DIRECTING LIGHT FOR THERMAL AND POWER APPLICATIONS IN SPACE

Abstract

Solar collectors can provide power for electricity, thermal propulsion, and material processing (e.g., mining asteroids). In one aspect, a rocket propulsion system is configured to produce thrust for a spacecraft and includes: one or more optical elements configured to receive solar energy. The optical elements include: a first window configured to allow energy to enter the rocket propulsion system and form a concentrated energy beam, and a second window positioned to allow the concentrated energy beam to pass to the heat exchanger. The second window is spaced away from the first window to form a pressurized plenum chamber therebetween. The system further includes: a heat exchanger configured to receive the energy and use it to heat and pressurize a propulsion gas, and a rocket nozzle configured to expel the pressurized propulsion gas.

Claims

1. A rocket propulsion system configured to produce thrust for a spacecraft, the system comprising: one or more optical elements configured to receive solar energy, the optical elements comprising: a first window configured to allow energy to enter the rocket propulsion system and form a concentrated energy beam; and a second window positioned to allow the concentrated energy beam to pass to the heat exchanger, the second window spaced away from the first window to form a pressurized plenum chamber therebetween; a heat exchanger configured to receive the energy and use it to heat and pressurize a propulsion gas; and a rocket nozzle configured to expel the pressurized propulsion gas.

2. The system of claim 1, wherein the first window is configured to function as a pressure window of the pressurized plenum chamber and further comprises one or more curved surfaces configured to provide a controlled amount of focusing of the energy.

3. The system of claim 2, wherein the controlled amount of focusing is selected to limit the concentration of the energy to temperatures within safe operating limits of system components.

4. The system of claim 1, further comprising a plenum injector tube configured to inject the propulsion gas into the pressurized plenum chamber.

5. The system of claim 4, wherein the second window comprises an array of holes that allow the propulsion gas to flow out of the plenum chamber into a pressure chamber containing the heat exchanger and toward the rocket nozzle.

6. The system of claim 5, wherein the array of holes is configured such that the propulsion gas flows through the holes to form a protective gas barrier within a defined distance of an output surface of the second window.

7. The system of claim 5, wherein the plenum injector tube is configured to introduce the propulsion gas into the plenum chamber prior to heating such that the gas: cools the plenum chamber; cools at least one surface of the first and second windows; and removes contaminants from at least one surface of the first and second windows.

8. The system of claim 7, wherein one or more surfaces of the first and second windows are coated with anti-reflection coatings.

9. The system of claim 4, further comprising a second injector tube, configured to inject a stream of propulsion gas directly into the pressure chamber.

10. The system of claim 9, wherein at least one of the plenum injector tube and the second injector tube is configured to receive and inject a mixture of molecules derived from an asteroid mining process.

11. The system of claim 1, wherein the heat exchanger is formed of a porous solid material configured to allow the flow of the propellant gas through the heat exchanger.

12. The system of claim 11, wherein the heat exchanger has an internal surface area having a size sufficient to transfer heat energy from the heat exchanger to propellant gas via heat conduction in order to produce the thrust for the spacecraft.

13. A method of manufacturing a spacecraft, comprising: positioning a first window toward an outer face of a spacecraft such that it can receive solar energy; aligning a second window with the first window such that the second window is spaced toward an interior of the spacecraft, forming a plenum chamber therebetween; configuring the first and second windows to focus an energy beam onto a heat exchanger; positioning a conduit for propulsion gas to flow into the plenum chamber; positioning the heat exchanger to receive the concentrated beam such that it heats and pressurizes propulsion gas located adjacent or within the heat exchanger; and providing a nozzle configured to expel the heated and pressurized propulsion gas after it is pressurized by the heat exchanger.

14. The method of claim 13, further comprising configuring the first window to maintain pressure within the plenum chamber.

15. The method of claim 14, further comprising providing for focusing of the energy beam using a curved surface on at least one of the first and second windows.

16. The method of claim 15, wherein the curved surface is configured to limit concentration of the beam to temperatures within safe operating limits of the heat exchanger.

17. The method of claim 15, wherein the second window comprises an array of holes configured to allow propulsion gas to flow from the plenum chamber into a pressure chamber containing the heat exchanger.

18. The method of claim 17, wherein one or more surfaces of the first window are coated with anti-reflection coatings.

19. The method of claim 18, wherein the array of holes are configured to allow propulsion gas to form a protective barrier at a second window surface that faces the pressure chamber.

20. The method of claim 18, wherein the array of holes is positioned and configured to facilitate cooling the first and second windows using the propulsion gas.

21. The method of claim 20, wherein the conduit is configured to introduce into the plenum chamber a propulsion gas comprising a mixture of molecules derived from asteroid mining processes, and the array of array of holes is configured to facilitate removal of contaminants from at least one surface of the first and second windows.

22. The method of claim 13, wherein the heat exchanger is formed of a porous material configured to allow the flow of the first propellant gas therethrough, and the porous material provides sufficient surface area for heat energy transfer to the propellant gas to produce the thrust for the spacecraft.

23. The method of claim 13, further comprising configuring a second injector tube to inject a stream of propulsion gas directly into the pressure chamber.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0068] FIG. 1 is a schematic illustration of an embodiment of a spacecraft installed within a launch vehicle, wherein the various light-weight structures of the spacecraft are folded into compact configurations in accordance with aspects of this disclosure.

[0069] FIG. 2 is a schematic illustration of the orbiting spacecraft, wherein lightweight structures have been unfolded and deployed to operational configurations in accordance with aspects of this disclosure.

[0070] FIG. 3 is a schematic illustration of the collection and concentration of solar energy by light-weight reflecting structures in accordance with aspects of this disclosure.

[0071] FIG. 4 illustrates a camera used to image the position of an energy beam on a polished optical surface in accordance with aspects of this disclosure.

[0072] FIG. 5 is a photograph of scattering of high intensity light from a polished surface in accordance with aspects of this disclosure.

[0073] FIG. 6 illustrates a protective hinge and cover plate that can interrupt the formation of an energy beam in accordance with aspects of this disclosure.

[0074] FIG. 7 illustrates a wire-grid sensor for measuring the centroid of a high energy beam of solar energy in accordance with aspects of this disclosure.

[0075] FIG. 8 illustrates an embodiment in which a turning mirror is configured to be rotated to direct the energy beam toward the rocket propulsion system in accordance with aspects of this disclosure.

[0076] FIG. 9 illustrates internal structures of the rocket propulsion system in accordance with aspects of this disclosure.

DETAILED DESCRIPTION

[0077] When launching spacecraft by rockets from the surface of the Earth, the spacecraft are typically carried within the aerodynamic nose cone of the launch vehicle. It is desirable to reducing launch costs by minimizing both the total mass of the spacecraft and its physical size. It is therefore useful to employ compact and light-weight structures wherever possible. As used herein, light-weight structures may generally refer to structures having a sufficiently low weight such that the structures can be launched from the Earth to escape from the gravitational influence of Earth using one or more rockets.

[0078] Aspects of this disclosure relate to systems and techniques for collecting solar energy and protecting against damage resulting from an energy beam formed form the collected solar energy. Although examples of such systems and techniques are described in connection with an asteroid mining spacecraft, aspects of this disclosure are not limited thereto. For example, the systems and techniques described herein can also be applied to solar collection systems that are installed temporarily or permanently on an asteroid or a moon. As described herein, there may be situations in which focused solar energy may have sufficient power to damage components which do not form part of the optical path for the solar energy, for example, due to misalignment of one or more components in the system. Thus, aspects of this disclosure can provide systems and techniques for preventing such damage from occurring in many different solar energy collection systems.

[0079] FIG. 1 illustrates an embodiment of an asteroid mining spacecraft, which can be positioned atop a launch rocket 5 (also referred to as a booster rocket) in accordance with aspects of this disclosure. The main body of the spacecraft 1, also called the bus, contains the fuel, electrical components, and communication equipment for a typical mission. The spacecraft 1 is further provided with one or more rocket propulsion systems 2 configured to provide maneuvering capability to allow the spacecraft 1 to reach and capture passing asteroids. The spacecraft 1 also contains an asteroid capture system 3 and a solar power collection system 4, both of which are shown in a folded and stowed configuration in FIG. 1. The booster rocket 5 is shown schematically by dashed lines.

[0080] FIG. 2 depicts an embodiment the asteroid mining spacecraft 1 in orbital flight after it has separated from the booster rocket 5 in accordance with aspects of this disclosure. The asteroid capture system 3 is shown in a deployed configuration. In certain embodiments, the asteroid capture system 3 may include a plurality of multiply-jointed rigid members or may include a plurality of hollow tubular polymer membranes which have been inflated by internal gas pressure. However, aspects of this disclosure are not limited thereto and in other embodiments the asteroid capture system 3 can include other structure member(s) configured to be deployed in order to capture and contain asteroids.

[0081] With continued reference to FIG. 2, the solar power collection system 4 is shown after it has been unfolded and deployed from its compact stowed configuration. In certain implementations, the solar power collection system 4 comprises a plurality of linear support elements 7, 8, and 9 which may be embodied as tubular polymer membranes that have been inflated and held rigid by internal pressurizing gas. The solar power collection system 4 further comprises a hollow lens-shaped or inflated lenticular structure 10 which provides a transparent upper surface 11 and a curved reflecting lower surface 12. The lower surface 12 is configured to serve as a curved mirror to collect and focus incoming solar radiation. The solar power collection system 4 further comprises rigid reflecting elements 13 and 14 and may further include additional optical elements (not shown) which can be used to direct a focused energy beam.

[0082] FIG. 3 illustrates an embodiment of the solar power collection system 4 configured to collect solar energy and focusing the solar energy into a powerful energy beam. When the collecting curved mirror 12 is pointed directly at the Sun, a portion of the incoming solar radiation 15 reflects from the curved surface 12 and converges toward a secondary mirror 13.

[0083] With an unobstructed view of the Sun, the collection of solar energy using the configuration illustrated in FIG. 3 can collect very large power beams. For example, in an embodiment in which the mirror 12 has a collecting area of about 1 square meter, the continuous power collected can exceed one kilowatt of thermal energy for spacecraft at Earth's orbital distance from the Sun. When this energy is focused to a spot of several centimeters in diameter, the beam can have sufficient power density to drill a hole into concrete or into a captured asteroid. In certain implementations, collecting mirrors 12 having an area larger than about 1 square meter and multiple mirror configurations can collect proportionally larger power beams.

[0084] The additional rigid reflecting elements 13 and 14 are configured to adjust the focus and direction of the concentrated energy beam 16. Energy beam 16 can be directed into the spacecraft 1 via the rigid reflecting elements 13 and 14. The solar power collecting system 4 can further include one or more additional optical elements 17 configured to direct the beam 16 to pass through the spacecraft 1 to the asteroid capture system 3 and to impinge upon a captured asteroid 18. The concentrated solar energy can be used to drill holes into the asteroid, to melt and vaporize volatile constituents of the asteroid material, and to cause spalling and excavation of the asteroid surface. Thus, as used herein, the high power energy beam 16 may refer to an energy beam having sufficient power to drill a hole into the captured asteroid, to melt and vaporize volatile constituents of the asteroid material, and/or to cause spalling and excavation of the asteroid surface.

[0085] It can be seen from the above description that it is desirable to carefully focus a powerful energy beam 16 and direct the energy beam 16 to avoid hitting and damaging parts of the spacecraft 1 structure. That is, the energy beam 16 may have sufficient power to damage the spacecraft 1 if a portion of the beam 16 is inadvertently misdirected and strikes components of the spacecraft 1 which do not form part of the solar power collecting system 4. There are several conditions that may affect a dynamic misalignment of the beam 16. A non-exhaustive list of these conditions includes, for example, errors in pointing reflecting mirror 12 directly at the sun, mechanical deformations of light-weight structures 3 and 7-12 due to spacecraft 1 maneuvering forces, and/or slowly changing dimensions of the structures 3 and 7-12 due to material fatigue and/or due to pressurization loss caused by micro meteor impacts.

[0086] Therefore, it is desirable to provide techniques for continually monitoring the position and focus of energy beam 16. It is also desirable to incorporate shutters which can safely block or prevent the formation of an energy beam 16 during solar alignment maneuvers. It is also desirable to quickly block the energy beam 16 should there be unexpected upsets in any of the spacecraft control systems, a so called “dead man” automatic safety shutdown system.

[0087] In various embodiments of this disclosure, the spacecraft 1 can include one or more sensors configured to generate signals indicative of the alignment, misalignment, and/or degree of alignment of the energy beam 16. For example, solar power collecting system 4 can use the signals generated by the sensor(s) to realign one or more of the optical components that redirect the energy beam 16 in order to ensure that the energy beam 16 is not directed at other portions of the spacecraft 1. The one or more sensors may be configured to provide data for pointing and tracking information for solar power collection and may also provide warning alerts and shutdown commands whenever the energy beam 16 points in potentially damaging direction.

[0088] The one or more sensors can be implemented by various different embodiments. For example, the one or more sensors can use or be embodied as video camera(s) 19 as described in connection with FIG. 4 below. In some implementations, the sensor(s) can use or be embodied as a grid of fine wires 26 as described in connection with FIG. 7 below.

[0089] Referring to FIG. 4, the spacecraft 1 includes a video camera 19 which may be placed in a convenient location upon the spacecraft body 1 or upon a supporting structure 7, 8, or 9. The camera 19 has a field-of-view 20, shown by broken lines in FIG. 4, which is directed to view the surface of reflecting element 13 while being simultaneously well away from the energy beam 16. If reflecting element 13 were a perfect mirror, none of the collected light in energy beam 16 would be scattered toward the camera 19. As a practical matter, virtually all polished mirrors diffusely scatter at least a small percentage of incident light into non-specular directions. Accordingly, the image seen by the camera 19 will be a diffuse glow located upon a region of the surface of element 13. If curved reflector 12 is perfectly aligned toward the incoming rays of sunlight 15, then the camera 19 image of the diffuse glow will be centered upon element 13. However, in the event that there is a small pointing error in the alignment of reflector 12, then the image glow will be displaced from the center of element 13. A displacement of the glow is an early warning measure of potentially damaging misalignment of the energy beam 16. Image processing software can readily detect the displacement of the glow image and provide signals to adjust the pointing direction of reflector 12. For example, in certain embodiments, the camera 19 may include one or more processors configured to run the image processing software in order to detect displacement of the glow image. In other embodiments, the spacecraft 1 can include one or more processors separate from the camera 19 configured to run the image processing software.

[0090] In a like manner, additional imaging cameras (not illustrated) may be placed to detect misalignments of the energy beam 16 on other reflecting surfaces along the path of the energy beam 16. Misalignment measurements at each imaged reflecting element can be used to correct the pointing accuracy of the preceding element in the optical path. Misalignment measurements by a plurality of imaging cameras can provide continuous adjustment of the high energy beam 16 during operation without requiring a system shutdown for calibration. In certain implementations, each of the plurality of cameras may include its own set of one or more processors configured to run image processing software to detect displacement of the respective glow images. In other implementations, each of the plurality of cameras may provide image data to one or more processors separate from the cameras which run the image processing software for each of the received video feeds.

[0091] FIG. 5 is a photograph of the diffuse glow image upon a polished mirror in accordance with aspects of this disclosure. In this image, concentrated sunlight from a lens 21 is reflected from a concave mirror surface 22. The camera taking the image has been positioned to avoid the strong specular reflection. The diffuse glow can be clearly seen as a spot of light 23 that is not centered on the polished mirror.

[0092] Referring to FIG. 6 in a second embodiment, the spacecraft 1 includes a hinge 24 and cover plate 25 affixed to cover the front surface of reflecting element 13. In this configuration, converging rays from curved reflector 12 are prevented from reflecting and combining to form energy beam 16. Depending on the surface reflectance of cover plate 25, the incident light will be partially absorbed and partially diffusely reflected. As in the first embodiment of FIG. 4, the camera 19 will image a diffuse glow upon the surface of cover plate 23. The image position of the glow spot on the surface of the cover plate indicates the accuracy of alignment of curved surface 12 to the incoming solar energy 15 without requiring a reflection from element 13. In this manner, the curved surface 12 can be accurately aligned to the direction of the sun before a powerful energy beam 16 is directed further into the spacecraft 1. Furthermore, the curved surface 12 may be accurately aligned in the direction of the sun without need for a separate sun tracker or sun observing device.

[0093] With continued reference to FIG. 6 with the cover plate 25 in the blocking position as shown, a significant amount of solar energy 15 can be absorbed by the cover plate 25 thereby causing its temperature to rise. As the plate 25 heats, it will radiate long wavelength thermal energy in all directions. Nearby structures, such as the lenticular structure 10 with the transparent upper surface 11 will intercept a portion of the radiated thermal energy and also begin to warm. In this manner, lenticular structure 10 and other surfaces can be warmed as needed to prevent moisture or ice build that could potentially degrade the optical performance of these components. Since moisture may be produced during asteroid mining operations, it is desirable to heat optical surfaces as needed.

[0094] In some implementations, the cover plate 25 may be replaced or supplemented with another type of physical blocking element. For example, the physical blocking element may include one or more of the following: optical shutter(s), blocking plate(s), absorber(s), diffuser(s), and/or optical element(s) configured to deflect or defocus the energy beam 16.

[0095] FIG. 7 illustrates a third embodiment of the solar power collection system 4 in accordance with aspects of this disclosure. As shown in FIG. 7, the solar power collection system 4 includes a grid of fine wires 26 placed near the reflecting surface of reflecting element 13. The wires 26 can be fabricated from materials which can tolerate high temperatures, such as tungsten or stainless steel. The wires 26 may have small diameters such as about 0.001 inch. The fine wires 26 intercept a small percentage of the converging solar radiation that is incident upon the reflecting element 13 and begin to warm. As the wires 26 are heated, their electrical resistance will increase. The wires closest to the center of the energy beam 16 will exhibit the largest resistance increases. The solar power collection system 4 can include one or more circuit configured to measure the resistance difference between adjacent wires 26, which provides information about directional misalignment of the energy beam 16. Very small resistance differences may be accurately measured by a common electrical circuit known as a Wheatstone Bridge. One or more processors may be configured to determine whether the energy beam 16 is misaligned based on the measured electrical resistances and provide control signals for repositioning and/or rotating the optical elements 10, 13, 16, 17 to correct the misalignment of the energy beam 16.

[0096] The solar power collection system 4 can include one or more actuators (not illustrated) configured to reposition and/or rotate one or more of the optical elements 10, 13, 16, 17 that generate and/or redirect the energy beam 16. Thus, the solar power collection system 4 can use the outputs received from the camera(s) 19 and/or the wires 26 as an input to control the alignment of the optical elements 10, 13, 16, 17 to ensure that the beam 16 is not directed onto any unintended components of the spacecraft 1, and thus, does not damage any components of the spacecraft 1.

[0097] FIG. 8 illustrates an embodiment in which a turning mirror 17 is configured to be rotated to direct the energy beam 16 toward the rocket propulsion system 2 in accordance with aspects of this disclosure. FIG. 9 illustrates internal structures of the rocket propulsion system 2 in accordance with aspects of this disclosure.

[0098] With reference to FIGS. 8 and 9, optical element 17 is embodied as a turning mirror 17 which can be rotated to selectively direct the energy beam 16 toward the rocket propulsion system 2 instead of towards the asteroid capture system 3. For example, the turning mirror 17 may be coupled to an actuator (not illustrated) configured to rotate the turning mirror 17 between a first orientation (e.g., as shown in FIG. 3) and a second orientation (e.g., as shown in FIG. 8).

[0099] The energy beam 16, represented in FIG. 9 within boundary rays, enters the rocket propulsion system 2 through two transparent windows 27 and 28. A first window 27 forms a pressure seal against components of the rocket propulsion system body 30 and 30a. The window 27 may have flat or curved surfaces forming a focusing lens. To reduce undesired optical reflections, one or both surfaces of the first window 27 may be coated with anti-reflection coatings.

[0100] The first window focuses and converges the energy beam 16 into a focused converging energy beam 16a, represented within boundary rays. The focused converging energy beam 16a passes through a second transparent window 28 and then into a heat exchanger 32.

[0101] The rocket propulsion system 2 is configured to heat a pressurized propellant gas to a high temperature and then expel the gas through a converging-diverging rocket nozzle 31 in order to produce useful rocket thrust. It is a general physical principle that most suitable propellant gasses are substantially transparent to solar energy. Thus, the gas will not efficiently absorb the wavelengths present in solar light.

[0102] The solid bulk material of heat exchanger 32 is designed to be an efficient absorber of solar energy, e.g., provided by the energy beam 16a. The incoming energy beam 16a heats the heat exchanger 32 to a high temperature, typically on the order of 2000 degrees Celsius. The heat exchanger 32 is also porous to allow the flow of propellant gasses through the heat exchanger 32. The heat exchanger 32 also contains a large internal surface area which, by heat conduction, efficiently transfers heat energy from the bulk material of the heat exchanger 32 to the propellant gas.

[0103] Highly concentrated solar energy in a small volume can produce temperatures far higher than the melting point of all known solid materials. Therefore, the focusing properties of first window 27 may be chosen to limit the concentration of solar energy to temperatures within the safe operating limits of the heat exchanger 32 and the surrounding structures 30a.

[0104] The second window 28 is spaced away from first window 27 by a predetermined distance. In some implementations, the distance between the first and second windows 27 and 28 may be on the order of 1 centimeter, however, in other implementations, the distance may be greater or less than 1 centimeter. In one illustrative embodiment, the distance may be 20 centimeters. In certain embodiments, the volume 40 contained between the two windows 27 and 28 forms a plenum configured to hold a pressurized propulsion gas. A first propulsion gas 41 can be injected into the plenum 40 through a first injector tube 42. In some embodiments, the second window 28 can include a plurality of small through-holes 29 which permit the first propulsion gas 41 to uniformly flow through window 28 into the volume of thrust chamber 50 in a two-dimensional pattern of linear jets.

[0105] One purpose of the first propulsion gas 41 is to provide cooling for both windows 27 and 28. The cooling gas reduces undesirable thermal stresses on both windows. Cooling reduces the probability of window fractures. Furthermore, cooling permits anti-reflection coatings to be applied to both windows, thereby increasing optical efficiency. Antireflection coatings generally have temperature limits that are lower than the temperature capabilities of solid materials and cannot be applied to excessively hot surfaces.

[0106] Similarly, a second propulsion gas 43 can be injected through a second injector tube 44 into the thrust chamber 50. Both of the gases 41 and 43 can be driven by their supply pressures to flow through the heat exchanger 32 and into the rocket nozzle 31. The temperature of the gases 41 and 43 and their thermal energy content (enthalpy) can be greatly increased by passage through the heat exchanger 32 (e.g., which provides heat from the energy beam 16a to the gases 41 and 43). The rocket nozzle 31 efficiently converts the random thermal motion and the relatively slow flow-velocity of hot gas molecules into linear motion in a high-speed exhaust plume, thereby producing rocket thrust.

[0107] The use of the two windows 27 and 28 and two injection gasses 41 and 43 provides a number of advantages over other thruster systems. One advantage is that many different types of gasses can be used to produce thrust in the rocket propulsion system 2. Examples of gases that can be used in the rocket propulsion system 2 include but are not limited to: H.sub.2, H.sub.2O, HF, Ar, Cl.sub.2, Br, CO.sub.2, N.sub.2, NH.sub.3, CH.sub.4, mixtures of these gases, and a variety of volatile hydrocarbon molecules. Such gases can be derived directly from asteroid mining operations.

[0108] Another advantage is that many gas mixtures, when subjected to intense photochemical and thermochemical processes, may dissociate and recombine to produce opaque compounds. It is undesirable to allow opaque compounds to settle on the transparent window surfaces where they can block the energy beam 16. By injecting a relatively cool first propulsion gas 41 through the multiple holes 29 in the second window 28, a gas barrier can be formed near the surface of the second window 28. The outflow of the gas 41 can prevent turbulent gas flow from the propulsion gases 41 and 43 and various contaminants due to heat exchanger 32 reactions from reaching and contaminating the second window 28.

[0109] When exposed to extreme temperatures, some gases including CH.sub.4 are known to deposit carbon compounds on nearby surfaces. This process is known as coking. According to aspects of the present disclosure, coking deposits may be removed by periodically injecting cleaning gasses such as H.sub.2, HF, or Cl.sub.2 singly or in combination through injection ports 42 and/or 44, providing yet another advantage to aspects of this disclosure.

CONCLUSION

[0110] Unless the context clearly requires otherwise, throughout the description and the claims, the words “comprise,” “comprising,” and the like are to be construed in an inclusive sense, as opposed to an exclusive or exhaustive sense; that is to say, in the sense of “including, but not limited to.” The word “coupled”, as generally used herein, refers to two or more elements that may be either directly connected, or connected by way of one or more intermediate elements. Likewise, the word “connected”, as generally used herein, refers to two or more elements that may be either directly connected, or connected by way of one or more intermediate elements. Additionally, the words “herein,” “above,” “below,” and words of similar import, when used in this application, shall refer to this application as a whole and not to any particular portions of this application. Where the context permits, words in the above Detailed Description using the singular or plural number may also include the plural or singular number respectively. The word “or” in reference to a list of two or more items, that word covers all of the following interpretations of the word: any of the items in the list, all of the items in the list, and any combination of the items in the list.

[0111] Moreover, conditional language used herein, such as, among others, “can,” “could,” “might,” “can,” “e.g.,” “for example,” “such as” and the like, unless specifically stated otherwise, or otherwise understood within the context as used, is generally intended to convey that certain embodiments include, while other embodiments do not include, certain features, elements and/or states. Thus, such conditional language is not generally intended to imply that features, elements and/or states are in any way required for one or more embodiments or that one or more embodiments necessarily include logic for deciding, with or without author input or prompting, whether these features, elements and/or states are included or are to be performed in any particular embodiment.

[0112] The above detailed description is not intended to be exhaustive or to limit the invention to the precise form disclosed above. While specific embodiments of, and examples are described above for illustrative purposes, various equivalent modifications are possible within the scope of the disclosed invention(s), as those skilled in the relevant art will recognize. For example, while processes or blocks are presented in a given order, alternative embodiments may perform routines having steps, or employ systems having blocks, in a different order, and some processes or blocks may be deleted, moved, added, subdivided, combined, and/or modified. Each of these processes or blocks may be implemented in a variety of different ways. Also, while processes or blocks are at times shown as being performed in series, these processes or blocks may instead be performed in parallel, or may be performed at different times.

[0113] The teachings provided herein can be applied to other systems, not necessarily the system described above. The elements and acts of the various embodiments described above can be combined to provide further embodiments.

[0114] While certain embodiments have been described, these embodiments have been presented by way of example only, and are not intended to limit the scope of the disclosure. Indeed, the novel methods and systems described herein may be embodied in a variety of other forms; furthermore, various omissions, substitutions and changes in the form of the methods and systems described herein may be made without departing from the spirit of the disclosure. The accompanying claims and their equivalents are intended to cover such forms or modifications as would fall within the scope and spirit of the disclosure.