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FROM:
AUTHORITY
THIS PAGE IS UNCLASSIFIED
ADB008182
Approved for public release; distribution isunlimited.
Distribution authorized to U.S. Gov't. agenciesonly; Test and Evaluation; NOV 1974. Otherrequests shall be referred to Air ForceMaterials Laboratory, Attn: NonmetallicMaterials Division, Elastomers and CoastingsBranch, AFML/MBE, Wright-Patterson AFB, OH45433.
AFWAL Notice dtd 3 Nov 1983
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AFMLTR75-17
ML-101 THERMAL CONTROL COATING SPACEFUGHT EXPERIMENT
i
cc ELASTOMERS AND COATINGS BRANCH NONMETALLIC MATERIALS DIVISION
AUGUST 1975
r
TECHNICAL REPORT AFML-TR-75-17 FINAL REPORT FOR PERIOD JANUARY 1972 - JANUARY 1974
Distribution limited to U.S. Government agencies only; (test and evaluation), November £974. Other requests for this document must be referred. 4e the AifTofce Materials Laboratory, Nonmetallic Materials Division, Elastomers and Coatings Branch, AFML/MBE, Wright-Patterson AFB, Ohio 45433.
AIR FORCE MATERIALS LABORATORY AIR FORCE WRIGHT AERONAUTICAL LABORATORIES Air Force Systems Command WrlghtPctterson Air Force Bcse, Ohio 45433.
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NOTICE
Whan Government drawings, specifications, or other data are used for any purpose other than in connection with a definitely related Government procurement operation, the United States Government thereby incurs no responribility nor any obligation whatsoever: and the fact that the government may have formulated, furnished, or In any way supplied the said drawings, specifications, or other data, is not to be regarded by implication or otherwise as in any manner licensing the holder or any other person or corporation, or <x>nveying any rights or permission to manufacture, use, or sell any patented invention that may in any way be related tliereto.
This report was submitted by the Elastomers and Coatings Branch, Nonmetallic Materials Division, Air Force Materials Laboratory, Air Force Systems Command, Wright-Patterson Air Force Base, Ohio under job order 73400703. Daniel E. Prince (AFML/MBE) was the laboratory project engineer.
This technical report has been reviewed and is approved for publication.
_3 *£U~JL £ 6 Daniel E. Prince Project Monitor
FOR THE COMMANDER
M. LAMinges ChiefAElastomers and Coatings Branch MonmetVllIc Materials Division
Copies of this report should not be returned unless return is required by security considerations, contractual obligations, or notice on a specific document. AIR lORCl - 12-12-75 - 200
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UNCLASSIFIED SECURITY CL ASSIFICATION Or THIS PAGEfWn»/! Dal« Enltrtd)
Block 20 (continued)
Based on selected data from over 5000 revolutions covering a period of one year, it was found that all the coatings initially degraded to a greater degree than expected, possibly due to contamination. The most stable coatings were optical solar reflectors and the least stable was a white q-AUO^
pigmented coating. An experimental fabric material showed greater stability than state-of-the-art white coatings.
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AFML-TR-75-17
FOREWORD
This report was submitted by the Elastomers and Coatings Branch,
Nonmetallic Materials Division, Air Force Materials Laboratory, Air Force
Systems Command, Wright-Patterson Air Force Base, Ohio. The work was
performed under Project 7340, Job Order 73400703. Daniel E. Prince
(AFML/MBE) was the laboratory project engineer.
Contributions by Lt. M. Blessing, AFML/MBE, to data reduction and
analysis were greatly appreciated.
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AFML-TR-75-17
TABLE OF CONTENTS
SECTION PAGE
I INTRODUCTION 1
II SATELLITE ORBITAL PARAMETERS AND EXPERIMENTS 3
III STP FLIGHT P72-1 SATELLITE SYSTEMS 5
IV EXPERIMENTAL PACKAGE 7
V EXPERIMENTAL THERMAL CONTROL COATINGS 9
VI OPTICAL PROPERTY DETERMINATION AND ENVIRONMENTAL EXPOSURE 12
VII COMPUTER DATA ANALYSIS 14
1. Heat Balance Equations 14
2. Runge-Kutta Integration Technique and Computer Program MB4 16
3. Assumptions 18
4. Development of Time-Temperature Profiles 19
VIII DISCUSSION OF RESULTS 20
IX CONCLUSIONS AND RECOMMENDATIONS 23
1. Conclusions 23
2. Recommendations 23
REFERENCES 24
APPENDIX I COMPUTER PROGRAM MB4 25
APPENDIX II COMPUTER PROGRAM DAN 30
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AFML-TR-75-17
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LIST OF ILLUSTRATIONS
FIGURE
1
2
3
11
12
13
14
15
16
17
ML-101 Sample Holder
ML-101 Sample Holders Attached to 72-1 Satellite
Beckman DK-2A Spectrophotometer with a Gier- Dunkle Absolute Integrating Sphere
Gier-Dunkle Model DB100 Portable Infrared R:flectometer
AFML In-Situ Exposure Device
ML-101 Simulation Experiment S-13G (500 EUVSH)
ML-101 Simulation Experiment Si0? Fabric (500 EUVSH)
ML-101 Simulation Experiment FEP/A1 (500 EUVSH)
ML-101 Simulation Experiment Zn?Ti0,/0I650 (500 EUVSH) ~ 4
ML-101 Simulation Experiment PV-100 (500 EUVSH)
ML-101 Simulation Experiment TiO?/RTV602 (500 EUVSH)
ML-101 Simulation Experiment .008" OSR (500 EUVSH)
ML-101 Simulation Experiment .050" OSR (500 EUVSH)
ML-101 Simulation Experiment Eu203/RTV602 (500 EUVSH)
ML-101 Simulation Experiment A1203/RTV602 (500 EUVSH)
Sensor Temperature vs Time for S-13G During Revolutions 46 and 4162
Sensor Temperature vs Time for SiOp Fabric During Revolutions 46 and 4162
PAGE
38
39
40
41
42
43
44
45
46
47
48
49
50
51
52
53
54
VI
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AFML-TR-75-17
LIST OF ILLUSTRATIONS (Continued)
FKURE PAGE
18 Sensor Temperature vs Time for FEP/A1 During Revolution: 46 and 4162 55
19 Sensor Temperature vs Time for Zn?Ti0./0I65] During Revolutions 46 and 4162 4 56
20 Sensor lemperatuie vs Time for PV-100 During Revolutions 46 and 4162 57
21 Sensor Temperature vs Time for 3M Black Velvet During Revolutions 46 3nd 4162 58
22 Sensor Temperature vs Time for Ti0?/Silicone During Revolutions 46 and 4162 59
23 Sensor Temperature vs Time for OSR (0.008) During Revolutions 46 and 4162 60
24 Sensor Temperature vs Tine for OSR (0.050) During Revolutions 46 anJ 4162 61
25 Sensor Temperature vs Time for Eu?0,/Silicone During Revolutions 46 and 4T62 6 62
26 Sensor Temperature vs Time for a-AI^CL/Silicone During Revolutions 46 and 4162 63
27 Sensor Temperature vs Time for Temperature Reference During Revolutions 46 and 4162 64
28 Resistance vs Time for Reference Resistor (499 Ohm) 65
29 Comparison of Simulated and Flight Data for S-13G 66
30 Comparison of Simulated and Flight Data for Si02 Fabric 67
31 Comparison of Simulated and Flight Data for FEP/A1 68
32 Comparison of Simulated and Flight Data for Zn2Ti04/0I650 69
33 Comparison of Simulated and Flight Data for PV-100 70
34 Comparison of Simulated and Flight Data for Ti0?/Silicone 71
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35
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Comparison of Simulated and Flight Data for OSR (0.00S)
omparison of Simulated and Flight Data for OSR (0.050)
Comp-ins ■).! of Simulated and Flight Data for Eu90 , 'Si :ioone
'.ompariscn of Simulated and Flight Data for -A'LO,/Silicons
PAGE
72
73
74
75
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AFML-TR-75-17
LIST Or TABLE:
TABLE
I Satellite C.bital Parameters
II Satellite Payloads
III Initial Data Inputs
IV Initial vs Space Solar Absorptance
TAGE
13
IX
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LIST Of SYMBOLS
SYMBOL
A p
S F e-m
Fe-s
G
H_
ne
TS
ref
as
,xm
e
(th
r ■n
f n
DESCRIPTION
Area of edge of sample
Cross-sectional area of sample
Specific heat of sample
Shape factor from edge of sample to Mylar holder
Shape factor from edge of sample LO space
Solar heating constant.
Incident planetary heating flux
Incident planetary albedo heating flux
Incident solar heating flux
Current
Thermal Conductance of sample
Mass of sample
Net heat exchange rate between sample and Mylar-
Total solar heat incident on sample edge
Temperature of sample
Temperature of adjacent Mylar
Temperature of adjacent samples
Temperature of sample holder base
Absorptance of sample
Absorptance of adjacent Mylar
Emittance of sample edge
Emittance of sample
Emittance of Mylar holder
Emittance of adjacent sample edges
Stephan-Boltzmann constant, 0.17 14x10
x
UNITS
ft'
ftc
BTU/lbm-°R
BTU/ftd -hr
BTU/ft2 -hr
BTU/ft? -hr
BTU/ft2 -hr
Amps
BTU/hr- R
lbm
BTU/hr
BTU/hr
"R
rR
f>R
°R
BTU/ft?-rR4
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AFML-TR-75-17
SECTION I
INTRODUCTION
The temperature control of spacecraft is one of the most challenging
technical problems confronting spacecraft designers today. The ultimate
objective of thermal design is to ensure that the spacecraft will operate
within the prescribed temperature ranges defined by the temperature
limitations of the spacecraft's materials and components, such as detectors,
batteries, solar cells, and electronics, which operate properly and
efficiently only if maintained within specific temperature ranges.
The surface temperature of a spacecraft depends primarily upon
the solar radiation absorptance (a ) of the surface, the radiation of
solar energy from the surface or surface emittance (< ), and the
generation of heat in *';(? body of the spacecraft. Secondary parameters
that influence the sun •■ temperature of a spacecraft are material
properties such as then . conductivity and specific heat of components.
Solar absorptance is defined as the ratio of solar energy absorbed by
a surface to that absorbed by a blackbody, while emittance is the ratio
of the energy radiated from a surface to that radiated from a blackbody
at the same temperature.
Two techniques are used for regulating spacecraft temperatures:
active systems and passive systems. Active systems usually employ
electrical power and moving parts which can be adjusted providing a
constantly varying surface in terms of optical properties. Passive
systems, in contrast, require neither electrical power nor moving parts
and rely on surfaces with appropriate thermophysical properties; namely,
solar absorptance, a and infrared emittance (c). Regardless of what
combination of systems is used, the thermophysical properties of the
entire spacecraft surface must be known since it is the average surface
properties which determine the spacecraft equilibrium temperature.
i i
Because of the large role surface coatings play in the passive
temperature control of spacecraft, extensive programs have been conducted
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AFML-TR-75-17
to develop space-stable, low a /e coatings. Future spacecraft requirements
call for the development of coatings with an initial solar absorptance
of less than 0.10 and initial emittance of more than 0.90, with long-term
stability (5-10 years) in the space environment; i.e., increase in ac
of less than 10i. Low a /c surfaces can be divided into three division: s
(1) those represented by diffusely reflecting white pigmented coatings,
or paints; (2) those represented by second surface mirrors that obtain
heir low solar absorptance from the metallic reflection of an aluminum
or silver mirror and their emittance from a selectively transmitting
substrate such as quartz or FEP Teflon; and (3) those represented
by newly emerging white fabric cloth materials.
Many of the previous materials have been extensively subjected to
in situ laboratory simulation exposures, however, most important to the
thermal designer is knowledge of actual spaceflight data and its
correlation with laboratory data. This report describes an Air Force
Materials Laboratory coatings spaceflight experiment wl.ose primary
objective was to determine the effects of the space environment on the
equilibrium-radiative properties of selected candidate temperature
control coating systems. Its secondary objective was to correlate the
actual space degradation with simulated laboratory data.
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AFML-TR-75-17
SECTION II
SATELLITE ORBITAL PARAMETERS AND EXPERIMENTS
The Space Test Program (STP) is a Department of Defense program
for the spaceflight of research and development payloads as well as
operational payloads. The STP P72-1 flight vehicle was launched from
\/ander,berg Air Force Base on an Atlas-F/BII combination. Liftoff
occurred at 2012.5 Z on 2 October 1972. The satellite orbit was sun-
synchronous with a nominal inclination of 98.34 degrees (polar) and a
nominal circular altitude of 400 nautical miles. The initial orbital
parameters and the orbital parameters as of September 30, 1973 are
compared to the planned nominal parameters for the 72-1 vehicle in Table I.
TABLE I
SATELLITE ORBITAL PARAMETERS
Planned Flight Parameters Nominal Actual (Initia 1) A ctual (Sep 197 3)
Maximum altitude (nm) 409.19 411.41 411.21
Minimum altitude (nm) ^98.13 400.37 399.27
Eccentricity 0.00041375 0.00039 0.00055995
Inclination (deg) 98.348 98.424 98.433
Period (min:sec) 99:40.158 99:39.465 99:30.217
Rev. No. 1 5191
The 72-1 satellite included a payload of six experiments. These
ore listed in Table II together with the type of experiment, principal
investigator, and appropriate organization.
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AFML-TR-75-17
TABLE II
SATELLITE PAYLOADS
Experiment No.
ML- ,\
RTD-802
ARPA-501
NNRL-114
SSD-988A
SSD-988B
Investigator fo Organization Dose ription
D. E. Prince Thermal Control Air Force Materials Laboratory Coatings
A. Y. Harper Ab'MDA-Huntsville
Radar calibration target
J. B. Reagen Gamma- spectrometer Lockheed Research Laboratory
C. S. Weiler Extreme UV radiation Naval Research Laboratory
D. R. Croley, Jr. Aerospace Corporation
P. Rothwell AFCRL: Hanscom Field
Low altitude particles
Low altitude particles
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AFML-TR-75-17
SECTION III
STP FLIGHT P72-1 SATELLITE SYSTEMS
1. ELECTRICAL POWER SYSTEM
The satellite electrical power is supplied by a solar cell array
that generates electrical energy required to operate the satellite sub-
system and experiments and to charge two nickel-cadmium batteries that
supply electrical power during solar eclipse. The electrical power is
distributees directly from the main power bus or switched by a relay control
box by uplink command.
2. TELEMETRY, TRACKING, AND COMMAND SYSTEM
The telemetry system transmits all data in PCM/PM at 8 or 64 Kbps
real-time and recorded data at 64 or 512 Kbps. A 2-watt SGLS reference
transmitter is used for the 8 Kbps rate and a 10-watt data transmitter
is used for the 64 and 512 Kbps data rate. The processor furnishes
multiplexing, analog to digital conversion (where required) formatting,
clocking, and synchronizing signals for the TT&C systems including
the programmer. Three tape recorders are supplied for reliability and
high data capacity for unusually high peak load periods. All other
systems are dual redundant. A command system which is also dual
redundant, includes the Command Receiver Demodulator and the Decoder/
Programmer which provides 375 MHz system real-time uplink command capability
and stored program routines for flexible payload and telemetry system
control.
3. STABILIZATION, ORIENTATION, AND ATTITUDE DETERMINATION
The STP P72-1 satellite operates in a sun-synchronous orbit with
the spin axis nominally normal to the orbit plane within five degrees
with a goal of one degree. The satellite spin speed is controlled
within 12+2 rpm.
The Satellite Attitude Control System (ACS), makes use of magnetic
torque coils to control the attitude and spin rate of the highly stable
spinning vehicle. Body fixed earth sensors provide data for primary
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AFML-TR-75-17
determination of attitude and rate. Ground commands control the torque
coil current to provide correction of observed errors. Gobble damping
is provided by a passive fluid loop damper.
Measurement of the earth's magnetic field is provided by two
three-axis magnetometers mounted on the forward antenna boom. The
associated Magnetometer Electronics Assembly provides proportional
outputs that are multiplexed into the telemetry stream. The outputs
are used as a polarity sense switch for spin control and provide data
for the backup attitude and spin rate determination.
A dual winding iron-core precession coil is used to maintain the
spin axis within five degrees of the orbit normal. The coil is operated
continuously to provide orbit regression rate. Periodic updates by
ground control commands to the muititap precession coil eliminate any
buildup of errors caused by random disturbance torques or bias in the
satellite magnetic field. Attitude information is provided by a
redundant horizon sensor system. A spin rate control coil, identical
to the precession coil, s provided to update the satellite spin speed
on command from the ground. The coil is automatically commutated by
radial magnetometer output signals. Backup capability is provided by
cross strapping a backup magnetometer through the PCM Processors
to the dual spin coil windings.
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AFML-TR-75-17
SECTION IV
EXPERIMENTAL PACKAGE
£
The complete basic design and initial evaluation data are contained
in References 1 and 2. Sample holders used on the 72-1 satellite are
Numbers 7 and 12. Each sample holder weighs approximately 11 ounces
(300 grams), occupies approximately 40 cubic inches, and will accommodate
five test samples and one reference sample. The samples or thermal
control coating specimens are 15/16 inch in diameter with the thickness
depending on coating applied and substrate. The sensor disks used for the
sample holders are Trans-Sonics T4086-3 platinum resistance thermistors
cemented in stainless steel sensor disks. Each platinum thermistor
has been calibrated to an accuracy of +0.5°C at temperatures of -193°C,
O'X, and 100' C. The constants for the temperature resistance relationship
of platinum are supplied with each sensor and are used in translating
resistance to temperature. A sensor (Trans-Sonics T4086-3) is attached
to the bottom surface of the holders as a temperature reference for
correction of conduction losses from the sample sensors.
The sensors are supported upon 50 layers of super-insulation (1/4
mil Mylar aluminized on one side) by a continuous Dacron cord under
tension. The individual samples and sensors are thus thermally isolated
from each other with the primary conduction loss through the thin sensor
leads. A temperature sensor is placed adjacent to the electrical
connector to permit correction of flight data for this loss. Ten pounds
of tension is placed on the Dacron thread used to bind the sensors in
place during fabrication. Each sensor is held at three places by the
cord and is cemented to the cord at these points. The sensors rest upon
a three mil sheet of Mylar (aluminized side down) which acts as a cover
for the Dacron cord.
The measurement of the resistance of the sensors is obtained by
supplying an external voltage (90 millivolts) to the sensors and measuring
the voltage drop across each sensor. The voltage drop across each
thermistor is in turn a linear function of resistance which in turn is
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AFML-TR-75-17
a linear function of temperature for this range. Through the calibration
data for each thermistor, the temperature can be determined. Room
temperature calibration was performed in both real-time and tape
recorder modes in the pre-launch satellite acceptance test phase. This
calibration data obtained prior to launch together with the installed
flight hardware is being used for the data reduction and then calculation
of the desired thennophysical values.
Figure 1 shows the ML-101 sample holder with six coatings attached
and Figure 2 shows the ML-101 sample holders attached to the partially
completed 72-1 satellite.
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AFML-TR-75-17
SECTION V
EXPERIMENTAL THERMAL CONTROL COATINGS
A total of 12 thermal control coating samples were used on two
sample holders for the ML-101 spaceflight thermal control coating
experiment. The sample holders are designated 7 and 12 on all experiment
references.
7-1, ZnO (Silicate Coated)/Methyl Silicone (S-13G)
Formulation, processing, and application was performed by IITRI*
in accordance with IITRI Report 4-6053-5-7; Batch No. C-133. The
S-13G was sprayed to a thickness of 8.0 + 1.0 mils.
7-2, Si02 Cloth Fabric
2024-T3 unclad 15/16 inch diameter 0.032 inch thick aluminum
substrate was prepared as follows: (1) faced on 240 SiC stationary
grinder, (2) prepolished with jeweler's rouge and buffed, and (3) final
polished using Buehler's Magomet (M0) on Microcloth. The silica cloth
fabric is a 12 harness double layer satin weave woven from 450/1/0
Suprasil yarn. Polyvinylalcohol sizing was removed by heating, and
filament diameter size was six microns. The fabric construction consists
of 150 ends/inch by 240 picks/inch. The fabric thickness is nominally 2 0.005 inch and weighs 4.65 oz/yd . The fabric was bonded to the polished
aluminum substrate with General Electric SR-585 silicone adhesive and
the entire system vacuum outgassed at 375°F. I j i :
7-3, Aluminized 5 mil FEP Teflon Second Surface Mirror I ? . ,—,
This material was manufactured by G. T. Schjeldahl, Northfield, i -
Minnesota, T>pe G400900, lot number 12-33 and was supplied to the AFML
throurh the courtesy of Mr. A. E. Hultquist, Lockheed Missiles and
' Space Company of Sunnyvale, California. The material was bonded to a
15/16 inch diameter 2024 NaOH etched aluminum disc in the following
manner. Both the front side of the disc and the back side of the vacuum
* ITT Research Institute, Technology Center, Chicago, Illinois.
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AFML-TR-75-17
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deposited aluminum were primed with General Electric SS4004 clear silicone
primer and then 3M Company's Number 465 adhesive transfer tape was used
on both the disc and the back side of the vacuum deposited aluminum
layer and pressure applied by hand rubbing with 0.125 inch Teflon sheet.
7-4, Zn^TiO./Methyl Silicone
Pigment synthesis, coating formulation, and application was performed
by the Illinois Institute of Technology using Owens Illinois 650 methyl
silicone resin as a binder material. Ccating thickness was 5+1.0 mils.
7-5, PV-100
This material is a rutile Ti0? pigment in a silicone-alkyd binder,
better known as Vita-Var heat absorption paint, formulation PV-100,
15966 white. The coating was prepared on 0.040 inch thick, 15/16 inch
diameter 2024 aluminum disks after they were cleaned by solvent wiping
followed with a sodium hydroxide etch and water rinse. Strontium
chromate primer was applied per MIL-P-23377B to a thickness of 0.0005 inch.
The PV-100 topcoat was then sprayed to a thickness of 0.002 inch and air-
dried.
7-6, 3M Black Velvet
This is a two component polyester-epoxy paint, type 401-C10
prepared by dispersing three parts of Part A, Lot No. VC-39-125 and one
part of Part B, Lot No. 16. This was then spray applied onto 0.032
inch aluminum discs which had been sanded, solvent wiped, and primed
with G. E.'s SS4044 primer (thickness less than 0.001 inch). Thickness
of the coating was 0.006 T 0.001 inch.
12-1, TiOo/Methyl Silicone
This coating consisted of a Ti02 pigment (Dupont R-992, Lot No. 9980)
in purified RTV 602 (G1638-46 EFG from Hughes Aircraft Company under
Contract F33615-71-C-1155) polydimethylsiloxane A pigment to binder
ratio of 150/100 was used. Thickness was 5.5 + 0.5 mils.
10
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AFML-TR-75-17
12-2, Optical Solar Reflector, 0.008
This is a second surface mirror with the front surface consisting
of 0.008 inch of Corning 7940 quartz, and the back surface coated with
vacuum deposited silver, overcoated with Inconel and a final dielectric
back cover. These samples were supplied to AFML through the cooperation
of Mr. E. N. Börsen, Aerospace Corporation and Mr. A. E. Hultquist,
Lockheed Missile and Space Company. This material was bonded to the
aluminum disc using the same procedure as sample 7-3.
12-3, Optical Solar Reflector 0.050
Same material as 12-2 except, the Corning 7940 quartz was 0.050
inch thick. The same bonding procedure used on sample 7-3 was utilized.
12-4, Eu20,/Methyl Silicone
This coating consisted of a europium oxide pigment (American
Potash and Chemical, Code 1014, Lot No. D1031, silicate treated) in
purified RTV 602 (G1538-46EFG from Hughes Aircraft under Contract
F33615-71-C-l155) polydimethylsiloxane. A pigment to binder ratio of
270/100 was used. Thickness was 8.0 + 1.0 mils.
12-5, aAl203/Methyl Silicone
This coating was made from an cx-alumina pigment (#1013 from
Martin-Marietta Company under Contract F33615-71-C-1410) in unpurified
commercial RTV 602, Lot #242, polydimethylsiloxane. A pigment to
binder ratio of 200/100 was used. Coating thickness was 8.0 + 1.0 mils.
12-6, 3M Black
A duplicate 3M Black per sample 7-6.
I s. v
7-7 and 12-7 Reference Temperatures
Junction inside each sample holder.
54-6 Calibrated Resistor
A 499 ohm resistor in sample holder number 7, used as a check on
voltage variations.
11
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SECTION VI
OPTICAL PROPERTY DETERMINATION AND ENVIRONMENTAL EXPOSURE
Table III provides initial data for each coating, including mass,
emittance, solar absorptance, and specific heat. The solar absorptance
(a ) value for each coating was determined utilizing a Beckman DK-2A
spectrophotomcter with a Gier-Dunkle Integrating Sphere and a xenon
energy source (Figure 3). Integration of solar absorptance was accomplished
using 25 points in conjunction with F. S. Johnson Solar Spectrum. In-air
infrared spectral reflectance of each coating was measured in-house
utilizing a Gier-Dunkle Model DB100 Portable Infrared Reflectometer
(Figure 4). Subtracting the infrared reflectance from unity gave the
total normal emittance. The theory employed by this instrument has teen
described by Nelson, et al (Reference 3).
12
-
AFML-TR-75-17
Figure 5 depicts the AFML In Situ Exposure Device for simulation
of ultraviolet-vacuum space conditions. The system was operated
generally in the 10" torr region during radiation exposure. A xenon
lamp ultraviolet source was used with an ultraviolet sun rate of
approximately 1.25. The coating reflectance values were measured in
situ before and after ultraviolet radiation exposure utilising a Beckman
DK-2A spectrophotometer with an integrating sphere inside the expos ire
chamber. Figures 6 through 15 present typical reflectance curves, btfore
and after 500 EUVSH, for the coating samples flown. The pre-test and
post-test solar absorptance values are also given for each coating. ■
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AFML-TR-75-17
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33
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AFML-TR-75-17
SECTION VII
COMPUTER DATA ANALYSIS
1. HEAT BALANCE EQUATIONS
The heat balance equations, usir.g the list of symbols, for a single
sample of a thermal control coating on this experiment are:
Q (input) = Q (direct solar) + Q (conduction) (1)
+Q (sensor dissipation) + Q (reflected so'lar)
+Q (albedo) + Q (earth emission)
Q (output) = Q (emitted) + Q (exchange with surroundings) {?.)
+Q (storage)
The individual terms of Equations 1 and 2 are:
Q (direct solar) = a AH
Q (conduction) = K(Ts"Tref)
9 Q (sensor dissipation) = I R loss in sensor
Q (reflected solar) = reflected sunlight from
the Mylar upper surface and other areas of
surroundings = Q
Q (albedo) = cxsAnHr
Q (earth emission) = t ..A H
Q (emitted) = e^ffTj
Q (exchange with surrounds)
Q, . = A ( F CTT + A F off 4 [ne) e e e-s s e e-m e m(T
s
+ I A F of ( (T4 . T4)
e e-m en s n
Q (storage) = MC dT /de
T
(3)
(4)
(5)
(6)
(?)
(8)
(9)
(10)
(ID
14
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AFML-TR-75-17
Combining Equations 1-11 results in:
«sA 11 +K(T -T ) + Q H asA „ , r A H n s s rei rs n r tl n [3
' , A aJ + A ( F <?T4 4 A F off (T4 - T4 ) th n s e c' p"-s s c e-m c m1 s m'
= ' , A oT H A ' F
(1?)
where I R and SA F fff f (T - T ) are considered nealioible e e-rn e n s n ^1/4
terms. Letting T_
tor MC dT
cl9
«m(H +H )/t Ml s r m {
results in the following:
dT
J •! and solvinn
MC p de A (H + H )Qs + A F ( am (H l-H )
n s r c e-m e s r
f(A F ( ( +A f )H +Q tKT -OIA , e e-m . m n th' p rs ref n'th
(13)
4 KTS +A F ( f +A t F ) T - —y±- e e-m e m e e e-s s vii
Dividing by MC and defining the following:
A A -- A (H 4 H n s ) as + A F ( nm (H +H
r e e-m e s r
P (14)
+ A F ( ( H + KT .+ Q Hi A ( e e-m e m p rej r s
MC p nth
BB - K
MC :i5i
cc ff(An'th ■' A F ' < + A f F P. £n e e-m cm e e e-s
MC
(If)
upon substitution into Equation 13 gives:
I i:
dT
dfi — AA - BBT - CCT (17)
! 15
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ArML-TR-75-17
From Equation 15, it can be seen that. BB is a constant dependent
only upon property factors. Likewise tC in Equation 16 is dependent
upon the property factor of the particular sample in question. The term
AA in Equation 14 contains the heat inputs, property factors, and the
solar absorptance term which is the quantity to be calculated.
The resulting differential equation (Equation 17) cannot be
solved analytically because it is nonlinear due to the T term. It can s A
be solved by La Place transform methods after the fourth order T term J s is linearized by an e-nansion such as the Taylor series expansion with
higher order terms being neglected. However, errors result by neglecting
these higher order terms.
Another more accurate method is to use the Runge-Kutta integration
technique with the aid of the computer which is the method chosen
and explained in Section 2.
2. RUNGE-KUTTA INTEGRATION TECHNIQUE AND COMPUTER PROGRAM MB4
The fourth order Runge-Kutta integration method solves the
differential equation (Equation 17) by giving a. solution set of values
of the dependent variable, T , for several increments of the independent
Variable, 6.
The method uses a variable step integration with the interval of
integration being specified. An interval of 0.001 was used in this
analysis. This varible step integration is designed to keep the maximum
local relative error of the dependent variable less than 1 x 10" . The
method integrates from one point to another. The initial point is assumed
to be at 9 = 0 for the independent variable while the dependent value
of temperature at 0 = 0 is given as T . so
A program to solve the differential equation of the form of
Equation 17 using Runge-Kutta integration techniques W5S written and
appears in Appendix I. The program integrates from some initial
condition (Lc.^nerature T at time T zero) to some final condition
16
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AFML-TR-75-17
(temperature T . at time T check). The program then checks to see if
the temperature found from the integration is the same as that given
from telemetry readings to within 0.1°R. If the temperature does not agree i
the parameter AA is then adjusted by 1.0 unit until the temperature
agrees. This adjustment of AA by 1.0 unit was found to be the best factor
to use. Should the function begin to oscillate the agreement interval ■
is increased to 0.4°R and AA adjusted until that interval of agreement
is met.
After the integration is completed and the equation fit to match
that received from telemetry data, the values of AA (fit parameter)
and corresponding values of temperature calculated from integration and
compared to T . are printed out. The solution set of temperatures SC K
found for the fitted differential equation from time zero to some later
time is printed out. Then these can be checked against further points
from telemetry for a good overall fit if desired.
The portion of the T versus 6 telemetry curve that is fit, is the
initial portion where the effect of H should be minimized and the points
chosen for the fit should be the same for all the curves in which the j
a are to be compared. The portion of the curve used in this analysis
begins with time zero and temperature T at 0.155 hours into the orbit 3 r so after just passing into the sun. The final point for which the curve is
fit is at time 0.47 hours later or 0.625 hours into the orbit. The I
temperature T . is the temperature from the T versus 6 curve which
occurs at this time and is used as a check temperature for fitting
the differential equation solution to the telematry curve. Although
this relatively large interval allows more time for the effect of
changing H to enter the system, it permits more smoothing of integrated
data.
The reason the initial point of 0.155 hours into the orbit was
chosen was that some of the telemetry curves to be compared did not
have telemetry data until that time in the orbit. This is due to
tracking acquisition that occurred on certain revolutions. I
17
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AFML-TP-75-17
During the actual calculation, values of initial temperature T
and check temperature T . for each sample for a given revolution are bCK
obtained. These are supplied to the program in the following order:
Sample 7-6, 7-1, 7-2, 7-3, 7-4, 7-5, 12-1, 12-2, 12-3, 12-4, and 12-5.
The program then uses the data from the black sample 7-6 to fit the
integrated equation to the telemetry data and obtain the amount of
heat incident upon each of the samples. This is done by calculating
the amount of heat from the equation for a fixed alpha (0.985). This
can Le done because the solar absorptance of the black samples should
not change appreciably with exposure time. This amount of total heat
input is then used to calculate the alphas for each of the nonblack
samples after they are fit to their corresponding telemetry curves.
The alpha values are then printed out along with coating property
information used in the calculations.
3. ASSUMPTIONS
(a) The heat capacity, C , was assumed to be constant and the
values given in Table III were used as the correct values for each
sample.
(b) The value of the total solar heat incident on the sample edge,
Qrs, was constant at 0.0375 BTU/hr.
(c) The shape factor from the edge of the sample to Mylar holder,
F , and the shape factor from edge of sample to space, F , were e-m 3 r r e-s constant at 0.5.
(d) The planetary heating flux, H , was 18.1 BTU/hr-ft2. r
(e) The solar heating flux, H , and the planetary albedo heating
flux, H , were constant and found from sample 7-6 for each revolution.
(f) The constant H assumption is made because only the initial
portion of the temperature vs time curve is used where increasing H
should have the smallest effect. 4 4
(g) The term 2 A F at < (T - T ) is zero, indicating
neglibible interaction between adjacent sample edges.
18
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AFML-TR-75-17
2 (h) Sensor dissipation or I R loss of the platinum resistors is
taken as zero since the term at 500 ohms and 100 mv results in a
temperature rise of about 1/4UF, which is well within our temperature
error limits.
4. DEVELOPMENT OF TIME-TEMPERATURE PROFILES
The computer program presented in Appendix II was utilized to
convert raw digital data tapes for each required revolution directly
into time-temperatur» curves. The program utilizes ground based
sensor calibration data and sample holder conduction coefficients to
convert millivolts to resistance to temperature. The program also plots
directly the reference temperature and the resistance of the reference
calibrated resistor. Typical plots for all sensors are presented in
Figures 16 through 28.
19
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11 AFML-TR-75-17
SECTION »'III
DISCUSSION OF RLSULTS
Table IV presents the solar absorptance changes for the coatings
investigated. Column one indicates the initial laboratory measured
solar absorptance, column two the initial in space calculated solar
absorptance (calculated during revolution number 46 after the satellite
had stabilized), and column three the calculated in space solar absorptance
after one year (5191 revolutions, 2160 EUVSH).
One of the most important features to be noted in Table IV is that
the initial in space solar absorptances of the two OSR's are twice those
measured in the laboratory. This indicates some form of contamination
on the surfaces. This contamination, from all indications, was present
even during revolution number one, indicating contamination during
liftoff as well as contamination during the initial revolutions from
outgassing of volatile constituents from the satellite materials.
The effects of contamination on the initial in space sclar absorp-
tance values for the other coatings appear much smaller with the greatest
change being for FEP/A1. The OSR coatings and the FEP/A1 coating were,
undoubtedly, affected the most because of their low initial solar
absorptance which produced low surface temperatures, therefore attracting
condensable contaminants. Also their low initial solar absorptance
values were the most vulnerable to changes caused by small amounts of
contaminant. The two coatings, EUpO^/m-silicone and a-Ai„0^/m-silicone
both had initial in space solar absorptances which were lower than those
measured in the laboratory. The reasons for this are not clear, but
these initial calculated values were used for determining the total
solar absorptance changes in Figures 37 and 38.
Figures 29 through 38 present a comparison of laboratory simulated
data and flight data for all the coatings flown, except for 3M black
velvet. These figures also indicate the rate of change in solar
absorptance over the one year exposure period. The flight L c
20
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AFML-TR-75-17
in these figures is the change in coating solar absorptance when compared
to the initial calculated in space solar absorptance, not the initial
laboratory values. For example, although the 0SR(0.008) increased by
0.087 from laboratory to initial in space, it only increased another
0.025 after one year exposure in space and this is the A a reported.
Each flight curve represents the compilation and analysis of 200
revolutions of data between revolution 1 and revolution 5191. The data
scatter was extremely small, primarily because of the satellite stability
in space and the large amount of data received per unit time (in this
experiment each sample was pulsed every 0.9 seconds), which allowed for
excellent computer data averaging. The actual plots of time-temperature
curves from the raw data tapes showed minimal data scatter and maximum
consistency from one revolution to the next. Over 800 data tapes were
received during the first year which allowed for the selection of
revolutions of primary interest.
From Figures 29 through 38 it can be seen that, in all instances,
after 500 hours, the flight coatings were degrading more rapidly than
the laboratory tested coatings. As indicated earlier, this can best be
explained by the synergistic effect of contamination. If, as it appears,
all surfaces were contaminated, the degradation of the coatings consisted
of a combination of coating solar degradation DIUS solar degradation of
the contaminant on the coating surface. Results from the Air Fc""ce
Materials Laboratory D024 thermal control coatings experiment aboard
SKYLAB, where samples were returned to earth from space and evaluated,
clearly indicated the detrimental effects of contamination on thermal
control coating stability. In all instances where contamination occurred
during the SKYLAB exposure, solar degradation of the coatings increased
much more rapidly than expected from laboratory tests.
Figures 30, 31, 35, 36, and 37 for Si02 Fabric, FEP/A1, two OSR's
and EUpO.,, all indicate an initial solar absorptance increase during the
early revolutions with a gradual leveling of degradation with continued
exposure. This result indicates a trend toward saturation of degradation
21
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AFML-TR-75-17
for these coatings, however, additional data collection beyond one year
is needed to confirm this as well as determine the degradation saturation
point for all of the coatings.
One further note relative to the figures is that in Figure 38 the
ordinate has been increased, because the degradation of the a-Al?CL
coating was much more severe than any other coating.
22
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AFML-TR-75-17
SECT'uN IX
CONCLUSIONS AND RECOMMENDATIONS
1. CONCLUSIONS
a. Some form of contamination, most likely from outgassing materials
from the satellite, caused flight degradation of the thermal control
coatings to be more severe than ground simulated degradation.
b. Of all the coatings flown, the Optical Solar Reflectors
(OSR's) were the most stable. No significant difference in degradation
was observed for the two OSR's of different thicknesses. This would
indicate that the solar degradation was primarily a surface phenomenon
rather than a bulk phenomenon.
c. FEP/A1 was very stable wi.h a Aa = 0.04 after one year at
400 nautical miles (EUVSH = 2,160 hrs).
d. Of the white coatings flown, EUpO^/m-silicone was the most
stable (Aa = 0.08) and Al^Oo/m-si1icone was the most unstable (Act = 0.20).
e. The silica fabric material flown was more stable than the white
coatings with a Aa of 0.08 after one year exposure.
2. RECOMMENDATIONS
a. This data could also be used to calculate changes in coating
emittance. This was not done due to time limitations.
b. Additional thermal control coating experiments should be
flown at or near synchronous orbit to more closely simulate the
environment of operational Air Force satellites.
c. The STP 72-1 satellite has continued to function satisfactorily
even after one year of orbit life. Data from the ML-101 experiment
should continue to be analyz^j to determine the effects of long-^erm
space exposure on these t'iermal control coatings.
d. Correlation of t,,e ML-101 degradation data with data obtained
from the AFML D024 experiment aboard the manned SKYLAB orbital workshop
should be attempted.
24
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AFML-TR-75-17
2.
3.
REFERENCES
RTD-TDR-63-4269, Design and Construction of Sample Holders for Orbital Temperature Control Coatings Experiment, Vol. 1, Design, Analysis, and Test Results, TRW Systems, April 1964.
RTD-TDR-63-4269, Calibration Data and Drawings, Vol. 2, TRW Systems, April 1964.
K. E. Nelson, E. E. Luedke, and J. T. Bevans, Journal of Spacecraft Rockets, Vol. 3, No. 5, p. 758, 1966.
25
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3
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APPENDIX I
COMPUTER PROGRAM MB4
XP( ) = Theta printout and storage array.
YP( ) = TS(1) printout and storage array.
TS (1) = Temperature of sample at time theta.
P (1) - Highest derivative value.
ACHECK ( ) = Storage array for AA values.
TSCHEK ( ) = Storage array for TS (1) values calculated and compared to TSCK.
AA = Constant in differential equation which is changed to fit different equation solution to telemetry curve in
(—- = AA - BBTs - CCTs4).
BB, CC = Other differential equation constants.
NSC = Sample identifier label check for sample 7-6.
TCHECK = Time value at which differential equation solution check is to be made with curve from telemetry data.
REV = Revolution number.
TREF = Reference sample temperature.
G = Solar heating constant.
NEXEC = Counter of number of times calculation loop is executed.
NS = Sample identifier label.
XMASS = Mass of sample.
CP = Heat capacity of sample.
ES = Emittance of sample.
TSO = Initial value of temperature of sample for integration.
TSCK = Temperature at which check is to be made of integrated equation temperature with telemetry data value.
AS = Solar absorptance.
HS = Solar heat input.
HR = Earth albedo heat input.
HP = Planetary heat from earth.
26 f
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TIMEE
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SUBROUTINE F(THETA, TS, P)
RKDF
IER
Upper limit on time for integration.
Total solar heat incident on sample edge.
Cross-sectional area of sample.
Area of edge of sample.
Shape factor edge sample to space.
Shape factor edge sample to Mylar.
Emittance of edge of sample.
Absorptance of adjacent Mylar.
Emittance of Mylar holder.
Thermal conductance of sample.
Stephan-Boltzmann constant.
Order of equation to be integrated.
Check variable for oscillation of function, if oscil *"ing, agreement interval increased to 0.4°F.
Time.
Increment for integration.
Counter.
Counter for printing out every 10th integration solution point.
Counter of number of solution points.
Function subroutine for differential equation.
Runge-Kutta subroutine.
Error flag.
27
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AFML-TR-75-17
APPENDIX II
COMPUTER PROGRAM DAN
This appendix lists the computer program utilized to convert
raw digital data tapes directly into time-temperature profile curves.
31
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