Information
-
Patent Grant
-
6216445
-
Patent Number
6,216,445
-
Date Filed
Wednesday, May 19, 199925 years ago
-
Date Issued
Tuesday, April 17, 200123 years ago
-
Inventors
-
Original Assignees
-
Examiners
- Thorpe; Timothy S.
- Rodriguez; William
Agents
-
CPC
-
US Classifications
Field of Search
US
- 060 2031
- 060 202
- 060 204
- 060 2001
- 315 11101
- 315 1112
-
International Classifications
-
Abstract
A pulse plasma thruster (50) utilizes a vapor producing solid (54) and a micro-sized heater (52) to produce a high pressure vapor that is directed into an ignition chamber (58) and to a thrust discharge chamber (70). The thrust discharge chamber (70) comprises two oppositely disposed electrode plates (72, 74) and oppositely disposed fuel propellants sources (60, 62). The passageway (56) leading from vapor producing solid (54) to the thrust discharge chamber (70) is configured to permit uniform feeding of the vapors to the thrust discharge chamber (70). A pair of electrode terminals (82, 84) extend from the electrode plates (72, 74) and through a housing (88). A power source (100) is coupled to the terminals (82, 84) and provides the ignition signals necessary to cause a spark and a breakdown to a useful plasma arc by controlling the voltage-current shape of the ignition signal.
Description
TECHNICAL FIELD
The invention relates generally to plasma thrusters and more particularly to a miniature pulsed plasma thruster capable of efficiently generating very small impulse bits at low levels of power and DC ignition voltages.
BACKGROUND OF THE INVENTION
Space vessels such as spaceships and satellites utilize thrusters to achieve motion in space. A thruster operates on the principle that a force generated in one direction generates an equal force in the opposite direction. By emitting a reaction-mass, a thruster accelerates a spacecraft in the opposite direction. A thruster may be used as a small rocket engine for orbit correction or as the main propulsion of the spacecraft.
Older conventional thrusters used chemical propulsion, which utilized liquid and/or solid propellants. Electric thrusters, which accelerate gases by electrical heating and/or by electric and magnetic field forces, can outperform chemical propulsion systems, in part, because of their high specific impulse (Isp) values. Advantages of electric thrusters include high efficiency and performance, low weight, increased spacecraft orbiting lifetimes, reduced overall costs, and a savings in fuel mass. Advances in onboard electric power sources and smaller more efficient electronic devices have expanded the use of electric thrusters in spacecraft applications.
Electric thrusters that convert electrical energy into kinetic energy may be grouped into three categories: electro thermal propulsion, electrostatic or ion propulsion, and electromagnetic propulsion. Within the electromagnetic propulsion category is the Pulsed Plasma Thruster (PPT), which accelerates the propellant plasma via interaction with an electric arc.
Multiple government and civil entities are developing small and micro sized spacecraft that can benefit from PPTs for space missions. Such spacecraft will require major reductions in thrust levels and/or impulse bits to ensure proper and precise control of the spacecraft. Many missions, in particular those that require significant mission propulsion energies and/or acceleration, will require specific impulses beyond those available from chemical rockets. Because present electric rockets cannot efficiently operate a very low level of power and impulse bits they are not well suited for such missions.
While PPTs are at a high state of development, they generally require high levels of voltage and power to initiate the plasma breakdown and are also very inefficient at low powers when operated at values of expelled propellant velocities of interest to space missions. For example, experimental PPTs have been operated at energy levels down to about 2 joules (J) per pulse requiring the use of high voltage charging supplies which can range from 2,000 to 8,000 volts depending on the design. Also, efficiencies of PPTs decrease with decreasing power and presently, are less than 10 percent efficient when operated at values of propellant velocities of interest to space systems. The inefficiencies result in significant increases in power to achieve desired levels of impulse bits.
An example of such a thruster is shown in FIG.
1
and denoted generally as
10
. The thruster
10
fits into the class of propellant devices that operates using an all gas propellent although an all solid solution could also be utilized. In particular, the thruster
10
utilizes a low atomic weight liquid propellant such as water or monopropellant hydrazine (N
2
H
4
) or a mixture of two liquids such as water and hydrazine which is stored in the tank
12
and flows through a conduit
14
leading to an opening
16
that forms the feeding mechanism of the thruster
10
. The liquid propellent within the tank
12
may be pressurized by high pressure helium in the tank
20
, in a manner well known to those of ordinary skill in the art.
The liquid propellent flows through the conduit
14
via the opening
16
and reaches a passage
18
within the thruster
10
. The passage
18
leads to a small opening
22
which is sized to provide the correct flow velocity for the liquid propellent and reduce back flow into the passage
18
. In the passage
18
, the liquid propellent is partially or fully atomized and partially evaporated, so that there is a two phase flow of liquid and gas into the thruster
10
. The liquid propellent is disassociated into low atomic weight elemental constituents thereof by an electric discharge that forms a plasma arc within the thruster
10
.
The liquid gas and plasma flow from an open end
24
of the passage
18
into the thrust nozzle
30
which, as shown, is shaped as a cone or bell having a curved confining surface, to provide high efficiency and conversion of the high pressure plasma into a directed supersonic flow having high momentum. This discharge of plasma is established primarily by the use of a high voltage DC (HVDC) power supply
32
which is coupled to electrodes
34
and
36
of the thruster
10
.
In particular, the thruster
10
operates when liquid from the tank
12
flows into the passage
18
and a high voltage ignition signal supplied by the HVDC power supply
32
is applied at terminals
34
and
36
at a predetermined frequency, such as 200 pulses per second, for example. This ignition voltage can vary but according to one design ranges from 2,000 volts to 8,000 volts. The ignition signal supplied by the HVDC power supply
32
causes a discharge to be established in the passage
18
between the electrodes
34
and
36
at a time when partially atomized fluid is entering the thrust nozzle
30
through the opening
24
. The velocity and mass flow rate of liquid flowing through the passage
18
and the repetition rate and energy of the plasma discharge between the electrodes
34
and
36
are matched to achieve optimum operation.
Typically, the HVDC power supply
32
raises the voltage of the thruster
10
until an electrical breakdown occurs between the electrodes
34
and
36
. The requirement, however, that the HVDC supply
32
generate high levels of ignition voltages makes the thruster
10
unsuitable for many propulsion applications where small spacecraft are involved. The HVDC supply
32
can be large and not well suited for such applications. Moreover due to its size, the HVDC supply
32
makes it difficult to achieve small and precise maneuvers for some spacecraft missions.
For many space mission applications, where small space systems are involved and which require extremely precise control, the use of high power and/or high voltage ignition circuits is impractical. Examples of such missions are those which require extremely precise ephemeris control and those which are otherwise penalized by high thrust, such as missions which require multiple acceleration and deceleration maneuvers. Thus a PPT that is able to efficiently operate without a high voltage ignitor system and at power levels several orders of magnitude less than prior art designs would be advantageous.
SUMMARY OF THE INVENTION
The present invention is a pulsed plasma thruster (PPT) capable of operating at low levels of power and impulse bits that is suitable for use in space applications where the space system is small and precise control of the spacecraft is required. The PPT of the present invention is capable of delivering reliable ignition of a spark breakdown at DC voltages less than 300 volts with reliable transfer of a spark to a useful plasma arc. The ablation, combustion and acceleration of the Polytetra Fluorethylene (PTFE) fuel propellent is precisely controlled with the use of miniaturized PPT and power processor components. The efficiency of the thruster is increased by the independent introduction of vapor (such a from a subliming solid) at optimal locations and times during the operational cycle.
According to one embodiment, disclosed is a PPT having optimally located solids capable of producing high vapor pressures for purposes of enhancing both ignition and efficiency. Heat generating elements, such as micro-heaters, are placed adjacent to the solids and configured to generate heat that causes the solids to sublime. The PPT includes an igniter section that forms a passageway from the solid to a thrust discharge chamber. In one embodiment, the ignition chamber includes a plurality of holes which are sized and spaced for optimally guiding vapors to the thrust discharge chamber for purposes of enabling arc ignitions at low voltages. In one embodiment, solids are also located within the thrust discharge chamber and, via the use of heat generating elements, independently introduce vapors into the thrust discharge chamber in order to enhance PPT efficiency at desired values of propellant velocities.
The thrust discharge chamber includes a set of properly spaced and shaped electrode plates which provide for transfer of an initial spark to a useful plasma arc in the gap defined by the electrodes plates. A solid propellent, such as PTFE, is provided within the thrust discharge chamber and arranged so that the plasma arc traveling through the thrust discharge chamber will ablate the PTFE and accelerate the plasma formed from ablated PTFE and the independently introduced vapor from high vapor pressure solids, as used.
A power processing unit provides the DC ignition voltage necessary to cause a spark to occur in the gap between the electrode plates. In one embodiment, the power processor unit has a variable output that operates in three segments: an open circuit to constant voltage segment, a constant voltage segment, and a constant current segment.
A high vapor pressure between the electrode plates is created when the heat generating means heats the solid to assist in ignition and transition of a spark to a useful plasma arc. Micro-heaters can also be embedded in, or at the edges of, the PTFE propellent and its temperature varied to control the amount of PTFE ablated to provide more control of the impulse generated by the PPT. Micro-heaters embedded in the solids, located in the ignitions section and/or the thrust discharge chamber, independently provide a source of vapor to the thrust discharge chamber to provide additional and independent control of the efficiency and impulse of the PPT. In one embodiment the electrode plates are equally spaced about a central axis through the thrust discharge chamber. In another embodiment, the PPT includes a means of varying the spacing between the electrode plates as a function of axial distance. In an other embodiment, slightly radioactive electrodes are used. In these ways ignition voltages and required power levels are achieved that are several orders of magnitude smaller than those previously obtainable.
Also disclosed is a method of operating a pulsed plasma thruster comprising the steps of heating a subliming solid to create a high pressure vapor and directing that high pressure vapor in the direction of a thrust discharge chamber through an ignition chamber. Next, a DC ignition signal is applied to electrodes coupled to the thrust discharge chamber that sparks a breakdown of a fuel propellent and causes a transition of the spark to a useful plasma arc. The DC ignition signal is applied in a way that its shape and magnitude are controlled. In one embodiment the DC ignition signal is controlled in three segments corresponding to an open circuit to constant voltage segment, a constant voltage segment and a constant current segment.
The high pressure vapor is directed to the thrust discharge chamber so that pressure is created between two electrode plates. The vapor can be fed uniformly to control ignition and breakdown of the fuel propellent. The spacing between the electrode plates may be adjusted to control the amount of the fuel propellent ablated. A source of ultraviolet radiation may be focused on the vapor to provide additional excitation energy that helps ignite the vapor from the subliming solid.
A technical advantage of the invention is the enablement of reliable ignitions at voltages more than an order of magnitude less that previously obtainable. This enables small and light-weight PPTs and power supplies and, therefore, much lighter PPT systems than previously obtainable.
Another advantage is the efficient enablement of impulse bits several orders of magnitude less than previously obtainable. This enables the deployment of PPTs suitable for space propulsion applications involving small spacecraft systems and for missions which require extremely precise control of the spacecraft.
BRIEF DESCRIPTION OF THE DRAWINGS
Other advantages of the invention including specific embodiments are understood by reference to the following detailed description taken in conjunction with the appended drawings in which:
FIG. 1
illustrates a prior art liquid thruster that uses a high voltage power supply to create thrust;
FIGS. 2
a
and
2
b
illustrate embodiments of a Pulsed Plasma Thruster (PPT) according to the invention;
FIG. 3
is a cross section view of the ignition chamber of the PPT of the invention;
FIGS. 4
a
and
4
b
illustrate the use and operation of the variable power processing unit that powers the pulsed plasma thruster according to one embodiment; and
FIG. 5
illustrate the micro-positioning of the electrode plates of the pulsed plasma thruster according to one embodiment.
References in the detailed description correspond to like references in the figures unless otherwise indicated.
DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS
The present invention provides a pulsed plasma thruster (PPT) that can be used as a rocket engine for small spacecraft. The PPT operates on a pulse basis where a spark is created at low voltage via the use of small separations of electrodes using micro-electromechanical systems (MEMS) technology, independent introduction of vapor from solids, and electrodes which are slightly radioactive and specially shaped. The spark is transferred to an arc via use of a power supply with three output sections. The arc creates a plasma consisting of constituents of PTFE, which is ablated by the arc, and the vapors from the solids.
Referring to
FIG. 2
a
, a pulsed plasma thruster (PPT) according to one embodiment of the invention is shown and denoted generally as
50
. The PPT
50
includes a heater
52
or other means of generating heat that is small enough to accommodate the framework of a small spacecraft. In one embodiment, the heater
52
is a micro-sized heater based on micro-electromechanical systems or MEMS technology.
The heater
52
is placed adjacent a subliming solid
54
. The purpose of the subliming solid
54
is to provide a vapor source so that, in combination with the heater
52
, the solid
54
generates a gas flow that assists ignition of an initial plasma arc in the spark region of the thruster
50
. Thus, the heater
52
increases the temperature of the subliming solid
54
which, in turn, generates vapor. The vapor flows through a screen
56
and into an ignition section
58
of the thrust discharge chamber
70
where a spark partially ignites the solid fuel propellant
60
as well as some of the subliming solid
54
that has been vaporized. The action of the subliming solid
54
and resulting vapor, coupled with the screen
56
and configuration of the ignition section
58
assist in igniting a spark that creates a useful plasma arc.
In general, the subliming solid
54
has the characteristic of being able to produce a vapor when heated. A low sublimation temperature of the solid
54
is desired so a large quantity of vapor gas is generated for relatively small incremental changes in temperature. This reduces the heat generating requirements of the heater
52
. While some gases provide better ignition sources than others, the requirement that the solid
54
produce easily ionized vapor restricts selection of the material to certain compounds. Candidates include carbonates (X(HCO
3
) and carbamates (S(CO
2
NH
2
)) which sublime into NH
3
, CO
2
, and H
2
O. The use of subliming solids enables independent addition of vapor into the PPT, eliminates the requirements for valves and seals, and assures long term compatibility with space environments.
FIG. 3
is a cross section of the PPT
50
taken along line
3
—
3
of
FIG. 2
a
and illustrating the arrangement of the ignition section of the thrust discharge chamber
70
in greater detail. As shown, the screen
56
contains a plurality of holes
80
which are spaced and sized to provide optimum feeding of vapor from the subliming solid
54
into the thrust discharge chamber
70
. The number of holes
80
depends on the size of the ignition section
58
and the requirement that plasma in the ignition chamber must be allowed to enter the chamber that holds the solid
54
Thus, the sizing, diameter and quantity of the holes
80
is influenced by the specific configuration of the PPT
50
.
Preferably, the velocity of the vapor into the ignition section
58
is kept relatively low. In general, many small holes are more effective than a few big holes. Also, the screen
56
is designed to separate the solid from the thrust discharge chamber
70
so that sparks and/or plasma does not interact with solid
54
.
As shown, the thrust discharge chamber
70
is comprised of the two oppositely disposed electrode plates
72
and
74
and two fuel propellants
60
and
62
. The fuel propellants
60
and
62
are preferably PTFE based, although other fuel sources may be utilized. In one embodiment, MEMS based micro-heaters (not shown) are embedded in the fuel propellents
60
and
62
and their temperature varied to control the amount of PTFE ablated and to provide more control of the impulse generated by the PPT
50
. In another embodiment, solids
54
are placed along the thrust discharge chamber
70
and nozzle
90
. The solids
54
contain micro-heaters which are independently controlled to allow the introduction of vapor into the thrust discharge chamber at optimum locations and times during the firing cycle. This vapor provides additional control of the efficiency and the impulse bits generated by the PPT
50
.
Referring again to
FIG. 2
a
, the PPT
50
also includes a set of electrode plates
72
and
74
. The electrode plates
72
and
74
correspond to the anode and cathodes of the PPT
50
, respectively. As shown, the distance “d” corresponds to the spacing between the electrode plates
72
and
74
. In one embodiment, the distance “d” between the electrode plates
72
and
74
is 50 micrometers or less. Additionally, the electrode plates
72
and
74
are positioned so that they are evenly displaced about the central axis “x” running through the thrust discharge chamber
70
of the PPT
50
.
An advantage of the PPT
50
is the ability to create a reliable breakdown within the thrust discharge chamber
70
using low levels of power. This is achieved, in part, by keeping the spacing “d” between the electrode plates
72
and
74
small so that a spark is more efficiently generated and ignition is achieved using less spark energy. Recent advances in MEMS technology enables the manufacture of small clearances between the electrode plates
72
and
74
. Thus, the fact that the PPT
50
incorporates MEMS technology provides a PPT
50
suitable for space missions where power is limited.
According to various embodiments, the electrode plates
72
and
74
are spaced anywhere from 1 micrometer to 50 micrometers apart. In general, the closer the electrode plates
72
and
74
are spaced, the lower voltage is required to a ignite a breakdown.
Coupled to the electrode plates
72
and
74
are corresponding electrode terminals
82
and
84
that extend through an insulating layer
86
and the housing
88
. The electrode terminals
82
and
84
are used to deliver the ignition voltage to the thrust discharge chamber
70
. The insulating layer
86
extends substantially over the thrust discharge chamber
70
and the thrust nozzle
90
. As is known to those of ordinary skill in the art, the insulating layer
86
can be configured to increase the local field strengths existing between electrode plates
72
and
74
.
A disadvantage of prior art thrusters is that they require very high ignition voltages to operate. For example, the thruster
10
requires a DC supply anywhere from 2000 volts to 8000 volts. Such high voltages have been used in PPTs for a long time since they result in greater thrust. The present invention contemplates the use of voltages less than 300 volts. In one embodiment, the spacing of the electrode plates
72
and
74
is such that 50 volts is sufficient to create suitable thrust levels. This permits the PPT
50
to be utilized in typical satellite applications where 50 volts is commonly available.
FIG. 2
b
illustrates another configuration of the PPT
50
according to the invention. Specifically, the PPT
50
is shown equipped with a means of adjusting the angle of the thrust nozzle
90
with respect to central axis “x”. The hinges
92
and
94
are provided for this purpose although other means of achieving the same function can be employed. In this way, the PPT
50
becomes a fuel dynamic device since the angle of the thrust nozzle
90
has some effect on the amount of fuel utilized for certain levels of thrust.
Referring to
FIG. 4
a
, therein is shown the PPT
50
driven by a power source
100
with terminals
102
and
104
coupled to electrode terminals
82
and
84
, respectively. In general, the power source
100
is capable of producing multiple volt-ampere signal forms that effect the shape and magnitude of the ignition signal used to spark the vapors in the ignition section
58
.
In one embodiment, the power source
100
comprises a flexible power processing unit that operates in the three segments: an open circuit to constant voltage segment, a constant voltage segment, and a constant current segment. The three segments are illustrated in the graph of
FIG. 4
b.
The open circuit voltage, Vo, from the power source
100
is applied to the electrodes. The vapor from solid
54
is also introduced into the ignition section
58
of the thrust discharge chamber. A spark occurs in the ignition section
58
. During the next segment, the voltage decreases to the constant voltage section of the power source
100
at current Ic. The current then increases at a constant voltage, Vc, to a constant current section where the current is held constant at Io. The values of Vo, Vc, Ic, and Io are preset to desired values dependent on the specific design and operating condition of the PPT. Designs of power supplies capable of such outputs are known to those of ordinary skill in the art.
With reference to
FIG. 5
, the PPT
50
is equipped with micro-positioning devices
110
and
112
operably coupled to the electrode terminals
82
and
84
, respectively. The purpose of the micro-positioning devices
110
and
112
is to adjust the positioning and spacing of the electrode plates
72
and
74
with respect to the central axis “x”. Preferably, the micro-positioning devices
110
and
112
are MEMS based so that they fit the framework of a small spacecraft and require only small amounts of power to operate. In this way, the spacing between each electrode plates
72
and
74
can be varied as a function of axial distance from the upstream end of the thrust discharge chamber
70
.
While the invention has been described in conjunction with preferred embodiments, it should be understood that modifications will become apparent to those of ordinary skill in the art and that such modifications are therein to be included within the scope of the invention and the following claims.
Claims
- 1. A pulsed plasma thruster comprising:vapor producing solids; heat producing means arranged adjacent said solids; a thruster housing having a thrust discharge chamber with a plurality of openings, a thrust nozzle, and a fuel propellant in said thrust discharge chamber; passageways leading from said solids to said thrust discharge chamber within said housing, the passageways arranged so that vapors from said solids are received through said openings of said thrust discharge chamber; first and second electrodes extending from said thrust discharge chamber through said housing; and a power source coupled to said first and second electrodes and configured to enable spark breakdown between the electrodes of said thrust discharge chamber, the power source configured to control the voltage-current shape of spark breakdown that results in a plasma arc capable of ablating said fuel propellant, and ionizing said solid vapors and fuel propellants within said thrust discharge chamber to create a thrust force outwardly directed from said thrust nozzle.
- 2. The thruster of claim 1 wherein said passageways are configured to separate said solids from said thrust discharge chamber so that sparks and plasma do not interact with said solids.
- 3. The thruster of claim 1 wherein said passageways includes a plurality of holes adjacent said solids.
- 4. The thruster of claim 3 wherein said holes are arranged so that vapors are optimally fed into said thrust discharge chamber.
- 5. The thruster of claim 1 further comprising a means of varying the angle of said thrust nozzle with respect to a central axis extending through said thrust discharge chamber.
- 6. The thruster of claim 1 wherein said heat producing means are MEMS micro heaters capable of independently producing sufficient quantities of heat to cause said solids to sublime at times and locations to reduce ignition voltages and increase the PPT efficiency at desired values of propellant velocity.
- 7. The thruster of claim 1 wherein said first and second electrodes are made of a slightly radioactive material.
- 8. The thruster of claim 1 wherein said power source delivers a maximum DC voltage signal of 300 volts.
- 9. The thruster of claim 1 wherein said power source is configured to control the shape and magnitude of said ignition signal in three segments corresponding to an open circuit to a constant voltage segment, a constant voltage segment, and a constant current segment.
- 10. The thruster according to claim 1 further comprising a means of varying the spacing as a function of axial distance between said electrodes.
- 11. The thruster according to claim 1 wherein the fuel propellant comprises PTFE.
- 12. A pulsed plasma thruster comprising:solids capable of producing a high pressure vapors; heat generating elements adjacent said solid and operably configured to generate heat that causes said solids to sublime; an ignition chamber forming a passageway from said solid to a thrust discharge chamber, said ignition chamber having a plurality of holes for guiding vapors from said solid to said thrust discharge chamber; said thrust discharge chamber having first and second ends, said first end coupled to said ignition chamber, said thrust discharge chamber further including two oppositely positioned electrode plates and two oppositely positioned fuel propellants, each of said electrode plates coupled to corresponding electrode terminals a nozzle coupled to said second end of said thrust discharge chamber; and a power processor unit coupled to said electrode plates through said electrode terminals and configured to provide an ignition voltage that causes a plasma arc to occur in the gap between said electrode plates; wherein said electrode plates are sized and shaped to assure transfer of said plasma arc in the direction of said nozzle to cause a predictable amount of said fuel propellants to be ablated and create a thrust force that exits said nozzle.
- 13. The pulsed plasma thruster according to claim 12 further including an insulating layer extending substantially over said thrust discharge chamber and said nozzle.
- 14. The pulsed plasma thruster according to claim 13 further including a housing surrounding said insulating layer with openings that allow access to said electrode terminals.
- 15. The pulsed plasma thruster according to claim 12 wherein the spacing between said electrode plates is less than 50 micrometers.
- 16. The pulsed plasma thruster according to claim 12 further comprising a UV variable intensity light source predisposed to provide an ignition signal that sparks vapors within said ignition chamber.
- 17. The pulsed plasma thruster according to claim 12 further comprising a means of varying the angle of said nozzle with respect to said thrust discharge chamber.
- 18. The pulsed plasma thruster according to claim 12 wherein said power processing unit is capable of producing multiple volt-amp signal forms that effect the shape of said ignition voltage.
- 19. The pulsed plasma thruster according to claim 18 wherein said power processing system produces an ignition voltage signal in three segments corresponding to an open circuit to constant voltage segment, a constant voltage segment and a constant current segment.
- 20. The pulsed plasma thruster according to claim 12 wherein said fuel propellants comprise PTFE propellants.
- 21. The pulsed plasma thruster according to claim 12 wherein said heat generating element comprises a micro-heater capable of heating said solid to create a vapor that travels into said thrust discharge chamber and exerts a pressure on said electrode plates.
- 22. The pulsed plasma thruster according to claim 12 wherein said holes of said ignition chamber are sized and located to provide uniform feeding of vapors subliming from said solid to said discharge thruster.
- 23. The pulsed plasma thruster according to claim 12 further comprising a plurality of heating elements embedded in said fuel propellants.
- 24. The pulsed plasma thruster according to claim 23 wherein said heating elements comprise variable temperature MEMS-based micro-heaters that can control the amount of said fuel propellants ablated.
- 25. The pulsed plasma thruster according to claim 12 wherein said thrust discharge chamber is arranged to prevent either of an ignition voltage and plasma arc from interacting with said solid.
- 26. The pulsed plasma thruster according to claim 12 wherein said electrode plates are evenly spaced about a central axis extending through said thrust discharge chamber.
- 27. The pulsed plasma thruster according to claim 12 further comprising first and second electrode positioning devices predisposed about said first and second electrode plates, respectively, for varying the spacing of said electrode plates as a function of the distance to a central axis extending through said thrust discharge chamber.
- 28. The pulsed plasma thruster according to claim 13 wherein said insulating layer is shaped to increase the local field strengths existing between said electrode plates.
- 29. The pulsed plasma thruster according to claim 12 wherein said ignition voltage less than 300V.
- 30. The pulsed plasma thruster according to claim 14 wherein said housing includes a means of mounting said thruster to a mass to be propelled.
- 31. A method of operating a pulsed plasma thruster having a heat generator means predisposed adjacent a subliming solid, a thrust discharge chamber formed of two oppositely positioned electrode plates and two oppositely positioned PTFE propellants, a device separating the subliming solid from the thrust discharge chamber, a power source coupled to the electrode plates and capable of providing an ignition voltage, the method comprising the steps of:heating the subliming solid to create a high pressure vapor; directing the high pressure vapor in the direction of said thrust discharge chamber; and applying a DC ignition signal to said electrodes to spark a breakdown of the PTFE propellants and cause a transition of the spark to a useful plasma arc.
- 32. The method of operating a pulsed plasma thruster according to claim 31 wherein the step of applying a DC ignition signal is performed by controlling the shape and magnitude of the output of the power source.
- 33. The method of operating a pulsed plasma thruster according to claim 32 wherein the shape is controlled in three segments corresponding to an open circuit to constant voltage segment, a constant voltage segment and a constant current segment.
- 34. The method of operating a pulsed plasma thruster according to claim 32 wherein the step of directing the high pressure vapor in the direction of said thrust discharge chamber is performed in a manner that creates pressure between the two electrode plates of the thrust discharge chamber.
- 35. The method of operating a pulsed plasma thruster according to claim 32wherein the step of directing the high pressure vapor in the direction of said thrust discharge chamber is performed so that the vapor is fed uniformly to the thrust discharge chamber.
- 36. The method of operating a pulsed plasma thruster according to claim 32 further comprising the step of transferring an initial spark to a plasma arc in the gap defined by the separation of the electrode plates.
- 37. The method of operating a pulsed plasma thruster according to claim 36 wherein the step of transferring is performed in such a manner that an amount of PTFE propellant is ablated.
- 38. The method of operating a pulsed plasma thruster according to claim 37 wherein the amount of PTFE propellant is ablated in a controlled manner.
- 39. The method of operating a pulsed plasma thruster according to claim 32 wherein the step of applying a DC ignition signal to said electrodes is performed by focusing an ultraviolet light source on the high vapor pressure.
- 40. The method of operating a pulsed plasma thruster according to claim 32 further comprising the step of adjusting the spacing of the electrode plates to effect the force delivered by the propellants.
- 41. The method of operating a pulsed plasma thruster according to claim 32 wherein the step of applying a DC ignition signal to said electrodes is performed by limiting the DC voltage signal to less than 300 volts.
US Referenced Citations (7)