Abstract:

AIRCRAFT NUCLEAR PROPULSION PROJECT QUARTERLY PROGRESS REPORT

Authors:

R. C. Briant Director, ANP Project

December 10, 1951

The search for a nonoxidative high- temperature fluid other than sodium which would be suitable as a reactor coolant has lead to the proposed use of fused fluoride salts containing uranium (Sec. 1). The resulting circulating-fuel reactor would have the important advantage of eliminating a heat-transfer stage within the reactor core. Preliminary design studies of such a reactor indicate that a 3,5-ft- diameter beryllium oxide — moderated circulating-fuel reactor will produce around 350 Mw at a maximum temperature of 1500°F. The design point of the aircraft incorporating this reactor is Mach 1.5 at 45,000 ft. Performance and weights of the airplane, reactor, shield, engines, and radiators are being explored.

Although the Aircraft Reactor Experiment (Sec. 2) was originally intended as a prototype of the sodium- cooled quiescent-liquid-fuel aircraft reactor, substantial portions of the equipment would be applicable to the circulating-fuel reactor. Consequently, construction, design, and procurement of equipment have proceeded on the original schedule. The building facility itself is 80% complete and most components are on their design or delivery schedule.

Reactor physics calculations (Sec. 3) have been devoted primarily to the statics, and to some lesser extent the kinetics, of the circulating-fuel reactor. The minimum uranium investment in the beryllium oxide—moderated circulating-fuel reactor has been determined as 69 lb of which 26 lb is in the core. (The uranium investment in a water-moderated circulating-fuel reactor is somewhat lower because of the better moderating capacity of water, but design considerations appear to favor the use of BeO.) The percent thermal fissions of this BeO-moderated reactor is 52, i.e., the reactor lies between epithermal and intermediate. Kinetic studies, which have shown the practicality of the conservation of delayed neutrons, imply that the design of the reactor must provide for as high a fuel volume as practical, possibly even at the expense of the uranium investment.

A second mockup of the General Electric “direct cycle” reactor has been assembled (Sec. 4). The critical mass of this assembly was 90 lb of uranium.

The cross-sections of iron and beryllium are being determined for refined reactor physics calculations (Sec. 5), Measurements of fast neutrons in the 5-Mev Van de Graaff give an average total cross-section for iron which varies from 2.5 barns at 0.6 Mev to 3.2 barns at 3.6 Mev. Upper limits for the (n,2n) cross-section has been determined in one case to be ^ 0.56 barn and in another to be < 0.26 barn. Final adjustment are being made on the time-of-flight neutron spectrometer with which resolutions of at least 1.2 /lxsec per meter are anticipated.

The development of reactor plumbing and associated hardware, while largely concerned with that for a sodium- cooled reactor, has been redirected to the requirements of the circulating- fuel reactor (Sec. 6). Valves, pumps, seals, heat exchangers, and related equipment are being developed for both fluids for operation at reactor temperatures, i.e., up to 1500°F. However, extensive development with the fluoride fuels has first necessitated a program for their manufacture, purification, and handling; hence, to date the experimental work with fluorides has been somewhat limited. A frozen- fluid-seal centrifugal pump which was successful with sodium is being modified for use with the fused fluorides. A centrifugal pump with a gas seal, however, has already been used to pump the fluorides satisfactorily for short periods. Valve tests in sodium systems show significant increase in torque with time although several combinations of metals appear to resist se1f-we 1ding. A NaK to NaK heat exchanger has now operated for 550 hr with a maximum temperature of 1200°F.

The objective of the Oak Ridge National Laboratory in the national Aircraft Nuclear Propulsion Program is the development of high-performance reactors, i.e., for supersonic propulsion. This implies the use of liquid- coolant systems which require small core sizes and have good heat-transfer characteristics. The specific objective of the ORNL-ANP project is, accordingly, the exploitation of nonoxidative high- temperature fluids. This line of research avoids not only the high pressures associated with some cycles but also the oxidation inherent to others. A sodium-cooled quiescent- liquid- fuel reactor was the Laboratory’s first considered proposal as a reactor with potentialities for supersonic flight. However, the almost prohibitively difficult task of assuring the safety of a sodium-cooled and water-shielded reactor, as well as the limitations of a unifunctional coolant, has enhanced the search for a more versatile, and less inflammable, coolant. As a result of this search, ORNL has turned toward the use of fused fluoride salts (generally a ternary or quaternary system composed of uranium fluoride and a mixture of two or three alkali fluorides or beryllium fluoride) — not for just the heat-transfer medium, however, but as the reactor fuel as well. In addition to the advantageous physical properties of the fused salts (although not so good as sodium from the standpoint of heat transfer) and their noninflammability in air and water, an important advantage of such a circulating-fuel reactor is that it eliminates a heat-transfer stage within the reactor core.

The advantage to be gained from separating core and heat exchanger cannot be overemphasized. Although it is perhaps possible to conceive of a small and at the same time a well- designed 500 Mw heat exchanger, the requirements of these entities are so different that it is difficult to design the heat exchanger into the reactor without severely penalizing both. Furthermore, whereas the sodium- cooled reactor required considerable departure from conventional design practice to remove approximately 200 Mw from a 3-ft core, it now appears feasible, from a fabricational and fluid-flow standpoint, to remove 350 Mw from a 3%-£t core employing circulating fuel.

The circulation of fissionable material external to the core directs attention to design studies of the entire system — reactor, radiator, engine, etc. Although the circulating- fuel system has not been studied sufficiently long to ensure the performance of the resulting aircraft, the more outstanding problems of this cycle have been appreciated. The preliminary design studies which have been initiated for the exploration of a supersonic airplane application df a circulating- fuel reactor are encouraging. However, performance and weights of the airplane, reactor, shielding, engines, radiators, and other components are so interrelated that it was found necessary to make studies of the overall arrangement. The studies have not yet been completed, and it is not possible to to draw conclusions at this time. However, the arrangement being studied may be cursorily described.

AIRPLANE AND OVERALL ARRANGEMENT

The airplane visualized is designed for a speed of Mach 1.5 at 45,000 ft, with a gross weight of approximately 350,000 lb, an L/D ratio of approximately 6.5, and a wing loading of approximately 70 lb/ft2. A divided shield is employed with gamma and neutron shielding about the crew compartment and neutron shielding about the reactor. Six turbojet engines are arranged in the fuselage in a circle aft of the reactor-shie1d assembly, and circulating fuel is ducted directly from the reactor to the engine radiators, thereby eliminating the weight and temperature loss associated with the use of intermediate heat exchangers.

REACTOR

Water-moderated and solid-moderated reactors have been explored, and tentatively, both types appear to be feasible. However, the water-moderated and -reflected reactor would, for this application, require a watercooling system capable of disposing of approximately 30,000 kw with a small radiator temperature difference and a small air temperature rise. The small temperature difference would require very large radiators, and the small air temperature rise would entail very large cooling air flow requirements. Further, the use of water as a moderator involves a potential hazard associated with rapid steam formation if a failed fuel tube should permit abrupt mixing of fuel and water. Accordingly, the studies outlined here are based on the use of BeO as a moderator and reflector. Moderator heat is removed by the circulating fuel and is employed usefully in the propulsion cycle. By proper allocation of coolant flows and coolant-tube surface areas, it is possible to maintain maximum coolant- tube wall temperatures only slightly in excess of the circulating-fuel maximum temperature. To permit removing reflector heat with a high radiator temperature difference, the design under study postulates cooling the reflector with a non-fuel-bearing fluoride mixture. This reactor conforms to the specifications given in Table 1.1.

Table 1.1

Features of the Circulating-Fuel Aircraft Reactor

Size 3.5-ft sphere
Moderator BeO
Moderator ~65%
Fuel ~ 33%
Structural ~ 2%
Design point
Power ~ 325, 000 Btu/sec
Reactor fuel inlet temp. ~1000‘F
Reactor fuel outlet temp. ~1500°F

 

A circulating-fuel type reactor should be inherently capable of supplying propulsive power by means of a turbojet cycle or either of several vapor cycles. The relative superiority of these cycles has not yet been completely established, but, pending comparisons, the use of the turbojet cycle has been presumed. Preliminary optimization studies have indicated that a compression ratio in the region of 6.1, with a turbine inlet temperature of approximately 1250°F, is desirable. Further increases in the compression ratio would diminish the radiator weight at the expense of engine weight, and further increases in turbine inlet temperature (with fixed reactor conditions), would favor turbojet weight to the detriment of the radiator log mean temperature difference and radi ator weight. More rigorous evaluations of these component weights will be required before the turbojet specifications indicated here can be accepted with any degree of finality.

Installational convenience favors the use of a small number of relatively large engines, whereas engine development and availability considerations favor the use of a larger number of smaller engines. It is difficult to predict at this time the size of engine that could be made available when an airplane of the type studied would require these engines, but it appears that the validity of the overall study is relatively insensitive to the number of engines presumed. Therefore the use of six engines conforming to the general specifications listed in Table 1.2 has been assumed.

Table 1.2Performance of Engines at Design Point, Mach 1.5 at 45,000 ft
Thrust 8200 lb
Maximum diameter 61 in.
Compression ratio 6.1
Turbine inlet temperature 12S0°F
Equivalent sea»level airflow 665 lb/sec

 

 

Calculations have indicated a speci fic impulse of approximately 30 and overall efficiency of 30%. This latter is defined as the ratio of net thrust horsepower to reactor thermal horsepower.

RADIATORS

Radiators with heat-transfer capacity per unit volume in excess of the best obtained to date will be desired for any liquid-cycle supersonic nuclear-powered airplane. Arrangements are being made to receive the advice and consultation of established heat- exchanger manufacturers relative to the design of radiators for such appli- cations. This advice will not be available for at least several months, however, and it was necessary to achieve a preliminary radiator design in order to permit related phases of the study to continue. The design conceived of, which may be far from optimum, involves a radial grouping of rectangular-shaped banks about the engine centerline between the compressor and turbine. Liquid-conveying tubes are passed through, and normal to , closely spaced sheet, fins. The flow pattern contemplated is counter- current , as dictated by the temperatures of the two fluids. The arrangement entails the use of dividin’g baffles between adjacent banks, and the installation of by-pass valves in these baffles will permit controlling turbine inlet temperature while a substantially constant liquid temperature is maintained.

SHIELDING

Shielding calculations currently are in process, and it is not possible to make a quantitative description of the shielding at this time. Initial studies have indicated (see “Circu- lating-Fuel-Reactor Shields” in Sec. 8), however, that the circulating fuel in the radiators and the associated plumbing may remain unshielded provided that sufficient neutron and gamma shielding is placed about the crew compartment, and also provided that the payload, and possibly other components, are protected by local shielding.

It currently is contemplated that the reactor shield will include only hydrogenous material.

ACCESSORY SYSTEMS

Accessory circuits are required to permit cooling the reflector and shield and to provide power for pumps and general aircraft accessory demands. As it appears that many of the accessories are favored by the use of variab1e-speed drives, the overall accessory system contemplated involves the use of individual pneumatic turbines for power supply. The energy is supplied, and reflector cooling is achieved, by passing compressor bleed- off air through the reflector radiators and then to the accessory turbines and to parallel propulsive nozzles. The accessory turbines are controlled by means of variable-area discharge nozzles, which provide some net thrust after imparting energy to the accessory turbines. Preliminary analysis of this system indicates that it not only provides reflector cooling and accessory power adequately, but that the reflector cooling system, when considered as an open Brayton cycle power plant, has a favorable cycle efficiency and power-weight ratio. The use of individual accessory drive turbines permits supplying the accessories with power and speed control without special mechanical, hydraulic, or electrical transmission systems.

The shield-water radiator is cooled by low-pressure compressor bleed-off air, and the air is then discharged through a propulsive nozzle.

 

 

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