COMPARATIVE STUDIES O F AN ELECTRICAL POWER GENERATOR FOR A MARS PROBEILANDER

by Monty Koslover

Avco-Corporation, Wi-lmington, Massachusetts

This paper describes the selection of a power generator for a conceptual design M a r s probe/ lander l The scope of generators considered is l imited to those expected to be at a p r a c - intended for the 1970 ' s .

t ical prototype level by November 1966 based on 1965 state-of- the-art evaluation. unusual design constraints such a s heat sterilization, high landing shock and thermal isolation, the following is concluded: i s selected as first choice. be used. since thermal and impact shock problems have no satisfactory solution at this t ime. generators a r e impractical for the mission. power-s'ource weight problem is likely to severely limit miss ion duration.

After application of

1) A battery is the mos t suitable power generator and a nickel-cadmium battery 2) With sufficient development allowed, some form of fuel-cell system may

3) The use of a radioisotope thermoelectr ic generator (RTG) with a hard lander is unlikely 4) Solar powered

5) More efficient generators need to be developed or the

Much

Nomenclature

Total heat supplied, watts Heat supplied to thermo elements, watts Insulation heat l o s s , watts Electr ical power out, watts Temperature C Area, cm2 Pa th length, c m Conductivity watts / " C- m figure of m e r i t = a 2 / k p Weight, gm Thermal conductance. watt/ "C Thermoelectric design constant, watts - c m 3 - O C ~ Thermoelectric design constant, watts - c m - OC

Element length, c m Power density, watts / f t 2 Volts Weight of radiatorlunit a rea ; gm/cmZ A r r a y efficiency, dimens ionl e s s Thermal efficiency, dimensionless Fuel power density, watts / g m Decay ra te , function of t ime dirnensionles s Stefan - Boltzmann constant watts/cmZ "K4 emissivity, dimensionless Heat of formation, Kcal/g-mole Functional expression for thermo element a r r a y Functional expression for heat relations Thermo element density, gm/crn3

I. Introduction

of the l i t e ra tu re on space power genera-

frequently leads to preliminary conclusions which m u s t b e discarded when the constraints imposed by an actual vehicle design a r e applied. of this paper is to show how to select a: first and second choice generator compatible with the probellander when all known applicable constraints, ar is ing from a conceptual design, a r e applied to a spectrum of possible power generators .

which much of this paper is based, was performed in 1965. In some a r e a s of power generators, de- velopment has been v e r y fas t and subsequently, the original information and conclusions have been updated.

approach is s o largely theoretical that no updating is needed. The theoretical weights, etc. represent minimums ra ther than actualities. Est imates of probable performance and weights a r e given in the text based on nonclassified information.

summary, it includes only those aspects pertinent to understanding the selection process which now follows.

The purpose

It mus t be borne in mind that the study, on

In other a r e a s such a s RTG design, the

Finally, since this paper i s essentially a

11. Discussion of Conceptual Design and Constraints

A. Conceptual Design

Figures 1 and 2 show a simplified vers ion of a conceptual probellander. It i s based upon an entry f r o m the approach trajectory, a weight allotment in the o rder of 2000 pounds, and a severe impact landing in any attitude, At thc required t ime and point i n the approach trajectory, the probellander (hereafter r e f e r r e d to a s the flight capsule o r FC) is detached f r o m the flight spacecraf t such as the -

t o r s is concerned with general capabilities, and Mariner 4 and enters into a landing trajectory. The

This paper presents results of mission studies, parts of which were performed under SASB Contract No. XAS 1-3221 for langley Research Center. The remainder of the work was supported by Avco Corp. IRAD funds. The author wishes to thank J.J. Nahoney and J.J. Dcwd Jr. of Srco fur their technical CQrnmenfS and assistance. Appreciation is also extended f O S.P. Bannerton, Allis-Chaimers Research Dirision and W. Patterson of Pram and Phitney Aircraft Division who submitted informarion about their most recenf work on fuel cells and gave permission for chis to he quoted.

60

trajectory is calculated to impact at a selected point on a southern latitude then experiencing reasonable sunlight period.

12 days during which the flight capsule i s virtually dormant. During the last day however, the com- plete entry, descent, and post-impact mission will be conducted. The probe measurements of the at- mosphere, etc. occur prior to entry and a re transmitted immediately to the flight spacecraft and Earth. impact a r e transmitted directly to Earth, the process being repeated after an approximate 23 - hour data-collection period.

A design of the landed capsule, shown in Figure 2, discloses thick layers of crush-up material for the purpose of surviving impact. fortunately, the crush-up material is also an ex- cellent thermal insulator. in removing unwanted heat that may be generated internally in the capsule during transit. promising solution may be a heat pipe2. 3, 4, 5 $ or several heat pipes which a r e connected to the inner landed shell and the outermost surface of the lander capsule. The outermost surface would be a very thin, light metal such as aluminum, coated to provide adequate emissivity (this is dis- cussed further in the section on RTG's). Imme- diately after impact, the crush-up material can be removed by one of several techniques leaving the inner shell frame to act as a radiator. or not this i s done would depend greatly on the power system chosen and the needs of thermal control resulting.

The landing trajectory phase may take up to

The lander measurements taken after

Un-

This causes a problem

A

Whether

B. Power Requirements

Before treating the constraints, a discussion of the power profile shown in Figure 3 follows. i s one of several profile options that were con- sidered and represents the lightest load require- ments. Only total power levels a r e shown rather than requirements for separate components. F o r example, the status check block represents power used by transmitters and instruments intermit-

It

36-hour, thermal-soak periods. Nothing may be added to the system after sterilization.

High-Impact Shock Survival. power system inside the landed capsule must survive landing impact in acceptable working order. The reference level i s l O O O g , triangular shape wave with base times of 3 or 5 milliseconds.

That portion of the

Thermal Isolation. launch to impact. partly countered by means such a s mentioned in the section on conceptual design. By whatever method, heat rejection has to overcome the following barr iers: the crush-up material, the entry shell, and possibly the steriiization canister a s shown in Figure 1.

This i s essentially from pre- The thermal isolation may be

Ambient Temperature. The temperature inside the capsule may be from -10 to perhaps 70°C. Outside, on Mars, assumed variations a r e -60 to t27 'C.

Zero G or Low-G Operation. primarily with respect to liquid flow conditions and mechanical movements.

This is a constraint

Nuclear Radiation Dose Rate. This is a s attributed to the power supply and may not exceed Neutrons and gammas of 2 x 10-1 m rads/hour a t 1 meter from the source. Earth background 1 eve1 .

It i s approximately 10 times

Magnetic Field Due to Power Supply. Not to ex- ceed 5 x 10-9 wb/m2 (5 gamma) at 1 meter from capsule . Volume Weight and Shape. limit i s taken a s 2 ft3; this includes all shape factors. made up of several convenient shapes that f i t into the payload module cavities.

Disposal of Waste Products. for disposal of waste products externally to the

An arbitrary volurpe

The total power generator volume may be

There is no provision

tently over the 12-day cruise period. Transmitters, both direct and relay, a r e the major power users accepting raw unregulated power. All other users Attitude. The power supply must operate in any require regulated power and condition it further to

capsule en route and none allowed after impact.

attitude. their particular needs. However, the power levels shown include allowance for regulator dissipation.

The total energy requirement i s only 749 watt-hours of which approximately two thirds a r e used after impact. only 150 watts, except for a few minutes a t pre- separation checkout. power ratio i s only 5 hours; this is an important index factor for many power generators.

The peak power levels a r e

The overall energy to peak-

C. Principal Constraints

In addition to the foregoing, there a re several constraints applying only to electrically recharge- able generators such as batteries. These are: Storage ability in the discharged and charged states, ability to be recharged by bus/orbiter/or external supply, charged stand-time and others. Several alternatives exist in these areas. plete discussion of these i s impossible in the space allowed. given candidate generators.

A com-

They a re discussed where applicable to

Table 1 shows the constraints with regard to the mission phase. As can be seen, not all the constraints apply all the time. straints (heat sterilization, shock, and thermal isolation) have been taken as the major criteria for acceptance or rejection.

The first three con- A l i s t of important constraints and pertinent comments follows:

Heat Sterilization. and complete sealed system a t +145"C for three

Performed on all components The renaair;np

61

constraints a r e important and must be complied with, but occupy a relatively minor role in the selection. The evaluation summary following is treated from this point of view.

111, Evaluation of Candidate Power Generators

The general categories of power generators are: 1) solar energy, 2) nuclear energy, and 3) chemical energy. They will be examined, each in turn, and the results summarized.

A. Solar Energy

Solar energy sources investigated consist of silicon solar cells, thermionics, plus concentrator, and heat'engines plus concentrator. the large (typically 9 centrators needed by the thermionics and heat engines7~ 8 in order to convert the low solar inten- sity at Mars, both thermionics and heat engines were eliminated from further consideration for flight-capsule.use. The use of silicon solar cells poses many problems. For example, installation of solar cells on the outer skin of the vehicle is not compatible with heat- shield requirements . Furthermore, attitude control for communications with the flight spacecraft and Earth conflicts with the sun seeking needs of the solar cells most of the time. intensity on Mars at a selected landing latitude of 30 degrees south. a rea correspond to an arrival period of from spring to early summer (14 November 1971 to 1 2 January 1 9 7 2 in that hemisphere).

Even if we assume an ability to track the sun at all times during the day, a large panel a rea of solar cells would still be necessary. shows calculated sizes of four solar-cell types and an approximation using the LASCA method pro- posed by Ray and Winicur9. considerations such as breakage and dust storms, etc. a battery would still be necessary since at best only 13.8 hours of the Martian day would be in sunlight. The constraints of volume and shape completely rule out any possibility of carrying the solar-cell erection mechanism inside the capsule while the high landing shock prohibits outside ap- pendages. Another possibility is that the capsule may land in a shaded site. Taking all constraints and likely possibilities into consideration, the use of solar cells for power appear impractical.

Because of t o 10-foot diameter) con-

Figure 4 shows the calculated solar

The intensities in the shaded

Table 2

Quite apart from

B. Nuclear Energy

Preliminary studies eliminated nuclear r e - actors from consideration because of nuclear radiation and the problem of thermal isolation which is evident from Figure 1. Radioisotope, thermoelectric generators (RTG's) on the other hand a r e worthy of consideration. Thermionic generators a r e not covered here because the state of the a r t is not considered to be far enough ad- vanced. For long landed missions of 30 days or

more after landing, an RTG/battery combination may be the only feasible low-weight power gen- erator. As a f i rs t approximation, a brief calcula- tion (see Appendix A) of required RTG electrical power gives 26 watts. An earl ier profile resulted in 30 watts required and this was used a s the basis for a parametric study of RTG weight versus cold- junction temperature. Figure 5 shows the results for chosen parameters and the use of plutonium (Pu 238) as a fuel in the form of (PuO2) plutonium dioxide.

1. The Radioisotope Thermoelectric Generator (RTG). Before applying the principal constraints to an examination of RTG character - istics, it i s necessary to make assumptions of generator geometry and to state other criteria. The process is to some extent iterative because a number of the assumptions originally stem from the constraints as will become apparent.

2. RTG Design Criteria, Assumptions, and Analysis. Much of the design analysis, particular- ly the parametric design section, is theoretical only and i s not meant to convey a full hardware design. A practical RTG of the sort described is likely to be 30 to 50 percent heavier in weight and possibly in size (according to Martin Company sources)l2. terms of redundancy and probable larger allow- ances for shielding and thermal insulation.

The difference is accounted for in

The analysis is performed by considering the following design areas: a ) fuel selection, b) neu- trons gamma radiation shielding, c ) selection of thermo elements, d) heat rejection and e ) safety analysis.

Generator Geometry: This is assumed to be a nearly rectangular block with the heat sink shaped to f i t the inside of the flight-capsule frame. The thermoelements a r e all arranged in a flat planar array, the hot-junction plane being parallel to the cold-junction plane.

sists of several double walled capsules containing the fuel and embedded in a block of Beryllium or graphite. The fuel capsules allow for 100 percent void volume. All surfaces, except that in contact with the thermo elements, a r e heavily insulated to ensure heat flow exclusively towards the thermo elements.

Radiation Criteria: Isotopes producing significant penetrating radiation a re to be avoided because of the extra shielding required and distri- bution of the heat load. Thus, pure gamma emitters a r e eliminated leaving decay modes of alpha or beta particles with alpha preferred. Maximum allowable doses for instrumentation in the vehicle a r e lo7 rads of gamma, l o 4 rads of fast neutrons and l o 4 rads of solar protonslo. Since the radioisotope is loaded sometime prior to launch, the effective maximurn allowable dose rate for instruments is based on approximately 1 year of life and is 107/8, 760 hours o r 1 . 1 4 x l o 3 rads/hour of gamma, 1.14 rads/hour of fast

Fuel-Block Geometry: The fuel block con-

62

neutrons and solar protons. ment instruments would have to individually shielded or removed from near the source. vicinity of the dose rate mentioned i s immediately around the RTG outer case. the dose rates would be reduced by at least s ix orders of magnitude using sufficient extra shield- ing. This governs the design, not the instrument tolerance. Half-life of the radioisotope should be 5 to 25 years for beta emitters with 1 mev particle energy and 5 to 100 years for alpha emi t te rs l l .

to survive reentry velocity, shock, aerodynamic heating, landing impacts on Earth or Mars or any credible accident.

flight capsule for radiator use exclusively, a r e to be charged to radiator weight.

element connection, no redundancy i s quantity i s allowed.

Heat Removal: all heat not transmitted through the crush-up i s transmitted through one or more heat pipes which a re disconnected only i f the crush-up i s removed.

Radiation measure-

The

For handling purposes,

Safety Assumptions: the fuel is to be contained

Radiator Weight: additions of weight to the

Redundancy: except for methods of thermo

Fuel Selection

Pu 238 was the fuel selected, based primarily on its low gamma radiation. isotope is 86 years. sidered to be a problem; however, the price i s more than $600 per thermal watt. Tables 3 and 4 review the isotope fuels and give data on selected isotopes, while Figure 6 shows isotope plus con- tainer weights for a 1000 thermal watt source after 15 months. a double-wall container and 100 percent gas -volume allowance. Typical container materials a r e Hastelloy X, Haynes 25 and TZ Molybdenum. It provides for intact re-entry andimpact speeds up to 300 f t / s ec l l without rupture. The container also acts a s a partial shielding for gamma rays. The shielding criterion leads to the heaviest capsule.

low shielding requirements, but its relatively short half-life of 2 .7 years mitigates against it; C m 244 would probably be the next actual choice. However, it needs relatively heavy shielding as compared with Pu 238 because of its hard gamma radiation. ground radiation i s assumed a s the maximum 1 meter from the source, the reasons for a choice of Pu 238 become evident.

The half-life of the Its availability i s not con-

The weights shown a re based on using

Pm 147 i s favored after Pu 238 because of its

If a value of 10 times the natural back-

Neutrons Gamma Radiation Shielding

Plutonium and curium produce neutrons and gammas, respectively. Thirty cm thick-water shielding is required to bring the neutron radiation level from a 1000 w(t) source down to rad/hour at 1 meter. This level cannot be met im- mediately onboard near the RTG. be found by a ) using a fuel having a reduced neutron flux, and b) use of local shadow shielding. A more

A solution may

pure Pu02 (without 017 and 0 1 8 ) would be almost free of alpha-neutron reactions responsible for most of the neutrons produced. Latest tes t results from Mound Laboratories show that a neutron count of 8200 to 9500 n/sec per w(t) has been obtainedl3. This i s an order of magnitude lower than previously considered figures.

reduced by a factor of 10 for each 18 cm layer of polyethylene. Scattering of incident fast neutrons is relatively minor and side shields can be com- paratively thin. For the shape and locations shown in Figure 2, any kind of extra shielding would pose a problem and tend to reduce the maximum RTG size and electrical-power output.

For shadow shielding, the radiation dose is

Selection of Thermo Elements

The thermo elements selected a re lead tellu- ride (Pb-Te) or silicon germanium (Si-Ge) with Si-Ge the preference for such reasons a s reliability, ruggedness, lack of dependence on pressure seals, and maximum hot junction temperature. compares the performance of a number of thermo elements. For the parametric design, Si-Ge is the thermo element.

Table 5

Heat Rejection and Parametric Design

The greatest problem i s that of heat rejection The three bar r ie rs of crush-up in transit.

material, heat shield, and sterilization canister shown in Figure 1 impede the flow of heat from the payload, and since the isotope cannot be shut off, it may overheat. It now becomes necessary to perform a parametric design in order to obtain bas ic heat - trans fer information. yields the level of operation possible without heat pipes, which a r e then discussed.

This information

Design of RTG. In an RTG, heat i s produced by decay of an encapsulated radioisotope. Most of this heat flows through the thermo elements which transfer approximately 5 percent into dc electrical energy. The remaining 95 percent has to be dis- sipated. The component parts are: isotopic fuel and container, thermo elements, thermal insula- tion, and a radiator or heat sink. The thermo element used i s Silicon Germanium, the thermal insulation assumed i s MIN-K, and the fuel is Pu 238 contained a s described earlier. No redundancy is allowed for, but the weight of shielding is taken a s producing a neutron gamma dose rate of approxi- mately 5 x 10-2 millirads/hour 1 meter from the source. millirads /hour immediately around the RTG case. As noted, special extra shadow shielding would be necessary for sensitive measuring instruments contained onboard to bring this level down to the 10-1 millirads/hour desired. A suggested design solution i s to extend the measuring instrument out on a boom to lower the radiation level.

This would produce approximately 2

Heat Supplied, Q. The total heat which must be supplied by the fuel is

63

where

QTE = heat supplied to junctions

QJ = heat leaked through insulation

Now QREQ(e)

VTE i s the a r ray efficiency and i s

QTE = - where qTE

Function 4 includes expressions for specific elec- tron heat dh , figure of meri t 2, area ratios

dT

, and assumptions of element design. Am

AP -

The insulation heat loss QJ i s a function of ma- ter ial conductivity(K), path length ( X ) and cross sectional a rea ( A ) a s well a s temperature dif- ferencesTH,T,. O r

QI = f [K, X, A (TH - T,)1

It i s also related to QTEand may be expressed as 9 = f ( g ) QTE

Thus Equation (1) becomes

where TH and Tc a r e the hot and cold junction tem- peratures, respectively. The required heat input i s plotted versus cold-junction temperature T, for QREQ(e) This graph is used in conjunction with Figure 5.

= 30 watts ( e ) and TH = 825°C in Figure 7.

RTG Weight.

( 3 ) FTOTAL = W~~~ + WFUEL BLOCK + ~ ' T E + WINS + FRAD

The total RTG weight i s then

When the equations a re derived in terms of the hot and cold junction temperatures TH and T,, the following results:

where f(r) = decay rate function

and pF = fuel power density

r2 [f x'QTOTALI F~~~~ BLOCK =

PF

( 3 . 1)

(3.2)

(QTOTAL) WINS =

(TH - T,j1/3

( 3 . 3 )

(3.4)

( 3 . 5)

The results a r e plotted in Figure 5 for TH = 825°C and using T, a s a variable. The top line represents the total generator weight and the addition of al l the increments. The weights of the fuel, the fuel block, and thermo electric a r r ay all increase with increasing T C , whereas the thermal insula- tion and radiator decrease. i s not optimum at 100°C but rather at near 300°C. However, decrease in radiator weight i s initially so sharp that it overcomes the other increasing component weights reaching a generator minimum weight at approximately 100°C. By far, the most important weight influence i s that of the fuel con- tainer, and fuel block. The results may be scaled fairly accurately for electrical output power in the range of 20 to 40 watts.

The radiator itself

Heat Rejection

It i s assumed that the outside skin of the landed capsule after removal of the crush-up layers, forms the radiator surface. From the parametric design of Figure 5, the best T, is 100°C. by the limitations on the electronic equipment and battery. cold junction T,and the radiator TR is low, TR will approach T, and be approximately isothermal. a battery will have to be used and will attain the same temperature a s TR, the upper limit on TR i s then 70°C for a NiCd battery. The equations for heat flow a re by stead unidirectional conduc - tion a s given by Fourier1<

The radiator surface temperature i s set

Assuming that the impedance between the

As

dT Q = - K A - dx AT KA

O r in terms of Watts Q =-KA- = - - AT

where L is the total path length in meters, K = the average conductivity Watts/"C x m and A = unit a rea m2. lent to thermal conductance Ki . in approximation is:

Ax L

The factor KA/L for any layer is equiva- Overall heat flow

the subscripts refer to surfaces shown in Figure 1.

of only 10 cm and an average crush-up conductivity of 0. 4 watt / O C -m produces the temperature distribution showninFigure 8. The available radiator a rea out- sidethe crushupis estimatedat 12 mz. Inpractice, however, the conductivity of the crushup material is

Assuming an overall crush-up path thickness

64

approximately an order of magnitude less (e. g . , .-$ 0. 04 watts/OC-m). The material i s a fiber-

glass phenolic honeycomb with a polyurethane foam filler. In addition, the design capsule path length i s an average of 3 times greater than that assumed. The combination of these two factors make i t quite impossible to use an RTG having an electrical output of more than a few watts (1 to 3) without using a heat pipe or similar device.

Heat Pipes. The heat pipe, a s briefly de- scribed here i s due to Grover et a l (Reference 2, 3 , 4, 5, 6) and i s illustrated in Figure 9. Ac- cording to its inventors and other investigators it has a thermal conductance higher than any known material and conducts heat with essentially no temperature drop. Quoted thermal conductivities a r e in excess of lo3 cal/sec-cm°K for a model using liquid sodium. The operation i s a s follows: As heat is supplied to the evaporator section, liquid in the material evaporates and flows in the core towards the region of low pressure which also i s the cold end. The liquid i s condensed here and returns by capillary action to the hot enc: after giving up i ts heat. why flexible pipes cannot be made, although none a r e referred to in the literature. Kemme5 of Los Alamos Scientific Laboratory have designed a model for flight test, but a t the time of this writingI5 they had not succeeded in getting it flown.

There i s no apparent reason

Deverall and

Conclusion on RTG Heat Dissipation, The heat pipe appears to be a most promising solution to a t least that part of the problem regarding the elimination of heat from the flight capsule. mitting i t through the entry shell and sterilization canister i s another matter. Solutions to these problems a r e believed to be possible and a r e currently being investigated a t Avco.

In short, the advent of the heat pipe reduces a major part of the heat dissipation problem but does not solve it. On this account, the practi- cality of an RTG in a hard lander remains very much in doubt.

Trans-

Safety Analysis

Assuming solution of the thermal dissipation problem, the safety problem with respect to Pu 238 resolves itself into the following categories: Handling difficulties on the ground (ingesting, shielding, etc. ); Accidental impact on Earth; Abort in the Earthatmosphere;and Failure of crushable material to protect capsule on Mars.

In addition, there i s the question of acciden- tally achieving criticality. Table 6 is a chart of the possible accidents that can happen a t any time fromproduction of the isotope until the the end of the mission. Only the hazards that affect the con- ceptual design a r e noted.

As far a s criticality of the fuel is concerned, the minimum for criticality estimatedat Avco is 1Oto 15kgahichisfarabovethatrequiredfor a given converter. Thus, assumingnormal shipping

precautions, criticality does not affect the con- ceptual design. problem areas exist that stil l have to be solved.

It i s evident f rom Table 6 that

Conclusions on RTG

1.

2. Shielding and safety problems have, a t

The thermal dissipation problem has yet to be solved.

present no satisfactory solutions compatible with landing capsule design.

3. The ability of the thermoelectric a r r ay to withstand high -g shock in any plane i s in consider- able doubt.

In view of the preceding and the less than a 2 to 1 weight advantage of an RTG battery combination a s compared with a battery alone, use of an RTG on a hard lander for a short-time (1 day) landed mission i s unlikely.

IV. CHEMICAL ENERGY

Introduction to Fuel Cells and Batteries

The fuel-cell systems investigated in this study a r e HZ02 and a lithium-chlorine system. Only closed systems a r e discussed because no waste product may be exhausted into the planet atmosphere.

a r e relatively well known and a r e treated first. Figure 10 shows the schematic for the Pra t t &

The essentials of the system, along with the com- plexity a r e well illustrated. cell system i s not the only type under current de- velopment. Some of the others a r e typified by the following categories : a) solid polymer electrolyte, b) capillary membrane, c ) carbon electrode, d) electrolytic regenerative.

H2-02 Systems.The hydrogen-oxygen systems

Whitney Aircraft modified Bacon Cell sys tem 16 .

The modified Bacon

Except for the last named, they all use the same basic system components and vary largely in cell design. pressures vary considerably. of each type follows.

duced by General Electric Company 17, 18, a solid electrolyte consisting of a sulfonated styrene poly- mer (approximately 10 mils thick) i s used. membrane permits hydrogen ions to migrate from one electrode to the other.

Thin films of platinum catalyst applied to both sides of the electrolyte act a s electrodes, support ionization of hydrogen on the anode side, and re- duction of oxygen a t the cathode side. titanium screen embedded into the platinum elec- trode reduces internal resistance to current flo\v and adds structural strength. away by wicks. cell system as used on the Gemini flights. Its major drawback is that it cannot be heat steri- lized19 due to the use of the solid polymer elec- trolyte.

Operating temperatures and A brief description

Solid Polvmer Electrolyte Fuel Cells. As pro-

This

A thin

Water i s carr ied This i s the familiar H2-Oz fuel

65

Capillary - hlemb r an e Cell - Alli s - Chalme r s . Differences between this type and the modified Bacon (PtWA) models center on the operating tem- perature, the method of holding the electrolyte, and removal of the water product. In the scheme used by Allis-Chalmers Companyz0, a vapor pressure differential i s used to exhaust the water vapor to a vacuum: or the water vapor may be con- densed and stored. Start-up and operating tem- peratures may be in the order of -20°C ( - l O ° F ) , although performance steadily improves a s the temperature i s increased to a typical 21 operating point of 90°C. These temperatures should be con- trasted with the t200 to 250°C operating tempera- ture of the modified Bacon Cells. The electrolyte i s retained in an asbestos capillary membrane. There a re no obvious physical reasons why the system cannot be heat sterilized, although a prob- able redesign of structure and auxiliaries a r e in- dicated. A water condenser would have to be used on Mars since it cannot be dumped into the atmos- phere.

Carbon Electrode - Union Carbide. Union Carbide makes a fuel-cell using carbon electrodes and a working temperature in the range of 50 to 55°C. coating with P. T. F. E.24An extended thermal soak at 145°C may cause some trouble here, Apparently i t has not as yet been tested at this temperature and there a re no reported plans to do so.

The electrodes a re wet-proofed by the

Electrolytic Regenerative Fuel Cell. A con- siderable amount of work has been done by Electro- Optical Systems, Inc. (EOS) on such a cell system. According to SoltisZ2 of NASA Lewis Research Center, the system has survived a heat sterili- zation soak at t150"C for 67 hours. operates basically in a manner similar to a battery. the system i s recharged by electrolyzing the water back into its hydrogen and oxygen com- ponents. 75°C after the 150°C thermal soak. Claimed capability for a six-cell system is given a s 17 to 20 watt-hours per pound for a 60-minute discharge. Cells have also been operated over a range of tem- peratures from 70 to 150°C. operation is not discussed here. cess to improve reliability and cycle life.

The system

After discharge and production of water,

Figure 11 shows cell performance at

Low temperature Work is in pro-

The Heat Sterilization Constraint. It i s appar- ent that a t least one of the H2-02 types (the re- generative type) can be heat sterilized. Probably both the Allis-Chalmers and P f W systems could also be designed to be heat sterilized, although it has not yet been done. apparently unsolvable engineering problem here. Sterilization, however, includes the fuel tanks making cryogenic storage impractical. There- fore, gaseous storage is necessary and the weights and volumes are greatly increased. A n exa iq l e of this is given in Table 7.

In any case, there is no

The Heat Dissipation Constraint. Since a fuel cell need not be activated until after removal of the sterilization canister, i ts prospects for heat dissipation are better than for the RTG. An approximation of the heat to be dissipated from H2-02 fuel cell i s obtained as follows:

Assuming 100 percent conversion of hydrogen and oxygen into water, and using Faraday's law for the weight of water produced in an electrochemical reaction, the total heat rejection Qin Btu/hr i s de- rived (see Appendix B) as:

Q = 3413 x P x 1 - 11 Btu/hr [ 0.675 Y (5)

where P = total power in kw at the cell voltage Equation (5) i s plotted in Figure 12. the minimum heat dissipation for the given cell voltages and not the actual. conjunction with a polarization curve. parasitic power operation have to be added to the resulting heat load of Btu/hr. An example using the latest informationz1 from Allis-Chalmers will serve to illustrate. According to line A in Figure 13, at a power density of 160 watts/ft ported performance at 88°C operation i s 0. 88 volts/cell. By the equation plotted in Figure 12, the heat rejection i s approximately 2300 Btu/hr. As much a s another 30 percent may be added to allow for operation of a water condenser, etc. producing 3000 Btu/hr per kilowatt electrical out- put. responds to the power profile) this is 450 Btu/hr or approximately 133 watts. Using the heat-flow analysis in the earl ier section on RTG's this i s only 73 watts above what can be dissipated through the crush-up material without the aid of a heat pipe. Operation at a lower power density to reduce the heat load to 60 watts would be the answer here and seems quite reasonable.

Dissipation of 60 watts of heat to space'after removal of the sterilization canister in flight appears quite feasible. On the ground, i t may be necessary to test the fuel-cell only at reduced power output levels for checkout, but no major objection exists here. Therefore, H2-OZ primary or regenerative (secondary)fuel cells having primary mode characteristics similar to the Allis-Chalmers fuel cell would operate within th, tieat-dissipation constraint.

It represents

This figure i s used in Losses for

2 the re-

At 150 watts electrical output (which cor-

Fuel Consumption. The fuel specific con- sumption expressed in terms of the operatingcell voltage is :

0.0829 lb,'kw-hr

v S.C. H2 =

and

0.663 lb/ka-hr

v s.e. o2 -

For an operating voltage of 0. 8 volts/cell, S. C. H2-02 0. 9 lb/kw-hr. dizer storage i s at cryogenic temperature, but estimated boil-off losses in the 9 or 10 months to

Normal fuel and oxi-

66

reach Mars would be prohibitive; therefore, gas- eous storage is required. times, the weight of the fuel and tanks i s not very large. However, for long times, it soon becomes impractical a s shown in Table 7 .

For low power and

The High-g Shock Constraint. All manufac- tu re rs consulted generally agreed that shock and vibration isolaters mus t be used., and a high impact shock of > 1000 g would be a major problem.

The Zero-g Constraint. This did not s e e m to be of much concern, possibly because of the lack of applicable data or even theories on the subject.

(or lack of them) remain an a r e a of doubt for any process involving liquid flow unless proven other- wise. This includes fuel-cell systems.

In the opinion of the wri ter , zero-g effects

Failure Modes. F r o m the reliability point of view, observations on likely failure a r e given. All of these do not necessar i ly apply to all H 2 - 0 2 systems discussed. For example, Platner of Allis-Chalmers 2 3 reports difficulties such as sticking of the purge valve and resonance of the reactant p ressure regulators. modes of failure were reported by him and possibly do not apply. simply by looking a t system diagrams.

None of the other

Many failure modes can be inferred

The failure modes l isted a r e a s follows:

a )

b )

c ) d)

pres sur e. e )

or sticking. f ) Fai lure of electrical heaters. g)

manifolds. h)

vibration. i) j ) k ) Water flooding.

Since a completely redundant fuel-cell sys tem

In view of

Failures of electrodes by excess electrode

Fai lure of electrolyte holder or separator

Fai lure of control solenoid valves. Leakage and loss of reference nitrogen

Failure of regulators by metal "freezing"

porosity causing "drowning" of the electrode.

membrane allowing possible mixing of gases.

Shorting of cells through parallel fuel-feed

Cracking of electrodes under shock o r

Tubing and other plumbing failures. Fan seizing or pump failures.

could not f i t in the flight capsule, high reliability would be required for a single system. the preceding, a good score in each is mandatory and has not yet been demonstrated, at l eas t in the open l i tera ture .

Conclusions on Hz-02 Systems

1.

2.

The solid polymer electrolyte system i s eliminated because of inability to be heat steril ized.

Heat steril ization of a t l eas t one of the current systems has been accoiiiplished ami i s ex- pected to be possible for a t l eas t one other with 3ome redesign.

Impact shock is a major problem area. 3 ,

4.

5.

The high weight and volume of gas storage

Improvements in cell performance have res t r i c t s the miss ion time.

lessened the heat dissipation problem to the point where this i s no longer an obstacle for low output power levels of 150 watts o r l e s s .

Insufficient data a r e available to demon- s t ra te and project reliability of cur ren t systems under zero-g, low-g operation or after high impact shock.

6 .

On the basis of the preceding, Hz-02 fuel-cell systems a r e available for u s e on the lander only if sufficient development is performed in t ime for use. They a r e not eliminated, but a r e not recommended as first or second choice.

Lithium-Chlorine Fuel Cell System. The Li-C1 fuel-cell system i s a relatively new addition to the fuel-cell family. he re i s that of Werth, Weaver, and KennedyZ5, 26 of G. M. Defense Research Labs. , Santa Barbara , California. a r e the same as for the H2-02 fuel cells, the re a r e enough important differences, particularly i n the mat ter of heat dissipation, to warrant an interes t .

The approach discussed

Although the basic system elements

System Description. The reactions at the anode electrolyte interface a r e

At the cathode electrolyte interface the reactions a r e

CI2 + 2 F - 2c1- .

The total reaction is:

at 62OU C 2 L i + C12 -

i n LiCl 2L iC l .

A schematic of the cell construction is shown in Figure 14. gaseous chlorine and the electrolyte i s fused lithium chloride (mp 613°C). Since the reaction product i s a lso lithium chloride, the electrolyte composition does not change with depth of discharge. Chlorine can be s tored a s a liquid at 100" F under 10 atmospheres of p ressure o r cryogenically. Heat steril ization at t145"C (293" F) does not cause into1 e r able tank pres sur e s . 60 to 80 percent with power densities on the o of > 5000 watts/ftZ. magnitude higher than the b e s t H t - 0 2 Riel cell. The total system weight of a i Obi -*\att si-stc111 lor 1 kw/hr energy would be only 7. pared with at l eas t doiiblc that fn7- projected H z - 0 2 fur l cell. However, a s i i c ~ I c d , thc present s ) steiil statu> is rei-> Iinrch that of a Id>- orator) tes t sct-up and not at a prototrpc At present, t h c t systciie I S x%htJ l ly primar)-, a r egeiitr at] x-e mn rie 1s conttnipl at t. c?.

The reactants a r e liquid lithium and

Typical discharge efficient-); is quoted a s 1 rani

This I S well ovcr an ort t ,>r of

p i s i n c i s a5 c o i t i -

It i s necessary to heat up the cell (i. e . , heat the Li-C1 and Li fuel) in order to s tar t it. ever, the exothermic reaction of about 19.7 kcal/mole or approximately 475 cal/gm of formed .Li-C1 a t 80 percent efficiency will easily keep it going. just to keep the reaction going, the following ad- vantages a r e apparent:

How-

Since much of the excess heat i s required

1. For long (e. g . , 12 days) periods of very low power, excess power levels must be used just to keep the cell warm.

i s no likely heat dissipation problem i f used in the conceptual lander design.

levels (as per (1)) tradeoffs a r e limited to periods of low power. At high power levels, some of the heat must be dissipated or the operating tempera- ture will rise.

2. At power levels of 150 to 200 watts, there

3. More insulation versus excess power

The High Impact Shock Constraint. High impact shock might cause a cell short circuit by forcing liquid lithium into contact with the chlorine or out of the cell. components seem likely and would necessitate shock mounting at the very least. state-of-the-art this remains a nebulous area until at least prototype models a r e built.

Cracking and damage of cell

At the present

Heat Sterilization. Heat sterilization should not be a problem here except with possibly the con- trol components where a design solution seems probable.

Zero-g (or upside down) Operation. This has not been demonstrated and will be investigated.

Cold Environments. Environments such as -10°C should not bother the system once it i s started. can insulate it. currently available.

This is mostly a matter of how well we Adequate insulation materials a r e

General Comments. A dc to dc converter would be desirable because individual cell output voltage i s about 3. 0 volts dc. designs indicate that few cells would be used and parallel operation is preferred for minimum com- plexity.

values for a nominal 50 watt and 100 watt Li-C1 system. It should be noted that redundant cells a r e used.

Figure 15 shows a comparison of RTG and Li-C1 weights versus operation time. Surprisingly enough, the Li-C1 system i s lighter than the RTG for operating times of less than 500 hours. If it is recalled that the RTG has neutron radiation shield- ing problems as well, and even if it had n6 heat dissipation problem, the Li-C1 system appears to be better.

Present conceptual

Table 8 shows calculated weights and volume

Conclusions on Li-C1 System. The Li-C1 systerx is the most promising and is a strong com- petitor with R'TG's for missions rip to 30 days.

Because of its short development history, however, it i s not likely to be available for use on the lander unless development along the lines indicated i s speeded up.

Chemical Energy - Batteries

Batteries fall into one of two categories, pr i- mary or secondary. For our purposes, a primary is defined as a "one-shot" or at best a 5-cycle battery. and discharged five times at the most and produce a reasonable output. Secondary batteries a r e r e - chargeable with a variable cycle life. Because of the long time it takes to reach Mars, the possibility of using primary-type batteries ar ises. batteries considered for use were silver-zinc (on an energy basis) and thermal (fused salt) types.

This means that it can only be charged

The

Primary Batteries

Thermal Batteries. Thermal batteries a r e highly cla s sified and performance characteristics cannot be discussed. However, long Avco experience with these batteries leads to the conclusion that they can only be used for power outputs of a few minutes duration. batteries and cannot be completely checked out electrically. some difficulty due to deterioration of the electrical igniter matches is likely particularly under pro- longed heating. sideration except to help out on power peaks.

These a re pure one-shot primary

It has never been heat sterilized and

This type i s eliminated from con-

Silver-Zinc Batteries - Primary. The pri- mary silver-zinc battery is simply a cell with less (and usually thinner) layers of separator material. The cells a r e dry charged and activated by addition of an electrolyte (KOH) to the cell. Activation may be performed manually or automatically. case of the lander, an automatically activated sys- tem would be mandatory; however, such a system has not been demonstrated to be heat sterilizable. A schematic of an automatically activated silver- zinc battery is shown in Figure 16. An objection to the use of this battery i s similar to that for the thermal types. It cannot be fully checked out be- fore heat sterilization. similar to those stated later for secondary silver- zincs.

In the

Other objections a re

Ammonia Cell - Primary. The liquid ammonia cell i s peculiar in that it likes to operate at low temperatures. For example, reported typical discharge temperatures have been as low as -63"C28. In view of the expected cold tempera- ture on Mars, this would seem to be an advantage. If this performance is added to a projected specific energy of 200 watt-hours/pound, then it i s very attractive indeed. A discussion with H. R. Smith and W. F. Meyers29 of Livingston Electronic Corporation (its chief developers) brought forth the opinion that the cell would probably not withstand heat sterilization a t +145"C. Livingston Electronics points up the developmental

A reportz8 from

68

nature of the battery. animonia cells, but generally i n the primarily very

Others30 have worked on only deliver a few cycles. development obviously needs to be done on this cell

A fair amount of

short discharge t ime field. type-

Secondary Batteries - P r i m a r y Nickel-Cadmium Cells . Hermetically sealed nickel cadmium cells have been evaluated by Avco and found to be heat sterilizable without large de-

as follows: a ) silver-zinc, silver cadmium and terioration in capacity and no apparent structural nickel-cadmium. The reasons for considering the damage. The cells were few in number ( l o ) , but silver-zinc and silver-cadmium cells are:

The secondary batteries considered here a r e

the finding i s significant in that it indicated that with sufficient selectivity, acceptable specific

1. They a r e the highest specific energy energy values a r e obtainable. Several cells con- sistently achieved outputs of greater than 12-watt

2. hours per pound at the 5-hour discharge ra te and one or two produced 14 watt-hours per pound. Figure 18 shows the range of performance for the smaller D cells. It should be noted that there was no effort to select the cells out of a production

no deterioration after heat sterilization, taken together with the fact that additional rounds of heat soaks did not produce further deterioration, signi- f ies that the heat sterilization process may be used as a selection device.

formed to date. to one volt for a number of cycles. Apart from the teniporary rejuvenation effect noted in cycle 64, and caused by deep discharging pr ior to this, there is no great variation. Perforniancc is about the same a s one would expect for a nickel-cadmium cell that had not been heat sterilized. In addition to heat sterilization, testing has bcen performed in the a reas of shock, vibration, high and Jow dis-

storage. Cells were heat soaked and discharged at t2000 with little apparent loss in energy.

At the other end of the scale, cells were discharged a t a 0 ° F (-17. 8°C) soak temperature, a typical example of

ance was much closer to the 7 5 ° F levels on an

500 and 2000 g, performed in 3 axes on charged cells.

damage of any kind was noted.

damage or deterioration.

previously, a number of cells urderwent a cold- soak tes t at minus 100" F for a month. The c r l f s were fully charged before the cold soak, an:! d i s -

Organic Separator Type. Vented Of charged imnicdiately upon warining up. Again n o

batteries (see Figure 17).

formance under varied conditions. A great deal is known about their p e r -

Unfortunately, the process of heat steril iza- tion changes the performance of so-called standard

be made. Since little work has been either done o r reported on the silver-cadmium cell (in the areaof heat sterilization), they will not be treated further. It should suffice to say that they possess much Of

the built-in disabilities of the silver-zinc cell pa r - ticularly with regard to deterioration of materials.

lized successfully a t Avco3l> 32 and will be dis- cussed further. Thus, the examination is reduced to s i lver- i inc and nickel-cadmium batteries.

The two types of

construction Many changes have to lot. The fact that several cells showed little or

Many charge-discharge cycles have been per- Figure 19 shows discharge times The nickel-cadmium cell has been heat s t e r i -

Silver-Zinc Batteries. silver-zinc battery cells discussed here a r e classified by separator type.

1. Inorganic separator 2. Organic separator charge temperatures, long timc charging, and

The importance of the separator c1assificatio.n centers on i ts part in the ability of the cell to sur- vive heat sterilization.

These are:

is shown in Figure 20.

Inorganic Type* This approach \%,hi& is shown in Figure 21. The 40" F perform- has been taken by Berger and others, 33, 34 apparently quite Successfully to date* According a\Terage ,-ell. TEe acceleration shock levels were to all indications, the cell looks good, but cell structural problenis remain. One of the diffi- culties is in the case material where polypropylenc tends to soften; Teflon is difficult to join and s o

forth. New such as polysulfone be Minuteman specification levels also produced no used, but to date, joining difficulties still abound. The cells a r e not hermetically sealed and require a sealed external case. Avco has not as yet evaluated these cells in the laboratory.

Base t imes were 5 and 0 . 6 milliseconds for triangular shape wave. No

Vibration a t

In addition to the tests nientioned briefly

this were by AvcO person- danlage or t[cter-ioratjon v,7as noted. This r~-lodc of riel. The results a r e described in an A ~ C O report35. storage is not recomrT1ended, b7.1t CR31 Briefly, all cells failed to deliver a reasonable and useful output after heat sterilization. catastrophically. It was concluded that cells using t3000 led to their destr,lctioll intcrllal silort- the construction tested a r e not capable of being he a t s t e rili z t: d s-,-ithor it s uf f e ring una cc ept ab1 e r i a ~ ~ ~ a g e . It transpired 37 that the cel ls tested were really pr imary in nature and could

be done if rc,quirecl. Some failed An attcrnpt t o discharge several cells at.

circuiting. This underlines the view held by the writer that heat sterilization should only bc pc>i-- i o rnic d on fit11 y - di s clia r g ti d p r IC r ably s l r o rS - circ:uit.ed (.(*I1 s to rcmcx-t: thc possi?Jilit>- of t ri-iodr* of failitre.

69

Testing on the remaining cells continues to date with no particular difficulty noted. uation program is also continuing on large (20 amp-hr) size cells.

An eval-

Magnetic Field Generation, A prime require- ment i s that of a minimum magnetic field genera- tion. show directional magnetic polarization of the individual cells which may depend on the state-of- charge of the battery or cell. reversing the physical positions of the battery cells or groups and balancing out the direction of current loop paths a s for example, shown in Figure 22, it should be possible to reduce the magnetic field to a tolerable level. The results in Table 9 a r e reported in terms of the equivalent magnetic dipole placed at the center of the battery or cell case giving a magnetic field pattern similar to that of the battery. dipole .are pole centimeters with one pole centi- meter equivalent to 0.001 ampere turn/meter2. The sign shows direction, a plus sign indicating a north seeking end. Four absolute dipofes produce a field of 1. 12 gamma a t 3 feet. On the basis of dipole per A-H, it would appear that silver-cads a r e best, the values for nickel cads being high as expected. *

Preliminary datal3 exhibited in Table 9

By successively

The units of the

Conclusions on Batteries. considered, the sealed nickel-cadmium cell appears the most likely candidate for f i rs t choice on the basis of heat-sterilizability, ruggedness, and simplicity of charging and operation. also one of the heaviest batteries and will tend to limit the mission time due to the required battery weight.

Of all the battery types

It i s

A secondary silver-zinc battery (probably of

the inorganic separator type) is second choice. This battery requires hermetic sealing and pro- tection of the vented cells by an inert pressurized atmosphere during heat sterilization. The effect of prolonged zero-g on the performance of liquid electrolyte (wet) cells i s open to question and needs to be explored further. Magnetic-field interfer- ence, particularly for nickel-cadmium cells may be a problem, but an acceptable solution seems to be in sight.

Comparison of the Power Generators - Conclu- sions. Figure 23 shows a summary of the con- clusions on all the power generators considered. The f irs t choice for the conceptual lander is a he rme ti cally - s e a1 e d, nickel - cadmium battery with a secondary silver zinc a s a backup. fuel cell group H2-02 systems (particularly regenerative types) a r e in the running but a r e low on the recommended list because of the sparsity of reliability data, large volume, and poor-shape factor, and a borderline heat transfer difficulty. The lithium-chlorine fuel cell i s a strong con- tender, being attractive from the point of view of heat dissipation, weight, and volume. However, it requires too much development to be in time. For the longer missions~ to come, the Li-C1 fuel - cell development merits considerable effort and i s competitive with the RTG.

In spite of improvements in heat transfer technology such as the heat pipe, the transfer of high heat quantities out of the spacecraft from pre- launch to landing remains a considerable problem having no satisfactory solution at this time. When added to the safety and contamination problems, particularly with respect to high impact shock, the use of an RTG on a hard lander i s unlikely. very short-time (i. e., a few days) landed missions, it is not worth it and is not recommended.

In the

For

7 0

APPENDIX A APPENDIX B

CALCULATION OF REQUIRED RTG POWER

a W z II

TIME

Derivation of Equation for WR or RTG power je lectr ical)

A represents power available for charging the battery plus allowance for efficiency of con- ver te r and charger ,

Wp = Peak power, watts

WR = RTGpower

W, = Minimum power

AT1 and I T 2 t ime periods

Area A = ( ITR- W, ) IT1

B = ( WR - W, ) .IT1 qcc = ($ - WR) &IT2

Solving for WR, we have

W I T 2 + Wo ATl qcc P WR = -IT1 9cc + IT2

Calculation of Required Power

Calculation of Heat Dissipation for H2-0 2 Fuel Cell -

Thermal efficiency of the cell may b e defined as vTH where

(8-1) Power Developed V X I

'ITH = Chemical Power Equiv. - A H Y WH2 x I

where A H = heat of formation of H20 = 68.4 kcal/g-mole

= 17. 3 kw-hr/lb H2

WH? = 0 . 0829 x 1:bjarnp-hr (%-J\;eig!it of hydr ogcn coiisumid,by F a rad3 :- Is latv).

Thus,

V ___- = 0.675 V (B-2) 'ITH = (17.8 103) (0.0820 10-3)

w-here v.= v d t s /cell .

Substituting

IT2 = 1. 22 hours

Bo = 7 watts

W = 132 watts P

( j 3 - 4 )

F r o m the profile period after impact we have

I T 1 = 22.7 hours (may be repeated n t imes) o r

Substituting in Equation (A-1) for WR we has-e

VR = 16. 9 watts

Allowing a 50 percent d e s i g n rcd~mdanry A gives 1. 5 x 16. 9 or 25. 5 watts electrical.

check of the period before impact shows that t l i c . period chosen requires the iiiaximuim e 1 r . c t 1 - i ~ al power from the RTG.

where v . is the open circuit voitaqc as take 7 at the irit H r c ( tpt . 01

Sohs t i tu+lng in Equation (I3 - 3 j w c have:

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1. Avco RAD, Wilmington, Mass . , FUD-TR-65-29 "ComDarative studies of con- 15. Telecon: Deverall , J. E., of Los Alamos

Scientific Lab. , N.M. on 8 June 1966 Re: "Flight test of heat pipe".

ceptual design and qualification procedures for a mars probe lander ' ' Volume IV,. Technical Analysis, Book 2 and 3, NASA Contract NAS 1-5224 (8 October 1965). 16. Connors, J. W., "Systems engineering of

fuel cell power plants" AIAA P a p e r No. 64-748, 3 rd Biennial Aerospace Power Sys- tems Conf. Phi la . Pa. (1 September 1964).

2. Grover , G. M . , Cot ter , T. P. and Er ickson, G. F., "Structures of v e r y high conductance", J. of Appl. Phys . (June 1964).

17. Johnson, T. K. "Ion exhange fuel batteries", P a p e r presented at 18th Annual Power Sources Conference, Atlantic City, N. J. (19 May 1964).

3 . Deverall , J. E. and 5. E. Kemme, "High thermal conductance devices utilizing the boiling of l i thium o r si lver" , Los Alamos Scientific Laboratory Repor t LA- 321 1 (October 1964).

18. General E lec t r i c Go. Handbook, "Fuel ce l ls for spacecraf t d i r ec t energy conversion operation, Lynn, Mass . (January 1964).

4. Cotter, T. P., "Theory of heat pipes", Los Alamos Scientific Repor t LA-3246-MS (23 February 1965).

19. Telecon: L . J. Nuttall, General E lec t r i c Co. Lynn, Mass . on 11 May 1965 Re: "Heat s ter i l iza t ion of fuel ce l l system" .

5.

6.

Deverall , J. E . , J. E. Kemme, "Satellite heat pipe", Los Alamos Scientific Laboratory, N.M. Report LA-3278-MS (1965).

20. P la tne r , I. L. , D. P. Ghere, and R. W. Opper thauser , "Capillary membrane hydrogen oxygen fuel cell sys tem for space vehicle application", Allis -Chalmer s Resea rch Div., P r o g r e s s Repor t on NASA Contract NAS 8-2696 (July 1964).

Arand, D. K., "On the performance of a heat pipe", Spacecraft and Rockets (May 1966).

7. Brosens , P. J . , "Solar thermionic generators for space power" ASME Paper No. 6 4 . WA/SOL-1. P resen ted at ASME Annual Meeting, New York, N.Y. (29 Xovember 1964).

21. Communication: N. P. Bannerton of Allis- Ghalmers Resea rch Div. Milwaukee Wis. to M. Koslover dated 1 July 1966, Subject: "Allis Chalmers , fuel cel l status.

8. Smith, H . , "Thermionics auxil iary power application fo r space" ASME paper No. 63-MD-54. ing Conference, New York, N. Y. (20 May 1963).

P resen ted at Design and Engineer-

22. Soltis, D. G., "Alternate approaches to s ter i l izable power sources", NASA TMX- 52137, Lewis Resea rch Center Cleveland, Ohio (1965).

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9. Ray, K. A., a n d D . H. Winicur, " A l a r g e a r e a cel l a r r a y (LASCA)", AIAA P a p e r No. 64-732, 3 rd Biennial Aerospace Power Sys- t e m s Conf., Phila. Pa. (1 September 1964). 24. Barak, M. "Fuel ce l ls - presen t posit ion and

outstanding problems" published i n Advanced Energy Conversion, Vol. 6, 1, pp. 29-56 (January 1966).

10. Arnold, E. D., "Handbook of shielding r e - quirements and radiation charac te r i s t i c s of isotopic power sources f o r terrestrial, marine and space application, ORNL-3476 (April 1964).

25. Kennedy, J . , J. Werth, and R. Weaver, "Lightweight l i thium chlorine battery" General Motors Defense Resea rch Lab. P a p e r (Septeniber 1964). 11. DeHaas, E. , "Radioactive isotope fueled

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27. Telrcon: J. flart inan of G. M. Allison Dis-. Iiiciianapolis x j t h M. Koslover Re: "Status of Li -CI Ce11s'I ( L O 3inle 1966).

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7 2

28. Meyers, W., "Development of a low tempera- tu re battery for space probe applications'' Livingston Electronic Gorp. Quar ter ly Report on NASA Contract NAS 3- 6009 (23 March 1965).

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30. Wood, R. L . , andD. J. Doan, "Ammonia bat ter ies , p a r t III" P a p e r presented at 17th annual Power Sources Conf., Atlantic City, N. J. (21 May 1963).

31. Koslover, M. "Performance of sealed nickel- cadmium cells after exposure to the NASA heat steri l ization environment", Avco RAD technical r e l ease AEDM-F510-340 (8 March 1965).

32. Koslover, M. , "Performance of nickel cadmium cells subjected to variable environ- ment", Avco RAD Technical Release AEDM-F540-477 (23 November 1965).

33. Berger , C. and F. C. Arrance, "Silver-Zinc Battery Capable of Thermal Sterilization According to J.P.L. Specification X 5 0 - 3 0 2 7 5 - TSTA", Douglas Aircraft Co. Report SM-48455 (March 1965).

34 . Telecon: Dr. C. Berger, Douglas Aircraf t Co. , Newport Beach Calif., 10 August 1965 Re: "Heat steri l izable si lver-zinc batteries".

35. Koslover, M., "Performance of Silver-zinc cells af ter exposure to the NASA heat sterili-

zation environment", Avco RAD Technical Release AEDM- F5 10-48 1 (November 1965).

36.

37. Telecon: G. Babb and J. R. Lucas; Eagle Picher Joplin, Mo. (8 October 1965) Re: "E. P. silver-zinc cells. ' I

38. Telecon: W. A. Chandler, NASAManned Spacecraft Center, Houston (9 September 1965) Re: space engineering (May 1963), "Cryogenic s torage fo r space e lect r ica l power, by W. A. Chandler.

ar t ic le i n astronautics and ae ro-

39. Pratt and Whitney Aircraft Communication SE 133; Adaption of Apollo for Space Station Applications, 9 August 1963 to J. F. N a t a l e , Avco.

40. "Maximum perniissiblc concentrations of radionuclides, U . S. Dept. of Commerce, Nat. Bureau of Standards Handbook, 69 Washington (June 1959).

31. Stafford, B. , I1Spacc p o u ~ r sitbs) stc-ni capabilities" ALAA Papclr No. 64-468. MAP, Sccond Annual Mcet ing, San Franciscu, CaliT. (26 July 1965).

42. Telecon: W. Pat terson o f Pratt-Whitnry Aircraf t Div. (UAC) Hartford, Conn., with M. Koslover, Re: "Most recent status of P?, WA fuel cells , I ' 30 Jiuiu 1966.

73

TABLE I

CONSTRAINTS AND REQUIREMENTS VERSUS MISSION PHASE

Prelaunch Postlaunch

to Preseparat ion

- Applies

A l l subsystems .

Mission Period Presepara tion

to Impact

Impact to end of

Mission Notes

\ Constraint ond Requirement

Sterilization Performed once on final assembly

Al l subsystems

t- Al l subsystems

~~

Degree of isolation varies with position and time

Thermal isolation Operating supply not necessary to thermally isolate

Operating supply t her- mally isolated

Ambient temperature t- All subsystems Degree of isolation varies with position and time

Stand time Applies Applies May be stored cold until this phase begins

Impact survival Applies After > 1000 G

Perform power profile Applies only to partial dis- charge depth

Applies to complet discharge ddptk

Both received and. generoted

A l l subsystems

Applies

Nuclear radiation minimum levels

Magnetic cleanliness Applies Applies Maximum of 15 gamma at 1 meter

Also any attitude throughout applies

Applies Zero G operation

Volume, weight and shape Maximum volume 2 ft3

t--- Al l subsystems Waste disposal

74

TABLE 2

ELECTRIC POWER FOR MARS LANDER SOLAR CELLS - WEIGHT AND SIZE

Isotope

Sr 90

Cs 137

Pm 147

Tm 171

T1 204

U 232

Th 228

Pu 238

Cm 244

* Reoson for I

. ~~ ~~

Assumed: a) 700 w/m2 (65.3 w/ft2) input a l l day

b) Tracking sun at a l l times within 5 degrees

c) 200 watts output required

d) No allowances for: radiation or other degradation, reliability, Atmosphere filtering or dust storms, erecting or tracking mechanisms.

Cell Type

Output a t

- 10°C W/f?

Silicon

Standard

S i I icon

Dendritic

Gallium Arsenide 3.27 Standard

P-0n-N

Thin-Film CdS. I 2.05

Weight

Panels (pounds)

I

Weight by LASCA*

method (pounds)

Panels (notes)

42.7 25 3' 29 + 2.22 1.6 Drum weight

40.8 24.2 27 + 2.1 1.6 Drum weight

61 I 37 I 6 7 + 1 5.0 1 1.6 Drum weight

98 57

1 2 3

0.6 lb/ft2 for A I panel excluding solar-cell array weight Panels assumed to be 0.6-inch thick Maximum panel width of 2 feet drum diameter of 12 inches

*LASCA9 TABLE 3

REVIEW OF ISOTOPE FUELS FOR RTG

Assumed Source Strength: 1000 w(t)

Chemicol Form

Oxide

Glass

Oxide

Oxide

Metal

Oxide

Oxide Matrix

Oxide

Oxide

nination '0, 40

Half - l i fe

(years)

28

26.6

2.7

1.94

3.9

74

1.9

90

18.4

Shielding Thickness to give gom a dose of 1.6 x lo-' mad/ hour at 1 meter.

(Cm Uranium)

14.5*

a*

5.2

1

4.1

24*

22*

3

11.5

Shielding Thickness ta give eutron dose of lO-'mrad/hour

at 1 meter

g/cm2 water

62

62

55

90

g/cm2 LiH

40

40

35

58

Remarks on

Radiation

High y Radiation

Radiation

B AEC Discourages i t s use

AEC Discourages i t s use No high temp. stable fuel form

High 3

High 3

B

<. c i"

7 5

TABLE 4

DATA ON SELECTED ISOTOPES

k w ( W year

Item

10 Not planned 12-15 8

Half Life

Combined weight o f fuel Producing 1000 w(t) which reduces gamma dose to 1.6 x lom2 m rad/hour a t 1 meter

Same a t 3.1 m distance

Availabil i ty starting 1969 (minimum required: 1-6 kw (t)/Year)

Estimated cost

Special treatment needed

Recommended for Lander

Item

Ma teria I

Hot junction temperature ("C)

Cold Junction temperature ("C)

Module efficiency (percent)

Module specific power (w/lb)

Output power flux (w/ft2)

Module specific cost ($/watt) (excludes isotope)

THERMOELECTRIC CONVERTER PERFORMANCE4'

Present Performance

Bi-Te Pb-Te Si-Ge MCC50

250 450 850 1200

80 180 350 700

4.5 5.0 6.5 4.0

15 to 20 15 to 20 15 to 20 10 to 15

--- --- --- 2.6

70 60 '100 100

Projected Performance

Bi-Te Pb-Te Si-Ge MCC50

250 450 950 1200

a0 1 80 350 700

5 .O 7.0 8.0 8.0

15 to 20 30 40 40

--- --- --- 6.5

15 to 20 20 to 30 10 to 20 10 to 20

76

TABLE 6 - POSSIBLE ABORT MODES AND SAFETY MEASURES

_I

Abort

1) Accidents, firer, explosions %route.

2) Abort on p d inrlLding fire, xplosion . 3A) Abort downronge. Vehicles x capsules impact with terminal relacity on land.

.3B) Same on water

:4) Vehicles olmst rwch Eorth arbit. Fuel capsule may burn up pr t ia l ly and be weakened suf- ficiently as to crack open on im- pact with the ocean surface.

(5) Third stage b i l s on second burn. Vehicles remain in low Earth orbit for somm days, then decoy, turn up wholly or par- tially. %me problems as in case (4) except that fuel con come down at any random loca- tion on Earth.

(6) Failure of separation or tm- jectory propulsion.

(7) Vehicle heat shield b i l r to slow vehicle down.

(8) Parachute foils. Same problem but to o lesser degree.

(9) Vehicle hits the ground with wrong ottitude.

Vehicle hits the ground a t 200 ft/sec with rigid attitude.

Normal Procedure

tronrprted to generator manu- facturer, then to launch p d .

Fuel arrives vrfely on p d and i s integrated with vehicle.

Vehicles ore launched okay.

Safety Measures

Capsules are made strong enougt to withstand a11 onslaughts.

Same as (1)

hme as ( I )

Fuel caprule made of materiol t b t doer not corrode or fuel should not dissolve in seawter.

See text

See texl

Vehicle bypasses Mors; okay from safety viewpoint.

Vehicle hits Mars ot 1900 fps -. perhop crushable material cushions foil enough that cop- sule survives (we text).

Same as (6)

h m e os (6)

Same problem os (8) but to o lesser degree.

Capsules must be designed to withstand lOOOg

Vehicles reach low Earth orbit.

Vehicles properly injected in Mors trajectay.

Vehicle in proper landing tro jectory

Vehicle heat shield operotes as planned

Parachute operates

Conceptual Design Status*

1

4

I I

TABLE 7

WEIGHT OF FUEL AND TANKS 38f 39 FOR H202 FUEL CELLS

Power Required Watt-hours

I-

Weight of Weight of H2 Fuel 0 2 Fuel Pounds Pounds i Tat01 filled Toto1 filled

Note 1 Note 2 -~

Total filled system weight high pressure

Qaseaus storage sterilizable

Note 3

* 100 watts overoge for 6 months

Note: 1 Storage pressure i s 300 psi fw H2 and 850 psi fw 0 2

2 Storage pressure i s loo0 psi for H2 and 7500 psi for 0 2

3 Allowance for heating to 1195°C. Assumes moterial strength not changed

77

TABLE 8

2.2

15.1

65

4.4

30

130

WEIGHT AND VOLUME OF L i - c l SYSTEM 25

5.4 5.5 8.3 x

18.3 19.1 2.9 x 10-1

68.2 71.4 1.07

10.4 10.6 1.6 x 10-1

36 37.5 5.6 x 10-1

136 142.5 2.1

-I Nominal

Power (watts)

50

50

Weight of1 Cells and

Auxiliaries

(pounds)

3.2

3.2

3.2

100 ’’ I 6.0

100

100

6.0

6.0

I

Operating Time

(hours)

24

168

720

24

168

720

Weight o f C1, + L i +

Tanks (pounds)

percent

(pounds)

1. Add 5 pounds for insulation

2. 1.8 Ib/KWH for total tanks and fuel.

3. Assumes no power being used 50 percent o f time except 10 percenf of nominal to keep fuel cel l (e.g., 100 watts i s peak power for 50 watts (average) ).

4. Excludes volume for controls such as values.

TABLE 9

MAGNETIC POLARIZATION OF BATTERIES AND CELLS13

Results

Yardney Silcad Cel l Model YS 5-6

Silver CFL

(8 cel l pack) PMO 5-7

(1 silver cadmium cell 5.0A.H. size)

+3.5 pole-cm (vertical. through top)

(8 silver zinc cells 0.5A.H. size)

-6.3 pole-cm, vertical

G. E. Battery (1 nickel cadmium cel l 12 A.H. size in stoinless steel - steel case)

+76 pole-cm, vertical +24 pole-cm, transverse

Cat. No. 42B012AB01 Ser. No. 261 1-6409

Cylindrical Cell (no designation)

Battery Single Cell (Gul ton)

S/N 173 1841 065-50 1

(1 nickel cadmium cell in plastic case 4 A.H. size)

+55.4 pole-cm, vertical +7.0 pole-cm, transverse

(1 nickel codmium cell in stainless steel case - 12 A . H. size)

+24 pole-cm, vertical +21 pole-cm, transverse +13 pole-cm, transverse

78

SEPARATION .L

IMPACT

ACCESS PORT

BlLlCAL CABLE

INCH FLIGHT

Figure 1 CONCEPTUAL PROEE/LANOER

CELEROMETERS

FIBERGLASS SHELLS

CRUSH-UP LAYERS BALSA WOO0

UMBILICAL CABLE

0 12 !.LuA-d

INCHES

2 PAYLOAD MODULE MOUNTING WE8

3 ANTENNA CAVITY 86.6275

Figure 3 FLIGHT CAPSULE POWER PROFILE

3 7 MINUTES- 23 SECONDS

86-6277 FRACTION OF SIDEREAL DAY

Figure 2 LANDER CAPSLLE

7 9

ISIDES ANDTOP)

RADIATOR SINK

I LTHERMOELECTRIC ARRAY ! y" 20-

g -

+ - P -

a -

LL z $l -- 0 - a

10- W

w I - w - COUPLE TYPE Si -Ge I

w + 5-

e - FUEL s -

2 '0 20 Id0 tl40 2AO 240 3AO Tcoc

NUMBER OFCOUPLES 2 N = 4 6 ACTIVE LENGTH EACH ELEMENT 1=2cm -

- o '

86 6278

FARAMETERS SELECTED

OUTPUT VOLTAGE V ' = 4 v

Figure 5 RTG WEIGHT VERSUS COLD JUNCTION TEMPERATURE

RADIATION WORKERS PERMITTED IHOUR/WEEK -

WORKMAN PERMITTED 40 HOUR/WEEK

W

86YEAR P u 2 3 8 z

0 IO 20 30 40 50 60 70 80 90 100 86.6279 KILOGRAMS

F i g ~ r e 6 #EIGHT OF ISOTOPE PNO CONTAiNE? VEPSUS RADIATION DOSE AT 1 METER FQOh5 THE SCURCE :rHICH PCTSAS SHIELD

4

86-6280 T ,cold junction temperature

L

3

W I-

.-200

loo! - -130 "C

Figure 8 TYPICAL TEMPERATURE DISTRIBUTION FOR 500-WATT LT) RTG HEAT SOURCE PROVIDED THE CONDUCTIVITY OF THE CRUSHABLE

MATERIAL K I S NOT LESS THAN 0.4 WPCxM

'=\'"''ORATOR VAPOR

WICK LIQUID FLOW

86-6287

Flgvre 9 GROVE'! HEAT PIPE

-< 02 TANK

ACCUMULATOR

Figore 7 JYPlCAL PELATE@NSHEP OF G AS A. FUNCTION CF COLD JUNCTION TE$2.;PE?ATUBE FCR A 30-:;4Ti 'E) 8TG

Figure 10 BACON FUEL CELL PGWEP SYSTEM

PERFORMANCE AT 7 5 O C

w

I) 80-

a

c 0

PERFORMANCE AFTER 67HR i 0.8 AT 150 OC

W

I 2 50 WATT RXG. PLUS BATTERY ,

60-

0.2c TEMP, 75% -I 1 0 20 40 60 80 100

TIME ,minuter 86-6284

Figure 1 1 EFFECT OF HIGH-TEMPERATURE STORAGE ON CELL PERFORMANCE, E.O.S. REGENERATIVE

t I

I I

1000 2000 3000 4000 5000 6000 1500 2500 3500 4500 5500 6500

HEAT REJECTION.Btu/ hr-kw 86-6285

Figure 12 HEAT REJECTION VERSUS CELL VOLTAGE FOR H 2 - 4 FUEL CELL

-1964-65(REFERENCE 18)

L W P 0.90

ij 0.80 0 > 0.70

060k

0;: Ib 2; 30 30 20 60 7b .30 910 too POWER DENSITY, w a t t r / f t 2

82-6286

LOA0

+ 0 . *:

POROUS IRON 0 0 ANODE

- POROUS GRAPHITE CATHODE

0 0

Li

2 Li IN

6,O INSULATION Li CLOUT

ATTHE ANODE ELECTROLYTE INTERFACE 2Li --zL~++zz

86-6287 THE TOTAL REACTION 2Li + C L ~ AT6200c-2Li CL

AT THE CATHODE ELECTROLYTE INTERFACE CL2+2Z 2 CL-

Figure 14 LiTHlUM CHLORINE CELL REPCTION

I O 0 I I I I I I I

I I I 1

Fiwre 15 TOTAL SYSTEM WEIGHTS FOR Li - CI AND RTG VERSUS TIME

ELECTRICAL INPUT

GAS GENERATOR

1

FILTERS SUMP TRAP

DIAPHRAGMS

STORAGE

DIAPHRAGMS

86-6289

Figure 13 H2-02 FUEL-CELL VOLTAGE VERSUS POWER DENSITY Figure 16 SCHEMATIC OF AN AUTOMATICALLY-ACT!VATED SILVER-ZINC BATTERY

81

::E 55

r - 2

SILVER ZINC

I I 1 . 1 I I 7

AT 75OF 3.776 watt- hours l 3 t AT ZOOOF 3.624 watt- hours -

- LEAD-ACID

0- 0 2 4 6 8 1 0

86-6290 DISCHARGE RATE,hours

Figure 17 TYPICAL ENERGYANIT WEIGHT CURVE FOR VARIOUS BATTERY SYSTEMS

TYPICAL SPREAD

L T

01 I I I I I I 0 8 16 24 32 40 48 5

DISCHARGE TIME.rnlnutes

86-6291 SCALE CHANGE

Fipre 18 OUTPUT PERFORMANCE RANGE FOR HEAT-STERILIZEC' NICKEL-CADMIUM CELLS

DISCHARGE CURRENT 0.8 AMPS, 3.5 A-H N i- CAD CELLS

I I I I I I I I 1 1 I T L 60 62 64 66 68 70 72 74 76 78 80 82

TOTAL NUMBER OF CHARGE- DISCHARGE CYCLES 86-6292

Figure 19 TIME TO 1 VOLT VERSUS NUMBER OF CHARGE-DISCHARGE CYCLES

$l.2b\>;;;F 2 ~

z g 1.1

w' 1.0

W c -I

0 I I I I I I 0 IO 20 30 40 50 60

86-6293 TIME.minuies

Figure 20 CELL TERMINAL VOLTAGE VERSUS TIME FOR 5 AMP-HOUR NICKEL-CADMIUM CELLS AT 75°F AND 2WDF AT 3 AMPS

CONSTANT CURRENT DISCHARGE

8 2

DISCHARGE CURRENT: 0.8 AMP

W (3

0 50 100 150 200 250 300 TIME ,minutes

86-6294

Figure 21 CELL TERMINAL VOLTAGE VERSUS TIME FOR A 3.5 AMP-HOUR NICKEL-CADMIUM CELL AT 0, 75, AND 150°F

Figure 22 MAGNETIC FIELD BALANCING CURRENT LOOP5

Figure 23 POWER SOUXCES AND APPLICATION OF CONSTRAINTS

8 3