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Produced by the NASA Center for Aerospace Information (CASI)
JPL PUBLICATION 84-56

(NASA -C -175b59) 3I,PChl Gb FINRL N85-224E4

RICOMM,uNLATIONS cCr IMPS fNG1NtF IN(;-SC1Z14CF;

PAYLOAD (Jet FrcFulzici, Lat.) 144 p

HC Avo/(7F sG1 C3CL 22b Uaclas

G.3/16 14799

Report on Final Recommendations

for IMPS Engineering /Science

Payload

^2i7^^+zS^^^^

ct"?

1 September 1984

Prepared for

Air Force Geophysics Laboratory I

U.S. Air Force Systems Command

Through an agreement with

National Aeronautics and Space Administration

by

Jet Propulsion Laboratory

California Institute of Technoloqy

Pasadena, California
JPL Publication 84-56

REPORT ON FINAL RECOMMENDATIONS

FOR IMPS ENGINEERING/SCIENCE

PAYLOAD

Henry B. Garrett

Project Scientist, IMPS

1 September 1984

Approved: Approved:

Henry B. Garrett

Project Scientist

IMPS

Frank L. Schutz

Project Manager

IMPS

Prepared for

Air Force Geophysics Laboratory

Air Force Systems Command

Through an agreement with

National Aeronautics and Space Administration

by

Jet Propulsion Laboratory

California Institute of Technology

4800 Oak Grove Drive

Pasadena, California 91109
The research described in this publication was carried out by the Jet

Propulsion Laboratory, California Institute of Technology, and was sponsored

by Air Force Geophysics Laboratory through an agreement with the National

Aeronautics and Space Administration.

Reference herein to any specific commercial product, process, or service

by trade name, trademark, manufacturer, or otherwise, does not constitute or

imply its endorsement by the United States Government, or the Jet Propulsion

Laboratory, California Institute of Technology.
ACKNOWLEDGMENTS

The following is a list of those figures previously published and their

sources: Fig. 2, Miller, N.J., and Brace, L.H., J. GEOPHYS. RES., 74, 5742,

1969; Fig. 3, Baron, M.J., Repr. DNA 3023F, Stanford Research Institute,

Stanford, CA; Fig. 4, Watt, T.M., Incoherent scatter observations of the

ionosphere over Chatanika, Alaska, J. GEOPHYS. RES., 78, 3000, 1973; Fig. 5,

Hoffman, J.H., and Dodson, W.J., J. GEOPHYS. RES., 85, 629, 1980; Fig. 6,

Iijima, T., and Potemra, T.A., Large-scale characteristics of field-aligned

currents associated with substorms, J. GEOPHYS. RES., 83, 612, 1978; Fig. 8,

Murphy, G.B., Picket, J.S., Raitt, W.T., Shawhan, S.D., Electron and ion

density depletions in the STS-3 orbiter wake, PROC. OF THE SPACECRAFT

ENVIRONMENTAL INTERACTIONS TECHNOLOGY CONFERENCE, USAF/NASA, Colorado Springs,

1982; Shawhan, S.D., Murphy, G.B., and Picket, J.S., Plasma diagnostics

package initial assessment of the shuttle orbiter plasma environment,

J. SPACECRAFT AND ROCKETS, in press, 1984, Copyright American Institute of

Aeronautics and Astronautics, reprinted by permission; Fig. 11, courtesy of

D. Smart and P. Shea, Air Force Geophysics Laboratory; and Fig. 12,

Kessler, D.J., and Cour-Palais, B.G., Collision frequency of artificial

satellites: Creation of a debris belt, SPACE SYSTEMS AND THEIR INTERACTIONS

WITH EARTH'S SPACE ENVIRONMENT, H.B. Garrett and C.P. Pike, eds., AIAA Press,

New York, 707-736, 1980.

iii
EXECUTIVE SUMMARY

In support of its activities under the Interactions Measurement Payload

for Shuttle (IMPS) project, the Jet Propulsion Laboratory (JPL) established an

Engineering/Science Working Group (ESWG) made up of engineering and scientific

specialists in the fields cf spacecraft interactions and space physics. The

ESWG was formed to assist JPL in providing its sponsor, the Air Force Geophysics Laboratory (AFGL), with a mission concept for the IMPS project. IMPS, as

originally conceived by AFGL, is intended to be a Shuttle-compatible complement of engineering experiments and environmetal sensors capable of defining

spacecraft interactions in the auroral/polar environment. The purpose of this

document, as prepared by JPL and the ESWG, is to provide an estimate of the

impact of these interactions on materials, equipment, and technologies of

future AF systems and, based on their importance, to develop for JPL a generic

payload capable of carrying out the IMPS mission so that an accurate mission

plan and cost estimate can be prepared. It is also intended that the report

will provide a broad base on which future IMPS planning can be conducted.

This document addresses what are perceived as the key scientific and engineering concerns for AF space missions due to the peculiarities of the auroral/

polar environment. These concerns have been combined into 6 general categories: (1) dielectric charging; (2) material property changes; (3) electromagnetic interference, plasma interactions, and plasma wake effects associated

with high-voltage solar arrays and large space structures; (4) radio frequency

distortion and non-linearities due to the enhanced plasma in the Shuttle ram/

wake; (5) Shuttle glow and contamination; and (6) plasma interactions with the

space-based radar. Other (lesser) areas of concern considered were: interactions associated with EVA; radiation and single-event upset effects peculiar

to the auroral/polar cap environment; and space debris. During a series of

meetings at which numerous experts in addition to the core ESWG panel were

invited to attend and present ideas, a consensus position on the information

most critical to an understanding of these categories of effects was developed. The report describes the measurements needed to adequately address the

concerns associated with each category and includes a list of generic instrument packages capable of making the required measurements.

Although the ESWG did recommend specific instrument categories to be flown

on the first mission, the emphasis in the report is on the spectrum of

measurements necessary to quantize the interactions in the auroral/polar

environment. The specific evaluation criteria used in selecting the relevant

instrumentation were as follows:

(1) Is the interacticn effect being studied unique to polar orbits or

different in polar orbit than other orbits?

(2) Is the interaction expected to be unique to or enhanced by structure

size?

(3) Is the interaction relevant to planned AF systems?

iv
(4) Is the effect being investigated by other rrograms?

(5) Is it appropriate to carry out the investigation from the Shuttle?

(6) Can useful information be made available by 1990 so as to have a

meaningful impact on the next generation of AF spacecraft?

The intent of these criteria is to provide the justification, in terms of AF

objectives, for the IMPS mission. Given the potentially broad impact of some

of the interactions which meet these criteria on AF missions, the IMPS mission

will require a large cross-section of technical expertise ranging from basic

research (6.1) to advanced development (6.3).

Given the great complexity and cost of individual AF systems, the report

attempts to identify specific components (such as a segment of the space based

radar antenna) that should be tested rather than to recommend that the actual

systems be flown. The identification of the "weakest link," such as dielectric materials in the case of charging, is a major theme of the report. As

the environments encountered in the auroral/polar regions are known to vary on

the order of minutes by orders of magnitude, it is unlikely that ground tests

will be able to duplicate the actual space conditions affecting these components. It is therefore likely that meaningful component tests will in fact

not be capable of being performed on many systems until the IMPS investigations suggested by the report are completed.

This report is written with the intent of making its recommendations

available to the widest possible technical community within the AF. The

instruments have been presented in a generic sense so that program offices can

readily suggest refinements that would specifically benefit their programs.

Obvious candidates for material testing would be optical windows for planned

AF sensor systems, where contamination, surface chemistry, plasma etching, and

ion-induced color centers would be of concern, and dielectric antennas where

insulator charging noise could be an issue. Thus the report is intended as a

starting point for assisting in the development of actual flight test programs. In support of this, the report includes key contacts (the members of

the ESWG) and a detailed appendix of references by topic for each of the

interactions discussed in the report.

A suggested ground test plan for the IMPS project has been included as an

integral part of the report. The report concludes with a description of

proposed follow-on ?MPS missions intended to provide a complete, integrated

spacecraft interactions program. As time and resources permit, AFGL plans to

extend the data base on spacecraft environmental interactions through a series

of workshops on specific topics.

V
F^

F

P '

C

CONTENTS

EXECUTIVE SUMMARY

. . . . . . . . . . . . . . . . . . . . . . . . . iv

I. INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . 1

II. BACKGROUND . . . . . . . . . . . . . . . . . . . . . . . . . 3

t

A. AF SYSTEM IMPACTS . . . . . . . . . . . . . . . . . . . 3

B. THE NATURAL ENVIRONMENTS . . . . . . . . . . . . . . . 4

C. THE INDUCED ENVIRONMENTS . . . . . . . . . . . . . . . 12

III. THE INTERACTIONS OF SPACECRAFT WITH POLAR AND INDUCED

ENVIRONMENTS. . . . . . . . . . . . . . . . . . . . . . . 19

A. MAJOR INTERACTIONS . . . . . . . . . . . . . . . . . . 19

B. OTHER INTERACTIONS OF CONCERN . . . . . . . . . . . . . 26

IV. INVESTIGATIONS SUGGESTED TO QUANTIFY AF SYSTEM DESIGN

GOALS FOR POLAR ORBIT . . . . . . . . . . . . . . . . . . 31

A. DIELECTRIC CHARGING, MATERIAL PROPERTY EFFECTS, AND

ELECTROSTATIC DISCHARGE (DME) . . . . . . . . . . . . 31

B. EFFECTS ON SPACE-BASED RADAR (SBR) . . . . . . . . . . 39

C. HIGH-VOLTAGE SOLAR ARRAY EFFECTS (HVA) . . . . . . . . 39

D. CONTAMINATION EFFECTS (CEM) . . . . . . . . . . . . . . 43

E. WAKE CHARGING EFFECTS (WCE) . . . . . . . . . . . . . . 47

F. ENVIRONMENT AND INTERACTION MONITORING (.3IM) 49

G. ELECTROMAGNETIC INTERFERENCE MEASUREMENTS (EMI) 49

H. ELECTRICAL PROPERTIES DEGRADATION (EPD) . . . . . . . . 49

I. RF TRANSMISSION DISTORTION (RFT) . . . . . . . . . . . 50

J. ELECTRON-ION BEAM-INDUCED INTERACTIONS (EIS). . . . . . 51

K. LOW-LIGHT LEVEL TELEVISION MONITOR (TVM). . . . . . . . 51

L. GLOW MEASUREMENTS (GLW) . . . . . . . . . . . . . . . . 51

M. LARGE SPACE-STRUCTURE INTERACTIONS (LSS). . . . . . . . 52

PRECL1)IN(i PA(xI± F3LANK NOT FJL.? ik D

vii
N. SINGLE-EVENT UPSET ( SEU) . . . . . . . . . . . . . . . 52

0. SPACE-DEBRIS DETECTION (SDD) . . . . . . . . . . . . . . 53

V. GROUND-TEST PLAN . . . . . . . . . . . . . . . . . . . . . . 55

A. INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . 55

B. PREFLIGHT TESTING . . . . . . . . . . . . . . . . . . . 55

.^ C. POSTFLIGHT GROUND TESTING . . . . . . . . . . . . . . . 56

D. GROUND BASED MEASUREMENTS . . . . . . . . . . . . . . . 57

VI. IMPS ADVANCED CONCEPTS PLAN . . . . . . . . . . . . . . . . 65

` A. INTRODUCTION . . . . . . . . . . . . 65

B. INFORMATION COLLECTION AND PLANNING . . . . . . . . . . 65

C. SIMULATION AND MODELING . . . . . . . . . . . . . . . . 67

D. FOLLOW-ON MISSIONS . . . . . . 68

E. ANALYSIS/DATA WORKSHOPS . . . . . . . . . . . . . . 7

F. SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . 72

VII. CONCLUSIONS . . . . . . . . . . . . . . . . . . . . . . . . 73

REFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75

BIBLIOGRAPHY. . . . . . . . . . . . . . . . . . . . . . . . . . . 77

APPENDICES

A. CONTRIBUTORS . . . . . . . . . . . . . . 111

B. ESWG MEMBERS . . . . . . . . . . . . . . . . . . 117

GLOSSARY. . . . . . . . . . . . . . . . . . . . . . . . . . 119

Figures

1. The Earth's environment . . . . . . . . . . . . . . . . . . . . . . 

2. Electron temperature (K) as a function of geomagnetic

latitude and local time at 1000 km over Grand Forks during

May-June, 1965 . . . . . . . . . . . . . . . . . . . . . . . . . . 

viii

J
3. Temporal variations in (a) electron and ion temperature and

(b) density at 3000 km for a 24-h period during the very

intense storm of August, 1972 . . . . . . . . 8

4. Temporal variations in electron and ion temperature and

electron density for a 24-h period February 11-12, 1972 . . . . . . 9

5. As a function of invaviant latitude, toe (a) H and 0

concentrations, (b) H flux, and (c) H flow velocity . . . . . . . 10

6. A summary of the distribution and flow directions of

large-scale field-aligned currents . . . . . . . . . . . . . . . . 11

7. A sample of PDP measurements indicating the pressure and

- plasma effects of thruster firings . . . . . . . . . . . . . . . . 13

8. Summary of PDP Langmuir Probe electron density and temperature

as functions of vehicle attitude . . . . . . . . . . . . . . . . . . 14

9. PDP measurements of Orbiter transmitter and subsystem electromagnetic interference on the STS- 3/OSS-1 mission . . . . . . . . . 22

10. SCATHA P78-2 measurements of contamination rates as a function

of surface bias potential . . . . . . . . . . . . . . . . . . . . . 25

11. Integrated high-energy particle fluxes for a 400-km circular

orbit as a function of orbital inclination . . . . . . . . . . . . 29

12. The man-made debris environment . . . . . . . . . . . . . . . . . . 30

13. Standard sample trays to test many samples using common support

instruments . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35

14. A typical electrode-dielectric sample . . . . . . . . . . . . . . . 35

15. Typical telemetry controlled switchable connections to sample

electrodes . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36

16. Possible solar array panel configuration for IMPS . . . . . . . . . 41

17. Layout for IMPS solar array experiment . . . . . . . . . . . . . . 42

G 18. Ambient ion density in the vicinity of a 2-m satellite flying

" in the Shuttle orbiter wake . . . . . . . . . . . . . . . . . . . . 48

s 19. Ground-test support plan for IMPS . . . . . . . . . . . . . . . . . 57

E

`r

20. Meridional chain of radars near 70° West longitude . . . . . . . . 59

21. Master time line for IMPS, 1980-2000 . . . . . . . . . . . . . 66

ix
Tables

1. IMPS environment near large surfaces . . . . . . . . . . . . . 15

2. Instrument (or investigation) categories for IMPS versus

measurement parameters . . . . . . . . . . . . . . . . . . . . . . 32

3. IMPS support instrumentation for investigation categories . . . . . 34

4. Incoherent scatter radar facilities . . . . . . . . . . . 60

5. Technical characteristics of incoherent scatter radars . . . . . . 61

6. Geomagnetic parameters for incoherent scatter radars . . . . . . . 62

7. Contacts at incoherent scatter facilities . . . . . . . . . . . . . 63

8. Ground simulation and analytic model requirements compared

with interactions and flight experimer_ts . . . . . . . . . . . . . 69

X
SECTION I

INTRODUCTION

This document contains recommended measurements and instrumentation for

the Interactions Measurements Payload for Shuttle ( IMPS). The recommendations

are defined in terms of the spacecraft interactions relevant to the polar and

auroral Shuttle environments. The Jet Propulsion Laboratory ( JPL) and its

advisory panel, the Engineering/ Science Working Group (ESWG), have reviewed

the IMPS objectives and have recommended a list of applicable zategories :or

investigation. Investigation categories proposed in this e r^ort are as

fellows: dielectric charging; material property changes; electromagnetic

interf o ronce, plasma interactions, and plasma wake effects associated with

high°-vol;age solar arrays and large space structures; radio frequency

distortion and nonlinearities due to enhanced plasma in the Shuttle ram/wake;

Shuttlz, glow and contamination; and plasma interactions with the space-'used

radar. Other concerns considered are the interactions associated with EVA;

the radiation and single-event-upset effects pecsliar to the auroral/polar cap

environment; and space debris.

This report first provides a review of the environmental concerns and

basic technology issues addressed and the impact of each of these on Air Force

(AF) systems. Second, this report lists the measurements required to

characterize these interactions and recommend:, instrument packages capable of

obtaining the measurements. Subsequent sectioixs include a suggested

ground-test program and possible follow-on missions.

1

s ^^ y 
SECTION II

BACKGROUND

A. AF SYSTEM IMPACTS

This section provides the background justification for why the AF should

be concerned with the IMPS mission, specifically:

(1) The AF's role in space clearly dictates the need for vehicles in

high-inclination orbits in order to achieve mission goals.

(2) Proposed AF missions require that large or high-voltage structures

be placed in this environment.

(3) At the military has relied more on sophisticated electronic

surveillance, communications, and navigation systems capable of

autonomous operation, these systems have become increasingly

sensitive to the space environment, a sensitivity that needs to be

aCCirately assessed.

There is a growing urgency to evaluate in detail the effects of the space

environment on the long-term operation of military space systems. IMPS is

intended to address this problem for the specific case of low-altitude,

high-inclination polar Earth orbit (PEO) and for the case of operations in

conjunction with large structures such as the Space Shuttle that are capable

of locally altering the natural environment.

The objectives of IMPS, as defined by the Air Force Geophysics Laboratory

(AFGL), are to determine the effects of the space environment on:

(1) Optical systems as represented by cooled infrared detection systems

and laser systems (low power). Such systems are particularly

sensitive to surface contamination and to the hazard of "Shuttle

glow" at PEO.

(2) Large military structures such as the Space-Based Radar. These

systems will be sensitive to a variety of environments, particularly

plasma density, turbulence, and radiation unique to the polar/

auroral regions.

(3) Large, potentially high-power systems such as the Space Based

Laser. High-voltage systems will be affected by the high-density

ionospheric plasma at Shuttle altitudes and, possibly, by spacecraft

charging during auroral arc passage. Contamination and aging of

structural and optical components of such systems are also concerns.

(4) Manned operations requiring EVA during passage through the auroral

region. Again, passage through the harsh radiation and charging

environments associated with the auroral and polar cap regions poses

potentially serious effects.

3 ► fi^EN1WNALL^ BW

1 1 RECINANCt PAGE. BLANK NOT FIT AILn

its r-.
(5) Comunication and radar systems requiring specialized antennas

sensitive to polar environments. Antennas with dielectric materials

are sensitive to auroral region energetic electrons.

These systems areas are derived from specific mission concepts callcd out

in the A1AA/AFSTC Military Space Sybtem Technology Plan (MSSTP). Typical

examples of proposed AF space systems that would be impacted are the IR

step-stare mosaic surveillance systems, space-based LIDAR, neutral particlebeam weapon systems, medium altitude surveillance radar, intermediate altitude

phased-array radar, and the space-based laser. As discussed at length in this

report, numerous, potentially &-structive interactions with the space

environment are possible for tuese systems. At the least, the proposed

effects, —ny of which are u-Aique to the auroral/polar region, will. degrade

system performance to the r.oint of failure on long-term (10 or morc years)

missions. Thus the threat to AF systems must he seriously considered and

quantized where possible--again, the purpose of IMPS.

The information gained from the 1MPS program is expected to have broad

A

Impact across many AF systems in space. The polar space environment and the

alteration of the environment by large structures or high-voltage surfaces

will cause effects on virtually all apace systems. Examples of the effects

indicate the generic AF system applicability for IMPS:

(1) Auroral fluxes of keV electrons are capable of surface charging,

resulting in electrical breakdown of exposed dielectrics.

(2) Auroral electrons negatively charge spacecraft causing positive ion

impacts to occur at high energy 0100 eV) on wake surfaces of

large structures. High-energy ions cause sputtering and severe

chemical effects on surfaces.

(3) Large spacecraft with charged surfaces or high voltages .:an "focus"

ions or electrons onto sensitive subsystems.

(4) In polar orbit, auroral electron fluxes are encountered four times

per orbit throughout the li f e of the satellite. Severe surface

effects (deterioration, discoloration, etc.) are possible due to

deep electron penetration with resulting electric fields and

enhanced ion sputtering effects.

(5) ChArged or high voltage spacecraft surfaces in concert with

clectromagnetic fields may be able to trap high density plasmas in

localized regions near a large spacecraft. Such plasmas can

interfere with radiowave propagation or with antenna arrays.

There are many other examples but the point is obvious. The IMPS information

is neccusary before we design large or polar spacecraft.

B. THE NATURAL ENVIRONMENTS

At altituces below 1000 ka, the Earth's natural environment is dominated

by the ionosphere and the upper reaches o` the neutral atmosphere. A.1though

4
this is a dynamic region of high plasma and neutral density, most mission

planners tend to ignore it. In Fig. 1, typical profiles for midlatitudes are

presented showing how the neutral and charged-particle environments vary with

' altitude between 100 and 1000 km. The main features of this region are that

the densities decrease roughly exponentially with altitude and that the

environment changes from one dominated by oxygen to one dominated by hydrogen

and hydrogen ions. In addition, this environment is influenced by diurnal and

seasonal variations, changes in solar EUV (extreme ultraviolet) and heating

due to geomagnetic activity, interactions between the charged particles that

make up the ionosphere and the Earth's magnetic field, coupling to the neutral

winds, and chemical reactions. Thus, the Shuttle environment is extremely

complex and variable, characteristics considerably enhanced over the polar

caps and in the auroral zones by the following:

(1) Electron temperatures (Figs. 2-4) can undergo variations on the

order of 1000 K over scales of 10° latitude during periods of

intense magnetic activity, and on the order of 300-500 K during

normal conditions. Temporal variations of 1000 K for Te (100 K for

Ti) over 1/2 to 4 hours are typical during active conditions.

(2) 0+ and H+ densities (Fig. 5) exhibit significant (order of

magnitude) spatial/temporal variations at high latitudes in

conjunction with such phenomena as the "plasma trough," "ionization

hole," "light-ion trough," and "polar wind."

(3) Auroral electrons with fluxes of 10-100 VA m_2 with mean

energies in excess of several tens of keV occur in auroral arcs

during disturbed periods. Figure 6 indicates typical precipitation

regions of electrons and protons as a function of average particle

energy:

(4) Magnetic fields are nearly vertical (as opposed to horizontal near

the equator) with rapid spatial variations over the geomagnetic

poles.

(5) Cosmic ray and solar flare fluxes are within a factor of 2 of those

seen in interplanetary space. (Although at lower latitudes the

Earth's magnetic field shields these particles, direct entry over

the polar caps is possible along the magnetic field lines.)

(6) Significant sudden enhancements in neutral density are produced

through joule heating by intensified currents flowing during

magnetically disturbed periods.

Whereas the average low-latitude Shuttle environment has been reasonably

well modelled, the preceding characteristics have precluded the development of

similar models for the auroral and polar cap regimes. Only a few large and

extremely complex programs are currently capable of approximating these

regimes. Thus the impact of the auroral and polar cap natural environments on

space systems represents a unique challenge for the AF mission planner and

spacecraft designer.

5
1000

i

E

500

tQ

(b)

MAXIMUM OF

SUNSPOT CYCLE

MINIMUM OF

SUNSPOT CYCLE

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F1

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r

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(o) - MAGNETOTAIL

DEFLECTED SOLAR WIND PARTICLES

POLAR CUSP INCOMING SOLAR WIND PARTICLES

EARTH'S ATMOSPHERE (0-100 km)

PLASMA SHEET

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NEUTRAL SHEET

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BOW SHOCK

(c) T-- T- ,

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TOTAL

^'•^`

0 0

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t 

102

103 104 105 ,105 107 102 106 1010

1014

ELECTRON CONCENTRATION, electrons cm-2 NUMBER DENSITY, cm-3

Fig. 1. The Earth's environment: (a) geospace/magnetosphere;

(b) ionosphere; and (c) neutral atmosphere

6
24

Fig. 2. Electron temperature (K) as a function of geomagnetic

latitude and local time at 1000 km over Grand Forks

during May-June, 1965

7

W.

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KCOLLEGE

3000

1000

E 0.5

0

0.4

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W

a

fW

F^

0 2 4 6 8 10 12 14 16 18 20 22 0 UT

14 16 18 20 22 0 2 4 6 8 10 14 AS

Fig. 3. Temporal variations in (a) electron and ion temperature and

(b) density at 300 km for a 24-h period during the very

intense storm of August, 1972

9

t

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pO
6

0

5

M 4

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t

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n

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FEB FEB

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2500

>000

M

7

1500

W

W

1000

500

0400 0800 1200 1600 2000 0000 0400 GOO UT

1 I 1 1 1 I I I

KCOLLEGE

Fig. 4. Temporal variations in electron and ion temperature and

electron density for a 24-h period February 11-12, 1972

9
i

v

z

0

U

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-10 < DST < 10 0 5: <2-

MLT 0800-1946 1972

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INVARIANT LATITUDE, deg

ISIS-11 IMS

Fig. 5. As a function of inva iant latitude, the (a) H + and 0+

concentrations, ( b) H flux, and (c) H+ flow velocity.

Each point on the graph is an average of data from 5

to 30 satellite passes. Dusk is to the left, and dawn

to the right. The time period over which the data were

averaged is 4 weeks around winter solstice.

10
6 18 6

1

18

► ALI < 100y JAL ► ;, 100,.

12 12

0 0

^ o) (b)

_ CURRENT INTO IONOSPHERE

i•::::;:::::•^

rrrrT"CURRENT AWAY FROM IONOSPHERE

Fig. 6. A summary of the distribution and flow directions of largescale field-aligned currents determined from (a) data

obtained from 439 passes of Triad during weakly disturbed

conditions ( AL < 100 y) and (b) data obtained from

366 Triad passes during active periods ( AL > 100 y)

11
18 6 18 6

IALI ` 100,. JALI 3100r

12 12

0 0

( a) (b)

- CURRENT INTO IONOSPHERE

CURRENT AWAY FROM IONOSPHERE

Fig. 6. A. summary of the distribution and flow directions of largescale field-aligned currents determined from (a) data

obtained from 439 passes of Triad during weakly disturbed

conditions ( AL < 100 y) and (b) data obtained from

366 Triad passes during active periods ( AL > 100 y)
C. THE INDUCED ENVIRONMENTS

In addition to the ambient environment encountered between 100 and 1000

km, spacecraft possess locally induced environments by virtue of their

interaction with the ambient environment. The induced environment is often

dominant in terms of the effects on spacecraft activities, especially for

large structures. Potentially important to the local environment are the

outgassing and thruster effluents, near-surface chemical interactions

(contamination), generation of plasma waves and turbulence, optical emissions

(the "Shuttle glow"), enhanced (depleted) gas and plasma densities in the ram

(wake) region, and local generation of electric and magnetic y ields. Much

useful experimental evidence concerning these interactions is available from a

variety of measurements aboard STS-3, results of which are reported by Murphy

et al., Shawhan et al., Murphy et al., Raitt et a1., and Narcisi et al.

The "Shuttle glow" has been observed in STS-3 and STS-4 by Banks et al. and

Mende et al. 2 , respectively. These observations are briefly summarized

below (see also Figs. 7 and 8 and Table 1).

1. Neutral Gaseous Environrir.t

(a) A ram/wake vodulation of two orere of magnitude in neutral

pressure (lU 5 T to 1077 T: 1071 T - ambient pressure at

240 km) has been observed within the payload bay on STS-3.

(b) The initial bay pressure on STS-3 was 10 -5 T. It took Aearly

twenty-four hours to outgas to the ambient level of 10 - T.

(c) Short-duration (approximately a few seconds) pressure increases

of an order of magnitude typically accompanied attitude-control

thruster firings on STS-3.

1G.B. Murphy et al., Electron and ion density depletions in the STS-3

orbiter wake, PROC. OF THE SPACECRAFT ENVIRONMENTAL INTERACTIONS TECHNOLOGY

CONFERENCE, USAF/NASA, Colorado Springs, 4-6 Oct. 1982; S.D. Shawhan et al.,

Plasma diagnostics package initial assessment of the Shuttle Orbiter plasma

environment, J. SPACECRAFT AND ROCKETS, in press, 1984; G.B. Murphy et al.,

Interaction of the space shuttle orbiter with the ionospheric plasma, PROC. OF

THE 17TH ESLAB SYMPOSIUM ON SPACECRAFT/PLASMA INTERACTIONS AND THEIR INFLUENCE

ON FIELD AND PARTICLE MEASUREMENTS (Noordwijk, Netherlands), 13-16 Sept.

1983a; G.B. Murphy et al., Perturbations to the plasma environment induced by

the Orbiter's maneuvering thrusts, PROC. SHUTTLE ENVIRONMENT AND OPERATIONS

MEETING (AIAA, Washington, D.C.), 1983b; W.J. Raitt et al., Measurements of

the thermal plasma environment of the space shuttle, PLANET SPACE SCI., in

press, 1984; R. Narcisi et al., The gaseous plasma environment around the

space shuttle, PROC. SHUTTLE ENVIRONMENT AND OPERATIONS MEETING (AIAA,

Washington, D.C.), 1983.

2 P,.M. Banks et al., Spa,:e shuttle glow observation, GEOPHYS. RES.

LETT., 10, 118-121, 1983; S. B. Mende et al., Observations of optical emissions

on STS-4, GEOPHYS. RES. I-F T., 10, 122-125, 1983.

12
T I T

/— EAVG

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

.. . ................... E.....

DIFF

+8V/M

10

J

L O

GC

N

Q

"LdQ

g -- 0

N 100

^E

3 >

N

U ^

W v

10-5

106

z N ^

M 0

101

106

W '^

> N Y Z

Ili

 D

^p

V

101

10-3

ui

wj

UfH

pW[

1

10-7

+12

0

ui

-8V/M

0 1 2 3 4 5 6 7 8 9 10

FPEG

VERNIER I I THRUSTERS PRIMARY

1800 1810

1982, DAY 86

Fig. 7. A sample of PDP measurements indicating the pressure and

plasma effects of thruster firings

13

-12

F --
fi

WL' 10

m J

00

d >

G = ^

Q

L CIO

g^ 0

^ _ 100

^E

Q >

3

v -- ^ N

H = U ^ WJ

W 

v 10 -5

106

N

U F-

. 

z

N ^

n 0U

101

106

U ^

n z

O

U

101

10-3

`o

w

W

oc

a

10-7

+12

0

w

-12

I I

EAVG

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

.

EDIFF

i I I

+BV/M

-8V /M

0 1 2 3 4 5 6 7 8 9 10

- FPEG

VERNIER

PRIMARY — THRUSTERS

1800 1810

1982, DAY 86

Fig. 7. A sample of PD: measurements indicating the pressure and

plasma effects of thruster firings

13
el

8 1 = PITCH ANGLE

B2 = AZIMUTH ANGLE

Pq

108

106

N

Z

104

102

u 6

u; 4

QQ

ezW 2

180

m

--r—

RAM

WAKE

0

360

N

CD

01 I 1 -%­ 1 1 1 1 

10-3

u.i

4A

N

N

W

CL

CL

10-7

83:20 83:21 83:22 8

TIME

Fig. 8. Summary of PDP Langmuir Probe electron density and temperature

as function of vehicle attitude. Neutral pressure measurements are included for reference. The cross-hatched areas are

where the probe sweep saturates and the routine used to calculate Ne underestimates density by as much as an order of magnitud

14
a,

82

01 1

10-3

0

w

N

N

W

OC

d

10-7

83:20 83:21 83: 22 83:

TIME

Fig. 8. Summary of PDP Langmuir Probe ele:tron density and temperature

as function of vehicle attitude. Neutral pressure measurements are included for reference. The cross-hatched areas are

where the probe sweep saturates and the routine used to calculate N e underestimates density by as much as an order of magnitude

14
Table 1. IMPS Environment near large surfacesa

Parameters Ram Wake Comment

Neutral density, 1075 1077 Measured

torr

Plasma density, As high as 5 z 106 As low as 10 Measured

cm 3

Plasma waves Q0 Hz --300 kHz Low Measured

( 22V/m ) 2/MHz at peak electrostatic

waves

Energetic Mean energy of electrons: Low Higher fluxes

particles 10-100 eV predicted;

Flux: -.108/ cm2 sec ster eV little numerMean Energy of ions: ical data

10-30 eV published

Glow, photons 107 - 108 Low Glowing layer

(cm3 s )-1 in Ram 10-20

cm thick

aReference: H. A. Anderson, Induced shuttle environments, in Minutes of the

IMPS ESWG, Feb. 14-15, 1984.

(d) The major neutral contaminants around the Shuttle are H 2O and

He. The major engine exhausts are N 2 , H2O, and H2.

(e) Apart from thruster firings and water dumps, the quantity of

water vapor detected versus time correlates directly with the

temporal variation of spacecraft temperature.

Note: Shawhan et al. 3 do not believe that the measured high pressures

represent the ambient bay pressure (which they say is likely to be

considerably lower) but that the measurements are more likely connected with a

surface near the instrument. Therefore, all of the higher-than -ambient

pressures quoted above may be upper-limit values except close to the

spacecraft surfaces normal to the ram direction. Even so, the neutral gas

density enhancements should be studied because they are beyond current

expectations.

3S.D. Shawhan et al., Plasma diagnostics package initial assessment of

the shuttle orbiter plasma environment, J. SPACECRAFT AND ROCKETS, in press,

1984.

^ 15
2. Plasma Environment

(a) Plasma densities of order 2 x 10 6 to 2 x 10 7 cm-3 are

observed in the ram direction of STS-3, whereas ambient values

are believed to be less then 106 cm-3.

(b) A very significant plasma wake (4-6 orders of magnitude

depletion) existed behind the STS-3 orbiter.

(c) Orbiter-produced H2O+ ions are observed with densities

comparable to the ambient ions, Troduced from the •:apid charge

transfer reaction 0+ + H2O+ H^0 + 0. H30+ ions

are also produced from the H2O + H2O • H30 + + OH

fast reaction.

(d) Plasma depletions of about an order of magnitude have been

observed during VCS firings with enhancements of NJ,

NO+ , and OHS densities created by ion-molecule reactions

between 0+ and the exhaust gases.

(e) Electrostatic background noise dominates the electric field

spectrum from 30 Hz to 170 kHz with a peak in the spectrum

between 300 to 500 Hz. The noise variability exhibits a marked

orbital periodicity indicating its origin to be Orbiter

induced. No theory for the source of the noise has been

confirmed, but it may be associated with plasma instabilities.

3. Shuttle Glow

Characteristics of the shuttle glow phenonmenon as inferred from

currently available data are as follows4:

(a) The glow emanates from a layer 5-10 em thick just beyond

spacecraft surfaces facing in the l irection of the velocity

vector.

(b) The glow brightness on AE-C decreased exponentially with

altitude, with a scale height of roughly 35 km, consistent with

that of atomic oxygen at a temperature of about 600 K.

(c) The brightness of the glow was about 10-100 kR on STS-3 and

100-300 R on STS-4, the discrepancy being due to the higher

altitude and larger angle with respect to the velocity vector

on STS-4.

40rigin: see T.G. Sl.anger, Conjectures on the origin of the sOrface

glow of space vehicles, GEOPHYS. AES. LETT., 10, 130-132, 1983; for glow

brightness on AE-C, see M.R. Torr, Optical emissions induced by

spacecraft-atmosphere interactions, GEOPHYS. RES. LETT., 10, 114-117, 1983,

and J . H. Yee and V.J. Abreu, Visible glow induced by spacecraft -environment

interaction, GEOPHYS. RES. UTT., 10, 126-129, 1983.

16
(d) the dimensions of the glowing layer are consistent with an

effective radiative lifetime of the emitting molecule(s) of

a^out 5 ms.

(e) The radiation is believe] to be emitted over a continuum (not

in discrete visible lines) extending throughout the visible and

reaching peak intensities between 6000 and 8000 A.

(f) The glow is enhanced after firing the Shuttle ' s attitude

thrusters, whose effluents primarily consist of H2O.

(g) There is evidence from AE - C data that a significant brightness

enhancement of unknown origin might occur in the vehicle wake.

In addition, recently analyzed data from STS-8 (Aug.-Sept. 1983;

S.B. Mende, private communication) showed that the glow brightness and

thicknes p viewed on the ram side of the manipulator hand varied depending on

the typ:: of material held up. A somewhat contradictory observation, though,

was that oxidation of the surface materials did not lead to perceptible

changes in glow characteristics for a particular material. A more complete

analysis of these data will be forthcoming O .B. Mende, private communication).

The above sections represent a brief compilation of natural and

induced environment characteristics. In the following, the interactions and

their effects on potential AY systems are examined, and where appropriate,

specific experiments and instruments are recommended for investigating the

interactions of greatest importance.

17
SECTION III

THE INTERACTIONS OF SPACECRAFT WITH POLAR AND INDUCEn ENVIRONMENTS

A. MAJOR INTERACTIONS

1. Dielectric Charging

Grou::l experiments and in-flight experience such as on SCATHA have

indicated that the buildup of potential on exterior surfaces of a spacecraft

may not be the only source of bpacecraf` charging. The buildup of charge 1a

dielectrics and on interior isolated conducting surfaces may also be a

Potential source. Energetic electrons (typically greater than a fee 100 eV)

can penetrate a finite distance into a material before stopping. In the case

of a dieaectric, the charge can become trapped. Eventually, the electric field

in the dielectric will build up until the field either repels all incoming

electrons or exceeds the breakdowr_ potential of the material and an arc

occurs. With high auroral fluxes, the breakdown strength of co,umon dielectrics

(typically 10 6 V cm-1) can he reached after only one orbit. In the case

of conductive materials, the penetrating electrons can reach interior surfaces

and cause charging within shielded areas. Internal charging, as it !s called,

can ca, vie serious problems for spacecraft systems when, as an illustratior.

the dielectrir insulation in a cable breaks down due to the gradual buildul of

charge in it. With the increasing size of AF satellites, huge dielectric

areas will be exposed to the space environment and can build up

charge--potentially greatly enhancing the threat from this interaction.

The principal surfaces of concern are those that employ large areas

of dielectric or have conducting electrodes at potential differences of >100

V mounted near the dielectric surface. The fact that such dielectrics will

typically be organic, not ceramic, insulators may make them especially

sensitive to bombardment by low-energy articles. in low-Earth orbit such

surfaces are immersed in a plasma of to 10 6 electrons cm-3 density

and with a temperature of a few thousand degrees (a fraction of an eV). The

plasma in the frame of the Shuttle is doubly anisotropic due to both the

magnetic field and the velocity of the plasma past the vehicle, a process

which produces ram/wake effects. Thermal electror curr!nts to a surface area

are several thousand pA M-2 while ram ion currents are cn order of magnitude less. In addition, photoelectron currer.cs are produced in sunlight.

Although such low-energy particles will typically not cause dielectric charging, particles with energies of hundreds of eV or more can. There exist at

least three different situations in which significant numbers of electrons

with energies of several hundred eV -an be generated:

(a) The auroral vane is crossed four times per pass for a polar

orbit. The flu: of electrons from hundreds of eV to a few tens

of keV regularly reaches 10 - 100 uA M-2 (ions have a

somewhat lower flux) in this orbit. This high-energy flux is

downcoming while the flux of electrons backscattered from the

atmosphere is of nearly equal magnitude but of lower average

energy.

19 'L- 1^11E, IICMQy 8Lk0

PRP.1'FnR`G PA r,F. BLANK NOT MMED
(b) The plasma is accelerlted by biased electrodes so that some of

the particles strike conductive surfaces while others strike

dielectric surfaces.

(c) acceleration due to electric fields can be created by

transmitting antennas or complex spacecraft/plasma interactions.

Dielectric charging begets further charging. Electrostatic fields

produced by one surface segment cause spaceborne charged particle trajectories

to change, thus inducing further charging on. adjacent surfaces. A negatively

charged surface prevents solar induced photoelectron currents from escaping

adjacent surfaces. These effects, interacting with the space plasma and polar

magnetic field, produce complex interactions that cause the dielectric cnarging process to be very difficult to predict for large structures in polar

orbit. Without knowledge of these effects, it is possible to have enhanced

charging in a location on the spacecraft where charging is not normally a

cause of failure.

2. Material Property Changes

An important concern for the AF is the effect of the auroral/polar

environments on the properties of materials to be used on future large or

high-powered spacecraft in polar Earth orbit. It is expected that the

energetic, precipitating auroral particle fluxes associated with polar orbit

will result in significant degradation of surface and bulk mechanical and

electrical properties. These changes must be taken into account in the design

of spacecraft structures and thermal control systems. Ionic sputtering and

surface contamination are processes that can cause these effects. The

differential charging of dielectric surf.•- s relative to the structure may

also enhance them. in addition, arc disc;,arges. resulting from charge stored

in dielectric/metal configurations, may degrade surfaces. The sputtering or

arc discharge by-products may then redeposit on other surfaces causing

contamination and thermal problems. Cooled surfaces and biased surfaces are

believed to be especially susceptible and should be studied.

Typical spacecraft coatings that might be altered by these effects

are metal coatings such as indium tin oxide (ITO) and gold or dielectric

coatings such as silicon dioxide. These surfaces are used for electrostatic

charge control and thermal control. Other types of surfaces that should be

considered, because of pitting, discoloration, and discharging, are aluminum,

kapton, and various paints. Second surface mirror properties and solar cell

surfaces, especially when biased, may be particularly susceptible to degradation. It is clear that the study of the effects of the environment on all of

these surfaces under different AC and DC voltage biases (on the order of

±500 V) should be an important IMPS objective. Tn addition, as indicated,

contamination of these surfaces under biased conditions is also a critical

concern. However, as the IMPS mission is of relatively short duration (9 days

or less), active measures, such as the use of electron and ion beams, may be

needed to enhance the effects of the natural environment and to simulate more

hazardous environments. Such hazardous environments are not likely to be

encountered during the short IMPS mission but may be encountered on longduration AF missions.

20

_ 
3. HV Solar Array/Large Structure Interactions

A major characteristic of future AF space systems is likely to be

increased power and, by implication, size. Several such missions are called

out in the MSSTP and similar long-range plans. These missions are projected

for low- or mid-altitude, near-polar orbits requiring power in the range of

5 KW to 30 kW continuous with peaks up to 10 times the continuous requirements

on the vehicle bus. To accomplish these missions, the AF needs to understand

the interaction of large, high-voltage arrays with the space environment. it

appears that the low-altitude polar orbit presents a worst case environment

for large, high-voltage solar arrays with its varying gravitational and magnetic fields, high-density ionospheric plasma, high-energy auroral particle

fluxes, and the need for the solar array orientation to be maintained normal

to the Sun's rays. Even missions projected for midaltitudes and higher will

have a higher reliability if they are designed for the low-polar-orbit case.

New solar array configurations are under investigation for hardening to

nuclear and laser weapon effects that employ materials and designs markedly

different from present systems. Examples include concentrator and thin,

hardened GaAs solar cell arrays. The concentration systems primarily expoae

metallic reflecting surfaces not part of the electrical circuit. The thin,

hardened arrays represent a very low thermal mass and employ thin substrates

and structures. Both of these, as well as rigid arrays, are expected to

operate at voltages in the range of 200 to 500 V so as to support an internal

volta;e in the range of 150 V. These concepts need to be evaluated in the

auroral/polar environments. Early testing (FY 66 or 87) would be desirable

since it would allow data early in the array development cycle currently

planned by the AF and would assist in developing lightweight, high-voltage,

high-power systems.

It is clear from these considerations that solar arrays with potentials of 200 or more volts will be required for future spacecraft. Such

arrays will have exposed interconnects that can attract current from the space

plasma. As electrons, for roughly the same mean energy, have 40 or more times

the mobility of the ambient ions, the negatively charged collection area (for 

positive ions) must be proportionately larger than the positively charged

collection area (for electrons) to assure current balance. For spacecraft

without onboard plasma sources, this will result in the array floating predominately negative with respect to plasma ground. Thus, exposed array

interconnections will collect ions. Since the solar cell coverglass is an

insulator, it will remain at a potential near plasma ground, creating a substantial differential voltage with respect to the solar array conductors.

This differential has been observed to cause arcing both in laboratory experiments and in space experiments 5 . However, the susceptibility of modern AF

5 For labora~ory experiments - K.L. Kennerud, High voltage solar array

experiments, REP. NASA CR- 121280 (Boeing Co., Seattle), March, 1974; N.J.

Stevens et al., Investigation of high voltage spacecraft system interactions

with plasma environments (Lewis Research Center), NASA Tech. Memo. 78831,

F 1978; for space experiment - N.T. Grier and N.J. Stevens, Plasma interaction

experiment (PIX) flight results, PROC. OF THE SPACECRAFT CHARGING TECHNOLOGY

CONFERENCE, AF Academy, Colorado Springs, Oct., 1978 (NASA CP-2071,

AFGL-TR-79-0082 295-314), 1979.

21
I'

systems to such phenomena is, as discussed above, unknown, and the relative

sensitivities of two different array designs to the same environment have

never been studied. In particular, the variation of arcing threshold with

ram/wake conditions, solar illumination, and high-energy auroral electrons has

not been determined. Since recent studies indicate surface effects of

conductors may be important in initiating arcs, it is necessary to examine

discharge rates as a function of exposure to atomic oxygen. An IMPS highvoltage solar array (HVA) experiment would provide information that will lead

to improved solar array performance. Experiments are also necessary to help

determine safe operating voltages for AF systems.

Electromagnetic and electrostatic interferences of various types are

anticipated that could increase with the size of the structure and that may

seriously impact sensitive sensor systems. Electrostatic noise has been

observed on the Shuttle between 30 and 180 Hz (Fig. 9). Amplitudes measured

180

SHUTTLE PRODUCED LIMIT

COMPOSITE PAYLOAD LIMIT

\ ` S BAND

e 4m

\

\

\ 1

UHF

i

P DP

^ MIN PALLET

400 MHz

O MAX PALLET

.^ M',X FPEG

I 1 I

0 102 103 104 105 106 107 108 109 1010

FREQUENCY, Hz dB

Fig. 9. PDP measurements of Orbiter transmitter and subsystem electromagnetic interference on the STS-3/OSS-1 mission (reproduced

with permission)

22

160

x

140

E

im

0 120

J

W

U

OC

J 100

W

0

7

Q

0

p 80

W

60

40
-sere a maximum of (22.4) 2 V m-1 MHz -1 at 300 Hz (for a receiver passband

of 100 hz, which represents a field strength of 0.22 V m - ). This noise,

observed both within and outside of the Shuttle bay when the detectors were on

the ram side of the Shuttle, decreased to near zero on the wake side, implying

that the phenomenon is associated with the ram/wake orientation. Recent

theories imply that the field strength of this noise would increase with the

size of the generating structure.

Data on the phenomenology of the EMI sources in the auroral/polar

environments, their location, amplitude, waveshape, modes, and spectrum, are

all important in designing effective countermeasures to electromagnetic

interference (EMI). (By phenomenology is meant the environmental conditions

such as plasma density [ram-wake], temperature, and ionic composition as well

as the physical configurations and materials involved in the EMI generation

process.)

4. RF Distortions

The Shuttle and similar large spacecraft will produce large ram/wake

effects. in addition, ram particles can sometimes be ionized by auroral

electrons producing transient "dense" plasmas. For antenna arrays on large

space structures, the unusual RF beam scattering properties of this extended,

nonuniform plasma may be a problem. Several specific concerns have been

identified that relate to these RF effects in the auroral/polar region. Three

such concerns are beam pattern distortion, EMI (discussed earlier), and

harmonic distortion. These concerns apply to communications links between the

space system and space- or ground-terminal points, to space-based radar

systems, and to systems with large reflectors and sensitive receivers such as

radioastronomy or RF surveillance of the Earth. Each is described below.

The first of the RF concerns is beam pattern distortion. Electron

density irregularities in the vicinity of an antenna and antenna feed system

can distort the far-field antenna patterns, reducing the main beam efficiency

and increasing the sidelobe levels. This effect is undesirable for either

receiving or Transmitting antennas. Severe destruction of the beam pattern

occurs if the plasma density leads to a plasma frequency comparable to the

wave frequency:

fplasma(Hz) = 9 x 103 N1/2

where N is electron number density in cm -3 . Densities of 10 8-10 12 cm-3 are

necessary to severely affect the 100 MHz to 10 GH range. Although natural

polar electron densities range only up to 106 cm-3, local ionization in

the vicinity of a large space structure may be significantly higher due to

auroral particle bombardment and other effects.

For densities expected in polar orbit the near-field phase pattern

can be modified so that the far-field pattern is distorted. If the density is

irregular in either space or time, the far-field pattern will cha3ge with

time. Ram/wake densities were observed to vary by a factor of 10 near the

Orbiter on STS-3 and STS-4. Such ram/wake variations would modulate the

far-field beam of any antenna-feed system to some extent. Even without the

large structure ram/wake, a spectrum of density irregularites exists in the

23

-r *.. . 
polar ionosphere due to auroral particle precipitation and auroral current

systems. These density irregularities hove scale sizes of centimeters to

meters.

The second RF concern, already discussed in some detail, is EMI.

EMI is a low-frequency RF problem that needs to be quantified so that the

threat can be avoided.

The third RF concerti is that due to harmonic distortion. The

presence of plasma and plasma irregularities in the vicinity of high-power

transmitters could cause nonlinear effects on the signals. If the plasma

tends to rectify, as at the surface of the spacecraft, then harmonic

distortions will occur. Also, irregularities can cause wave energy to become

trapped in a localized volume leading to nonlinear effects. Such effects

cause spectrum spreading and beam envelope spreading, which lead to

interference and reduced privacy.

5. Shuttle Glow and Contamination

Three facets of spacecraft contamination for IMPS to evaluate in the

auroral/polar environment are (1) contaminant modification of the

electrostatic charging and discharging of vehicle and payload surfaces;

(2) electrostatic charging and discharging effects on contamination; and

(3) optical contamination by the so-called "Shuttle glow" phenomenon. Common

contaminants can decrease the photoyield and can change the secondary electron

yield of surface materials. These changes will affect the response of the

vehicle and payload to the transient charging environments found at

auroral/polar latitudes. Further, since many contaminants are dielectrics,

they may serve as sites for discharges and cause extensive differential

charging.

Work by Clark ar..d Ha11 6 and others has indicated that charged

spacecraft surfaces may buildup contamination at a faster rate than uncharged

sufaces (Fig. 10). Dielectrics tend to have the largest differential

potentials relative to the spacecraft ground. Since many optical and thermal

surfaces consist of dielectrics, the implication is that these surfaces would

be rendered even more sensitive to contamination as a result of their ability

to charge and attract contaminant ions. This phenomenon is not well

characterized and requires study by IMPS at Shuttle altitudes in polar orbit.

As discussed in Section II.C, a major source of background optical

contamination may be the glow that has been detected on Shuttle surfaces

exposed to the ram environment. One apparent source of this "Shuttle glow" is

the interaction between the neutral oxygen environment and the leading

surfaces of the Shuttle (a plasma interaction may also be involved). It is

currently postulated that an actual chemical reaction takes place that leads

to the degradation and aolation of surfaces (particularly organic compounds

such as kapton). Aside from the obvious surface changes thus introduced, the

6 D.M. Clark and D.F. Hall, Flight evidence of spacecraft surface

contamination rate enhancement by spacecraft charging obtained with a quartz

crystal microbalance, SPACECRAFT CHARGING TECHNOLOGY, 1980, NASA CP-2182, 1981.

24
9

8

7

r

0

N

C

•5

E

T

OV

N

E 2

V

a ^

4 I I I

a

Z^ QUESTIONABLE DATA

O GOOD DATA

3 – M - -0.0019V + 3.08

_ — r

M - -u.0009V + 2.44 (ALL DATA)

1

0.

-i00L_-L0 100 200 300 400 500

RPA GRID BIAS, V

M ,

Fig. 10. SCATHA P78-2 measurements of contamination rates as a function

of surface bias potential (from Clerk and Hall, 1981). ML12-6

TQCM mass accumulation rates at various RPA grid bias potentials during: (a) winter experiment and (b) during summer

experiment.

I

25

a -!,j 4,
glow apparently associated with the process is quite bright (upwards of 100

kR, as mentioned earlier). This glow potentially threatens optical systems

operating in this environment. As the effect may be related to size and

neutral and plasma density, it will be necessary to carefully study trade-offs

between orbital altitude (at higher altitudes, because of the exponential

decrease of oxygen, it is assumed that the problem will disappear) and size.

Further, it is unknown how variations in the environment found at high

latitudes will affect the phenomenon. A key to understanding this effect

would be spectroscopic measurement of the glow during flight operations.

6. Plasma Interactions

One candidate system for future AF missions is the space-based

radar. This system and its large antenna will employ dielectric surfaces and

high-voltage AC fields that can induce adverse interactions. Several interactions have been proposed that are unique to this type of system. For example,

EMI due to dielectric discharge pulse noise initiated by high-energy (E > 1

keV) electrons encountered in the polar orbit and RF noise at the operating

frequencies such as observed on the Shuttle, would likely be enhanced for

large antenna surfaces and could cause serious interference with sensitive RF

detectors. Likewise, as discussed in previous sections, dielectric breakdown

on large surfaces, =vitiated by the high-energy polar electron fluxes and

powered to completion (full breakdown) by a high antenna voltage, could damage

the antenna structure itself and other supporting subsystems.

Solar panels, powered antennas, and large space structures will have

exposed potential surfaces that can cause interactions with the space plasma.

Such interactions can accelerate the plasma particles to energies greater than

100 eV. The process is not simple and the resulting current flows are

difficult to calculate. Theoretical predictors need to be developed and

compared with experiment. For example, a 1000 volt dipole 10 cm long will

collect a few milliamperes of current (magnitude very uncertain). Some of

this current will be collected by other materials adjacent to the dipole

causing damage to these materials. Codes to predict resulting plasma particle

energies/currents/trajectories need to be validated before biased elements can

safely be designed for large space structures.

B. OTHER INTERACTIONS OF CONCERN

1. EVA Interactions

Although no adverse interactions between astronauts and the space

environment have been observed, it is expected that as the EVA life-support

systems become more complex and astronaut activities more involved, a

potential hazard may exist. Specifically, as life-support control systems

change from those controlled by analog electronics to electronically complex

and active real-time processor systems, the potential failure and the effects

of that failure increase dramatically. It has been postulated that an

astronaut on EVA in the depleted plasma wake of the Shuttle could experience

severe spacecraft charging. The reasons for this are twofold. First, a

passive body in orbit at Shuttle altitudes normally charges (at most) to a few

26
volts relative to the ambient plasma. This is due to the low energy of the

ambient plasma. The high density of the plasma tends to "short" any

differential potentials that might be generated by shadowing or anisotropics

in the flux. As structures become larger, however, a large plasma void is

created in the antivelocity vector direction (wake). An isolated body such as

an astronaut on EVA in this void region may charge to a potential independent

of the ambient cold plasma. As only relatively high-energy (!:eV and higher)

particles can easily penetrate this void, under some conditions kV potentials

might be expected. At low latitudes, the flux of such particles is diminishingly small. At high latitudes, the auroral fluxes in this energy range can

be quite high. Further, these fluxes are along the magnetic field direction

(roughly perpendicular to the velocity vector at these latitudes) so that

indeed a high-energy-charging source may exist.

It is believed that an actual example of low-altitude charging in the

auroral zone was observed on the Defense Meteorlogical Satellite Project

(DMSP) satellite. Voltages on the order of 100-200 V were observed. Although

no such events have been observed during EVA, it is also true that manned

missions have yet to encounter a significant auroral flux. Further, the

simplicity of current EVA systems has rendered them virtually immune to such

effects. Plans for new, digitally controlled life-support systems and the

requirement for frequent EVAs in the auroral and polar cap environments may

change this, however.

2. Radiation and SEU Effects

Recently, much attention has been focused on the phenomenon of the

single-event upset (SEU). SEUs are the result of the sudden generation of

sufficient free carriers in an IC to cause an electronic upset. The increase

in free carriers is produced by the passage of a proton or heavier fast

particle. The upsets can be divided into two categories:

(1) OFF to ON (Memory Reset) caused by particles in the energy

range E : 100 MeV/nucleon

(2) Latch-up caused by particles with E greater than a few hundred

MeV/nucleon

Characteristic of these interactions, the effects are at a maximum near the

stopping point energy of the nucleon (typically a few 100 MeV). With the

increasing sophistication and the movement toward spacecraft autonomy of AF

space systems, it is clear that multitudes of ICs with potential sensitivity

to SEUs will be employed. Thus SEUs, which could disrupt or confuse

sophisticated sensors and autonomous control systems, are a serious threat for

AF space systems.

I.n low-Earth orbit, high-Z galactic cosmic rays cause the first type

of interaction, while any cosmic rays of over a few tens of MeV energy, as

well o g cosmic ray secondaries produced in massive parts of a spacecraft and

in the atmosphere below, will cause nuclear collisions. Secondaries are both

charged and neutral particles including neutrons (cosmic ray albedo

neutrons). These effects occur in any orbit, but the cosmic ray fluxes will

be greater in polar orbit because of the so-called Stormer cutoff variation

27
with latitude (the magnetic field of the Earth causes asymmetries in the

incoming cosmic ray flux; see Fig. 11):

R min - 15 008 4 (LAT) (in GV)

with

LAT s latitude

R z magnetic rigidity, momentum/Z

Above about 60° latitude, the cutoff goes to approximately zero because the

polar cap magnetic flux is into the magnetotail. Further, even though a great

many measurements of cosmic rays have been made in the past 25 years, it will

be worthwhile to make an accurate calculation of the fluxes expected in polar

orbit on IMPS. As an example, theoretical calculations 7 imply flip rates of

10-8 upsets per bit- ay at low inclinations while at high inclinations, flip

rates as high as 10- may be possible. Thus, SEUs in the auroral/polar

regions represent an important concern for IMPS.

3. Space Debris

Kessler, Dour-Palais, and others 8 have systematically analyzed the

rate of buildup of man-made debris in low-Earth orbit. Their results

(Fig. 12) indicate that there is an increasing likelihood of collisions

between the debris itself and active satellites. Such collisions will be at

hypervelocity so that the resultant damage will lead to further fragmentation. Under sime scenarios, once a few collisions a year become commonplace,

the growth rate could become exponential so that within a decade space operations in the affected orbital regimes (typically in and near the Shuttle

orbit) could become exceptionally difficult. Given the additional factor that

the cross sectional area of space structures may also increase by orders of

magnitude in the same time frame, the problem of space debris--its size distribution, its orbital distribution, and its growth/decay rate--will be a

major concern by the early 1990's. IMPS, with its emphasis on environmental

impacts, makes an excellent platform for studying such space debris. As the

SPAS-mounted IMPS will be physically separated from the Shuttle, IMPS will

allow measurement of microscopic debris as a function of position relative to

the Shuttle as the Shuttle is maneuvered around it. This will make possible a

classification of the microscopic debris encountered into Shuttle-induced and

"environmental" debris.

7 P.A. Robinson, SEU rates for IMPS, Minutes of the IMPS ESWG, Feb.

14-15, 1984.

8D.J. Kessler and B.G. Cour-Palais, Collision frequency of artificial

satellites: Creation of a debris belt, SPACE SYSTEMS AND THEIR INTERACTIONS

WITH EARTH'S SPACE ENVIRONMENT, H.B. Garrett and C.P. Pike eds. (AIAA Press,

New York), 707-736, 1980.

28

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SECTION IV

i INVESTIGATIONS SUGGESTED TO QUANTIFY AF SYSTEM

i

DESIGN G 1- `.LS FOR POLAR ORBIT

Based on the requirements described in the previous sections, fifteen

t,.4neric categories of investigations capable of obtaining the required

measurements have been defined. In the following, the purpose of each of

these investigations is briefly described followed by a suggested list of

support instruments (,.he investigations are listed approximately in order of

importance as estimated by the ESW(;). At the end of this Section, for

reference, the investigations are listed in Table 2 along with the types of

measurements supported; Table 3 lists the various types of support

instruments. As should be clear from this latter table, the total list of

instruments necessary for a comprehensive IMPS mission is not as large and

diverse as might have been initially assumed. This finding should be of value

to mission planners in defining the IMPS payloads since the number of

instruments that must be considered is significantly limited.

A. DIELECTRIC CHARGING, MATERIAL PROPERTY EFFECTS, AND ELECTROSTA'T'IC

DISCHARGE (DME)

As discussed earlier, an important investigation for IMPS would be the

monitoring of changes in the bulk properties and surfaces of materials. In

addition, the pac,cage should allow the study of the effects of charge deposition in dielectrics and associated arc discharges. To analyze the effects of

spacecraft potentials on charge deposition, surfaces should be capable of bias

to t5UU V relative to the SPAS ground. Cryogenic cooling, to simulate

effects on IR sensors, would also be desirable. Typical sample surfaces would

be dielectrics, optical surfaces, mirrors, and solar cells. A number of

dielectric samples should be flown to look for discharge pulses. Since exposure time to the auroral fluxes will be relatively short and since the parameters of interest are numerous, a large number (>IUU) of samples are

required. Auroral electrons are likely to induce discharges in the dielectrics. Althoueh the passage through one auroral arc may be insufficient to

initiate a pulse, the effect is cumulative, with subsequent passes producing

discharges. It should be noted, however, that leaky dielectrics, produced by

sunlight shining on the material, would be less likely to break down by this

process.

The purpose of this DME investigation would be to measure several material properties which are expected to be important in polar orbit over a broad

class of materials. the experiment will be designed to simultaneously measure

a number of effects on many samples. Figure 13 shows how many samples can be

exposed to the Shuttle environment simultaneousl y . Simply exposing materials

to polar shuttle orbits will usually not be a sufficient test of the material. Some of the effects, which occur either slowly or rarely but cause a

significant effect when they do occur, need to be artificially enhanced during

a short duration mission like IMPS so that they can be studied; the experiment

described in Figs. 13, 14 and 15 is designed to do this.

jl
m t.

Table 2. Instrument (or investigation) categories for IMPS

versus measurements parameters

Instrument Categorya

Measurement

Parameter

DME SBR HVA CEM WCE EIM EMI EPD TVM RFT EIS GLW SEU LSS DDE

e

Material Properties

Bulk conductivity P S

Bulk charge P

Secondary emissiou S P

Photoelectron

emission S P

Surface ablation P S P

Contamination

Surface build-up P P S

Absorptance P P

hmittance P P

Reflectivity P P

Vehicle glow P P S

Induced contamination P P S S S

RF Contamination

Arc noise P S S P S

hMl background S S P S P

Propagation

effects P P P

hV Effects

Pcuer loss P P

Arc breakdown S P P S P

Single :'vent Upsets

Event detector ? P

Plasma hnviror.mevt

'_vuospheric plasma P S

Auroral fluxes P

Radiation fluxes P

Cosmic ray ilu.cs ? P

Atmospheric

Environment

Neutral density S P

Neutral

composition S P

Neutral pressure S P S

Ram/Wake Environment

Plasma density S P P

EMI noise S S P

Charging P S S S P

32
Table 2 (cont)

Instrument Categorya

Measurement

Parameter

DME SBR HVA CEM WCE EIM EMI EPD TVM RFT EIS GLW SEU LSS DDE

E&M Environment

DC magnetic fields

AC magnetic fields

DC electric fields

AC electric fields

Photon Flux

Visible

LUV

IR

X-ray

Neteoroid Environment

Particulates

P

P

P

S P S

? S S

? ? S

? ? S

S P

ap Primary experiment

S Secondary experiment

? Inclusion uncertain

The nlaema retarding ring/grid shown in Fig. 14 may be used to accelerate

ions and reject electrons from the plasma in order to study the effects of

icns on the sample. The ring/grid would be negatively biased to voltages of a

kilovolt or more under the restriction that the current drawn by the ring/grid

does not exceed a predetermined value and that the Shuttle does not change its

potential with respect to the space plasma. If higher energy ions are to be

studied, the high voltage can be applied directly to the sample while the

ring/grid remains closer to ground pc_enti.al. In this configuration the

number of ions collected will be lens but at a higher energy. Ring/grid

biases would be adjusted while considering the effect on adjacent samples (a

biased ring can affect the plasma particle trajectories near adjacent samples).

1. Pulsing in Insulators Caused by Auroral Electrons

It is estimated that one or a few passes through the auroral zones during

electron precipitation events would be sufficient to initiate electrical

pulsing noise in highly insulating materials. This needs to be studied in

both ram and wake configurations. The ring/grid can be used to simulate the

wake condition by biasing out the positive ions while letting the electrons

continue to bombard the sample. Ram conditions can not be simulated when in

wake. Poising would he monitored en the sample using electrodes which are

wired tc pulsed current detectors such as Lhe IDM detectors being developed by

JPL for the AFWL. In addition, radiated electromagnetic noise should be

measured by RF sensors placed as much as a few feet away from, and in front

33

o
34

{

1

fable 3. IMPS support instrumentation for investigation categories

W,.

Supporting

Instrument

Investigation

Category

Supporting

Instrument

Investigation

Category

Artificial Arc Source EMI Monitor-Albedo CEM

Beam-Electron EIS Neutral Pressure Monitor EIM

EPD RFT

Beam-Ion EIS Plasma Diagnostics Package EIM

WCE EIS

Biasing Supply CEM LSS

DME RFT

HVA WCE

SBR Samples-Dielectric DME

Contaminant Release Samples-IC SEU

Canister CEM Samples-Optical Surfaces DME

Current Monitor EMI Samples-SBR Antenna Segment SBR

EPD Samples-Solar Cells HVA

Detector-Arc DME Simples-Surfaces DDE

HVA EPD

WCE GLW

Detector-EMI EMI Spectrometer-EUV C.?.W

LSS Spectrometer-IR GLW

SBR Spectrometer-Neutral Mass CEM

WCE EIM

Detector-Faraday Cup EPD Spectrometer-Visible GLW

Detector-Particulates DDE Sun Sensor EPD

Free Flyer WCE HVA

I- E ger-IR DDE Surface Potential Monitor EIS

Impact Monitor DDE EMI

LSS TQCM CEM

Langmuir Probe HVA LSS

SBR. Telescope-Cosmic Ray SEU

Low-Light TV DDE Transient Pulse Monitor DME

EIS Transponder or Reflector RFT

GLW Variable Frequency

LSS Transceiver RFT

TVM
10 cm 25 c^

— -^• } bcm

RAD

SHIELD

POTENTIAL EMC

ACA DC AMPEpES PULSE mONITOR SHIELD

0-I keV 0-IA MONITOR

BASE POWER I TELEmETRY

Fig. 13. Standard sample trays to test many samples using common

support instruments. Each tray may contain up to 100

bins; each bin contains one sample.

PLASMA RET<AD RING/GRID

-----------------------^

I AT LEAST b FEED THRUS (PER SAMPLE) AVAILABLE FOR USE

Fig. 14. A typical electrode-dielectric sample. The feed throughs are

necessary to pass through the thick plate that protects the

electronics from space radiation as well as structurally

supports the samples and apparatus.

of, the sample trays shown in Fig. 13. The RF sensors should be looking for

rare pulsEs which last from 10 nanoseconds to 10 microseconds. The RF signal

level should be measured and if possible the frequency spectrum should be

determined (see the EM1 experiment for RF noise measurements).

Several electrode configurations will be necessary for this experiment

because large structures will have a large variety of dielectric materials

associated with them such as antenna elements mounted on insulators or glasses

mounted on meta] holders. These configurations should be tested using

appropriate electrode configurations, and in some cases bias should be applied

to the electroces. AC bias could be especially interesting because of the

unusual fields that could be created within the dielectric. Biased electrodes

should be wired with current limit i ng resistors so that a discharge pulse does

not overload the power supply.

s5
V

N Cu.i

Q <

K r ti J

V Q

J

Fig. 15. Typical telemetry controlled switchable connections to sample

electrodes

Pulsing should be monitored through several paths, between

(1) An electrode and ground

(2) Two electrodes on the same surface

(3) Two electrodes on differing surfaces of the same sample

(4) An electrode and the ring/grid

The pulsing should be monitored with and without bias applied to one of the

electrodes.

2. Surface Charging in Wake and by Auroral Electrons

It will be important to know the extreme potentials achieved by surfaces

in wake in the dark. One may do this by measuring the potentials of DME

sample electrodes or of conductive DME samples using the potential monitor

shown in Figs. 13-15. However, the validity of such data is strongly

dependent on experiment designs and on effects such as the "boot strap" which

ctiuse the potentials of nearby surfaces to be similar. The whole DME

apparatus may have to be electrically floated to do this experiment well. It

may turn out that the multisample trays shown in Figs. 13-15 are not

appropriate for determining the surface potentials on insulated surfaces of a

:Large structure. In order to properly design such an experiment, one should

36

4 
first decide exactly what information is needed and then use a surface

charging computer code such as POLAR or NAS(:AP to sce if an experiment can be

designed to get the data. It will often turn out that the surface potential

will depend mostly on interactions with other porti-ins of the spacecraft

through plasma particle trajectories and depend only weakly on the properties

of the sample in question.

It is of interest to measure differential voltages between small adjacent

samples on a large structure. Such data are presently lacking and the

computer codes are not yet reliable on this case. Thus in Fig. 15 is shown a

potential monitor which is basically an electioatatic probe to sense the

potential on any chosen electrode. The probe must be of very high input

impedance (infinite), and the telemetrically cor.tr . 11ed switches must also be

highly insulated for this measurement. It is nr n bably only necessary to have

one probe in each sample tray, and it will only be necessary to sense a few of

the electrodes; these electrodes should br cbosen based on analysis of

expected surface potentials, and the most extreme crises would be the logical

choices.

It should be noted that the surface potential monitors developed for

geosynchronous satellite!, are not appropri•:ce here because the surface

potentials are expected to be much smalle_.. However, such detectors could be

used to look for the expected extreme : -ia di g levels of a kilovolt in the

wake. This information will not be very material specific and will depend

greatly on the average over many materials on the wake side of the satellite.

Further, geosynchronous potential monitors measure the potential on the back

surface of an insulator and assume that the potential drop through the

insulator is small relative to the measured potential. Such an assumption is

not valid for much of the IMPS investigations.

3. Material Degradation (Surfaces)

Both the chemically reactive and the relatively high-energy species in

space are expected to have an adverse effect on sample surfaces. Large

structures are expected to alter the nature of this environment by introducing

chemical species as well as changing the trajectories of native species.

Surfaces charged by auroral electrons will experience bombardment by higher

energy positive ions. Some surfaces will experience enhanced ion bombardment

not because of charging but because of inadvertent focusing of the space ions

by electric fields generated by the structure itself. Thus, a test sample

with U bias (and with ± grids to keep out ions and electrons) should be

flown to be compared with an adjacent series of samples at increasingly more

negative bias. This will highlight contamination/damage by ions relative to

neutrals. Similarly, one may test for high energy "ram" neutral effects by

exposing one surface to the ram while a similar surface, facing down into

shuttle bay at all times, is only exposed to slow neutrals, never to energetic

neutrals.

A series of experiments using bias and sample covers/position can be used

to distinguish effects by various spaceborne species. However, Air Force

problems associated with large structures and high voltages should be related

to accelerated charged species bombardment effects. These studies should be

emphasized for AF studies as NASA and others look at nonaccelerated cases.

37
M

The following diagnostics will be used as appropriate for specific

samples:

(1) Surface resistivity

(2) Volume resistivity

(3) Surface voltage

(4) Secondary electron emission (see EPD)

(S) Dependence on temperature, comparing hot sample to cold sample

(b) Optical absorptivity, reflectivity (before and after flight)

(7) Mass loss/gain (before and after flight)

(8) Contamination (before and after flight)

(9) Morphology (before and after flight)

The above arguments concerning high- and low-energy ions and neutral particle

effects, as controlled b y bias and exposure, will also apply to experiments

using the above nine diagnostics.

Often, material degradation will depend on sample temperature. At the

least, the trays' temperatures will be monitored. If possible, a tray will be

designed to heat and/or cool its samples for a long time during the flight.

With the tray configuration shown, it is possible to measure the net

electrical current to the entire tray. The tray should be biased relative to

the shuttle ground. This provides a method for validating spacecraft charging

codes and current collection codes if current versus voltage is measured for

the entire tray. Applied bias is not expected to harm individual sample

experiments as long as it is applied only for a short time. These trays are

of sufficient size to validate the charging codes. Monitoring the (zero bias)

current allows a check on the incident space currents (auroral electrons,

ions, etc.) during the flight at the tray location. Thus the trays should be

grounded through an ammeter.

When devisiug an experiment scenario, the effect of a biased electrode

upon the fluxes to a sample in an adjacent bin should be determined. Because

the bin widths are larger than a debye length, the effect is not expected to

be large. Measurements should be performed to confirm this, however.

A number of specialized measurements should also be made. Such

experiments would be dedicated to a particular AF system component sample

requiring detailed testing. For example, reflectivity vs a or vs

accumulated exposure during flight could be measured on mirror materials using

laser reflectometer. Surface chemistry effects could be studied on a SIMS

apparatus (Secondary Ion Mass Spectrometry) as well as by measuring sputtered

species. Such specialized experiments should be limited to important

materials where the effects of the polar environment are expected to be

significant.

38

Is 
Exposed electrodes should be biased by capacitors fed by a high impedance

(>104 ohms) source. This prevents a discharge from creating a permanent

arc which would drain the power source. The capacitor (>1.0 μF) provides

stored charge sufficient to cause larger arcs than occur in dielectrics

alone. A monitor can be provided in the 10 4 ohm line to indicate when a

pulse has occurred and can correlate with the EMI experiment.

B. EFFECTS ON SPACE-BASED RADAR (SBR)

To investigate the effects of plasma interactions with the Space-Based

Radar (SBR) as functions of plasma density and Shuttle orientation, an actual

sample of a SBR-antenna segment should be flown on IMPS. fo characterize the

environment directly in front of the SBR sample, a Langmuir probe or similar

low-energy plasma probe is required. To characerize any electrostatic

discharges or EMI generated by the operation of the SBR, an ESD-EMI detection

system is suggested. Finally, to study the effects of different biases on the

operation of the SBR, the sample should be biased relative to the Shuttle

using DC and radar RF.

It is recommended that IMPS not only fly but also operate the sample

section (1 m2 ) of the RADC/OC space radar antenna and expose the antenna

during its operation to the ram and wake. The antenna is a structure composed

of an aluminum sheet surrounded by an array of dipoles mounted on kapton

membranes. When it is operating as a transmitting antenna, it will have

roughly lUO W peak radiated power. During the IMPS flight, the segment should

be transmitting as much as possible (at least during 1/4 of the flight) using

pulses typical of radar. During nontransmission periods, the antenna should

be either in a pure receiving mode listening for EMI noise or in a lowfrequency to DC mode looking for dielectric discharges within its own

structure. It should also have up to 1 kV bias on its elements at times.

Because the de-phasing of SBR beams by the space plasma is so important,

the SBR investigations should be correlated with the RFT investigation as

discussed in Section 1.

C. HIGH-VOLTAGE SOLAR ARRAY EFFECTS (HVA)

Although it would be preferable to fly a large, high-power solar array to

investigate the interaction between exposed nigh-voltage (SOU V) surfaces

and space plasmas, this may n_L be possible during the early IMPS missions.

Therefore, it is recommended that on early IMPS missions some small array

segments ut about 112 ft 2 of 2 x 2 cm GaAs solar cells (70 cells) would be

flown to generate 70 volts open circuit and about 55 or 60 volts under load.

Flying a silicon cell panel of the same 112 ft 2 size would generate about.

5U volts at open circuit. The latter would also be biased to observe changes

in leakage to the plasma as a function of negative and positive bias up to

!lODU volts. Corona effects in current between parts of the array itse'f

would also occur at voltages of 300 t,) 500 volts. This would be especially

true at times of contaminant release when local pressure could approach

1U -1 Torr and the surrounding vapor/plasma is at least partially ionized.

This effect could be measured with a bare conductor held at vehicle ground,

1-1 i

39
instrumented and monitored to measure current vs. array bias voltage. The

silicon cell segment should be in two parts: one with interconnects not

shielded and one with the interconnects coated with inorganic oxides as

accomplished by the Plasma Activated Source (PAS) coating technique under

development by the Aero Propulsion Lab of AFWAL. The GaAs solar cell segment

could also employ the integral inorganic coating if its development status is

sufficient. At present the technique is only suitable with Si segment

technology. The current voltage curves should be obtained when the solar cell

array segments are illuminated. The cell temperature and sun angle should be

measured during testing. The combination of biasing the silicon cell

segment(s), measuring EMI, leakage currents, and corona between different

solar array parts, and correlating these measurements with the natural and

induced environments should provide a first look at high-voltage array

interactions with the PEO environment. Other critical parts of solar cell

arrays are the materials of which the arrays are made. These include

lightweight substrates such as mylar (treated) Kevlar and various laminates

for stiffeners and light weight. For the newer hardened concentrator array

systems, reflector components should be included and the effects of the

Shuttle and space environment on the reflective surfaces should be measured.

It is anticipated, however, that a qualified concentrator module could not be

prepared in time for the first flight and that achieving orientation of the

SPAS 1° normal to the sun would pose severe constraints on the SPAS

housekeeping stores.

In subsequent missions, the proposed approach is to deploy a 20-m2

array (i.e., 2-m x 10-m deployed length), with at least 2 m of active solar

cells, extending out of the Shuttle bay (it may be desireable to unfold this

array in two different shapes). Figures 16 and 17 show a possible

configuration. The segment should also contain a solar concentrator segment as

f •

well as provisions for testing candidate materials for solar array components

and other spacecraft components. The array experiment should be capable of

generating 500 V at about 200 W when oriented normal to the solar radiation.

The array segments should include the following experimental technologies:

(1) Two concentrator concepts hardened to laser effects

(2) Hardened, thin GaAs solar array segments

(3) Thin silicon solar array segment (both [2] and (3] would be used in

single and series segments to build up the voltage)

(4) Comparison of array segments with and without grounded conductive

covers

(5) Provision for biasing the experiment in increments to ±500 V

(6) Conductors arranged such that magnetic torques can be assessed

(7) A current generator so that currents up to 5 A can be imposed

(8) TV monitoring capability

(9) Ability to evaluate in sunlight and eclipse

110
W DEPLOYMENT MECHANISM

CONCENTRATOR MODULE 112 m2

J CONCENTRATOR MODULE 1/22

THIN HARDENED SOLAR CELL SEGMENT

Go k

THIN OR CONVENTIONAL Si

CELL MODELS

itI

OTHER MATERIALS

SOLAR CELL BLANKET

TYPE UNDEFINED

Fig. 16. Possible solar array panel configuration for IMPS

41
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42

Cam.
(10) Ability to evaluate in ram and wake and edge on

(11) Orientation and deployment mast structure

(12) Open areas for other material experiments

For studying (and, if necessary, simulating) EMI and related plasma

effects on the small or large arrays, electron and ion sources, ±500 V bias

supplies, Langmuir probes, differential chargeup monitors, arc discharge

detectors, spectrum analyzers, and electric and magnetic field sensors would

be required. Electric and magnetic fields should be recorded in the 1 kHz to

1 GHz range, while electric discharge current and voltage wave forms should be

measured in the 1-ns to 100-us range. A search coil magnetometer capable of

determining the absolute orientation of the array relative to the magnetic

field should also be flown. Sensors should be emplaced near and around the

EMI sources and at remote locations to define the applicable coupling or

transfer functions (see EMI). Finally, allowance should be made for

measurements of flexible body effects and array short circuits.

D. CONTAMINATION EFFECTS (CEM)

The need to characterize contamination in the vicinity of the STS Orbiter

in the polar auroral environment arises from the expected complexity of the

interactions. Ultimately, of course, any interest in contamination

measurements is based on the effects of contamination:

(1) Damage to the sensitive surfaces of certain detectors, optical

elements, and special materials

(2) Degradation of the field-of-view of certain instruments

Plasma/contamination interactions will affect the magnitude and nature of

these effects.

The rapid changes in material charge states (arcs) that occur in the

polar auroral environment will enhance the sources of contaminants. The

material itself and its breakdown products will be directly released by

arcing. Additionally, pre-existing particulates will be indirectly released.

The transport of all contaminants may be profoundly affected by the plasma

through ionization (of gases) and charging (of particulates) and their

subsequent motion in the induced and ambient electric and magnetic fields.

The deposition of contaminants on victim surfaces will be quite complicated by

the plasma environment. Plasma collisions will tend to remove contaminants.

At the same time, plasma-induced chemical reactions may change the

contaminants to forms that tend to persist longer or that have a tendency to

more readily collect more contaminants. The chemical changes would also

affect the physical (e.g., optical) properties of the contamination on the

surface. Finally, the altered physical properties of the surfaces due to

contamination, whether changed through further plasma interactions or not, may

interact through plasma charging to, in turn, affect sourceP and transport.

43
1. Requirements

The desired capability of a contamination monitor instrument for IMPS

includes the following:

(1) Measurement of auroral plasma-induced enhanced contamination sources

(2) Characterization of contamination transport in the auroral

plasma/STS Orbiter environment

(3) Measurement of contamination deposition as affected by the plasma

for a catalog of contaminants

(4) Measurement of the alteration of the physical properties of a

catalog of deposited contaminants by the plasma

(5) Determination of the feedback interaction (e.g., dielectric

contamination causing arcing that produces more contamination or

charging that enhances deposition at other specific sites)

(b) Measurement of the rate of contamination as a function of surface

potential.

Several difficult requirements follow from the need to characterize the

Orbiter/auroral plasma/contamination interaction so as to permit the use of

generic models and predictions for specific unflown arrangements. The data

should provide for the separation of plasma-interactive effects from the

effects obtained in the absence of the plasma. IMPS specific effects should

be separable from generic Orbiter effects. The competing effects than

comprise net contamination deposition should also be separable.

Two conclusions may be drawn from the discussion to this point. The

contamination monitor package will require a combination of instruments and

sensors and will likely not accomplish everything desired. If the additional

constraint of a 1967 availability is imposed, the possibilities are even more

limited. However, a very useful package is feasible and will. be discussed

below.

2. Approach

Any approach to contamination measurement may be cl,aracrerized as either

a technique for measuring contaminants in the vicinity of the sensor or a

technique for measuring contaminants striking and/or depositing on the

sensor. Most approaches are either better for determining molecular (gas)

contaminants or for determining particulate contaminants. For the pi.irposes of

this discussion, contamination sensors may be categorize ,' ^pt'_cai imaging

sensors, gas mass spectrometers, deposited mass sensors, thermal-optical

effects sensors, impact sensors, o r inactive collectors.

-- 
Optical imaging sensors are the principal means for the measurement of

contaminants in the field-of-view. In p '.Iuments in this category typically

resemble those threatened by such contamination. Examples are television and

film cameras, coronagrapns, laser scattering instruments, and telescopes. All

of these are especially sensitive to particles. however, direct depositio-.

on the external optical element can confuse the data interpretation. 'these

instruments also have limited sensitivity to small particles (e5 j,m).

Typically the measure m ent of contamination threats to an infrared telescope

requires the use of an infrared telescope. The measurement of gas species in

the field-of-view, in contrast, requires the techniques of absorption spectroscopy (i.e., a stable well-characterized light source at a known distance

from the detector), which is extremely difficult.

(gas mass spectrometers exist that can very accurately measure either

neutral contaminants or ionized contaminants. Since the source cannot be

pumped in vacuo, the measured quantity is by necessity the flux. Without a

separate determination of species temperature, however, the density cannot be

ascertained. Furthermore, the measured flux may be very specific to the

location and orientation of the spectrometer where orders of magnitude effects

are common. On the positive side, mass spectrometers provide precise species

identification.

Deposited mass sensors are represented by various forms of quartz

oscillators whose frequencies change with contaminant (mass) deposition. The

prominent types are quartz crystal microbalances (QCM) and tapered element

oscillating microbalances (TEOM). The limitations of these sensors are a

saturation mass load, temperature effects, no species information, and

difficulty in particle collection. The advantages include small sensor size

(permitting several measurement locations), very good sensitivity, simplicity,

and dependability. QCMs resolve temperature effects by pairing the exposed

sensor with an unexposed control crystal; TEOMs, on the other hand, have a

respectable dynamic range. A simple existing adaptation of the temperature

controlled QCM (TQCM) -rovides a partial solution to mass saturation and

species identification for deposited molecular contaminants by periodic

bakeouts which can remove most of the deposition and restore dynamic range.

The stepped bakeout after a fixed collection temperature provides information

on the activation energy of the deposit which aids in the species identification. Another adaptation of the QCM, its use as the detector element of a

retarding potential analyzer, permits species mass identification with an

auxiliary species temperature measurement. Furthermore, this technique has

been used successfully to quantize the effects of vehicle potential on mass

accretion. Finally a soft, sticky coating on either a QCM or TEOM can be used

to accomplish particulate collection.

Tnermal-optical effects sensors span the range from temperature measurements on thermal control surfaces to narrow-band optical transmission or

reflection measurements cif targets. The appeal of this approach is the

direct determination of the effect of interest. However, except in its most

complex form (i.e., full optical spectroscopy), it cannot determine species

and quantity. Fortunately, calorimeters exist that can directly measure a/e

(absorptivity to emissivity ratio) for a selected thermal control coating at

little cost in terms of data requirements, size, and weight. These sensors

are an effective compromise.

45

0
Impact sensors, such as Y-cut quartz crystals, can provide count rates

for the impact of large particles (such as space debris). With the development of a preflight calibration procedure, size distributions may a'so be

possible. Count rate saturation and "ringing" are problems with this approach.

Inactive collectors are designed to return collected contaminants for

postflight study. The most common form is the passive sample array. This

approach provides the opportunity to use materials of direct interest and

special materials chosen for selective contamination collection. Particulates

are best retained upon reentry by the use of diaphragms over the collecting

surface and traps with doors. To facilitate the separation of auroral environment effects from those of the rest of the orbit, consideration should be

given to synchronized shutter mechanisms. This additional feature is very

important for the integrating sensors (e.g., deposited mass, thermal-optical

effects, and inactive collector types).

Auxiliary measurements, if not available from other packages on the

spacecraft, snould include solar flux (orientation), ambient pressure, and

plasma temperature. An active release and moveable sensor package would alsu

De of value to allow the study of the evolution of artificial contaminant

clouds in the vicinity of the Shuttle and in the auroral zone. By the careful

timing of releases and placement of the free flyer instrument package, it

should be possible to spatially and temporally map the expansion of a contaminant cloud. by careful election of the contaminants released, the ef f ects of

unusual contamina;it sources such as laser by-products on representative material surfaces could be studied.

To summarize, the principal recommendation that arises from the preceding

is for a package both on the SPAS and in the STS Orbiter bay for the partial

separation of Orbiter effects. The SPAS-mounted contamination package sho;.id

have its active contamination sensors (at least) very near the dielectric

charging-material property package (DCM). An active release source should be

integrated into the Day package.

Each h:lf of the contamination mcnitor package should comprise a low temperature (-50°C) TQCM, an "ambient" temperature TQCM (+10°C), and a coated

TQCM (for particulates). Each half should also have two identical calorimeters and two identical passive sample arrays. the two calorimeters and the

two sample arrays should have a clock-driven shuttering system to expose one

at high latitudes (only) and one ac low latitudes (only). At least one of

each passive sample array's positions should be dedicated to a diaphragm or

trapdoor particulate trap.

This recommendation consists of existing designs, the only exception

being the synchronized shutter. As discussed, the development of the shutter

seems essential to IMPS purpose. The recommended contamination monitor

package and active release source for the first flight is a compromise that

could only be enhanced by the addition of other sensors as described above.

Measurements of contaminants that are ionized by aurorz or that are ionized

normi.11y when emitted are especially interesting since ionized contaminants

may remain with spacecraft for long times.

46
E. WAKE CHARGING EFFECTS (WCE)

To determine the likelihood and effects of charging of an electrically

detached body, such as an astronaut, in the Shuttle wake region during auroral

passage or during artificial charging, wake charging should be investigated.

Specifically, a recent survey of DMSP data has uncovered several cases where

precipitating auroral electron fluxes are both sufficiently intense and energetic to charge the spacecraft materials to high potentials. The usual

mechanism that limits the voltage attained by a spacecraft is the increase of

collected ambient ion current with increased negative surface potential. In

the wake of a large object it has been estimat e d that the collection of ion

current requires a much greater potential than for an isolated satellite

(Fig. 18). These high voltages could cause arcing during EVA, docking, or on

electrically isolated subsystems on a large space structure. However, the

theory of this phenomenon has not been fully developed and calculations are

very difficult to perform.

With the IMPS payload on SPAS, there -..ould be the opportunity to measure

the ion current voltage characteristic of a large object in the wake which

will help greatly in the validation of computational. models. The power necessary to maintain high negative voltages on the SPAS should be quite small as

the ambient (positive) current is very low in the wake. Assuming an effective

collection area of 10 m`, it is estimated that the maximum positive ambient

:urrents to the SPAS (ion current, secondary current, and photoelectron current) would be of the order of 10 mA and that in the dark wake region (where

the ion current and photoelectron current would be minimal) as little as 0.1

mA. Thus the electrical power required is less than 100 watts. The simplest

method for biasing the SPAS would be through a conducting wire tether to the

Shuttle. This would also allow for direct measurement of the Shuttle-SPAS

differential potential and current while using the least electrical power. As

currently conceived, however, the IMPS will be a free-floating object, electrically and physically disconnected from the Shuttle. Alternatively, therefore, the SPAS can be bombarded with particles from ion and electron accelerators on the Shuttle. Use of an accelerator allows the free flyer to be at

some distance but places constraints on the allowed orientations with respect

to ti.e geomagnetic field. In addition, if the current in the particle beam is

too large, a beam-plasma interaction may occur which complicates interpretation of the charge neutralization data. Interference wi t h incoming ambient

ion measurements could be minimized by using ions in th, dun whose atomic mass

is very different from oxygen.

The data rate for this experiment would be quite low, consisting of current, voltage, and ambient density and temperature measurements. Since the

results are a function of position in the wake, one second resolution seems

appropriate. Also required is the SPAS location relative to the shuttle

orbiter and the location of the Mach and sun vectors. Tne experiment should

be performed in darkness and in the shadow of the vehicle. To make comparison

with analytical models simple, the structure of the SPAS should be covered to

a great extent with conducting material tied to a common ground. if need be,

much of this material might consist of a transparent mesh. [his would provide

a collecri_-n area That is easily resolvable and geometrically simple enough to

be modeled accurately.

47

I a
E 25

V

Z

Q

0 20

45

40

35

30

15

10

5

0 5 10 15 20 25 30 35 40 45

GRID UNITS, m

Fig. 18. Ambient ion density in the vicinity of a 2-m satellite flying

in the Shuttle orbiter wake. How much ion current such a

satellite collects for a given voltage is the issue addressed

by the wake-charging experiment.

48

h 
The total duration of such an experiment would be on the order of a single orbit and during this time, to prevent interference, there should be no

otner exposed high voltage surfaces on the SPAS.

F. ENVIRONMENT AND INTERACTION MONITORING (EIM)

An important function for the IMPS mission will be to provide in situ

measurements of the charged particle, neutral particle, magnetic field, and

electric field environments within the Shuttle bay, in the environment perturbed by the Shuttle, and in the undisturbed ambient environment. Measurements of this type are also required for correlation with many of the other

instrument packages. Such an instrument package would consist of ambient

plasma detectors for low-energy ions and electrons (<10 eV), a magnetometer

(AC/DC), an electric field monitor (AC/DC), charged particle detectors capable

of measuring fluxes in the energy range of 10 eV to 4U keV, a mass spectrometer capable of measuring ion and neutral composition up to 64 AMU, and a neutral pressure monitor. The package should be placed above or outside the

bay. A detailed description of this package is currently being developed by

AFGL.

G. ELECTROMAGNETIC INTERFERENCE MEASUREMENTS (EMI)

The purpose of the EMI investigation will be to characterize the signals

due to electrostatic discharge and wake-induced electromagnetic interference

in the vicinity of the Shuttle. This interference could significantly impact

sensitive AF sensor systems. An arc detector and EMI measurement system would

form the base of such a system. To characterize the current flow through and

from surfaces and the potential changes associated with an arc discharge, a

current monitor and surface potential monitor should be included. As

envisioned, the system would be used to locate the arc by ranging techniques

and multiple sensors. An artificial arc source could be used for calibration.

Low-frequency EMI due to plasma interactions has already been discussed.

Large structures with biased or aurora-exposed dielectrics are expected to

experience occasional breakdowns producing high frequency EMi pulses. Detectors and pulse counters should register these events. Pulses are expected to

be of duration 1 to 100 ns and of amplitude 10 -3 to lU amperes over a range

from lU to ^UUU volts. The larger currents, voltages, and durations occur in

the same pulse, producing S X lU -3 watt sec pulses of 100 ns duration. Various '^F and EMI detectors are currently available for measuring these effects.

These detectors should be placed near the HVA and DME experiments where

dielectrics are being tested.

H. ELECTRICAL PROPERTIES DEGRADATION (EPD)

Given the fundamental importance of the electrical properties of

materials in the spacecraft charging computer models, it is necessary to

accurately characterize these properties in space. To determine the temporal

49

j* 
variations in secondary emission, photoelectron emission, and backscatter flux

as functions of solar incidence angle, plasma variations, and contamination,

an instrument capable of measuring these properties in real time should be

flown. Aside from selected sample surfaces, a Sun sensor (to characterize the

photoelectron emission rate), a Faraday cup detectDr (to characterize the

secondary fluxes), and a surface c • :rrent monitor should be included. A

low-energy electron beam would also serve as a calibration source relative to

the ambient electron flux.

I. RF TRANSMISSION DISTORTION (RFT)

To investigate the distortions in radio-wave propagation through the

Shuttle plasma sheath, it is suggested that a variable frequency transceiver

be used to measure radio-wave phase delays propagating through the sheath.

One possible configuration would place th r transponder at one end of tho bay

and a radio-wave reflector or transponder for returning the signal at the

other end. By properly orientating the ShutL!e (i.e., bay into the ram), the

signal could be transmitted alternately through the ram-plasma enhancement or

through the wake-plasma void. As significant phase shifts are expected below

1 GHz, the transmitter should be at this frequency and less. Alternatively, a

two-path phase interferometer could be constructed, one path being in the ram

plasma.

t In support of this instrument, several studies should be attempted prior

to the IMPS flight. On Spacelab --- 2, the PDP is to map the S- and Ku-Band

near-field beam patterns. Distortions of these patterns can be detected along

with density irregularities in a 50° inclination orbit that will sometimes

penetrate into the auroral zone. For EMI, the incidence of ground EMI should

be assessed on every Orbiter flight by monitoring the S- and Ku-band

receivers. On Spacelab-2, the PDP will have directional S- and Ku-band

receivers with the beam looking toward the Earth once per spin period. Any

ground-generated EMI could be detected by correlating the PDP and the Orbiter

receivers. For harmonic distortion, nonlinear effects in the range of 50 to

600 MHz can be studied in a systematic way with ground radars such as at

•Jicamarca, Peru, and the EISCAT in Norway. Although the transmitter is not

located in the ionospheric plasma, these radars can create V/m fields at the

ionosphere level.

Based on the pre-IMPS findings, more definitive measurements could be

defined for future IMPS missions. Such instruments would utilize the unique

characteristics of the the IMPS mission, namely the changing plasma densities

and magnetic fields in polar orbit. A possible suggestion for studying

beam-pattern distortions is to map in detail the near-field antenna pattern

for different irregularity conditions. To do this, two receivers at Ku- or

S-band could be used: One receiver would be placed in the main beam as a

phase and amplitude reference while the other would be manipulated through the

sidelobe pattern under different density and irregularity conditions.

Simultaneously, the electron density and irregularity spectrum should be

measured at both receivers. From the phase and amplitude plots, the far-field

beam could be constructed. However, a more viable alternative would be a

50
simple phase interferometer. For EMI, IMPS could carry sensitive, wideband

receivers to monitor a wide frequency spectrum for large signals of ground

origin. Probes to measure both the density and the irregularity spectrum are

also necessary. These measurements should be made in the auroral region since

the auroral plasma tends to be most unstable. Results from this investigation

would be critical to SBR studies.

J. ELECTRON-ION BEAM-INDUCED INTERACTIONS (EIS)

Although a major objective of the IMPS will be to investigate the charging and material properties changes resulting from the natural environment and

especially the aurora, it is quite possible that during the short duration of

a mission no very intense aurora will be encountered. The energetic auroral

particle bombardment of surfaces could be simulated or enhanced by particles

from an electron and ion accelerator of modest size. In contrast with the

situation achievable in the laboratory, the test surfaces would be bombarded

by kilovolt energy particles (whether from the accelerator or the aurora) and

the neutral and plasma components of the ionosphere which are impacting the

surface at the Shuttle orbital velocity of \,7 km/s. In both the natural and

simulated cases, the flux reaching the test surfaces should be monitored with

suitable detectors. A more detailed assessment of this experiment is

required, however.

K. LOW-LIGHT LEVEL TELEVI jN MONITOR (TVM)

To observe low-light level visual phenomena such as Shuttle glow,

thruster firings, electron-beam operations, arcing, and aurora, the IMPS

instrument complement should include a low-light level TV. This would allow

identification of the spatial location of the phenomena, estimates of their

amplitudes, and, when necessary (as in case of the electron beam), real-time

changes in experimental configurations to better observe a given phenomenon.

L. GLOW MEASUREMENTS (GLW)

To determine the variations in Shuttle glow during passage through the

auroral zone and polar cap, it is suggested that EUV, IR, and visible

spectrometers be flown on IMPS. A low-light TV or similar imaging system

should be used to identify "glowing" surfaces in real time. As this implies a

very high data rate, it is probably preferable to keep such an assembly in the

Shuttle bay during early missions. The glow should be measured as a function

of ambient plasma and neutral variations, materials, and Shuttle orientation.

This would involve measuring the glow at different locations around the

Shuttle during the mission. Induced conditions at a glowing surface should be

measured as well. The glow issue is not addressed in detail here since it is

being studied by the NASA Shuttle office.

51
M. !.ARU E SPACE-STRUCTURE INTERACTIONS (LSS)

Future space missions have requirements which necessitate very large

deployable or erectable structures. These new spacecraft will have dimensions

ranging up to kilometers and will use light weight materials to achieve the

required low density. There are several types of interactions particularly

relevant to large structures:

(1) Mechanical interactions: Large structures are subject to a variety

of torques including atmospheric drag, solar radiation pressure,

gravity gradient forces, and magnetic field-induced forces. The

large area-to-mass ratios of large structures will especially

require a careful evaluation of their distortion, deformation, and

misalignment, and ways in which these effects can be controlled.

The movement of large conducting surfaces across magnetic fields

will induce electric_ fields, currents, drag, and surface heating.

(2) interactions connected with use of on-board high-voltage systems are

considered elsewhere in this report and are not discussed here.

(3) The plasma wake leads to asymmetries in the charge distribution

around a structure. This asymmetry produces potentials on the

surface and differential charging. The wake can also induce plasma

waves which may interfere with RF communications.

(4) Contamination of the structure environment originates from at least

four sources: (1) exhaust gases or ions emitted by the attitude

control bystem; (2) gas leakage; (3) surface erosion by meteoritic

ablation and ion sputtering; and (4) cold photoelectron emission.

Some effects of contamination are degradation of solar cell and

thermal radiator efficiencies.

On later 1MPS missions, it is suggested that a large (>25-100 m in

characteristic length) structure be flown to measure the effects of such a

structure on the environment and the effects of the environment on it. The

structure should be biased and should include a plasma diagnostics package and

EMI and RF pulse detector to monitor plasma interactions. A TQCM, low-light

TV, and impact monitor should be incluued to study enhanced contamination,

glow, and debris impact effects. The changes in plasma and neutral densities

at various locations caused by erection of the structures should be especially

studied.

N. SINGLE-EVENT UPSET (SEU)

To summarize the SEU investigation, the IMPS should

(1) Fly samples of ICs likely to be sensitive to the effects of

i+ cosmic-ray-caused SEUs over the polar caps at Shuttle altitudes,

C^

M (2) Identify the incident cosmic ray fluxes responsible for these upsets

by including a cosmic ray telescope or a simple cosmic ray track

detector

52
To compensate for the short duration of the Shuttle mission, it would be

sensible to expose a large number of junctions (such as a mass memory) with

various amounts of shielding to measure the number of SEUs. It is also

possible to measure the charge deposition rate by an array of electronic

detectors (proportional and solid-state counters) or using a nuclear

emulsion. This might be done with the test memory as part of a coincidence

telescope with the counters to identify which cosmic ray caused an $EU. The

detectors should cover the relativistic region (E > 10 MeV nucleon 1 ) to

provide the necessary correlations between the environment and the

experimental results. A dosimeter would be a useful adjunct to this

experiment to help separate total radiation exposure effects from short-term

SEU effects. This study should be coordinated with existing DoD SEU programs.

0. SPACE-DEBRIS DETECTION (SDD)

Given the increasing concern with space debris at Shuttle altitudes, it

has become necessary to measure the size and distribution of such particles in

the vicinity of the Shuttle and along the Shuttle orbit. As demonstrated by

C. A. Maag (Private Communication, 1984), even microscopic particles can cause

significant damage to optical surfaces. It would be valuable therefore to

include an impact detector and returnable samples on the IMPS carrier. An IR

imager and a low-light level television would provide size and distribution

information of larger particles. Samples in the DME, CEM, LSS, SBR, EPD, HVA,

and EMI should be inspected after the flight to see if debris impact has

occurred. Such impact should cause measurable effects, and the cause should

be ascertained.

53
1JUT..EDING, 1 1 AGL BLAidK NUT IlhAfi?7

PAGE ^L INTN''TIONALLY BLANK 

E':

SECTION V

GROUND TEST PLAN

A. INTRODUCTION

A major consideration missing from the preceding IMPS instrumentation

description is that of a ground-test program to be carried out in conjunction

with the actual flight(s). Aside from the obvious requirements for preflight

calibration and payload integration testing, IMPS can be tested postflight

because of the inherent "returnable" nature of Shuttle missions. In conjunction with a well-thought out preflight test plan, such postflight testing will

provide a unique opportunity to compare inflight and ground-test techniques.

The intent here, therefore, is to describe in general terms a ground test program that would enhance the usefulness and would increase the understanding of

the IMPS data.

B. PREFLIGHT TESTING

Several examples in the following illustrate the value of ground testing

during preflight planning for the evaluation of the IMPS data return. The

major thrust is that the instruments planned for IMPS would be of value in

studying space interactions in general--both on the ground and in space. In

this concept, the Shuttle flight becomes an extension of the laboratory rather

than a separate entity. The laboratory used for this work must be a fullscale space plasma simulation chamber. Currently, large vacuum facilities

capable of holding the entire ?MPS/SPAS exist at NASA/JSC, NASA/JPL, and

AFDC. Both NASA/MSFC and AFDC should have plasma simulation capabilities by

the time of IMPS that could be used to test the IMPS components.

As a first example, the EMI and ESD instrument packages could be used for

several experiments before the IMPS flys. It would be very useful prior to

launch in conjuction with plas ►aa simulation studies to characterize arcs and

plasma noise to simplify that classification during the mission. The instrumentation should also be placed in the flight configuration and used to refine

the arc-location technique proposed for the flight.

As a second example, the EPD and DME could be employed to catalog Shuttle

material properties prior to the first mission. Not only would this significantly enhance the data return from the flight, but it would also be of general value in the understanding of Shuttle materials and how they interact

with the environment--currently a topic of very real concern. This is doubly

important as, to date, laboratory efforts at characterizing spacecraft charging properties have been minimal.

Development of the EIS electron- and ion-beam systems would complement

ongoing AFGL efforts at designing a charge control system for high-voltage

arrays. Charging sources similar to the proposed EIS have been utilized in a

number of chamber tests such as NASA tests of solar panel charging. These

sources are currently used to study the so-called beam plasma discharge phenomenon in vacuum chamber tests. The charging characteristics of the EIS

should make it useful in tests with the VOLTS hardware or the IMPS HVA in G trumentation. Much still remains to be learned from such chamber tests, and the

55
existence of the HVA prior to the IMPS flight would allow the study of a number of interesting plasma interaction phenomena.

Although there is at present no plan to define specific WCE instrumentation, several interesting experiments could be performed even at this date

with existing ground-based equipment. For one, the MMU and suit materials

should be characterized electrically, and the ICs being considered for incorporation in future versions should be evaluated for electrostatic sensitivity. This information, aside from influencing the planning of a WCE

experiment, would aid in the design of new, safer EVA systems. Likewise, the

potential danger that the astronaut might pose to a satellite subsystem due to

his own static charge should be investigated on the ground in vacuum chambers

prior to launch.

Many experiments require analytical modeling in order to determine the

meaning of the results as well as to properly design the investigation/

experiment. Wake charging and plasma currents to biased elements are obvious

examples. For the RFT, another example, the magnitude of the phase delay and

the narmonic distortion need to be estimated to determine the parameters for

the apparatus prior to design. Ionization of contaminants and E x B effects

on the transport of ions around shuttle need to be estimated in order to

develop a successful experiment. Development of such models needs to begin

now before poorly designed experiments are generated and meaningless data

compiled.

C. POSTFLIGHT GROUND TESTING

In addition to the preflight tests described above, there are unique

postflight ground-test opportunities afforded by the IMPS mission. Principal

among these are the opportunities to recalibrate the instruments and test

assumptions about how an event occurred by conducting chamber simulations with

the actual flight hardware. In particular, if an arc were postulated to have

occurred at a specific point and to have, as a result, certain electrical

characteristics, it would be feasible to test such assumptions by setting up

the configuration, synthesizing the arc source, and comparing the results with

the original observations. Ideally, wich an experiment would permit an unambiguous test of the assumptions.

Material samples can be retested to determine the effects of the space

environment on their properties (see Table 2 for a list of properties). If

the materials are properly handled, the effects of rentry could be studied

systemaLically. Testing of the material properties such as tensile strength

and reflectance of small portions of the Shuttle itself prior to launch and

after return would be of additional value. As already noted, several of the

IMPS instruments are capable of accomplishing this testing. Additional

testing using standard laboratory equipment would complement these studies.

This latter type of testing would be valuable in determining the actual

I

sensitivities of the IMPS instruments.

A final t y pe of postflight testing that would be of value is that

involved in reconfiguring the system. The initial flight in any series always

indicates ways to improve the basic design. With IMPS, as it is intended to

56

^.

. . .--t.0.
be reflown, the recovery of the payload will permit rapid redesign. Testing

of the new payload will benefit from the flight data and the postflight ground

testing. Given better knowledge of the effects considered critical, the

reconfiguration testing can concentrate on those areas.

In Fig. 19, a possible ground-test schedule is presented that incorporates the preceding ideas. The schedule centers on the launch date and indicates prelaunch and postlaunch activities. Prelaunch experiment calibration

and systems integration testing have been left out as these will be included

in the detailed IMPS mission plan that accompanies this report.

D. GROUND-BASED MEASUREMENTS

Numerous complementary observations in conjunction with 1MPS would be of

value to the mission. Besides the obvious value of magnetometer, riometer,

DMSP auroral photographs and electron precipitation measurements, all-sky

auroral photographs, and other measures of the gross features of th! magnetosphere during the mission, incoherent scatter radar measurements offer a particularly fruitful source of information on the ambient environment at IMPS.

The possibilities implied by such measurements are explored below.

FISCAL YEAR

1983 1984 1985 1986 1987 1989 1989 1990

PREFLIGHT TESTING

ARC SIMULATION

MATERIAL PROPERTIES l

VOLTS 1 C

CRRES C

ASTRONAUT SUIT CHARGING

( CHAMBER TESTS)

IMPS FLIGHT 1 C

POSTFLIGHT TESTINr

RECALIBRATION

ARC SIMULATION

MATERIAL PROPERTIES

RECONFIGURATION

IMPS FLIGHT 2 C

Fig. 19. Ground-test support plan for IMPS

W
The polar ionospheric plasma is characterized by considerable spatial and

temporal variability, as compared with lower latitudes, due to the strong

influences of convection electric fields and precipitating particles of magnetospheric origin. Implications for IMPS are that if the spacecraftenvironment interaction is to be understood, we must require that

(1) Near simultaneous monitoring of the ambient plasma environment (Te,

Ti, Ne, and composition) be provided for

(2) The measurements scenario allows for separating spatial variations

in the plasma "disturbance zone" from temporal variations due to

movement of the spacecraft through a highly structured and varying

medium

A single instrumented subsatellite cannot provide unambiguous separation

of the sources of variability of the measured "disturbance zone" plasma characteristics. A preferred approach might be a set of environment sensors

strung along the length of a moveable or even a stationary boom or an array of

small satellites orbiting in formation. In any case, on the initial and

future IMPS missions, both of these options may be precluded by budget considerations. For this reason, the possible contributions of ambient plasma

parameters as measured by incoherent scatter radars should be closely

examined. Incoherent scatter radars provide measurements of Te, Ti, Ne, and

plasma drifts between nominal altitudes of 100 and 1000 km and complement the

capabilities of satellites in that they are capable of investigating temporal

and sometimes spatial behavior from a fixed geographic location. Resolutions

of 40 km and 10 minutes are fairly typical; however, these can be reduced considerably, at the expense of spatial coverage and/or accuracy. The latitudinal

coverage afforded by the American sector meridian chain of incoherent scatter

radars is illustrated in Fig. 20. basic data relating to currently operational incoherent scatter radars are provided in Tables 4, 5, and 6. A list

of key personnel (contacts) at these facilities is included i- Table 7.

It is specifically recommended that Thomson Scatter ground support be

included as part of the overall IMPS ground support plan to provide information

on ambient plasma properties and to provide a context for interpretation of

onboard IMPS diagnostics. The meridian chain of radars at Sondrestrom, Millstone Hill, Arecibo, and Jicamarca would provide the capability to monitor Te,

Ti, and Ne from low to polar latitudes in the American longitude sector during

the mission and to overlap data with IMPS on selected passes. Due to the

steerable capability of the Millstone Hill radar, it is anticipated that this

facility would play the major role in such a ground support activity.

58 
Fig. 20. Meridional chain of radars near 70° West longitude. The radar

fields of view in altitude and latitude are shown in relation

to the size of the Earth. While these fields of view are

typical for most experiments, altitudes above 1 R E

 have

been obtained at both Arecibo and Jicamarca. Simultaneous

measurements can be obtained over an extended region of the

atmosphere between the polar cap and the magnetic equator.

59
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Radar

Dip Angle,

deg

Declination,

deg

Invariant

Latitude

deg

L Value,

RE

B

T

'

Jicamarca 2.2 2.2 18.2 1.1 0.23

Arecibo 48.4 -10 34.7 1.5 0.35

St. Santin 59.9 -4 42.6 1.8 0.39

Millstone 70.4 -14 55.3 3.1 0.47

Chatanika 77.1 27 65.7 5.9 0.49

EISCAT 77.6 1 56.9 6.5 0.46

Sondrestrom 80.4 - 39 74.4 13.9 0.48

a Calculated for 35U km altitude from the 1GRF (1980) model updated to 1983.0

62
Table 7. Contacts at incoherent scatter facilities

Facility Contact

Sondrestrom or Chatanika Dr. Vincent B. Wickwar

Radio Physics Laboratory

SRI International

Menlo Park, CA 94025

Ph. (415) 859-4782

EISCAT Dr. Murray J. Baron

EISCAT Scientific Foundation

Box 705

S-981 27 Kiruna

Sweden

Ph. 46 980 18740

Millstone Hill Dr. John Foster

Haystack Observation

Route 40

Westford, MA 01886

Ph. (617) 692-4761

St. Santin Dr. Michael Blanc

C.R.P.E.

4, ave:nve de Nep^un-2

94107 Saint-Maur CEDEX

France

Ph. 33(1) 886 84-66

Arecibo Dr. Jules Fejer, Director

Arecibo Ionospheric Observatory

P.O. Box 995

Arecibo, Puerto Rico 00612

Ph. (809) 878-2612

Jicamarca Bela G. Fejer

School of Electrical Engineering

Phillips Hall

Cornell University

Ithaca, N.Y. 14853

Ph. (607) 256-6471

+i

t

t

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1
a ..A. .. 

SECTION VI

IMPS ADVANCED CONCEPTS PLAN

A. INTRODUCTION

IMPS does not exist in a vacuum. Several other missions (see Fig. 21)

directly related to the 1MPS program goals are currently planned for the same

time frame. Likewise, the IMPS advanced program concept is not limited to a

single flight. It is necessary, in order to obtain the maximum value from

IMPS, that IMPS be integrated into other planned efforts and that the IMPS

program incorporate the results of these missions into its long-range plan.

Such a long range plan is developed below with emphasis on future IMPS payloads and missions into different space environments. Although not intended

as a detailed plan, the phased approach presented should provide the skeleton

for such a program.

In planning a long-range space program of the scope of IMPS and its companion flights, a phased approach is a necessity. Here, the long-range plan

has been divided into four phases: an information-gathering phase, a simulation phase, the actual flight phase, and the analysis phase. Each can exist

concurrently with the others (i.e., information will undoubtedly be gathered

throughout the program), and the process will be repeated for each flight.

B. INFORMATION COLLECTION AND PLANNING

The purpose of the first phase is to gather data. There are numerous

methods of accomplishing this, and indeed, the ESWG has had this as its primary objective. The four principal means that can be employed are

(1) Collecting documentation

(2) Conducting workshops/conferences

(3) Visiting key facilities

(4) utilizing a panel. of experts (i.e., the ESWG)

As an illustration of the first method. numerous literature searches were

carried out for AFGL on specific IMPS concerns by the ESWG and JPL. An extensive bibliography of papers on IMPS-related material has been prepared for

AFGL under this effort. Several reviews of spacecraft and plasea interactions

were prepared. In reference to method two, a workshop was held in December

19bl and a joint hF/I;ASA conference in October of 1983. Numerous facilities

such as NASA headquarters and AFWL were visited with the assistance of the

ESWG members, AFGL, and JPL, and data on li•;PS issues were collected.

Building on the IMPS data base, future flights should concentrate on specific interaction concerns. If funding permits, workshops on these interactions should become an integral part of the IMPS long-term program. In concert with these topical meetings, every two years a general conference should

be held (e.g., such a conference was last held in October 1983). Based on the

workshops and the conferences, she data base of references on spacecraft

interactions developed by the ESWG can be expanded and made permanent. The

65 _ 1^1IF+1'1^1Ot'18LL^
1980 1985 1990 1995 200

WORKSHOPS ♦♦ ♦ O 00 00 00 00 O

CONFERENCES ♦ ♦ O O O O O

INTERACTION MODELS

SHIELDING/ DOSAGE O(UPGRADE)

CHARGE DEPOSITION O

SEU

NASCAP

LEO ♦

PEO ♦

MATERIAL PROPERTIES O

GROUND SIMULATIONS

LEO

PEO O

AURORA O

GEO ♦

PAPS FLIGHTS

1 O

2 O

3 O

4 O

CDAW

CONTAMINATION O O O

ARCING O O

CHARGE CONTROL O O

MATERIAL PROPERTIES O O O

COMPANION MISSIONS

PIX 2 ♦

VOLTS 1 O

VOLTS 2 O

CRRES O

SPACE PLASMA LAB O

♦ COMPLETED O TO BE COMPLETED

Fig. 21. Master time line for IMPS, 1980-2000

F,: 66

Y ^ ^
rudiments of such a data base now exist. The material has been divided by

interaction effects and is currently being cross-referenced to the specific

systems affected.

Blue-ribbon panels, such as represented by the ESWG, should be organized

on a permar,^nt basis to advise AFGL on progress in mitigating individual problems and on future research. For example, panels should advise or, each of the

following:

(1) Space plasmas

(2) Ionosphere/atmosphere

(3) Radiation effects

(4) Charging/plasma interactions

(5) Contamination

(6) Material properties

(7) Astronaut safety

Furthermore, a master technology road map in the area of spacecraft

interactions should be developed (the rudiments of such a plan actually exist

within the joint AF/NASA technology program) based on the findings of these

panels.

C. SIn'LATION AND MODELING

For the second phase, the main thrust is to improve the capability to

simulate interactions. Giver the existence of the data base on spacecraft

interactions developed in phase 1, the adequacy of the existing models and

experimental data associated with the different interactions can be evaluated.

This information can be used to determine where simulation capabilities need

to be improved and where more data are required. Again, several approaches

are necessary and, as indicated in Fig. 21, this too is a continuing process.

Two approaches are considered here:

(1) Theoretical modeling

(2) Ground simulation

As in any scientific activity, the ability to control a given phenomenon

depends on the adequacy of the theoretical constructs used to define it. In

studies of spacecraft interactions, an adequate understanding of the phenomenon includes an understanding of the source (the environment), of the victim (the space system), and of the interaction (spacecraft charging, radiation

damage, etc.). The model then attempts to simulate the effects of the source

on the system. Currently, although fairly adequate models of the space environment exist and systems can be modeled to some degree, interaction models are

at a very rudimentary level in general (dosage and shielding calculations are

an exception). Thus the development of adequate models is a primary concerti.

An invaluable adjunct to such models is actual experimentation. Even

with the advent of the Shuttle, ground testing remains in most cases the

cheapest and easiest way to study many phenomena associated with spacecraft

r

'' I

67

i
interactions. It is suggested in this report that such ground simulation be

given at least as high a priority as the modeling efforts. The primary problems to date with ground testing have been difficulties with scaling plasma

phenomena and with simulating the space plasma characteristics. In a departure from previous studies, it is recommended here that specific facilities be

developed and dedicated to simulating each of the principal space plasma

environments (namely, the plasma associated with low-earth orbit, polar Earth

orbit, the aurora, and the energetic particles associated with the radiation

environment). LiKewise, adequate simulations of particular phenomena are also

necessary (launch conditions, rocket-plume effects, arcing, high-voltage

surfaces, etc.). Various examples of the required simulations and analytic

models in terms of the interactions and flight experiments are presented in

fable d.

D. FOLLOW-ON MISSIONS

Figure 21 shows several possible follow-on flights for IMPS. The intent

is to modify the IMPS payload so that the interactions typical of each of the

key space regimes (ionosphere, auroral zone, and polar region) discussed in

this report are emphasized. It should also be emphasized that the individual

!MPS instruments should be viewed as evolutionary so that follow-on versions

of the initial engineering experiments would be improved after eacn mission.

In this way, the basic IMPS complement could evolve into an increasingly more

complex and sophisticated sensor system capable of characterizing a variety of

critical interactions. Those missions in the actual flight phase, in

chronological order, are as follows:

(1) IMPS 1--Polar Earth orbit/auroral zone: This is the principal IMPS

mission now envisioned and outlined in this report. The SPAS will

probably be the carrier for this mission.

(2) IMPS /VOLTS (IMPS 2)--Polar earth orbit/auroral zone: A joint mission with the NASA VOLTS array would mutually benefit botn programs. It would afford IMPS the possibility of flying with a large,

high-voltage structure. For VOLTS, the IMPS diagnostic capabilities

would be of great value in studying interactions with the auroral

and polar regions. As IMPS has been designed with such a mission in

mind, no modification to the basic IMPS 1 package should be necessary.

(3) IMPS 3--Low-latitude plasmasphere/ionosphere-large structure:

Although the primary IMPS mission will pass through the low-latitude

regime, the mission is not optimized for this region, nor will it

necessarily fly with a large structure (100 m or larger). An actual

large structure (as opposed to the samples on IMPS 1), such as the

prototype of the space-based radar or the space station, should be

available by the time of this launch. Depending on the size and complexity of the structure, multiple EIM packages could be deployed to

simultaneously monitor the environment around the structure.

(4) IMPS 4--Polar cusps/magnetosheath: Although currently not really

accessible to the Shuttle, improvements in the Shuttle or the free

68

M.

A.
Ground Simulation

Develop "natural"

radiation simulation

Develop techniques fcr

simulating longduration exposure

Develop shuttle,

geosynchronous, and

polar simulation

environments

Develop arc simulation

techniques

Sheath and wake

simulation

Analysis

Charge deposition

modeling

SEU prediction

Astronaut EVA dosage

models

Develop NASCAP and

similar models

Develop low altitude

plasma sheath models

Develop material

response models

Chamber testing

Test shielding methods

Compare NASCAP and

similar models at

geosynchronous orbit

Test low altitude

plasma sheath models

Simulate launch

environment

Simulate space

contamination/

degradation

Develop thruster

simulations

Rocket plume

measurements

Radar cross-section

studies

Chemical reaction

studies

Compute simulation

of deposition

Test atmospheric

reactions models

Compare chemical

releases with theory

Simulate ionospheric

changes

Confirm magnetospheric

models

s

69

Cable 8. Ground simulation and analytic model requirements

compared with interactions and fligtl experiments

Interaction Experiment

Radiation High time resolution crass

Effects spectrometer on IMPS

Samples of high density

microcircuits on IMPS

Charge deposition

measurements

Spacecraft Charged beams experiment

charging

Sheath and wake

measurements

Photoelectron, backscatter,

and secondary emission

properties

Arc discharge simulation

and monitoring

Surface potential

monitoring

Systems response study

Large, High Exposed, high potential

Voltage surfaces of var y ing size,

Structures construction, and

potential

Sheath measurements

Test shielding methods

Contamination Surface contamination

monitors

Measure artificially

induced contamination

Environmental Launch exhausts measurements

Effects

Wake-induced wave

measurements

Meteoroid/debris

measurements

Chemical release ;tudies

Particle depletion

measurements

Microwave heating/turbulence

measurements

Measure ambient environment

E!
flyer should my ke a flight into this region possible in the indicated time frame. The IMPS package should be modified to allow

better detection of high time variations in the ambient electric

fields and plasmas in this region as the primary environmental issue

is the rapid fluctuation of these parameters. As AF systems will

occasionally encounter these rapid variations, it will be important

to characterize their engineering effects.

(Note: There will be a continuing need on each successive flight to incorporate updated AF hardware concepts and concerns within the IMPS generic

philosophy.)

E. ANALYSIS/DATA WORKSHOPS

The most critical phase for IMPS will be the actual analysis of the

data. Although as already indicated, invaluable data can be gained from

ground testing, analysis of actual flight data is the ultimate step in gaining

a real understanding of interactions. Further, for the IMPS program to be of

any lasting value, that understanding has to be documented. As turn-around is

a crucial issue in adequately disseminating the IMPS data, a carefully conceived data analysis plan incorporating real-time analysis, data workshops,

and quantifiable outputs such as MIL-STDS is a necessity. Each of these

issues will be addressed for the IMPS and its companion missions in this

section.

Real-time analysis of the IMPS data will be a necessity for some of the

instruments. In particular, much of the glow data will have to be returned in

real time as the operation of imaging equipment, although perhaps automatic,

will require careful monitoring when particle sources (thrusters, etc.) are

turned on or the imagers are moved to another position. Further, the status

of the aurora will need to be monitored in real time in order to predict the

encounter of IMPS with an auroral arc. It is hoped, in fact, that the IMPS

package can be configured in specific modes so as to optimize data collection

when passing through auroral features, thus making the data available for

real-time analysis. It is recommended that at least one such optimized

real-time run take place each day. Several candidates for such runs are

listed below:

(1) Auroral arc encounter: All instruments capable of recording rapid

variations should be in their highest time-resolution modes and,

where possible, the data should be broadcast back to Earth in real

time.

(2) Thruster firings/beam operations: Specific experiments to observe

the results of thruster firings or, if available, charged particle

systems should be developed. As was learned from SCATHA, such

operations can induce rapid plasma variations.

(3) EMI events: If the ESD/EMI detectors on IMPS report peculiar activity, such intervals would be logical candidates for quick analysis.

This analysis requires simultaneous EIM data (see Section IV.F).

70

a
(4) Contaminant releases: Past Shuttle flights have indicated that the

Shuttle-induced environment can significantly change over short

periods.

(5) !Major changes in Shuttle orientation: Changes in Shuttle attitude

relp tive to its velocity vector, the Sun, and the Earth's magnetic

field can all generate interesting variations.

(6) Movement of the SPAS: Real-time data analysis of the SPAS during

its movement would help to indicate locations of interest for

further study during the flight and for future flights.

Such real-time analysis will require the principal investigators to commit to a rigorous schedule during flight. Even so, as evidenced by previous

Skylab and Shuttle flights, the ability, based on real-time data, to reconfigure the experiments in real time is crucial. An integral part of the program should be a data management system capable of handling real-time needs.

Within the first year following (and, indeed, in the months preceding)

launch, a series of data workshops should be held, similar to the NASA Goddard

Coordinated Data Analysis Workshops (CDAWs). At these workshops, the IMPS

data would be available through the data-management system so that the experimenters could rapidly compare their results. This approach argues for a large

central processing unit such as a dedicated VAX and a number of interconnected

terminals. By the IMPS launch time (1966-88), such systems should be common,

and by the time of the later launches, standard. The key to the workshops,

However, is the selection of key topics such as "contamination," "arcing,"

"charge control," etc. By limiting each workshop to a key topic, it should be

possible to generate a report concentrating on that topic as the output of the

workshop. These reports should be directed toward improving the relevant MILSTDs and Guidelines.

A major conference such as held in 1983 should be timed to occur within

out to two years of each IMPS mission. These conferences should represent the

culmination of each mission and should have several sessions devoted to

summarizing the results. In particular, the parallel results from the

ground-test programs should be incorporated into the mission reports at this

time. The output from these conferences should be comprehensive missionanalysis reports.

Based on the conference reports and the workshop results, the updating of

the MIL-STDs and Guidelines should begin in earnest. A time table, spanning

the two decades of the IMPS missions, should be established for updating these

documents. A tentative time line is presented in Fig. 21. These updates

represent the primary goal of the IMPS program and should be given the highest

priority of any items considered thus far.

71

-6 .. ^ 
F. SUMMARY

The steps necessary for taking the IMPS and its sister missions from

concept to utilization have been documented in this section. The major value

of this presentation is that it organizes the IMPS mission into a logical

sequence of events. It should be remembered, However, as indicated in

Fig. 21, that the steps overlap and repeat. Even so, the progression is clear

and valuable for future planning efforts.

72
C SECTION VII

CONCLUSIONS

t This report has defined a baseline IMPS program. The description of this

program has provided both the justification (in terms of AF mission needs) and

the rationale for the proposed IMPS investigations. The environments and the

interactions of potential concern to AF missions through the auroral zone/

polar caps have been described. Key investigations and the necessary instrumentation for these investigations have been identified. A coherent, long-

' ranSe plan for multiple IMPS flights has been developed. It is concluded that

R' IMIS offers a significant opportunity to

(1) Improve the reliability of AF missions

(2) Improve the survivability of AF missions

Given the great expense of future systems, the IMPS program outlined in this

report is readily justifiable.

73
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Sharp, R. D., E. G. Shelley, and R. G. Johnson, "A search for helium ions in

the recovery phase of a magnetic storm," J. GEOPHYS. RES., 82, 2361, 1977.

Su, S.-Y., and Konradi, A., "Average Equatorial Radiation-Belt Environment at

Low-Earth Orbit," Draft for May 1980 AGU Meeting., 1980.

Theoretical Models

Spjeldvik, W. N., and T. A. Fritz, "Energetic ionized helium in the quiet time

radiation belts: Theory and comparison with observation," J. GEOPHYS. RES.,

83, 654, 1978a.

Spjeldvik, W. N., and T. A. Fritz, "Theory for charge states of energetic

oxygen ions in the earth's radition belts," J. GEO PHYS. RES., 83, 1583, 1978b.

Observations/Data

Fritz, T. A., and W. N. Spjeldvik, "Observations of energetic radiation belt

helium ions at the geomagnetic equator during quiet ccnditions," J. GEOPHYS.

RES., 83, 2579, 1978.

Plasmasphere

General References

Carpenter, D. L., and C. G. Park, "On what ionospheric workers should know

about the plasmapause-plasmasphere," REV. GEOPHYS. SPACE PHYS., 11, 133, 1973.

Chappell, C. R., "Detached plasma regions in the magnetosphere," J. GEOPHYS.

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85
Chappell, C. R., K. K. Harris, and C.W. Sharp, "The morphology of the bulge of

the plasmasphere," J. GEOPHYS. RES., 75, 3848, 1970.

Chen, A. J., and J. M. Grehowsky, "Plasma tail interpretations of pronounced

detached plasma regions measured by Ogo 5," J. GEOPHYS. RES., 79, 3851, 1974.

Chen, A. J., and R. A. Wolf, "Effects on the plasmasphere cf a time-varying

convection electric field," PLANET. SPACE SCI., 20, 483, 1972.

Chen, A. J., J. M. Grebowsky, and H. A. Taylor, Jr., "Dynamics of mid-latitude

light ion trough and plasma tails," J. GEOPHYS. RES., 80, 968, 1975.

Gringauz, K. I., and V. V. Bezrukikh, "Plasmasphere of the earth" (review),

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Nishida, A., "Formation of plasmapause, or magnetospheric plasma knee, by the

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J. GEOPHYS. RES., 71, 5669, 1966.

Analytic Models

Chiu, Y. J., J. G. Luhmann, B. K. Ching, and D. J. Boucher, Jr., "h ► :

equilibrium model of plasmaspheric composition and density," J. GEOPHYS. RES.,

84, 909-916, 1979.

Cowley, S. W. H., "Energy transport and diffus'on," in PHYSICS OF SOLAR

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Gringauz, K. I., and V. V. Bezrukikh, "Asymmetry of the earth's plasmasphere

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TERR. PHYS., 38, 1071, 1976.

Lemaire, J., "Steady state plasmapause positions deduced from Mcllwain's

electric field models," J. ATMOS. 'TERR. PHYS., 38, 1041, 1976.

Theoretical Models

Kavanagh, L. D., Jr., J. W. Freeman, Jr., and A. J. Chen, "Plasma Clow in the

magnetosphere," J. GEOPHYS. RES., 73, 5511, 1968.

Observations/Data

Ahmed, M., and Sagalyn, R. C., "Thermal Positive Ions in the Outer Ionosphere

and Magnetosphere from OGO 1 9 " J. GEOPHYS. RES., 77, 1205-1220, 1972.

brinton, H. C., J. M. Grebowsky, and H. G. Mayr, "Latitude variations of ion

composition in the mid-latitude trough region: Evidence for upward plasma

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E6
Burke, W. J., and Donatelli, D. E., "Injun 5 Observations of Low-Energy Plasma

in the high-Latitude Topside Ionosphere," J. GEOPHYS. RES., 83, A5, 2047-20565

1978.

Carpenter, D. L., "Whistler studies of the plasma pause in the magnetosphere,

1, Temporal variations in the position of the knee and some evidence on plasma

motions near the knee," J. GEOPHYS. RES., 71, 693, 1966.

Chappell, C. R., "Recent satellite measurements of :he morphology and dynamics

of the plasmasphere," REV. GEOPHYS. SPACE PHYS., 10, 951, 1972.

Gringauz, K. N., and Zelikman, M. Kh., "Measurement of the Concentrations of

Positive Ions Along the Orbit of an Artificial Satellite," The Russian

Literature of Satellites, Part II, 133-147, International Physical Index, 1958.

Knudsen, W. C., "Evaluation and Demonstration of the Use of Retarding

Potential Analyzers for Measuring Several Ionospheric Quantities," J. GEOPHYS.

RES., 71, 4669-4678, 1966.

Park, C. G., D. L. Carpenter, and D. B. Wiggin, "Electron density in the

plasmasphere: Whistler data on solar cycle, annual, and diurnal variations,"

J. GEOPHYS. RES., 83, 3137, 1978.

Rich, F. J., W. J. Burke, P. J. L. Wildman, and R. C. Sagalyn, "Electron

temperature profiles measured up to 8000 km by S3-3" (abstract), EOS TRANS.

AGU, 59, 338, 1978.

Sagalyn, R. C., M. Smiddy, and J. Wisnia, "Measurement and Interpretation of

Ion Density Distributions in the Daytime F Region," J. GEOPHYS. RES., 68,

199-211, 1963.

Shawhan, S. D., and Murphy, G. B., "Plasma Diagnostics Package Assessment of

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Instrumentation

Generel Interest

Burrous, C. N., Lieber, A. J., and Zaviantseff, V. T., "Detection Efficiency

of a Continuous Channel Electron Multiplier for Positive Ions," REV. SCI.

1NSTR., 38, 1477-1481, 1967.

Chutjian, A., "Geometries and Focal Properties of Two Electron-lens Systems

Useful In Low-Energy Electron or Ion Scattering," REV. SCI. 1NSTRUM., 50,

347-355, 1979.

Conley, J. M., "Conceptual Design of IMPS Electrical Properties Degradation

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Knott, K., "Payload of the GEOS Scientific Geostationary Satellite," FSA--ASE

r

SCI. TECH. REV., 1, 173-196, 1975.

v^

yam. r 8i

ti.
Masek, "Satellite Positive Ion Beam System," AIAA International Electric

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Sellers, B., et al., "A High-Time Resolution Spectrometer for 0.05 to 500 keV

Electrons and Protons," in Spacecraft Charging by Magnetospheric Plasmas,

PKUGRESS IN ASTRONAUTICS AND AERON., 47, 31-42, 1976.

Whipple, E. C., "The Problem of Low Energy Particle Measurements in the

Magnetosphere," UCSD, San Diego, NASA Contract 8-31145, 1978b.

Effects on Space Systems

General

General References

Durrett, J. L., and Stevens, J. R., "Description of the Space Test Program

P78-2 Spacecraft and Payloads," Spacecraft Charging Technology--1978, NASA

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Ehlers, H. K. F., "Modeling Correlation with :'light Data," NASA, The Shuttle

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A147-A15b, 1983.

Garrett, H. B., and Pike, C. P., eds., "Space Systems and Their Interactions

with harth's Space Environment," Prog. Astronaut. Aeronaut., Vol. 71, AIAA

Press, New York, 1980.

Grard, R. J. L., ed., Photon and Particle Interactions With Surfaces in Space,

D. Reidel, Hingham, Mass., 1973a.

Halverson, W. D., "Modifications of Ionospheric Simulation Capability," NASA

Jt,hnson Space Center, Houston, Progress Briefing #1, 24 May, 1979.

Pike and Lovell, eds., Proceedings of the Spacecraft Charging Technology

Conference, AFGL-TR-77-0051--NASA TMX-73537, 1977.

Potter, A. E,, "Effects of Shuttle Environment on Instrument Performance,"

NASA, The Shuttle Environment Workshop, Calverton, MD, Oct. 5-7, 1982,

Contract NAS5-27362, A157-A162, 1983.

Shawhan, S. D., et al., "Subsatellite Studies of Wave, Plasma and Chemical

Injections from Spacelab," AlAA-82-0008 5, AIAA 20th Aerospace Sciences Meeting,

1982.

Singer, S. F., ed. "Interactions of Space Vehicles With an Ionized Atwosphere," Pergamon, New York, 1965.

Whipple, E. C., Jr., "The equilibrium electric potential of a body in the

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88
Wilkerson, T., "Space Shuttle: A View of What We Have Done So Far," NASA, The

` Shuttle Environment Workshop, Calverton, MD, Oct. 5-7, 1982, Contract

NAS5-273b2, A295-A299, 1983.

Theoretical Models

Parker, L. W., and Whipple, E. C., Jr., "Theory of a Satellite Electrostatic

Probe," ANNALS OF PHYS., 44, 126-161, 1967.

Observations/Data

Rosen, A., "Spacecraft Charging," TRW Systems Croup, internal document, 1973.

Spacecraft Charging

General References

Adamo, K. C., and Nanevicz, J. E., "Spacecraft Charging Studies of Voltage

Breakdown Processes on Spacecraft Thermal Control Mirrors," AGU 1975 Spring

Annual Meeting, Paper SA42, 1975.

Agamy, S. A., and Robinson, J. E., "Energy Spectra and Charge Fractions For

KeV Hydrogen and Helium Backscattered From Silicon," SURF. SCI., 90, 648-bb0,

1979.

Al'pert, J. L., Gurevich, A. V., Pitaevskii, L. P., Sp p ,:e Physics With

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Baragiola, R. A., Alonso, E. V., and Florio, A. Oliva, "Electron. Emission from

Clean Metal Surfaces Induced by Low-Energy Light Ions," PHYS. REV. B, 19,

121-129, 1979a.

Baragiola, R. A., et al., "loa-Induced Electron Emission From Clean Metals,"

SURFACE SCI., 90, 24U-255, 1979b.

Beattie and Goldstein, "Active spacecraft Potential cuntrol system for the

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Technology Conference, AFGL-TR-77-0051--NASA TMX-73537, edited by C. P. Pike

and K. R. Lovell, 143-165, 1977.

Beers, B. L., First principles numerical model of avalanche-induced arc

discharges in electron-irradiated dielectrics," NASA-C-168 (Rev. 10-75), 1979.

Bohm, D., Burhop, H. S., and Massey, H. S. W., "The Use of Probes for Plasma

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Bunney, R. E., "Review of Literature, Secondary Electron Emission," NASA CR

54366, CEK65VAo-KEB12a, 1964.

v

39

^.1 A
Cambou, R., Sagdeev, R. Z., and Zhulin, A., "Reviewing Artificial Radiation

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Cauffman, D. P., "Recommendations Concerning Spacecraft Charging in the

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Chen, F., "Electric probes," Plasma Diagnostic Techniques, edited by R. H.

huddlestone and S. L. Leonard, Academic, New York, 113-200, 1965.

Chung and Everhart, "Simple calculation of energy distribution of low-energy

secondary electrons emitted from metals under electron bombardment," J. Appl.

Phys., 45, 707, 1974.

"'hutjian, A., "Spectra and Cross Sections for Electron Autoionisation in Low

Energy Mg+ -Ne Collisions," J. PHYS. B., 1980.

DeForesr, S. E., "Spacecraft charging at smchronoLs orbit," J. GEOPHYS. RES.,

77, b51, 1972.

Dekker, A. J., "Secondary electron emission," Solid State Physics, Prentice

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Dorozhkin, A. A., et al., "Ion-Bombardment Auger Spectroscopy for Studying

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Everhart, T. E., "Simple Theory Concerning the Reflection of Electrons from

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F'euerbacher, B., Willis, R. F., and Fitton, b., "Electrostatic Charging and

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1973.

Frank, D. E., "ESD Phenomenon and Effect on Electronic Parts," Avionics

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Frederickson, A. R., "Electrostatic Charging and Discharging in Space

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Vacuun, 1983.

Freeman, J. W., Jr., Fenner, M. A., and hills, H. K., "The Electric Potential

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in Space, ed. by R. J. L. Grard, 363-368, 1973.

Garrett, H. B., "The Charging of Spacecraft Surfaces," REV, GEOPHYS., 19,

577-616, 1981.

Ghosh, S. N., and Khare, S. P., "Secondary Electron Emissions from Metal

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_ 
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Godfrey, D. J., and Woodruff, D. P., "A Simple Ion Scattering Spectrometer for

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Goldstein, R., and DeForest, S. E., "Active Control of Spacecraft Potentials

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PROGRESS IN ASTRONAUTICS AND AERON., 47, 169-181, 1976.

Grard, R. J. L., "Spacecraft Potential Control and Plasma Diagnostic Using

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Grard, R. J. L., "Effect of the Ambient Medium Upon the Electric Properties of

the Spacecraft Surface and Environment," Proceedings, Conference on Lightning

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NASA TMS-73537, edited by C. P. Pike and R. R. Lovell, 203-221, 1977a.

Grard, R. J. L., et al., "The influence of photoelectron and secondary

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by R.J.L. Grard, D. Reidel, Hingham, Mass., 163-192, 1973.

Grard, R., Knott, K., and Pedersen, A., "Spacecraft Charging Effects," SPACE

SCI. REV., 34, 289-304, 1983.

Hachenberg and Brauer, "Secondary electron emission from solids," ADV.

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Halverson, W. D., "Techniques for the Simulation of the Space Plasma

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Halverson, W. D., "Simulation of Spacecraft Charging Environments by

Monoenergetic Beams," USAF--NASA Spacecraft Charging Technology Conference

III, 1980b.

Jung, H., "The Origin of Solid Particles in lnterstellsr Space," ASTRON.

NACIT,. , 263, 426, 1937.

Kenaall, B. R. F., "Mass Analysis of Neutral Particles and Ions Released

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Surfaces," Photon and Particle Interactions With Surfaces in Space, ed. by

R. J. L. Grard, D. Reidel, Hingham, MA, 3-21, 1973.

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Mizera, P. F., et al., Spacecraft Charging Handbook, The Aerospace Corp., 1983.

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Purvis and Bartlett, "Active control of spacecraft charging," Space Systems

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Kosen, A., ed., Spacecraft Charging by Magnetospheric Plasmas, Prog.

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1963.

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s:► tellite charging," Spacecraft Charging by Magnetospheric Plasmas, Prog.

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bl-7b, 1976.

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J
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107, 6-33, 1948.

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Sternglass, E.J., "Theory of secondary electron emission by nigh speed ions,"

PHYS. REV., 108, 1-12, 1957.

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Stevens, N. J., "Space Environmental Interactions with Spacecraft Surfaces,"

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Stevens, N. J., Kamen, R. E., and Holman, A. B., "Design Guidelines for

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NASA Lewis Res. Center, Cleveland, Ohio, 1980c.

Tautz, "SSPM Experiment Simulation Study for the April 24," AFSC--AFGL, 1979.

Whipple, E. C., Jr., "Modeling of spacecraft charging," Proceedings of the

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Wilkenfeld, J., C. Mallon, J. Horne, "Conduction and Charge Storage in

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Willis, R. F., and Skinner, D. K., "Secondary Electron Emission Yield

Behaviors of Polymers," SOLID STATE COMMUN., 13, 685, 1973.

Analytic Models

Besse, R. L., and Rubin, A. L., "A Simple Analysis of Spacecraft Charging

Involving Blocked Photoelectron Currents," J. GEOPHYS. RES., 85, 2324-2328,

1980.

Garrett, H. B., and Rubin, A. G., "Spacecraft Charging at Geosynchronous

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1978.

Garrett, H. B., et al., "Prediction of spacecraft potentials at geosynchronous

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1979.

9 3
Tsipouras, P., and H. B. Garrett, "Spacecraft charging model, 2 Maxwellian

approximation," Rep. AFGL-TR-79-0153, Air Force Geophys. Lab., Bedford, Mass.,

1979.

Theoretical Models

Albers, N., "Computer Simulation of a Spherical Langmuir Probe," REP. SU-IPR

449, Stanford University, Stanford, CA, 1973.

Allen, J. E., Boyd, L. F., and Reynolds, P., "The Collection of Positive Ions

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Cauffman, D. P., "Ionization and Attraction of Neutral Molecules to a Charged

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Angeles, 1973a.

Chang, J. S., Prokopenko, S. M. L., Goddard, R., and Laframboise, J. G.,

"Prediction of Ion Drift Effects on Spacecraft Floating Potentials,"

Spacecraft Charging Technology--1978, NASA CP-2071,'AFGL-TR-79-0082, ed. by

R. C. Finke and C. P. Pike, 179-187, 1979.

Derksen, F., and Knott, K., "Effects of a Photoelectron Cloud on

Electric-Field Measurements From a Geostationary Satellite,"

ELDO-CECLES--ESRO-CERS SCIENT. AND TECH. REV., 6, 3-16, 1974.

Evdokimov, 0. B., and Tubalov, N. P., "Stratification of Space Charge in

Dielectrics lrradiated With Fast Electrons," SOV. PHYS. SOLID STATE, ENGL.

TRANSL., 15, X869-1870, 1974.

Grebowski, R., and Fischer, T., "Theoretical Density Distribution of Plasma

Around a Cylinder," PLANET. SPACE SCI., 23, 287-304, 1975.

Katz, I., Parks, D. E., Wang, S., and Wilson, A., "Dynamic Modeling of

Spacecraft in a Collisionless Plasma," Proceedings of the Spacecraft Charging

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Pike, 101-122, 1977.

Katz et al., "The capabilities of the NASA charging analyzer vrogram,"

Spacecraft Charging Technology-1978, NASA CP-2071---AFGL-TR-79-0082, edited by

R. C. Finke and C. A. Pike, 101-122, 1979.

Konemann, B., and Schroder, H., "The Irfluence of Capacitor Effects on the

Surface Potential of Satellites with Partially Insulating Surfaces in the

Solar Wind," PLANET. SPACE SCI., 22, 321-331, 1974.

Kurt, P. G., and Moroz, V. I., "The Potential of a Metal Sphere in

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(Translated in PLANET. SPACE SCI., 9, 259-268, 1962).

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Laframboise, J. G., "Theory of Spherical and Cylindrical Langmuir Probes in a

Collisionless Maxwellian Plasma at Rest," REP. 100, Univ. of Toronto Inst.

Aerosp. Stud., 'Toronto, Ont., 1966.

Laframboise, J. G., "Calculation of Secondary-Electron Escape Currents From

Inclined Spacecraft Surfaces in a Magnetic Field," Proc. USAF--NASA Spacecraft

Environmental Interactions Technology Conference, 1984.

Laframboise, J. G., and Prokopenko, S.M.L., "Numerical Simulation of

Spacecraft Charging Phenomena," USAF--NASA Spacecraft Charging Technology

Conference, 1976.

Laframboise, J. G., et al., "Results From a Two-Dimensional SpacecraftCharging Simulation, and Comparison with a Surface Photocurrent Model," in

SPACECRAFT CHARGING TECHNOLOGY 1980, NASA Conf. Pub. 2182, AFGL-'TR-81-0270,

709-716, 1981.

Laframboise, J. G., Goddard, R.,and Yrokopenko, S. M. L., "Numerical

Calculations of High-Altitude Differential Charging: Preliminary Results,"

Spacecraft Charging Technology--1978, NASA CP-2071/AFGL- TR-79-0082, ed. by

R. C. Finke and C. A. Pike, 188-196, 1979.

Laframboise, J. G., R. Goddard, M. Kamitsuma, "Multiple Floating Potentials,

Threshold-Temperature Effects, and 'Barrier' Effects in High-Voltage Charging

of Exposed Surfaces on Spacecraft," Proc. Inter. Symp. on Spacecraft Materials

in Space Environment, ESA, 1982.

Lai, S. T., Gussenhoven, M. S., and Cohen, H. A., "The Concepts of Critical

Temperature and Energy Cutoff of Ambient Electrons in High Voltage Cnarging of

Spacecraft," Proc. ESLAB Symp. on 'Spacecraft--Plasma Interactions and their

Influence on Field and Particle Measurements', ESA SP-198, 1983.

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CP-2071/AFGL-TR-79-0082, ed. by R. C. Finke and C. A. Pike, 31.1-375, 1979.

Parker, L. W., "Plasma sheath-photosheath theory for large high-voltage space

structures," Space Systems and Their Interactions With Earth's Space

Environment, Prog. Astronaut. Aeronau t-., vol. 71, edited by H. B. Garrett and

C. P. Pike, AIAA Press, New York, 477-522, 1980x.

Parker, L. W., "Effects of Plasma Sheath on Solar Power Satellite Array,"

J. ENERGY, 4, 6, 241-244, 1980b.

Parker, L. W., and Lxframboise, J. G., "Multi-electrode Plasma Probe for

Orbit-limited-current Measurements. II. Numerical Verification," PHYS. FLUIDS,

21, 4, 58b-591, 1978.

Parker, L. W., and Murphy, B. L., "Potential Buildup on an Electron-Emitting

Ionospheric Satellite," J. GEOPHYS. RES., 72, 1631, 1967.

Parker, L. W., and Whipple, E. C.. Jr., "Theory of Spacecraft Sheath

Structure, Potential, and Velocity Effects on Ion Measurements by Traps and

Mass Spectrometers," J. GEOPHYS. RES., 75, 4720, 1970.

Parks and Katz, "Charging of a large object in low polar earth orbit,"

Proceedings of the USAF--NASA Spacecraft Charging Technology Conference ?II,

1981.

Reiff, P. H., Freeman, J. W., and Cooke, D. L., "Environmental Protection of

the Solar Power Satellite," Space Systems and Their Interactions With Earth's

Space Environment, Prog. Astronaut. Aeronaut., vol. 71, ed. by H. B. Garrett

and C. P. Pike, AIAA Press, New York, 554-576, 1980.

Samir, U., and Willmore, A.P., "The Equilibrium Potential of a Spacecraft in

the Ionosphere," PLANET. -S PACE SCI., 14, 1131-1137, 1966.

Stevens, N. J., "Space environmental interactions with biased spacecraft

eurfaces," Space Systems and Their Interactions With Earth's Space

Environment, Prog. Astronaut. Aeronaut., vol. 71, edited by H. B. Garrett and

C. P. Pike, ALAA Press, New York, 455-476, 1980.

10')

♦ - 
Tsien, H. S., "Super Aerodynamics, Mechanics of Rarefied Gases," J. AERONAUT.

SCI., 13, 653-664, 1946.

Walker, E. H., "Plasma Sheath and Screening Around A Stationary Charged Spt.-re

And a Rapidly Moving Charged Body," Interactions of Space Vehicles With an

Ionized Atmospher,., ed. by S. F. Singer, Pergamon, New York, 61-162, 1965.

Walker, E. H., "Plasma Sheath and Screening of Charged Bodies," Photon and

Particle Interactions With Surfaces in Space, ed. by R. J. L. Grard,

D. Reidel, Hingham, MA, 73-89, 1973.

Whipple, E. C., Jr., "Theory of the Spherically Symmetric Photoelectron

Sheath," J. GEOPHYS. RES., 81, 601-607,.1976a.

Williamson, P. R., Denig, W., Banks, F. M., and Raitt, W. J., "Vehicle

Potential Measurements Using a Mother-Daughter Tethered Rocket" (abstract),

EOS TRANS. AGU, 61, 1068, 1981.

Zonov, Yu. V., "On the Problem of the Interaction Between a Satellite and the

Earth's Atmosphere and the Earth's Magnetic Field" (in Russian), ISKUSSTV.

SPUTNIKI ZEMLI, 3 (NASA TECH. TRANSL. TTF-37), 1959.

Observations/Data

Bernstein, W., et al., "Electron Beam Experiments: The Beam Plasma Discharge

at Low Pressures and Magnetic Field Strengths," GEOPHYS. RES. LETT., 5,

127-130, 1978.

Bernstein, W., et al., "Further Laboratory Measurements of the Beam-Plasma

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Bourdeau, R. E., Donley, J. L., Serbu, G. P., and Whipple, Jr., E. C.,

"Measurements of Sheath Currents and Equilibrium Potential on the Explorer

VIII Satellite," J. ASTRON. SCI., 8, b5-73, 1951.

Carignan, G. R., "Neutral Gas Mass Spectrometer on the IECM," NASA, The

Shuttle Environment Workshop, Contract NAS5-27362, A139-A146, 1983.

Grier, N. T., "Experimental kesults on Plasma Interactions with Large Surfaces

at High Voltages," Lewis Research Center, NASA Technical Memorandum 81423,

1980.

Hanson, W. B., McKibbin, D. D., and Sharp, G. W., "Some Ionospheric

Measurements With Satellite-Born Ion Traps," J. GEOPHYS. RES., 69, 2747-2763,

1964.

Henderson, C. L., and Samir, U., "Observations of the Disturbed Region Around

an Ionospheric Spacecraft," PLANET. SPACE SCI., 15, 1499-1513, 1967.

Hendrickson, R. A., "The Electron Echo Experiment, Observations of the Charge

Neutralization of the Rocket and Analysis of the Echo From Electrons

Artificially Injected into the Magnetosphere," TECH. REP. CP-160, Sch. of

Phys. and Astron., Univ. of Minn., M{nneapolis, 1972.

106
Hess, W. N., "Generation of an Artificial Auro-a," SCIENCE, 164, 1512, 1969.

isensee, U., "Plasma Disturbances Caused by the Helios Spacecraft in the Solar

Wind," J. GEOPHYS., 42, 581-589, 1977.

Jacobsen, T. A., and Maynard, N. C., "Evidence for Significant Spacecraft

Charging By an Electron Accelerator at Ionospheric Altitudes," PLANET. SPACE

SCI., 28, 291-307, 1980.

Jacobsen, T. A., and Maynard, N. C., "Polar 5--An Electron Accelerator

Experiment Within an Aurora. 3. Evidence for Significant Spacecraft Charging

by an Electron Accelerator at Ionospheric Altitudes," Planet Space Sci., 28,

291-307, 1980.

Kennerud, K. L., "High voltage solar array experiments," REP. NP-SA CR-121280,

Boeing Co., Seattle, Wash., March 1974.

Konradi, A., McIntyre, B., and Potter, A. E., "Expezimental Studies of Scaling

Laws for Plasma Collection at High Voltages," ,;. SPACECRAFT ROCK., 21,

287-292, 1984.

Krassovsky, V. I., "Exploratirn of the Upper Atmosphere With the Help of the

Third Soviet Sputnik," PROC. IRE, 47, 289-296, 1959.

McCoy, J. E., et al., "Current Leakage for Low Altitude Satellites," AiAA

79-0387, AIAA 17th Aerospace Sciences Meeting, 1979.

McCoy, E., Konradi, A., and Garriott, 0. K., "Current Leakage for Low

Altitud. -tellites," Space Systems and Their Interactions With Earth's Space

Environment, Prog. Astronaut. Aeronaut., vol. 71, ed. byH. B. Garrett and

C. P. Pike, AIAA Press, New York, 523-553, 1380.

Narcisi, R., Trzcinski, E., Federico, G., '..'lodyka, L., and Delorey, D., "The

Gaseous and Plasma Environment Around Space Shuttle," A Collection of

Technical Papers, AIAA Shuttle Envirorment and Operations Meeting, CP838,

183-190, 1983.

O'Niel, R. R., Bien, F., Burt, D., Sandock, J. A., and Stair, A. T.,

"Summarized Results of the Artificial Auroral Experiment PrecQde," J. GEOPHYS.

RES., 83, 3273, 1978a.

O'Niel, R. R., et al., "Excede 2 Test, an Artificial Auroral Experiment:

Ground-Based Optical Measurements," J. GEOPHYS. RES., 83, 3281, 1978b.

Reasoner, D. L., et al., "Relationship Between ATS-6 Spacecraft-Charging

Occurrences and Warm Plasma Encounters," Spacecraft Charging by Magnetospheric

Plasmas, PROGRESS IN ASTRONAUTICS AND AERON., 47, 89-101, 1976.

Sagalyn, R. C., and Burke, W. J., "INJUN 5 Observations of Vehicle Potential

Fluctuations at 2500 km," Proceedings of the Spacecraft Charging Technology

Conference, AFGL-TR-77-0051/NASA TMX-73537, ed. by C. P. Pike and R. R.

Lovell, 67-80, 1977.

107
Samir, U., "Charged Particle Distribution in the Nearest Vicinity of

Ionospheric Satellites--Comparison of the Main Results From the Ariel I,

Explorer 31, and Gemini-Agena 10 Spacecraft," Photon and Particle Interactions

With Surfaces in Space, ed. by R. J. L. Grard, D. Reidel, Hingham, MA,

193-220. 1973.

Samir, U. and Jew, H., "Comparison of Theory With Experiment for Electron

Density Distribution in the Near Wake of an Ionospheric Satellite,"

J. GEOPHYS. RES., 77, 6819-6827, 1972.

Samir, U. and Wrenn, G. L., "The Dependence of Charge and Potential

Distribution Around a Spacecraft on Ionic Composition," PLANET. SPACE SCI.,

17, 693, 1969.

Samir, U., Brace, L. H., and Brinton, H. C., "About the Influence of Electron

Temperature and Relative Ionic Composition on Ionic Depletion in the Wake of

the AE-C Satellite," GEOPHYS. RES. LETT., 6, 101-104, 1979a.

Samir, U., Gordon, R., Brau, L., and Theis, R., "The Near-Wake Structure of

the Atmosphere Explorer C (AF-C) Satellite: A Parametric Investigation,"

J. GEOPHYS. RES., 84, 513-525, 1979b.

Samir, U., Kaufman, Y. J., Brace, L. H., and Brinton, H. C., "The Dependence

of Ion Density in the wake of the AE-C Satellite on the Ratio of Body Size to

Debye Length in an (0+ )-Dominated Plasma," J. GEOPHYS. RES., 85, 1769-1772,

1980.

Shawhan, S. D,, and Murphy, G. B., "STS-3/OSS-1 Plasma Diagnostics Package

(PDP) Measurements of the Temperature, Pressure, and Plasma," NASA, The

Shuttle Environment Workshop, Contract NAS5-27362, A207-A232, 1963.

Whipple, E. C., J. M. Warnock, R. H. Winkler, "Effect of Sate111Le Potential

on Direct Ion Density Measurements Through the Plasmapause," J. GEOPHYS. RES.,

79, 179-1b5, 1974.

Williamson, P. R., Banks, P. M., and Raitt, W. J., "Vehicle Charging and

Potential on the STS-3 Mission," NASA, The Shuttle Environment Workshop,

Contract NAS5-27362, A187-A206, 1983.

Win(Kler, J. R., "A Summary of Recent Results Under the 'Echo' Program for the

Study of the Magnetosphere by Artificial Electron Beams," COSMIC PHYS. TECH.

REP. 168, Sch. of Phys. and Astron., Univ. of Minn., Minneapolis, 1976.

Winckler, J. R., "The Application of Artificial Electron Beams to

Magnetospheric Research," REV. GEOPHYS. SPACE PHYS., 18, 659-682, 1980.

Torques/Drag

General References

Hohl, F., "The Electromagnetic Torques on Spherical Earth Satellites in a

Rarefied Partially Ionized Atmosphere," Langley Research Center, NASA

TR R-231, 1966.

;' 108
Theoretical Models

Chang, H. H. C., and Smith, M. C., "On the Drag of a Spherical Satellite

Moving in A Partially Ionized Atmosphere," J. BR. INTERPLANET. SOC., 17,

199-205, 1960.

Observations/Data

Jastrow, R., and Pearse, C. A., "Atmospheric Drag on the Satellite,"

J. GEOPHYS. RES., 62, 413-423, 1957.

Radio Frequency

Observations/Data

Shawhan, S. D., and Murphy, G. B., "STS-3/OSS-1 Plasma Diagnostics Package

(PUP) Measurements of Orbiter Transmitter and Subsystem Electromagnetic

Interference," NASA, The Shuttle Environment Workshop, Contract NAS5-27362,

A233-A250, 1983.

Smiddy, M., Sullivan, W. P., Girouard, D., and Anderson, P. J., "Observation

of Electric Fields ; Electron Densities and Temperature from the Space

Shuttle," A Collection of Technical Papers, AIAA Shuttle Environment and

Operations Meeting, CP838, 118-126, 1983.

Contamination/Debris

General References

Jemiola, J. M., "Proceedings of the USAF/NASA International Spacecraft

Contamination Conference," REP. AFML-TR-78-190/NASA-CP-2039, 1978.

Jemiola, J. M., "Spacecraft Contamination: A Review," Space Systems and Their

Interactiuns With Earth's Space Environment, Prog. Astronaut. Aeronaut.,

vol. 71, ed. by H. B. Garrett and C. P. Pike, AIAA Press, New York, 680-706,

1980.

Kruger, R., Triolo, J., and McIntosh, R. , "OSS-1/Contamination Monitor," NASA,

The Shuttle Environment Workshop, Contract NAS5-27362, A163-A178, 1983.

Liemohn, H.B., et al., "Charging and Contamination During Geosynchronous Orbit

Insertion," Society of Photo-Optical Instrumentation Engineers Meeting, SHE

216-09, 1980.

Oberg, J. E., "Control of RCS Plume Effects During Proximity Operations on STS

Missions," A Collection of Technical Papers, AIAA Shuttle Environment and

Operations Meeting, CP838, 68-71, 1983.

Theoretical Models

Cauffman, D. P., "Ionization and Attraction of Neutral Molecules to a Charged

Spacecraft," AIR FORCE REPORT NO. SAMSO-TR-73-263, AEROSPACE REPORT NO.

TR-0074(9260-09)-1, 1973a.

109
Observations/Data

Barengoltz, J., Maag, C., and Kuykendall, F., "STS-3 'Snowflake' SL,:dy," NASA,

The Shuttle Environment Workshop, Contract NAS5-27362, A289-A294, 1983.

Bunner, A., Bartoe, J. D., and Triolo, J., "Test for Contamination of

MgF 2-Coated Mirrors," NASA, The Shuttle Environment Workshop, Contract

NAS5-27362, A179-A186, 1983.

Cofer 111, W. R., Bendura, R. J., Sebacher, D. I., Pellett, G. L., Gregory,

G. L., and Maddrea Jr., G. L., "Airborne Measurements of Space Shuttle Exhaust

Constituents," A Collection of Technical Papers, AIAA Shuttle Environment and

operations Meeting, CF838, 36-41, 1983.

Fraundorf, P., et al., "Erosion of Mylar and Protection by Thin Metal Films,"

A Collection of Technical Papers, AIAA Shuttle Environment and Operations

Meeting, CP838, 131--137, 1983.

Leger, L. J., "Oxygen Atom Reaction with Shuttle Materials at Orbital

Altitudes," NASA, The Shuttle Environment Workshop, Contract N:.S5-27362, 1983.

Leger, L. J., Spiker, I. K., Kuminecz, J. F., Ballentine, T. J., and

Visentine, J. T., "STS Flight 5 LEO Effects Experiment--Background Description

and Thin Film Results," AIAA Shuttle Environment and Operations Meeting,

AIAA-83-2631-CP, 1983.

Miller, E. R., "Induced Environment Contamination Monitor Ascent/Entry,

Optical and Deposition Measurements," NASA, The Shuttle Environment Workshop,

Contract NAS5-27362, Al23-A138, 1983.

Murphy, G. B., Shawhan, S. D., and Pickett, J. S., "Pertut)ations to the

Plasma Environment Induced by the Orbiter's Maneuvering Thrusters," A

Collection of Technical Papers, AIAA Shuttle Environment and Operations

Meeting, CP838, 59-65, 1983.

Weinberg, J. L., Giovane, F., Schuerman, D. W., and Hahn, R. C., "OSS-1/STS-3

Shuttle Induced Atmosphere Experiment," NASA, The Shuttle Environment

Workshop, Contract NAS5-27362, A251-A266, 1983.

IMPS

ESWG Minutes

Reports

Johnston, A. R., and Bergman, L. A., "Fiber Optic Experiment for the Shuttle

Long-Duration Exposure Facility," SPIE Vol. 296, Fiber Optics in Adverse

Environments, 125-133, 1981.

110
APPENDIX A

I CONTRIBUTORS

Dr. Hugh R. Anderson

SAI

13400B Northrup Wy #36

Bellevue, WA 98005

Lt. Col. Rick Babcock

AFGL-PH

Hanscom AFB

Bedford, MA 01731

Dr. Jack Barengoltz

JPL (MS 157-507)

4800 Oak Grove Dr.

Pasadena, CA 91109

Dr. Joe Binsack

M1T Center for Space Research

Bldg. 37-241

Cambridge, MA 02139

Mr. Robert M. Broussard

The Aerospace Corp.

P.O. Box 92957

Los Angeles, CA 90009

Dr. Rebecca Chaky

TRW 135-3367

1 Space Park

Redondo Beach, CA 90278

Lt. Ray Cull

AFWAL-MLBT

Wright Patterson AFB

Dayton, OH 45433

Dr. William E. Denig

AFGL-PHK

Hanscom AFB

Beford, MA 01731

Mr. Robert Ebbett

JPL (MS 233-307)

4800 Oak Grove Dr.

Pasadena, CA 91109

111
Dr. Jeffery M. Forbes

Dept. of Elec., Comp., and Sys. Eng.

Boston ITniversity

110 Cummingron St.

Boston, MA 02215

Capt. David Founds

AFWL-ARTC

Kirtland AFB

Albuquerque, NM 37117

Mr. A.R. Frederickson

ESD-ESR-RADC

Hanscom AFB

Bedford, MA 01731

Mr. Robert E. Freeland

JPL (MS 157-410)

4800 Oak Grove Dr.

Pasadena, CA 91109

Dr. Henry B. Garrett

JPL (MS 144-218)

4800 Oak Grove Dr.

Pasadena, CA 91109

Dr. Donald A. Guidice

AFGL-PHK

Hanscom AFB

Bedford, MA 01731

Mr. William N. Hall

AFGL-PHK

Hanscom AFB

Bedford, MA 01731

Dr. David Hall

Aerospace (M2-271)

P.O. Box 92957

Los Angeles, CA 90009

Dr. Moshe Harel

JPL (MS 144-218)

4800 Oak Grove Dr.

Pasadena, CA 91109

Dr. Michael Heinemann

AFGL-PHK

Hanscom AFB

Bedford, MA 01731

112

...... 
Capt. Terry Hinnerichs

AFWL-ARTC

Kirtland AFB

Albuquerque, NM 87111

Mr. Curtis Horton

JPL (MS 144-218)

4800 Oak Grove Dr.

Pasadena, CA 91109

Di. George Inouye

TRW, STG M2-1145

1 Space Park

Redondo Beach, CA 90278

Dr. I. Katz

Systems: Science and Software

P.O. Box 1620

La Jolla, CA 92038

Mr. Clyde King

JPL (MS 158-224)

4800 Oak Grove Dr.

Pasadena, CA 91101

Mr. Gene K. Komatsu 

TRW 135-3972

1 Space Park

Redondo Beach, CA 90278

Ms. Cindy Lebel

JPL (MS 168-222)

4800 Oak Grove Dr.

Pasadena, CA 91109

Dr. Carl R. Maag

JPL (MS 157-507)

48UO Oak Grove Dr.

Pasadena, CA 91109

Lt. Patricia Morse

AFWL-ARBE

Kirtland AFB

Albuquerque, NM 87117

Mr. Rory Ninneman

AFWL-ARBC

Kirtland AFB

Albuquerque, NM 87117

113

- A
Charles P. Pike

. , L-PHK

i Hanscom AFB

Bedford, MA 01731

Dr. Carolyn K. Purvis

'ewis National Res Cntr (MS77-4)

21000 Brookpark Road

Cleveland, OH 44135

Dr. D.L. Reasoner

NASA Marshall Space Flt Cntr

Code ES 53

Huntsville, AL 35812

Capt. Dana Rush

AFGL-PHK

Hanscom AFB

Bedford, MA 01731

Mrs. Rita C. Sagalyn

AFGL-PH

Hanscom AFB

Bedford, MA 01731

Mr. Frank Schutz

JPL (MS 180-600)

48UO Oak Grove Dr.

Pasadena, CA 91109

Mr. Ron Sessions

AFWL-ARBC

Kirtland AFB

Albuquerque, NM 87117

Dr. Stanley D. Shawhan

Solar lerr. and Astrophysics Div.

NASA Hdqts. (CODE EE-8)

Washington, D.C. 20546

Dr. Jim Slavin

JPL (MS 169-506)

4800 Oak Grove Dr.

Pasadena, CA 91109

Dr. M. Smiddy

AFGL-PHG

Hanscom AFB

Bedford, MA 01731

114

E

/E f

a
Mr. N. John Stevens

Space and Comm. Grp., Hughes

P.U. Box 92919

Los Angeles, CA 50009

Mr. Edward W. Taylor

AFWL-NTCO

Kirtland AFB

Albuquerque, NM 87117

Mr. Joseph F. Wise

AFWAL-APL(POOC)

Wright Patterson AFb

Dayton, OH 45433

Dr. Michael J. Wiskerchen

Solar Terr. and Astrophysics Div. (EE)

NASA Hdqts.

Washington, D.C. 20546

Maj. S. Peter Worde n

OUSDRE-ADEW

Rm. 3E-1034, Pentagon

Washington, D.C. 20301

115

a
t

APPENDIX B

ESWG MEMBERS

Dr. H.R. Anderson

Charged particle beams, wake charging, SEUs

206-747-7152

Lt. R. Cull

' Cryogenic systems, contamination, surface erosion

513-255-7377

Dr. J.M. Forbes

Ionospheric physics, glow

617-353-4782

Mr. A.R. Frederickson

Material properties, space-based radar, dielectric charging

617-861-4033

Mr. R.E. Freeland

Large structures

818-354-3540

Mr. H.B. Garrett

Space environment, spacecraft charging

818-354-2644

Capt. T. Hinnerichs

Space-based weapons requirements

505-844-2986

Dr. G. Inouye

Arcing and SC design, high-voltage arrays, EMI

213-535-8448

Dr. I. Katz

Interaction modeling, wake charging

619-453-0060

Dr. D.L. Reasoner

Magnetospheric physics

205-453-3037

Dr. S.D. Shawhan

Engineering measurements - EMI, radio frequency transmissions,

induced environment

202-453-1676

Dr. M. Smiddy

Magnetospheric physics, induced environment

617-861-2431

PKW,E01N(; PAU BLANK NOT F-I

117

^^"TENTIONALIY RJ eur
Mr. J.F. Wise

High-voltage array3

513-255-6235

Maj. S.P. Worden

Natural environment, space-based weapons requirements

202-695-02U8

118

I^

7
GLOSSARY

AF Air Force

AFGL Air Force Geophysics Laboratory

AIAA American Institute of Aeronautics and Astronautics

AFSTC Air Force Space Technology Center

AEC Atomic Energy Commission

CEM Contamination affects Monitor

CRAW Coordinated Data Analysis Workshops

DME Dielectric charging, material proper effects, and electrostatic

discharge

DMSF Defense Meterological Satellite Project

EIM Environment and interactions Monitor

EIS Electron-ion beam-induced interactions

EMi Electromagnetic interference

EPD Electrical properties degradation

ESD Electrostatic discharge

ESWG Engineering/Science Working Group

EUV Extreme ultraviolet

EVA Extra-vehicular activity

GLW Glow measurements

HVA High-voltage array

IMPS interactions Measurements Payload for Shuttle

1R Infrared

ITO Indium tin oxide

JPL Jet Propulsion Laboratory

LIDAR Laser detection and ranging system

119
LSS Large space structure

MIL-STDS Military Standards

MMU Manned Maneuvering Unit

MSSTP Military Space System Technology Plan

NASA National Aeronautics and Space Administration

PLY Plasma Diagnostic Package

PEU Polar Earth Orbit

QCM Quartz crystal microbalance

RADC/0C Rome Air Development Center, Section OC

RTF RF transmission distortion

SBR Space-based radar

SCATHA The Spacecraft Charging at High Altitudes Program

SDD Space debris detection

SEU Single-event upset

TEUM Tapered element oscillating microbalances

TQCM Thermally controlled quartz crystal microbalances

TVM Low-light TV monitor

WCE Wake charging effects

12U

a

l I I •.