A starter circuit particularly suitable for a plasma of an ion engine for a spacecraft includes a power supply having an output inductor with a tap. A switch is coupled to the tap. The switch has a control input. A pulse control logic circuit is coupled to said control input, said pulse control logic circuit controlling said switch to an off state to generate a high voltage discharge

Cardwell, Jr., Gilbert I.

Phelps, Thomas K.

NASA Technical Reports Server

I11111 111ll Il1 Il11 III 11ll III III III III 1ll111ll1 
US006369521B3 
(12) United States Patent (io) Patent No.: US 6,369,521 B1 
Cardwell, Jr. et al. (45) Date of Patent: Apr. 9,2002 
STARTER CIRCUIT FOR AN ION ENGINE 
Inventors: Gilbert I. Cardwell, Jr., Palos Verdes 
Peninsula; Thomas K. Phelps, 
Torrance, both of CA (US) 
Assignee: Hughes Electronics Corporation, El 
Segundo, CA (US) 
Subject to any disclaimer, the term of this 
patent is extended or adjusted under 35 
U.S.C. 154(b) by 0 days. 
Notice: 
Appl. No.: 09/935,189 
Filed: Aug. 22, 2001 
Related U.S. Application Data 
Division of application No. 091352,011, filed on Jul. 12, 
1999, now Pat. No. 6,304,040. 
Int. C1.7 H05B 37/00 
60139.03 
Field of Search ................................. 3151289, 290, 
3151291, 224, 307, 111.21, 111.51, 362; 
3611247, 268; 60139.03 
U.S. c1. ....................... 89; 3151291; 
References Cited 
U.S. PATENT DOCUMENTS 
3,491,250 A * 111970 Shoh ........................... 363171 
4,533,836 A * 811985 Carpenter et al. ............ 363117 
4,695,933 A * 911987 Nguyen et al. 
4,825,646 A * 511989 Challon 
5,269,131 A * 1211993 Brophy 
5,352,861 A * 1011994 Steigenvald et al. ... 2191121.54 
5,369,953 A * 1211994 Brophy ............... 315 111.01 X 
5,434,770 A * 711995 Driefuerst et al. 363171 
5,451,962 A * 911995 Steigenvald ................. 363117 
5,561,350 A * 1011996 Frus et al. .............. 3151209 R 
5,576,940 A * 1111996 Steigenvald et al. .......... 363117 
4,733,137 A * 311988 Dunham ........ 
5,610,452 A * 311997 Shimer et al. 
5,657,217 A * 811997 Watanabe et a 
5,666,278 A * 911997 Ng et al. ...................... 363171 
5,825,139 A * 1011998 Nuckolls et al. ............ 3151307 
5,852,555 A * 1211998 Martin ........................ 363171 
5,862,041 A * 111999 Martin ........................ 363171 
5,862,042 A * 111999 Jiang ........................... 363117 
5,875,103 A * 211999 Bhagwat et al. .............. 363171 
5,923,549 A * 711999 Kobayashi et al. 
5,930,122 A * 711999 Moriguchi et al. 
5,949,668 A * 911999 Schweighofer .............. 363171 
5,991,179 A * 1111999 Schweighofer 
6,154,383 A * 1112000 Cardwell, Jr. .. 
6,181,585 B1 * 112001 Cardwell, Jr. et al. ........ 363171 
6,295,804 B1 * 1012001 Burton et a1 601203.1 
OTHER PUBLICATIONS 
Thomas A. Bond et al., “NSTAR Ion Engine Power Proces- 
sor Unit Performance: Ground Test and Flight Experience”, 
S A E  Paper 99APSC47, Apr. 1999.* 
Thomas A. Bond et al., “The NSTAR Ion Propulsion Sub- 
system for DSl”, AIAA Joint Propulsion Conference, AIAA 
Paper 99-2972, Jun. 23, 1999.* 
John A. Hamley et al., “The Design and Performance 
Characteristics of the NSTAR PPU and DCIU’, AIM 
ASMEISAEIASEE Joint Propulsion conference & Exhibit, 
34th, Cleveland, OH, Jul. 13-15, 1998, AIAA Paper 
* cited by examiner 
Primary Examiner4a issa  Philogene 
(74) Attorney, Agent, or Firm-Terje Gudmestad 
(57) ABSTRACT 
A starter circuit particularly suitable for a plasma of an ion 
engine for a spacecraft includes a power supply having an 
output inductor with a tap. A switch is coupled to the tap. 
The switch has a control input. Apulse control logic circuit 
is coupled to said control input, said pulse control logic 
circuit controlling said switch to an off state to generate a 
high voltage discharge. 
98-3938. * 
6 Claims, 3 Drawing Sheets 
https://ntrs.nasa.gov/search.jsp?R=20080005014 2019-08-30T02:49:34+00:00Z
U S .  Patent Apr. 9,2002 Sheet 1 of 3 US 6,369,521 B1 
U S .  Patent Apr. 9,2002 Sheet 2 of 3 US 6,369,521 B1 
U S .  Patent Apr. 9,2002 Sheet 3 of 3 US 6,369,521 B1 
I- 
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1 
I 
I 
I 
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I 
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1 
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FIG. 4 
US 6,369,521 B3 
1 
STARTER CIRCUIT FOR AN ION ENGINE 
This appl. is a divisional appl. of appl. Ser. No. 091352, 
011, filed Jul. 12, 1999, now U.S. Pat. No. 6,304,040. 
“This invention disclosure herein was made in tie per- 
formance of work under NASA Contract Number NAS3- 
27560 and is subject to the provisions of Section 305 of the 
National Aeronautics and Space Act of 1958 (72 Stat. 435; 
42 U. S. C. 2457).” 
TECHNICAL FIELD 
The present invention relates generally to an ion propul- 
sion system, and in particular to a method and apparatus for 
igniting a plasma in an ion propulsion system. 
BACKGROUND OF THE INVENTION 
For over thirty years, ion engines have been proposed for 
propulsion of vehicles in space. Outside of space propulsion, 
ion generation may also be applied to various types of 
materials processing systems involving ion sources, such as 
for ion beam etching or micromachining. Ion engines use 
movement of ions to provide thrust. 
Generally, an ion engine has an ion accelerator system 
that uses an anode, a cathode, a screen grid and an accel- 
erator grid coupled within a thruster housing. Generally, an 
ion engine works by generating an inert gas plasma within 
the thruster housing. Xenon is an example of a suitable gas. 
A charge within the plasma between the anode and cathode 
forms ions. The inert gas ions leave the thruster through the 
charged screen and accelerator. The net force from the ions 
leaving the thruster housing generates a thrust. A neutralizer 
is located outside the thruster housing and generates elec- 
trons. The electrons are attracted to the ions so the ions do 
not re-enter the thruster housing as they otherwise would in 
space. 
To initiate a breakdown of the xenon to form ions in the 
thruster or electrons at the neutralizer a high voltage break- 
down must occur between the anode and cathode. 
Previously, it was thought that separate power supplies must 
be used to initiate the high voltage breakdown at both the 
thruster and the neutralizer. 
In spacecraft design, it is desirable to eliminate parts and 
complexity when possible. More parts increases weight of 
the spacecraft. More parts and complexity inherently 
reduces reliability. 
It is therefore an object of the invention to provide a 
power supply system that operates reliably and reduces 
overall weight and complexity. 
SUMMARY OF THE INVENTION 
It is therefore one object of the invention to provide a 
starter circuit that operates reliably and reduces overall 
weight of the spacecraft. 
In one aspect of the invention, a starter circuit includes a 
power supply having an output inductor with a tap. A switch 
is coupled to the tap. The switch has a control input. Apulse 
control logic circuit is coupled to said control input, said 
pulse control logic circuit controlling said switch to an off 
state to generate a high voltage discharge. 
In a further aspect of the invention, a method of starting 
plasma includes the steps of  
emitting a gas; 
charging an inductor having a tap and an output; 
coupling a starter circuit to said tap; 
2 
controlling the starter circuit to initiate a high voltage 
producing a current through the gas; 
establishing a plasma; and 
igniting the plasma. 
Another advantage of the invention is that the because a 
pulse input is used rather than a continuous source a high 
voltage rectifier and regulation control circuit are not 
required. 
One advantage of the invention is a separate power supply 
for the starter circuit has been eliminated from the space- 
craft. This reduces weight and complexity. 
Other features and advantages of the invention are readily 
1s apparent from the following detailed description of carrying 
out the invention when taken in connection with the accom- 
panying drawings. 
discharge; 
BRIEF DESCRIPTION OF THE DRAWINGS 
20 FIG. 1 is a perspective view of a spacecraft having a 
FIG. 2 is a cross sectional view of an ion thruster having 
FIG. 3 is a block diagram of a power supply system 
FIG. 4 is a block diagram of a starter circuit. 
power supply circuit according to the present invention. 
a power supply according to the present invention. 
2s according to the present invention. 
DESCRIPTION OF THE PREFERRED 
EMBODIMENTS 
30 In the following description, identical reference numerals 
are used to identify identical components in the various 
figures. The present invention is particularly suitable for use 
in a spacecraft. The power supply circuit of the present 
invention is also useful in other applications that have a wide 
3s dynamic range of system operability including a varying 
load or input. The present invention is also applicable to 
other systems that include ion sources such as materials 
processing equipment like ion beam etching or microma- 
chining. 
Referring now to FIG. 1, a spacecraft 10 has a thruster 
housing 12 that houses an ion thruster 14. Spacecraft 10 
further includes solar panels 15 as a source of electrical 
power. In the present invention, spacecraft 10 is powered by 
xenon ions which are generated in ion thruster 14. Space- 
craft 10 includes a xenon feed subsystem 16 supplying 
xenon to thruster 14. A digital interface and control unit 
(DCIU) 18 is also coupled to the thruster housing 12. 
Aneutralizer 20 is also coupled to thruster housing 12 and 
so xenon feed subsystem 16. As will be further described 
below, neutralizer 20 generates electrons to neutralize the 
positive ions emitted by thruster 14. 
Thruster 14 generally includes an anode 24 and a cathode 
26. Neutralizer 20 also includes an anode 28 and a cathode 
5s 30. Cathodes 26, 30 each have a respective heater 32, 34. 
Thruster 14 and neutralizer 20 also include a respective 
xenon source 36,38 that are part of xenon feed subsystem 
16. A keeper 40, 42 for concentrating the stream of xenon 
(ions or electrons) may also be provided near respective 
Thruster 14 further includes a screen grid 44 and an 
accelerator grid 46. Both screen grid 44 and accelerator grid 
46 are formed of an electrically conductive mesh material. 
Aplasma screen 48 may be used to enclose thruster 14 on 
65 sides other than where screen 44 and accelerator 46 are 
positioned. Plasma screen 48 is used to capture and prevent 
spalling of ion sputtered grid material. 
40 
4s 
60 cathodes 26, 30. 
US 6,369,521 B1 
3 4 
Apower supply circuit 50 is incorporated into spacecraft starter circuit 78 except that the feedback current threshold 
circuitry. Power supply circuitry 50 is coupled to anodes 24, is adjusted downward as will be further described below. 
28, cathodes 26, 30, heaters 32, 34, screen grid 44 and Sufficient power to generate a high voltage pulse to 
accelerator 46. initiate an arc is obtained from a power supply that is 
At a high level of operation, xenon sources 36,38 are used 5 currently used in the present invention. By using a power 
to generate a plasma of xenon adjacent to cathodes 26, 30, supply already available new components for providing 
respectively. Heaters 32, 34 are used to heat the xenon power to starter circuit 76 are not required. Discharge power 
plasma upon start up. An arc starter circuit shown in FIG. 3 supply 64 is suitable because the circuitry includes a 
is used to ignite the xenon plasma. Thruster 14  uses the smoothing inductor 80 as part of the output of circuitry. 
xenon ions for thrust. As the xenon ions pass through screen Current is established between positive output 64P and 
44 and accelerator grid 46, thrust is created. Neutralizer 20 negative output 64N of discharge power supply 64. Dis- 
generates a xenon plasma as well. However, the goal of charge power supply 64 also has a primary winding 82 and 
neutralizer 20 is to generate electrons that are used to a secondary winding 84. Secondary winding is coupled to 
electrically balance the xenon positive ions in space to rectifier diodes D1 and D2. Secondary winding 84 may also 
prevent the xenon ions from being attracted back to the have a tap 86 extending therefrom. Tap 86 is coupled to the 
spacecraft. thruster cathode and inductor 80 through diode D3 and 
Referring now to FIG. 3, power supply circuit 50 is through capacitor c1 and diode D4. 
illustrated in greater detail, A central spacecraft bus 52 Starter circuit 76 includes control logic 88 that controls 
couples the base components of power supply circuit 50 the initiation of a high voltage. Control logic may comprise 
together. Spacecraft bus 52 includes a bus input 54 and a bus a Plurality of logic circuits Or may be microProcessor-based. 
return 56. 20 Control logic is coupled to a transformer 90 having a 
winding is coupled to a resistor R1 and a voltage clamp 96 reduce electrical noise. Input filters 58 may take the form of that is comprised of a pair of zener diodes 98 and A 
capacitors or other circuit components as would be evident second resistor R2 is coupled in parallel with voltage clamp 
to those skilled in the art. 
The control ofthe Power supply circuit50iscontrolledbY Control logic 88 controls a switch 102. Switch 102 has a 
DCIU 18. DCIU 18 is also coupled to bus 52. AhousekeeP- control input 102C that is coupled to control logic 88 
ing supply 60 may also be incorporated into power supply through transformer 90, 
circuit 50. Housekeeping Supply 60 may be used for other Switch 102 is coupled between a tap 104 on inductor 80 
functions besides a centralized system and may not be 30 through an isolating diode D5. Inductor 80 has a discharge 
coupled to bus 52. output 106. Current at discharge output is monitored through 
Power supply circuit 50 includes a plurality of application a sensor 108. Sensor 108 is coupled to control logic 88 
specific power supplies. The application specific power through a feedback input 110. Control logic 88 may also 
supplies are sized in terms of current and voltage based on have a discharge onioff input 112. Discharge onioff input 112 
the specific components to which they are connected, The 35 may be derived from other controllers within the spacecraft 
specific power supplies may include a discharge heater such as DCIU. 
supply 62, discharge supply current source 64, screen supply In operation, the starter circuit 96 generally operates as 
voltage source 66, an accelerator supply voltage source 68, fOllOWS. When switch 102 is turned on, Current increases in 
a neutralizer supply current source 70, and a neutralizer inductor 80 to store energy therein. When switch 102 is 
heater supply 72. Discharge heater supply 62 is coupled to 4o turned Off 
heater 32 and is disposed within thruster 14, Discharge 
Input filters 58 may be coupled to spacecraft bus 52 to primary winding 92 and a secondary winding 94. Secondary 
2s 96. 
a high spike is generated 
which appears at discharge Output lo6. Dis- inductor 
supply current source 64 has a positive output 64P coupled 
to anode 24. Discharge supply current source 64 also has a 
negative output coupled to cathode 26. Negative output may 
also be coupled to screen grid 44. Screen supply voltage 45 
charge Output be coup1ed to the 
thruster anode 24 described above where the high voltage 
will generate ions which conduct discharge current to form 
the plasma. 
To determine whether a high voltage discharge is to be 
may2 for 
Source 66 has a positive Output 66p that may 
to anode 24. 
be coup1ed 
supply-voltage Source 68 has a applied to output 106, control logic 88 monitors current at current Sensor 108, If the plasma is maintaining ion genera- 
negative terminal coupled to accelerator 46. Neutralizer tion the current flowing between the cathode and anode is 
supply current Source 70 has a positive Output 70p coup1ed greater than 1.5 amps. In the present example, if the current 
to neutralizer anode 28. Neutralizer supply current source so monitor indicates less than 1.5 amps which is a level 
has a negative output 70N coupled to neutralizer cathode 30. corresponding to no ion generation, the starter operation is 
A f i h  capacitor 79 and a voltage clamp 77 may be coupled initiated as described in the following: A pulse is released 
to negative output 77 of neutralizer supply 70. Neutralizer through transformer 90 to drive the switch 102. Inductor 80 
heater supply 72 is coupled to heater 34. Neutralizer heater acts as an auto-transformer that boosts the voltage to about 
supply 72 has a positive output 7OP and a negative output 55 200 volts for about 20 microseconds. When the arc is 
70N. established between cathode 26 and anode 24, current from 
Anegative output 66N of screen supply voltage source 66, discharge supply 64 maintains the plasma. When the dis- 
a positive output 68P of accelerator supply voltage source charge current exceeds 1.5 amps, control logic 88 inhibits 
68, a negative output 70N of neutralizer supply current further pulses and thus enters a standby mode to conserve 
source 70 and negative output 72N of neutralizer heater 60 energy. If the discharge current drops below 1.5 amps 
supply 72 may all be coupled together at the same electrical (indicating that the arc has been extinguished), control logic 
potential. Discharge arc starter circuit 76 and a neutralizer 88 pulses switch 102 to generate high voltage pulses at 
arc starter circuit 78 may be coupled to cathodes 26, 30 output 106. 
respectively. As described above, arc starter circuits 76, 78 The same circuitry as starter circuit 76 may be used for 
are used to ignite the ion plasma. 65 neutralizer starter circuit 78. However, the threshold to 
Referring now to FIG. 4, starter circuit 76 is illustrated in initiate a high discharge output by control logic 88 need only 
further detail. Starter circuit 76 is identical to neutralizer be 0.5 amps. 
US 6,369,521 B1 
5 6 
While the best mode for carrying out the present event has 
been described in detail, those familiar with the art to which 
this invention relates will recognize various alternative 
designs and embodiments for practicing the invention as 
defined by the following claims. 
controlling said switch with a control signal; 
opening the switch; and 
generating a high voltage discharge in response to the 
opening of the switch. 
3. A method as recited in claim 1 wherein the step of 
high voltage discharge between an anode and cathode of a 
thruster. 
4. A method as recited in claim 1 wherein the step of 
generating a high voltage discharge comprises generating a 
high voltage discharge between a neutralizer anode and 
neutralizer cathode of a neutralizer. 
5 .  A method as recited in claim 1 further comprising the 
6. A method as recited in claim 1 wherein the step of 
monitoring comprises the steps of sensing a current at said 
output, and generating a control signal in response to said 
current, 
What is claimed is: 
emitting a gas; 
charging an inductor having a tap and an output; 
coupling a starter circuit to said tap; 
controlling the starter circuit to initiate a high voltage 
producing a current through the gas; 
establishing the plasma; and 
igniting the plasma. 
2. A method as recited in claim 1 wherein the step of 
coupling a starter circuit to said tap comprises the steps of  
coupling a switch to said tap to store energy in the 
1, Amethod for starting a plasma comprising the steps of: generating a high discharge comprises generating a 
discharge; 
1s step of monitoring a continuance of the plasma. 
inductor; * * * * *  


Starter circuit for an ion engine

I 11111 111ll Il1 Il11 III 111ll I1 11111 Il1 11111 ll1111ll1 
US006304040Bl 
(12) United States Patent (io) Patent No.: US 6,304,040 B1 
Cardwell, Jr. et al. (45) Date of Patent: Oct. 16,2001 
(54) STARTER CIRCUIT FOR AN ION ENGINE 5,666,278 911997 Ng, et al. ............................... 363171 
5,852,555 1211998 Martin .................................... 363171 
(75) Inventors: Gilbert I. Cardwell, Jr., Palos Verdes 5,862,041 111999 Martin .................................... 363171 
5,862,042 111999 Jiang ...................................... 363117 
5,875,103 211999 Bhagwat et al. ....................... 363171 
5.923.549 711999 Kobavashi et al. 363117 
Peninsula; Thomas K. Phelps, 
Torrance, both of CA (US) .................... 
(73) Assignee: Hughes Electronics Corporation, El 
Segundo, CA (US) 
Subject to any disclaimer, the term of this 
patent is extended or adjusted under 35 
U.S.C. 154(b) by 0 days. 
( * ) Notice: 
(21) Appl. No.: 09/352,011 
(22) Filed: Jul. 12, 1999 
(51) Int. Cl? ..................................................... H05B 37/00 
(52) U.S. C1. .......................... 315/290; 3151289; 3151291; 
60139.03 
(58) Field of Search ..................................... 3151291, 224, 
3151307, 111.21, 111.51, 289, 290, DIG. 5; 
60139.03; 3611247, 263; 363141 
(56) References Cited 
U.S. PATENT DOCUMENTS 
3,491,250 111970 Shoh ...................................... 363171 
4,533,836 811985 Carpenter et al. ..................... 363117 
4,695,933 911987 Nguyen et al. ........................ 363171 
4,733,137 * 311988 Dunham ............................... 3151291 
4,825,646 511989 Challoner, et al. 601202 
5,269,131 1211993 Brophy ........................ 3151111.01 X 
5,369,953 1211994 Brophy ........................ 3151111.01 X 
5,434,770 711995 Driefuerst et a1 
5,561,350 * 1011996 Frus et al. ...... 
5,576,940 1111996 Steigenvald et al. 
5,610,452 311997 Shimer et al. ......................... 307189 
5,657,217 811997 Watanabe, et al. .................... 363171 
5,451,962 911995 Steigenvald ... 
5:930:122 711999 Moriguchi et al. .................... 363117 
5,949,668 911999 Schweighofer ........................ 363171 
5,991,179 1111999 Schweighofer ........................ 363171 
6,154,383 * 1112000 Cardwell, Jr. .......................... 363171 
6,181,585 * 112001 Cardwell, Jr. et al. ................ 363171 
OTHER PUBLICATIONS 
Thomas A. Bond et al., “NSTAR Ion Engine Power Proces- 
sor Unit Performance: Ground Test and Flight Experience ”, 
S A E  Paper 99APSC47, Apr. 1999. 
Thomas A. Bond et al., “The NSTAR Ion Propulsion Sub- 
system for DS1 ”, AIAA Joint Propulsion Conference, AIAA 
Paper 99-2972, Jun. 23, 1999. 
John A. Hamley et al., “The Design and Performance 
Characteristics of the NSTAR PPU and DCIU ”, AIM 
ASMEISAEIASEE Joint Propulsion Conference & Exhibit, 
34 th , Cleveland, OH, Jul. 13-15, 1998, AIAA Paper 
* cited by examiner 
Primary Examiner4aissa  Philogene 
(74) Attorney, Agent, or Firm-Terje Gudmestad 
(57) ABSTRACT 
A starter circuit particularly suitable for a plasma of an ion 
engine for a spacecraft includes a power supply having an 
output inductor with a tap. A switch is coupled to the tap. 
The switch has a control input. Apulse control logic circuit 
is coupled to said control input, said pulse control logic 
circuit controlling said switch to an off state to generate a 
high voltage discharge. 
98-3938. 
7 Claims, 3 Drawing Sheets 
https://ntrs.nasa.gov/search.jsp?R=20080004648 2019-08-30T02:43:12+00:00Z
U S .  Patent Oct. 16,2001 Sheet 1 of 3 
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US 6,304,040 B1 
U S .  Patent Oct. 16,2001 Sheet 2 of 3 US 6,304,040 B1 
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U S .  Patent Oct. 16,2001 Sheet 3 of 3 US 6,304,040 B1 
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Discharge 
Supply Power 
Stage I 
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FIG. 4 
US 6,304,040 B1 
1 2 
STARTER CIRCUIT FOR AN ION ENGINE producing a current through the gas; 
establishing a plasma; and 
igniting the plasma, 
Another advantage of the invention is that because a pulse 
input is used rather than a continuous a high voltage 
rectifier and regulation control circuit are not required. 
One advantage of the invention is a separate power supply 
for the starter circuit has been eliminated from the space- 
craft. This reduces weight and complexity. 
Other features and advantages of the invention are readily 
apparent from the following detailed description of carrying 
out the invention when taken in connection with the accom- 
panying drawings. 
“The invention described herein was made in the per- 
fOrmance of Work under NASA Contract Number NAs3- 
27560 and is subject to the provisions of Section 305 of the 5 
National ~eronaut ics  and Space Act of 1958 (72 Stat. 435; 
42U.S. C 2457).” 
TECHNICAL FIELD 
ne present invention relates generally to an ion propul- lo 
sion system, and in particular to a method and apparatus for 
igniting a plasma in an ion propulsion system. 
BRIEF DESCRIPTION OF THE DRAWINGS BACKGROUND OF THE INVENTION 15 
For Over thirty years, ion engines have been proposed for FIG, 1 is a perspective view of a spacecraft having a 
propulsion of vehicles in space. Outside of space Propulsion, 
ion generation may also be applied to various types of 
materials processing systems involving ion sources, such as 
for ion beam etching or micromachining. Ion engines use 20 
movement of ions to provide thrust. 
Generally, an ion engine has an ion accelerator system 
that uses an anode, a cathode, a screen grid and an accel- 
erator grid coupled within a thruster housing. Generally, an DESCRIPTION OF THE PREFERRED 
ion engine works by generating an inert gas plasma within 25 EMBODIMENTS 
A charge within the plasma between the anode and cathode are used to identify identical components in the various 
forms ions. The inert gas ions leave the thruster through the figures, ne present invention is particularly suitable for use 
charged screen and accelerator. The net force from the ions in a spacecraft, The power supply circuit of the present 
leaving the thruster housing generates a thrust. A neutralizer 30 invention is also useful in other applications that have a wide 
is located outside the thruster housing and generates elec- dynamic range of system operability including a varying 
trons. The electrons are attracted to the ions so the ions do load or input, The present invention is also applicable to 
not re-enter the thruster housing as they otherwise would in other systems that include ion such as materials 
space. processing equipment like ion beam etching or microma- 
thruster or electrons at the neutralizer a high voltage break- Referring now to FIG, 1, a spacecraft 10 has a thruster 
down must Occur between the anode and cathode. housing 12 that houses an ion thruster 14. Spacecraft 10 
Previously, it was thought that separate Power supplies must further includes solar panels 15 as a source of electrical 
be used to initiate the high voltage breakdown at both the power. In the present invention, spacecraft 10 is powered by 
thruster and the neutralizer. xenon ions which are generated in ion thruster 14. Space- 
In spacecraft design, it is desirable to eliminate parts and craft 10 includes a xenon feed subsystem 16 supplying 
complexity when possible. More parts increases weight of xenon to thruster 14. A digital interface and control unit 
the spacecraft. More parts and complexity inherently (DCIU)18 is also coupled to the thruster housing 12. 
reduces reliability. Aneutralizer 20 is also coupled to thruster housing 12 and 
It is therefore an object of the invention to provide a xenon feed subsystem 16. AS will be further described 
power supply system that operates reliably and reduces below, neutralizer 20 generates electrons to neutralize the 
overall weight and complexity. positive ions emitted by thruster 14. 
Thruster 14 generally includes an anode 24 and a cathode 
26. Neutralizer 20 also includes an anode 28 and a cathode 
power supply circuit according to the present invention. 
a power supply according to the present invention. 
according to the present invention. 
FIG. 2 is a cross sectional view of an ion thruster having 
FIG. 3 is a block diagram of a power supply system 
FIG. 4 is a block diagram of a starter circuit. 
the thruster housing. Xenon is an Of a gas. In the following description, identical reference numerals 
To initiate a breakdown of the xenon to form ions in the 35 chining. 
40 
45 
SUMMARY OF THE INVENTION 
It is therefore one object of the invention to provide a 30. Cathodes 26, 30 each have a respective heater 32, 34. 
starter circuit that operates reliably and reduces overall Thruster 14 and neutralizer 20 also include a respective 
weight of the spacecraft. 
In one aspect of the invention, a starter circuit includes a 
power supply having an output inductor with a tap. A switch 
is coupled to the tap. The switch has a control input. Apulse 
control logic circuit is coupled to said control input, said 
pulse control logic circuit controlling said switch to an 
state to generate a high voltage discharge. 
plasma includes the steps of  
xenon source 36,38 that are part of xenon feed subsystem 
16. A keeper 40, 42 for concentrating the stream of xenon 
55 (ions or electrons) may also be provided near respective 
cathodes 26, 30. 
a Screen grid 44 and an 
accelerator grid 46. Both screen grid 44 and accelerator grid 
46 are formed of an electrically conductive mesh material. 
Aplasma screen 48 may be used to enclose thruster 14 on 
sides other than where screen 44 and accelerator 46 are 
positioned. Plasma screen 48 is used to capture and prevent 
spalling of ion sputtered grid material. 
A power supply circuit 50 is incorporated into spacecraft 
65 circuitry. Power supply circuitry 50 is coupled to anodes 24, 
28, cathodes 26, 30, heaters 32, 34, screen grid 44 and 
Thruster l4 further 
In a further aspect of the invention, a method of starting 6o 
emitting a gas; 
charging an inductor having a tap and an output; 
coupling a starter circuit to said tap; 
controlling the starter circuit to initiate a high voltage 
discharge; accelerator 46. 
US 6,304,040 B3 
3 4 
At a high level of operation, xenon sources 36,38 are used currently used in the present invention. By using a power 
to generate a plasma of xenon adjacent to cathodes 26, 30, supply already available new components for providing 
respectively. Heaters 32, 34 are used to heat the xenon power to starter circuit 76 are not required. Discharge power 
plasma upon start up. An arc starter circuit shown in FIG. 3 supply 64 is suitable because the circuitry includes a 
is used to ignite the xenon plasma. Thruster 14  uses the 5 smoothing inductor 80 as part of the output of circuitry. 
xenon ions for thrust. As the xenon ions pass through screen Current is established between positive output 64P and 
44 and accelerator grid 46, thrust is created. Neutralizer 20 negative output 64N of discharge power supply 64. Dis- 
generates a xenon plasma as well. However, the goal of charge power supply 64 also has a primary winding 82 and 
neutralizer 20 is to generate electrons that are used to a secondary winding 84. Secondary winding is coupled to 
electrically balance the xenon positive ions in space to rectifier diodes D1 and D2. Secondary winding 84 may also 
prevent the xenon ions from being attracted back to the have a tap 86 extending therefrom. Tap 86 is coupled to the 
spacecraft. thruster cathode and inductor 80 through diode D3 and 
Referring now to FIG. 3, power supply circuit 50 is through capacitor C1 and diode D4. 
illustrated in greater detail. A central spacecraft bus 52 Starter circuit 76 includes control logic 88 that controls 
couples the base components of power supply circuit 50 the initiation of a high voltage. Control logic may comprise 
together. Spacecraft bus 52 includes a bus input 54 and a bus a plurality of logic circuits or may be microprocessor-based. 
return 56. Control logic is coupled to a transformer 90 having a 
Input filters 58 may be coupled to spacecraft bus 52 to primary winding 92 and a secondary winding 94. Secondary 
reduce electrical noise. Input filters 58 may take the form of winding is coupled to a resistor R1 and a voltage clamp 96 
capacitors or other circuit components as would be evident that is comprised of a pair of zener diodes 98 and 100. A 
to those skilled in the art. 2o second resistor R2 is coupled in parallel with voltage clamp 
The control of the power supply circuit 50 is controlled by 96. 
DCIU 18. DCIU 18 is also coupled to bus 52. A housekeep- Control logic 88 controls a switch 102. Switch 102 has a 
ing supply 60 may also be incorporated into power supply control input 102C that is coupled to control logic 88 
circuit 50. Housekeeping supply 60 may be used for other 25 through transformer 90. 
functions besides a centralized system and may not be Switch 102 is coupled between a tap 104 on inductor 80 
coupled to bus 52. through an isolating diode D5. Inductor 80 has a discharge 
Power supply circuit 50 includes a plurality of application output 106. Current at discharge output is monitored through 
specific power supplies. The application specific power a sensor 108. Sensor 108 is coupled to control logic 88 
10 
supplies are sized in terms of current and voltage based on 3o through a feedback input 110. Control logic 88 may also 
the specific components to which they are connected. The have a discharge onloff input 112. Discharge onloff input 112 
specific power supplies may include a discharge heater may be derived from other controllers within the spacecraft 
supply 62, discharge supply current source 64, screen supply such as DCIU. 
voltage source 66, an accelerator supply voltage source 68, In operation, the starter circuit 96 generally operates as 
a neutralizer supply current source 70, and a neutralizer 35 follows. When switch 102 is turned on, current increases in 
heater supply 72. Discharge heater supply 62 is coupled to inductor 80 to store energy therein. When switch 102 is 
heater 32 and is disposed within thruster 14. Discharge turned off rapidly, a high voltage spike is generated across 
supply current source 64 has a positive output 64P coupled inductor 80 which appears at discharge output 106. Dis- 
to anode 24. Discharge supply current source 64 also has a 
negative output coupled to cathode 26. Negative output may 
also be coupled to screen grid 44. Screen supply voltage 
source 66 has a positive output 66P that may also be coupled 
to anode 24. Accelerator supply voltage source 68 has a 
negative terminal coupled to accelerator 46. Neutralizer 
supply current source 70 has a positive output 70P coupled 
to neutralizer anode 28. Neutralizer supply current source 
has a negative output 70N coupled to neutralizer cathode 30. 
A filter capacitor 79 and a voltage clamp 77 may be coupled 
to negative output 77 of neutralizer supply 70. Neutralizer 
heater supply 72 is coupled to heater 34. Neutralizer heater 
supply 72 has a positive output 70P and a negative output 
charge output 106 may, for example, be coupled to the 
4o thruster anode 24 described above where the high voltage 
will generate ions which conduct discharge current to form 
the plasma. 
To determine whether a high voltage discharge is to be 
applied to output 106, control logic 88 monitors current at 
45 current sensor 108. If the plasma is maintaining ion genera- 
tion the current flowing between the cathode and anode is 
greater than 1.5 amps. In the present example, if the current 
monitor indicates less than 1.5 amps which is a level 
corresponding to no ion generation, the starter operation is 
so initiated as described in the following: A pulse is released 
through transformer 90 to drive the switch 102. Inductor 80 
70N. acts as an auto-transformer that boosts the voltage to about 
Anegative output 66N of screen supply voltage source 66, 200 volts for about 20 microseconds. When the arc is 
a positive output 68P of accelerator supply voltage source established between cathode 26 and anode 24, current from 
68, a negative output 70N of neutralizer supply current 55 discharge supply 64 maintains the plasma. When the dis- 
source 70 and negative output 72N of neutralizer heater charge current exceeds 1.5 amps, control logic 88 inhibits 
supply 72 may all be coupled together at the same electrical further pulses and thus enters a standby mode to conserve 
potential. Discharge arc starter circuit 76 and a neutralizer energy. If the discharge current drops below 1.5 amps 
arc starter circuit 78 may be coupled to cathodes 26, 30 (indicating that the arc has been extinguished), control logic 
respectively. As described above, arc starter circuits 76, 78 60 88 pulses switch 102 to generate high voltage pulses at 
are used to ignite the ion plasma. output 106. 
Referring now to FIG. 4, starter circuit 76 is illustrated in The same circuitry as starter circuit 76 may be used for 
further detail. Starter circuit 76 is identical to neutralizer neutralizer starter circuit 78. However, the threshold to 
starter circuit 78 except that the feedback current threshold initiate a high discharge output by control logic 88 need only 
is adjusted downward as will be further described below. 
Sufficient power to generate a high voltage pulse to While the best mode for carrying out the present event has 
initiate an arc is obtained from a power supply that is been described in detail, those familiar with the art to which 
65 be 0.5 amps. 
US 6,304,040 B1 
5 6 
this invention relates will recognize various alternative 
designs and embodiments for practicing the invention as 
defined by the following claims: 
What is claimed is: 
1. A starter circuit comprising: 
a power supply having an output inductor, said inductor 
a switch coupled to said tap having a control input; and 
a pulse control logic circuit coupled to said control input, 1o when said current output is below said threshold. 
said pulse control logic circuit controlling said switch 
to an on state to store energy and then to an off state to 
generate a high voltage discharge. 
2. Astarter circuit as recited in claim 1 further comprising 
a current sensor coupled to said output inductor for sensing 
3. A starter circuit as recited in claim 1 further comprising 
a discharge current sensor, said current sensor coupled to 
said pulse control logic circuit. 
4. A starter circuit as recited in claim 1 wherein said 
5 switch comprises a field effect transistor. 
5 .  A starter circuit as recited in claim 1 wherein said pulse 
control logic comparing said current output to a current 
threshold and controlling said switch to an on and off state 
6. A starter circuit as recited in claim 5 wherein said 
7. A starter circuit as recited in claim 5 wherein said 
having a tap; 
current threshold is 1.5 amps. 
current threshold is 0.5 amps. 
a current output of said inductor. * * * * *  


Starter circuit for an ion engine

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