NASA Technical Reports Server (NTRS) 19920001884: Pellet bed reactor for nuclear propelled vehicles: Part 2: Missions and vehicle integration trades

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N9 2- 1 [ 02
PELLET BED REACTOR FOR NUCLEAR
PROPELLED VEHICLES:
II. MISSIONS AND VEHICLE INTEGRATION TRADES
V. E. (Bill) Haloulakos
McDonnell Douglas
Huntington Beach, CA
As Mohamed said, I will be discussing the mission and vehicle integration trades and so I
am not going to say anything about reactors, neutronics or anything else. The issue here
is that you can make a reactor or an engine, but unless you can hang it into a vehicle it
won't go anywhere. So I would like to address some of these issues.
You have to go through all of these factors (Figure 1) before you know if the vehicle can
fly. You have to look at the whole vehicle. You can have all kinds of efficiencies you
want in the reactor, but if it doesn't fly, it won't go anywhere.
Here are some of the trades done back then during the NERVA program (Figure 2).
What shape is your tank and where do you put your rocket engine and your reactor?
You go in with some distance to avoid the radiation (this will cause feed system
problems), then you begin to play with geometry; the optimum that came out is a 15
degree cone angle.
Figure 3 shows the mass and radiation breakdown for the shielding from the previous
chart that I showed you. The 15-degree cone angle gives you the lowest radiation for a
given shield mass. So, based on this chart it was decided that we would pick the 15-
degree cone angle as the bottom of the tank.
There were many other trades that were done. Here is what the problem looked like;
you are not going to Mars and get rid of the reactor, you are going to fire it, shut it
down, and then you have to cool it. When you use propellant as coolant, you lose
specific impulse. The trades done back then show what happens to your specificimpulse
as you cool the reactor down (Figure 4). So you have to go through these trades as well.
As to radiation maps (Figure 5), I am not a radiation expert, but these were done back
for the NERVA engine. You have neutron flux, you have gamma radiation, a reference
point up there and we are talking about a 1575 megawatt reactors operating for 53
minutes and so on. So all these factors have to be addressed.
Then as to what happens after shut down (Figure 6), you have a decay which goes as
shown, and here is the radiation versus distance, which continues on, and so on.
255
PRECEDING PAGE BLANK NOT FILMED
In our present studies (Figures 7-9) we are moving from the 1960's to the 1980's and
1990's via computer programs. We had a very good correlation between the calculations
from the old NERVA data that we got out of the design handbooks. The same thing
was found for a small engine that was supposed to operate an ROTV out of the space
shuttle, (if you can believe that) (Figure 7).
For a pellet bed reactor mission to Mars, just the other day one of our guys gave me
these numbers (Figure 10). If you fly on May 11, 2018, taking 250 days for the total trip,
with 30 days stay, these are your Delta-V breakdowns. So on the basis of this, we can
take a thrust, an engine, and hang it on the vehicle and start calculating some system
masses and see what happens.
This is what happens when you plot Delta velocity versus mass (Figure 11). The way we
break things down is shown in Figure 12. We have a Delta velocity and a specific
impulse of 1,000 seconds when we calculated with our program. We come up with a
payload of 36 metric tons, the thrust is 315 kilo-Newtons. That's about 70,000 pounds or
so, including the mass of the shield. This is the output. I must say this mass ratio is not
payload fraction. Payload fraction is shown in Figure 13. This is for the top curve, the
heaviest vehicle that we got and that's almost a half a million kilograms there. Pretty big
stuff!
Looking at it parametrically in terms of payload fraction, we show that, as you demand
more and more velocity out of a fixed performance, your vehicle becomes almost like the
chemicals we have today, which have something like three to four percent payload
fraction. This says that what you want to do is increase the specific impulse. And by the
way, if you go to a single stage Delta V, which is like nine to ten kilometers per second
with a nuclear vehicle, you begin to approach 25 percent of payload fraction.
I was talking to airplane people who design airplanes being flown for money and they
say that of their takeoff weight, fuel is something like 40 percent. What we would like to
do is drive the space vehicles in that direction.
We didn't do anything on cost for this workshop, but we did a lot of work on cost back
in the 1970's. There is a whole bunch of reports that I sent NASA, and one written on
February 1973 cost data, 1973 dollars. Oh, do they look good. I suggest that you take
that to Congress when you go and talk to them.
256
BIBLIOGRAPHY
Bill Haloulakos
PCBR fgr Electric and Thermal Nuclear Propuhion
Acknowledgments
The development of the PeBR for electric and thermal nudear propulsion missions has been
performed by the University of New Mexico's Institute for Space Nuclear Power Studies. The System Trades
and Performance Studieswere performed atMcDonnell Douglas Space Systems Company, Huntington
Beach, CA; thecontributionsofT.M. Millerand J.F.l.aBararehereby acknowledged.
io Alcoufee, R.E., F.W. Brinkley, D.R. Mart and R.D. O'DeU, "Users guide for TWODANT: A Code
Package for Two-Dimensional, DifiMsion-Accelerated Neutral-Particle Transport', Report No. LA-
10049-M Rev. 1, UC-32, Los Alamos National Laboratory, Los Alamos, NM (1984).
. Altseimer, J.H. et al "Operating Characteristics and Requirements for the NERVA Flight Engines',
J. Spacecraft, 8(7):766-773 (1971).
, Bennett, G. "Nuclear Thermal Propulsion Program Overview," NASA Nuclear Thermal Propulsion
Workshop, Cleveland, OH 10-12 July, 1990.
4o Durham, F.P. "Nuclear Engine Definition Study Preliminary Report', LA-5044-MS, Vol. I-Ill, Los
Alamos Scientific Laboratory (1972) Los Alamos, NM.
. E1-Genk, M.S., A.G. Parlos, J.M. McGhee, S. Lapin, D. Buden and J. Mires, "System Design
Optimization for Multimegawatt Space Nuclear Power Applications', J. Propulsion and Power,
(2)'194-202, 1990a.
o E1-Genk, M.S, N.J. Morley, and V.E. (Bill) Haloulakos, "Pellet bed Reactor for Nuclear Propelled
Vehicles", NASA Nuclear Thermal Propulsion Workshop, Cleveland, OH 19-22 June 1990.
. EI-Genk, M.S. "Pellet Bed Reactor Concept for Space Propulsion', NASA Nuclear Electric
Propulsion Workshop, Pasadena Convention Center, Pasadena, CA 19-99 June 1990.
o Gordon S. and B.J. McBride. "Computer Program for Calculations of Complex Chemical
Equilibrium Composites, Rocket Performance Evident Reflected Shocks and Chapman-Jouguet
Detonations', NASA SP-2T3, NASA Lewis Research Center, Cleveland, OH, March 1976.
. Haloulakos V.E. and C.B. Coehmer. "Nuclear Propulsion: Past, Present, and Future', McDonnell
Douglas paper No. MDC H2642, Trans. 5th Symposium on Space Nudear Power Systems, CONF-
880122-SUMMS, Albuquerque, NM, 11-14 January 1988.
10. Nabielek, H. et al. "Fuel for Pebble-Bed HTRs', J. Nuclear Engineering and Design, 78:155-166,
(1984).
II. Stansfield, O.M., F3. Noman, W,_. Simon, and R.F. Turner. "Interactions of Fission Products and
SiC in TRISO Fuel Pardclea: Limiting HTGR Design Parameter', Report No. Ga-A17183-UC-7"/,
General Atomis Technologies, San Diego, CA, U.S. Dept. of Energy, San Francisco, CA, 1983.
257
DESIGN TRADE STUDIES
• Propellant Tank Geometries
• Weight
• Volumetric Efficiency
• Radiation Considerations
• Skirts and Interlaces
• Handling, Transportation and Launching Factors
• Reusable vs Expendable
• Refueling, Refurbishing
• Start, Shutdown, Restart Factors
• Fluid Transients
• Heat Soak Back
• Post Shutdown Cooling
Performance Loss/Recovery
Figure 1
CONVENTIONAL TANK CONFIGURATION
f--.,,
f--._
I
_N ELLIP$OIO
ELLIPSOID
V3.1 i
LLIPSOID
_36415 IN._
L200 tN. 200 IN. _k_
e
Figure 2
258
SHIELD WEIGHT REQUIREMENTS FOR
CONVENTIONAL TANK CONFIGURATIONS
104
Z 103 300.000 LH 2 CAPACITY
I<- _ELLIPSOID. 356 IN. 3.300-L8 INTERNAL SHIELD
U 7.500-LB LH 2 RESIDUAL
0.J 3-ZONE DISK SHIELD
ALL SOURCES IN MAY 1969 CRAM RPT
uJ
¢U¢ 200-IN. ENGINE TANK SEPARATION
)- [UNLESS OTHERWISE NOTED}
<
.-_-.-._.... .___ ._
O
0 %/2 ELLIPSOID. 500 IN._ -
15° CONE HALF-ANGLE /
101 I I 1 [ ! I [ 1
3 5 7 9 II 13 15 17 19
TOTAL SHIELD WEIGHT. I1.000 LB1
EFFECTIVE SPECIFIC IMPULSE Figure 3
1.00
.: _-...:....t...,.,,.,ll,i1. !.,_."_,°,:oo_,.o.o_i':,!,i!liii_
.'" c' ;,-xr ;'l :r.: ,:!- ''''t'::'.°t .-'=':L:U
o.,ofii i_! l:!_I!_!:l_;_;:!]t_-1,_,,,Iocoo,o,,,............
H °-qs :;:. :=:-,:÷,+; ..:1"_
,o.,_..,... _,.:!-,:_- .-2
'- " " " _--___-.o
0.92 _-.... _" ;,_rrl".;,'_i.,,L
,.= 0.90 ./...'."_. ""4_--_ _-_- _ '-_
.... _ CC'OLO(WN: -'FFE(':TIVEI'JESS FAC;'OR--
a. : ::-:':-_ _-_ .... I----4- ._--t-:_-- i-_ ---
0.88
!,:
...... . , .... , ,,,._,__:_-
h.i_o.ee//// -_; ......
IX: _:,-_i- ! _ +i1-_ " " :'"
l¢1 •_ " " _ i i _ I i i i . 1 . I I 1 1 [ -tI.1! ,I,,,,:_• . :,,, ,_,__..
<> 0.84
4; .... ..L .11_'_ |'|1 i i|[ !1_: ;i: i' '|J
0.82
.... .-T'.IZ -LI-L- ._4-.L _J.. :.L;
500 I000 1500 2000 2500 3000
FULL POWER RUN TIME, sec
259 Figure 4
RADIATION MAP
ASSUMES
DISKSHIELDAND
SINGLE$3MIN BURN
xlo6
OISTANCE --- OlSTANC[#"[)
NZUI'RCNFLUXg_IK;M2S"lEC>O3,A_r'VI /
DURING|. 375-MWOPERATION
GA,MAO,OSE RATE(REM/HR)
DURING1.$75-MWOPERATION
Figure 5
DOSE RATE AFTER SHUTDOWN
0.1
0.01
• 0.0_1 I I I ,
10 lOJ 1o000
TIMI AftER 114q,ITOOWII. 1141111
m'FT asrr toe II II_ _ _ff
I
PmIIQN PROOqJ¢'Ir _ RAT|
I NIl AIFTIR (LI,HROPIEIliATII(_I AT II.IrJ'II,,MW AFTllq O.J*HII. I.I')I.MW Ol_U1rloN
260 Figure 6
CLASS 1 SINGLE-MODULE HYBRID RNS
COMMANDB_AFOFFRLWEASRLCODOASNNHTDR_OLuMMBOIIDULE (6 FT) 7
INTEGRAL WAFFLE--
._ °--
-"_
TUNNEL
INTERMODULE STRUCTURE
DOCKING
LIQUID LEVEL S
\ APSNOZZLE
SPRAYNOZZLE
REFILL
_-- PROPELLANT MODULE
(103.4 FT}
\ PROPELLANT THERMAL ANDMETEOROID
FEEDSYSTEM PROTECTION
PROPULSIONMODULE(60FT)
NERVA
Figure 7
ii i
SPACE TRANSFER VEHICLE DESIGN DATA
LEO to GEO With Return Peylcad Mau 36,000 kg
Million Data : Velocity Increment 9,000 m/s
CRYOGENIC (O2/H2) NUCLEAR FISSION FUSION
_m 6 RL - 10's J - 2 LANL ALPHA 2 hE
Thrust (kN) 400.3 902 71.7 100 I
Specific Iml_lie (s) 450 429 860 3.500 1
Bum Time (=) 3,675 1,850 12,835 5.766 I
(kg) 333,291 396,423 109,120 16,792 99.
Fu_ (t.H2) 51,275 60,988
Oxkllzer (LOX) 282,015 335.435
PROPEU.ANT TANK(S)
ToIal Volume (i'n3) 970 1,154 1.748 269 1,e
i (kg) 8,156 9,701 11,746 1,808 10,7
_TION
(tie system) (kg) 1,374 1,635 1,979 305 1.8
mOWS (k9) 792 1,579 2,567 17,168 80.9
(kg) 3,411 4,058 4.913 756 4,4
.. ......... _..°.°._..°.. ......
TOT/_LVEHICLEMASS (ks) 383,024 449,394 166,326 72,828 233
Figure 8
26!
SPACE TRANSFER VEHICLE DESIGN DATA
LEO-GEO-LEOMiss_on ; Mp1,,36.0001_ ; DelV-gkm/s ; Burn Time=3675s
ROCKEr ENGINE CRYOGENIC ,6 RL-10" I NUCLEAR. 4ALPHA 2"l FUSION. Me-12 (Isp)
Thrust (_) 4OO 278 208
Sp_ncImpulse(=) 450 860 2500
PROPELLANTS (kg) 333,291 134,548 34.685
F.e (LH2) 51,275
Od(Izer (LOX) 282.015
PROPELLANT TANK(S)
Total Volume (m"j) 970 1,937 499
Me. (kO) 8.154 10.849 2.797
PRESSURIZATION
He Syslem (kg) 1.374 1,828 471
ENGINE(S) (kg) 792 10.270 30.000
MISCELLANEOUS (_g) 3,411 4,538 1,170
TOTAL VEHICLE MASS 383,024 198.033 105,123
Figure 9
MAR.s P_l$ _to_ _V %UMhA A P-Y
TR_N5- l AE5 iN 3"E C.TI OM _."71
IP
C..nP'TU E E" ,_.o_ tl
I "TP.i_MS- E_I2-rH IhJ3EC'l'lOki ee
*/,_o
E_ p...-l"_ ca PTUP..E tf
TOTI_ L "3 0._5"_
• LAUMCI-I "OP,'TF"_ II mn_ ;lo[8
• TOTAL TI2_IVEL "TttPC.' 2 50 "Z)I_t5
• _nP.s 5"7-_q.Y"//r_& : 30 ,,
262 JAdllnceCPm_& Pm_,------- Figure lO
TOTAL OI'V MASS vs. VELOCITY INCREMENT
PELLET BED RUCTOR NTR
500000
Mpl = 36 m
II 400000
.._
i
Mpl - 20 mt
I 300000
IMpl- I0 mt
200000
0
>
I
Ill 100000
0
I " I " I " I " I " I ,
2000 4000 6000 8000 I0000 t2000 '14000 16000
Detta V Im/sl
4 Advanced Propulsion &Power=mmm Figure 11
m J_=84_A_B. _4dmi_UD
Ell.LIT lid NUCLEAR O1V MASS BREAKDOWN
Input PIrmmeters
Delta V (&V) 15,000 m/s
Sped_lmpulse 1000 s
Payload Mass 36,000 kg
Thrust 315 kN
Engine Mass 1,875 kg
Shield Mass 4,000 kg
Calculated Parlmeter=
Mass Ratio R 4.611
Propellant Fraction (Mp/Mo) 0.857
Payfi)ad Fraction (Mpl/Mo) 0.086
Tank Volume 5,249 m=
Bum Time 170 min
Component Miss Breakdown
Propellant (H=) 364,568 kg
Prop_lant Tank 34,703 kg
Thrust Structure 649 kg
Pressurization System (He) 4,365 kg
Meteoroid/Thermal 9,164 kg
Total Vehide Mass 455,324 kg
263 Figure 12
4e-_vNar=_ _OUOLAI , Space Exp/oraffon InlUaUve m
PAYLOAD FRACTION VS. VELOCITY INCREMENT
PELLET BED REACTOR NTR
06
0.5
36 metric ton payload
20 metric ton payload
j O.d
E
_ 0.2
spec,,c_mp.ls+IsD - 1000s __
Engine Mass Meng -1875 kg _"",_l
i'l_rust - 3 15 kN "_
NTR Shield Mass - ,4000 kg
0.0, J I ...... I ! ! | I
2000. 4000 6000 8000 10000 12000 14000 16000
Den= V (nVs)
Figure 13
II I I Admnced Propu/s/on S Pow_ _
264