Patent Application
20110233343
The invention relates to satellite systems. In particular it relates to position control such as attitude control of a satellite.
For purposes of this invention the three dimensions of movement of a satellite will be referred to in this application as attitude, whether it includes pitch, yaw or role of the satellite. Traditionally attitude control for satellites has been achieved through the use of rockets generating a thrusting force thereby adjusting the attitude of the satellite by Newton's third law of motion: For every action there is an equal and opposite reaction. In other words by exerting the force of thruster rockets in one direction, the rocket is caused to accelerate in the opposite direction. The use of rockets, however, severely limits the accuracy with which the attitude can be adjusted. Typically, liquid fuel rockets are used for this purpose in which nozzles emit the rocket fuel and can be opened or shut off. While this allows the duration of the force to be roughly adjusted it does not allow for either fine control of the duration, or control of the amount of force generated by the rockets.
A monopropellant liquid fuel thruster relying on valve control typically relies on a monopropellant such as hydrazine or hydrogen peroxide that has to be contained and then selectively released into a catalyst containing decomposition chamber. These thrusters have the disadvantage that they are subject to engine wear; make use of high pressure poppet valves with limited cycle life; occupy a significant volume from aluminum/titanium fuel tanks, high pressure valves, and pipes; and provide limited resolution (several thousandths of a second fuel pulse) being limited by the speed of mechanical poppet valves.
For example, in one micro-chemical thruster for micro satellites based on a hydrogen peroxide thruster the finest resolution was limited to bursts of 80 μNs pulses. In a 20 cm×10 cm nano satellite with a mass of 5 kg this would create an end-over-end rotation speed in excess of 4 degrees/second making it impossible to finely adjust the attitude.
Another prior art attitude adjustment device is the Momentum Wheel or Reaction Wheel, which is based on an electric motor spinning a flywheel to achieve reactive angular momentum. While it can provide small angular adjustments, it cannot provide translational movement. It also tends to build up stored momentum that needs to be canceled, requiring supplemental attitude control systems.
Yet another attitude control system is the Control Moment Gyroscopes in which rotors are mounted on gimbals and spun at a constant speed. This provides a higher torque capability than a Reaction Wheels but is also more costly and heavier. Its complexity also makes the Control Moment Gyroscope more prone to failure (for instance, the International Space Station uses a set of four CMGs to provide dual failure tolerance). Yet another attitude control system involves the use of solar sails, which use the reaction force created by incident solar radiation allowing attitude and translational movement. Solar sales are beneficial in eliminating fuel and useful on long missions but are not suited to maintaining geostationary orbits.
A light, small and highly accurate micro-adjustment position controller for satellites would therefore be highly desirable to adjust for satellite drift or to position the satellite to perform a particular task, or simply to fine-tune attitude adjustments as provided by rockets or other systems currently used in the art of satellite attitude control
According to the invention there is provided a satellite control system operable in a low pressure environment, comprising at least one substrate structure having a distal surface and a proximal surface with multiple holes extending at least partially into the substrate structure from the proximal surface, at least one heating element arranged at the bottom of the holes or at a predefined distance from the proximal surface, a liquid that is thermally ejectable from the holes by the at least one heating element, and at least one valve or cover, for selectively sealing the liquid from the low pressure environment. The liquid may be a non-volatile liquid. The system may include an electrical circuit that includes at least one controllable switch for controlling current flow to the at least one heating element. The control system may include a liquid supporting reservoir in flow communication with the holes in the substrate structure. The liquid is preferably a high-density liquid such as mercury. The liquid may also contain particulate matter. The at least one valve may comprise at least one micro-valve. The liquid may include ferrous particles and the at least on micro-valve may include a non-mechanical ferro-fluid valve. At least one of the micro-valves may comprise a piezoelectrically actuated micro-valve. The system may include pressure exerting means for exerting pressure on the liquid in the reservoir. The pressure exerting means may comprise a flexible balloon arrangement or plunger arrangement using a gas under pressure, e.g., nitrogen. The pressure is preferably controlled so as to limit the volume liquid flow rate into each hole due to the pressure differential and capillary action to one pre-defined ejection volume between ejections. The plunger or balloon arrangement may be connected directly or indirectly in flow communication with the reservoir. The substrate structure is preferably made from a material having a high operating temperature and low coefficient of thermal expansion and providing high thermal conductivity, a high heat capacity and a high thermal shock parameter, e.g. silicon carbide or any of its poly types (different atomic arrangements). These may for example include atomic arrangements such as cubic (4C), hexagonal (4H and 6H), or rhombohedral crystal lattice arrangements. The holes formed in the substrate structure may have a ratio of diameter to the depth of the heating elements from the proximal surface of between 1 to 1 and 1 to 10. The holes may be 74 μm in diameter with silicon carbide streets between the holes that are for example 12 micrometers wide to provide a center to center distance between the holes of 86 μm.
The substrate structure and electric circuit may be implemented as a MEMS device (micro electromechanical system).
The control system may, further include a processor or controller for determining which holes, the number of holes, and/or the number of firings for such holes that is required for a particular attitude adjustment of the satellite in a particular time period. The control system may further include a radio receiver for providing signals to the processor defining an attitude adjustment or desired orientation. Preferably each hole is provided with a separate heating element formed at or near the distal end of each hole or around each hole and defining part of an electrical circuit that includes at least one switch. All of the switches may be controlled by the processor or controller.
Further, according to the invention, there is provided a method of controlling the position of a satellite, comprising ejecting a non-volatile liquid from a channel by thermal ejection. Typically the position control comprises an attitude adjustment of the satellite. The method may include ejecting from multiple channels. The channels may comprise holes formed in at least one substrate structure, e.g. a SiC substrate structure. The liquid may comprise a high density liquid such as mercury. The ejection of the liquid may be controlled by a processor. The holes may be pre-filled with the non-volatile liquid or filled prior to ejection. The holes may be filled from a reservoir and may be refilled one or more times after liquid has been ejected from the holes. The processor may control which holes to eject from, and the number of holes from which to eject, and may define the duty cycle of the ejections if any hole is required to eject liquid more than once. The holes may be formed by MEMS technology in a SiC substrate, the method comprising ejecting the non-volatile liquid from the holes in the substrate structure.
The present invention proposes a method and a means for generating controlled, small amounts of thrust in defined directions. Thus the invention relies on Newton's third law of motion: “For every action there is an equal and opposite reaction” to turn a satellite in space by generating a thrust in an opposite direction. This is achieved by thermally ejecting small amounts of liquid such as mercury, from one or more channels, in defined directions in a controlled manner. Due to conservation of momentum the momentum of the liquid droplet that is ejected (mass×velocity of the droplet) is reflected as an opposite momentum of the satellite. Thus, although the volume of the droplet is rather small, the combination of a high density liquid and a substantial velocity with which the liquid droplet is expelled translates into an appreciable momentum for the satellite.
In one embodiment, a semiconductor substrate structure is formed e.g. by MEMS technology as shown in
MEMS (Micro Electromechanical Systems), also referred to as micro machines or micro systems technology, is a modified semiconductor device fabrication process that makes use of molding, plating, wet etching (KOH, TMAH) and dry etch (RIE and DRIE) and electro discharge machining (EDM) techniques to produce systems on a substrate in the micrometer range (typically 1-900 μm). While thin films can be thinner than 1 micron, in practice the structures that have a mechanical function need a minimum mass, and a minimum area.
The upper end of the thickness range is determined by the thickness of standard wafers from which the die are made. MEMS fabrication lines make use of wafers that are up to 200 mm in diameter (referred to as 8″ wafers), having a thickness of 675 micron. Some MEMS substrate structures have, in the past, made use of die from polished glass wafers at a thickness of as much as 1 mm, which typically marks the upper limit of what is commonly called Micro Technology.
The fabrication of devices using MEMS technology typically involves the deposition of layers of materials, patterning of the layers by photolithography, followed by etching.
As indicated above, MEMS devices can be manufactured from a variety of material. Probably the most popular material for MEMS devices is silicon due to its inherent ability to incorporate electronic functionality. In its mono-crystalline form it displays almost no hysteresis when flexed, and thus virtually no energy dissipation. Also, unlike most metals it suffers virtually no fatigue when repeatedly stressed. Polymers can also be used in these processes and are suited to injection molding, embossing, and stereolithography. As alluded to above, metals can also be used but have physical limitations. On the other hand metals can be deposited by electroplating, evaporation, and sputtering processes. Commonly used metals include gold, nickel, aluminum, copper, chromium, titanium, tungsten, platinum, and silver.
MEMS devices may be provided with a central unit or microprocessor that communicates with peripheral units such as micro-sensors.
The present invention, in a preferred embodiment, makes use of MEMS technology to produce attitude control devices in accordance with the present invention. The embodiment of
Bulk micromachining is the oldest technique for forming silicon based MEMS. In this approach the whole thickness of a substrate structure, e.g., a silicon substrate structure, is used for building the micro-mechanical structures. The silicon is machined using various etching processes, and additional structures, e.g., silicon structures or glass plates are added by anodic bonding to create features in the third dimension for hermetic encapsulation. This technique is also used for high performance pressure sensors and accelerometers.
Another technique involves surface micromachining which uses sacrificial layers deposited on the surface of a substrate as the structural materials, rather than using the substrate itself. Surface micromachining was created in the late 1980s to render micromachining of silicon more compatible with planar integrated circuit technology, with the goal of combining MEMS and integrated circuits on the same silicon.
Currently both bulk micromachining (of the order of 10-900 micron thick structures) and surface silicon micromachining (of the order of 1 micron thick structures) are used in the industrial production of sensors, ink-jet nozzles, and other devices. However, in many cases the distinction between these two processing techniques has diminished. A new etching technology, deep reactive-ion etching, has made it possible to combine good performance typical of bulk micromachining with comb structures and in-plane operation typical of surface micromachining. Reactive-ion etch is a form of high aspect ratio (HAR) silicon micromachining. In HAR silicon micromachining the aspect ratios are of the order of 1:10 to 1:100, which are a function of the precision of the physical effect, and of the chemical nature of the process. Reactive Ion Etching (RIE) can thus produce side wall angles better than 6 degrees [=tan(0.1)].
However, the critical dimensions of a structure are not limited only by the theoretical parameters that a process is capable of but are dictated also by transport phenomena, e.g. bringing etched material out of the hole.
As indicated above, MEMS technology borrows many of the process techniques used in semiconductor manufacturing. However, the consensus of the industry currently favors separate manufacturing of the mechanical and electronic components, with the option of subsequently bonding the two structures to one another. The flexibility and reduced process complexity obtained by having the two functions separated currently outweighs the small penalty in packaging.
In the embodiment shown in
The heating of the liquid for thermal ejection is achieved by one or more heating elements, which in this embodiment is formed around each hole. The heating elements 108, in this embodiment, are shown at the bottom of the channels on the distal surface. However, in another embodiment the heating elements are located at a predefined depth from the proximal surface, thereby defining the droplet size as defined by the depth of the heating element and the area of the hole or channels 102. In one such embodiment the heating elements are formed by depositing SiN rings and doping the SiN to form integrated high resistance structures. Intervening wafer material, referred to herein as roads 110, are formed between the channels 102.
In the above embodiment the holes or channels extend all the way through the wafer and the heating elements are formed around the holes. However, other configurations can be used e.g., channels extending into the substrate material or wafer from a proximal surface of the wafer to a predefined depth, with heating elements at or near the bottom of the hole. The device may be designed to be fired once only from each of the holes (either individually or simultaneously, but most commonly by calculating the number of holes that have to be fired in order to achieve a desired attitude adjustment of the satellite). The devices may also be designed to be refillable, for multiple ejections or firings from each hole. In the embodiment of
While much of MEMS technology is based on fabrication using silicon, the present invention proposes the use of materials having a much higher operating temperature and lower coefficient of thermal expansion and providing high thermal conductivity, a very low heat capacity and a high thermal shock parameter. In particular the present application proposes the use of silicon carbide or any of its poly types (different atomic arrangements). In the present embodiment the wafer is made from silicon carbide having a 6H crystal lattice configuration.
Silicon carbide (SiC), also known as carborundum, is a compound of silicon and carbon with chemical formula SiC. The grains of silicon carbide can be bonded together by sintering to form very hard ceramic plates, and SiC is widely used in high-temperature/high-voltage semiconductor electronics. As mentioned above, while SiC always involves a combination of silicon and carbon, the crystal lattice structure may vary and includes structures such as 3C (cubic) atomic arrangements with the atoms located at the corners of cubes forming a lattice structure, or a hexagonal (4H or 6H) arrangement that repeats every four or six layers, or a rhombohedral arrangement. A comparison of the arrangements and properties of 3C, 4H and 6 H are given in the table below.
While silicon is used in one embodiment of the invention, SiC offers several benefits that make it the preferred material for the MEMS substrate material.
SiC is in many ways more robust than silicon, both thermally and mechanically:
A. Thermally
SiC offers a higher thermal shock parameter resulting in slower development of crystalline fault formation, macroscopic cracking and migration pit formation. This allows it to withstand higher temperature cycling, allowing it to provide for greater mass ejection speed of the liquid.
SiC has approximately 3 times higher thermal conductivity (depending on the crystal lattice arrangement of the SiC), and about a 16 times lower thermal capacitance, than silicon. This provides for rapid heat dissipation after ejection, ensuring greater control over the duty cycle (rate of firing). It also provides improved thermal spreading to achieve greater thermal uniformity for maintaining the device at the desired viscosity temperature for the liquid being ejected (typically liquids display different viscosities at different temperatures and thus the ease with which it ejects from a channel is a function of its temperature).
B. Mechanically
The physical robustness of SiC ensures slower development of crystalline fault formation, macroscopic cracking and migration pit formation than Si, and thus suffers less degradation due to the outward pressure-pulse shock fronts, and less side-wall erosion due to the outward flow that follows, and perpendicular turbulence resulting from the thermal ejection process. Since SiC suffers less pitting, it is less vulnerable than silicon to the formation of local bubble nucleation sites within the initial vaporization disk (one result of which is increased turbulence), and will also develop less additional wall interface friction, and consequently less age degradation of the ejection speed and colinearity. The physical robustness of SiC also provides additional resilience against ballistic damage from micro-meteorites.
Other embodiments of the invention makes use of SiN, AlN (Aluminum nitride), GaN, AlGaN, GaAs, or other single or poly crystalline materials, with a preference for single crystalline material to form the substrate structure. Control devices of the invention may include combinations of materials, e.g., Si, SiC, SiN. For example, the substrate structure that supports the ejection holes may comprise SiC while any other elements of the control system, e.g., covers, lids, reservoirs, micro-valves, could be made from a different material such as Si or SiN. Even an individual element of the control device may comprise more than one material, e.g., SiC or SiN can be epitaxially grown on Si, or SiN; or SiC can be epitaxially grown on SiC to form the substrate structure or one of the other elements of the control device.
In one embodiment AlN was grown on SiC to form substrate material for the substrate structure. In another embodiment SiC was grown on Si and in yet another embodiment SiC was grown on SiC to form the substrate material of the substrate structure.
Different embodiments were tried with holes ranging in diameter from 30-100 um and with the depth of the fluid column being of the order of 50 um to 100 um. Embodiments are not however limited to these hole configurations. In one embodiment, discussed in greater detail below, the holes formed in the wafer have a radius of 37 μm and are formed in a wafer that has been micro-machined to a thickness of 74 μm. The holes are formed with intervening streets of 12 μm for a center-to-center distance of 86 μm.
In order to determine the ideal hole aspect ratio for any particular liquid, empirical data is required for the various liquid parameters (density, viscosity, latent heat, and boiling point) and the power supply available. For example, for water based substances such as ink, it has been found that holes with aspect ratios of 1:1 to 1:3 (channel diameter to channel depth) work well in thermally propelling the liquid from the holes.
It will be appreciated that viscosity and channel diameter will determine refill rates due to capillary attraction and differential pressure across the column of liquid. Once the liquid is in the channel, its specific gravity and viscosity (which defines how easily the fluid flows in the channel) will determine the force needed to eject a droplet of a particular size (due to the varying mass with varying specific gravity). The requisite force generated is, in turn, related to the power source that is available and the resistivity of the heating element since the power P dissipated in a resistive element is related to the resistance R of the resistive element and the current I flowing through it, according to the formula P=I2R. However the amount of heat energy required in order to expel a droplet of liquid by thermal ejection depends also on the rate with which the liquid can be brought to its vapor phase. For example, even though mercury has a boiling point more than 3 times that of water, its latent heat is only about 1/10 of that of water and therefore heats up to phase transition far more rapidly than water, making it an ideal candidate for purposes of the present invention. Also, mercury has a viscosity that is only about 1/10 of that of water, making it far easier to eject from a channel. The benefits of mercury over water as a non-volatile ejection liquid may be summarized as follows:
Considering again the embodiment with holes having a diameter of 74 um and a depth of 74 um, and assuming that droplets of thrust-producing liquid e.g. mercury are emitted from each of the holes using heating elements located at the distal ends of the holes, the volume of material in each hole will be JI r2×t=JI (372×74)×10−18=3.18×10−13 m3=318 pl (pico liter). The amount of area (hole and surrounding street area) for each unit or hole is thus (37+12+37)2 μm2=7.396×10−9 m2. Thus in a wafer of 6 inch×6 inch=6.45×3.6×10−3 m2 this provides for a total of 4.867×105=3,140,000 holes for a total Mercury volume in the holes of approximately 10−6 m3=1000 μl=1 ml.
It will be appreciated that droplet volume will vary depending on the hole diameter and length, and will depend also on the surface tension, density, and viscosity of the liquid being ejected. By way of example, for hole diameters chosen to be equal to the hole length (ratio of 1:1), hole diameters of 5 μm, 15 μm, and 38 μm, provide droplets with a volume of 0.1 picoliter (one millionth of a microliter), 2.7 picoliter and 44 picoliters, respectively. The choice of hole depth (or location of the heating element from the proximal surface of the wafer) and hole diameter will depend on the density and kinematic viscosity of the liquid. As the length of liquid that is propelled out of the hole increases, the mass increases for the same amount of force exerted by the heating element thereby reducing the velocity of the propelled droplet. On the other hand the hole cannot be made too wide since the propulsion of the liquid requires vaporization of a disk of material beneath the liquid that is to be expelled.
Modeling software has been developed from empirical data for the ejection of ink droplets.
Similar software models can be developed for the various other liquids contemplated for the attitude control system, based on empirical data using different hole widths and depths, different hole aspect ratios (side-wall slope), and heating element resistances for the power supplies available in the various satellites in which the device is to be implemented.
One embodiment of the invention makes use of a capacitor to charge up from a low voltage power supply. The low internal resistance of the capacitor allows a large current to be released to the heating elements for a short period of time as the holes are fired (i.e., when the liquid in the holes is to be thermally ejected). For a capacitor, Q=CV where Q is charge, C is capacitance and V is voltage potential of the charged capacitor. The capacitor allows a current I to be discharged over a period t according to the equation I=Q/t
It will be appreciated that larger heating elements and larger energy sources can be used insofar as larger energy sources are available. In one embodiment a nichrome resistor was chose as the heating element and a voltage of 12V was applied to the nichrome wire as energy source. Nichrome at 38 gauge (0.004 inch diameter) has a resistance of 42.2 Ω/ft. In practice a heating element could be deposited with much smaller dimensions and correspondingly higher linear resistance. In the case of mercury, in order to avoid the mercury reacting with the heating element a substance that will not form an alloy with mercury is desirable, such as Tungsten or silicon nitride doped to provide the desired conductivity.
As will be discussed in greater detail below the mercury or other liquid in the holes may be replenished from a reservoir to permit multiple firings from each hole.
In the above embodiment a hole diameter of 74 μm was chosen, which provides for reasonably large drops of liquid. Such larger holes are particularly suited to the use of mercury as the liquid to be expelled. The high density (specific gravity of 13.57) allow for small, heavy droplets. Also, the high surface tension of mercury requires that the hole size cannot be too small. As is mentioned above, the additional benefit of mercury is that its viscosity is much lower than water, and at about 0.11 centistokes makes it much more slippery and easier to expel. Thus, as will become clearer from the discussion below the use of large holes requires more energy in order to expel or shoot out the droplets, however this is helped by the low wetting coefficient of mercury allowing it more easily to be released from surfaces that it is in contact with. The use of larger holes has the advantage of larger drops with higher mass and therefore higher momentum. In order to ensure that the velocity of ejection is not too severely curtailed by the larger mass, a higher resistance heating element and a larger power source (or the addition of a capacitor as discussed above) may be provided to achieve greater heat dissipation into the liquid. Thus mercury offers some clear benefits in the device of the invention that seeks to produce thrust in order to achieve a reactionary force for purposes of attitude adjustment. As discussed above, by the conservation of momentum, the larger the mass and velocity of the emitted droplets, the larger the momentum and consequently the larger the velocity with which the satellite will be propelled in the opposite direction.
It is proposed that droplet sizes of the order of 50 to 300 pl be expelled from a substrate structure with ratio of hole diameter to substrate structure thickness of, for example, 1:1 to 1:10. Some embodiments of the control system may include die with different hole sizes or the matrix of die can have different hole sizes, with each die dedicated to a particular hole size. For simplicity the embodiments discussed below show substrate structures with only a few holes, each of the same depth and diameter.
A cross-section through one such embodiment is shown in
The use of the device in the outer space environment brings with it additional difficulties, apart from the near vacuum environment, which seeks to suck the liquid out of the channels.
The low pressure also causes liquids exposed to the environment (for example the mercury in the channels to evaporate. Since a satellite may have a life-span of the order of 30 years it is therefore desirable to seal the mercury in its holes until it is ready to be fired or seal the mercury in a separate reservoir or chamber prior to filling the channels or holes. The sealing may be achieved by providing a plug over the hole openings or providing a layer of SiC over the proximal surface to seal in the liquid.
It will be appreciated that such an embodiment will be useful only where all of the mercury is to be fired once the plug or cover is removed e.g., all holes are single fire holes or are fired repeatedly until the mercury is depleted. Insofar as each hole is provided with its own plug, this approach allows the holes to be fired individually or in groups as needed. One such embodiment is shown in
In another embodiment, shown in
It will also be appreciated that in the embodiment of
In a preferred set of embodiments, micro-valves are used to seal off the liquid from the outer space environment. The micro-valves, in one embodiment seal off the liquid in a separate chamber or reservoir shortly before filling the firing holes or ejection channels, e.g., less than a minute prior to ejection (firing) of liquid from the channels, to minimize loss of mercury due to evaporation. Since the vacuum of outer space will tend to suck the liquid out of the channels, the differential pressure across the liquid in the channels has to be controlled and preferably is controlled to ensure that the ejection takes place once the channel has been filled. Thus the differential pressure and corresponding speed with which the channels or holes are filled has to be controlled to allow a processor to time the ejections of the fluid from the channels.
Several such micro-valves have been developed in the art and some of these are discussed below with respect to
Micro-valves can be categorized as active (in which their open/closed configuration is manipulated by an external actuator) or can be categorized as passive.
The micro-valves can best be considered as falling into three groups: (a) mechanical, (b) non-mechanical and (c) external, based on their actuation mechanism.
(a) Mechanical micro-valves are typically surface or bulk micro-machined MEMS devices with a mechanically moveable membrane or micro-ball that is coupled to a magnetic, electric, piezoelectric or thermal actuation mechanism. An example of a magnetic actuation mechanism is a solenoid plunger.
(b) Non-mechanical micro-valves are actuated by virtue of their smart materials, e.g., phase changing or rheological.
(c) External micro-valves are actuated by external modular or pneumatic means.
The covers, shutters or mechanical micro-valves used for sealing the liquid from the low pressure environment may be formed using MEMS technology and may be made of Si, SiC, SiN, or other suitable materials, and may be bonded to the substrate structure.
An example of a non-mechanical micro-valve is an electromechanical valve in which a flexible membrane is deflected by generating oxygen gas by electrolysis in a chamber bounded by the membrane.
In another non-mechanical micro-valve a smart hydrogel volume is changed by inputting a change in pH, temperature, electric field, light, carbohydrate, antigen, or glucose.
In yet another non-mechanical micro-valve the phase change nature of a paraffin material is used to create a reversible or irreversible seal by melting away a plug blocking a channel. It can be used together with external air or vacuum system to make the seal reversible, i.e., turn it to liquid and remove it as a plug and then use external air or vacuum to move the paraffin material back into place and change it back to a solid. The phase transition may be activated by thermal heating. The advantage of this micro-valve was that even at a pressure of 1725 kPa no leakage was detected over a 15 minute period, thus making it good candidate as a seal or at least as supplemental seal together with a main seal such as a piezoelectrically actuated mechanical seal.
A similar micro-valve to the above paraffin micro-valve is an electro-rheological fluid which changes viscosity under the influence of electric fields. Another non-mechanical micro-valve that could be used with mercury is a ferro-fluid type device in which ferromagnetic particles of 10 nm size are suspended in a carrier fluid, e.g., mercury (since mercury does not react with iron). Two embodiments of such a micro-valve are shown in
The choice of closure member for the device may therefore vary depending on the type of liquid, the hole and reservoir arrangement, and the duration for which the device is to be used. For example some satellite adjustments or re-positioning may be severe and require a large burst, while others may require only fine-tuning or micro-adjustments to the satellite's attitude. Also the lifespan of satellites may vary, which affects the amount of attitude adjustments over its lifetime and the amount of evaporation of the liquid.
In one embodiment, shown in
In the above embodiment, in which mercury is used, with its low viscosity, it will be appreciated that the pressure exerted on the mercury cannot be too high in order to avoid it being squeezed out of the holes at too high a rate (a rate that exceeds the duty cycle or firing rate of the holes) as a result of the near vacuum conditions on the proximal end of the holes.
In an atmospheric environment, the height of liquid in a tube due to capillarity or capillary attraction can be expressed as
h=2σ cos θ/(ρgr)
where:
h=height of liquid (ft, m)
σ=surface tension (lb/ft, N/m) (which for mercury is 0.465 N/m)
θ=contact angle
ρ=density of liquid (lb/ft3, kg/m3) (which for mercury is 13.5×103 kg/m3)
g=acceleration due to gravity (32.174 ft/s2, 9.81 m/s2)
r=radius of tube (ft, m)
Thus the height h will be infinite when acceleration due to gravity (g) is zero. Thus, in space where the force of gravity is matched by an equal and opposite centripetal force due to the angular velocity of the satellite, a sense of weightlessness is experienced which fails to contain or pull down the mercury or other liquid in the tubes or holes. Therefore only a slight pressure differential is required across the tubes or holes to cause the mercury, in effect to be sucked out of the holes.
In order to address this issue, the present invention provides for the cover over the top of the structure or a micro-valve controlling the flow of liquid into the channels, as is discussed in above. Thus the cover or micro-valve arrangement serves not only to avoid evaporation of the liquid in a near vacuum environment but also addresses the problem of the liquid being sucked out of the channels by the differential pressure. In the case where the channels are filled just prior to firing, the timing of the firing can be synchronized to correspond to the channel fill time, taking into account the pressure differential across the channels and the kinematic viscosity of the liquid.
It will be appreciated that in order to address the issue of pooling of the liquid channels, reservoirs or tanks in which the liquid is kept initially have to be entirely filled. As the satellite is launched into space, gravitational acceleration and the acceleration of the space craft carrying the satellite act upon the liquid and cause it to accumulate in one area unless it is air gaps are eliminated e.g. by containing the liquid in a stretchable bladder.
Yet another embodiment of the invention is shown in cross-section in
As discussed above, in order to eject mercury droplets from the holes or channels in the substrate structure the heating elements such as the elements 108 shown in
As mentioned above, the present invention makes use of a substrate structure material such as SiC, with a high thermal shock parameter. The thermal shock parameter is given by the equation:
RT=(HσT(1−μ))/αE
Where H is the thermal conductivity, σT is the maximum tension the material can resist, μ is Poisson's ratio, α is the thermal expansion coefficient, and E is Young's modulus.
As discussed above, each heating element will heat a section of the liquid in the hole to define a disk of heated liquid that will be turned into its gaseous phase to define a bubble. In order to effectively eject the droplet of liquid the heating of the layer of liquid to its vaporization point (nucleation) has to take place extremely quickly (more than 1 million degrees C./second) to form a bubble within a 10 μs time-frame.
One embodiment of an electrical circuit controlling the firing of a mercury droplet is shown in
A typical satellite is shown in three dimensions in
It will be appreciated that in another embodiment the attitude control structures or individual attitude control systems may be mounted directly on the satellite without the use of outriggers.
As discussed above, mercury may be used as the liquid to be ejected from the holes. While other liquids may be used, it should be borne in mind that mercury has a much higher density than water and has a surface tension of 450 dyne per centimeter compared to 72 dyne per centimeter for water. As mentioned above, the density of mercury is 13.5 kg per liter=13.5×103 kg per cubic meter. For a satellite having a mass of 5 kg and a radius R of 0.1 m and a length L of 0.2 m, and using a droplets radius of 37 μm ejected at a droplets velocity of 10 m/s, the moment of inertia, droplet volume, droplet mass, droplet momentum and angular velocity can readily be calculated.
Thus assuming the use of mercury and a satellite defined by a solid cylinder of radius R=0.1 m and length L=0.2 m, mercury droplets fired from the surface at 90° will have a momentum of P and generate a rotational velocity ω as given by the equation P*R=I*ω where I is the moment of inertia of the cylindrical satellite.
I for end over end rotation is given by I=0.25 MR2+ 1/12 ML2 and for rotation about its longitudinal axis I=0.5 MR2.
Thus I for end over end rotation is 0.029 kg m2
I for rotation about the longitudinal axis is 0.025 kg m2
Assuming a mercury droplet size of 212 pl fired at 10 m/s, its mass mdrop (given by multiplying the volume by its density), is 2.864×10−9 kg
Therefore the droplet momentum Pdrop=mdrop*v=2.864×10−8 kg m/s
Thus the angular velocity of the satellite (end over end) as a result of firing one droplet is ωdrop=Pdrop(L/2)/I=9.821×10−8 rads/sec
For a rotation of 1 degree (end over end) in 10 minutes the required ωreq is 2π/(360×10×60) rads/second.
The number of drops required is therefore ωreq/ωdrop=296 drops
In a control system comprising a matrix of die or substrate structures with a total of 5×106 holes this allows 5×106/296 groups of firings. Every adjustment requires an opposite firing to stop the satellite, the total number of adjustments is 5×106/(296×2)=8.44×103.
Over a lifespan of 30 years=10950 days this gives an average of 0.77 end over end adjustments per day.
For rotation about the longitudinal axis, the angular velocity as a result of firing one droplet is ωdrop=Pdrop(R)I=1.146×10−7 rads/sec
Therefore a rotation of one degree about the longitudinal axis requires 254 drops thus provide a total of 9.847×103 adjustments or an average of 0.9 adjustments per day over a 30 year period.
If more adjustments have to be made on average, either multiple control devices could be affixed to the satellite or additional mercury could be provided, e.g., from a reservoir, as discussed above.
In the above embodiment the satellite was depicted as a cylindrical structure 20 cm in length and 10 cm in diameter and weighing 5 kg.
New generation satellites, including pico satellites (in the range of 100 g-1 kg), nano satellites (in the range of 1-10 kg), and micro satellites, often have a cubic configuration. It will therefore be appreciated that the example above of a cylindrical satellite was for illustrative purposes only.
In one embodiment, in order to create greater flexibility, multiples of holes e.g., 296 or 254 holes each firing once (depending on the location of the device—whether for rotation end over end or about the longitudinal axis) or fewer holes firing multiple times in succession, are grouped together to be fired as groups.
In another embodiment small wafer elements e.g. individual die or small groups of die may share a common reservoir and each hole may have its own heating element or multiple holes may share a heating element, i.e., the holes in a die may share a heating element or each have their own heating element, the die defining a partially autonomous module. Such modules with their own reservoir thus include a reservoir dedicated to replenishing only the holes of that module. This allows ejection devices to be built up to the desired size by simply securing the desired number of modules to a frame. For practical reasons the modules that are combined are preferably centrally controlled from one processor.
It will be appreciated that in the embodiments making use of a substrate structure in which the heating elements are provided on the distal surface of the substrate structure that is in contact with an underlying reservoir of mercury or other liquid, heating of the liquid in the substrate structure holes will produce conductive heat loss into the underlying liquid reservoir. In order to minimize this large liquid area in contact with the heating elements, another embodiment of a substrate structure is shown in
As mentioned above, one variation of the embodiment (such as the one depicted in
In the above example in which holes or ejection channels of 37 um diameter and 37 um depth were use, the energy to be generated by the resistive element to eject a mercury droplet can be calculated by combining the energy required to heat the droplet from ambient of say 20 degrees C. to the boiling point of mercury at 356.73 degrees C., with the energy required to turn the mercury to vapor. The specific heat of mercury (C) is 0.14 J/g degree C.
The temperature to heat from a droplet to mercury from 20 to 356.73 degrees C. (delta T of 336.73 degrees C.) can be calculated from the mass of the droplet.
Thus Energy Q=droplet mass×specific heat×delta T=2.5319×10−5 J
Total energy to vaporize mercury droplet from 20 degrees C.=Q+Vapor, which is approximately 1.837×10−4 J
However, the above calculations are based on vaporization of the entire droplet. In practice only a thin layer of the order of 1 μm is vaporized.
So instead of a depth of liquid of 37 um, only a layer thickness of 0.1 um has to be heated, which requires only 1/370 of the energy=4.96×10−7 J
A 1 μm layer of the mercury has a volume of 1.08×10−16 m3 and a mass of 1.45×10−12 kg. Since the mercury has an atomic mass of 200.59 g/mol, the layer of mercury that is vaporized contains 7.2287×10−12 mol. When in the vapor phase, the adiabatic volume of mercury is 22.4 liters/mol. Therefore each hole produces 1.62×10−10 liters of mercury vapor to generate the ejection pressure for the droplet of mercury that is ejected.
Empirical data relating the adiabatic volume of mercury vapor to the emission velocity for different diameter holes is given in Table A below. This allows the a preferred droplet velocity to be related to a hole diameter to optimize hole size.
For purposes of this application, the thermal creation of a bubble of vapor within a channel to eject a volume of liquid from the channel will be defined as thermal ejection.
The other concern mentioned above when dealing with attitude control devices in outer space is the issue of temperature. Extreme temperature variations exist in space due to the lack of atmosphere, and may range from −243 degrees C. when shielded from the sun's radiation to over 100 degrees C. when exposed to the sun. Mercury, for example, has a boiling point of 356.9 degrees C. and a freezing point of −38.8 degrees C. Thus the devices, which are typically mounted at various locations on the satellite, may be exposed to temperatures that exceed the boiling point of mercury or drop below the freezing point of mercury. The devices therefore have to be temperature controlled.
One embodiment of the invention makes use of a thermo-electric device (thermogenerator) operating on the Seebeck effect, wherein the temperature gradient from a hot surface to a cold surface across a pair of junctions induces a corresponding logarithmic mobile charge carrier gradient. The most usual of these are electrons—for example, across metal. This temperature gradient induced current from the thermogenerator can be stored in a battery. In addition to absorbing heat at the hot junction and giving off heat at the cold junction during the battery charging process, current from the battery can be used at another time to heat (or Peltier cool) either junction to maintain the correct temperature of the mercury. Similarly current from the battery can be used to warm the mercury by heating a resistive element within the substrate, or even to power the heating elements that are used for thermal ejection. Thus the charged battery from the Seebeck effect can, in one embodiment be used at a later time for vaporizing the mercury. DC to DC conversion electronics may be used to increase the voltage level from the battery in order to provide the desired voltage across the heating elements to create an appropriate amount of heat. Such temperature control allows the temperature of the liquid that is to be ejected to be controlled prior to ejection, to avoid the liquid from boiling or freezing and preferably keeping it within a defined range that optimizes the liquid's viscosity. The temperature control device, such as a thermo-electric device is, in one embodiment, connected to the control structure, e.g., attached to the liquid reservoir to control the temperature that is fed to the ejection channels of the control device.
In the above embodiments, various configurations were discussed in which a reservoir was secured to the distal surface of the substrate structure or the channel was sufficiently long to facilitate multiple firings from a single channel. However, in one embodiment, the substrate structure is configured without a reservoir, allowing a single firing of liquid from each hole, either simultaneously or individually. One such embodiment is shown in
In another embodiment, as shown in
The present invention provides a satellite attitude control system that permits much finer control over the attitude adjustments compared to monopropellant thrusters. For instance in the example given above for one monopropellant micro-chemical thrusters for micro satellites based on a hydrogen peroxide fuel the finest resolution was limited to bursts of 80 μNs pulses. As discussed above, in a 20 cm×10 cm nano satellite with a mass of 5 kg this would create an end-over-end rotation speed in excess of 4 degrees/second making it impossible to finely adjust the attitude. In contrast, the embodiment of the invention discussed above that made use of 37 μm diameter holes to eject mercury at 10 m/s, a 2857 times finer adjustment granularity can be achieved. This allows for a 1 degree adjustment over a 10 minute period by firing 296 holes from its MEMS substrate. As discussed above, the present invention also has the flexibility of increasing or decreasing hole size in a substrate structure or providing a substrate structure with a range of hole sizes, and can control the number of firings to accommodate different granularity requirements.
The invention is also much lighter than other prior art attitude adjusters such as Momentum/Reaction Wheels or Control Moment Gyroscopes.
While much of the discussion above relates to attitude control of a satellite, it will be appreciated that the control system is not limited to attitude control of satellites but allows for any position change, including translational movement of the satellites. It will be appreciated that the nature of the movement achieved depends on the positioning of the control systems on the satellite and the selection of the control systems to activate, including, which holes to fire.
While the present application has been described with reference to specific embodiments it will be appreciated that the invention could be implemented in different ways without departing from the scope of the invention as defined by the claims.
This is an application that claims priority from provisional application 61/341,121 to Charles E. Hunter et. al filed Mar. 26, 2010 and provisional application 61/342,649 to Charles E. Hunter et. al filed Apr. 16, 2010.
| Number | Date | Country | |
|---|---|---|---|
| 61341121 | Mar 2010 | US | |
| 61342649 | Apr 2010 | US |
