Author(s)
Morris, James H.
Title
Space Experiments Aboard Rockets : SPEAR III
Publisher
Monterey, California. Naval Postgraduate School
Issue Date
1994-03
URL
http://hdl.handle.net/10945/28554
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MONTEREY CA
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Space Experiments Aboard Rockets:
SPEAR III
by
James
H
Morris
Lieutenant, United States
Navy
B.S. University of Maryland, 1986
Submitted
in partial fulfillment
of the requirements for the degree of
MASTER OF SCIENCE IN APPLIED PHYSICS
(SPACE SYSTEM ENGINEERING)
from the
NAVAL POSTGRADUATE SCHOOL
Monterey, California
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SPACE EXPERIMENTS ABOARD ROCKETS: SPEAR
TITLE
4.
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III
AUTHOR(S)
6.
James H. Morris
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REPORT NUMBER:
Naval Postgraduate School
Monterey
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13. ABSTRACT (maximum 200 words)
The SPEAR HJ experiment was conducted in an effort to better understand and compensate for the effects of satellite charging, at levels
up to KV. This experiment was designed as a lower-ionosphere test to both record vehicle charging and the effect of neutral-gas
grounding systems. Prelaunch tests were conducted at the NASA-Plum Brook facility; launch took place at the NASA- Wallops facility
Electrostatic analyzer data provided a record of the rocket body potential, and indications of ion production, or energy-angle scattering
within the plasma sheath. Plasma wave information was extracted from floating probe data and skin current probe data. Both provided
sampling to resolve signals up to 10 KHz; the skin current probe also provided burst-mode sampling up to 500 MHz. There were no
obvious signals in the 0-10 KHz data, other than a diffuse, low-frequency noise. The burst-mode data, acquired at the initiation of each
5-second charging sequence, showed a strong signal at around 100 KHz. This roughly corresponds to the lower-hybrid resonance
frequency. It is possible, that LHR waves are responsible for energy-angle scattering of the ion flux accelerated to the charged rocket
body.
14.
SUBJECT TERMS
SPEAR m
17. SECURITY CLASSIFICATION OF REPORT
Unclassified
NSN
7540-01-280-5500
15.
NUMBER OF
PAGES:
18. SECURITY CLASSIFICATION OF THIS PAGE
Unclassified
19. SECURITY CLASSIFICATION OF ABSTRACT
Unclassified
92
CODE
16.
PRICE
20.
LIMITATION OF
ABSTRACT
UL
Standard Form 298 (Rev. 2-89)
Prescribed by ANSI Std. 239-18
ABSTRACT
The
SPEAR
compensate for the
III
experiment was conducted
effects
of satellite charging,
designed as a lower-ionosphere
test to
in
an effort to better understand and
at levels
both record vehicle charging and the effect of neutral-
gas grounding systems. Prelaunch tests were conducted
launch took place
at the
up to 2 KV. This experiment was
NASA-Wallops
facility.
at
the
NASA-Plum Brook
facility;
Electrostatic analyzer data provided a
record of the rocket body potential, and indications of ion production, or energy-angle
scattering within the plasma sheath.
Plasma wave information was extracted from
probe data and skin current probe data.
10
KHz;
Both provided sampling
the skin current probe also provided burst-mode sampling
were no obvious
burst-mode
signals in the
KHz data,
data, acquired at the initiation
strong signal at around 100
frequency.
0-10
It is
up to
to resolve signals
up to 500
MHz
There
other than a diffuse, low-frequency noise.
The
of each 5 -second charging sequence, showed a
KHz. This roughly corresponds
possible, that
floating
to the lower-hybrid resonance
LHR waves are responsible for energy-angle
ion flux accelerated to the charged rocket body.
in
scattering of the
TABLE OF CONTENTS
INTRODUCTION
I.
II.
1
HISTORY
2
A.
INTRODUCTION
2
B
CHARGE II
4
C
THE SPEAR PROGRAM
7
D
SPEAR SERIES EXPERIMENTS
7
1.
SPEAR I
7
2
SPEAR II
9
3.
SPEAR III
9
SPEAR III
A
B
11
SPACECRAFT BODY
12
1
Sphere Bias System
13
2.
High-voltage (HV) System
14
GROUNDING DEVICES
1.
Hollow Cathode Plasma Contactor
14
14
2.
Gas Release System
14
3.
Thermionic Emitter
15
iv
.
DUDLEY KNOX LIBRARY
NAVAL POSTGRADUATE
MONTEREY CA
Field Effect Device
4.
C
III
15
INSTRUMENTATION
16
1
Charged
2.
Langmuir Probe
16
3
Floating Probe
16
4.
Skin Current Probes
17
5.
High-Speed Data System
17
SPEAR
III
Particle Detector
PRELAUNCH TESTING
18
B
THE PLUM BROOK TESTING CHAMBER
18
C
RESULTS OF THE TESTING
19
Vacuum
2.
Plasma Results
20
Results
20
OPERATIONS
21
V OBSERVATIONS
VII.
18
INTRODUCTION
IV FLIGHT
VI
16
A
1.
23
DATA ANALYSIS
29
CONCLUSION
35
v
SCHOOL
93943-5101
APPENDIX
LIST OF
A:
COLLECTED ILLUSTRATIONS
REFERENCES
36
75
INITIAL DISTRIBUTION LIST
77
VI
LIST
OF TABLES
Table
1
SPEAR III
Table
2.
Grounding Events
24
Table
3.
Flux Ratio Table
25
Table
4.
Standard Constants
29
Table
5.
Gyro Frequencies
30
Table
6.
Plasma Frequencies
31
Table
7.
Lower Hybrid Frequencies
32
Launch Time-line
vn
22
LIST
OF ILLUSTRATIONS
1
Charge
II
Figure 2
Charge
II
I-V Curve
37
Figure
Charge
II
Sheath Position
38
Rocket Configuration
39
Figure
3.
36
Figure 4
SPEAR
Figure
Magnetic Field Orientation
5.
Figure 6
I
SPEAR I
40
Current Potentials
41
Figure 7
SPEAR ES
42
Figure 8
SPEAR III
43
Figure 9
Sensor and Emitter Longitudinal Locations
44
Gauge
45
Figure
10.
I
Configuration
Neutral Pressure
Position
Figure 11
Sensor and Emitter Orientations
Figure
12.
Schematic of HV Capacitor Discharge Modes/Quiescent
Figure
13.
SPEAR III
Figure
14.
Simplified High-speed Data Schematic
49
Figure
15.
Plum Brook Plasma Chamber
50
Figure
16.
Mockup
51
Figure
17.
Plum Brook Chamber Suspension
Figure
18.
Floating Probe
Waveform
-
Shot 322
53
Figure
19.
Floating Probe
Waveform
-
Shot 324
54
Timing of High-speed Data System
Physical Configuration
viii
46
Mode
.47
48
52
Figure 20.
SPEAR III
Original Timeline
55
Figure 21. Wallop's Digisonde
56
Figure 22.
SPEAR III Electron Density
57
Figure 23.
SPEAR III SAIC
-
100 Seconds
58
Figure 24.
SPEAR III SAIC
-
120 Seconds
59
Figure 25
SPEAR III SAIC
-
140 Seconds
60
Figure 26.
SPEAR III SAIC
-
160 Seconds
61
Figure 27.
SPEAR III SAIC
170.770
62
Figure 28.
SPEAR III SAIC
170.962
63
Figure 29
SPEAR III SAIC
170 994
64
Figure 30
SPEAR III SAIC
171 634
65
Figure31.
SPEAR III SAIC
171.666
66
Figure 32.
SPEAR III Floating Probe
Figure 33
SPEAR III Floating Probe Line Plot
68
Figure 34.
SPEAR III Floating Probe
69
Figure 35.
SPEAR III
Skin Current Spectragram
Figure 36.
SPEAR III
Skin Current Probe
@
134.714
71
Figure 37.
SPEAR III
Skin Current Probe
@ 137.226
72
Figure 38.
Figure 39.
SPEAR III
SPEAR III
Spectragram
Spectragram of Entire Flight
Skin Current Probe
@
196.753
Skin Current Probe Spectragram of Entire Flight
IX
67
70
73
74
LIST
OF EQUATIONS
(1)
Charge Decay Calculation
13
(2)
Non-charged Flux
25
(3)
Charged Peak Flux
25
(5) Calculation for
Plasma Frequencies
(6)
Lower Hybrid Frequency
(7)
Geometric
(8) Ion
Mean Gyrofrequency
Sound Velocity
(9) Electrostatic Ion
Wave
30
30
31
32
32
ACKNOWLEDGEMENTS
I
would
and expertise
like to express
my
in the data analysis
appreciation to Professor Chris Olsen for his guidance
and graphical computer programming.
assistance the writing of this paper
would not have been
possible.
I
would also
express appreciation to John Antoniades, Paul Rodriguez, Carl Seifring, and
scientists at the
the
SPEAR
Without such
like to
Dave Walker,
Naval Research Laboratory, Space Plasma Division. Their assistance
EQ Plum Brook data analysis during
my
six-week experience tour gave
me
in
the
foundation to analyze the flight data.
I
would also
like to
Corporation (SAIC).
thank Douglas Potter
His assistance
at the
Science Applications International
in interpreting the electrostatic analyzer data
permission to use the data were paramount in the findings of this paper.
XI
and
INTRODUCTION
Space electronic systems operate
affect their operation.
when
Degradation of
in a
plasma environment which can adversely
command and
control systems
is
more
noticeable
operating high-power and high-voltage systems compared to lower electrical
activity systems.
The phenomenon
degrades the lifetime of
and systems
is
systems
higher
powered systems,
is
integrity.
which
The power
spacecraft-to-environment interaction.
systems
termed "spacecraft charging," which
Arcing, due to the accumula-
one of the main causes of phantom commands.
tion of charges,
require
satellites
that occurs is
in
turn
levels
Space defense
increases
for
the
at
some high
repetition
of
space-based defense
estimated to be 100 kilowatts to 10 megawatts, and this level
generation
level
may
require
Accomplishing the goal of effectively
rate.
grounding a spacecraft requires an understanding of the space environment and the
development of some effectual grounding scheme.
The
effect of
understood.
high-power discharge
The environment
in
ionosphere below 300 kilometers.
plasma.
As
which
in
this
the upper ionosphere
experiment
is
is
still
not well
designed to occur
This portion of the atmosphere
is
is
the
a weakly ionized
such, this environment offers the advantage of producing effects in "macro
proportion" to those that might be seen in true space.
In this respect,
measurements are
comparatively easy to obtain and analysis in a straightforward exercise.
I.
A.
INTRODUCTION
Dwight Duston,
the Strategic Defense Initiative Organization
Science and Technology director, has
tion
HISTORY
programs moving forward.
He
worked
(SDIO) Innovative
to keep the United States space experimenta-
has been one of the key leaders in the development of
space systems such as Thunderbolt and the space-based rail-gun systems.
"We're the only organization
in
DOD has specified space as our battleground.
weapons systems environment
is
not a big deal but
it
is
for us because
we
He
has stated
In general, the
have got to deal
with space. The limitations and potential leverage of environmental factors must be clearly
understood to increase existing systems
design of
new systems
capabilities
[Ref. 1]." Dwight's outlook
and performance and to optimize the
was responsible
for
weapons systems
environmental research, such as Queen Match, the Janus experiment, Red Gemini, the
Sounding Rocket Measurements Program, and many
others.
Information from these
experiments has been collected to build a strong knowledge-base for further study of the
ionosphere [Ref.
2].
With the advancement of Strategic Defense
Ballistic
Initiative (SDI),
now known
as Theater
Defense (TBD), the charging phenomenology of the space environment has become
an area of major interest and study. Even with the realignment of SDI, ionosphere research
for future strategic systems continue.
There are many reasons for
2
this interest.
First,
not only
does the problem of spacecraft charging impede the progress of development of space-based
weapons
systems, but
it
will eventually
propulsion and space-station design.
become
begin to stunt research into forms of high-energy
Secondly, information about spacecraft charging will
increasingly important to space-support industries that
activities
such as manufacturing of critical
civilian
wish to enter into space-based
and military components.
In the 1980's several programs were initiated to build a catalogue of information
relating
space-based weapons and the space environment.
Some of the programs
that
have
contributed to a better understanding of this environment are:
*
Infrared Background Signature Survey (IBSS)
shuttle missions that provided information
These experiments
— One of the most complex
on the space background and plumes.
utilized a cryogenically
cooled infrared radiometer and
spectrometer along with ultraviolet and visible spectragraphs and images to study
chemical and rocket plumes [Ref.
*
3].
Cryogenic Infrared Radiance Instrumentation for Shuttle (CIRRIS)
set
of experiments studied the energetic
Northern and Southern Lights. This
emissions during an aurora
Charge II Experiments
in the
is
particles
around the aurora of the
of interest because the
may emulate
— This
spatial structured
reentry vehicles [Ref. 4].
— Studied the spacecraft charging and sheath formation
ionosphere. Designed to investigate the electrodynamic interactions of a
tethered system with the ambient ionosphere plasma and
phenomena associated
with the injection of electron beams from spacecraft [Ref.
5].
The Defense Nuclear Agency (DNA)
—
Is also
funding research into high-
voltage power sources and power conditioning elements for space
utilization.
Their interest has been the development of high- voltages from 50 kilovolts to
several megavolts with high repetition rates for generation.
requirement that led to
[Ref
The
feasibility studies
DNA has additionally funded
Experiment Aboard Rockets
power systems
conducted
B.
at
the
this
of high-voltage systems
in
DNA
space
in
experiments that investigated the phenomenon of
(HV) discharges
(SPEAR) was conceived
DNA
LEO.
funded
NASA-Lewis Plum Brook
SPEAR
in
low earth
orbit
(LEO). Space
to explore the feasibility of operating
I,
II,
Station [Ref.
III,
and vacuum chamber
tests
7].
CHARGE II
A
experiment.
Missile
called
is
6],
spacecraft charging due to high- voltage
HV
It
forerunner of
The Charge
Range
in
II
SPEAR
series that deserves further
mention
is
the Charge
II
payload was launched on a Black Brant IX from White Sands
New Mexico
in
December 1985. The Charge
Mother and Daughter. The prime purpose of the
flight
was
II
consisted of two sections
to study the interactions of
the tethered system within the ambient plasma, and study the effects of electron
beam
emissions from the spacecraft.
The
spacecraft consisted of two sections as
consisted of the electron
beam
shown
in
Figure
1.
The Mother
section
accelerator, floating probe array, current monitor for the
four, gold-plated spheres
mounted on
a plastic telescope
and 100 centimeters from the rocket
The Daughter
The
skin.
Mother
floating probe to estimate the
electron density.
beam photometers. The
supply and electron
tether, high voltage
potential and as a
The plasma
compared
electrically
to provide a reference plasma potential
One model
through an insulated
potential for the
spacecraft.
area of the
Charge
In the
on the Mother. The Mother
tether.
were measured throughout the experiments and then
in particular that
describing the structure of the potential sheath
model was created to
(PLP) was used as a
section consisted of the communication equipment and the
current and potentials
to models.
of 25, 50,
Langmuir probe to measure the
outside the disturbed regions generated due to the experiments
and Daughter were connected
at radial distances
floating probe array
The purpose of the Daughter was
charge probe.
boom
floating probe consisted of
is
the
appeared to have been successful
NASCAP/LEO
in
model. The code for this
calculate current collection, surface charging, and plasma sheath
II
The model allows
experiment.
code the model
is
for the
non-symmetric shape of the
considered a 12 sided cylinder with a nose cone. The
model was 4.63 meters square verse the 4.68 square meters of the
The
original
surfaces of the model were considered to be perfect conductors for the purpose of ion
collection.
The
grid used in the
model would have a resolution of 21.5 cm. and 10
close to the rocket. In the area of the
The payload
points for the period
down
leg
characteristic
when the
of apogee and black
PLP
I-V plot
the resolution
spacecraft
is
shown
in
is
circles are the
model
5
when
2.
The squares
to apogee.
predictions.
normalized with the ion thermal current to the Mother
cm
about 2.7 cm.
Figure
was on the up-leg
8
The
White
are the data
circles are the
currents in the plot are
stationary in a plasma rest frame
The
calculations for this plot included computation with a plasma density of 4 * 10 5 cm" 3
Simulations were completed with and without the effects of the magnetic
observed that the
field
had a very small effect on what
to the size of the payload and
model
flight
is in
was
large
compared
We
was
large
Mother payload
experiments and that predicted by the model.
agreement with the observation.
in the
It
was
model and
compared
to the plasma sheath generated by the
Figure 3 shows the sheath position from the
section.
from the
observed both
A possible reason was that the Larmor radius
from the data observation.
Mother
is
field.
.
In this plot
also note that the sheath
is
skin estimated
we
see that the
not of a constant
size but varies with the potential.
Though
approximately
potentials, the
the
15%
to
20%
at predicting
shown
as
in
Figure
most observations,
2.
With an increase
it
deviated by
still
in the
negative bias
observed current increased more rapidly than the model predicted.
explanation for this
skin as ions
model was good
is
scattering
bombard the
by some mechanism or electron emission from the payload
spacecraft surface [Ref 8].
We may conclude from the Charge II
The magnetic
The
A possible
field
had
little
experiment that the following are
effect
true:
on the flow of current
velocity of the rocket gave an increase of about twenty percent to the ion
current, relative to the
The dimensions of the
Mother payload.
potential sheath
do
fluctuate with the applied potential to
the payload.
*
Secondary electrons are emitted from the spacecraft body, or there exists some
other mechanism of electron generation.
6
C.
THE SPEAR PROGRAM
The
DNA funded a series of experiments that would investigate the effect of a
highly charged,
HV
started in 1987,
had the primary goal of examining the
in space.
the
The
system
SPEAR I
SPEAR II
and
III
in the
space environment.
D.
SPEAR
feasibility
in the
of experiments,
IX sounding
experiments rode on a sixty-foot Black Brant
NASA facility.
series
of high-power operations
experiment rode on a sixty-foot Black Brant
experiments were launched from the Wallops Island
payloads were tested
The
X
rocket, and
sounding rocket All
Prior to each launch, the
NASA-Lewis B2 vacuum chamber.
SPEAR SERIES EXPERIMENTS
1.
SPEAR I
SPEAR I was
of experiments to
applications.
launched December 13, 1987.
test the physics
The payload
of developing and designing
consisted of two,
of thirty-nine inches (see Figure
4).
SPEAR I was the first
HV
in a series
systems for space
20 cm. diameter spheres separated by a distance
The spheres were constructed from aluminum and
plated
with gold over nickel. The spheres were separated by a gradient ring section of eighty-two
inches to ensure a uniform potential drop from the main portion of the
spheres.
boom
to the
two
The main body consisted of six
The
sectors,
Low-Light TV. (LLTV), and the
section contained the photometers,
first
each performing independent functions.
The second section contained the high-voltage power supply and capacitor
pressure gauge.
This section generates the bias voltage applied to the spheres.
circuits
neutral
contained the measuring devices, Langmuir probe,
wave
Section three
receivers and the particle detector.
After the measuring devices there are the attitude control system and finally the plasma
contactor.
The
magnetic
attitude control system's
field as
shown
Figure
in
5.
primary purpose
is
to align the spacecraft with the
Changing the spacecraft alignment allows
for the
collection of data and analysis of interactions between the geomagnetic field and the high-
voltage system.
magnetic
field
The V-plane of the booms was oriented
during the
first
portion of the
the spacecraft so that the V-plane of the
orientation
flight.
booms was
maneuver placed one sphere-boom
The
second
tests
test
were conducted
was conducted
at
attitude control system positioned
parallel to the
parallel to the
Two vacuum chamber tests were conducted
chamber
to be near perpendicular to the
as a
magnetic
geomagnetic
mock up
test
field.
The
last
field.
before launch
The
the University of Maryland plasma chamber facility and the
at the
NASA-Lewis Plum Brook B-2 chamber. The mock-up
tested the circuit, equipment robustness and survivability to arcing and high-voltages in a
plasma environment. The integrity of the spacecraft high- voltage
were validated and proven to be sound [Ref
circuits
and experiments
9].
Observation of the chamber test showed a glow around the spheres suggesting plasma
breakdown.
It
was
generally believed that this
8
breakdown would occur when more than a
few
kilovolts
expected
were applied to exposed electrodes
spheres during the entire
time.
The decay
in
flight.
LLTV
were found
from
Van Horn,
in all current
charging peak, there
formed or scattered
is
1
MQ
data plots.
SPEAR
on
I
Figure
plots at the
system impedance was nearly
ions accelerated into the (negatively
SPEAR I at the Naval
with such data.
a diffuse spectra below the peak.
in the sheath (see
The I-V
[Ref. 10].
SPEAR I showed
in his thesis
analyzed the charging behavior of
the
spikes observed are considered to result from a
The plasma impedance was about
charged) rocket body.
Note
Postgraduate School,
that in addition to the
These were considered to be ions
7).
SPEAR H
2.
SPEAR II was
first
The
was no plasma breakdown during
linear relationship, indicating the
Electrostatic analyzer data
the
implies that there
data.
oscillatory current and
bottom of Figure 6 show a
constant.
did not detect any arcing between the
these plots are typical of a capacitor discharging into a linear resistance.
confirming the
damped
LLTV
Figure 6 shows plots of current and potential versus mission
The exponential decay shown here
flight
The
above 100 km.
at altitudes
plasma chamber The effect was also
in a
designed to test
rail-gun into space.
new
high-voltage components and also carried
After the launch, the rocket guidance system failed and the
booster was destroyed thirty-five seconds into the launch. Even though the
disaster, information
later flights.
was gained about
the design of
The chamber experiments of the new
SPEAR experiment.
flight
ended
in
HV circuits that would be utilized on
HV designs would be employed in the next
SPEAR ffl
3.
SPEAR
III,
the most recent experiment in the sounding rocket series,
designed similar to
SPEAR
shown
This figure illustrates the Weitzmann
in
Figure
8.
was attached and
However, instead of using two
Boom
spheres, only
one was used as
onto which the floating probe
the sphere extended.
Some of the
(i)
I.
was
scientific objectives
of
SPEAR III
were:
Test several grounding mechanisms
when
the
body reached
several
negative kilovolts.
(ii)
Diagnose physical mechanisms of the grounding techniques
(iii)
Study the dispersion of gaseous effluent emitted from a space platform,
(iv)
Test the effectiveness of the grounding techniques
differential
(v)
in
charging on a diagnostic probe and on solar
Monitor the undisturbed and disturbed plasma and
sections which follow address the data obtained in this mission.
10
cells,
neutral gas environ-
ments of the pay load.
The
reducing local
SPEAR
II.
SPEAR
SPEAR
the
I
the most recent experiment in the sounding rocket series,
III,
in design.
As
in
15, 1993, at
after launch
was
similar to
previous experiments, testing prior to launch was conducted
NASA-Lewis B2 vacuum chamber.
March
III
SPEAR
21:13 EST.
The
III
SPEAR
III
rocket
at
was launched
reached an apogee of 289 km., 278 seconds
from NASA-Wallops.
SPEAR
III
science
unanswered questions from
the control thereof.
fluxes observed
on
objectives,
SPEAR
I,
per the
and
previous
new concerns about
The primary purpose of this work
SPEAR
I,
section,
is
largely
arose
from
differential charging,
and
to further address the return ion
and attempt to explain the (apparent) scattering process
observed there.
To
achieve the
SPEAR
III
science objectives,
designed to be implemented during the space
SPEAR
I.
These methods, discussed below,
flight
are:
(i)
Ambient Gas
(ii)
Thermionic emitter
(iii)
Field Effect Device
(iv)
Gas Release System
(v)
Hollow Cathode Plasma Contactor
11
new grounding schemes were
along with circuitry similar to that of
Another objective of
SPEAR
was
III
HV
to study the effect of slow versus fast
charging, and again address questions as to the effect of orientation relative to the
geomagnetic
A.
field.
SPACECRAFT BODY
The rocket used
stages of solid fuel
the rocket
which
was seventeen
Figure 9
illustrated in
were the nose cone,
was a Black Brant
for the launch
are, Terrier
inches.
The
five
The
It is
composed of three
Booster, Black Brant 5c, and Nikha. The diameter of
scientific
modules
HV module,
X rocket.
that
Science
1
equipment was housed
were important to the
in the
SPEAR
rocket body as
III
experiment
package, Science 2 package, and the Science 3
package.
The nose cone consisted of the
HV boom and
sphere assembly, the
Light Level Television Camera (LLLTV), and forward viewing
Module was
supply,
HV
fifty-four inches long
LLLTV
aft
viewing
camera
and contained primary battery power, the
Low
The
HV
HV
power
capacitors, switching network, current and voltage monitors, and the signal
conditioning unit.
Science
1
was
thirty-eight inches long
particle detectors, solar cell system,
two
and contained the two energetic
skin current probes, differential charging device,
electron emitter, field-effect device, magnetometer, and a three-channel fast data unit.
Science
2's
length
was
fifty-four inches
and contained another
fast
channels, the neutral pressure gauge system, transient pulse monitor,
data unit with three
two
additional skin
current probes, differential charging device, gas release system, hollow cathode plasma
12
contactor, Langmuir probe electronics, and floating probe electronics.
Science 3 was
eighteen-and-a-half inches in length and contained the Langmuir probe and boom, and the
floating
probe and boom.
Figure 10 shows the location of the neutral pressure gauge, hollow cathode plasma
contactor,
Langmuir probe, and the
neutral gas release system jets
sensor and emitter orientations viewed
The two charging
the required potential.
circuits
The two
aft
shown
in
Figure
Figure
1
2 were designed to charge the sphere to
high- voltage circuit-charging devices, a switched fast and a
A charging
designed to charge the sphere for each grounding device. Each charging event
over a five second interval [Ref.
device
was
was conducted
11].
Sphere Bias System
The sphere was pulsed
circuit the
details the
of the payload.
ramped, were designed to charge the sphere to the same potential.
1.
1 1
plasma
in
positive with respect to the payload.
To complete
the
space provides the ions or electrons, depending on the sphere potential
Charging the sphere positive forces the body negative. The expected charge decay
described for positive voltage
as:
Vc = V
(1)
e
RC
Charge Decay Calculation
This type of charge decay was seen in the
flight.
is
SPEAR I
data, per Figure 6,
This similarity validates this conceptual approach to the
13
from the
SPEAR
SPEAR experiments.
I
These
SPEAR I chamber results as well as validating
figures validate the
III
data
set.
The type of decay seen
system, the plasma
2.
in the
would not break down
for ionospheric systems [Ref. 12].
High-voltage (HV) System
a differential bias
and the vehicle to drive the vehicle to negative potentials.
to charge a capacitor
drawing of the
illustration.
HV
circuits.
which
body negative
A twenty kilovolt
power supply
connected to the deployed sphere. Figure 12
The incorporation of
RC
components
is
also
shown
It is
is
a
in that
mode of
these circuits that quickly charged the sphere, which in turn forced the
[Ref. 13].
GROUNDING DEVICES
1.
Hollow Cathode Plasma Contactor
The plasma contactor
generating a dense plasma.
volts
is
between the deployed sphere
Also shown are the differences between the ramped, fast and quiescent
the circuit design.
B.
SPEAR
Figure 6 indicates that for a properly designed
The high-voltage system provided
was used
the accuracy of the
It
is
a hollow cathode gas discharge that
generates 0-10
The plasma contactor was designed
amps
is
capable of
current at bias levels greater than -100
to produce a high plasma density in the vicinity of
the vehicle which results in an electrically
low impedance path between the vehicle and the
ionosphere [Ref. 14].
14
Gas Release System
2.
The gas
nozzles.
The flow
release system
was
rate
The torque induced upon
was designed
to release non-corrosive gases through four
a max. of one gram/second to a lower limit of .02 grams/second.
was designed not
the spacecraft
torque generated by the ACS.
The gas was discharged
to exceed 1/1 00th of the
tangential to the cylindrical skin of
the payload, and perpendicular to the longitudinal axis of the payload.
This provided a high
density neutral gas cloud in the vicinity of the vehicle so that additional plasma
by
collisional ionization,
minimum
is
generated
thereby creating an electrically low impedance between the vehicle
and the ionosphere. The plume formations were two opposed
pairs,
with each pair on
opposite sides of the rocket body [Ref. 15].
Thermionic Emitter
3.
The Thermionic
emitter
is
basically a device similar to a heating element.
element was three percent rhenium-tungsten wire (0.008" diameter). The
maximum
The
operating
temperature expected was 3073 degrees Kelvin. The design lifetime was ten hours.
A tungsten metal cathode protruded into the plasma sheath where when heated would
emit electrons into the plasma sheath
in current.
transfer
The concept here
is
Emissions were limited to a
maximum of one ampere
to utilize the negative charge of the spacecraft
of electrons to the plasma. The intended
result
of this emitter
is
body
for the
a rapid ground for
the spacecraft [Ref. 16].
4.
Field Effect Device
The Field
Effect Device
is
an array of collected fine points
be the focus of charge accumulation.
These
fine points will
This procedure utilizes negative charge of the
15
spacecraft
The
body and the
electric field to
net result here will be a reduction of the negative charge of the spacecraft after being
discharged through a plasma [Ref
C.
cause emissions through the sharp tips of the device.
17].
INSTRUMENTATION
Charged
1.
Particle Detector
The charged
detector
particle
was developed by Science Applications
The analyzer
International Corporation (SAIC).
positive ions in a predefined energy bandwidth.
different
Two
curved plates biased to collect
pairs
of sensors were used, with
geometric factors. The electrostatic analyzer (ESA) generates an energy spectrum
of the ions
is
utilizes
in the
plasma sheath. The geometric factor used here for the high sensitivity
3.4* 10" 5 E*seer-cm 2 [Ref.
18].
The energy
resolution
is
ESA
seven percent and the energy
ranges are 12 to 897 electron volts and 0.39 to 29 kilo electron volts. The time resolution
was 32 ms. per sweep,
2.
1
ms. per sample [Ref. 19].
Langmuir Probe
The Langmuir probes were mounted
flush with the vehicle skin.
measured the distribution of return currents to the vehicle
to the grounding systems.
3.
with respect
Figure 10 shows the location of the probe [Ref 20].
Floating Probe
The
spacecraft
at different locations
The probes
floating probe
was
a spherical device deployed three meters from the
body on a boom The purpose of the probe was
16
to
measure the potential difference
between the vehicle and the sphere when the spacecraft body
that the
is
negative.
was expected
It
probe would be beyond the plasma sheath when the vehicle was -300
volts.
The
floating probe also gave an indication of the effectiveness of each grounding mechanism.
shows how
Figure 8 shows the position of the probe during deployment, and Figure 12
HV circuit in the modes of operation [Ref
connected to the
4.
it
was
21].
Skin Current Probes
The
vehicle skin.
skin current probes were plane current collectors
2
The range of the instrument was 500mA/m
to
mounted
50mA/m 2
designed to measure the distribution of return currents to the vehicle
near and remote to the vehicle [Ref. 22],
The output from
flush with the
The probes were
.
at different locations
the skin current probes provided
one of the inputs for the high-speed data system.
5.
High-Speed Data System
A
store
it
high-speed digitizer was incorporated
into a cache
memory. The sample
were used to prevent
KHz
low-pass
filter
aliasing.
Figure 13
was used on
the
interval
is
in
the design to capture the data and
was one microsecond.
pass
a simplified high-speed data schematic.
Rogowski
KHz
low-pass
with the Nyquist sampling theorem, the largest observable frequency
filter.
is .5
a diagram of the digitizer circuit and the sampling cycle [Ref. 23].
acquired in a 16 ms. "snapshot" beginning just before the high-voltage
17
filters
A 300
current coil and the sphere voltage
measurement. The skin current was filtered with a 150
is
Low
In accordance
MHz.
Note
is
Figure 14
that data are
switched.
SPEAR III PRELAUNCH TESTING
III.
A.
INTRODUCTION
The NASA-Lewis Plum Brook
Station
vacuum chamber was used
to test
SPEAR
III
prior to launch.
B.
THE PLUM BROOK TESTING CHAMBER
NASA
The
Plum Brook
Station B-2
vacuum chamber
is
a large, cylindrical chamber
configuration measuring approximately 13 meters in diameter and twenty meters high.
The chamber
x
of 2.5
is
6
pumped by
1CT Torr.
the rocket
twelve, 90 cm. diffusion
pumps
to a
minimum
base pressure
During the experiments, the high-voltage was applied to booms while
body was held
at
ground
potential.
This chamber has been used for testing of payloads for various rocket projects.
facility
provides the capabilities of testing payloads in
pressure and ion mixtures that
atmosphere.
may be
vacuum or
set to parallel (or
in
varying degrees of gas
model) different conditions
These conditions usually involve height and
This
partial effects
in the
of vacuum on
payloads.
Figure 15 shows the physical configuration of the spacecraft
chamber.
The
total length
in the
plasma
testing
of the spacecraft body was 300 inches and an extender arm
18
exceeded
this length
the spacecraft
inside the
by another eighty
body perpendicular
chamber
to
to the extender
in turn
the insulated
the fiber optic cable
video cameras were mounted
Fiber-optic cables were connected to the pulser,
Into the base of the spacecraft (see Figures 16 and 17)
was connected
as
were the high-voltage ground
mockup were connected
boom. The suspension
C.
Two
from
connected to the relays and power controls, which also came through
chamber bulkhead.
suspension for the
arm.
forty inches
Their cables were fed through an insulated
observe the sphere.
junction in the wall of the chamber.
which were
The sphere was extended
inches.
bolts did not
to the base structure
Support
cables.
of the spacecraft and the
have any effect upon the experiment.
RESULTS OF THE TESTING.
In the process of testing, the charging circuits were tested and skin potentials
measured.
floating
vacuum
Figures 18 and 19 areplots of the potentials during the
test.
The
probe and the skin potential decay have approximately the same decay constant
after charging.
These
results
conformed
operation and design of the spacecraft and
and validated the
to scientific expectations
Weitzmann boom. The
probe measurements with the vehicle skin
potential.
figures
During the
show
the floating
test there
indications of plasma breakdown.
The
results of the testing
may be summarized
19
in the following
manner:
were no
1.
Vacuum
Results.
Boom
orientation perpendicular to the magnetic field increases the
breakdown threshold
10" 5 Torr.
4 x
for the sphere
Oblique orientation
above lOkV for pressures below
in this
chamber
for the
SPEAR
I
mockup was 4kV by comparison.
Body
sheath threshold
was above lOkV.
4
For pressures above 2 x 10 Torr the threshold drops below 5kV.
The
produce an observable change
disrupter does not
characteristics for voltages
2.
up
in the
breakdown
lOkV.
to
Plasma Results.
For pressures
in the
1
x
10" 5
Torr range or higher, the sphere sheath
broke down for voltages down
range of
1
x
4
10
The maximum
-
2
x 10 5 cm" 3
to
200
V
with plasma densities in the
.
voltage held by the negative body sheath was 1.5kV,
possibly due to enhanced ion flux from the breakdown at the sphere.
These
results
were considered repeatable.
20
FLIGHT OPERATIONS
IV.
SPEAR
MK70
III
was launched on
motor, consisting of the
a Black Brant 10 rocket.
MK12
motor case
The booster was
a Terrier
The rocket
for better performance.
booster had a case diameter of 18 inches and a length of 170 inches. The sustainer was the
The dimensions of
standard Bristol four-fin motor assembly.
the sustainer
were
a case
diameter of 17.26 inches and body length of 208 inches.
SPEAR
III
Flight Facility.
was launched from
a
Langmuir probe. Note
the mission.
Figure 22
The time
instruments and
line
all
is
a
The
at
21:13
list
at the
NASA-Wallops
Figures 20 and 21
latter figure includes the density inferred
below 2
*
5
show
from
3
10 cm" throughout
of selected events pertaining to the writing of this paper
as planned for the entire flight as well as activation of
attitude maneuvers.
field.
EST
flight time-line.
that electron densities remain
was executed
respect to the magnetic
processed by
launcher
Figure 20 illustrates the planned
the data from the Wallops Digisonde.
the
rail
PCM
NASA at Wallops Island
All three science attitudes
data
was
[Ref 24].
21
were achieved with
collected for the entire flight and later
Table
1
SPEAR III
Launch Time-line
MET(s)
MET(s)
Launch
200
N. G.
69.0
payload separation
220
T E
71.0
eject
240
FED
71.4
instrument deployment
260
PC
76.0
science attitude
280
no ground
890
maneuver complete
281
science 2
91.0
deploy
HV boom/FP
290
N G
95.0
charge
HV capacitors
310
T E
95.1
turn on
ESA HV
330
FED
event/no ground
350
PC
0.0
nose cone
1
1000
first
1100
neutral gas
370
no ground
130.0
T. E.
371
science 3
150.0
FED
400
N. G.
170.0
Plasma Contactor
420
FED
1900
no ground
460
no ground
470
N G
22
OBSERVATIONS
V.
Observation and data analysis from wave and particle instruments were reduced and
analyzed for this work.
and studied to see
The
particle data
if substantial
ion flux. Figures 23 thru 26
energy angle scattering was again occurring
show
The
ten kilovolts.
the
initial
The data
interesting features in the data.
vehicle charged to slightly
were examined to determine charging behavior
more than one
potential drops
80 seconds of operation, and the majority of
are
shown
were exponential
at
in time,
on
is
raised to
We
with some exceptions.
SPEAR
I
(see Figure 6).
Figure
110 seconds shows a dropout associated with a momentary discharge of the rocket
body, probably due to a gas release. The lack of measurement or detection
is
The
as energy time spectragrams.
kilovolt negative as the sphere bias
also observed the exponential decay previously seen
23
in the return
unexplained.
The
ESA
at
1
14 seconds
data did not detect any difference between fast and slow ramp
charging of the sphere.
Figure 24 two sequences are shown which include operations of the thermionic
In
emitter.
The
altitude (220
ESA
emitter
is
on
for the shots at 130
and 135 seconds. Observe that
km.) the emitter worked extremely well
shows no ions above the 10
EV energy threshold.
in
at this
grounding the spacecraft
This appears to be due to the
low
The
TED
operation.
The grounding events
spectragrams
that
were
in
operation during the time-frame of the four
are:
23
SAIC
Table
2.
Grounding Events.
MET
It is
the
the
Event
100-110
No
110-130
Neutral Gas
130-145
Thermionic Emitter Device (TED)
145-150
No
150-170
Field Effect Device
170-180
Hollow Cathode
TED event at
Hollow Cathode
Event
Event
130 seconds that completely grounds the spacecraft. For flux analysis,
is
chosen
Data from the shot
at
at
170 seconds.
170 are shown
in
Figure 26. Line plots were
made from
of data (see Figures 27 thru 31) to analyze the phase space density and the
function.
fitted
As seen
in the figures, the
(LSF), and from this
peak was modeled (but not
we
obtained the density and thermal temperatures
fitted) as
parameters are also included
low energy portion of the spectrum was
in the figures.
with that of the charging peak. In the table,
and "Charging Peak" refers to what
flux and flux'
Table 3 summarizes what
is
"Low Peak"
is
refers to that
least
squares
The charging
found
peak The "Ratio"
for calculating the flux are:
24
in
The
the data.
peak and flux'
under the low energy
observed under the charging peak
kT corresponds to the potential
The equations
distribution
a Maxwellian, accelerated through a potential drop.
In the equations below, flux correlates to data not associated with the charging
corresponds to flux as
this set
is
kT is
energy that
the ratio between
kT q
N
:=
flux
x
\
Non-charged Flux
(2)
~
flux'
j
—
kT'q
N'
i
t
(1+—
JcT'
Charged Peak Flux
(3)
table representing the results of these calculations
Table
Timing
18
2.923 x 10
480
170.77
1000
4.270
x
171.09
820
4.144
171.122
820
4.515 x
171.666
580
2.820
From the table we
•
m'
2
x io 18 .
m"
2
18
x io 18
•
as follows:
Charging Peak
1
•
sec'
•
sec"
«
sec"
lO^-m'^sec'
see that the
some process which can cause
2
m"
io
is
Flux Ratio Table
•
m'
2
•
m"
8.
2
•
Ratio
sec"
lO^-m^-sec
1
5.499 x
1
17
2
9.174 x 10 Tn- «sec-
1
9.174
-1
•
17
8.700 x io
sec
x
10
17
17
8.700 x io
Tn" 2 »sec•
m"
low energy portion of the spectrum
charging peak by a factor of 3 to
to energy and angle.
3.
Low Peak
Potential
171.634
in the
4>
•
2nm
\
The
2nm
2
•
is
1
3.241
1
4.922
1
4.517
1
7.766
1
sec"
3
36
larger that the flux
This observed fact motivates a careful search for
the unusual distribution of ions to be scattered with respect
The observation here
is
similar to
25
what was found
in the
SPEAR I
data.
The question which
is
is
motivated by the electrostatic analyzer data
the source of the scattering
answer
finding an
skin current
mode
is
probe
which
is
in the electric field
Data sampled
at
apparently occurring here?
is,
therefore:
The most
likely place for
(wave) data, particularly from the floating probe and
KHz
20
from both were studied, and the
1
MHz burst
data were studied.
The
floating
probe should have detected any significant signals below 10
Figure 32 shows the fourier-transformed floating probe data for a 20 second interval
here constant line frequency signals throughout the time of the shot.
frequencies here as mechanical artifacts.
during the entire
effect
that
The
flight
would have been
The
field.
here
was
at 2, 4, 6,
typical
and 8
is
in the
of the
may
of all the floating probe
probe data
imply that either the signals were masked by noise,
place that did not allow signal detection.
line plot data.
correspond to the frequencies
spacecraft.
classifies the
effect in the floating
a line plot of the electric field versus the frequency in
these signals throughout the entire
artifacts
see
range of the cyclotron frequencies that are dependent on the
lack of any signal
KHz
We
three orientations in the magnetic field did not have any apparent
some other phenomenon was taking
Figure 33
The author
KHz
This floating probe plot was typical for the probe
on any of the data analyzed. There was no noticeable
magnetic
or
what
flight
The peaks
in the
that
KHz. The
you see
line plot
in the line plot
spectragram. The consistency of
supports our conclusion that they are mechanical
Floating probe data for the entire mission are
shown
in
Figure 34.
This reinforces the lack of evidence for resonant signals which mirror physical processes
low-frequency, diffuse spectra
is
apparent throughout the mission.
26
The
The
time.
is
skin-current probe plot Figure 35
mode
were acquired
down
broken
at
at
data transformed into the frequency domain.
MHz,
1
into
shortly after the initiation
32 segments of
resulting spectra are plotted as
the time the snapshot
the snapshots.
The 100
physical process.
Some of the
Figure 35
is
the
The
was
5
frequency and mission
signal
is
a
The
shown
strongest spectra
These measurements are
body remained below 10V
grounding devices, but
were acquired
in
bands plotted are used to separate
the plot appear to be characteristic of
is
peak
is
in the spacecraft
MET.
line
found around 100 kilohertz.
showed no charge peak and where
It
real
always present.
Figure 36 and 37 are skin current
A broad
magnitude.
some
events from 130-145 seconds
at
ESA spectragram of the time frame.
between current flowing
Sixteen milliseconds of data
of a sphere bias
vertical black
in
intensity varies with
TED event where the ESA data
the rocket
This
range.
though they occurred over a 4.5 second interval beginning
acquired.
KHz
KHz
ms. each, and transformed into the frequency domain.
plots that correspond to the 120 time event.
This
vs.
This spectragram, a gray scale, shows that there are signals in the 100
the burst
The
shows the amplified
seems there
is
the potential of
a profound relationship
environment system and the oscillations observed
here.
Somewhat more
typical are data
from 180 to 200 seconds Figure
broad peaks from wave data acquired
interest
is
characteristic
Figure 39. This figure
is
we
Again
the start of each probe bias sequence.
the lack of signal intensity during the
were generally a
in
at
38.
ramp charge. The strong 100
Also of
KHz
of fast charging. Another characteristic of fast charging
a frequency vs. time spectragram of the entire flight.
27
find
signals
is
THe
seen
black
vertical lines that appear
from
to
1
MHz
seem to be associated with the
Occasionally they appear in the ramped case, but infrequently.
28
fast
charge
VI.
For calculations
DATA ANALYSIS
following notation
in this chapter, the
is
assigned:
N - Particle Density.
M
'-*
The
that
is
table
below
is
would be expected
Mass of Ion.
m
- Mass of Electron
B
- Magnetic Field Strength.
a
list
in the
of constants
that are
ionosphere.
The
used to calculate characteristic frequencies
variation
of the magnetic
field
during the
flight
not strong enough to have a noticeable effect.
Table
4.
Standard Constants.
Magnetic Field (B)
5
3.20 x 10"
Particle Density (nominal)
4
5.00 x 10
tesla
*cm 3
30
0.91 x 10-
«£#
Proton Mass (M)
27
1.67 x \0'
»kg
Oxygen Mass
2.67 x
Helium Mass
27
6.68 x 10'
Eo
8.854 x
Electron
Mass (m)
m
26
»kg
•
\V n
19
1.67 x 10"
q (Electron Charge)
The
•
.
kg
* farad/
meter
cotd
particles that are expected to generate frequencies in the ionosphere are the electron,
proton, helium, and the oxygen ions.
Using the values
frequencies:
29
in the table
we
calculate the gyro
qB
/(4) Calculation
of the Gyro Frequencies
Using the above equation where q
magnetic
field strength
we
1%M
the charge of an electron mass and
is
B
is
the
arrive at the following nominal gyro frequencies:
Table 5
Gyro Frequencies
Electron
0.890
MHz
Proton
487.0
Hz
Helium
122
Hz
Oxygen
30
The plasma frequency
is
Hz
calculated by:
/=
Nq 2
(5) Calculation for
1
Plasma Frequencies.
These calculated plasma frequencies of the
particles in the
4
3
a density of 5 x 10 cm." per the Wallops Digisonde, are:
,
30
plasma environment using
Table 6
Plasma Frequencies
Electron
2.000
MHz
Proton
0.048
MHz
Helium
0.023
MHz
Oxygen
012
MHz
Figure 20 shows the densities obtained from the Wallops Digisonde and aboard the LP,
The lower-hybrid frequency (LHR)
and the associated range for plasma frequencies
generally close to the geometric
mean gyrofrequency:
+
00
LHR
Q
c
co
c
«
Qr\
gyrofrequency,
is
then
Q c u>,
w um
defined
a>
= —— and Q
M
Q
~
p
\|
€o
2
M
Lower Hybrid Frequency
Q w c and/u/R
,
c
-
fgmr Where fgmr
{qW
Mm
J_
2
7i
Geometric Mean Gyrofrequency.
Assuming the following masses,
Nq
,
as:
/.g>»g
(7)
where
,
(6)
If
is
the
LHR
frequencies are:
31
the geometric
mean
Table
The
first
last
wave that
find the ion
sound
is
7.
Lower Hybrid Frequencies
Proton
20.9
KHz
Helium
10.5
KHz
Oxygen
5
possible
velocity.
is
\
this velocity
To
calculate this
we
Assuming one dimensional compression, y = l,we have:
M
(8) Ion
Using
KHz
the electrostatic ion sound wave.
JcT+ykT
Vs =
2
,
where
_.
Vs = 620
m
src
Sound Velocity
of the sound wave
we
calculate the electrostatic ion
sound wave
frequency to be:
/
2ta,2
-ft+k'Vs
(9) Electrostatic Ion
Assuming
a wavelength
a frequency in the range
The
floating
1
2*
Wave
of ~3 meters, a reasonable bound on the sheath
size,
we
find
of .6 MHz.
probe data were previously analyzed using standard fourier transform
technique. Signals at 4 and 8 kilohertz (see Figure 27)
32
were observed continuously during
The constant frequency suggests
the entire
flight.
be
generated
6
locally
KHz at
artifacts.
100 seconds,
rising
that these signals should
be considered to
There was one additional monochromatic signal beginning
KHz
monotonically to 9.3
at
at
500 seconds. These variations do
not correspond to any obvious physical parameters such as magnetic field strength or
orientation, electron density, or altitude
during
discharge sequences, in the 0-2
all
previously
filters
shown
in
therefore tentatively identified as an
KHz
signal,
produced
frequency range.. Note that the band pass
Figure 14 limits the sensitivity
The thermionic
transformation.
is
mentioned signals there was a broadband
In addition to the three
artifact.
This signal
the higher end of the
at
emitter operations generated narrow interference lines at
-2300 Hz, -4500 Hz, -6800 Hz, and -8000 Hz.
There was no noticeable effect
in the floating
probe data that would have been
range of the ion cyclotron frequencies that are dependent on the magnetic
any signal
may
infer that either the signals
was taking place
The above
calculations
show
that the
in the skin current
the electrostatic ion sound wave.
would
does not
in the
The lack of
some other phenomenon
lower hybrid frequency
probe
at
100
KHz
sound wave.
same range of the lower hybrid and would be
classify the signals as
waves could generate the
lower hybrid or
within the range of
in the calculations is
of the plasma sheath then
If such a
wave
is
difficult to detect.
generated
it
it
The author
electrostatic, but notes that those types
signals observed in the data.
33
falls
Also shown
If we restrict the dimensions
possible to generate the electrostatic ion
fall
noise, or
the
that did not allow signal detection.
what was observed
may be
were masked by
field.
in
of
The
skin current probe yielded the high-frequency data that appears to harbor
frequencies of interest.
signals
The
analysis
of the data was similar to that of the floating probe. The
of the probe vary with events of the
flight,
but this
may
not be due to any of the
grounding techniques.
The
three orientations in the magnetic field did not have any apparent effect
the data analyzed.
34
on any of
CONCLUSION
VII.
The
many
SPEAR
aspects of
III
it.
data has not been fully explored.
The
waves
finding of
that
There are opportunities to explore
may be
the lower hybrid or electrostatic,
warrants further investigation into the mechanism of their generation
these
waves may have to one of the grounding experiments needs
detailed analysis
in the
is
low energy
needed on the
range.
The
ESA
The
association
further study.
A
more
data to understand the mechanism of the large flux
finding of characteristic frequency in the range of the lower
hybrid, and possibly electrostatic ion
sound wave, may be the key to solving the spacecraft
charging problem.
The
floating probe
characteristic
ESA were
extremely useful for this analysis. The frequency ranges of the
were too low to be of
100
KHz
The
skin current probe revealed a
signal.
The success of this
is
significant use.
flight clearly
demonstrates that current theory and engineering
capable of achieving the objectives of the sponsor. Engineers were capable of developing
the hardware and demonstrated
robustness of both
at the
SPEAR I
and
its
survivability in the
SPEAR HI flights may be attributed
NASA-Plum Brook B-2 chamber and the
All future space experiments
plasma chamber.
would
benefit
The success and
to rigorous testing
done
space chamber at the University of Maryland
from
35
this
type of testing prior to launch.
APPENDIX
A:
COLLECTED ILLUSTRATIONS
^IJeam Photometer
Mother
•jixgS^jpv Electron Beam
WM&V
Right Direction
^x/nSetm Camera
PLP Probe Amy
^%
Charge Probe
VjTether Camera
Sbetth Photometer
*
Strobe
r
Tether
^
(A^^^vLP AatenrM
Charge Probe
Daughter
Figure
1
Charge
36
/
II.
HP
Antenna
40
o
2
UT
30
DSQ3
O SQ4
_ • Simulation
20
_
I
10
M «9.3jiA
580 m/s
ne
—
T
i
200
100
-<>HV
Figure
2.
Charge
37
II
-4xlO l0 m-3
(V)
I-V Curve.
300
400
1.5
S
a
1.0
O
Experiment
•
Simulation
-
a
o
O
CO
0.5
0.0
-<&HV(V)
Figure
3.
Charge
II
38
Sheath Position.
39"
20 cm
diameter
spheres
115"
82"
t
Photometers
TV Camera
Neutral Pressure Gauge
High-Voltage Power Supply
and Capacitors
Langmulr Probe
Wave
Receivers
140"
Particle Detectors
Telemetry
Attitude Control System
Plasma Contactor
17"
Figure 4
SPEAR I Rocket
39
Configuration
r
^i
Position
1
.
Near Perpendicular
•>
\t
^T
v
\t
Position 2. V-PIane of
Booms
Parallel
^*i
«\\^
Position 3.
Boom
Figure
5.
of Sphere
1
Parallel
Magnetic Field Orientation.
40
Sphere
i
i
i
i
|
.
o"
rj—
i
———
n
1
1
-\
— ——— ———————————
|
i
i
i
i
i
230
MET
—r~r
T
I
I
i
i
i
c
(seconds)
I
—
I
I
I
F
—— —— —————
1
1
1
1
1
1
1
——
1
T
I
1
<
E
c
.^
v_
a>
°eg
-t
i_
km
3
o
/.**'
<D
1—
0)
SZ
Q.
CO
i
«,
0)
o
i
i
232
E
^
Zi
|
BIS2-S1
303km
-*
i
|
231
(seconds)
Altitude
—
i
232
231
MET
.
i
|
i
i
t
230
<
Sphere 2
1
——————— ——————
i
i
|
,;.)»
a
.««"•"
„
1
1
1
1
10
I
I
I
L_l
1
III
l—L
-i
1
I
I
6.
CO
L
10
Potential (kV)
Figure
1
30
20
Potential (kV)
SPEAR I
Current Potentials
41
SPEAR10000
>
£
1000
en
CD
c
100
290
288
286
M. E.
Ion
Figure 7
T.
I
42
(s)
Electrostatic Analyzer
SPEAR ESA.
292
4C
o
m
H>
O
>
r- ^^
o
m
0)
O
2
en
33 T>
XI
m
<
m
£
CD
vQ
Figure
8.
SPEAR in
43
Configuration.
F
—
'
HV
54"
Forward
Sensor,
System
^
1
LLLTV
Aft LLLTV
Optical Spectrometer
—
—
MEJEA-BM
and Emiffig Longitudinal
Thermionic Emitter
Skin Current Probes
I
Lacaiiflns
(2)
Science* 1
9"
Differential Charging Device
Field Effect Device
I
'
Solar Cell Panel
•»<
Energetic Particle Detectors (2)
Vehicle
Support Module
T
30"
i
—
—
1
Neutral Pressure
Differentia)
Gauge
Charging Device
Skin Current Probes
(2)
Science-2
*
_j
Hollow Cathode Plasma Contactor
1
Release System
-Gas
I
J
ACS IT
*
Scicnc«-3
Figure
9.
T
1*3*
I
—
—
I
Lungmuir Probe Deployer
Floating Probe Deployer
Sensor and Emitter Longitudinal Locations
44
Figure
10.
Neutral Pressure
45
Gauge
Position.
<4
C
r*
s
01
I
<2>
o
<
< <
>
Z
o X
5 <
s
-J
AC
t
C
eg
<
8
«
©
8
2
si
u
CO
O)
C/2
< g
z
UJ
5
Q- U.
©
CO
<
c <
u.
< Ss
Q
Z
<
o
S
\
<?
(Z) saqoxd
I#J0U319a
luaimo upis
aistuB,}
Dtwioug
UJ
Figure
11.
Sensor and Emitter Orientations.
46
&
&
<A>
(
•
SPEAR-3 SCIENCE MEETING, W.
J.
SCHAFER ASSOC., ARLINGTON VA AUG
6-7 1991
CENTER FOR ATMOSPHERIC AND SPACE SCIENCES UTAH STATE UNIVERSITY
.v
WW
30kfl
L
I
X
coil
0-
,700fcQ
o.5»if;
m
-WW^
1MO
Roftmki
SO
1
Ramped
HV Mode
• Denotes hi|h speed sampling capability
/mm/
VWV^
1MQ
Rtucwii.
col
2kQ
0v:
'700kQ
E
1
Switched
HV Mode
* Denotes high speed sampling cspsbility
-©30kn
Rogovski
Cttl
•:
2kfl
"
VWV^
IMfl
2kn:;
©,
••\700kn
2fiF
E
Quiescent
Mode
* Denotes high speed sampling capibilily
Simplified Schematic Diagrams of
Figure
12.
Schematic of
HV Capacitor Discharge Modes and Quiescent Mode
HV Capacitor Discharge Modes/Quiescent Mode
47
o~
y
•
O)
o
CO
/
*2
55 Z
2
/
/
r^
Lit
C/)
>If)
co
<
h<
=°
*"
CD
E
•5.
o
CC lu
F
0)
—
-SPE
"*
PEA
0)
If
X
o
X
LL
o
o
z
I
H
CO
-*
a.
/
CTJ
LU"""'
"u*
-----co
*-
/
Cfl
9
/
E
-Q.
E
3
o
(0
Ram|
Netw
CM
^*
C^
o
E
£•
3
Q.
CO
2
CO
«
— .--____
- -'
Q
O
\_
>
j£
o
sBbjioa ejeijds
Figure
13.
SPEAR III
Timing of High-speed Data System
48
SPEAR-in
Simplified High-Speed Data Schematic
lMfl
300 kHz
Lowpass Filter
Rogowski
300 kHz
Lowpass Filter
Sphere
Voltage
Current
2kft
700 kfl
±:
(
V
2\if
/77
>
150 kHz
Lowpass
Filter
n7
Figure
14.
Simplified High-speed
49
Data Schematic.
SCP
Current
2.2 in
7.9
Figure
15.
m
Plum Brook Plasma Chamber
50
V.dlllCI
CI
Chamber
Wall
NRI
Boom
video
00'
Camera
J
\
I""
*l
I
Relay
Co
(
1
1
MM
o
|
Pulsci
I
!
in
*1
7\
Cables &
Powei
f
C 01
it
i
a
300'
lange
ol
Powei &
1
|
.
Ibei
Optk.
nsti
umenl
L/_
7\"
Tiber
Optics
I
at ion
lliijh
Voltage
DC.
Supply
LIT
A
Cable &
r
lange
Video
fiber
optics
^\
These items must have Insulation on outer Jacket sufficient
withstand the full potential of the high voltage supply
Figure
16.
Mockup
Physical Configuration.
51
lo
1
<
a:
Q
u
D,
I*
o
O
s
E
c
.2
c
Q
CD
C/3
immiuumnwhunQ
Figure 17
Plum Brook Chamber Suspension
52
O
b
Uj
0>
O
E
>
a>
E
^OT
re
O
CO
ID
Q.
o
£Q. LO
s>
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ex.
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p
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a:
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on
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00
4->
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CM
i—
CO
m
Q.
<\J
(XI
v*—
o
,
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o
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^J-
2
CL
1)
^3
O
id
o
O
*
o
1^
4N
lO
C\J
=3
CO
>
ID
_J
r\3
co
o
-:?
Figure
18.
Floating Probe
53
Waveform
-
Shot 322
c
o
o
E
t
-
^
s
•
CD
O
v0)
1/
a
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n.
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LU
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rr
m
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O
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1
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£
o
m
CM
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<XJ
o
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Figure
19.
Floating Probe
54
Waveform
u,
-
Shot 324.
i<5
CM
4—
co
Cl
> o
o
G
a
to
CO
4—
*-J
0)
to
to
E
E
&
—
o s
CM
1
o>
3
<
j?
H
-
O
o
o
in
i
I O Z
U-
?UJQOW
3
CM
1
1
1- LU Q. DC
LU
CD
O
2
CO
§
<
z
o
£
O
-
o
m
~
o
o
O
X
X
li.
li.
hLAJ
Z
-J
Hi
O
2
H
_j
<
Z
o
o
in
(3
CO
X
0*
x
0)
tf).
li.
8
co
»-
E
p
*•*
JZ
o
g>
o
o
MSB
^^"»
o
m
X
X
oc
<
CM
u_
LU
Q.
0)
li.
h-
o
o
CM
h-
z
CD
o
X
o
m
X
li.
li.
h-
o
o
z
1
1
1
1
D
o
in
CM
o
o
CM
o
in
O
O
c *>
< =>
<
(ium)
Figure 20
i-
epmmv
SPEAR III
55
Original Timeline
(
i
3
n
Ll
Digi sonde
Wallop's
350-
i
1
1
—
——
i
i
1
t
V
V
o
a
2045
2100
o
<#
o
D*
fe§
21 15
300-
JBB
§
2130
*
E
«6»B
250
G
4
n
P<a
„J
** «A
•*
o
Oo
o
O
o
200
_1
I
I
10
I
I
J
l_
10
_
Density (cm
3
)
Figure 21. Wallop's Digisonde.
56
I
L
350
iii—
i
———SPEAR
—
1
i
i
i
1
i
1
1
1
Deduced from LP Data
—
2115 Digisonde Data
O
300
i
1
1
1
1
250
«
200
13
150 -
100
-
O
O
50
_i
'
i
i
1
1
O O o
_i
1
i
10'
i
i
10*
_i
i
'
'
i
i
i
(cm
i_
i
'
10'
10
Electron Density
'
i
10
)
_i
i
i
i
i
i
i_
1000
100
Electron Plasma Frequency (kHz)
-i
i
'
>
'
i_i_
_i
i
i
i
i_i_
i
10
100
Proton Plasma Frequency (kHz)
_i
i
i
i_
_i
i
i
i
i
i
i_
_i
10
1
Oxyqen Plasma Frequency (kHz)
Figure 22.
SPEAR III Electron Density
57
i
i
i_
s
SPEAR-3
SAIC
>
CD
S
1
05
110
2.20
1.94
1.68
1.41
>
en
CD
&
100
-fi
1.15
o
0.89
0.62
T^
1
I
0.36
mm
T
105
115
110
M.E.T.
Figure 23
58
0.10
(s)
SPEAR III SAIC
-
100 Seconds
SPEAR-3
SAIC
i
i
i
i
>
03
en
i
—
i
i
CD
C
Ld
k
120
i
125
I
130
I
135
j
'.
i
2.20
1.94
1.68
1.41
>
in
0>
c
& 1000)
c
O
0.89
0.62
0.36
i
120
125
130
M.E.T.
Figure 24.
_L
t
1
1
135
(s)
SPEAR III SAIC
59
1
-
120 Seconds.
1
1
r
0.10U
SPEAR-3
SAIC
>
en
l.
CD
C
140
145
150
155
i
i
i
i
2.20
1.94
i;
1.68
1.41
>
CO
CD
I 1.15*
& 100L-
en
CD
c
_o
LJ
0.89
0.62
0.36
j
140
145
(s)
SPEAR III SAIC
60
0.10
r
155
150
M.E.T.
Figure 25
mil
r^—
-
140 Seconds
I
SPEAR-3
160
165
J
:
i
SAIC
170
175
170
175
L
>
CD
£
100
C
i
160
165
1
r
M.E.T. (s)
Figure 26.
SPEAR III SAIC
61
-
160 Seconds
10
-1
SPEAR-3
c
r
——
i
~i
-
i
170.770
SAIC
1
1
1
1
1
1
1
—
I
1
1
r
"1
1
1
3
•—• Low Energy Channel
0—0 High Energy Channel
10
-2
Range = 23.72 - 367.82 eV
-3
dens = 8.24e + 05 cm
LSF: Energy
kT =
95.98 eV
Charging Peak
= " 1000.0 V
4>
dens = 7.00e + 03
kT
200
400
800
600
Energy (eV)
Figure 27.
SPEAR III SAIC
62
170.770.
=
-2
crrr
55.00 eV
1000
1200
10
SPEAR-3
-1
1
1
n
r
1
r
T
t
170.962
SAIC
r
1
t
1
r
—
I
1
1
r
-i
1
i
—
•—• Low Energy Channel
°~° High Energy Channel
LSF: Energy
Range =
dens = 8.37e+05
kT =
cm
23.72 - 367.82 eV
-3
94.93 eV
Charging Peak
=
880.0 V
'
<t>
10
6
u
L
200
400
600
Energy (eV)
Figure 28
SPEAR III SAIC
63
170 962.
800
1000
la_
1200
10
-1
—
c
SPEAR-3
i
1
170.994
SAIC
1
r
3
1
1
i.
lo_
•—• Low Energy Channel
t>—°
High Energy Channel
LSF: Energy
23.72 - 367.82 eV
Range =
dens = 8.80e + 05 cm"
kT =
83.76 eV
3
Charging Peak
"
$ =
880.0 V
V)
c
V
dens = 2.00e + 03 cm"
T3
kT
a>
=
25.00 eV
u
O
Q.
W
<D
V)
D
s:
200
400
600
Energy (eV)
Figure 29
SPEAR III SAIC
64
170.994
800
1000
1200
10
-1
—
c
SPEAR-3
i
1
r
-i
1
171.634
SAIC
r
t
1
r
-i
1
1
—
•—• Low Energy Channel
°~° High Energy Channel
Range = 23.72 - 367.82 eV
-3
dens = 8.45e + 05 cm
kT =
42.76 eV
LSF: Energy
Charging Peak
4>
=
"
dens =
kT
10
=
580.0 V
1
.00e+04 cm
-3.
13.00 eV
_6
u
200
400
600
Energy (eV)
Figure 30
SPEAR III SAIC
65
171.634
800
1000
1200
10
-1
—
c
SPEAR-3
i
1
r
171.666
SAIC
i
i
r
~i
1
1
—
•—• Low Energy Channel
0—0 High Energy
Channel
LSF: Energy
Range =
dens = 7.76e + 05
kT =
47.19 eV
cm
23.72 - 367.82 eV
-3
Charging Peak
= ' 580.0 V
<t>
en
c
0>
dens = 1.00e+04
d>
kT
D
=
cm -3.
13.00 eV
U
D
Q.
CO
(/)
D
-C
a.
10~ 6 ^
L
200
400
600
800
1000
Energy (eV)
Figure31
66
SPEAR III SAIC
171 666
k>-
1200
SPEAR-3
Floating Probe
10000
8000
-£
—
6000
u
c
3
cr
4000
—
2000
—
wl'iMi
0.000
Novol Postgroauate School
0.002
0.004
0.006
0.008
14-Mor-1994 19:22:16.00
Figure 32.
SPEAR III
Floating Probe Spectragram.
67
SPEAR-3
Floating Probe
10. OOF
Time = 175.462
1.00
N
X
E
>^
0.10
<L)
0.01
:-
_i
i
i_
4
6
Frequency (kHz)
Figure 33
SPEAR III
Floating Probe Line Plot
68
10
SPEAR-3
Floating Probe
10000
8000
X
6000
u
c
cu
D
cr
(D
4000 —ftHm-t:
2000
_
0.000
Naval Postgraduate School
0.002
0.004
0.006
0.008
0.010
0.012
23-Mor-1994 17:47:07.00
Figure 34.
SPEAR III Floating Probe
69
Spectragram of Entire
Flight.
SPEAR-3
125
Skin Current
ME T
0.0000
Novoi Postgroduote School
0.0002
135
130
0.0004
(s)
0.0006
0.0008
i4-Mor-1994 17:17:05.00
Figure 35
SPEAR III
Skin Current Spectragram
70
0.0010
SPEAR-3
0.0100
Skin Current Probe
~Z
Time = 136.714
CD
T3
3
0.0010
0.0001
J
i
I
I
)0
I
1—1
I
I
1
'''''
0.10
I
I
l_J
I
l_l
0.20
I
L
0.30
0.40
Frequency (MHz)
Navol Postgraduate School
14-Mor-1994 17:41:49.00
Figure 36.
SPEAR III
Skin Current Probe
71
@ 134.714.
0..
SPEAR-3
Skin Current Probe
0.0100 ~j
0.0001
0.00
0.10
0.20
0.30
0.40
Frequency (MHz)
Nova: Postgroduote School
14-Mar-1994 17:41:51.00
Figure 37.
SPEAR III
Skin Current Probe
72
@
137.226.
0.50
SPEAR-3
Skin Current Probe
0.0100
0.0001
0.00
0.10
0.20
0.40
0.30
Frequency (MHz)
Novo! Postgraduate School
14-Mar-1994 18:53:54.00
Figure 38.
SPEAR III
Skin Current Probe
73
@
196.753.
0.50
SPEAR-3
~
0.50
0.40—
I
N
0.30
I
I
I
I
:
Skin Current Probe
I
I
I
I
i
I
I
l
l
I
I
l
l
I
:
_ — —«
X
I
I
—
silttfr-.
Q
-•-¥
o
c
CD
3
cr
0.20
BliP-H
rt.
i
.
.,f.
r
S
.;
;
-^"(8
get
-
«r-**«
;r
«
~
l»
= 11
.
s
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-
rp-
i
-I
*
!!
:
?r
r £ p ^.
r*r
,i,,
.,
t
iiif
ri-§ e?S
-F-
?
-'
MET
x
0.20
Naval Postgraduate School
Figure 39.
0.30
.-}
¥>
!-
,
.
'
10
•I
PP~
ifffr;
*
{
-Si-
«*
».
,*
M 1
.
S*«X)fe
ahvp
* ««>(
400
(s)
0.50
0.60
23-Mar-1994 19:08:33.00
SPEAR III
Skin Current Probe Spectragram of Entire Flight
74
I
Ft;
&«§:»» -
-3
0.40
!
f-H-^t
300
200
100
tf^Sl*
I
Jf-
1
l;, ,,
i
il
1*1-
.
1".
".
0.10
0.00
.
.^.*>.
.
(a
i:l
kr$m
p
il
-
_..
..,*..
.
500
LIST OF REFERENCES
George E "Science and technology of directed energy weapons," Chairman Review
American Physical Society, 1987 pp. 21-22.
Committee Washington
'Pake,
,
DC
.
2
3
4
5
:
Ibid.
Pake, pp. 23-34.
Ibid
Neubert,
Raitt,
W.
J.,
T
,
Mandell,
M.
J.,
Sasaki, S
Meyers, N. B., Oyama, K.
I
,
,
E Banks,
Gilchrist, B.
& Katz,
I.,
,
P.
M,
Williamson, P
R
"the sheath structure around a negatively
charged rocket payload," Journal of Geographic Research
95(A5),
May
1,
1990.
pp. 6156-6158.
6
7
Pake, pp. 23-34.
Allred,
D B
ionosphere,"
8
9
,
et. al.,
IEEE
"The
SPEAR I
experiment high-voltage effects on space charging
Transactions on Nuclear Science
35(6)
December
6,
1988
p.
in
the
1386.
Neubert, pp. 6157-6164.
Allred, p.
10
Spear HI Preliminary Design Review Data Package
June
11
13
14
15
16
3,
3,
.
Logan, Utah: Utah State University
1991.
SPEAR HI
June
12
1386
Preliminary Design Review Data Package
1991
SPEAR HI
Preliminary
...
.,
pp. 4-5.
Ibid.
Ibid.
Ibid.
Ibid.
17
Ibid.
75
.
Logan, Utah: Utah State University,
18
From
19
20
21
22
23
24
a conversation
between the author and Douglas W. Potter,
May
12,
1994
Ibid
Ibid
Ibid
Ibid
Ibid
SPEAR HI
Post Flight Data Review Meeting
76
Logan, Utah: Utah State University
INITIAL DISTRIBUTION LIST
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SAIC
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SAIC
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