Patent Number: 042467519
Section: summary

The present invention relates to thrust engines and propellant exhaust arrangements therefor and more particularly to such systems which utilize a nuclear reactor as a power source and which further utilizes a portion of the reactor assembly itself in nozzling the exhausting propellant for desired or specified thrust. A nuclear thrust or rocket device engine as described herein is one in which energy generated by a controlled chain nuclear reaction is transferred to a propellant which then develops device accelerating thrust when it is exhausted from the rocket device. The chain nuclear reaction is propagated by means of fissile material normally interstitially distributed in a modular structure, with the amount and geometric distribution of the fissile material and other factors providing critical reactivity conditions. Further, the modular structure can be elongated with longitudinal openings for flow of the propellant therethrough and consequent energization of the propellant by transfer of fission generated heat thereto. The propellant is then exhausted from the device and in order to obtain maximum specific impulse or thrust per pound of propellant it is necessary that the propellant be in the form of a gas having the minimum possible molecular or atomic weight. For this reason, hydrogen has been used as the propellant in early ground test nuclear engines and will be used in the first flight nuclear engines now undergoing final development. It is noted, however, that other gases can be used as a nuclear engine propellant if the advantages associated with such use outweigh the disadvantage of lower specific thrust. During exhaust of the propellant, the acquired heat energy is transformed into velocity energy by means of pressure induced acceleration of the propellant particles along the exhaust flow path. According to applicable flow phenomena, at subsonic velocity (compressible) gas particle acceleration is obtained by a converging exhaust flow path whereas, at supersonic velocity, (compressible) gas particle acceleration is obtained by a diverging exhaust flow path. Therefore, where supersonic propellant exit velocity is desired or necessary, a convergent-divergent flow exhaust path is required for the propellant from the reactor engine to the propellant exit area. On this basis, a so-called convergent-divergent nozzle is well known for use in providing this required exhaust path for the total propellant flow. One limiting factor, however, is that nuclear or other thrust engines are normally designed to operate at extremely high temperatures for efficiency reasons, for example the propellant as it departs from the engine reactivity region in a nuclear thrust engine can have a temperature of 4000.degree. F. or more. This factor, coupled with the fact that the hydrogen or other propellant can be or is characterized with a very high convective heat transfer coefficient as determined by the propellant velocity and pressure, can and normally does lead to problems in heat dissipation along the exhaust flow path. The heat dissipation problem is particularly acute adjacent the throat section of a convergent-divergent nozzle where the hydrogen or other propellant acquires velocity and pressure parameters which cooperatively effect a relative maximum amount of convective heat transfer from the propellant. The resultant temperature rise of the surrounding nozzle throat section which accepts such heat transfer leads to the necessity of using structure and materials in a manner that enables both required heat dissipation and specified mechanical nozzle strength to be obtained. One manner in which the exhaust system can be engineered to resolve the problem is to provide a single convergent-divergent nozzle for directing the entire propellant flow and further to utilize materials of suitable heat capacity in a throat section structure which meets both heat dissipation and strength specifications. An example of such structure is disclosed in a copending application entitled Thermal Barrier for Thrust Vehicle Nozzle and Method of Making Same filed by D. Thomas on August 17, 1962, Ser. No. 217,698, and assigned to the present assignee. There are, however, applications where structural materials are not available for specified temperature and heat conditions or where it is more advantageous to avoid the nozzle throat structure problem altogether if the avoidance does not produce unacceptable penalties in other respects. Thus, according to the broad principles of the present invention, a nuclear or other engine is provided with a plurality of elongated nuclear fuel bearing or other heat generating modules assembled together to form an elongated propellant energizing or heating region. The propellant exit of preferably all of the fueled or other modules is provided with a convergent-divergent nozzle structure so that velocity development in the total propellant flow is accomplished over parallel flow paths rather than over a single main flow path. In this manner, the overall exhaust nozzle flow path is shortened and convective transfer of heat from the propellant is accepted by a nozzle structure which can be made of the same temperature bearing base material of the engine modules and the problem of dissipating heat in the nozzle throat sections is substantially eliminated while structural integrity of the modules does not become problematical. Further, if additional velocity development is desired beyond that obtainable over the modular nozzle flow paths, a unitary divergent skirt can be provided on the engine so as to extend outwardly from the propellant modular exit area and thereby serve as a single total propellant flow path for the added propellant velocity development. It is therefore an object of the invention to provide a novel nuclear or other rocket engine and propellant exhaust system therefor wherein propellant nozzling is accomplished with improved heat transfer efficiency. It is a further object of the invention to provide a novel nuclear rocket engine and propellant exhaust system therefor wherein propellant exhaust flow is accomplished at least partly through parallel nozzle paths so as to provide improved heat transfer efficiency. An additional object of the invention is to provide a novel nuclear or other rocket engine and propellant exhaust system therefor wherein exhaust nozzle structure is comparatively shortened. A further object of the invention is to provide a novel nuclear rocket engine and propellant exhaust system therefor wherein the engine comprises a modular fissile fuel bearing arrangement and the propellant exhaust system includes nozzle exhaust structure provided adjacent the propellant exit end of each module in the engine arrangement so as to provide improved heat transfer efficiency and so as to enable the nozzle structure to be shortened comparatively. Another object of the invention is to provide a novel nuclear or other rocket engine and propellant exhaust system therefor wherein improved heat transfer efficiency is obtained through employment of a main divergent nozzle to which propellant is directed at supersonic velocity by other convergent-divergent nozzle means.