Author(s)
Werner, Paul W.
Title
Ion gun operations at high altitudes
Publisher
Monterey, California : Naval Postgraduate School
Issue Date
1988
URL
http://hdl.handle.net/10945/23060
This document was downloaded on May 04, 2015 at 22:58:19
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KNOX LIBRARY
VAL POSTOFuADTJATE SCHOOL
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MOJJTEBEY, CALIFORMIA 93943-6002
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,
California
THESIS
ION GUN OPERATIONS AT HIGH ALTITUDES
by
Paul
W.
Werner
June 19 88
Thesis Advisor:
Richard
C.
Olsen
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ION GUN OPERATIONS AT HIGH ALTITUDES
PERSONAL
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i
WERNER, PAUL W
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SPACECRAFT CHARGING, SCATHA
ION GUN,
(SCATHA) SATELLITE AS PART OF A
JOINT AIR FORCE /NASA PROGRAM ON SPACECRAFT CHARGING AT HIGH ALTITUDES. EXPERIMENTS WITH THE
SCATHA ION a'N WERE MONITORED BY CHARGED PARTICLE DETECTORS AND THE ELECTRIC FIELD EXPERIMENT. IT WAS FOUND THAT THE ELECTRIC FIELD EXPERIMENT COULD BE USED TO MEASURE SATELLITE
POTENTIAL DURING ION BEAM EMISSION IN SUNLIGHT AND ECLIPSE. UNNEUTRALI ZED TON BEAM EMISSION
IN HIGH ENERGY M- 7 KpV) AND HIGH CURRENT ( 1-? mA) MODES RESULTED IN THE SATELLITE MOMENTARILY CHARGING TO A NEGATIVE POTENTIAL NEAR THE MAGNITUDE OF THE BEAM VOLTAGE AND THEN
RISING TO SOME LESS NEGATIVE VALUE, TYPICALLY -500 TQ -800 V. THE NET EMITTED CURRENT WAS
APPARENTLY LIMITED BY the FORMATION OF A VIRTUAL ANODE. LOW CURRENT (20-30 ^A), HIGH VOLTAGE
(1 kV) RESITTED IN - 10 TO
- 50 V SATELLITE POTENTIALS.
TRICKLE MODE ( 20-80 ^A, NO ACCEL
VOLTAGE) OPERATIONS RESULTED IN SATELLITE POTENTIALS NEAR ZERO VOLTS. IN SUNLIGHT THE SPACE
CRAFT POTENTIAL EXHIBITED A SPIN MODULATION, ATTRIBUTED TO VARIATIONS IN THE NEUTRALIZATION
OF THE BEAM AS IT PASSED THROUGH THE PHOTOELECTRON CLOUD. RAPID FLUCTUATIONS OF THE SPACE
CRAFT POTENTIAL OCCURED WHICH MAY BE EXPLAINED QUALITATIVELY BY SPACE CHARGE INSTABTLITES
Distribution /AVAILABILITY of abstract
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Approved for public release; distribution is unlimited.
Ion Gun Operations at High Altitudes
by
Pau W We r ner
Lieutenant, United States Navy
B. S., University of Utah,
1960
I
.
Submitted in partial fulfillment of the
requirements for the degree of
MASTER OF SCIENCE
f
IN
PHYSICS
rom the
NAVAL POSTGRADUATE SCHOOL
June 1968
,)
ABSTRACT
Experiments in charge control were conducted on the P782
satellite
(SCATHA)
program
Experiments
charged
with
the
particle
experiment.
It
joint
as part of a
spacecraft
on
charging
SCATHA ion
detectors
altitudes.
monitored
were
gun
and
Force/NASA
Air
high
at
electric
the
by
field
was found that the electric field experiment
could be used to measure satellite potential during ion beam
emission
in sunlight and eclipse.
Unneutra ized
beam
ion
1
emission in high energy (1-2 KeV) and high current (1-2
resulted in the satellite momentarily charging
modes
negative
potential near the magnitude of the
the
high
a
voltage
beam
typically -500
and then rising to some less negative value,
to -800 V.
mA
to
The net emitted current was apparently limited by
formation of a virtual anode.
voltage
potentials.
(1
kV)
Trickle
Low current (20-30
resulted in -10 to
mode
(20-80
>iA,
-50
no
acce
uA )
satellite
V
I
voltage)
operations resulted in satellite potentials near zero Volts,
In
sunlight
modulation,
of
the
spacecraft potential
exhibited
the beam as it passed through the
a
spin
neutralization
attributed to variations in the
photoe ectron
1
cloud.
Rapid fluctuations of the spacecraft potential occured which
may
instabi
explained
be
1
i
t
ies.
qualitatively
by
space-charge
TABLE OF CONTENTS
I.
II.
III.
AND BACKGROUND
INTRODUCTION
1
A.
INTRODUCTION
1
B.
PROBLEM OF SATELLITE CHARGING
2
C.
HISTORY OF ACTIVE EXPERIMENTS
3
D.
THEORY
9
THE SCATHA PROGRAM
13
A.
SATELLITE
13
B.
UCSD PLASMA DETECTORS
15
C.
ELECTRIC FIELD MONITOR
17
D.
ION GUN
19
-
25
Low Energy, Low Current Mode
- 32
OBSERVATIONS
A.
DAY 200
1.
2.
B.
a.
Observations
b.
Analysis
High Voltage, High Current Mode
DAY 47
1.
2.
C.
26
32
32
-
40
50
High Voltage, High Current Mode
50
a.
Observations
50
b.
Analysis
54
High Voltage, Low Current Mode
DAY 293
55
56
.
':.-
D.
IV.
V.
DAY 295
1.
Observations
2.
Analysis
DISCUSSION
CONCLUSIONS
APPENDIX
r^
J 43-5003
69
69
71
75
77
80
LIST OF REFERENCES
83
INITIAL DISTRIBUTION LIST
86
.
ACKNOWLEDGMENT
author expresses his appreciation to Professor
The
C.
R.
whose unwavering patience and scholarly advice led
Olsen,
him through the obstacles of research.
Dr.
Olsen's work on
this project was supported by the Naval Postgraduate
Foundation
research
program
and
by
NASA/Lewis
School
Research
Center
The author also wishes to thank Dr.
Dr.
of
Dr.
E.
the data;
T.
S.
E,
Mcllwain and
Whipple (UCSD) for providing support in analysis
C.
L.
Ms.
W.
Aggson
experiment; and Mr.
H.
W.
Li
for help in processing the data;
(NASA/GSFC),
the
Cohen,
for the SCA experiment.
the PI
PI
for
the
SCIO
.
I
A.
INTRODUCTION AND BACKGROUND
.
INTRODUCTION
This
thesis is concerned with the equilibrium and
potentials of a spacecraft in
dependent
the
and the effect of active ion particle emission on
understanding
of
spacecraft charging is
time
magne tosphere
them.
An
because
important
charging can directly affect satellite survival. Charging is
also
great
of
concern to space
scientists
studying
the
satellite environment primarily because of its effect on the
measurement
of ambient particles.
This thesis is based
collected by the UCSD particle detectors and the
data
emission
was
accomplished
by
the
GSFC
The active
electric field monitors on the SCATHA satellite.
ion
on
AFGL
ion
gun
exper iment
The
problem of spacecraft charging has been an
concern since the early days of satellite flights
1965).
With
the advent of geosynchronous
ongoing
(Whipple,
spacecraft,
the
observations showed the existence of even higher potentials,
up to kilovolts in eclipse (DeForest,
this
which
size
may
(McPherson
can produce
damaging
1972).
electrostatic
result in spacecraft anomalies
et
al.,
1975).
Potentials of
The
or
remainder
discharges
malfunctions
of
this
introductory chapter presents the foundations and principles
of charging and active spacecraft potential
B.
control.
PROBLEM OF SATELLITE CHARGING
charging
Spacecraft
to high
levels
first reported by DeForest
orbit
was
1972).
Charging
effects were identified
geosynchronous
at
(DeForest,
1972
in
characteristic
by
signatures in the charged particle data. Accelerated ambient
ion fluxes result in a peak in the low-energy ion spectra at
the
spacecraft
several
hundred
(Gussenhoven
Charging
levels
1986).
1983;
Olsen
and high densities.
to
sunlight
in
1983;
in general,
been a problem at low altitudes because of
temperatures
of
Purvis,
and
Satellite charging,
tens
levels
to
volts negative were observed
and Mullen,
Mullen et al.,
not
potential.
negative were observed in eclipse and
kilovolts
low
Exceptions to this
have recently been found in DMSP data (Yeh and
has
plasma
trend
Gussenhoven,
1987).
In
addition to charging relative to the environment,
the
spacecraft may also differentially charge due to the various
conducting
and insulating characteristics of the
materials of which the spacecraft
charging
satellite
is
composed.
different
Differential
sets up a "barrier" that traps electrons near
and
inhibits current flow.
This may
cause
the
the
satellite mainframe to charge more negatively than otherwise
possible due to the environment alone.
Large spacecraft potentials give rise to the possibility
electrostatic
of
discharges
damage
could
that
the
spacecraft. Anomalies and malfunctions and negative charging
events
local
strongly concentrated in the midnight
are
time sectors,
anomalies
These
failures
that
as shown in Figure
dawn
to
(McPherson,
1
1975).
were generally logic upsets
or
might
electrostatic
be
attributed
to
switching
discharges.
Because
accurate
analysis of data from a satellite
knowledge
existence
of
incentive
control
C.
be.
strongly
satellite
requires
potential
affects
survival,
for
design
the
the
charged
there
is
methods
subjects
These
an
and
also
of
have
high-voltage
of
large-scale structures which are planned
spacecraft and the
to
and
potentials.
implications
important
satellite
to understand the causes and
those
of
the
large potential
a
measurements
particle
strong
of
built in space in the near future.
HISTORY OF ACTIVE EXPERIMENTS
The
positive
rocket
test
of
a
ions was SERT
mercury ion
artificially
involving
American experiment
first
generated
I
in
thruster
This
1964.
demonstrated thrust and beam neutralization.
A
bridge".
is
In
often
called a "plasma
November
contactor"
1969 and August
1970,
a
plasma source
potential
which is used primarily as a means for spacecraft
control
was
successfully
and
or "plasma
the
Soviets
V
O
Figure
DSP LOGIC UPSETi
OiCi II RCA UPSETS
INTELSAT
IV
O INTELSAT
III
1.
Occurrences of satellite anoaalies plotted as a
function of local time in geostationary orbit
The radial distance is not significant.
launched
rockets containing ion engines
two
engine
study
conditions.
The
ionosphere
were
part
efficiency
characteristics situated at
(Gavrilov et al.,
was
the
neutralizers
of
different
distances
1973)
launched on February
II,
surface
studied
of
3,
It
demonstrated
varying the potential of the neutralizer with
that
the
spacecraft
caused
variations
first
was the
1970,
successful orbital flight of an ion engine.
to
to
automatic
of
used
and
Also
order
near-natural
under
laboratories
neutralizer
from the beam.
SERT
engines
ion
cesium on tungsten.
of
on
different
operations
space-flight
ionization
effect
systen
in
the
in
respect
beam
to
spacecraft potential, and therefore the spacecraft potential
respect
with
the
to
distant
bridge
Plasma
plasma.
neutralizers provided an electron current to neutralize
beam.
(Jones et al.,
ATS-4,
-5,
the
1970)
and -6 each carried a cesium ion
thruster.
ATS-4 and -5 both suffered malfunctions and complete testing
of
the ion engines was not possible.
1968,
low
altitude elliptical orbit.
that
the
ATS-5
spin
rate
launched August
suffered a launch vehicle failure and entered
10,
Bartlett,
ATS-4,
ion
engine was
Limited
working
testing
properly.
a
indicated
(Hunter
and
1969)
spun up in the wrong direction and at too high
(100
rpm)
when ejected into
its
final
a
orbit.
seriously
limiting
operation.
thruster
limited
The
operations with the ion engine were successful and indicated
no change in the spacecraft potential,
energy
limit
of
the
UCSD
particle
within the 50 eV low
Electron
detectors.
experiments with the hot filament
emission
neutralizer
reduced large negative potentials to near zero
eclipse
differential charging of the insulating surfaces
in
but
ultimately
suppressed electron emission. This reduced the effectiveness
of
the emitter in controlling the spacecraft potential.
6
ion engine operation clamped the spacecraft potential
-5
suppressed
ATSto
differential
charging.
Operation of the plasma bridge neutralizer alone
maintained
about
the
volts
and
potential to within a few volts of the
ambient
plasma
potential for all plasma conditions observed. ATS-6 provided
that a low energy plasma discharge can provide
evidence
effective
(Olsen.
The
gun
means
for
controlling
an
potential.
1985)
West German Porcupine rocket contained a Xenon
which directed a 200 eV
the Earth's magnetic field.
and
spacecraft
associated
Xe-*-
ion beam
processes
wave
perpendicular
The dynamics of the plasma
were
studied
via
ion
to
jet
an
instrumented central pay oad and on subpayloads ejected from
1
it.
(Kintner and Kelley,
Most
ions
of
1981)
the space experiments performed with
so far have used plasma beams rather than
energetic
non-neutral
.
beams.
ion
plasma
The
Table
1
(Banks et al.,
gives a
list
of
beam space experiments since the beginning of
1975.
ATS satellites and SCATHA are not listed because
their
plasma
sources were considered too small to be
ionospheric modification.
used
table
the
1987)
rockets.
experiment,
'Meteor'
useful
for
Most of the experiments listed in
The exceptions
are
Soviet
the
performed from a satellite,
and
the
Japanese SEPAC experiment, part of the Spacelab-1 pay oad on
1
the NASA Space Shuttle.
pos i
t
i
The energies quoted are those of the
ve ions
The Araks pay oad had an electron gun as the main active
1
device
of
with the plasma source acting primarily as a
source
neutralizing ions rather than a source of energetic ions.
The
plasma
ejected
was
perpendicular
to the
local
near the payload.
remain
readily
ionosphere.
and
more
device.
In
from
the
ambient
the
the
The plasma sources on ATS-6
and
the
in
in
table are
similar
to
that the plasma source was the
most
cases the plasma source was
the
source
active and
with
it
remaining
Porcupine
main
active
operated
pulses and not continuously in order to observe the
with
more
This allowed the payload to
SCATHA were quite successful in this respect. The
experiment
would
The data are not clear on how successfully
was controlled.
experiments
less
or
magnetic field so that it
neutralizing electrons
collect
potential
continuously
switched
in
effects
off.
The
of most of
these experiments
scientific
objective
study
propagation of plasma beams
through
was
to
the
space
plasma in the presence of the Earth's magnetic field.
(Banks
et al
the
.
,
1987).
TABLE
I
SPACE EXPERIMENTS WITH POSITIVE ION BEAMS, SINCE 1-1-75
A t
I
(1)
(2)
1-26-75
2-15-75
(1)10-29-77
(2)10-30-77
(3)11-30-78
(4)11-18-79
(1)
(2)
i
Energy
tude
190 km
185 km
115-160
115-160
110-145
110-145
km
km
km
km
Caesum
50 eV
Barium unknown
unknown
unknown
Barium
METEOR
1977-79
850-900 km
Xenon
AELITA
10-06-78
(2)10-25-79
100-145 km
Lithium
3-19-79
3-31-79
196-464 km
191-451 km
(1) 1-27-80
(2)11-14-82
120-220 km
120-451 km
Argon
25 eV
33 eV
Porcupine
(1
)
(1)
(2)
Current
unknown
130 eV unknown
4-10 eV
300 A
O.
1
A
SEPAC
11-28-83
240 km
Argon
110 eV
1. 6
ARCS 3
2-10-85
129-406 km
Argon
ZOO eV
0.2 A
kA
D.
THEORY
in
a plasma.
A
spacecraft in the magnetosphere
theory
are
of
anisotropic
the
photoemi ss ion,
is
effectively a probe
Elements which must be added to
current
basic
probe
introduced
insulating surfaces to the probe so that it is no
an equi potent ia
longer
surface.
1
equilibrium,
In
by
secondary electron emission, and the addition
magnitude
the
potential
the
of
is
defined by the requirement that the currents to and from the
satellite
current,
sum
and
to zero.
A
non-zero sum
net
is a
spacecraft will charge till
the
it
charging
reaches
equilibrium. A spacecraft will discharge and charge with two
capacitive
time
modulation
of
constants.
the differential
other
One
is
the mainframe or ground
potential
the
for
potential
and
the
surfaces.
charging of the insulated
The major terms in the current balance equation are: ambient
electrons, ambient ions,
impact,
The
photoe ectrons.
1
secondary electrons due to electron
electrons
secondary
due
environment
to
impact,
ion
surface
and the
of
satellite are therefore important because of the effects
ambient electrons and ions,
electrons.
At
pho toe ectrons
1
geosynchronous
electron current,
that is,
altitudes,
,
and
the
and
the
of
secondary
high-energy
the current from that portion of
the electron population with energies greater than about
30
KeV,
in both sunlight
drives the spacecraft frame potential
and eclipse (Gussenhoven and Mullen,
Photocurrent
1983).
spacecraft,
the major current from the
is
the exact level of which depends on the work function of the
irradiated material.
one to two orders of magnitude larger than the
^A/m=,
most
the ambient electron
important term,
emission
electrons
incident
per
electron)
an insulated surface
is
various
surfaces.
For
rarely equi potent ia
determines
important
if
insulated
an
since
this
because
1
regimes.
As
change.
The
shadow,
in
electrons
the
is
which
term
a satellite
plasmasphere or plasma sheet,
boundaries
is
its
the surface will charge negatively.
At or near geosynchronous orbit,
the
surface
primary
emission due to
electron
particularly
(
for
spacecraft
A
the different secondary emission characteristics of
secondary
in
unity
exceed
can
with
emitted
The yield (the number of
primary particle energies of a few hundred eV.
with
next
Secondary-
flux.
caused by the impact of primary particles
is
sufficiently high energy.
of
10-100
Typical values in sunlight are
the
move
level
and
solar
a cold
of
usually
is
different
very
activity
the characteristics
plasmasphere,
10-100 cm~^) plasma,
of
two
varies
(several eV)
the
regions
the
and
dense
extends outward from its source,
the
ionosphere, and corotates with the Earth. The outer boundary
is
not well defined and varies with local time and
magnetic
activity.
spacecraft
The
is
most likely
be
to
the
in
plasmasphere between the hours of lOOO LT and 2000
The
LT.
plasmasphere may move completely inside geosynchronous orbit
some days following
on
high magnetic
activity.
following prolonged magnetic quiet,
days
the
other
On
plasmasphere
may expand to envelope most of the geosynchronous orbit. The
equatorial
average position of the plasmapause shown in the
plane
(Chappel
1
,
1974)
plasma
The
dominated
magne tosphere
the
of
is
shown
Figure
in
sheet begins where corotation
effects
by the cross-tail magnetospher ic electric
and
cni~='
typically
both
ions
and
inside
the
periods
The satellite
the plasmasheet between 2100 LT
of
electrons
Ambient currents are high in
the vehicle is in the plasma sheet.
normally
LT.
exceed 10 KeV.
for
sheet and spacecraft charging is limited to
plasma
when
temperatures
are
field.
Densities in the plasma sheet are typically on the order
1
2
.
and
is
0900
Figure
2.
The average position of the plasmapause shown
the equatorial plane (Chappell, 1979).
I
A.
I
.
THE SCATHA PROGRAM
SATELLITE DESCRIPTION
The Air Force P78-2 SCATHA (Spacecraft Charging AT
Altitude) satellite,
designed
charging.
part of a joint program with NASA, was
study the causes and
to
was
It
dynamics
launched January 30,
1979,
satellite
length
booms.
cylindrical in shape,
is
and
1
is
It
.75 m diameter,
spacecraft
of
and reached its
near geosynchronous orbit on February
final
High
The
1979.
2,
approximately 1.7 m
and has
seven
in
experimental
spin stabilized and rotates with a period of 59
seconds about an axis which lies in the orbital plane and is
nominally perpendicular to the earth-sun line.
elliptical,
period
and
of
the
satellite
with perigee of 5.3 Re,
23.5 hours.
apogee of 7.6 Re,
The orbital plane was inclined
satellite drifts eastward about 5.1=
is
The orbit is
day.
The
eclipsed for about one hour periods every
day
for 40 days around the spring and fall
a
equinoxes.
Many of the instruments are contained around the
of
the
and
7.8*^
cylinder in an area called the
belly
middle
which
band,
contains both insulating and conducting materials. The "top"
of
the satellite is a conducting surface and the "bottom" is
composed
of
insulators.
Solar cells make up most
of
the
remaining parts of the cylinder. Figure 3 shows the location
UCSO PLASMA DETECTOR
FORWARD END
Figure
3.
The P78-2 SCATHA satellite.
.
of
several of the experiments on the satellite.
The
experiments used for this work are the ion gun,
detectors, and the electric field monitor.
B.
the
(Fennel,
P78-2
plasma
1982)
UCSD PLASMA DETECTORS
University
The
experiment
particle
analyzers,
California
at
San
(
SC9 ) consist
of
five
Diego
ESA's are mounted
on
charged
electrostatic
Pairs
three for ions and two for electrons.
electron
and
ion
of
top
the
of
of
the
satellite in rotating heads which sweep in orthogonal planes
over
220"
a
range
shown in Figure
4.
radially outward.
NS
spin
rotating
axis
and
radially
angles
geometry and rotation
are
The third ion detector is mounted to look
One of the rotating assemblies (the HI or
has an energy range of
detector)
other
through the
detector
The
outwards.
1
assembly (the LO or EW
eV to
81
detector)
KeV.
The
and
the
single ion detector (FIX detector) have an energy range of
eV
to
1800
eV.
Energy resolution is
channels overlapping in coverage.
is 5 X 7 degrees.
sampled
the spin ax
i
1
adjacent
The angular field of view
The outputs of the rotating detectors are
is
sampled
Pitch angle distributions are obtained
by parking the detectors so that they
has
with
4 times a second and the fixed detector
at 8 times a second.
Using
20%
look perpendicular
to
s
particle detectors to determine vehicle
drawbacks.
They
are not positioned on
the
charging
satellite
K««
LO DETECTOR
W WnriM
HflCTM
-20^
200*
LO "DETECTOR
200 HI DETECTOR
Figure
4.
Charged particle detector
(
SC9
)
geometry.
.
with detection of charging as their prime objective. Energy
widths
channel
optimal.
and sampling intervals are
aligned
field
and if the detector is not looking
right direction or energy range,
be observed.
C.
therefore
not
the ambient ion flux is sometimes magnetic
Also,
in
the
a charging peak will
not
1982; SCATHA Handbook,
(Fennel,
1980)
ELECTRIC FIELD MONITOR
NASA
The
Goddard Space Flight
experiment
monitor
electric
Center
(SCIO) experiment
consists
antennas (cylindrical double floating probes) 100 m
tip in
uses
field
dipole
of
tip-to-
length and projecting into the spin plane. One of the
the
of
charging
experiment
relative
potential
was
to
measure
sensor's
the
to
events (Aggson et al.,
the
1983).
The
satellite
spacecraft
during
antennae
are
beryllium copper tubbing insulated except for the last 20
with a kapton coating,
m
providing an 80 m baseline. The booms
are electrically isolated from the satellite body.
A
schematic
of
the
two
experiment is shown in Figure
is
5.
different
modes
the common mode,
which
antennas
The data used in this paper is from
measures the voltage of one antenna
relative to the spacecraft ground with a time resolution
O. 5
seconds
this
The ambient electric field
obtained from differential signals between the
(the differential mode).
of
of
—
J
1
/'
^
1
/
/
insulated
1
1
1
\
—
50 meters
—
"'~-
,
1
/
\
/
+
\
\
y
—<^/ differ en
tiaJ
common mode
Figure
5.
Schematic showing SCIO coiaiDon
mode outputs.
and
output
outpjt
differential
been shown that the
has
It
experiment
SCIO
measures
spacecraft potential very accurately in sunlight (Mullen
al.,
1986).
This
because
is
et
which
beryllium,
copper
constitutes the outer 20 meters of the boom, has such a high
secondary
coefficient
electron
abundance
(meaning
secondary emission) that significant negative charging
not
occur in most plasma environments (Lai et
Hence,
antennae
the
a
SCIO
the
of
data
shows
that
reasonable
vehicle
the
During
changes very rapidly with sun angle.
potential
1986).
al.,
to
to be at independent floating potentials.
approximation,
Graphs
can be considered,
of
does
the
periods when the antenna points directly at the sun or is in
the
shadow
somewhat
the satellite,
of
due
measurement
to
the
absence
actually
then
the
antenna
of
charges
also
the
potential
difference between the charged boom and the charged
and
(Lai
not between the charged vehicle and the ambient
et al.,
charged
the
1986).
during
important
The charging of the antennae
periods in which the
to high voltages.
spacecraft
Part of this paper will
ejection,
interaction of ion beam
The
pho toemi ss ion.
represents
vehicle
plasma
becomes
is
not
address
photoemi ss i on,
and
satellite charging.
D.
ION GUN
The
Satellite Pos i t i ve- on-Beam System (SPIBS)
I
is
the
AFGL ion gun experiment (SC4-2) developed by Hughes Research
y
Laboratories.
It
is
located on the bottom of the
satellite
beams of positive
and is used to emit electrons,
ions,
or
The SPIBS instrument was designed to
beams containing both.
create several different spacecraft ground- to-ambi ent plasma
techniques
potential differences and to investigate various
maintaining spacecraft ground at or near the
for
of
potential
the ambient plasma.
The
system consist of a
basic
power
electrostatically
accelerated
discharge
to
operated at the beam potential to
is
discharge
the
plasma
Electrons
electrons.
collisions
by
restricts
emitting
electron
collisions.
filament)
accelerating
all
to
to control
plasma.
flow
atoms
in
and
The
An axial magnetic
radially
(
downstream
neutralizer can
or a fraction of
increases
and
thermionical
of
Figure
be
satellite potential relative
6
the
the beam and can be
The neutralizer and the ion source can be
independently.
the
conventional
a
neutralizer
A
located
is
structure.
neutralize
+_1000 V
allow
between
plasma by means of the discharge voltage.
electron-atom
The
The electrons are then accelerated into the
hollow cathode.
field
ion
Ions are formed
generated using
are
an
velocity.
high
a
to exit at near ground potential.
ions
supply,
Positive Xenon ions are
and an expel lant assembly.
source,
is a block
1
ion
used
to
biased
to
the
operated
diagram of the ion
gun
UIT-.^
21
and Figure
system
is an
7
the general electrical
Unneutra ized
electrical
ion beam current
1
mA to 2 mA at energies of
0.3
the normal operating modes,
keeper
The
on.
maintaing
and 2 KeV.
from
set
addition
In
to
there is a "trickle mode" which
is
the
means
discharge and all Xenon
the
cathode and the keeper.
In
trickle mode,
80 microamps and of
is 20 -
1
levels can be
the beam voltage is switched off and the keeper
occurs when
remains
indicating
scheiaatic
interconnections.
for
starting
and
through
the
flows
the beam discharged
low (essentially thermal,
""1
eV)
energy.
The beam current is measured in several ways.
beam
The
telemetry value is the current passing through
current
the
beam power supply. The net beam current is the measured beam
current
supply
currents.
reduced by the sum of the acce
Therefore,
spacecraft
common
the
and
1
decel
measured
SP BS net current,
I
as shown in the schematic of
Figure
at
8,
more accurately defines the current leaving the ion gun. The
current is measured by a
SP BS
net
whose
output is accurate to
1
positive
or
Neutralizer
+^10% of
bipolar
current
the true
negative currents greater than
emission
electrometer,
and net currents (ion
2
and
(for
microamps).
electron)
between the SPIBS instrument and the satellite ground can be
detected down to a level of 2 microamps.
1978; Masek,
1978))
(tiasek and
Cohen,
Figure 7.
SC4-2 ion gun electrical schematic.
f^
TO HEATER CURRENT
ERROR AMPLIFIER
(NEUTRALIZER
HEATER SUPPLY)
I
SENSE
CI-TOrtrLiCTCDC
ELECTROMETERS
SENSE
I
^
"
^^
?
''^
POWER
SUPPLY
S/C
GROUND
1
i
1
/T7
Figure
8
.
Electrometer block diagrao
r
Inet
tm
III.
this thesis,
In
will
OBSERVATIONS
results from four sets of observations
be presented.
Data from the UCSD
particle
detectors
(SC9) and the GSFC electric field monitor experiment
(SCIO)
will
be used to investigate the effects of emitting an
ion
beam
from
The
spacecraft
a
in
geosynchronous
observation days to be presented were chosen to
various
modes
of
ion gun operation in
both
orbit.
demonstrate
sunlight
and
ec ipse.
1
The
sunlight.
first set of observations concern the satellite
The
ion gun is operated in a high
in
energy,
high
current mode which induces a large negative potential.
Also
discussed
mode"^.
is
a
low energy,
low current mode termed "trickle
The SCIO experiment is examined to determine if
it
accurately measures spacecraft potential during sunlight ion
emission.
The
second set presents ion gun operations in
a
different environment and builds on work already reported by
Olsen (Olsen,
1987).
The gun is operated in sunlight at
a
^NASA's Polar
satellite,
part of
the
NASA Global
have
to
is planned
(GGS) program,
Geospace Science
hollow cathode source which will
produce an ion
a
be
due
to
It is
flux
equivalent to trickle mode.
launched in 1992.
lower current.
slighty
of
observations
spacecraft's
currents
is
SCIO
The
of
also
examined and compared to
Lastly,
measurement
the
voltages
beam
the
and
operations.
sunlight
validated
also
is
Here
eclipse.
in
different
to
for
response
satellite's capacitive
termination of ion emission
A.
low current mode^ is
kV,
1
the spacecraft
response
potential
eclipse.
A
This section concludes with two similar sets
examined here.
to
discussed.
is
DAY. 200
from operations on 19 July,
Data
1979 (Day
200),
are
presented first to illustrate two modes of operation of
the
gun
ion
in sunlight.
The first example
ion
emission
The satellite located in
begins shortly after 2140 UT.
plasmasheet
of
near local dusk (1954-2042 LT) between L
=
the
7.6
and 8.0, at an altitude of 7.5 Re and a magnetic latitude of
6 degrees.
The
particle
data
count
day
this
for
spectrogram format in Figure
9.
rate (or particle flux)
In a
is
are
summarized
spectrogram,
plotted as a
in
instrument
function
of
time on the horizontal axis and energy on the vertical axis.
^Low current mode is applicable to the European Space
This is
Agency Cluster satellite liquid metal ion gun.
a 10 - 20 microamp, 7 kV device to be flown in 1995.
c^
-20
t
L
:
\-'''r'''p'^n'-:'i=yrrWfMff-^^^
3
Higure9
.
Spectrogram for ions and electrons,
Day
idOO
27
of
1979.
Detector
Angle
The flux magnitude is depicted using a gray scaie« with high
fluxes
appearing
fluxes
showing
dark gray or black and
as
up
light gray
as
to
low
displays data from the two high-energy detectors,
electron data on top.
downwards
figure
with
the
The energy axis is zero in the middle
electrons
the figure and increases upwards for the
of
zero
or
This
white.
The lower horizontal
and
line plot
in
the center is the pitch angle of the measured particles.
The
angle
pitch
is
the angle between the
and the magnetic field line.
vector
velocity
particle's
The upper
horizontal
plot is the detector head angle (see Figure
line
the
for the ions.
plasma injection at 2149 UT,
increase
particle
in
over
flux
wide
a
Note
4).
abrupt
recognized by the
energy
range.
Injections are the sudden appearance of hot plasma and occur
at
least
more frequently during
daily,
magnetic
activity,
Whipple,
1986).
(count
rate)
is
variable
with
periods
spectra
of
(Mcllwain
high
and
The peak in the differential electron flux
in the
1
to 10 KeV range,
typical
for
the
plasma sheet. The upper bound on the electron flux ("10 KeV)
is
the result of magnetospher ic convection processes and
termed the Alfven boundary.
of
It
is
must exceed a critical energy
15 to 20 KeV for negative charging to
The
occur.
high
energy ion flux extends up to 80 KeV.
Unneutra ized ion gun operation at
1
setting of
1
1
kV and
a
current
mA begins at 2141 UT and terminates at 2301 UT.
This
results
spectrogram
high
The
in satellite charging,
as seen in
as a black band representing high
the
fluxes are due to ambient low energy
2200-2203
voltage),
correspond
to the times
band
flux
ion
potential
There are two periods of trickle mode operation (no
energy.
accel
being
ions
accelerated into the spacecraft at the spacecraft
ion
fluxes.
ion
is
Electron
and
UT,
which
black
UT
and
This
absent.
potential magnitude of
2227-2232
in the spectrogram where the
suggests
spacecraft
a
less than 10 Volts.
ion data for the 0-80 KeV
energy
range,
taken from the HI detectors at 22:01:56 UT, are presented as
distribution
functions
Figure
in
characterize the environment.
phase
divided
space
density,
is
order
in
to
The distribution function, or
proportional to
the square of the energy.
by
10,
In
count
the
this
rate
format,
a
Maxwell ian distribution would appear as a straight line. The
density
and
calculated
temperature
the
of
observed
particles,
from the distribution functions using
wise least squares fit method,
is
given in Table
a
piece-
II.
Density
and energy can also be obtained by taking the moments of the
distribution function.
density
This is also shown in Table
estimate obtained by integrating
function can have several flaws.
the
II.
The
distribution
I
I
I
I
I
I
I
\
SCATHA
HI Electron Detector
Day 200 of 1979
%
22:01:56
M.
'>•,
'
I
r
-]
L
1
1
1
\
\
\
1
1
1
1
1
1
SCATHA
HI Ion Detector
Day 200 of 1979
22:01:56
i
-
r
-]
V%U
t
'
t
::**;,
i
1
Figure 10.
1
1
1
1
1
1
1
1
Electron and ion distribution functions, Day 200.
::
,
1
For ions:
positive satellite potential
will
repel
low
energy ambient ions,
resulting in an underestimate
of density.
a)
a
b)
pitch angle distributions are typically anisotropic,
but the assumption of isotropicity is made.
Most of
the
measurements are biased towards perpendicular
ions
c)
the mass assumption made in calculating the distribution function from the count rate is that m = 1
amu (proton).
This
results
in
an underestimate
the
of density
if
by up
to
a
factor of 4
0"ions are 0^.
Typically,
the
is about 20-50% of
factor
ions,
so the density is underestimated by a
of
2.
For electrons:
a positive satellite potential
results
in
trapping
integration
locally generated photoe ectrons.
The
stopped to avoid this contribution, but density can
is
overestimated.
Note that exp(e$/kT)
still
be
greater than one for electrons for a positive
satel i te.
a)
1
TABLE
I
I
DAY 200 PIECE-WISE MAXWELL AN PARAMETERS
I
Energy Range
El
100-1200
ectrons
eV
KeV
80 KeV
50 eV-80 KeV
4-10
10
Moments
I
ons
:
10 -500
eV
KeV
80 KeV
eV-80 KeV
1-25
20
Moments
-
1
-
Density (cm-3)
Temperature
0.
69
1.
15
306 eV
2.6 Kev
8.3 KeV
2.2 KeV
,
0. 13
1..43
0. 17
0..008
0.,0028
,
1..34
eV
1 KeV
167 KeV
3.6 KeV
139
14.
mode operation will be examined first and
Trickle
of
the results applied to the observations of
1
some
gun
kV ion
operation.
1
Low Energy, Low Current Mode
.
Observations
a.
Figure 11 is an example of the SCIO potential
trickle mode operation.
In
in
trickle mode the beam voltage is
switched off and the keeper remains on, allowing the ions to
diffuse
the
out
ion
of
gun
the gun.
current,
monitor telemetry,
Two spin periods are
taken from the
SPIBS
with
shown
net
current
plotted on the bottom of the figure. The
data indicates the mainframe is within a few volts
SCIO
the antenna potential, and hence near plasma potential.
the
probe
modulated,
voltages
and
the
and by inference,
net
ion
gun
current
of
Both
are
so is the spacecraft mainframe
potential. This may be explained by examining the effect
of
solar radiation on the satellite.
b.
Analysis
Since the photocurrent is the major current from
the spacecraft at geosynchronous orbit,
properties
potential.
greatly
influence
of
the surface material
spacecraft
ground
Especially important are the conducting surfaces
grounded to the spacecraft frame.
bottom
the
The normal
to the top and
the satellite are close to perpendicular
to
the
sun and thus their materials do not contribute significantly
2201
2202
TIME (HHMM)
2203
IJT
Figure 11. SCIO and SPIBS net current during trickle mode.
33
to
photocurrent
the
.
Most of the conducting
material
is
contained in the band around the middle of the cylinder that
contains the instruments.
of
indium oxide coating,
There is also a significant
a conductor,
area
cells
on the solar
under the SC9 instrument.
The
total
photocurrent of illuminated
exterior
grounded conducting surfaces from the SCATHA satellite at
O
Volt potential has been calculated by Craven (Craven et al.,
The
1987) as a function of the rotation angle from the sun.
calculated photocurrent varies from 8 to 20 microamps and is
shown in Figure 12.
between
and
This figure also shows the relationship
this total photocurrent,
the ion gun net
the SCIO potential as a function of the
the SClO-2 boom and the sun.
potential
current,
between
angle
This illustrates how the small
modulations are strongly linked to
the
material
properties of the conducting surface areas grounded directly
to the spacecraft frame and can be understood by considering
the balance of currents to the spacecraft.
In
the
steady state,
satellite
potential
is
governed by the balance of incoming and outgoing current:
where I. is the ambient electron current (less secondaries),
I^r,
is
the photoe ectron current from the spacecraft,
1
the ion beam current,
and
I
r-
is
the returned beam
1^
is
current.
o SO""
Boom-Sun Angle
Figure 12. photocurrent
,
ion gun net current, and SCIO potential as a
function of boom-sun angle.
35
term
additional
One
that
can
significant
be
between the booms and the
photocurrent
small positive potential on the satellite,
photoe ectrons
from
1
1987).
the
Conversely,
for
For
a
the effect of the
booms is negligble
a small
the
is
spacecraft.
et al.,
(Lai
negative potential on
the
the fraction of the photoe ectron current going
spacecraft,
1
the booms is negligible.
The ambient ion
currents
are
usually much smaller than the ambient electron currents
and
to
may
be
ambient
The
ignored.
electron
currents
(less
secondary emission) are generally important in charging, but
for
200 in sunlight are small enough relative
Day
others to neglect.
as in this case,
If
the spacecraft potential
the return current
Ir-
to
disregarded.
can be
For potentials less than the beam energy the return
is
negligible because all the ions from the
escape (unless scattered for some reason).
result
of
current
source
should
the major
Hence,
terms are the photocurrent and the ion source
satellite potential
current.
(and hence the SCIO measurement)
the balance between these two.
In
the
positive,
is
is
trickle
the net emitted ion current is determined by the
The
the
mode,
spacecraft
potential and is probably proportional to expCe'^/kT].
Using the SCIO data,
spacecraft
potential
off can easily be found.
mode voltage,
first order values for the
in trickle mode and with the
ion
gun
The expected behavior of the common
as derived from probe theory
during
periods
of no appreciable charging,
in angle
(0)
varies as the satellite rotates
The oscillatory variations
like ln(sin 0).
twice the satellite spin period are due to changes in
at
probe
illumination. The negative spikes every half spin period are
due
to
the singularity in the
boon potential. Physically,
suppression
ln(sin0) dependence
the singularities are due to the
the photocurrent when the antennae
of
the
of
are
at
grazing angles relative to the sun vector. When the antennae
are directed along the sun-satellite line,
minimuD
or
most
satellite. At this point,
potential.
maxima
to
is
respect
the
to
the antennae are closest to plasma
At or near normal
sun angles,
the boom is close
The one volt modulation
its maximum potential.
to
they are at their
negative potential with
caused by a change in the spacecraft
of
the
ground
due
the various exposed conducting surfaces directly grounded
to
the
spacecraft frame and not from a difference
illumination of the antennae (Craven et al.,
An
maximum
of
5.4
the
Volts
for
the
(Lai
et
Using this value for the maximum boom potential,
1986).
the
theoretical
trickle
value
SCIO boom potential was determined by Lai
al.,
along
approximate
in
1987).
Figure
13,
with the measured SCIO potential for ion gun off
and
boom potential
mode operation,
approximate
magnitude
is
plotted
as a function of
and variation
in
in
sun
angle.
spacecraft
The
frame
potential can be obtained by taking the algebraic difference
—
-
6°
4.
Theoretical boom potential
n^n hP)t\
VV
'
2"
1'
^.y^^ SCIO potentia,!
/
\ (gun off) /
>
1
^
1
^^
\
SCIO potential
(gun on)
0-
-2-
SCATHA
Day 200 of 1979
H
1
\
180
\
180
—
1
\
180
Boom-Sun Angle (Deg.)
Figure 13. SCIO potentials for ion gun on and off.
SCATHA
Day 200 of 1979
Spacecraft potential
(gun off)
Spacecraft potential
H
180
(gun on)
-h
-t
-4
180
180
1
180
Boom-Sun Angle (Deg.)
Figure 14. Estimated spacecraft potentials for ion gun on and off.
between
the boom potential and the SCIO measured
is shown in Figure 14.
and
potential
The gaps represent the
periods
when the antenna was pointing directly at the sun or in
shadow of the satellite.
potential modulation but reduced the spacecraft
Volt
the
The ion gun had no effect on the
1
frame
potential to just slightly positive.
The
distribution of the emitted
approximately
photoe ect rons
I
a
most
probable energy on the order of 1.5 eV (Grard et al.,
1983).
is
Since
spacecraft
a
effectively
isotropic
an
potential
Maxwellian
of
about
off the photocurrent
shut
will
potential
just
requirement
that
a
,
volts
+5
slightly
positive is consistent with the
most
the photoe ect rons escape in order
of
with
1
to
cause
the
observed potential modulation.
Because the injected ions and the photoe ectrons
1
are of
Since T
2.
low energy,
"
1
eV and
the currents are modulated as expCe-5/kT3.
§
"
1
seen in the net current.
the
Volt,
expCe-J/kT] has a value near
This agrees with the modulation from 40 to 80
magnitudes
of
There is however,
the photocurrent and
since the two should be roughly equal.
either
SPIBS
in their
ion
Thus it appears that
calibration
vary
estimations of photoyield values)
is off
in
current,
the calculated photocurrent (laboratory values
considerably
the
the
microamps
a disparity
by a factor of
2
to
4.
or
It
should also be noted that this does not affect the
validity
of Craven's results.
When
the
Current Mode
High
High Voltage.
2.
gun is activated
ion
2141
at
the
UT,
satellite is driven to a potential between 600 and 800 volts
negative,
This
which is significantly less than the beam energy.
indicated
15
repiesenting
the
the spectrogram
in
ambient
energy ions being
low
by
black
a
batid
The ion flux peak is due
high ion fluxes.
accelerated
into
to
the
spacecraft at the spacecraft potential energy. Potentials of
less
than about
this
method,
+^10
Volts are difficult to determine
seen by the complete absence of
as
charging peak during the two trickle mode
using
any
operations.
accurately
ion
The
determined
spacecraft
potential
converting
flux (count rate)
functions.
The peak at about 700 eV in the ion distribution
function of Figure
to
is
most
to
the
particle
taken at 22:10:44 UT,
15,
be at the spacecraft
potential.
by
distribution
is
interpreted
Representative
energy
channel widths are shown using bars on the data points.
Note
that the energy channels overlap.
Figure 16 is a plot of the potential as measured
the SCIO experiment.
A
minimum of two points (the high
low values) are plotted per spin period.
spin
period
is
immediately obvious.
A
by
and
modulation at the
There is
transient, also seen in the SC9 data (not shown),
a
turn
on
indicating
P^
(T>
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0)
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11
11
SCATHA
SCIO Potential
Day 200 of 1979
Time (UT)
Figure 16. SCIO potential. Day 200.
Time UT
Figure 17. Ion gun beam current.
(
ill
I
spacecraft momentarily charges to a negative
the
corresponding
spacecraft
potential
voltage.
beam
the
then rises (becomes
potential
The
negative).
less
behavior is seen each time the gun is switched on
This
is
the magnitude of
to
counter- 1 ntui
t
i
the beam current,
ve as
actually increases during the transient.
17,
current
beam
beam
net
which stays constant,
current,
spacecraft
Note that the
measured at the beam power supply
is
and
plotted in Figure
and
the
measured
is
at
common.
Figure
shows greater resolution
a
period
after the turn on transient and is representative of
normal
18
ion gun operations for Day 200.
every
plotted
points
between
It
for
shows the potential,
measured
SCIO
2210 UT and 2215 UT or for approximately five
spin
plotted
Also
periods.
in
as
this figure
the
is
spacecraft
potential as determined by SC9 charging peaks for the
Energy
detectors.
LO
representative
between
the
experiment
SC9
SC9
channel
data points.
and SCIO values
The resolution of
measurements above 300
channel
I
The
excellent
suggest
of
V.
HI
shown
are
the
the SCIO experiment
and
for
agreement
SCIO
the
that
ion-emission induced charging
during
sunlight.
sraal
widths
provides a valid measurement of
potential
the
with
by
two seconds,
spacecraft
events
is
40
V
in
for
This and the difficulty in picking
the ion peak may account for
differences between SC9 and SCIO values.
some
of
the
«
!
'
900
1
O
SCATHA
Magnitude of SC9 and SCIO PotentiaJ(^
Day 200 of 1979
"•
A
•
1'.'
'
HI Ions
LO Ions
/•
'J
r|
500
2210
11
12
2215
14
13
Time
UT
(HHMM)
Figure 18. SC9 and SCIO potentials, Day 200.
1
1
i
1
1
1
1
1
1
SCATHA
HI Ion Detector
**
Day 200 of 1979
-
22:06:28
1— 4—
-
\
*—
I
\**
/
*
.
charging
peak
-
-
-
1
Figure 19.
1
1
.
1
1
1
1
Ion distribution function,
44
1
1
Day 200.
Normal operation of the ion gun for this observation
period
is
characterized in the SCIO data
noise-like
fluctuations,
trickle mode operations,
that
rapidly
fluctuating.
is
constant
outputs
the
spacecraft
the
off
frame
potential
Since
period.
of
the
caused
by
+^10%
beam
or
spin
The SC9 data appear
the
true
to
indeed
is
1
during this
in
gun
The SP BS Net Beam Current
are accurate to
variations
rapid
large,
and the previously mentioned
modulation peaking once per minute.
confirm
by
present during
not
telemetry
electrometer
current,
sputtering
gross
other
or
mechanisms within the ion gun do not appear to be the cause.
Given
in
failure of the satellite to charge to
the
energy,
the
it
beam
the
seamed reasonable to search for signs of the beam
kV channel.
1
Figure 19 shows a distribution function taken during
gun
ion
operation at 22:06:28
UT.
Representative
energy
channel widths are shown using bars. The significant feature
of
than
this figure is the second peak at an energy greater
the spacecraft potential
secondary
peak appears to be the
into the UCSD detector.
distribution
three
and near the ion gun
function
minute
period.
angle
KeV beam ions
scattered
Figure 20 shows a portion of the ion
for
ten energy sweeps taken
Of
note is
distribution function around the
detector
1
This
energy.
1
the
variation
KeV energy
and pitch angle are given for
over
a
of
the
channel.
The
each
1
KeV
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sample.
The
strong
satellite potential
secondary peak at
statistical
analysis
1
is
-0.7 to -0.9 kV with
kV most evident at 2220
or correlation ot
UT.
second
the
a
No
higher
energy peak to the detector head angle or the particle pitch
angle was attempted.
distribution
operation
This type of temporal behavior in the
function
typical
is
of
(excluding trickle mode).
periodically,
significant
It
Day
200
ion
gun
may
indicate
that
fluxes of ions at or
near
beam
energy are being returned to the satellite.
The
peaks
once
angle
two
potential modulation seen in Figures 16 and
is
per spin period.
not
a
This indicates the
injections.
peaks
of
During electron beam injections
spin
per
period
are
due
electron
the
I
operations
295
The
1987).
is
spin
seen
during
et
ion
gun
not observed in eclipse (see Day 293 and
Day
suggesting
observations),
influence.
modulation
two
boom- genera ted
to
pho toe ect ron currents flowing toward the satellite (Lai
al.,
18
sun
significant factor since this would introduce
peaks per spin period as seen in the case
beam
boom
the
To investigate further,
sun
is
the
driving
the orientation of
the
satellite relative to the sun was determined.
The singularities observed in trickle mode occur
a
boom-sun angle of
around
mode,
the
0*=*
and
180*^.
ion gun transition from full
the phasing of
at
By examining the SCIO data
on
into
trickle
the spin modulation can be determined.
the minima occur at a boom sun angle
into trickle mode,
about
Thus the satellite's orientation
SCIO potential minimum corresponds to the
the
emitting
of
The second transition into trickle mode at 2227
300*=*.
UT yielded the same result.
at
transition
which depicts the 2200 UT
As seen in Figure 21,
sunward
the
in
direction.
This
ion
gun
suggests
two
possible explanations:
1)
The net ion current is much smaller in magnitude
than
the telemetry indicates.
The electron density near the
spacecraft
the sunward side
is
photoemi ss ion.
higher on
because
of
The beam current is being balanced
by
pho toe ect ron
the
1
current,
reducing
thus
the
net
current escaping the satellite.
2)
Neutralization
of
the
space-charge
beam
as it passes through the
place
is
taking
pho toe ect ron
cloud,
1
shown schematically in Figure 22.
increased electron density,
might also reduce the net current
This
sate
1
I
i
beam's
To
escaping
the
te.
Insufficient
problem.
This is a region of
on the order of 100 cm-3.
do
parameters
information
is
available to resolve
so would require better knowledge
and
behavior
and
more
accurate data on the ambient environment.
of
complete
this
the
and
—
,
SCATHA
SCIO Potential
Day 200 of 1979
2157-2203 UT
(
,<
I
n
.
I
r
.
'
r.
I
I'
•
'
r.
,V'.
.
:
J''
'
I
!»
'
II
'•
i
'
I
'
I,'
TT
ii
-I
1
1
180
1
180
1
1
1
\
\
180
180
»
180
Boom-Sun Angle (Deg.)
Figure 21. SCIO potential during transition into trickle mode.
Boom-sun angle
SClO-2
\,
Photoelectron
cloud
Figure 22. Schematic of spacecraft's orientation.
B.
DAY 47
Observations
for
February,
16
1979
(Day
are
47),
presented to illustrate the effects of ion gun operation
the
spacecraft
similar
(Day
operations
was used.
of
In
a different
in
the ion gun an experimental
this mode,
the keeper operated.
microamps.
environment
During
parameters.
beam
200)
plasma
on
using
initial
low current
mode
the main discharge went out and only
This resulted in a beam of
Unfortunately,
1
kV and 25
the SCIO booms had not yet
been
deployed so those data are not available.
On this day the spacecraft was located near an
of 5.5 Re between L
in
the
plasmasheet
=
shows
the
detector at 1447
UT.
Figure
at local dawn.
distribution functions taken by the
The
altitude
5.7 and 6.5. Data used here were taken
HI
calculated ambient particle densities and
23
temperatures
The electron density moment
are displayed in Table III.
is
slightly higher than on Day 200.
1.
High Voltage, High Current Mode
a.
Observations
In
Figure 24,
the spacecraft potential,
on
the energy peak,
is
which is close to
plotted versus time.
transient at 1449 UT again seems to show
charging
the
The turn
satellite
to near the beam energy and then stabilizing
near
-500 Volts after one minute. This is 200 volts more positive
than the Day 200 operations.
Figure 25,
taken from the ion
.
I
\
I
\
SCATHA
HI Electron Detector
Day 47 of 1979
14:47: 34
r
"1
\
\
!
\
1
;
1
1
1
1
1
1
I
SCATHA
HI Ion Detector
Day 47 of 1979
14:47: 34
1
******
%
*
*
'-«
V*
.
*
I
1
Figure 23.
1
1
1
1
1
1
1
1
Electron and ion distribution functions. Day 47.
ENERGY FOR PEAK ION COUNT RATE
SCATHAAJCSO
DAY 47 OF 1979
ri
DETECTOR
o
HI
+
LO/FIX
:i-.i
iji^h
TIME (HHMM) UT
Figure 24. Energy peak vs time,
52
Day 47,
SCATHA
Ion Gun Beam Current
Day 47 of 1979
1 .1-
/Nv^-^'^H^
en
/
^
/]
V..y\W
/
a
"0.9(
0.8.
:
^
-
1
*
r
30
w
a
\y^
B
22a
"
O
e
10
^
Time
Figure 25.
1
1
\
1450
51
(HHMM)
5
UT
Ion gun beam current.
53
1
53
Day 47,
1454
gun
telemetry
for the same period,
also
shows
the
beam
current increasing during the transient. After half a minute
the beam steadies at current value near 1.05 mi
1
1
iamp.
This
suggests that although the potential does vary with the
current,
transient
the
is
the response of
ion
satellite
the
system to a change in the input parameters.
TABLE
I
I
I
DAY 47 PIECE-WISE MAXWELL IAN PARAMETERS
Energy Range
Electrons:
20
200
2
20
- lOO
- 1500
20
eV
eV
KeV
80 KeV
80 KeV
-
10 eV-
50
-
1
-
5 KeV
5 -
80 KeV
80 KeV
10
Moments:
b.
50
500
eV-
1
eV
eV
Density (cm"^)
Temperature
1.3
17,1
0.24
0.30
0.02
2.04
370 KeV
8.9 KeV
11.8 KeV
503 eV
0.12
15.8 eV
140 eV
1.3 KeV
14.8 KeV
5.5 KeV
O. 19
O. 13
O. 19
0.58
Analysis
The beam current is 0.5 mA lower than that
on
Day
200,
satellite
eV
which
potential.
could account for
the
The ambient environment
used
more
positive
is
slightly
denser and hotter but the effect on the spacecraft potential
during
ion
gun
operations is less
obvious
and
in
this
probably
case,
suggests
minimal.
that the SCATHA
changes
evidence
spacecraft
in current balance during
relatively
is,
The
presented
is
so
far
very sensitive
operation.
ion gun
small changes in current
cause
to
That
relatively
large changes in potential.
On this day the spin modulation can be discerned
directly
from the plot of the ion particle data
This
24.
other
fluctuations
fluctuations
seen
on
are
Day
magnitude.
of
in
not
200,
the
Although
potential.
as prevalent in this data
they are probably present
at
to
the
as
those
a
lower
This would make them difficult to detect because
the previously mentioned limitations of the SC9
detectors.
particle
There was no strong evidence of a second, higher
energy peak above the charging peak.
appear
Figure
in
was difficult to do with the Day 200 data due
be
to
more stable
and
well
the beam
Thus,
would
This
behaved.
also
suggests that the fluctuations in the spacecraft's potential
and
the
appearance of a second peak
in
the
distribution
function near the beam energy are related.
2.
High Voltage, Low Current Mode
At 1451 UT,
-10
the gun drops into the low current mode
to -20 Volt potential
net
beam
current is not modulated as on Day 200 because in this
mode
and
a
is
seen.
The
the ion current is determined by the gun perveance and
accelerating voltage.
As on Day 200,
1
kV
however, net currents
in the 10 to 100 microamps range result in relatively small
($oyc
the
50V)
<
ambient
plasma,
negative
the
potentials.
Once the emitted current
exceeds
current levels available through photoemi ss ion or
beam
should
the
potential,
energy.
spacecraft
will charge
to
from
large
a
at times approaching the magnitude
The magnitude of
critical
the
be roughly the same as the photocurrent
in
of
current
sunlight
as this is the dominant positive current.
C.
DAY 293
Data from operations on 20 October,
now
presented
Eclipse
eclipse.
provide
a
1979 (Day 293), are
to illustrate operation of the
events
are of great value
variation of one of the dominant
ion
gun
because
terms
in
they
the
in
spacecraft current balance equation, the photoe ectron flux.
1
The
effects of differential charging are therefore
and
potential
electrons
(Olsen,
barriers
between
1983).
and
spacecraft
the
transport
surfaces
can
When these data were taken,
secondary
be
neglected
the spacecraft
was in the plasmasheet at an altitude of 6.4 R^ and
L = 7.0 and 7.2.
minimal
of
between
The spacecraft enters penumbra at 2042
UT
and umbra at 2043 UT. Umbra exit is at 2141 UT. The data are
summarized
in the spectrogram of Figure
26.
contains several gun on/off transitions and
This
example
current- vo tage
1
combinations. The gun off transients at 2114, 2119, and 2129
are particularly interesting because of the way the ion peak
Fiaure 26. Spectrogram for ions and electrons,
Day 293 of 1979.
::
off over a period of several minutes.
trails
ion distribution functions,
2045
density
and
relatively
and
taken from the HI detectors
are shown in Figure 27.
UT,
Electron
calculated
The
Table
temperature are presented in
quiet day provides substantially
at
ambient
This
IV.
plasma
higher
densities than the previously presented experiments.
TABLE
IV
DAY 293 PIECE-WISE MAXWELL AN PARAMETERS
I
Energy Range
ectrons
50
200
-
-
2 10 10 eV-
Moments
10 -
Ions:
100
-
2 10 -
Moments
The
to
2130
Figure
1
eV-
150
lOOO
eV
eV
6 KeV
80 KeV
80 KeV
50
500
eV
eV
8 KeV
well
Temperature
1.2
2.9
5.7
0.036
5.5
59.5 eV
586 eV
655 eV
11.4 KeV
707 eV
0.063
20.6
eV
145
eV
6.9 KeV
12.5 Kev
11.0 KeV
0. 17
0.69
80 KeV
80 KeV
1.2
1.5
SCIO potential and SC9 data are shown for the
UT period of unneut ra
28.
1
i
zed ion
gun
shown in Figure 29.
in
shown
the SC9 and SCIO data agree quite
and show the satellite potential
corresponding
2100
operation
Representative energy channel widths are
with bars. As can be seen,
The
Density (cm~^)
fluctuating
beam voltage and net
beam
rapidly.
current
are
There is an obvious correlation between
I
1
1
1
1
1
1
1
1
1
SCATHA
HI Electron Detector
\
*
Day 293 of 1979
*
t
t
*
i
1
t
^
i
[III
1
1
1
1
1
1
1
1
J
1
1
1
SCATHA
HI Ion Detector
Day 293 of 1979
:
^
*
*
!
Figure 27.
1
1
1
1
1
1
1
1
Electron and ion distribution functions, Day 293.
59
I
SCATHA
T'
10
SC9 and SC ^^
T
Potential
Magnitudes
lit
I
i
),
,
".
II'
O"
l'")
I
Lit
I
'•'
'I'
op
Day 293 of 1919 Gil
I
•D
1
p.QO-'
I
1
II
Oi
6 6
Time UT
Figure 28. SC9 and SCIO potentials, Day 293.
3eam Vo 1 tage
u
i^
SPIBS Net Beam Current
r-
2100
Figure 29.
2120
2110
Time UT
Ion gun beam voltage and beam current.
beam
the
parameters
potential,
and the
summarized in Table
behavior
of
satellite
the
V.
TABLE V
DAY 293 BEAM PARAMETERS AND SATELLITE POTENTIAL
NET BEAM
CURRENT
BEAM
VOLTAGE
1.0 mA
IkV
0.3 mA
1
kV
SATELLITE
POTENTIAL
FLUCTUATIONS
-800 V
+
100 V
-900 V
+
100 V
0.5 mA
2 kV
-1500 V
+
500 V
1.5 mA
2 kV
-1800 V
+
100 V
During the time period from 2109 to 2114 UT,
is
operating
current
of
at
.5 mA.
a beam
voltage of 2 kV
and
the ion gun
Examining the SC9 data for
beam
net
a
this
period
closely showed the presence of secondary peaks in
more
Figure
30
21:12:45
to
distribution function at the beam energy (2 kV).
shows
the
21:17:17.
the ion distribution function from
These
data were taken from the HI
the
ion
detector.
Note that the gun was off from 21:13:49 to 21:14:37, as seen
by
the absence of any peaks in the
distribution
The most obvious 2 kV peaks are at 21:13:01,
21:16:45.
The
detector
function.
21:14:53,
head angle was fixed at
near
degrees and the corresponding particle pitch angle is
in the table.
During the 2125-2130 UT time period,
and
lOO
given
the ion
iii
r^
(J)
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t
I
gun
operated at 2 kV and a net current of
is
magnitude
of
fluctuations in the SCIO
the
smaller than seen with the 2 kV,
1.5
21:23:57 to 21:26:21.
distribution
.5 mA beam.
functions
The top plot shows the
Shifting the
for gun off.
HI
ion
unacce erated
ion
1
distribution curve up 800 eV (invoking Liouville's
or phase space
indicating a satellite potential
when the gun is on at
Volts at 21:24:29,
subsequent
Theorem,
invariance) brings the first and second plots
agreement,
into
much
is
Figure 31 shows a series of four distribution
from
The
mA.
data
energy scan,
plot three,
is
-800
of
kV and
1
mA.
1
same
for the
The
gun
parameters and is similar to the second plot. The difference
is
the boxed data point at
returned
1
potential
again
.
of
1988).
HI
-1800 Volts.
indicates
The two boxed data
beam ions.
of
represents
The circled
spacecraft
data
generated
beam.
satellite
a
points
interpreted as a result of spreading of the
distribution
likely
suggests
This data point
KeV.
last plot for 2 kV and 1.5 mA,
The
the
1
KeV ions from the gun or scattering of the
velocity-
point
If
our interpretation of
most
(Norwood,
ions
Figure 32 shows the corresponding distributions
electrons.
are
variations
for
in
satellite potential are correct, equivalent shifts in energy
should make it possible to bring agreement to these
This works for
is
1
kV but not for 2 kV,
curves.
where a Z^ of 1.5
needed for the electrons vice 1.8 kV for the
ions.
kV
The
9.0
.
SCATHA
Ion Detector
Day 293 of 1979
HI
21 :23:57
gun off
2.0
Energy (KeV)
Figure 31.
Ion distribution functions.
Day 293.
3.0
^
SCATHA
^
2.0
HI Electron Detector
Day 293 of 1979
A\
21:23:57
gun off
-^v
^.
\ ^-/
2.0
\\
\
-\^E= 800 ev
^1
C
E
Tn
2^^^^^^^
1
kV,
1
mA
0.0
2.0-
^
^ AE= 800 eV
21
1
\
A A
\
:24:45
1mA
kV,
>,
\
AE=1.5
\^
\
KeV
21 :26:21
2 kV,
1 .5
mA
00,
-2.0
2.0
3.0
Energy (KeV)
Figure 32. Electron distribution functions. Day 293.
65
explanation of this last discrepancy is uncertain.
may be
It
due to a change in the ambient electron density. A change on
the order of 10% would be sufficient.
In
comparing the data taken on this observation day
that taken on Day 47 for a
1
kV,
1
mA beam,
the spacecraft stabilized at a higher
that
obvious
is
it
(less
to
negative)
Table VI compares the beam parameters for Day 47
potential.
and Day 293.
TABLE VI
ION BEAM PARAMETERS.
Parameter
DAY 47 AND DAY 293
Day 47
(1450 UT)
Beam Voltage
1
kV
Beam Current
1
e- Density
2.04 cm~'
e- Temperature
.
05 mA
503 eV
Spacecraft Potential
Day 293
(2100 UT)
-500 V
1
1
.
kV
08 mA
5.50 cm~^
707 eV
-800 V
The ion gun parameters are the essentially the same; the
calculated
electron
densities differ by a factor
of
two.
Changes in the instrument response with time will affect the
validity of this comparison,
differences
days.
in
local
but it appears consistent with
time and magnetic activity
on
The other obvious difference between is that the
66
these
Day
293
operations
photoe ectrons
take place in eclipse.
likely
is
It
play a significant role in
1
that
determining
the
spacecraft potential during ion gun operations. One possible
effect
that in sunlight,
is
partially
space-charge
the ion beam
neutralized by the electrons of the
is
photosheath.
We would therefore expect the satellite to charge nearer
magnitude
excluding
extended
very unstable modes
the
survey
eclipse
logical
the beam voltage in eclipse than in
of
of
similar
at
unneut ra
1
i
operation.
of
zed ion
beam parameters
gun
would
the
sunlight,
more
A
operations
in
next
the
be
step in this study.
The additional feature which appears in eclipse data
a
is
delay in the return to the equilibrium potential. The SCIO
potential
33.
This
is
plotted for the gun off transition
the gun is switched off.
in the SCIO potential,
minute
Initially,
and hence,
long
decay towards zero.
when
there is a rapid decay
the spacecraft potential,
by an overshoot and a small
followed
Figure
in
segment is typical of the times for Day 293
and
rise,
finally
This can be seen
on
a
the
spectrogram as the trailing off of the ion peak to near zero
energy
when the ion gun is switched off.
decay,
a
From the
maximum value for the spacecraft frame
second
time
constant on the order of
last
decay has a time constant of 10 to
67
1
was
20
initial
capacitive
obtained.
The
seconds.
The
SCATHA
SC10 Potential
Day 293 of 1979
L 100
T
2119
2120
rime(HHMM) UT
Figure 33. SCIO gun off transition.
2121
last
set of observations show an excellent example of
this
effect.
DAY 295
E.
1
Observations
.
Data
(Day
295),
from ion gun operations on
are
presented
October,
22
illustrate
details
1979
of
the
eclipse.
The
spacecraft enters penumbra at 2002 UT and umbra at 2004
UT.
to
spacecraft response to ion gun transitions in
Umbra
exit
function
were
at 2058UT.
is
to 80 KeV are shown
for
taken
distribution
electron
and
Ion
from the HI detectors at
in
Figure
19:54:21.
34.
Table
Data
VII
gives the calculated ambient densities and temperatures.
on
Day
this is a quiet
293,
with
day,
relatively
As
high
ambient density.
TABLE VI
I
DAY 295 PIECE- WISE MAXWELL AN PARAMETERS
I
Energy Range
Electrons:
20 - 100
100 -lOOO
eV
eV
KeV
80 KeV
80 KeV
1-10
10
-
20 eV-
20 - 400
400 -lOOO
eV
eV
KeV
80 KeV
80 Kev
1-10
10
Moments:
1
-
eV-
Density (cm~^)
Temperature
2.0
2.7
4.8
0.038
6.0
25.2 eV
420 eV
940 eV
15.2 Kev
883 eV
0. 17
0. 12
111
eV
735 eV
8.7 KeV
13.2 KeV
10.1 KeV
0.74
1. 1
1.3
I
I
I
I
I
SCATHA
't
HI Electron Detector
Day 295 of 1979
1
1
I
\
i
I
\
r
I
SCATHA
Ion Detector
Day 295 of ly 79
HI
4
1
\
1.0
16.0
r
"1
1
1
r
24.0
Figure 34. Electron and ion distribution functions, Day 295.
Figure
The
SC9 data are summarized in spectrogram form
in
35.
The
is
decay spacecraft potential at gun
off
clearly seen in the trailing off of the ion peak at 2037 UT.
energies for the ion charging peaks for
The
three
minutes
after the ion gun is switched off are plotted in Figure
along
with the SCIO potential.
that found on Day 293.
36,
The response is similar
to
This behavior is seen repeatedly
in
this data set.
Anal ysis
2.
The
time
behavior
indicates differential charging.
two
different
structure
changes
The
action.
in
rapidly in
cessation of ion gun operation.
by
charging
Figure 36 illustrates
charging processes
potential
satellite
the
of
response
This process is
process
differentially
associated
sequence
is
more
determined
capacitance
is
The
suggested
Initially
37.
the mainframe is near -1300 Volts,
on),
Volts,
their eclipse equilibrium. Typically,
potential.
zero
hundred
several
and
At
volts less negative
gun off,
is
The
zero
these surfaces
than
the
the frame potential drops to
the the insulator potential
(gun
observed.
as
insulated surfaces are slowly discharging back to near
are
the
The
surfaces.
micro-Farads.
is shown in Figure
plasma.
involves
and
non-conducting
charged
of events
complex
the
the
to
the pico-Farad capacitance of the spacecraft to
second
the
First,
pushed
frame
near
towards
a
Figure 35. Spectrogram for ions and electrons,
Day 295 of 1979.
72
-
SCATHA
Maanitude of
SC9 and'sciO Potential
160
Day 295 of 1979
A,
G HI
A LO
OA-..
Detector
Detector
o...
A-.
OS-
—
'r^n—.
71
2036
37
38
2039
Time (HHMM) UT
Figure 36. SC9 and SCIO gun off transient.
m
l/^
-200
"S
f
N
t
Spacecraft Potential
;:
-1300
V
300
V
L
T
Insulator Potential
;
S
/^
A<P
Figure 37. Spacecraft and insulator potential.
positive
value.
The positive surfaces are not
attract a large negative current.
equilibrium,
negative
responds
again.
The
by
The system,
swinging
differential
the
stable
frame
potential
and
still seeking
potential
then
decays
exponentially with a time constant on the order of 10 to 100
seconds, allowing the mainframe to approach zero Volts.
DISCUSSION
IV.
has been suggested
It
SCATHA
unneut ra
limited.
The
space-charge
1
i
zed
(Stannard et al.,
referenced
1980)
extremely
ion beam is
analysis was
based
that
the
space
charge
upon
simple
limited diode theory and brings the
predicted
satellite voltages into rough agreement with experiment. The
predicted
net
beam currents escaping
ion
from
near
the
satellite region are on the order of 50 to 60 microamps. The
photoe ectron
1
observed
in this
magnitude.
the
current could have a significant
Also,
satellite
thesis,
increasing when
potential
limiting
the ion currents are
as
this
of
the data in Table V show the magnitude of
current is decreased from
charge
if
effect,
of
1
mA to .3 mA.
the ion beam is
the
consistent
beam
kV
1
Therefore,
space-
with
the
observations made in this thesis.
When
limiting
There
is
the
injected current
current,
exceeds
the
space-charge
the system becomes inherently
no simple analytical
framework that
adequate description of the complex time-dependant
bunching and reflection that occur.
studied
1963),
fluctuations
seen
an
particle
This phenomena has been
using numerical particle simulation codes
and Birdsall,
unstable.
provides
(Bridges
The instability predicted resembles the
during
operation of
75
the
ion
gun
and
extends
the
space-charge
limited
diode
one
model
step
further.
what
Qualitatively,
current
injected
a virtual
limit,
back
from
may be happening is that
the gun exceeds
as
the
charge
space
anode is formed. Some charges are reflected
the injection plane.
to
the
This
reflection
creates
a
charge bunch that further traps and reflects charge injected
at
Simultaneously,
times.
later
the
charges
were
that
injected prior to the first reflection continue to exit. The
electrostaic potential of the virtual anode region decreases
because charge is exiting the region in both directions. The
charges injected at still
and propagate out,
then repeats.
frequency
does
frequency.
The
exiting
from
may
trapped
The oscillations are not sinusoidal and their
not
scale linearly with
the
resulting oscillations in the
the
satellite region
potential fluctuations.
function
later times are no longer
increasing the potential. The oscillation
may
beam
net
account
plasma
current
for
the
The second peak in the distribution
be explained as ions from
the
gun
randomly
intercepting trajectories back to the spacecraft because the
beam never comes to an equilibrium state.
CONCLUSION
V.
Examinations
that
trickle
a
microamps
(no acce
effective
is
relatively
daylight ion emission
of
1
potential
for
reduce the small
not
control.
(about
frame
the exposed grounded conducting surfaces.
mode
For
successful
was
in reducing
80
the
thesis,
Trickle mode, however,
Volt)
1
potential
modulation caused by the different photoemi ss i ve
of
suggest
than
less
typical observation day studied in this
the potential was held to near zero.
did
events
voltage) current
properties
The low current
satellite
the
potential
magnitude to less than 50 Volts.
potentials
also
appears
responsible
be
potential.
close
at
1
and
kV
2
induced
less than the magnitude of
This is attributed to space-charge
energy.
may
operation
gun
Ion
satellite
beam
A
the
for
current of .3 to
to be close to optimum;
to
magnitude
the
fluctuations are small
NASCAP
operation
(see
as
of
beam
the
which
effects,
fluctuations
in
the
mA for a
kV
beam
1
the induced
beam
the
negative
1
potential
energy
and
is
the
in magnitude.
Appendix)
examined
in
modeling
this
of
thesis
similar
resulted
ion
gun
in
the
spacecraft charging to a value equal to the beam energy down
to
a
beam
current
near
25
microamps
(Olsen,
private
communication,
NASCAP has been thoroughly validated
1988).
as an accurate model of spacecraft charging response but
take into account space charge effects
not
does
it
other
or
aspects of charged particle beam physics. This also suggests
that the beam itself is the source of the effects
described
in this thes is.
the spacecraft charging system appeared very
In general,
sensitive
Operations
at
deviations
in
the
current
balance.
kV in high and low current modes suggest
1
current above which the spacecraft will charge
critical
near
small
to
beam
energy.
critical current,
If
the beam current is
less
than
a
to
the
the spacecraft will charge to a much lower
(less negative) potential.
NASCAP code (Olsen,
This was also indicated using the
private
communbicat i on,
which
1988),
yielded a value near 25 microamps for the critical
current.
This is consistent with current balance theory in that it is
of
the
same order as the estimated photocurrent
from
the
spacecraft.
A
fundamental
limitation of this thesis is the lack
accurate data concerning the composition,
of
the ambient particle environment-
the behavior of
impossible.
energy,
Detailed analysis
the ion beam is greatly complicated,
This
exemplifies the need for better and
complete instrumentation on future satellite flights.
78
of
and mass
if
of
not
more
The
existence
differential
of
charging
was
inferred
during ion gun operation in eclipse and may be expected
any induced charging event.
most
using
NASCAP.
While
direct
no
also
This was
observations
of
for
suggested
barrier
formation can be made for the observation times used in this
thesis,
existance of one
the
is
typical
for the
estimated
differential
charging
this thesis.
The formation of a barrier would increase
density
their
of
and environments
electrons near the spacecraft by
escape.
environment
level
would
affect the strength of the
value
dependendence
analyzed
by
spacecraft
the
the spacecraft potential
of
on
of
charging
differential
examining
the
data
in
the
allowing
ambient
barrier
interaction of
ion beaa with the barr ier- trapped electrons
operations
not
Since the density and energy of the
hence the density of trapped electrons,
equilibrium
observed
may affect
and
the
the
potential.
The
during
gun
could
collected
be
by
ion
further
the
SCI
experiment surface potential monitors.
In
emitted
conclusion,
the
interaction of
beam,
and
the ambient
ion
a
spacecraft,
environment
is
an
very
complex and difficult to unravel. More accurate knowledge of
the environment and better beam diagnostics are required
fully analyze the physics.
which
However,
may provide at least a fundamental understanding
suggestions for future experiments.
79
to
much data already exists
and
c
APPENDIX
NASCAP
The
NASA Charging Analyzer Program (NASCAP)
in
a 16
X
currents
16
spacecraft
to
Material
environments.
spacecraft
can
Particle
be traced,
and
magnetospher i
in
on the
modeled as well as the
is
connections.
surfaces
placement
object
potentials
32 grid and solving for the
X
three
is a
dimensional code which is capable of establishing an
surface
trajectories reaching the
including those from onboard
the
of
electrical
interior
detectors
positive
and
in order to estimate escaping flux and to
negative emitters,
simulate active control of satellite potentials. The complex
three-dimensional
effects can more easily be studied
using
the codes extensive graphic capability.
Simple
used
as
Particle
shapes,
building
such as cubes or slices of
blocks to model
more
cubes,
complex
fluxes of the ambient environment can be input
single or double Maxwell ians or as actual observed
The
code
are
objects.
accounts
for
electrons
incident
backscattered and secondary electrons,
and
secondary
as
spectra.
protons,
electrons
due to protons, and photoe ectrons.
1
The
NASCAP
code
computes
alternate
solutions
Poisson's equation and calculations of particle motion and
80
to
charge
deposition in a series of time steps appropriate
specific
the
process
problem.
proceeds
NASCAP
and
is
It
assumed
that
through a series of
its underlying physical
equilibrium
principles
to
charging
the
states.
been
have
validated by comparing prediction to experimental and actual
spacecraft
NASCAP
data.
has proven itself a
very
useful
predictive and analytic tool.
that particle trajectories may be calculated out
So
series of nested grids.
spacing
Each successive grid has twice
the previous grid with the spacecraft
of
to
NASCAP computations are carried out in
distances,
large
the innermost grid.
a
the
being
in
This thesis uses the "one-grid"
NASCAP
model which uses considerably less computer time and
memory
but
shows a similar charging response to the more
versions.
The
allows the model to reproduce SCATHA'
1
1
Each
in
s
a
which
features
geometrical
.
of
the 15 distinct exposed surface materials
this model
Experimentally
otherwise
as
Grid resolution is 19.6 cm
right octagonal cylinder.
very we
detailed
spacecraft's main body is represented
is
specified by the value of
measured
estimates
values
based
were
on the
used
14
if
properties
used
parameters.
available,
of
similar
materials were used.
The NASCAP code can recognize plasma spectra in the form
of a single and a double Maxwellian.
In
general,
the double
double
Maxwellian
is a
better representation of
non-
the
Maxwell ian actual environment.
Particle
emitters are modeled in NASCAP
by
specifying
the location of the emitter on the spacecraft and the energy
and
angular
width
of
the beam.
The
the satellite sheath field
tracked
in
through
a distant boundary or
beam
until
1980; Stannard et al,
82
they
1980)
are
escape
return to a specific cell
the spacecraft.
(Rubin et al,
particles
on
LIST OF REFERENCES
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T.
L.
B.
G.
Ledley,
A. Egeland, I. Katz,
"Probe Measurements
of
DC
Electric
Fields,"
Proceedings of the 17th E5LAB Symposium on Spacecraft/
Plasma
Interactions and Their Influence on Field and
Particle Measurements. ESA SP-198, pp 13-17, 1983.
,
Banks,
P.
M.
A. C. Fraser-Smi th, B. E. Gilchrist, K.
Harker,
R.
L.
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Storey,
P. R. Williamson, "New
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in
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Engineering, Space,
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J.
Bridges,
W.
B.
and C.
K.
Birdsall,
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Instabilities
Journal
of
in
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Chappe
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D.
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1
,
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Craven,
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D.
R. C. Olsen, J. Fenne
D. Croley, T.
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Journal of Spacecraft and Rockets
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I
,
1
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,
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77, pp.
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,
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65-
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A. Myasnikov, G. Zhadan, G. Orlova, M.
Gavrilov,
F.
Ion"Some Results of Flight Tests of an
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Engine Model Using Surface Ionization of Tungsten,"
v.
edovaniya
Kosmicheskie
ss
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from
Translated
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I
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and Pedersen, A.
K.
Knott,
R.
Charging Effects," Space Science Reviews
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Grard,
,
,
,
,
"Spacecraft
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34,
pp.
and E. G. Mullen, "Geosynchronous
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M.
S.
for
Severe
Spacecraft
Environment
Charging,"
Sapcecraft
and Rockets
v. 20,
Journal
of
pp. 26-34,
January-February 1983.
,
,
R.
0.
Bartlett, "Cesium Contact Ion
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R.
E,
Experiment
Aboard
Applications
Microthruster
Technology Satellite (ATS)-IV,"
Journal of Spacecraft
and Rockets v. 8, pp. 968-970, 1969.
,
,
S. ,
J. Staskus, D. Byers, "Preliminary Results
Jones,
II
Spacecraft Measurements Using Hot Wire
SERT
of
Emissive Probes," NASA TM X-2083 November 1970.
.
"Differential Charging of HighMande
The Equilibrium Potential
of
Voltage Spacecraft:
Research
Insulated Surfaces," Journal of Geophysical
v. 87, pp. 4533-4541, June 1982.
Katz,
I.,
M.
1
1
,
,
Kintner,
P.
M.
M.
C.
Kelley,
"Ion Beam Produced
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