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Hollow cathode plasma source for active spacecraft charge controlWilliam D. Deininger, Graeme Aston, and Lewis C. Pless Citation: Review of Scientific Instruments 58, 1053 (1987); doi: 10.1063/1.1139607 View online: http://dx.doi.org/10.1063/1.1139607 View Table of Contents: http://scitation.aip.org/content/aip/journal/rsi/58/6?ver=pdfcov Published by the AIP Publishing Articles you may be interested in Self-heated hollow cathode discharge system for charged particle sources and plasma generatorsa) Rev. Sci. Instrum. 81, 02B305 (2010); 10.1063/1.3258033 Electron beam generation in a diode with a gaseous plasma electron source II: Plasma source based on ahollow anode ignited by a hollow-cathode source J. Appl. Phys. 94, 55 (2003); 10.1063/1.1577229 Characterization of the interaction between a hollow cathode source and an ambient plasma J. Appl. Phys. 71, 4709 (1992); 10.1063/1.350661 Summary Abstract: A hollow cathode for ion beam processing plasma sources J. Vac. Sci. Technol. A 1, 258 (1983); 10.1116/1.572108 Hollow cathode discharge as a plasma source for H— production Rev. Sci. Instrum. 52, 1459 (1981); 10.1063/1.1136475
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Hollow cathode plasma source for active spacecraft charge control William D. Deininger, Graeme Aston, and Lewis C. Pless
Jet Propulsion Laboratory, California Instilute a/Technology, Pasadena, California 91109
(Received 4 June 1986; accepted for publication 24 February 1987)
A prototype plasma source spacecraft discharge device has been developed to control overall and differential spacecraft surface charging. Such charging phenomena have contributed to many satellite operating anomalies and systems malfunctions. The plasma source is based on a unique hollow cathode discharge, where the plasma generation process is contained completely within the cathode. This device can be operated on argon, krypton, or xenon and has a rapid cold start time ofless than 4 s. The discharge system design includes a spacecraft-discharge! net-charge sensing circuit which provides the ability to measure the polarity, magnitude, pulse shape, and time duration of a discharging event. Ion currents of up to 325 /-lA and electron currents ranging from 0.02 to 6.0 A have been extracted from the device, In addition, the spacecraft discharge device successfully discharged capacitively biased plates, from as high as ± 2500 V, to ground potential, and discharged and clamped actively biased plates at + 5 V with respect to ground potential during ground simulation testing,
INTRODUCTION
The tendency of spacecraft and spacecraft surfaces to charge has been noted since the beginning of satellite flight. Spacecraft charging appears to be particularly severe at geosynchronous orbit, Spacecraft frame potentials of thousands of volts negative with respect to the ambient space plasma have been observed during eclipse on the ATS-5 and ATS-6 spacecraft 1-6 and the SCATHA 1,7,8 satellite. The largest potential observed to date was - 19 kV on ATS-6 during eclipse, but even in sunlight potentials as high as - 2.2 kV have been observed.5
,9 These high potentials result from a natural balance of the charged particle fluxes to the spacecraft surfaces from the ambient electron and ion populations fonowing a geomagnetic substorm. uo Spacecraft potentials are generally negative since electrons have higher mobilities as compared to ions. Overall spacecraft frame charging enhances surface contamination which can cause the degradation of thermal and optical surfaces. In addition, charging interferes with science measurements of the ambient space environment.
Of more concern than overall spacecraft charging is differential charging of adjacent surfaces, or differential charging resulting from charge deposition in dielectrics and insulators, Differential charging and its subsequent electrostatic discharge (ESD) is believed to be responsible for much of the anomalous behavior seen on various satelIites,\),ll-lfJ The transient electrical impulses produced as a result ofESD can couple into the spacecraft electronics and cause problems ranging from logic upsets to complete system failures, In addition, ESD can cause mechanical damage to spacecraft surfaces and enhance surface contamination,
As a result of the connection between anomalous satellite behavior, differential charging, and ESD, considerable effort has gone into developing methods for preventing spacecraft charging using both passive and active means. Passive means include the use of conductive materials on spacecraft surfaces, use of proper grounding techniques, and
filtering of electronics. 17 Active spacecraft charge control encompasses the use of a wide range of charged particle emitters, induding electron, ion, and plasma emitters. Active charge control devices which have been most successful in discharging overall and differentially charged spacecraft emit a neutral plasma, Furthermore, only devices emitting a neutral plasma have been found to maintain spacecraft potentials near the ambient space plasma potentia! and significantly suppress differential charging.2
-7
•17
This paper describes a hollow cathode, neutral plasma source which was designed for use as a plasma emitter on the Air Force geophysics laboratory, beam emission rocket test-1 (BERT -1) sounding rocket experiment, 18 The device was also going to be used as an electron emitter during one part of the sounding rocket flight. The BERT -1 rocket experiment was to examine the extent and cause of spacecraft charging created by the ejection of charged particle beams at low altitude and the ability for a system to automatically discharge space vehicles. The sounding rocket experiment environment determined the instrument requirements which included a start time ofless than 5 s, a total lifetime of at least 450 s (including an ability to start 36 times), and an ability to extract at least 20 f-lA of ion current and 0,02 A of electron current. Launch window constraints coupled with rocket subsystem testing requirements prevented final integration of the plasma source into the BERT -1 sounding rocket.
I. PRINCIPLE OF OPERATION
A charged spacecraft or spacecraft surface cannot generally bleed its charge back to the surrounding low-density space plasma since the electrical conductivity of the intervening plasma sheath region between the spacecraft and space plasma is too low, In order to reduce the potential of the spacecraft, it is necessary to artificially increase the electrical conductivity between the spacecraft and surrounding space plasma. An effective way of doing this is to form a relatively high-conductivity plasma bridge to allow the pas-
1053 Rev. Sci. Instrum. 58 (6), June 1987 0034·6748/87/06 t 053· t 0$01.30 @ 1987 American Institute of Physics 1053
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sage of positively and negatively charged particies through the sheath. The plasma blidge provides an additional source of charged particles for current balance to reduce and prevent spacecraft-surface potential buildups. This process may be thought of conceptually by visualizing the spacecraft and surrounding space plasma as the charged plates of a capacitor and the sheath as the medium between the plates. The plasma bridge acts as a shorting strip across the capacitor. The plasma bridge must be established in a controlled way with an adequate particle flux and spatial extent to ensure that the spacecraft, or portion thereof, is returned to a potential near zero without experiencing an uncontrolled and potentially damaging arc discharge (ESD).
A precedent has been established in the plasma generation technique used for active spacecraft charge control by the successful charge control experiments on the A TS-6 and SCATHA satellites which used small ion engines and hollow cathode plasma bridge ion beam neutralizers.3
-7 During
tests with these satellites it was observed that the hollow cathode plasma bridge neutralizer was sufficient to control satellite charging under a variety of space plasma environments. These results are consistent with previous experiments where hollow cathode plasma bridge neutralizers were used to neutralize the positive ion beam emerging from an ion engine spacecraft propulsion system. 19
-21 For that
application, the hollow cathode created a plasma bridge for electron flow between itself and the ion beam edge which was generally several centimeters away. The plasma bridge process can be particularly important in suppressing daylight differential charging by providing an additional source of electrons to compensate for the suppressed photoelectron emission due to potential barrier formation. 22
A hollow cathode discharge typically produces plasma densities on the order of 102°_1021 m -3 just downstream of the cathode orifice. Electrons in this plasma have an average temperature of about 1.0 eV while ions have the cathode thermal temperature of approximately 0.1 eV. The much higher electron speed means that, in the plasma bridge coupling process, the electrons are carrying most ofthe current. Furthermore, ambipolar diffusion causes ion flow by spacecharge coupling of the ions to the emitted hollow cathode electrons so that a significant ion population is always present in the plasma bridge.
II. DESIGN CHARACTERISTICS
The field enhanced refractory metal (FERM) hollow cathode was selected for the plasma source since it had previously demonstrated an ability to start in less than 5 s as well as provide resistance to contamination. Only some basic features of this hollow cathode will be described here since it is well documented elsewhere. 23
-25 A schematic of the plasma
source, delineating the basic features of the hollow cathode, containment vessel, control electrode, and the power processor is shown in Fig. 1. The hollow cathode is started by applying 300-400 V between the central emitter tube and surrounding cathode barrel, as shown in Fig. 1. In this arrangement, the cathode barrel acts as an anode and the working gas, which is flowing through the cathode, breaks
1054 Rev. Sci. Instrum., Vol. 58, No.6, June 1987
r-----,(±)
~~-------------------, KEEPER SUPPLY
,----,1-)
START GA TI NG SUPPLY 1+) DIODE
8.08 ohms
FIG. I. General schematic of plasma source system.
KEEPER
VESSEL
SPACECRAFT CHASS! S
down resulting in a glow discharge to the end of the central cathode emitter tube. Ion bombardment of the emitter tube rapidly heats the tube end to thermionic electron emitting temperatures. The much larger heat capacity of the cathode barrel prevents this component from heating significantly during the starting process. When the tube reaches thermionic emission temperatures, the cathode transitions from a glow to an arc discharge, resulting in a low coupling voltage between the emitter tube and surrounding cathode barrel. Once this arc is established, the cathode is said to be ON and is a stable, high-density plasma source. The entire starting sequence takes less than 5 s.
Typical volt-ampere starting characteristics for the plasma source hollow cathode are shown in Figs. 2(a) and 2 (b). The power consumed in the hollow cathode discharge is shown in Fig. 2 (c). The most critical parameter controlling the turn-on time and hollow cathode power consumption is the heat capacity of the end portion of the emitter tube where the plasma discharge occurs, Figure 3 shows the hollow cathode emitter tube assembly. The FERM hollow cathode is significanily different from most other cathode designs in that the plasma generation process is contained completely within the hollow cathode barrel in the sense that no external sustaining electrodes are needed to maintain the discharge. It was found from previous studies of the FERM cathode24
•25 and during this development program that only
the last 2-3 mm of the emitter tube emitted thermionic electrons and participated in the plasma discharge. As a result, the emitter tube tip was optimized for minimum heat capacity while providing good mechanical strength. This optimization process resulted in the particular emitter tube design shown in Fig. 3.
The lifetime of the plasma source is dictated by the robustness of hollow cathode emitter tube. Ion bombardment of the emitter tube tip, which enables thermionic electron emission, also causes sputter erosion of the tip, reducing the
Spacecraft charge control 1054
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0 , la)
-4+ -1--\---r-----r-,----------,-r~.....r'"' .. _J
~ -80~ g -120~ -' ::: -160~ ..... z w I-
-200r 0 "-w
'" 0 -240~ :r: I-< u
~l -320 ~
-360 3.2
2.8 -
on 2.4 ~ Q
"'-E 2.0 co ,...: Z UJ 1.6 a:: a:: :::> u w 0 0 :c 0-< U
es 200 5: o 0.. 100-
(bJ
TIME, seconds
FIG. 2. Typical plasma source starting characteristics.
~
1
~ J I I l 1
cathode lifetime. This is, in fact, the Hfe-limiting mechanism. The optimized emitter tip, shown in Fig, 3, has a lifetime of hundreds of start and run cycles with a demonstrated total operating time in excess of 10 000 s.
FIG. 3. Flight version of cathode emitter tube assembly.
1055 Rev. Sci. Instrum., 1/01. 58, No.6, June 1987
..•.••• ; •••••••• ' •.••• " ................ -•••.• :.~.: ••••••••••••••.••• ~.:.:.~ ••••••• ;-•••.• :-.:.: ••••••• ; •• ~."7: .. ;-:.:.; •••••••••• --; .... :.:.:' •• : •••••••••••. -. .•
No rare-earth oxide impregnates are used in the FERM cathode emitter tube to lower its electron work function, thus eliminating the poisoning problem which often occurs with low work function impregnates upon exposure to air and moisture. In the past, controlled vacuum enclosures were required before and during launch to prevent contamination of the rare-earth oxide hollow cathodes.6.7.19-21.26.27 The pure tantalum emitter tube of the FERM cathode can be exposed to humid air after use and then placed under vacuum again and restarted with no significant change in its start or run characteristics.
An important parameter in successfully starting the plasma source was ensuring that the hollow cathode internal gas stagnation pressure was in the proper range, given by the Paschen breakdown curve, for a minimum potential glow discharge to occur. For the plasma source cathode barrel and emitter tube geometry, low Paschen breakdown voltages are realized, for most gases, with a pressure distance product of about 1.0 Torr em (1.33 Pa m).28 Various tests were performed to determine the emitter tube to cathode orifice plate separation that gave the most reliable cathode startup and the lowest run power requirements. Eventually, a separation distance of 0.0048 m was selected. Consequently, the stagnation pressure within the cathode for most reliable startup was about 2.0 Torr (266 Pa). The plasma source has been operated on krypton, argon, and xenon. Stagnation pressure measurements were taken (with the cathode off) at a variety of flow rates for each gas tested, Cross plotting these data for the optimum cathode internal pressure of 2.0 Torr (266 Pa) gives the required gas flow rate as a function of the molecular weight of the flowing gas as shown in Fig, 4. This curve is for the particular cathode orifice diameter of
lOOr-----~--~--~-------,-------.-------,
80
~ «: .... : :r: <.::> 20 ~ a:: ..; -' :::0 U
"" JD -' 0 :2' VI is « <.!)
6
~
FIG. 4. Plasma source gas flow requirements as a function of gas molecular weight.
Spacecraft charge control 1055
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TABLE 1. Plasma source characteristics.
Parameter Value
Startup time: Cold 3--4 s Hot < 1 s
Working gas krypton Startup energy 1904 W s Run power 196 W' Particle currents:
Extracted electron 6.0Ab Extracted ion 3.S2X 10- 4 Ab Electron leakage Ux 10- 5 A
Mass 12.2 kg Lifetime
Demonstrated 10 800 SO
Duty cycle (max.) 1.0 Starts 350b
• Value for highest power electron collection mode, state 7 (seC Table VI). b Highest value seen during characterization testing, higher values possible.
0.001 m and orifice plate thickness of 0.0018 m used in the plasma source cathode developed under this program. Krypton was chosen as the working fluid at the sponsors' request so that a mass spectrometer could be used to differentiate between plasma source efflux and argon and xenon particles from other devices during the rocket flight.
Some general characteristics of the plasma source spacecraft discharge device are summarized in Table 1. Plasma flows with either predominantly electron or ion currents may be extracted through the cathode orifice by applying either a positive or negative bias, respectively, to the keeper electrode shown in Fig. 1. A negative keeper bias inhibits electron flow while a positive keeper bias actively draws electron flow. Ambipolar diffusion ensures that, for the few tens of volts keeper electrode bias, a plasma will be drawn from the cathode. Electron currents ranging from 0.02 to 6.0 A and ion currents up to 352 /LA have been coupled to external electrodes.
POWER PROCESSOR STRUCTURAL PACKAGES,
--'-
PLASMA SOURCE POWER LEADS
III. IMPLEMENTATION
The plasma source system is outlined in Fig. 5. The basic system consists of the structural vessellhollow cathode chamber, gas-handling system, and the power processor. The major components are described in detail below.
A. Structural vessel and hollow cathode
The plasma source structural vessel/hollow cathode chamber serves as the mounting structure for the hollow cathode/keeper assembly. This vessel is fabricated from 304 stainless steel and its overall length is approximately 0.2 m with a maximum width of approximately 0.15 m.
The keeper electrode, mounted in an insulating, boron nitride block, is a tantalum plate containing a 0.OO3-m-diam center hole, The hollow cathode barrel consists of a molybdenum tube with an 80% dense, tungsten plate welded to one end which contains a O.OOl-m orifice. The odfice serves as the exit for the low-energy, hollow cathode plasma.
The emitter tube assembly shown in Fig. 3 consists of a commercially available vacuum, high-voltage feedthrough welded to a stainless-steel mounting flange. A stainless-steel tube is welded to the threaded end of the feedthrough which supports the tapered, tantalum emitter tube, threaded onto the other end of the stainless-steel tube. The stainless-steel tube and the first 0.005 m of the tantalum emitter tube are covered by a mullite insulator.
B. Gas system
The gas system stores, regulates, and delivers krypton gas to the plasma source hollow cathode. The gas-handling system is comprised of a reservoir, fill-fitting and pressure transducer, latching valve, pressure regulator, flow-metering valve, and the interconnecting cabling and tubing. Table II summarizes the main characteristics of the gas system. The components in this assembly were provided by the spon-
EM IHER TUBE
/- CATHODE BARREL
---'1 CONNECTORS FOR: - INPUT POWER 1
1 POWER I
f-'+--Io------,,.:--..:~..L.., I-"~_}- KEEPER
- TELEMETRY - COMMANDS
PROCESSOR 0-----\ ELECTRONICS
L_'----~
FILL ~ FITTING, / PRESSURE TRANSDUCER
LATCHING VALVE
1056 Rev. Sci. !nstrum., Vol. 58, No.6, June 1987
\-'-, 'I :
" I ~ '--- -~--~ ~'SOLATORS \
-- PRESET \STRUCTURAL NED OLE VESSEL
\ VALVE GAS LINE
PRESSURE REGULATOR
Spacecraft charge control
FIG. S. Block diagram of plasma source system.
1056
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TABLE II. Gas system characteristics.
Parameter
System type Reservoir
Ca) Maximum operating pressure (b) Volume ( c) Typical capacity (Kr) ( d) Material
Latching valve (al Type (b) Operating current
Pressure regulator (a) Type (b) Outlet pressure (c) Outlet pressure adjustable range (d) Minimum inlet pressure
Pressure transducer Ca) Type (b) Mounting
Flow metering valve Ca) Type (h) Flow metering range
(c) Maximum inlet pressure
Value
Regulated, high pressure
1000 psi O.Sl 50 standard litcrs Steel
Solenoid latching 1 A-open; 0.1 A-close
Aneroid 5 -+ 0.3 psia 5-10 psia
20psia
Seminconductor Suih into a screw and attachcd to fill fitting
Diaphragm 6.0-60.0 sccm with 50-psig inlet pressure (for He) 250 psig
sor and aside from the flow-metering valve were spare parts from the satellite positive ion beam system (SPIBS) instrument (a small ion source26
,27 ) used on the SCATHA satellite. The metering valve outlet is connected to the hollow cathode gas inlet via Tygon tubing to electrically isolate the gas system from the hollow cathode. Porous, metallic plugs
+28V , LOW ~ VOLTAGE KEEPER CONVERTER -+5, tI2 CONVERTER iHO[;SEKEEP1NGI
V-
~ TELEMETRY
L START
BUFFER CONVERTER
V-
t CURRENT FEEDBACK
PRESSURE SIGNAL
TEMPERATURE
L.e.- RUN
are slip fitted into the Tygon tubing near the hollow cathode gas inlet to force plasma recombination should plasma back up in the flow tube. This prevents the plasma from grounding to the gas system and extinguishing the arc discharge. A krypton inlet pressure of 5 psia was provided to the flowmetering valve by the pressure regulator, resulting in a flow range ofO.S-9.S sccm.
C. Power processor
The primary functions of the power-processor assembly include operating and controlling the plasma source, providing telemetry data, and accepting commands from the microprocessor controller. The plasma source power processor consists of three source-operating, dc-de converters (START, RUN, and KEEPER), a low-voltage dc-dc converter which provides the converter control voltages, and the associated command and telemetry circuitry. A functional block diagram of the overall power processor is shown in Fig. 6. The power processor assembly has been used to start the plasma source several thousand times and operate it for a total time in excess of 45 h.
1. Primary converter design and operation
The operating characteristics for the three primary converters are given in Table III. Current-regulated, dc-dc converters are used since the loads are plasma discharges and the converters must limit their output currents. The plasma source is started by activating both the START and RUN converters, which are connected in series, after the gas flow
P R 1+) TO KEEPFR 10 E ELECTRODE IL V I A E i R R II 5 i A ~ I Y L
TO CATHODE EM ITIER TUBE
1-)
't~ ~- GATING DIODE
1+) FIG. 6. Block diagram of power processor.
~
CONVERTER
COMMANDI (+) TO CATHODE ISOLATOR BODY BUFFER
<2J NET 8.08 51 CURRENl MONITOR
SPACECRAFT FRAME
1051 Rev. Set Instrum., Vol. 58, No.6, June 1981 Spacecraft charge control 1051
••••••••• :.-•• ; •••••••••• ".,. ••• -.~.: ••••• , •• ' ••••• > ••••• ".--•• ; ........ ' ••••• "; •••• " .. :.;.;-•••••••• , •• :: •• : ... ; ••• -•••••• .. ····.:.··.·~.:.·:·.·.·t.·.·.···,···,··.,-· ••.••.
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TABLE III. Characteristics ofthc three primary dc-de converters.
Output Current potential characteristics
Converter (V) (A)
START" 100--300 LO RUN 15-60 2.0 KEEPER 10--60 0.3-2,0
a 30-8 maximum operating time with a 0.16 duty cycle.
has begun. The open circuit voltage of the RUN converter is about 88 V, and that of the START converteris about 390 V. The sum of the output voltages of these two converters is sufficient to initiate a glow discharge in the hollow cathode. The minimum power required to maintain the glow discharge and start the source in less than 5 s is 400 W, When this occurs, the start converter limits the glow discharge to its current set point of 1.0 A. Since the RUN supply set point is 2.0 A, the RUN converter is switching at the maximum pulse width providing about 88 V at 1.0 A. As the cathode emitter tube temperature rises (due to heating from ion bombardment) the cathode terminal voltage drops. The START converter maintains the current at about 1.0 A until cathode heating results in thermionic emission of electrons and the cathode transfers to the arc discharge mode. At this point, the cathode terminal voltage drops below the RUN converter output voltage of 1.0 A allowing the RUN converter to conduct more current. This allows the cathode plasma voltage to drop along a negative resistance characteristic and forces the ST ART converter to shut down, since the cathode current is higher than the START converter set point. At a cathode current of about 1.2 A, the START converter is forced completely off since the START converter pulse-width modulator is forced to zero-duty cycle by the current feedback signal (see Fig. 6) . At this point, the gating diode in the START converter provides a bypass path for current around the START converter. The negative-resistance characteristic of the arc discharge causes the cathode current to rapidly increase to 2.0 A, which is the control set point for the RUN converter. The system is now on with the RUN supply providing 2.0 A at the cathode arc-discharge operating potential, typically 40--50 V, and the START sup~ ply is off. The system is a stable, high-density plasma source in this mode. The entire starting process takes less than 5 s. Typical start characteristics for the plasma source are shown in Fig. 3. The efficiency of these supplies, during both the start and run modes, is approximately 80%. The run supply is regulated to within 1 % ofthe cathode run current. Typical turn-off characteristics for the plasma source are shown in Fig. 7. Cathode shut-down times are generally on the order of2.0ms.
The KEEPER converter is a constant-current converter with a reversible-output polarity. The KEEPER converter output voltage is referenced to the cathode emitter tube potential (Fig. 6). This converter is used to bias the keeper electrode to preferentially extract either positive or negative particles from the cathode plasma. In the electron collection mode (keeper positive), the KEEPER has three operating set points. The set points are defined in terms of electron
1058 Rev. Sci. Instrum., Vol. 58, No.6, June 1987
2.3
2. 0 b-"--"v--....r""\.. ~
~ 1. 7[-.: 1.4' 5 r ~ l.l~
~ ::t ~:t~_-,--_-,--
+2 -i-----r--.--.----.
-3
-8
;;f -13 >= z U,j
5 0-W C>
-18
-23 o ~ -28 '-'
-33 J\) -38>--
-42~ -4Bt'-Y~ i
2.53.03.54.04.55.05.5 TIME, milli,econds
FIG. 7, Plasma source shutoff characteristics,
l ]
t t I ~ 6.0 6.5 7. a 7.5
current collected by the keeper electrode, and are 0.3, 1.2, and 2.0 A, at keeper potentials of 30-50 V. The keeper supply is regulated to within 1 % of the emitted electron currents. There is only one set point with the keeper electrode biased negatively, corresponding to collected ion currents of tens of microamperes. When biased negatively, the keeper electrode acts to inhibit electron flow from the cathode. Since the collected ion current is well below the first KEEPER set point, when biased negatively, the converter output potential remains at its maximum (open circuit) value of approximately 88 V.
2. Support function circuitry
The support functions are those circuits which do not directly operate the plasma source but which are essential for the control and monitoring of the source. The major support circuitry consists of the command buffer and isolation, housekeeping processor, and telemetry.
There are six commands available to operate the plasma source, as indicated in Table IV. A command is comprised of an 8-bit input word defining the desired source state. The command isolatorlbuffer provides the command interface for the source inverters. Input commands are optically isolated, stored, and then buffered. The command buffer stores the commands issued by the controlling source and shifts the voltage level from the TTL (5 V) to the high-level (12-V) signals required by the converters. The command isolator
Spacecraft charge control 1058
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TABLE IV. Plasma source command capability.
Command
(1) Run on/off
(2) Start on/off
(3) Keeper on/off
(4) Keeper medium current enable
( 5) Keeper low current enable
( 6) Keeper polarity enable
Function
Activates/deactivates RUN converter Activates/deactivates START converter Activates/deactivates KEEPER converter in high current (2.0.A collected electrons) collection mode. Selcct the medium current (1.2-A collected electrons) collection mode. Selects the low-current (O,J-A collected electrons) collection mode. Selects positive (electron collection) or negative (ion collection) keeper electrode bias.
consists of two sets of optical isolators which transmit signals between the control point and the command buffer while electrically isolating the two grounds. The stored commands are isolated and transmitted to the command source for display and/or verification.
The telemetry circuits are provided to buffer. scale, and clamp the analog signals which show the operating condition of the plasma source and electronics. The signals are all scaled at 0-5 V and are clamped at approximately - 0.5 and + 6 V. The 12 telemetry outputs are summarized in Table V
and can be grouped into two types: plasma source operation monitors and plasma source output monitors, Eight of the output telemetry functions are used for source-operation monitoring. These include the hollow cathode voltage and current, the keeper voltage, current and polarity, gas reservoir pressure, input battery voltage, and the temperature on the RUN converter baseplate.
TABLE. V. Analog telemetry outputs.
Description
Hollow cathode voltage" Hollow cathode current Keeper voltage Keeper current Integrated net currentb
Amplified net current Gas reservoir pressure Temperature of RUN converter baseplateC
Battery voltage Keeper polarity Net current integrator polarity Amplified net current polarity
Actual value for 5 V ±." 5% output
400 V 2.0A 88 V 2.3A 0.44 As 25mA 480 psia -IO'C
48V positive See Note 1 See Note 2
Note 1: Used to monitor the sign of the charge being accumulated by the integrator. 5 V ±: 5% means negative charge is being accumulated.
Note 2: Used to monitor the sign of the net charge leaving the plasma source. 5 V ± 5% means electrons are flowing from the source to spacecraft ground, positive charge emitted.
aIn two ranges: 0-70 V, 0--400 V. b Saturates at 4.4-V output and provides a measure of the total emitted charge (coulombs).
C Inverted scale, 0.0 V ± 5% corresponds to 90 'C
1059 Rev. Sci. Instrum., Vol. 58, No.6, June 1987
.... : •••••••••••• .: ••••• ;. ••• -." •• ;.:.;.-;.~ ••••• / •• ; ••• ; •••••••• T •••• :.: •••• :.:.:.:.: ••• .> •••••••• ~ •• ;: ••• 7" ••.•••• :.:." •• :O:.;.;.; •• -.l-•••••• ;O •• / •••••••••••• ,.:< .... ;-: ..•.........•.•. ,.-......... -........ ,. ',' ...... ',' .~ .•...
The entire power processor system is isolated from spacecraft ground through a common 8.08-n resistor (see Fig. 6) which acts as a current path for charged particles flowing into and out of the plasma source. The voltage drop across the resistor provides a means of measuring the plasma source output current, During a discharging event, a plasma with a net negative or positive charge will leave the area of the spacecraft depending on the nature of the original biased surface. A corresponding amount of electrons will have to flow into or out of the plasma source through the common resistor from or to spacecraft ground. Monitoring the current flow through the resistor can provide information on the amount of charge involved, the duration of the event, and polarity of the plasma cloud. The voltage drop across this resistor is the input signal to the net current monitor, sourceoutput monitoring circuitry, This circuit provides output for the amplified net current signal and its polarity, and the integrated net current signal and its polarity.
IVo SYSTEM TESTING
A number of tests have been performed during the plasma source system development process to evaluate the performance of various assemblies, verify design techniques, and determine the system characteristics, These tests included electrical performance, endurance, startup cycling, contamination resistance, integration, shock, vibration, thermal cycling, functional evaluation of all circuitry, and overall system performance. The results of the system operational and functional testing are described below. These tests were conducted in a cylindrical, stainless-steel vacuum chamber, 2 m in liameter and 4 m long. The chamber was pumped by two O.8-m intake diameter and one 1.2-m intake diameter oil-diffusion pumps. A background pressure of approximately 3,OX 10- 5 Torr (4.0X 10-3 Pa) was maintained during plasma source operation. During system testing, the plasma source was electrically tied, through the 8.0S-!} common resistor, to a steel support rack which was tied to laboratory ground,
A. Operational characteristics
The plasma source system has ten operational states which are determined by the status of the hollow cathode and keeper, and are defined in Table VI. A typical plasma source system run exhibiting states 10, 1, 2, 5,4, 3, 7, 8, and 6 (in order of appearance) through the output telemetry signals, is shown in Fig, 8. The plasma source startup procedure causes the hollow cathode to go from an OFF condition, state 10, to a high-voltage, low-current glow discharge, state 1, and then to the ON condition, state 2, which is a lowvoltage, high-current arc discharge. The startup time for this sequence was about 2 s. The small cusp at the beginning of the start sequence in the cathode voltage telemetry output resulted from the activation of the RUN supply. When the START supply was activated, the cathode voltage increased to about 340 V [telemetry output (TIM) 3.3 V] and the cathode current increased to 1 A (T 1M 2.5 V) to support a glow discharge. The cathode potential dropped to about 35
Spacecraft charge control 1059
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TABLE VI. Plasma source operational states.
System System Cathode Keeper powcr state status mode" (W)
I START OFF 476 2 ON OFF 112 3 ON (e-) HC 196 4 ON (e ,.,) MC 154 5 ON (e-) I.e 126 6 ON u+) 118 7b OFF (e )HC 115 8" OFF (e- )MC 48 9" OFF (e-) LC 2
10 OFF OFF 0
• (e .. · ) HC~High-current electron collection mode; (e - ) MC-Mediurn-current electron collection mode; (e - ) LC-Low-current electron collection mode; (i + ) ion collection mode.
I> Cathode self-heating modes; STATE 8 is unstable, STATE 9 not mpported.
V (TIM 2.5 V) and the cathode current increased to 2 A ( TIM 4.9 V), turning off the START supply and placing the hollow cathode in the ON condition once the cathode emitter tube was hot enough to support thermionic emission of electrons. The keeper voltage increased during the start sequence since the KEEPER supply is electrically floating, with the cathode emitter tube acting as a common point for it and the START and RUN supplies. The currents corresponding to the low current, state 5, medium current, state 4, and high current, state 3, keeper modes are approximately 0.3, 1.2, and 2.0 A of collected electron current, respectively. In the self-heating modes, states 7 and 8, the RUN supply is
>-, 5, or ~ "'-~ ~ 2.5t' ",0
~~ >- >-<) a
~ s.ar <r_
g ~ 2.5[' ",,0
~~ t:::>-~ a
~--~~,-'~~~-e8~6--~lG~a--~1~~O-TIME IS)
FIG. 8. Plasma source operational states (T /,t[ outputs).
1060 Rev. Sci. !nstrum., Vol. 58, No.6, June 1987
off and the KEEPER supply is on so that ion bombardment through the cathode orifice keeps the cathode emitter tube hot enough to thermionically emit electrons. There is no selfheating mode corresponding to the low-current keeper (state 9) mode because the current is too low to sustain an adequate flow of backstreaming ions to self-heat the cathode. Ion currents collected by the keeper in the reverse-bias mode (keeper negative, state 6) are on the order of, at most, tens of microamperes, and as a result are not detectable on the keeper current telemetry output signal.
So Functional characteristics
Tests were conducted to evaluate how well the plasma source could discharge biased surfaces. An aluminum plate, 0.3 m on a side, was placed 0.12 m downstream of the plasma source opening. The plate could be biased two ways: by charging a 500-pF capacitor attached to the plate or by applying a bias directly to the plate with a variable voltage power supply. The potential and current collected by the plate were monitored to evaluate the plasma source system performance. An effective sounding rocket capacitance was estimated using the equivalent sphere method whereby the rocket is assumed to be a sphere with a surface area equal to that of the actual sounding rocket. The capacitance is then calculated using29
(1)
where R is the radius of the equivalent sphere (3 m), A. D is the Debye length, and Co is the permittivity constant. The expected sounding rocket capacitance was 20 uF. A highvoltage power supply was used to charge the capacitor through a 22-MH charging resistor. The plate biases could be varied from 0 to ± 2500 V in steps of ± 100 V. A typical discharge pulse for a capacitively biased plate is shown in Fig. 9, which has a pulse duration of approximately 0.035 s with a peak discharge current of approximately 0.017 A. The plasma source was turned on in system state 3, the high current keeper mode. The initial plate potential was 590 V. The oscillations following the discharge peak resulted from EMI feedback into the sensing circuitry.
24: -~- ',- -,-:- 1-'" -r --,- ., -'l 20~ j
i M~ i j ~ 12~ \ I g aL \}' I
~ :l VN\Y1 (v~VWAl V1 -81 j
'12L.--1. ___ 1-,-L._,-L-_,_' _,~' _.l_ .... ~ __ J....-.i o 0.02 0.04 0.06 0.08 O. lC D. 12 O. 14 O. 16 O. 18 0.20
TIME. seconds
FIG. 9. Typical nct current discharge pulse for a capacitively biased plate.
Spacecraft charge control 1060
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The current which could be extracted from the plasma source was evaluated by biasing the plate directly with a variable voltage power supply. Then, the plasma source was activated in system state 3 to examine the initial discharge pulse and steady-state output. A typical plate discharge pulse is shown in Fig. lO(a) for a directly biased plate whose potential is shown in Fig. lO(b). The pulse width is approximately 0.01 s with a peak height of 0.023 A. The plate potential was held to about 5 V, while drawing a constant 0.008 A. The oscillations following the discharge peak resulted from EMI feedback into the sensing circuitry. When the biasing power supply output was increased, the plate potential remained constant and the collected current increased,
The plasma source, net current monitoring system was also evaluated using a surface tied to a variable voltage power supply. Typical telemetry outputs are shown for the plasma source net current integrator and amplified net current monitor in the curves of Fig. 11 for current collection by a positively biased plate. These curves were generated while operating the plasma source in the medium current keeper mode (system state 4). These curves show how, as the biased plate collected electron current increases, the slope of the
26
'" 22 ::: ., Cl.
E '" 18 E ,....-
14 z UJ a:: §§ <..) 10 I-L.cl :z L.U
6 <..)
"" => 0 V1
50 (a) TIME, milliseconds
9Qa
800
700 .-
'" 600 "0 > wi 500 (::; ..:
~ !:J
400 0 > w >- 300 <Z: ...J a..
200
100
0 0
(b) TIME, milliseconds
FIG. to. (a) Typical plate discharge pulse and (b) the plate potelltial durillg discharge pulse.
1061 Rev, SCi.lnstrum., Vol. 58, No.6, June 1987
PLATE BIAS,';'-- ···-l.QOTIlJ2IITIOC ur COllEqED CURRENT,:nA~oi...l_u:r~).LiI._·i9. 7 21.~
5.0
INTEGRATED ~ NET CURRo 1;5 2 5 SIGML o· TIM OUT. §§
AMPLIFIED 2: NET CURRo SIGNAL TIM OUT. ~
I I I I I , I I I I
I I I : : : I SATURATED
TIME, arbitrary units
I I I
FIG. 11. Current collection by a positively biased plate during plasma source operation (T / M outputs).
integrator gets steeper. The plate bias and actual values of the collected electron current are given for each step of the curve for the amplified net current signal output. The integrator saturates when about 0.44 C have been emitted from the plasma source. A set of curves showing the electron current collected by the biased plate as a function of plate poten-
140
120
V> lOO~ J
g -J-
:::;
~~ !-Z w I-
~ ~ w
! I-
:5 CL
oor o LOW CURRENT KEEPER MODE
~ ,
o MEDIUM CURRENT KEEPER MODE o HIGH CURRENT KEEPER MODE A SELF HEATING, HIGH CURRENT
KEEPER MODE '" SELF HEATING, MEDIUM CURRENT
KEEPER MODE
I 20
oL-- I I I 0 10 15 20 25
COLLECTED CURRENT, milliamperes
Fre;. 12. Collected current as a function of plate potential for various plasma source operating modes.
Spacecraft charge control 1061
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ION CliRRENT COLLECTED BY NEGATIVELY B IASFD PLATf
INTEGRATED NET CliRRENT SIGNAL TIM OUTPUT RFSPONSE
AMPUFIED NET CURRFNT SIGNAL TIM OUTPUl RFSPONSE
2 s 00
'" >=
:::t 2J ~--- ---------._---------
0.0 __ .. _________ . __________ ' ____ _
5T
ZT 0.0 "
----L. __ "-1-,--..L __ ~~~ D -100 ·2DO -300 ~400 -500
PLATf POTENTIAL, volt,
FIG. 13. Current collection by a negatively biased plate during plasma source operation.
tial for the five source states where the keeper was also set to collect electrons (states 3-5,7, 8) is shown in Fig. 12. Operation of the plasma source in the high-current keeper mode provides the greatest electron emission capability. The data for the low-current keeper mode end at 100 V since the discharge extinguished itself at this point repeatedly.
Typical net current monitor telemetry outputs are shown in Fig. 13 for current collection by a negatively biased plate (ion collection) located 0.1 m downstream of the keeper electrode with the plasma source in state 3. The integrator slope is slightly smaller than the integrator slope corresponding to the accumulation of current due only to electron leakage from the source indicating that electron flow is inhibited in this mode. The "hash" marks on the amplified output signal correspond to small, low-energy arcs between the biased plate and the plasma source. The average ion current collected by the biased plate is given in the top curve of Fig. 13 for several plate potentials and is in the hundreds of microamperes regime. At low plate biases, between 0 and - 40 V, the collected ion current was in the tens of mi-
croamperes range.
ACKNOWLEDGMENTS
The authors wish to thank Robert L. Toomath for assistance in the electronics layout and fabrication, Allison G. Owens for assistance in the mechanical layout and fabrication, and Col. W. T. McLyman for electronics design support. The research described in this paper was performed at the Jet Propulsion Laboratory, California Institute of Tech-
1062 Rev. Sci. Instrum., Vol. 58, No.6, June 1987
nology, under contract to the United States Air Force through the National Aeronautics and Space Administration.
'H. B. Garrett, Rev. Geophys. Space Phys. 19,577 (1981). 2R. C. Olsen, J. Spaccc,r. Rockets 18, 527 (1981). 'R. Goldstein, Proceedings of the Spacecraft Charging Technology Confer. ence, edited by C. P. Pike and R. R. Lovell, Paper 1-6, NASA TMX-73537 (Air Force Geophysics Lab., Hanscom, AFB, MA. 1977), pp. 121-130,
'c. K, Purvis, R. O. Bartlett, and S. E. Deforest, Proceedings afthe Spacecraft Charging Technology Conference, edited by C. P. Pike and R. R. Lovell. Paper 1-5, NASA TMX·73537 (Air Force Geophysics Lab., Hanscom AFB, MA, 1977), pp. 107-120.
5E. C. Whipple and R. C. Olsen, AIAA 12th Fluid and Plasma Dynamics Conference, AIAA Paper No. 79-J506 (1979).
OR. C. Olsen, J, Spacecr. Rockets 18, 462 (1981). 7H. A. Cohen and S. Lai, AIAA 20th Aerospace Sciences Meeting. AIAA Paper No, 82-0266 (1982).
"P. F. Mizcra, AIAA 18th Aerospace Sciences Meeting, AIAA Paper No. 80-0334 (1980).
"D. L Reasoner, W. Lennartsson, and C. R. Chappell, Spacecraft Charging by Magnetospheric Plasma-Progress in Astronautics and Aeronautics edited by A. Rosen (AIAA/MIT. Cambridge, MA, 1976), Paper SA-36A, 47, pp. 89-102.
;oS. E. Deforest, J, Geophys. Res. 77,651 (1972). !lA. Rosen, J. Spaeecr. Rockets 13,129 (1976). 12K. G. Balmain, M. Cuchanski, and P. C. Kremer, Proceedings of the
Spacecraft Charging Technology Conference, edited by C. P. Pike and R. 1'. Lovell, (NASA. 1977). Paper 1II-7, TM X-73537, pp, 516-526.
;:lJ. E. Nanevicz and R. C. Adamo, Space Systems and Their Interactions with Earth~' Space Environment-Progress in Astronautics and Aeronautics, edited by H. B. Garrett anel C. P. Pike (AIAA, New York. 1980), 71, pp.252.
"G. T. Inouye, Proceedings of the Spacecraft Charging Technology Conference. edited by C. R. Pike and R. R. Lovell, Paper No. V-lO, NASA TM X-73537, pp. 829-852 (1977).
15 A. Robbins and C. D. Short. Proceedings of the Spacecraft Charging Technology Conference, edited by C. P. Pike and R. R. Lovell, Paper No. V-II, NASA TM X-73537, pp. 853-863 (1977).
16R. R. Shaw, J. E. Nanevicz, and R. C. Adamo, Spacecraft Charging by Magnetospheric Plasmas, Progress in Astronautics and Aeronautics, edited by A. Rosen (AIAA/MIT, Cambridge, MA, 1976). Paper SA-41, 47, pp. 61-76.
PC. K. Purvis, H. B. Garrett, A. C. Whittlesey, and N. J. Stevens, NASA TP-2361, 1984,
!RH. A. Cohen, Air Force Geophysics Laboratory, 21 Junt' 1983 (private communication) ,
19V. K. Rawlin and E. V. Pawlik, J. Spacecr. Rockets 5,814 (1968). 20J. W. Ward and H. J. King, J. Spacecr. Rockets 5,1161 (1968). 21W. R. Kerslake, R, G. Goldman, and W, C. Neirberding, J. Spacecr.
Rockets 8, 213 (1971). 22R. C. Olsen and C. K. Purvis, J. Geophys. Res. 88, 5657 (1983). 23G. Aston, J. Vac. Sci. Techno!. AI, 258 (1983). 2"'G. Aston, Proceedings JSASS/ AIAA/DGLR 17th International Electric
Propulsion Conference. Paper IEPC 84-87, pp. 647-651 (1984), 25G, Aston and W. D. Deininger. NASA CR-174623, 1984. 261'. D. Masek and H. A. Cohen, J, Spacecr. Rockels 15. 27 (1978). 27T. Masek. Air Force Geophysics Laboratory, Final Report No. AFGL
TR-78-0l41, 1978. 2"J. M, Meek and J. D. Craggs, Electrical Breakdown of Gases (Oxford Uni
versity, Oxford, 1953). 29D. J. Fitzgerald. AIAA 11th Electric Propulsion Conference, AIAA IJaper
No. 75-404 (1975).
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