Power plants – Reaction motor – Electric – nuclear – or radiated energy fluid heating means
Reexamination Certificate
1999-05-19
2001-04-17
Thorpe, Timothy S. (Department: 3746)
Power plants
Reaction motor
Electric, nuclear, or radiated energy fluid heating means
C060S202000
Reexamination Certificate
active
06216445
ABSTRACT:
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 s
Byers David C.
Lewis, Jr. David H.
Rodriguez William
Thorpe Timothy S.
TRW Inc.
Yatsko Michael S.
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