Author(s)
Yoon, Sang Il.
Title
Definition study and model for a tethered sounding rocket.
Publisher
Monterey, California. Naval Postgraduate School
Issue Date
1988
URL
http://hdl.handle.net/10945/23008
This document was downloaded on May 04, 2015 at 23:45:26
NAVAL POSTGRADUATE SCHOOL
Monterey, California
V
-6 >1^-S
MODEL FOR
A TETHERED SOUNDING ROCKET
DEFINITION STUDY AND
by
Yoon, Sang
11
December 1988
Thesis Adviser:
Approved
for public release;
Richard C. Olsen
distribution
is
unlimited
T242452
Unclassified
OF
.UHIIY CLASSIFICATION
fHIS
PAGE
REPORT DOCUMENTATION PAGE
REPORT SECUHIIY CLASSIFICA llOt^
RESTRICTIVE
lb
MARKINGS
Unclassified
SECURITY CLASSIFICATION AUIHOHIIY
DISTRIBUTION/AVAILABILITY OF REPORT
3
Approved for public release;
DECLASSIFICATION
/DOWNGRADING SCHEDULE
distribution is unlimited
PERFORMING ORGANIZATION REPORT NUMBER(S)
NAME OF PERFORMING ORGANIZATION
OFFICE
6b.
Naval Postgraduate School
and
{City, State,
SYMBOL
NAME OF MONITORING ORGANIZATION
;a.
applicsble)
{If
ADDRESS
MONITORING ORGANIZATION REPORT NUMBER(S)
5.
61
Naval Postgraduate School
•
ZIP Code)
ADDRESS
7b.
ADDRESS
(City, State,
and
8b. OFFICE
{If
and
ZIP Code)
Monterey, California 93943-5000
Monterey, California 93943-5000
NAME OF FUNDING /SPONSORING
ORGANIZATION
(City. State,
SYMBOL
9.
PROCUREMENT INSTRUMENT IDENTIFICATION NUMBER
applKable)
ZIP Code)
10.
SOURCE OF FUNDING NUMBERS
PROGRAM
WORK
TASK
NO.
PROJECT
NO.
ELEMENT NO.
UNIT
ACCESSION NO.
TITLE (Include Security Clamlication)
Definition Study and Model for a Tethered Sounding Rocket
PERSONAL AUTHOR{S)
I.
TYPE OF REPORT
Master's
13b
Thesis
TIME
COVERED
FROM
14.
DATE OF REPORT
(Year.
Month, Day)
15
PAGE COUNT
109
1988 December
TO
SUPPLEMENTARY NOTATION
The views expressed in this thesis are those of the author and do not
fleet the official policy or position of the Department of Defense or the U. S. Government
COSATI CODES
FIELD
GROUP
18.
SUB-GROUP
SUBJECT TERMS (Continue on reverse
if
necessary
and
identify by block
number)
Tether deployment system, Plasma contactor Hollow Cathode,
Flight program, Model of system electrodynamic behavior study
ABSTRACT {Continue on reverse
t
if
i^ecessary
and
identify by block
number)
.,...„,.,.„,„„,„...,..„„.,.,...
l„.„„.„_„....
\) [roin a satellite into the ambient plasma, and llie return patli through the plasma. The devices to be
jseJ to establish electrical contact with the plasma are hollow cathode plasma sources, using Xenon
Hopeilent. Two satellite sections will be connected by a l()()-m cable, which can be used to bias the two
satellites with respect to each other. The current through the connecting cable will be monitored, along
return
kvith the particle fluxes to the satellites. A fiber-optic technique will be used to measure the
:urrent flowing through the ambient plasma. Electric fields, densities, and plasma waves will be
iionitored with floating probes. We intend to launch this payload from Wallops Flight Facility two years
ilter
funding begins.
I
DISTRIBUTION/AVAILABILITY OF ABSTRACT
El UNCLASSIFIED/UNLIMITED
SAME AS RPT
;». NAME OF RESPONSIBLE INDIVIDUAL
21.
22b TELEPHONE {Include Area Code)
Richard Christopher Olsen
Professor
83 APR edition may be uied
l>FORM1473,8nMAR
All
ABSTRACT SECURITY CLASSIFICATION
DTIC USERS
(408)
until
exhausted
other editionj are obsolete
22c.
646 - 2019
Unclassified
SYMBOL
61 Os
OFFICE
SECURITY CLASSIFICATION OF THIS PAGE
A
U.S.
0«>ifnm«nl
Prlnllni
onicd
Unclassified
If it—•Ot-143
Approved
for public release; distribution
is
unlimited.
Model for a Tethered
Sounding Rocket
Definition Study and
by
Yoon, ^ang II
Major, Republic of Korea Army
B.S., Republic of Korea Military Academy, 1979
Submitted
in
partial fulfillment of the
requirements for the degree of
MASTER OF SCIENCE
IN PHYSICS
from the
NAVAL POSTGRADUATE SCHOOL
December 1988
ABSTRACT
The
coupling
HOCAT
of
large
experiment
currents
is
(up
a
to
sounding
1
from
A)
a
plasma, and the return path through the plasma.
establish electrical contact with the
Two
Xenon
cable,
which can be used to bias the two
The
current
through
the
the
return
densities,
to
launch
current
flowing
and plasma waves
this
satellite
sections
will
satellites
connecting cable will
particle fluxes to the satellites.
A
The
the
into
devices
to
ambient
be used to
be connected by a 100
the
be monitored,
along with the
ambient
plasma.
Flight
begins.
ni
Facility
measure
Electric
be monitored with floating probes.
payload from Wallops
m
with respect to each other.
fiber-optic technique will be used to
through
will
sattelite
measure the
to
plasma are hollow cathode plasma sources,
using
propellent.
payload
rocket
two years
We
after
fields,
intend
funding
TABLE OF CONTENTS
I.
INTRODUCTION
A.
A HISTORY OF TETHER CONCEPTS
1
B.
ELECTRODYNAMIC TETHERS
6
C.
D.
H.
1
1
Basic Concept
6
2.
Electrodynamic Tether Applications
7
3.
Tether Experiments
8
a.
Tethered Satellite System-1 Experiment
b.
US-JAPAN
c.
MAINHK
Charge Rocket Experiment
Rocket Experiment
PLASMA CONTACTORS
8
10
11
13
1.
Basic Concept (Requirements and Future Mission)
13
2.
Plasma Contactor Flight History
15
a.
Balloons
b.
Electron
c.
Neutral Gas Releases
d.
Ion
Beams
20
e.
Ion Engine
21
f.
Hollow Cathodes-
15
Gun
17
18
••
29
PROPOSED EXPERIMENT
30
TETHER DEPLOYMENT
31
BASIC DYNAMICS
31
A.
IV
B.
III.
IV.
V.
HOCAT DEPLOYMENT OBJECTIVES
1.
US-JAPAN
2.
MAIMIK
3.
CANADIAN OEDIPUS
4.
Tethered Satellite System-1 Deployment System
Joint Experiment
32
Deployment System
34
Rocket Experiment Deployment System
38
Experiment Deployment System
39
42
PLASMA CONTACTOR HOLLOW CATHODES
43
A.
LABORATORY STUDIES
43
B.
THEORETICAL STUDIES
50
C.
FLIGHT EXPERIMENT DATA
52
EXPERIMENTAL FLIGHT PROGRAM
54
A.
ENVIRONMENT
54
B.
LAUNCH VEHICLE
60
C.
DEPLOYMENT SYSTEM
62
D.
FLIGHT SCENARIO
63
E.
PAYLO AD
•
63
1.
Hollow Cathode
2.
Experiment Control and
3.
Electric Field Instruments
66
4.
Thermal Plasma Analyzer (Ions and Electrons)
67
5.
Direct Current Measurements (Current Loop)
71
6.
TV Camera
73
7.
Magnetometer
74
64
PCM
Encoder
MODEL OF SYSTEM ELECTRODYNAMIC BEHAVIOR
A.
•
PROBE THEORY
65
75
75
B.
C.
VI.
LIST
MODIFICATIONS DUE TO PLASMA CONTACTOR
86
1.
Electron Emission
86
2.
Election Collection
87
SUMMARY
89
CONCLUSIONS AND RECOMMENDATIONS
OF REFERENCES
90
92
INITIAL DISTRIBUTION LIST
97
VI
LIST
OF FIGURES
Fig.
1.1
Gemini-Titan Space Vehicle Configuration
2
Fig.
1.2
Agena Target Vehicle Configuration
3
Fig.
1.3
Gemini Spacecraft
3
Fig.
1.4
Tethered Spacecraft/Target Vehicle
3
Fig.
1.5
Effect of
Damping on the
4
Fig.
1.6
Effect of Relative Velocity
Fig.
1.7
Electrodynamic Tether System
7
Fig.
1.8
TSS-1
9
Fig.
1.9
Payload Configuration of fourth
Fig.
1.10
Individual Liquid Metal Ion Emitter
14
Fig.
1.11
Collector Screen in the Deployed Configuration
16
Fig.
1.12
Configuration of
ATS-5
Ion Thruster
22
Fig.
1.1.3
Configuration of
ATS-6
Ion Engine
23
Fig.
1.14
Spacecraft Plasma Current Diagram of
Fig.
1.15
Distribution of Emission Currents
27
Fig.
1.16
Excess Neutralizer Emission
28
Fig.
2.1
Dumbbell Tether Model
31
Fig.
2.2
Hollow Cathode Configuration
33
Oscillation of Tethered System-
on Motion of Gravity-Grad-
System
Satellite
TPE
Experiment
SERT-II
Wire Extension Velocity vs Time
•
•
•
5
11
26
Fig. 2.3
Plot of Tether
Fig. 2.4
Plot of Tether Separation Distance vs
Fig. 2.5
Deployment Velocity
Fig.
2.6
Plot of Separation Velocity vs
Time
40
Fig. 2.7
Plot of Separation Distance vs
Time
41
of fourth
Vll
Time
US-JAPAN
36
37
Experiment
38
Fig. 2.8
TSS-1 Reel and Boom Mechanism
Fig.
3.1
Perturbing Effect of
Fig.
3.2
Contactor Anode-to Ambient Plasma Potential Difference
vs Electron Emission Current
46
Plasma Potential Contour around Hollow Cathode
47
Fig. 3.4
Hollow Cathode Ignited mode Configuration
48
Fig.
3.5
Current to Contactor vs Bias Voltages
49
Fig-
3.6
Bias-to-Plasma Potential Difference
49
Fig.
3.7
Current vs Contactor Potential
50
Fig. 3.8
2-Dimensional Cut
52
Fig.
Ionospheric Electron Density (daytime)
56
Fig. 4.2
Ionospheric Electron Density (nighttime)
56
Fig- 4.3
SPEAR-1
Neutral Pressure Data
57
Fig. 4.4
Electron Temperature and Thermal Velocity
58
Fig- 4.5
Ion Temperature,
Fig. 4.0
Debye Length
Fig- 4.7
Electron Collision Frequency and
Fig. 4.8
NASA
Fig- 4.9
HOCAT
Fig. 4.10
Hollow Cathode Design
64
Fig- 4.11
Dimensions of Hollow Cathode
64
Fig. 4.12
UAH
66
Fig. 4.13
Electric Field and Density
Fig.
Fig.
3.3
4.1
4.14
Fig. 4.15
42
Vacuum Tank Wall
45
Mass and Thermal Velocity
59
Mean
Sounding Rocket Performance
Black Brant
58
X
Payload
Fast Event System
Data
Free Path
59
60
61
68
Electrostatic Analyzer
69
ESA
69
Included Mass Resolution
viii
Fig. 4.16
ATS-6 ESA
Fig- 4.17
Spectrometer Results from
Fig. 4.18
Principle of the Current
Fig. 5.1
Plot of Electron Current Density vs Positive Potential
77
Fig. 5.2
Plot of Ion Current Density vs Negative Potential
79
Fig. 5.3
Plot of Bias Potential vs Current Density
80
Fig. 5.4
Schematic Diagram of System
81
Fig. 5.5
Sphere Current Data from
Fig.
Plot of -r- vs Positive Potential
5.6
70
SPEAR-1
Loop Measurement
SPEAR-1
72
73
83
84
10
Fig. 5.7
Plot of Radius vs Bias Potential
85
Fig. 5.8
Plasma Cloud Expansion Model
88
IX
ACKNOWLEDGEMENTS
wish
I
to
Richard
Professor
Gnanalingam
study.
I
for
my
Christopher
for the
and
gratitude
Olsen
and
instruction, guidance
appreciation
my
second
to
thesis
reader,
advisor,
Professor
and friendly advices throughout
S.
this
'
wish to thank the numerous scientists
this
John
express
Figures and data
work, particularly, Dr. Ira Katz, S-Cubed; Dr. Paul Wilbur and Mr.
Williams,
Alabama
who contributed
Colorado
in Huntsville;
State
University;
Mr. John Patterson,
Dr.
Roy Torbert,
NASA/LeRC; and
University
of
Dr. John Raitt,
Utah State University.
Finally,
their
many thanks
to
my
wife,
Hye-Soog and
my
daughter, Ye-Ji, for
love and being healthy and patient for two and half years in Monterey,
California.
I.
INTRODUCTION
A HISTORY OF TETHER CONCEPTS
A.
The concept
of using a long thin structure in space
was
first
documented by
Tsiolkovslcy in 1895 in the form of a tower reaching from the Earth's equator to
beyond geostationary
the
altitudes. This grandiose idea has received further attention in
modern space era
for the
purpose of
have been developed under the
some
[Ref.
scaled
down but
still
titles
payloads into space, and concepts
lifting
"space elevator", and "orbital tower", with
ambitious variations proposed in the 60's and
70's.
1]
The Smithsonian Astrophysical Observatory (SAO) proposed and
rigorous
investigation
of
the
Shuttle— borne radio physics
feasibility
facility
of
tether
that included
thin—wire antenna was proposed by Mario D. Grossi
recognizing the potential of
more general uses
systems.
a
vertical,
G. Colombo,
of the shuttle—borne tether, provided
He
feasibility.
additional applications, including a scheme for the
measurement
heights.
long,
in 1972. In 1974,
with a rigorous analysis the proof of dynamic
at orbital
example,
For
km
a 20—100
initiated the
Since 1974, a series of investigations has
identified
several
of gravity gradients
made
it
possible to
explore several critical aspects of feasibility in such areas as tether dynamics, tether
electrodynamics,
tethered
subsatellite, etc. [Ref.
The "Space
constellations,
tether—induced
dynamic noise
in
the
2]
elevator"
capable of crawling in a controlled manner along a
deployed tether appears to have a number of potential Space Station applications.
These include
a variable microgravity laboratory, transportation unit
between two
tethered bodies, system center—of—gravity manageinpiit, tether inspection and repair
device, and carrier foi re—entiy probe.
The
space
first
occurred
undertaken
for
actual demonstration of a thin tether connecting two vehicles in
the purpose
ol
during the Gemini program.
mode
Atlas— Agena
D
The Gemini program was
Two
the
of
were explored
rotating tethered vehicles.
The other
separate missions conducted experiments
manned Gemini
spacecraft
to
a
spent
booster stage.
The Gemini XI
flight
was launched on September
Gemini—Titan Space
configuration of the
1.1.
One mode was
attachment
tethered
196(3.
of tethered space vehicle operations
studied gravity gradient effects.
requiring
in
space—flight capabilities during the period between
Two modes
Mercury and Apollo.
program
Gemini
the
in
12,
vehicle and Spacecraft
The launch
1966.
is
shown
in
Figure
Figure 1.2 shows the Agena target vehicle configurations. Figure 1.3 shows the
Gemini spacecraft
module. [Ref.
configuration
which
is
composed
adapter
of
and re—entry
3]
SPACECRAFT
• WIIGHT
SPACECIAET
.ArflOX 1000 US
LAUNCH VEHICLE
• IINCIH 90 ri
• FUtL
.
.•UNO OF UDMH
ANOHTOKAZINE
lAUNCH
109 II
VEHICIE
• OXIDIZER
STAGE
*NIIROCfH IITIOXIOE
• IHIUSI
.StACE
I
OXIDIZER
71
I
430,000
US
It
M
•EUEl
-STAGE U
100.000
US
SPAdE VEHICLE
• WEICHI AT llTTOrr
-
n
m
OVER 340.000 US
Fig. 1.1
10
Gemini—Titan Space Vehicle Configuration[Ref.
2
3]
AGE^M
^-^
• paiMAKY noruiSiON
- run .UDMH
-OKIDI2II .IRfNA
- IMIUST 16000 IIS
ACtNA
3
n
3; FT
.
• SICONDARY PIOPUlStON
- rutL UDMH
- 0)IIOI7ER MON
-
TMtUST
2
2
ATUAS SLV
104 fT
fNGINlS
tNCIN(S
'u«
ACH
lACH
200 II
(^ 16 II
f
3
• riOPUlSlON
- full if \ irnosiNii
- OKlDt2EI IIOUIO OXYCtN
- tHRUST 390.000 II
TARGET SPACE VEHICLE
• WDGHI AT llfT.OM
-
OVU
2/3.00r IIS
Agena Target Vehicle Configuration[Ref.
Fig. 1.2
3]
-tOUIfMf NT SfCllON.
ADAriER
-lETROCRAOE SECTIOK
ADArUR
CAIIN
ADAflER
BffNtay MQPuif
Fig. 1.3
Gemini Spacecraft
Figure
vehicle.
1.4
siiows
The unstretched
Fig. 1.4
the
tethered
configuration
tether length was 100
Dacron webbing with a breaking strength
of 13 seconds
gave
rise to
both vehicles was 10
Tethered Spacecraft /Target Vehicle
of
ft.
of
The
tether was constructed from
800 pounds. The total
a final spin rate of 0.9 deg/sec.
ft.[Ref. 4]
3
Gemini spacecraft /target
lateral thrusting
The maximum diameter
of
On
Geiiiiiii
XI,
a
maneuver was
s|)iri— iii)
evidence of siguilicani, cable—dynamics
were not
behavior
critical in the rotational
siiown to be feasible).
effects.
The procedure
for
(i.e.
successfully
conducted with no
This confirmed that cable dynamics
a spin—stabilized configuration was
spinning up the tetheied spacecraft /target
vehicle system consisted of backing the spacecraft
away from
the target vehicle until
the tether was almost taut, then firing the translational thrusters to provide thrust
on the spacecraft normal to the
line
between the vehicles. The spin—up moment on
the system was supplied entiiely by the spacecraft translation—control system.
I'^igure
damping
is
1.5 illustrates the
quite effective
in
spin— up behavior with and without damping. The
eliminating a slack/taut tether oscillation.
half of Figure 1.5 presents a time history of tension in the tether,
of the spacecraft relative the target vehicle.
It
The bottom
and the yaw angle
can be seen that the tether tension
was strongly affected by damping.
Wilhoul damping
With damping
600
]T- \
.411(1
1.,
.
,l
£;
200
imm^3.0
100
150
200
lime, sec
Fig. 1.5 Effect of
Damping on
the Oscillation of Tetheied System[Ref. 3]
The
Gfiiiiiii
XII
was launched on
experiinoiit
flight
11
november,
(Jemini XII sljuiied Uie feasibility of gravity gradient stabilization.
initialized
by
was followed by an active stabilization
spacecraft. This
the target— vehicle attitude control system.
of
a slightly taut tether,
circular
orbit,
and
a
initial
The proper
the
of the target vehicle using
starting conditions consisted
alignment along a local vertical, approximately
velocity
relative
The system was
and attitude thrustering maneuvers by
translational
various
1966.
of
ft/s
0.1.38
(or
a
100
tethered
ft
spacecraft/vehicle combination. Figure l.G illustrates typical result obtained from
the point mass analysis on the sensitivity of the system motion to initial relative
velocity
between
orientation
be
point
the
masses.
The
vertical.
The main requirement was
result
of
this
experiment
stabilization were that with a i)roperly functioning control
for
that
the initial
gravity
gradient
system operated by the
crew, the initial conditions necessary for the tethered system could be accomplished
with relative ease.
demonstrated
in
The gravity— gradient
stabilization orientation
was successfully
Gemini XII.
•
0.015 (l/sec
Velocity error*
.060 (Usee
Velocity error
Velocity error
cr
•
.
150 It/sec
- JOr
fe
>
20
\^
//
//
^>
8
\
\'
Pertecl
start
^
™
-20
2^-40
<
1000
2000
3000
4000
.
5000
6000
Time, sec
Fig. l.G Effect of Relative Velocity
This
coiicppt
has
evolved
into
on Motion of Gravity-Grad[Ref.
the
set
of
Tethered
experiments. Elertrodynamic applications will be tested with the
Satellite
3)
System
upward—deployed
TSS— 1
experiment
current
interest
for the first sliuttle
include
borne tether experiment. The general areas of
experimentation,
scientific
transportation
applications,
tethered platforms, and propellant storage and transfer. Tether shows particular
promise
for resolving the
problem of de—orbiting
station—a task shuttle will not be able to meet.
in
wasting mass from the space
Many
of these areas are
now
receiving serious attention, and the concepts are progressing from the initial idea
generation stage to the detailed analysis and proof-of—concept testing stage.
There
is
a gravity gradient for a vertically aligned tether system.
satellite experiences
a greater gravitational force than the higher one, which causes
tension on the tether.
vehicles
In
require proper
an experimental or operational system, the tethered
starting
conditions so that
limited in amplitude. In particular,
about
local
established, a
B.
The lower
be
vertical
number
limited
in
will
be
desirable that oscillations of the system
is
it
subsequent motions
amplitude.
Once the system
is
properly
of electrodynamic applications are possible.
ELECTRODYNAMIC TETHERS
1.
Basic Concept
A
connected
tethered payload system in which two separate payloads in space are
with
an insulated
conductive wire forms
an electrodynamic tether.
Electrodynamic tethers make use of interactions between a moving conductor (the
tether wire), the earth's magnetic field, and the ambient
plasma
for propulsion
and
braking, power generation, science and communications.
A (vxB)-L
potential difference develops across the tether due to
motion through the geomagnetic
thrust electrical contact
field.
In order to generate electrical
must be made with the ambient plasma
The
to close the circuit to the environment.
conduction of electrons along
B—field
tether circuit
lines in the
6
is
its
power or
at each tether
end
made complete by
environment. [Ref.
5j
Figure 1.7 shows the elcctrodynamic tether system with elect rodynamic
drag. Typical ioiiosplieric values, for an orbiting system, result in induced voltage of
several kilovolts lor a
5— lU km system.
ELECTRONS
•
•
vx B
VOLTAGE = TETHER E*dS
E=
J
=
«
©VECTOR SENSE
iu
X B
^^
~
""'^ vector along tether)
dS
OF B-FIELD
POWER
•
TETHER = RATE OF SYSTEM
TOTAL MECHANICAL ENERGY LOSS
•
•
V ~ 8 km/sec
B ~ 0.45 G
•
E ~ 200 V/km
IN
DRAG VECTORS
ELECTRONS
Fig. 1.7
Electrodynamic Tether Systern[Ref.
5]
Typical ionospheric densities of lU'^— 10^ electrons/cm^, and collecting
areas
of
a few square meters, suggest
collected, to give a net
emitted
2.
at the
power
same current
that
of kilowatts.
currents of a few amperes can be
Of course,
level at the opposite
end
this requires that electrons be
of the tether.
Electrodynamic Tether Applications
A number
of possible applications exist.
By applying a voltage between
two payloads connected with a conductive wire, the response
of ionospheric
plasma
to potential differences
between collector and plasma can be studied. By regulating
the current collected (emitted) at the ends, the tether current can be controlled.
By
modulating the tether current, low frequency radio waves with long wave length
such as Whistler and Alfven modes, can be excited.
An
electrodynamic
system
tether
can
be
used
for
electrodynamic
propulsion and braking by using the interaction force between the tether electrical
current-
and the ambient magnetic
electrical
energy
power generator up
by the orbiting system
lost
science,
to
an
electrodynamic
field.
MW
is
tether
In principle
it
can be applied as an
mode, the kinetic
In this (braking)
level.
converted to electrical energy. For ionospheric
system
can
be
used
to
induce
controlled,
large—scale perturbations which can be tracked and measured.
Tether Experiments
3.
Besides the early Gemini experiments, relatively few experiments have
been conducted. The major planned experiment
a.
for
TSS— 1 which
is
described next.
Tethered Satellite System— 1 Experiment
Space shuttle engineering
planned
is
the early
1990s
using
tests
the tether
of the space tether concept
satellite
TSS
developed jointly by Italy and the US. The
goal
is
system which
is
are
being
to test the feasibHty of
deploying, controlling and retrieving a tethered satellite from space shuttle, as well
as
to
demonstrate the system's usefulness
for scientific
The TSS
research.
will
provide a unique, reusable facility for conducting research based from the Space
Shuttle and for conducting proof—of—concept demonstrations of tethered satellite
systems in space.
TSS-1
mission, planned
now
for late 1992, will
be an engineering
verification flight. It will also perform limited electrodynamic science. Figure 1.8
shows
the
configuration
of
TSS— 1
satellite
8
system
in
space.
A
20
km—long
electrically
conductive
tetlirr
be deployed above the shuttle (spaceward) to
will
study the electrodyiiamic properties of the ionosphere as well as conduct plasma
physics experiments.
This mission
will test
the feasibility of generating electricity with
the insulated conductive tether. This process could be an important source of
for
future spacecraft and space station.
As
the shuttle
moves
power
in earth orbit,
the
tether will cut through the planet's magnetic field and interact with the ionospheric
plasma.
TSS-Sa.^flife
r.«*v
•''ih-^
Fig. 1.8
.
TSS-1
Satellite
9
System
[Hef. 6]
The
positive deployed
electrically
from the ionosphere. These
will
plasma source on the shuttle
satellite
will
collect
electrons
be emitted back into space by an electron gun or
This causes a net current to flow upward
orbiter.
through the conductive tether. The intended current
is
0.75 A, voltage of
up
to 3
power of a few kW. The tether resistance
is
about 0.2 fi/m which
will
kV,
for a
limit the conducted current in the 20
b.
US—JAPAN
km
cable to the above values. [Ref.
6]
Charge Rocket Experiment
Tethered sounding rocket experiments have been underway since
1980 as part of a joint
Its
objective
is
US-JAPAN
to obtain technical
program (Tethered Payload Experiment; TPE).
and
scientific
data in support of the space shuttle
electrodynamic tethered subsatellite experiments on the space shuttle Tethered
Satellite
active
System— 1 mission. The
goal of the rocket
program has been
to perform
plasma experiments, by ejecting an electron beam from the tethered
mother—daughter payload system.
the
In
third
and fourth rocket
flights,
the
conductive tether wire was successfully deployed more than 400 m. Figure 1.9 shows
the payload configuration of the fourth Tethered Payload Experiment. These latter
pay loads have been dubbed charge
The
in
DC
1
and charge
2.
electron gun generated a narrow electron
beam
of
1
kV, 80
mA
and multi—pulse modes. The floating/Langmuir probe array consisted of 4
cylindrical probes installed on a rod every 25 cm.
sensitivity color film were installed.
firings to
observe the
beam
Two
35
mm
cameras with high
One was operated synchronized with
trajectory.
the
beam
Another was synchronized with flashing of a
strobe light to illuminate the reflective tape attached to the daughter rocket.
Two
photometers
filtered at
emission from the interaction of the electron
charge sheath around the rocket surface.
3914
X
were used to detect the light
beam with
the atmosphere and the
The charge probe was
10
installed to apply a
high voltage up
installed
\.o
500
V
between the two payloads. The deployment system was
with deployment monitoring system of 0.1
m
was found that the tether wire acted as an antenna. Current
in
on the daughter pay load
resolution.
It
the tether wire apparently resulted
the generation of waves
beam emission up
range. Vehicle charging due to
In
in
the series of the experiments.
Beam
to 80
mA
energies up to 500
in
the
MHz
was repeatedly measured
eV were
used.
emission experiments displayed clear evidence for the ignition of a
discharge, as measured by photometers and
wave
frequency
receivers. [Ref.
The 80
mA
beam plasma
7]
MOTHER PAYLOAD;
391* I
PHOTOMETER
3SMM CAMERA
PLP PROBE
ELECTRO STATIC ANALYZER
MOTHER CHARGE PROBE SENSOR
STROBE
TED SENSOR
3SMM CAMERA
X
3SI4 X
FLIGHT
DIRECTION
PHOTOMETER
TETHER
TETHER DEPLOYMENT SYSTEM
VLP ANTENNA
DAUGHTER CHARGE PROBE SENSOR—
DAUGHTER PAYLOAC
HF ANTENNA
Fig. 1.9
c.
Payload Configuration
MAIMIK
of fourth
TPE
Experiment[Ref.
7]
Rocket Experiment
The "mother—daughter" payload Maimik was launched
Andoya Rocket Range November
lO 1985 on Terriei/Black Brant
11
VC
out
of
rocket motor.
The two payloads were connected by
a tether, and the "daughter" payload carried
an accelerator which emitted pulse of electrons at energies of approximately 8 keV.
The main aims
beam and
of this
program were to study the interactions between the electron
the environment for various boundary conditions, to conduct studies of
various physical processes have been conducted, including plasma modifications,
wave generation, and
optical excitation studies,
and to study the physical processes
associated with the neutralization of electrically charged vehicles in an ionospheric
plasma.
was launched when the geomagnetic conditions were exceptionally quiet
It
which eliminated the need to disentangle
artificial
from natural processes and during
the period with very low background plasma density.
In this experiment, the tether wire
and daughter payload independently.
Apparently,
was
to be cut on the
the tether
mother
was cut on the
daughter 112.6 sec after launch at an altitude of 192 km. The wire was cut
conjunction with the injection of an 800
mA
electron pulse, and the daughter
have charged to a very high potential due to
this pulse.
have been due to arcing on the daughter. [Ref.
8]
The daughter payload was charged
the
beam
current was equal to or higher than 80
The
early squib firing
mA. This
earlier
experiments,
and
probably
,
may
result apparently differs
ECHO
that electron collection at altitudes near (or above) 200 km,
may
may
to several kilovolts every time
from previous sounding rocket experiments, such as the
previous experience suggested. This
in
is
series.
This suggest
not as effective as
indicate improper analysis of the results of
indicates
contactors.
12
the
need
for
improved
plasma
C.
PLASMA CONTACTORS
Basic Concept (Requirements and Future Mission)
1.
The ongoing need
to establish electrical contact
between
satellites
and
the ambient plasma environment has resulted in studies of devices generically
termed
"PLASMA CONTACTORS"
which enable low impedance
with the ionosphere. This type of device
operation of an
electrical contact
needed for bidirectional electric current
is
electrodynamic tether and stable spacecraft
electric
reference
potentials.
A plasma
contactor might be a large conducting surface, such as an
aluminized mylar balloon. Charged particle emitters such as electron and ion guns
are
traditionally
enhanced
thought of
this
in
of conductivity.
levels
context.
The most
Neutral gas releases can produce
effective
techniques typically utilize
neutral plasma emission.
Plasma contactor performance can be characterized by plotting the
contactor current versus the contactor potential measured relative to the local space
plasma. Ideal performance
appropriate species)
is
is
approached
if
the effective collecting area (for the
equal to or greater than the associated tether current divided
by the ambient space current density (Ne ^KT/m). The collecting area can be
physical, or simply the enhanced
plume area
of a
plasma source or neutral gas cloud.
Plasma sources have been extensively studied
charge control.
early 1970's,
The need
when
it
for
for the
purpose of active
charge control technology has been apparent since the
was recognized that
satellite
charging at geosyncronous orbit
reached kilovolt levels, and that such charging might be related to satellite failures.
In this regime, current levels of
microamps are typically found. [Ref.
The European Space Agency
launched
in the
mid— 1990's
will
CLUSTER
spacecraft
9]
scheduled
to
be
have conducting surfaces so that the behavior of
13
tlif'ii
[intpptial slionUl
GEOS
ix-
sidiilar to
TjEOS and ISEE
of tlir
t.lint
and ISEE experience has been that the potential
balance betweeji iiholdrloction
the spacecraft
is
and
eiriission
r>— 1(J
V
The
determined by the
aiiibiont electron collection as long as
outside the plasniasphere and
expected positive potentials of
is
spacecrait.
in
sunlight.
In
order to reduce
to near zero, a liquid— metal ion gun
is
included part of the payload. Currents of lU—2U /tA will be emitted at 7 kV. Indium
ions are emitted
1.10.[Ref.
lU]
from needle shaped sources
The Liquid Metal
Ion
in
this design
Source (LMIS)
microfocused ion beams and has wide application
This method has advantages
low
of
as
suited
is
shown
for
in
formation of
microelectronic technology.
in
power consumption, high mass
compactness and low mass. Unfortunately,
is
it
unipolar,
that
is,
it
efficiency,
only emits
positive charge.
ION-EMITTEI?|
SINGLE LIQUID
'»•
/
\
7""""— ION beam(Ti)
@-c:
Y
(9)
FOCUSING
ELECTRODE
i'
3
(D-c
t
I
3
;
NEEDLE
INDIUM
C
-
^
3
c
<;
A
@
/^ ^.
5KV
(D
CERAMIC (D
HEATER
O.SWatr
BEAM CUgRENT( /;
'
SC
5
14
j?CUNO
mm
Fig. I.IU Individual Liquid
Figure
Metal Ion Emitter
[Ref. 10]
The
scientific objectives
of
many sounding
rocket active experiments
have been to study the propagation of plasma beams through the ionospheric plasma
in the
presence of the earth's magnetic
field.
Recent interest has been stimulated by
active experiments on the space shuttle orbiter, such as the former
SEPAC,
payload,
as part of the
a large electron gun experiment (1 A, 10 kV), which will be reflown
ATLAS
payload.
The 1983 SEPAC experiments included
To
shuttle orbiter.
or
more
are
large positive potentials on the
prevent such charging, net positive
transmitting an electron
The
(e.g. ion)
currents of ampere
would allow the fundamental mission objective of
This
desired.
to be achieved.
SPACELAB
beam through
the ionosphere to the top of the atmosphere
joint Italian— American
Tethered Satellite System (TSS— 1)
scheduled for a late 1992 flight on the shuttle orbiter. In the case of
sources are needed to collect and emit the estimated
1
A
is
TSS— 1, plasma
current which will flow
along the tether wire to the orbiter.
It
is
anticipated that charge control technology will be needed for the
space station, particularly
active plasma experiments are to be conducted from
if
that platform. Orbital systems which emit large fluxes of charged particles presently
being considered for defense purposes will also require this technology.
2.
Plasma Contactor Flight Historv
a.
Balloons
The
free space capacitance of a typical rocket
that ejection of even a
than
1
1
mA
beam
msec. Consequently,
if
will
payload
is
charge the vehicle to more than 10
the electron
beam
current
is
so small
kV
in less
not neutralized by
ejection of an equal positive ion current, sustained ejection of electrons will require
the collection of an electron return current through the ionosphere plasma.
15
The assurance
addiessed
by
a
deploy uig
of an
large
adequate collection of return current can be
collecting
For
surface.
this
a large
purpose,
inflation—deployed electron collection screen was developed and fabricated for the
first
US rocket— borne
electron
accelerator
The experiment
experiment.
carried
electron guns capable of up to 49(J inA at 9.5 keV. This artificial aurora experiment
was launched on an Aeiobee 350 from Wallops Island
m
payload included an a.luininized mylar disk, 26
hub and rim and
four inflatable spokes.
The
regulator. Nevertheless, the technique
was
in diameter,
collector
The deployment was apparently incomplete, due
January, 1969. The rocket
in
is
with
inflatable
aji
illustrated in Figure 1.11.
to a malfunction in a pressure
at least partially successful. [Ref. 11]
HLLIUM
DIfl
iMFLATEO RIBS
|7"
(4,^0
RIMS
85
DIA.
ff
1/4 mil
MYLAR
COLL ECTOR
(CONDUCTIVE COATEO AREA ?SIDES
"^-^Bn '^niiflRF FFET EACH
1
Fig. 1.11 Collector Screen in the
ECHO
1
Deployed Configuration[Ref.
was launched on
13
August,
1970.
This
11]
experiment
apparently showed that the conducting rocket body could collect a sufficient return
current for
beam
neutralization without a special collector disc or other device3.[Ref.
12]
16
b.
Electron
Gun
Electron guns are a primary form of a plasma contactor. Electron
guns have been used in space at current levels from microamperes to amperes.
Examples
of flight experiments are as follows:
The
first
This experiment was a
was the above mentioned
first feasibility test in
study of magnetospheric
configurations through the controlled injection of
field
and beam currents up to 490
was
aluminized mylar
lines.
were used. In
by
collecting
this
an
ECHO
kV and
series of
beam
keV
experiment, the vehicle
ionospheric
field.
current
[Ref
using
11]
sounding rocket experiments, electron guns
current levels of 100
mA
and above have been used
study magnetospheric plasma phenomena. Beginning with
that sufficient
Energies up to 9.5
deployed perpendicular to the magnetic
In the
with energies of 40
mA
accomplished
foil
aurora experiment.
the development of a system for the
monoenergetic electron beams on to magnetic—field
neutralization
Artificial
ECHO
I,
it
to
was found
electrons escaped to the top of the atmosphere to generate an
artificial aurora. In this sense,
the 'contactor' was adequate to couple lO's— lOO's
mA
to the environment. [Ref 13]
Space
Experiments
with
Particle
Accelerator
(SEPAC)
was
designed to perform an active experiment in space by ejecting a high power electron
beam up
to 5
kV, 300
failed in this function,
mA
(e.g.
kilowatt power) from shuttle orbit.
It
apparently
due to shuttle charging, and the occurrence of a beam plasma
discharge, which disrupted the beam. [Ref 14]
A
second (French) payload carried on the same spacelab mission
was dubbed PICPAB. This lower power electron gun (10 mA, 8 kV) was designed to
study wave generation. Diagnostics were inconclusive, but appear to indicate the
beam
resulted in low potentials, and hence the
17
beam
escaped. [Ref. 15]
A
electron guns of
The
CHARGE
away from the
1
series
kV, 80
virtual
DC
in
and multi—pulse modes
experiments
used
as described previously.
beams
vehicles.
20-800 mA. The
beam
mA
rocket
tethered
rockets have been at least partially successful in propagating
The
resulted in
US/JAPAN
the
of
MAIMIK
MAIMIK
experiment used an electron gun with 1—3.2
experiments, were also successful. These experiments
some observations indicating that the
vehicle charged to near or above
energy, however, limiting the emitted current. In fact,
%
cathode formed, and only 1—10
electrons were accelerated to 110—120
The ATS— 5 filament
microampere current
kV and
levels
at
%
of the
of the
beam
beam
energy.
it
appears that a
escaped.
The escaping
neutralizer was used as an electron emitter at
ATS—5
geosynchronous orbit.
electron
emitting
operations succeeded in reducing the magnitude of the negative eclipse charging
potentials
on
the
satellite.
However,
the
was
spacecraft
rarely
discharged
completely. This was the result of differential charging on the satellite surface
limiting the emitted current.
This
means that on high
altitude
satellites
with
substantial
portions of the satellite covered with insulators, electron emitters are generally
ineffective as
plasma contactors.
levels ranging
from
1
^A
to 13
SCATHA
mA, and from
showed that even with high energy beams
experiments used an electron gun at
50
V
to
3 kV.
differential
SCATHA
experiments
charging prevented
beam
emission. [Ref. 16)
c.
Neutral Gas Releases
The
the spacecraft.
injection of a
plasma or a neutral gas
The mechanisms which
is
one way to neutralize
are important are those which result in
substantial ionization of the neutral gas cloud. This
18
may
be due to the
beam
itself,
the return (collected) electron flux, or perhaps even secondary electron emission
from the vehicle surface. Under some conditions,
believed that as the ionization
it is
process proceeds, electrostatic waves are set up, which increase the ionization rate,
instigating a
Beam— Plasma
Discharge (BPD).[Ref.
17]
Neutral gas releases were used on
effective vehicle neutralization
ECHO
was obtained. Indeed, the
I.
ECHO
ECHO
I
results
IV experiment considered again the
The 40 kV, 80
neutral gas release.
direct evidence of
mA
beam was
fired
have been
body provided adequate
traditionally interpreted with the thought that the rocket
return current collection.
No
effect of
a
through an N2 plume which
had densities were obtained by means of a photometer calibrated
at
3914 X. The
return current was enhanced by the neutral N2.[Ref. 18]
Possible occurrences of
are described
by
J.
R. Winckler.
The
BPD
BPD
conditions are that:
ylIc
where
V
is
the accelerating voltage,
scale length of the experiment. Since
it
is
clear that
BPD
during sounding rocket experiments
5
= C go-7pL
B
the magnetic
(1)
field,
beam perveance
beam perveance determines
is
pressure and L the
generally of the form laV'"^,
the occurrence. Winckler concludes that
has been observed. [Ref. 19]
This technique was also used in the
beam energy
of 2.9
keV and
current of 200
mA. The
emitted 10^3 molecules of nitrogen (N2) in a 100
ms
SEPAC
experiments with
neutral gas
pulse.
in a neutral gas pressure increase in the shuttle bay,
Torr.
P the
from
The
10"^
plume (NGP)
gas release resulted
Torr to 2
—
3x10-6
This technique was apparently successful, but appears to have involved
substantial interactions with the ambient neutrals and plasma.
19
One
curious aspect
was that the
of these experiments
mA PICPAB
8 keV, 10
may
substantial ionization of the emitted neutrals. This
PICPAB
low shuttle potentials induced by
disagreement
the
in
over
literature
beam
be related to the relatively
There
(~ +10).[Ref. 20]
occurrence
did not result in
BPD
of
is
in
substantial
the
SEPAC
experiments. [Ref. 21]
Ion
d.
Beams
A non—neutral
series
ion
beam
is
a form of plasma contactor.
The ARCS
experiments used Argon ion beams to study the electrodynamics in the
ionosphere.
A
mA,
100
25
eV Ar*
ion
beam was used on experiments beginning with
a sounding rocket
ARCS— 1,
to be accelerated
toward the beam emitting payload during ion beam operations.
This
launched on 27 January, 1980. Electrons were observed
was ascribed to the creation of an
effect
geomagnetic
electric
parallel
field
to
the
There also appear to have been substantial electron heating,
field.
apparently due to wave turbulence associated with
ARCS— 3
showed that the injection
beam
operations. [Ref. 22]
was launched on 10 February, 1985. This experiment
of particles parallel or perpendicular to the
field consistently resulted in
geomagnetic
the appearance of a population of ions at low energy (~
15 eV) and at 90o pitch angle. These ions apparently are scattered out of the beam.
These experiments indicate that the sounding rockets are able to draw neutralizing
currents
to
beam and
the
rocket
thermalize the ambient plasma.
It
via complex processes
appears that the beams do escape. [Ref. 23]
Another form of ion beam
Electron Drift).
abandoned
thruster
satellites.
ACED
in the
initiated in the
mid-60's. Subsequently,
was used
The
was
which accelerate and
is
ACED
(Accelerator with Closed
United States but appears to have been
it
was taken up
in the
USSR. An
ACED
to adjust the near synchronous orbit of one of the "Meteor"
ACED
ion
beam
source delivered 4
24]
20
A
of
Xenon
ions at 200 eV.[Ref.
was supplied by the Soviet participants
It
for the
West— German
"Porcupine" program of rocket experiments in the auroral ionosphere. Porcupine
rockets
F3 and F4 were launched
March, 1979 from the European Space Research
in
Range. The Xenon ion beam was injected perpendicular to the ambient magnetic
field into
the collisionless ionospheric plasma at altitudes ranging from 190 to 450
km. These experiments showed that the beam propagated nearly undistorted across
the plasma.
The beam was not
e.
Ion Engine
An
charge
and
current neutralized. [Ref. 25]
ion engine
is
balancing
current
a device which combines an ion
electron
source.
current. The first major ion engine flight experiment
10,
1968.
ATS—4
was intended
failed to achieve a second burn.
booster,
and remained
in a
There
is
beam with
generally
was .\TS—4 launched on .August
ATS-4 remained
low parking
attached to
its
Centaur Stage
Two
large
current
330
/;A,
booms
prior to the last of five ion engine test
ion currents available from the relatively dense ambient
to 760
km, density 10 lo/m^, ambient pressure 10"9-10"6 Torr,
spacecraft
potential
—132 V)
precluded
the
achievement
appreciable neutralizer emission current except for a few brief periods.
exception of the emission limited behavior of the neutralizer during test
engines
performed
as a
26]
ram
plasma (altitude 218
of
ion thruster systems were on board, and a
The boom was deployed
The
number
orbit. In spite of this failure, a
spacecraft potential monitor that employed one of the gravity gradient
periods.
net
little
to be a geosynchronous satellite but the booster
successful ion engine tests were run.
Langmuir probe. [Ref.
a
normally
in
all
respects.
Particularly
gratifying
5,
With
of
the
both ion
were
the
complete absence of detectable electromagnetic interference, and the lack of arcing.
21
^3 5
2 31
S
—S
tZ
O^
<I
UJ
3,1-
t'l
Q^ h—
o:
VECTO
I—
z
UJ
•<
UJ
MENTED
ODES
a
m
»—
_J
.
•/'o^
i^S
EL
PLIT
ELE
Kn
cj
£
to
*s ac
UJ fr H^
,
_
Si
:3
lO
UJ
CJ
>
o
o
o
CM
o
a.
tn
»—
_i
o
>
o
o
o
'
a:
cr cr
UJ UJ
UJ
0. en
««
u
>
.
—
»-
'
•
=E
2
UJ
CO UJ
UJ u.
«J
zJ ^
2 -4
_ 3E
CO OL
o
3
= CO
»
UJ
O
>
I"
.^'^
UJ
O U.
:3
o
5o
I
UJ
oc
UJ =J
in
=> CO
o
o
ac UJ
UJ
O
a. u.
UJ UJ
O (—
to
Fig. 1.1'J Configuration of ATS-v^) Ion TluustcrfHfl.
'
y
!(.]
ATS—^^
neutralizer.
stationed at
a
carried
cesium
conlact,
ATS-v) was launched into synchronous
10!")
^
W
orbit on
is
shown
in
August
longitude. Again, there was a launch failure.
Figure 1.12. The ion engine worked well
problem. Coupling potentials of
less
than 5U
V
resulted
in
the effectiveness of neutral
beam
beam
filament
12,
1969 and
ATS^
was to be
ATS-^
ion
spite of the spin
from engine operation.
Induced charging experiments (no neutralization) with the ion
resulted in induced charging to near the
and
lUO rpm. The
gravity gradient stabilized, but ended up spinning at
engine
engine
ion
beajii
in
sunlight
energy. These experiments indicate
emission. [Ref. 16]
BOUNDARY ANODE RINGS
PLASMA GRID USSOV)
ACCELERATING GRID 1-550V)
RING MAGNETS
DISCHARGE CHAMBER
TANTALUM COIL
*'.r
I.
.'."a:
-ll
Ml
,i-.ir.
,
jLi.J.'L'.'.ii. '
• •••
HI.
Jll'
1
I
i:^
:
"
^".'..*J
••••
=mv.', u;
CATHODE
VAPORIZER
'.
". itti
f.L".ritinTTri
SHEATH HEATER
•ANODE
NEUTRALIZER
PROBE
VAPORIZER
CATHODE
FEED
SYSTEM
^S
^S
ANODE
FEED
SYSTEM
Vp
NEUTRALIZER
VAPORIZER
NEUTRALIZER
FEED
SYSTEM
'^--^
Fig. 1.13 Configuration of
ATS-6
Ion Engine[Hef. 16]
ATS-^i carried twin cesium thrusters designed to
technology
and their usefulness
for
stationkeeping on
23
the
test
ion engine
three-^axis-stabilized
satellite,
It
was launched
schematic with neutralizer.
1974.
in
ATS-6
Figure
The plasma bridge
ion
engine
mainframe and
differentially charged
all
A
92 hour operation of the ion engine at 160
mA,
3
successfully conducted.
A
geophysical rocket was used to launch an automatic ionosphere
space—flight laboratory (SFL). Experiments with
USSR
ATS-6
neutralizer alone could also discharge large negative
potentials in sunlight or eclipse.
kV was
shows the
engine operations were successful, and had the
beneficial side effects of discharging the
surfaces.
1.13
in
this rocket
were conducted by the
November, 1969 and August, 1970. These experiments made use
of an ion
engine using surface ionization of cesium on tungsten to study the operation of the
various systems of this engine.
mA
(240
sec)
and the
The maximum value
effective
layer surrounding the
SFL was about
SERT—II
voltage
accelerating
magnitude of the potential was about —1700
V
of the ion
beam
current was 100
was about 2400 V. The
and the thickness of the space charge
7 m.[Ref. 27]
(Space Electric Rocket Test
II),
carrying dual
cm
15
diameter mercury electron bombardment ion thrusters, was launched on February
1970 into a polar, sun—synchronous orbit at 1000
of
SERT— II
was a long duration (6—month)
km
test
altitude.
of a
3,
The primary purpose
mercury ion thruster.
A
secondary objective was investigation of the interaction between the ion thruster,
the spacecraft, and the ambient space environment. [Ref 28]
Emissive probes flown on the
SERT—II
spacecraft in conjunction
with the prime ion thruster experiment allowed an investigation of the interaction
between the spacecraft, the ion thruster, and the ambient space plasma. One
thruster
thrusters
operated
failed
for
5—months and the other operated
for
3—months. Both
due to sudden shorts between the high voltage
grids.
It
was
determined that the cause was sputtering of grid surfaces. The specific impulse was
24
I
about 4200
and thrust was a 28
sec,
main objectives were
for
mN
largely reached.
for a thruster input
power of 850 W. The
The engines were operated
at 253
mA,
kV
3
600 hours. They demonstrated thrust, and an absence of harmful interactions
with the vehicle. [Ref.
29]
SERT—II
The mean
with the
equilibrium potential
spacecraft
engine off was —6 to —8 V. This relatively high negative potential was due to the
presence of exposed solar array interconnections at high positive potentials (36
The space probe measurements
Volts).
indicated
detectable effect on spacecraft potentials.
difference
was possible
It
had no
photoemission
that
to control the potential
between the spacecraft and the space plasma, using the neutralizer
bias.
This effect could be used to minimize electrostatic influence on the space plasma of
beam
an ion thruster bearing spacecraft. The ion
constant for
all
neutralizer potential difference
was
neutralizer biases (with respect to the spacecraft ground) with a
constant neutralizer emission current. [Ref. 30]
In
new
1973,
spacecraft which included,
storage,
ii)
i)
goals were added for the
still
functional
SERT—II
demonstration of thruster restability after long space
study of factors limiting thruster restarting,
iii)
demonstration of the
space lifetime of thruster system components, such as propellant feed systems,
closed
loop
measurement
control
of
main
systems,
insulator
shields,
and power processor
solar array degradation after years of space exposure,
verification of the compatibility of the spacecraft
efflux.
units,
Minor components
of the spacecraft
had
and
iv)
,
v)
system with sustained thruster
failed,
but this did have not interfere
with the functional status of the spacecraft. The short between the engine grids
removed
itself,
apparently due to mechanical
In 1974, the spacecraft
stress. [Ref. 31]
was placed into a spin-stabilized mode
in
order to be ready for the next period of complete solar— array illumination. In 1979,
25
additional goals wcic addfd
tlirustci
systcin
performance of
years
!)
this
agfl
1)
:
dcnionstratioii of sl.rady stale opeiatioii of an ion
alter
launch,
tjiinstei
vvitli
2)
operation of the ion thiuster system to determine
engines vvorkc^l well after 9 years
2
was operated
thrustei
for neaily t)OU
was found
hours at 85
unchanged
be
to
orbit.
in
new
of a
tjiat
in
beam
comparison
and
thrnstei,
3) long
the
term
limiting factor. Again, the
its life
Dining the
mA
and
nicasnreineiit
cpiailer of 1979, thnister
first
current.
The performance
The
the 9 years.
spacecraft
of the
remained
functional and had only experienced minor failures of a battery in the telemetry
system and the
loss
01
SERT— II
one of one tape recorder. [Ref. 32]
particular
expeiiments,
are
interest
the
peirticularly
the
beam
plasma bridge
neutralizers were hollow cathode,
plasma
contac:tor
neutralization.
neutralizers. Figure
-25
-10^
".
-r
m
"N'MO^/cm^
\
(AMBIENTL
'
I
V
^--
N|
i
I
aspects of the
The
SERT—II
J.H shows a
M
= 10''/cm^
'*''^^'^'"
^ -^ ' TcT'o'^rc ^ ~ELECTRONS,
EXCESS
/
CURREMT NEUTRALIZATION
FROMT/S-1 (UPT0 85mA)
AMBIENT
-^
-1
PRIMARY
\ BEAM
-"^ \ N|-lo8/cm^
ELECTRON
FLUX; ~<1 mA/m \lONS~85mA
OF BEAM LENGTH
s.
j^r-
/'
m
I
EXCESS ELECTRONS
FOR SPACE
NEUTRALIZATION
r
NEUTRALIZER
Ni~10^°/cm3
/
~
IONS-
A
CHARGE
\ ^i
EXCHANGEt- \6x10''
~
N;
r'
\
2
^\
mA
I
^MAIN
k
.^-
^
IONS~9mA'v
) IS CHAR
DISCHARGE*
Nj~10l^
IONS- 40 mA
T/S-1
NEUTRALIZER
Nj~10l0
IONS-
2
mA
SPACECRAFT
Fig. 1.14 Spacecraft
Plasma Current Diagram
2fj
of
SERT-II[Ref.
33]
spacecraft plnsiiia cuiiciii
SEHT—11
tliiuslcis
(I'uigraiii
conii)ari3()u of flight
data
beam
wcie conducted.
which showed
neutralizer
currents
willi
(pialilal ive
emissiu!i
of
experiiiiciil.
Operation
of the
two
one grounded and the other biased positive permits a
vvil.li
neutralization
SEllT— II
of
thctjretical
agiceiiHMit
.SEilT— II
with iirgallvf neulrallzei
coupling chai<w.tcristics. Experiments in
A plume
with
penetration model was developed
data. Tiie average of the total
flight
data was
72
inA.
The SEllT— 11
neutralize!
biases were up to about twice the theoretical
ground screen. [Kef.
|)redictions for election collection by the
33]
OCCD
100
CM
O
3
C
E
o
SERT n
e
^ 50
•
Data
eV
5
s
•
MM
E
UJ
+20
+30
Neut.
+10
+10
Neut. 2
I
positive
Neutralizer
bias,
+
20
V
Plume penetration model
Fig. 1.15 Distribution of
Emission Currents[llef. 34]
Figuie 1.15 shows the shilt
two neutralizeis
for
thruster (No. 1) was
mode
(IIV on).
vaiious
in
positive
discharge only
The discharge
neutralization current between the
in
biases
for
mode and
plume penetration model. One
the other was
in
normal operating
only thruster emitted sufficient discharge—chamber
electrons to neiitializc the associated ion beam.
27
The experimental data show a
shilt
Iroiii
2
iieulializci
leiiiissioii at
a
at
niiiissioii
a iicutralizer
2 pu.silive bias of
iiciil.iali7.(:r
TJio (lata tiom
SERT— 11
(j
1
12— IG V
cxpeiiiiienl iiidirate lliat a hollow cathode
only takes a few volts to extract
it
A. Tlir data nlso iiidicalc thai an ion
beam
neutialized by a pliysicaily close electron source. Electrons
away botwrrn
SEllT—11
Uie
clcclroii
nonlralizer curroiil,^. Figuie
electron
nenlralizer
current
space need not be
in
may
be emitted
and the ion beam. The neutrallzets
1
meter
of
the
were operated negative of the spacecraft showing large excess
lhrusl(;r3
negative
source
to all neutializer
V.[Ref. 34]
provides laige eledruii ciniasion captibilily and
cuiieiils of 0.1
bias of
to
bias
l.lfj
for
siiow excess nculralizer emission as a function of
noimal operation.
ground screen may exist
tiic
indicates
It
the
in
a significant
that
absence of a negative
neulralizer bias (5 eV).
0.2r
O
SERT n Data
Te, eV
c
01
u3
c
o
o
o
2
k-
5
0.1
0)
"o
o
Neutralizer bias,
Fig. 1.16
is
struck
V
Excess Neutralizer Emissionfllef. 34]
A modified form
magnetoplasmadynamic
-50
-30
-20
-10
arcjet
is
of an ion engine
is
the
MPD
arcjel. Essentially
a hollow—cathode dischaigc source,
between a cylindrical rod anode on the
cathode forming the outer wall of the device.
28
A
axis,
in
an
which an arc
and a hollow cylindrical
strong azimuthal magnetic field
created by current flowing along anode yields the propulsive force along the axis
which
is
proportional to the square of the current. Thus an
at high current
MPD
arcjet
because
was used
it
expels
to neutralize the
from the designs flown on the
continuous current. The
during a
1
ms
main features
pulse.
much
The
SERT
SEPAC
MPD
arcjet
works best
neutral gas along with the jet of plasma.
SEPAC
and
ATS
beam Spacelab-1.
electron
series in that
arcjet generated
lO^^
it
It differs
does not produce a
argon ion-electron pairs,
neutralization effect persisted for tens of milliseconds.
of this neutralization
A
The
have been explained by existence of a cold
secondary plasma near the orbiter. The cold secondary plasma was produced by
charge exchange processes between the released neutral gas atoms and injected high
speed plasma ions. This result suggested that the production of plasma that will
remain near a vehicle
will effectively neutralize the
charged spacecraft. This limits
the usefulness of this design as a plasma contactor. [Ref. 35]
f.
Hollow Cathodes
One
successful
form of a plasma contactor
Hollow cathodes are effective sources of dense, low energy
particularly effective as low impedance electron sources.
for
hollow
application,
cathode
technology
was
in
association
(1
is
eV) plasma. They are
The
with
the hollow cathode.
original satellite use
ion
engines.
In
this
hollow cathodes are utilized as electron sources inside ion engine
discharge chambers, and as electron emitters to balance the charge and current
contained in the emitted ion beam. They have worked well in this regard on the
SERT-II and ATS-6
missions.
Hollow cathode technology has been extensively considered
purpose of charge (potential) control on
satellites.
Hollow cathodes operating
for the
at low-
current levels (microamperes) were effective in controlling satellite charging at high
altitudes (goesynchronous orbit). [Ref. 36]
29
Hollow cathodes have been operated as high current sources (100
mA)
in association
with the
SERT—II
ion engine experiments as described in the
previous section.
D.
PROPOSED EXPERIMENT
We
propose to test hollow cathode technology for application as plasma
contactors, using a tethered system, deployed from a sounding rocket.
In this study,
deployment systems
of different tethered rocket experiments will
be investigated in order to choose a proper deployment system for the hollow
cathode rocket experiment. Modeling of tethered deployment system
will
be studied
using the numerical integration method. This modeling work will be compared with
our experimental configurations.
30
II.
A.
TETHER DEPLOYMENT
BASIC DYNAMICS
The dominant
tether
dynamic behavior, when
rigid—body in—plane libration.
If
we assume
tip
masses are present,
that the tether
is
is
shown
in
Figure
2.1.
The
tether length
1,
is
the
massless, then the
system behaves as a dumbbell subject to only trajectory dynamic
geometry
is
effects.
The
allowed to vary as a
controlled function of time, since the deploying rate will be used as a control input
for precise positioning of the tether tip.
Fig. 2.1
The nature
first is
Dumbbell Tether Model
of the tether dynamics suggests
two
different control inputs.
The
tether reeling acceleration. This can be used to achieve the desired tether
length and reeling rate.
tip vehicle.
The second technique
These can be cycled on or
is
to use small thrusters on the tether
off to control the tether pitch angle
31
and
pitching
Only
rate.
one
dimensional
dynamics
separation
the
of
tether
are
considered here, thus allowing reeling motions along the tether length.
B.
HOCAT DEPLOYMENT OBJECTIVES
The configuration we
The
are proposing for
HOCAT
would be deployed gradually to a
subsatellite
is
illustrated in Figure 2.2.
maximum
extent of about 100
meters. This separation distance has been selected for practicality
(e.g.
a reasonable
m
distance in the available time), and as a logical step up from the 1—10
chambers currently
The
ISEE— 1
in use for testing hollow cathodes.
intention for
satellite.
HOCAT
is
to utilize the
Newton. Several other versions
that
It
return
The conductors have a
resistance of about one
tether link can be used to carry
command
verification
has a tensile strength of about 600
of higher strength are available with only a small
we would power the sub—payload with
The
same tether used previously on the
This material consists of a Kevlar strand with 8 inner conductors,
shielded with a conducting outer sheath.
weight penalty.
vacuum
ohm
per meter so
separate, local batteries.
commands
to the daughter payload,
and
and instrument parameters to the main payload, where
the telemetry systems would reside.
The
tether will provide us the capability to
electrically isolate, connect or drive the potential of the
sub—satellite and the tether
independently, in order to study the operational characteristics of an electrodynamic
tether, particularly the
Two main
plasma contactor aspects.
options were initially considered. These were a spinning, or bolo
configuration versus an inertial system. In order to be capable of measuring the
conductor's effects at a variety of aspect angles to the magnetic
configuration
should be rotating about a
common
center of mass.
strength largely determines the spin rate of the satellite system.
32
field,
the entire
The
tensile
/\.
DETECTORS
/ %/
/
n^ /
E-N/PROBE ^^xn/
/
PARTICLE PLASMA
DC
HOLLOW
CATHODE #
1
/ 1//lL
/O /
HOLLOW
CATHODE #2
\vV
/nr^/^^^^-^
3^^
/I/
^3
CURRENT LOOP
STAGE MOTOR
Fig. 2.2
Hollow Cathode Configuration
33
2nd DC-E
PROBE
.
An
spin
initial
rate
of
'M)
rpin,
will
result,
in
maxiiiiuin doplovineiit (For an initial separation of
kg, a centripetal force of
about
50(J
Newton
1
a
of near zero
spin
final
at
meter, and end masses of lOU
and hence a reasonable tension).
results,
This configuration does put a substantial strain on the deployer, however, and
complicates that design. For simplicity, therefore, an inertial system
is
focused on in
the remaining work.
order
In
minimize meclianical design
in
experience obtained by other experiments.
•Japanese
OEDIPUS
1.
experiments.
Two
US— lAPAN
.Toint
Experiment Deployment System
The
wire
was deployed
I
ether
to
maJce
use
of
are the
future experiments
and TSS-l
at
mother—daughter separation by a
multiple spring system. Separation speed was set at
first,
we wish
The major pievious experiments,
CHAR/JE, and Norwegian MAIMIK
are Canadian
costs,
(J.5,
1.0,
second, third and fourth experiment respectively. In the
1.5
and 1.05 m/s
first
in
the
two experiments,
the deployment s[)eed gradually decreased with time, and finally stopped abruptly
at
38
m
systein.
and
(>5
m
However,
respectively, due to a frictional force in the tether
in
(he last two experiments, an improved deployment system,
called a Reaction (Jontrol
operated
witli
System (KCS) was used on the daughter rocket. The
RCS
gas jets every K) sec to compensate for the friction.
This
system
is
characterized
Mechanically, the deployer design
McCoy
deployment
is
by
a non—constant
deployment speed.
similar to a spinning reel (e.g. a fishing reel).
has adopted a similar strategy for his Plasma Motor Generator
test.
To
model the deployment system, without a RCS, from the above data, the equation
motion
for
the
sepaiation
of
the
expanding the separation distance
in
tethered
payload
system
can
be
solved
a Taylor series where the separation distance
expressed as a function F(t):
34
of
by
is
F(t4-At)
The
error.
where x
is
of
motion
the separation distance,
m
(e.g.
to be (J.OOOli,
(2)
system can also be expressed:
mass
is
(3)
^
^
of the
daughter payload, and
values
— —
4-t
-
of
=
the
—
).
and time from the
velocity,
initial
The
which gives a separation distance
To model
making the
1
a
m/s
coefficient
of 35.4
m
c/m was determined
and 71
m
for
initial
respectively. In the second experiment, the
not linear
is
term was
differential equation;
=
C|X
+
C2 x^
(5)
Trial and error solutions yielded C| 0.0001817
quadratic equation, giving
rise to
34.7
m
and
the cubic equation, the analytical values of
the separation distances of 34.93
very
two
first
this nonlinearity in the drag, a quadratic or cubic
X
agree
c is
In anticipation of less linear systems, numerical
separation was 65 m, which indicates that the acceleration
with velocity.
This
^
-Fnlt)
solved analytically to obtain the constant c/m, by
first
separation velocities 0.5 m/s, and
included,
•
=
-^i
m
integration was used to model the separation.
maximum
^+
of the different frictional constants.
known
experiments
F"(t)
of the separating
+
Equation (3) was
tlie
+
F'(x)At
x
combination
using
+
F(t)
neglected term provides an estimate of the local discretization
first,
The equation
=
well
with
m
the
ci
and 65.1
and
m
practical
35
fi5.4
m
and
for 0.5
C2
0.0000599 for the
m/s, and
C2 are 0.000196,
for 0.5
data
m/s and
of
1
m/s
.
For
0.0000499 giving
m/s
respectively.
US— JAPAN
experiment
1
«•
••
Time
Fig. 2.3 Plot of Tether
8«
(sec
!••
)
Wire Extension Velocity vs Time
36
IM
—— — —— — —— —— — —
1
I
I
1
1
I
I
1
1
1
et
r
1
1
\
—
ei
B
u
.9
Q
a
o
"(3
4»
Eh
8*
J
I
I
I
I
I
I
8«
I
I
1
I
4t
I
I
et
Time
Fig. 2.4 Plot of
I
I
I
••
(
sec
I
I
I
I
!••
)
Tether Separation Distance vs Time
37
I
I
I
18t
without ROS. FigiiK^s
J.
change with ifspcct to
similar to the t^S—
3 and
timt-
'J.
lor
show how
a system
lAl'AN experiment.
tether deployed to a length o( 418
resjiectively.
4
Figure
J.'S
m
In
tlip
witli
the
volofity
initial
third
velocity
aiid
for
1
louith
during 283 second and 426
shows the wire deployment
separation distaiire
<\ii(l
m
the
in/s,
and drag
experiment, the
during 29U second
velocity
of fourth
experiment with the Reaction Control System on board.
Fig. 2.5
2.
Deployment Velocity
MAIMIK
The
some 62 second
US-JAPAN
Experiment[ref.
7]
Rocket Experiment Deployment System
MAIMIK
eifter
of fourth
rocket payloads were separated at an altitude of 86.5
km
launch with relative speed of 0.8 m/s. During the separation,
the payloads axes were tilted at an angle of 23 ± 2 degrees relative to the local
geomagnetic
field
vector. Hence, the relative cross—field daughter velocity
m/s.
38
was 0.3
CANADIAN OEDIPUS
3.
A
Canadian experiment team has developed a deployment system which
appears optimal
HOCAT. By
for
The
"trolling" reel.
The
control jets.
velocity
The
approach
satellites
to
maximum
where the
0.6(W0 m/s^.
A
separation
distance
of
negligible
friction,
which
in
the
to 'JO ft/s.
is
a
lU sec with attitude
first
then decelerates until the
It
separation distance of
our system,
of
accelerated
up
Japanese approach, this
tjie
takes 328 sec with a constant deceleration rate of U.0819
'J'liis
Canadian's
is
inrit-ased
is
to
contrast
payload
daiigiiter
velocity drops to zero.
in/s^.
Ex[KTiiiicni DeployiiK-nl. Sysleni
acceleration
initial
We
km.
can apply this
would be the same as
deceleration rate of U.0U5 m/s^ would give a
m
lUO
1
at
means
apogee
the chosen
for
constant
acceleration
conditions.
and
maximum
If
there
deceleration,
is
the
acceleration and deceleration can be easily calculated using;
x=-4-xt2
(6)
an(
x
For
O.GO'X)
iii/s-
velocity
1
the
x
(7)
t
(JBDIPUS deployment system,
which gives
m/s, and
=
is
valid for our system.
in
The
m
for the velocity to
separation to apogee. VVc
different
at
the
was
maximum
deceleration sh(Mild be different,
maxiuium separation distances and
velocities. It takes just 1.64 sec to reach the required velocity of
200 sec
acceleration
initial
the separation distance of 0.8
rise
because of the differences
the
initial
1
separation
m/sec, and takes
drop to zero. This gives a total time of 201.64 sec from
know
that the eicceleration
is
linear
with velocity which
is
from I'S— .IAPy\N experiment. This de[)loyer model meet our experiinental
39
rcqiiirnncnls.
with
I
Fif^nip
show
2.7
;ui(l
'.'.Ti
Uio vrlfx
Ui<it.
soparalion distance
niid
it.y
('Spec t time.
——————————— ——— —————— — —
-|
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
t.e
-
03
I
-
•G
>
d
o
ex
en
• .a
I
I
I
I
4t
I
I
I
I
I
I
I
I
I
L
I
lit
et
Time
I
ie«
(
sec
)
Fig. 2.6 Plot of Separation Velocity vs
40
I
Time
I
J
I
•••
L
•4a
—r— — — — — —
I
"1
I
r~~i
r
i
T
——————
I
I
I
I
I
!••
••
B
V
u
8*
.9
Q
S
ft
13
41
a;
CO
at
J
I
I
I
4*
J
I
I
I
I
I
I
«
!•
••
Time
(
I
I
sec
'
I
att
let
)
Fig. 2.7 Plol of Separation Distance vs
41
I
Time
I
I
J
C4t
TeUicicd Satellite System— 1 Deploy ineiit
4.
To complete
TSS— 1
system
the consideration o[ past and luture tether systems, the
illustrated.
is
Sysleiii
Deployment
tether and
of the
stable maneuver, requiring
5—8 hours
the desired altitude region
upward direction from the
satellite at the desired
through the
tetliei
the case of a 20
in
deployment position
tension.
stretching motions, which are
course of the experiment,
it
is
lUO
km
satellite
a
is
tether to reach
orbiter. Stabilization of the
accomplished by introducing damping
Swing motiona
damped
—
TSS— 1
of
the
tether
can
be converted to
by the reel control mechanism. During the
should be possible to change the length of the tether to
permit observations at various altitudes. The system should have a capability for
controlling
dynamic
orbiter. Figure 2.8
TSS— i
trajectories of the satellite in the ]—[()
shows the tethered
satellite
system
reel
km
zone surrounding the
and boom mechanism
for
mission.
so
m EXTENSIBLE
BOOM
PYROTECHNIC
ILIOTINE
THER & BOOMI
LECTRICAL SUPPLY
TO BOOM TIP
BOOM
DRIVE
HOUSING
SAttLLITE
SUPT.
STRUCTURE
GUILLOTINE #7
(TETHER ONLYI
MOTOR «
J
(VARIABLE SPOI
TAKEUP
TETHER
1
SATELLITE
SUPT.
Fig. 2.8
TSS-1
ELWINO
SPOOL
X UP TO 100 km
& TENSION CONTROL
mm TETHER
Reel and
STRUCT
y/
'
(TYPI
MOTOR
Boom Mechanism[Ref.
42
n
1
(VARIABLE SPEED)
(3]
PLASMA CONTACTOR HOLLOW CATHODES
III.
Because
the
of
need
ongoing
to
understand
the
process
of
electrically
connecting rockets and satellites with the ambient plasma, a concerted effort has
been conducted by
NASA/ Lewis
Research Center in coordinated laboratory and
theoretical studies, along with analysis of existing flight data.
LABORATORY STUDIES
A.
Much
of the
work done
in this area
has been conducted at Colorado State
University, under the direction of Dr. Paul Wilbur. This work has been
in
annual reports as noted
and approximately 3.5
here.
In the
CSU
m
in the Reference.
of the axial length
The CSU chamber
is
utilized for the
is
documented
m in
1.2
diameter,
experiments described
work, hollow cathodes have been demonstrated to be effective
electron sources at current levels of
1
A
or more. Extraction voltages of a few tens of
volts are generally required for such currents.
Of greater concern
is
the matter of
emitting a positive current.
Laboratory tests of hollow cathode systems as positive current sources
emission
or
electron
understood, and
absorption)
may be important
have
shown
in space. Ion
effects
which
is
possible at high gas flow rates. If an "ambient"
currents of 50-100
levels, recent
mA
totally
emission current levels of lO's of
milliamperes are typically observed with hollow cathodes, and as
ion current
not
are
(e.g. ion
much
plasma
as 100
mA
exists, electron
are collected at bias voltages of 20—100 V.
Above these
laboratory results typically have shown an operating regime termed
"ignited mode".[Ref. 37]
43
mode, a discharge
In this
observed in the chamber, which appears to be
is
with ionization of the neutral gas in the chamber.
associated
structure which results
to occur. It
The
ionization
a double layer, which allows a space-charge limited flow
is
argued that the transition to the ignited mode results from plasma
is
phenomena, and not from electron currents being drawn from the vacuum tank
walls. [Ref. 38]
The observed luminosity
ionization
is
due to atomic or ionic excitation, implying that
is
occurring and that
it
could be the cause of the improved contacting
performance accompanying the transition to ignited mode electron collection. Figure
3.1
shows a typical comparison of the electron current collected by the contactor
(Jce)
w^ere
compared
to the electron emission current
referenced
initially
dependent variable
Figures,
the
the
to
tank.
from the simulator
Figure
3.2
(Jse).
Such plots
shows that a more natural
the anode to ambient plasma potential difference. In both
is
transition
to
ignited
mode
represents
a
substantial
increase
in
performance.
The most
recent
work by Wilbur has focused on understanding the potential
distribution around the electron collecting cathode, in combination with modeling
work described below. Figure
3.3
shows the plasma potential contours around a
hollow cathode in electron collection
The high plateau
features
mode
(Wilbur, private communication, 1988).
(near anode potential) close
in,
and the 10—20
V
drop are typical
found in this work. One relative weakness of the work so
relatively poor diagnostics of the
Figure 3.4 shows the ignited
Similar work
Mike Patterson
is
mode
m
x
19
is
the
plasma characteristics (density and temperature).
configuration of hollow cathode by Wilbur.
being conducted at
in the 5
far
NASA/Lewis Research Center (LeRC) by
m chamber used by
the ion thruster group.
Work
in
the facility has been focused on high current (1—10 A) collection, in an environment
44
X
p
<
ro
o-
•
ON
<
o
-3
-
-.
.
1
o
o
«
E
u
u
vt
r^
CJ
N
<
o
«
g o
o
X 2
m <
H
E
u
•E
qT
CVJ
t
•
o
o
CJ
o
o
o
o
I
I
I
ELECTRON CURRENT ImA)
Fig. 3.1
Perturbing Effect
of
Vacuum Tank
45
VVall[Ref. 37]
»
o
o
lO
I
ELECTRON
EMISSION CURRENT
rIOOO
9
JCA =0-9 A
VcA • 4 TO 14 V
-750
2.7 seem (Xe)
fti
»
P,
«3.5 X \0-* Torr
12
cm
DIA
ANODE
-500
250
CONTACTOR ANODE TO
PLASMA POTENTIAL
50
•50
100
I
I
DIFF. (V)
150
—I
TRANSITION TO
IGNITED
MODE
250
500
•750
Fig. 3.2
Contactor Anode— to Ambient Plasma Potential Dilference
vs Electron
Emission Current[Ref. 37]
46
Jcoih
••
O. 6
3. 7
A
seen
Vco- 23 V
(Xa>
7»10'' Torr
12 cm ANOOe OlA
JcE" -600 KiA
Po -
3.
z
LU
o
Q.
<
<
Q.
NUMBERS INDICATE POTENTIALS IN VOLTS
RELATIVE TO POTENTIAL AT POINT A
AXIAL POSITION fZl (cm)
Fig. 3.3
Plasma Potential Contour around Hollow Cathode[H,ef.
47
37]
Fig. 3.4
of relatively high
Hollow Cathode Ignited mode Configuration
ambient neutral density. Major
effects
were found due to volume
ionization caused by the collected electrons. Recent modifications of the
will
chamber
enable experiments at more reasonable neutral background levels.
Patterson's work focused on high current levels (higher than those studied by
Wilbur). Multiampere level (10 A) currents were extracted from the simulated space
plasma environment. Figure
without
power on
for
the
3.5
shows the current to the plasma contactor with and
contactor
voltages. Relatively little difference
is
discharge,
at
comparable flows and bias
found. [Ref. 39]
Figure 3.6 shows the dependence of the collected current on the discharge
mode. Substantially higher currents are drawn
experiments of the
CSU
and LeRC
the results on the chamber
size.
in ignited
facilities indicates
mode. Comparison of
the substantial dependence of
Figure 3.7 shows the large variation in results for
similar operating conditions. Substantially large currents are possible in the
48
LeRC
cc>nt»c tor
f»{ III tv
<»»t
flow r«»«
pref»vir«
no.
-
StCCA >'2.««-« (olid-tt
!.««-« op»n-lt
1.4«-«
IS
«.««-«
«*
0-i»
-
-
Current to
plasna contactor,
100
200
Bias Uoltage, v
8Dac» Plasna
Sinulator
Tatt
no.
txpell^n'
Tvt>«
a a
11
12
li/^" holw.
1 ca«i>o<>«
1
"
erp»I Ian*
Typ«
-
M
Fig. 3.5 Current to
Sinulator
R«f ervnca
Contactor
Bias Pt.
rower
•
4nk
gro«jri<J
ftnode
H
M
6*f
ye
S.9e-«
C.»«-t
39
300
er>
xe
COLLECTED ELECTRON CUnHENT
-SS
L
Plasna Contactor
ftr
ft
A
Contactor vs Bias Voltages[H.ef. 39]
IJoJ.
A
e
UN-IGNireO MODE
a
50
100
150
?00
BIAS-TO-PLASMA POTENTIAL Dirr.
IGtflTEO-nOOE
300
250
(Vb-Vp).
V
Fig. 3.0 Bias-t(>-Plnsrna Potential Difference[Ref. 39]
49
CONTACTOR POTtWTIAL [Va-V^] |V)
40
BO
120
teo
—too
JcD'OS TO OS A
•
\^0''0 TO
04
•h,
12
•
ELECTRON
COLLECTION
CURRENT «A
•IS 7
14
V
teem (X«)
em ANODE OIA
08
CSU rACILITY
P.'BxK)-* Torr
LeRC
CSU
He
He
1.2
'>'
-"
<5 X 18E3/cn'^3
8
X 10E8/cn'^3
L«RC
rACILITY
i.e
P,»
2x10-*
Torr
•t.o
Fig. 3.7
chamber, even
been
tliougli the electron density (bias off) is substantially lower. Tiiis
attributed
secondary
Current vs Contactor Potential
to
electrons
the
has
enhanced area of chamber walls capable of emitting
(Katz,
private
communication, 1988). The work
the program proposed below
at
to
is
communication,
CSU and LeRC
is
1988.
continuing.
Cohen,
private
One element
of
conduct complementary experiments at Naval
Postgraduate School.
B.
THEORETICAL STUDIES
Substantial
theoretical
work has been conducted under the
NASA/LeRC
program, particularly by Dr. Dan Hastings (MIT), and the S-<:ubed group. Hastings
(1988) has addressed the plasma contactor problem at low earth orbit, as
it
might
apply to tethered systems, from an analytic viewpoint. This work focuses on the
micro— physical
])rocesses
which
will
determine how well hollow cathode systems will
50
micro— i)hysical processes
work
at high current
An important
(
1
determine how well hollow cathode systems
vvliich will
A
or
more)
will
levels.
conclusion from this work
is
that large
amounts
of turbulence
to effective operation of these devices. If such levels of
(^'eff—0-4u.'pe) are essential
turbulence can be achieved, a plasma contactor can exhibit gains of 2—6 for ion
Ampere
current in the
gain
gain
is
level
and
for potential
drops
the range lUO—500 V. Here,
in
The
defined as the ratio of collected electron current to emitted ion current.
is
a decreasing function of the total current, so high current devices are less
efficient than
lower—current device.
This suggests that plasma sources with low mass flow rate and high ionization
fraction
may
importance
be the most effective for use in plasma contactor. Of fundamental
for collecting large
currents
plasma, by the current flowing
in
multi—ampere systems
work that
is
the level of turbulence established
the plasma. There are
will
be
ineffective,
in
some indications
that
is,
in
excessively
the
this
large
potentials will develop because insufficient current can be collect ed.fRef. 40]
This analytic work
at
S—Cubed
is
(a division of
complemented by numerical modeling
efforts
conducted
Maxwell Laboratories). This work has been focused on
three-dimensional simulations of the hollow cathode, and of the overall systems
response. This work
has been successfully
applied to interpretation of Wilbur's
S—Cubed
laboratory experiments. Figure 3.8 shows a 2 dimensional cut from an
simulation
data
of
communication,
experiments
is
1988).
for
will
that
to
Excellent
emerging from the
The challenge
models
similar
shown
agreement
NASA/LeRC
in
Figure
between
3.3
theory
(Katz,
and
private
laboratory
effort.
be to apply these models to space, and to validate the
space by obtaining flight experiment data, and making comparisons. [Ref.
41)
61
CNTR-LEVELS
O.OOtt-i^OO
(0)
5.00«4^00
I.DOa+OI
1. 5004^01
2.000401
2.50»401
3.00«401
3.50e401
4.009+01
4.50e401
S.OOe+OI
X
<
5.50O+01
>-
6.009401
6.509401
7.009401
7.509401
8.009401
8.509401
8
10
12
14
16
18
20
22
Z-AXIS
24
26
28
30
32
34
36
Minimum Poierrtial - 6 41©-03 Maximum F^olential • 8 00«+01
<2< 36.00. -9.00 <Y< 27.00. CU7PLANE OFFSET X- 9.50
0.00
Fig. 3.8
C.
2-Dimensional Cut[Ref.
41]
FLIGHT EXPEIUMENT DATA
Hollow cathode systems have been successfully operated
The
SERT-11
(
all
previous
chapter.
mA
ATS-(i
demonstrated success
missions
satellite
L970-l!)79),
Under
the
in
(1974-1976),
their
NASA/LcRC
relative
and
in
space for 2 decades.
SCATHA
domains, as presented
program,
data from
have been considered. Electron and ion gun experiments
level
were conducted on
were available
for study.
SCATHA
(1979-1983)
the
in
a
SCATHA
at the 10
^A —
1
from 1979—1983. Particle and wave data
Electron gun experiments showed evidence that at current
52
above 10
levels
beam was
—
100 //A (roughly equal to the ambient thermal electron flux), the
disrupted by space charge effects. [Ref. 42]
Charged particle data showed that the beam was causing heating of the
local
electron population, and was failing to escape the satellite region. Recent studies
have shown that electron cyclotron waves
implications
of
this
may
be the mediating factor.
work are that electron beams are
Contactor". This conclusion
partially supported
is
ineffective
"Plasma
as
by shuttle observations, though
remains somewhat controversial. [Ref. 43, 44]
this
Ion gun experiments on
SCATHA
showed a similar character,
in that
found that the beam rapidly became space charge limited when high current
high voltage
the
The
beam
region.
kV) experiments were attempted. The
energy, indicating only 1—10
The
operated
(1
in
ion
gun was
"trickle
effective as a
mode". In
this
%
emitted, at coupling voltages of less than
beam escaped
it
all
mA)
charge to
the near satellite
this
1
operates without accelerator voltage,
mode, currents of
lO's of //A could be
V.[Ref. 21]
These "Plasma Contactor" experiments, and the
data
(1
was
low energy plasma source, however, when
mode,
and resembles a large hollow cathode. In
of the
satellite failed to
it
earlier
ATS—6
and
SERT—II
indicate high levels of success with hollow cathode systems. There remains,
however, a need to further
current regime
(Ampere
question remains:
cathode in space
Can
1
test
level).
hollow cathode systems, particularly in the high
From an
ampere
operational point of view, the critical
of positive current
?
53
be emitted from a hollow-
IV.
The
EXPERIMENTAL FLIGHT PROGRAM
natural extension of the ground—program conducted thus far
to conduct
is
a flight experiment which will extend our understanding beyond the limits of the 10
m
chamber
as a
NASA/LeRC. We
at
plasma contactor with the
intend to test the hollow cathode for
specific application of a
1
A
its
viability
tether system.
This test will be a tethered, mother—daughter rocket payload, with hollow
cathode plasma sources on both ends. The experiment has been dubbed
The payloads,
as illustrated in Figure 2.2,
would be instrumented
"HOC AT".
in a
way
that
would allow us to determine the nature of the coupling between the hollow cathodes
and the ambient plasma, and the
effects of driving positive
and negative currents
through the system.
The mother payload
will
be recovered, and could be reflown with alternate
plasma contactor designs. Our goal
mid
1991.
Subsequent
flights
at this time
could occur at
is
to launch our first experiment in
1
year intervals,
at
substantially
reduced cost. In the following sections, the rocket, payload, and launch scenario for
this experiment, are described.
A.
ENVIRONMENT
The plasma environment
sub—auroral
at
altitudes
for the space shuttle
and proposed space station
from 250—500 km. This environment
is
is
a cold dense
plasma, with temperatures of a few tenths of an electron volt (1000—3000 oK), and a
plasma density which ranges from 10^ particles/cm^
in sunlight.
The lower ambient plasma
in
shadow
densities provide the
54
to 10^ particles/cm^
most severe
test of the
plasma contactor performance, and we wish to focus the experiment on that
The launch
would be Wallops
site
Island, at the
limit.
Wallops Flight Facility (WFF),
order to insure a quiet geomagnetic environment.
in
This eliminates the need to
disentangle artificial from natural processes, and provides access to substantial
ground resources (particularly the Millstone
Hill incoherent scatter radar facility).
For the experimental region, the electron densities of daytime and nighttime
are
shown
Figure 4.1 and
in
4.2.
lowest ambient plasma density
It
is
Ne we
apparent from these two Figures that the
are likely to encounter in the 1991—1992
The
period (solar
maximum)
this region
10"/cm3, primarily Oxygen. This value can be compared to the neutral
is
is
the 10^ /cm^ level found at night.
density of 3.5xl0iVcm3 found in the
CSU
NASA/LeRC
and
Torr. In the ionosphere, the neutral density
is
neutral density in
chambers
at lO'^
-
lO'S
almost hundred times larger than the
electron density.
The
neutral
environment
gas
N(neutral)>N(ambient) because
pressure data from
SPEAR— 1.
of rocket
Its
a
for
sounding
km
local (rocket)
density
and measured minimum
payload pressure was about lO"^ Torr. Until very late in the
dominates the
has
Figure 4.3 shows neutral
outgassing.
apogee was 370
rocket
outgassing
flight,
environment. [Ref 45]
Typical electron temperature and thermal velocity profiles are shown in Figure
4.4.
These are important
determining
for
the
ambient electron current.
temperature, average mass, thermal velocity are shown in Figure
result in a
and mean
the
Debye—length
free paths as
potential
maximum,
illustrated in Figure 4.6,
shown
distribution
night, 500
km
in
and
and electron
4.5.
Ion
These values
collision frequencies
Figure 4.7. These are important for determining
collected
altitude, are a
currents.
Debye length
Expected
of 0.3
about 600 m, and collision frequency of about 350 Hz. [Ref. 46]
55
values
cm, mean
at
free
solar
path of
lono
800
PAVTIMt
SOLAR MIN
SOLAR MAX.
son
Km
\
400
\
200
J
1
I
I
I
ml
I
I
I
I
I
ml
K'J
I02
J
I0«
1
1
1
1
1
—
J
III
I
I
I
10'
,
ELECTRON DENSITY
1
III
I
I
I
I
I
10''
III
10'
(rm'^i
F\g. 4.1 Ionospheric Electron Density (daytirne)[Ref. 46]
1000
10^
10*
(cm-J)
ELECTRON DENSITY
I
10^
I
I
I
I
Mill
10
I
10
I
I
I
1
1
III
10
I
II
I
I
I
I
ml
10
L
17
I
1
1
ml
10
I
13
I
I
I
1
1
III
10
14
2,
TOTAL COLUMNAR CONTENT (cni'M
Fig. 4:2 Ionospheric Electron Density (nighttinie)[Ref. 46]
56
I
T-n
—
I
1
f
1
m
I
I
"1
iini
—
I
I
I
1
i
m
i
t
n—
|rni
I
I
I
—
iiri
I
\
I
I
—
CD
O
CD
in
^^
in
Q
o
o o
12;
CO
^
CJ
en
o
o Q
W
m
3
o W
o
00
C/1
o
C\2
o
CD
I
o
PRESSURE (TORR)
Fig. 4.3
SPEAR-1
Neutral Pressure Data[R.ef. 45]
r
'^
IDOO
— — —— — —— ——
I
I
:
Si,.-
I
1
I
I
I
I
I
1
I
I
1
I
I
I
I
1
I
I
I
1
I
I
'
'
'
'
MOO
??oo
I*).,
OtGRflSK
•
1
08
1
Fig. 4.4 Electron
inno
1
1?
10
3?
1
1
1
II
1
16
0??
70
18
46]
\
VVt,
/
A,
\
/
\K<--'
6X -
-
1
'
—
'^
\
\
—
/
-^
\
?00
1
on
Temperature and Thermal Velocity[Ref.
\
%
100
1
0?<
"^
\
km
1
0?0
1'
1?
1
Oi
800
1
1
1
mil
y
/\
/
SOLAR "IN
1
\l^^
^^""^^^^^^r^..
«
»
n
10
»
60
V»/Vt
'
1
1
•
1
1
n
i
1
.
.
1
1
.
1
1
.
60
so
10"" K|
.
i
.
1
.
1
3000
?200
1100
600
1
JO
H
1
1
OeCREESK
1
001
1
0?
I.I
0)
kT, /•
1
a
1
?00
400
1
1
01
1
OS
006
1
oo;
I»V)
.
(00
1
1
1
too
1000
i?uo
1
IWO
Vti (V5)
Fig. 4.5 Ion
Temperature, Mass and Thermal Velocity[Ref.46]
58
uuu
/
/
/
-
!
'
/
/
/
—
800
1
\
/
/
//v-^l/
/
/
/
,^
600
/
/
/
/
/
-
^'
DAYTIME SOLAR
CJ"
/
\
"
200
/'
/
//
/
/
/
/
/
—
400
//
/
/
km
/
JLMIN.
'v<r
r^nzizMAX.
^<\-
-
_t
n
I
1
1
1
1
lij
.
,
i
1
i
1.
1
ml
DEBYE LENGTH
I
I
'
'
'
'
'
III
r
I
I
I
I
Ill
III
I
10^
NUMBER OF ELECTRONS
Fig. 4.6
100
(Aq) cm
105
10*
i.LJLLL
i.-l
...1
10
l.O
0.1
Debye
OEBYE SPHERE
A
IN
lO'
Length[R,ef. 46]
1000
/
800
SOLAR MIN
/
vg
/
fmlp
/
/
/
SOLAR MAX
600
e^ip SOLAR MIN
km
400
-
200
J
I
I
I
I
ml
10'
I
I
I
I
I
m
l
I
I
I
I
I
II il
10^
W^
I
t
I
i
I
ml
I
t
10'
m & Hi
Fig. 4.7 Election Collision Frecjuency
69
and Mean Free Path[Ref. 46]
I
I
I
I
II
10'
LAUNCH VEHK'LE
B.
A
Black Biaiit
HOC AT
X
experiirient.
lia.s
been selerterl as the appropriate vehicle
This decision
with an estimated payluad of 010
payload
comiionenls
based
experiment. As illustrated
60-70
is
lb (e.g.
about 10
10 %).
feet.
is
lb.
on
for
launching the
based on the desire to reach 5U0
Table
previous
1
shows the rough allocation
experience
with
km
altitude,
of weight to
UAH-CRIT
tlie
1
Figure 4.8, this gives us a weight growth margin of
in
The Black Brant X outer diameter
The configuiation
of the Black
Brant
X
is
17.26 inches and length
payload
shown
is
in
Figure
4.9.
1200
NASA SOUNDING ROCKET PERFORMANCE
-1
1100 -
1000 -
•00
-
•00
-
UJ
700
-
Z
d
eoo
-
v>
K
O
o
o
a.
<
«» TOMAHAWK
400 -
TAURUS
\TOMAMAWK
TOMAHAWK
300 NIKC
Ames
TEMKie
200 -
BMANT
TAUMUSOMtON
100
• lACK lltANT V
Sl>»tH
ANCAS
1
100
»»ACMt
NIKE CAJUN
ONION
1
1
300
300
MIKE OniON
1
1
400
SOO
n
r
TOO
600
1
1
1
1
1
800
900
1OO0
1100
1200
PAYLOAD WEIGHT IPOUNDS)
I
100
300
—
300
\
400
NASA
—T"
SOO
Sounding Rocket Performance
60
3000
.
PAYLOAD WEIGHT(KC)
Fig. 4.8
y
4
5S0
•00
o
I**
E-riELO
K>OU SCT
Kcmznt
ACS
T
MACNCTOMETen
3^
rRON
ELECTf
SPECTROMETER
STOWtO
MV SUPPLIES
MASS
SPEC.
OEPLOTEO
MJ
t"* t-riELO
BOOM SET
SEALED SECTK5N:
E-riELO ELECTRONICS
fLCMT COMPUTER
ANTCHNAS
HOLLOW CATHODE ELECTRONICS.
tR*NSMITTE»«S
BATTERIES
TANK, REOULATOftS
POWER CONTROL
HOLLOW CATHODE
BATTERIES
ft
ELECTRONICS
NIHKA MOTOR
B ICNITION
Fig. 4.9
HOCAT
Black Brant
61
X
Payload
TABLE
1
ALLOCATION OF WEIGHT AND POWER
Weight and Power
Weight(Pounds)
Section
Nose Cone
Boom
#1
Set
Power(Watts @28
20.
0.
VDC)
12.
0.
Recovery
89.
O.(self)
RCS
72.
O.(self)
Magnetometer
Plasma Analyzers
Scaled Section
Flight Computer
Transmitters(2)
1.4
18.
2.4
12.
18.
4.
112.
Transponders
2.
12.
TV
1.
20.
6.
5.
4.
1.
Batteries
12.
0.
Hollow Cathode # 1
electronics k plumbing
21,
E
Camera
field electronics
Power Control
Altitude Switches
(G—switch, relays, timers)
Hollow Cathode
Current Loop
#
Hollow Cathode
1
#
2
&
Power system
Deployment Mechanism
Batteries
Nihka ignition
k
_
25.
10.
Structure, Cabling
sep
Total
C.
1.
8.4
1.
0.
8.
4.
225.
0.
22.
25.
8.
2.
20.
0.
43.
0.
610.
244.
DEPLOYMENT SYSTEM
For the
HOCAT
experiment, the Canadian
system has been adopted
velocity of
1
OEDIPUS
as nearly ideal for our purposes.
m/s can be obtained by
firing
62
a gas
Canadian deployment
An
initial
jet for 10 seconds.
separation
The maximum
separation will
distance
in
\>c
in
l(Jl)
about
at,
km
.S(JO
apogee. This
a reasonable separation
is
the available time.
FLIGHT SCENAIUO
D.
This rocket
flight
motor burnout occurs
would begin
should provide about 11 minutes o{ flight time. The Nihka
at lUU second, as the lUU
at this point,
tether deployment.
km mark
is
passed.
Our operations
with deployment of the booms, and the beginning of the
The Nihka motor
will
be jettisoned
to
reduce neutral
gas
contamination.
We
100
m
at
have tentatively selected a goal
of
reaching
apogee, at about 8 minutes into the flight.
would then be b minutes
time the tether would
of
maximum
On
data before crossing the 100
be
the
m
downward
leg,
there
altitude mark, at which
and the recovery
jettisoned,
extension of about
effort
(for
the
mother
pay load) would begin. During the operation period, the currents to each satellite
would be determined, both
for passive conditions,
biased with respect to each other.
magnetic
their
of
mass
E.
allow
will
perpendicular to the magnetic
tiie
tether line with respect to the
studies
of
current
collection
parallel
and
field line.
PAYLOAU
The proposed payload, and
UAH,
attitude of
vary with time. Rotation of the tethered pair about
field will, of course,
center
The
and with the packages actively
with assistance from
VVFF
and acquiring standard VVFF
satellite provides
satellite,
in
would largely be
built
and integrated
at
key areas (In particular, machining deck plates,
flight
hardware such as transmitters). The main
power, a basic telemetry system
63
(i.e.
transmitters and receivers),
at.titucle roiitrol,
mid
tests, launch, recovery,
Hollow
1.
The
niagiicloinptrr.
n
ion
VVFF
services include environmental
and ground telemetry-
(,'atliode
core of the experiment
to use the standard hollow
ATS-6
Basic
engine
is
of course, the
plasma source.
Figure
4.1U and
features and size of the Hollow cathode respectively.
4.11
illustrates
Gas (we
will use
propose
SERT
cathode geometry, used previously on the
experiments.
We
II
and
the standard
Xenon)
is
fed
SWAGED HEATER
BARIUM ALUMINATE IMPREGNATED
POROUS TUNGSTEN INSERT-
VAC
r
v
CATHODE ORIFICE^
'k
=
12
V
KEEPER ELECTRODE
-hli(^
Fig. 4.10
Hollow Cathode Design
HEATER WITH FOIL INSULATION
ROLLED TANTALUM FOIL
INSERT (R500 TREATMENT)
7
'
''^
'""^
RING ANODE
mm
5
0.7
Fig. 4.11
Dimensions
64
of
Hollow Cathode
mm
mm
ORIFICE DIA
througli the erul of the rnUiode, i)a.ssing through Uie cylindrical device, with an arc
The porous tungsten
struck at the exit aperture.
insert provides a low
work function
electron emitting surlarc, which
is
typically operated at about lUUO oC, with heat
supplied by an external heater.
A
typical cathode
One problem
inches long.
barium
insert
blown
is
The
source.
for the
1/4"
hollow cathode systems
vvitli
diameter, and a few
in
contamination
is
after
off
supplier
SPIBS
SRS
is
the
Two
launch.
enclosure around the cathode,
designs have been considered for plasma
Hughes Research Laboratory design recently implemented
of the
first
vacuum
resolved by a
is
AF/Flight Model Discharge System (FMDS). This design
successful
of
by ambient atmosphere, and long startup sequences.
This problem
which
is
flown on
design
SCATHA
the
derived from the
The second
satellite.
which provided a cathode system
Inc,
is
Ihe recent
for
possible
SPEAR—
sounding rocket experiment.
Experiment Control and P(JM Encoder
2.
The encoder system developed
collect
and eiuode
all
at
UAH
provides the facility not only to
data (analog and digital) into a high
bit
rate
steam
(751)
kbps), but also to intelligently control complex experiments, to store the data at
Mbps) and
even higher
bit rates
during
interesting intervals.
less
system (about 5U
%
(up to
of a
VAX
Instrument power
of
the
encoder
compatible with
directly
all
1'2
Significant
780) so that
at 28,
drives
to telemetry
NASA
Wallops groimd
back down to the ground
processing power exists within
on— board
+/— 15,
it
processing
and +/->5 volts
standard
facilities.
successful space flights, the last being the
most complex that Wallops has yet flown.
bb
is
S—band
is
this
convenient.
provided.
The output
transmitters
and
is
This system has now completed three
CRIT
I
campaign
whicli
was one
of the
A
is
tlie
than
"burst
1(J
Mil/
iiartinil'irly
valuable aspect of the fast—event—analyzer (Figure 4.12)
mode" capability—the
lor
J— J seconds.
to
ability
Tliis
be particularly
will
transient response of the system to changes
sample instruments
in
for
studying the
the bias voltage on the tether wire.
PROCESSOR 32COI6
48 KWORDS E»PROM
32
useful
at rates greater
RS232 LINK
< »
SERIAL LINK
KWORDS RAM
<
>
OlOrTAL
,52
12-BIT
CONTROL
H MEGABTTES
STATIC RAM
NS 1024
DMA /SEQUENCER
lA) BUSS (8 BIT)
SLOW AMAJ.00
5«0 KSAMPLES/SEC
MEMoirr
BUSS
(IS
'
:a
ANALOO CHANNELS
DIGITAL BITS
I6«BIT ACCUMULATORS
>
BIT)
C^^
FAST ANALOO
.IS
S330 KSAMPLES/SCC
,21
SUPER- FAST ANALOG
.'6
SIS
<:=:>
MSAMPUS/SEC
POWER CONTROL
B
,.2'
LK—-P
SPACECRAFT CONTROL
—-
-«
ANALOO
ACCUMULATORS
BIT ANALOG
ACCUMULATORS
16-BIT DIGITAL
CONTROL
SERIAL LINKS
'."«
7
OR 9
IS- BIT
*/?
DACS. DIGITAL
c:
12-BIT
16-BIT
.
-
-
DAC OUTS
SERIAL LINK
SERIAL LINK
DIGITAL
(<2)
CONTROL
6 BIT
Fig. 4.12
3.
m
Fast Event System
Electric Field Instruments
The
on 3
UAH
electric field
instrument consists of a pair of spherical double probes
boom. The objectives
density and temperature,
tlie
of this
experiment are to measure thermal electron
plasma wave spectrum,
05
electric fields,
and
to provide
a
refprcnre
placed as
far
experiment
payload
iridirating
j)r)iiit,
iiol.eiit.inl.
Two
desip,ii
been siuxessfully used on numerous rocket experiments over
lias
When
the instrument
spherical sensors are
biased
is
operated
is
potentials
fixed
at
current collected by the spheres
obtained,
be
measured. High speed sampling of the electric
NHlz, allowing us
The
measurements
will
watch
to
plasma and the
the
to
and
electric
local
(0—15
kHz)
can
fields
be
data (using the super—fast
field
from 15 kHz to
at frequencies
high frequency waves
for
waves
fiequency
low
relative
computer) allow measurements
flight
the Langmuii probe mode, the
to float with respect to the vehicle, information
about the payload potential can
channels of the
in
measured.
By allowing the probes
waves).
be llown,
apart as practical on the main payload, oriented orthogonally. This
the last decade.
10
pairs of l)Oonis will
electron
(e.g.
electric
for
fields
plasma
or
()n/n
be sampled directly, allowing for observations of turbulence in
the plasma, and key
signals such as the
lower hybrid resonance.
One telemetry
transmitter would be dedicated to these "wideband" measurements.
Electric
Held
Stromfjord, Greenland
is
and
data
density
shown
in
from
a
recent
objective of this pair of instruments
the distribution function of plasma ion
as practicable.
km
apogee.
This
will
is
to obtain
components and electrons
measurements
in as
much
of
detail
require the use of instruments having differential response
angle and energy, with energy coverage from a few
primary tasks are to determine the
nonthermal characteristics
interesting
Sondre
Thermal Plasma Analyzer (Ions and Electrons)
The
in
from
Figure 4.13. As can be seen, the plasma density
ranges from '2*10^ to about lU^ on a typical flight to 5UU
4.
flight
of
eV
to keV.
satellite potential accurately,
the local
The instruments
and to watch
for
plasma which indicate heating, or other
plasma interactions. The detectors can characterize the density, and
G7
29.023 Ful
Doto
Fl ight
I
Dans \f
I
ttltM.M _.
n
^^v\A 11* M^^*^^
— ————
I
I
I
I
T
————————————————
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
Plaaito Fraquaner
ii'^ii
I
I
—————— —
————
I
-t
n^.Syl, » »ww»^ »Vi
I
T
I
I
I
I
I
I
T
———————
I
I
I
I
I
I
Zonal E Field
t
MM
*<
tl.M
.
^^VrivYA^-^^^t.^^'
«
^^ A J^iJ
———
———————————
e
• isa.M
1
I
I
I
I
I
I
I
I
I
I
I
T
I
T
——————
I
I
I
I
I
U*r idlonol t Plaid
Fig. 4.13 Electric Field
temperature
the
of
supra-thermal
low
features
energy
such
angular distributions which
The
which the
UAH
core
of
and Density Data
the
as anisotropic
may
electrons
and
toroidal
electron detector (Figure 4.14)
is
group has successfully used on several rocket campaigns.
at
resolution to separate
the exit
H
,
the
an electrostatic analyzer (ESA)
is
aperture of the ESA.
0"^, and
Xe
.
68
to 3
It is
a
2—
keV, with 10
%
based on a modification of the
order to include ma.ss resolution (Figure 4.15). This
magnet
also
arise.
energy resolution over that range. The ion detector
in
and
heating or acceleration and conical
imaging spectrometer, covering the energy range from rear zero
ESA,
ions,
It
will
is
done be adding a
have sufficient mass
Fig. 4.14 Electrostatic
Analyzer
TOROID
MAGNET
MASS
SPECTROSCOPIC
Fig.
The unusual
4.1,"^)
ESA
ESA
Included Mass Resolution
capability of resolving extremely heavy ions will allow us to
separate locally generated ion distributions from the ambient ions. This
for
making accurate determinations
of negative
heating of locally emitted ions, such as
underway
at
UAH.
experiment. This
is
Figure
4.16
Xe
shows
,
result
from
a giey scale presentation of particle data,
69
important
payload potentials, and inferring
The modification
ESA
is
for
mass analysis
ATS—6
ion
is
engine
where the horizontal
UCSD-ATS-6
Day 292 of 1974
NS IONS
0845-1000 UT
80,000-
10,000 -
>
^
1000 -
>-
o
u
UJ
100-
10
-
200
-20
ROTATION ANGLE
Fig. 4.16
ATS-6ESA
'0
aocis is
detector rotation angle on
axis
ion kinetic energy.
is
The
;^nxis—stabilized sparecraft, and the vertical
tlie
third axis, Ilux,
is
presented as shades of grey, with
black corresponding to zero, or low counts. White, or light grey shaxdes correspond
to high [lux. I'his is seen at high energies
cornor(—!20o to +2(Jo,
(10—8U keV), and
in the
lower
left
to 10 eV). This demonstrates the separation of local
(j
hand
and
distant plasma by energy and arrival direction. Ions from the ion engine (charge
exchange) are found
in
the lower
left
hand corner. [Fief.
SPEAR— 1
Spectrometer results from the
mass analysis) design are
illustrated
experiment using the basic (no
Figure 4.17.
in
36]
The data
arc from a period
where a small sphere deployed on a boom was biased 16 kV with respect to the
sounding rocket body. As a
One— hall
second
the ion flux from 3 to 4
—3
to
potentials
—a
hiter, the
data from Figure 4.17 are found. The peak
A
indicate the detector (and rocket body) are charged to
illustrate
how
fundamental element
the
ESA
design can be used to infer satellite
of the mission. [R«f. 47]
major question to be addressed with this experiment
the currents flow from the satellite, and parallel and
magnetic
in
Direct Current Mecisurements (Current Loop)
6.
how
keV
—4 kV. These data
body was pushed 5 to 10 kV negative.
result, the rocket
field
lines,
'i'raditionally,
ambient particle distributions.
Genrially,
We
however,
is
the question of
jierpendicular to the
currents are measured by using
will
tiie
observed
be making such measurements.
important
function are easily missed, either due to
portions of
.satellite
the
ambient distribution
charging, or gaps on the angular
coverage. Hen( e energy and angle space are inadequately covered. In the proposed
would be highly desirable
mission,
it
density,
several
meters
measurements has
beiMi
Iroin
the
developivl
to
have a direct measurement of current
satellite.
at
A new
UAH. The
technique
principle
of
for
making such
the
current
loop
ENERGY P3 R)
INT.
10^
—
T
1
M
I
I
r
n
I
(230.66.1)
lEBfl
I
I
1
1
n
1
—
1
I
I
I
1
1
r—r-rrTTT
I
1|
xlO'-
X
bJ
10'-
——
10'
T
I
I"
i
T
loi
i
1
1
r
1
1
1
I
M Tl
I
I
I02
I
I
I
IT
iffi
SHEEP ENERGY (ELECTR0MVOLTS)
Fig.
117 Spectrometer
Results from
SPEAR—
ineasureincut, illustraiccl in Figure 4.18, determines the (uireni tiirough a loop of
opticed
fiber
around the
by
measuring
tlie
P'araday
rotation
of
polarized
light
propagating
fiber.
This
tec
hni(jue allows
measurement
kilohertz and a sensitivity of a few
of currents at frequencies
microamps per square meter.
It
is
up to a few
impervious to
electromagnetic interference from the spacecraft, and measures the entire current
independent
of the characteristics of the distribution function or the species that are
72
cnriying
it,
and
is
rclalively
light, \v(Mglit
and rugged. Tlic niirrnt loop
will
be
deployed appioxiuiately J— in iioni the niother paylo-ul.
POLARIZED
INPUT
a
=
v/HdT= v/j-
Current Loop Measurement
Fig. 4.18 Principle of the
6.
TV
dS
Caineia
A low— light
level television
two reasons. The primary reason
is
camera has been included
to
watch
73
for
in
the design (or
visual aspects of the
"ignited
mode", and
collecting
ns(^condarv
to cJi'teiiiiinc
snbsatcllitc.
means
of
'1
Uic oLciit, of
lie
second
iiieasuiing
tlit*
reason
is
to
verily
rockel—satellite
the
"off— the—shelf" caiiieia has been specified. Tlie size
and weight
7.
is
11
('
giaiiis
withnut lens.
'I'lu-
aiound
cHsrliaigc region
is
lOU
suppli(M will he
deplo\
jiient,
separation.
iiiiii
x 47.5
EEV,
tlie
electron
and provide
A
mm
standard
x 50
Inc.
Magnetomete r
Wallops Flight Facility
will
provide a (lux— gate magnetometer with a
resolution of about ](JU iiT that can be directly sampled on three—axes by the
telemetry
mm
systein.
UAH
This should be sufficient for attitude determination and the
detection of any large magnetic disturbances.
74
MODEL OF SYSTEM ELECTRODYNAMIC BEHAVIOR
V.
A.
PROBE THEORY
The design and operation
of the
system described thus
far
depends on the
currents and potentials that will be developed. These parameters are addressed in
the following sections, beginning with elementary probe theory.
and
of electrons,
ions, with, in our case, a substantial
The plasma
consists
background. The interesting
quantities to be determined using a probe are the particle temperatures, particle
densities,
and random current
densities.
These are the quantities that determine the
current flow through the contactors and the tether wire, which
is
the
main theme
of
this study.
Langmuir and Mott— Smith showed that a small metal electrode
inserted
into
the
or "probe"
plasma can be used to determine some of these quantities
experimentally, by applying various potentials and measuring the corresponding
collected
currents.
ignoring for
When
the
random
because
now
theory, the
is
at
ion current.
electrons
is
apply this basic probe theory to the
the extent of the
a probe
the
distribution
We
Debye
have
it
electron current
much
system,
sheath. [Ref. 48]
plasma potential,
The
HOCAT
larger
collects
is
both the random electron and
much
average
greater than the ion current
velocities.
If
the
electron
characterized by a temperature Te, then according to the kinetic
random
electron current density
Je
= No
e Vrandom
is:
= No
6
|^-^
2 nif
TT
75
(8)
where Nq
density, e
is
electron charge, and vth
is
the electrons. Typical ionosphere values give J
(electron) current requires a conducting balloon
If
Debye
we
the average thermal velocity of
is
=
1
mA/m^
whose radius
is
so to collect
12.6
1
m.
ignore the effect of the potential on the ambient plasma distribution
shielding),
we can determine how
A
(e.g.
the particle flux varies with potential. For
a negatively charged body, flux for the repelled electrons can be expressed as
follows:
Flux
where
e(p
>
0.
The
= No |^^^~^"It;^ - No
"^
where e0 >
0.
(1
- -^)
(9)
flux of electrons for an attractive potential can be expressed as
follows:
Flux
l^^
= No
\J^
--
—
e
e
^^f;^ = No
Ir^^ (1 +
-^)
(10)
This expression for a collector with spherical geometry can be
expressed as follows:
Je-Nevth(l+-||^)
where Je
(f)
is
The
is
current density, e
spacecraft potential,
K
is
is
electron charge, vth
is
(11)
the average thermal velocity,
Boltzmann constant, and Te
total current collected to the outer surface of spacecraft
would be the main term
contactor
operations.
for the positive
Figure 5.1
is
is
electron temperature.
just
1= Je*Area. This
end of the tether system, excluding plasma
shows a plot of the electron current
76
density
which would bp rollertod
end
at the positive
of
a tether,
in
the absence of a plasma
The model ionosphere
source device, at thiec different altitudes.
values previously
illustrated were used here.
300_km
408 km
41
-T
.
.
1
1
.
1
1
1
1
r
1
1
1
I
1
1
\
5^
t.e
C4
09
a
t!
u
a
S
(J
V
• .a
^
J
I
I
I
I
L
I
I
I
1
I
I
L
I
I
L
It
Positive Potential
(
V
)
Fig. 5.1 Plot of Electron Current Density vs Positive Potential
77
km
Ion collection at the negative end of the tether system has a similar functional
form. For the ion collector, the equation can be expressed as follow:
Ji
=
where Vi
KTi/M
\
M
,
=
NqVi(l--|^)
The
ion mass.
is
(12)
ion current
much
is
smaller than the
electron current because of the greater mass. Figure 5.2 shows the plot of the ion
current density versus negative tether end potential.
If
we
we can determine
specify the collected current,
0+,
<^-,
and hence the
necessary bias voltage to be applied. Equilibrium requires that the collected electron
current equal to the collected ion current for the system.
moment
small
positive
opposite end
potential
(e.g.
\(f)^\
A/m2, the potential
current density
difference
5.3
—
that collecting areas are equal
is
is
we assume
the
for
system equilibrium should be relatively
one end, relatively large negative potential at the
at
<<
If
|0-|).
about
3, 6,
we assume
and 15
V
that the current density
is
for the three different altitudes.
0.2
This
applied to the ion collection for equilibrium condition. Potential
between the electron collection and ion collection
shows plot of current density versus bias potential
Up
If
to this point,
we have ignored the Debye
we assume Debye approximations can be
the required bias potential. There
is
applied,
is
bias potential. Figure
for those altitudes.
shielding effect
and
collisions. If
we can obtain an opposite
limit
on
a practical limit to the collected current for
large potentials. This can be estimated by considering the current collected at the
edge of the Debye sheath.
If
we
use the
and assume cylindrical geometry, the
I
=
Debye lengths
inferred in previous sections,
total collected current can be expressed as:
Jthermal
78
'
2
;r
r2
(13)
30 0_km
400_km
-f
T
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
r
5^
frj
'
n
a
a
-f.
o
-^
J
i_
J
I
I
I
I
I
I
I
I
I
I
I
-4
Negative Potential
(
V
)
Fig. 5.2 Plot of Ion Curient Density vs Negative Potential
79
L
km
380 k
4ee k
41
r
1
"T
r
r
1
"T
"T
r
5'0l0
la
i«
o
o
>
a
o
J
I
L
J
• .a
I
L
I
'
'
t.s
• .4
Current density
(
A/m^
)
Fig. 5.3 Plot of Bias Potential vs Current Density
80
••
k
where
r
Debye length
is
Debye length
spacecraft and
and
1.1
cm
The thermal
0.78,
length, or
Debye
and 500
km
mA
r
sum
is
of the radius of
Debye lengths are
length.
0.52, 0.75,
which are much smaller than diameter
effect limits the collected current to the
current density Jthermai
maximum
and 0.39
Debye
for small
Hence the Debye
equation (13), the
is 1.4,
large
at altitudes 300, 400,
of the rocket.
values.
for
is
4.7,
and
2.6,
1.3
thermal
mA/m2. Using
current that can be collected in a coUisionless plasma
to the surface perpendicular to the
magnetic
field line (e.g.
saturation current ignoring conductivity—but with turbulence to enhance transport
when
there
is
a magnetic
field).
Obviously, these extremes do not adequately define the problem. For our case
of J
>>
Jth,
</>
>> KT,
alternate approaches
must be found. Theories developed by
Beard and Johnson, Parker and Murphy, and Linson can be applied. These works,
summarized by Linson, estimate the current
the ionosphere.
They
differ
mainly
collection to a highly charged
body
approach to dealing with magnetic
in their
fields.
Ignoring magnetic effects, the vehicle as cylinder along magnetic field with radius
meters, and a collecting area of radius
field.
which
'b'
is
\
<< C = 200 km/3 /
ION
^
Jo
enc
1
rSyrcracJius.
Detye
length
<<
\
EXCLUDED
\
Xl
a
HIGH POTENTIAL REGION
0=0
\
/
Fig. 5.4
B
*
\
Schematic Diagram of System
81
'a'
perpendicular to the magnetic
Figure 5.4 shows the schematic diagram of physical system. [Ref. 49]
»th«on} <«. Usdielllte
in
The
ratio
between
I
and
of the collecting area
lo
can be expressed in terms of the ratio between the radius
and the radius of the rocket as
'
where
lo
thermal current, and
is
experiment.
The
I
follows:
=(^)^
is
(14)
collected
current which
potential can be calculated in terms of
—r- for
is
1
A
this
for
the cases of spherical
space charge limit by equation (15) as follows:
=
(-L
34 a3
)^
(15)
Ao
Figure
shows
5.5
measured data from
the
the
SPEAR— 1
This
experiment.
ionosphere experiment shows the relationship between bias voltage and sphere
current,
bias,
up
the
to 40 kV, 0.2 A.
rocket
body
is
The
(positive) sphere potential
about
—0.25Vbias- These
is
data
about 75
suggest
%
of the
that
the
(approximately) linear relationship of equation (15) can be applied to ionospheric
current collection.
If
the magnetic field effects are partially included, as developed by Parker and
Murphy, the
ratio
"dynamic upper
—i-
can be expressed by equation (16), which gives
gap
=
178*a2 for a typical magnetic
between
a
limit", as follows:
-[;<l +
where 0o
rise to
the
magnetic
field
(^)'
field
limited
82
(16)
strength of 0.45 Gauss.
collection
and
space
To
bridge the
charge
limited
T
1
1
r
r
1
V(kV) = 244
1(A)
O
c
0)
—
i_
1—
13
O
o
o
i_
Q.
CO
9
o
MET
286s
350km
J
I
u
I
J
I
I
alt.
1_
30
40
Potential (kV)
Data from
Fig. 5.5 Sphere Current
collection, Linson has developed a theory
which deals with the magnetic
ratio —I- can be expressed in terms of the potential
is
1/4
—
1.
field.
The
radius
different altitudes above.
'b'
dynamic upper
limit,
of the collecting area
(17).
IJ
—y-
versus bias potential for
and the cylindrical model with
is
5.8, 7.8,
and
11
is
100
For ion collection,
that the area must be
—
m
for three
Figure 5.6 shows the relation between radius and bias
potential so that the radius can be expressed as a function of potential.
that the potential
The
(17)
-
qc0o[ln(20/qc0o)
Figure 5.6 shows the plot of
spherical space charge limit,
magnetic
by equation
field.
200
I
1^-
where qc
SPEAR—
200 V, the radius
first,
much
is
3.5
— 5.7
The
radius
83
is
45
m
we assume
m.
the space charge limited current problem
larger.
If
at
400
km
is
similar so
altitude,
which
in
!••••
.*
+
SPHERICAL SPACE CHARGE LIMIT
(Beard and Johnson)
!•••
w/ B
field(
€
Cylindrical Model)
If
DYNAMIC UPPER LIMIT
(Parker and Murphy)
!•
Potential
Fig. 5.6 Plot of
itett
ite*
!••
r- vs
—"T^
84
(
V
)
Positive potential
t«tt*t
"1
24
1
1
(
(
1
1
1
1
f
\
\
1
————
1
r
1
_
'
'
'
£•
le
12
J
I
1
1
I
I
I
I
I
I
I
I
I
L
J
L
If
Potential
Radius
=
(
V
(X !••>
)
2.253*10-2- (Bias Potential)
Fig. 5.7 Plot of
+
1.173
Radius vs Bias Potential
85
turn gives a higher potential of bout 8 kV. This does not include the important
of
effect
electron
emission
from rocket
surfaces,
due
secondary
to
electron
production from the surface or nearby neutral gas (outgassing). This leads to greatly
reduced potentials. For the
40
B.
kV and
SPEAR— 1
experiment, the illustrated bias potential was
the reduced potentials 0+, 0- were ± 10 kV.
MODIFICATIONS DUE TO PLASMA CONTACTOR
The introduction
of the plasma contactor has two effects.
effective electron emitter. Therefore, the ion collection
comparison with Wilbur's data (presented
problem
is
earlier), potentials of less
is
an
eliminated.
By
V
are
First,
it
than 100
needed to extract Amperes of electrons. Second, the electron collection process
is
enhanced. For the tethered satellite system, the potential difference between the
electron collector and electron emitter will be the potential that
through the tether
we have
to apply
for the tethered rocket system.
Electron Emission
1.
The theory
for electron transport has
electron emission to the space plasma.
When
electron emitter, the plasma density at distance
been applied
in simplified
the hollow cathode
R
from the
orifice
is
form to
used as a
can be expressed
as equation (lS):[Ref. 50]
R=
where Uq
orifice
is
about
ro
(^rVP
produced electron density (about lO^^/cm^),
1
cm, and Uamb
effective collision frequency
is
(18)
ro
is
distance from the
lO^/cm^. For the collisionless plasma, there
which determines the potential-current
be expressed as follow:
86
is
an
relation. It can
U
where a
about
is
0.1,
Wp
be expressed in terms of
is
= Q
(19)
the plasma frequency {^Amc'^fm).
known parameters
+
M^o) = KTe ln(-i\)
namb
A
Up
as function of
9xl0'0
The
potential can then
To:
-^^-^
i lii(-^)
io«^pV^o;-^
*'amb
hollow cathode operating in the spot
mode
at
For those values, the impedance
V
1
for
a
is
mA
a flow rate of 100
equivalent, produces an electron density of about lO'^/cm^ at about
orifice.
(20)
cm from
1
23 Q. Therefore, the potential drop
the
is
23
Ampere system.
Electron Collection
2.
Electron collection can again be considered as a problem of establishing a
conducting plasma cloud,
of sufficient radius to collect the required current.
Figure
shows the plasma cloud expansion model. The three density regions are defined
5.8
here as the inner, collisional, higli density region (a few cm), a transition region (a
few
m
in
radius),
and "ambient" low density
the collecting cloud
potentials.
is
collisionless plasma.
not changed from our previous discussion
The space charge
The diameter
—only
of
the necessary
limited condition has been relaxed by the addition of
positive space charge from the contactors. [Ref. 51]
The current
is
no magnetic
effect,
potential relation can be estimated
and the system
is
if
we assume
operating at constant current.
that there
The plasma
ions stream out across the sheath, canceling a portion of the electron space charge,
formed by the inward streaming electrons. The ratio of the
which
is
in this
case
is
just that
due to the relative mobility.
87
theriiicd fluxes
o
It
3
o
o
o
r.
HI
T
u
a
o
I
o
t/i
o
.0^"
J
At
V)
c
I
o
I
o
K
o
»«
^
-
I
i-l
I
CC
v»
UJ
O
-
§
Q H
C/1
-J
i
9
XT
Fig. 5.8
Plasma Cloud Expansion ModelfPtef.
88
51]
.J.
This relation
is
known
stability condition for a strong
required
sustain
to
a fixed
"Langmuir condition"
^4h-J-
as the
layer.
current
electron
(21)
"Langmuir condition" which
plasma double
current
ion
=
is
At
the basic
this ion current, the voltage
reduced
by
one—third.
If
the
<
le,
the
exceeded where -^m/M*Ie
is
is
<
Ip
generated plasma remains quasineutral and the ions expand hydrodynamically. This
will lead to substantially
reduced potentials. Theories developed by Parks and Katz
and Hastings indicate potentials
1
Ampere
of
V
about 20-30
current. Theoretical works by Dr.
D
for
a
of 0.2 are necessary to
Hastings have been investigating the
radius and gain for the ambient plasma to develop the function
C.
draw
R=
R((^).[Ref40]
SUMMARY
Relatively
driven to the
large (kilovolts)
Ampere
level.
potentials are developed
in
ionospheric systems
For both the electron emitting and electron collecting
ends of the tethered satellite system, locally generated plasmas eliminate the space
charge sheath. The high voltages necessary to transport charge across a space charge
sheath makes the sheath
regions
the
highest
impedance portions
of
the
tether
system. Reducing this impedance by local plasma sources will greatly enhance the
effectiveness
predicted
of
a tethered
satellite
system.
to develop potentials of the order
Systems with hollow cathodes are
10 V. This value
laboratory results which indicate that potentials of order 100
collect
1
estimates
Ampere
is
designed for
electron
current.
The conclusion
that the primary energy range that
is
the
1
—
lOO's of
eV
which
89
V
in
contrast
to
were needed to
results
from
these
the particle detectors should be
range, with at least
kilovolts range.
is
some
capability in the
VI.
CONCLUSIONS AND RECOMMENDATIONS
This study has defined the the Hollow Cathode sounding rocket experiment
(HO CAT). The purpose
currents (up to
1
of this experiment
A) from a
satellite into the
return path through the plasma.
m
Two
is
to
measure the coupling of large
ambient plasma, and the nature of the
bodies will be deployed to approximately 100
separation with Hollow Cathode subsystems mounted at both ends.
will
be aligned along the geomagnetic
practical
deployment system
OEDIPUS
field line
for this rocket
during the
initial
The system
deployment. The
payload will be based on Canadian
deployment system.
Laboratory and theoretical studies of Hollow Cathode performance have been
reviewed. Analysis suggests that for the
bias potential of about 30
V
will
(e.g.
1
to 100 V.
system to derive
A
current, a
power system should designed to cycle the bias
Once the current reaches
1
A, the bias potential will be constant
current limited operation).
Instrumentation
will include particle detectors
keV, and mass analysis
for H"^,
0*, and
Xe"".
with with coverage up to
Hence,
it
will
measurement can be made through the current loop
Faraday rotation of polarized
measurement
light
—
2
Direct current
of optical fiber
propagating around the
of currents at frequencies
1
be possible to infer
negative potentials, and the nature of local plasmas accurately.
few
1
be required. This bias estimate should be within a
factor of 3 of being correct. Hence,
from
HOCAT
fiber.
by measuring
This allows
up to a few kilohertz and a sensitivity of a
microamps per square meter. The langmuir probe
will
provide
us
with
measurements of thermal electron density and temperature, the plasma wave
90
spectrum, electric
wave data
signatures in the
The
Hollow
make
will
and a reference point indicating payload potential. The
fields,
VLF
possible
and
results of this
Cathode
emitting/collecting
works
1
A
HF
to
a
will
current
turbulence,
and plasma heating
how
enable us to understand
source
and
sink.
If
we
the the
succeed
in
current, such subsystems can be applied to long tether
systems which can produce power up to
for
for
frequency range.
experiment
as
search
MW
level. Similarly,
they can be employed
experiments with large electron guns which require current
balance.
This
experiment should demonstrate the validity of theories previously developed
for
new
to
laboratory work, or indicate which
properly model charged particle
physical processes
beam and electrodynamic
91
must be included
tether system.
LIST
1.
2.
OF REFERENCES
G., 4 Guidance Algorithm for Cooperative
/Orbital Rendezvous, pp. 31—45, MIT, 1987.
Stuart, D.
M.
Tether-Mediated
Background Leading to the Tethered Satellite
AIAA
System (TSS)"
24th Aerospace Sciences Meeting, Reno. 6—9 January,
Grossi,
D., "Historical
1986.
3.
Lang, D.
L.,
and Nolting, R.
K.,
Operation with Tethered Space Vehicles
Gemini Summary Conference, Manned Spacecraft Center,
NASA SP— 138,
1967.
4.
Piland, R. O., and Penrod, P. R., Summary of the Gemini Program and Its
Effect on Manned Space Science, The seventh International Symposium on
Space Technology and Science, Tokvo. pp. 27—53,
Publishing Inc.,
1967.
AGNE
5.
1545—03—88—TGM,
Research
Corporation
Report
Spring
Electrodynamic Tether h'orkshop, by K. Kolecki, Section No. 2, 18—19 May,
General
1988.
6.
Team, Utah State University, Report NASA Contract
N ASS— 33383, The Tethered Satellite System, by P. M. Banks, pp. 7—75, 23
Facility
Definition
April, 1980.
7.
Kawashima, N., Hirao, K., Obayashi, T., Raitt, W.
J., Williamson, P. R., Bamks, P. M., and Sharp, W. F., "Results from a Series
of US/Japan Tethered Rocket Experiments", MSA/AIAA/PSN International
Conference on Tethers in Space, 1986.
Sasaki, S.,
Oyama, K.
Maehlum, B.
I.,
W.
8.
Egeland, A. A., Friedrich, M., Hansen, T.,
Holmgren, G. J., Maseide, M., Maynard, N. C, Narheim, B. T., Svenes, K.,
Torkar, K., Troim, J., and Winningham, J. D., "A High Current Electron
Beam Experiment on a Sounding Rocket from Andoya Rocket Range", The
8th ESA symposium on European Rocket and Balloon Programs and Related
Research, (EAS SP-270) SWEDEN, 17-23 May, 1987.
9.
Obayashi, T., Kawashima, N., Kuriki, K., Nagatomo, M., Ninomiyo, K.,
Yanagisawa, M., Kudo, L, Ejiri, M., Roberts, W. T., Chappell,
C. R., Reasoner, D. L., Burch, J. L., Taylor, W. L., Banks, P. M.,
Williamson, R.R., and Garriott, P. M., "Space Experiments with Particle
Accelerators", Science, Vol. 225, pp. 195-196, 1984.
N., Denig,
F.,
Sasalci, S.,
92
10.
Riedler, W., "Proposal for an Active Space Potential Control Experiment for
the STSP CLUSTER Mission Part 1: Scientific/Technical Plan", IWF,
Austira, July 1987.
11.
Hess, W. N., Trichel, M. C, Davis, T. N., WiUiam, C, Geggs, G. E., Kraft,
E. S., Stassinopoulos, E., and Maier, E. J.. "Artificial Aurora Experiment and
Principle Results" Journal of Geophysical Research, Vol. 76, No. 25, 1971.
12.
Winckler,
J. R.,
"An
13.
Winckler,
J. R.,
"A Summary
Investigation of Wave Particle Interactions and Particle
Dynamics Using Electron Beams Injection from Sounding Rockets", Space
Science Reviews 15 (1974), pp. 751-780, 1974.
of Recent Results under the
ECHO
Program
for
the Study of the Magnetosphere by Artificial Electron Beams", Comic Physics
Technical Report 168, 1 September, 1976.
14.
Kawashima,
Sasaki, S.,
"Vehicle Charging Observed in SEPAC Spacelab— 1
Spacecraft, Vol. 23, No. 2, 1985.
15.
Beghin,
and Obayashi, T.,
Experiment", Journal of
N., Kuriki, K., Yanagisawa, M.,
C, "Phenomena Induced by Charged
Particle
Beams", Science, Vol.
225, pp. 188, July 1984.
16.
Olsen,
R.
C,
Orbit— ATS— 5 and
"Experiments in Charge Control at Geosynchronous
ATS—6", Journal of Spacecraft and Rocket, 1985
W.
D., and Smullin, L. D., "Beam— Plasma Discharge: Buildup of
Oscillations", Journal of Applied Physics, Vol. 34, pp. 3421. 1963.
17.
Getty,
18.
Israelson,
G, A., and Winckler,
J. R.,
Electrical Charging of an Electron
ECHO
"Effect of a Neutral
Beam— Emitting
Rocket
N2 Cloud on the
in the Ionosphere:
IV", Journal of Geophysical Research, Vol. 84, No. A4,
1
April.
1979.
19.
Grandal, B., '^Artificial Particle Beam in Space Plasma Studies", pp.
3-33, Plenum Press., 1982.
20.
Burch, J. L., "Space Plasma Phvsics Results from Spacelab-l", Journal of
Spacecraft, Vol. 23, No. 3, May-June 1986.
21.
Watermann,
J., and Wilhelm, K., "Space Shuttle Charging or Beam— Plasma
Discharge: What Can Electron Spectrometer Observations Contribute to
Solving the Questions ?", Journal of Geophysical Research, Vol. 93. No. A5.
pp. 4134-4140,
99
May
1988.
Kaufmann, R. L., Arnoldy, R. L., Moore, T. E., Kintner, P. M., Cahill Jr,
L. J., and Walker, D. N., "Heavy Ion Beam— Ionosphere Interactions: Electron
Journal of Geophysical Research, Vol. 90, No. AlO, pp.
9595-9614, October 1985.
Acceleration",
93
23.
Moore, T. E., Arnoldy, R. L., Kaufmann, R. L., CahiU Jr, L. J., Kintner,
P. M., and Walker, D. N., "Anomalous Auroral Electron Distribution Due to
an Artificial Ion Beam in the Ionosphere", Journal of Geophysical Research^
Vol. 87, No. A9, pp, 7569-7579, September 1982.
24.
STAR
University, Report D131— 1, Neiu Concepts
in
Ionospheric Modification by Storey, L. R. O., Bank, P. M.,
Eraser—Smith, A. C, Gilchrist, B. E., Barker, K. J., and Williamson, P. R.,
pp. 83-133, 1987.
25.
Bausler, B., Tremnann, R. A., Bauer, O. H., Baerendel, G., Bush, R.,
Carlson, C. W., Theile, B., KeUey, M. C., Dokukin, V. S., and Ruzhin, Yu.
Ya., "Observations of the Artificially Injected Porcupine Xenon Ion Beam in
the Ionosphere", Journal of Geophysical Research^ Vol. 91, No.Al, pp.
287-303, 1 January, 1986.
26.
Himter, R. E., and Bartlett, R. O., "Cesium Contact Ion Microthruster
Experiment aboard Applications Technology Satellite (ATS— IV)", Journal of
Spacecraft and Rockets^ Vol. 6, 1969.
27.
Gavrilov, F. V., Myasnikov, A. S., Zhadan, G. G., Orlova, G. S., and Strokin,
M.V., "Some Results of Flight Tests of an Ion— Engine Model Using Surface
loniztion of Cesium on Tungsten", Kosmicheskie Issledovaniya, Vol. 11, No.
1, pp. 140—144, January— February 1973.
28.
Laboratory,
Stanford
Byers, D. C, and Staggs, J. F., "SERT-II Thruster System Ground Testing",
7th Electric Propulsion Conference, Paper 69—235, 3—5 March, 1969.
AUA
29.
Goldman, R. G., and Neiberding, W. C, "SERT-II Mission,
Thruster Performance, and In— Flight Thrust Measurements", AIAA 8th
Electric Propulsion Conference, Paper 70—1125, 31 August—2 September,
Kerslake,
W.
R.,
1970.
30.
G., Staskus, J. V., and Byers, D.
SERT-II Spacecraft Potential Measurements
Probe", NASA
X-2083 1970.
Jones,
S.
C,
"Preliminary Results of
Using Hot Wire Emission
TM
31.
32.
Kerslake, W. R., and Finke, R.
Restarts in Space", AIAA 10th
October-2 November, 1973.
33.
Electric
Propulsion
Multiple
Conference, 31
and Finke, R. C, "SERT-II Thruster Space Restart", AIAA
Electric Propulsion Conference, 19—21 March, 1975.
Kerslake,
nth
W.
C, "SERT-II Hollow Cathode
R.,
Kerslake, W. R., and Domitz, S., "Neutralization Tests on the SERT-II
Spacecraft", AIAA 14th International Electric Propulsion Conference,
Princeton, N. J. AIAA paper 79-2064, 30 October-1 November, 1979.
94
SERT-II Operation",
NASA
34.
Kaufman, H. R., "Plasma Physics Analysis
CR-159814, January, 1980.
35.
Sasaki, S., Kawashima, N., Kuriki, K., Yanagisawa, M., and Obayashi, T.,
"Neutralization of Beam—Emitting Spacecraft by Plasma Injection", Journal
of Spacecraft, Vol. 24, No. 3, May-June, 1986.
36.
"Modification of Spacecraft Potential by Plasma Emission",
Journal of Spacecraft and Rockets, Vol. 18, pp. 462, 1981.
37.
Lewis Research Center Annual Report, Advanced Electric Propulsion and
Space Plasma Contactor Research, by P. J. Wilbur, Colorado State
University, NASA CR 175119, January, 1986.
38.
Wei, R., and Wilbur, P. J., "Space— Charge— Limited Current Flow in a
Spherical Double Sheath", Journal of Applied Physics, Vol. 60, pp. 2280,
Olsen, R.
of
C,
1986.
39.
M.
"Plasma Contactor Laboratory Characterization",
Electrodynamic Tether Meeting, General Research Corporation, 18 Mav,
Patterson,
J.,
1988.
40.
Hastings, D. E., "The Theory of Plasma Contactors Used in the Ionosphere",
Journal of Spacecraft and Rockets,
41.
in press, 1988.
Davis, V. A., Katz, I., Mandell, M. J., and Paris, D. E., "Three— Dimensional
Simulation of the Operation of a Hollow Cathode Electron Emitter on the
Shuttle Orbiter", NASA/AII/PS^ International Conference on Tethers in
Space 17—19 September, 1986.
42.
Olsen, R.
C,
"Electron
Electrostatics, Vol.
Beam Experiments
at
High Altudes", Journal of
20, pp. 43-57, 1987.
McPherson, D. A., and Schober,
W.
43.
"Spacecraft Charging at High
Altudes: The SCATHA Satellite Progrm", Progress in Astronautics and
Astronautics, Vol. 47, "Spacecraft Charging by Magnetospheric Plasma",
edited by A. Rosen, pp. 15-30, AAIA, NY, NY, 1976.
44.
Katz,
45.
Raitt,
46.
Michael, Kasha., Ionosphere and Its Interaction with Satellites, Gordon
and Breach, pp. 7-71, 1969
R.,
I., Jongeward, G. A., Parks, D. E., Reasoner, D. L., and Purvis, C. K.,
"Energy Broadening Due to Spacecharge Oscillations in High Current Electron
Beams", Geophysical Research Letter, Vol. 12, pp. 64, 1986.
W. J., Burt, D. A., Berkey, F. T., Jost, J., Kelley, M. C, Pickett, J.,
Torbert, R., "A Sounding Rocket Pay load to Study High Voltage Current
Collection from the Ionosphere at Low Earth Orbit (LEO) Altitude", AIAA
26th Aerospace Sciences Meeting, 13 January, 1988.
95
47.
Foley, T. M., "SDI Suborbital Launch Yields Data for High-power Platform
Design", Aviation Vetk and Space Technology, 1987
48.
Reitz,
49.
Linson, L. M., "Current—Voltage Characteristics of an Electron— Emitting
Satellite in the Ionosphere", Journal of Geophysical Research, Space
Physics Vol. 74, No. 9, 1 May, 1969.
50.
Parks, D. E., and Katz, I., Theory of Plasma Contactors for Electrodynamic
Tethered Satellite Systems", Journal of Spacecraft, Vol. 24, No. 3,
J.
R., Miiford, F. J., and Christy, R. W., Foundations
Electromagnetic Theory, Addison—Wesley Series in Physics, June 1980.
May-June
51.
McCoy,
of
1984.
"Plasma Motor/Generator Reference System Designs for Power
and Propulsion", MSA/AIAA/PSN International Conference on Tethers in
Space, 18, September 1986.
J. E.,
I
96
INITIAL DISTRIBUTION LIST
No. Copies
1.
Defense Technical Information Center
Cameron Station
Alexandria, VA 22304-6145
2.
Library,
2
Code 0142
2
Naval Postgraduate School
Monterey, CA 93943-5002
3.
Physics Library Code 61
Department of Physics
Naval Postgraduate School
Monterey,
CA
1
93943-5000
WH
4.
Department Chairman, Code 61
Department of Physics
Naval Postgraduate School
Monterey, CA 93943-5000
4.
R. C. Olsen Code 61 OS
Department of Physics
Naval Postgraduate School
Monterey, CA 93943-5000
5.
Professor
1
10
Professor
S.
Gnanalingum Code
GM
61
1-
Department of Physics
Naval Postgraduate School
Monterey, CA 93943-5000
6.
Maj. Yoon, Sang II
Postal Code 138-150
2
Gangdong Gu, Bang-i Dong, Samik
Seoul, Republic of
7.
8.
APT
201-1103
Korea
Maj. Han, Hwang Jin
SMC 2402
Naval Postgraduate School
Monterey, CA 93943-5000
1
Maj. Kim, Jong Ryul
SMC 1659
Naval Postgraduate School
Monterey, CA 93943-5000
1
97
9.
LT. Commander. Youn, Duk Sang
SMC
1254
Naval Postgraduate School
Monterey CA 93943-5000
10.
Lee,
Yong Moon
Ma-Po
Yun-Nam
Dong, 382-17
Seoul, Republic of Korea
11.
Gu,
Cpt. Song,
Tae
Ik
SMC
2686
Naval Postgraduate School
Monterey, CA 93943-5000
12.
Mr. Joe Kolecki
NASA/Lewis Research Center
MS 302-1
21000 Brookpark Rd
Cleveland, OH 44135
13.
Dr.
14.
Cpt. Ryu, Jong Su
Roy Torbert
Physics Department
Universitv of Alabama
Huntsville, AL 35899
__^
^
SMC
2039
Naval Postgraduate School
Monterey, CA 93943-5000
15.
Library, P.O Box 2
Korea Military Academv
Postal Code 130-09
Gong—neung Dong, 556—21 Dobong Gu
Seoul, Republic of
16.
Korea
Library, P.O Box 2
Postal Code 140-01
Yongsan Gu, Yongsan Dong
Seoul, Republic of Korea
98
3^y-s^?
='"dy and
Thesis
Y5273
c.l
Yoon
Definition study and
model for a tethered
sounding rocket.