US3530009A

US3530009A - Deployable radioisotopic thermoelectric generator

Info

Publication number: US3530009A
Application number: US676851A
Authority: US
Inventors: Guy F Linkous; Paul J Dick
Original assignee: Teledyne Inc
Current assignee: Teledyne Inc
Priority date: 1967-10-20
Filing date: 1967-10-20
Publication date: 1970-09-22
Anticipated expiration: 1987-09-22

Description

Sept. 22, 1970 G. F. LINKOUS ETAL. 3,530,009

DEPLOYABLE RADIOISOTOPIG THERMOELECTRIC GENERATOR Filed Oct. 20, 1967 2 Sheets-Sheet l ,g PATH 0F FLIGHT INVENTORS GUY F. LINKOUS PAUL J. DICK ATTORNEY Sept. 22, 1970 LINKOUS ET 3,530,009

DEPLOYABLE RADIOISOTOPIC THERMOELECTRIC GENERATOR Filed 001' 20, 1967 2 Sheets-Sheet 2 FIG. 4

INVENTORS' I GUYF. LINKOUS PAUL J. DICK ATTORNEY United States Patent 3,530,009 DEPLOYABLE RADIOESOTOPIC THERMO- ELECTRIC GENERATOR Guy F. Linkous, Glen Arm, and Paul J. Dick, Lutherville,

Md., assignors, by mesne assignments, to Teledyne,

Inc., Los Angeles, Calif., a corporation of Delaware Filed Oct. 20, 1967, Ser. No. 676,851 Int. Cl. H01v 1/00 US. Cl. 136-202 20 Claims ABSTRACT OF THE DISCLOSURE A thermoelectric generator for use with space vehicles and satellites. The generator uses a flat, rectangular shaped radioisotopic heat source which is positioned between two parallel banks of interconnected thermoelectric elements. The paddle shaped assembly is deployable from a thin arm extending from a satellite. Heat deriving from the radioisotopic heat source passes through the inward facing ends of the thermoelements to be dumped from radiating panels forming the opposed outer sides of the paddle shaped structure.

BACKGROUND OF THE INVENTION Field of the invention This invention relates to power supply systems for artificial satellites and more particularly to an arrangement of deployable radioisotopically heated thermoelectric generators for supplying power to such satellites and space traversing vehicles.

Description of the prior art The advent of highly sophisticated artificial satellites and space traversing vehicles has witnessed a catalysis within the scientific community of efforts for developing a broad spectrum of satellite carried technical missions. This emphasis in exploiting the capabilities of orbiting vehicles and the like has been observed to range from astronomical and biological experimentation to systems of immediate practical utility, as evidenced in communications relay weather data collections and mapping missions.

With each new technical advance and correlative suggestion of further utility for the space vehicles, there is generally introduced a requirement for the launch of more complex, bulksome and heavy mechanisms. Additionally, as satellite functions and their related systems become more complex and costly, practical economic considerations often dictate that the orbiting missions be of relatively longer duration.

The design complexities encountered in accommodating all of the advancing technical desiderata have led to the general and practical requirements for highly eflicient satellite instrumentational or functional systems having higher power capabilities as well as enhanced reliabilities.

An improvement in any of the satellite functional systems will permit the advantageous maximization of the effective payload capacities of existing launch vehicles.

The effort of developing enhanced system efiiciencies has, in particular, emphasized the need for overall design flexibility. Such flexibility is advantageous inasmuch as in many cases, technical advances made in the course of developing new systems are seen to render obsolescent related space traversing devices already in their final stages of manufacture.

Design flexibility is particularly desired in the field "ice of artificial satellite powering systems. These systems have long introduced design burdens and restrictions resulting from their relatively heavy weight or lower power densities, their relatively bulksome size and shape, presenting high profile drag areas, and their somewhat limited reliability and effective operational lifespans. It follows that the industry would consider the development of a modular power supply technique as advantageous to the broadening of the utility of satellite systems. The industry would be particularly receptive to a modular power supply offering low profile drag area and high reliability which retains the capability of ready incorporation within launch systems now in design, production and use. A capability for somewhat immediate insertion within extant launch system would permit the expansion and extension of missions presently near a completion status without incurring prohibitive redesign and modification costs.

Power systems now considered conventional fail to lend themselves to the modular flexibility requisite to such redevelopment, for the most part, as a result of their inherent physical characteristics. These characteristics are discussed in connection with certain of the more basic power system concepts in the paragraphs which follow.

Conventional batteries Almost universally considered in the selection of antificial satellite and space vehicle powering techniques is the battery. These devices, while affording a relatively stable power output, impose an oppressively high weight penalty upon the launch vehicle. This weight factor necessarily must detract from the mass allowance allocated to the instrumentation payload. Further detracting from the use of batteries is their short operational lifespan. The latter disadvantage precludes battery use where a space vehicle is slated to perform terrestrial servicing functions as in communications networks and the like. The coupled characteristics of comparatively higher weight along with a lower operational lifespan serve to minimize any design flexibility which might otherwise be realized from battery systems.

Solar cell panels Efforts to expand the operational lifespans of space devices have also devolved upon the use of solar cells. These devices, operating to photoelectrically convert light energy into electrical energy, are assembled within large planar banks to form panels. The panels, necessarily having relatively large surface or sail areas, are extended in orbit to collect solar radiation. While retaining some advantage of lower weight or higher power density, the solar energized power systems have encountered undesirable operational restrictions. For instance, the individual power cells of the panels have been found to be overly sensitive to various of the solar radiation wave lengths. As a result, the cells are prone to degenerate during use, thereby, lowering the operational lifespans and reliability of the panels. Additionally, the build-up of heat within the panels as a result of the impingement of solar radiation has been observed to cause their structural warping which, in turn, tends to destroy the integrity of protective coatings and the like. These coatings would otherwise serve to isolate the photocells from damage.

Solar cell energized power systems are also characterized in requiring means for properly orienting their surfaces with the sun. Generally, this orientation is accomplished by extending a plurality of the solar panels from a satellite, each or pairs of which are positioned to optimize the reception of impinging radiation for a given series predetermined Vehicle orientations. As a consequence of this deployment, the extended panels will function efficiently only during a portion of a flight program. The presence of the relatively extensive sail areas of the panels also is considered undesirable. By necessarily presenting a larger profile area to the direction of satellite orbit, the panels will tend to undesirably contribute to orbital decay.

In applications wherein solar cell power systems are utilized during earth orbit, it is necessary to install a supplemental power supply within space vehicles to accommodate them during their movement within the earth shadow. This accommodation generally is provided by additionally incorporating supplemental batteries within the power system. To promote longer lifespans, the batteries are charged during solar cell activation and load discharged during earth shadow orbit. Unfortunately, this repeated charge and discharge cycling has been found to adversely affect the reliability of the batteries. Of course, the addition of batteries penalizes the Weight-load capabilities of a space device.

The modular flexibility now desired for satellite power systems generally may not be realized with the use of solar panels. Inasmuch as the panels are of large dimension, the storage of their bulk during launch must be reckoned with. Further, in View of the precise orienting required of them during flight, there remain few alternatives to their mode of attachment to a vehicle.

Radioisotopic batteries Another approach investigated as a source of operational power for artificial satellites has been that of attaching a radioisotopically heated thermoelectric battery as generator to the devices. In general, the batteries comprise a relatively small quantity of a heat generating radioisotope which serves to heat one end of a number of interconnected thermoelectric elements. The thermoelectric elements, formed of certain semiconductive materials, are joined to form thermocouples, which when heated at a selected end serve to statically generate an electric current. In order to function efliciently, the thermocouples must be maintained Within a certain ambient environ and must be heated in a manner maintaining a preselcted differential of temperature across their individual lengths. The designs for radioisotopically heated thermoelectric units heretofore presented generally have assumed a somewhat cylindrical shape wherein a central heat producing core is surrounded on as many sides as possible by closely fitted clusters of thermocouples. By so clustering the thermocouple arrays, a degree of maximized consumption of the radioisotope heat energy is though to be realized. In order to establish and maintain a requisite diiferential of temperature across the thusly arrayed thermocouples, it is necessary to introduce and interconnect heat conducting and disposing systems from the cold ends of the thermocouples to ambient surroundings. This disposal arrangement is usually provided by somewhat elaborate banks of radiative fins. To further inject a degree of heat distribution control, various forms of insulation are inserted about the thermocouple arrays and a protective inert atmosphere is introduced into portions of the generator housing. Thusly deployed about the central heat source, the assemblage of thermoelements in most instances becomes structurally elaborate, close tolerances and difliculties of installation being the rule rather than the exception. To further add to their bulk and complexity, radiation shielding must also be incorporated within the device housings.

Assembled under the thusly described conventional design approach, the radioisotopic generators have been characterized as bulksome, heavy and intricate, requiring elaborate fin structures for heat dissemination as well as regulated safety procedures exposure. a

When adapted for attachment to an artificial satellite or the like, the difficulties attendant With utilizing the radioisotopic devices become more complex. Radiation hazards to launch personnel must be accommodated for both in shielding and duration of the presence of the devices in the launch-complex. During flight, the functional electronic circuitry or payload instrumentation of the satellite must be protected both from radiation and heat emanating from thermal conduction through the union of the thermoelectric battery and satellite framing. The added weight of shielding detracts from payloads otherwise available to the functional systems of the satellite. Additionally, the necessarily bulksome' structure of the heat dissipating and radiating fins of the devices presents an undesirably large profile drag area to the direction of satellite movement. The latter will shorten satellite longevity by encouraging orbital decay.

A further design problem is interjected by the necessity for protecting or isolating radioactive materials in instances of launch abort. In such situations, it is desirable to incorporate a stationary or low altitude ejectment procedure. Should the generator be complex, heavy and bulksome, ejectment schemes become unduly elaborate and impractical. Similarly, it is often desirable or necessary to impose a control over the radioactive device during its atmospheric re-entry following orbit. Should the generator design be unweildy, control at re-entry is diflicult to attain. Also, complexity is added by virtue of a necessity for suitable ejection devices and the like. Of course, the above-enumerated design constraints have been seen to derogate from a now desired design flexibility.

for avoiding radiation Thermoelectric solar panels The art has also suggested an adaptation of conventional solar panels wherein the conventionally used photoelectric cells are replaced by arrays of thermocouples. To provide a heat input into the thermocouples, one side of each panel is fabricated so as to act as a planar solar energy collector, while the opposed surface serves as a space radiator. A temperature differential is thusly provided for the energization of the thermoelements.

While the problems encountered from the thermal breakdown of photoelectric cells are ameliorated with the suggested technique, new constraints present themselves to the designer with this approach. Particularly where earth orbiting is contemplated, solar energy can provide only a limited heat flux. The use of thermoelectric elements requires that this heat be first collected and then dumped by radiation and, as a consequence, relatively low temperature thermoelements operating under a minimal temperature differentialwould be called for with the panels. The resultant low power densities obtainable from such an arrangement would again interpose the re quirement for expansive panels having undesirably large sail areas. Additionally, the thermoelectric elements would impose a higher weight burden than would the use of conventional photocell panels. Further, the large heavy panels of the proposed assembly would be inherently fragile and diflicult to protect during launching procedures. As may be evidenced from the foregoing, the use of thenmolectric solar panels with artificial satellites contributes more constraint than advantage over photocell devices.

SUMMARY OF THE INVENTION From the above review of present day approaches to designs for artificial satellite power supplies, it will be apparent that all fail in one measure or another to afford a design flexibility now demanded for the purpose of accommodating continually altering mission requirements. While proffering this highly desired modular flexibility the present invention additionally offers power supply improvements which eliminate or minimize the design constraints extant in the prior systems.

The power system of the invention is characterized in providing a radioisotopic powered thermoelectric generator which, through its incorporation within deployable panels allows significant modular versatility. In general, the present power generator employs a flat, panel shaped structure retaining a centrally situate planar radiosotopio heat source. This planar heat source is disposed between two somewhat planar banks or arrays of the thermoelectric elements. The entire assembly is interconnected with an artificial satellite or space vehicle through depolyable supporting arms of relatively minor dimension. Heat disposal requisite to the establishment of a temperature differential is evolved through the use of inelaborate reflective sheets. The latter reflective sheets serve the added function of providing a portion of the structural support for the panel form.

By virtue of its panel form, the generator arrangement of the invention enjoys the advantage of being depolyable away from the vehicle it powers during orbit or space travel. In such position the inherent isolation of the panels provides a much sought immunity of payload instrumentation and the like from interference caused by radiation and thermal conduction from the radioisotopic heat source of the generator. In the same light, requisite heat dissipation from the radiating portion of the generator is optimized as a result of the relative isolation of each individual radiating surface from other radiators or metal surfaces which may be attached to the satellite.

The panel structure of the invention provides further advantage by virtue of its facile attachment to a satellite structure. For instance, since the panel supporting arm may be of relatively minor diameter, the available points of attachment are increased and, additionally, may be of very simple design. This latter aspect of the inventive arrangement allows for simple additions of extra power supplies in late stages of satellite or space vehicle fabrication. 1n the same vein, the panels of the instant invention may readily be incorporated within space devices otherwise designed for solar cell powering with nominal re-engineering investment. The panel structures offer the further advantage of contributing to satellite stability and of being amenable to a simple and rugged design. By fashioning the generator units from groupings of standardized shapes and sizes of planar thermopiles and heat sources, the panels themselves may enjoy modular design flexibility.

The use of a radioisotopic heat source within the panel shaped generator offers the advantage of a long term span of operation on available lifespan. By affording a relatively high powered source of heat, the generator design will be seen to require a much smaller sail area as compared with solar photocells. This diminuation of profile drag area affords considerable advantage in lessening a tendency to orbital decay. Similarly, the inventive panel shaped generator will be seen to be characterized in having a lower weight per unit of power output in addition to its low bulk. The latter advantage is to be compared with the high weight and bulk of conventional radioisotopically powered thermoelectric-generators. Further advantageously distinguishing from radioisotopic generators heretofore in the art, the present device will be seen to service, handle and install in the terrestrial launching arena with relative ease and simplicity. The design is readily amenable to safe ejection procedures both at launch abort and during reentry procedures. Of course, by virtue of having a continually active heat source, the generator is not hindered by shadow during earth orbit.

The inventive generator is additionally characterized in the use of planar arrays of interconnected thermocouples which are disposed in parallel fashion opposite both faces of the aforedescribed planar radioisotopic heat source. Thusly disposed, the thermopiles are seen to maximize the output of the heat source while not introducing extraneous weight and bulk. The planar mounting of the thermoelements of hte device is non-complex, thereby not only lowering the cost of generator fabrication but also enhancing overall generator reliability.

Another object of the invention is to provide a deployable, panel shaped radioisotopic thermoelectric generator characterized in having a high power density and low profile drag area.

These and other objects and advantages of the invention will become apparent from the following detailed description and drawings.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a pictoral representation of an artificial satellite showing the power generators of the invention in deployed position, and, in phantom, in position for launch stowage.

FIG. 2 is a perspective view of a deployable radioisotopic thermoelectric generator fabricated in accordance with the invention with portions cut away to reveal internal structure.

FIG. 3 is a sectional view of a portion of the generator structure of the invention showing, in larger scale, the deployment of its thermopile structure.

FIG. 4 is an enlarged sectional view of a portion of the generator structure showing the attachment of thermoelectric elements into the external radiative surface.

DESCRIPTION OF THE- PREFERRED EMBODIMENT Looking to FIG. 1, an artificial satellite having a shape which generally may be encountered in the art is shown pictorally at 10. Satellite 10 is illustrated in a launched and orbiting mode, its assumed path or direction of flight being depicted by arrows 1 2. De ending from the body of the satellite 10 are radioisotopically heated thermoelectric generator panels shown generally at 14 and 16. These panels supply electrical energy to the instrumentation of the satellite.

Generators

14 and 16 are illustrated in their deployed or extended positions for orbit, their extending support being provided by

respective arm members

18 and 20. Inasmuch as relatively minor cantilever stresses are imposed upon the supporting arms during gravity-free orbit and since the arms need only serve as a simple conduit for electrical lead attachment with the satellite, their dimension may be advantageously small. In order to accommodate the short-lived stresses encountered at launch, however,

arms

18 and 20 are generally provided with a form of hinging union. These hinged unions, which are common in the art and consequently not pictured, permit both facile stowage of the panels =14 and 16 within the confines of a rocket vehicle at lunch and an adequate diminution of moment stresses otherwise encountered within

arms

18 and 20 by virtue of their cantilever structure. An illustrative positioning of the panels at launch is shown in phantom in the figure where the stowed orientation of panel 14 is shown at 14a and a stowed orientation of panel 16 is shown at 16a. The outward portion of

arms

18 and 20 which move with the generator panels when the latter are manipulated are shown respectively at 18a and 20a. Numerous techniques for lifting the generator panels into orbital position have been introduced into the art, generally, a simple hydraulically or pneumatically actuated piston and lever arrangement has been found satisfactory. Inasmuch as the panels are positioned when the vehicle is within a weightless environ, only a small amount of positioning force is required. Of course, springs or electromotor devices have also been found adaptable to the panel positioning function.

Arms

18 and 20 are fastened to the structure of the vehicle 10 through a simple coupling flange as at 22. It will be apparent to those skilled in the art that numerous structural couplings will serve the instant purpose. Of particular note, however, is the intrinsic simplicity and broad selection of the point of attachment inher nt with such small and inelaborate couplings.

Also to be evidenced from the figure is the relatively small size or bulk of the

generator panels

14 and 16. By virtue of this advantageous dimension and the above described mode of attachment to the satellite 10, the profile drag area as observed along the flight path direction is minimized. As a result, satellites employing the instant powering arrangement will enjoy a lengthier freedom from orbital decay. The advantageous shape and relative dimension of panels 14 and 116, deriving from a structural arrangement elaborated upon hereinafter, may also serve the general purpose of affording an improved stabilization for the vehicle to which they are attached. Generally, two generator panels are employed when this added stabilization function is desired about a roll axis such as shown at 21.

Turning to FIG. 2, the general structure of the generator panel is more elaborately portrayed. In the panel 14, there are disposed three modular generating

units

24, 26 and 28 aligned in side-by-side relationship so as to develop the fiat rectangular panel shape. Retaining these units in the noted position is a thin metal rectangular frame 30 functioning to support the units about their outwardly exposed peripheries. The frame 30 is fixed to the arm member 18 and serves the additional function of retaining the multilayered basic components of each of the units 24 through 28 in appropriate parallel spaced relationship.

Each of the

units

24, 26 and 28 are structured in the general fashion more clearly illustrated at unit 24. In that unit, a flat, rectangularly shaped radioisotopic heat source 32 is centered or sandwiched between two planar thermopile units. These planar thermopile units are depicted generally as

layers

34 and 36 in the figure and will be seen to comprise an array or network of interconnected thermoelectric elements.

Thermopile units

34 and 36 are spaced or separated from the flat heat source 32 by small gaps respectively at 3-8 and 40. The

thermopile units

34 and 36 are connected at their outwardly disposed faces to flat or planar radiating surfaces shown respectively at 42 and 44. Connection will be seen to be effected by countersunk nuts or the like shown typically at 46 which are distributed over the outward faces of the radiating surfaces 42 and 44. As will be evidenced in connection with later figures, each of the thermopile structures may be conveniently structured in identical fashion.

In assembling the panel as illustrated, numerous techniques will occur to those skilled in the art. It is desirable, however, to maintain the heat source 32 in relative isolation from spurious thermal conduction with the frame 30. Consequently, it will be found convenient to maintain the source 32 in appropriate central position by means of non-conducting brackets or the like which, while connected to the frame 30, will minimize the degree of thermal conduction into it. Should a particular application necessitate it, thermal insulation may be installed within the inward facing surface of the frame 32. The presence of such insulation will, of course, serve to promote an enhanced temperature differential across the thermopile layers 34 and 36.

Considering the broad aspects of its operation, panel 24. generates electrical energy by virtue of the heat input from source 32. This heat input from the planar source emanates outwardly from both surfaces of the source to impinge upon the hot sides of the

thermopiles

34 and 36. By virtue of their planar orientation, heat from the parallel source 32 must traverse the thickness of both thermopiles and, as a result, establishes a requisite temperature differential across them. An electric current is thereby generated, to be directed by suitable interconnective circuitry across the panel and through the arm 18 for supply to satellite instrumentation.

To afford efficient operation, the above-mentioned differential of temperature extant across the width or thicknesses of the thermopiles must be maintained throughout the lifespan of generator operation. Consequently, upon traversing the array of thermoelectric elements with the thermopiles, the heat must be dumped into the ambient environ. In the present invention, adequate dumping is effected by the remarkably simple expedient of thermally and mechanically interconnecting the thermopile cold sides to simple, planar reflective surfaces. For instance, the cold side of thermopile 34 is connected to the radiative surface 42, while the cold side of thermopile 36 is connected to a radiative surface 44. Each of the

radiative surfaces

42 and 44 may be interconnected to the frame member 30 by conventional technique. In some instances, it may be found desirable to interpose a degree of thermal insulation between the frame 30 and the radiative panels.

As a result of their simple configuration, the

radiators

42 and 44 enjoy unique advantage. When deployed, the radiating surfaces are not juxtaposed to any other radiator or, of similar impart, any thermally conductive body such as the surface of the satellite. Parlance of the industry describes this orientation as a condition where the radiators do not face each other. This isolated orientation serves to greatly enhance the available efficiencies of the generators.

Returning to the drawing, three generating units, 24, 26 and 28 have been illustrated as making up one generator panel 14. It will be apparent that a single, or a wide variety of combinations of the units may be em ployed to make up a generator. This structural aspect further illustrates the advantageous modular flexibility of the concept. Such flexible design variations may also encompass the direct attachment of small size units to the satellite surface and consequent elimination of deploying arms as at 18.

The general structure of the deployable panels also will be seen to simplify certain prevalent space device flight programing difliculties. For instance, the inherently simple method of attachment of the arm member 18 to a vehicle body will permit the use of quick fastening devices and the like. As a result, the radioactive generators may be attached at later stages of a launch vehicle countdown, thereby simplifying the difficulties presented by radiation safety restrictions. In the same light, it is often desired to retain such sources of radiation in an intact or relatively undamaged condition during re-entry from earth orbit. The flat, rectangular shape of the source 32 of the instant device may be of advantage for such situations. Aerodynamically, the low ballistic parameter characteristics of such shapes afford decreased re-entry heating and advantageously lower impact velocities.

Further with regard to the aspects of flight programing when using the instant generator panels, the relatively simple structure and, more particularly, the small dimension or diameter of the panel arm member 18 lends itself to simplified ejection procedures in the instance of launch abort. It is generally desirable that the radioactive ingredients be recovered in an intact state from such occurrences, consequently ejection procedures may be mandatory. Also, when extended for use during earth orbit, upon re-entry the arms 18, fabricated from aluminum or magnesium, will automatically sever from the space vehicle. Their small cross-sectional dimension will cause the arms to burn away during re-entry, leaving the panels to descend under the aforedescribed advantageous ballistic conditions.

Looking to FIGS. 3 and 4, the structural orientation of the components of a modular generating unit is pictured in more detail. At the center of the assembly, the radioisotopic heat source 32 provides heat input from both of its faces into

thermopiles

34 and 36. The source may be formed of a suitable isotopic fuel having a reasonable half-life such as Sr-90, Pu-238 or the like. Heat from source 32 radiates across the

gaps

38 and 40 to impinge upon the hot sides of the thermoelectric elements.

Gaps

38 and 40 serve to accommodate thermally induced expansions and contractions within the thermopile-radiator assembly. The incorporation of these gaps within the structure results in several additional advantages. For instance, interface complexities otherwise encountered with conductive sheets or the like used for evening heat distribution are eliminated. Further, a weight savings is realized and design flexibility is again enhanced with the elimination of tolerance problems.

The

thermopile structures

34 and 36 are formed of an array of spaced thermocouples shown at 50. Thermocouples 50 may be fabricated in numerous shapes and sizes, however, for the present illustration, they are shown having P and N elements of half-cylinder shape. The half cylinders of each couple are separated longitudinally by strips of insulating material 52 and are bonded to hot shoes or hot side collectors 54. The latter are seen to face the heat source 32 from across the

gaps

38 and 40. Typically, the cold sides of the thermoelectric elements are provided with half circular bonded cold shoes 56, to which are, in turn, bonded in layered fashion conductive half circular stress compensation wafers 58, thence the assembly is bonded to electrical connector straps 60. The straps 60 are conveniently fashioned from copper and serve to electrically interconnect the thermocouple outputs in series, parallel or combinations thereof. Above the connector straps 60 there are bonded electrically insulating wafers 62, on top of which may be bonded a second stress compensating wafer 64. Each of the wafers and connectors bonded upon the cold sides of the thermoelectric elements is selected having relatively high thermal conductivity so as to permit the facile passage of heat into

radiators

42 and 44. Where the wafers must be electrically insulative but thermally conductive, the metal wafers may be flame sprayed with an insulative oxide coating. An additional oxide coating useful for this purpose is described in United States Pat. No. 2,692,851, entitled Method of Forming Hard, Abrasion-Resistant Coatings on Aluminum and Aluminum Alloys under the inventorship of C. F. Burrows. Connection into the radiators is effected by bonding the upper portion of the thermocouples to conically shaped stud attachments 66. Threaded at their apex, the metal studs serve to provide both thermal conduction for heat dumping purposes, as well as to retain the thermopile assemblies in appropriate position establishing the

gaps

38 and 40. The studs are held in place withint he counterbores in the radiators by virtue of a threaded connection with nuts 46.

To provide an enhanced temperature distribution within the thermopiles, it may be found advantageous to form the entire array of thermocouples within an insulating medium such as that indicated at 68. A rigid-foam'aceous product often utilized for tihs purpose is a product having the brand name Min-K manufactured by the Johns-Manville Corporation of Manville, NJ. To further enhance thermal control within the generators, the material selected for fabricating the

radiators

42 and 44 should retain a relatively high thermal conductivity and sufiicient strength. Beryllium and similar materials will be found adequate for the purpose. The selection of thermoelectric materials for incorporation within the thermopiles will be determined from a number of design parameters including their design temperature of operation and optimum form of operating environment. It will be apparent from the drawings that the elements are easily assembled within the structural arrangement of the inventive generator panels. No close interface integration or tolerance problems are presented.

The presence of a thermopile structure on both sides of the fiat heat source will be seen to not only afford a more desirable and non-complex structure, but also will allow a realization of greatly enhanced power densities. Evidence of these improved power densities and of the relatively small requisite panel sail areas may be drawn from the accompanying tabulation of design performance for a typical generator:

In the tabulated examples it will be seen that when incorporating conventional lead telluride thermoelectric elements, a panel power density of 10.0 watts per square foot is realized. Similarly, when incorporating germanium-silicon elements, a power density of 25.0 watts per square foot may be obtained. It has been determined further, that power generation efiiciencies are not hindered by virtue of the exposure of solar radiation to any one of the radiating surfaces. Adequate heat dumping may be maintained under such influences.

It will be apparent to those skilled in the thermoelectric and satellite design arts that many variations may be made in the detailed disclosure set out herein for illustrative purposes, without departing from the spirit or scope of the invention.

We claim:

1. A thermoelectric generator for a space traversing vehicle comprising:

(a) At least one radioisotopic heat source;

(b) At least one thermopile formed having opposed heat collecting and heat dumping surfaces;

(c) Frame means for retaining said thermopile heat collecting surface in position adjacent said heat source;

(d) Heat radiating means in connection with said thermopile heat dumping surface for disposing of thermal energy emanating through said thermopile and establishing a differential of temperature thereacross; and

(e) Connector arm means in connection between said frame means and said vehicle for extending said generator a select distance outwardly from said vehicle.

2. The thermoelectric generator of claim 1 in which said thermopile heat collecting surface is spaced from said adjacent heat source surface a select distance so as to define a separation for permitting a substantial transfer of thermal energy thereacross.

3. The thermoelectric generator of claim 1 in which said frame means is disposed about the outer peripheries of said heat source and said at least one thermopile and is thermally insulated therefrom.

4. The thermoelectric generator'of claim 1 in which said heat source and said at least one thermopile are configured as rectangular parallelepipeds.

5. The thermoelectric generator of claim 1 wherein said connector arm means include pivot means for selectively positioning said generator about said supporting body.

6. The thermoelectric generator of claim 1 in which at least one of said thermopiles is operatively positioned adjacent each of said heat source surfaces.

7. The thermoelectric generator of claim 1 in which said heat radiating means comprises a planar thermally conductive surface.

8. The thermoelectric generator of claim 1 wherein said frame means is adapted to retain a plurality of said heat sources and said thermopiles.

9. The thermoelectric generator of claim 1 wherein said thermopile comprises:

(a) a plurality of interconnected thermoelectric elements characterized in having hot ends and cold ends;

(b) heat collecting hot shoes selectively bonded to the said element hot ends and positioned within said thermopile so as to establish said heat collecting surface;

() thermal insulation means disposed along the lengths of said thermoelectric elements; and

(d) circuit means electrically interconnecting said thermoelectric elements.

10. The thermoelectric generator of claim 9 including stud means connected to said thermoelectric element cold ends for. mechanically connecting said thermopile to said radiating means and providing thermal energy transference thereinto.

11. The thermoelectric generator of claim 1 in which:

(a) said thermopile heat collecting surface is spaced from said adjacent heat source surface a select distance so as to define a gap permitting a substantial transfer of thermal energy therefrom; and

(b) said frame means is disposed about the outer peripheries of said heat source and said at least one thermopile and is thermally insulated therefrom.

12. The thermoelectric generator of claim 11 in which said heat source and said at least one thermopile are configured as rectangular parallelepipeds.

13. The thermoelectric generator of claim 12 in which said heat radiating means comprises a planar thermally conductive surface.

14. The thermoelectric generator of claim 13 in which at least one of said thermopiles is operatively positioned adjacent each of said heat source surfaces.

15. A power supply for a space traversing vehicle comprising:

(a) at least one radioisotopic heat source formed substantially in the shape of a fiat rectangular parallelepiped having two opposed planar and parallel first and second surfaces;

(b) at least one first thermopile formed in the shape of a rectangular parallelepiped having opposed parallel heat collecting and heat dumping surfaces, said heat collecting surface being disposed adjacent said first heat source surface;

(c) frame means disposed about the peripheries of said heat source and said at least one thermopile and adapted to support said thermopile heat collecting surface in spaced relation from said heat source surface;

(d) a planar thermally conductive radiative surface in connection with said thermopile heat dumping surface;

(e) deploying means in connection between said space traversing vehicle and said frame means for effecting the flight positioning of said generator; and

(f) circuit means interconnecting said at least one thermopile and said space vehicle.

16. The power supply of claim 15 including thermal insulation means intermediate said frame means and said heat source and thermopile.

17. The power supply of claim 16 including at least one second thermopile formed in the shape of a rectangular parallelepiped having opposed parallel heat collecting and heat dumping surfaces, said heat collecting surface being disposed adjacent said second heat source surface.

18. The power supply of claim 17 wherein said thermopile heat collecting surfaces are spaced from said adjacent heat source surfaces a distance selected so as to define a gap permitting a substantial transfer of thermal energy between said surfaces.

19. The power supply of claim 18 wherein each said thermopile comprises:

(b) heat collecting hot shoes selectively bonded to the said clement hot ends and positioned within said thermopile so as to establish said heat collecting surface;

(c) thermal insulation means disposed along the lengths of said thermoelectric elements; and

(d) circuit means electrically interconnecting said thermoelectric elements at the said cold ends.

20. The power supply of claim 19 in which said frame means is adapted to retain a plurality of said heat sources and said thermopiles.

References Cited UNITED STATES PATENTS 3,167,482 1/1965 Katz 136-202 XR 3,277,827 10/1966 Roes 136202 XR 3,347,711 10/1967 Banks, et a1. l36-202 RODNEY D. BENNETT, Primary Examiner M. F. HUBLER, Assistant Examiner