ORNL-1287.txt

 

 

 

 

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AEG JESEAREH A, e Pt g

AIRPLANE IN WHICH CIRCULATING FUEL
IS PIPED DIRECTLY TO THE
ENGINE AIR RADIATORS

   

 

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T

OAK RIDGE NATIONAL LABORATORY

OPERATED BY

CARBIDE AND CARBON CHEMICALS COMPAN

A DIVISION OF UNION CARBIDE AND CARBON CORPORATION

POST OFFICE BOX P
OAK RIDGE. TENNESSEE @
s i

 

 

 
 

 

ORNL-1287

This document consists of 6§ pages.

Copy é-;of 170 copies. Series A.

Contract No. W-T7405-eng-26

A DESIGN STUDY OF A NUCLEAR-POWERED AIRPLANE IN WHICH
CIRCULATING FUEL IS PIPED DIRECTLY

TO THE ENGINE AIR RADIATORS

R. W. Schroeder and B, Lubarsky

DATE ISSUED:

MAR 31 1953

 

O0AK RIDGE NATIONAL LABORATORY
Operated by
CARBIDE AND CARBON CHEMICALS CORPORATION
A Division of Union Carbide and Carbon Corporation
Post Office Box P
0ak Ridge, Tennessee

 
 

3 U456 03L0ALY b
 

e

 

 

 
R AT R 9 ot £ L e e

 

O R I U WA
. . - * - . . - .

64-74.
75.
76-80.
81.
82-84.
85.

86-87.
88 -93.

9.

POMEOEr P =P = UE00NDEIPEIMOED

 

 

kORNL-1287
ReactorjFResearch and Power

JERNAL DISTRIBUTION

I

ubarsky

    
       
  
   
 
 
   
 

G. Affel . 28. B

S. Bettis 29. R.N. Lyon

S. Billington 30. WD, Manly

F. Blankenship | 31. L. Meem

P. Blizard L 32. F J. Miller

C. Briant L 33. . Z. Morgan

B. Briggs 1 34. . F. Poppendiek

H. Buck Y 35.4P. M. Reyling

W. Cardwell 1 36/ H. W. Savage

E. Center : 3% E. D. Shipley

H, Clewett A [ A. H. Snell

E. Clifford 3§ f. F. L. Steahly

B. Cottrell p 0. R. W. Stoughton

D. Cowen ! 21. C. D. Susano

B. Emlet (Y-12) 42. J. A. Swartout

K. Ergen F 43. E. H. Taylor

P, Fraas ' 44, F. C. VonderLage

R. Gall 45. A. M. Weinberg

R. Grimes 46. G. C. Williams
Hollaender 47, C. E. Winters

S. Householder 52. ANP Library

B. Humes (K-25) 53. Biology Library

P. Keim Central Files

T. Kelley 60. Health Physics Library
M. King 61. Reactor Experimental
E. Larson Engineering Library
S. Livingston Central Research Library

EXTERN A

Argonne National Lal
Armed Forces Speciall
Atomic Energy Commjly
Battelle Memorial
Brookhaven Natio
Bureau of Ships %
California Rese

Carbide and Cajili
Chicago Paten

ect (Sandia)
o on

SR Company
- (Y-12 Plant)

et

 

1ii
 

 

iv

95.

9.
97-101.
102~ 104.
105-108.
109.
110-116.
117.
118-121.
122-123.
124.
125-127.
128.
129.
130-131.
132-133.
134.
135.
136.
137.
138.
139.
140-141.
142.
143.
144-147.
148-155.
156-170.

 

Chief o Naval Resear
Departme§t of the Navfifi; Op 36
duPont C-«pany i

General ERectric COuVany (ANPP)
General El§ctric Cop Pany, Richland
Hanford Opefations Ufflce

Idaho Operan-ons 0
Iowa State C8l1leg i’

Knolls AtomicQop¥r Laboratory

Los Alamos Sci *.lflc Laboratory

Massachusetts gtitute of Technology (Kaufmann)
Mound Laborato/
National Adv1s
National Advig
New York Oper
North Americg

   
 
 
  
  
     
     
   
     
      
 

-mm1ttee for Aeronautics, Cleveland
_y ! =m1ttee for Aeronautics, Washington
FionsQffice

¥ Aviat) fi&, Inc.

Nuclear Devel§ pment Al ‘:c1ates, Inc.

Patent Bran-f Washlng
Rand Corpor ,1on

San Francis Operatlons f1ce

Savannah Rr fer Operations?y Mf1ce, Augusta
Savannah ”z er Operations @fice, Wilmington
University / of California Rafitation Laboratory
Vitro Co ?ratlon of Americ § 3

    
 
  
    
 

  

-sg-x
)

 
  
  
  
  
    
 

Walter Kf#de Nuclear Laborat 1es, Inc.

Westingh fice Electric Corporj ffion

Wright Air Development Centen\

Technig  Information Servicel}§ Oak Ridge
i”

AP E

 

 
 

AR BT S g i S e Ve

 

CONTENTS

INTRODUCTION AND SUMMARY .

DESIGN OF AIRPLANE AND POWER PLANT FOR MACH 1.5 AT 45,000 FEET .

Reactor Core . . . e e e e e e e e e e e e e e e
Physical descrlptlon
Power distribution . . . . & & v « ¢ 4 ¢ ¢ v v o e e e e .

Engines and Accessories . .
General descr1pt10n of the power plant .
Main engine system . .
Shield-cooling system . . . . . . . . . . . . . .
Reflector-cooling system . . . . . . « ¢« « . .

Accessory system . . . e et e e e e e e e e e e e
Over-all power plant performance . .
Physical arrangement of power plant . . . . . . . .

Power plant weight . . . . . . . . . . . ¢ ¢« ¢ v+« ..

Power Plant Radiators . . . . . . . . ¢« &« v v ¢ « o o
Physical description .
Radiator design relat10nsh1ps
Fuel-to-air radiator .
Auxiliary radiators . . . . . . . . i v 4 e 4 e e e e e

Airplane . . . . . . . . L i i e et et e e e e e e e e e e
Airplane configuration . . .
Airplane lift-to-drag ratio . . « « ¢« v ¢ ¢« ¢ « « ¢ o o
Airplane pitch control .
Airframe weights . . . . . . . . ¢ & v ¢ v e e e e e e e

SEA-LEVEL PERFORMANCE

SHIELDING ANALYSIS .
Assignment of Radiation Contributions . . . . . . . . . .
Configuration to be Shielded . . . . . . . . . . . .
Basic Data for Shield Design . . . . . . . . ¢« « ¢ « « « &

Calculation of Shield Dimensions . . . . . . . . . . . . .
Delayed neutrons into crew compartment rear
Delayed neutrons to crew compartment sides . . . . . . . .
Delayed neutrons into front . . . . . ¢« ¢« ¢ ¢ ¢ o o o @
Gamma rays from the exposed fuel e e e
Gammas from radiators into rear of crew compartment . . .

T
W

. 25

 

VOO e

13

16
16
17
17

19
19
21
23
25

25

28
31
32

33

35
35
35
36

37
37
37
37
37
38
 

 

Gammas from radiators to sides .
Radiator gammas into front . .

Specification of Reactor Shield Thickness
Reactor neutrons into crew shield rear .
Reactor neutrons into crew shield sides
Reactor neutrons into front of crew shield .
Reactor gamma rays into crew shield rear .

Gamma rays from reactor to crew shield sides .
Gamma rays from reactor to crew shield front .

Special Shielding Considerations .
Crew shield sides near the rear
Slanting front wall

Physical Description of Shield .
STATIC CHARACTERISTICS OF THE REACTOR
REACTOR CONTROL

Control Features Determined by Simulator Study .

vi

Pressure in Fuel Tubes .

 

 

 

38
39

39
39
40
40
41
41
41

42
42
43

43
45

56
ST
58
AR W, ©

 

 

ACKNOWLEDGEMENT

The authors are indebted to E, P, Blizard and F. H. Murray who prepared the
section entitled “Shielding Analysis,” to W. K. Ergen and C. B, Mills for the
section on “Reactor Statics,” and to E. R. Mann for the section on “Reactor
Control.” The advice and assistance of J. Y. Estabrook, B. L. Greenstreet, E.
L. Hutto, J. D. Jackson, and A. B. Longyear* materially contributed to the
completion of the calculations and drawings contained herein. Special recog-
nition is due R, C, Briant, whose criticisms and suggestions have substantially
improved the technical content of this report.

 

*On loan from Aerojet Emgineering Corp.

vii

 
 

A DESIGN

STUDY OF A NUCLEAR-POWERED

AIRPLANE IN WHICH CIRCULATING FUEL
IS PIPED DIRECTLY TO THE ENGINE
AIR RADIATORS

R. W. Schroeder

INTRODUCTION

The search for a nuclear power
plant capable of propelling an airplane
at supersonic speeds at high altitudes
has led to a close study of circulat-
ing- fuel reactors, One of the ad-
vantages of such a reactor is that the
heat developed in the fuel may be
transmitted to the air stream 1n
several ways, The heat might be
employed in a vapor cycle so that use
of a compressor-jet engine would be
possible, or the heat might be trans-
ferred to a liquid coolant that would
be used in a turbojet engine.

In the divided-shield concept, all
parts of the aircraft except the crew
compartment are subjected to thoroughly
uninhabitable radiation conditions.
Ground handling of such an airplane
imposes problems that are perhaps not
even now thoroughly appreciated. How-
ever, 1f 1t 1s assumed that these
problems are soluble in a practical
manner, then 1t 1i1s not only prudent
but necessary to investigate the
extreme of such a system.

The inherent adaptability of the
fluid fuels being developed permits
the study of a high-powered system
wherein the heat is transmitted directly
to the air in the engine, The first
asset of such an arrangement is that
the liquid-to-liquid heat exchanger is
eliminated., The first difficulty is,
of course, shielding. In this arrange-
ment, the intensely radiocactive fuel
would have to be carried through a

(I)On loan from Lewis Flight Propulsion Labora-
tory, National Advisory Committee for Aeronautics.

B. Lubarsky(1)

AND SUMMARY

large space between the reactor and
the engine radiators. The shield
would be, then, 1n some sense, the
opposite extreme of a unit shield. The
notion must of necessity exploit shadow
shields to the utmost. Since the air-
plane and the surrounding air would
be subjected to more radiation than in
any other scheme, theair and structure
scattering are of maximum importance,
as would be expected.

In most nuclear airplane proposals
it 1s i1mpossible, really, to separate
power plant and airframe studies. In
this instance, any such separation
would be completely impossible; there-
fore this report covers in an initial
way the design of a circulating-fuel-
direct-to-air tactical airplane operat-
ing at Mach 1.5 and 45,000 feet.

The reactor, fluid circuit, heat
exchangers, shielding, and airplane
studied are described and illustrated
in the body of this report. However,
a brief description of the entire
system is presented at this point to
orient the reader.

The reactor investigated includes
beryllium oxide as a moderator and
reflector, Inconel as a structural
material, and fused fluoride salts
combined with uranium tetrafluoride as
the fuel. The fuel, which is in the
liquid state at operational tempera-
tures, 1s pumped through Inconel fuel
tubes that pass through the moderator.

The fuel leaves the reactor at a
temperature of 1500°F and i1s routed to
fuel-to-air radiators located in each
 

 

DESIGN STUDY

of six turbojet engines. After being
cooled to 1000°F in the radiators, the
fuel 1s pumped back to the reactor by
axial-flow pumps driven by air turbines.

The system postulated is not predi-
cated on any specific radiator design;
however, the radiator designs studied
included Inconel tubes (with Inconel
fins) through which the fuel passes.
The designs studied were such that the
heat exchanger frontal area require-
ments exceeded the engine frontal area
by a large factor. Accordingly, the
heat exchangers shown have been divided
into rectangular banks and placed
parallel to the engine longitudinal
axis. Compressor-discharge air flows
parallel to the engine axis, makes a
right angle turn to pass through the
radiator, and then 1is directed toward
the turbine nozzle box.

The turbojet engines employed were
designed for a turbine inlet tempera-
ture of 1250°F and acompressor pressure
ratio of 6.1 while operating at Mach
1.5 at 45,000 feet. They are similar
in principle to current turbojet
engines except for deletion of the
chemical burners and addition of fuel-
to-air radiators,

A divided shield with water sur-
rounding the reactor and lead and
hydrogenous plastic around the five-
man crew compartment is employed. The
shield has been designed for a maximum
dosage of 1 r/hr within the crew
compartment at design-point operation

(Mach 1.5 at 45,000 ft),

No mechanical control system has
been shown. As discussed more fully
in the body of the report, 1t 1is
expected that the negative temperature
coefficient of reactivity of the
reactor described will cause the
reactor to behave as a slave to the
external heat-removal system (engines
and radiators). If this premise is
valid, the primary control requirements
may be satisfied by a fuel-enrichment
shim for start-up purposes and fuel
drainage provisions for shut-down. The

 

ANP Aircraft Reactor Experiment will,
it is hoped, clarify the validity of
these premises.

The airframe has a delta-wing con-
figuration, The empennage includes a
triangular planformrudder and elevator.
The center of lift and center of
gravity, which coincide, are forward
of the reactor and engines because of
the crew-compartment moment. The bomb
load has been located at the center of
gravity to avoid changes in trim con-
current with bomb release, The engines
are located behind the reactor-shield
assembly, but as close toit as possible
to minimize fluid-piping length. The
engines are also located as close to
the airplane center line as their size
permits to minimize fuselage diameter
and to obtain maximum shadow shielding
by the reactor shield assembly. The
engine air intake 1i1s located forward
of the wing leading edge and is in the
form of an annulus surrounding the
fuselage. ‘

The descriptions and discussions
contained in the body of the report
have been prepared as concisely as the
complexity of the subject matter per-
mits, and no attempt has been made to
summarize this material. Comments
regarding the ultimate feasibility of
the cycle described, or comparisons
between this cycle and other cycles,
would be premature because much more
detailed study, experimentation, and
advancement of the related arts are
needed, It may be said, however, that
the studies made to date indicate a
high performance potential and have
not revealed the presence of inherent
limitations or obstacles that are
believed to be insurmountable. It is
expected that the Aircraft Reactor
Experiment and parallel research and
development being conducted by the
Oak BRidge National Laboratory may
clarify many of the premises and
suppositions included in this study,
and, in addition, advance the tech-
nology of high-temperature circulating-
fuel reactors.

 
 

Problems such as airplane operation,
flight stability, ground handling,
maintenance, and repair are not dis-
cussed in detail. These matters re-
quire exhaustive study and are regarded
as being beyond the scope of this
report. However, with regard to ground
handling and maintenance, any nuclear-
powered airplane with a so-called
“divided shield’’ will require sup-
plementary shielding for airplane
access during ground operation or
after shut-down., The amount of such
supplementary shielding required will
depend on the power history of the
reactor, the distribution of sources
of radiation within the airplane, and
the amount of shielding permanently
installed about these sources. The
configuration discussed here will
require a greater thickness of supple-
mentary shielding than one in which
the fuel circuit is more deeply sub-
merged in the airplane shielding.
The extent to which this will compli-
cate the ground-handling problem would
require very detailed investigations.
Also, with regard to airplane operation,
flight stability, and other such con-
siderations, it should be recognized
that only a few experimental airplanes
have to date achieved supersonic
speeds, and none of these approach in
size the airplane discussed here.
Determination of the optimumaerodynamic
configuration, stability criteria,
incidence angles required for take-off
and landing, etc. will involve further
aerodynamic research and airframe
design studies. The airframe con-
figuration illustrated should there-
fore be regarded as highly tentative,
These studies deal primarily with the
power plant and the shielding. Changes
in the airframe will have little
effect on these studies unless the
reactor-to-crew separation distance or
power requirements affected
significantly,

The calculated performance of the
system studied is summarized as follows:

are

NUCLEAR-POWERED AIRPLANE

AT

AT
SEA LEVEL 45,000 FEET
Speed Take-off Mach 1,5
Total net thrust (1lb) 165,600 53,850
Take-off distance (ft) 2,500
Total air flow (lb/sec) 4,137 1,751
Turbine inlet temperature
(°F) 1,125 1,250
Fuel temperature (°F)
reactor inmlet 1,000 1,000
Fuel temperature (°F)
reactor outlet 1,500 1,500
Fuel flow (lb/sec) 3,130 1,650
Maeximum reactor tube
temperature (°F)
Inside surface 1,583 1,554
Outside surface 1,608 1,567

A summary of the weights of the
various portions of the aircraft is
given in the following:

WEIGHT (1b)
Airplane
Wing 46,000
Tail 9,200
Fuselage 29,900
Landing gear 18,900
Controls 2,100
Total 106,100
Power Plant
Engines 59,900
Auxiliary system 5,000
Inlet and exhaust ducting 10, 300
Rediators
Core 17,900
Baffles, structure, headers,
contained fuel, etc. 6,000
Total 99,100
Shielding
Crew shield
Lead 30,800
Plastic 25,900
Reactor shield assembly
Reactor assembly 10, 000
Water 28,200
Structure, insulation, etc. 10,200
Total 105,100
 

DESIGN STUDY

 

Payl oad
Crew (5 at 250 1b) 1,250
Furnishing 850
Pressurizing and oxygen 550
Communicating equipment and
jamming radar 600
Bombing and navigating equipment 1,700
Photographic equipment 50
Instruments 400
Bomb load 10,000
Firepower (tail turret and
ammunition) 3,000
Contingencies peculiar to
shielded cockpit 1,600
Total 20,000
Contingency 19,700
Total airplane weight 350,000

A summary of the fuel holdup in
the various portions of the power

 

plant is given below (there are 3.14
1b of U235 per cubic foot of fuel).

FUEL HOLDUP
(£t

Reactor

Core 7.96

Headers 3.65
Radiators

Core 6.9

Headers 8.4
Piping between reactor and

radiators

Common inlet piping 2.5

Common ocutlet piping 2.3

Individual piping between
lines and radiator (including
pumps, etc.) 4.0

Total

35.711

 

DESIGN OF AIRPLANE AND POWER PLANT FOR MACH 1.5 AT 45,000 FEET

REACTOR CORE

A general discussion of a reactor
intended to provide sufficient power
to operate an airplane at Mach 1.5 and
45,000 ft is presented in this chapter.

The decision to explore the po-
tentialities of circulating- fuel
reactors necessitated the review of
several broad classes of moderators:
(1) low-temperature hydrogenous
liquids (such as water) used with
double-wall construction or insulation
between the fuel and the moderator,
(2) high-temperature hydrogenous
liquids used with single-wall con-
struction, and (3) solid moderators,
such as beryllium oxide. FEach of
these possible moderator arrangements
appears to offer some advantages and
some disadvantages, but it is not
possible to make an irrevocable
decision at this time as to which one

should be used.

Use of the first moderator would
involve the difficult problem of

rejecting the moderator heat from a
low-temperature source to a relatively
high-temperature sink. The required
air-flow rates would be large, inas-
much as the permissible air tempera-
ture rise would be limited and the
driving temperature differences would
be low. Furthermore, the double-wall
construction within the reactor appears
to involve serious problems because of
differential expansion between the
cold tubes and the hot tubes, tube
sheets, headers, etc. Accordingly, it
was decided to avoid this approach for
the present, The second moderator
appears to be attractive 1n many
respects, At present, however, there
are no combinations of high-temperature
hydrogenous fluids and structural
materials that are known to be com-
patible at the operating temperatures
of circulating-fuel reactors. There-
fore active consideration of this
possible moderator must be deferred.
The third arrangement has been employed
in the design studies outlined here
because it appears to involve no major

 
 

material uncertainties and permits a
relatively simple core design.
Inasmuch as the heat of the fuel is
not transferred within the core,
incorporation of a heat exchanger
lattice within the core is not neces-
sary, and relative coarseness of core
geometry is permitted. As the fuel-

tube surface area is diminished, how-

ever, two constraints appear that
influence the required tube diameter,
tube surface area, and fluid velocity.
First, the moderator heat inflow to
the fuel stream causes a film tempera-
ture drop, 6, which increases the
fuel-tube temperature. Second, the
lower velocities of the fuel particles
adjacent to the walls lead to greater
fuel residence times and higher wall
temperatures. In the geometry achieved
after several iterations, the first
effect was found to dominate. The
film drop associated with moderator
heat inflow may be expressed as

 

0.2
,.QL_ oDt
A h A Vo8
where
@ = temperature difference, °F,
é%== heat flux, Btu/sec. ft?,
h = heat transfer coefficient,
Btu/sec*°F- ft?,
D = tube diameter, ft,
V = fluid velocity, ft/sec.

If it is desired to achieve a maximum
wall temperature of approximately
1550°F with a fluid inlet temperature
of 1000°F and a fluid outlet tempera-
ture of 1500°F, the permissible &
will be 550°F at the inlet end and
50°F at the outlet end. The high
permissible inlet & can be employed
advantageously by using a two-pass
arrangement in which the cold inlet
fluid is passed first through the
region of highest power generation - the
central portion of the core. Since
wall temperatures in this region were
found to be readily controllable,
relatively low flow velocities and
large tube diameters could be used.

NUCLEAR-POWERED AIRPLANE

Wall temperatures near the outlet end
of the reactor tended to become more
critical as the fuel temperature in-
creased. This tendency was alleviated
by the reduction in specific power
generation as the fuel approached the
unreflected end of the peripheral pass.
Further alleviation was provided by
decreasing the tube size and increasing
the number of tubes, which also in-
creased the surface-to-volume ratio,
and by increasing flow velocities in
the second pass. After several
iterations, a geometry was achieved
that resulted in maximum fuel-tube
wall temperatures, in each pass, of
approximately 1550°F.

The core (Fig. 1) consists of a
series of parallel tubes, arranged in
two series passes, that convey circu-
lating fuel through a beryllium oxide
block lattice. A beryllium oxide
reflector adjacent to all core surfaces
except the fluid inlet and outlet end
has been provided and is to be cooled
by circulation of nonuranium-bearing
fused fluorides.

Physical Description, The reactor,
as shown in Fig., 1, can be considered
as being contained in a 55-in, -dia
sphere if the fuel inlet and outlet
lines and reflector coolant (salt)
lines are excluded., The reactor core
consists of parallel tubes arranged
in concentric circles and contained
ina40,4 -in. ~-dia cylinder with conical
and truncated -conical ends. Each core
tube is surrounded by a moderator in
the form of hot -pressed beryllium
oxide, Specific design features are
presented in the following:

1. The cylindrical core has mani-
folds on the ends to provide for two-
pass flow of the fuel,

2. The fuel, metal tubing, and
moderator volume fractions are held
constant throughout the core. The
cylindrical core contains approximately
34% fuel, 2% metal tubing, and 63%
moderator.

3. The fuel used for the calcu-
lations of this study is a molten
 

DESIGN

 

STUDY

DWG. 17662

 

PLAN

    
 

REFLECTOR

~> FUEL IN

 
  

SALT ouT

  

== SALT IN, SIX PLACES

VERTICAL SECTION ON ¢

Fig. 1. Beryllium Oxide-Moderated, Circulating-Fuel Reactor.
b LA bt . mie 2L s e e e

 

mixture of fluoride salts, one of which
is uranium tetrafluoride in a low
concentration,

4, The moderator and reflector are
beryllium oxide blocks.

5., The reflector is situated about
the core as shown in Fig. 1.

6. The core shell is perforated
around the cylindrical section to
permit the influx of reflector coolant
to fill the core-moderator interstices.

Six small tubes connect the core to
the reflector through the crossover

header to augment filling the inter -
stices. The coolant will be maintained
at an absolute pressure above that of
the fuel circuit to prevent the
accumulation of stagnant fuel in the
moderator interstices in the event of
an internal leak in the fuel circuit.

7. The tube sheet, at the un-
reflected end of the core, is separated
between the fuel inlet and outlet to
permit the differential expansion that
occurs because of the temperature
rise in the core,

8. Minimum pressure loss and mini -
mum volume (uranium holdup) were con-
sidered in designing the inlet, outlet,
and crossover headers. The 1inlet
header is a single 9,5-in, line that
feeds all core tubes in the first
pass through a single header., This
inlet line extends 5 ft from the
reactor to a collector manifold that,
in turn, receives all fuel returning
from the engine radiators. The outlet
is a l.5-in, annulus that receives
all outgoing fuel from the second -
pass core tubes and transmits the fuel
to a common, annular manifold that,
in turn, feeds all engine radiators.
The reactor outlet annular header is
concentric with the reactor inlet
line. This header arrangement elimi -
nates any adverse flow conditions that
may arise i1f one or more engines are
shut down as a result of malfunction
or battle damage.

9. All metallic parts that come
in contact with either the fuel or

NUCLEAR-POWERED AIRPLANE

the coolant are Inconel, which has
been shown to have the best corrosion
resistance to molten salts and also
good high-temperature strength charac-
teristics.

Power Distribution. The six turbojet
engines require a reactor power output
of 321,000 Btu/sec and a fuel flow
rate of 14.7 c¢fs, The freezing point
of the molten salt mixture dictates
a minimum, reactor -inlet, mixed-mean
fluid temperature on the order of
1000°F., The strength of the materials
of the reactor core and pressure shell
dictates a maximum, reactor -outlet,
mixed-mean fluid temperature of 1500°F,

The physical properties of the
molten salt mixture used in the calcu-
lations of mixed-mean fluid temperature
and fuel-tube wall temperatures are
given in the following:

0.39 Btu/lb-°F
112 1b/fc?
0.5 Btu/hr- ft2 (°F/fv)

Specific heat, Cp
Density
Thermal conductivity

Viscosity 8.3 to 2.1 centipoises

The power distribution within the
core 1s determined in the section
entitled “Static Characteristics of
the Reactor” and is shown in Fig. 29.
Five per cent of the total power
generated was assumed to be generated
in the moderator. This power is trans-
mitted to the fuel via heat conduction
through beryllium oxide, interstices
filled with molten salt, and the tube
wall, and then by convection to the
fuel. The power distribution in the
moderator was assumed to be the same
as the fuel power distribution,

Fuel-tube wall temperatures based
on these power distributions were
calculated for various tube stations
in both the first and second pass, as
shown in Fig. 2. A temperature profile
through a typical core section 1is
shown in Fig. 3,
TEMPERATURE (°F)

ed
OWG. 17663
1600

FUEL-SIDE TUBE-WALL FUEL-SIDE TUBE-WALL
TEMPERATURE TEMPERATURE

1500

1400

1300

[] MIXED-MEAN FUEL TEMPERATURE
m BOUNDARY LAYER TEMPERATURE RISE DUE TO MODERATOR HEAT FLUX

BOUNDARY LAYER TEMPERATURE RISE DUE TO VOLUMETRIC POWER GENERATION
1200

L= LENGTH FROM UNREFLE END

1100

1000

© [REFLECTED END
REFLEGTED END

o Q.2 0.4 0.6 a8

o

o8 0.6 0.4 0.2 o

L/L,, CORE TUBE LENGTH RATIO L/L,,CORE TUBE LENGTH RATIO
FIRST PASS - ROW NO 11 SECOND PASS-ROW NO1{5

Fig. 2. Longitudinal Temperature Pattern in Fuel Tubes in Reactor Core.

   

AQNLS N9ISHd
 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

D’! 1!564
1800 p —
v Lo
[ —MOLTEN SALT
1700 P
OUTER ROW, FIRST PASS, /
Ll o=0.576 //
1600
§ "} —-MODERATOR
c & (Be0)
¥ d
w 1500
§ C
@
¢
=
E 1400 '
N
r &
=1
-
1300 j}
1200
) —at b= TUBE WALL
0 0.2 0.4 0.6 0.8 1.0

DISTANCE FROM FUEL TUBE CENTERLINE (in)

Fig. 3. Temperature Profile Through
Fuel Tube and Moderator.

ENGINES AND ACCESSORIES

A turbojet cycle, in which fuel -to-
air radiators are substituted for
the conventional chemical burners,
is employed to provide sufficient
thrust for operating the design air-
plane. Compressor bleed -off air is
used for the reflector- and the shield-
cooling systems; some of this air is
then expanded through turbines to
furnish power for accessories, and the
remainder is expanded through adjust-
able nozzles to give propulsive thrust,

Once the total air-flow require -
ment was established, the total com-
pressor inlet area needed was deter -
mined on the basis of NACA develop -
mental experience., The number of
engines necessary to accommodate the
total air flow (or to provide the
total inlet area) will depend on the
size of engine that can be made

very arbitrary.

NUCLEAR-POWERED AIRPLANE

available when an airplane of the
type described is constructed. At
present, any determination of the
number of engines to be used will be
The use of six
engines has been postulated because
of the convenience from the standpoint
of installation. The use of a different
number of engines, within reason,
would have only secondary effects on
the over-all airplane weight and
performance.

General Description of the Power
Plant., Figure 4 is a schematic diagram
of the main engines and accessories.
The main engines are turbojets with
circulating -fuel -to -air -radiators
instead of the conventional combustors.
Heat is generated in the circulating
fuel as it passes through the reactor
and is then transferred to the engine
air flow in the circulating-fuel -to -
air radiators. Air is bled from the
compressors of the main engines to
auxiliary radiators to remove heat
from the reactor -shield coolant and
the reflector coolant. A portion of
the air passing through the reflector-
coolant radiator is used to operate
a number of air turbines that drive
all the liquid pumps in the power
plant. All the air bled from the
main compressors is eventually dis-
charged rearward and provides some
additional thrust. With an airplane
gross weight of 350,000 1b and an
airplane lift-to-drag ratio of 6.5,
the power plant is required to produce
a total thrust of 53,850 1lb at design
flight conditions.

The power plant may be considered
as consisting of four principal
portions: the main engine system, the
shield -cooling system, the reflector -
cooling system, and the accessory
system,

Main Engine System. The air for
all four of the systems enters the
inlet duct of the airplane and passes
through the diffuser. It is then
carried in ducting around the reactor
shield and into the compressors of the
 

 

 

 

DESIGN STUDY

o
DWG. 17665

   
  
  
  
  

 
 
  

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o =

  

 

 

 

 

 

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==
ONE OF TWO GENERATORS AND AUXILIARY
= TURBINES

 

 

 

5 19 20

 

 

 

 

18

 
  

_”';/
ONE OF SiX AUXILIARY TURBINES
ONE OF SiIX PUMPS

 

 

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et

3

 

 

- ONE OF S

 

 

 

 

 

 

 

 

 

 

A
TOTAL TOTAL AL TOTAL TotaL | wovar
LOCATION FLUID | PRESSURE | TEMPERATURE FLOW LOCATION FLUID ; PRESSURE | TEMPERATURE| "g ow
(peia) °p) (1b/sec) (psia) C°F) | (1b/aec)
A, Radiator Imlet Line Fuel 105 1500 1646 8. First-Stage Bleed Air 8.68 145 BS5.6
Radiator Outlet 9. Radiator Inlet Line | Air 8.46 145 85.6
Line Fuel 25 1000 1646 10, Radiator Outlet Line| Air 7.60 3c0 85.6
C. Pump Outlet Line Fuel 175 1000 1646 11. Jet Pipe Air 1.38 300 85.6
D. Reactor Inlet Line Fuel 160 1000 1646 12. Auxiliary Jet Air 2.142 85.6
E. Reactor OQutlet Line Fuel 120 1500 1646 13. Exght-Sta ¢ Bleed Air 29.0 430 110.7
F., Radiator Inlet Line | Water| 200 350 61.8 14. Radiator Inlet Line | Air 28.3 430 104.9
G. Radiator Outlet Line | Water 167 300 61.8 15. Radiator Outlet Line| Air 25.4 1000 104.9
H. Pump Outlet Line Water| 211 300 61.8 16. Auxiliary Turbine
J. Shield Inlet Line Water 206 300 61,8 Inlet Line Air 24.7 1000 7.3
K. Shield OQutlet Line Water| 205 350 61.8 17. Auxlliari Turbine
L. Radiator to Pump Outlet Line Air 7.24 T17 7.3
Line Salt 137 1000 195.2 18, Auxiliary Jet Air 2.142 7.3
M. Radiator Qutlet Line | Salt 137 1000 10.8 19, Jet Pipe Air 24.7 1000 97.6
N. Radiator Inlet Line | Salt 165 1200 206 20. Auxiliary Jet Air 2.142 97.6
P. Reflector Outline 21. Radiator Inlet Line | Air 28.3 430 5.8
Line Salt 170 1200 206 22, Radiator Qutlet Line| Air 25.4 1000 5.8
3. Reflector Inlet Line | Salt 175 1000 206 23. Auxiliary Turbine
. Pump Outlet Line Salt 180 1000 206 Inlet Line Air 24.7 1000 5.18
0. Aircraft Ambient Air 2.142 -67 24, Auxiliary Turbine
1. Compreasor Inlet Outlet Line Air T.24 T17 5.18
ine Air T.24 108 1948 25. Auxiliary Jet Air 2.142 5.18
2. Compresaor Outlet . 26. Auxiliary Turbine
ine . Air 43.5 544 1751 Inlet Line Air 24.7 1000 0.17
3. Radiator Inlet Line [ Air 42,4 544 1751 27. Auxiliary Turbine
4, Radiator OQutlet Line | Air 38.0 1250 1751 Outlet Line Air T.24 T17 0.17
5. Turbine Inlet Line Air 36.9 1250 1751 28. Auxiliary Jet Air 2.142 0.17
6. Turbine Outlet Lime | Air 10.6 824 1751 29, Jet Pipe Air 24.17 1000 0.62
1. Jet Air 2.142 1751 30, Auxiliary Jet Air 2,142 0.62

 

 

 

 

 

 

 

 

 

 

Fig. 4. Schematic Diagram of Power Plant.

10

 

 
 

 

six main engines. The air required
for the shield-cooling, the reflector-
cooling, and the accessory systems
is bled from various stages of the

main compressors, as will be described.

The air for the main engine system
passes through the compressors and
enters the fuel -to-air radiators, where
it is heated by the fuel circulating
from the reactor. The air then expands
through the turbines that drive the
compressors and is exhausted rearward
through variable -area exhaust nozzles.

A thermodynamic calculation was
carried out to determine the specific
impulse and cycle efficiency of the
turbojet engines for various values
of compressor -pressure ratio, turbine

NULCEAR-POWERED AIRPLANE

inlet temperature, and pressure drop
in the radiators and associated ducting
between the radiators and the com-
pressors and turbines., The following
efficiencies were used for the various
components:
Diffuser and inlet ducting
pressure recovery factor
(actual total pressure per

ideal total pressure) 0,92
Compressor efficiency, total -

to-total adiabatic 0.85
Turbine efficiency, total-

to -total adiabatic 0.90
Exhaust nozzle velocity co-

efficient 0.97

Figures 5 and 6 show the specific
impulse and cycle efficiency of the

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

UNCLASSIFIED
OWG. 17666
a0
—-"——___
__-‘"“‘*--\‘ 1380 "—_-___-hh“*-.‘\_‘\\\\‘\“‘!350
30 —] — =
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2 50 \\\‘\\\ 1150
s 20 . q
§ 4¢) - g 48 N
3 ( 2 )RAD——O 05 \ 050 & ( > )MD—O.IO 0so
-4 & x
B g 5
£ 10 & & —
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g 40 & ]
$ g z
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"i" 30 ___———"—'-\ 1350 5 \ t350 — E —
o ] e T ——
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-—-—-__-.§-“5~‘\\“ =0 fih‘"“-~. 1250
20 - 1150 T~ 150
G%q =019 \\\\\\\ A )
£y
RA0 1050 ( )mo 0.20 10850
10
2 4 6 8 2 4 6 8
COMPRESSOR PRESSURE RATIO
Fig. 5. Variation of Specific Impulse with Compressor Pressure Ratio for

Turbojet Engines.

Alticude, 45,000 ft; Mach, 1.5; diffuser efficiency, 0.92;

compressor efficiency, 0.85; turbine efficiency, 0.90; nozzle efficiency, 0.95.
All efficiencies are total-to-total adiabatic.

11
 

DESIGN STUDY

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

UNGCLASSIFIED
DWG. 17667
40
2
-
[&]
=
5 b T
S 30 ’fl_ i
& B
- AP
W Q—- =0.05 Ap
3 £ lanp ~p ) =0.t0
= RAD
[&]
20 I
TURBINE INLET TEMPERATURE (°F)
1350
—————— 1250
——1150
——— —1050
40
&
P
O
=
il
o —_—
o 30 e —— e —
W S —— e ¢ —— ——
—
H — —— — —
0 Ap =1 T~
6 s (7’" =045 (~A—P) =0.20
RAD | P lanp i
20
2 4 6 8 2 4 6 8
COMPRESSOR PRESSURE RATIO
Fig. 6. Variation of Cycle Efficiency with Compressor Pressure Ratio for

Turbojet Engines.

turbojet engines for compressor -
pressure ratios from 2 to 8, turbine
inlet from 1050 to
1350°F, in the
radiators and associated ducting of
5 to 20 per cent,('? These curves,
together with the radiator data in-
cluded in the ‘“Power Plant Radiators’
section, permitted selection of the
following design-point conditions:

temperatures
and pressure drops

Compressor-pressure ratio 6.0:1
Turbine inlet temperature 1250°F
Pressure drop in radiator

and radiator ducting 15%

It will be noted that the engines
alone would be favored by lower com-
pression ratios, higher turbine inlet
temperatures, and lower radiator

(l)ln the actual power plant, the specific
impulse and cycle efficiency will be reduced some-
what by compreassor bleed-off. More exact specific
impulses and cycle efficiencies are presented in
a later gection on "Over-All Power Plant Per-
formance. '

 

12

Conditions and efficiencies same as in Fig. 5.

pressure drops and that the radiators
alone would be favored by higher com-
pression ratios (greater densities),
lower turbine inlet
(greater driving forces), and higher
pressure drops (greater velocities),
Several preliminary engine and radiator
design studies, in which various
combinations of the controllable
variables were used, indicate that the
design-point conditions selected are

close to optimum.

temperatures

With the use of the efficiencies
and other factors given, thermodynamic
calculations were made of the air
circuit of the main engine system.
Allowance was made for the guantities
of bleed air needed for the other
systems. Pertinent values of air
pressure, temperature, and weight
flow at various stations in the main
engine system are given in Table 1 and

 
e R AR S 0 0wt sl iic'a e

Fig. 4. The values of weight flow are
for all six engines combined.

The thrust produced by the six
main engines is 48,690 pounds. This
is approximately 90.4% of the required
thrust, the remaining 9.6% being
produced by the other systems. The
specific impulse of the main engine
air is 27.8 1lb of thrust per pound of
air per second. The amount of power
that must be generated in the fuel and
moderator of the reactor is 321,000
Btu/sec.

The heat generated in the reactor
core (fuel and moderator) is trans-
ferred to the main engine radiators
by the circulating fuel. The maximum
fuel temperature leaving the reactor
was set at 1500°F. Higher temperatures
would, of course, be desirable but
would make the problem of designing
the various components appreciably
more difficult, The temperature
entering the reactor was chosen as
1000°F, This, again, 1s a compromilse
between conflicting requirements,
Higher reactor inlet temperatures would
reduce the size of the radiators but
would increase the fuel flow rate and
hence the duct sizes for the same
pressure drop, thus increasing the

TABLE 1.

NUCLEAR-POWERED AIRPLANE

amount of fuel in the system. Lower
values of temperature would, in
addition to increasing the radiator
size, increase the danger of freezing
the fuel. Therefore the value of
1000°F was selected as a reasonable
compromise., The total weight flow of
fuel for all six engines is 1646 lb/sec.,
Values of fuel temperature and pressure
at various stations in the main engine
system are listed in Table 2 and
Fig. 4.

The properties used in the analysis
of the circulating fuel are:

112 1b/fe?

0.39 Btu/1b*°F

0.5 Btu/hr- ft? (°F/ft)
4.84 1b/hr-ft

Density
Specific heat
Thermal conductivity

Viscosity

Shield-Cooling System. A conserv-
ative estimate of the rate of heat
generation in the reactor shield is
1% of the core heat generation rate, (1)
Therefore 3210 Btu/sec must be removed
from the reactor shield, This 1is
accomplished by circulating the shield

(2} peport of the Shiclding Board for the Air-
craft Nuclear Propulsion Program, ANP-53 (Oct. 16,
1950).

AIR PRESSURES, TEMPERATURES, AND WEIGHT FLOWS AT

VARIOUS STATIONS IN THE MAIN ENGINE SYSTEMS

 

 

 

TEMPERATURE PRESSURE WEIGHT FLOW
(°F) (psia) (l1b/sec)

Station 0, ambient conditions -67 2.142

Station 1, compressor inlet 108 T1.24 1948
Station 2, compressor exit 544 43.5 1751*
Station 3, radiator inlet 544 42.4 1751
Station 4, radiator outlet 1250 38.0 1751
Station 5, turbine inlet 1250 36.9 1751
Station 6, turbine exit 824 10,6 1751
Station 7, exhaust jet** 1751

 

 

 

 

*Air (197 1b/sec) is bled from various stages of the main engine compressors for the shield-

cooling,
**Velocity = 2345 {fps.

the reflector-cooling, and the accessory systems.

13
 

DESIGN STUDY

water through a radiator. The temper-
ature that can be maintained in the
shield without boiling the shield
water 1s, of course, dependent on the
pressure maintained. To keep the
pressure reasonably low, a shield-
water temperature of 350°F and a
pressure of 200 psia were selected.
The shield-water temperature is reduced
50°F in the radiator. The weight flow
of shield water required 1s 61.8
1b/sec. The temperatures and pressures
of the shield water at various stations
in the shield-cooling system are given
in Table 3 and Fig. 4.

Air for the shield-water radiator
is bled after the first stage of the
main engine compressors, because at
that point the shield water i1s at a
low temperature; bleeding at a later
stage would increase the temperature
of the bled air., It .would be possible
in flight to use ram air to feed the
shield-water radiator, but it seems
more desirable to design the system

TABLE 2.
IN THE MAIN

 

to use air bled after the first com-
pressor stage, since this will permit
cooling of the shield water while
stationary on the ground without the
use of auxiliary equipment external
to the airplane. The temperature of
the air entering the shield-water
radiator is 145°F, and the temperature
leaving the radiator is 300°F., The
weight flow of air is 85.6 lb/sec,.
After the air passes through the
shield-water radiator, it is exhausted
through a variable—area nozzle and
produces some thrust. Air temperatures
and pressures at various stations in
the shield-cooling system are given
in Table 4 and Fig. 4. The thrust
produced by the jet is 425 1b, and the
specific impulse is about 4.96 1b of
thrust per pound of air per second.
Reflector-Cooling System. It is
estimated that the rate of heat
generation in the reflector will be
about 3% of the core heat generation
rate. Therefore 16,050 Btu/sec must

FUEL TEMPERATURES AND PRESSURES AT VARIOUS STATIONS
ENGINE SYSTEM

 

 

, radiator inlet
Station

A
B
Station C,
D
E

, radiator outlet
pump outlet

Station
Station

, reactor 1inlet

, reactor outlet

 

TEMPERATURE (°F) PRESSURE (psia)
1500 105
1000 25
1000 175
1000 160
1000 120

 

 

 

 

TABLE 3. WATER TEMPERATURES AND PRESSURES AT VARIOUS STATIONS
IN THE SHIELD-COOLING SYSTEM
TEMPERATURE (°F) PRESSURE (psia)

Station F, radiator inlet 350 200
Station G, radiator outlet 300 167
Station H, pump outlet 300 211
Station J, shield inlet 300 206
Station K, shield outlet 350 205

 

 

 

14

 
g bR BPRER S BT i 17 T st § el 2

 

be removed from the reflector. This
is accomplished by circulating a
molten mixture of fluoride salts
(containing no uranium tetrafluoride)
through the reflector and then through
a radiator, where the heat picked up
by the salt in the reflector i1s re-
moved. The reflector inlet temperature
of the salt was set at 1000°F and the
outlet temperature at 1200°F, The
weight flow of salt regquired is 206
1b/sec. The temperatures and pressures
of the salt at various stations in the
reflector-cooling system are given in
Table 5 and Fig. 4. The properties
used for the circulating-fuel analysis
could be used for this salt analysis
because the fuel has a low uranium
concentration,

Air for the reflector-coolant
radiator is bled after the eighth
stage of the main engine compressors.

NUCLEAR-POWERED AIRPLANE

This bleedpoint is a compromise
between the conflicting requirements
of over-all engine performance (which
favor bleeding at an earlier stage)
and radiator and duct size (which
favor bleeding at a later stage)., No
attempt has been made to optimize the
bleed point, but one possible com-
promise was selected. The air temper-
ature entering the reflector-coolant
radiator is 430°F and the outlet
temperature is 1000°F. A weight flow
of 110.7 1lb/sec is required. After
the air has passed through the re-
flector-coolant radiator, a portion
of it (12,48 1lb/sec) is used to operate
a number of air turbines that drive
the power plant accessories. The air
that is not diverted to the accessory
system is exhausted through a variable-
area exhaust nozzle. Values of air
temperatures, pressures, and weight

 

 

TABLE 4. AIR TEMPERATURES AND PRESSURES AT VARIOUS STATIONS
IN THE SHIELD-COOLING SYSTEMS
TEMPERATURE PRE SSURE
(°F) (psia)
Station 8, bleed point after first compressor
stage 145 8.68
Station 9, radiator inlet 145 8.46
Station 10, radiator outlet 300 7.60
Station 11, exhaust nozzle entrance 300 7.38
Station 12, exhaust jet®

 

 

 

*Velocity = 1610 fpa.

TABLE 3. SALT TEMPERATURES AND PRESSURES AT VARIOUS STATIONS
IN THE REFLECTOR-COOLING SYSTEM

 

 

TEMPERATURE PRESSURE
(°F) (psia)
Stations L and M, radiator outlet 1000 137
Station N, radiator inlet 1200 165
Station P, reflector outlet 1200 170
Station Q, reflector inlet 1000 175
Station R, pump outlet 1000 180

 

 

 

15
 

 

DESIGN STUDY

flows at various stations in the
reflector-cooling system are given 1in
Table 6 and Fig. 4. The thrust
produced by the jet is 4530 1lb, and
the specific impulse is about 46 1b
of thrust per pound of air per second.
Accessory System. Power must be
provided to drive the liquid pumps in
the power plant and to drive the
electric generators that furnish
electrical power for the airplane.
This is accomplished by using a portion
of the air coming out of the reflector-
coolant radiator to operate a number
of air turbines that drive the power
plant pumps and the electric generators.
By assuming a pump efficiency of 80%
and a total generator capacity of
about 425 kw, the required pumping
power for the power plant at design
flight conditions has been calculated
to be about 1310 horsepower. (At
sea level the required pumping power
is much greater, and the pumps and
air turbines must be designed to
handle this greater load; also, a

greater portion of the reflector-

cooling system air flow must be di-
verted to the accessory system. This
has been provided for and is described

 

in the chapter on “Sea-Level Per -
formance,” )

The weight flow of air required for
the air turbines has been calculated.
For the calculation, it was assumed
that the turbine exit pressure was
equal to the ram pressure (7.24 psia)
and that the turbine efficiency was
70%, The weight flow required 1is
12,48 lb/sec. Values of air tempera-
tures and pressures at various stations
in the accessory system are given in
Table 7 and Fig. 4. The various jets
produce a thrust of about 210 1lb, and
the specific impulse is about 16.8 1b
of thrust per pound of air per second.

Over-All Power Plant Performance.
The combined thrust of the shield-
cooling, the reflector -cooling, and
the accessory systems is about 5160
pounds. This thrust, added to the
main engine thrust of 48,690 lb, gives
a total power plant thrust of 53,850
l1b, the required value, The average
specific impulse of the power plant
is about 27.64, and the over-all cycle
efficiency (with the power generated
in the reflector and shield included
in the power input) is about 29.63%.

 

 

TABLE 6. AIR TEMPERATURES, PRESSURES, AND WEIGHT FLOWS AT
VARIOUS STATIONS IN THE REFLECTOR-COOQLING SYSTEM
TEMPERATURE PRESSURE WEIGHT FLOW
(°F) (psia) (lb/sec)

Station 13, bleed point after

eighth compressor stage 430 29.0 110.7
Stations 14 and 21, radiator

inlet 430 28.3 110.7 (total)
Stations 15 and 22, radiator

outlet 1000 25.4 110.7 (total)
Stations 19 and 29, exhaust

nozzle entrance 1000 24.7 98.22 (total)
Stations 20 and 30, exhaust

nozzle* 98,22 (total)

 

 

 

 

*Velocity = 2930 {ips.

16

 
 

 

TABLE 7.

NUCLEAR-POWERED AIRPLANE

AIR TEMPERATURES AND PRESSURES AT VARIOUS STATIONS

IN THE ACCESSORY SYSTEM

 

 

TEMPERATURE (°F) PRESSURE (psie)

 

Stations 15 and 22, reflector-cooling
radiator outlet

Stations 16, 23, and 26, turbine
inlet

Stations 17, 24, and 27, turbine
outlet

Stations 18, 25, and 28, exhaust jet*

1000 25.4
1000 24,7
117 7.24

 

 

 

*Veloeity = 1990 fps.

Physical Arrangement of Power Plant,
One possible layout of the required
power plant equipment is shown in
Fig. 7. The six turbojet engines are
arranged circumferentially around the
cowl and as far outward as they would
go. (There is space left in the
bottom of the cowl where there is no
engine, because it was originally
thought that the main wing spar might
come through at that location. It is
apparent from Fig. 7 that for this
particular airplane configuration the
spar will not be at that location, and
therefore the engines could actually
be spaced differently.) The main
engine fuel -to-air radiators occupy the
space normally occupied by the com-
bustors of the turbojet engines. The
shield-coolant radiator is located
Just behind the reactor in the central
hole between the engines. The re-
flector -coolant radiator is divided
into seven parts. One part, located
in the central hole, is of sufficient
size that the air handled by it is
adequate to operate the air turbines
that drive the shield -water pumps and
the electric generators., These
turbines, pumps, and generators are
also located in the central hole. The
remainder of the reflector -coolant
radiator is divided into six equal
parts that are located in the triangular
spaces between the engines, outboard

of the engine center -line circle. A
portion of the air flow from each of
these reflector -radiator sections is
used to operate six air turbines that
drive six fuel pumps and six reflector -
coolant pumps. The air turbines are
located in the triangular spaces
between the engines, outboard of the
engine center -line circle, The fuel
and reflector -coolant pumps are located
in the central hole, the power being
transmitted by gears and shafting
from the air turbines. Space has been
left in this section of the fuselage
for the installation of the rear
landing gear, which is shown dotted.

Power Plant Weight. The turbojet
engine weight was calculated by each
of three methods: (1) the empirical
method described in the TAB report, ¢’
(2) the method of Rand Corporation,¢*?
and (3) by using the specific weight
data (pounds of engine weight per
pound of sea-level air flow), published
by the manufacturer, for an advanced
turbojet model. Since the engine is a
proprietary model, its identity will
not be divulged.

The last method yielded the highest
estimated weight and was employed in

 

(3)Repart of the Technical Advisory Board,
ANP-52 (Aug. 4, 1950).

*)R. S. Schairer, R. B. Murrow, amd C. V.
Sturdevant III, Bomber Capabilities - Turboprop
and Turbojet Power Plants, R-143 (Aug. 1, 1949).

17
81

i
DWG.17668
REFLECTOR COOLING RADIATOR (7 PARTS)

/ /‘MAIN ENGINE TURBINE (6}

FUEL PUMP (AND PIPE) (&)
MAIN ENGINE COMPARTMENT 6y -

B A — — ——_ | ™ FUEL AND REFLECTOR COOLANT 7
SHIELD WATER RADIATOR (1) ~— =" PUMP AIR TURBINE (AND DRIVE
SHAFT) (6) \{
—REFLECTOR COOLANT
, H PUMP (6) /‘Q\.fifi
‘ >

________ : - @}g

2\
. , X
5
I COOLING RADIATOR

T
Tz
&7 \. .
{ 7 PARTS)
\ i
REACTOR SHIELD ASSEMBLY {{) e ;g
e _ - \4‘©

SECTION A-B MAIN TURBOJET ENGINE (6)

 

 

 
   
   
 
 

 

 

 

 

 

 

  
   
  

ELECTRIC GENERATOR AIR TURBINE (2)

ELECTRIC GENERATOR (2)
____\ _ 444 1 E - S
NN : 1

*__WATER PUMP AiR
TURBINE (2)

WATER PUMP (2}

 

 
  
  
 

 

 

 

‘ REFLECTOR

 

 

 

 
    
  
   
 
 
    
   

MAIN ENGINE COMPARTMENT (6

 

SHIELD WATER
RADIATOR {1}

 

REACTOR SHI
ASSEMBLY (1)

MAIN ENGINE TURBINE (6}
FUEL-TO-AIR RADIATORS (6)

     
   
 
    

1ELD WATER RADIATOR (1)

 

REAR LANDING GEAR /C 1
(APPROXIMATE POSITION) . /
I
SECTION A-A SECTION C-C
NOTE: NO. iIN PARENTHESES INDICATES

NUMBER OF SIMILAR ITEMS IN
POWER PLANT,

Fig. 7. Power Plant Arrangement.

AANLS NOISHId
 

the weight summaries (a specific weight
of 15.8 1b per pound of sea-level air
flow or 33.2 1b per lb/sec design-
point air flow). It was assumed that
the weight increase associated with
the longer shaft, which was needed
because of the heat exchangers, was
compensated for by omission of the
combustion chambers. The air flow of
the six engines of the aircraft con-
sists of three parts: 1751 lb/sec
passes through the entire engine;
approximately 111 lb/sec is bled at
the eighth compressor stage; approxi-
mately 86 1lb/sec is bled at the first
compressor stage. The engine weight
was calculated by assuming 33.2 1b of
engine per lb/sec of air flow as the
flow that passes through the entire
engine; 40% of this value was assumed
for the air flow bled after the eighth
compressor stage; and 10% was assumed
for the air flow bled after the first
compressor stage. The weight of the
engines, less radiators, therefore is
59,900 pounds. The weights of the main
engine radiators and the reflector-
and shield-coolant radiators are
presented and discussed in the section
on ““Power Plant Radiators.” The total
weight of the main radiators, including
baffles, structure, headers, circu-
lating fluid, etc., is 23,900 pounds.

The weight of the auxiliary radi-
ators, pumps, air turbines, electric
generators, and liquid piping, in-
cluding circulating liquid for all the
radiators, is estimated at 5000 pounds.
The weight of the inlet and exhaust
air ducting was calculated by a method
similar to that used in the TAB re-
port®) and found to be 10,300 pounds.
Therefore the weight of the entire
power plant is 99,000 pounds,

POWER PLANT RADIATORS

All the nuclear powered aircraft
studied to date require heat transfer
equipment with surface-to-volume and
surface -to-weight ratios beyond those
required innormal industrial practice,

NUCLEAR-POWERED AIRPLANE

To achieve the ratios required, close
surface -to -surface spacing and thin-
walled surfaces must be used. These
design criteria, coupled with the high
operating temperatures and the strong
incentives to minimize pressure loss,
create heat exchanger design problems
without precedent. Various heat
exchanger lattices have been explored,
and, as might be expected, an improve-
ment in performance or compactness
would increase fabricational diffi -
culties and probably decrease dura-
bility., Determination of the best
compromise between these conflicting
considerations will require a con-
siderable amount of fabricational
development and functional testing by
a competent heat exchanger manu -
facturer, The radiators described
here are believed to be in the proper
surface area, size, and weight range,
but it is not intended to imply that
any radiators ultimately developed for
this application will resemble in
detail those illustrated (Figs. 8, 9,
10).

Physical Description. Figure 8
shows a representative fuel -to -air
radiator; Figs. 9 and 10 show the
reflector - and shield -coolant radi -
tors. The three different types of
radiators in the power plant are of
the same general design, that is, the
tube and fin type with the liquid
passing through the tubes and the air
across the tubes. Each tube is bent
into a serpentine coil and aligned so
that the air flows across the tube
along the axis of the tube coil. This
arrangement permits the combination of
a counterflow log mean temperature
differential and a crossflow heat
transfer coefficient., The various
serpentine coils are arranged in the
over -all lattice so that the individual
tubes form a conventional, triangular
pattern.

The over -all radiator dimensions
resulting from this design are generally
of the order of several inches thick,
2 to 18 in. high and 50 to 350 f¢

19
0¢

_——
DWG, 1 7669

by SECONDARY DISCHARGE HEADER
DISCHARGE HEADER

  
       

SUPPLY HEADER

SECONDARY SUPPLY
HEADER

  

 

INSIDE SHELL

  

 

 

 

DEVELOPED SECTION A-A

  
   
  

DISCHARGE HEADER
TYPICAL ENGINE
GENERAL ARRANGEMENT

SRR
SUPPLY HEADER wR N { 0 1FT

¢ SECONDARY DISCHARGE HEADER

€ SECONDARY SUPPLY HEADER
e . == - = s
50.20 x 3.24 x 0.010-in. T4 _‘. §

 

 

 

 

 

 

 

 

 

 

c e

_ PLATES; SPACED ON 9= o

TUBING, 0.08-in. 0D, 0.042 CENTERS. o ©
0.01-in. THICK WALL. 0.4732 in.

500 TUBES PER

RADIATOR SECTION B-B

=

7 315 in.

TYPICAL RADIATOR UNIT

PLAN (fimfi
INCH

 

Fig. 8. Fuel-to-Air Engine Radiator.

Ad0LS NI9ISAd
 

 

RADIATOR FOR: REQ'D | DiM. “A" [ DIM. "B’
WATER PUMP AND GENERATOR 12 6% | 25k
FUEL PUMP AND REFLECTOR COOLANT PUMP 2 8l | 17Y%

 

 

 

 

 

 

 

 

    
 
 
    
     

  

TUBES 0125 0D » Q10 |

HEADER
! 7 T3 -0010 THICK

e FINS 0042 ON ¢ ~
8 27, —————‘l

Fig. 9.

        

1%

long. Obviously, some method of
dividing the radiator into sections
and arranging these sections into a
somewhat more compact space 1is needed.
Accordingly, each fuel -to -air radiator
has been divided into a number of
sections of equal length and these
sections grouped cylindrically like
the teeth of a spur gear (Fig. 9).
There are from one to three sections
of radiator (stacked one above the
other) to each “tooth of the gear,”
and there are 12 teeth in all, The
faces of the radiator sections are
parallel to the normal path of the
air flow, and therefore the air must
be turned 90 deg to enter the radiator
and then turned back 90 deg upon
leaving the radiator. This 1s accom-
plished by dividing the space between
the “ gear teeth” into two parts with a
reinforced sheet that connects the
front of one “tooth” with the rear
of the next., The space between two
teeth therefore acts as the inlet air
duct for one tooth and the outlet air
duct for the other. To permit control
of the turbine inlet temperature in
relation to fixed reactor temperatures,.
provision has been made for a control -
lable by-pass in the reinforced sheet
that will permit the engine air to
by -pass the radiator, 1f desired.

  

NULCEAR-POWERED AIRPLANE

DWG 1!570
ALL DIMENSIONS ARE IN INCHES
A —

 

 

B
FIN LENGTH

 

   
  
 

1% 0D x Vg WALL-TAPERED TO 3 0D

Refl ector-Coolant Radiator.

The fuel is brought to the radiator
from the reactor in a 3 -in. pipe and
distributed to the 12 teeth by a
tapered ring header; short, constant-
diameter lines perpendicular to the
ring header lead to the radiator
sections, and long, tapered tubes
parallel to the “gear teeth” feed the
individual serpentine coils. The
outlet headers are similar to the
inlet headers described above.

The precise division and arrangement
of the auxiliary radiatorsis different
from that of the fuel -to-air radiators,
but the principle is similar.

Al]l radiators were designed with
the tubes having both large, common,
sheet fins (Figs. 8 and 9) and indi -
vidual round fins (Fig. 10). Either
of these alternate methods of con-
struction would result in approximately
the same radiator performance. The
former is probably preferable from a
fabricational viewpoint.

Radiator Design Relationships. The
following relationships for heat
transfer and pressure drop were used
in designing the radiators.

Air-Side Heat Transfer.
lation was made froma curve by Kern($)
that was based on the data of Jameson,

A corre-

 

(S)D. Q. Kern, Process Heat Transfer, McGraw-
Hill, New York, 1950,

21
¢e

   

TYPICAL OUTLET OR (NLET
NOTE: CUT-AWAY SHOWN /s
WITHOUT FINS.

  
    
 

     

 

 

 

  
  
    

          

  

 

 

 

 

 

 

 

 

 

     
  

 
 

 

=
=3, NOTE: FINS 0010 THICK AT 0.042
o= e ON §:534 FINS PER UNIT
: FINS 3291 WIDE x 23,625

END VIEW — LOOKING AFT
NOTE: 14 RADIATOR UNITS

 

 

 

TYPICAL TUBE PATTERN
ALTERNATE CONSTUCTION

DWG.1767

63 HOLES THIS ROW
61 HOLES THIS ROW

 

 

 

 
  

 

 

 

 

 

 

 

   
  
  

D - = = —
_% % o:9?°°°°0°OOO°O°0°0°0°o°o°o°o°u°o°o°o°o°o"d°c°c°o°o°o°0° T‘ %=
0.375+ =% oF sYm, o s
=l l=0375  TYPICAL HEADER DETAIL 3
—=— FORWARD AFT — - |-— 1109
INLET HEADER J’C_’f .
1
! ' AR OV Y 0324R| -1 --—% OUTLET HEADER
. L 2
- I — ::.:__‘;-_ _q’- BANK r 1
- : OUTLET HEADER 8 Tl ®©
AR ‘ \ st I
‘ RADIATOR 85'% 1D g
__________ 4 EE [l e 34 DIA. DISK 0010 THICK
A E— i OUTER SHELL o | B 22 AT 0.042 ON € - 534 DISKS.
i - T Y
) MR OUT \ N N
t 1 i{r
— == £ BANK ik
- (- - ]{fjr
= 1 L
= \N | ! I’
MR bil INLET HEADER
INLET
OUTLET HEADER % INLET HEADER
3-in. 1D FEEDER LINE FEEDER RING
! SECTION A-A
SECTION A-A 1
INCHES INCHES o914 0D OF TYPICAL RADIATOR UNIT
INNER SHELL
4 8 12 0O t 2 3
-
8
w7
o8
3‘291—-L ' or
0648 & INLET HEADER &8s
£ COLLECTOR RIN ! =0.324 58
LE & ol w0187 I 0,324 Io
=" A ;
°

SEE TYPICAL HEADER
DETAIL

SECTION B-B —TYPICAL TUBE PATTERN

NOTE: ALL DIMENSIONS IN INCHES

Fig. 10. sShield-Water Radiator.

NOTE: SEE ALTERNATE CONSTRUCTION.

AdNLS N9IS3a
 

Foster-Wheeler, and Tate and Cartinhour:
Nu = 0.092 Beo- 723 Pr0.33 ,

2 surface of fin and tube
DH=

 

7 projected perimeter of fin and tube ’

where Nu = Nusselt number, Re =
Reynolds number, Pr = Prandtl number,
and D, is used for the diameter term.

Air-Side Pressure Drop. The data
of Gunter and Shaw'®’ were used.

Fin Efficiency. The curves of
Gardner(’’ for circular fins of uniform
thickness were used.

Liquid-Side Heat Transfer. The
calculation was made by using the
following relationship:

Nu = 0,023 Re® ® Pr’-* .,

Liquid-Side Pressure Drop. The
calculation was made by using the
following equations:

L v?
a“) = 4F{f T »
D 2g
_ 0.046
Beo. 2
where
AP = pressure drop,

liquid density, 1b/ft?,
friction factor,
= equivalent length-to-diameter
ratio,
liquid velocity, fps,
gravitational constant,
ft/secz.
An allowance of 75 equivalent diameters
was taken for the pressure drop in an
180-deg bend.

The following liquid properties
were used both for the fuel and for
the reflector-cooling salt.

32,2

112 1b/ft3
0.39 Btu/1b.°F
0.5 Btu/hr- ft2 (°F/ft)

Density

Specific heat
Thermal conductivity
Viscosity 2 centipoises

Fuel-to-Air Radiator. A number of
fuel -to -air radiators were designed and

(6)A. Y. Gunter and W. A. Shaw, Trans. ASKE 67,

643 (1945).
(7)K. A. Gardner, Trans. ASME 67, 621 (1945).

NUCLEAR-POWERED AIRPLANE

a complete tabulation of the geometry
and performance of these radiators is
contained in Table 8., All values
listed in the table and mentioned
below are for the radiator for one
of the six engines. The radiator
design actually used in the power
plant is presented in column 1. It
is designed to transfer 53,500 Btu/sec;
the fuel enters at 1500°F and leaves
at 1000°F; the air enters at 544°F and
leaves at 1250°F; the air flow 1s
292 1b/sec and the fuel flow is 275
lb/sec; the inlet air pressure 1is
42.4 psia.

The radiator geometryis as follows:
The tubes are 0.06 in. ID with 0.020
in. walls., The fins are 0.25 in. in
diameter, 0.010 in. thick, and spaced
24 to the inch. The tubes are arranged
ina triangular pattern with a 0.25-in,
center-to-center spacing. The tube
material is Inconel, and the fins
are type 430 stainless steel., The
radiator face area is 67.2 ft?; the
radiator height is 4 in, and there are
18 banks longitudinal to the flow.

The radiator was designed to have
an air-side pressure drop of 10% of
the inlet pressure and a liquid-side
pressure drop of under 75 psi. The
liquid-side pressure drop for the
radiator of column 1 is 23 psi. The
fuel volume contained in the radiator
core is 1,15 ft?, A manifold system
was designed for the radiator of
column 1 (but not for any of the
other radiators) that contained 1.4

ft3 of fuel and caused a pressure
drop of 57 psi., The radiator, 1in-
cluding core, structure, headers,

baffles, contained fuel, etc. weighed
about 4000 pounds.

Columns 2 to 15 of Table 8 indicate
the effect of variations in geometry
and performance of the fuel-to-air
radiators. Columns 2 to 11 1llustrate
changes in geometry. Columns 2 and 3
show the effect of varying the number
of banks while holding the air-side
pressure drop constant, Column 4 is
similar to column 1, except that the

23
¥e

 

 

TABLE 8. FUEL-TO-AIR RADIATORS
1 2 3 4 H 6 7 8 9 10 11 12 13 14 15
Heat transfer, Btu/sec 53,500 53,500 53,500 53,500 53,500 53,500 53,500 53,500 53,500 53,500 53,500 57,900 49,300 56,100 55,600
Air flow,1b/sec 292 292 292 292 292 292 292 292 292 292 292 273 253 305 292
Fuel flow, lb/sec 275 275 275 215 275 275 275 275 275 275 275 297 269 288 283
Air inlet temperature, °F 544 544 544 544 544 544 544 544 544 544 544 430 544 544 544
Air outlet temperature, °F 1250 1250 1250 1250 1250 1250 1250 1250 1250 1250 1250 1250 1350 1250 1275
Fuel inlet temperature, °F 1500 1500 1500 1500 1500 1500 1500 1500 1500 1500 1500 1500 1500 1500 1500
Fuel outlet temperature, °F 1000 1000 1000 1000 1000 1000 1000 1000 1000 1000 1000 1000 1000 1000 1000
Air inlet pressure, psia 42.4 42.4 42.4 42,4 42.4 42.4 42.4 42,4 42.4 42.4 42.4 28,3 42.4 42,4 42.4
Tube ID, in. 0.060 0.060 0,060 0,060 0,060 0.080 0.10 0.10 0.10 0.10 0,10 0.060 0.060 0,060 0,060
Tube wall thickneas, in. 0.020 0,020 0,020 0.010 0.0125 0,0125 0.0125 0.012§ 0,0125 0,0125 0,0125 0.010 0.010 0.0125 0.0125
Fin diameter, in. 0.25 0.25 0.2§ 0.20 0.17 0.21 0.2§ 0.286 0,3125 0.25 0,25 0.20 0.20 0,17 0,17
No, of fins per in. 24 24 24 24 24 24 24 24 24 24 24 24 24 24 24
Fin cthickness, in, 0.010 0.010 0.010 0.010 0.010 0,010 0.010 0,010 0.010 0.010 0.010 0.010 0.010 0.010 0.010
Tube spacing, in. 0,25 0,025 0.025 0.20 0,17 0,21 0,25 0.286 0,3125 0.286 0.219 0,20 0,20 0.17 0.17
Tube material Inconel Inconel Inconel Inconel Inconel Inconel [nconel Inconel Inconel Inconel Inconel Inconel Inconel Inconel Inconel
Fin material 430 S8 430 S5 430 58 Inconel Inconel Inconel Inconel Inconel Inconel Inconel Inconel Inconel Inconel Inconel Inconel
No. of banks {(longitudinal to flow) 18 20 16 18 18 16 16 16, 18 18 20 10 16 26 20 20
Frontal area, ft? 67.2 71.6 63.2 57.2 [ 65.6 T0.5 36.6, 67.6 62.0 58.6 85.6 69.4 59.2 57.1 57.1
Radiator height, in. 4.0 2,52 1.6 6.0 7.2 5.0 9.0 13.8, 6.85 9.75 6.6 18 6,0 3.0 9.0 10.2
Air inlet velocity, fps 83.0 78.2 88.5 97.1 99 102 96 96, 89 91 97 102 92.7 80.7 111.6 117.6
Liquid velocity, fps 6.6 3.7 15.2 9.3 7.7 15.6 4.7 12.5, 5.0 T.6 6.8 7.3 7.0 4.9 7.4 8.4
Air-side heat transfer coefficient,
Btu/sec* ft2-°F 0,0288 0,.0276 0.0303 0,0346 0.0355 0.0340 0.0311 0.0307, 0.0293 0,0290 0.0318 0.0330 0,0270 0,0307 0.0402 0, 0402
Fuel+side heat tranafer coef.
ficient, Bru/sec: ft2:°F 0,60 0.38 1.12 0.82 0.47 0,99 0.40 0,78, 0.40 0.62 0,47 0.58 0.67 0,47 0.46 0.52
Air-aide pressure drop, psi 4.24 4.24 4.24 4,24 4.24 4.24 4.24 4.24, 4.24 4.24 4.24 4,24 2.83 4.24 6.36 6.36
Fuel-side pressure drop, psi
Core 23 7.6 147 44 49 29 12,5 68, 12.3 35.5 16.8 17.4 34.3 18,6 42 70
Headers 57
Fuel volume, ft?
Core 1.1§ 1.34 0.954 1.222 1.68 2.08 3.0 2.19, 2.75 2,15 2,712 2.54 1.32 1.82 1.59 1,59
Headers 1.4
Radiator weighe, 1b 3980 4540 3520 3040 3110 3160 3550 3570, 4110 3940 3340 3070 3220 3840 3020 3020

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

AdALS NOISId
 

tube wall thickness has been changed
to 0,010 in,, the fins to 0.020 in.,
and the fin material to Inconel.
Columns 5, 6, and 7 indicate the
effect of varying tube inner diameter.
The tube wall thickness in columns 53,
6, and 7 is 0.0125 in., and the ratio
of fin diameter to the tube outer
diameter has been held constant at 2,
Columns 8 and 9 are similar to column
7, except that the ratio of fin diameter
to tube outer diameter has been varied.
Columns 10 and 11 are also similar
to column 7, except that while the
fin diameter has been held constant,
the tube spacing has been changed.
(This is possible only if the tubes
are individually finned. The large
sheet fin type of construction will
not permit this variation.) In the
radiator described in column 11, the
fins are actually interlocking.
Columns 12 to 15 indicate the effect
of variations in radiator performance.
Since variations in radiator per-
formance will cause changes 1in power
plant performance, these radiators
were all designed so that the thrust
of the power plant remained constant.
Column 12 is similar to column 4,
except that is is designed for air
inlet conditions that correspond to a
compressor pressure ratio of 4 instead
of a compressor pressure ratio of 6,
as 1s found in the actual power plant,
Column 13 is similar to column 4,
except that the air outlet temperature
has been raised from 1250 to 1350°F.
Columns 14 and 15 are similar to
column 5, except that the air-side
pressure drop has been increased
from 10 to 15% of the air inlet
pressure. In column 14, the air
outlet temperature was maintained at
1250°F, but in column 15 it was
raised to 1275°F, This temperature
was selected so that the thrust per
pound of air handled by the power plant
is the same in columns 5 and 15,
Auxiliary Radiators. The designs
of the reflector - and shield -coolant
radiators are quite similar to that of

NULCEAR-POWERED AIRPLANE

the fuel -to-air radiators, A de-
cription of their geometry and per-
formance is given in Table 9.

AIRPLANE

In accordance with the general
premises of the ‘“Introduction,” an
airplane 1s presented that preliminary
studies indicate will meet the re-
quirements for flight at Mach 1.5 at
45,000 ft with the designed power
plant. No attempt has been made to
present a final design; the aim 1is,
rather, to present a reasonably
plausible design that may serve as a
starting point for more detailed
study. The general configuration of
the airplane, an aerodynamic calcu-
lation of the airplane lift-to-drag
ratio, a brief consideration of the
sea-level performance of the airplane,
and an estimate of the weights of the
various components of the aircraft
structure are presented.

Airplane Configuration. Figure 11
shows the general configuration of the
airplane, and Fig. 12, a longitudinal
section, shows the location of the
crew, reactor, and power plant. The
reasoning governing the location of
the various items in Figs. 11 and 12
is presented in the following.

The center of lift and center of
gravity of the aircraft, which, of
course, coincide, were taken as the
reference point. The wing and tail
were placed suitably, forward and
aft of the center of lift, so that the
resultant of the lift of the wing and
horizontal tail surface occurred at
the center of lift, and the center of
lift of the horizontal tail was 85 ft
from the airplane center of lift.
(It may be noted in Fig. 11 that a
triangular planform is used for the
wing and horizontal tail surface. It
1s normal practice 1n current tri-
angular-wing aircraft to have no
horizontal tail surface but, rather,
to use elevons in the wings to provide
control in the pitch direction. The
moment of inertia of this aircraft,

25
 

DESIGN STUDY

 

TABLE 9. REFLECTOR- AND SHIELD-COOLANT RADIATORS

 

 

REFLECTOR-COOLANT SHIELD-COOLANT
RADIATOR RADIATOR

Heat transfer, Btu/sec 16,050 3210
Air flow, lb/sec 113 82.6
Liquid flow, lb/sec 201 59.4
Air inlet temperature, °F 430 145
Air outlet temperature, °F 1000 300
Liquid inlet temperature, °F 1200 350
Liquid outlet temperature, °F 1000 300
Air inlet pressure, psia 29 8.44
Liquid inlet pressure, psia 200
Tube ID, in. 0.10 0.10
Tube wall, in. 0.0125 0.0125
Fin diameter, in, 0.25 0.375
No. of fins per 1in. 24 24
Fin thickness, in. 0,010 0.010
Tube material Inconel Aluminum
Fin material Inconel Aluminum
No. of banks longitudinal to flow 10 10
Frontal area, ft? 30.2 49,4
Radiator height, in. 7.2 23.1
Liquid-side pressure drop, psi 28 33
Air-side pressure drop, psi 2.9 0.844
Radiator weight (including baffles,

headers, structure, contained

liquid, etc.), 1b 1400 800

 

 

 

however,

1s going to be quite large

In order to avoid changes in the

because of the heavy crew shield in
the noseof the airplane, and therefore
a horizontal tail was added to secure
a longer lever arm for the control
forces in the pitch direction. Whether
this 1s actually necessary is not
known; the problem of control in the
pitch direction is considered further
in a subsequent paragraph.) The size,
shape, and proportions of the wing
and tail surfaces were determined from
aerodynamic considerations and are
discussed i1n the following subsection.

26

balance of the airplane when the bomb
load 1s dropped, this load was located
at the center of lift. The reactor
and power plant were grouped suf-
ficiently aft of the center of lift
to balance the moment caused by the
crew shieldin the nose of the aircraft
and the resultant moment caused by
the weights of the various components
of the airplane structure., The engines
were placed behind the reactor to
afford some shadow shielding of the
fore portion of the aircraft; and,

 
LG

100

 

200 400 600
SCALE IN INCHES

800

1000

Fig.

11.

 

 

 

 

 

 

 

 

 

 

 

Airplane Configuration.

T
DWG. 17672

 

ANVIdUIV dIdIn0d-HVATONN

 
 

DESIGN STUDY

 

 

OWG. 17!73

 

 

 

 

 

  

 

 

 

 

 

 

 

 

 

 

 

 

Fig. 12.

furthermore, they were placed as close
as possible to the reactor to minimize
the fuel volume in the ducts to and
from the main engine radiators.

The cowl, the central portion of
the fuselage, was made large enough in
diameter to permit the passage of the
engine air flow around the reactor
shield; and it was extended rearward
to the engine exhaust nozzles and
forward far enough to permit the air
intake to be ahead of the wing. The
engines were placed as far out in the
cowl as they would go. The crew and
crew shield were placed in the nose
ogive, forward of the cowl, which gave
a separation distanceof 120 ft between
the reactor and the crew compartment
(center to center). The diameters and
proportions of the nose ogive and tail
boom were chosen to meet the spatial
and structural requirements and to
give low aerodynamic drag. (The nose
ogive 1s located on the center line of
the airplane; there 1s therefore
considerable air inlet area above the
ogive. Recent NACA aerodynamic studies
indicate that at high angles of attack
this portion of the air intake may be
““smothered” by a very thick boundary
layer. This difficulty could be
alleviated by raising the nose ogive
with respect to the air inlet until
the upper surface of the ogive was
actually an extension of the cowl.)
The diameter of the nose ogive and the
inlet air-flow area requirements are
such that they permit the cowl to be
tapered in the manner shown in Fig. 11,

28

Longitudinal Section of Airplane,

Airplane Lift-to-Drag Ratio: The
various aerodynamic formulas in this
section were taken from the following
references:

1. Eugene S. Love, Investigations at
Supersonic Speeds of 22 Triangular
Wings Representing Two Airfoul
Sections for each of Eleven Apex

Angles, RM L9DO7, May 10, 1949,

2. Generalized Lift and Drag Charac-
teristics at Subsonic, Transonic,
and Supersonic Speeds, Consolidated
Vultee Aircraft Corporation,
Fort Worth, Texas, FZ AO4la,
November 27, 1950,

3. NACA Conference on Aircraft
Propulsion Systems Research, Lewis
Flight Propulsion Laboratory,
Cleveland, Ohio, January 18 and
19, 1950.

4. Notes and Tables for use in the
Analysis of Supersonic Flow,
NACA Technical Note 1428, December
1947 .

The airplane wing is of triangular
planform with a 60-deg sweep and a 3%
thickness-to-chord ratio. The wing
profile 1s that of a circular-arc
airfoil with an elliptic leading edge.
The nose ogive, tail boom, and cowl
are parabolic bodies of revolution,
and the nose ogive and tail boom are
pointed at the ends. The following
lift and drag formulas were used for
the calculation of the airplane
lift-to-drag ratio.

 

enT e LSS A
 

 

Wing
Wave drag:
CD_
= mB tan € - 0.65,
(TR)?
where
C, = wave drag coefficient based on
v exposed planform area,
TR = thickness ratio = ratio of

maximum thickness to chord,

m= 4.9 at Mach 1.5,
B =M} -1

M = flight Mach number,
€ = 90 deg minus the sweep angle.
Induced drag:

Aspect Ratio CD_/Ci

4 0.341

3 0.342

2.31 (e = 30°) 0.352

where ’
C, = induced drag coefficient based
' on exposed planform area,

C, = lift coefficient based on

exposed planform area.

Friction drag:

 

0.0306
C, = ,
YV - 1 5/7
Re'/7 | 1 + M
4 0
where
C, = friction drag coefficient

f based on total wetted surface,
= ratio of specific heats of air,
Reynold’s number, Vicpo/pu ,

forward velocity, f{ps,

average wing chord, ft,

ambient density, 1b/ft?,

= ambient viscosity, lb/sec*ft.
ptimum lift coefficient:

 

where

G, =

opt

lift coefficient at maximum
lift-to-drag ratio (based on
exposed planform area),

NUCLEAR-POWERED AIRPLANE

AW
CDo = CD' + CDf X o
A= exposed planform area

airplane gross weight ,

Gy * P Vg

opt

 

Aw = total wetted surface ™~ 24.
Total wing lift:
L = ( Ag ,
ept
where

L_ = total wing life,
q = %pV:.
Total wing drag:

D

v = |:CD° + CD] Aq ’

¢
where

D

total wing drag.

Tail

The horizontal tail is geometrically
similar to the wing and has an area
equal to 20% of the wing area. The
vertical tail has a 45-deg sweep
angle, a 3% thickness ratio, and an
area equal to 15% of the combined
wing and horizontal tail area. The
lift and drag of the tail surfaces
were calculated by using the same
formulas as those used for the wing.

Nose Ogive and Tail Boom

 

Wave drag:
! 10.7
D ?
v (FB)2
where
C; = wave drag coefficient based on

maximum frontal area,

fineness ratio (ratio of length
to maximum diameter).

Friction drag:

 

 

0.0306

C; = 1.05 X%

f 7_1 /17

Re!/ 7|1+ M2
4 0

where
C, = friction drag coefficient based

f on total wetted surface,

Re = Reynold’'s number, V,Lp/pu,

L

length, ft.

29
 

DESIGN STUDY

Total nose ogive and tail boom
drag:

 

 

 

ratio of outlet
station of

Cowl area ratio =
area to area at
maximum diameter,

 

D' = [C;' A"+ Cfif A:] q. L = length of the aft-portion of the
“cowl.
wh?re _ Friction drag: The friction drag
D" = total‘drag of the nose ogive coefficient for the cowl may be
' and.tall boom, calculated from the formula used for
A" = maximum frontal area, obtaining the fuselage friction drag
A, = total wetted surface. coefficient if L is defined as the
Cowl length of the cowl.
Wave drag of fore-portion: A table Total cowl drag:
of wave drag coefficients for the
D" = {-[CB J + [CB ] }, A"q * Cp Alq
v fore-portion *laft-portion f
portion of the cowl forward of the  jere
station of maximum diameter follows: D" = total drag of cowl,
COWL AREA RATIO L/D C; VM, A" = maximum cowl frontal area,
. A’ = total cowl wetted surface.
0.4 10 0.0040 It is assumed that there 1s no
8 0.0065 increase in drag due to the inter-
6 0.010 ference of wing, fuselage, cowl, and
4 0.0185 tail; therefore
0.6 10 0.0020 L L, * Ly,
8 0.0025 D . D + D + D Dt + Dn
6 0.0040 airplane ht vt
4 0.0080 ‘"E%e L liteted
) = -to-
0.8 10 0.0010 airplane ::zgoane ' o-drag
8 0.0015 L,, = lift of the horizontal
6 0.0025 ¢ tail
4 0.0040 D,, = drag of the horizontal
1.0 All 0 tail,
where D , = drag of the vertical
C, = wave drag based on maximum tail.

v frontal area,

Cowl! area ratio = ratio of inlet area
to area at station of maximum
diameter,

L = length of fore-portion of the
cowl,

D= maximum diameter of cowl
section,

Wave drag of aft-portion: The wave
drag coefficient of the portion of
the cowl aft of the station of maximum
diameter may be taken from the same
table as the fore-portion by using the
following definitions:

30

Calculations made by using the above

formulas for the airplane of Fig. 12
give:
Wing
Cp, = 0.00227
2 .
CDi/C2 = 0,352
Cp = 0,00182
f
CDo = 0.00591
Cy = 0.1298
opt

 
 

CDs = 0.00591
A = 4620
L = 291,100
D = 26,540
Horizontal tail
Cp = 0.00227

Cy /CF = 0.352

Cp = 0.00190

 

NUCLEAR-POWERED AIRPLANE

Cowl

The cowl 1s evaluated as 1f the
nose ogive and tail boom were not
present, since their drag has already
been accounted for. The inside surface
of the cowl 1s actually engine ducting
and engines and its drag has already
been accounted for by the efficiency
of the various engine components.

[CB ] = 0,0100 (approx.)
YIfore

 

f
cDo = 0.00608 [cg ] = 0,0100 (approx.)
Yiaft
C; = 0,1315
opt CB = 0,00180
= f
CD‘ - O. 00608 AII = 269
Ay, = 924 A’ = 6330 (approx.)
L,, = 58,900 D" = 8150
_ The airplane lift-to-drag ratio is
Dy, = 5480 therefore
L 2 Lift _ 291,100 + 58,900 . 350,000 7,03
D/iiiotiame = Drag 26,540 + 5,480 + 3,350 + 6,280 + 8,150 49,000 T

 

Vertical tail
CD = 0,00435

CDf = 0,00198
CDo = 0.00831
A = 832
= 3350

vt
Nose Ogive and Tail Boom

The nose ogive and tail boom com-
bined are assumed to be similar to a
parabolic body of revolution of about
95 ft in length and about 10 ft in
maximum diameter,

CB' = 0.1185
Céf = 0,00181

A' = 78,5

A; = approx. 2000
D' = 6280

For the sake of conservation and be-
cause of the uncertainties present in
the lift-to-drag ratio calculation,
a value of 6.5 was used for the lift-
to-drag ratio; this leaves a con-
tingency of 0.53 in the ratio.

A calculation was made for a
rectangular wing of 3% thickness, 25%
and an aspect ratio of
3, and the lift-to-drag ratio was
about 10.55, as compared with about
10.98 for the delta wing.

Airplane Pitch Control. The problem
of controlling the aircraft in the
pitch direction may become acute be-
cause of the heavy weight of the crew
shield far forward in the airplane,.
For this reason it was decided to have
a horizontal tail surface with an 85 -ft
lever arm to provide this control
rather than elevons in the wings as
in the normal practice. A calculation
showed the mass moment of inertia of
the airplane in the pitch direction to

taper ratio,

31
 

DESIGN STUDY

be about 3.8 x 107 slug ft?2., If the
entire horizontal tail surface was
movable, as it is 1n some recent air-
craft, it would be possible to exert
a torque of 3.70 x 107 ft.1b (assuming
a maximum lift coefficient of 1.1 for
the surface). This would provide an
angular acceleration of 55.0 deg/sec?
to the aircraft, which would probably
be more than ample. The angular
acceleration in the pitch direction
required of a large airplane of this
type is not known at this time. If
this requirement were established,
some other arrangement of control
surfaces might prove more desirable,

Airframe Weights. The following
formulas for the weights of the various
components of the airplane structure
are taken from the TAB report(3) and
from various Rand reports, primarily
R-143, (%)

Wing Weight. The wing weight was
calculated in the same manner as in the
TAB report.

 

 

KfizSa
.KEA +'(772)‘4 [W f&(AJ _lfli fa()\;k)]
W, = 1.15 ’
1+ K,n S8 £.00
(TR) A "2
where
w_ = weight of wing,
W = lifting force provided by wing
= 291,100 1b,
K, = a constant = 4.0 1b/ft?,
K, = a constant = 12,5 X 10-% fe!,
A = wing area = 4620 ft?,
n = load factor = 4.0,
TR = thickness ratio = 0,03,
f, = 0.113¢%)
fa = 0.064(%)
fs = (8)
W, = distributed weight in wings = 0,
k = portion of span over which ¥,
is distributed = 0,
A = taper ratio = 0,
S = structural span (length of span

measured along the midpoints of

the chords) = 136.5 ft.

(S)Schairur, Murrow, and Sturdevant, op. cit.;
fl' fp, and f3 plotted on p. 120.

 

32

 

By the above formula,
is 46,000 pounds.

Tail Weight. The weights of the
horizontal and vertical tail surfaces
were estimated by two methods. The
method of the TAB report,(s) which
assumes that the total tail weight 1is
20% of the wing weight, resulted in a
total tail weight of 9200 pounds. The
method of the Rand report,(*’ which
gives relations for the weight of the
horizontal and vertical tail surfaces
similar to the wing relationship
above, resulted in a total tail weight
of 8100 pounds. The more conservative
estimate of 9200 pounds was used.

Fuselage Weight. The fuselage
weight was estimated by the same method
in the TAB report and in the Rand
report,

the wing weight

15.0 nL (W, + W, )

Wf = D

 

L 4.0 +
-1 108 D?

Wf = weight of fuselage,

fuselage maximum diameter,
Lf length of fuselage,

load factor = 4.0,

ch= weight of fuselage contents.

n o

3
"

This equation gives a weight of
8200 1b for the nose ogive and tail
boom and a weight of 21,700 1b for the
cowl. The total fuselage weight 1is
therefore 29,900 pounds.

Landing Gear Weight. The weight
of the landing gears was estimated by
the method of Rand{*’ which assumes
that the landing gear weight is 5.4%
of the gross weight of the aircraft,
This is slightly more conservative
than the TAB method, ¢(3) which assumes
the landing gear weight to be 5.0% of
the airplane gross weight. The weight
of the landing gear i1s therefore
18,900 pounds.

Controls Weight. The weight of the
airplane controls was estimated by
the method of the TAB report,(3) which

 
 

assumed the controls weight to be 0.6%
of the airplane gross weight. For
this airplane, the controls weight is
therefore 2100 pounds.

NUCLEAR-POWERED AIRPLANE

Total Airframe Weight. The total
weight of the airframe, including wing,
tail, fuselage, landing gear, and
aircraft controls is 106,100 pounds.

 

SEA-LEVEL

The optimization of engine and
radiator per formance was based on
design-point operation, but liquid-
line sizes and pump capacities were
based on the higher flow rates that
would be required at sea level., 1In
studying design-point performance, the
engines and radiators were sized to
permit the attainment of a stipulated
thrust., In considering sea-level
static performance therefore the
design is constrained by the geometry
selected to meet design-point (45,000
ft) conditions. These constraints
still permit broad operational latitude,
however, and additional operational
constraints were established to permit
solving for sea-level performance, as
follows:

Engine rpm: take-off engine speed was
selected as equal to design-point
engine speed.

Engine air flow: take-off air flow
was selected as equal to design-
point air flow on a corrected air
flow basis, that is, constant wN8/$,
where w 1s the mass air flow, &
varies as the compressor inlet
temperature, and & is the compressor
inlet pressure.

Reactor inlet and outlet temperatures:
take-off reactor inlet and outlet
temperatures were selected as equal
to those at design point. Since the
mean reactor temperatures are there-
fore substantially the same as at
design point, there is no need for
shimming. A higher reactor AT would
entail exceeding the established
metallurgical limits; a lower reactor
AT would entail large increases in
pumping power for a given power
abstraction,

PERFORMANCE

With these specifications, aspecific
solution for reactor power, fuel flow
rate, turbine inlet temperature and
engine thrust can be obtained. De-
creasing the operational altitude in-
creases the engine mass air flow,
which in turn increases the radiator
heat-removal capacity and therefore
demands increased reactor power, Were
the entire system tooperate at design-
point temperatures, the power flow
would increase directly with fluid
flow rates, and it would be necessary
for heat transfer coefficients to vary
with flow rate to the first power.
Actually, however, the heat transfer
coefficients will vary as flow rate to
some fractional power. Consequently
an increase in driving temperature
difference is required to permit the
higher sea-level powers. This re-
quirement for a higher temperature
difference causes the system to
stabilize at a lower turbine inlet
temperature than was attained at
design point (1125°F at sea level;
1250°F at design point). However, the
greatly increased air flow permits a
total thrust of 165,000 1b at sea
level, compared with 53,850 1b at
design point. This take-off thrust
appears to be adequate, since it
permits a calculated take-off ground
roll of approximately 2500 feet.

A reactor power of 640,000 kw is
required, This wil] increase the crew
radiation dosage but has not been
considered in connection with shield
design because of the presumably short
duration of operation at this power
level. Sea-level performance 1is
summarized in Table 10.

33
 

 

DESIGN STUDY

TABLE 10. SEA-LEVEL STATIC PERFORMANCE

 

 

 

PRESSURE (psia, TEMPERATURE (°F) WEIGHT FLOW
approx.) (1b/sec)
Fuel Circuit
Radiator inlet 292 1500 3130
Radiator outlet 25 1000 3130
Pump outlet 525 1000 3130
Reactor inlet 475 1000 3130
Reactor outlet 342 1500 3130
Air Circuit
Aircraft ambient 14.7 59
Compressor in 14.7 59 4137
Compressor out 87.7 461 3720
Radiator in 85.5 461 3720
Radiator out 78.0 1125 - 3720
Turbine in 76.1 1125 3720
Turbine out 24.1 728 3720
Jet 14.7 Jet vel. = 1360 ft/sec 3720

 

 

 

Air flow to shield coolant radiator, 182 lb/sec

Air flow to reflector coolant radiator, 235 lb/sec
Portion to air turbines, 88 lb/sec
Portion directly to jet, 147 lb/sec

Thrust from main circuit, 155,900 1b

Thrust from portion of reflector coolant airflow that goes directly to jet
assuming radiator outlet air temperature = 875°F, which comes from assuming

(6, /6 e, /6 9700 1b

Thrust from portion of reflector cooling circuit through air turbines and from
shield cooling circuit assumed = 0

f
Total thrust = 165,600 1b

De:)-ain rad, Des)refl. cool. rad. =

Maximum reactor tube wall temperatures

DESIGN POINT SEA LEVEL
Inside tube 1554 1583
Outside tube 1567 1608

Take-off ground roll to 110% of stall speed

stall speed = 165 mph, assuming all lift from wing
Cy = 1.1

ground roll = 2480 ft

 

34

 
 

SHIELDING

The shield design for the aircraft
requires an extension of the methods
described in the report of the Shielding
Board(!) to take account of the
delayed neutrons and fission-product
gamma rays from the exposed part of
the circulating fuel.

The first step in the shield design
was to assign fractionsof the radiation
tolerance to the several radiation
sources. This was done on the basis
of an approximate estimate of the
weight penalty for shielding each
component, After careful analysis,
the dose distribution can presumably
be revised, with some weight reduction,
but the analysis will not be made at
this time. Next, the crew shield was
designed to provide protection from
the delayed neutrons and gamma rays
from the unshielded circulating fuel.
Finally, the reactor shield was
chosen so that in conjunction with
the crew shield the primary radiations
would be approximately attenuated.

Structure scattering was calculated
separately and treated as aperturbation
on the design determined without it,
The crew shield thicknesses were then
slightly increased to take account of
the structure scattering.

ASSIGNMENT OF RADIATION CONTRIBUTIONS

A total -gamma-dose to total -neutron-
dose ratio of 3 was chosen, since
neutron shielding is accomplished
with less weight than gamma shielding.
Another reason for adhering to this
ratio is that much less is known about
the relative biological effectiveness
of neutrons, and by keeping this
contribution to a small part of the
total, the over-all uncertainty is
correspondingly reduced,

For both neutrons and gamma rays
the contributions through the crew
shield rear and sides were taken to
be the same. The reasoning in this
case was that although the area of

(1) geport of the Shielding Board for the

Aircraft Nuclear Propulsion Program, ANP-53
(Oct. 16, 1950).

NUCLEAR-POWERED AIRPLANE

ANALYSIS

the rear is much less than that of
the sides, the difference is almost
offset because the radiation entering
the rear is mostly unscattered and
hence harder than that incident on
the crew shield sides.

For radiation entering the front
of the crew shield, an appreciably
smaller contribution is assigned,
since not only is the radiation
scattered, and hence comparatively
soft but the area of the front shield
slab is small.

In distributing the contributions
between reactor and exposed fuel in
the radiators, account was taken of
the relative hardness (energy of 1
photon) of the radiations. For both
neutrons and gamma rays the radiator
radiations are more easily shielded
and hence these are assigned a smaller
contribution. The results of these
deliberations are given in Table 11,
In the following sections the numbers
Ia, etc. refer to the contributions as
listed in Table 11.

The biological toleranceis specified
as the maximum at any location in the
crew shield; the calculation of the
radiation level at all points of the
interior is beyond the scope of this
report. As an estimate, the maximum
is takento be the sum of contributions
from front, all four sides, and rear.
In Table 11, ‘““sides’ means total
contribution from four sides.

CONFIGURATION TO BE SHIELDED

The reactor is a sphere 3/5 ft in
with a 6-in. beryllium oxide
reflector. It releases heat at the
rate of 325 megawatts. The reactor-
to-crew separation distance is 120 ft,
and the radiator-to-crew separation

diameter,

distance is 132 feet. The fuel 1s
divided as follows:
Radiators and headers 15.6 ft3
Pipes, etc. 10.4 ft?
Reactor 8 fed
Total 34 £t}

35
 

DESIGN STUDY

 

 

 

 

 

TABLE 11. ALLOWED CONTRIBUTIONS TO THE TOTAL DOSE
DOSE 2
COMPONENT FLUX (neutrons/cm“*-sec)
rem/hr rep/hr
I. Neutrons
a. Radiators to rear 0.02 0.002 29,4
b. Radiators to sides 0.02 0.002 1/4 X 29,4 per side
c. BRadiators to front 0.005 0.0005 7.25
d. BReactor to rear 0.10 0,010
e. Beactor to sides 0.10 0.010
f. RBReactor to front 0.005 0.0005
Total 0.25 0,025
IT. Gamma Rays
a. Radiators to rear 0.250 0.250 5 X 10* (hard photons)
b. BRadiators to sides 0.300 0.300 See reference 1
¢c. Radiators to front 0.025 0.025 1,38 x 10* Mev/cm?.sec
d. Reactor to rear 0.100 0,100
e. Reactor to sides 0.05 0.05
f. Reactor to front 0.025 0.025
Total 0.75 0.75

 

 

 

 

The total circulation time for

fuel 1s 2.45 sec, of which 0,54
1s spent in the core, 0.25 sec in
headers inside the shield, 0.05
sec in reaching the shield exterior,
and 1.61 sec in the radiators and
external and return pipes.

the
sec
the

BASIC DATA FOR SHIELD DESIGN
Gamma ray equivalents:
1r 2 X 10° Mev/cm?,
1r/hr 5.5 X 10° Mev/cm?*sec,

= 2 X 10% hard photons/cm?*sec.
Neutron equivalents:

1 rep/hr = 10 rem/hr (biological dose),
1 rem/hr = 6930 fast neutrons/cmz'sec,
1l rem/hr = 14,700 delayed neu-

trons/cm?* sec (from Snyder’ s
BBE curves(?2)),
Delayed neutrons per neutron formed
in fission, 7.3 % 1073,
Neutrons per fission, 2.5,
Fissions per watt'sec,

3 x 10!°,

 

(20w, S. Snyder and J. L. Powell, 4 Joint
Project of the ORNL Health Physics Division and
the ORNL Suamer Shielding Session, ORNL.421,

36

Mean free paths for neutrons in
air ( from reactor):

E = 3 Mev, A = 130 meters,
E=0.5 Mev (delayed neutrons),
A = 40 meters,

Mean free paths for gammas in air
(from reactor):

E~2 to 3 Mev, A ~ 210 meters,
E~0.5Mev, A\ ~ 90 meters.
Beryllium oxide reflector:

= 2.8 g/Cms.
= 7.8 cm,

11.8 cm (for 3 Mev/photon),

Density
A

vent

My

Attenuation of beryllium oxide for
reactor neutrons = 1/6.8.
Relaxation lengths:

Prompt neutrons

In water 10 cm

In polyethelene 8.4 cm

In lead 9 cm

In iron 6 cm
Delayed neutrons

In water 2.7 cm

In polyethelene 2.26 cm

 
 

Gamma rays

In water 23 cm
In polyethelene 24.8 cm
In lead 2.2 cm
In 1ron 4 cm

The values for polyethelene are
based on its density of 0.93 for
gamma rays and its hydrogen density
of 8 X 10°%% cm~3, as compared with
water.

CALCULATION OF SHIELD DIMENSIONS

Delayed Neutrons into Crew Com-
partment Rear (Ia). The delayed-
neutron source strength is obtained
from the product of the power of the
reactor, the fissions per sec per
unit power, the total neutrons per
fission, and the delayed fraction.

S, = 3,25 % 108 x 3 x 10'°
x 2,5 x7,3x 103

= 1.78 X 10!? neutrons/ sec.
Of these, 15.6/34 are produced in the
radiators, one half are intercepted
by the reactor shield, and a further
factor of 1/3 is introduced to take
account of self-absorption in the
radiators.

Scattering calculations based on
single collisions in the fuselage
indicate that the number of delayed
neutrons arriving on the crew shield
rear will be increased by about 14%
because of the scattering.

The flux incident on the crew
shield rear is thus

[1.78 x 107 x (15.6/34) x 1/2 x 1/3
x 1.141/47(132 * 30.5)% = 7.6 X 107

neuntrons/cm?*sec.

The allowed tolerance for this component
is14,700 x%0.002 neutrons/cm?-sec=29.4
neutrons/cm?+sec. The lead in the
rearwall acts primarily asa scatterer;
so its attenuation is taken to be
only 1/2.

The water thickness for crew shield
rear 1s thus
t = 2.7 In[(7.6%107)/29.4]

39.8 cm of water

39.8 X (6.7/8) = 33.4 cm of
plastic (polyethelene).

dr

8.2 x 10!°

NUCLEAR-POWERED AIRPLANE

Delayed Neutrons to Crew Compartment
Sides (Ib). The delayed-neutron
source for scattering into the sides
includes neutrons produced in the
exterior piping, as well as one-third
of those produced in the radiators.

S,, = 1.78 x 10'7 x (15.6/3) X 34
+(10.4/34) = 8.2 x 10'°

The simple isotropic scattering
formula is used to obtain the flux
incident on the crew shield side
walls:

S = 2 X 107 neutrons/cm?*sec,
where d is the separation distance,
132 ft, and A is the mean free path
in air, 40 meters.

Scattering from the cowl into the
sides increases the flux by about
14%, as determined from single-
scattering calculation. The wing
contributes onlya negligible fraction.
The side wall thicknesses are, if only
one-quarter of the dose is allowed to
enter each wall,

2.7 1n [2 % 10% x (1.14/14,700)
X 0.002 %X 0.25] = 46.5 cm of water
= 39 cm of plastic .

Delayed Neutrons into Front (Ic).
The scattered neutron flux into the
front face based on isotropic scattering
would be (7/2) - 1 times that incident
on the side.

The flux on the front is thus

[(m/2)- 1] x 2 x 10°

= X 8 2,
Thicknesseg'gie 10® neutrons/cm®-sec.
tgg= 2.7 In [1.14 X (10%/14,700)

x 0.0005]
44.5 cm of water

= 37.3 cm of plastic.

Gamma Rays from the Exposed Fuel.
The time required for fuel to travel
from the reactive region to the
exterior is about 0.3 sec, and therefore
periods shorter than 0.3 sec can be
ignored. This is fortunate because
no data exist for delay times less
than about 0.25 sec. The available
data, however, are not by any means

37
SRR A ot SRR R

 

DESIGN STUDY

adequate for the present purposes;
so the numbers used represent con-
servative estimates rather than well-
measured values. The data include
the work of Bernstein et al., (3:4)
who measured the number of gammas of
energy sufficient to photodisintegrate
deuterium and beryllium from the
fission products of U?3%, All numbers
discussed will be in terms of photons
per fission. Bernstein et al. obtained
a value of 2.5 hard gammas per fission
on the basis of deuterium disinte-
gration, but the value is quite un-
certain because 1.58 of this quantity
is attributed to a gamma of 2.25 Mev,
an energy which is so close to the
threshold that it is subject to
considerable cross-section uncertainty
for even slight energy variation.
A new determination made by Bell and
Elliott,¢(5) which gives a threshold
value of 2.237 Mev, further emphasizes
the uncertainty. The energy de-
termination for the fission-product
gamma rays was not good enough to
make Bernstein’s value meaningful .

Accordingly, reliance must be
placed on the data of Sugarman et
al., (%) who report that in the in-
terval 10 sec to 2 hr there are
0.8 photons of 2,2 Mev. By integration
of their extrapolated curves, it is
deduced that there is, at most, 3 Mev
of gammas per fission in the period
from 0 to 10 sec. For the present
purposes, 1t will be assumed that
there are 1.5 3-Mev photons per
disintegration. Ergen,(’) by inde-
pendent analysis, arrived at a value
of 0.5 hard gammas per fission; so it
appears that the value 1.5 is quite
conservative.

(3)
(1947).

(4)5. Bernstein et al., Yield of Photoneutrons
from U235 Fission Products in Be, AECD-1833
(Feb. 20, 1948).

S)R, E. Bell and L. G, Elliott, Phys. Rev. T4,
1552 (1948).

G)N. Sugerman et al., Radiochemical Studies:
The Fission Products, Book I, Paper 37, p. 371,
NNES IV, 9, McGraw-Hill, New York, 1951,

¥, K. Ergen, private discussion.

 

S. Bernstein et al., Phys. Rev. 71, 573

38

 

Gammas from Radiators into Rear
of Crew Compartment (IXa). The total
hard fission product calculation is
made in nearly the same manner as was
that for the delayed-neutron source.

S = 3,25 X 108 x 3 x 101!'% x ].5

nr

= 1.46 x 101°9 photons/sec .
Of these, 15.6/34 are produced in the
reactor, one half are intercepted by
the reactor shield, and a factor of
1/2 is taken for self-absorption,
As in the case of the neutrons, 14% is
added for structure scattering.

The gamma flux incident on the
rear of the crew compartment shield
is thus
[1.46 > 10'® x (15.6/34) x 1/2 x 1/2
X 1.141/47(132 % 30.5)2 = 9.4 x 109

hard photons/cm2+sec .
The rear crew shield plastic will
attenuate by a factorof exp (33.4/24.8)
or exp (1.35).

The compressor, forward of the
radiators, constitutes a shield
equivalent to about 1 in. of Fe and

gives an attenuation of exp (2.5/4)
or exp (0.625). The lead thickness
must therefore be

tp, (rear) = 2.2 In [(9.4 x 10°)/(0.25

X 2 X 10%) - 1.35 ~ 0.625]
= 22.4 cm of lead .
Gammas from Radiators to Sides (IIb).
For the radiator gammas scattered in
air, the pipes are included in the
source, and self-absorption is taken
as 1/4. The source is then

1.46 x 10'% x (26/34) % 1/4

= 2.8 X 10'% photons/sec .
In order to use the curvesin ANP.53,(8)
1t is necessary to convert this to
the equivalent source for a 50-ft
separation. This is

2.8 x 101 x (50/132)

= 1.06 X 10'® photons/sec .
Structure scattering is neglected
here because of the slant penetration
of the shield by the structure-
scattered gammas. Slant penetration
is probably more effective in the

®op. cit., aNP-53, p. 134,

 
 

attenuation of gammas than neutrons
because of the energy degradation
accompanying turning of gammas in the
shield.

According to the reference,(?)
6.25 cm of lead are required to reduce
the dose to 0.3 r/hr. Since some of
the lead is replaced by plastic, the
lead thickness 1is

6.25 - 39 x (2.2/24.8)
2.75 cm of lead.

tohe

u

Radiator Gammas into Front (Ilc).
For this calculation, gamma scattering
is assumed isotropic, with a cross
section equal to the average over-
scattering angles from 7/2 to 7, that
is, about 0.4 % 10°2% cm? per electron
per steradian,(?)

The electron density of air 1s
approximately
0.602 x 10%* x 14.4

¢ 22,412
= 3.87 x 102%° ¢cm-3

 

The effective mean free path is then

 

1
4m X 0.4 % 10-2% x 3,87 x 10?°
= 5.1 X 10%* e¢m = 510 meters .

The flux incident on the front face,
obtained by using the previous source,
1s

7
2.8 x 108 C—-— l) =
2

A
d

3.07 x 108,

5.1 % 10% cm ,
132 X 30.5 cm .

nwon

It will be reasonable to assign to
these gamma rays the energy for a
scattering through an angle of 7,
since for smaller angles of scattering
the slant penetration of the shield
will compensate for the higher energy.
The energy is 0.24 Mev, for which the
relaxation lengths of plastic and lead
are 8.4 and 0.147 cm, respectively.

(9)R. Latter and H. Kahn, Gamma-Ray Absorption
Coefficients, R-170, p. 14 (Sept. 19, 1949),

NUCLEAR-POWERED AIRPLANE

The required number of relaxation
lengths 1is

(3.07X108xo.24
ln =
0.025%x5.5% 10%

= 8.6 relaxation lengths .

The lead thickness required, with some
of the lead replaced by plastic, 1s

In (5.36 x 10%)

 

tppy = 8.6 X 0.147-37.3 % (0.147/8.4)
= 0.62 cm of lead.
SPECIFICATION OF REACTOR SHIELD
THICKNESS
In the following sections, a

reactor shield is presented that in
conjunction with the radiator-con-
trolled crew shield will attenuate
the reactor sources to the levels
specified in Table 11,

Reactor Neutromns into Crew Shield
Rear (Id). For comparison with Bulk
Shielding Facility (BSF) data, it is
necessary to make some comparison of
the relative leakages of the BSF
reactor and the circulating-fuel
reactor. When the mean free path is
much less than the average chord
length of the core, the leakage should
be inversely proportional to the
latter., A fair approximation for the
average chord length is

4v

s

= 11.7 in. for the BSF reactor

28 1in.
reactor,
where vis the volume and s the surface
of the core. The comparison factors
must also include the ratio of the
circulating-fuel reactor power,
3.25 x 10% watts, to the BSF reactor
power, which is normalized to 1 watt,
In addition, the attenuations of the
beryllium oxide reflector (1/6.8)
and of the iron shells (1/1.53) are
included. The factor by which the
BSF data must be multiplied in order
to obtain the expected value in the
circulating-fuel reactor configuration
is thus

F=3,25x 10% x (11.7/28) %X (1/6.8)
x (1/1.53) = 1.3 x 107

for the circulating-fuel

39
 

DESIGN STUDY

With the inverse square attenuation
for reactor-crew separation included,
the governing expression 1is

0.01 rep/hr = D{ig®™) x 1.3 x 107
X (ry/120)2

or
D{reer) = 1.1 x 107°/r} rep/hr

= 1,1 % 10"/r§ D units,
where o 1s the outside radius of the
reactor shield in ft (ro = 5).

The equation is satisfied for
130 cm of water¢!?) of which 40 cm,
or 1ts equivalent in plastic, are
located effectively at the crew
compartment. The lead at the crew
compartment can be counted on for
further attenuation, since the reactor
neutrons are of high enough energy
so that inelastic scattering will be
important. On the other hand, the
lead is not backed up by hydrogenous
material and therefore cannot be
allowed 1ts usual 9-cm relaxation

length. A conservative value of 18 cm
1s chosen, which gives an attenuation
of

exp (22.4/18) = exp (1.2) ,
or 12 ¢cm of water. The resultant
reactor shield thickness becomes

t (reactor front) = 78 cm

Reactor Neutrons into Crew Shield

Sides (Ie). The ratio of flux incident
on crew shield sides to that on the
rear 18, according to simple first-
scattering calculations,

 

 

SO
87Ad ) d
S, 2\
4d?

For this case, the ratio 1s

(120/2) x 130 x 3,28 = 0.141 ,
where 3.28 is the number of feet per
meter, 130 1s the mean free path of
neutrons in air, and 120 is the
separation distance in feet,

 

(lo)fi. P, Blizard, Introduction to Shield
Design -~ II, ORNL CF-51-10-70 (March 7, 1952).

40

The allowed dose into one side 1is
one-fourth the total side dose or
one-fourth of the dose in the rear,
since the allowances for sides and
rear are the same (Table 11). The
dose to be measured 1in the BSF to
correspond to the proper thickness of
water for attenuating the side neutrons
will thus be that which attenuates by
a factor of 4 X 0.14 more than the

thickness chosen for the rear. Thus
, 1.1 x 10°% 1
(on sid = -
DL S% ¢) = rep/hr

4 X 0,141 r2
0

(1.95 X 10'5)/r:rep/hr

The thickness corresponding to this
condition 1s 122 cm of water, for ro
equal to 5 feet. Of this thickness,
46.5 cm of water equivalent 1s sup-
plied at the crew compartment, and
thus 76 cm 1s required at the reactor.
The lead at the crew compartment is
ignored, since it 1s not very thick
and 1s not backed up by hydrogenous
material. To allow for some structure
scattering, a total reactor shield
thickness of 78 ¢cm is specified.
Since this value agrees with that for
the reactor shield thickness calculated
in the previous paragraph, a uniform
shield thickness is chosen,

Reactor Neutrons into Front of
Crew Shield (If). The scattered
neutron flux i1nto the front face on
the basis of 1sotropic scattering
is (7n/2) - 1, or 0.57, times that
incident on the side. The allowed
flux is one-fifth that for one side
(one-twentieth of the dose from four
sides) .

The ratio of the attenuation
required of the front shield to that
of the side shield 1is

5 X 0,57 = 2.85 .
On the other hand, the front shield
is thinner than the side shield by
about 2 cm, which corresponds to
a factor of about 1.22. The over-all
dose entering the front is greater
than the tolerable dose by

1.22 x 2,85 = 3.5 ,

 
 

There are two factors that tend to
minimize this excess: (1) the air
scattering is not isotropic but rather
strongly forward for the high-energy
neutrons; and (2) the neutron beam
1s attenuated in air. This attenuation
1s certainly important for the neutrons
entering the front with the present,
large, reactor-to-crew separation
distance. These two effects will more
than compensate for the factor of
3.5. Note that the forwardness of
scattering is not characteristic of
the delayed neutrons, so that this
saving could not be used for delayed
neutrons.

Reactor Gamma Rays into Crew Shield
Rear (IId). For gamma rays, the
relative escape probabilities in the
circulating-fuel reactor and the
BSF reactor are

Aerr \ S/BSF 12 em  11.7 in.
X x

 

= 0.33

Nesp [4Y\ 15 cm 28 in.
$ JcFR

A more exact calculation, made by
using the method of Murray,{!!) gives
a ratio of 0.46, which will be used.

The beryllium oxide reflector gives
an attenuation of 3.65; so the effective
factor of comparison is:
f,=1(0.46/3.65) x3.25% 10%=4.1 x 107,

For 78 cm of water, the BSF data
show 0.25 r/hr. In addition to the
water, there are 1 in. of iron, 24.4
cmof lead, and 33.4 cm of polyethelene
plastic. Thus the total attenuation is
exp (2.54/4) + (22.4/2.2)

+t (33.4/24.8) = exp (12.2)

The gamma dose contribution in the
crew compartment is accordingly

0.25 r/hr x 4.1 x 107

X exp (~12.2) (-5/120)?%2
= 8.8 X 10°%? r/hr .

(II)F. H, Murray, Fast Effects, Self-Absorption,
Fluctuation of Ion Chanber Readings, and the
Statistical Distribution of Chord Lengths in
Finite Bodies, CP-2922, p, 15 (Apr. 6, 1945).

NUCLEAR-POWERED AIRPLANE

The allowed quantity, 0.100 r/hr is
thus almost exactly correct.

Gamma Rays from Reactor to Crew
Shield Sides (IIe). As in the previous
section, 78 cm of water corresponds to
0.25 X 4,1 %X 107 exp (-2.54/4)

= 5.4 x 10% r/hr

for the gamma dose measured at the
shield exterior. The equivalent point
source of 3-Mev gamma rays is thus

5.4 ¥ 10% r/hr X 5.5
X 10% Mev/cm?+sec/(r/hr) /(3 Mev/photon)
X 4m (5 % 30.5)% em?® = 2.9

X 10'7 3-Mev photons/sec.

At a separation distance of 50 ft,
this would correspond to

2.9 x 10'7 x (50/120)
= 1.2 X 10!'7 3-Mev photons/sec .

For a dose of 0.05 r/hr, the curve
in ANP-53(%) gspecifies a thickness
of 6.25 cm of lead. This is the same
as the basic amount calculated for the
radiator gamma rays; so the sides are
adequate.

Gamma Rays from Reactor to Crew
Shield Front (IIf). This calculation
1s carried out in a manner similar to
that for IIc¢. The flux incident on
the front face is thus:

29_m_<w .
87Ad 2 "

3.5 X 107 photons/cm?+sec
5.1 X 10 cm,
120 X 30.5 cm.

wou

A
d

The required number of relaxation
lengths 1is

3.5 x 107 x 0.24
In
0.025 x 5,5 x 105

 

)= In (6.14% 102)

= 6.4'

Previously the requirement was for
8.6 relaxation lengths to take care
of the radiator gamma rays. The
present design is thus safe. It is

4]
 

DESIGN STUDY

inadvisable to reduce the lead on the
front of the crew compartment below
the 0.62 cm previously specified,
since this will ensure that no large
number of soft gamma rays will enter
this area through the plastic.

The basic thicknesses of lead and
plastic required for the crew shield
for a reactor with 78 cm of water on
all sides are listed in the following:

LEAD PLASTIC

{cm) (cm)
Rear 22.4 33.4
Sides 2.75 39
Front 0.62 37.3

SPECIAL SHIELDING CONSIDERATIONS

Crew Shield Sides Near the Rear,.
In this region it is possible that
radiation entering the rear will be
scattered in the plastic side-walls
and penetrate the sides. To take
care of this eventuality, the lead
must be thickened in this region and

tapered off to the side-wall thickness
specified in Table 1l.

To estimate this effect, it 1s
assumed that the optimum angle of
scattering is 7/4, for which a 3-Mev
gamma ray would be degraded to about
1 Mev. The effective solid angle to
be considered is about 1 steradian.
The electron density of the plastic is:

(0.602 x 10%* x 0,95 % 8)/14

= 3.3 X 1023 electrons/cm?
The cross section per steradian about
an angle of m/4 is, from the Rand
report,(??
1.361 x 3.3 x 1023 x 10-2°

= 4.4 x 103 cm*?

The total cross section is

0.1136 % 10-24 x 3.3 x 1023
= 3,75 X 10°2? cm*! ;

so the fraction scattered near the
proper angle is 0.44/3.75 = 0.117.
Equating the attenuation along the
paths through the rear lead disk to
the plastic plus the slant lead paths,
after scattering,

 

 

 

 

 

 

— — — — X — — i PahII
Plastic /4 / |
/i
_I (I
===
/
f*"—— — e e 4 __IEfihI
Crew Lead | Plastic Reactor
_—
-22.4/2.2 . . ]
e / ]Pnth I, = 0.117 e =/ 24.8 e 1.4y/1,25]

42

Path II *

 
 

 

where 1.25 is the relaxation length in
the lead for 1-Mev gammas, and 1.4 is
the secant of 7/4.

22.4 x

— = 2,14 +
2.2 24.8

10.18 = 2.14 + 0.04x + 1,12y
y=17.1-0.036 x .

At the inside corner of the crew
shield, x 1s 22.4 cm, so
y=17T.1 -0.036 x 22.4
= 6,3 cm of lead.

The lead thickness never will be
below 2.75 cm on the sides, for other
reasons. The value of x at which this
value is here specified for y is:

2,75 =7.1~0.036 «x

 

+ 1.12y

7.1 - 2,175
X t—
0.036
= 120 cm.
Thus the side lead is decreased

linearly in thickness from 6.3 cm
at the rear corner to 2.75 e¢m at a
location 120 cm forward of this.
From that point forward, the 2.75-cm
thickness is specified constantly.

ALL DIMENSIONS ARE IN INCHES.

 

kel L

 

e — 1229 0D

STAGNANT WATER —7

SCIRCUL ATED WATER{

REACTOR

 

2912-XC
R 9
Y-
™~
/ “\REFLECTOR
0250 INsuLATION— \L_
) WATER PRESSURE

Y
SHELL 0.50 THICK |
54=10,000 psi

SUPPORTS

 

 

 

 

 

 

 

 

LONGITUDINAL SECTION

Fig. 13.

NUCLEAR-POWERED AIRPLANE

Slanting Front Wall, The shield
for the slanting front wall must have
about 38 cm of plastic, as can be seen
from inspection of Table 11, and about
1.5 cm of lead.

PHYSICAL DESCRIPTION OF SHIELD

A five-man crew shield and a reactor
shield assembly were designed by using
the thicknesses of lead and plastic
prescribed in the preceding paragraphs.
Figures 13 and 14 show the reactor
and the crew shield assemblies, re-
spectively., It is important to note
that the thickness of water in the
reactor shield is greater than the
78 c¢cm mentioned in the preceding
paragraphs. This 1s due to the fact
that the thickness of 78 cm was
calculated for waterof normal density,
that is 62.4 1b/ft®. The water in the

reactor shield is actually less dense
than this (due to its temperature),
and the thickness of the water was
increased to compensate for the lower
density.

DOWG. 17!-15

 

VERTICAL SECTION

Reactor Shield Assembly

43
 

vy

4725 ———=

 

 

 

 

 

 

 

 

—
| | E
O c©o o |
’ T \Z e 8
b P

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

—e=le=- 0.244 LEAD

 

 

 

150

 

059 LEAD

14.96 PLASTIC

 

 

 

 

186.904
FLOOR PLAN

   
 

 

 

 

 

 

 

 

 

 

Fig. 14.

 

ALL DIMENSIONS IN INCHES

-
DWG. 17674

 

 

 

 

 

Five-Man-Crew Shield Assembly.

SECTION A-A

AANLS NIISId

 

 
 

The weights of the resultant crew
shield and reactor shield assembly
are as fol lows:

Five-man crew shield, 1b

Lead 30,800
Plastic 25,900
Total 56,700
Reactor shield assembly, lb
Outside H,0 layer at 145°F 14, 500

NUCLEAR-POWERED AIRPLANE

Inside H,0 layer at 350°F 13,700
Outside Boral shell 2,300
Inner H20 pressure shell 4,500
Insulation and its canning 1,600
Reactor assembly 10,000
Reactor and shield 46, 600
Structure within shield 1,800
Total shielded package 48, 400

Total shield weight, 1b 105, 100

 

STATIC CHARACTERISTICS OF THE REACTOR

The reactor has the appearance of
a gourd, with the circulating-fuel-
coolant passing through the stem in
two concentric tubes. A third con-
centric tube, enclosing the fuel,
contains an inert molten salt that
cools the beryllium oxide reflector
that surrounds the cylindrical core.
The reflector is separated from the
core by a 1/2-in.-thick Inconel
pressure shell. A second, spherical,
1/2-in.-thick, Inconel pressure shell
encloses the entire assembly and
merges smoothly into the stem. From
the physics point of view, this
reactor is relatively homogeneous in
the core because of the small self-
shielding effect of uranium and the
thin beryllium oxide and Inconel
structural members. The 1/2-in.-thick
pressure shell between core and
reflector is a strong absorbing layer
for thermal neutrons, and a thick
fuel-coolant layer at the closed end
of the core introduces an important
neutron source discontinuity, Both
of these effects can be evaluated,
but not easily; the latter effect 1is
especially troublesome. 1In addition,

TABLE 12.

PHYSICAL DIMENSIONS

the unreflected stem of the reactor

permits high neutron leakage from the

core and results in a region of low
importance for uranium in the vicinity
of the core.

The reactor was divided into four
sections, each of which has, essen-
tially, a different reflector. The
sections are (Fig. 1):

A. the cylindrical sides,

B. the beryllium oxide reflected end,

C. the section of the fuel-reflected
end adjoining the cylindrical
sides (where there is considerable
structure in the reflector),

D. the section of the fuel-reflected
end farthest from the cylindrical
sides (where there is no structure
in the reflector).

Each section was then separately

treated as if i1t were part of a

spherical reactor with geometry

similar to that of the pertinent
section, Table 12 indicates the
data used for the four sections of the
reactor. Table 13 summarizes some of
the results of the reflected-reactor

calculations. Figures 15 through 26

present the power distribution, fission

AND CONSTITUENTS OF

THE FOUR SECTIONS OF THE REACTOR

 

REACTOR Ar* B* B!+

CORE CONSTITUENTS (vol %)}

REFLECTOR CONSTITUENTS (vol %)

 

 

SECTION Fuel Structure | Imert Salt Moderator Fuel Structure Inert Salt Moderator
A 2,138 | 24 33 35.00 2,176 5,00 57.24 10 3 85
B 2,039 | 26 37 34.07 1.97 3.20 60,76 3 95
C 2.039 | 26 13 34,07 1.97 3.20 60,76 84 16
D 2,039 26 37 34.07 1,97 3.20 60,76 100

 

 

 

 

 

 

 

 

 

 

 

 

*Core radius is BAr, cm; extrapolated radius to reflector boundary is B'Ar, cm.

45
 

 

DESIGN STUDY

 

 

 

 

 

 

N are
1.7 1 I 1 ]
—
e \\ COMPOSITION IN VOLUME FRACTIONS
: N CORE REFLECTOR
FUEL-COOLANT 35.0%
1.5 STRUCTURE (INCONEL)  2.76% STRUCTURE | 10%
MODERATOR {B0) 57.24% MODERATOR | 85%
L INERT SALT 5.00% INERT SALT | 5%

 

 

¥ NORMALIZED TO ONE SOURCE NEUTRON PER
CUBIC CENTIMETER OF CORE

 

 

\ Ar=2.138cm

09 N
N

0.8

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

CORE — REFLE|GTOR -;4[

o7 |
o 2 4 6 8 10 12 14 16 t8 20 22 24 26 28 30 32 34

uNITS OF A~

 

NUMBER OF NEUTRONS PRODUGED PER CUBIC CENTIMETER OF CORE (¥}
ro

Fig. 135. Spatial Power Distribution for Reactor Section A — The Cylindrical
Sides.

D!G. {7877

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

X

 

1.8

5 47 LN

x \ COMPOSITION IN VOLUME FRACTIONS

& N CORE REFLECTOR
§ 8 N FUEL-COOLANT 34.07% ]
o \ STRUGTURE (INCONEL) 1.97%  STRUCTURE 5%
E 1.5 MODERATOR (BeO) 60.76%  MODERATOR 95 %
o \\ INERT SALT 3.20%

=

E 1.4

w \ * NORMALIZED TO ONE SOURGE NEUTRON PER
Q \ CUBIC CENTIMETER OF GORE

g 1.3 \

B, N

§ Ar=2.039 cm

o {4

O

@

a

¥ 4.0

5 \

[

s}

w

=

uw

O

e

w

o

-

=

2

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

0.8 \\
0.7 N
— CORE REFLECTOR—TO 374
0.6 | | | I
0 2 4 6 8 10 12 14 16 18 20 22 24 26 28 30 32 34
UNITS OF Ar

Fig. 16. Spatial Power Distribution for Reactor Section B - The Beryllium
Oxide Reflected End.

46

 

 
 

NUCLEAR-POWERED AIRPLANE

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

wan—
DWG 17678
¥ 20 I | | | |
W COMPOSITION IN VOLUME FRACTIONS
g 18 CORE REFLECTOR
W — FUEL-COOLANT 3407%  FUEL 84%
. 16 — STRUCTURE (INCONEL) 1.97%  STRUCTURE 16%
L T MODERATOR (Be0) 60.76%
E \\ INERT SALT 30.20%
1.4
@ N *NORMALIZED TO ONE SOURCE NEUTRON PER
© \ CUBIC CENTIMETER OF GORE
a .2 h, | !
3 \ A‘ |
o \\\- _ r=2039cm
¥ oo N
0
i
S
2 o8
Q
&
0 06 N, !
=
5 N\ |
-
g os N\ |
z \ \ |
S
|
= 0.2 < |
@ CORE ~.REFLECT0R‘-;::
0 2 4 6 8 10 12 14 16 18 20 22 24 26 28 30 32 34
UNITS OF Ar

Fig. 17. Spatial Power Pistribution for Reactor Section C - The Fuel-
Reflected End Adjoining the Cylindrical Sides.

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

— DWG. (TETS

* 1.8

@

8 16 COMPOSITION IN VOLUME FRAGTIONS

u ’ CORE REFLECTOR

- FUEL-COOLANT 34.07% FUEL 100%

1.4 STRUCTURE (INGONEL)  1.97%

y MODERATOR (8e0) 60.76%

R INERT SALT 3.20%

—

Q M~

g 1.0 \ *

© NORMALIZED TO ONE SOURCE NEUTRON PER

x \ GUBIC GENTIMETER OF CORE

& os >

o TN

S \\\ l

2 0.6 M \ Ar=2039cm i

g N |

© 04 \‘ {
I

o

£ |

@ 0.2 \\ '

w \ |

O \\I

x O Y

5 _

g CORE -— REITLECTOT

2

=z

0 2 4 6 8 10 12 14 16 8 20 22 24 26 28 30 32 34 36 38

UNITS OF Ar

Fig. 18. Spatial Power pistribution for Reactor Section D - The Fuel-
Reflected End Farthest from the Cylindrical Sides.

47
 

 

 

 

DESIGN STUDY

S
DWG. 17680

 

043

 

012

COMPOSITION IN VOLUME FRACTIONS
CORE REFLECTOR
FUEL - COOLANT 35.0%
STRUCTURE {INCONEL) 2.76% STRUCTURE 0%
MODERATOR (BeO) 57.24% MODERATOR 85%
INERT SALT 5.00% INERT SALT 5%

010

0.09

0.08

0.07

0.06

0.05

0.04

0.03

0.02

FISSION PER UNIT LETHARGY PER % INITIAL NEUTRONS

0.0t

LETHARGY ()

Fig. 19. Fission Spectrum vs. Lethargy for Reactor Section A - The Cy-
lindrical Sides.

s

013

 

012 COMPOSITION N VOLUME FRACTIONS

0.1 CORE REFLECTOR
FUEL-COOLANT 34.07%

010 STRUGTURE (INGONEL) 1.97%  STRUCTURE 5%
MODERATOR (Be0) 60.76%  MODERATOR 95%

0.09 INERT SALT 3.20%

0.08
0.07
0.06
0.05
0.04
0.03

0.02

FISSION PER UNIT LETHARGY PER Y. INITIAL NEUTRONS

0.01

        

19 118 17 16 15 14 13 12 4y 10 9 B 7 6 5 4 3 2 { o
LETHARGY («)

Fig. 20. Fission Spectrum vs. Lethargy for Reactor Section B - The Beryl-
lium Oxide Reflected End.

48

 

   
 

 

NUCLEAR -POWERED AIRPLANE

 

JRn
DWG. 1T682

0.13

0.2
o
3
€ oM COMPOSITION IN VOLUME FRACTIONS
>

R

& o0 CORE REFLECTOR
3 FUEL - COOLANT 34.07% FUEL 84%,
£ 0.09 STRUCTURE (INCONEL)  1.97% STRUCTURE 16%
= MODERATOR (Be0) 60.76 %
L 008 INERT SALT 3.20%
&
W 0.07
>
& 0.06
I
I
& oos
W o
=
Z 004
o
& o003
z
o
@ 0.02
N
W

0.01

0
0 19 1B 17 16 (5 14 13 12 4 40 9 B8 7 6 B 4 3 2 1 0

LETHARGY (¢}

Fig. 21. Fission Spectrum vs. Lethargy for Reactor Section C - The Fuel-
Reflected End Adjoining the Cylindrical Sides.

SR
OWG. 17683
013

ote

011 COMPOSITION IN VOLUME FRACTIONS

CORE REFLECTOR

0.10
FUEL - COOLANT 34.07% FUEL 100%

0.09 STRUCTURE (INCONEL} 1.97%

MCDERATOR (Be0) 60.76%
0.08 INERT SALT 3.20%
0.07
0.06
0.05
0.04

0.03

.02

FISSION PER UNIT LETHARGY PER ¥, INITIAL NEUTRONS

0.01

 

0 f¢ 18 17 16 15 14 13 12 W 10 9 8 T 6 5 4 3 2 1 0o
LETHARGY (v}

Fig. 22. Fission Spectrum vs. Lethargy for Reactor Section D - The Fuel-
Reflected End Farthest from the Cylindrical Sides.

49
 

DESIGN STUDY

LEAKAGE PER UNIT LETHARGY PER v, INITIAL NEUTRONS

0012

0.011

0.010

0.009

0.008

0.007

0.006

0.005

0.004

0.003

0.002

0.001

Fig.

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

S
DWG. 17684
COMPOSITION IN VOLUME FRACTIONS
- CORE REFLECTOR
FUEL- COOLANT 35.0 %
- STRUCTURE (INCONEL) 2.76% STRUCTURE 10%
MODERATOR (BeO) 57.24%, MODERATOR 85%
. INERT SALT 5.00% INERT SALT 5%
Y fl-
‘LL.H__
jy
e
0.1305
THERMAL
B ESCAPE AT
v=18.6 [~ _‘
20 18 16 14 12 10 8 6 4 2 0
LETHARGY (¢)
23. Leakage Spectrum vs. Lethargy for Reactor Section A - The Cy-

lindrical Sides.

50

 
 

NUCLEAR-POWERED AIRPLANE

 

 

 

 

 

 

 

 

SllRaF
DWG. 17685
COMPOSITION IN VOLUME FRACTIONS
0.008 |—
CORE REFLECTOR
0.007 |— __| FUEL-cOOLANT 34.07%
STRUCTURE (INCONEL) 1.97 % STRUCTURE 5 %
MODERATOR (BeO) 6076% - MODERATOR 95%
0.006 — — INERT SALT 3.20 %
0.005 }—

 

 

 

 

0.004 |—

 

0.04350 o

THERMAL H_ha._
0.003 — ESCAPE AT

¢ =18.6 ]

. - :

0.001 ] l
0

20 18 16 14 12 10 8 6 4 2 0
LETHARGY (¢)

 

 

 

 

 

 

LEAKAGE PER UNIT LETHARGY PER v, INITIAL NEUTRONS

 

 

 

 

 

 

 

 

 

 

 

 

 

 

Fig. 24. Leakage Spectrum vs. Lethargy for Reactor Section B - The Beryl-
lium Oxide Reflected End.

51
 

 

DESIGN STUDY

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

S

0.045 . DWG. 17686
o L L
2 COMPOSITION IN VOLUME FRAGTIONS
o e
E 0.040— CORE REFLECTOR
o FUEL- COOLANT 34.07%  FUEL 84 %
f STRUCTURE (INCONEL) 1.97% STRUCTURE 16 % |
< 0.035— MODERATOR (BeO) 60.76 %
£ INERT SALT 3.20%
© 0.030—
vl
L
T 0.025(+—
» 0.00195
& THERMAL
F 0020— | Eescape AT
3 U= 18.6\
-
5 0.015}— f
o
vl
& I
T 0.0i0F—
W
< T
¥
< 0.005[— /_r,_f
- _fl

0 \t'
20 18 16 19 12 10 8 6 3 2 0
LETHARGY (u)

Fig. 25. Leakage Spectrum vs. Lethargy for Reactor Section C - The Fuel-
Reflected End Adjoining the Cylindrical Sides.

52

 
 

spectra, and the neutron-leakage
spectra for the four sections of the
reactor,.

The effect of uranium self-shielding
in the fuel-coolant tubes is indicated
in Fig. 27. The change in effective
multiplication constant with tube
size and the corresponding uranium

NUCLEAR-POWERED AIRPLANE

weight in the reactor core are given
as a function of fuel-coolant tube
diameter, These were computed by the
bare-reactor method for survey pur-
poses. Uranium weight in the core
vs. k, is given in Fig. 28.
Reactivity coefficients for the
reactor, which have been evaluated
approximately, are given in Table 14,

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

L. I

DWG. {7687
0040 T I T ] [

2 COMPOSITION IN VOLUME FRAGTIONS| l

& CORE REFLEGTOR

E 0.035 —

2 FUEL-COOLANT 34.07% FUEL 100%

< STRUCTURE (INCONEL ) 1.97 %

3 0.030}— MODERATOR (BeO) 60.76 %

'.-:: INERT SALT 3.20%

& 0.025 19—

o 0.00293

L

a THERMAL rHT

% 0020 lescape AT

g v= 18.6\

Y 0.015

- .IJ

': PR

Z 0.010 —

m \

w

s o

w 0.005— /,r

<

% ig

w 0

-

20 i8 16

14 12 {0 8 6 4 2 0

LETHARGY (u)

Fig. 26.

Leakage Spectrum vs. Lethargy for Reactor Section D -~ The Fuel-

Reflected End Farthest from the Cylindrical Sides.

TABLE 13. SOME RESULTS OF THE REFLECTED-REACTOR CALCULATIONS

 

 

o PERCENTAGE OF REFLECTOR A A )
REACTOR SECTION keff THERMAL FISSIONS THICKNESS (cm) (Bk/k)/AT (°F)
A 0.914 62.0 17.8 1.6 x 1073
B 0.972 60.9 204 ~0.13 x 10-3
c 0.949 50.9 12 ~0,03 x 10-5
D 1,103 46.6 20 ~0.45 x 10-5

 

 

 

 

 

*For 22.5 1b of U235 in reactor core,
core is 25 pounds.

kcff = 0,963 by area weighting.

Critical uranium mass in

53
 

 

 

DESIGN STUDY

The net(!) temperature reactivity

coefficient due to thermal expansion
and thermal base change is -0.66 x 104

pe
in

r °F; the critical uranium weight
the core is 25 1lb; the uranium

requirement for 27.7 ft? of fuel out-

side the core 1is 87 1b,

we

The uranium
ight is that of U235 in a 93.4%

enriched uranium fuel.

in
ha

The power density in watts/cm?
the moderator and the fuel-coolant
s been evaluated for this reactor

 

(1)

Expanaion coefficients asasumed: BeO,

 

 

 

 

 

 

 

 

 

 

 

 

 

and is shown in Fig. 29. The values
given are a first approximation, since
the fuel reservoir at the blind end
of the reactor 1s omitted and the
effect of the 1/2-in., absorbing layer
(Inconel) at the core boundary is
included by addition to the reflector
material, The integrated neutron flux
normalized to one fission per unit
volume (cm?) of core per second is 512
neutrons/cm?+sec. The average flux
at full power is 8 x 10'%, with a peak
value of 2 X 10'%® pneutrons/cm?*sec at
at the center of the reactor.

 

 

 

 

 

 

 

 

 

 

 

 

-6 -4 o
14,8 16 ; fuel- 1 , 1.47 10 F. e e .
" eireoolfant " per The possibility of decreasing
—— critical mass significantly by using
100 DWG. 17688 beryllium instead of beryllium oxide
E because of its larger density of
E 0.99 moderating nuclei has been evaluated.
8 26.31b
O 0.98 |p———
5 0.97 z DWG. 17689
£ O 2 1 ,
< 0
S oe6 \ _291b z £2: 00020308 ]
T N z . ASSUMED REFLECTOR SAVING,15.24cm ___—
5 095 g —
= \ & /
w 5 09
> 0.94 2
= w
W 0.93 \\\\\ g
- 3 |b/"] g os
ul L
0.92 ui
0 ' > 20 25 30 35
FUEL-COOLANT TUBE DIAMETER (in.} URANIUM MASS IN THE CORE (Ib)
Fig. 27. Effect of Uranium Self- Fig. 28. Uranium in Core vs. Ef-
shielding in the Fuel-Coolant Tubes. fective Multiplication Constant.
TABLE 14. REACTIVITY COEFFICIENTS FOR THE REACTOR
REACTIVITY RANGE OF
COEFFICIENT VALUE
Moderator (BeO) density [(Ak/k)/(Ap/p)] 0.321 99 to 100% of
quoted density
Coolant density [(Ak/k)/(Do/p)] 0.011 90 to 100% of
quoted density
Structure (Inconel) density -0.147 100 to 125% of
[(Ak/E)/(Dp/p)] quoted vol. %
Thermal base (reactor temperature)
[(Ak/k)/ AT(°F)] -1.20x%10"8 1283 to 1672°F
Uranium weight [(Ak/k)/(AM/M)] 0.348

 

 

 

54

 
 

NUCLEAR-POWERED AIRPLANE

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

o,
2.2
CURVE A: NORMALIZED RADIAL AND AXIAL {REFLECTED END)
2.0 POWER DISTRIBUTIONS. TO OBTAIN POWER DENSITY — |
N FUEL-COOLANT, MULTIPLY BY 2030 watts/cm3,
A CURVE B: NORMALIZED, AXIAL (BARE END), POWER DISTRIBUTION.
1.8 e — TO OBTAIN POWER DENSITY IN FUEL-COOLANT
B \ MULTIPLY BY 2310 watts/cm3
1.6 ~y CURVE C: RADIAL POWER DISTRIBUTION IN MODERATOR AND
REFLECTOR. TO OBTAIN POWER DENSITY, MULTIPLY
~ BY 700 watts /¢cm3 FUEL TRANSPORT ADDS A
e 1.4 ‘~\\\ FACTOR OF 0.76.
Q
S NG CORE VOLUME, 22.0 ff°
w2 N 34.1 % FUEL-GOOLANT
z \ 80.7 % BeO
= 1.0 t.97 % INGONEL
N N
-
3 \\\\‘
m O-B \\
o
=z —
0.6 z
g \\\\ \\
0.4 — N
\-\ \ B
0.2 —
\ \
. \;,_
o 10 20 30 40 50 60 70
DISTANGE FROM CORE CENTER (cm)
Fig. 29. Power Density in Fuel-Coolant and Moderator of Circulating-Fuel
Reactor.
A net reduction of only 6.5% in increase in k of 10% will be re-

critical mass would result from this
change. At the present time a flat
increase 1n keff of 5% by means of
control apparatus is a maximum value
to over-ride maximum transient xenon
and other fission products and loss
in delayed neutrons for the operating
temperature range. The presently
available data indicate that a net

quired to raise the reactor from room
to operating temperature, The uranium
requirement per aircraft in flight
will thus be a maximum of approxi-
mately 75 1b in core, plumbing, and
heat exchanger. This value will not
be changed significantlyif some other,
nonpoisoning, fuel-coolant solution
is used.

 

REACTOR CONTROL

An ANP power plant electronic
simulator was set up with design-point
values for the various reactor param-
eters. By means of this simulator,
the following time variables were
determined as the system response to
various stepped changes in reactor
excess reactivity: power level (p/p,),

rate of change of power level (p/p,),
mean fuel temperature (£,), and rate
of change of mean fuel temperature
(Qf). Figures 30 through 33 show
these quantities plotted as a function
of time for steps in excess reactivity
(Ak/k) of 0.002, 0.004, and 0.006.
The fuel temperature coefficient of

55
 

DESIGN STUDY

P/A,

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

L
OWG, 17691
8
LN
\Akaooos
1N
L
/ /\*——Akflr:oooo.l
|
3 \ ]
// A K7k =0.002
1 S —
0
O 10 20 30 40 50 80 70 BO SO 100
TIME {msec)
Fig. 30. Power Level After Various

Step Changes in Excess Reactivity.

P/P,

700
600
500
400
300
200

100

-100
-200
-300

-400
-500

-600

Fig

DWG. 17692

Ak/k 20006

 

k/k =0004

k/k=0002

0 10 20 30 40 50 &0 70 80 90 100 YO

- 31.

TIME {msec)

Bate of Change of Power

Level After Various Step Changes in
Excess Reactivity,

56

_—_—
DWG. 17693
1390

1365 k=
1340
1315
1290
1265

k/k=0004

k=

TEMPERATURE (°F)

 

0 10 20 30 40 50 sC 70 B8O 90 100 110 120
TIME (msec)

Fig. 32. Mean Fuel Temperature
After Various Step Changes in Reactiv-
ity.

o
DWG. 17694

6500
6000
5500
5000
4500
4000
3500
@ 3000
2500
2000

1500

Ak/k 20004

1000

500
k/k=0002

 

-500
O 10 2C 30 40 50 60 70 80 90 100
TIME {msec)

Fig. 33. Rate of Change of Fuel
Temperature After Various Step Changes
in Reactivity.

reactivity, (Ak/k)/AT (°F), used in
obtaining Figs. 30 through 33 is
-8 X 10°%, This value was determined
by taking the temperature coefficient
of reactivity due to the volumetric
expansion of the fuel and dividing by
2 to allow for other forseeable and
un forseeable fast effects. The
temperature coefficient of reactivity
 

due to the moderator, structure, etc.
was neglected as being too slow to
affect the fast transients.

The delayed-neutron steady-state
contribution, usually designated as

6
z a;B,, for this circulating-fuel
i

for
that is, non-circulating fuel,

reactor was taken as 0.00172;
a, =1
i ?

6
E a.lfi:. = (,0073. The mean neutron
i=1

life-time was taken as 2.65X%10-% sec.

CONTROL FEATURES DETERMINED
BY SIMULATOR STUDY

A few control features of signifi-
cant merit are characteristic of the
circulating-fuel type of reactor. When
fuel expansion with temperature rise
provides a negative reactivity temper-
ature coefficient, the power demand
signal from the turbojet engines 1is
transmitted through the circulating-
fuel coolant and provides an over-all
power plant in which the reactor is a
slave to the external loading system,
that is, the engines. Since the major
portion of the reactor power 1is
generated in the fuel, the fuel has the
fastest temperature response time of
any element of the reactor, Accord-
ingly, the reactor power follows the
load demand more closely for the
circulating fuel than for acirculating
moderator, or for that matter, it
follows more closely the load demand
than is possible for any other type of
coolant,

As a consequence of the above-
described feature, external coupling
in the form of servo control loops
with sensors, actuators, control rods,
etc. are not necessary 1in a reactor
such as this. In fact, studies made
of this reactor by using an electronic
power plant simulator indicate that
only the self-stabilizing features
inherent in the negative-reactivity
temperature coefficient make possible
control of this power plant in which

NUCLEAR-POWERED AIRPLANE

the power density is so high and such
a large portion of the normally static
fuel delayed-neutron contributions
are not effective in the control.
This comes about because of the
severe servo system response times
needed for control. These simulator
studies indicate that servo frequency
responses of as high as 500 cps would
be required to control a similar but
nonself-stabilized reactor in any
manner comparable to the control
provided by the negative-reactivity
temperature coefficient. Furthermore,
regulating rods would require acceler-
ations comparable to impacts to provide
comparable control. Such servo systems
lie considerably beyond the presently
known control art,

Surge pressures derived from fluid
temperature transients constitute the
limiting features of ANP controlla-
bility. The first and second time
derivatives of the mean fuel tempera-
ture were determined by using the
simulator, and from these data the
pressure surges were calculated and
found to be tolerable.

Relatively slow control of the
reactor can be provided by some
shimming means, either rods or enrich-
ment, The shimming provided by rods
takes care of the system poisoning,
and either rods or enrichment would
provide for fuel depletion. Essen-
tially, rod motion merely changes the
mean fuel temperature and does not
control the load power.

PRESSURE IN FUEL TUBES

The increase in temperature of the
fuel in the core, owing to a step
change in reactivity, causes an increase
in fuel volume. Inasmuch as the fuel
is incompressible, the incremental
fuel volume must be transferred from
the reactor core to the surge tank
as rapidly as it is generated, that
is, in approximately 40 milliseconds.
The rapid acceleration of the fuel
required to transfer the generated
volume results in appreciable inertial

57
 

forces in addition to the frictional
forces involved. These inertial and
frictional forces have been evaluated,
both in the reactor core and in the
piping to the surge tank (the surge
tank is 2 ft from the reactor inlet).
Figure 34 1s a plot of the variation
with time of the incremental pressure
at the reactor core outlet, at which
peint the incremental pressure 1is at
a maximum, for a step change 1in
reactivity of 0.0025. Figure 35 is a
plot of the maximum i1incremental
pressure at the reactor core outlet
for various step changes of reactivity.
For a step change in reactivity of
0.006, the maximum incremental pressure
at the reactor core outlet is 100
psi. For 0.65-in. -0D tubes with
0.025 -in, walls, the incremental

BETess

20

@

PRESSURE PULSE (Ib/in2)
H

Q

-4

 

0 0.02 0.04 0.06
TIME, (sec}

Fig. 34. Pressure Pulse vs. Time

(Ak/k = 0.0025).

58

pressure of 100 psi will give a hoop
stress of approximately 3250 1b/in.?2.
The tensile yield point of Inconel 1is
about 13,000 1b/in.? at 1500°F. The
maximum incremental pressure allowable,
based on the tensile yield point, 1s
400 1b/in.?. The apparent factor of
safety is 4,

Basing the mechanism of failure on
the maximum shear stress theory gives
a maximum incremental pressure of
800 1b/in.?; basing the mechanism
of failure on the distortion-energy
theory gives a maximum incremental
pressure of about 900 1b/in.?. 1In any
case, the least apparent factor of
safety is 4.

o e

 

10

 

100

S0

 

 

80

70

 

60

 

MAXIMUM PRESSURE PULSE {Ib/in2)

50

40 //
30 /////
o /|

0.002

 

 

 

 

 

 

 

 

 

 

 

0.004
Fa¥ 77 4

0006

Fig. 35. Pressure Pulse at Core
Outlet for Various Changes in Reactiv-
ity.