Patent Application
20190315498
Not Applicable.
A portion of the disclosure of this patent document contains material that is subject to copyright protection. The patent owner has no objection to the facsimile reproduction by anyone of the patent document or the patent disclosure, as it appears in the Patent and Trademark Office patent file or records, but otherwise reserves all copyright rights whatsoever.
None.
This disclosure relates to rocket engines, and more specifically, to rocket engines which utilize nuclear fission as the source for thermal energy in the creation of motive force to create specific impulse sufficient for lifting objects to earth orbit, or for insertion into interplanetary flight.
A continuing interest exists for improvements in rocket engines, and more particularly for designs that would provide a significant increase in propulsive power, as often characterized by the benchmark of specific impulse, especially as might be compared to conventional chemically fueled rocket engines. Such new rocket engines might be useful in a variety of applications. Launch operational costs might be substantially reduced on a per pound of payload basis, by adoption of a new nuclear thermal propulsion rocket engine design that provides significant improvements in the specific impulse, as compared to existing prior art rocket engine designs. Further, from the point of view of overall mission costs, since the mass of most components of rocket vehicles are proportional to the mass of the propellant, it would be desirable to develop a new rocket engine design that reduces the mass of consumable components necessary for initiating lift off and acceleration to orbital velocity. Such an improvement would have a major impact on the entire field of rocket science from a launch weight to payload ratio basis. Such an efficiency improvement would also facilitate the inclusion of wings and recovery systems that would enable an economic fully reusable launch system with airliner type operations, for example, as described in U.S. Pat. No. 4,802,639, issued Feb. 7, 1989 to Richard Hardy et al., entitled HORIZONTAL TAKEOFF TRANSATMOSPHERIC LAUNCH SYSTEM, the disclosure of which is incorporated herein in its entirety by this reference. And, for missions beyond earth orbit, it would be advantageous, from the point of view of mission duration, to provide a new rocket engine design that increases not only the payload to launch weight ratio, but also reduces the transit time to the mission objective, by providing high specific impulse, so as to minimize fuel required to achieve high vehicle velocities necessary to accomplish a selected interplanetary mission in a specific time frame. And, it would be desirable to provide such an improved rocket engine that includes components which have been reused and identified as comparatively reliable and cost effective, and thus, minimizes design risk and thus minimizes the extent of testing that may be necessary, as compared to many alternate designs which are subject to stress and strain from temperature and pressure in rocket engine services. Thus, it can be appreciated that it would be advantageous to provide a new, high efficiency rocket engine design which provides a high specific impulse, thus minimizing the launch weight to payload ratio.
In general, the efficiency of a rocket engine may be evaluated by the effective use of the consumable propellant, i.e. the amount of impulse produced per mass of propellant, which is itself proportional to the velocity of the gases leaving the rocket engine nozzle. In nuclear thermal rocket engine systems, the specific impulse increases as the square root of the temperature, and inversely as the square root of the molecular mass of the gases leaving the rocket engine nozzle. Consequently, in the design of a nuclear thermal rocket engine, efficiency is maximized by using the highest temperature available, given materials design constraints, and by utilizing a fluid that has a very low molecular mass for generation of thrust.
A variety of fission based rocket engines have been contemplated, and some have been tested. An overview of the current status of such efforts, and suggestions as to suitable configurations for various missions, was published on Oct. 16, 2014, at the Angelo State University Physics Colloquium in San Angelo, N. Mex., by Michael G. Houts, Ph.D, of the NASA Marshall Space Flight Center, Huntsville, Ala., in his presentation entitled Space Nuclear Power and Propulsion; a copy of which is available at: http://ntrs.nasa.gov/search.jsp?R=20140016814. As he notes, the Rover/NERVA program (Ser. No. 19/554,973) tested a fission rocket engine design. Further, the most powerful nuclear rocket engine that has been tested, to date, was the Phoebus 2a, which utilized a reactor that was operated at a power level of more than 4.0 million kilowatts, during 12 minutes of a 32 minute test firing. However, it is clear that the various nuclear fission rocket engine designs currently available have various drawbacks, such as excessive gamma radiation production of retained core components, which requires extensive and heavy shielding, if used on manned missions.
One of the more interesting disclosures of a fission based rocket engine was provided in U.S. Pat. No. 6,876,714 B2, issued on Apr. 5, 2005 to Carlo Rubbia, which is titled DEVICE FOR HEATING GAS FROM A THIN LAYER OF NUCLEAR FUEL, AND SPACE ENGINE INCORPORATING SUCH DEVICE, the disclosure of which is incorporated herein in its entirety by this reference. That patent discloses the heating of hydrogen by fission fragments emitted from a thin film of fissile material, such as Americium metal or a compound thereof, which is deposited on an inner wall of a cooled chamber. However, that device generally describes the use of fissile material in critical mass conditions, and although it mentions the contemplation of sub-critical mass fission arrangements, details of such a condition are scant, if indeed present at all in the description thereof.
Additionally, an improved design for a nuclear thermal propulsion rocket engine was provided in U.S. Pat. No. 9,180,985 B1, issued on Nov. 10, 2016, to Hardy et al., which is titled NUCLEAR THERMAL PROPULSION ROCKET ENGINE, the disclosure of which is incorporated herein in its entirety by this reference. That patent, by two of the same inventors as the present design, includes specific various design elements useful for providing plutonium to a reactor using hydrogen as a carrier. However, subsequent work has revealed that it would be desirable to more precisely regulate the fission reactions occurring, to regulate the energy output from such a rocket engine.
Thus, there remains a need to provide an improved design for a high specific impulse nuclear thermal propulsion rocket engine that simultaneously resolves two or more of the various practical problems, including (a) reducing the reaction chamber pressure required to provide the required density of the moderator, as compared to the case when hydrogen atoms alone are used in the carrier fluid (b) providing for power control, especially as related to power generation amounts at any given time, by providing for throttling of the fission reaction; (c) minimizing the weight of consumables (such as chemical fuel constituents) on a per payload pound basis; (d) avoiding excessive radiation shielding requirements when the design is used in manned missions, by avoiding use of retained radioactive hardware that generates large gamma ray emissions; and (e) providing a high specific impulse, as compared to prior art rocket engines for space vehicles.
A novel fission based nuclear thermal propulsion rocket engine has been developed, which, in various embodiments, simultaneously provides a high specific impulse propulsion system, yet enables the regulation of the power output by moderating the neutrons produced during fission. The rocket engine design utilized a source of fissionable material such as plutonium in a carrier, such as methane, or ethane, or a combination thereof. A neutron source is provided from a neutron beam generator. By way of engine design geometry, various embodiments may provide for intersection of a neutron beam from the neutron generator with the fissionable material injected by way of a carrier fluid into a reactor. Impact of the neutron beam on the fissionable material results in a nuclear fission in sub-critical mass reaction conditions in the reactor, resulting in release of heat energy to the materials within the reactor. The reactor is sized and shaped to receive the reactants and to receive a low molecular weight expandable fluid such as hydrogen, and to confine heated and pressurized gases for discharge out through a throat, into a rocket engine expansion nozzle for propulsive discharge therefrom.
An advantage of the novel fission based nuclear thermal propulsion rocket engine design disclosed herein is that when such an engines is utilized in a second or subsequent stage of a launch system, the radioactive fission products would be exhausted into the vacuum of space, rather than at the launch site.
Importantly, advances in neutron beam generator technology are believed to make possible the development of a nuclear thermal rocket engine in which the process of production of neutrons can be partially separated from the process of absorption of neutrons by fissionable material, so that the fission process can be initiated and maintained while utilizing less than a critical mass of fissionable material. In this manner, a design has been developed in which radioactive fission products ejected out of the rocket nozzle into space with other exhaust gases, while amounts of fissionable material consumed are replenished with new fissionable material only as necessary to support continued fission, to obtain the necessary heat release for operation.
The present invention(s) will be described by way of exemplary embodiments, using for illustration the accompanying drawing in which like reference numerals denote like elements, and in which:
The foregoing figures, being merely exemplary, contain various elements that may be present or omitted from a final configuration for a an embodiment of a nuclear thermal rocket engine using sub-critical mass fission of fuels, or that may be implemented in various embodiments described herein for a rocket engine. Other variations in nuclear thermal rocket engine designs may use slightly different mechanical structures, mechanical arrangements, solid flow configurations, liquid flow configurations, or vapor flow configurations, and yet employ the principles described herein and as generally depicted in the drawing figures provided. An attempt has been made to draw the figures in a way that illustrates at least those elements that are significant for an understanding of exemplary nuclear thermal rocket engine designs under sub-critical mass fission conditions. Such details may be quite useful for providing propulsion for a high specific impulse space vehicle, and thus, reduce cost of payloads lifted to earth orbit, lunar, or interplanetary missions.
It should be understood that various features may be utilized in accord with the teachings hereof, as may be useful in different embodiments as useful for various sizes and shapes, and thrust requirements, depending upon the mission requirements, within the scope and coverage of the teachings herein as defined by the claims.
Attention is directed to
A selected actinide fuel F which provides a fissile material may be supplied from storage container 50 for mixing with the first fluid 28. In an embodiment, a selected fuel F may be provided in various forms. In an embodiment, the selected fuel F may be provided in a very fine particulate, or more specifically, in a finely powdered form. In an embodiment, the selected fuel F may be provided in chemical solution form. In an embodiment, the fuel F may comprise a selected actinide compound. In an embodiment, the fuel F may comprise a substantially pure metallic actinide. In an embodiment, the fuel F may be supplied in a form including of one or more plutonium (Pu) isotopes. In an embodiment, the fuel F may be supplied in as a fissile material in the form of plutonium 239 (239Pu). In an embodiment, the fuel F may be supplied as a fissile material in the form of uranium 235 (235U). In various embodiments, the selected fissile material providing fuel F, before injection into the reactor 12, may be provided in form compatible with the first fluid 28.
In an embodiment, the first fluid 28 from the first fluid storage compartment 26 may be mixed with a selected amount of fuel F, before injection into reactor 12. In an embodiment, the first fluid 28 and a selected amount of fuel F may be mixed to create a rich fuel mixture 52, before passage of the rich fuel mixture 52 (i.e. a mixture of fuel F and first fluid 28) through control valve 53 and then into a fuel turbopump 54, which pumps the fuel rich mixture 52 into reactor 12 via fuel supply line 56, fuel header 58, and a first set of fuel injectors 60 which confine and direct passage of fuel rich mixture 52 into reactor 12. In an embodiment, control valve 53 may provide on-off capability. In various embodiments, control valve 53 may additionally provide throttling capability to regulate the quantity of flow of the rich fuel mixture 52. In an embodiment, at time of injection, the fuel rich mixture 52 may be in fluid form while carrying an actinide fuel F therein. As shown in
In various embodiments, a rocket engine 10 may operate with fission of the fissile material of fuel F under sub-critical mass conditions. Under various embodiments, the fissile material may include plutonium 239. In an embodiment the amount of plutonium 239 (239Pu) provided may be between about between ten percent (10%) and fifteen percent (15%) by weight in said first fluid 28. In an embodiment the amount of plutonium 239 (239Pu) provided may be between about between ten percent (10%) and twenty percent (20%) by weight in said first fluid 28. In an embodiment the amount of plutonium 239 (239Pu) provided may be between about between ten percent (10%) and twenty five percent (25%) by weight in said first fluid 28. In an embodiment the amount of plutonium 239 (239Pu) provided in said first fluid may be between about one percent (1%) and thirty percent (30%) by weight.
In various embodiments, the first fluid may be a commercially available methane, having less than zero point one (0.1) mole percent of C5+ hydrocarbons thereon. In an embodiment, the first fluid 28, may further include one or more isotopes of hydrogen. In an embodiment, the first fluid 28 may additionally include deuterium. In an embodiment, the first fluid 28 may also include deuterium (2H). In an embodiment, the first fluid 28 may include at least some tritium (1T3). In an embodiment, the first fluid 28 may additionally include both deuterium and tritium. In an embodiment, the presence of tritium may induce secondary fusion reactions in the center of the fluid flow while being directed out through the nozzle, thereby increasing specific impulse without significantly increasing engine wall temperature.
By detailed investigation using a Monte Carlo nuclear reaction (neutron transport) modeling program, when varying mixtures of hydrogen and other fluids are utilized in combination with varying amounts of plutonium as fissionable material, it has been found that the generation of neutrons using the first fluids disclosed in our prior U.S. Pat. No. 9,346,565 B1 (noted above), which were limited to hydrogen and its isotopes, resulted in excess neutron generation. Thus, the results obtained under analysis via Monte Carlo simulation described a “banks full” result, indicating excess energy and neutron production. See TABLE 1, below.
Consequently, we have now discovered that it would be advantageous to use a first fluid containing a high hydrogen density molecule, such as simple hydrocarbons, and especially fluids containing carbon in molecular combination with hydrogen, such as methane and/or ethane, in order to provide hydrogen and/or carbon at densities in the reactor 12, in order to moderate the neutron emissions (i.e. slow down neutrons resulting from fission) in the reactor 12, to enable the reaction chamber to operate at or near conventional rocket engine design pressure ranges. In an embodiment, a suitable first fluid may be provided in a selected hydrocarbon composition that includes an alkane in the C1 to C5 range, inclusive. In an embodiment, a suitable first fluid may be provided including a selected hydrocarbon composition that consists essentially of methane (CH4), or ethane (C2H6), or a mixture of methane (CH4) and ethane (C2H6).
With the just described hydrocarbons as first fluids, and more particularly, when using high hydrogen density molecular substances, namely first fluids with molecules containing carbon and hydrogen (e.g, methane or ethane), the required operating pressure for the reaction chamber 12 may, in an embodiment, be operated in the range of from about two thousand pounds per square inch absolute (2000 PSIA) to about four thousand pounds per square inch absolute (4000 PSIA). In an embodiment, the required operating pressure for the reaction chamber 12 may be operated in the range of from about three thousand pounds per square inch absolute (3000 PSIA) to about four thousand pounds per square inch absolute (4000 PSIA). In an embodiment, the fuel turbopump may operate at about six thousand (6000) pounds per square inch (PSI) discharge pressure, plus or minus about fifteen percent (15%). As a prior art example, the Space Shuttle Main Engine, operated NASA in the United States, operated at a specified reaction chamber pressure of 2994 PSIA (pounds per square inch absolute) per numbers published by Aerojet Rockedyne. See: http://www.rocket.com/space-shuttle-main-engine.
The data provided by way of
To provide thrust, by way of heating and expansion in the reactor 12 and resultant expulsion out thru expansion nozzle 20, a low molecular weight fluid such as hydrogen (H2) is provided as the second fluid 32. A second fluid 32 may be stored in a second fluid storage compartment 30, and on demand is delivered by line 70 to the thrust fluid turbopump 44. The thrust fluid turbopump 44 receives the second fluid 32 from the second fluid storage compartment 30 and provides (generally indirectly) the second fluid 32 under pressure to the reaction chamber 12.
In an embodiment, the second fluid 32 may be sent under pressure from thrust fluid turbopump 44 via second fluid supply line 72 to a distribution ring 74 located at or near the exit plane 77 of expansion nozzle 20. The second fluid 32 may be supplied via distribution ring 74 to nozzle coolant passageways 76 located on the exterior 78 of expansion nozzle 20. In this manner, an extremely cold fluid, e.g. liquid hydrogen, may be utilized as a coolant for the expansion nozzle by passage of the second fluid 32 through the nozzle coolant passageways 76. Likewise, as also seen in
Once second fluid 32 reaches the upper end 90 of reactor 12, a collection header 92 may be utilized to accumulate the second fluid 32 from the reactor coolant passageways 86. In an embodiment, from collection header 92, the second fluid 32 may be directed to a second set of injectors 94 which are configured for confining the passage of the second fluid 32 during injection into the reactor 12. By way of injectors 94, the second fluid 32 may be directed toward or injected into a mixing zone 96, which mixing zone 96 is located downstream of the reaction zone 62. In mixing zone 96, the second fluid 32 is heated and expanded, in order to provide thrust by ejection through throat 14 and outlet 16 of reactor 12. Also, the first fluid 28 is heated and expanded, in order to provide thrust by ejection through throat 14 and outlet 16 of reactor 12.
As mentioned above, in order to provide power for the thrust fluid turbopump 44, a gas generating chamber 38 may be provided to generate combustion products in the form of a hot gas 40 that drives a turbine 42, which in turn drives a pump impeller 100. Consequently, when oxygen 36 is supplied for combustion with hydrogen as second fluid 32, water vapor is formed, and the resultant low pressure water vapor stream 46 is discharged overboard. Likewise, hydrogen as second fluid 32 and oxygen 36 may be supplied to a second gas generating unit 102 to generate hot gas 104 that drives turbine 106 which in turn drives fuel pump impeller 108 in fuel turbopump 54.
In another embodiment for rocket engine 10′ as seen in
In various embodiments for a rocket engine 10 or 10′ or the like, using nuclear thermal heating of a low molecular weight fluid such as hydrogen as described herein, a rocket engine may be provided that has a specific impulse in the range of from about eight hundred (800) seconds to about twenty five hundred (2500) seconds. In various embodiments using nuclear thermal heating of a low molecular weight fluid such as hydrogen as described herein, a rocket engine may be provided that has a specific impulse in the range of from about one thousand (1000) seconds to about twelve hundred fifteen (1215) seconds.
To summarize, in order to facilitate supply of hydrogen to the reactor 12 for heating, a thrust fluid turbopump 44 or 144 or the like may be provided as generally described herein above. In an embodiment, liquid hydrogen, i.e. a cryogenic liquid, may be provided to the rocket engine 10 or 10′, by way of a thrust fluid turbopump that is driven by a turbine which is rotatably energized by high temperature gases. In an embodiment, the high temperature gases may be provided by way of combustion products, such as by way of combustion of hydrogen and oxygen in a gas generating chamber GG to generate a high temperature combustion gas, which after passage through the turbine 42 or 162, as the case may be, may be exhausted overboard in the form of a water vapor stream 46 or 46′. The tradeoff of loss of efficiency due to loss of propellant (hydrogen) expended in the gas generating chamber GG, in view of the usual weight savings and simplicity of design (and lack of radioactive contamination), as compared to additional weight and complexity in view of any additional specific impulse contribution in designs that might avoid such combustion losses, may be evaluated for a specific space vehicle design and attendant mission profile, as will be understood by those of skill in the art. Various configurations for drive of a suitable thrust fluid turbopump for feeding hydrogen to the reaction chamber may be provided by those of skill in the art using conventional liquid turbopump system design principles, and thus, it is unnecessary to provide such details. In general, the thrust fluid turbopump must avoid cavitation while pumping liquid hydrogen at relatively low inlet pressure, and deliver the hydrogen to the reaction chamber (and in an embodiment, via distribution ring and cooling passageways) at very high pressure, and preferably, with capability to provide a relatively wide throttling range. In various embodiments, the selected thrust fluid turbopump 44 or 144 design may be optimized for minimizing weight while providing necessary performance while at the same time minimizing the thrust fluid turbopump package size, in order to minimize necessary space in a selected space vehicle design. Selection of suitable bearing sand seals are of course necessary, and various design alternatives are known to those of skill in the art. More generally, those of skill in the art will understand that turbopumps for supply of cryogenic liquids to rocket engines require designs that provide maximum performance at minimum weight.
Similarly, to facilitate supply of the first fluid carrying plutonium or other actinide to the reactor 12 for fission of at least some of the plutonium, a fuel turbopump 54 may be provided. In various embodiments, liquid methane, or liquid ethane, or combinations thereof (i.e. cryogenic liquids), may be provided to the rocket engine 10 or 10′, by way of a fuel turbopump 54 or 160, that is driven by a turbine (106 or 162) which is rotatably energized by high temperature gases. In an embodiment, the high temperature gases may be provided by way of combustion products, such as by way of combustion of hydrogen and oxygen to generate a high temperature combustion gas. Various configurations for drive of a suitable fuel turbopump for feeding reactants to the reaction chamber may be provided by those of skill in the art using conventional liquid turbopump system design principles, and thus, it is unnecessary to provide such details. In general, the fuel turbopump (54 or 160) must avoid cavitation while pumping liquids at relatively low inlet pressure, and deliver the liquids to the reaction chamber at very high pressure, and preferably, with capability to provide a relatively wide throttling range. In various embodiments, the selected fuel turbopump design may be optimized for minimizing weight while providing necessary performance while at the same time minimizing fuel turbopump package size, in order to minimize necessary space in a selected space vehicle design.
Further, in order to generate electricity for a selected neutron beam generator 22, an electrical generator 146 may be combined with a turbopump 144, so that a hot gas driven turbine 162 in the turbopump 144 also provides shaft power for an electrical generator 146. In an embodiment, the high temperature gases may be provided by way of combustion products, such as by way of combustion of hydrogen and oxygen in a gas generating chamber GG to generate a high temperature combustion gas, which after passage through the gas turbine 162, may be exhausted overboard via a water vapor exhaust tube 46. Alternately, a stand-alone electrical turbine generator may be provided, with its own fluid or combustion gas driven turbine, in the manner as generally described above.
In an embodiment, a deuterium-deuterium (“DD”) type neutron generator 22 may be utilized. As an example, high yield neutron generators are currently available for various applications with variable neutron output between 1×1011 and 5×1011 neutrons per second (n/s). It is an advantage of a DD type neutron generator design that because no tritium is utilized, radiation shielding and accompanying safety concerns and regulatory burdens are significantly reduced. Thus, such designs may be more suitable for manned space vehicles.
In an embodiment, a deuterium-tritium (“DT”) type neutron generator may be utilized. As an example, extremely high yield neutron generators based on DT design principles are currently available with variable neutron output between 1×1013 and 5×1013 neutrons per second (n/s). Such designs may require appropriate shielding and regulatory approvals for manned spaceflight applications, but may be especially suitable for high payload unmanned spaceflight vehicle applications.
Neutron generators of either deuterium-deuterium design or of deuterium-tritium design have been developed by Phoenix Nuclear Labs, 2555 Industrial Drive, Monona, Wis. 53713, with a web page at http://phoenixnuclearlabs.com. Other vendors currently provide different designs. For example, Gradel Group, 6, Z.A.E. Triangle Vert, L-5691 ELLANGE, Luxembourg (see http://gradellu/en/activities/neutrons-generators/products/14-1-mev-neutrons-dt/) currently provides a 14 MeV neutron generator of deuterium-tritium design, with basic functionality as follows:
1D2+1T3->2He4(3.5 MeV)+0n1(14.1 MeV)
It is currently anticipated that any selected neutron beam generator design may require adaptive configurations to various structures and components to make them suitable for the rigors of a rocket launch and subsequent spaceflight environment. However, the fundamental principles described herein for creation of a fission based rocket engine may be achieved by provision of a suitably adapted neutron beam generator device.
In the foregoing description, for purposes of explanation, numerous details have been set forth in order to provide a thorough understanding of the disclosed exemplary embodiments for the design of a nuclear thermal rocket engine operable in sub-critical mass fissile conditions. However, certain of the described details may not be required in order to provide useful embodiments, or to practice selected or other disclosed embodiments. Further, for descriptive purposes, various relative terms may be used. Terms that are relative only to a point of reference are not meant to be interpreted as absolute limitations, but are instead included in the foregoing description to facilitate understanding of the various aspects of the disclosed embodiments. And, various actions or activities in any method described herein may have been described as multiple discrete activities, in turn, in a manner that is most helpful in understanding the present invention. However, the order of description should not be construed as to imply that such activities are necessarily order dependent. In particular, certain operations may not necessarily need to be performed precisely in the order of presentation. And, in different embodiments of the invention, one or more activities may be performed simultaneously, or eliminated in part or in whole while other activities may be added. Also, the reader will note that the phrase “in an embodiment” or “in one embodiment” has been used repeatedly. This phrase generally does not refer to the same embodiment; however, it may. Finally, the terms “comprising”, “having” and “including” should be considered synonymous, unless the context dictates otherwise.
It will be understood by persons skilled in the art that various embodiments for novel nuclear thermal rocket engine designs utilizing sub-critical mass fission of a selected actinide fissile material have been described herein only to an extent appropriate for such skilled persons to make and use such nuclear thermal rocket engine. Additional details may be worked out by those of skill in the art for a selected set of mission requirements and design criteria, such as whether the mission is manned or unmanned, (e.g., whether any necessary radiation minimization or radiation shielding may be required). Although only certain specific embodiments of the present invention have been shown and described, there is no intent to limit this invention by these embodiments. Rather, the invention is to be defined by the appended claims and their equivalents when taken in combination with the description.
Importantly, the aspects and embodiments described and claimed herein may be modified from those shown without materially departing from the novel teachings and advantages provided, and may be embodied in other specific forms without departing from the spirit or essential characteristics thereof. Therefore, the embodiments presented herein are to be considered in all respects as illustrative and not restrictive or limiting. As such, this disclosure is intended to cover the structures described herein and not only structural equivalents thereof, but also equivalent structures.
Numerous modifications and variations are possible in light of the above teachings. Therefore, the protection afforded to this invention should be limited only by the claims set forth herein, and the legal equivalents thereof.
