NASA/CR--2001-211411

Testing and Optimization of Electrically

Conductive Spacecraft Coatings

R.J. Mell

AZ Technology, Inc., Huntsville, Alabama

G.E. Wertz

Marshall Space Flight Center, Marshall Space Flight Center, Alabama

Prepared for Marshall Space Flight Center

under Contract NAS8-98212

and sponsored by

The Space Environments and Effects Program

managed at the Marshall Space Flight Center

National Aeronautics and

Space Administration

Marshall Space Flight Center • MSFC, Alabama 35812

December 2001
NASA Center for AeroSpace Information

7121 Standard Drive

Hanover, MD 21076- ! 320

(301) 621-0390

Available from:

ii

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22161

(703) 487--4650
TABLE OF CONTENTS

1.0 INTRODUCTION ............................................................................................... 1

1.1 Background .................................................................................................. 1

1.1.1 Surface Charging ...................................................................................... 2

1.2 Coatings ....................................................................................................... 2

1.3 Purpose ........................................................................................................ 4

1.4 Program Objectives ...................................................................................... 4

2.0 RESEARCH AND EVALUATIONS ................................................................... 5

2.1 Theory and Design ....................................................................................... 5

2.2 The Experimental Setup and Procedures ...................................................... 6

2.2.1 Test Chamber and Equipment ................................................................. 6

2.2.2 Test Procedure ........................................................................................... 12

3.0 Test Results ...................................................................................................... 16

3.1 Spectral Solar Absorptance .......................................................................... 16

3.2 Surface Voltage ........................................................................................... 17

3.3 Conductivity ................................................................................................ 20

4.0 Electrically Conductive Ceramic Pigments ........................................................ 23

4.1 Evaluation of Coating Additives and Binders ............................................... 23

4.2 Optimization of Electrically Conductive Ceramic Pigments ......................... 26

5.0 Conclusion ......................................................................................................... 28

iii

LIST OF FIGURES

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1. Diagram of Conductivity Measurement System ........................................................... 6

2. Electron and Enhanced UV Sources ............................................................................ 7

3. Spectral Output of the Enhanced UV Source ............................................................... 7

4. Movable Test Sample Holder ...................................................................................... 8

5. Sample Coupon ........................................................................................................... 8

6. Conductivity Test System Control & Data Acquisition ................................................ 9

7. General Layout of Conductivity Test System ............................................................. 12

8. Probe in Reference Position ....................................................................................... 14

9. Probe in Sample #1 Position ..................................................................................... 14

10. Probe in Sample #2 Position .................................................................................... 15

11. Coating Surface Volts Measured During Electron Beam Exposure .......................... 18

12. Coating Surface Volts Measured During Combined Electron Beam

and Solar UV Exposure ............................................................................................ 18

13. Surface Voltage (greater than 2 volts) Versus Exposure Dose .................................. 19

14. Surface Voltage (less than 2 volts) Versus Exposure Dose ....................................... 19

15. Conductivity Values Derived from Electron Beam Exposure Data ........................... 20

16. Conductivity Values Derived from Combined Electron Beam and Ultraviolet

Light Exposure Data ................................................................................................ 21

17. Conductivity Versus Exposure Dose (surface voltage greater than 2 volts) .............. 21

18. Conductivity Versus Exposure Dose (surface voltage less than 2 volts) ................... 22

19. Transmittance of Boeing Binder Through the Solar Spectrum .................................. 24

20. Solar Absorptance of Potassium Silicate Binder ...................................................... 25

21. Solar Absorptance Zinc Oxide Pigmented Proprietary Binder .................................. 26

22. Synthesized Electrically Conductive Whickers ........................................................ 28

LIST OF TABLES

Table

Table

Table

Table

Table

Table

Table

Table

1. Representative Properties for Two Types of Thermal Control Coatings ........................ 3

2. Conductivity Research Equipment .............................................................................. i 1

3. Pre-Exposure Test Measurements ............................................................................... 12

4. Sample Exposure Procedure ....................................................................................... 13

5. Solar Absorptance Values for Pre and Post Irradiated Test Samples ............................ 16

6. Surface Voltage Measurements ................................................................................... 22

7. Current Candidate Metal Atoms and Their Use in This Program ................................. 23

8. Support Materials ....................................................................................................... 27

V

APPENDIX A

Figure 23.

Figure 24.

Figure 25.

Figure 26.

Figure 27.

Figure 28.

Figure 29.

Figure 30.

Figure 31.

Figure 32.

Figure 33.

Figure 34.

Solar Absorptance of AZ-70 .................................................................................. A-2

Solar Absorptance of AZ-93 .................................................................................. A-2

Solar Absorptance of AZ- 100 ................................................................................ A-3

Solar Absorptance of RM-400 ............................................................................... A-3

Solar Absorptance of AZ-1000 .............................................................................. A-4

Solar Absorptance of AZ-2000 .............................................................................. A-4

Solar Absorptance of AZ-2100 .............................................................................. A-5

Solar Absorptance of AZX-S 1 ............................................................................... A-5

Solar Absorptance of AZX-SG1 ............................................................................ A-6

Solar Absorptance of AZX-SG2 ............................................................................ A-6

Solar Absorptance of TMJ-810 .............................................................................. A-7

Solar Absorptance of TMS-800 ............................................................................. A-7

APENDIX B

Figure

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Figure

Figure

35. Surface Voltage Measurements of AZ-70 ............................................................ B-2

36. Calculated conductivity values for AZ-70 ............................................................ B-2

37. Surface voltage measurements of AZ-93 ............................................................. B-3

38. Calculated conductivity values for AZ-93 ............................................................ B-3

39. Surface voltage Measurements of AZ-100 ........................................................... B-4

40. Calculated Conductivity Values for AZ-100 ........................................................ B-4

41. Surface Voltage Measurements of AZ-400 .......................................................... B-5

42. Calculated Conductivity Values for AZ-400 ........................................................ B-5

43. Surface Voltage Measurements of AZ- 1000 ........................................................ B-6

44. Calculated Conductivity Values for AZ-1000 ...................................................... B-6

45. Surface Voltage Measurements of AZ-2000 ........................................................ B-7

46. Calculated Conductivity Values for AZ-2000 ...................................................... B-7

47. Surface Voltage Measurements of AZ-2100 ........................................................ B-8

48. Calculated Conductivity Values for AZ-2100 ..................................................... B-8

49. Surface Voltage Measurements of AZX-S 1 ......................................................... B-9

50. Calculated Conductivity Values for AZX-S1 ....................................................... B-9

51. Surface Voltage Measurements of AZX-SG1 .................................................... B- 10

52. Calculated Conductivity for AZX-SGI .............................................................. B-10

53. Surface Voltage Measurements of TMJ-810 ...................................................... B-11

54. Calculated Conductivity Values for TMJ-810 .................................................... B- 11

55. Surface Voltage Measurements of TMS-800 ..................................................... B- 12

56. Calculated Conductivity Values for TMS-800 ................................................... B- 12

vii

1.0 INTRODUCTION

This is the Final Report discussing the work done for the Space Environmental Effects

(SEE) Program. It discusses test chamber design, coating research and test results on electrically

conductive thermal control coatings. These thermal control coatings are being developed to have

several orders of magnitude higher electrical conductivity than most available thermal control

coatings. Most current coatings tend to have a range in surface resistivity from 10 _ to l013

ohms/sq.

Historically, spacecraft have had thermal control surfaces composed of dielectric

materials of either polymers (paints and metalized films) or glasses (ceramic paints and optical

solar reflectors). Very seldom has the thermal control surface of a spacecraft been a metal where

the surface would be intrinsically electrically conductive. The poor thermal optical properties of

most metals, have in most cases, stopped them from being used as a thermal control surface.

Metals low infrared emittance (generally considered poc. for thermal control surfaces) and/or

solar absorptance, have resulted in the use of various dielectric coatings or films being applied

over the substrate materials in order to obtain the required optical properties.

During the 1970's, surface charging of spacecraft began to become a concern to the

industry. In the late 1980's and early 1990's, this problem became significant as electronic

systems became more sensitive to electrical anomalies and spacecraft surfaces or systems were

damaged through electrical arc discharges. Electrical discharges can be generated from

spacecraft interaction with the electrically neutral environment. Interaction between spacecraft

and environment can cause charge segregation. This has resulted in charged spacecraft surfaces

in GEO and LEO orbits, depending on the external biasing of the spacecraft. Indium tin oxide

has been used to alleviate surface electrical resistivity problems for many applications.

However, this and similar materials are typically deposited as thin films and tend to be

susceptible to damage from cracking or de-bonding when deposited onto thin or flexible

dielectric polymer substrates. Such damage to the electrically conductive thin film can cause a

loss of electrical conductivity and an increase in surface charging.

A contract was awarded to AZ Technology by NASA through the SEE program and

monitored through the Marshall Space Flight Center for the optimization and testing of

electrically conductive spacecraft coatings. The purpose of this effort was to test both white and

black spacecraft coatings in vacuum and optimize these coatings to have electrical conductivity

values higher than currently available thermal control coatings. In addition, these coatings are

designed to be deposited onto large complex geometry surfaces with minimal difficulty or

environmental impact. Optimization of these coatings with higher conductivity will be a

significant step forward in materials and spacecraft technology.

1.1 Background

A significant and recurring spacecraft problem is the build-up of surface charge resulting

from interaction of the spacecraft with its orbital environment of charged particle radiation or

space plasma. Charging of a spacecraft surface can further be intensified as spacecraft increase

in size and/or with increased power requirements. In addition, payload sensitivity to electrical

noise caused by electrical discharges are increasing for many items on spacecraft. Other factors

like the electrical biasing of a spacecraft power generation system can also contribute to this
problem. SpaceStationwill be the largestandhighest-poweredspacecraftput into orbit. It is

anticipatedthe structurewill be drivento approximately140Vnegativeto the ambientplasma

andwill requireoneor moreplasmacontactorsto maintaina nearneutralsurfacecharge.

I.I.I Surface Charging

Basically, spacecraft surface charging results from several different kinds of charging

phenomena and how the environment interacts with the various surfaces of the spacecraft. The

different kinds of charging phenomena are typically the following: 1) electrons from space

plasma, 2) photoelectron emissions from spacecraft sucf_,zes, 3) ions from space plasma, 4)

secondary electron emissions generated by electron impacts, and 5) secondary electron emissions

from ion impacts. Charge transfer typically occurs in one of the following ways: 1) from the

surounding environment to the spacecraft, 2) from the spacecraft to the surrounding

environment, or 3) between different spacecraft surfaces. The interaction of these factors is

highly complex. Though interactions of spacecraft and the space environment together cause

much of the charging problems, one can not neglect the _;,',eraction that can occur between

different materials on or in the spacecraft as well as electronic and electrical grounding circuits

which also contribute to the formation of surface charging. Even with our current level of

understanding of schemes for controlling spacecraft surface charge build up, there continues a

need to develop and certify greater varieties and ranges of electrically conductive thermal control

coatings for flight hardware utilization.

Spacecraft surface charging is believed to be responsible for the loss or damage of a

number of satellites and the production of anomalous instrument data. With the development of

spacecraft like the Tethered Space Satellite (TSS) and others carrying instrument packages far

more sensitive than previous systems, the problem of conducting away surface charge has

become even more difficult to master. Tethered Satellite's function was to study the space

plasma environment surrounding the earth. To perform this task, the exterior surface coating not

only was required to act as an effective thermal control surface, but also as a good electrical

ct,tlductor. Because of this electrical conductivity requirement, current state of the art thermal

control coatings were evaluated based not only on their thermal optical properties but also more

importantly on the coating's intrinsic electrical conductivity. Through extensive testing and a

few serendipitous events, it was found from testing that the coating originally selected and

applied to TSS was ineffective because of poor electrical conductivity in a vacuum. Marshall

Space Flight Center developed an effective and successful coating (RM-400) to meet the needs

of the short duration TSS mission. This coating had only a short-term stability of its thermal

optical properties in the space environment. This electrically conductive thermal control coating

only had a required thermal optical stability requirement of about two weeks of space exposure

in LEO. However, what is actually needed is long-term, space stable coatings for a variety of

space and potentially aircraft applications.

1.2 Coatings

Currently, only a few thermal control coatings have been produced that have the potential

to dissipate surface charge build-up for specific applications or conditions. Little is known and

understood of these with the possible exception of a few with solar absorptance values of (0.18
to 0.30). These coatings are produced by Space Craft Incorporated (SCI) and Jet Propulsion

Laboratory (JPL) and have been tested and used on a few spacecraft. Although effective at

mitigating charge buildup they required long lead times to produce and require great expertise to

deposit onto a surface. In addition, current designers in many cases want coatings that that have

better thermal optical properties, ease of application, space environment stability, while

maintaining good or having better electrical properties. Much newer coatings and films from AZ

Technology, Triton, ITTRI and Boeing have limited testing and flight history. Some are just

now being considered for use on flight hardware. Of the high (0.85 to 0.98) solar absorptance

coatings, most are based on polymer binders and or graphite and carbon pigments that are not

appropriate for all uses or orbits. Therefore, the problem of damage from electrical arcing or

discharge still cause many problems for the spacecraft industry. The coatings developed under

AZ Technology's research program are completely inorganic, and are comprised of soluble

glasses and ceramic pigments potentially making them useful for most if not all orbits.

Each coating has its strong areas of performance. -_s shown in Table 1, the ceramic

coating, AZ-93, has poor performance in this category which is the most critical for it's intended

mission. The RM-400 performs well in the areas of conductivity and thermal emittance but falls

short in the category of solar absorptance (Ors), as a radiator coating, with a value of 0.49.

Typically for radiator coatings a solar absorptance value of 0.25 or less is required.

Table 1. Representative Properties for Two Types of Thermal Control Coatings

MATERIAL SOLAR THERMAL VACUUM

DESIGNATION ABSORPTANCE EMITTANCE a. ELECTRICAL

% RESISTIVITY D./sq.

Conventional TC

Coating 0.15 0.91 1013_1014

AZ-93

Tether satellite

coating RM-400 0.49 0.89 5 X 105

Extensive research has taken place over the last few years to develop a variety of

spacecraft coatings with the unique property of being able to conduct surface charge to a

substrate or grounding system. The ability to conduct surface charge to a safe point, while

maintaining optical properties and performance, is highly advantageous in maintaining

operational space based systems. Without this mechanism, surfaces of a spacecraft can

accumulate charge to the point that a catastrophic electrical breakdown can occur, resulting in

damage to, or failure of the spacecraft. Ultimately, use of this type of coating will help mitigate

many of the concerns that NASA and the space industry still have for their space based systems.

When these spacecraft coatings are found to meet stability needs, they can be used to control

electrical charge build up and possibly conduct sufficient electrons for power generation from

the space plasma, as was demonstrated with the TSS. The unique coatings that have been

developed and studied fall into two specific categories: 1) broadband absorber (black as= 0.94 to

0.97), and 2) selective absorber (white as=0.16 to 0.30). These coatings have controllable solar

absorptance and electrical surface resistivity values over designated ranges. These coatings were

developed under a program which focused on the development of constituents and coatings to

meet the study goals.
1.3 Purpose

The purpose of this research program was to test and to optimize electrically conductive

thermal control coatings developed under a previous program.

This project focused on simulated space environmental effects testing with the intent of

using this data to help optimize the stability and initial properties of these coatings. It was

originally planned to expose candidate coatings to a new test that consisted of exposing the

candidate coatings (independently) to an environment of low energy charged particles (generally

10Kev electrons) and/or UV with in-vacuum electrical resistivity measurement. Then, if

possible, expose these same materials to a combined environment of VUV, UV, protons, and

electrons with in-vacuum optical solar absorptance measurements. Through the utilization of

such testing, the understanding of these coatings' space environmental stability would provide

sufficient data to determine the coatings stability. Then, if sufficiently stable, the tests would

provide data needed for their acceptance by the aerospace community and their potential use on

current and future space programs.

Also included, was the design, procurement, assembly, debugging, procedural and

methodology maturation, of a completely new space environmental testing chamber. Also

discussed are the syntheses, formulation, and deposition of semiconductor spacecraft TC

coatings, and their subsequent testing on custom .configured test samples. The development and

production of these types of coatings will be very useful for the management of electrical charge

that can accumulate on the surface of a spacecraft. Also, the developed coatings could be used

for current collection, storage, and conduction in environments where metals would not be

effective because of poor thermal-optical properties or material reaction with the environment.

1.4 Program Objectives

• Determine suitable reference coatings that have significant ground testing, history, and if

possible flight data;

• Determine effect on reference coatings and then baseline experimental coatings when

exposed to the environment of UV, low energy e-, and vacuum) with in-vacuum electrical

conductivity measurements;

• Determine from space environmental effects data if test coatings need to be modified to

meet stability requirements;

• Develop and implement a plan for optimizing test coatings (if needed);

• Scale-up material volumes to be capable of effectively meeting industry needs for at least

one white and black coating with the following requirements:

Initial properties

White Coating Black Coating

- o_s<0.30 - cts>0.85

- eT >0-85 - _T >0.85

- Resistivity tailorable between 102 to 109 ohm/sq.

- Contamination potential

> Collectable Volatile Condensable Materials CVCM < 0.1%
- Environmental stability when exposed to the space environment for one year

> Change in solar absorptance Ao_s < 0.10

> Change in resistivity < lxl02

- Mass loss <0.5% when exposed to an AO fluence of l xl020 atoms

2.0 RESEARCH AND EVALUATIONS

2.1 Theory and Desi2n

The standard approach used to measure the conductivity of a material is to bring

electrodes in mechanical contact with the surface of the material to be measured. The surface

voltage and current measurements, made relative to the contact area, determines the surface

resistivity. However, variables such as non-uniform surface structure and contact pressure can

introduce error to this measurement. Therefore a more desirable approach, both for simulating

the space environment and measuring conductivity, is t_, leplace the mechanical contacts by a

beam of electrons. Since sunlight also plays an important role by enhancing surface charging, a

combined electron and simulated solar spectra was used in the investigation of material

conductivity.

The calculation of material conductivity used is made on the assumption that the material

coatings obey Ohm's law. Ohmic materials demonstrate a linear behavior between current

density and the electric field. This ratio of the current density and electric field is a constant _,

which is independent of the electric field producing the current. That is,

J = cE,

Where J is the current density, E is the electric field and the constant of proportionality,o,

is called the conductivity of the material.

The electron irradiation of the coating produces a potential difference AV across the

thickness, d, of the coating. The potential difference is related to the electric field through the

relationship

V=Ed.

Therefore, we can express the magnitude of the current density as

J = t_E = c(V/d).

Assuming the material to be uniform, the current density J is defined as current per unit

area,

J- I/A.

Therefore, as illustrated in Figure 1, we can express conductivity as

= Id/VA.

Where I is the current through the coating, d is the thickness of the coating, V is the

voltage on the surface of the coating relative to ground and A is the cross-sectional area of the

coating being measured.
OV

Figure 1. Diagram of Conductivity Measurement System

2.2 The Experimental Setup and Procedures

The George C. Marshall Space Flight Center designed, fabricated, and operated a unique

space environmental effects test chamber incorporating the ability to measure the electrical

conductivity of various spacecraft coatings. Its purpose was to record what effect different forms

of energy (radiation) have on external spacecraft materials performance, function, and

survivability (stability or lack of change) over time or radiation dose. As is shown and discussed

in detail in the following paragraphs a custom designed vacuum chamber was fitted with both an

electron flood gun and a thousand watt near ultra-violet (NUV) radiation source. These radiation

sources can be used simultaneously or individually as needed to understand a particular

phenomenon or material effect.

2.2.1 Test Chamber and Equipment

The test system, shown in Figure 2, consists of an energetic electron source and enhanced

UV source. Each source impinges with the sample coupons at an angle of 5.5 degrees off sample

normal. The test system was maintained at a base vacuum of 5 x 10 -7 Tort. A Kimball Physics

50 Key electron source was adjustable from 0 to 50 Kev and capable of generating currents up to

1mA/cm 2. For testing purposes, the source was operated at a constant energy of 10 Kev. The

electron beam current density was variable from 1 to 10nA/cm 2.

6
Figure 2. Electron and Enhanced UV sources

The Spectral Energy enhanced ultraviolet (UV) source uses a xenon lamp to simulate the

UV output of the sun. The output of this assembly was mea,;ured with an Optronics Laboratories

Model 752 Spectoradiometer. The photon flux, integrated from 200 to 400 nanometers, was

measured to be 25.5 mw/cm 2 nm. UV source spectral output is shown in Figure 3.

3 00fz-O4

2.50E-04

E" 2 OOE-e4

c

,.

E

_ 1 50E-O4

Q

u

c

,m_

500E*05

O00E+OG

2_ 220 240 260 280 300 320 340 360 380 400

Wavelength (nanom_em)

Figure 3. Spectral Output of the Enhanced UV Source

The sample holder, shown in Figure 4, allows two samples to be tested simultaneously

and has one solid disk in the top position used as the reference source for the electrostatic probe.

This sample holder assembly is movable within the vacuum chamber for fine adjustment of the

separation distance between the voltage probe and the sample surface.
Figure 4. Mobile Sample Holder Assembly

Two Faraday cups, shown in Figure 4, are mounted on the sample holder between the

two sample coupons. Each of these Faraday cups have a 0.07cm 2 aperture used to measure the

electron beam current density. A digital current integrator monitors one Faraday cup. Digitized

output (electron counts) from the current integrator is read by the data acquisition system and

used to control events. A second Faraday cup is monitored by a sensitive electrometer for

continuous logging of current density data.

Guard Rin._ .........

Figure 5. Sample Coupon.

The sample coupon shown in Figure 5 was made of 6061-T6 one-half inch thick aluminum alloy.

A schematic side view is provided in Figure 1. Sample coupons consisted of an outer (guard)

ring and inner (test) disc. The test disc was secured within the guard by means of a rear mounted
0.06-inch thick Teflon TM disc and mounting screws. The test disc sample area is 2.85cm 2 and is

isolated from the guard ring with lmm thick Kapton TM tape. The total sample area is 15.5 cm 2,

an aspect ratio of approximately 5 to 1. Each sample coupon was secured to the sample holder

by threading the test disk to an insulated mounting bolt. This was wired via a vacuum electrical

feed-through to the current monitoring device. Sample surface to voltage probe separation was

adjusted and the guard ring connected to chamber ground.

Surface voltage of the coating is measured by means of an electrostatic probe mounted on

a pneumatic linear motion vacuum feed-through. A Trek Model 344 Electrostatic Voltmeter has

a measurement range from 0 to 2kVDC with an accuracy _-H).1% of full scale. The voltage probe,

attached to a linear motion vacuum feed-through, translates the probe to the required sample

position for a surface voltage measurement.

Initially sample coupons were coated across and over the Kapton separator located

between the inner and outer sections of the sample coupon (see Figure 5). This was changed

when it was found to result in a charge collecting area larger than expected. As a result of these

findings, sample coupons were coated before assembly to ensure that the gap between the inner

and outer coupon sections was not coated. This change in coupon coating configuration helped

to ensure that a charge collecting area was confined to the inner disc of a coated sample coupon.

VACUUM

Sotenoid corarol

Posilion

McroSwitch

O-2kVdc

J _;ourcerneter [ Sample

I I Currenl

Cotl3terSource InputICounter Gale Control

0-40Vdc

Data

Labwew Data Acquisition Software

COMPUTER

Figure 6. Conductivity Test System Control & Data Acquisition
The data acquisition system layout, shown in Figure 6, is custom designed to implement

system controls and perform data storage operations. Utilizing National Instruments hardware

and Labview Software this system is programmed to perform:

• Translation of pneumatic linear motion devices for blocking the electron beam and

moving the electrostatic probe into position for a surface voltage measurement,

• Shuttering of the ultraviolet light,

• Suppression of the electron beam,

• Communicating with GPIB devices,

• Event timing,

• Data-logging operations,

• System error handling.

During an exposure test, Faraday cups (see Figure 6) mounted near the samples produce a

current proportional to that of the electron beam current de_psity. The digital current integrator

measures this coulomb charge and produces a digitized output relative to the number of incident

electrons per unit area. This digitized output (electron counts) is used to control timing events.

Table 2 provides a listing of the test equipment used to provide the electronic components

used to operate and acquire data from this unique space simulation test chamber. Use of dual

radiation sources allowed AZ Technology and NASA investigators to evaluate the effects of two

typical but very different energy sources naturally occurring in space. Invacuo electrical

measurements during space simulation irradiation testing provide insight into possible different

degradation mechanisms and how they effect the new generation of semiconductor thermal

control coatings.

10
Table 2. Conductivity Research Equipment

Equipment Manufacturer / Model Description

Electron Source Kimball Physics

Model EFG-9 Electron Flood Gun

UV Source Spectral Energy Xenon Lamp

Dichloric Lense

Electrostatic

Voltmeter

Current Measuring

Devices

Data Acquisition

System

Trek Model 344

Beam Energy: lkeV to 50keV

Beam Current: 10gA to 10-4A

Voltage measurement range: 0 to •

2kVDC

Measurement accuracy: • 0.1% of full

scale

Keithley Model 6430

Keithley Model 6415

EG&G Ortec Model 439 Digital Current Integrator

EG&G Ortec Model 996

Sub-Femtoamp Remote Sourcemeter:

100aA sensitivity

Electrometer: 100aA sensitivity

Counter L Timer

Pentium .;OOMhz Computer System

Microsoft Windows 95

National Instruments Labview Software Version 5.0

_lational Instruments NI-6025E PCI multifunction I/O Card

National Instruments NI-4060 PCI Digital Muitimeter

National Instruments SC-2062 Electromechanicai relay board

National Instruments SC-2050 Cable Adapter Assembly

National Instruments CB-50LP Connector Block

Figure 7 provides a general configuration layout of the test chamber. The electron flood

gun and the NUV source are each offset by 5.5 degrees in the horizontal (Z) axis from normal

and are normal to the samples in the vertical (Y) axis. This minimizes the offset angle to the

sample, while providing sufficient separation.

NUV source was equipped with a dichroic infrared (IR) energy filtering system to

minimize sample heating. Both sources were located at a sufficient distance from the samples

that resulted in coverage of all test samples and detectors with a uniform beam. A UV grade

quartz window was used to allow the externally mounted NUV source radiation to pass into the

vacuum test chamber. The entire system was pumped down using a vacuum turbo pump.

Samples were exchanged through a 10-inch diameter quick access door airlock combines with a

viewport.

11
Figure 7. General Layout of Conductivity Test System

2.2.2 Test Procedure

Each sample coating was tested in exactly the same manner. Thermal-optical property

evaluation and thickness measurements were performed on each coating sample. Additional

samples were produced from the same paint batch for a surface resistivity measurement to be

made with the Hewlett Packard high resistance test meter and resistivity cell. This measurement

was performed at atmospheric pressure under an argon purge. Results from these tests are shown

in Table 3.

Table 3. Pre-Exposure Test Measurements

Sample Batch

II

AZ-70 1-041

!AZ-93 H-209

AZ-100 1-039

AZ-400 1-040

AZ-1000 H-215

Emittance Alpha

0.892 0.122

0.913 0.155

0.911 0.150

0.902 0.295

0.963

0.916 0.238

0.893 0.153

0.876 0.352

0.890 0.242

0.852 0.144

0.892 0.352

0.895 0.376

AZ-2000 H-212

AZ-2100 1-043

TMS-800 1-048

TMJ-810 1-044

AZX-S 1 1-055

AZX-SG1 1-053

AZX-SG2 1-054

Thickness Thickness

(mils) (cm)

9.3 0.023622

5.1 0.012954

5.4 0.013716

3.4 0.008636

2.9 0.007366

3.21 0.0081534

6.91 0.0175514

3.82 0.0097028

3.79 0.0096266

5.09 0.0129286

3.61 0.0091694

3.75 _ 0.009525

12
Solar absorptanceand thermal emittance of test samples were measuredat AZ

Technologyand receivedat MSFC/ED31usingthe LaboratoryPortableSpectroreflectometer

(LPSR) instrumentandthe Temp2000emissometer.Eachsamplehad a lmm KaptonTM tape

placed in the spacing between the guard ring and the test disc and then was assembled. The

samples were then positioned as either as "Sample 1" or "Sample 2" on the movable sample

holder within the vacuum chamber. A vacuum environment of less than 1 x 10 .6 Torr is required

for testing. Exposures began only after a one-hour warm-up of the electron source, UV source,

and electrostatic meter. Each sample exposure follows the procedure described in Table 4.

Table 4. Sample Exposure Procedure

Run Exposure Description

I 10keV electrons

2 10keV electrons

3 10keV electrons

4 10keV electrons

5 10keV electrons

6 10keV electrons

7 10keV electrons + UV

8 10keV electrons + UV

9 10keV electrons + UV

10 10keV electrons + UV

11 10keV electrons + UV

12 10keV electrons + UV

Current Setting

1na/cm 2

1na/cm 2

5na/cm 2

5na/cm 2

10na/cm 2

10na/cm 2

1na/cm 2

1na/cm 2

5na/cm 2

5na/cm 2

10na/cm 2

10na/cm 2

Counts Discharge Data

2500

!

2500

2500 Sample #1

2500 Sample #2

Sample #1

5000

5000

2500

2500

2500

2500

Sample #2

Sample #1

Sample #2

Sample #1

Sample #2

Sample #1

Sample #2

5000 Sample #1

5000 Sample #2

A snapshot method was used to take measurements of the sample coating's surface

voltage and current flowing through the coating. Snapshot measurement was obtained during an

exposure run and the sequence begins only when the data acquisition system has logged the

required interval of electron counts. Snapshots occur approximately every I00 electron counts

with the following sequence:

1) The voltage probe is in the reference position, as shown in Figure 8. A measurement is

made of the reference disk that has been biased to a known voltage potential. This

measurement is then compared to the known bias voltage and the resulting offset used as a

correction factor for the upcoming surface voltage measurements.

13
2)

3)

Figure 8. Probe in Reference Position

The current passing through sample #1 is logged to file.

Probe is remotely controlled to move to sample position #I, as shown in Figure 9.

surface voltage measurement of sample #1 is made and logged to file.

A

Reference Disc ]

Sample #1

I Voltage Probe

Sample #2 ]

Figure 9. Probe in Sample #1 Position

14
l) The current passing through sample #2 is logged to file.

2) Probe is remotely controlled to move to sample position #2, as shown in Figure I0. A

surface voltage measurement of sample #2 is made and logged to file.

Reference Disc ]

3)

Figure 10. Probe in Sample #2 Position

The probe is remotely controlled to move back to reference position, where it remains

until the next snapshot sequence is triggered.

Exposure runs one through six were electron beam exposures only. Exposures seven

through twelve were combined UV and electron exposures. For each exposure run:

4) The electron beam current density was stored to file and integrated into electron counts

that controlled exposure events,

5) The current through the sample coatings were continuously measured (ten

samples/second) and stored to file,

6) Surface voltage measurements (snapshots) were periodically made for each sample.

Exposure runs are made concurrently with the elapsed time between exposures varying

from 1 to 30 minutes. Each sample received an exposure of 40000 electron counts and 6.25

equivalent sun hours (ESH). One electron count is equal to 1 x 10 -10 coulombs or 7.9 x 10'_

electrons/cm 2. Each sample received a total fluence of 3.16 x 1014 electrons/cm 2.

15
3.0 TEST RESULTS

Thefollowing isa descriptionof eachcoatingthatwastested:

AZ-70

AZ-93

AZ-100

RM-400

AZ-1000

AZ-2000

AZ-2100

AZX-S1

AZX-SG1

AZX-SG2

TMJ-810

TMS-800

Inorganicwhite,nonspecularthermalcontrolcoating

Inorganicwhite,nonspecularthermalcontrolcoating

Developmental inorganicwhite,electricallydissipativethermalcontrolcoating

Epoxywhite,electricallyconductivethermalcontrolcoating

Inorganicsemi-conductiveblackthermalcontrolcoating

Inorganicsemi-conductivewhitethermalcontrolcoating

Developmental inorganicwhite,electricallydissipativethermalcontrolcoating

Developmental inorganicwhite,electricallydissipativethermalcontrolcoating

Developmental inorganicwhite,electricallydissipativethermalcontrolcoating

Developmental inorganicwhite,electricallydissipativethermalcontrolcoating

Inorganicyellownonspecularsemi-conductivemarkercoating

Inorganicyellownonspecularmarkercoating

Table 5. Solar Absorptance Values for Pre and Post Irradiated Test Samples

Sample Batch Pre-Al_ha Post-Alpha Percent Chan_ge

AZ-70 1-041 0.122 0.123 -0.82%

AZ-93 H-209 0.155 0.162 -4.52%

AZ- 100 1-039 0.150 0.149 0.67%

RM-400 H-213 0.295 0.329 -11.53%

AZ- 1000 H-215 0.963 0.967 -0.42%

AZ-2000 H-212 0.238 0.250 -5.04%

AZ-2100 1-043 0.153 0.148 3.27%

TMS-800 1-048 0.352 0.359 - 1.99%

TMJ-810 1-044 0.242 0.249 -2.89%

AZX-S 1 1-055 0.144 0.145 -0.69%

AZX-SG1 1-053 0.352 0.365 -3.69%

AZX-SG2 1-054 0.376 0.377 -0.27%

3.1 Spectral Solar Absorptance

Solar absorptance measurements were obtained using the LPSR-200 laboratory portable

spectroreflectometer instrument both before and after the sample exposure procedure (see Table

4). The resulting solar absorptance (alpha) for each coating is shown above in Table 5. Spectral

solar absorptance curves data for each of the coatings are separately plotted in Appendix 'A'.

Solar absorptance measurements were performed after the chamber was purged with nitrogen.

This significantly reduces the rate of bleaching of the test samples. Experiments done by James

Zwiener and David Edwards at MSFC have shown that purging of the space simulation chamber

with nitrogen gas for repress, reduces sample bleaching. In addition, Michel Michishnek, of the

Aerospace Corporporation, has demonstrated the same effect in his space simulation test

chambers. Bleaching of a test sample is the reduction or elimination of optical changes (usually

16
damage)causedby exposureto oneor moreradiationsourcesduringa simulatedor actual space

environment test. As a result, the nitrogen purging technique was used for this program to

significantly reduce the rate at which a test samples bleach since the chamber was not configured

for insitu optical measurements. These optical changes typically occur in seconds or fractions of

seconds with most materials but can continue for months or years as shown by Don Wilkes and

James Zwiener in the Optical Properties Monitor (OPM) science report to NASA. The OPM

materials flight experiment performed solar absorptance measurements as well as other types of

optical measurements on materials while on orbit. OPM transmitted stored data back to earth

typically on a weekly basis.

3.2 Surface Voltage

Periodically, through each exposure run, a snapshot of each sample surface voltage was

made. The custom designed data acquisition (Labview Version 5.1) software controlled the

snapshot event sequence. Every one hundred electron counts, the electrostatic probe was

positioned above the center of the sample coupon to measure the surface voltage. These surface

voltage measurements were stored to a file, designated for a given exposure run and sample.

Measurements were logged and referenced to the accumulated number of electron counts.

A total of twelve exposure runs were performed. The first six exposures were performed

with only the electron beam operating, while the following six exposures were with the

combined electron beam and ultraviolet light. The effect of changing electron beam current

density was investigated at three energy levels. Each had two exposures at lnA/cm 2, 5 nA/cm 2

and at 10nA/cm 2 current densities. For each exposure run, the median (the median is the number

in the middle of the set of numbers, half the data set numbers have values greater than the

median and half of the numbers have values that are less) data point was determined for surface

voltage for each sample. As a representative surface voltage or "exposure" current density, the

two median values of the repeated current density exposures were averaged. This average

surface voltage potential versus the exposure sequence, is listed in Table 6. Materials are listed

by the magnitude of surface voltage.

The measured surface voltage data from the electron beam exposures are plotted in

Figure 11. The surface voltage data from these combined (electrons and UV) exposures are

plotted in Figure 12. A comparison of the surface voltage measured during the electron beam

exposures and the combined exposures are charted in Appendix B.

17
Io00

100

lo

I

<>.,.--'"

5000 10000 IS000 20000

Accum ull_ivl El_ct_on C ountl

- - o - -AZ.,,o

- - ,o. - -Az.m_

- - _. - -Az._oo

o Az.4oo

• - II. - -AZ-_O00

- - 4" " -AZ-_000

• -,_ - -AZ-2_00

Azx.sl

_AZX_SO_

_Azx.so2

-.......,,_,--- TM J -a 10

_Tus.aoo

Figure 11. Coating Surface Volts measured during electron beam exposure

1000 ]4 ..........

,oo t

l

10

1

m |

to,,.m_ ................... .e ...... 5o,,om2......... t ..................... 0o^,om'.......................

....... • .............................................. _,_, - - o - -Az,7o

......... --_>-.AZ-g3

&,-" '-_ jf

G..._/ ._ ....................................... • ......... •

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

O, . ....................... Ill

I

!

oo,[ i

20000 25000 30000 35000 40000

Accumulative Electron Counts

- _ - "AZ._O0

O " kZ.,O0

- Ill - -AZ-IO00

• _" -AZ-2000

- _ - "AZ _100

- _ - "AZXmSl

Cl AZX.SG1

.O, AZX.SG 2

----i_------ T M J *a t O

_r MS-eO0

Figure 12. Coating Surface Volts Measured During Combined Electron Beam

and Solar UV Exposure

The accuracy of the electrostatic voltmeter is 0.1% of full scale with a measurement

range of + 0 to 2kVdc. Therefore, data files have been grouped into those that measured greater

than two volts and those that measured less than two volts. These surface voltage charts are

shown in Figures 13 and 14.

18
}

J

i

I E+03"

1E+02"

1E÷O1

I E+O0

. r-I .... O .......... o .......... O

Cr _..,m. ........ A

_;: :'- _.'" ./o .... o .................... o

"'- // ..'" 8c...._._""

i

..o .... o .......... o .......... 0

.".... " ...... io--------_

• .... .&.." /C> .... O _

,..,O. ;::(_

// _ .... ,_ ...... -

--O--AZ.7O

---O---_Z-93

|

................. E_.clron Expolu_es .......... _4_ ...... Enh&nced UV & EleelF_ Ezposureg -

....... I ...... ' I ....... ' I ....... I ....... i ....... z

5000

-, = , .,6 Az._oo

O... - &',"

AZ-2_oo

I ----o-- T_S Bo0

10000 15000 20000 25000 30000 35000 40000

Accum ulated Electron Counts

Figure 13. Surface Voltage (greater than 2 volts) Versus Exposure Dose

1E+01 -

IE+OO

1 E -Ot

I E-O2

-' "-m,

- - ,111.

' "-'l .........

o "1.

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

.,e

0 -''*" "

0 .... • Jl, -" ,

'. ,'* "" .-'" 0

""" "0"" '%,

• i

0 5000 10000 15000 20000 25000 30000 35000 40000

Accumulated Etectron Counts

. - 4_. -Az.4oo

- - II - -AZ I.oO

- - "$" " "AZ-2OOO

•. 41". -AZX $l

-----o---- k ZX SG i

_AZX SG_2

.----,=,_ r k_J a_o

Figure 14. Surface Voltage (less than 2 volts) Versus Exposure Dose

19
3.3 Conductivity

The conductivity values computed from data measured during the electron beam

exposures are plotted in Figure 15. During the combined (electrons and UV) effects exposures, a

phenomenon occurs that reduces the measured surface voltage, thereby increasing the possibility

of error in the calculation of conductivity. The phenomenon may be caused by changes in the

coating, photoelectron emission, or other causes not yet determined. These conductivity values

computed during the combined effects exposures are plotted in Figure 16.

1 E -08

1 E -0g

-E 1 .E-10

E

!

1 E-I 1

1.E-12

_E,13

............. + ............. ' ....... + .................................... ' ......................... t-]

l

• • o - -AZ-PO

-. ¢_ * -AZ-¢a

= • _, - -AZ-9OO

O AZ.tOO

- - Q - -^Z.+O00

O ^Z-ZOO0

-. _ _ .AZ-?I0O

• • _ - "AZX-S_

----B--- ^ ZX,S G i

---_*--- _Z X-SG 2

,,---,,l_ T M J .810

.,-..,_o.-.-- T M $. ii o0

Figure 15. Conductivity Values Derived from Electron Beam Exposure Data

20
t E -08

1.E-09

tE*lO

1E-11

1 E-12

IE-13

20000

,....... ,o, oo ......... ÷..,°,,oo=............ I.................. 1oo,,o ............................

!

i

!

........................... • .... ,,,,*

=oooO..°°'

&...--i." .- _° z

E3-'" i

i

25000 30000 35000 40000

Accumulated Eleclron Counts

* - C} - -AZ 70

• - .o- - -Az.s3

-.._ - -Az,loo

o Az.4oo

- - 41- - -AZ Ioc_

.o AZ _ooo

• • &- - kAZ _100

- - O - AZX-S_

"---t3_A ZX-SG 1

_AZX'SG2

_TMJ.aIO

----O---- TM S°DC'O

Figure 16. Conductivity Values Derived from

Combined Electron Beam and Ultraviolet Light Exposure Data

The accuracy of the electrostatic voltmeter was 0.1% of full scale with a measurement

range from 0 to +_2 kVdc. Therefore, coatings with measured voltage potentials greater two volts

are plotted separately in Figure 17. Coatings that measured voltage potentials less than or equal

to two volts are plotted in Figure 18. These coatings displayed conductivitites in excess of the

measurement range of the voltage probe and are shown for reference only.

E

"T

8

IE-10

1 .E-1 1

1 E-12

t,E°13

....... I ......, ! ........ ] ......I ....... ,_, ........ ,i

................. Electroe Expo|mel ................. _,_ ............... Enhanced UV & [_c_ofl EI_olurel ................ _ _

i

-" .:_: .........

.o ..... o ..... " .::-

. ,:.

-o"_-o'" ,..- _v u

o_:L.e........ .... °- • ........ •

A**" •,J

[] .... -_1"

5000 10000 15000 20000 25000 30000 35000 40OOO

Accumulated Electron Counls

..,1_.., _ : : : ': "-: ;_ -- o-*Az.7o

.. [

i ,t= - -Az.2_oo

- - --_--- T M S-a<_o

°.-A" ."

A--

[] .... -(3J

Figure 17. Conductivity versus Exposure Dose (surface voltage greater than 2 volts)

21
E

P

o

_- ...................... Electron Exposures ...................................... T ................. Enhanced UV & Electron Eq)osur_

1E-Og J .0-. "''''-'''0

1 E-IO

1 .E-11

r

i.E-12 ]

• i

°" "'I .. .... :._:,--:::-:

p.. ,-'" . ........... • P-.. o-" ...-

_...." "'o'" .,.r'_ .:,._',':-" ^ .... "°'*'. I"_D-_

,..:/ ..

.. G..AZ.400

• " _ " -AZ-20_O

" " _. "AZX-SI

---C_--'AZ X-SG I

-_O'-'-A ZX-SG 2

.... _--_ AZX-SG 3

----z_---- T MJ-81 o

0 5000 10000 15000 20000 25000 30000 35000 40000

Accumulalod Electron Counts

Figure 18. Conductivity versus Exposure Dose (surface voltage less than 2 volts)

A comparison of the computed conductivity values measured during the electron beam

exposures and the combined electron beam and solar UV exposures are charted in shown in

Appendix B.

Table 6. Surface Voltage Measurements

Material Coating Surface Voltage

AZ-70 -101 -141 -153 -102 -123 -138

AZ-2100 -46 -89 -108 -32 -59 -81

TMS-800 -19 -57 -70 -14 -38 -59

AZ-100 -35.8 -82.6 -84.7 -6.2 -16.1 -20,4

AZ-93 -15.4 -37.8 -44 • -6 -15,7 -20

AZ-2000 -2.3 -3.4 -3.5 -1.3 -2 -1.7

AZ-1000 -2.9 -2.1 -1.6 -0.7 -1.3 -1.6

TMJ-B10 -0.3 1 1.2 1.1 1.4 1.5

AZX-SG1 -2.2 -1.2 -0.8 -0.65 -0.66 -0.5

AZX-Sl -1 -1 -0.27 -0.17 0,24 0,34

AZX-SG2 -0,38 -0.1 0.08 0.2 0.22 0.3

RM-400 -0,2 -0,2 -0.1 0.2 0.15 0.2

Electron Exposure Currenl OensliV (nA/cm t)

increasing Elec(ron Exposure

(electron counts)

_lamw_ m m _attiBlm _ _

5000 10000 20000 25000 30000 40000

Jncremelng UV Expotlure _,_:_._._,¢z . ii_.;,_;

(equifxlent sun hours) 4.5 5.5 6.5

22
4.0 ELECTRICALLY CONDUCTIVE CERAMIC PIGMENTS

AZ Technology has identified and done initial evaluation on a number of new inorganic

pigments. These materials have lower solar absorptance values than those identified in previous

research. These compounds had metal atoms that were specifically chosen because of their

known or theoretical potential to produce low absorptance metal oxides in the UV through the

NIR from previous research. Therefore, AZ Technology took the approach of producing oxides

and/or alloying known high reflectance substances to other metals (starting as organo-metals) or

metal oxides. In so doing, AZ Technology could potentially solve or decrease the impact of

several problem areas that reduce the usefulness of thermal control coatings. The first is

decreasing the solar absorptance of the TC coating to be more efficient in rejecting solar energy.

Second is stabilizing the chemical structure against ion or free radical formation from solar

radiation, hence decreasing degradation of these materials. Third is providing surface charge

buildup protection through the formation of a semi-conductive pigment or additive.

Table 7. Current Candidate Metal Atoms and Their Use in This Program

Metal Atom

Rhenium

Chemical Designation

Re

Rodium Rh

Ruthenium Ru

Rubidium Rb

Application

4.1 Evaluation of Coating Additives and Binders

AZ Technology continues to acquire new coating materials where possible, such as

potential binder additives potassium hexafluorosilicate and potassium perchlorate. Each one of

these compounds can be used to aid in the stabilization of a thermal control coating. Potassium

hexafluorosilicate could be used as a means to impede the loss of oxygen from the primary

pigment. This is hypothesized to being possible because of the dense tightly bound electron

cloud structure that resided around this compound. We intend to investigate the possibility that

the use of such a compound will tend to inhibit the depletion of oxygen from the pigment that is

used in a thermal control (TC) coating. Because of the dense electron cloud of this compound, it

may help or at least impede an ionized oxygen atom from diffusing out of the coating matrix by

electron donation or simple charge repulsion.

Another approach is the use of a perchlorate compound, specifically potassium

perchlorate and ammonium perchlorate. The purpose of using this type of compound as an

additive in a thermal control coating, is to act as a supplemental source of oxygen, since the for

most white TC coatings degrade through the loss of oxygen. Presently, there has been no

success introducing this material directly into a silicate binder. When the perchlorate came into

contact with the silicate solution, precipitation of the silicate occurs and forms a solid. The

precipitated silicate is rendered useless as a binder in this state.

23
A few other alternatives still can be evaluated. One is to purposely mix the perchiorate

with silicate while mixing at high speed, thus potentially producing a homogenous solid that then

could be ground and added to a coating. Second is to blend the perchlorate with the pigment and

determine if the pigment will act as a buffer slowing down or stopping the rapid solidification

that now occurs.

A new type of binder material was also evaluated. It is currently a Boeing proprietary

material that is supplied to AZ Technology to determine it's utility for space applications. The

reason for doing some evaluation of it for this program is its transparency, Figure 19 vs. that of a

potassium silicate as shown in Figure 20. This new material has promise, because of the fact that

not only does it have very good transmittance, more importantly, it under goes significant

shrinkage during the curing process. The shrinkage characteristic of this binder material is very

important because of several aspects.

0. I0

O.(X)

25t) 4(10 55(I 7(1(I 850 L(XKI 115(I 1341(I 145(I |_1(| t75(I 19(1(I 2(15(I 22(R) 2.35(I 25(R) 265(I 2_(I

WAVELENGTH (rim}

Figure 19. Transmittance of Boeing Binder Through the Solar Spectrum

First, in this limited study, the binder does not degrade mechanical bonding to an

aluminum substrate as a consequence of shrinking. Two, this binder does not seem to loose

physical adhesion, nor does it appear to crack or fracture during this process, even as a neat film

compound. Third, a binder that shrinks as part of an electrically conductive coating curing

process without fracturing, forces the conductive particles to be pulled together. Binder

shrinkage without crack propagation should result in a more efficient conductive coating.

Hence, AZ Technology hopes that the hypothesis relates to the physical and mechanical

performance of this new binder is proven to be accurate.

24
I

09

0.8

07

O6

(l.5

O4

O3

0.2

o.I

o

_o

fJ _'\

J

/ \

\\\

\

\

\

K=_il Pmud_ in Pr_,cd Pellet. Alpha = 0.227

,-....

_1 550 7IX) 850 10(K) tiff) 1"_00 14.¢g) iffiX) 1750 lch_O

WA VIEI,.IE_If;TH _nm)

, . , , . , , . ,

Figure 20. Solar Absorptance of Potassium Silicate Binder

To demonstrate the potential of this binder AZ Technology currently has produced a

coating using this potential binder and calcined zinc oxide, shown in Figure 21. With a pigment

loading less than that used for AZ-93 and a similar cured coating thickness of 4.5 mils, we

achieved a comparable solar absorptance value of about 0.155. For a first attempt, these results

were quite promising.

25
0.7

_().5

0.3

0.2

0,!

o

2Sl • g) 550 70(1 $50 I_X) 1150 IWIO 14511 1600 1750 ltJO0 2050 2200 235f1 250fl 2650 2_X)

WAVELENGTH (rim)

Figure 21, Solar Absorptance Zinc Oxide Pigmented Proprietary Binder

4.2 Optimization of Electrically Conductive Ceramic Pigments

During this program, AZ Technology was engaged in the development of whisker type

material that had good conductivity. Some initial development on such material was conducted

under the SBIR program, but little progress could be made because of our focus on basic

semiconductors that could be produced having a color of white or black.

The synthesis and production of a whisker like semiconductor material will provide a

much better mean free electrical pathway for electrons to flow than the relatively poor one

produced by the typical particle to particle contact. AZ Technology did discover in a previous

study that very white and electrically conductive whiskers could be produced. But the yield was

very low and the whiskers were almost impossible to separate and collect from the

nonconductive residue left behind after the reaction was finished.

The approach taken in this program is to use what we refer to as supports. The term

supports as we use it, means that the semiconductor precursors are deposited on to a whisker

(short fiber). The supports that we are using are commercially available by products of the

composite industry and are available in a variety of materials. The types of materials available,

range from carbon to zirconium oxide. Processing starts by producing a solution that will react

to form the semiconductor. In AZ Technology's current process, the solution of semiconductor

precursors overcoat the supporting whisker with the material. The overcoated support material is

then dried and heat-treated to produce the final semiconductor having acceptable optical

properties e.g. white, black or transparent.

26
Although this processing appears easy, and has been successful so far, there still is a need

to investigate the various coating parameters and detailed techniques need to be worked out, it is

hoped that it will. However, at the present time, the knowledge base in this process is limited.

AZ Technology is currently working with four support materials given in Table 8.

Table 8. Support Materials

Material

Carbon /

Graphite

Quartz/

Silica

Zirconium

oxide

Silicon

carbide

Rational for use

Oz Reactive leaves hollow core, and is low

mass impact on coating

Transparent if mixed with many binders,

little effect on solar absorptance

Good reflectance throughout solar

spectrum. May help to enhance Orsof some

coatings

Good support material for black coatings,

good documentation in ceramic applications

Color

gray

Clear to

white

white

Gray to

black

Dimensions

(length by diameter)

3-8 mm x 8-12 _t.

10-14 mm x 20-30 IX

1 mm or less x 8-12 Ix.

3-8 mm x 8-12 Ix.

The primary challenges are to determine which technique is going to be better to use with

supports, and to determine coating material by mass or volume percent. Because we are dealing

with whisker materials that have a wide range of densities, determining a good ratio of whiskers

to solution is very difficult since we have a wide range of parameters to satisfy. Some of these

parameters are solution concentration, viscosity, coating technique, and drying. Even with all of

these variables to consider, this seems to be a viable approach as the results of a trial run shows

in Figure 22.

As is seen in Figure 22, using the previously discussed whisker coating method,

formation of the desired type of material can be achieved. These whickers were produced using

a carbon support and a fine particle dispersion of tin oxide. The resultant product after heat

treating is a white whisker that should provide a good electrical pathway with little effect on the

optical properties of the thermal control coatings. The measured resistance was 10 3 to 10 4 _'_]1-'I

in air using only a multimeter and straight probes. It is not likely that we will be able to obtain a

more accurate resistance value using the Hewlett-Packard electrical resistance cell, because of

the volume of material required, until these whiskers can be incorporated into a coating.

27
Figure 22. Synthesized Electrically Conductive Whickers

5.0 CONCLUSION

To judge the performance of the test chamber system, it was decided early in this

program to use known coatings that ranged from dielectric to a good semiconductor as reference

materials. The AZ-70 was the dielectric, AZ-93 was considered the poor semiconductor and

RM-400 a good semiconductor. Preliminary data recorded during the checkout of the test

chamber was found to be order(s) of magnitude or decades greater in surface charge than those

presented in Table 6 of this report. Early data did follow an expected logical trend, that being, as

one tests coatings from dielectric to a good semiconductor the surface charging of the coatings

decreases. The AZ-70 (dielectric) coating accumulated a surface charge (electron exposure only

at l na) in the order of approximately 2500 to 2700 volts, AZ-93 (a natural, but poor

semiconductor when used in this configuration) accumulated a surface charge in the order of

approximately 175 to 225 volts and RM-400 would not accumulate a surface charge during the

initial testing. In later tests these same coatings as shown in Table 6 have surface charges of -

101 to -153, -15.4 to -44 and -0.2 to -0.1 respectively. This dramatic difference in surface

voltage values demonstrates the importance of chamber and test design. It provides scientists

and engineers with material data tested using the best up to date system and test methodology

known at the time.

Even more import is how materials should be classified for use. Depending on the way

materials are classified whether by charge voltage, resistivity or conductivity could have an

impact on the cost and time in constructing a spacecraft. If resistivity is used, a much more

expensive material may be chosen with a long lead time, instead of a different, readily available

material that meet the charging criteria but not the resistivity requirements.

28
The phenomenon of a reduction in surface voltage upon exposure to UV radiation after

being exposed to only electron charged particles is poorly understood. In addition, it is believed

that this is the first time that it has been demonstrated using spacecraft thermal control coatings

as test samples. Although the magnitude of surface charge reduction is likely to be material

dependent, the effect of introduction and removal of UV radiation from an exposed surface is an

important factor to be considered by designers and could be used as an advantage for future

missions.

Evaluation of the spectra provided in Appendix A, exhibit very little change in the solar

absorptance of the coatings tested in this program. If the method and technique for purging the

chamber with nitrogen were successful, it demonstrates that all of these coatings have potentially

good resistance to the space environment. The RM-400, which had an epoxy binder, was an

exception to this and was only used because there was historical data from TSS. It was also

known to have and retain good electrical properties over several years.

The experimental coatings tested are all organo-metallic polymers based on nanotechnology to produce them. In addition, they have very good conductivity and seem to have

potentially very good optical stability. Such experimental coatings may be the way of the future

if test chambers like the one are designed and built with other radiation sources and in vacuum

solar absorptance measurement capability. In the case with optical measurements, this would

eliminate any possible unknowns concerning bleaching of the test samples. With an in-vacuum

solar absorptance measurement capability, investigators could also potentially correlate the

formation of mobile electron and any change in the materials solar absorptance. It is this

investigators theory, that if a coating cannot efficiently conduct away electrons, absorption bands

will occur in the near UV, near infrared and the mid-infrared. This will cause an increase in the

spacecraft's solar absorptance and potential heating up of the craft. In addition, if the electron

charge continues to build up, electrical arcing will occur and damage to or failure of the entire

system may result.

Much further work needs to be done using the type of test chamber developed during this

program. It could provide invaluable data for many types of materials used on or in a spacecraft.

Although there was little opportunity to optimize the coatings tested, this program has provided

insight as to how to proceed. The experimental coatings tested during this program will be aided

by the data obtained. With further work and understanding, it may be possible to record how

current and new materials charge and discharge and eliminate such problems from occurring.

29

APPENDIX A

A-1
0.9

0.8

0.7

1 i-

-E

1-

0.6-[

t

o.5-E

0.4-[

0.3 -_

0.2 -_

0.1-_

0 .r

250 400

AZ-70 prior to exposure, Alpha = 0.122

...... AZ-70 after exposure, Alpha = 0.123

550 700 850 1000 1150 1300 1450 1600 1750 1900 2050 2200 2350 2500

Wavelength (nm)

Figure 23. Solar Absorptance of AZ-70

0.9

0.8

0.7

o 0.6

0.5

_ 0.4

0.3

0.2

0.1

AZ-93 prior to exposure, Alpha = 0.155

...... AZ-93 after exposure, Alpha = 0.162

250 400 550 700 850 1000 1150 1300 1450 1600 1750 1900 2050 2200 2350 2500

Wavelength (nm)

Figure 24. Solar Absor _tance of AZ-93

A-2
1

0.9

0.8

0.7

0.6

0.5

_. 0.4

0.3

0.2

0.1

F

0

250

I

AZ-100 prior to exposure, Alpha = 0.150 j

I

...... AZ-100 after exposure, Alpha = 0.149

T i i i i i r i i

400 550 700 850 1000 1150 1300 1450 ! _,q0 1750 1900 2050 2200 2350 2500

Wavelength (nm)

Figure 25. Solar Absorptance of AZ-100

1

0.9

0.8

0.7

0.6

0.5

_ 0.4

0.3

0.2

0.1

0

I_ AZ-400 prior to exposure, Alpha = .295 ]

i

...... AZ-400 after exposure, Alpha = .329 I

250 400 550 700 850 1000 1150 1300 1450 1600 1750 1900 2050 2200 2350 2500

Wavelength (nm)

Figure 26. Solar Absorptance of RM-400

A-3
0.5

0.45

0.4

0.35

0.3

0.25

_, 02

0.15

0.1

0.05

0

250

_AZ-1000 prior to exposure, Alpha = .963

...... AZ-1000 after exposure, Alpha = .967

II

, , , , , i , , ,, _ , ,, , , _ , • ,, , , ,, , , ; , , ,, , , ,, , , ,, , , ,, , , ,, , , i , j

400 550 700 850 1000 1150 1300 1450 1600 1750 1900 2050 2200 2350 2500

Wavelength (nm)

Figure 27. Solar Absorptance of AZ-1000

1

0.9

0.8

0.7

0.6

0.5

,e, 0.4

0.3

0.2

0.1

0

250 400 550 700 850 1000 1150 1300 1450 1600 1750 1900 2050 2200 2350 2500

Wavelength (rim)

Figure 28. Solar Absorptance of AZ-2000

A-4
0.9 :

0.8 -

0.7 -

0.6 .

0.5 "

0.4

f

AZ-2100 prior to exposure, Alpha = O, 153

...... AZ-2100 after exposure, Alpha = 0.148

0.2

0.1

0 ........... ','' *,'' ','' ','' ','' ','' ',' _ i'' ', _ ' I ' ' ', ' '

250 400 550 700 850 1000 1150 1300 1450 1600 1750 1900 2050 2200 2350 2500

Wavelength (nm)

Figure 29. Solar Absorptance of AZ-2100

l °

0.9

i0.8 -_

0.7 -_

0.6_:

"_ o.5 .t

= [

0.4{

i0.3 J_

1:

o.2 -t

0.1-_

f

_ AZX-S 1 prior to exposure, Alpha = 0.144

...... AZX-S1 after exposure, Alpha = 0.145

250 400 550 700 850 1000 1150 1300 1450 1600 1750 1900 2050 2200 2350 2500

Wavelength (nm)

Figure 30. Solar Absorptance of AZX-S1

A-5
0.9 :

o.8

0.7 -:

E

0.6--

F

o.5_

o.4_

0.3 -E

0.2 -[

O.l*j F

0

250

...... _-s01 a_e,exposu,o, Alpha-- 0.365I

400 550 700 850 1000 1150 1300 1450 1600 1750 1900 2050 2200 2350 2500

Wavelength (nm)

Figure 31. Solar Absorptance of AZX-SG1

0.9

0.8

0.7

0.6

0.5

0.4

0.3

0.2

0.1

0

250 400

...... AZX-SG2 after exposure, Alpha = 0.377

550 700 850 1000 1150 1300 1450 1600 1750 1900 2050 2200 2350 2500

Wavelength (nm)

Figure 32. Solar Absorptance of AZX-SG2

A-6
0.9

0.8

0.7

0.6

0.5

0.4

0.3

0.2

0.1

0

250 400 550

_TMJ-810 prior to exposure, Alpha = 0.242 I

..... TMJ-810 after exposure, Alpha = 0.249

700 850 1000 1150 1300 1450 1600 17_0 1900 2050 2200 2350 2500

Wavelength (nm)

Figure 33. Solar Absorptance of TMJ-810

0.9

0.8

0.7

0.6

0.5

_ 0.4

0.3

0.2

0.1

0

250

_TMS-800 prior to exposure, Alpha = 0.352

...... TMS-800 after exposure, Alpha = .359

,

400 550 700 850 1000 1150 1300 1450 1600 1750 1900 2050 2200 2350 2500

Wavelength (nm)

Figure 34. Solar Absorptance of TMS-800

A-7

APPENDIX B

B-1
200

150

100

5O

lnA/cm 2 _ 5r_A/cm 2 J_

1 T

10nAlcm 2

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

2500 5000 7500 10000 125'_0 15000 17500

Accumulated Electron Counts

20000

Figure 35. Surface Voltage Measurements of AZ-70

1.E-10

'_ 1.E-11

1.E-12

L)

1.E-13

lnA/cm 2 _ 5nA/cm 2 + 10nA/cm 2

,.0 i E, ................. - - - - _l[

2500 5000 7500 10000 12500 15000 17500 20000

Accumulated Electron Counts

Figure 36. Calculated conductivity values for AZ-70

B-2
100

u_

o

>

75

5O

25

44------- lnA/cm 2 + 5nA/cm 2 _ 10nA/cm 2 I_

UV÷e- e-

• . ° . _o. • •

2500 5000 7500 10000 12500 15000 17500 20000

Accumulated Electron Counts

Figure 37. Surface voltage measurements of AZ-93

1 .E-10

'_ 1.E-11

_ 1E-12

0

1.E-13

lnA/cm 2 _ 5nA/crn2 _ 10nA/cm2

25 O0 5000 7500 10000 12500 15000 17500

Accumulated Electron Counts

20000

Figure 38. Calculated conductivity values for AZ-93

B-3
1 O0

I

tol

Ir ....... "-'-O ...................... " O- " ................... 41I

./

O-, o'"

_o ,q

• -0 -*1-

.... ....+

_e- and UV

. lOnA/cm z

1

0 5000 10000 15000 20000

Accumutative Electron Counts

Figure 39. Surface voltage Measurements of AZ-100

1 E-08

.........InA/cm2 ......... ._'_,F_, .........5nAlcm2 ..........,_. ..................... 10nA/cm2

1 E-09

=E

o

1.E-10

I E-11

.._ ..-0 ....................

1 E-12

• .......... 0"*-

1 E-t3

5000 10000 15000 20000

Accumulated Electron Counts

Figure 40. Calculated Conductivity Values for AZ-100

B-4
1 -

¢0

_0.1

0.O1 -

| |

10nA/cm 2

....... i n_Jcm 2 ......... _--5nMcm 2 ............ _#---_

_e- a_'ld UV

5000 10000 15000

Accumulatwe Electron Counts

20000

Figure 41. Surface Voltage Measurements of AZ-400

1 .E-O8

1.E-09

"o

g

1.E-10

.......... lnA/cm2 ............. ,_4- ...... 5nA/cm ......... _ .......................... 10nA/cm2 ............................. :'i

i

1 .E*I 1

5000 10000 15000 20000

Accumulated Electron Counts

Figure 42. Calculated Conductivity Values for AZ-400

B-5
10

1

_- ...... lnNcm 2 ........... ,_._---, 5nAJcm z ............. :_ ........................ 10nA/cm 2 _ ....................... ,:_

..oo.e-

•--.e.--- e- andUV

5000 10000 15000 20000

Accumulative ElectrOn Counts

Figure 43. Surface Voltage Measurements of AZ-1000

1 .E -O8

1 .E-09

E

¢J

=Eo

_, 1 .E-IO

.=

g

1.E-11

1 .E-12

..... ,, , ....... d

5000 10000 15000

Accumulated Electron Counts

20000

Figure 44. Calculated Conductivity Values for AZ-1000

B-6
100

=n

10

=p ...... InA/cm z ...........,_-- ..... 5nA/cm .............. "_

,Q'-'"• O ...........

+

lOnMcm z .......................

.... • ...................

----,(k--- e- and UV

1

5000 10000 15000 20000

Accumulative Electron Counts

Figure 45. Surface Voltage Measurements of AZ-2000

E

I E-IO

1 .E -11

--.e.-- +41- + +.l+ =ha uv l

+

• - .,. .... .Q ,"

t.E-12

5000 1 OOOO 15000 20000

Accumulated Electron Counts

Figure 46. Calculated Conductivity Values for AZ-2000

B-7
1000

100

I0

_------lnA/cm 2

...........4,_.----- 5nA/cm a ....... :_,_

lOn_cm 2

,.,, ..... • .....................

11

0 5000 10000 15000 20000

Accumulative Electron Counts

Figure 47. Surface Voltage Measurements of AZ-2100

I E-O8 -

I .E-O9

1 .E-lO

_o

o

1.E-11

1 E-12

1.E-13

5000 1OO00 15000 20000

Accumulated Eiectron Counls

Figure 48. Calculated Conductivity Values for AZ-2100

B-8
10

6')

0.1

._ ...... lnNcm2 ..... _*---- --SnA/cm2 -- f. ........................ tenA/cm2 ..............

¢, ,,_

,,--.0-,-e- and UV

5000 10000 15000 20000

Accumulative Electron Counts

Figure 49. Surface Voltage Measurements of AZX-S1

1.E-08 "." ............................. 10nA/c ................. ]

I nA/cm2- ........ +--- 5nA/c +

1.E-09

I.E-I0

1.E-11

0

/

==='_'_e- and UV

1.E-12

5000 10000 15000 20000

Accumulated Electron Counts

Figure $0. Calculated Conductivity Values for AZX-S1

B-9
10

_ ,<_---.--- I nA/cm z ..... -_-_ 5nA/cm 2 ......... +

10nA/cm 2

"e ....... •

• " " "'''O ............... " ...... 4

• I

0.1

0 5000 10000 15000 20000

Accumulative Electron Counts

Figure 51. Surface Voltage Measurements of AZX-SGI

I E-08

I E.09

E

1 E-10

1 E-_t

1_E-1

5000 10000 t 5000 20000

Accumula;ed Elecqron Counls

Figure 52. Calculated Conductivity for AZX-SG1

B-10
10

e

$

• .......... Q--"

0.1

0 5000 10000 15000 20000

Accumulative Electron Counts

Figure 53. Surface Voltage Measurements of TMJ-810

1 .E-08

1 .E-09

=Eo

_, 1.E-10

>

1.E-t1

t .E-12

............ Ncm" ......... t4_ ....... Atom a _[_ ..................... A/c mz ..... ,_

5000 10000 15000 20000

Accumulated ElectTon Counts

Figure 54. Calculated Conductivity Values for TMJ-810

B-ll
100 -

10

4 ...........lnA/cm 2 .........._ ............. 5nA/cm z ................ + ................................ 10nA/cm _ ........................................

t

0 5000 10000 15000 20000

Accumulative Electron Counts

Figure 55. Surface Voltage Measurements of TMS-800

1 .E-08

I E-09

E

1 IF-tO

1.E-1 1

1E-12

1 E-13

• m

5000

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

tO000 15000 20000

Accumulated Electron Counts

Figure 56. Calculated Conductivity Values for TMS-800

B-12
Form Approved REPORT DOCUMENTATION PAGE OMB No. 0704-0188

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1, AGENCY USE ONLY (Leave Blank) I 2. REPORT DATE

I

December 2001

4. TITLE AND SUBTITLE

Testing and Optimization of Electrically Conductive

Spacecraft Coatings

6. AUTHORS

R.J. Mell* and G.E. Wertz

7. PERFORMING ORGANIZATION NAMES(S) AND ADDRESS(ES)

AZ Technology, Inc.

7047 Old Madison Pike, Suite 300

Huntsville, AL 35806

9.SPONSORING/MONITORING AGENCYNAME(S) ANDADDRESS(ES)

George C. Marshall Space Flight Center

Marshall Space Flight Center, AL 35812

3. REPORT TYPE AND DATES COVERED

Contractor Report (Final)

5. FUNDING NUMBERS

NAS8-98212

8. PERFORMING ORGANIZATION

REPORTNUMBER

M-1035

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA/CR--2001-211411

11. SUPPLEMENTARY NOTES

Prepared for NASA's Space Environments and Effects (SEE) Program

*AZ Technology, Inc.; Technical Monitor: D.L. Edwards

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited

Subject Category 88

Standard Distribution

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

This is the final report discussing the work done for the Space Environments and Effects (SEE) Program.

It discusses test chamber design, coating research, and test results on electrically thermal control coatings.

These thermal control coatings are being developed to have several orders of magnitude higher electrical

conductivity than most available thermal control coatings. Most current coatings tend to have a range in

surface resistivity from 1,011 to 1,013 ohms/sq.

Historically, spacecraft have had thermal control surfaces composed of dielectric materials of either

polymers (paints and metalized films) or glasses (ceramic paints and optical solar reflectors). Very seldom

has the thermal control surface of a spacecraft been a metal where the surface would be intrinsically

electrically conductive. The poor thermal optical properties of most metals have, in most cases, stopped

them from being used as a thermal control surface. Metals low infrared emittance (generally considered

poor for thermal control surfaces) and/or solar absorptance, have resulted in the use of various dielectric

coatings or films

properties,

"i4. SUBJECT TERMS

being applied over the substrate materials

surface charging, coatings, thermal control

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

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18. SECURITY CLASSIFICATION

OF THIS PAGE

Unclassified

in order to obtain the required optical

19. SECURITY CLASSIFICATION

OF ABSTRACT

Unclassified

15. NUMBER OF PAGES

60

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Standard Form 298 (Rev. 2-89)

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