Abstract:
A nuclear thermal propulsion rocket engine. A reactor is provided to receive a fissionable fuel and a propellant fluid. Fuel may be transported and injected using a carrier fluid. Carrier fluid for fuel may be hydrogen or an isotope thereof. Fuel may be plutonium or selected actinide. A neutron generator is provided, and utilizes an ion generator and a target container which extends into the reactor to hold a target material therein. Neutrons may be emitted almost omni-directionally by impact of ions from the ion generator on target material. Cooling of the target container may be provided by a cooling sleeve that receives and circulates a cooling fluid, and discharges the cooling fluid after it has been heated. Fuel injectors provide fuel through fuel injector valves regulated to cycle on and off to pulse output power of the rocket engine, by timing frequency and duration of fuel feed, and regulating injected fuel amounts, to regulate the energy released over discrete time periods. The reactor receives reactants and an expandable propellant fluid such as hydrogen, and confines heated and pressurized gases for discharge out through a throat, and into a rocket engine expansion nozzle for propulsive discharge.

Description:
STATEMENT OF GOVERNMENT INTEREST 
     Not Applicable. 
     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 
     This application is a Continuation-In-Part of pending U.S. patent application Ser. No. 14/692,349, filed on Apr. 21, 2015 entitled NUCLEAR THERMAL PROPULSION ROCKET ENGINE, the disclosure of which is incorporated herein in its entirety, including its specification, drawing, and claims, by this reference. 
     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, entitled 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 fluid 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 neutron generating reaction conditions in a reactor, resulting in release of heat energy to a propellant fluid such as hydrogen which is supplied to the reactor for heat expansion and discharge therefrom. 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. In an embodiment, a nuclear reaction may be provided in an intermittent, pulsed fashion, by intermittent engagement between a fissionable material and neutrons from a neutron beam generator. By control of frequency of the pulses of energy creation from the nuclear reaction, and by control of the amount of energy released per pulse, a workable reaction containment structure can be designed and maintained to withstand the high energy pulsed nuclear reactions, as well as to maintain a sufficiently cooled containment structure, to enable the rocket engine to work during sustained power operations as necessary for provide an energetic boost to a selected payload package. 
     An advantage of the novel nuclear thermal propulsion rocket engine design disclosed herein is that use of such a design in a second or later 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. 
         FIG. 7  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 via controlled frequency fuel injectors in a first fluid, and also showing injection of a second fluid which to be expanded and thereby to provide thrust due to heating in the reaction chamber, as well as diagrammatically depicting use of the second fluid as coolant for a target chamber and target material of the neutron beam generator, where a portion of the second fluid is discharged out through a perforated wall in the target chamber, so as to maintain inflow of additional coolant by continued supply of the second fluid. 
     
    
    
     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 nuclear reactions for power generation, 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 for a rocket engine, plutonium 239 ( 239 Pu) may be provided at about sixty parts per million (60 ppm), by weight, in the first fluid  28 . In an embodiment for a rocket engine plutonium 239 ( 239 Pu) may be provided at up to about one thousand (1000) parts per million (1000 ppm), by weight, in the first fluid  28 . In various embodiments for a rocket engine, plutonium 239 ( 239 Pu) may be provided at somewhere in the range from about ninety (90) parts per million (90 ppm) and about one thousand (1000) parts per million (1000 ppm), by weight. In an embodiment fissionable fuel F may be provided at somewhere in the range from about ninety (90) parts per million (90 ppm) and about one thousand (1000) parts per million (1000 ppm), by weight, depending on the efficiency at which the fuel F is consumed in the reactor  12 . In various embodiments, the first fluid  28  may be provided as deuterium (may be shown as either  1 D 2  or  2 H). In an embodiment, the first fluid  28  may be provided as hydrogen (H 2 ). 
     In various embodiments, the first fluid  28  may include one or more isotopes of hydrogen. In an embodiment, the first fluid  28  may be primarily hydrogen. In an embodiment, the first fluid  28  may be essentially hydrogen. In an embodiment, the first fluid  28  may include deuterium. In an embodiment, the first fluid  28  may primarily be deuterium ( 1 D 2 ). 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 ,  10 ′,  200 , 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 excess of about 800 seconds. In various embodiments, specific impulse may be 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. 
     As a further example of adaptive design configurations for a neutron generator, attention is directed to  FIG. 7 , which provides details for an embodiment of a nuclear thermal rocket engine  200  which is configured for controlled pulse reaction between atoms of fissionable fuel F (diagrammatically depicted as diamonds  202  in  FIG. 7 ) and neutrons  204  generated (omni-directional, in an embodiment) from target material  206  in neutron generator  208 . In an embodiment, a neutron beam generator  208  may include a ion generator  210  and a target container  212  containing a target material  206 . In an embodiment, a target container  212  may include a target container reactor portion which extends into the reactor  12 . The neutron beam generator  208  is configured to accelerate ions, such as deuterium ions, or a combination of deuterium and tritium ions, toward the target material  206 , normally containing deuterium. When deuterium is utilized for creation of an ion beam, two deuterium ions fuse, resulting in D-D type fusion. If tritium is used for creation of an ion beam, a deuterium and a tritium ion fuse, resulting in D-T fusion. In both cases, neutrons are the by-products of the fusion reaction. Various suitable target materials are known in the art for manufacture of neutron generators. One commonly utilized design technique for selection of a target material  206  is to use, or at least coat, a target material  206  with palladium or some other metal that readily forms hydrides, or since deuterium is used, deuterides. In an embodiment, a palladium coating might be used, since palladium may store nearly one deuterium atom for each palladium atom. However, a typical target material  206  may need to remain cool in order to maintain an acceptable density of deuterium target atoms, since excess heating may cause the actual targeted atoms for neutron generation to be desorbed from a metal target in which target atoms are initially saturated. In an embodiment, in order to cool the target container  212  and the target material  206  inside, second fluid  32  such as cryogenic hydrogen H2 may be utilized. In an embodiment a second fluid  32  such as cryogenic hydrogen H2 may be injected to header  214  and thence along a cooling passageway  215  defined by cooling sleeve  216  that houses an ion beam pathway  218 . As shown in  FIG. 7 , the cooling sleeve  216  may have mounted at or near a distal end  219  a target material  206  in which target atoms (not shown) are contained. Cooling sleeve  216  may include outlets  220  from which second fluid  32  is ejected as a heated second fluid  33 , thus allowing fresh cooling of the target material  206  via incoming cool second fluid  32 , as may be supplied, in an embodiment, via turbopump  44 . In an embodiment for a fifty thousand pound thrust engine, initial simulations have found that inserting the target material  206  a distance of about ten (10) inches down into the reactor  12  should be effective in producing millions of fissions using impacts with the fissionable fuel atoms  202 . Neutron generation rates of 10 11  neutrons per second, as may be produced from current technology deuterium-deuterium (“D-D”) neutron generators would thus be sufficient to produce enough energy to power rocket engine  200 . Further, manufacturers of neutron beam generators have indicated that 10 13  neutrons per second should be obtainable with currently available technology. Consequently, a combination of 10 7  fissions per neutron, and 10 13  neutrons per second from a neutron generator, would be adequate to produce fission rates greater than about 10 19  as required to produce specific impulses over 1000 seconds. 
     In order to maintain inner sidewalls  230  of reactor  12  at a structurally workable temperature, pulsed timing of nuclear reactions maybe desirable, so that the amount of energy released over a selected time interval does not exceed the capability of cooling of second fluid  32  (as expanded and heated from the nuclear reaction) when injected via second set of injectors  94  according to a selected injection, mixing, and thermal behavior profile. To pulse the release of energy from nuclear reactions, energetic neutrons and fissionable atoms in the fuel F may be brought together intermittently. In an embodiment, use of controllable fuel injectors  260  may be desirable. In an embodiment, one or more of the fuel injector valves  262  may be utilized to control an on-off supply of fuel rich mixture  52  (e.g., via line  56 ) to reactor  12 . In an embodiment, one or more, or each of fuel injector valves  262  may be utilized for modulated control of supply of fuel rich mixture  52  (e.g., first fluid  28  having fuel F therein, via line  56 ) to reactor  12 . By intermittent supply or by modulated supply of fissionable fuel F to reactor  12 , the rate of energy release, and thus heating of propellant second fluid  32 , e.g., hydrogen, as well as heating of adjacent components (such as target container  212  and inner sidewalls  230  of reactor  12 ) is controlled. Likewise, by adjusting the energy released in each pulse, the structural integrity (e.g, yield under stress/strain at high temperature) may be set at workable limits, which will of course depend on selected materials of construction and applicable design techniques for such structural materials. 
     As depicted in  FIG. 7 , in an embodiment, the first fluid  28  from the first fluid storage compartment  26  may be hydrogen (H 2 ). The first fluid  28  may be mixed with a selected amount of fuel F, before injection into reactor  12  via injectors  260 . 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  through a fuel header  58  (not seen in  FIG. 7  but functionally the same as shown and described above), and thru fuel injector valves  262  to a set of fuel injectors  260  which 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 an actinide fuel F therein. In an embodiment, fuel injectors  260  may be oriented at a selected angle that directs a rich fuel mixture  52  stream toward a reaction zone  263  wherein energetic neutrons  204  from neutron beam generator  208  collide with atoms  202  of fissile material in fuel F as found in the rich fuel mixture  52 , to cause fission of atoms  202  of fuel F, with resultant heat release. 
     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.