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NERVA Development Status The Space Congress® Proceedings 1967 (4th) Space Congress Proceedings Apr 3rd, 12:00 AM NERVA Development Status C. M. Rice NERVA Program Manager, Aerojet-General Corporation W. H. Esselman Deputy Project Manager, NERVA Project, Westinghouse Astronuclear Laboratory Follow this and additional works at: https://commons.erau.edu/space-congress-proceedings Scholarly Commons Citation Rice, C. M. and Esselman, W. H., "NERVA Development Status" (1967). The Space Congress® Proceedings. 3. https://commons.erau.edu/space-congress-proceedings/proceedings-1967-4th/session-18/3 This Event is brought to you for free and open access by the Conferences at Scholarly Commons. It has been accepted for inclusion in The Space Congress® Proceedings by an authorized administrator of Scholarly Commons. For more information, please contact [email protected]. NERVA DEVELOPMENT STATUS by C. M. Rice NERVA Program Manager Aerojet-General Corporation Sacramento, California w. H. Esselman Deputy Project Mana ger, NERVA Project Westinghouse Astronuclear Laboratory Pittsburgh, Pennsylvania INTRODUCTION producing the required engine thrust. During the past few years significant accomplish­ A small portion of the core exit hot gas is ments have been made in nuclear rocket development. drawn off at the hot bleed port located in the It is the purpose of this paper to review this progress convergent section of the nozzle and mixed with and to highlight the present status of the NERVA engine cold diluent extracted from the dome end of the development. NERVA is part of the ROVER nuclear pressure vessel. This fluid is used to power the rocket engine program which was initiated at the Los turbine that drives the turbopump. The hydrogen Alamos Scientific Laboratory in 1955. Figure 1 traces mass flow rate to the turbine is controlled by the key accomplishments of this development program the turbine power control val ve. Turbine flow is from the beginning through the demonstration of feasi­ exhausted through the two exhaust lines . bility, to the present phase of advancing the tech­ nology and extending performance. Under operating conditions, the temperature of the hydrogen entering the reflector is approxi- The initial progress achieved by Los Alamos mately 130°R. The reflector consists of 12 · on the conceptual reactor design and fuel element beryllium sectors, each containing a control drum. development was rapid and, by 1959, the KIWI series Boral plates in these drums supply the poison for of reactor tests demonstrated the significant per­ reactor control. T11 e hydrogen cools the various formance and potential of nuclear rocket engines parts of the reflector, and the gas temperature and stimulated interest in the development of a reaches approximately 220°R before passing through flight-type engine. The NERVA (Nuclear Engine for the core. The NRX-A reactors are right circular Rocket Vehicle Applications) Program was initiated cylinders and produce 1120 Mw of power at nominal in 1961 . This effort, under the direction of the full power. In the core, nuclear energy increases Space Nuclear Propul sion Office of J\TASA and the AEC, the gas temperature by more than 4ooo0 R. The core is being performed by the Aerojet-General Corporation consists of graphite fuel elements combined into as the prime contractor and the Westinghouse Electric clusters. Corporation as the principal subcontractor for the nuclear subsystem development. The KIWI development The NRX reactor tests and the first engine program demonstrated feasibility and proof-of­ system test (NRX/ EST) were performed at the Nuclear principle of the nuclear rocket reactor. The NERVA Rocket Development Station (NRDS) with the system Program is intended to extend these principles to in the upfiring position shown in Figure 3. The practical application in the development of a system test assembly is mounted on the test car as shown that would withstand the loads, environment, and and is connected to the test cell. For the NRX operating requirements of flight. The KIWI and NERVA tests the turbopump is auxiliary gas driven by a reactor programs have been closely coordinated to facility installed turbopump and for the NRX / EST, provide a continuing, logical development, and the the turbopump was mounted on the test car and was chronology of progress clearly highlights the note­ driven by the hot b leed gas from the nozzle. worthy advance that has been achieved in our basic technological understanding of the operating All future engine system tests starting with potentials and characteristics of the nuclear rocket the testing of the first experimental engine (XE-1) engine. in the fall of 1967 will have their components in flight oriented configuration and will be tested Figure 2 shows a schematic of the technology in the down-fired position in the Engine Test Stand engine which is designed as a ground test system at NRDS. An important capability of this test stand that delivers approximately 55,000 pounds of thrust. is altitude simulation and consequently a significant Liquid hydrogen (LH ) is delivered to the turbopump 2 step forward in the knowledge of nuclear rocket by dewar pressure. T1-. e turbopump increases the LH 2 engine operation in space will be achieved in this pressure to 940 psia and provides approximately 75 testing. Performance margin evaluations wi ll be pounds/second through the pump discharge line to the conducted with the two experimental engines pre­ nozzle inlet torus. At this point a small fraction sently scheduled in the NERVA technology develop­ of the LH is by-passed a.round the nozzle coolant 2 ment program. tubes to cool the nozzle to pressure vessel closure bolts . The remainder of the LH flows through the 2 Any appraisal of progress implies a comparison nozzle U-tubes, through the outer reflector of the with the developmental status at some previous key reactor core and into the dome end plenum of the date. The period of early 1964 has been selected pressure vessel . Here the hydrogen is reversed in as a base point for this paper. Reflecting back direction and passed through the reactor fueled to this date we can recall a great many feasibility section where it is heated and into the thrust questions. Some of these were of second order, but chamber formed by the convergent section of the a number of key questions existed. These are: nozzle. The hot hydrogen from the thrust chamber accelerates through the nozzle throat thereby 18-55 1) Can a reactor be constructed that will of 10.5 minutes. This test was the first of the produce optimum propellant temperatures Phoebus series of tests directed toward a higher and maintain structural integrity and thrust and performance reactor. At the present reliability for useful operating periods? time, the experience on ROVER systems tested since 1964 includes over 100 minutes of operation 2) Will the reactor have multiple restart at or near full operating power and temperature. capability? 3) Will the control of the engine be inherently RESTART CAPABILITY stable, simple and reliable? In addition to the endurance tests which 4) Will engine performance be predictable by were conducted, the table in Figure 4 shows starts use of analog and digital models? to power conditions that have been achieved. The first restart was achieved on KIWI B4E and, since 5) Will non-nuclear hardware such as the turbo­ that time, experience has been gained on 21 starts-- pump, nozzle, valves, lines, instrumentation, 1 on KIWI B4D, 2 on KIWI B4E, 2 on NRX-A2, 3 on etc. , be reliable for useful operating NRX-A3 , 1 on Phoebus lA, 10 on NRX/EST and 2 on periods? NRX-A5. The prime purpose of the restart capability REACTOR ENDURANCE requirement was for ground testing so that extended endurance data could be obtained with limited In early 1964 the structural integrity of the facility hydrogen supplies. In addition, various reactor had not been demonstrated . The core vibration engine system tests were required and it was questions introduced by the KIWI B4A test were not essential that a single reactor be used for those completely eliminated. The cold flow tests on KIWI experiments. and NRX-Al were designed to demonstrate that the problem was understood and corrected. At that time A number of these restarts were made under no power test had been conducted on the reactor conditions worthy of particular note. The shutdown design principle selected for the NERVA development. on NRX-A3 was very severe because flow to the reactor was inadvertently lost while the reactor Figure 4 lists the key tests that have been was at full power. The reactor was scrammed and conducted since that period and the cumulative time the temperature transient to which the test article at nominal full power. In 1964 the power tests was subjected was very large. A thorough analysis conducted on KIWI B4D, NRX-A2 and KIWI B4E showed indicated that the reactor integrity, while some­ that the structural problem with the reactor had been what impaired, was capable of a restart and that corrected and operation was achievable at high chamber no impairment to the nozzle or other system temperatures for significant periods of time. It components was detected. Restart was demonstrated then became necessary to show that the system would on 20 May 1965. operate for useful mission times. A restart of interest is Run 2 of EP-IIB on Early estimates of .required engine operating NRX/EST. During previous restarts, the engine times for useful missions varied up to twenty minutes. component material temperatures were ambient and Later mission studies indicate times up to 40 minutes the hydrogen was heated by the stored energy in for the more ambitious missions ; however, the nominal these components prior to entering the core. The operating time for a favorable Mars mission is in the question arose, could the engine system be subjected 20-30 minute range and the operating time for a very to an excessive temperature transient similar to useful lunar mission is 10 minutes for the large size that expected in space? To investigate this point, (200-250 K thrust) NERVA engine and 20 minutes for a the outer reflector was cooled to 6o0 R prior to 55,000 pound thrust engine.
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