A simulation facility was established to determine the response of typical spacecraft materials to the geomagnetic substorm environment and to evaluate instrumentation that will be used to monitor spacecraft system response to this environment. Space environment conditions simulated include the thermal-vacuum conditions of space, solar simulation, geomagnetic substorm electron fluxes and energies, and the low energy plasma environment. Measurements for spacecraft material tests include sample currents, sample surface potentials, and the cumulative number of discharges. Discharge transients are measured by means of current probes and oscilloscopes and are verified by a photomultiplier. Details of this facility and typical operating procedures are presented

Berkopec, F. D.

Stevens, N. J.

Sturman, J. C.

English

A simulation facility was established at the NASA-Lewis Research Center to determine the response of typical spacecraft materials to the geomagnetic substorm environment and to evaluate instrumentation that will be used to monitor spacecraft system response to this environment. Space environment conditions simulated included the thermal-vacuum conditions of space, solar simulation, geomagnetic substorm electron fluxes and energies, and the low energy plasma environment. Measurements for spacecraft material tests included sample currents, sample surface potentials, and the cumulative number of discharges. Discharge transients were measured by means of current probes and oscilloscopes and were verified by a photomultiplier

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https://ntrs.nasa.gov/search.jsp?R=19770010165 2020-03-22T10:23:56+00:00Z
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NASA TM X-73602NASA TECHNICAL
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(NASA-TM-X-73602) THE LLWIS ELSEARCH CENTER
GEOMAGNETIC SUESTCRM SIMULATICN FACILITY
(NASA) 8 p HC A0[/MF A01
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THE LEWIS RESEARCH CENTER GEOMAGNETIC
SUBSTORM SIMULATION FACILITY
by Frank D. Berkopec, N. John Stevens, and John C. Sturman
Lewis Research Center
Cleveland, Ohio 44135
R
FEB 1977
RECEIVE
	
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NASA 9TI 
NOr4 1NPpT g
TECHNICAL PAPER presented at the
Spacecraft Charging Technology Conference cosponsored by the
United States Air Force and National Aeronautics and Space Admimstration
Colorado Springs, Colorado, October 27-29, 1976
THE LEWIS RESEARCH CENTER GEOMAGNETTC SUBSTORM SIMULATION FACILITY*
by Frank D. Berkopec, N. John Stevens, and John C. Sturman
National Aeronautics and Space Administration
Lewis Research Center
Cleveland, Ohio 44135
ABSTRACT
A simulation facility has been established at the NASA-Lewis Research
Center to determine the response of typical spacecraft materials to the geo-
magnetic substorm environment and to evaluate instrumentation that will be
used to monitor spacecraft system response to this environment. Space envi-
ronment conditions simulated include the thermal-vacuum conditions of space,
solar simulation, geomagnetic substorm electron fluxes and energies, and the
low energy plasma environment. Measurements for spacecraft material tests
include sample currents, sample surface potentials, and the cumulative number
of discharges. Discharge transients are measured by means of current probes
and oscilloscopes and are verified by a photomultiplier. Details of this
facility and typical operating procedures are presented.
1.0 INTRODUCTION
Geosynchronous spacecraft have experienced anomolous electronic switching
in the midnight-to-dawn region of their orbits. l Environmental measurements
have shown that energies of transient particle fluxes are higher than expected
in this region. 2 > 3y4 Spacecraft anomalous behavior correlates well with the
occurrence of geomagnetic substorms. 5,6 Differential charging of spacecraft
surfaces can occur 7	and breakdown of charged dielectric materials can follow.
Breakdown can result in electromagnetic interference, degradation of thermal
control surfaces, and surface contamination.8
A joint technology program has been implemented by NASA and the U.S. Air
Force to investigate the spacecraft charging 	 One objective of the
joint program is to determine the charging behavior of spacecraft materials in
a substorm environment and what effect configuration has on this behavior.
This information will be used in future spacecraft design practice.
The approach to materials characterization i.s both experimental and ana-
lytical. The results of survey tests for a wide variet of spe.^ecraft surface
materials have been summarized and have been published. 10-14 , An analytical
program has been developed in parallel with the experimental effort. 15 The
experimental work has been performed in a facility specifically developed to
simulate the substorm environment. This substorm simulation facility is the
subject of this paper.
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*Published in the proceedings of
Technology, NASA TM X-73537, 1977.
STAR Category 18
the Conference on Spacecraft Charging
to
22.0 FACILITY DESCRIPTION
The simulation facility was developed to characterize the behavior of
spacecraft materials exposed to a simulation of the geomagnetic substorm envi-
ronment. A schematic diagram of the spacecraft charging test facility is pre-
sented in figure 1.
2.1 Test Chamber
The facility test chamber is a stainless steel vacuum chamber 1.8 m in
diameter and 1.8 m in length. A 1.5-m diameter thermal control shroud lines
the chamber interior. The shroud temperature is controlled by gaseous nitrogen
which can be set to any temperature in the range from -185 0 to +1200 C. The
shroud is aluminum and is painted with a black electrically conductive paint
providing a grounded boundary for all tests. The test chamber is pumped by a
0.9-m (36-in.) diameter oil diffusion pump and typically operates in the range
from 6x10-8 tc 21x10-7
 torr. Pumpdown time is on the order of 90 to 120 min-
utes but generally testing is delayed until samples have sufficiently out-
gassed.
2.2 Simulation
The substorm environment is simulated in discrete increments. The aspect
of the substorm environment that is of most interest is the electron environ-
ment. It is simulated with a monoenergetic electron beam operated at a voltage
in the range from 0 to 30 kV and at a current density in the range from 0 to
5 nA/cm2 . The divergent electron beam is generated from a hot wire filament
by means of a spherical segment accelerating grid kept at ground potential.
The cathode and the beam-forming grids are biased negatively relative to this
accelerating grid. The electron beam current density is uniform to about
30 percent over a diameter of 0.5 m at the test plane. The test plane is ap-
proximately 1 m from the accelerating grid.
Solar situlation is used when photoeffects are to be determined. A 3/4-
sun intensity xenon lamp is used; intensity is measured at the test plane.
The solar simulator is located outside the chamber and the radiation is passed
through a quartz window. The spectral distribution, with the quartz window of
the chamber, is within 10 percent of that recently published 16 for solar radi-
ation.
Low energy plasmas are simulated by means of a gaseous nitrogen electron
bombardment plasma source. Nitrogen gas is admitted into a discharge chamber
-ontaining a hot wire filament cathode and a cylindrical shell anode. A mag-
netic field coil is spirally wound around the anode to increase the path
length of the bombarding electrons from the cathode to the anode and thereby
enhance the ionization efficiency. Plasma densities from about 10 particles
per cm3_up to 10 6 particles per cm3 can be simulated. The plasma source is
routinely used to discharge samples after testing.
32.3 Sample Accommodation
Figure 2 is a photograph of the test chamber ?nterior. Samples to be
tested are mounted on a three-position sample rotator. Up to three different
samples can thus be tested during each pumpdown of the facility. Vacuum can
be maintained for several weeks for survey tests of three samples. Samples
up to 30 by 30 cm in size can be accommodated. The sample under test is lo-
cated on the test chamber centerline as is the electron source. The electron
source is mounted on the chamber door seen partially on the right in figure 2.
2.4 Instrumentation
2.4.1 Electron Environment Measurements
The electron current density at the test location is measured with a
Faraday cup. The Faraday cup is mounted to a 30 by 30 cm metal plate which
shields the test sample. The Faraday cup-sample shield assembly is positioned
in front of the sample only while the current density is being set. The en-
trance area of the Faraday cup is 10 cm2 . The suppression grid is operated at
-40 volts. Stationary current probes are mounted around the sample. These
probes are plain metal disks, 5 cm 2 in area, that are used to monitor the cur-
rent density at the test plane throughout testing. the current density profile
of the test plane is obtained by sweeping a rake of 5 current probes across the
test chamber. The Faraday cup, shield, stationary current probes, and rake can
be seen in figure 2.
2.4.2	 Sample Measurements
r Two basic measurements are made when a sample is subjected to the simu-
lated substorm environment.
	
The first of these is the sample electron current-
to-ground.	 Samples are generally mounted on a metal plate substrate with the
dielectric surface facing the beam. 	 During test the substrate is grounded
through an electrometer. The current through the sample is then obtained as a
function of time.
Sample surface potential is the second basic measurement made.	 Surface
potential is measured by sweeping an electrostatic voltmeter probe across the
sample surface.	 The electrostatic voltmeter is a noncontacting capacitance
coupled device.	 'The electrostatic voltmeter operates on a null balance prin-
ciple whereby the surface potential probe is brought to the potential of the
sample surface by a high voltage power supply. 	 This design provides accurate
measurement and'minimizei voltage gradients in the measurement location. 	 This
measurement is made in the electron beam.	 Since the probe and the sample are
nearly the same potential, the probability of arcing between the probe and the
surface under measurement is minimal. 	 The response time of the device is
20 milliseconds to change 10 kV; this is faster than typical charging times
being measured.	 The probe-to-sample surface spacing is generally maintained
' at 2.5 mm; resolution is within 5 percent at this spacing for spots larger
than 9 mm in diameter or strips wider than 6.5 mm.
4When arc discharges occur, some additional data is taken. Loop antennas
are used to sense and quantify discharge activity, The loop antennas are
15 cm in diameter and the plane of the loop intersects the plane of the sample
within the sample area. The antenna-to-sample spacing is about 0.7 m. The
signals received by the antennas are amplitude discriminated such that all
sensed pulses of greater than several specific magnitudes are counted. The
cumulative number of discharges of amplitude greater than 1, 2.5, and 5 volts,
for example, at the input to the discrimination circuitry then becomes the
basic discharge data. When discharges occur, the sample current measuring
electrometers are shorted out of the measurement circuitry and the sample cur-
rent directly grounded. Inductively coupled current probes and fast oscillo-
scopes (100 and 250 MHz) are used to measure the-arc-discharge currents. A
photomultiplier tube is used to sense the visible emission portion of the dis-
charges. The photomultiplier is also used to confirm the current probe data.
A vacuum spark gap in the test chamber is used to periodically verify the
functioning of the discharge monitoring circuitry.
One of the most frequently used pieces of test chamber apparatus is a
Polaroid camera which is used to photograph discharges. Discharge locations
as well as some visual discharge characteristics are documented. Time expo-
sures are made for varying periods, depending on the frequency of discharging.
3.0 TEST PROCEDURE
3.1 Initial Conditions
Prior to any testing, all instrumentation is calibrated. The test chamber
is then evacuated to a pressure of less than 5x10 -7 torr before any equipment
is operated. Samples are generally maintained in vacuum for up to 16 hours
before any high voltage testing is performed. Outgassing for this period has
been found to be good practice. Before any testing is performed the sample
surface potential is measured and discharged with the plasma source. The state
of the sample surface is determined from measurements by the sample surface po-
tential probe.
3.2 Establish Electron Substorm Conditions
w„
The electron beam is established by bringing the Faraday cup-shield as-
sembly to its position in front of the sample shielding the sample from the
electron beam. The proper electron beam conditions are then seta These con-
ditions are typically a beam voltage of 2, 5, 8, 10, 12, 14, 16, 18, or 20 kV
negative at a current density of 0.5,-1, or 3 nA/cm 2 . Testing is performed by
starting at the lowest beam voltage and current density and increasing these,
in steps, as the test progresses,
3.3 ,Testing
Testing is typically performed by setting the beam conditions and step-
ping through increasing beam voltages at a given current density, increasing
5the current density, and then again stepping through increasing beam voltages.
The sample is discharged with the plasma source before the beam voltage is
changed. In this manner, conditions from -2 kV at 0.5 nA/cm 2 to -20 kV at
3 nA/cm2 are imposed upon the sample. If the test is a survey test, each con-
dition is maintained for 20 minutes or until equilibrium is attained, whichever
is longer. When long term effects are under investigation, the specific condi-
tions of interest are imposed on the sample for periods of days or weeks as
appropriate.
Testing is routinely done in the dark and at ambient temperature. When
photoeffects are to be determined, testing is repeated with the solar simulator
illuminating the sample. Simulation of solar eclipse conditions can be done by
testing with and without solar simulation for given periods of time. Eclipse
testing might be performed, for example, with a -20 kV beam at 1 nA/cm 2 for
30 minutes alternating periods of solar simulation and darkness.
4.0 CONCLUDING REMARKS
The LeRC substorm facility is in continuous, reliable operation. Charac-
terization of spacecraft materials is in progress and some results have been
rn_ported. The facility is modified to incorporate new techniques of measure-
ment and simulation as they are required or as they are available. Independent
development of instrumentation is continuously maintained and, when significant
instrumentation advances are achieved, they are incorporated into the facility.
REFERENCES
1. Fredricks, R. W. and Scarf, F. L. (1973) Observations of Spacecraft Charging
Effects in Energetic Plasma Regions in Photon and Particle Interactions
with Surfaces in Space. R.J.L. Grard, Ed., D. Reidel Publishing Co., pp.
277-308.
2. DeForest, S. E. and McIlwain, C. E. (1971) Plasma Clouds in the Magneto-
sphere, J. Geophys. Res. 76 (No. 16): 3587-3611.
3. DeForest, S. E. (1972) Spacecraft Charging at Synchronous Orbit, J. Geophys.
Res. 77 (No. 4): 651-659.
4. Bartlett, R. 0., DeForest, S. E., and Goldstein, R. (1975) Spacecraft
Charging Control Demonstration at Geosynchronous Altitude. AIAA Paper,
75-359.
5. Pike, C. P. (1975) A Correlation Study Relating Spacecraft Anomalies to
Environmental Data, in Spacecraft Charging by Magnetospheric Plasmas,
Progress in Astronautics and Astronautics, vol. 47, A. Rosen, Ed., Am.
Inst. Aeronaut. Astronaut./Mass. Inst. Tech. Press, pp. 45-60.
4
^	
A
y
y
r
Y
'r
.	 a
6S. Shaw, R. R., Nanevicz, J. E., and Adamo, R. C. (1975) Observations of Elec-
trical Discharges Caused by Differential Satellite Charging, in Space-
craft Charging by Magnetospheric Plasmas, Progress in Astronautics and
Aeronautics, vol. 47, A. Rosen, Ed., Am. last. Aeronaut. Astronaut./Mass.
Insc. Tech. Press, pp. 61-76.
7. [dhipple, E. C., Jr. (1975) Observation of Spacecraft Generated Electro-
static Fields in the Vicinity of the ATS-6 Satellite. AAS Paper 75-220.
8. Stevens, N. J., Lovell, R. R., and Gore, V. (1975) Spacecraft Charging
Investigation for the CTS Project, NASA TM X-71795.
9. Lovell, R. R., at al. (1975) Spacecraft Charging Investigation; A Joint
Research and Technology Program. Paper presented at Spring Annual Meet-
ing of the American Geophysical Union, Washington, D.C.
10. Stevens, N. J., Klinect, V. W., and Berkopec, F. D. (1976) Environmental
Charging of Spacecraft Surfaces - Tests of Thermal Control Materials for
TT.... -- .. I. - lI_L_I n- -441...J e.. 6.......... V11n1,M C-.	 17a1,4.1n - PavM 1.
11. Stevens, N. J., Berkopec, F. D., and Blech, R. A. (1976)
f!h avn{no of Snananrafh Cnrfanec — Teafa of Thermal Cor
X-
12. Berkopec, F. D., at a1. (1976) Environmental Charging Tests of Spacecraft
Thermal Con trol Louvers, NASA TM X-73517.
13. Berkopec, F. D. and Stevens, N. J. (1976) Testing and Evaluation of Solar
Array Segments in Simulated Geomagnetic Substorm Charging Conditions.
Presented at the IEEE 12th Photovoltaic Specialists Conference, Baton
Rouge, La.
14. Stevens, N. John, et al. (1976) Testing of Typical Spacecraft Materials
in a Simulated Substorm Environment. Paper presented at the USAF/NASA
Spacecraft Charging Technology Conference, Colorado Springs, Colo.
15, Purvis, C. K., Stevens, N. J., and Oglebay, J. C. (1976) Charging Charac-
teristics of Materials: Comparison of Results with a Simple Analytical
Model. Paper presented at the USAF/NASA Spacecraft Charging Technology
Conference, Colorado Springs, Colo.
16. (1971) Solar Electromagnetic Radiation, NASA SP 8005.
x
0I
Figure 2. - LeRC Substorm Simulation Facility test chamber interior.(-) '?IGINAI " PA61,,, IS
()b' "0()R QUALITY
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Figure 1.
	
Schematic diagram 0! the LeRC substo ► m simulation facility.
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PREGMING PAGE BLANK NOT FILMED 


The Lewis Research Center geomagnetic substorm simulation facility

t
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goht_to +
1. introduction 423
2. Facility Desc=iptibn. 424
3. TeSt ProCedure 428
4. Co_cluding RemarkS+ 429 LI
FteferenceS 429
1. The Lewis Research Center Geomagnetic
Substorm Simulation Facility
l_mnkD. Be&ope_ N. JohnStevenS,andJohnC. Stun_an.
HntionnlA_fiauflcs anL.dew_iseRes_lrd_AdmihistraflOncenter
Cleveland,Ohio +
t
Abstract
A Simulation facility has been established at tha NASA-Lewis Rese_trel_ Center
to det_rrnine the response of typical spacecraft materi_ls to the georna_netic sub-r
' storm environment and to evaluate instrumerttatton that will be usefl to monitor
spacecraft system response t_ this environment. Space environment conditions
simulated inclUde the thermal-vaeuttm conditions Of space, solar simulation, geo-
rnagllette sUbstorm electron fluxes And energies, and the 10_ energy plksma
envirot_ment. Meas_ren_ents for _pacecraft mtLteria1 tests include _am_le eu=rentS,
stunple stir, face potentials, and the cumulative number of discharges. Disch_tz_ge
tPan.sients are measured by means of _ttrz_ent probes And OsciliOeco_eS and are
verified by a photOmulHplier. Detttils of this ftteility and typical operating proee-
dttres _tre p_esented.
/
1. INTRODUCTION
GeosynChr_nous spacecraft have experienced anomai0uk electt-Onic _witcliing
/
in the mittnight-tt_-dgwn region of their OrbitB. 1 Environrnentai measurement_ ._t_k_J rllI
_, .,¢ge . *''
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https://ntrs.nasa.gov/search.jsp?R=19780002212 2020-03-20T12:38:43+00:00Z
, f I l ! t 1 ! l
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l_ve sh()wn that ene,-_les of transiet_t particle fldxes al-e higher than ex_.ected hi
this region. 2, 3, 4 _ ace_l-_ft _nomalotts behaviOr _orrelates well with the occur-
rence of geomaghetic sttbstormS. 5, (] Differehtial oh_trging o_ spa_eenaft Surfaces
can occur, "/ and bl_eakdO_n of cha_'ged dielectric materials can follow. Breakdown
can result in electromagnetic inter_e_'ence, degradation Of thermal control sur-
face.% and surface contamination. 8
A joint technolog_ p_-ogram has beet_ implemented b_"NASA and the USAF-to
investigate the Spacecraft charging pheru)menon, 9 One objective of the Joint _,_ _
program is to determine the charging behavior of spacecraft materials in a sub-
storm environment and what effect configuratiun has on this behavior. This in-
formation will be used ir_ future Spacecraft design practice.
The approach to materials characterization is both experimental and analytical.
The reStilt_ of survey tests for a wide Variety of spacecraft surface materials
have been sttmmarlzed anct have been published. 10-14 An analytical program haS
been developed in parallel with the experimental effort. 15 Th-. experimental work
has been performed in a facility specifically developed to simulate the subStorm
; environment. This SUbstorm _imulation fatality is the Subject of this p£per.
2. FACILITY DESCRIPTION
The simulation facility w_s developed to Characterize the behavior of space-
c_'aft m_tterials exposed to a simulation of the geomagnetic SubStorm environment.
A schetnatie diagram of the sp/tcecr_ft charging test facility is preSe.ted in
Figure 1.
2.1 Te_t Chamber
The facility test chamber iS _ st_ifiless steel vacuum chamber 1.8 m in
diameter and i. 8 m in length. A L 5-m diameter thermal Contx_ol shroud lines the
chamber interior. The shroud temperature is Controlled by gaseous nitrogen
which Can-be set to any temperature in the range _rom -185 ° to +120oc. The
ShrOud i_ aluminum and is painted with a black eie_tri_ally conductive paint prb-
_,tding [ grounded bouhdary for all tests. The test chamber is pumped by & 0.9 m
(_6oin,) diL,neter oil diffusion punip and typically operates in the range from
' 6 × 10"8 to 2 × 10"7 tOrx% Pumpdown time is on the order of _0 to 120 minute_ but
generally testing is delayed until s_mples have sttfficiently out_as_ed.
2.2 Simulation
The Sub_rm er,virOnmet_t is bimui_teti in discrete increments. The a_pect
of the substorm envi:_onment that i_ t_f most interest i/_ the _lectr0n enVironment.
424
00000005-TSD03
I ! I 1 ........... ! ! ill
: s
j '
,j ,
_'_2 ll,.mo. _. cuii[m Ill L L
I,] llUlilPill TKI_ llllOltisY_nVlllOllr I" I | --II!cll li_ |
_ -- Iillilli ltll/ S_lCllll. 1"1_'" | | J I
r_ iil,t,ci .,i, i. L I I
-'-_-"_"-: VOtllilf 1i,. ....
------+ _ --" "'-- ,, r [ [ [
_k,'i , [ _., " ] , ,, --i--l--r-I
_++- ,_++ ,++
_+_. + t ililil# if
-' _ _ _ --, _ I''<-+'<'-I _I_','- ;,
_!!_ <_ _. IPt" ll'tll.ll'""
_ Figure I. Schematic Diagram of the LeRC SublitormSimulition
_I" F acility
_, iti._simuiatedwitha monoenergetic electronbeam operated at a voltageinthe
_:_i, range ffrOm 0 to 30 kV and at a eurre_ density in the range t_'om 0 to 5 nA/cm 2.
_ The divergent electron beam it generated from S hot wire filament by mebns Of a
_ spherical Segrdent accelerating grid kept at grotmd potential The cathode and the
_}i! beam-_orrning grids are biaseti negati,_ely relative to this accelerating g_id. The
_!i electron beam current density iS uniform to about 30 percent over a diameter of
_,, 0.5 m atthe testplane.The te_tplane is approximately 1 m _rom the accelerating
_ gri_.
_'i_i Solar simulation is useti when photoeffects are to be determined. A 3/4-sun
intensity xenon l_mp is used; intensity is measured at the test plane. The solar
_-_',_- simulator is located outside the chamber _nd the radiatto_ is passed through a
,, quart_ window. The llpectral distribution, with the quart_ window of the ehamber,
:+ iS within 10 percent of that recently published 16 _or Solar radiation.
_, LOWenergy plasmas are simulated by means of a +_aseouS nitrogen electron
bOmbardfnent plaSm_ _ource. Nitrogen gas is adrilitted into a discharge chan_beP
+_! containing a hot Wll+e filament cathode anti a cylindrical shell anode. A magnetic
i field toil iS spirally wound around the atiOde to incz'ease the p_th length of the<,.+ bombarding electrons from the cathode to the anode and thereby enhance the iOniza-
, tion efficiency. PldSma dtnsttles from about 10 particles per cfn3 up to 10e par-
_°++ ticies per cm 3 can be simulated. The plasma Source iS routinely used to discharge
;++ samples after testing.
.....- 425
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+::,_".+-- .......................... :.+......................... +......... .+,....... _U+",.,........ f: .... .... I +_+i...................•_L.:]+. o:+- :_ I_ :+ " +:l _ j : : " + _+
?'++"+ ,_.+ : +M + 2,.---_0++-++..............r+-+-_a,_-_,+-'_++_+'?' 0_'+o++ + ° ................
......_" ...... ++,_ - ...... 00000005-TSD04
2.3 Sample,_,_,eommbdat|on
Figure 2 is a phOto/traph of the test chamber ir_terior. Samples to be teated
are mounted on a three-p0sitton Sample rotator. Up to three different Samples
can thus be tested during each pumpdOWn of rite facility. Vacuum can be maintained
for several weeks for Survey tests of three samples. Samples up to 30 by 30 cm in
size can be accommodated. The sample under test is located on the-test chamber
eenterltne as is the electron Source. The electron source is mounted on the Cham-
ber door S_en partially on the right in Figtlt'e 2.
Figure 2. LeRC Sttbstorrn _Imulatlon Facility Test Chamber
Interiol"
2.t [nstruntentatitm
2.4.1 ELECTRON ENVIRONMENT MEASUREMENTS
The electron currerlt density at the test location iS measured with a Faraday
cup. The Faraday cup is moul_ted tO a 30 by 30 cm metal plate wlitch shields the
test sample. The Faraday cup-Sample Shleid assembly ia positioned in front of
the sample only while the currerit density iS beiflg Set. The entraiice area of the
Faraday cup is I0 cm 2. The st_ppr_salOn grid is ope_'ated at -40 volts. Staiionary
426
i I I ......I ! i' I I
clu'rent _r0bes at'e mortared bt,ound the Satnple. These probes ate l_'_:n metal
disks, 5 _r/_2 tn.ai, ea0 that are used to monitor-the cu'rrent density s: the test plane
throughout testing, The current densRy profile of the test plane is obtained by
sv_eepi_g h rake of 5 c_rre,t prObeS acrOSs the test• chamber. The Faraday crop,
shield., stationary Curt"eat pi-obes0 and _ake can be seen in Figure 2.
_. 4.2 SAMPLE M_ASUREMEN-TS
Two basic measurements _re made when a sample is Subjected to the siluu-
i:' lated subStOrm environment. The ft,-st of these is the sample elec+,'_n current-
i _ to-ground. SanlpleS are generally mottnted on a metal plate stibsti. : _vith the
i dielectric Surface facing the beam. During test the substrate is grounded through.
ir
an electrometer. _e current-through the sampl_ iS then obtained aS a funct_n o_ .......
'_ " time.
i_ Sample surface I_otential is the second basic measurement made. Surface
potevtial iS measured by sv_eeping an electrostatic voltrfleter probe across the
i sample gurface. The el,_ctroStatic voltmeter is a noncontac_ing c_pacitance Coupled
device. The electroSt_.tic Voltmeter operate_ On a null balance principle whereby
i- the stirface potential, probe is brought to the potential of the sample surface by ai.
i high voltage power supply. This desi_=n provides _ccurate measurement and
. minimizes volhAge gradients i_ the measurement location. This measurement is
made in the electro_ beam. Since the probe alid th_ sample are nearly the same
potenti_d0 the probability of arcing between the probe and the Suz-face under mea-
Surement is minimal, The response time o_ the device is 20 msec to change 10 kV.;
this is faster th_n typical charging times being measured. The prObe-to-sample
Surface Spacing is generally maintained at 2, 5 ram; resolution is within 5 percent
'' at this Spacing _or spots larger than 9 mm in diameter or strips wider than 6.5 ram.
i :" When arc discharges occur, some additional data is taken, Loop antennas are
! " used to SenSe and quantify discharge activity. The loop antennas are 15 cm in
! ,: diameter and the plane of tile loop intersects the plane of the sample within the
i_- sample area.- The antenna-to-s_mple spacing is about 0. 7 m. The signal_ re-
ceived by the antennas are amplitude discrimlnated-such that all sensed pulses of
i " gre_ter than several SpeCific magnitudes are counted, The cumulative number of
i - dtscharge_i of amplitude greater, than 1, 2, 5, and 5 voltS, for example, _t the input
i_ to the discrimination ci_cttttry then becomes the bathe discharge data. When dt_-
: charges occur, the sample current measuring electrometet_S are ShOrted out Of the
! . tfleasurement circuitry and the Sample current dire_tiy grounded. Irtductively
i "" coupled current probes and fast oscillo_cope_ (I00 and 250 MHz) are used to tries-
_ sur_ fl_e _re-tliScharge currents. A phot_muitipiier tube iS Used to SenSe the
i ;
___ Visible emission portion Of the dischargeS. The photomuliiplier Is ai_o used tO
periodically verify the functioning of the diScltarge monitoring circuitry.
i-_!." 427
I :f
00000005-TSD08
i/
One of the in, St frequently used pieces of test chamber apparatus iS a Polaroid
camet'a wttich iS used to photograph discharges. DtSehat_ge lOcatiotis aS well aS
s0tne visual discharge characteristics are documented. Time exposures are made
for varying periodic, dependin_ on the freOuenc_v of d|s_h_t,_tn_ ....
3.._..TESi" PROCEDUI_,.
3.1 Initial Condition-........ rob,,
Prior tOany testing,allinstrumentatiot_is calibrated.The testchamber i_
thenevacuatedto a pressure of lessthan 5 × 10-7 torrbeforeany equipment is
operated. Samples are generallymaintainedinvacuum for up to 16 hr beforeany
highvoltagetestingisperformed. OUtgassingfor thisperiodhas been found tobe-
good practice. Before any testingiSperformed the sample surfacepotentialis
measut*edand dischargedwith theplasma source. The stateof the sampie surface
iS.determinedfrom meaSurements by the sample surfacepotentialprobe.
3.2 Establish Electro_.SubstoFmConditions
The electron beam-iS established by bringing the Faraday cup-shield assembly
tO its position in front Of the sample shielding the Sample n-ore the electrott beam,
The proper electron beam conditions are then Set. These conditions are typically i
a beam voltage of 2. 5,. 8, 10, 1_.. 14, 16. 18, or 20 k_"negative at a Current 1
density of 0. 5, 1, or 3 nA/cm 2. Testing is performed by starting at the lt_west i
beam voltageand currentdefmityand increasingthese,inStepS,aS the test ]
progresses.
3.3 Testing
Testing is typically performed, by setting the beam conditions and stepping
through increasing beam voltages at a given current density, increasing the cur-
rent density, and then again Stepping through increasing beam Voltages. The
sample is discharged with the plasma SOtlrce before the beam voltage is changed.
in this manner, conditions from -2 kV at 0. 5 nA/cm 2 tO -20 kV at 3 nA/cm 2 are
imposed upon the Sample. If the test is a _urVey test. each condition l_ main-
tained for _0 minutes Or until eqtiilibriutrt is attained, wlticheVer is longer. When
long term effect_ are under investigation, the specific conditions of ititerest are
imposed on the sample for periods of days or weeRs as appropl'iate.
TeSting iS routinely dotte in the dark and at ambient tempel'ature. When
pho{oeffectS are to be determined, testing i_ i'ep_ated with the solar simulator
illuminating the sanip|e. SlrhtiiatiOn of solar eclipse conditions can be done by
428
00000005-TSD07
testing With anti without s0iar simulation for given periOdS of time. EclipSe
testing might be pet'Eormed, for e_tanipleo. With a -20 kV beam at,.1, _/cm 2 fbt"
30 rnit_ttt_alternatingp_riOds of Solttrsimulationak_.darltkmss.
I. CONCLUDINGHE_IAI_KS
The LeRC _Ubstorm facility is in continuous, reliable operati6n, Characteri-
zation of spacecraft materials is it_ progress and some reSultS have been reported.
The facility iS modified to incorporate new techniques of measurement and simu- _"
lation as they are required or as they are availabie. Lndepvndent development of
inStrumentatiOn is continuously maintained and, _hen Significant instrumentation
advances are achieved, they are incorporated into the facility.
References
I. FredrickS, R.W., and Scarf, F.L. (1973) ObServationS of spacecraft Charg-
ing effects in energetic plasma regions in Photon and Particle Interact, ions
with_urfa.ces in Space, B. J. L. Gr_rd, Editoro D. _eidel PUblishing Co.,
pp. 277-308.
2. DeForeSt, S.E., and McIlwairh C.E. (1971) Plasm_ clouds {n the magneto-
sphere, J. Geophy. Re_. 7_6fNo. i6):3587-3S11.
3. DeForest, S.E. (1972) Spacecraft charging at SynchrOnous orbit, J. GeophyS. _i
Res._.....7_(No. 4):651-659.
4. Bartlett, R.O., DeForest, S.E., and G01dstein, R. (19_5)SpaCecrdh !
Charging Control DemonStration at GeOS_n_hron0uS Altitutfe, AI_A' Paper7S-359. • "l
5. Pike, C.P. (1975) A correlation StUdy relating Spacecraft _tflomalieS to
environmental data, in Spacecraft CharglnB by Ma_netospherlc Pi_m,as°
Progress in AStrOrtautiCS anti A.stronaUtiC$° voi. _7, A. Hosen, t_di_or. ':
Am. Inst. Aeronaut. AstrOnaUt./Mag$. inst. Te_h. Press, pp. 45-80.
_. Shag,, R.R., Nanevicz, J.E., and Adamo, R.C. (1975) Observati0ris of
electrical d!scharges caused by differentiae satellite charging, in _pace,
craft Char_m_ bY Magnet0Spheric Pi_sma, Progress in Ast_0rtatJ¢'_d
/_eroneutics, VoL 47, A. Rosen, Editor, Am. Tn_t. Aeronaut. AStronaut. /
_daSs. InSt. Tech. PresS, pp. 61'-7_.
7. Whipple. E.C., Jr. (1975) Ob_ervation of Spacecraft Gene_ate_i Electt0static
Ft,.._,Ids in the Visinity O_the A TS;'6 Satellite, AAS Paper 75,220.
8. Stevefls, N.J., Lovell, R. _., and Gore, V. (1975) Sp._cecraft Char_Ing
Investlgation for the CTS P_oject, NASA TM X-71#95.
9. Lov@llo 1t. R., et ai. (i975) Spacecraft chargirl_ t_Ve_t|gation: A joint f'e-
search and t_chnology prog_'a_n, Paper presented _t Spring AnnUal Meetin_
of the American Geophysical Union# Washlrigt0ri, D.C.
, 429
• , _ . ¢ tt t ! l i
.............. 00000005-TSD08
i!.
i \
!i: 10. Stevefis, N.J., Kltnect, V.W., and BerRopec, F.D. (1978) Environmental
t: C.hargttlK.of Spacecraft $.urfaceS: Tests ..of ThermbirCoptr_pl M.$te,rlals for
i_: L!de 0n the GI_0b01ppoitiotting..sy;o.te._m Flight space Vehicle - Part 1: '
_ Specimens 1, to 5, NASA TM Xo73467.
_" 11. SteVenS0 N.J., Berltopec, F. D. 0. ar_ Blecho R.A. (19_61 ,F-nviro_mental
_ Ch_trgtng of Spaceeraft Surfaces: T_st$ of Thermal Cotttrol Materials for
_: Use On the Global F.ositi0n[nK system Flight space vehieie - _art 2:
i .... Specimen 6 tO 9, NAS_ TM X-7"3436.
j, 12. Berkopec, F.D., et al. (19761 En_tronmentgl _h.argtng Tests of SpacecraftThermal Control Lo_tVers, NASA TM X-735
i _, 13..... Berkopec, F.D., and StevenS, N.J. (1976) Testing and evaluation of Solar
i _i_, array segments in simulated geomagnetic SubStorm charging cOndition_,i_
_' Presented at the IEEE 12th PhotoVoltaic Specialists Conference, Baton
! _ l_ouge,_ La.
i-_ 14. StevenS, N. John° et al. (1976)Testingof typicalSpacecraftmaterialsin a
;-'_ Simulatedsubstorm environment, Paper presented atthe USA F/NASA
!_ SpacecraftCharging"Technology Conference, Colo=acloSpringS,_Coto.
_ 15. Parvis, C. L. 0 Stevens, N.J. ° and Oglebay, J.C. (197_) Charging character-
i!i: iSticsof materialS:.CompariSon of resultswitha simple allalyticalmodel,Paper present d at the USAF/NASA _pacecraft Charging Technology Con-
ference, Colorado Springs, Colo.
i l'' 16. (1971)Solar ElectromagnetlcRadiation,NASA SP 8005.
ii
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i si_"
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00000005-TSD09


The Lewis Research Center geomagnetic substorm simulation facility

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