Power plants – Reaction motor – Interrelated reaction motors
Reexamination Certificate
2000-11-10
2002-04-30
Kim, Ted (Department: 3746)
Power plants
Reaction motor
Interrelated reaction motors
C060S225000, C060S250000, C060S257000, C060S039823, C244S169000
Reexamination Certificate
active
06378292
ABSTRACT:
FIELD OF THE INVENTION
The present invention is related generally to micro electrical mechanical systems (MEMS). More specifically, the present invention is related to MEMS based microthruster arrays which find one application in satellite propulsion.
BACKGROUND OF THE INVENTION
Satellites orbiting the earth range in size from Sputnik to the Hubble Space Telescope and beyond. Attitude maintenance control has typically been carried out using conventional attitude control thrusters, for example, controllable compressed gas release thrusters. There has been interest in replacing single, large satellites with clusters of small satellites. Each satellite in a cluster may be as small as a deck of playing cards, but collectively the cluster could function as a single satellite having the diameter of the cluster. Building a cluster of small satellites may be cheaper and more versatile than building a single large satellite. In particular, one application may include arranging a number of small satellites as a sparse aperture radio wave antenna for imaging distant objects. Sparsely positioned small satellites may also function as different elements of an interferometer having a large aperture.
One difficulty with using numerous small satellites is controlling the inter satellite distance accurately. The satellite-to-satellite distance is preferably maintained within limits for many applications. Small satellites may require a very small rocket motor for maintaining the inter satellite distance.
Very small rocket motors have been difficult to design and/or manufacture. One proposed design uses cold gas thrusters, which can operate by opening gas valves intermittently. This requires a moderate amount of hardware, tubing, and valves, which are often difficult to scale down and do not scale down far enough to be of use in very small satellites. Digital propulsion rocket chips have been proposed and<e prototypes designed. Some prototypes have arrays of individually addressable explosive pixels. Existing designs have drawbacks. One current problem is thermal fratricide.
Thermal fratricide presents at least two problems. First, an individually addressable and ignitable explosive element in an array requires ignition. Ignition typically requires heating an igniter element to a temperature sufficient to cause an explosion of a material disposed close to the igniter. As the igniter temperature increases, heat may be dissipated away from the igniter element, preventing the element from ever approaching a temperature sufficient to cause the explosion or combustion of a propellant or explosive. Many times, the igniter does generate sufficient heat, which is retained, causing the propellant to explode or vaporize. When the temperature is sufficiently high, the heat may be conducted into adjacent individually selectable and ignitable elements, causing them to explode as well. This thermal fratricide thus can cause the explosion of pixels adjacent to the pixel for which the only explosion is desired.
What would be desirable is a microthruster array which provides for individually addressable and selectable microthrusters that do not cause the unwanted ignition of adjacent microthruster cells.
SUMMARY OF THE INVENTION
The present invention provides a MEMS microthruster including a plurality of propulsion cells, wherein the propulsion cells can be deployed in an array, each cell being individually addressable and ignitable. In one embodiment, each propulsion cell has a first cavity having an explosive igniter disposed, and preferably suspended within the first cavity, and a second cavity separated from the first cavity by a first diaphragm. The first diaphragm is preferably cooperatively dimensioned together with the igniter to break the diaphragm after the explosion of the explosive igniter. In one embodiment, a propellant is disposed within the second cavity, where the second cavity is disposed on the opposite side of the first diaphragm from the first cavity. Upon the breakage of the first diaphragm, the propellant can expand rapidly in response to the igniter exploding through the first diaphragm, thereby causing the rapid expansion of the propellant within, and ultimately out of, the second cavity. In. a preferred microthruster cell, the explosive igniter is suspended within the first cavity, and substantially surrounded by a thermal insulator such as a vacuum. The individual propulsion cells are preferably individually selectable and ignitable.
In one embodiment, the propellant is provided as a single component disposed within the second cavity, and can be a plastic explosive such as a nitrocellulose or nitrocellulose acetate. In another embodiment, the propellant is provided as two components. The first propellant component can be a fuel and the second propellant component can be an oxidizer. The second cavity in this embodiment can be divided into a first portion and a second portion, separated therebetween by a second diaphragm. The first diaphragm can be broken by the explosive igniter which also breaks the second diaphragm, thereby causing mixing of the fuel and oxidizer. The fuel and oxidizer then can generate sufficient force to break a third diaphragm disposed toward the exterior of the microthruster, thereby allowing the propellant exhaust gas to be ejected from the microthruster cell.
In one illustrative embodiment, the first set of cavities is formed upon a silicon substrate. The silicon substrate may have supporting electronics formed in the top surface thereof. A silicon dioxide layer may then be grown on the top of the wafer and supporting electronics. To produce a suspended igniter element, a cavity may be etched into the silicon dioxide layer, and filled with a polymer or other sacrificial layer. A silicon nitride layer or the like may then be deposited over the silicon dioxide and sacrificial layer, and an electrically resistive layer may be put on top of the silicon nitride. The sacrificial layer may then be removed, leaving a suspended igniter structure. The igniter element is preferably a heatable metal resistor coated or otherwise coupled to an explosive compound.
In a preferred embodiment, the first cavity is filled with a thermal insulator, such as a vacuum. The first set of cavities can have a first diaphragm forming the ceiling of a cavity, which, in one embodiment, is formed of silicon nitride. The second set of the cavities can be formed on top of the silicon nitride diaphragm, and may have walls formed of silicon dioxide. In one embodiment, the top of the second cavity is open, and serves as an opening through which an explosive compound is packed into the second cavity. In another embodiment, a plastic explosive is poured into the second cavity through the top orifice and cured. In yet another embodiment, the second set of cavities is further capped by yet another silicon nitride diaphragm.
In one microthruster embodiment, the second cavity has a ceiling or second diaphragm separating the second cavity from a third cavity, or, alternatively, dividing the second cavity into a first portion and a second portion. In this arrangement, the first portion may be filled with a fuel and the second portion may be filled with an oxidizer. The explosion of the igniter can break both the first diaphragm and the second diaphragm, thereby allowing mixing of the fuel and oxidizer, which causes rapid expansion of the propellant formed by the combined oxidizer and fuel. In one embodiment, the top of the third cavity is capped by a silicon nitride third diaphragm.
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http://www.spie.org/web/oer/august/aug98/sunny.html, “Pyrotechnic Materials Ingetrate with Micromachined Silicon to Produce Space Probes”, dated Aug. 1998, 3 pages.
http://www.design.caltech.edu/micropropulsion/sunday-times.html, “Jet Engines Smaller Than Penny Will Prop
Fredrick Kris T.
Honeywell International , Inc.
Kim Ted
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