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NASA TECHNICAL nasa tm x- 62,272 

MEMORANDUM 


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{NASA-TH-;X-62272) AH EVALUATION OP SOME N73-27S43 ' 

SPECIAL TECHNIQUES FOR NUCLEAR WASTE- 
DISPOSAL IN SPACE (NASA) 25 p HC $3 a 25 

CSCL 06.1 Unclas 

G3/05 ^9888 J 


AN EVALUATION OF SOME SPECIAL TECHNIQUES FOR 
NUCLEAR WASTE DISPOSAL IN SPACE 


John S. MacKay 


Ames Research Center 
Moffett Field, Calif. 94035 


August 1973 



ABSTRACT 

This note presents a preliminary examination of several special ways 
for space disposal of nuclear waste material which utilize the radioactive 
heat in the waste to assist in the propulsion for deep space trajectories. 
These include use of the wastes (or an extract of the 90 Sr or 137 Cs com- 
pounds contained in the waste) in a thermoelectric generator (RT6) which 
operates an electric propulsion device and a radioisotope - thermal 
thruster which uses hydrogen or ammonia as the propellant. These propul- 
sive devices are compared to the space tug and the space tug/solar electric 
propulsion (SEP) combination for disposal of waste on a solar system escape 
trajectory. Such comparisons indicate that the waste-RTG approach has con- 
siderable potential (disposing of perhaps four times as much waste) pro- 
vided the combined specific mass of the waste container - RTG system does 
not exceed approximately 150 kg/kw fi . 

Although this study stresses the solar system escape destination, 
several exploratory numerical calculations have been made for high Earth 
orbit and Earth escape destinations. These show that some care must be 
exercised in selecting an Earth escape path in order to avoid future near 
encounters with the Earth or Venus. In general, it is believed that useful 
calculations are possible using numerical integration which could help in 
an orbit or trajectory selection process. 



Table of Contents 


Introduction 

Waste Material Form 

Disposal Destinations 

Radioisotope Waste RTG System 

Solar-Electric Propulsion 

Waste Thermal Thruster 

Combined Systems 

Other Heat Sources 

Summary 

References 



AN EVALUATION OF SOME SPECIAL TECHNIQUES FOR 
NUCLEAR WASTE DISPOSAL IN SPACE 

John S. MacKay 
Ames Research Center 
Moffett Field, California 

Introduction 

If the projected future United States power demands are to be 
partly satisfied by stationary nuclear power plants, then there will 
be an associated increase in the amount of nuclear waste material 
which results from the reprocessing of the spent reactor fuel elements 
to recover the unused fuel. While such reprocessing is an inherent 
part of the economical operation of such nuclear power stations, it can 
lead to surprising amounts of radioactive residue. Reference 1, for 
example, contains some projections which lead to an annual output of 
over 5 x 10 5 kg of waste by the year 2000. This, of course, depends on the 
electrical power demands continuing to increase as they have in the past 
and also that no new power producing methods (such as nuclear fusion 
reactors) emerge to meet the increasing demand. 

As a result of such nuclear waste possibilities, the Atomic Energy 
Commission has asked NASA to study the feasibility of disposing of the 
waste products in space. Reference 1 and its associated documents consti- 
tute a direct response to the AEC request. This memorandum has been 
stimulated by that effort but is not an official part of the response 
to the AEC. 

The purpose of this paper, then, is to consider several concepts by 
which the energy still contained in the nuclear waste material could be 
utilized to augment or complete the space disposal process of such waste. 

As indicated in figure 1 (taken from ref. 1), it can be seen that the 
energy output per unit mass (shown on the ordinate in figure 1) is about 
300 w/kg if the waste is obtained at one year from the time of reprocessing. 



- 2 - 


This is a specific energy value equivalent to that of 238 Pu. However the 
nature of the waste is such that (see ref. 2) it is much less dense (in 
its solid form) and much more difficult to shield. The radiation is due 
to the high percentage of short lifetime radioactive elements in the 
waste. As indicated in figure 1, the energy output eventually becomes an expo- 
nential function of time. This represents a transition to activity from 
a few dominant, longer half life elements (i.e. 139 Cs and 90 Sr). This is 
shown more clearly in figure 2 (taken from ref. 3). 

Thus, it would appear that there is at least an interesting amount 
of energy in the waste if it can be obtained early enough (i.e. the 
specific energy output is similar to that of isotopes usually used in 
space applications). Another possibility is the separation of special 
high heat output elements out of the waste. This will be considered as 
an alternate but more expensive way of utilizing the heat in the waste. 

Two propulsive techniques will be considered which use the waste 
heat. One is the direct conversion of the heat into electricity by a 
thermoelectric generator (RTG) and the other is heating of some working 
fluid such. as liquid hydrogen or ammonia and subsequently expelling the 
fluid at high velocity to produce thrust. This is similar to the "poodle" 
thruster concept described in reference 4, and will be referred to here 
as an isotope thermal device. Another somewhat related concept — solar 
electric propulsion (SEP) plus waste (RTG)— will also be considered but 
not evaluated in quantitative terms. These various concepts are illus- 
trated in figure 3. 


Waste Material Form 

Considering the energy output only, it would follow that early 
acquisition and containment of the solid waste would be desirable. This 
would utilize the high energy output of the short half-life elements before 
they decay to less active states. However, it is recognized that the 
existing processing facilities may restrict the acquisition time to one 
or two years after reprocessing of the fuel elements. Thus, one form of 



- 3 - 


the waste products which will be considered here will be the solid state 
one and two years after reprocessing. 


Of the various solid forms which are currently being considered (see 
ref. 2) their density varies between 1.33 and 2.8 gm/cm 3 . In reference 
1, the spray melt solidification process was selected as a desirable form 
for packaging and heat conductivity purposes. The properties of this 
type of solid waste are listed in table 1. 


TABLE 1. SOME PROPERTIES OF SPRAY MELT SOLID WASTE MATERIAL 

(from refs. 1 and 2) 


Heat output (one year) 

Heat output (two years) 

Density 

Thermal conductivity 

Maximum (center line) temperature 


300 watts/kg 
150 watts/kg 
3.0 gm/cm 3 
1 .8 watts/cm 3 /°K 
1170°K 


These properties will be used here as those of typical solidified but 
otherwise unprocessed waste material. 

The second waste form which will be considered here is the separa- 
tion of 137 Cs and/or 90 Sr from the main waste stream. This could be done 
at about five years or more after initial reprocessing operation. At 
the present time, facilities for separating these elements out of the 
waste stream do exist (see ref. 5); thus the cost of increasing output 
or purity may not be excessive. However, the additional cost of perform- 
ing the separation must still be included. 

The purpose of separating out these isotopes is that they have a high 
specific energy output combined with a rather long half life. Some prop- 
erties of these isotopes in their common (usually oxide) forms are given 
in table 2. 



- 4 - 


TABLE 2. SOME 

PROPERTIES OF l37 Cs and 
(from ref. 6) 

90 $r ISOTOPE 

COMPOUNDS 

Isotope 

* 

Shield Density Density 
cm of Uranium gm/cc 

watts/gm 

Half-life 

years 

90 $r (SrO) 

2.3 4.7 

0.453 

27.7 

137 Cs (CsF) 

5.5 3.586 

0.134 

30 


The shield densities shown in this table are computed for a 
spherical 1 kw. source and are for 10 rem/hr at 1 meter from 
the center of the source. 

90 Sr, for example, has an energy output similar to 238 Pu which is often 
used as a heat source for flight RTG power supplies, but has a much more 
severe shielding problem and other safety disadvantages. However, both 
isotopes are relatively good heat sources if separated from the rest of 
the solid waste. Also, they could be bothersome to store on Earth because 
of their relatively long half lives. 

Another waste form which also results from special processing is 
the separation of the actinide compounds from the waste material. These 
isotopes have very long half lives which implies extensive ground storage 
time if they are left in the waste. However, the heat output is very low 
and as such, does not constitute an interesting heat source for propulsive 
or power generating purposes. Rather it probably represents the most 
compact form the waste can take without utilizing some form of nuclear 
transmutation. However, the actinides are only the most troublesome part 
of the waste material and facilities for the storage or use of the rest 
must be provided. 

Because the actinides represent only a fraction of the total waste, 
they are probably best considered along with more conventional space 
disposal techniques such as the shuttle/tug or shuttle/centaur. Use of 
such launch vehicle systems is being examined by the NASA, Lewis Research 
Center (see ref. 1) and will not be considered here. 



- 5 - 


Disposal Destinations 

A number of destinations for the proposed waste containers are 
currently being studied. They include high Earth orbit. Earth escape, 

0.99 AU and 1.1 All circular orbits about the sun, solar system escape and 
solar impact (see figure 4). A comparison of the propulsive energy 
requirement of these destinations can be found in reference 1. 

The nearby destinations such as high Earth orbit or Earth escape are 
of interest because the propulsive energy expenditure required to achieve 
such orbits is low and would probably be most attractive as a destination 
for chemical rocket systems such as the space tug or the centaur. How- 
ever, they may also be of interest for the waste heated fluid concept as 
such devices may have low specific impulses, depending on the working 
fluid used (e.g., ammonia). However, while such destinations may be easy 
to reach, they create another problem regarding whether or not the waste 
is actually disposed of in such a case. This is particularly true of the 
Earth escape case where there is some chance, however small, that the 
waste container may someday return. Several exploratory numerical inte- 
grations were carried out in order to illustrate some of the problems that 
can arise in certain cases. Specifically, the cases so far studied are 
Earth escape and high Earth orbit. 

All numerical integrations have been performed on a CDC 7600 
computer using a version of the LeRC N-Body program (see ref. 7). The 
CDC 7600 has a 60 bit word length which allows single precision 14 digit 
arithmetic. Thus, very accurate numerical integrations are possible 
without the usual need for double precision arithmetic or accumulation. 

Considering first the high Earth orbit case, it was first deter- 
mined that the important perturbations were those due to the moon, sun 
and the Earth's oblateness. Inclusion of Jupiter and several other 
planets had little noticeable effect after several years of integration. 
The predominant changes in the orbit's elements were precession of the 



- 6 - 


line of nodes and the argument of pericenter. The orbit chosen for study 
was circular at 50,000 nautical miles altitude and inclined at 28.5° to 
the equatorial plane. These calculations for the Earth orbit case were 
very time consuming, requiring about 25 seconds of computer time per year 
of orbit time. 

The Earth escape case was less expensive, using less than 50 seconds 
of computer time for 500 years of interplanetary flight. The orbit under 
investigation was an Earth escape trajectory with a perigee altitude of 
100 nautical miles and an eccentricity of 1.1. 

The main results for the Earth escape case were that some care must 
be taken to keep the orbit as much inside the Earth’s orbit as possible. 
That is, Earth departure conditions should be such that the trajectory 
enters heliocentric space at aphelion. Otherwise it was found the 
trajectory would re-enter the Earth’s sphere of influence several times 
within a 100 year span. On one other occasion a trajectory passed 
through the sphere of Venus at 273 years during a 514 year integration 
even when special care is taken to inject at aphelion. This indicates 
that care must also be exercised in selecting the trajectory perihelion. 
In the Earth escape cases, it was found necessary to include all the 
planets out to Jupiter. More planets could not be included because of 
present limitations of the program. 

In all cases the planets and the moon were included with fixed orbit 
elements chosen from some particular epoch. This is a serious omission 
only in the case of the moon, which precesses around the Earth at the 
rate of about 18°/year. However it has become clear that useful calcula- 
tions can be made which can very likely be of value in orbit selection 
and simulation. 

Unlike the low energy cases, there are at least two other destina- 
tions which probably constitute true disposal. They are solar impact 
and solar system escape. Of these two, solar system escape is perhaps 



- 7 - 


the most preferable because of the generally lower energy requirements. 
One objection to solar system escape is that it may become someone else's 
problem in due time. However, the time to reach the nearest star is 
enormous and could leave the package no more harmful than a meteorite. 

(A detailed treatment of the probable hazards associated with these 
destinations can be found in ref. 10.) 

Thus, it would appear that the most preferable destination is solar 
escape with solar impact a second choice. Of the other destinations, 
high Earth orbit or solar orbit are perhaps the least likely to return 
to Earth. However, the preliminary numerical calculations which have 
been completed for the Earth orbit case have indicated that the orbit 
will precess (not unlike the moon) due to solar, lunar and Earth oblate- 
ness perturbations. Thus, it may be difficult to track the waste 
containers for the hundreds or perhaps thousands of years which may be 
required by safety considerations. 

Consequently, this section will consider only solar escape and 
impact as likely destinations for the propulsion systems under considera- 
tion herein. The other destinations will be given ample consideration 
(in ref. 1) and need not be considered here in any further detail. 

It has been shown in reference 1, direct solar impact requires a 
velocity relative to the Earth of 30 km/sec. This stops the package 
relative to the sun and it falls straight down on a radial line to impact. 
Very few propulsion systems presently under consideration (with the 
possible exception of the laser ignited fusion devise described in 
reference 8) could accomplish such a mission. For example, the waste- 
RTG and solar electric propulsion systems can simulate such a mission 
only by a slow spiral into the sun. Unfortunately, the effective velocity 
change for such a maneuver is the difference between the circular orbit 
speeds at the different radii. Thus, to reach the surface of the sun 
(a radius of 0.698 x 10 6 km) would require: 



- 8 - 


aV = V C>Q - V c>0 s 436. - 30 = 406 km/sec (1) 

where V c is the orbital speed at the indicated distance, which is pro- 
bably beyond the capabilities of any ion thruster system. 

A more optimal approach, even for chemical rocket systems, would 
be to first proceed outward to some high aphelion and then nullify the 
velocity at aphelion and drop into the sun. This, as well as the direct 
method, is illustrated in figure 4. The limit in this process is' to 
first essentially escape the solar system (i.e., very high aphelion) 
and then apply a very small correction and return to solar impact. 
Unfortunately, the time involved in such a maneuver is excessive and 
some compromise must be made between the time and a V involved in the 
maneuver. Figure 5 illustrates the interchange between time of flight 
and aV. 

Thus, it is clear that low energy {i.e. low aV) solar impact missions 
and solar escape are closely related and considering one is equivalent 
to considering the limiting case of the other. For this reason, only 
solar escape will be considered in the following sections. 

Radioisotope Waste RTG System 

As noted previously, it is best to use the waste early. Suppose, 
for example, that we obtain the waste at one year; then, from figure 1; 

P/m =0.3 kw/kg 

“min = P7HT^ = 7TTT55~ = 67 k S /kw ( 2 > 

where 


a min = mimnium specific mass of the power supply 
o c = efficiency of thermoelectric converter 



- 9 - 


As indicated, this assumes that the RTG converter efficiency is 5 per- 
cent, which is typical of present day technology. 

However, since the value of P/m falls off so rapidly in figure 1, 
some average value should be chosen. To do this, it will be assumed 
that the electro-static thruster system can operate no longer than 20,000 
hr. (Again, this is typical of current estimates from test and flight 
data. A general description of electrostatic thruster developments and 
operations can be found in reference 11.) Thus, the value of P/m taken 
from figure 1 should be between one and three years. This gives an 
average a . of about 134 kg/kw. 

Assuming the thruster efficiency to be of the form: 

B 0.842 

^ 0 ) 


where 

C = ion exhaust velocity, km/sec 
B = propellant utilization efficiency 
D = ionization loss factor, km/sec, 


and that the propulsion time (t Q ) is limited to 20,000 hr, a value of C 
can be found which gives the highest initial acceleration. This is given 
by the relation: 


C 


opt 


.y 1 


2000 B t 


min 


+ (Dx 1000) 2 


= 34,200 m/sec 

The payload ratio, for optimum C, can be shown to be: 

_ t a o a min C opt 
y L. " 1 B 


( 4 ) 



- 10 - 


where a Q is the initial thrust/mass ratio. 

Since no payload is to be carried in addition to the waste-RTG 
package y^ - 0 and, 


B 


a n = r 

0 a L. 


= 1.84 x 10" 4 m/sec 2 


mi n opt 


Therefore, the propellant fraction, y is 

r 

a t 

y = n 2 — 1 1 = 0.388 
p C opt 


Thus the A V capability of the system is 


(5) 


( 6 ) 


AV = -C opt In (1 - y p ) = 16,800 m/sec (7) 

Using the same criterion noted before, (see equation 1) it follows that 
the package will spiral out into the asteroid belt before it runs out of 
propellant. (This includes Earth escape which requires an additional 
aV of about 8 km/sec.) 


The above example illustrates that some additional velocity may be 
required at Earth departure in order to escape the sun's gravity field. 
To investigate this possibility, some computer calculations were made to 
determine the value of a Q required to reach solar escape starting from 
various values of velocity relative to Earth (supplied by some chemical 
rocket stage such as the Centaur). These are shown in figure 6. This 
has been done with a limit on t of 20,000 hr. and a fixed value of 
C = 30,000 m/sec (this is the lowest practical value based on current 
thruster technology work. Lower values of C develop difficulties in 
accelerator grid spacing required). 

Given the data shown in figure 6, it is then possible to determine 
what values of a Q and V w ^ are required to escape the solar system for 



- 11 - 


any chosen value of . These are determined from the following 
equation: 


a n C 

= 0 = 1 - y - -x a • 

P 2 n t ^ min 


( 8 ) 


where ^ and can be determined from previous relations (equations 
(3) and (6)). 


Assuming the use of the shuttle/centaur, we have the following 
relation between V w ^ and m Q (at Earth escape): 

TABLE 3. SHUTTLE/CENTAUR EARTH DEPARTURE MASS CAPABILITY 


V.j (km/sec) m Q (kg) 


0 10,400 
3 8,500 
6 4,900 
9 2,050 


For this launch system and using an ^ of 134 as before, it is found 
that V <| - 3.0 gives y^ = 0 and leads to an ejected system mass (final 
mass) of 4,850 kg (10,700 lb). This is considerably better than the 
direct solar escape payload of 1,230 kg. given in reference 1 for the 
shuttle/tug (expended). Continuing this process for other values of a 
other than a ^ leads to the results shown in figure 7 where the mass sent 
to solar escape is shown as a function of the propulsion system specific 
mass. As indicated in this figure, all cases above a = 60 kg/kWe will 
require some assistance (from a Centaur, tug or some other chemical rocket 
stage) during Earth escape. These results indicate that the best case of 
a = a min 1S interest * n 9 and that more work should probably be done to 
better define a. 


$ 



Solar-Electric Propulsion 

Another set of calculations has been made for the case of a 20 kw 
solar electric propulsion (SEP) system as the ejection stage. These were 
performed for the same shuttle/centaur departure mode but used an a of 
30 kg/kw (typical for SEP stages) and a typical solar cell profile of 
power as a function of distance from the sun. In this case, it was 
found that V , = 6.0 was required (with I = 3000 sec. as before) which 
gave an injected mass (excluding the SEP system) of 1320 kg at solar 
escape. As indicated in figure 7, this is essentially the same as the 
shuttle/ tug system. Thus it would appear that the SEP approach would 
not have any great advantage over a simple expended tug. Also, the 
combined cost of both the SEP stage and the Centaur would probably exceed 
or equal that of a single tug. 

Waste Thermal Thruster 

Another device examined was one in which the heat in the waste is 
transferred into some fluid which is ejected to produce thrust. A pre- 
liminary set of calculations for such a device was made assuming that 
liquid hydrogen could be heated to a maximum temperature of =200G°F. 

(This is near the maximum centerline temperature of most solid waste 
forms given previously. ) x Assuming complete expansion into a vacuum, an 
exhaust velocity of about 7 km/sec (I = 700 sec.) is theoretically 
possible. However, the amount of waste material (or any isotope) needed 
to produce a sizable thrust is critical. For example, to achieve a 
thrust/mass of .10 requires: 

P ~ F C _ .10 x 7000 _ ocn watts 

m7 ~ nT ? 2 350 TT" 

0 0 

From figure 1, it is clear that this is about as much heat output as can 
be expected from any radioisotope heat source. 

In order to escape the solar system with an initial acceleration of 
.10 m/sec 2 , more aV than the 8,5 km/sec required with very high values of a 



- 13 - 


must be supplied. This is due to the energy expended in lifting the unused 
propellant through a gravity field. Such "gravity loss" factors can be 
found in such documents as reference 9. Specifically, it is found from 
reference 9 that AV must be increased by 1.75 to overcome the "gravity 
losses" associated with an initial acceleration of 0.10 m/sec 2 . 

At this point it is appropriate to try and size a stage which will 
escape the solar system. Assuming that the, liquid H 2 can be contained 
in tank with a mass of about 10 percent of the contained propellant, it 
can be shown that the mass ratio for the maneuver (excluding tanks) is: 


jf = 0 + o)e" AV/C - a = .0314 
0 

where a = tankage factor = 0.10. 


(9) 


If the transfer began in low Earth orbit with m Q = 29600 kg, then 
the propulsion system could be no more massive than 

m eng " m o ( iif) = °- 0314 x 29600 = 927 kg 
Therefore, the thrust, F,' can be no more than: 


_ 2P _ 2 x 927 x 300 
C 7TO3 


79 N 


Thus, 


a 0 ~ 296 — 2.68 x 10" 3 m/sec 2 

At this low a value of a Q , it would require a aV of about 8 km/sec just 
to escape the Earth (see equation 1). This indicates that the system 
can't escape from low Earth orbit without some high thrust assistance. 



- 14 - 


For the case of high thrust assist, new data similar to figure 6 
have been generated. Using the data and table 3, it has been found 
that V = 6 gives the highest mass of ejected waste isotope. This 
mass is shown in figure 8 where the values for shuttle/tug and SEP (20 
kWe) are also shown for comparison. 

In figure 8 it can be seen that a has no effect on the injected 
mass over the range shown. This results from the higher values of a Q 
(to the right in figure 8) that result when the Centaur is used for 
Earth escape. Thus, a has very little effect on aV (through changes in 
a ) until very high values of a are assumed. 

It is therefore concluded from figure 8 that there is insufficient 
heat energy in the waste (or some of its components such as 90 Sr) to be 
of much interest as a thermal thrust producing device. However, all of 
this has assumed only one shuttle launch. If more launches are used, 
the system may be made to compete with a single shuttle/tug. However, 
considering the optimistic assumptions about the propulsion system size, 
it would not seem likely that the device could have much economic 
advantage. 


Combined Systems 

It may also be possible that some combination of the systems so far 
discussed could be a better choice. This is perhaps most true of an SEP 
waste- RTG system. As the power from the solar cells drop off, the RTG 
power would remain to give a more uniform power distribution throughout 
the flight. However, the same solar cell cost argument stated before 
still applies here. Thus, this would appear to be another system worthy 
of further investigation along with the pure waste-RTG devices. 

Other Heat Sources 

As noted earlier, the possibility of using some extracted compound 
of 137 Cs or 90 $r should also be considered. From tables 1 and 2 it is 



- 15 - 


clear that 90 $r would very likely be a much better heat source than the 

waste or 137 Cs. The 450 watts/kg output plus the long half life would 

combine to give a value of a m ^ n = 44 kg/kw. This clearly gives superior 

performance (shown in figure 7) to the shuttle/tug combination. However, 

the cost of separating the 90 Sr compounds from the waste must be 

included. Also, it must be recognized that the compounds may be mixed 

and of a type giving lower energy output than the SrO value given in the 

table. Most important, however, are containment, shielding, and other 

safety considerations which will probably increase a much above a . . 

mm 

This is a detailed design problem which is beyond such a preliminary 
survey. Our purpose here is to try to narrow the alternatives for more 
intensive future study. 


Summary 

This investigation, although very preliminary in nature, has indi- 
cated that there could be some useful ways in which the heat in the 
nuclear waste can be used to augment space disposal of the waste. In 
particular, it appears that an RTG system operating on waste or 90 Sr 
compounds separated from the original waste could be used together with 
an electric propulsion system to reach solar system escape. In each case 
the results are very attractive in the extreme case of no-containment 
weight estimates. This does not mean the scheme will be ultimately use- 
ful, but does indicate that further consideration could be worthwhile. 

Other approaches, such as SEP only and a nuclear-thermal thruster 
using the waste heat and liquid hydrogen, do not appear attractive, even 
with the aid of optimistic assumptions. 

Although this study stresses the solar systems escape destination, 
several exploratory numerical calculations have been made for the high 
Earth orbit and Earth escape destinations. These show that some care 
must be exercised in selecting an Earth escape path in order to avoid 
future near encounters with the Earth and Venus. In general, it is 
believed that useful calculations are possible using numerical integra- 
tion which could help in an orbit selection process. 



- 16 - 


REFERENCES 


1. Hyland, R. E.; Wohl , M. L.; Thompson, R. L.; and Finnegan, P. M.: 

Study of Extraterrestrial Disposal of Radioactive Wastes > Part II. 

NASA TM X-68147, October 1972. 

2. Schneider, K. J.: Solidification and Disposal of High-Level Radio- 

active Wastes in the United States. Reactor Technology, Vol. 13, 
No. 4, Winter 1970-71. 

3. Fox, C. H.: Radioactive Wastes. AEC Series on Understanding the 

Atom, 1969. 

4. Romero, Jacob B.: Radioisotope Propulsion . AIAA J. Spacecraft, 

Vol. 1, No. 5, September-October 1964. 

5. Siting of Fuel Reprocessing Plants and Waste Management Facilities. 

Oak Ridge National Laboratory Report #0RNL-4451 , July 1970. 

6. Properties of Selected Radioisotopes. NASA SP 7031, 1968. 

7. Strack, W. C.; and Huff, V. N.: The N-Body Code — A General Fortran 

Code for the Numerical Solution of Space Mechanics Problems on 
an IBM 7090 Computer. NASA TN D-1730, November 1963. 

8. Boyer, Keith: Laser— Initiated Fusion — Key Experiment Looming . 

AIAA, Astronautics and Aeronautics, January 1973. 

9. Willis, Edward A., Jr.: Finite-Thrust Escape from and Capture into 

Circular and Elliptic Orbits. NASA TN D-3606, September 1966, 

10. Evans, Lawrence C.: Terrestrial and Stellar Contamination from 

Space Disposal of Nuclear Wastes. Proposed NASA TM X- 50, 000 

series. June 1973 (approximate). 

11. Richley, E. A.; and Kerslake, W. R.: Bombardment Thruster Investi- 

gations of the Lewis Research Center. Journal of Spacecraft and 
Rockets , Vol. 6, No. 3, March 1969, pp. 289-295. 



SYMBOL TABLE 


a acceleration, m/sec 2 

B constant in thruster efficiency expression 

C exhaust velocity, km/sec 

D constant in thruster efficiency expression, km/sec 

F thrust, N 

I specific impulse, kg-sec/kg 

a p 

m mass kg 

P power, watts 

t time, sec 

V velocity, km/sec 

a propulsion system specific mass, kw/kg 

aV velocity increment, km/sec 

n efficiency 

y mass ratio 

o tank mass/propellant mass 

Subscripts 
c converter 

eng engine 

f final 

L payload 

min minimum 

opt optimum 

p propellant or propulsion 

th thruster 



- 18 - 


0 initial (t = o) 

1 Earth departure 

00 indicating conditions on the asymptote of a hyperbolic orbit 
® Sun 

® Earth 



HEAT PRODUCED IN WASTE, W/kg 


- 19 - 



YEARS AFTER REMOVAL FROM REACTOR 

Figure 1. - Fission product thermal power as a function of time after 
reprocessing of spent fuel elements. (From NASA TMX-68147) 



Figure 2. - Heat production in high-level wastes from spent fuel processing. 

(After five years, Strontium and Cesium account for most of the 
heat production. Note that removal of these isotopes before 
several years of aging would have little effect on heat producti 
of the remaining mixture.) 




- 20 - 


ISOTOPE- THERMAL 


PROPELLANT 
(LH 2 OR 
AMMONIA) 



ISOTOPE 

WASTE 

HEAT 



H 2 

C= 7 km/sec 


ISOTOPE- ELECTRIC 


POWER 

ONDITIONING 


ISOTOPE 

WASTE HEAT [PROPELLA 
SOURCE ) \ Hq 


NEUTRALIZER 
rZzT* \ ELECTRONS 


ION BEAM 


RTG CONVERTE 



Hg + 

— c= 30 km/sec 

"ELECTROSTATIC 
THRUSTERS 


Figure 3. - Isotope waste heat propulsion concepts. 



Figure 4. - Nuclear waste disposal space trajectories. 




- 21 - 



Figure 5. - Propulsive velocity increment and flight time for solar impact 
missions; H p = 556 km. 



HYPERBOLIC VELOCITY AT EARTH DEPARTURE, 

Vco } |, km/sec 


Figure 6. - Initial acceleration required for solar system escape mission 
t p = 20,000 hr; I = 3000. 


- 22 - 


20000 r 


a> 

~ 15000 
c o 

CO 

< 


10000 


Q 

LU 

|__ 

O 

LlI 

“5 


90 Sr / 


EARTH ESCAPE 


7] r - 0.05 

0.842 

^ th " I + (16/C) 2 
a (SEP) = 30 Kg/KWe 

LOW 
EARTH ORBIT 

SHUTTLE/ 

CENTAUR/ WASTE - ELECTRIC 

SOLID WASTE 


SHUTTLE/ 
CENTAUR/ 

5000 [- SEP (20 KW e ) 



SHUTTLE/TUG (EXPENDED) 


0 50 100 150 200 

PROPULSION SYSTEM SPECIFIC MASS, a, Kg/KW e 

Figure 7. - Mass delivered to solar system escape by various propulsion 
system concepts; shuttle payload = 65,000 lb. at 100 n.mi.; 
t « 20,000 hr. 


2000 


O' 

^ 1500 

CO 

CO 

< 


Q 
LU 
I — 

o 

LU 

-> 


1000 h 


500 


SHUTTLE/CENTAUR/WASTE-THERMAL 


SHUTTLE/ 
CENTAUR/ 
SEP (20 KW e ) 


SHUTTLE/TUG (EXPENDED) 


0 50 100 150 

PROPULSION SYSTEM SPECIFIC MASS, a, kg/KW e 


Figure 8. - Mass delivered to solar system escape; comparison of some 

alternative systems; I = 700 (isotope-thermal); a(SEP) = 30 kg/kw, 

sp