Power plants – Reaction motor – Ion motor
Reexamination Certificate
2000-02-04
2001-02-06
Kim, Ted (Department: 3746)
Power plants
Reaction motor
Ion motor
C060S203100, C060S204000
Reexamination Certificate
active
06182441
ABSTRACT:
BACKGROUND OF THE INVENTION
This invention relates to electric drive circuits for powering spacecraft thrust engines including arc-jet engines and ion engines and, more particularly, to a drive circuit adapted for employing voltages supplied by solar panels of a spacecraft for application directly to the electrodes of the engine.
The engines which drive a spacecraft are in the nature of a thruster emitting a stream of high-velocity particles which accelerate the spacecraft in a desired direction as may be required for adjustment of the spacecraft position or orientation during station keeping, as well as for adjustment of the spacecraft orbit. An arc-jet engine and an ion engine are examples of typical thrusters in common use.
Both the arc-jet and the ion engines employ a gas which is electrically charged by an input power source to produce a plasma of high speed electric particles. In the arc-jet engine, the intense heat of the plasma results in an acceleration of particles out of the mouth of the thruster with a corresponding force being developed against the spacecraft to drive the spacecraft in a desired direction. In the ion engine, an electrode grid structure applies an accelerating voltage to the charged particles to accelerate the particles out of the mouth of the thruster to develop the force which drives the spacecraft in the desired direction.
The operation of such thrust engines requires both high voltage and a relatively low voltage. The high voltage is required to initiate electric conduction through the gas of the engine to produce the plasma. Thereafter, electrical conduction through the plasma can be maintained at the low voltage. The solar panels of a spacecraft are employed for generating the electric power necessary for operation of the thrust engine.
A problem arises in that the voltage outputted by a solar panel is too low for operation of the thrust engine. In the past, this problem has been addressed by interposing a DC-to-DC converter between the power source of the solar panels and the thrust engine to increase the voltage, at which the direct current (DC) is supplied by the solar panels, to a higher voltage suitable for operation of the thrust engine. This has the disadvantage of increasing the weight and the complexity of the electrical equipment carried by the spacecraft in contradiction to the general principle of minimizing the weight and the complexity of spacecraft.
SUMMARY OF THE INVENTION
The foregoing problem is overcome and other advantages are provided, in accordance with the invention, by a spacecraft power system for powering an engine of a spacecraft, wherein the engine is a thrust engine including an arc jet engine or an ion engine. The thrust engine is operated electrically and has a start-up voltage threshold and a running voltage substantially less than the start-up voltage. The spacecraft carries an array of solar panels for applying electric power for operation of the thrust engine.
In accordance with a feature of the invention, in the array of solar panels, at least a plurality of the solar panels are connected by series connection to output a voltage equal to the sum of voltages provided by the series connection of said solar panels. The sum of the voltages is less than the start-up voltage but sufficient to maintain steady-state operation at the running voltage of the thrust engine.
The electric circuitry which drives the thrust engine includes the foregoing series connection of the solar panels plus a start-up circuit which is connected between the array of solar panels and the thrust engine to provide the star-up voltage. The start-up circuit has input terminals for connection across terminals of the array of solar panels and output terminals for connection with the thrust engine. Also included in the start-up circuit is a switch which, upon an opening and a closing of the switch, temporarily boosts the sum of voltages beyond the start-up voltage threshold and allows the sum of the voltages to pass directly from the array of solar panels to the output terminals of the start-up circuit for activation of the thrust engine.
In accordance with the invention, the electric switch is placed in parallel with an electric load presented by the electrode assembly and plasma of the engine. The parallel combination of the switch and the load is connected via an inductor serially between first and second terminals of the input terminal pair of the drive circuit. This enables current to flow from the array of solar panels to the output terminals of the start-up circuit. Preferably, the switch is a semiconductor device such as a transistor which is operated by a pulse generator which places the switch in alternate states of conduction and nonconduction.
Initially, prior to generation of plasma in the engine in the operating procedure, the switch is closed by the pulse generator to be in a stage of conduction, essentially a short circuit, for conducting current through the inductor via the switch. Since there is not yet any plasma, the engine load appears essentially as an open circuit and draws no current. After initiation of the inductor current, the switch is opened by the pulse generator to discontinue conduction of the inductor current. The inductor then generates sufficient voltage to drive the inductor current into the engine load and strike an arc for generating the plasma. The plasma reduces the voltage drop across the load to a value which can be sustained by an array of solar panels. This enables the array of solar panels to continue to supply current via the inductor to the load to maintain the engine thrust.
The plurality of solar panels may be connected in a series-parallel array to output a sufficient magnitude of voltage to sustain the state of plasma in the engine, while providing sufficient current to power other circuits of the spacecraft. A capacitor is connected across the input terminal pair of the start-up circuit to protect the solar panels, as well as other spacecraft circuitry from any voltage spikes which may develop during initiation and termination of the state of plasma in the engine. The inductor also serves, in combination with the capacitor to filter out any electrical noise generated within the engine load. A further switch is connected between the inductor and the array of solar panels to deactivate the start-up circuit upon completion of an engine thrust interval.
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patent: 4996407 (1991-02-01), Traxler
patent: 5142861 (1992-09-01), Schlicher et al.
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patent: 5359180 (1994-10-01), Park et al.
patent: 5605039 (1997-02-01), Meyer et al.
patent: 5626315 (1997-05-01), Flament et al.
patent: 5720452 (1998-02-01), Mutschler, Jr.
patent: 6029438 (2000-02-01), Hosick
Kim Ted
Perman & Green LLP
Space Systems Loral, Inc.
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