts H- Acceptance Night ~ Analyzer Power 1. 69 watts I • Operate Power 7. 7 watts

Lunar Surface 250°F i Lunar Surface -3000y " . . .. \

"\ r ~ 1\ - ....

v ' \ A ~

\ r JV I

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\ J ' ~ ~ Mlr I&'

100 120 1110 160 180 200 220 2110 260 280 ELAPSED TIME - HOURS

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Figure 9. 5/ /, r <<'

/

BENDIX AER~SPACE SYSTEMS OIVISICN ALSEP ARRAY 'E' THERMAL VACUUM ACCEPTANCE TEST·

ZERO TIME • 000001 OF 07/01/72<

. Page 42 o£;51

m0~8 LKS ANALYZER SIMULATOR TEMPERATURE NO. 1 ~0~9 L"S ANAlYZER SJ"ULATOA TE"PERATUAE NO.2

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.

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ATMlllll


PAGE

DATI!

43 OF sy

resulted in a baseplate temperature 4Z oF warmer than no power dissipation. During the Flight lunar night operating mode, this dissipation took place in a simulator removed from the experiment. Removing this dissipation and the associated colder baseplate results in a radiator plate l0°F colder than with the actual baseplate.

9. 3 Comparison of Data and Correlated Temperatures

Tables 6 and 7 include the test temperatures of critical modes for both Qual and Flight tests. Radiator plate temperatures were correlated to within 3 oF for all conditions. The Flight Test radiator plate temperature was not actual data, but was extrapolated from Qual data of the radiator plate thermocouple and HK-41 thermistor. This temperature difference was +6 °F during lunar noon and +3 oF during lunar night.

I f'age 44"Xof 5 I

TAB ..... 9. 3

LUNAR MASS SPECTROMETER QUALIFICATION THERMAL VACUUM TEST DATA AND CORRELATION

Test Condition

Design Limit Noon Acceptance Noon

Location Test Analytical Test Analytical Data Data Data Data

Radiator 120°F 117°F 111 °F 109 °F Plate

HK-41 127° 123° 117 ° 115 ° Radiator Plate Internal

Analyzer 140° 152° 121° 129 ° Baseplate

Analyzer 158° 153° 136° 129 ° Cover

Electronics 165° 152° 143° 129 ° Cover End

Electronics 1170 2) 151° 960 2) 129 ° Cover End

Thermal 145° 151° 124° 127° Cover End

Electronics 224° 3

) 279 ° 184° 3}

249 ° Cover Bottom

Thermal t Cover Bottom

203° 3 > 272° . 184 ° 3 ) 243°

""

1 )These temperatures were still changing at a rate of 3° F /hr or greater.

2)T /C 08 ran consistently low, T /C 06 is more representative of the cover end temperature.

// 3

)The cover bottom was resting on an a1urninwn box standoff.

Qual Night

Test Analytical Data Data

0°F 0°F

30 30

-108° 1 > / - 99 °

-200 ° -230°

-194 ° 1 ) -252°

-200° 1 ) -258°

-101° 1}

-191°

-230° -269 °

-127° l) -204°

i

I I

Page 4.5,of 5:1 ,-

'- TABLE 9.4

LUNAR MASS SPECTROMETER FLIGHT ACCEPTANCE THERMAL VACUUM TEST DATA AND CORRELATION

Test Condition

Acceptance Noon Acceptance Night

Location Test Analytical Test Analytical Data Data Data

HK-41 111 °F 108°F - uoF Radiator Plate Internal

(1) Radiator Plate 105 ° 102° - 14°

Analyzer 140° 129 ° -120 ° Baseplate

!.....___ ____ ~

( 1) This temperature was not an actual test point - it was extrapolated from the radiator plate to HK-41 temperature gradient of Qual T/V.

(2) An analyzer simulator was utilized eliminating 1. 69 watts of power dissipation on the analyzer baseplate - expected temperature with analyzer dissipation will be -4°F.

Data

-11°F

...

(2) - 14°

-131°

·-

~--_Aerasp ace / --9pterne Dlvl•lan


SECTION III

LUNAR PREDICTIONS

MO.

ATM 1111

, ... DATI

This section summarizes the lunar surface analysis of the correlated thermal model.

ltiV. NO.

, .. OP .2!.:_

_Aamspaae -~DIWIIon


10.0 LUNAR PREDICTIONS

1 o. 1 Nominal Operation of LMS with Clean Thermal Surfaces

ATM 1111

PAGI

DATI

The operating temperatures of the 40 in. 2 radiator plate of an LMS dissipating 7. 7 watts internally is l10°F during lunar noon and -4°F during lunar night. This assumes all surfaces are clean and non-degraded. The radiator plate temperature during the survival mode dissipating 8 / watts is -17 °F.

Figure 16 indicates the day/night operating temperature of a nominal experiment as a function of power dissipation and radiator area.

10.2 Degraded or Dust Covered Radiator and Cover Su:z:faces

Figure 17 indicates the radiator temperature as a function of solar angle for various degraded surfaces. These plots assume the sun is incident on the side of the analyzer cover. The first plot indicates temperatures of an experiment with all surfaces totally dust covered with the exception of those initially covered by the dust cover. The parametric studies include various states of mirror/radiator plate degradation from the nonti.nal a= 0. 08 to a= 0. 3. Dust coverages (percent areas covered) corresponding to these Qt

1s are indicated in Figure 18. For example an Qt = 0. 3 is comparable to 22.5% dust coverage of mirrors and masking.

The second series of plot indicates the effects of radiator surface degradation assuming all other surfaces clean of dust.

Using +125 °F as maximum allowable radiator temperature, a "clean" experiment could have its mirror degraded from Ql = 0. 08 to Ql = 0. 133 before exceeding the allowable temperature. An experiment with non-protected surfaces 100% dust covered could have mirror degradation

OF SJ/

from Ql = 0. 08 to a= 0. 105 before exceeding the +125 oF allowable temperature limit.

10.3 Lunar Noon Non-Operating Conditions

1 o. 3. 1 Dust Cover On

The equilibrium temperatures of the LMS with dust cover intact

I -· ' LUNAR1 MASS SPECTROMETJ!;k RADIATOR TEMPERATURE AS A FUNCTION OF SOJ.AR ANGLE

;:

,.J

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:3 l 100 ·ll. ~

2 :s l:l

75 ~-...

90

Figure

-.' I •

'ALL SURFACES EXCEPT RADIATOR PLATE TOTAL.L.Y LJ)US]' DEGRADED (P' •.1.0)

RADIATOR AREA 40 SQUARE INCHES 1

RADJATOll PLATE IS DEGRADED AS INDICATED

EFFECTI VE MIRROr< ABSORBT NCE

..,..__ - ..,. ·n lA --r-- A. .0. 0· .. ~ Cl 0.2 lo·

-~ "7 o. 0 ' ~ -- -- -........ "' ~,

I

' ........, _""-,. '

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~ "" ~ -· ..

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~(), 30 ·-e

'ALL-SURFACES ARE CLEAN SURFACES r .•. ~

IAN~

,..__ 0

I .,

~0 '0 0

':' ;I' SOL.AR ANOLE (!>F.OlU;~:S)

'to.:t' , ,)

48 Page of' 51

1.0

0.8

0.6

f,:t;.l 0.4 u z ~ E-4 ~ ~ 0 ~ 0.2 ~

0

EFFECTS OF DUST ON THE SOLAR ABSORBTANCE OF Figure 10.2

THE LMS RADIATOR COATINGS

,-----

AIJUMINIZED EFLON

--- ····-··---T-·-----

I

I I

·--·------1- ~ -~-.--L-- ~-------

0 20 40 60 80

PERCENT DUST COVERAGE (o/o)

100

Page 49 of S')r· .. ~· /' I

s


ftUo

ATM 1111

PAGI

DATI

OF

are; for clean surfaces the radiator plate will attain + 107 • F, for dust cover and

~ I 31·

all surfaces 100% dust covered the radiator plate will attain a +22s•F temperature.

10.3.2 Radiator Exposed - no Power Dissipation (Lunar Noon)

A nominally clean experiment at equilibrium, with no internal power dissipation, will attain an +27 oF radiator plate temperature at lunar noon.

10. 3. 3 Deployment at a 20° Latitude

Deployment of the LMS at a 20° latitude has the effect of decreasing the lunar noon operating temperature from+ llQOF to 105°F. This results from the combination of the lunar surface decrease in temperature and the sun no longer normal to the radiator plate.

,~ .. --/ ~-lamaDiulliiDn


REFERENCES

ATM 1111

51 PAGI

DATI

1) AL 900132 "Performance and Design Requirements for Lunar Mass Spectrometer Experim@t Subsystem for ALSEP Package System" Bendix Aerospace Systems Div.

2) TP 2365582 "ALSEP Array E Deployed System Design Limit Thermal Vacuum Test Procedure", Bendix Aerospace Systems Div.

3) TP 2365575 "ALSEP Array E Flight System Deployed Thermal Vacuum Test Procedure", Bendix Aerospace Systems Div.

4) ATM-1097 "Lunar Mass Spectrometer Design Verification Thermal Vacuum Test", Bendix Aerospace Systems Div.

5) LED-520-lF "Design Criteria and Environments for the LM", Grumman Aircraft Engineering Corp.

OF 51

ATM 1111 J PAGE op_L

. Attl!Jita IP IIDe ~'~\~Dhllllaft


DATI

APPENDIX A

THERMAL MODEL NODES

1 Second Surface Mirror 72 Support Structure 2 Second Surface Mirror 73 Baseplate 3 Second Surface Mirror 74 Ribbon Cable 4 Second Surface Mirror 75 Backup Heater 5 Second Surface Mirror 76 Baseplate 6 Second Surface Mirror 77 Baseplate 10 Masking 78 L-Shaped Clip 11 Radiator Plate 79 Baseplate 12 Support Structure Isolator 80 Analyzer Cover 20 Masking 81 Analyzer Cover 21 Radiator Plate 82 Analyzer Cover 28 L-Shaped Clip 83 Analyzer Cover 30 Masking 84 Analyzer Cover 31 Radiator Plate 85 Electronics Cover 32 Support Structure Isolator 86 Electronics Cover 40 Masking 87 Electronics Cover 41 Radiator Plate 88 Dust Cover 42 Support Structure 89 Electronics Cover 50 Masking 90 Analyzer Cover 51 Radiator Plate 91 Analyzer Cover 55 Thermal Cover 92 Analyzer Cover 56 Thermal Cover 93 Analyzer Cover 57 Thermal Cover 98 Connector 58 Thermal Cover 100 Mother Board 59 Thermal Cover 101 Data Compressor Board 60 Masking 102 Data Compressor Board 61 Radiator Plate 103 Signal Conditioner Board 62 Support Structure 104 Signal Conditioner Board 63 High Voltage Power Supply Support 105 Housekeeping Multiplexer Board 64 High Voltage Power Supply Support 106 Housekeeping Multiplexer Board 65 Thermal Bag 107 Preamp/Discriminator Board 66 Thermal Bag 108 Preamp/Discriminator Board 67 Thermal Bag 109 Preamp/ Discriminator Board 68 Thermal Bag 110 Control and Monitor Board 69 Thermal Bag 111 Emission Control Board 70 Baseplate 112 Emission Control Board 71 Support Structure 113 Programmed Sweep High Voltage

Board


APPENDlX A (CONT.)

114 Programmed Sweep High Voltage Board 115 Low Voltage Power Supply 116 High Voltage Power Supply Base 117 Flat Ribbon Cable 118 Flat Ribbon Cable 119 Flat Ribbon Cable 120 Mother Board 121 Interconnect Board 122 Interconnect Board 127 Analyzer Connector 128 Relay Mount 129 Analyzer Wire Connector 130 Mother Board 131 Relay Mount 140 Mother Board 143 Aluminum Card Support 144 Aluminum Card Support 161 High Voltage Power Supply 162 High Voltage Power Supply 163 High Voltage Power Supply 164 High Voltage Power Supply 165 High Voltage Power Supply 166 High Voltage Power Supply 167 High Voltage Power Supply 168 High Voltage Power Supply 199 Lunar Surface 200 Space 201 Radiator Plate Cable Clip 202 Analyzer Plate Cable Clip 204 Analyzer Cable 206 Analyzer Cable 208 Analyzer Cable 210 Analyzer Cable

ATM 11111 PAGI 2 0, 2

DATI

A

l .. j

Appendix B

------------------------

--, Node '

-------------,-------------------~-------------

Initial Temperature

Table Number (Heat Input)

Product of (Abeortrtanc:e I< Abeorbing Area)

-----·-- ·-----

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