Pushing the Envelope for Space Nukes

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THE IDEAL SPACE RADIATOR FOR AN ORION STYLE
3-DOME BOMBLET "BLAST FURNACE"
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In a discussion of radiator systems, a method for calcu-
lating an 'optimum radiator design' has been investigated by
Gunther(2):

The optimum radiator design minimizes mass, while minimizing
the total number of fin heat pipes required. Minimizing mass
reduces launch costs, while minimizing the total number of
heat pipes minimizes fabrication costs. A computer model was
developed to determine the fin heat pipe and fin design that
would lead to the optimum radiator design. The independent
parameters a Number of Header Heat Pipes, Header Heat Pipe
Diameter, Header Heat Pipe Shape, Header Heat Pipe Wall
Material, Number of Fin Heat Pipes per Header Heat Pipe,
Fin Heat Pipe Cross-Sectional Shape, Fin Heat Pipe Height,
Fin Heat Pipe Width, Fin Material, and Fin Thickness. The
computer model includes the complete set of pressure drop
and heat transfer equations for both the header heat pipe
and fin heat pipes. The effective temperature of the fin for
radiation is based on the temperature of the Na and K in the
heat exchanger, as well as the following temperature drops:
header pipe evaporator envelope conduction, header pipe eva-
porator wick conduction, header pipe vapor temperature drop,
header pipe condenser wick conduction, header pipe condenser
envelope conduction, fin pipe evaporator wick conduction,
fin pipe vapor temperature drop, fin pipe condenser wick con-
duction, fin pipe condenser envelope conduction, transverse
conduction into the fin, and conduction along the fin, with
radiation from the fin outer surface. For each set of para-
meters, the computer model determines the radiator mass if
a solution exists. A solution exists when the heat rejected
meets the designed goal, and the pumping capability of the
header and fin pipe wicks is sufficient to return fluid from
the condenser to the evaporator.

Using the above approach, a few parallels can be drawn be-
tween the heat removal system of the Space-R Radiator(2) and
Dome Structures. Both systems use W as a 'cladding' material,
except that on the dome structures, the W cladding is open to
the environment of space, rather than being closed. The number,
diameter, shape, and material of heat pipe headers vary accord-
ing to the geometry of the domes and radiator support structure.
The primary lithium coolant loop is powered by a thermoelectric
electromagnetic pump (TEM) similar to the one used for the
SP-100 reactor(3). The primary heat loop includes a gas gap
heat exchanger, test assembly, expansion tank, TEM pump,
lithium charge tank, system heaters (required for startup),
volume compensator, level detector, magnetic flowmeter, valves,
gauges, and Nb 1% Zr piping from the reactor dome(6). The de-
scription of the heat removal system will focus on the primary
header pipe and heat loop.

During startup operations, frozen Li is uniformly preheated
to 120 degrees centigrade. Header pipe heaters are then ac-
tivated to gradually increase the melt-out performance through
the headers toward the maximum (1250W) of heat input. The heat
exchanger, or gas gap IHX located between the first heat loop
and heat fin piping is analyzed using parameters to solve for
heat exchanger outlet temperature (of the shell and tube tem-
peratures). The design is optimized for a counterflow condition
under the temperature-cross limit envelope(5). Some of these
parameters include:

1. Total heat transfer area for the shell and crossflow tubes
2. Crossflow, tube, and shell inlet temperatures
3. Crossflow, tube, and shell outlet temperatures
4. Overall heat transfer coefficient for shell and crossflow
5. Tube - Shell capacity rate ratio
6. Capacity rates for tube, shell, and crossflow
7. Tube - Shell no. of transfer units
8. Equivalent number of transfer units for crossflow - shell
heat transfer.
9. Ratio of crossflow - shell inlet temperature difference to
tube - shell inlet temperature difference.
10. Dimensionless constants

The heat exchanger effectiveness (E) is defined as the actual
temperature change divided by the maximum possible temperature
change. According to (1), the effectiveness is a "tangible
control" over the "normally expected conditions (tube size,
fluid flow rate, and fin geometry)" as well as the "duct length
and duct section lengths." Computer analysis is used to solve
for fin sizing in the primary heat loop. Parameters such as heat
exchanged, fluid and duct temperatures, friction and pressure
drop are configured utilizing round tubes and symmetrical fins.

DEFINITION OF TERMS:

DATA FILE POSITION
COMPUTER PROGRAM SYNTAX
VALUE IN UNITS

DA(1)
DO
OUTSIDE DIAMETER IN FT. = 0.2395

DA(2)
DI
INSIDE DIAMETER IN FT. = 0.20575

DA(3)
ENT
NO. OF DUCTS IN HEAT EXCHANGER = 1.0

DA(4)
WDOT
TOTAL FLUID WT. FLOW IN LB. PER HOUR = 216.63

DA(5)
FINLH
EXTENDED SURFACE FIN LENGTH IN FT. = 5.0725

DA(6)
FINTH
THICKNESS OF EXTENDED SURFACE AT ROOT = 0.0208

DA(7)
FINTC
FIN THICKNESS AT FAR EDGE = 0.0208

DA(8)
RHOFM
DENSITY OF FIN MATERIAL = 480.384 LB / CU FT

DA(9)
RHOTM
DENSITY OF TUBE MAT'L IN LB / FT3 = 480.384

DA(10)
NOT USED
NOT USED

DA(11)
FNK
THERMAL CONDUCTIVITY OF FIN MAT'L =
30.623 BTU PER HR-FT-OR

DA(12)
CP
SPECIFIC HEAT OF FLUID = 0.3105 BTU PER LB OR

DA(13)
TF1
FLUID TEMPERATURE AT ENTRANCE = 2880 OR

DA(14)
HA
HEAT TRANSFER COEFFICIENT SIDE A = 312.0

DA(15)
HB
HEAT TRANSFER COEFFICIENT SIDE B = 312.0

DA(16)
TAA
AMBIENT TEMP SIDE A = 1170.0 OR

DA(17)
TAB
AMBIENT TEMP SIDE B = 1170.0 OR

DA(18)
ALPHAA
ABSORPTIVITY OF SURFACE FACING SUN = 0.3

DA(19)
ALPHAB
ABSORPTIVITY OF SURFACE AWAY FROM SUN = 0.3

DA(20)
EPSA
EMISSIVITY OF EXTENDED SURFACE FACING
SUN = 0.8

DA(21)
EPSB
EMISSIVITY OF EXTENDED SURFACE AWAY FROM
SUN = 0.6

DA(22)
EPSX
EMISSIVITY OF EXTERNAL BODY 'X' = 0.7

DA(23)
FA
RADIATIVE FORM FACTOR BETWEEN HEAT EXCHANGER
AND BODY 'M' = 0.7

DA(24)
FAX
RADIATIVE FORM FACTOR BETWEEN HEAT EXCHANGER
SURFACE FACING SUN AND A SECOND SURFACE NEAR
HEAT EXCHANGER = 0.07

DA(25)
FB
RADIATIVE FORM FACTOR BETWEEN HEAT EXCHANGER
BODY 'M' FOR SURFACE FACING SUN = 0.07

DA(26)
FBX
RADIATIVE FORM FACTOR BETWEEN HEAT EXCHANGER
SURFACE AWAY FROM SUN AND SECOND SURFACE NEAR
EXCHANGER = 0.07

DA(27)
RHOM
SURFACE REFLECTIVITY OF BODY 'M' = 0.7

DA(28)
RHOX
REFLECTIVITY OF SECOND SURFACE 'X' = 0.7

DA(29)
THETAP
ANGLE BETWEEN SUN'S RAYS AND NORMAL TO FIN
SURFACE = 40O

DA(30)
THETAM
ANGLE BETWEEN SUNS RAYS AND NORMAL TO BODY 'M'
IN DEGREES = MINUS 30O

DA(31)
THETAX
ANGLE BETWEEN SUNS RAYS AND NORMAL TO SECOND
SURFACE = 30O

DA(32)
DM
SURFACE TEMPERATURE OF BODY 'M' = 1170 OR

DA(33)
TX
SURFACE TEMPERATURE OF BODY 'X' = 1170 OR

DA(34)
EPSM
EMISSIVITY OF BODY 'M' = 0.7

DA(35)
ITLT
MAXIMUM NUMBER OF ITERATIONS ALLOWED TO
REVISE INITIAL DZ / DW = 4.0

DA(36)
ITER
NO. OF INTEGRATION STEPS FOR A STARTING
VALUE DZ / DW = 0.0

DA(37)
SC
SOLAR CONSTANT = 424.77 BTU / FT2HR

DA(38)
BIGE
HEAT EXCHANGER EFFECTIVENESS = 0.8

DA(39)
FMESH
NUMBER OF SECTIONS IN WHICH DUCT LENGTH
IS DIVIDED = 1.0

DA(40)
P
PRESSURE = 720 PSI

DA(41)
CKH
NUSSELT NUMBER = 3.65

(References available on request - contact )

Long range distributed fortran programming is used for
prototype orbital and deorbital diagnostic programming for
establishing range-rate data is provided after establishing
the actual position and degrees latitudinal periastric.

Range-rate data is supplied by the DFP as well as Romotar
(Range only measurement of trajectory and recording). In
addition to DFP and Romotar, a star tracker and laser
ranging system are used for asteroid transfer trajectory,
determining the spacecraft's orientation, and geosynchro-
nous mapping of the asteroidal surface.

The asteroid transfer trajectory is maintained by inter-
planetary transfer orbit, recalling that the Earth's
orbital speed represents the speed at aphelion or peri-
helion of the transfer orbit, and the spacecraft's velo-
city merely needs to be increased or decreased in the tan-
gential direction to achieve the desired transfer orbit.
It is actually interplanetary trajectory that must be
timed to occur on the proper side of Earth, towards or
away from the sun. During transfer trajectory, additional
command sequences are uplinked and loaded aboard for exe-
cution, to replace the command sequence exercised during
orbital launch. These take the spacecraft through its rou-
tine cruise operations, such as tracking Earth with its
HGA and monitoring celestial references for attitude con-
trol. Safing using FP (Fault Protection) algorithms that
request the CDS (Command and Data Subsystem) to help re-
establish Earth-pointing and regain communications in case
of malfunction. For example, CLT (Command Loss Timer)
fault protection response issues commands for actions
such as swapping to redundant hardware in an attempt to
re-establish the ability to receive commands.

Raw data is dependent upon the tuning of sensor electron-
ics used to establish range, trajectory, orientation, and
tracking of the spacecraft. Typical inputs to the DFP,
Romotar, tracker, and ranging system are in ascii code
that is read from artificial input for diagnostic testing
and software analysis of the teleoperated system. The
dynamic simulator described below is used to create data
sets for artificial inputs to an orbital simulator.

http://server6.theimagehosting.com/i...=simulator.gif

Multiformat Dynamic Simulators are used to test the
Multiformat Decommutator and database defining the data
stream. They are also used to provide both static and
dynamic simulation capabilities for virtually all telemetry
formats in both open and closed-loop applications. Both
models MDSxxx and XDSxxx exceed the requirements of chapter
4 of IRIG 106-86 Class II Telemetry Standards. The MDSxxx
model generates a serial data stream in any selected code
at rates up to 20Mbits/sec and can instantaneously switch
formats at major frame boundaries for up to 16
different formats.