Patent Number: 056365128
Section: summary

FIELD OF THE INVENTION This invention relates to nuclear rocket engines generally and more specifically to a nuclear rocket feed system that incorporates an auxiliary power cycle. BACKGROUND OF THE INVENTION Conventional nuclear rocket engines employ a nuclear fission reactor to heat the rocket propellant, typically hydrogen, to extremely large temperatures. The hot hydrogen is then expelled from a nozzle into space at supersonic speed to create thrust for the rocket. To conserve nuclear fuel and propellant, space mission operations will typically only require short duration engine firings. The reactor is turned on for a brief period to generate thrust to propel the rocket to a cruise velocity in space and then the reactor is shut down. Shutting down nuclear rocket engines during the space mission has presented many design challenges. One challenge results from the fact that nuclear reactors cannot be immediately turned off. Delay neutrons and daughter products of the fission reaction generate power long after the reactor ceases to operate. This energy or heat must be removed from the rocket to prevent overheating and destruction of the engine. In addition, the engine feed system (pumps and turbines) must be shut down as the reactor power decays to throttle propellant flow and to prevent the pumps from surging. Shutting down the feed system, however, makes it makes it extremely difficult to remove the reactor heat from the engine. Another challenge created by shutting down rocket engines in flight is cooling the nuclear reactor. Byproducts of the nuclear reaction (waste heat) continuously heat the components of the reactor during engine firing and long after the engine has been shut down. To solve this problem, liquid or gaseous hydrogen (apart from the actual propellant) is typically used to cool the reactor. The hydrogen is directed through the reactor, which transfers some of its heat to the hydrogen, and is then expelled from the rocket into space. This process continues until the reactor temperature has been brought down to a safe level. One problem with this method is that cooling the reactor can take a long time (from a few hours to a few days). Thus, an enormous amount of hydrogen must be stored in the rocket to cool the reactor. This large volume of hydrogen increases the weight of the rocket which decreases mission performance (payload/initial mass) and increases mission cost. To decrease the amount of hydrogen needed to cool the reactor, existing systems have attempted to alternate between undercooling and overcooling the reactor. In this scheme, the reactor is allowed to heat up until it reaches a very high temperature and is then quickly cooled down with extremely cold propellant. Hydrogen is conserved because it is not continuously pumped through the reactor and into space. Alternating between undercooling and overcooling the reactor, however, can create thermal shocks that damage reactor components and create flow instabilities, thereby decreasing the life of the nuclear engine. SUMMARY OF THE INVENTION The present invention solves these problems of shutting down the engine during flight and cooling the reactor throughout the space mission. To accomplish this, a nuclear rocket engine includes a primary feed system for pumping rocket propellant from a propellant source to a nuclear reactor and an auxiliary feed system coupled to the primary feed system. The auxiliary feed system can be configured into a high thrust mode for withdrawing heat from the engine when the reactor is operating at full power, a low thrust mode for throttling propellant flow and radiating heat into space during reactor shutdown and a zero thrust mode for cooling the nuclear reactor and generating electricity for the rocket's auxiliary power requirements during the remainder of the mission. In the high thrust mode, the auxiliary feed system includes a bypass line with an inlet coupled to a recycling port in the primary feed system, an outlet coupled to the propellant source and means for withdrawing heat from propellant flowing along the bypass line. A recuperator is coupled to the primary feed system for transferring heat from the hot propellant in the reactor to the cool propellant from the propellant source. A portion of this now heated propellant is bled into the auxiliary feed system after passing through the recuperator to withdraw heat from the engine. This heat can be converted into electricity to power other operations on the rocket or discharged into space to release heat from the engine or control the attitude of the rocket. Preferably, the auxiliary feed system is a Brayton power cycle having a turbine coupled to a compressor for pumping propellant through the feed system. The heat withdrawing means comprises a space radiator for withdrawing heat from the warm propellant that has passed through the recuperator and discharging this heat into space. The heat withdrawing means may also include a motorgenerator coupled to the turbine for translating the mechanical energy of the turbine into electricity. The cooled gaseous hydrogen is then recycled back into the propellant source to maintain a suitable pressure within the propellant source. In the low thrust mode, the auxiliary feed system has an inlet coupled to a recycling port in the nuclear reactor between the fuel assemblies and the nozzle and an outlet coupled to the reactor between the reactor inlet and the fuel assemblies. In this mode, a portion of the hot propellant exiting the fuel assemblies is bled into the auxiliary feed system to discharge heat (generated by the neutron delay reactions in the fuel assemblies) into space. Some of this heat is used to drive the turbine so that the auxiliary feed system can pump the propellant through the primary feed system. In this manner, the main pumps and turbines in the primary feed system can be shut down to facilitate throttling of the propellant flow and to prevent the pumps from surging during reactor shutdown. Similar to the high thrust mode, electricity is generated from the waste heat with the motorgenerator and the cooled propellant maintains a suitable pressure in the propellant source. In the zero thrust mode, the auxiliary feed system has an inlet coupled to the reactor between the fuel assemblies and the nozzle and an outlet coupled to the primary feed system. In this mode, the nozzle is bypassed so that all of the hot propellant from the fuel assemblies flows into the auxiliary power system to discharge heat through the space radiator. The cooled hydrogen is then recycled back through the reactor to continue the cooling process. Thus, the reactor can be completely cooled without losing any hydrogen. This results in a substantial decrease in the amount of hydrogen needed for a space mission and, therefore, a substantial decrease in the weight of the rocket. Another advantage of the present invention is that reactor coolant gas is preheated before cooling the fuel assemblies in the low and zero thrust modes. Preheating the reactor coolant gas increases flow stability and proper cooling of the fuel assemblies even at low flow velocities. This increases the life of the system because thermal shocks and flow instabilities are avoided. The above is a brief description of some deficiencies in the prior art and advantages of the present invention. Other features, advantages and embodiments of the invention will be apparent to those skilled in the art from the following description, accompanying drawings and appended claims.