Abstract:
A fission based nuclear thermal propulsion rocket engine. An embodiment provides a source of fissionable material such as plutonium in a carrier gas such as deuterium. A neutron source is provided, such as from a neutron beam generator. By way of engine design geometry, various embodiments may provide for intersection of neutrons with the fissionable material injected by way of the carrier gas, while in a reactor provided in the form of a reaction chamber. Impact of neutrons on 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 an 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.

Description:
COPYRIGHT RIGHTS IN THE DRAWING 
     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. 
     RELATED PATENT APPLICATIONS 
     None. 
     TECHNICAL FIELD 
     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. 
     BACKGROUND 
     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 reduces not only the payload to launch weight, but also 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; available at website: ntrs.nasa.gov/search.jsp?R=20140016814. As he notes, the Rover/NERVA program (1955-1973) 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 gas 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. 
     Thus, there remains a need to provide a design for a high specific impulse nuclear thermal propulsion rocket engine that simultaneously resolves two or more of the various practical problems, including (a) minimizing the weight of consumables (such as chemical fuel constituents) on a per payload pound basis; (b) 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; (c) providing for power control, especially as related to power generation amounts at any given time, by providing for throttling of the fission reaction; and (d) providing a high specific impulse, as compared to prior art rocket engines for space vehicles. 
     SUMMARY 
     A novel fission based nuclear thermal propulsion rocket engine has been developed, which, in various embodiments, significantly enhances the specific impulse provided by the propulsion system. The rocket engine design provides source of fissionable material such as plutonium in a carrier gas such as deuterium. 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 gas into a reactor provided in the form of a reaction chamber. 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 an 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 use of such a design in a second stage of a launch system would result in all radioactive fission products being exhausted into the vacuum of space. 
     Moreover, recent developments in neutron beam generators has made 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. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWING 
       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: 
         FIG. 1  is a partial cross sectional view for an embodiment of a rocket engine, showing a reactor in the form of a reaction chamber with a restrictive throat forming an outlet which leads to an expansion nozzle, and a neutron beam generator provides a beam of neutrons into the reactor to intersect with actinides injected into the reactor with a first fluid, and also showing injection of a second fluid which is provided to provide thrust for expansion due to heating in the reactor, as well as diagrammatically depicting turbopumps for providing both a first fluid and a second fluid to the reactor. 
         FIG. 2  is a partial perspective view for an embodiment of a rocket engine, showing the components just mentioned with respect to  FIG. 1  above showing use of a neutron beam generator that provides a beam of neutrons into the reaction chamber to intersect with actinides injected into the reaction chamber with a first fluid, and also showing injection of a second fluid which is provided to provide thrust for expansion due to heating in the reaction chamber, as well as conceptually depicting turbopumps for providing both a first fluid and a second fluid to the reaction chamber, and also showing use of a gas generator for developing high pressure combustion gases for driving a fuel turbopump and a thrust fluid turbopump. 
         FIG. 3  is a top view of an embodiment for a reaction chamber, showing a location for a neutron beam generator, and also showing coolant passageways which run along the outer surface of the sides and top of the reaction chamber to a header which collects the heated second fluid and from which the second fluid is injected into the reaction chamber. 
         FIG. 4  is a bottom view of a an embodiment for a rocket engine, taken looking up at line  4 - 4  of  FIG. 1 , showing the second fluid distributor at the outlet of the expansion nozzle that is used to distribute the second fluid to coolant passageways along the walls of the expansion nozzle and the reaction chamber, and also showing the outlet of the reaction chamber. 
         FIG. 5  is a perspective view of an embodiment for a rocket engine, showing a neutron beam generator mounted to a reaction chamber, and an expansion nozzle for receiving heated gases from the reaction chamber, as well as showing coolant passageways on outer surfaces of the reaction chamber and on the outer surfaces of the expansion nozzle. 
         FIG. 6  is similar to  FIG. 1  above, showing a partial cross sectional view for an embodiment of an rocket engine, depicting a reaction chamber with a restrictive throat forming an outlet which leads to an expansion nozzle, and a neutron beam generator provides a beam of neutrons into the reaction chamber to intersect with actinides injected into the reaction chamber with a first fluid, and also showing injection of a second fluid which is provided to provide thrust for expansion due to heating in the reaction chamber, as well as diagrammatically depicting use of turbopumps on a common shaft or gear box for providing both a first fluid and a second fluid to the reaction chamber, and also having, driven by the same turbine system, an electrical generator for generating electrical power for supply to the neutron beam generator. 
     
    
    
     The foregoing figures, being merely exemplary, contain various elements that may be present or omitted from a final configuration for 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. Further, like features in various nuclear thermal rocket engine designs may be described using like reference numerals, or other like references, without further mention thereof. 
     DETAILED DESCRIPTION 
     Attention is directed to  FIGS. 1 and 2  of the drawing.  FIG. 1  shows in a partial cross sectional view an embodiment for a nuclear thermal rocket engine  10 , showing a reactor  12  having a tubular portion  13  and a restrictive throat  14  forming an outlet  16  which leads to an expansion nozzle  20 . A neutron beam generator  22  is provided to direct a beam of neutrons  24  in the reactor  12 . A first fluid storage compartment  26  for storage of a first fluid  28  such as deuterium (may be depicted as  1 D 2  or as  2 H) is provided. A second fluid storage compartment  30  is provided for storage of a second fluid  32  such as hydrogen H 2 . A third fluid storage compartment  34  is provided for storage of third fluid  36  such as oxygen (O 2 ). In an embodiment, the third fluid  36  may be used for reaction with a second fluid  32  such as hydrogen (H 2 ) in a gas generator  38  (also marked as GG in  FIGS. 1 and 2 ), to generate a high pressure fluid  40  (e.g., combustion gases) for driving a turbine  42  in a turbopump  44 , as further discussed below. In such case, after pressure reduction through turbine  42 , remaining low pressure water vapor may be discharged overboard as indicated by reference arrow  46 . 
     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 a particulate form. 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 powdered fuel may comprise a selected actinide compound. In an embodiment, the powdered fuel 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 ( 239 Pu). In an embodiment, the fuel F may be supplied as a fissile material in the form of uranium 235 ( 235 U). In various embodiments, the selected fissile material providing fuel F, before injection into the reactor  12 , may be provided in particulate form. 
     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 gaseous form, while carrying a particulate actinide fuel F therein. However, as shown in  FIGS. 1 and 2 , in various embodiments, at time of injection, the fuel rich mixture  52  may be in liquid form, while carrying a particulate actinide fuel F therein. In an embodiment, a first set of fuel injectors  60  may be oriented at a selected inwardly directed angle alpha (α) that directs a rich fuel mixture  52  stream toward a reaction zone  62  wherein energetic neutrons  24  from neutron beam generator  22  collide with atoms of fissile material in fuel F as found in the rich fuel mixture  52 , to cause fission of atoms of fuel F, with resultant heat release. In any event, a neutron beam generator  22 , which is further discussed below, is configured to direct neutrons  24  to collide with at least some of the fuel F fissile material in the reaction zone  62 , wherein the neutrons  24  and the fissile material interact to thereby effect fission of at least some of the atoms of the fissile material in fuel F and release heat. 
     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 ( 239 Pu) provided may be between about thirty parts per million (30 ppm) and about one hundred and twenty parts per million (120 ppm), by weight, in the first fluid  28 . In an embodiment the amount of plutonium 239 ( 239 Pu) provided may be between about sixty parts per million (60 ppm) and ninety parts per million (90 ppm), by weight, in the first fluid  28 . In an embodiment of rocket engine  10 , plutonium 239 ( 239 Pu) may be provided at about sixty parts per million (60 ppm), by weight, in the first fluid  28 . In various embodiments, the first fluid  28  may be provided as deuterium (may be shown as either  1 D 2  or  2 H). 
     In various embodiments, the first fluid  28  may include one or more isotopes of hydrogen. In an embodiment, the first fluid  28  may include deuterium. In an embodiment, the first fluid  28  may primarily be deuterium ( 2 H). In an embodiment the first fluid  28  may include essentially only deuterium ( 1 D 2 ). In an embodiment, the first fluid  28  may include at least some tritium ( 1 T 3 ). In an embodiment, the first fluid  28  may 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. 
     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 (H 2 ) 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 send 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  FIG. 3 , the reactor  12  includes reactor coolant passageways  86  on the reactor external surface  88 . In this manner, an extremely cold fluid, e.g. liquid hydrogen, is utilized as a coolant for the reactor  12  by passage of the second fluid  32  through the reactor coolant passageways  86 . Thus, the rocket engine  10  may utilize the second fluid  32  as a coolant by way of the passage of second fluid  32  through the nozzle coolant passageways  76  and through the reactor coolant passageways  86 , before injection of the second fluid  32  into the reactor  12 . 
     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  FIG. 6 , a different design for a thrust fuel turbopump  144  may be provided. In such design, the thrust fuel turbopump  144  may provide pumping of second fluid  32  by pump impeller  145 , while also additionally providing an electrical generator  146 . In an embodiment, the electrical generator  146  may be configured to generate electrical power, and supply the same via electrical power lines  148  and  150  to neutron beam generator  22 . In an embodiment a thrust fluid turbopump  144  may further include a fuel turbopump  160 , for receiving first fluid  28  from the first fluid storage compartment  26  and providing the first fluid  28  under pressure to reactor  12 . In an embodiment, the thrust fluid turbopump rotor  145 , the fuel turbopump rotor  161 , and the electrical generator  146  may all be driven by a gas turbine  162  on a common shaft  164  or via gearbox from a common shaft  164 . 
     In various embodiments for a rocket engine  10  or  10 ′ or the like, using nuclear thermal heating of a low molecular weight gas such as hydrogen as described herein, a rocket engine may be provided that has a specific impulse in the range of from about 800 to about 2500 seconds. In various embodiments using nuclear thermal heating of a low molecular weight gas such as hydrogen as described herein, a rocket engine may be provided that has a specific impulse in the range of from about 1000 to about 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 plutonium carrying deuterium gas to the reactor  12  for fission of at least some of the plutonium, a fuel turbopump  54  may be provided. In an embodiment, liquid deuterium i.e. a cryogenic liquid, 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 deuterium (and plutonium carried therewith) 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 liquid deuterium at relatively low inlet pressure, and deliver the deuterium 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 hydrogen gas 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 1×10 11  and 5×10 11  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. 
     However, 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 1×10 13  and 5×10 13  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 phoenixnuclearlabs.com. Other vendors currently provide different designs. For example, Gradel Group, 6, Z.A.E. Triangle Vert, L-5691 ELLANGE, Luxembourg (see website at gradel.lu/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:
 
 1   D   2 + 1   T   3 → 2   He   4 (3.5 MeV)+ 0   n   1 (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.