Author(s)
Melvin, Michael Edward
Title
Design and evaluation of ion source for satellite charge control.
Publisher
Monterey, California. Naval Postgraduate School
Issue Date
1992
URL
http://hdl.handle.net/10945/23620
This document was downloaded on May 04, 2015 at 22:47:16
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THESIS
design and evaluation of ion source
for Satellite Charge control
by
Michael Edward Melvin
June 1992
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DESIGN AND EVALUATION OF ION SOURCE FOR SATELLITE CHARGE CONTROL
12
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Melvin
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Michael
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Master's Thesis
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The views expressed in this thesis are those of the author and do not reflect the
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ABSTRACT (Continue on
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Satellite Charging, Satellite Charge Control,
Differential Charging, Lithium Ion Source, Potassium
Ton Source
necessary and identify by block number)
This thesis describes the design of a new spacecraft charge control device that
incorporates a solid ion source made of Lithium or Potassium salt impregnated into a
porous tungsten plug. The ion source was configured with a reentrant thin-wall heat
shield to reduce heat loss and an experimental thermocouple imbedded in the plug to
accurately measure emission temperature. The initial design of the charge control
device included an extraction grid, deceleration grid, and an electron filament source
Experiments were conducted on the charge control device and results were used to modify
the design for optimization of current out of the device versus power used. Incremental
testing and modifications resulted in the deceleration grid being removed and the
extraction plate's wire mesh being removed to allow a clear path for the ions. With
these changes the requirement of 10 microamps was achieved at 13 watts with the
Potassium ion source.
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Design and Evaluation of Ion Source for
Satellite
Charge Control
by
Michael Edward Melvin
Commander, United States Navy
B.S., University of Colorado, 1979
Lieutenant
Submitted in partial fulfillment of the
requirements for the degree of
MASTER OF SCIENCE
IN
PHYSICS
from the
NAVAL POSTGRADUATE SCHOOL
June 1992
Woehler, Chairman,
Department of Physics
K. E.
ABSTRACT
This thesis describes the design of a
that
a
incorporates
impregnated
reentrant
into
source
ion
solid
new
made
spacecraft charge control device
of
Lithium
or
Potassium
a porous tungsten plug. The ion source was configured with a
thin-wall
heat
thermocouple
imbedded
temperature.
The
initial
shield
in
to
the
reduce
plug
to
heat
loss
and an experimental
measure
accurately
were conducted on the charge
and an electron filament source.
control device
and
results
Incremental testing and subsequent modifications resulted
removed and the
for the ions.
was achieved
at
extraction plate's wire
mesh
Experiments
were used
the design for optimization of current out of the device versus
a clear path
emission
design of the charge control device included an
extraction grid, deceleration grid,
grid being
salt
in
to
modify
power used.
the deceleration
being removed to allow
With these changes the requirement of 10 microamps
13 watts with the Potassium ion source.
in
/A
.
i
TABLE OF CONTENTS
I.
INTRODUCTION
1
II.
Spacecraft charging
3
A.
ANOMALIES
3
B.
Charge Development
4
1
Charging Equation
5
2.
Charging Currents
6
3.
C.
Ambient Plasma Currents
6
b.
Photoelectric Effect
7
c.
Secondary Electron Emission
8
Differential
Charging
RESULTS OF CHARGING
1.
ill.
a.
Timing
of
9
10
Discharges by Location
Spacecraft charging control
11
12
A.
passive Control
12
B.
ACTIVE CONTROL
13
C.
1.
Electron Emission
13
2.
Ion Emission
14
3.
Neutral
Plasma Emission
RESULTS FROM SATELLITES AND ROCKETS
1.
Space
2.
ECHO
Electric
Rocket Test (SERT)
15
16
16
17
IV
8
DUDLEY KNOX USBARY
hlAVAt POSTGRADUATE SCHOOt
MONTEREY GA §3043-5101
3.
Porcupine
4.
Auroral Rocket for Controlled Release
5.
Applied Technology Satellite (ATS)
V.
VI.
1
18
b.
ATS-5
18
c.
ATS-6
19
GEOS
and ISEE Series
Theory
B.
17
ATS-4
a.
A.
(ARCS)
a.
Spacecraft Charging at High Altitude
6.
IV.
17
(SCATHA)
Satellites
20
20
22
ION EMISSION
22
1.
Gas Discharge
22
2.
Surface Ionization
23
ELECTRON EMISSION
24
DESIGN OF THE DEVICE
26
SOURCE
26
A.
ION
B.
EXTRACTION GRIDS
28
C.
ELECTRON SOURCE
29
EXPERIMENT
30
A.
Vacuum Chamber
30
B.
Electrical Setup
31
C
experimental Procedure
31
1.
Initial
32
Design
a.
Power Sweep
b.
Extraction
Sweep
32
33
2.
c.
Deceleration
d.
Plate
e.
Comparison
f.
Results
Sweep
34
Sweep
35
of Deceleration Grid
Voltages
36
36
Removal Of Deceleration Grid
37
a.
Power Sweep
b.
Extraction
c.
Plate
d.
Results
39
3.
Grid Spacing Effects
39
4.
Modified Extraction Grid
40
5.
37
Sweep
38
Sweep
38
a.
Grid Spacing at 0.10 inches
40
b.
Grid Spacing at 0.31 inches
42
c.
Grid Spacing at 0.18 inches
42
d.
Comparison
of
Spacing with
New
Extraction
Grid
43
Electron Source
44
VII.
DISCUSSION
45
VIII.
CONCLUSION
49
APPENDIX A
51
APPENDIX B
52
APPENDIX C
95
A.
GRID TRANSPARENCY
95
B.
Grid Spacing
95
VI
C.
RESULTS
96
APPENDIX D
A.
LIST
99
THERMOCOUPLE
99
OF REFERENCES
INITIAL
101
104
DISTRIBUTION LIST
VII
OF FIGURES
LIST
Figure
Figure
1.
2.
Figure
3.
Figure
4.
between charging events and satellite
anomalies (McPherson and Schober, 1976)
52
Charging of a satellite by ambient plasma and
photoemission (Grard, etal, 1983)
53
Correlatiion
Illustration of current flow to
5.
Figure
6.
7.
(JPL Report,
1
989)
a satellite experiencing
charging (Grard, 1983)
Potential barrier
satellite
Figure
satellite
developed by
differential
55
charging on a
(Davis and Katz, 1989)
Emission of a neutral plasma
(JPL Report, 1989)
ATS-6 hollow cathode
56
to control differential
charging
57
ion engine (Olsen
and Whipple,
58
1978)
Figure
8.
SCATH A
Figure
9.
Solid state ion source (Spectr-Mat,
hollow cathode ion engine (Olsen, et
1
al, 1
990)
Electrical setup of charge control device in vacuum
chamber. Components of device: (a) thermocouple,
source,
(c)
source,
(f)
59
60
980)
Figure 10. Solid state ion source emitter current versus filament power(
Heinz and Reaves, 1968)
Figure 11.
54
Illustration of current flow to
differential
Figure
a
61
(b) ion
extraction grid, (d) deceleration grid, (e) electron
collection plate, (g) screen
Figure 12. Engineering drawing
of
62
charge control device (Bob
Berggren, Spectra-Mat, Inc)
63
Figure 13. Charge control device, side view
64
Figure 14. Charge control device, oblique view
65
VIII
Figure 15. Charge control device, top view
66
Figure 16. Lithium ion source
initial design
67
Figure 17. Lithium ion source
design
1,
1
,
2,
Figure 19. Lithium ion source
design
2,
1
,
current versus source power,
initial
current versus extraction voltage,
initial
71
Figure 22. Lithium ion source
initial design
1,
Figure 23. Lithium ion source
initial design
2,
Figure 26. Lithium ion source
current versus source temperature,
70
2,
Figure 25. Lithium ion source
design
initial
69
Figure 21. Lithium ion source
design
Figure 24. Lithium ion source
design
current versus source power,
68
Figure 18. Lithium ion source
initial design
Figure 20. Lithium ion source
design
current versus source temperature,
current versus extraction voltage,
initial
72
current versus deceleration voltage,
73
current versus deceleration voltage,
74
1
,
current versus plate voltage,
initial
75
2,
current versus plate voltage,
initial
76
1
,
comparison
of current
versus
extraction voltage for 3 different deceleration voltages,
initial
77
design
Figure 27. Lithium ion source
2,
comparison
of current
versus
extraction voltage for 2 different deceleration voltages,
initial
78
design
Figure 28. Lithium ion source
deceleration grid
2,
current versus ion source power,
79
removed
Figure 29. Potassium ion source, current versus ion source power,
deceleration grid
removed
80
IX
Figure 30. Lithium ion source
deceleration grid
2,
current versus extraction voltage,
removed
81
Figure 31. Potassium ion source, current versus extraction voltage,
deceleration grid
removed
Figure 32. Lithium ion source
deceleration grid
2,
82
current versus plate voltage,
removed
83
Figure 33. Potassium ion source, current versus plate voltage,
deceleration grid
removed
84
Figure 34. Potassium ion source, comparison of current versus ion
source power for different extraction grid spacings,
deceleration grid
removed
85
Figure 35. Potassium ion source, current versus ion source power with
new extraction grid at spacing of 0.10 inches
86
Figure 36. Potassium ion source, current versus ion source power
while varying voltages in tandem, new extraction grid at
spacing of 0.10 inches
87
Figure 37. Potassium ion source, current.versus ion source power with
new extraction grid at spacing of 0.31 inches
88
Figure 38. Potassium ion source, current versus extraction voltage,
new extraction grid at spacing of 0.31 inches
89
Figure 39. Potassium ion source, current versus ion source power with
new extraction grid at spacing of 0.18 inches
90
Figure 40. Potassium ion source, current versus ion source power
while varying voltages in tandem, new extraction grid at
spacing of 0.18 inches
91
Figure 41. Potassium ion source,
for
new
total
current versus ion source
power
extraction grid at 3 different spacings
92
Figure 42. Potassium ion source, current out of the charge control
device versus ion source power for new extraction grid at 3
different spacings
93
Figure 43. Electron source, current versus filament power
94
Figure C1. Current versus extraction voltage
for different extraction grid
97
transparencies
Figure C2. Screen current versus extraction voltage
extraction grid spacings
for
2 different
Figure D1. Calibration of thermocouple, thermocouple voltage versus
source temperature
XI
98
100
I.
Spacecraft charging
charged
a natural
is
and
particles
Introduction
high
energy
man-made
photons
Numerous
natural
and
phenomena
but
primarily determined
it
is
a
result of the interaction of
sources
space
the
in
contribute
by the collection
environment.
the
to
of
satellite with
charging
ambient ions and
electrons and the emission of photoelectrons and secondary electrons from the
satellite surface.
Typical values for spacecraft potential relative to the ambient
plasma can exceed
-1
kV
in
sunlight
recorded on the ATS-6 spacecraft
measures
to control the
caused anomalous
electronic
and
potentials as high as -19
eclipse (Olsen, 1987).
in
satellite
behavior including problems with telemetry, spurious
to satellite surfaces,
Adams,
Additionally
and
the absence of
charge accumulation, the subsequent discharges have
commands, damage
(Nanevicz
In
kV have been
1980).
on
and even
satellite failure
spacecraft
designed
for
experimental work the buildup of charge can impact dramatically on results and
even
limit
the types and degree of experiments.
Control of large negative potentials has been partially achieved onboard
satellites
same
by emitting electrons which drives the spacecraft potential
result is
ion source.
differential
properties.
satellite
accomplished
for positively
However, since most
charge
If
forms
satellites are not
across
an electron source
charged
these
is
used
satellites with the
made
materials
to
of
in
use
The
of
an
of uniform material,
different
a
conductive
discharge a negatively charged
the charge on insulated sections of the satellite
the electron source and the difference
to zero.
will
not be modified by
charge between the sections
of the
satellite will
grow.
At
some
point this differential charge
will
cause an arc or
discharge which has been found to correlate with observed anomalies. (Olsen,
1985)
It
has been found that discharging both electrons and ions together
control the spacecraft potential
between the sections
and
at the
same
will
time reduce differential charging
Present methods to create a neutral
of the satellite.
A
plasma emission incorporate a heavy and bulky gas discharge system.
simpler and smaller ion and electron charge control device
solid ion emitter rather than
The purpose
of the
gas or
desired that uses a
liquid.
research described
control device that incorporates
is
in this
thesis
was
to
design a charge
improvements over the gas discharge system
presently used.
Once designed
vacuum chamber
for the
the charge control device
optimum modifications
that
out of the charge control device at the lowest power.
was
tested
in
a
would maximize the current
The
results found at
each
step of the experiment were then used to steer the design process towards
better designs
and
further testing.
The
and discussed and recommendations
results of the
for further
experiments are presented
improvements are given.
Spacecraft charging
II.
Early rockets
and
satellites
energy plasma environment.
launched
into
low earth
orbit
experienced a low
Basic (Langmuir) probe theory indicated the
ambient plasma currents would be dominated by electron current which would
up
build the spacecraft potential
energy plasma
to - -1.0 V.
at higher altitudes
it
was
As
later satellites
probed
into
higher
predicted that spacecraft potentials
would reach higher values. (Grard, 1983)
Subsequent
test
confirmed
flights
this
hypothesis
and
additionally
demonstrated that large potentials were most frequently observed on the
side
in
geosynchronous
Low
orbit.
negative potential on the order of
-1
earth orbit satellites
V
A.
in
were charged
to
a
with respect to the environment, while
large electrostatic potentials of the order of tens of
geosynchronous spacecraft
night-
kV have been measured on
the earth's magnetosphere. (Whipple, 1981)
ANOMALIES
As
satellite
flights
have become commonplace a pattern
behavior onboard the spacecraft has been observed.
defined as any behavior that
satellite,
its
unordered
is
equipment, loss
of data,
to
a few
components,
anomalies are
directly affects the
Anomalies can include
disruption
of
electronic
and numerous other unwanted behavior.
operation and success of a
for
measurements.
satellite
This anomalous behavior
anomalies
unordered or unexplained that
control, or experimental
commands
Satellite
anomalous
of
is
distracting at best but
satellite.
satellites.
Table
Because
1
lists
can be very serious
a summary of
of the potential for failure of
to the
some known
a
satellite or
cause
loss of control, the
of interest to the
Numerous
of
anomalies and
space community
studies of
all
correlation observed
A
time.
Figure
time
1
behavior at
between the anomaly and where the
shows a concentration
This
geosynchronous
is
of
its
a discharge.
orbit
of
indicates
1
soon
orbit or
in
In fact,
frame
time
Figure
satellites.
suggests that sudden changes
trigger
anomalous behavior
the
also
prominent during the night side
during
this
(Whipple,
satellite
some
1981).
time with a
satellite is in local
compilation of data from several geosynchronous satellites collected
frame.
may
have correlated the
of satellites
charge on the
of
have been affected by
satellites
has been
for years.
anomalous behavior
unwanted action with a buildup
Almost
their elimination or control
in
the 2000 to 1000 local
significant
that
in
charging
are
more
sunlight.
Data
anomalies
after entering
for
the electrical environment of the spacecraft
the
provides a change
movement
of the satellite
from night to day
environment that makes charging and
in
subsequent discharge possible.
B.
CHARGE DEVELOPMENT
Typical values of geosynchronous satellite potentials range from
kV
in eclipse,
satellite
must
and
and ~0
its
balance.
photoemission,
to -1
kV
in sunlight.
space environment
Important
plasma
is
a
-1
to -20
This potential that develops between a
result of the
contributors
bombardment,
of
the
charging currents which
charging
secondary
backscattering electrons, and other charging mechanisms.
equation
electron
are
emission,
1.
Charging equation
Many charge and
current sources contribute to the satellite
buildup of charge and eventual balance of currents.
The
overall
satellite's potential is
governed by the charging equation
=
'tot
'e
+
'/
+
'se
+
'si
+
+
'bsce
'/tv
+
'exp
+
Uth
)
where
= ambient electron current
= ambient
se
ion current
= secondary electron current
= secondary
ion current
bsce
= backscattered electron current
hv
= photoemission current
exp
= active current sources such as electron or
oth
ion
beam experiments
= other current source.
dV
At equilibrium fjr = 0), the total current
The most important
(l
tot )
is
zero.
plasma
factors are the net flux of ambient
the secondary emission of electrons, and photoelectric emission
Additional
currents
include
impacting electrons and ions.
by the movement
energy
Finally
to
(>
backscattered
More
of the satellite
10 keV
)
satellite
fluxes
to sunlight.
associated
with
subtle sources consist of current generated
across an ambient magnetic
field
and by high
electrons which deposit charge inside insulating surfaces.
any onboard components such as
exposed
electron
due
current,
ion thrusters or
induced current flows
surfaces with high potentials also contribute (Garrett, 1980).
The
rate of
charge
transfer, positive or negative, is
and the operating environment (Whipple, 1981).
characteristics of the satellite
depends on charge already
Specifically,
it
the
the design of the
satellite,
2.
satellite,
plays no
and by
general the equilibrium charge
role,
such as
in
local
regions where photoemission
positive. (Whipple,
The major
will
on the vehicle, the motion
magnetic and
not be zero.
eclipse, the equilibrium
of the higher flux of electrons to
be
residing
of
electric fields.
Charging Currents
In
will
dependent on the
is
charge
Where photoemission
will
be negative because
an uncharged surface compared
to ions.
In
the dominant process, the equilibrium charge
1981)
natural sources of high voltage potentials are discussed next:
ambient space plasma interaction with the
satellite,
secondary electron emission,
and photoelectric emission.
a)
Ambient Plasma Currents
One
of the underlying principles of
"quasi-neutrality".
a plasma
is
the assumption of
This charge neutrality requires that, on average, electron and
ion densities are generally equal.
However, assuming the simplest case
of
equal
temperature, the ion and electron velocities are then quite different. Given equal
temperature, the ion and electron thermal energies are equal and given by
equations
(2)
and
(3)
E» = j>KT
=
1
(2)
mp Vp2
2
" i^e^2
(3)
therefore
vp
\j
£
m - 43
p
(4,
assuming an H+ plasma. Then with equal charge and density
for
both ions and
electrons the higher electron velocity term changes the current density equation
J = qnv
(5)
Therefore the current density for electrons
for electrons.
is
about 43 times
higher than for ions which causes a net negative charge buildup on the
This negative potential increases on the
satellite
surface
satellite.
the repulsive force
until
on the incoming electrons produced by the electrons residing on the
satellite's
surface reduces the electron flux to a balance with the ion flux as
shown
in
Figure 2(a).
In
midnight to
temperature
the plasma sheet, the hotter plasma distributions are found
dawn
increases
Additionally, the
(DeForest and Mcllwain,
region
the
degree
satellite potential is also
to
and
velocity
1971).
density
current
in
the
As the plasma
also
increase.
which ambient plasma bombardment affects the
determined by the design and structure of the individual
satellite.
b) Photoelectric Effect
Photoemission
geosynchronous
(Grard,
et
al,
orbit
an important source
photoemission
Indeed,
1983).
photoemission current
is
will
at
is
to
plasma densities below -1000 e'/cm 3
dominate (outside L~2) (Olsen, 1989).
H Lyman-alpha)
is
In
,
the absence
develops a positive charge as
of sufficient
surface material and knock electrons free as
buildup of positive charge
At
the major current from the spacecraft
of differential charging the spacecraft surface
photons (mostly due
of current for satellites.
shown
energy
in
strike the satellite's
Figure 2(b).
The
actual
influenced by the ability of the photoemission
induced current of 10-100 microamps/m 2
to leave the satellite.
This can be
affected by the design of the satellite
and the formation
of potential barriers
near
the satellite surface.
Secondary Electron Emission
c)
Secondary emission
The impact
the magnetosphere.
an important charging concern
of electrons is
of
ambient plasma on a negatively charged
When a
spacecraft produces secondary electron emission.
spacecraft
it
may
then escape the spacecraft at an energy of about
This process can cause a spacecraft to charge positively
spite of the
supposed dominance
The
function
number
actual
of the
of the
one
for incident electron
than one at higher energies. The
plasma current can be
average energies
ratio of
related to the
of less than
ambient electrons resulting
in
is
for
1980).
-
The
in
plasma temperature. For
net positive current.
yield
to
ambient
will
exceed the
For temperatures above ~5
-90%
of the incident current is
the magnetosphere the incident
the absence of photoemission, equation (5) then
8
is
distributions with
a few keV, secondary electrons
In
a
1000 eV, and less
secondary emission current
by the secondary emission.
is
considerably less than the electron flux and therefore less important.
In
-
angle of incidence, and the
energies of 10
keV, the ambient electrons dominate though
compensated
eclipse
by the impact
of electrons that are emitted
incident electron's energy, the
typically greater than
in
1
ambient electron currents.
spacecraft material's emission characteristics (Garrett,
ion flux
particle hits the
loses energy and a portion of this energy can be used to "excite"
other electrons which
2 eV.
in
becomes
and
an electron energy
for
(6) yields
a current density
The nature
shown
Figure
in
Barriers
of
~7 microamps/m 2
actual
In
many
photoemission
is
On
the current
electron
and
equilibrium
In
is
to
ion
flow
is
less
than
and
ions.
charge negatively when simple
incident
purely conductive surfaces
As shown
in
in sunlight,
Figure 4(a) the plasma ion
on both the shadowed and illuminated
is
added
to the current balance.
currents the satellite charges to a positive potential
As
until
reached.
most
satellites are not built with purely
made
Even
conductive surfaces
of different materials (probes, solar panels).
and conducting surfaces the current flow
of insulating
distributed over the sunlit
can charge
current
satellite is
photoemission exceeds the difference between the ambient
but consist of sections
charging.
and from the
should buildup positive charge.
the dominant current.
practice
mosaic
it
the sunlit side photoemission
due
the
satellite to
satellite with
and electron currents are
surfaces.
.
Charging
For a high altitude
this
practice,
can be created which cause a
Differential
,
factors affect the actual flow of electrons
current considerations indicate
3.
density of n=10 6/m 3 equation
of the various current flows to
3.
straightforward as
KTe = 10 4 eV and a
of
and the dark sides
spin-stabilized satellites
of the vehicle
causing
is
With
unevenly
differential
have permanently shadowed areas that
differentially.
When
a
satellite is
configured with conductive and insulating surfaces the
charging problem becomes quite complicated. As can be seen
charging equation on the sunlit side of the
However on the shadow
satellite is
in
Figure 4(b) the
unchanged from
before.
side of the satellite, unlike before, the incident plasma
electron current
is
unable to conduct across to the
by the photoemission current.
The
negative potential on the dark side.
potential barrier
potential barrier
the
calculate
seen
in
Figure 5
experiencing differential charging.
which
is at -1
barrier
is
kV
charge
will
a highly
to
show a
An example
to the sunlit side.
The
of this
which a computer model was used to
in
on
photoemission
of
effect
surface
and be balanced
Solutions of Laplace's equations
must form adjacent
is
satellite
sunlit side
a
simulated
spherical
satellite
sunlit side is the left side of the satellite
The
with the rest of the satellite at -5 kV.
such that photoelectrons are not allowed
and subsequently the whole spacecraft
will
to
effect of this potential
escape from the
charge negatively,
in
sunlit side
spite of the
nominally large photocurrent.(Grard, etal, 1983)
C.
Results of Charging
The
build-up of charge on a satellite
discharge.
The charge by
itself
has
is
a benign event
little
effect
on a
in
the absence of a
satellite
except for
experimental satellites attempting to conduct low energy particle measurements.
However, as the potential difference between parts
breakdown threshold
,a
discharge or arcing
will
of
a
satellite
exceed a
occur across these components.
Typically potential differences on the order of
500 V are needed
to
produce
discharges that are significant to an operating system (JPL Report, 1989).
this
voltage differential
behavior on
satellites.
and subsequent discharge
All
•
spurious electronic switch activity of components
•
breakdown
thermal coating
10
is
causes anomalous
discharges don't always cause anomalies.
even a weak discharge can cause
of vehicle
that
It
However,
and solar
•
amplifier
•
optical
•
unplanned
•
unplanned downlinking
1.
degradation
cell
sensor degradation
orbital
maneuvers
of telemetry
Timing of discharges by location
Different satellites
will
environment depending on
to the
charge
to different potential levels in the
their surface materials, size,
sun (Gussenhoven and Mullen, 1983).
different levels
During
as magnetic
shape, and orientation
Also, satellites
will
charge
to
activity varies.
magnetic substorms, the removal
photoelectric current
same
caused by eclipse passage
shifts in satellite potential (Purvis, et
al,
is
and reinstatement
observed
to result in
At geosynchronous altitude the
1983).
eV and as
sunlight the vehicle charges
a negative potential roughly equal
to
the
dramatic
electron temperature can reach 10 4
up
of
the spacecraft traverses out of
to the
electron temperature, (KT/e) = 10 kV.
The dynamics
of
discharge have been observed to correlate with the
build-up of spacecraft potential.
As was seen
in
is
also dependent on orbit location.
discharge
is
more common between 0400 and 0600
the quiet time injection events and the preferred
Early evening yields the
minimum
(Deforest, 1972)
11
1
the timing of the
During normal magnetic
discharge
(East).
Figure
local.
drift
This
may be due
for injected
probability of
activity,
to
electrons
a discharge event.
ill.
The stimulus
Spacecraft charging control
for
controlling spacecraft potentials
and
on low-energy
particle
concern related
to the ability to
Ejection of an electron
was
beam
and
eventually
The
experiments (Whipple, 1981).
use electron beams
for experimental
purposes.
drives the spacecraft to a positive potential
As the problem
understood the control
changing
for
the interfering effect of the satellite charge
electric field
thereby reduces the energy of the
the electrons.
methods
investigating
beam and can
of satellite
of spacecraft
charging
and
prevent the further escape of
discharges have
was deemed
become
better
important to control
the occurrence of satellite anomalies associated with charging.
Spacecraft charging control
is
accomplished
in
two basic ways:
passive techniques which involve the design of the
configuration,
emitters to
and
2) active
satellite's
1)
with
materials
and
techniques which involve the use of charged particle
vary the charge
of
the
spacecraft and
its
distribution
on the
spacecraft (Whipple, 1981).
A. PASSIVE
CONTROL
The simplest method
to control spacecraft charging is to
employ proper
design techniques that modify the characteristics of the charging equation term
that
is
causing the charging.
In addition,
the judicial use of conducting surfaces
wherever possible and proper grounding techniques
will
significantly
reduce
differential charging.
In
practice complete elimination of insulators
satellite
is
practically impossible
as any
requires certain isolated elements from the satellite ground such as
12
antennas, particle collectors, and solar arrays.
carefully select satellite materials that
have high secondary and photo-emission
properties to reduce negative charging.
cavities that contribute to
shadows
actions that have been tried
to provide
is
Additionally the design should avoid
that accentuate differential charging.
used on the
)
entire satellite.
GEOS
Unfortunately the procedure
However, not
all
negative charging
charge control on some
To
satellites
was
NASA
nonconducting
satellite potential
was
successfully
quite expensive (Grard, et
eliminated requiring
1983).
al,
some form
of active
using conductive coatings.
assist the satellite designer
computer program, the
was
When
series of geosynchronous satellites
were coated with conducting indium oxide the
reduced.
Further
the coating of insulators with a conductive coating
a conduction path across the
surfaces (solar panels
Therefore the designer should
and experimenter,
Charging
NASA
has developed a
Analyzer Program
(NASCAP),
to
evaluate a design for possible charging sites.
B.
ACTIVE CONTROL
Another method
to
modify and control spacecraft potential
spacecraft with a plasma source that
creates
new
currents
in
artificially
and ions
will
to configure the
enhances ambient
the equilibrium charging balance.
emitting a neutral cloud of electrons
is
Utilizing
fluxes or
a source
effectively increase the
charge
density around the satellite and equalize the currents to the satellite body and
surfaces.
1.
Electron Emission
It
has been suggested
that the large variations in satellite potentials
during eclipse passage could be eliminated by finding a suitable replacement for
the photoelectron current.
The operation
13
of
an electron source emitting a
beam
of electrons
of the
would be an
effective
balance
Though
photoemission current.
charging equation for the loss
to the
it
may seem
that adding
an electron
emitter would produce a current that would effectively reduce large negative
potentials the
spacecraft's
resulting
and
conducting
The emission
potential.
be undesirable.
effect could
insulating
of electrons
surfaces
will
a hot plasma, the
In
charge
to
a negative
from the spacecraft frame can create large
charging between the insulators which are at the unchanged plasma
differential
ground and the conducting surfaces
(Davis and Katz,
1989).
at the
new
elevated spacecraft potential
This sudden increase of differential charging can
accelerate or accentuate hazardous arcs.
As was seen previously
Figure 5 illustrates the
differential charging,
caused by the emission
with
most surfaces
electron emitter on
kV.
The
with photoemission
of electrons.
at -5 kV.
one side
figure
is
satellite
result of
experiencing
a potential barrier
a simulated spherical
The NASCAP program was used
of the satellite driving the
to
satellite
model an
conducting surface to
-1
potential barrier or saddle point in front of the conducting surface
prevents electrons from leaving the
2.
The
same
on a
satellite.
Ion Emission
Ion emission
can be used
to
reduce positive potentials, or induce a
negative charge on a satellite (Werner, 1988).
charged
potential
positively
back
be a return
charged
an
ion emitter
to zero.
In
periods
would be a useful method
when
to
the satellite
reduce the
is
satellite
For a low energy ion source, an additional effect would
flux of ions to the
spacecraft to discharge the insulators which have
differentially.
14
Neutral
3.
Plasma Emission
Neutral plasma sources provide the necessary currents to control and
vary the
satellite
ATS-5
frame
While electron emission reduced potential on
potential.
from several thousand
to several
hundred
plasma sources on ATS-6 maintained the
satellite at
a near zero potential
the
satellite
observed plasma conditions, both
in
sunlight
and
volts negative, the
eclipse. (Purvis
and
for
all
Bartlett,
1980)
Experiments and spacecraft data show that a plasma source designed
control spacecraft charging
must provide a
to
sufficient current of thermal ions to
hold the insulated surfaces at spacecraft ground and a sufficient current of
thermal electrons to vary the spacecraft potential. The combination of electrons
and ions being discharged by the
satellite is
shown
low energy ions that return to the
satellite to
discharge the insulated surfaces
paramount
to
effective
control
electron currents of about 10
of satellite
in
Figure
The
addition of
is
Modeling shows that
potentials.
microamps should be
6.
sufficient to control the
spacecraft potential (effectively replacing photoemission), with similar levels of
ion current required to discharge the insulators.
potential
and the most optimum discharging
The
control of spacecraft
of insulators is
accomplished by
biasing the plasma source relative to the spacecraft potential. (Olsen, 1981)
Additional research
is
needed
to optimize the ion
source with the goal of
reducing power requirements, weight of the fuel and possible contamination of
the spacecraft by the emitted material.
method
at
all
to continuously emit
times (Purvis and
A
further step
an appropriate current
Bartlett,
1980).
15
would be
to devise
a
to control satellite potential
C.
RESULTS FROM SATELLITES AND ROCKETS
Spacecraft charging affects most satellites but
most
is
readily
plasma detectors such as those flown on the experimental
ATS-6, SCATHA, and ISEE. These
measure spacecraft
additionally configured with
ATS-5,
satellites
satellites carried special instrumentation to
and the ATS, ISEE, and
potential
observed by
SCATHA
onboard experiments that influenced
satellites
were
this potential.
Experimentation on the relationship between charging and active plasma
experimental
al
been
has
emission
conducted
satellites.
while creating
Early work on
ARAKS
,
sponsored by the
on
effects
rockets,
1.
1971).
al,
Air Force.
the
was
Additional
Electric
SERT
1
and the
project,
collected on the
Hess,
et
of active
EXCEDE
rockets
plasma emission
ECHO, PORCUPINE, and ARCS
Rocket Test (SERT)
a mercury
test of
flight
demonstrate thrust and beam neutralization.
SERT
2
was a
ion
beam
of ions that
neutralizer neutralized the
thruster to
satellite flight test of
the ion engine at low altitude orbit with limited instrumentation.
plasma bridge
and
satellites.
was a rocket-borne
successfully emitted a
rockets
work with poorly reported data
The best documentation
SCATHA
Space
sounding
using
beam experiments was done by
a French and Russian
satellites
ATS and
1969
auroras with an electron accelerator flown on a
artificial
sounding rocket (Hess, et
are the
since
The
thruster
The
escaped from the spacecraft.
beam and demonstrated
that
it
was
able to vary the spacecraft potential by varying the neutralized potential relative
to the spacecraft. (Olsen, 1980)
16
ECHO
2.
Winkler conducted an extensive sounding rocket program
beam
This included an
emission.
study the charging problem on a
effort to
sounding rocket. He found that the rocket does not charge
beam
emitted
in
space plasma or
3.
locally
generated plasma (Winkler, 1980).
Porcupine
project,
comprised of two rockets launched
was conducted by researchers from Germany,
Two
United States.
beam
in
1979,
France, the Soviet Union, and the
rockets were launched with multiple payloads of which one
configured with a 200
approximate 4
the
to the voltage of the
general but rather draws a substantial return current from the
The PORCUPINE
was
electron
in
amp beam.
eV Xenon
ion
beam
emitter capable of producing an
Detectors on the other sub-payloads determined that
carried a current across magnetic field lines after
it
propagated
at least
several meters from the source and a return current from the plasma
was
generated outside the beam. (Pollock, 1987)
4.
Auroral Rocket for Controlled Release (ARCS)
The ARCS
capable
of
level of
1
mA beam.
in
1980 with a single Ar+
At the initiation of the ion
payload charging was observed from a
in
ARCS
that
it
2 rocket
ion
was launched
carried two ion
the diagnostic payload.
and Ar+
was launched
-1
V
ion
gun
beam, evidence
pre-experiment level to
approximately -5 V. (Pollock, 1987)
The
ARCS
rocket
producing a 100
of transient
a
1
beam on
Little
beam
in
November 1982 and
differed from
generators on a separable payload from
information
is
the spacecraft potential.
17
available of the effects of the
He +
The
ARCS
payload with
3 rocket, flown
some
February 1985, was a
in
reflight of
the
ARCS
changed
modification including both ion generators
2
to only
Ar+ emission. Data collected during the numerous experiments indicate that the
sub-payload was charged to
at least
3
V
negative during operation of the ion
generator aimed perpendicular to the magnetic
5.
field.
Applied Technology Satellite (ATS)
a)
ATS-4
The ATS-4
low altitude
orbit
due
satellite
to
was launched August 1968
launch vehicle
failure.
thruster which operated successfully.
current of 100-400
The
satellite carried
beam
During
but only entered into a
a Cs + ion
operation the escaping
microamps nearly balanced the photoelectric and ambient
plasma bombardment current leaving the spacecraft during
sunlight.
During
eclipse the neutralizer emission current did not balance the charging equation
and drove the spacecraft
b)
to
V
a 100
negative potential. (Hunter, etal, 1969)
ATS-5
The Applied Technology
1969
into
a geosynchronous
orbit.
Satellite
The
(ATS-5) was launched
satellite
was
in
August
cylindrical with solar arrays
covering most of the exterior except for a bellyband containing instrumentation.
Located
at the
endpoints were cavities containing a mixture of conducting and
insulating surfaces.
The
satellite
carried experimental
onboard with a separate electron beam filament
The
filament neutralizers on the ion engines
electrons
in
an attempt
during eclipse.
to
discharge the
The operation
-1
cesium
for neutralizing the ion
were designed
to -10
kV
beam.
to emit thermal
potential of the satellite
of the electron emitter did
18
ion thrusters
reduce the large
was unable
negative potentials but
charge completely
to eliminate the spacecraft
(Whipple, 1981).
Olsen showed that because
100 V, less than
1% of the
charging
differential
was evidence
was
of the
of differential potentials
on the order
emitted electron current escaped the spacecraft.
sufficient to explain the equilibrium potentials
effect
differential
of
The
seen and
charging would have on spacecraft
configured with insulators. (Olsen, 1985)
C)
ATS-6
The ATS-6
1974.
for
It
satellite
was designed
keeping.
station
was launched
to carry
A
to
a geosynchronous
two ion thruster engines
hollow
cathode
plasma
orbit in
to test their
bridge
May
usefulness
neutralizer
was
incorporated to provide charge and current balance for the main ion beam.
Particle data from
ATS-6 showed
that the satellite
largest recorded potential to date, -19 kV,
when
charged
in
eclipse up to the
the plasma sources were off
(Olsen, 1987).
Considerable data was also obtained on the use of ion emission and
electron emission on both spacecraft potential
and
operation of the ion thruster and plasma neutralizer
in
differential charging.
The
various environments had
major effects on the spacecraft potential with respect to ambient plasma and on
surface differential charging.
The
large spacecraft potential
operation of either the ion engines or the neutralizer.
was reduced by
Differential
charging
was
eliminated by operation of the ion engine and reduced by operation of the
neutralizer
showed
when operated
that
neutral
in
ion
mode.
These
tests carried out
plasma emission could be used
19
to
control
on ATS-6
spacecraft
charging and did not create surface
differential
charging as electron emission
does.(Olsen, 1985)
Spacecraft Charging at High Altitude (SCATHA)
6.
SCATHA satellite was
The
The
procedures.
satellite
carried
including a mixture of particle detectors
detectors.
electron
Xe+
Additionally a
gun capable
of
50 eV
in
1979
conduct a complete study
to
near-geosynchronous altitudes and
of satellite charging effects at
control
launched
gun capable
ion
to 3
and magnetic
electric field
of
1
-
keV were configured
charge
packages
experimental
thirteen
and both
test
field
2 keV emission and an
for active
charge control
experiments.
Review
charging,
of
SCATHA
discharge,
data indicates a clear linkage between
and anomalous behavior (Koons,
et
al,
satellite
1988).
In
experiments with the electron gun, results similar to ATS-5 were observed when
the electron emission discharged the satellite
caused by
differential charging.
differentially
1988).
charged
satellite
Results from the
The
when
SCATHA
ion
until
of ion
limiting point
gun was very
was reached
effective in controlling
using a neutralized ion
beam
(Olsen, et
experiments indicate that not only was
possible to reduce large negative potentials but
satellite to either
a
it
was
a
al,
it
possible to charge the
negative or positive potentials by the appropriate combination
and electron beam currents (Whipple, 1981).
a)
CEOS and ISEE series satellites
Both the
GEOS
and ISEE series
1970's and were different than the
specifically
designed
to avoid the
satellites
ATS and SCATHA
problem
20
were launched
in
satellites in that
they were
of differential charging. Their
the late
surfaces
were made
entirely of, or
covered
in
conducting material such as indium oxide to
better facilitate the study of magnetospheric
plasmas (Norwood
Accordingly, almost no differential charging
spacecraft kept a positive potential
in sunlight.
et
al,
1988).
was experienced and
the
The exception occurred when the
spacecraft entered a relatively cold and dense electron environment where small
negative potentials were observed (Grard, 1983).
21
theory
IV.
Active control of spacecraft charging relies on the emission of plasma
The
containing electrons and ions.
known
for
many years and
principle of thermionic emission
the basis for any active emitter design that
is
deliver the required current for discharging spacecraft potentials.
can be accomplished by numerous methods
in particular,
meet the requirements
accomplished
in
of
various
has been
charge
ways
will
Ion emission,
which only a few
of
This technique of ion emission can be
control.
with two of these methods,
gas discharge and
surface ionization, discussed further.
A. ION EMISSION
Gas Discharge
1.
Gas discharge systems have been shown
tendencies
The system
for ion emission.
to
have several excellent
capable of long
is
life
high ion densities by an electrical discharge through a gas vapor.
1
to
10
mA
are easily obtained
cathode and the anode (Moore,
The gas discharge,
engine onboard the ATS-6
liquid
Cs
is
when a
is
satellite
arc.
applied between the heated
or hollow cathode emitter,
further heated
anode which causes an
is
Ion currents of
et al, 1983).
heated and vaporized
cathode tube
voltage
and produces
and
in
and a
is
was used
shown schematically
in
for the
Figure
7.
both the cathode and anode tubes.
potential
After the arc
is
is
The
The
applied across the cathode and
struck, the discharge
by passing a current through the ionized gas. The stream
22
ion
of
is
maintained
Cs+ vapor
is
then
accelerated out of the discharge chamber by the accelerating grids. (Moore,
1983)
A
for
similar hollow
cathode system was used onboard the
experimental purposes and
is illustrated in
propellent at nominal currents of 0.3, 1.0
Figure
and 2.0
Designed
8.
mA
SCATHA
it
satellite
to emit
Xenon
based on the same
is
technology as the ATS-6 ion engine. (Werner, 1988)
Though the hollow cathode system has
ion
streams
charge
it
has some severe drawbacks when
control.
currents
liquid or
gas
considered for
relatively
satellite
high
ion
Finally
satellite.
experience
operation of
due
system
is
heavy and voluminous
bottles are discharged the satellite could experience
problems with weight balance and
to the
electromagnetic
some science
stability,
especially for a spin-stabilized
arcing required for ionization the satellite can
which
interference
can
adversely
affect
the
instruments.
Surface Ionization
The technique
that
is
requires a high level of power, up to 20 watts, to achieve this steady
it
and as the
2.
it
Even though the hollow cathode emits
Additionally the entire hollow cathode
output.
excellent properties for producing
when an impure
for production of ions
material
evaporate (Cobine, 1958).
presence
of
is
used here
is
based on the theory
placed on a heated filament, positive ions
Additionally,
when
the filament
is
a vapor whose ions can escape from the metal
heated
in
will
the
of the filament
a
copious amount of ions are produced.
In
a study
of thermionic
emission of positive ions Blewett and Jones
observed that Lithium with a Beta-Eucryptite coating gave roughly twice the
emission as the next best mixture (Blewett and Jones, 1936).
23
It
has been shown
that current densities of
to
1
mA
5
per square centimeter of coated filament
surface can be drawn continuously (Johnson, 1962).
Numerous designs have been employed
concept but the
utilizing this
compact Lithium emitter described by Heinz and Reaves
experiments
is
most useful
control device (Heinz
The
is
shown
in
low energy
design of a low energy spacecraft charge
and Reaves, 1968). This emitter
by Spectra-Mat, Inc and
Mat documents
for the
for
Figure
9.
It
is
is
commercially produced
further described in Spectra-
as:
emitter consists of an indirectly heated, highly porous, tungsten
The molybdenum body
machined with a solid partition for complete
isolation between the emitter and the heater cavity.
The three rhenium
120°
support struts are brazed at a
spacing with a moly/ruthenium eutectic
in
hydrogen,
yielding
a ductile and versatile mounting tripod.
at 2100°C
The heater is a noninductive wound bifilar coil with heliarc welded rhenium
leads solidly potted into the body cavity. The high purity Al 2 3 potting mix is
plug into which the emitter material has been fused.
holding the tungsten plug
H2
fired at
is
1900(°C) which completely immobilizes the heater. The emitter
a specially prepared, extremely porous, tungsten disc with a density
30% (70% porosity) is heliarc welded to the moly body. (Spectra-Mat, Inc.
matrix,
of
1980)
When
this emitter is sufficiently
the negative potential helps ions
emitter
for
in
the presence of an electric
overcome the surface vapor pressure
and accelerates them outward.
versus filament power
B.
heated
The observed
a 0.6 inch diameter source
is
of the
emission current
total
shown
field,
in
Figure
1
0.
ELECTRON EMISSION
Electrons are emitted by surfaces at high temperature
thermionic emission.
a process called
Depending on the material properties and the temperature
of the surface the electrons are emitted
bombardment,
in
electric fields,
as a result
of electron
bombardment,
ion
chemical effects, or photoemission (Cobine, 1958).
24
It
has been found that electropositive metals such as thorium emit much
larger electron current than metals with larger
heated filament
of thorium in the
work functions.
presence of a very high
effective at producing large electron currents in the
a stronger
source
is
to
material than thorium
a
practical
mA
method
electric field is
very
range. Since tungsten
of fabricating
mix tungsten and thorium together to form a filament.
25
Therefore a
is
an electron
V.
design of the Device
Active spacecraft charge control has been experimentally tested by the use
of ion
engines and neutralizes onboard
satellites
and sounding
To date
rockets.
an operational active charge control device has not been flown though they are
being
built for
design
an extension
is
SCATHA
satellites.
limitations
and
the upcoming
NASA/POLAR and
of the successful ion
ESA/Cluster mission. The
engine results from the ATS-6 and
However, the hollow cathode technology used
and improvement
in its
NASA
to date
has
parameters would provide a more effective
feasible charge control device.
A. ION
The
SOURCE
ion
source of the active charge control device was the principal design
challenge with the other components designed around
certain design requirements
and other properties
source must emit a minimum
of
The
it.
ion
source has
that are desirable.
about 10 microamps
at
The
ion
no greater than 20 watts
power. Any improvement of these parameters, higher current or lower power,
highly desired.
Additionally the emitter should
provide control for the entire
As discussed
requirements but
at
life
previously,
a cost
desired.
The basic
the
of weight,
ion
life
to
of the satellite.
hollow
cathode
system
volume, and power.
that incorporates the output of the hollow
was
have a long operational
is
A
achieves
these
different ion
source
cathode but with lower requirements
source described by Heinz and Reaves and
commercially produced by Spectra-Mat, Inc was chosen.
It
is
a
1/4"
diameter
Lithium or Potassium impregnated tungsten plug and a slight variation of the
26
surface ionization emitter previously discussed and
the emitter to be used
is
parameters
lifetime
in
Figure
Note that
9.
shown.
1/2 the size
This ion source has been investigated
and
shown
for different
depth by Gant (1991) for current
in
impregnate material; Lithium, Cesium, and
Potassium. His results indicate that either Lithium or Potassium hold promise for
our purpose of achieving high current output
emitter tested by
achieved a
microamps
In
Gant produced currents
of
lifetime
at
93 hours.
of
The K
15 watts and achieved a
low power input.
at
10 microamps
at
The
1/4" Li
27 watts and
emitter produced currents of
lifetime of
10
44 hours. (Gant, 1991)
the design of an active charge control device each component; the ion
source, ion extraction system, and electron source must be optimized to provide
The basic
the required plasma at the least power.
shown
uses the ion source
(b)
as the foundation
extraction grid (c)
mounted above
it
with the emitter
and extraction
grid.
component
of the
now
Disregarding the electrical connections for
Figure 11.
in
of the
electron
charge control device and
is
is
the design
charge control device with the
and the deceleration
The
components
layout of the
mounted
grid (d)
source
(e)
is
mounted above the
in-line
the
final
grid
and
emitter system so as not to interfere with them.
Using
this basic
concept a
engineering drawing
in
flight
Figure 12.
prototype
This
is
was designed and
a side view
of the
device and includes 3 extra copper plates at the base to assist
of the various electrical leads.
inches
in
Figures
diameter and 2 inches
13 through
imbedded
in
15.
The
tall
in
actual charge control device
and
is
Additionally
shown
an
at
is
shown
charge control
the connection
measured
1
.6
3 different angles of view
in
experimental
thermocouple
the heater potting for accurate temperature measurements.
27
the
in
was
In
an attempt
to
reduce power requirements a new method of placing a
reentrant thin-wall heat shield around the ion source to reduce heat loss
The
attempted.
shield to prevent
is
emitter surface
space charge
designed with 4 leads, 2
for
unit.
and support
They are
plate are
situated
base plate as shown
B.
in
in
same
plane as the plate of the heat
effects from the heat shield
power leads
leads for the thermocouple imbedded
shield,
the
is in
in
was
to the
moly
The
itself.
source
heater and 2 other
bifilar
The
the heater potting.
ion
ion source, heat
welded together and must be replaced as a single
the charge control device as the fourth plate from the
Figures 12 through 15.
Extraction Grids
The
extraction grids provide for acceleration
once they are emitted. Each
and deceleration
grid is electrically isolated
and
its
of the ions
potential
can be
independently varied to provide for optimum extraction and then deceleration to
the required ion energy level.
effectiveness of slowing the ions
The design
The deceleration
down
to
added
to
test the
mesh and spacing was a
result of
grid
is
lower energy levels.
of the grid's apperture, wire
reviewing previous experiments with extraction grids (Rovang and Wilbur, 1982),
(Homa and
Wilbur, 1982),
done (Appendix C)
and
(Haskell, et
al,
1966) and from additional work
with the experimental setup that
was used by Gant.
spacing of the grids between each other and to the ion emitter
electric
field
needed
to
extract ions
and the transparency
will
The
influence the
of the
grid
will
determine the current through the grids and out to the plasma.
Both of the grids are similar and situated exactly
can be seen
device.
The
in
in-line with
each other as
Figures 12 through 15 as the top 2 plates of the charge control
grids are manufactured from
28
molybdenum
with
HT moly
wire
interspersed at 0.078 inches
in
the 1/4" inch diameter center hole.
between the acceleration and deceleration
grid
and the
ion emitter
grids
Spacing
and between the acceleration
can be varied by using various combinations
of
ceramic
spacers.
C.
ELECTRON SOURCE
The
electron source
is
an
integral part of the spacecraft
charge control
A
0.010
purpose.
The
device and provides for a neutral plasma discharge from the spacecraft.
inch diameter thoriated (1.5%) tungsten wire
filament
its
was
further treated with
a mixture
of
used
is
for this
barium carbonate from
RCA
to test
emission properties. The electron source design has 2 filaments that can be
interchanged
if
1
filament
was
to burn out.
The
filaments are positioned at the
top of the charge control device and above both extraction and deceleration
grids.
During operation the high energy electrons
the space plasma to assist
in
will boil off
discharging the spacecraft frame.
29
and discharge
to
VI.
The spacecraft charge
environment with
different
Experiment
control device
was
completed
by
with
a simulated space
power
supplies,
optimum design.
a copper mesh screen and
The experimental setup
by a variable power supply.
various
in
modifications to determine the
Space was simulated by a vacuum chamber
collection plate biased
tested
and
potential
current
is
measuring
equipment, and an optical pyrometer to accurately measure emitter temperature.
Vacuum chamber
A.
The vacuum system used
large
(22"
18"
by
tall
space environment consists
to simulate the
diameter)
cylindrical
glass
bell
jar
with
of
a
electrical
connections through vacuum feedthroughs on the bottom base-plate and the
glass top plate.
torr
by a combination
control device
In
The chamber was maintained
of turbo
was mounted
at
an experimental vacuum
The spacecraft charge
and mechanical pumps.
horizontally
in
of 10 7
the bell jar facing the cylindrical wall.
addition to the charge control device the experimental setup consisted of a
copper wire mesh screen wrapped around the
and
isolated from
was connected
10 inches
to
all
a
in front of
isolated from
all
other components.
BNC
feedthrough
internal sidewalls of the
Additionally
a 4 by 6 inch copper plate
at the top of the
chamber and positioned
and
electrically
was added
during the
the emitter face on the charge control device
other components.
experiment to provide a clean surface
The copper
plate
for current collection.
30
chamber
B.
ELECTRICAL SETUP
On
the spacecraft charge control device, the extraction and deceleration
grids are electrically isolated from both the ion
ceramic spacers allow the varying
effect of
are
Extra
emitters.
spacing to the emitter to measure the
of grid
spacing on plasma emission.
A power supply
grids,
and electron
is
assigned
to
each
emitter, both extraction
and the copper mesh screen and
shown schematically
copper mesh screen and
space and
in
collection plate.
Components labeled
Figure 11.
collection plate respectively
collect the emitted current.
Components
charge control device and are the electron
grid, ion emitter,
and the thermocouple
electrical circuits
(a)
and
and are used
(c)
through
(g)
(b)
are the
to simulate
make up
the
emitter, deceleration grid, extraction
respectively.
To ease power supply requirements
ground. As verified previously, the
The
and deceleration
same
all
voltages are applied relative to
results could
have been obtained by
varying the emitter voltage relative to the extraction grid which would be at
spacecraft ground on an actual spacecraft.
C.
Experimental Procedure
Experimental measurements were
made on
various configurations of the
spacecraft charge control device for both a Lithium and Potassium ion emitter.
Changes were made
to the
initial
design of the charge control device after
experimental results indicated modifications might improve the operation of the
charge control device.
operated
The
initial
design of the charge control device
at various configurations of
power and biasing
pertinent data that indicated design modifications
Additionally, since
more Lithium
ion
of the grids.
was
Only
were warranted are presented.
sources were available than Potassium for
31
testing, the majority of early
experiments were conducted with the Lithium ion
source.
1.
Initial
Design
The
design of the charge control device
initial
shown schematically
previously and
in
Figure 11
without any modifications
except for the addition of a copper collection plate for
extraction grid
was spaced 0.10 inches from
deceleration grid
was
as described
is
some
The
experiments.
the ion emitter face and the
positioned another 0.05 inches from the extraction grid.
This design
charge control device
was operated
for
with two different Lithium ion sources
comparison
of the ion
in
the
sources and verification of the
design.
a)
Power Sweep
an
After calibration of the thermocouple, using
power and temperature sweep was conducted on the
The
V,
extraction grid
was biased
to
and the screen was biased
approximately <50
eV
and being attracted
measured from the
emitted current
-100 V.
to
device
watts.
is
In
principle,
a surface 50
show
V
this
should give
negative with respect to the deceleration
the results of these sweeps.
extraction grid, the deceleration grid,
is
Lithium ion source.
-100 V, the deceleration grid was biased to -50
also plotted for comparison.
be seen that the emitter produces a
and 23
first
a
ions (emitted kinetic energy) leaving the deceleration grid
Figures 16 and 17
grid.
total
to
optical pyrometer,
However, the
total
32
is
microamps
at
is
The
and the screen.
Figures 16 and 17
current of 10
ion current that
only about 2 microamps.
In
The current
it
can
1100 °C
emitted from the charge control
A second
after
a
Lithium source
was
installed into the
range of experiments was conducted on the
full
were obtained. After the
similar results
first
and
Lithium source
series of experiments a 4 by 6 inch
first
copper plate was positioned 10 inches
charge control device
front of the
in
charge control device
The
emitter face to provide a clean collection point for current measurement.
was biased
extraction grid
the copper plate
A
shown
at
in
-150 V, the deceleration grid was biased to -100 V,
to
was biased
to -1
50 V, and the screen was biased
representative temperature and
power sweep
to -1
1050 °C and 24
With
watts.
this
out of the charge control device
screen and
is
is
chamber
the
sum
source
for this
Figures 18 and 19. This ion source emitted 10 microamps
5 V.
is
total current
configuration, however, the current
of the current to the plate
and the
approximately 6 microamps.
The
first
Lithium ion source
was unable
to
achieve 10 microamps
output current to the screen while the second Lithium ion source achieved 10
microamps out
power
level is
of the
much
charge control device
effect that the extraction grid potential
current the extraction grid
first
was
varied from
to
-200
V and
had on emitted
This
means
Lithium ion source, the deceleration grid
there
is
no further acceleration
of the
deceleration grid with a nominal kinetic energy of 100 eV.
the total current emitted rose exponentially
and then
steadily
until
was biased
As shown
1
to
100
leaving the
ions
in
the extraction grid
Figure 20
was -15 V
rose while the magnitude of the extraction voltage
33
was
the current
-100 V, the screen was biased to -100 V, and the ion source was heated to
°C.
this
Sweep
To measure the
For the
28 watts power. Unfortunately
too high.
b) Extraction
measured.
at
was
The screen
increased.
The excess
voltage reached -8 V.
microamps
current stabilized at 3
was
current
after the extraction
collected by the extraction grid on
the charge control device.
With the second Lithium ion source the only difference
the previous example
was biased
to -15 V.
rose exponentially
before.
The
was
that the plate
As can be seen
in
was biased
of the voltage
was
total
the screen
steadily as
collected on the extraction grid
was increased beyond 40
voltages significantly increase the
V and
was -40 V and then rose
majority of the excess current
as the magnitude
-150
setup from
Figure 21 the total current emitted again
extraction voltage
until
to
in
V.
Higher extraction
emitted current, but the current collected
by the grids on the charge control device collect the majority of any marginal
current increase.
c)
Deceleration
The
Sweep
next series of experiments
The deceleration
of the deceleration grid.
energy
of the emitted ions leaving the
The
voltage from
first
to
Lithium ion source
-200
V
was meant
grid
to
measure the influence
was designed
to
manipulate the
charge control device.
was setup
for
a sweep
of deceleration grid
with the extraction grid biased at -100 V, the screen
biased to -150 V, and the source heated to a temperature of
1
shows
total ion
that the deceleration voltage
of the source.
As the deceleration
does not determine the
grid voltage is
deceleration grid increases at practically the
extraction grid decreases.
The
100 °C. Figure 22
production
decreased the current
same
level
as the current
to the
to the
current out of the charge control device
is
approximately 5 microamps for negative deceleration voltages, but for potentials
34
> -20
V
little
be emitted
or no ion current leaves the source.
to the
screen with
The second
first
at least
This
means
20 eV energy.
Lithium ion source
was again configured
source except that the plate was biased to -150
biased to -25 V. As shown
in
same
Figure 23 the
V and
similarly
was one
experiments run on the ion source and the emitted current rose
is
was
general pattern can be seen
This experiment on the second Lithium source
several experimental runs. This
as the
the screen
except the currents measured are significantly lower than with the
source.
must
that the ions
first
of the
ion
first
significantly after
not considered important for this comparison
since the nature of the ratio of the currents
was
of interest
and not the
initial
amplitude.
d) Plate
Sweep
To measure the
sweep
of plate potential
The
sources.
installed
first
effect of
from
Lithium
to
source potential versus plasma potential a
-200
was heated
and the extraction
decrease
driven
in
more
at
the
1
the 2 Lithium ion
100 °C with the copper collection plate
-100 V, the deceleration grid biased to
Figure 24
shows the
total
emitted current
12 microamps while the plate current increased
extraction grid
and deceleration
grid current
in
step with a
as the plate voltage was
negative.
When
at
to
grid biased to
-50 V, and the screen biased to -50 V.
was constant
V was conducted on
a similar experiment was run on the second Lithium ion source
same temperature
with
the
extraction
grid
biased to
in
Figure 25.
Again
with plate current increasing as plate voltage
35
total
the
V
similar results
emitted current
was constant
deceleration grid biased to -100 V, and the screen biased to -15
were obtained as shown
-150 V,
was decreased
to
-200 V.
Comparison of Deceleration Grid Voltages
e)
To determine
extraction of ions
the effect of the deceleration grid potential on the
a series
were conducted with
source was heated
sweeps
of extraction grid
different deceleration voltages applied.
The
1100 °C with the screen biased
-100
to
deceleration grid potential biased at -10, -50, and -100
was
varied from
sweeps.
The
to
total
-200 V.
Figure 26
sources
of the 2 Lithium
V
at
first
Lithium
V and
the
while the extraction grid
shows the comparison
of these
3
emitted current and the current leaving the charge control
device increased as deceleration voltage and extraction voltage decreased.
However the increase between -50 V and -100 V on the deceleration
grid is very
small.
The same setup was used
with the 2
source except that the plate was biased to -150
25
V.
sweeps
V and
the screen
Again Figure 27 shows that the results are similar
source with
total
current emitted
of the
second Lithium
was biased
to
-
to the first Lithium ion
and the current leaving the charge
control
device (plate and screen current) both increasing as the extraction voltage
decreases.
f)
Results
The
design charge control
initial
device
was configured
with
a
deceleration grid to control the energy level of the ions leaving the spacecraft.
However
energy
the results of these experiments indicate that controlling the kinetic
of the emitted ions
the net ion emission
in
would be
difficult
most cases. Since the goal was
current out of the charge control device at the
collection of
and the deceleration
a large percentage
minimum
of the current
36
to
grid
reduced
reach 10 microamps
possible
power
level,
the
by the deceleration grid was
deemed a
hindrance and the advantage of having control of the emitted ion
energy was unwarranted.
The
was
deceleration grid
removed
therefore
for
further testing.
2.
Removal Of Deceleration Grid
A series
new potassium
grid
A
on the
direct
of
experiments were run on the second Lithium ion source and a
ion source to
total
gauge the
effect of the
removal of the deceleration
emission and the emission that leaves the charge control device.
comparison between the charge control device with deceleration
and then without can be made using the second Lithium
a)
grid
ion source.
Power Sweep
The second
microamps
Lithium ion source with deceleration grid reached 10
total current at
about 23 watts as shown previously
in
Figure 19 and
the total emission does not change upon removal of the deceleration grid.
Lithium source
was setup as
for the
previous power
sweep
(section
The
l,a)
discussed with the extraction grid biased to -150 V, the plate biased to -150 V,
and the screen biased
to -15 V.
As can be seen
emits 10 microamps at about 23 watts.
in
Figure 28, the source
However, a comparison
still
of Figures
19
and 28 shows
that without the deceleration grid the current from the extraction
grid increases
by the amount that the deceleration grid contributed previously.
The
current out of the charge control device increases with
power
until
it
steadies at about 7 microamps at 24 watts.
The Potassium
ion source
and a power sweep was conducted
was
installed in the
charge control device
with the extraction grid biased to -100 V, the
plate biased to -110 V,
and the screen biased
power sweep are shown
in
to
-100 V.
The
results of the
Figure 29 and differ significantly from the Lithium ion
37
source. Almost
all
by the extraction
at
of the total emitted current is
The
grid.
concentrated
in
current collected
current out of the charge control device steadies out
approximately 3 microamps at 17 watts power.
b) Extraction
An
extraction
the plate biased to
grid
Sweep
50
-1
sweep
V and
of the
charge control device was conducted with
the screen biased to
-1
5 V. With the deceleration
removed the second Lithium source had an increase
extraction grid current increased to 10
extraction voltage
decrease
was decreased
to
microamps then
-200
V as shown
of total current but
decreased as
steadily
in
Figure 30.
With
this
extraction voltage the current out of the charge control device
in
increased steadily.
Compared
to the
results with the
second Lithium source showed an increase
extraction current as the voltage
and
was
deceleration grid
of total current
varied to -200
V
in
place, the
and an increase
as shown
in
of
Figures 21
Current out of the charge control device steadied out at about 8
30.
microamps
after -40
V
extraction voltage.
The Potassium
as the magnitude
was almost
totally
Figure 31.
The
ion source also exhibited
of extraction voltage
due
to
an increase
was
in
an increase
increased.
of total current
However the increase
extraction current as can be
seen
in
current out of the charge control device remained constant at
about 5 microamps throughout the power sweep.
c) Plate
A
-150
V and
operated
at
Sweep
plate
sweep was conducted
the screen biased to -15 V.
1
100 °C.
Figure 32
shows
with the extraction voltage biased to
The second Lithium
that while total current
38
ion source
was
remained steady
at
18 microamps, the current moved from the extraction
was adjusted from
plate voltage
to
-100 V.
experiment with the deceleration grid
was
extraction current
what
is
wanted
not as great as
for application
on
in
as the
grid to the plate
While similar results occur
the
in
place the gradient of the decrease
shown
in
differentially
Figure 25. This result
charged
is
in
roughly
satellites.
With the Potassium ion source very different results were obtained as
all
The
currents remain essentially unchanged.
microamps
with the extraction grid current
as shown
total
in
total current is
steady
encompassing 65 microamps
at
70
of the
Figure 33.
d) Results
With the removal of the deceleration grid
more current
is
was
it
not entirely clear that
being emitted by the charge control device.
It
appears that the
majority of the current previously collected by the deceleration grid
collected by the extraction grid, at least for Lithium.
out
of
the
To improve
is
now
the net current
charge control device a way must be found to increase the
transparency of the extraction grid while maintaining the electric
extraction of ions.
field
Before pursuing this problem the effect of grid spacing
for
was
researched.
3.
Grid Spacing Effects
To measure
the
influence
that
extraction
extraction of ions the Potassium ion source
set at 0.05, 0.18,
face.
The
was
varied
V, the plate biased to -110 V,
grid
set
and 0.31 inches distance between
ion source
changing the
was
in
power
grid
in
39
has on the
up with the extraction
grid
mesh and
grid
ion emitter
with the extraction grid biased to -100
and the screen biased
spacing can be seen
spacing
to
Figure 34 which
-100 V.
The
shows the
result of
plot of total
6
current emitted
and current out
grid spacings.
As the
total current
plot
of the
shows,
and current out
of the
for
charge control device
power
settings
up
to
for the
3 different
about 12 watts the
charge control device have similar gradients
but above 12 watts the currents steady out at significantly different current levels.
The
closer that the extraction grid
is
positioned to the emitter the greater
the total current emitted but the less net current emitted from the charge control
device.
At a spacing of 0.18 inches
and above 12 watts the
total current is
between the 0.05 and 0.31 inch currents but the current emitted from the charge
control device
maximized and
is
is
a greater percentage
of the total current than
at the other spacings.
4.
Modified Extraction Grid
As a
result of the large
extraction grid a modification
percentage of
was attempted
total
to
reduce
the current out of the charge control device.
current
was being
collected by the wire
extraction grid previously described, a
not have any wire mesh,
7/1
H
.
new
current being collected by the
mesh
It
this current
was thought
of the grid
extraction grid
so
was
and increase
that too
in
place of the
fabricated that did
and the center hole diameter was increased by 50
Even though the wires helped provide a more even
appears that the net current improved. The new extraction
on the charge control device without the deceleration
grid.
much
grid
% to
electric field,
it
was configured
The Potassium
ion
source was used and the spacing of the extraction grid from the emitter face was
adjusted for 3 different spacings; 0.10, 0.18, and 0.31 inches.
a) Grid
Spacing at 0.10 inches
The new
biased to -140
V
grid
was spaced 0.10 inches from
with the plate biased to -150
40
V and
the emitter and
was
the screen biased to -140 V.
A power sweep was conducted and
power
to the ion
source
is
until
microamps. Current out
13 watts which
is
a
in
As the
Figure 35.
increased the extraction current remains at zero
However the
12 watts and then rises slowly.
device rises steadily
the results are plotted
current out of the charge control
15 watts and then becomes constant
of the
significant
until
at
about 13
charge control device reaches 10 microamps
at
improvement over the previous designs with the
old grid.
When compared
to previous results of the
the original extraction grid installed as plotted
in
Figure 29 the current out of the
charge control device has increased dramatically.
extraction grid has
dropped
significantly
setting
has improved. Note
a
misleading since Figure 29
little
Potassium source.
and the
Potassium ion source with
The
total
current collected by the
same power
current at the
that the difference in overall current levels could
is
based on a very early sweep
be
of the
Ion emission from these sources increased gradually over
time.
This success motivated a look at
desired
mode
charged
satellite surfaces.
was conducted
screen
in
of satellite
operation; emitting
A
the device would behave
same
potential
was biased 10 V below them
the ion source.
The
in total
and the
was powered
sweep are shown
in
and the
grid (satellite
plate (a differentially
while the source
results of the combination
indicates an increase
the
combination sweep of the charge control device
tandem. The screen (space plasma) and the extraction
The
in
ions to space or differentially
that varied the potential of the extraction grid, the plate,
ground) were biased to the
surface)
how
at
charged
1 1
watts.
Figure 36 which
current as the potentials are decreased relative to
current out of the charge control device increased steadily
41
and
is
a larger percentage
of total current
Surprisingly the extraction current rises
and then
falls off to
b) Grid
was biased
to
V
in
was
extraction grid
biased at -50
In this
series of
sweeps however the
extraction
-100 V, the plate was biased to -110 V, and the screen was
A sweep
the
Figure 37.
power
to the ion
of current
source was increased and the results are
As previously seen, the
majority of the current
is
emitted out of the charge control device except at higher power settings.
extraction current
at this
is
makes up a much smaller percentage
spacing than was seen at 0.10 inches but the
also reduced from 10 microamps to 3.5
To
of the
test the influence of the
new
at
the Potassium ion source
38 the extraction
was
14 watts.
extraction grid potential on current out
set to 15 watts power.
grid potential
produced and the current out
is
biased to -150
has an almost
of the
V
c)
Grid Spacing at
0.
to -1
50
As can be seen
linear effect
on the
charge control device.
the current collected by the
does not increase very much beyond
The
emitted overall
total current
microamps
now
of total emitted current
charge control device the extraction grid was biased from
extraction grid
V
next configured at a spacing of 0.31
(midpoint settings for previous sequence).
was conducted as
plotted
is
Spacing at 0.31 Inches
inches from the ion emitter face.
biased to -100
the extraction grid
zero.
The modified
grid
until
as the potentials are decreased.
new
in
V while
Figure
total current
Even as the
extraction grid
zero.
18 inches
Since a spacing of 0.18 inches between the extraction grid and the
emitter face
was
previously found to maximize current out of the charge control
device with the old extraction grid the charge control device
42
was configured
with
.
the
new
grid
-140
was biased
V
to
-140 V, the plate biased
sweep
with the results of the
with the
power sweep. The new
extraction grid at this spacing for a
new
to
extraction
-150 V, and the screen biased to
As
plotted in Figure 39.
previous
in
was
extraction grid the majority of total current
sweeps
current that
was
able to leave the charge control device. The goal of 10 microamps current out of
the charge control device
increases more than
is
reached
was seen
at
at 0.31
The
14 watts.
extraction current
inches spacing and less than at 0.10
inches spacing.
A
combination sweep of the extraction
and screen was
grid, plate,
conducted as previously described with the ion source powered
current out of the charge control device increases linearly.
drops steadily and plate current steadies out
-120
is
at
-120 V.
The
Extraction current
increasing screen
the only contribution to current out of the charge control device after
V which
is
different
combination sweep shown
from the results found
in
at the
0.10 inches spacing
Figure 36.
Comparison of Spacing with New Extraction Grid
d)
The
total
current emitted
is
extraction grid from the emitter face.
increases as the extraction grid
movement
current.
15 watts. The
40 and show that as potentials are decreased the
results are plotted in Figure
current
at
is
of the extraction grid
As the
plot
shows the
affected significantly by the spacing of the
As shown
moved
in
closer to the emitter.
does not appear
current
Figure 41 the total current
However, the
to linearly affect the total
change from 0.10
to
0.18
is
not as
dramatic as from 0.1 8 to 0.31
Even more important than
the charge control device.
total
A comparison
43
current
is
the current emitted out of
of this current at the 3 different grid
spacings
is
shown
in
spacing linearly but
Figure 42. Again the current does not appear to follow the
falls off
sharply as grid distance from the emitter face
is
increased.
5.
Electron Source
The
The
electron filament source
electron source
was
initially
was
tested for current emitted to the screen.
brought up to a temperature of 2500 °C for 2
minutes to flash the thoriated tungsten filament. The results after flashing were
not as dramatic as expected so a
carbonate mixture
configured on the
in
Figure 43.
to increase thermionic emission.
initial
electron source
is in
The
results for this filament
charge control device and heated
The screen was biased
and the deceleration
grid
to
+50
was biased from
V, the plate
to
+100
to
power requirement
is
easily obtained.
(2.5 V, 6.4 A,
44
and 16
The
1700 °C are shown
was biased
The
V.
the milliamp range and the required
out of the charge control device
substantial
second filament was treated with a barium
1
to
V,
current off the
milliamp of current
only problem
watts).
+150
is
a
fairly
discussion
Vll.
The purpose
of the
experiments conducted was to optimize the charge
control device to achieve 10
microamps current out
The experiments were designed
at the lowest possible
to test incremental
power.
design modifications to the
charge control device and as results were analyzed the charge control device
was
modified and the next set of experiments were conducted.
Results of these
various experiments raise a variety of interesting points.
The
initial
design of the charge control device was tested using 2 different
A
Lithium ion sources.
felt
plate
was added
to the
a clean collection surface was needed
device to provide a clear electric
current.
Though
at the
same
and the goal
of
and extraction
for emitted
power
10 microamps out of the charge control device
The
current to the grids
was
of both Lithium
a percentage
grids took too large
of total
was
not
benefit of having the deceleration grid
available for precise control of emitted ions
much
case
was
charge control
and clearer measurements
for both emitters. In the
possible at a reasonable power.
diverting too
it
temperature, the emission of 10 microamps
same power
emitters the deceleration
current
directly in front of the
the 2 Lithium sources were found to require different
settings to reach the
found to be
field
experimental setup after
was
felt
to
be too costly
and the deceleration
grid
in
terms
of
was removed
from the design.
When
the deceleration grid
was removed
for
a series
Lithium and Potassium ion source the results were not what
current that
was
of
experiments on a
was expected. The
previously collected by the deceleration grid did not
45
all
go out
of
the charge control device, rather a sizable percentage
extraction grid.
appears that the removal
It
was
collected by the
of the deceleration grid affected the
electric field sufficiently that the majority of the ions that
used
to collect
on the
deceleration grid did not have the energy to exit the charge control device
Though the removal
collected on the extraction grid instead.
grid
improved the results
is insufficient
An
of the deceleration
charge control device the increase
achieve 10 microamps at low power.
to
interesting result
on the Potassium
total
of current out of the
and
was obtained
ion source.
with the removal of the deceleration grid
While the Potassium source emitted a much larger
current than the Lithium source the current emitted by the Potassium
source went predominantly to the extraction
control device with the
though
relative
is
of the
Lithium
depend
to
larger than the Lithium
Potassium atom encounters more
This result can be
and Potassium atoms.
some degree on
atom and
difficulty in
The
structure of the Beta-Eucryptite.
salt is
current out of the charge
reached 70 microamps.
masses
extraction of the ions appears to
Potassium atom
The
Potassium source never exceeded 5 microamps even
total current consistently
related to the
grid.
it
is
their
The
masses. The
possible that the larger
being extracted from the
lattice
extraction of ions through the Potassium
helped by higher extraction voltages.
This
is
an important
result
as
emission from the Potassium ion source can be significantly improved with
higher extraction voltages at no cost of higher power to the ion source.
The
between the extraction
variation of the spacing
grid
found to affect the current emitted.
The
the extraction grid and the emitter
determined by a factor of
moved
closer the electric
field
is
and the emitter
strength of the electric field formed by
1/d.
strengthened and ion extraction
46
is
As
is
the grid
is
increased.
However, the current out
As the
relationship.
current extracted
on the
of the
grid is
charge control device does not clearly follow
moved
closer a greater percentage of the
electric field at very close distances
plate than before.
The
collected by the extraction grid.
is
It
was found
that
must
a spacing
divert
was moved
When
the
more
of 0.18 inches
of the ions to the
maximized current
new
wire-less extraction grid
effects that
was
installed in the
the
of the wires
superseded much
formed by the new
and enlarging the center hole
of the previous results found.
grid diverged out of the emitter
charge control device and as the
grid
charge control
the charge control device
total current that exited
improved dramatically. The removal
electric field
if
closer.
device the percentage of
had dramatic
new
effect of the grid plate
out of the charge control device even though total current would be higher
grid
this
was moved
The
and through the
closer to the emitter an
increasingly greater percentage of the total current went to the extraction grid.
At a spacing of 0.10 inches the current collected by the extraction grid
significantly greater
total
than at the other spacings.
current increased
more
significantly
current out of the charge control device
spacings.
Therefore
the
charge
However,
at this
was
spacing the
over the other distances such that the
was
control
still
greater than at the other 2
device
configured
without
the
deceleration grid and with the extraction grid spaced at 0.10 inches from the
emitter face provides the greatest current out which can be used for charge
control.
The
electron
source exhibited the
purposes and was operated
for
required
some experiments
47
in
charge control
output
for
tandem
with the ion source
without any adverse effects.
be carburized
at
The
electron source requires a coating or
needs
manufacture to provide the necessary emission currents
active charge control.
48
to
for
VIII.
CONCLUSION
The experiments conducted on
its
suitability
the charge control device were to determine
charge control device design was modified
current out at 10
The
The
as a replacement spacecraft charge control device.
microamps
extraction
in
succeeding steps
to optimize the
low power.
at
and deceleration
grid
the extraction grid
was
was determined
design
inadequate given the purpose of maximizing current
was removed and
initial
modified.
out.
The
to
be
The deceleration
grid
design achieved
final
the required current out of the charge control device by manipulating the
and the spacing
extraction grid's center hole
the charge control device of 10 microamps
which
is
a
significant
improvement over our
hollow cathode designs.
Additional
A
to the emitter face.
was achieved
initial
at
current out of
a power
design and
is
of
13 watts
competitive with
improvements over other designs
is
the low
weight and volume of the charge control device and the elimination of possible
electromagnetic interference.
The problem
of lifetime
remains however, and further reductions
consumption should be possible with
by further improvements
on the extraction
grid.
in
An
method
The
lifetime
power
can be improved
the ratio of current emitted versus current collected
operational design for a charge control device would
be provided by mounting multiple
compact sources mounted
this device.
in
in
ion sources in
an array.
to replenish Lithium or
a matrix, with
A second
alternative
20-100
of the
would provide
a
Potassium by diffusion from the back end of the
emitter.
49
The
power
radiated
oT 4
(
)
of the ion
source
only accounts for about 1/2 of the power used.
is
therefore lost
in
the production of ions.
A
on the order
of 6 watts
The remaining power
and
7 watts
of
review of the charge control device
some recommendations on
design provides
is
reducing this loss factor.
The
ion
emitter appears to lose substantial heat out of the bottom of the emitter-heat
shield unit.
plug, the
a method could be devised
If
to hold this
heat
in,
power requirements should drop. Additional heat
the emitter could be reduced by the addition of
reduce the power requirements by 10
-
more heat
such as a ceramic
loss from the side of
shielding which could
20 %.
Additional improvements to the design of the charge control device would be
movement
of the
electron filaments
and
their
posts further
away from
the
This would preclude the emission from the electron filament
extraction grids.
from coating the ceramics on the filament posts and providing a conduction path
to the extraction grid.
The present method
further
improvement.
connection
is
of connecting leads to the various
For testing
purposes a
solid
but
required. Instead of the 3 copper plates at the
control device for connecting to the emitter
1
components needs
quick
base
method
of the
charge
plate separated into 3 isolated
sections would ease connections and lower the weight and volume further.
actual connection to the
attachment point
removal
of the
10 microamps
for emitter leads
would ease change-out
test
of the emitter
and
chamber.
control device with design modifications
at
For
leads from the emitter a quick connect tab with
charge control device from the
The charge
of
met the required goal
of
low power and further enhancements would improve the device
considerably.
50
)
Appendix
Table
Summary
of
A
1
some known anomalies (JPL
Report, 1989)
Anomaly
Satellite
Voyager
Power-on resets
1
SCATHA
34 Pulses detected
DSP
False flag from star sensor
Thermal control degradation
Sensor data noise
Control circuit switching
DSCS
Spin up
II
Power system
Clock
False
GPS
INTELSAT
III
and
IV
t
failure (#1
shift
command
Unexplained spin up
Skynet 2B
Telemetry problems
ANIK
Power downs
CTS
Short
circuit
power
noise bursts and
inverter
shutdown
Meteostat
Status changes
GEOS
Upsets and loss
Solar
4 and 5
Max
Navstar
Telesat
1
mission
of
GEOS
10 upsets/year
Solar array hold
mode
Telemetry logic switching
51
4
,
APPENDIX B
•«?*
BO HOtlil
|
».
local lime distribution of AIS-6 spacecraft charging
events.
II
V
M
II
II
DIP IOCIC UPJt tl
OVT.%
II
DO
T inllllll
O INFdJlT
uf!HI
IV
III
Occurrences of melliie opertiionil snomiliei ploned
orbit
Figure
1.
10
II
(McPherson tnd Sehober.
1976).
n
function of loctl lime in the geosntioniry
The ndul dimnce hit no tifnificince
Correlation between charging events and satellite anomalies
(McPherson and Schober, 1976).
52
SJPFACE
SURFACE
•CNS PEPfLLEO
SURFACE
~2ZSZD©
9
VMM m Pl*SM>
9 PL ASM a
AS"! ELECTRONS
'I
E'.E:*r;ss pepelis:
N£jAT".!
SjP c ACE
I
ty.V.WM"
PLAS M » iGNS
By Th{
POSHivE CMAQGE
9t
the
I
ELECTRO
ESCAPING
PHCTOEIECTPONS
charge
a
ions
SOLAB PHOTONS
b
I
P U 3T0E'.ET30NS
lTTCiCTia BACK BY THE
CKAP&E
SUPFACE
Qualitative illustration of the charging of
a
surface by a plasma.
currents collected and emitted by
a
surface element
is
The width of
the arrows
is
is
reached when the sum of the
zero, (a) Surface in
shadow: the current balance
proportional to the flux of each particle species; the equilibrium potential
requires equaliu between the flow of the plasma ions and that of the plasma electrons impinging on the
surface, (b) Surface in sunlight; equilibrium
to the difference
is
achieved w-hen the flow of escaping photoelectrons
between the incoming flows of plasma electrons and
Figure 2. Charging of a satellite by ambient
(Grard.efa/, 1983).
53
is
equal
ions.
plasma and photoemission
Plasma environment
Surface interaction
Sut face currents
Secondary
particles
Backscatlered
Figure
I.
electrons,
(after
In surface eliarpinp, current*:
and pliotoclectrons
from the movement of ambient
result in a net current
electron?;, ions,
on the external surface of
secondary
(lie satellite
body,
Robinson, 1989)
satellite (JPL Report, 1989).
Figure 3. Illustration of current flow to a
54
PLASMA
POTENTIAL
RAOIAL DISTANCE
Schematic representation of particle flous to and from a satellite Tor the ease of (a) a conductive
The lower portion gives a qualitative plot of the associated potential
surface and (b) an insulator surface.
profiles in a hot plasma.
Figure
4.
Illustration of current flow to a satellite
charging (Grard, 1983).
55
experiencing differential
12000
11
000
10 000
9 0000
8 0000
70000
eoooo
5 0000
Figure
5.
6 0000
7 0000
90000
B 0000
10000
11
000
Potential barrier developed by differential charging
satellite (Davis and Katz, 1989).
56
12 000
on a
XENON PLASMA
SPACE
PLASMA
np o
Figure
1
6.
cm
Emission of a neutral plasma to control
(JPL Report, 1989).
57
differential
charging
ATS-6 ION ENGINE
BOUNDARY ANODE RINGS
PLASMA GRID {*550V)
ACCELERATING GRID I-550V)
RING MAGNETS
DISCHARGE CHAMBER
TANTALUM COIL,
CATHODE
VAPORIZER
SHEATH HEATER
NEUTRALIZER
PROBE
CATHODE
FEED
SYSTEM
|
NEUTRALIZER
VAPORIZER
Figure
7.
ATS-6 hollow cathode ion engine (Olsen and Whipple, 1978).
58
NEUTRALIZE**
DECEL ELECTRODE
ACCEL ELECTRODE
SCREEN ELECTRODE
HOLLOW CATHODE
ANODE
KEEPER
ION
SOURCE_
POWER
LINES
XENON
EXPELLENT
nnnaa
LINE
DISCHARGE CHAMBER
n
VACUUM ENCLOSURE
SC4
-
2
ION GUN
Ion gun block diagram.
Figure
8.
SCATHA
hollow cathode ion engine (Olsen, et a/, 1990).
59
Porous tungsten containing
emitter material.
He liar c weld
1
.100
.002 thick moly heat shield
(3
o ^
,530
dimpled layers)
•AI2O3 heater potting
ox
.020 dia. moly bifilar heater
wrapped with .010 moly at
exit leads.
"£-
Re Support struts brazed to
moly body. 3 places at 120°.
500 Min
u-
.025 Dia
-.040 Dia
Standard 600 Ion Source.
Figure 9. Solid state ion source (Spectra-Mat, 1980).
60
900
1000
1100
EMITTER TEMPERATURE
38
1200
*C
35
76
FILAMENT POWER
votlt
103
Total emission current as a function of power and temperature:
Insert
shows experimental arrangement.
Figure 10. Solid state ion source emitter current versus filament
(Heinz and Reaves. 1968).
61
power
S
XXXX
o-
t
€>
•t
<u
J
I
<&
1
I
f
r
Figure 11. Electrical setup of charge control device in
vacuum chamber.
Components of device: (a) thermocouple, (b) ion source, (c)
extraction grid, (d) deceleration grid, (e) electron source, (f)
collection plate, (g) screen.
62
THORIATED TUNGSTEN FILAMENT
EMITTER SUPPORT
ASSEMBLY
MOUNTING
PLATE
ION
GUN ASSEMBLY
Figure 12. Engineering drawing of charge control device
Spectra-Mat.lnc).
63
(Bob Berggren,
Figure 13.
Charge control device, side view.
64
Figure 14.
Charge control device, oblique view.
65
Figure 15.
Charge control device, top view.
66
Lithium Source
1
Current versus Source Temperature
100 -r
10
--
CO
Q_
E
o
o
1
—
0.1
—
a?
0.01
850
900
950
1050
1000
1100
1150
1200
Source Temperature (C)
Figure 16. Lithium ion source
1,
current versus source temperature,
design.
67
initial
Lithium Source
1
Current versus Source Power
100
-r
10
—
CO
Q.
E
o
o
E
1
=3
0.1
—
0.01
10
12
14
16
18
20
22
24
26
Source Power (W)
Figure 17. Lithium ion source
1,
current versus source power,
design.
68
initial
28
Lithium Source 2
Current versus Source Temperature
100
10
—
1
—
CO
Q_
E
o
2
E
0.1
--
0.01
1000
1020
1060
1040
1080
1100
1120
1140
Source Temperature (C)
Figure 18. Lithium ion source
2,
current versus source temperature,
design.
69
initial
Lithium Source 2
Current versus Source Power
100 -r
10
--
CO
CL.
E
o
p
0.1
0.01
15
17
19
21
23
25
27
29
Source Power (W)
Figure 19. Lithium ion source 2, current versus source power,
design.
70
initial
Lithium Source
1
Current versus Extraction Voltage
12 -r
10
--
8
—
CO
E
o
2
o
E
6
4
—
2
-
50
150
100
Extraction Grid Voltage
Figure 20. Lithium ion source
1,
( -
V)
current versus extraction voltage,
design.
71
200
initial
Lithium Source 2
Current versus Extraction Voltage
30 -r
A
50
O
O
Q
6
Q
Q
C*—^(—C>-
150
100
Extraction grid voltage
( -
V)
Figure 21. Lithium ion source 2, current versus extraction voltage,
design.
72
200
initial
Lithium Source
1
Current vs Decel Voltage
-o
20
40
60
80
100
Deceleration Grid voltage
Figure 22. Lithium ion source
initial
1,
120
(
-
o
140
o
o
160
V)
current versus deceleration voltage,
design.
73
Lithium Source 2
Current versus Decel Voltage
20
40
80
60
Deceleration grid voltage
( -
100
V)
Figure 23. Lithium ion source 2, current versus deceleration voltage,
initial
design.
74
Lithium Source
1
Current versus Plate Voltage
CO
Q_
E
o
p
E
6
50
Plate Voltage
( -
V)
Figure 24. Lithium ion source 1, current versus plate voltage,
design.
75
200
150
100
initial
Lithium Source 2
Current versus Plate Voltage
30 -r
Plate voltage
Figure 25. Lithium ion source
2,
( -
V)
current versus plate voltage,
design.
76
200
150
100
50
initial
Lithium Source
Different Decel Voltage
50
1
Sweeps
100
Extraction Grid Voltage
150
( -
200
V)
Figure 26. Lithium ion source 1, comparison of current versus extraction
voltage for 3 different deceleration voltages, initial design.
77
Lithium Source 2
Different Decel Voltage
Sweeps
8 -r
50
150
100
Extraction Voltage
( -
200
V)
Figure 27. Lithium ion source 2, comparison of current versus extraction
voltage for 2 different deceleration voltages, initial design.
78
Lithium Source 2
Current versus Source Power
100 -r
10
--
CO
Q-
E
o
p
1
--
0.1
--
Z3
0.01
15
17
19
21
23
25
27
Source Power (W)
Figure 28. Lithium ion source 2, current versus ion source power,
deceleration grid removed.
79
29
Potassium Source
Current versus Source Power
100
T
10
--
CO
Q.
E
o
s
o
s>
«—
o
=3
1
--
0.1
12
17
22
27
Source Power (W)
Figure 29.
Potassium ion source, current versus ion source power,
deceleration grid removed.
80
Lithium Source 2
Current versus Extraction Voltage
18 -r
50
100
Extraction voltage
Figure 30. Lithium ion source
2,
150
(
-
V)
current versus extraction voltage,
deceleration grid removed.
81
200
Potassium Source
Current versus Extraction Voltage
70 -r
20
40
80
60
Extraction voltage
Figure 31.
120
100
( -
140
V)
Potassium ion source, current versus extraction voltage,
deceleration grid removed.
82
160
Lithium Source 2
Current versus Plate Voltage
-«
>
»
i«
>
200
100
Plate voltage
Figure 32. Lithium ion source
grid removed.
2,
( -
V)
current versus plate voltage, deceleration
83
Potassium Source
Current versus Plate Voltage
80 -r
60 --
-50
E
"
Extraction
-n
Plate
-
Screen
-O
Total
-*
Out
o
'E
40
<v
CJ>
30
20
--
10
--
—o^
-<t>
50
Plate voltage
Figure 33.
150
100
(
-
V)
Potassium ion source, current versus plate voltage,
deceleration grid removed.
84
200
Potassium Source
Comparison of Grid Spacing
1000 -r
100 --
10
--
CO
E
s
1
--
0.1
--
0.01
--
0.001
H
9
1
1
13
11
h15
17
19
21
Source Power (W)
Figure 34.
Potassium ion source, comparison of current versus ion source
power for different extraction grid spacings, deceleration grid
removed.
85
Potassium Source
Power Sweep
at 0.10 in
ZU ~
18 -
16
-
14
-
E
o
o
b
12
-
'E
10
-
8
-
6
-
Vi
a.
4 -
2 -
7.00
9.00
11.00
13.00
15.00
17.00
19.00
21.00
Source Power (W)
Figure 35.
Potassium ion source, current versus ion source power with
new extraction grid at spacing of 0.10 inches.
86
Potassium Source
Combination Sweep at 0.1
20
40
60
100
80
Plate Voltage
( -
in
120
140
160
V)
Figure 36. Potassium ion source, current versus ion source power while
varying voltages in tandem, new extraction grid at spacing of
0.10 inches.
87
Potassium Source
Power Sweep
14
16
at 0.31 in
18
20
Source Power (W)
Figure 37.
Potassium ion source, current versus ion source power with
new extraction grid at spacing of 0.31 inches.
88
Potassium Source
Extraction Voltage
Sweep
at 0.31 in
8 -r
20
40
60
80
Extraction Voltage
Figure 38.
120
100
(
-
140
160
V)
Potassium ion source, current versus extraction voltage, new
extraction grid at spacing of 0.31 inches.
89
Potassium Source
Power Sweep
at 0.18 in
16 -r
11
13
15
17
19
21
Source Power (W)
Figure 39.
Potassium ion source, current versus ion source power with
new extraction grid at spacing of 0.18 inches.
90
Potassium Source
Combination Sweep at 0.18
in
80
130
12 -r
10
--
CO
Q_
E
o
2
CJ
E
6
Si
O
4 --
2 --
60
70
90
100
110
Plate Voltage
Figure 40.
( -
120
140
150
V)
Potassium ion source, current versus ion source power while
varying voltages In tandem, new extraction grid at spacing of
0.18 inches.
91
Potassium Source
Comparison of Current Total
20
18
16
14
--
V)
§- 12
a
2
o
E
Total (0.18
in)
10
Total (0.31 in)
<w
o
Total (0.10
8
in)
6
4
2
10
15
20
25
Source Power (W)
Figure 41.
Potassium ion source,
for
new
versus ion source power
3 different spacings.
total current
extraction grid at
92
Potassium Source
Comparison
of Current
Out
14 -r
12 --
10
--
Out (0.18
CO
in)
CL
§
O
Out
8
(0.31 in)
Out (0.10
QJ
in)
6
4 --
2 --
10
15
20
25
Source Power (W)
Figure 42.
Potassium ion source, current out of the charge control device
versus ion source power for new extraction grid at 3 different
spacings.
93
Electron Source
Current versus Decel Voltage
10000
1000 --
CO
Q_
E
o
o
k_
o
'E
100 --
o
10
20
40
80
60
Decel Voltage
( -
V)
Figure 43. Electron source, current versus filament power.
94
100
.
APPENDIX C
To
the design of the charge control device to be used
facilitate
in
this
experiment, previous designs were investigated (Rovang and Wilbur, 1982),
(Homa and
Wilbur, 1982),
and
(Haskell, et
al,
These designs varied
1966).
in
the utilization of the ion source and used different methods to pull the ions off of
the emitter face.
our preliminary design the question
In
transparency of the
effect the type of grid, the
between the extraction
preliminary test
of the final
A. Grid
grid
was conducted
charge control device and the
biased to -200 V.
center hole.
On
was
A
in his
installed
grid
to finalize the
to optimize
this plate
and heated
a set
to
its
A
design
output.
for
with a
of crossing
grooves were machined. By placing
different
grooves a mesh pattern was formed.
in
sweep was then conducted
in
A
removable plate around the
of the
spacings between the wires giving various transparency
the results are plotted
these tests.
1100 °C while the screen was
and groove arrangement allowed the movement
extraction grid
B.
program
overall results.
experiments was modified
was manufactured
equal length stainless steel wire
different
test
actual spacing
Transparency
Lithium ion source
plate
answer these questions
to
raised as to what
and the
and the emitter face has on the
The device used by Gant
The
grid,
is
at
wires to
to the grid.
An
these different transparencies and
Figure C-1
Grid Spacing
A
set of 2 different sized ceramic spacers allowed the extraction grid to be
spaced
at
a distance
of 0.25 or 0.41 inches
95
from the emitter face.
A sweep
of
the extraction voltage
was conducted
the results are plotted
in
C.
with the ion source heated to
1
100 °C and
Figure C-2.
RESULTS
The
results of Figure C-1
indicate that grid transparency influences the
current out of the emitter and to the screen.
As transparency increases the
current out of the emitter-grid assembly to the screen also increases.
Figure B-2 indicates that spacing has an effect on the current emitted but
not as great as the transparency.
when
the extraction grid
is
moved
The differences
closer
is
96
about 10
in
current out of the emitter
% of total current produced.
Effect of
Changing Grid Mesh Size
100 -r
10
--
E
o
2
"E
22
O
=3
1
--
0.1
50
100
Extraction Grid Voltage
Figure C-1
.
150
( -
200
V)
Current versus extraction voltage for different extraction grid
transparencies.
97
Effect of Grid Spacing
on Current
18
16 --
14
|
2
E
12
10
CD
I
8
S
6
Grid at .25'
4 -Grid at .41*
2
50
150
100
Extraction Grid Voltage
( -
V)
Figure C-2. Screen current versus extraction voltage for 2 different
extraction grid spacings.
98
200
Appendix D
A.
Thermocouple
An experimental thermocouple made
imbedded
in
each
of
Tungsten and Rhenium (5%) was
The
ion source's heater potting to assist in data collection.
thermocouple was calibrated against the ion source's temperature readings
using an optical pyrometer.
accurate and
and
when used
easily found.
Each
The thermocouple reading was found
with the calibration table the temperature
ion source
was
-10%
the thermocouple voltage versus
from source
to source.
measured temperature
source.
99
was
quickly
calibrated for thermocouple voltage
versus temperature prior to conducting any experiments.
given temperature varies by
be very
to
Output voltage
Figure D-1
is
a
for
a
plot of
for the first Lithium ion
Lithium Source
Calibration of
1
Thermocouple
21
20 --
19
--
18
1
17
-
16
--
15
—
14
—
13
--
O
"
12
750
800
850
900
950
1000
1050
1100
1150
Temperature (C)
Figure D-1. Calibration of thermocouple, thermocouple voltage versus
source temperature.
100
.
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