Structurally embedded intelligent power unit Number:7,150,938 from the United States Patent and Trademark Office (PTO) owispatent

Title: Structurally embedded intelligent power unit

Abstract: An electrical power system for a spacecraft or other vehicle or structure comprising a structural member containing an integrated solid-state power source is disclosed. The power source includes a solar cell system and an energy storage system that are combined into a single unit or package, and in certain embodiments, are configured into the shape of a flexible aerobot balloon or a rigid nanosat surface panel. The power components of a preferred lightweight power unit, include thinly layered photovoltaic cells, a rechargeable lithium solid polymer electrolyte battery, a capacitor, electronics and thermal management capability.

Patent Number: 7,150,938 Issued on 12/19/2006 to Munshi,   et al.

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Inventors:	Munshi; M. Zafar A. (Missouri City, TX), Longhi, Jr.; Alfred J. (Alvin, TX)
Assignee:	Lithium Power Technologies, Inc. (Manvel, TX)
Appl. No.:	10/107,565
Filed:	March 27, 2002

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Current U.S. Class:	429/162 ; 429/120
Current International Class:	H01M 10/02 (20060101)
Field of Search:	429/111,120,162

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References Cited [Referenced By]

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Foreign Patent Documents

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Other References

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Primary Examiner: Ruthkosky; Mark
Attorney, Agent or Firm: Conley Rose, PC

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Parent Case Text

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CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims the benefit under 35 U.S.C. .sctn. 119(e) of U.S. Provisional Patent Application No. 60/280,087 filed Mar. 30, 2001.

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Claims

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What is claimed is:

1. An electrical power system for a spacecraft, vehicle or building, said system comprising: a structural member of a spacecraft, vehicle or building to be powered by the system, said structural member selected from the group consisting of structural panels, housings, coverings, supports and trusses and containing at least one integral solid-state power unit, each said power unit comprising a solar power device and an energy storage device, said power unit including a laminate comprising: at least one film photovoltaic cell, at least one thermal management device that is additional to any other components of the system; at least one lithium battery having an energy density of at least 125 Wh/kg, optionally, at least one polymer film capacitor, optionally, at least one film electrochemical capacitor, and electronic circuitry in electrical contact with said at least one photovoltaic cell, thermal management device and lithium battery, and polymer film capacitor, if present, and film electrochemical capacitor, if present.

2. The power system of claim 1 wherein said laminate comprises said at least one photovoltaic cell, thermal management device, battery, polymer film capacitor; and film electrochemical capacitor, layered in the order stated.

3. The power system of claim 2 comprising at least one electronic control in electrical contact with said electronic circuitry.

4. The power system of claim 3 wherein at least one said electronic control comprises a preprogrammed microprocessor.

5. The power system of claim 4 wherein at least one said electronic control comprises a dc-dc converter.

6. The power system of claim 1 wherein said electronic circuitry electrically interconnects at least two said photovoltaic cells in series or in parallel.

7. The power system of claim 1 wherein at least two said batteries are connected in series or in parallel and said electronic circuitry electrically interconnects said photovoltaic cells to said batteries.

8. The power system of claim 1 wherein said electronic circuitry electrically interconnects said photovoltaic cells to said batteries, and, if said at least one polymer film capacitor and/or at least one film electrochemical capacitor is present, interconnects said batteries to said at least one capacitor, to form a single function power module.

9. The power system of claim 1 wherein, said at least one polymer film capacitor and/or said at least one film electrochemical capacitor is/are present, and said electronic circuitry electrically interconnects said photovoltaic cells to said batteries, and interconnects said batteries to said at least one capacitor, to form a multi-function power module.

10. The power system of claim 1 wherein said electronic circuitry connects said at least one battery and/or said at least one capacitor, if said at least one polymer film capacitor and/or said at least one film electrochemical capacitor is present, to a defined electrical load.

11. The power system of claim 1 wherein each said thermal management device comprises at least one heating element.

12. The power system of claim 1 wherein each said battery comprises at least one electrochemical cell containing a film amorphous carbon anode, a film lithium ion polymer electrolyte gel containing a liquid organic solvent, and a film lithiated metal oxide cathode.

13. The power system of claim 12 wherein said metal oxide is chosen from the group consisting of the oxides of manganese, cobalt and nickel, and combinations thereof.

14. The power system of claim 12 wherein said film lithium ion polymer electrolyte gel contains no more than about +wt % liquid organic solvent.

15. The power system of claim 1 wherein said battery comprises an electrochemical cell containing at least one solvent-free solid state lithium polymer electrolyte cell.

16. The power system of claim 15 wherein said solvent-flee solid state lithium polymer electrolyte has a conductivity greater than 1.times.10.sup.-4 S/cm at 25.degree. C. or below, and comprises a mixture of: (a) a base polymer material comprising at least one ionically conductive polymer and having an initial conductivity of at least 1.times.10.sup.-5 S/cm at 25.degree. C. when combined with a metal salt; (b) a lithium salt; (c) an inorganic filler having an average particle size less than 0.05 micron in diameter and a surface area of at least about 100 m.sup.2/g; and (d) an ionic conducting material having an average particle size less than 0.1 micron in diameter and an initial ionic conductivity of at least 2.times.10.sup.-3 S/cm at 25.degree. C.

17. The power system of claim 15 wherein each said cell comprises a film lithium metal anode and a film lithium insertion cathode.

18. The power system of claim 1 wherein each said film battery has an energy density at least in the range of about 200 250 Wh/kg.

19. The power system of claim 1 wherein said at least one polymer film capacitor is present and each said polymer film capacitor has an energy density at least in the range of about 5 10 J/cc.

20. The power system of claim 1 wherein said at least one film electrochemical capacitor is present and comprises at least one high surface area carbon electrode and an aqueous or non-aqueous electrolyte.

21. The power system of claim 1 wherein said at least one film electrochemical capacitor is present and is capable of supplying 1 3 volts.

22. The power system of claim 1 wherein said at least one film electrochemical capacitor is present and comprises at least one electrode containing a valve-metal oxide.

23. The power system of claim 22 wherein said valve-metal oxide contains 50 95 wt. % ruthenium oxide and the balance is vanadium oxide.

24. The power system of claim 1 wherein said structural member comprises a solar panel containing electronic circuitry laminated between two power units, each said power unit comprising: at least one film photovoltaic cell, a thermal management device, at least one lithium battery, at least one polymer film capacitors and at least one film electrochemical capacitor.

25. The power system of claim 24 wherein said laminate comprises, in the order given: at least one film photovoltaic cell, a thermal management device, at least one battery, at least one polymer film capacitor, at least one film electmochemical capacitor, electronic circuitry, at least one polymer film capacitor; at least one film electrochemical capacitor, at least one lithium battery, a thermal management device, and at least one film photovoltaic cell.

26. The power system of claim 1 wherein said laminate is conformable to a predetermined three-dimensional shape.

27. The power system of claim 1 wherein said laminate is flexible.

28. The power system of claim 1 wherein said laminate is substantially impervious to gaseous species.

29. The power system of claim 1 wherein said laminate is at least semi-rigid.

30. The power system of claim 1 wherein said laminate is resistant to thermal and electromagnetic radiation.

31. The power system of claim 1 wherein said solid-state power unit is embedded in said structural member.

32. The power system of claim 1 wherein said structural member comprises an exterior surface of a spacecraft, vehicle or building.

33. The power system of claim 1 wherein said structural member is an interior support of a spacecraft, vehicle or building and said solid-state power unit is conformable to the shape of said structural member.

34. The power system of claim 1 wherein said solid-state power unit is laminated onto a structural member of a spacecraft, vehicle or building.

35. The power system of claim 1 wherein said solid-state power unit provides structural support for said structure.

36. The power system of claim 1 wherein said structural member comprises a structural panel of a satellite.

37. The power system of claim 1 comprising at least two electrically interlockable modular panels.

38. The power system of claim 1 wherein said at least one thermal management device comprises a temperature sensor.

39. The power system of claim 38 wherein said at least one thermal management device further comprises a heating element and/or a thermoelectric cooling element.

40. The power system of claim 1 wherein said at least one polymer film capacitor is present and also serves as a radiation shield.

41. The power system of claim 1 wherein said solid-state power unit comprises at least one thermally insulating material.

42. The power system of claim 1 wherein said battery and said at least one polymer film capacitor, if present, and/or said at least one film electrochemical capacitor, if present, are operational at a temperature of about -20.degree. C. to about +150.degree. C.

43. A vehicle comprising the solid-state power system of claim 1 wherein said solid-state power unit is integral with a structural member of said vehicle.

44. The vehicle of claim 43 chosen from the group consisting of satellites, nanosats, aerobots, balloons, high altitude or space platforms, terrestrial vehicles and boats.

45. The vehicle of claim 43 wherein said solid-state power unit comprises at least one operationally interconnected laminate containing: at least one thin film photovoltaic cell layer, at least one layer containing a thermal management device, at least one thin film lithium battery layer, each said battery being capable of a multiplicity of discharge/recharge cycles, optionally, at least one film capacitor layer, optionally, at least one film electrochemical capacitor layer, and electronic circuitry.

46. The vehicle of claim 43 wherein said thermal management device comprises at least one heating element and said power unit is capable of operating under cryogenic conditions.

47. The vehicle of claim 46 wherein said solid state power system is capable of electrolyzing, compressing, liquefying or freezing a vehicle-transported gas or of electrolyzing, compressing, liquefying, freezing, sublimating or boiling a frozen or liquid phase of an atmospheric gas comes into contact with said vehicle when said vehicle is used for its intended purpose.

48. The vehicle of claim 43 wherein said power unit is capable of operating for up to about 15 years.

49. A spacecraft, vehicle or building comprising the power system of claim 1 wherein said at least one power unit is integral with a structural member of said spacecraft, vehicle or building.

50. The structure of claim 49 wherein said structural member is flexible.

51. The structure of claim 49 wherein said structural member is at least semi-right.

52. The power system of claim 1 wherein each said thermal management device comprises at least one thermoelectric cooling element.

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Description

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BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention generally relates to integrated solid state electrical power sources and to structures and vehicles containing such integrated power sources embedded in or formed as a structural member. More particularly, the invention relates to solar power systems containing flexible integrated photovoltaic solar cells and advanced energy storage systems in which ultra thin-film batteries, capacitors and other components are formed as a single unit and embedded into a supporting or covering element of a structure.

2. Description of Related Art

Integrated Power Sources. Several integrated electrical power sources have been devised over the past two decades for a variety of applications, and in some cases a battery and/or a solar photovoltaic collection device has been integrated or embedded into a support or covering for a structure or device. The term "integrated power source" refers to a power source that is structurally combined as a single unit. For example, U.S. Pat. No. 4,740,431 describes an integrated solar cell and battery primarily for use in portable electronic devices such as radio transceivers, portable computers and emergency lights, as integrated power generation and storage modules. Thin film lithium-aluminum alloy anodes, polyethylene oxide/lithium salt solid electrolyte, and a molybdenum selenide cathode are described as exemplary battery components and the solar cells may contain amorphous silicon.

U.S. Pat. No. 5,626,976 describes a certain flexible energy storage device with integral charging unit (photovoltaic device). Conventional aqueous or non-aqueous electrolyte or polymer gel electrolyte material, or a solid state electrolyte is employed, and enclosed in a polymeric vapor barrier package.

U.S. Pat. No. 5,180,645 describes a certain integral solid state embedded power supply for a self-powered portable electronic product, such as a radio. The elimination of the outside metallic case of each cell and the elimination of the outer packaging for the overall battery provides a considerable reduction in volume and weight of the product.

U.S. Pat. No. 5,644,207 describes a renewable modular integrated power source that is bonded to a housing or structure or is molded into a desired shape using the battery material itself. In one application solar cells backed with a thin film polymer battery supply a lightweight fully integrated power source. Conventional solid state or ionically conducting gel polymer electrolyte batteries, such as that described in U.S. Pat. No. 5,637,421 are employed.

U.S. Pat. No. 5,695,885 describes a battery configured in the form of a flexible wrist band for a watch or personal radio, television or communication device, with a plurality of photovoltaic cells disposed on the outer surface of the battery. Conventional alkaline or Ni--Cd type batteries are described along with presently known flexible anode and cathode materials.

U.S. Pat. No. 6,224,016 describes a flexible energy producing envelope material or covering for a balloon used in high altitude and stratospheric applications. The covering includes a flexible solar cell layer, a flexible substrate that matches the shape and size of the airship gore as well and an electrically conductive conduit disposed in a flexible electrically non-conductive adhesive connecting the flexible solar cell substrate to the airship substrate. No energy storage means is integrated into the covering with the photovoltaic cell layer, however.

For powering satellites, aerobots (balloons) and other spacecraft applications, such as altitude control systems, communications and various payloads, a conventional integrated power system (e.g., solar power system) that might be satisfactory for powering a small, personal device such as a radio or a wristwatch would not be adequate, particularly spacecraft intended for long missions in space.

Space Power Systems. A number of advanced propulsion and energy storage technologies with reduced mass and volume, long service life, higher reliability, thermal and radiation resistant and low cost for use in spacecraft such as satellites, nanosats and aerobots (balloons) are required for future space missions, and for military, surveillance, scientific research and commercial applications. For instance, there is a great need for a longer lasting power source for use in deep space exploration away from the sun, and no suitable power source is presently available using today's technology. In the cryogenic conditions of deep space, aerobots or balloons will require more battery capacity and higher power capabilities in a lightweight design than is presently possible with existing power sources. Typically, nanosats and aerobots call for the incorporation of photovoltaic devices on the surface of panels to charge the auxiliary on-board batteries for a variety of usages. Another application for lightweight sources of electrical energy is in the construction of lightweight high altitude or spaceborne platforms (e.g., for missile defense use).

Further complicating the problem, today's satellite designers continually strive to raise the level of onboard power generation while at the same time endeavoring to lower the cost of such increased power capability. Therefore, satellites that are now envisioned for various future anti-missile roles, for example, have an ever increasing demand for power. Although the goals and operational requirements of satellite missions can vary widely, they are in almost every case constrained by the type, level, and duration of the on-board power source.

Space power systems designed for continuous high-power-output applications have most commonly been based on the use of photovoltaic panels that regenerate the battery. In all orbits, satellites are subjected to a greater or lesser number of eclipses, and thus have to rely on internal energy to continue their missions. This internal energy is provided by rechargeable batteries, which store the surplus energy generated by spacecraft solar arrays. In geosynchronous orbit (GEO), the satellite is in eclipse for only a short time: for approximately 4 months, the satellite is in the sun at all times; for the next 50 days, the satellite is in eclipse for periods up to 72 minutes per day (it builds from 0 minutes to 72 minutes in the first 25 days and then drops down to 0 minutes); then, the satellite is in the sun for approximately 4 months. GEO applications require approximately 100 charge/discharge cycles (C) per year and as such the requirement is not so strenuous. In low earth orbit (LEO) missions, the battery typically provides power for 30 minutes at the 1 C rate, followed by a one hour charge. LEO applications usually require about 6000 cycles per year.

Ni--H.sub.2 Batteries. In the past decade, nickel-hydrogen batteries have been the technology of choice for both commercial and defense-related satellites in both GEO and LEO applications. They have inherent advantages over their predecessor nickel-cadmium batteries. These include superior energy density, longer cycle life, and better tolerance to overcharge and reversal. The goal of increasing cell energy density provided the original impetus for the development of the nickel-hydrogen system. In a battery, the useable energy density of a nickel-cadmium battery is typically in the range of 20 35 Wh/kg while that of a nickel-hydrogen battery it is about 40 45 Wh/kg.

Present day satellites can be categorized into four groups: large, small, mini and micro. The large satellites have mass greater than 1000 kg; the small category is 500 1000 kg; the mini in the 100 500 kg range and the micro less than 100 kg. Table 1 shows satellite battery-power systems data for some previous space missions. The power requirements for GEO satellites are in the 10 15 kW range for extremely long-term missions lasting up to 15 years.

TABLE-US-00001 TABLE 1 Specifications of Some Previous Satellite Battery-Power Systems Array Battery Satellite Power Battery Size Satellite Program Payload Mass (kg) (W) Type (Ah) HESSI NASA Small Solar Imager 290 505 Ni--H.sub.2 15 Explorer CPV MightySat AFRL Research & 121 330 NiCd 10 2.1 Imaging Deep Space 1 JPL New Adv. Tech. Ion 486 2500 Ni--H.sub.2 30 Millenium Prop. CPV ARGOS AF Experiments 2100 1074 Ni--H.sub.2 45 IPV Stardust NASA Discovery Sample returns 380 Ni--H.sub.2 16 CPV NEAR NASA Discovery Asteroid 805 1800 Super 9 Rendevous NiCd EO-1 GSFC New Earth 529 600 Super 50 Millenium observations NiCd Terriers NASA STEDI Earth atm. 125 31 NiCd 4.8 observations HETE-2 NASA Gamma ray 124 168 NiCd 1.2 detection.

For longer missions requiring high cycle life and cryogenic operating conditions, low energy density Ni--H.sub.2 batteries are now employed. These batteries are not only bulky and constructed in heavy steel pressure vessels, but also offer lower current capabilities. For shorter missions and low cycle life requirements in which high power is required, NiCd batteries are typically used. These batteries and solar panels occupy between 15 25% of the total weight and volume of a present day satellite. The volume of the satellite is a function of the surface area for the solar panels and hence the power availability. They vary from about 90 m.sup.3 for satellites weighing over 2000 kg to about 2.5 m.sup.3 for those weighing about 35 kg.

Lithium Ion Batteries. Space power systems of the future are projected to require power levels that may extend far beyond the current levels of demand. Thus, there is an increasing need for lightweight, high energy density batteries with long active and cycle lives beyond what can be delivered by nickel-hydrogen batteries. Toward that end, lithium ion battery systems are at this time undergoing intense investigation. For Example, U.S. Pat. No. 5,456,000 describes one type of lithium ion rechargeable battery containing polymeric film composition electrodes and separator membranes. An advantage of a lithium ion battery system is that the useable specific energy is two to three times greater than that of nickel-hydrogen batteries. This represents a significant launch cost savings or increased payload mass capabilities. However, the energy density and cycle life of a conventional lithium ion battery system is typically only about 150 watt hours per kilogram and 500 cycles, respectively, under deep discharge conditions.

Solid Polymer Electrolyte Batteries. Commonly assigned U.S. Pat. Nos. 6,645,675 and 6,664,006, and PCT International Publication Nos. WO 01/17051 and WO 01/17052 describe all-solid-state electrochemical cells and batteries with a flexible, ionically conductive polymer membrane as the electrolyte For example, a rechargeable all-solid-state lithium polymer electrolyte battery comprises an ultra thin lithium anode, which may be either a metallic lithium element or a lithium metal layer about 0.1 to 100 microns thick, over the metallization layer of a metallized polymer substrate. The metallized polymer substrate has an inactive polymer layer about 0.5 to 50 microns thick and a metallization layer 0.1 1.mu. thick on top of the inactive polymer layer. This battery also has an ultra thin-film cathode layer containing a metallized polymer substrate. The metallized polymer substrate has an inactive polymer layer about 0.5 to 50 microns thick and a metallization layer about 0.01 1.mu. thick on top of the inactive polymer layer, and has a layer of active cathode material 0.1 100.mu. thick on top of the metallization layer. The battery also has a polymer electrolyte layer 0.2 100.mu. thick placed between the above-described anode and cathode layers. This polymer electrolyte has a conductivity greater than 1.times.10.sup.-4 S/cm at 25.degree. C., or may even conduct as well below 25.degree. C. The polymer electrolyte comprises a mixture of a base polymer material comprising at least one ionically conductive polymer and having an initial conductivity of at least about 1.times.10.sup.-5 S/cm at 25.degree. C. when combined with a lithium salt. The mixture also includes the lithium salt, an inorganic filler having an average particle size <0.05 micron in diameter and a surface area of at least about 100 m.sup.2/g, and a lithium ion conducting material having an average particle size <0.1 micron in diameter and an initial ionic conductivity of at least 2.times.10.sup.-3 S/cm at 25.degree. C. In some embodiments, the inorganic filler is 0.1 20% (by volume of solid polymer electrolyte) high surface area filler having an average particle size .ltoreq.0.01 micron and chosen from the group consisting of fumed silica and alumina. In some embodiments, the lithium ion conductor material is 0.1 80% sulfide glass (by volume of solid polymer electrolyte). In some embodiments, the lithium ion conductor material is a ceramic ion conductor chosen from the group consisting of lithium beta alumina and silicates. Ion mobility is achieved through coordination of electrolyte ions by sites on the polymer chain, thus promoting electrolyte dissolution and salt dissociation. Such a battery design overcomes the disadvantages inherent in liquid electrolytes and provides better long-term storage stability. By also employing thin, flexible electrodes, such batteries can be made into virtually any shape and size, are reasonably rugged and leakproof, have high specific energy (Wh/kg) (gravimetric) and energy density (Wh/L) (volumetric), high cycle life, low self-discharge, high current drain capability, lower resistance, and wider operating temperature range.

Two major factors that drive satellite design are launch costs per kilogram of satellite and instruments and power availability. Since small and moderate mission costs can run from $100 million to greater than $1 billion, reliability, efficiency, and density of power in the satellite design and components is necessary. Moreover, future satellites for missile defense applications will require not only greater longevity from the power source but also higher power, lower weight and volume, greater degree of power management, and significantly lower cost, so that more firepower can be packed into these satellites. Other applications such as surveillance, deep space exploration or terrestrial use will have similar requirements. The existing space power technologies are only capable of providing power sources that are rigid, bulky, heavy and costly, and which are monofunctional, i.e., they serve as a power source only. Present day satellites also need, in addition to the power source components, a special `skin` enclosure that is sturdy enough to protect the power source from the sun's heat and from space debris. This is necessary since the conventional batteries contain a liquid electrolyte that can evaporate away at high heat or cause a dangerous situation. At the same time, the electrodes need to be contained in a rigid environment to prevent them from disintegrating due to constant bombardment by space debris. Furthermore, the conventional liquid electrolyte batteries, and even the newer lithium ion batteries, do not provide satisfactory high current drain capabilities for use in many of the latest applications being developed, such as anti-missile systems, aerobots or balloons for planetary exploration.

Another problem with existing photovoltaic systems is that the solar cells at optimum levels of performance get hot, which reduces the efficiency. Solar cells operate better at cooler temperatures.

Despite the technological advancements provided by prior art devices, known integrated power sources still suffer from limitations of excessive size and weight, lack of sufficient flexibility or conformability, insufficient power density (watts/volume) and specific power (watts/weight), particularly for use in the new and future high altitude and space-related military programs, for space exploration, and for various other scientific research and commercial applications.

BRIEF SUMMARY OF PREFERRED EMBODIMENTS

The present invention overcomes many of the shortcomings of power source technologies presently available for use in space exploration and earth orbital operations. In accordance with certain embodiments of the present invention, an electrical power system for a structure is provided which comprises a structural member containing at least one integral solid-state power unit. The terms "integral or integrated solid-state power unit" means that the component parts of the power unit are laminated together or formed as a single unit. The integrated power unit is preferably embedded or integrated into a structural member (e.g., a truss, housing, or structural panel) of a spacecraft, vehicle, building, apparatus or other structure, or the laminated power unit itself is configured as a dual or multi-functional structural member for use in constructing or assembling a structure as well as serving as a power source for the structure. The structurally integrated power units disclosed herein are not limited to merely wrappings or coatings for structural members, in contrast to previously existing power units that are sometimes referred to as being integral to a structure. The new integrated or embedded power unit is sometimes called a Structurally Embedded Intelligent Power Unit ("SEIPU") in this disclosure. For example, a lightweight structural panel comprising an integrated solid-state power source may be configured in the shape of a balloon, nanosat, satellite or air/space platform building material in which the laminated components of a solar power system, including photovoltaic solar cells), and advanced energy storage system, including batteries and capacitors, and auxiliary devices such as heaters, microelectronics and cooling elements, are directly embedded into the structural panels of the satellite, or configured into the shape of the balloon structure. Benefits of using a SEIPU instead of a conventional power unit for such structures include increased capacity and high power capability of the system, both of which features have not previously been available simultaneously in the same satellite. As discussed further in the Detailed Description of Preferred Embodiments, SEIPU-containing systems allow greater latitude in power management and provide reliable power under a wide range of conditions, reduce the weight of the satellite, or other structure, and provide improved charged acceptance capability.

The term "solid-state" as used herein generally means non-liquid containing. A solid-state polymer electrolyte device can contain a small or limited amount of non-aqueous liquid electrolyte, preferably no more than about 30 wt. %, entrapped in a polymer. Accordingly, the term "solid-state" can also refer to devices that contain a soft or gel-like semi-solid polymer component.

Accordingly, in certain embodiments the power unit provides structural support for a structure, and in some embodiments the structural member is an interior support in a structure, such as a truss framework of a satellite. In certain embodiments the power unit is conformable to the shape of the structural member. In some embodiments the power unit is laminated onto a preexisting structural member, or constituent part of a structure. In some embodiments the structural member comprises a panel, or a group of interlockable modular panels, in which the power unit is embedded.

In certain embodiments the power system comprises a laminate of the following components: at least one thin film photovoltaic cell, optionally, a thermal management device (e.g., a heating element or a thermoelectric element), at least one thin film battery, at least one high voltage thin film capacitor, at least one thin film electrochemical capacitor, and flexible electronic circuitry that is in electrical contact with each of the above layers. Each of these components are preferably, but not necessarily, layered in the order listed. In still other embodiments, the thermal management device includes a temperature sensor, a heater, such as one or more resistive heating elements, or a thermoelectric cooling element, or any combination of those.

In certain embodiments, the battery system is based on a thin film lithium ion polymer electrolyte battery (utilizing a carbon anode) having an energy density of at least 125 Wh/kg, or utilizes a lithium polymer electrolyte battery (with a lithium metal anode) having an energy density of at least 200 Wh/kg, that is capable of operating over a wide temperature range, including cryogenic conditions and higher temperatures, e.g., -20.degree. C. to +150.degree. C. or more. In fact, preferred lithium polymer electrolyte batteries operate more efficiently at higher temperatures than at room temperature. In highly preferred embodiments, the thin film solid state lithium polymer electrolyte has a conductivity greater than 1.times.10.sup.-4 S/cm at 25.degree. C. or below, and comprises a mixture of: (a) a base polymer material comprising at least one ionically conductive polymer and having an initial conductivity of at least 1.times.10.sup.-5 S/cm at 25.degree. C. when combined with a metal salt; (b) a lithium salt; (c) an inorganic filler having an average particle size less than 0.05 micron in diameter and a surface area of at least about 100 m.sup.2/g; and (d) an ionic conducting material having an average particle size less than 0.1 micron in diameter and an initial ionic conductivity of at least 2.times.10.sup.-3 S/cm at 25.degree. C.

In preferred embodiments, the SEIPU-containing power system also contains a thermal management system for adapting the system to extreme temperature use. This can include thermal management devices such as heating elements and/or thermoelectric element, for example.

In certain embodiments, the active photovoltaic power source is on a rigid support such as stainless steel. In preferred embodiments, the active photovoltaic power source is on a flexible support such as Kapton.TM., polyester, polypropylene, polyvinylidene fluoride or polyethylene naphthalate.

In preferred embodiments the power system also includes one or more electronic controls, such as preprogrammed microprocessers and dc-dc converters, connected to the electronic circuitry. In some embodiments at least two photovoltaic cells are connected in series or in parallel. In some embodiments at least two of the batteries are connected in series or in parallel and are interconnected to the photovoltaic cells by the laminated electronic circuitry. In some embodiments the photovoltaic cells, batteries and capacitors are interconnected to form a single function power module, and in other embodiments they are electrically interconnected to form multi-function power modules. In some embodiments the electronic circuitry connects one or more battery layer and/or one or more capacitor layer to an electrical load, to provide either high energy or high power, as desired.

In certain embodiments the battery layer of the power unit comprises one or more cells containing a thin film amorphous carbon anode, a thin film lithium ion electrolyte gel containing a liquid organic solvent (up to 30 wt. %), and a thin film lithiated metal oxide cathode where the metal is manganese, cobalt, nickel or a combination of one or more of those metals.

In certain other embodiments the battery layer comprises at least one rechargeable thin film solid-state lithium polymer electrolyte cell that has a thin film lithium metal anode and a thin film lithium insertion cathode. In preferred embodiments the film battery has an energy density in the range of 200 250 Wh/kg. In preferred embodiments the high voltage film capacitor system is based on a thin film metallized polymer capacitors having an increased energy density over existing polymer film capacitors, and in certain embodiments the film capacitor has an energy density of at least 5 J/cc, more preferably in the range of 5 10 J/cc. The capacitor provides the high drain current as well as pulse power capabilities, which the battery may or may not be able to provide.

In certain preferred embodiments, the electrochemical capacitor system is based on high surface area carbon electrodes and either aqueous or non-aqueous electrolyte. In other preferred embodiments, the electrodes are based on battery active valve-metal oxides such as ruthenium oxide. In some embodiments, the electrodes contain 50 95 wt. % ruthenium oxide and the balance of the valve-metal oxide is vanadium oxide and is capable of supplying 1 3 volts.

In certain embodiments, the power unit is configured as, or integrated with a structural member, and comprises a solar panel containing a laminate of the following layers: at least one thin film photovoltaic cell, optionally, a thermal management device, a thin film battery, a thin film capacitor, a thin film electrochemical capacitor, electronic circuitry, another thin film capacitor, another thin film electrochemical capacitor, another thin film battery, optionally, another thermal management device, and at least one additional thin film photovoltaic cell, the layers being operationally interconnected by the electronic circuitry.

In accordance with certain embodiments, the power system includes a structural member that comprises an exterior surface of a structure, and in some embodiments the laminate is flexible. In some embodiments the laminate is substantially impervious to gaseous species, e.g., a defined gas is retained by and/or chemically benign toward a balloon formed by said laminate to a sufficient extent and for a sufficient time to allow the balloon to be used for a desired purpose. In certain other embodiments, the laminate is at least semi-rigid. In some embodiments at least some components of the laminate are resistant to thermal and electromagnetic radiation, i.e., the component does not transmit heat within that component or deteriorate when exposed to radiation. In some embodiments the thin film capacitor also serves as a radiation shield, and in some embodiments the power unit contains one or more thermally insulating materials.

In certain embodiments, the thermal management device may include a heating element to either heat the power source for functionality or to heat other devices on the spacecraft, if necessary. For example, this option will be useful when collecting cryogenic samples and heating to liquefy for storage or analysis. In certain embodiments, the thermal management device may contain thermoelectric cooling elements to keep the solar cell arrays cool while providing waste heat (resulting from the extraction of heat from the solar cells) to the battery or other parts of the structure. The thermoelectric elements would function similarly to thermoelectric-based household freezers, which keep the inside housing cool and also dissipates the extracted heat to the outside.

In addition to the many applications of the present invention to solving the energy needs for many space exploration and earth orbital operations, the present invention also overcomes many of the shortcomings of terrestrial solar power technologies now employed or previously proposed for military and non-military equipment, in both mobile and non-mobile structures. For example, certain embodiments of the new power system are also applicable to automobiles, homes, office buildings, and especially to isolated structures and those situated in remote locations.

In accordance with certain embodiments of the present invention, a vehicle is provided that includes an above-described solid-state power system in which the solid-state power unit is integral with a structural member of the vehicle, such as an exterior covering or body. In other embodiments the structural member on the vehicle is interior to the vehicle, such as a support or a truss framework. In some embodiments, the vehicle comprises a spacecraft such as a satellite, nanosat, aerobot or hot air balloon, and in some embodiments the spacecraft is designed for deep space exploration and is capable of operating under cryogenic conditions for an extended period of time. In certain embodiments, the vehicle's power system is capable of electrolyzing, compressing, liquefying or freezing a transported or atmospheric gas or of sublimating or boiling a frozen or liquid phase of an atmospheric gas. In certain embodiments the vehicle is an airborne or spaceborne platform, and in certain other embodiments the vehicle is a terrestrial vehicle.

In accordance with still other embodiments of the present invention, a structure comprising an above-described solid-state power system is provided in which the power unit is integral with a structural member of a structure, such as a residential or commercial building or a piece of equipment, especially in an isolated or remote location. In certain embodiments the structural member is flexible, and in others it is rigid or semi-rigid.

Some of the features and advantages of certain embodiments of the present invention include the following: A structurally embedded intelligent power unit (SEIPU) can provide continuous and pulse power to a spacecraft under all environmental conditions. Using a SEIPU instead of a conventional electrical power source offers a way to reduce spacecraft payload and thus, launch cost and at the same time provide a lightweight and longer lasting ultra-thin film power source. A preferred solid-state power source can be assembled that does not deteriorate (i.e., compromise to the cycle life or longevity of the battery) under extreme temperature conditions, and can offer higher cycle life than existing liquid-based systems.

In a preferred embodiment, a SEIPU comprising a rigid photovoltaic power source, a thin film high energy battery (a lithium ion or lithium polymer), a thin film high voltage metallized polymer film capacitor as the energy storage power sources, a layer of heating element, a layer of thermoelectric cooling elements, and a layer of power electronic circuitry that manages various charging and power management schemes, is embedded or is a part of the space platform or satellite to be used for anti-missile launchers, where a source of continuous power is required, but also high pulse power demands for launching missiles.

In another preferred embodiment, a SEIPU comprising a flexible photovoltaic power source, a thin film flexible lithium polymer electrolyte battery, a thin film flexible electrochemical capacitor, thermal management components comprising a heater and/or thermoelectric cooling elements interposed between the solar cells and battery, and a layer of power electronics circuitry that manages various charging and power management schemes, forms the actual flexible inflated balloon or aerobot structure that could be used for atmospheric evaluation or surveillance purposes. A preferred new solid-state power source can undergo fast-charge and discharge without any compromise on the cycle life or performance and has the ability to provide continuous power to on-board instruments and other electronic devices. By including a thermal management system (e.g., embedded heating elements, temperature sensors, and "intelligent" temperature control circuitry), a SEIPU can provide heat to either the power source itself or to predetermined instruments in cryogenic conditions. Still other features of preferred embodiments of the new power system include the ability to provide high power on demand. By careful selection of the SEIPU components and materials, protective heat and radiation shielding for the power sources are obtained without adding separate shielding elements. Another feature of some embodiments of the invention is that a flexible power source can be laminated directly onto the existing structures of a spacecraft; or, alternatively, modular panels of the power units can be readily interconnected through electrical interlocks to form larger panels.

The new power system makes possible for the first time the design and manufacture of modular, high energy power supplies that are comparatively maintenance-free, lightweight, and offer significant space savings. The preferred embodiments of the present system are self-regenerating, with no additional charging cost. A key advantage is that the preferred ultra-thin film components will allow fast-charge capability and will also provide improved charging efficiency, compared to other methods of charging. Consequently, higher cycle life is possible for prolonging satellite or other application usage, thereby reducing the requirement for frequent replacement and lowering overall cost. These and other embodiments, features and advantages of the present invention will become apparent with reference to the following description and drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic cross-sectional view of the SEIPU components according to certain embodiments of the invention.

FIGS. 2A B depict a satellite having solar panels, heaters, batteries, capacitors, and electronic circuitry, in accordance with certain embodiments of the invention.

FIG. 3 is a drawing of a balloon and base structure containing flexible SEIPU panels, in accordance with some embodiments of the invention.

FIG. 4 is a drawing of a SEIPU-containing aerobot that is suitable for use in exploring a planetary atmosphere.

FIG. 5 is a drawing of a miniature satellite (nanosat) as modified according to certain embodiments of the present invention to include SEIPU exterior panels.

FIGS. 6A B depict a lithium polymer electrolyte cell of a lithium polymer electrolyte battery, in accordance with certain embodiments of the present invention A is a perspective view, and B is a cross-sectional view taken along line A--A.

FIG. 7 is a drawing of an enlarged longitudinal cross-section of a metallized film capacitor according to some embodiments of the invention.

FIG. 8 is a partial cutaway view of a rigid square structural panel in accordance with an embodiment of the present invention.

FIG. 9 is an exploded view of the structural panel of FIG. 8, showing the individual components.

FIG. 10 shows a rectangularly designed structural panel in accordance with an embodiment of the present invention.

FIG. 11 is an exploded view of the structural panel of FIG. 10.

FIG. 12 shows several of the panels of FIG. 10 joined together to form a panel array, in accordance with an embodiment of the present invention.

FIG. 13 illustrates a still larger array formed by joining together several of the panel arrays of FIG. 12.

FIG. 14 depicts a spacecraft housing in accordance with an embodiment of the invention, formed by joining together the structural panels according to FIG. 13.

FIG. 15 is a block diagram illustrating one embodiment of the electronic control circuitry for a satellite panel containing an embedded SEIPU.

FIG. 16 is a conceptual drawing of a building, the structural members of which has been modified to include a structurally embedded intelligent power unit according to certain embodiments of the present invention.

FIGS. 17A B illustrate conceptually how an automobile is modified to include a SEIPU according to certain embodiments of the present invention. FIG. 17A shows a partially rolled flexible SEIPU sheet, and FIG. 17B shows two SEIPU sheets according to FIG. 17A, rolled and unrolled, over the windows of an automobile.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

A lightweight, integrated structural panel solid-state power source, called a "Structurally Embedded Intelligent Power Unit" or "SEIPU," can have the shape of a balloon, nanosat, satellite, air/space platform building material, or any of a variety of other structures. For example, the laminated components of a solar power system (photovoltaic solar cells) and advanced energy storage system (batteries and capacitors) are directly embedded into the structural panels of a satellite or configured into the shape of the balloon structure, thereby increasing capacity and rate capability of the respective systems. Such features, which were previously not available simultaneously in the same satellite, allow greater latitude in power management, provide reliable power under any conditions, reduce satellite weight, and provide improved charged acceptance capability. Layers of advanced film solar cells, film batteries (with energy densities in the range 200 250 watt-hours/kilogram (Wh/kg)), film capacitors (with specific energy in the range of 5 10 Joules/cubic centimeters (J/cc)), electrochemical capacitors (with specific energy in the range of 15 100 J/cc, flexible electronics, and ancillary components such as thermal management devices (heating and cooling devices), all configured in an ultra-thin film flexible laminate design, are formed into uniform and flexible structural panels that can be shaped into the outer shell of a space device body or other structure. The outer shell would then generate power to supplement the main solar panel array. Such material will also do the same for mobile terrestrial solar power in military as well as non-military equipment such as automobiles, homes, office buildings, and especially structures situated in remote locations.

Structurally Embedded Intelligent Power Unit (SEIPU). Referring to FIG. 1, a schematic cross-section of a preferred structurally embedded intelligent power unit (SEIPU) 10 is provided, in which the laminated components are shown. The outer layer of this SEIPU consists of the solar array 20 that may or may not have concentrators (i.e. materials or devices having the ability to concentrate the sun intensity). These solar panels are preferably mounted on Kapton.TM. (DuPont, Wilmington, Del.) polyimide flexible sheets or another suitable plastic substrate such as polyester, polypropylene, polyvinylidene fluoride or more preferably polyethylene naphthalate, which offers improved tensile strength and thermal properties. The second layer (underlying layer) is the flexible heater and/or thermoelectric cooling system 30 whose lining is preferably made out of polyimide. The heating means is not limited to resistive-type heating only, but can also be provided by microwave, ultrasonic, infrared or electromagnetic induction-type heating. The heater system, constituting all or part of a thermal management device, provides increased protection to the underlying battery and capacitor power sources as well as the associated electronics from space debris bombardments as well as intense heat. The thermoelectric cooler removes the heat from the solar cells, cools the solar cells for more efficient operation, and at the same time provides the removed heat to the battery. The next layer is the solid-state thin film battery 40 with high energy content. The layer under the battery is the laminated metallized film capacitor 50, which is preferably made of polyvinylidene fluoride (PVDF)-based material (as described in U.S. Pat. No. 6,426,861 and PCT Application No. PCT/US00/11883, which are hereby incorporated herein by reference) with energy densities exceeding 5 J/cc. For example, the film capacitor may comprise a thin coat of a material of high dielectric constant and relatively low electrical properties, such as PVDF, onto a capacitor grade polymer film of lower dielectric constant but higher electrical properties, such as PP, PET, PEN, PPS, PC or PTFE, or copolymers or hybrid polymers formed from such blends. The coating material thickness ranges from 0.1 micron to 25 microns, and the coated substrate thickness ranges from 0.5 micron to 25 microns. The coating can be solvent cast directly onto the polymer substrate, or vapor deposited in an atomized manner, or melt cast directly onto another melt cast substrate, or heat laminated. The coating can be applied to either MDO or TDO substrate polymer film. If an MDO substrate is used, the coated film could be stretched subsequently in the TDO direction, to achie