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BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates in general to plasma generators and, in
particular, to plasma generators utilizing Penning-discharge type
configurations using dc discharges and magnetic fields to enhance electron
confinement.
2. Description of the Related Art
Orbiting spacecraft travelling through space plasma build up high
potentials on their outer surfaces which are often made of insulating
materials, due to bombardment by both negative and positive particles.
Such potentials are often localized on a particular portion of the
spacecraft, which results in different areas of the spacecraft's outer
shell being charged to different voltage potentials. For example, in a
nonspinning spacecraft in solar orbit, one side faces the sun and one
faces darkness. The side of the spacecraft which faces the sun will not
become charged, because radiant solar energy causes photoelectrons to be
emitted from the spacecraft's surface, and this photoemission compensates
for the incident electrons, thereby limiting charge buildup on that side
of the spacecraft. The dark side, however, can become charged to very
large negative potentials, e.g., -10 kilovolts.
The difference in potential between various areas on the spacecraft's
surface can cause an electrical discharge therebetween, which can damage
electronic equipment on the spacecraft.
Alternatively, particle bombardment can result in the entire spacecraft
being charged to a different potential than space plasma, which could, for
example, have an adverse effect on scientific satellites which measure the
ambient environment around the spacecraft.
Plasma sources, which produce positive ions and electrons, have been used
to discharge the surface of a spacecraft and to clamp the spacecraft to
space potential. Such plasma sources discharge the spacecraft because the
emitted electrons will be attracted to the positively charged side of the
spacecraft while the emitted ions will be attracted to the negatively
charged side of the spacecraft, thereby bringing the spacecraft's surfaces
to the same potential. In addition, the plasma source clamps the
spacecraft frame to space potential because the emitted plasma provides a
conductive bridge to the space plasma.
Prior plasma sources have suffered from the disadvantages of being slow to
ignite and of providing a low ion emission current, in addition to being
relatively large devices In prior hollow-cathode type Penning-discharge
devices, the cathodes require constant heating, to effect thermionic
emission of electrons, each time the plasma source is to be used. This can
result in the power consumption of the plasma source being on the order of
20 watts and several minutes being required to heat the cathode to
ignition temperature.
Accordingly, it is the primary object of the present invention to reduce
the size, ignition time, and gas and power consumption of a
Penning-discharge type plasma source, while achieving large electron and
ion currents.
SUMMARY OF THE INVENTION
The present invention, in a broad aspect, is a Penning-discharge type
plasma source including a cathode for thermionically emitting electrons,
an electron emission means disposed inside the cathode for thermionic
electron emission, anode means for accelerating electrons emitted by the
cathode and by the emission means to a discharge space defined by the
anode means, means for supplying gas to be ionized into the discharge
space, as well as a heater for heating the emission means and magnet means
for providing a magnetic field in said discharge space to increase
ionization of the gas. Initial heating of the emission means by the heater
causes electrons to be emitted therefrom. The emitted electrons are
accelerated by the positively-charged planar anode to ionize the gas in
the discharge space, with some of the ions being accelerated out of the
source and with other of the ions impacting the cathode to effect heating
thereof to cause thermionic emission. As a result, the ionization of the
gas continues and further heating of the emission means is unnecessary.
In accordance with one feature of the invention, the anode is a planar
anode. The planar anode and the magnetic field provided in the discharge
space to increase the ionization of the gas therein, cause a large
fraction of the plasma production to occur near the exit orifice of the
plasma source. It is a purpose of the present invention to provide a
plasma source particularly suited for applications such as spacecraft
charging control. In such applications, no ion-beam-accelerating component
is used with the plasma source, and therefore the ions exit from the
plasma source by diffusion or under the influence of the weak electric
fields associated with spacecraft charging. The present invention produces
most of the plasma near the exit orifice of the plasma source, and thereby
makes a larger percentage of the plasma produced available for spacecraft
charging control purposes even without additional accelerating components.
In accordance with another feature of the invention, the cathode is a
hollow cathode and the emission means is a foil insert of tantalum
material containing a barium compound disposed within the cathode for
ignition of the device.
In accordance with yet another feature of the invention, a plurality of
samarium-cobalt magnets encircle the discharge space to increase
ionization of the gas and to effect the Penning discharge.
The present invention also provides a novel method of igniting a
Penning-discharge plasma source of the type containing a thermionic
cathode for emitting electrons, and an anode for accelerating the
electrons into a discharge space receiving a gas flow, to effect
ionization of the gas. The method includes placing an electron emissive
cathode insert surface into the cathode, applying a discharge potential
across the discharge space to initiate ionization of the gas, and then
admitting a brief burst of high-pressure gas into the interspace between
the cathode and keeper-electrode, when a keeper electrode is used, which
is optional, or between the cathode and the anode. This precipitates an
arc-discharge between the emissive surface of the cathode insert and
either the keeper electrode or the planar anode. The arc-discharge rapidly
heats the cathode insert to thermionic emission temperature, at which
point normal hollow-cathode operation is established. This method of
ignition enables the source to be brought into full operation within
approximately one second.
It is a purpose of the present invention to provide a plasma source which
is capable of producing a relatively large ion current with very modest
input power requirements.
Another purpose of the present invention is to provide a plasma source
which achieves very rapid ignition when started-up from a cold condition.
Furthermore, the plasma source of the present invention has a long
life-time, and the plasma produced is low energy plasma.
The present invention, by virtue of the features and purposes enumerated
above, is particularly well-suited for spacecraft charging control.
However, the usefulness of the present invention is by no means limited to
spacecraft charging control, but extends also to other applications
requiring efficient, low energy plasma production.
Other purposes, features, and advantages of the present invention will
become apparent from the consideration of the following detailed
description and from the accompanying drawings.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 shows a schematic view of a Penning-discharge plasma source
according to the present invention; and
FIG. 2 shows characteristics of the plasma source shown in FIG. 1 in
operation with argon and xenon gases.
FIG. 3 shows the magnetic field provided inside the plasma source shown in
FIG. 1.
FIG. 4 shows in greater detail the cathode portion of the device shown in
FIG. 1.
DETAILED DESCRIPTION
Referring more particularly to the drawings, FIG. 1 shows a compact
Penning-discharge plasma source 10 according to the present invention. The
plasma source 10 has the characteristics of low power consumption, high
gas efficiency, long cathode lifetime, as well as being compact and low in
mass. The plasma source 10 comprises a cylindrical cathode 12 which may be
on the order of 3 mm in diameter. The cylindrical cathode 12 conducts gas
admitted into it via a cathode tube 13 to a discharge space 23. The
discharge space 23 is defined on one side by planar anode 7.
Although the preferred embodiment of the present invention features a
planar anode, shown in FIG. 1 and discussed below, a cylindrical anode can
also be used in an alternative embodiment. When a cylindrical anode is
used, discharge space 23 is defined by the portion of the cylindrical
anode which is above the cathode 12.
Magnets 16 in a ring configuration are located at two axial and several
azimuthal positions. Magnets 16 may be samarium-cobalt (SmCo.sub.5)
magnets or other permanent magnets or electromagnets with pole-pieces
placed in a ring configuration, and are placed so that one ring is
adjacent to planar anode 7 and is downstream with respect to the gas flow,
whereas the other ring of magnets 16 is upstream with respect to the gas
flow, that is, closer to the tip 30 of cathode 12 and the point at which
the gas is introduced into cathode 12. These magnets 16 provide a strong
divergent axial magnetic field to increase the iongeneration capability of
the source 10. Magnetic flux is returned through an iron shield 20 which
also retains the magnets and reduces the stray magnetic field leaving the
plasma source 10. The iron shield 20 is in turn surrounded by an outer
enclosure 18 which is at cathode (common) potential. (Common potential is
typically connected to spacecraft ground through current detectors which
measure the plasma-source emission current.)
Iron shield 20 is secured to and electrically isolated from outer enclosure
18 by insulating mountings 44. In the preferred embodiment an annular
seating ridge 46 is used to assure rigidity in the shield and enclosure
structure for the rigors of spacecraft use.
The plasma source enclosure 18 comprises a sealed cylinder with a plasma
exit-orifice 24, which is the only unsealed opening. This orifice 24 may
be covered during launch by a blowopen cover (not shown) to protect the
source from contamination. A vacuum seal can also be made to the flange of
the orifice 24 to permit evacuating the plasma source and operating it on
the ground for pre-launch checkout.
Internally, the plasma source 10 may be supported by ceramic insulators or
other means known in the art which allow the anode 7, the outer enclosure
18, and spacecraft ground to be at different electrical potentials.
The cathode 12 includes a tip portion 30 having an orifice 32 of smaller
diameter than that of the cathode 12 as shown in FIG. 4. A cathode heater
36, including a plurality of heating coils 34 and a radiation shield 38 is
used for first-time-only "conditioning" of the cathode. The cathode heater
36 is connected to a cathode heater supply 26. A discharge supply 28 is
connected between the cathode 12 and the anode 7 to ignite the source.
A cathode insert 40, which may be a tantalum-foil insert, is utilized in
the ignition of the source 10, as explained in more detail below.
In the preferred embodiment, insert 40 is a tantalum foil coiled up in
layers about a central axis so as to form a cylinder. There are spaces
between the adjacent tantalum layers of insert 40 to provide egress for
escaping electrons. The tantalum foil of insert 40 is attached to a
support 42 which secures insert 40 to the walls of cathode tube 13. Insert
40 is covered with a barium-containing compound to reduce its work
function, thereby allowing electron emission at relatively low
temperatures, on the order of approximately 900 degrees Celsius. This
insert 40 eliminates the need for heating of the cathode by an external
supply as employed in the prior art.
The cathode tip 30 may be made of impregnated porous tungsten material
which is capable of thermionic emission. The cathode insert 40 can also
comprise an impregnated porous tungsten matrix insert with a tantalum foil
extension having the barium coating. The anode 7 can be made of a
molybdenum material, although the anode material is not especially
critical.
Concerning the cathode heater supply 26 and discharge supply 28, both are
conventional. The heater supply 26 can be an AC on a DC supply providing
the necessary current to the heating coils 34. The discharge supply is
designed to provide from several hundred to 1000 volts at extremely low
current to light the initial discharge, and then to provide a lower
voltage, e.g., 20-30 volts, and a somewhat higher current, on the order of
200-500 mA, to maintain the discharge. The keeper power supply 11 and
discharge supply 28 are designed to limit both the energy and current
during the gas-burst ignition to avoid arc damage to the cathode insert
40.
In normal steady-state operation, xenon or argon gas is admitted by a
gas-feed system 51 into the cathode tube 13 for passage through the
cathode orifice 32 into the discharge space 23 where ionization occurs.
Electrons are emitted thermionically from the cathode insert 40. Electrons
leaving the cathode orifice 32 are accelerated to the potential of the
plasma in the discharge space 23, which is near anode potential. The
electrons then oscillate axially between the cathode tip 30 and the outer
shield 18, both of which the electrons are energetically unable to reach.
The electrons are confined radially by a strong magnetic field which is
produced by the two magnetic rings 16, the magnetic flux of which is
linked externally by the iron magnetic shield 20. This magnetic field is
illustrated in FIG. 3 by dashed lines 70, and exhibits a cusped-shaped
null point 72 and an axial maximum 74 near the aperture 8 in anode 7. This
magnetic field geometry causes most of the plasma produced by the source
to be generated in the region between points 72 and 74. Therefore this
plasma has a high likelihood of exiting before recombination through exit
orifice 24 which is at ground potential. Thus, this arrangement results in
plasma being formed near the exit aperture in the anode, thereby resulting
in a significant increase in the ratio of the plasma-production-rate to
the input power. This is a significant improvement over prior devices
wherein the magnetic-field and anode geometries cause the region of
highest plasma production to occur deep inside the source and further away
from the exit aperture. This increases the probability of the plasma
undergoing wall recombination before leaving the source, and thereby
decreases the ratio of plasma production to input power of the source.
This causes such sources to consume excessive power for the ion current
produced. Moreover, in the present invention, this improvement in
efficiency of production of plasma flux is achieved without the use of ion
accelerating extraction grids near the exit aperture. Use of ion
extraction grids typically increases power and gas flow rate requirements
and results in a high energy ion beam which is not suitable for spacecraft
charging control applications.
Thus, in the discharge space 23, the electrons are trapped by the magnetic
and electrostatic mechanisms until they undergo collisions with the gas
that fills the discharge space 23. These collisions have a high
probability of producing new electron-ion pairs, resulting in a relatively
high ion-generation rate.
Some of the ions which are formed in the discharge space impact the cathode
tip 30, and maintain its temperature at a level consistent with thermionic
electron emission. The electric fields produced by the charged surfaces
cause the other ions to leave the plasma source via the exit aperture 24
in the outer shield 18, and these ions are attracted by and neutralize the
negatively charged spacecraft surfaces. Electrons also leave by the exit
aperture 24, and they are accelerated away from the spacecraft until
charge neutrality is accomplished.
The cathode 12 is ignited by a unique `gas-burst` method. A conventional
gas feed system 51 is used. Gas from a tank 60 passes through a
pressure-reducing regulator 57 through a high pressure valve 58, a bypass
valve 56 and into a burst reservoir 54. With low-pressure valve 50 closed,
when the burst reservoir 54 is filled to equilibrium pressure, bypass
valve 56 is closed. To ignite the cathode 12 and thereby activate source
10, a high voltage, on the order of about 1000 volts, is applied to keeper
electrode 4 by keeper power supply 11. Valve 50 is then opened. High
pressure gas then rushes from reservoir 54 into the cathode 12--keeper
electrode 4 interspace and initiates an arc discharge between the cathode
insert 40 and the keeper electrode 4. The arc discharge rapidly heats the
cathode insert 40 to a temperature near 1000.degree. C., at which point
the discharge changes from an arc discharge to a thermionic hollow-cathode
discharge. When the transition is complete, the cathode-to-keeper voltage
falls to a small value, typically 15 V. During this time, the supply
high-pressure gas from the reservoir 54 is exhausted, and the gas flowrate
reaches equilibrium at a low rate which is determined by the
characteristics of the flow impedance 52, and is typically on the order of
8.times.10.sup.-4 Pam.sup.3 s.sup.-1 or 0.5 standard cm.sup.3 per minute.
Thermionic hollow-cathode operation produces a partially ionized plasma in
the cathode-keeper interspace. This plasma is the source of electrons
which flow through the orifice 5 in the keeper electrode 4 into the
discharge space 23 when a positive voltage, typically about 20 V, is
applied to the anode 7 from the discharge supply 28.
First time ignition of the plasma source is initiated by briefly heating
the heating coils 34 in the cathode heater 36 using the cathode heater
supply activates the emissive compound in the insert 40 and causes
electrons to be emitted from it. A large (approximately 1 kilovolt)
potential difference is then applied by the discharge supply 28 between
the anode and the cathode, and a brief burst of higher than normal gas
flow is admitted by the gas feed system 51 as described earlier. The high
potential difference from the discharge supply 28 initiates the Penning
discharge. The discharge power supply characteristic is chosen to cause a
rapid transition to low discharge voltages (approximately 20 volts) as the
discharge ignites. Subsequent ignitions of the plasma source are
accomplished by the foregoing procedure, but without the need to heat
cathode heater 36 with cathode heater supply 26.
As seen from the foregoing, the cathode heater 36 is merely used to
condition the cathode (i.e., thermally reduce the emissive compound) in
preparation for the first ignition in space. It is not used afterwards,
except in the event of cathode contamination.
The plasma source 10 just described is an exceptionally compact,
fast-starting, low-power plasma source which is capable of delivering
relatively large electron and ion currents. The plasma source 10 has the
unique advantages of high ion-emission current capability (greater than 1
mA), low power and gas consumption (less than 15 watts and less than
8.5.times.10.sup.-4 Pa.multidot.m.sup.3 s.sup.-1 of gas), and rapid
ignition (less than 1 second). These attributes make plasma source 10
especially well-suited to the spacecraft charge-neutralization application
because it places a minimum power and mass burden on the host spacecraft.
Additionally, the expected lifetime is greater than 20,000 hours, with an
expected restart capability of greater than 10,000 starts.
Furthermore, the lower operating discharge voltage, on the order of 25 V of
the present invention significantly reduces sputtering. Sputtering is a
problem when conventional ion sources with typically higher discharge
voltages on the order of 40 volts are used in spacecraft. At the higher
voltages there is generally a significant amount of sputtering of the
cathode and other surfaces at the cathode potential.
FIG. 2 shows the characteristics of the plasma source 10 in operation with
both argon and xenon gases. The advantages of operation with xenon are
apparent in FIG. 2. Lower discharge voltages (and consequently, lower
power consumption) are achieved at a given gas flow rate than with argon.
This advantage is associated with the higher atomic mass of xenon. An
operating point near the "knee" of the voltage-flow characteristic will
afford a low discharge voltage and a relative insensitivity to small
changes in flow without the high-flow rate penalty which is associated
with operating on the flat portion of the curve. Hollow cathode type
sources are most often operated in this "knee" region of the voltage-flow
rate characteristic.
The gas feed system 51 referred to above, which is not part of the present
invention, is also conventional and provides the following two
characteristics. First, feed system 51 provides a low constant gas flow
rate during plasma source 10 operation, for example, a xenon flow rate of
approximately 8.5.times.10.sup.-4 Pa.multidot.m.sup.3 s.sup.-1 . Second,
in order to accomplish reliable ignition of plasma source 10, a brief
initial burst of gas at a higher pressure than is needed during normal
sustained-running operation can be injected by feed system 51.
Lastly, by using electromagnets as the magnets 16, the magnetic field
surrounding the anode 7 can be adjusted and therefore act as throttle
control for large ion emission current capability. The separate keeper
electrode 4 in the ion source shown in FIG. 1, and adjustment of the
ion-emission current by varying the discharge current may also be used as
a throttle mechanism in the present invention. If adjustment of the
discharge current is used as the throttling mechanism, then the keeper
electrode is needed to sustain cathode operation when the discharge
current is reduced. In alternative embodiments, the separate keeper
electrode can be omitted. When a keeper electrode is not included, to
ignite the source using the gas-burst method, the high voltage on the
order of about 1000 volts earlier described as being applied to keeper
electrode 4 is applied instead to the anode 7. The ignition steps remain
basically the same. When the transition is complete, the cathode-to-anode
voltage falls to a low value, typically 25 volts.
To illustrate the peformance characteristics of the plasma source 10,
typical operating parameters in steady-state operation are given below.
These parameters are being provided merely by way of example and not as
limitations. These values correspond to operation with the preferred
operating gas, which is a 90%-xenon and 10%-hydrogen mixture. With zero
cathode-heater power, a keeper electrode current of about 0.25 A and
voltage of about 19.0 V, a discharge current of about 0.2 A and discharge
voltage of about 23.5 V, an ionemission current of about 0.001 A can be
obtained with modest total input power of about 9.5 W and a low gas flow
rate of approximately 8.times.10.sup.-4 Pam.sup.3 s.sup.-1 .
In the foregoing description of the present invention, a preferred
embodiment of the invention has been disclosed. It is to be understood
that other mechanical and design variations are within the scope of the
present invention. Accordingly, the invention is not limited to the
particular arrangement such as has been illustrated and described in
detail herein.
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