Y

AlAA 90-1954 Design of Second-Generation Nuclear Thermal Rocket Engines

S. V. Gunn

Rockwell International Rocketdyne Division Canoga Park, CA

AI AAlSAElASM ElASEE 26th Joint Propulsion Conference

July 16-1 8, 1990 / Orlando, FL

For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics 370 L'Enfant Promenade, S.W., Washington, D.C. 20024

DESIGN OF SECOND-GENERATION NUCLEAR THERMAL ROCKET ENGINES

Stan Gunn Rockwell International, Rocketdyne Division

Canoga Park, California 1

Ah s t r a c t

Thedesignsofproposedflight nuclearmket cnginesevolving in the 1960-1970 time period featured bleed cycles, rnodcrate nozzlc expansion ratios,projected core exit gas temperatures in the range of 4100"R to 4500"R, and reactor-core shield con- figurations intended to protect theengincsystem and propellant tankagc from the attendant high radiation fields. With the advances in fuel element operating temperaturcs and power densities (as demonstrated in the Phocbus, Pewee and Nuclear Furnace reactor test programs), the evolution of topping-ex- pander cycles, and the development of radiation-resistant con- trol components and high-speed, zero-NPSHpump inducers; it is now possible to dcsign advanced nuclear rocket engines fcaturing specific impulses in excess of 900 s, engine specific weights (excluding reactor core shields) approaching 8 lb of thrust/pound of engine weight, higher intrinsic system rcliabil- ity, and capable of operating, or being independently shut down, when incorporated into a multiple engine clustcr. A representative second-generation nuclear rwket engine design conccpt is prcsented. The assumed thrust requircmcnt is 100,000 lb, which not only provides a basis of direct compar- son to the earlier designs but may also k of interest to NASA and vehicle contractors studying optional space transfer ve- hicles for fast trajectory, allpropulsive, manned Mars missions.

Assumed Propulsion Requirements-. Manned Mars Mission

Thedelineation of the propulsion requirements associatcd with currcntly envisioned mission scenarios, for the accomplish- ment of manned Mars missions in the 2014-2018 time period, is strongly depcndent on trip times and upon the availability of aerobnking for Mars capture. Nuclear Thcrmal Rocket (NTR) propulsion systems offer specific impulses in excess of 900 s, which affords the mission planner an all-propulsive mcans of rcalizingrelatively short total trip times (-1 yr) with rcasonable initial mass in low earth orbit. For the purposes of this paper, the following propulsion system requirements (see Table 1) havc been assumed to typify the requirements anticipated for the aforementioned high AV missions.

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Typical Mars Mission NTR Propulsion Requirements

Selection of the engine cycle and the nozzle expansion ratio may be dictated by the maximum, steady state, core exit gas tcmpcrature. For solid core reactors, cmploying composite v

Table 1. Typical Manned Mars NTR Propulsion Requirements

* Thrust, nominal 100,MM Ibl * Performance 2900 s * Maximum weight - Full performance operating range

<18,MM Ib (withoul shield) 110% 4 50%

Emergency thrust (with one pump cut) 70% NPSH(min) One velocity head lrom

saturakd liquid in tank * Maximum operating time 2 h

Number of restatis 26 Transition, flow initiation to lull thrust

Maximum core temperature, alter heat

30 S

30s i8W'R

* Transition, 50% thrust lo cut-off

(uranium carbide, zirconium carbide inagraphiternatrix) fuelele- mcnts, coated with zirconium carbide, a mixedmean, core exlt gas ternperatureof486OoR canbeassumed. The mission rcquirc- ment of a specific impulse of at least 900 s drives the selection of an expander cycle and a nozzlc expansion ratio (E) greater that 100. However, the utilization of an E greater than 800 (see Figure 1) buysvery liuleimprovemenlinperformance. For the purposes of this paper, an E design value of 500 has been selected for the complete nozzle assembly, yielding a predicted specific impulsc of 922 s and a requisite %flow rate of 108 Ib/s.

880 920 960 1.000 1,040 Specific Impulse (5)

Figure 1. Performance of NTR Operating with High Pressure Hydrogen

Within the category of candidate expander cycles for NTR en- gines,thereexistsavariety ofchoicesofenergy sourceforheating the turbine drive gas to the desired turbine inlet temperature. For core designs incorporating internal flow loops to cool elements of the core support structure, there exists the potential of extracting

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significant energy at moderate temperatures. In addition, energy deposited in the reactor's reflector, provides an addi- tional energy source for turbine drive gas heating. Finally, the cooling of at least the convergent and throat sections, as well as a portion of the divergent section of the nozzle, affords the engine cycle designer with potential supplemental sources of beat and temperature. Figure 2 presents two candidate NTR expander cycles; to facilitate the potential for minimizing the loss in specific impulse, through heat transfer to the nozzle coolant system, the concept shown in Figure 2, configuration B has been adopted for the engine design presented herein.

Configuration A Configuration B

Figure 2. NTR Expander Cycles (Typical)

ForexpandercycleNTR engines, the powerand, bence,thesize of the reactor, required to heat the total flow of propellant to the requisite48WR, isa function ofthecore inlet temperature. For an optimized design, the selection of the turbine inlet tempera- ture is a complex itcrative process involving considerations of (1) the targeted chamber pressure, (2) the energy and tempera- ture available'in the core support Structure, and (3) the energy and temperature available in the reflector and (if required) the nozzle coolant systems.

For the purposes of this paper, it will be assumed that (1) for reasons of envelope constraints, the selected chamber pressure will be loo0 psi, (2) the core support structure can bc designed to provide the requisite turbine-drive propellant heating, and (3) a turbine inlet temperature of 1W"R can be realized.

The resultant turbine exit temperature for an assumed turbine pressure ratio of 0.8 would be expected to be approximately 950OR. However, this turbine flow mustjoin theturbine bypass flow (lOOOOR)andthenozzleandreflectorcoolantflow. Until 'a the heat pickup by the nozzle and reflector coolant flow can be calculated (see below), it will be assumed that the reflector exit temperature will be approximately 50% of the core support structureexittemperature, or500'R. Further,asafmtapproxi- mation, it will be assumed that 50% of the total propellant flow will be routed through the core support structure and 50% will beemployedas nozzleandreflectorcoolantflow. Thus, atthis point in the analysis, the core inlet temperature would be estimated tobeapproximately735'Randthecore-fuel element thermal power would be 1880 M W . If the solid core reactor to be employed in this design example of a 100,W Ibf NTR engine were to be based on the proven RoverlNerva (see Figure 3) 35 in. diameter cores (with 1870 loaded fuel elements), the resultant thermal power generation would be approximately 1 MWhaded fuel element. This value matches the power density demonstxated by both Phoebus IB and Phoebus 2A.

I' . \ Nuclear Subsvsterr

Fuel Element Internal Shield I

Assembly

Figure 3. Roverh'erva Core-Reactor

The core support structure of the Phoebus 2 reactor was com- prised of a top core support plate, counter-flow tie tubes and core support blocks. The tie tubes were installed in unfuclcd, 3/4 in. hexagonal center support elements located in the center of a cluster of six 3/4 in. hexagonal fuel elements. This cluster design offers the potential of satisfying the turbine inlet tem- perature requirements, but the requisite flow areas down the inner tie tube and back up the annular space between the inner and the outer tie tube, would take up a large percentage of the available cross sectional area of the unfueled center element. For the 1880 MW fuel element thermal power assumed herein,

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'with 325 center element, tie-tube assemblies, it appears that the requisite turbine flow can be heated to l W o R , by the transfer of200MW withacceptable pressuredrop, butthat thediameter of the outer flow tube will be of the order of 0.7 in. Thus the remaining volume between each outer flow tube and the six inner surfaces of the surrounding fuel elements is significantly reduced, which limits the effectiveness of substituting zircc~ nium hydride for the unfueled composite center element, to improve moderation and reduce the uranium loading rcquire- ments.

The heat transferred from the 7 in. thick reflector to the nozzle- reflector coolant flow is projected to be 36 MW, with an attendant increaseintemperature from 310"R to 500'R. Below the reflector, additional heat will be transferred from the total rcactor core exit gas flow to the walls of the converging- diverging nozzle. This heat loss comes at the expense of the realizable kinetic energy of the exhaust jet at the exit section of the nozzle, and thereby represents a source of Isp loss. Accord- ingly,itisdesirabletolimitthcnozzlewall heattransfer,butthe propcrties of appropriate nozzle fabrication materials force the utilization of cooled walls down the converging section and at Icast through the throat area.

There is a performance benefit to employ an adiabatic nozzle skirt, from the final exit section back toward the throatregion, but the stagnation temperature in thenozzleboundary laycrfsee Figure 4) creates increasingly severe wall temperatures, which may become a concern (radiative heat loads) in applications cnvisioning fairly close clustering of NTR engines. In any event, therc is a likelihood of the utiliyation of an extendible nozzle skirt which would cause the designer to favor an inter- face between the cooled and uncooled divergent sections of at l e a ~ t ~ = 65. This selection couldbeacceptable if carbon/carbon composite skirts, up to 42 ft in length and 25 f t in diameter, can bc fabricated. However, ifrefractory metals are sclccted for the uncwled skirt, it is believed prudent to limit the wall tempera- ture to 2600'R. For these reasons, an E = 150 has been selected as the design choice for the study presented herein. Further, in this design study a choice has been made, based on fabrication considerations,tolimitthe~ofthcchannel walldesignto6,and to utilize tubular wall consmction for the portion of the nozzle between E = 6 and E = 150.

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The operation of an unshielded 35 in. core, 2000 MW reactor will gencrate an intense gamma ray and neutron flux field around the reactor equivalent to approximately 20 MW of power. The intensity of this radiation hadbeenanalyzed for the case of a hypothetical 1500 MW Nerva engine, and simple power extrapolations from the referenced study lead to the conclusions that (I) the cooled portions of the nozzle will rcccive gamma radiation between 104 to 1W rad/s and neutron fluxes bctwecn 5x10" to 5x10" ncutrons/cm2-s, and (2) that -

5000 0.35 inch Thick Wall

cc K = 3.47 Btuhn sec "R 4000 EmissiviV -0.92

10001 I I 0 100 200 300 400 500

Area Ratio AIA'

Figure 4. NTR Thrust Nozzle - Carbodcarbon Composite

without a reactor top shield, the turbopump, control compo- nents and the propellant delivery and thrust smcture subsys- tems will be exposed to gamma radiation between lxlod to 5 x W rad/s and neutron fluxes between 2 ~ 1 0 ' ~ to 1x10" neu- trons/cm2-s. For the non-nuclear componenu involved, these radiation fields impose formidable, but not insurmountable design challenges.

In general, the design solutions involve (1) avoidance of the use of materials which have moderate-to-low radiation dam- age thresholds, (2) design of all components with the view of adequatemeansof disposingofradiation induced heat, and (3) the design of the pumping system to accommodate the inges- tion of boiling (2-phase) liquid hydrogen without a loss in pump performance. Basedon thelevel of technologybeingde- veloped during the latter phases of the Rover program, it is believed that the basis for the aforementioned design solutions exist. In thecomponent discussions to follow, indications will be made of the nature of these dcsign solutions.

The design of the cooled nozzle has been indicated above, and pertinent temperature and pressure distibutions, derived from preliminary analyses, are presented in Figure 5. The reactor- nozzle interface is assumed to be similar to that employed in the Phoebus I nozzles. The flange design, the sealing details, as well as thecooled bolls,closelyresembletheconfigurations successfully demonstrated in the Phoebus 1B test. However, in recognition of the anticipated higher heat loads in the con- verging and throatsections, anappropriatechannel walldesign has been adopted which utilizes NARloy Z and design criteria proven in the Shuttle main engines.

At E = 5.86, a transition is made to a tubular wall design, utilizing 0.012 in. thick wall A286 tubing. This design con- figuration (seeFigure6), whichextendstoe= 150,isbasedon well-established rocket nozzle configurations. The design of the uncooled, and possibly extendible, n o d e skin has not yet

50

40 0 %

N

E 30 2 < $ 20 0

- LL

m a, I

c

10

0

1600-

1400 LT *, 1200-

,g1000-

g 800- e

3 t-

= .- 600-

z 3 400-

200

n -

-

- ----____ -40 -20 0 20 40 60 80 100 120 140

Axial Distance From Throat, inches

Figure 5. lOOK NTR Thrust Chamber Thermal Design Parameters

diameter, surrounds the reflector. The overall length of the reactor assembly, including space for the nozzle plenum, the Area

Ritin

Figure 6. IOOK NTR Nozzle Assembly

been addressed. A preliminary estimate of thc nozzle weight is 970 Ib.

The assumed 2000 MW reactor design is based on the solid experience base of the Phoebus lB, Phoebus 2A and NRX-A6 dcvices, and their test results. As indicated earlier, the core is assembled from 1870 uranium-loaded, composite fuel ele- ments, and 325 modified center element, tie tube assemblies. The resultant core, including 325 core support blocks, is 35 in. in diameter and 54 in. long. The installation of graphite filler, thermal expansion and conlainment elements, external to the outer-most fuel elements, brings the entire core assembly to 39 in. in diameter. Next, a 7 in. thick beryllium reflector, contain- ing 12, equally-spaced, 4.24 in. control drums, surrounds the core. Finally a titanium pressure vessel, 55 in. in outside

inlet flow distributionandcoresupportplate,and space for atop shield (if required) is 71 in. The estimated reactor weight, including all internal accessories, is 12,500 Ib.

The design of the liquid hydrogen pump is based on an exten- sive data base developed on the pump utilized in the Phocbus IB, Phoebus IIA, and the NRX-A6 reactor experiments, and also employed as a test bed for the development of hybrid- hydrostatic bearings and “negative NPSH inducers. This pump, designated by Rocketdyne as the Mark 25, was an axial- flow machine, featuring a mixed-flow, axialexit inducer fol- lowed by four stators and rotors. Its design speed was 34,ooO rpm, and it had a flow capacity approximately 40% grcstcr than required for the 100,OOO Ibf engine assumed herein.

If a single pump were to be specified for this engine, the Mark 25 pump could be derated by trimming the rotor and startor blades to approximately 75% of their current design Icngth, by modifying the inducer to match the new inlet flow and exit flow requirements, and by incorporating the hybrid-hydrostatic bearing assemblies previously evaluated. However, the as- sumed propulsion system reliability requirements set forth carlierindicateatlcastadual turbopump-equipped feed system.

Balancing the pumping rates of feed systems incorporating more than one pump is intrinsic, if the pumps exhibit only a negative slope H/Q curve. Thus the utilization of axial flow pumps in a dual turbopump feed system facilitates the design of an inherently stable pumping system; therefore, no balance controls are needed in either the pump delivery lines or in the

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turbine inlet lines. In the event of a turbopump failure, the inoperable unit needs tobe isolated, but this configuration can be provided by equipping each pump and turbine with an inlet shut-off valve and a discharge check valve.

The selection of the operating maps of each of the pumps in a dual turbopump feed system, required to (1) operate over a flow range of 50% to 110% of design, while supporting full reactor ouUettemperature,and(2) underemergencyconditions (pump outconfigured) to deliver sufficient flow andpressure,andalso support full reactor outlet temperature operation, requires care- ful selection of the pump’s hydrodynamic design parameters. Figure 7 sets forth a candidate pump operating map for such a dual tubpump configuration.

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lOOK NTR Engine 70, I , ’ , I 1

Flowrate, GPM x

Figure 7. Candidate Dual Pump Operating Map

The judicious selection of the pump’s design spccific speed line, relative to the nominal 100% thrust operating point, can allow throttled operation (at full Ep) down to 50% of rated flow without encountering stall. Further, by providing a modest operating speed margin above the nominal operating speed, either pump will respond to shuttling down the other pump by shiftingtoancwopcratingpointwhose(1)flowrateisapproxi- mately 40% higher, (2) discharge pressure is approximately 30% lower, and (3) pump operating speed is approximately 10% higher. Figure 8 presents a conceptual design of a candidate turbopump configuration, based on MK 25 technol- ogy, engineered to satisfy the aforementioned requircmcnts. Its weight is estimated to be 185 Ib.

e Pump Flowrate = 54 Ib/s

RPM = 37.500 Pump Inlet Diameter = 5 71 in

Figure 8. Scaled, 5-Stage Mk 25 Pump, Single Stage Expander

The design of a zero or negative NF’SH inducer was first addressedin 1967 when, undertheRoverTechnologyprogram, development of an advanced high speed, direct drive, inducer was undertaken. Analysis had suggested that the suction performance of the MK25 pump could be improved if (1) the inlet flow area were increased 50%, and (2) the inlet flow incidence angle at the design specific speed line, were reduced toapproximately 1.5 deg. Accordingly a new titanium inducer was designed, fabricated and then tested at theNuclear Rocket Development Site. The results of these tests established that this test inducer was capable of ingesting up to 30% vapor of 9 before the pump discharge pressure started to decay. These results lend support to the assumption, adopted for this study, thalthepumpconfigurationspresented herein wouldbecapablc of accepting mixed phase hydrogen flowing from a tank of saturated liquid hydrogen.

The impact of the pumpout operation, upon the ability of its assumed direct-drive inducer to ingest saturated liquid hydro- gen,needs furtherevaluation. If the flow incidenceangleonthe leading edge of the inducer is caused to go negative, the NPSH performance can be expec ted to deteriorate. One solution to this emergency operating problem would be the installation of a separate, hydraulic-driven boost pump whose speed could be independently controlled to provide a more advantageous flow incidence angle.

The operational control of the candidate 100,ooO Ibf NTR engine considered herein requires (1) a balance of the reactor power generated with the thermal power absorbed by the propellant,allinaccordance with the thrustperformance sched- ule for the mission, (2) the accommodation of emergency operational requirements. such as pump-out operations, (3) the

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on-demand supply of pulsed coolant flow, for the after-heat removal, on a flow schedule that is intended to maximize lo' average I,, from this flow, and (4) the implementation of re- -10s

medialactionstosatisfyenginehealthandsafetyrequirements. g l @ b If a design objective is adopted to employ a control system that 104

does not require shielding, the consideration of an all-pneu- 1@ matic/fluidics control system becomes of interest. Figure 9 3 102 presents a candidate NTR engine system control logic intended 5 = 10' to provide at least the capabilities necessary to satisfy the first (0

ti loo

alo-' three (see above) operational modes. c

10;: Time Afler Shutdown (hr)

Figure 10. Heat Generation After Shutdown

as a maximum acceptable steady-state condition. On the bais of after-heat control system simplicity, the latter systcm is assumed herein.

Figure 9. NTR Integrated Pneumatic - Fluidics Control System

During the start-up, high thrust and thrust dccay mode, after criticality has been established and e-folding of reactor power (as measured by a low power flux sensor) has commenced, the imputed reactor power and thrust schedule would be translated to commands for control drum rotation, propellant and turbine inlet valves opening and turbine bypass valve adjustmcnt. The dircct sensing of reactor cxit gas prcssure and temperature (assuming an appropriate pneumatic tcmpcrature sensor is available) provides the requisite inputs fordctermination of the match of the developed reactor power/cngine thrust to the scheduled inputs. Trimming of the angular rotational position of the control drums and of the turbine thmttlc valve will provide the required adjustment in reactor power and pump spccd/flow. Should the nced for shutdown of one of the turbopumps be sensed, the systems control logic could com- mand the closure of the propellant inlct valve and the turbine inlet valve indicatcd, and the remaining turbopump will adjust to the new operating requirements associated with the control- ler-commanded pump-out rcaclor powcr/cngine thrust.

The removal of after heat (sce Figure IO), in a performance (Isp) efficient manner requires monitoring the temperature of some representative region of the core, and then commanding low- level propellant flow to either match the heat removal to the aftcr-hcat generation or provide pulses of propellant flow to intermittently reduce the core temperature to a level prescribed

The state-of-the-art for designing appropriate radiation rcsis- tant LH,shutoff and check valves, pneumatic rotary actuators for hot gaq turbine throttle valves, pneumatic actuators for control drum rotation was fairly well established at the time the Rover-Nervaprograms were broughttoaconclusion. Also, the technology for designing appropriate pneumatic gas tcmpcra- ture sensors, pyrometers, neutron flux sensors and relevant fluidics logic circuitry and components had advanced to the 'L/ point that prototypes could be engineered for development testing. Figures 11 and 12 show examples of these types of components. For thcsc circumstances, and for the wcight advantages of eliminating shielding specific to the enginc control system, an all-pneumatic/fluidics, radiation-resistant control system has been assumed in this design study. If, for reasons of increased confidence in the overall reliability of the NTR engine, redundancy of the control system is desired, it shouldbepossible toengincerandintegratea separate, but fully compatible, shielded elcctrical control system.

The high-radiation field attendant to the operation of an un- shielded, 2000 MW reactor deposits heat in any matcrial located in the field. Thus, the engine thrust structure will experience a heat input, and since the temperature increase of any givensegmentofthethruststructureisdependenton (])the material of the thrust structure, (2) the thickness of thc material experiencing the radiation, and (3) the local intensity of the radiation, it cannot bc assumed that the induced heating will be uniform. Therefore direct thrust structure elemcnt cooling, to stabilizing the alignment of the thrust structure, is prudent and the integration of the propellant delivery ducts with the thrust structure affords a direct method of accomplishing the aforc- mentioned temperature stabilization. Figure 13 prcsents the selected design configuration for the IOOK NTR engine and

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lwimorsion Pyrometer Typical installation

Pneumaiic Louic

x

v

P = 1.2005 T = 855.1 w = 53.55

ChecH Valve P = 1.626.5 T = 52.84 W = 108.84 H = 25.83 s = 233

TruQlne

con1m valve

P = 1.500.6 T = 1,000 w = 54.42 H = 3.441.1 S = 1309

Bypass

w = 54.42 H = 1.863.1

Rellfflor and Relleclor a l i i

Throat Plane

W = 108.835 H = 18.936.82 S = 20.40

N O Z Z I ~ cwiea 10 E = 1 5 0 1

Design Values:

Pump Flowrate 109 Ibis Pump Discharge Pressure 1,627 psia Turbopump rpm 37,500 rpm Turbopump Power 12,319 hp

Turtine Pressure Ratio 1.25 Turbine Flow Rate 53.5 Ibis Reactor Thermal Power 2,163 MW Fuel Element Thermal Power 1,856 MW Tie Tube Thermal Power 199 MW

Nozzle Chamber Temperature 4860"R Chamber Pressure (Nozzle Stagnation) 1,000 psia Nozzle Expansion Area Ratio 500:l

919.1 s

Turbine Inlet Temp IOOOOR

Vacuum Specific Impulse (Delivered)

Figure 14. lOOK NTR, Expander Cycle, Dual Turbopump, Engine

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