Abstract:
A method of making a solar cell assembly includes placing backsides of multiple solar cells in contact with a substrate. The solar cells are electrically connected to each other. Heat and pressure are applied to the solar cells and the substrate to simultaneously impress the solar cells into the substrate and bond the solar cells to the substrate.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims priority under the provisions of 35 U.S.C. §119(e) to U.S. Provisional Application No. 61/844,786, filed Jul. 10, 2013. 
     
    
     TECHNICAL FIELD 
       [0002]    The present technology is directed generally to high altitude aircraft with integrated solar cells, and associated systems and methods. 
       BACKGROUND INFORMATION 
       [0003]    Unmanned air vehicles (UAVs) have proliferated recently because they can perform a variety of valuable tasks without incurring the costs and risks associated with a piloted aircraft. Typical UAV tasks include surveillance and communication tasks. However, one drawback with many existing UAVs is that they have limited endurance and can accordingly remain on-station for only a limited period of time. As a result, it can be difficult to provide the foregoing tasks consistently for an extended period of time. 
         [0004]    One approach to addressing the foregoing endurance issues is to provide solar power to a UAV, potentially enabling the UAV to remain on-station for extended periods of time because it generates the power it requires while in flight. However, such systems tend to be heavy and/or expensive, both of which factors make it difficult to provide surveillance, communications, and/or other services at a competitive price. Accordingly, there remain unmet needs for providing long endurance, unmanned air vehicle services at competitive rates. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0005]      FIGS. 1A and 1B  are partially schematic, isometric illustration of unmanned air vehicle systems having solar cells integrated in accordance with an embodiment of the present technology. 
           [0006]      FIG. 2  is a partially exploded illustration of an aircraft substrate having solar cells integrated in accordance with an embodiment of the present technology. 
           [0007]      FIGS. 3A-3D  illustrate a representative process for integrating solar cells with an aircraft substrate in accordance with the present technology. 
           [0008]      FIGS. 4A-4C  illustrate cross-sectional views of substrates having integrated solar cells in accordance with embodiments of the present technology. 
       
    
    
     DETAILED DESCRIPTION 
       [0009]    The present technology is directed generally to high altitude aircraft with integrated solar cells, and associated systems and methods. Specific details of several embodiments of the disclosed technology are described below with reference to particular aircraft configurations and solar cell configurations, but in other embodiments, representative systems can include aircraft and/or solar cells having different configurations than those described below. Several details describing structures and/or processes that are well-known and often associated with high altitude unmanned aircraft and/or solar cell arrangements, but that may unnecessarily obscure some significant aspects of the present technology, are not set forth in the following description for purposes of clarity. Moreover, although the following disclosure sets forth several embodiments of different aspects of the presently disclosed technology, several other embodiments of the technology can have configurations and/or components different than those described in this section. Accordingly, the presently-disclosed technology may have other embodiments with additional elements, and/or without several of the elements described below with reference to  FIGS. 1A-4C . 
         [0010]      FIG. 1A  is a partially schematic side elevation view of an aerospace system  100  that includes an aircraft  120  configured in accordance with an embodiment of the present technology. The aircraft  120  can be particularly configured to fly at very high altitudes (e.g., just within the earth&#39;s atmosphere) under power produced by multiple solar cells  110 . In a particular embodiment, the aircraft  120  includes a narrow, elongated fuselage  121  coupled to one or more wings  122  or other lifting surfaces. The wings  122  can have a high aspect ratio with a significant dihedral or polyhedral e.g., suitable for low-speed, high altitude, long endurance flight. The aircraft  120  can further include a stabilizer boom or empennage  124  carrying one or more stabilizers  123  that provide stability and control functions. The aircraft  120  can still further include a propulsion system  126 , which can in turn include one or more nacelles  129 , each of which houses an electric motor  128  that powers a corresponding propeller  127 . 
         [0011]    Power for the propulsion system  126  is provided by the solar cells  110 . The solar cells  110  can be arranged in a plurality of panels or arrays  111  that are carried by the aircraft  120 . Elements of the aircraft (e.g., the wings  122  and/or the stabilizers  123 ) can include one or more substrates  150 . In particular embodiments, as will be discussed in greater detail below, the solar cells  110  (individually, in panel form, and/or in other forms) can be integrated with the underlying substrate  150  to improve the manufacturability, cost performance, and/or structural efficiency of the aircraft  120 . 
         [0012]      FIG. 1B  is a partially schematic side elevation view of an aerospace system  100  that includes an aircraft  120  having a different configuration than that shown in  FIG. 1A . For example, the propulsion system  126  includes a single nacelle  129  and a single propeller  127 , and the stabilizers  123  are oriented orthogonal to each other. The solar cells  110  are integrated with the aircraft structure in a manner identical to or generally similar to that of the aircraft  120  shown in  FIG. 1A .  FIGS. 2-4C  illustrate further details of representative integration techniques, as described below. 
         [0013]      FIG. 2  is a partially exploded, isometric illustration of a portion of a solar cell assembly  156  configured in accordance with a representative embodiment of the present technology. The assembly  156  can include the substrate  150 , which carries multiple solar cells  110 . Optionally, the assembly  156  can further include a protective cover  153 . The cover  153  can be placed over the outwardly facing first or active surfaces  112  of the solar cells  110  to protect the active surfaces  112  from erosion and/or other environmental degradation factors. The solar cells  110  have an oppositely-facing second or back surface  113  that is attached to the substrate  150 . The substrate  150  can have a composite construction. In a particular embodiment, the substrate  150  can include first and second sheets  151   a ,  151   b  positioned on opposite sides of a central core  152 . In other embodiments, the composite structure of the substrate  150  can have other physical arrangements within the scope of the present technology. 
         [0014]    In a representative embodiment, the solar cells  110  are integrated with the substrate  150  as part of the manufacturing process for forming the substrate  150 . The resulting structure can be a unitary structure, e.g., one for which the solar cells  110  are bonded intimately with the substrate  150 . For example, as shown in  FIG. 2 , the elements forming the assembly  156  can be built up in a stacked fashion on a mold element  140 . The mold element  140  can accordingly include a smooth inner mold surface  141  that receives the stacked components forming the assembly  156 . The surface of the assembly  156  positioned against the inner mold surface  141  forms an outwardly-facing, aerodynamic surface  160  of the aircraft  120  shown in  FIG. 1 . For example, the outwardly-facing surface  160  of the assembly  156  can form the upwardly or outwardly facing exposed portion of the wings  122  and/or the stabilizers  123  shown in  FIG. 1 . A representative process for forming the assembly  156  is described further below with reference to  FIGS. 3A-3D . For purposes of illustration, only portions of representative larger assemblies  156  (e.g., sized to form an entire wing surface or a large section of such a surface) are shown in  FIGS. 3A-3D . 
         [0015]    Beginning with  FIG. 3A , the mold element  140  (and in particular the inner mold surface  141 ) can be shaped in a manner that corresponds to the desired shape of the external surface carrying the solar cells  110 . For example, the inner mold surface  141  can have a profile that produces the curved surfaces of the wings  122  or stabilizers  123  shown in  FIG. 1 . The illustrated inner mold surface  141  can be shaped to produce an upper wing surface, and a complementary mold element (not shown in  FIG. 3A ) can have an inner mold surface for forming the lower wing surface. The resulting wing halves can then be bonded together (e.g., at the leading and trailing edges) to form a complete wing. 
         [0016]    In other embodiments, the surfaces can be formed in accordance with other techniques. For example, the inner mold surface  141  can be flat, and the resulting assembly  156  can be flexible enough to be laid over a set of curved ribs or other supports. The assembly  156  (or multiple bonded assemblies  156 ) can then be post-cured at elevated temperatures to harden the assemblies  156  into the final curved shape. 
         [0017]    In a particular embodiment in which the protective cover  153  is used to protect the solar cells  110 , the cover  153  can be laid face-down on the inner mold surface  141 . In particular embodiments, the mold surface  141  is polished and/or waxed or otherwise treated to create a surface that the cover  153  and/or other elements forming the assembly will not stick to. The protective cover  153  can include a first adhesive  154  facing upwardly, away from the inner mold surface  141 . The protective cover  153  and the first adhesive  154  can be transparent, at least at the wavelengths that activate the electrical current generation function of the solar cells  110 . The solar cells  110  can then be laid face down onto the protective cover  153 , with the active surfaces  112  in contact with the first adhesive  154 . 
         [0018]    The back or second surfaces  113  (facing upwardly in  FIG. 3A ) of the solar cells  110  can carry one or more contacts  115  (e.g., pads, terminals, or other structures) for conveying the electrical current generated by the solar cell  110 . The manufacturing process can include attaching electrical connectors  114  to the back surfaces  113  at the contacts  115  (e.g., via a solder joint or other suitable joint) to transmit current from one solar cell  110  to another. The solar cells  110  can be connected electrically in parallel or in series, depending upon the target voltage and current requirements for a particular array of solar cells  110 . In a typical installation, some solar cells  110  may be arranged in parallel, and others in series. The overall electrical system, which includes the solar cells  110  and the electrical connectors  114 , can also include blocking diodes and/or other circuit elements to control the flow of electrical current among the solar cells  110 . 
         [0019]    In  FIG. 3B , a second adhesive  155  has been applied to the back surfaces  113  of the solar cells  110 . Depending upon the embodiment, the second adhesive  155  can be applied with a brush, and/or can take the form of a film, and/or a spray. In a particular embodiment, (for example, when the second adhesive  155  is sprayed on to the solar cells  110 ), the process of applying the second adhesive  155  can include placing a mask  180  over the assembly  156 . The mask  180  can include mask apertures or openings  181  that align with the back surfaces  113  of the solar cells  110 , allowing the adhesive to deposit on the back surfaces  113 , without depositing on other surfaces (e.g., the inner mold surface  141 ). In other embodiments, the mask  180  can be eliminated and the second adhesive can be applied to the back surfaces  113  and possibly the exposed first adhesive  154  as well. 
         [0020]    In  FIG. 3C , the remaining elements of the assembly  156  have been positioned on the exposed first adhesive  154  (carried by the protective sheet  153  and visible in  FIG. 3B ) and the exposed second adhesive  155  (carried by the back surfaces  113  of the solar cells  110  and visible in  FIG. 3B ). The remaining elements can include a first sheet  151   a , a second sheet  151   b , and a core  152  positioned between the first and second sheets  151   a ,  151   b . The first and second sheets  151   a ,  151   b  can include a fabric, for example, a glass ply fabric or a carbon ply fabric. Each sheet can include a single ply or multiple plies. The fabric can be woven or unwoven, and can be pre-impregnated with an epoxy or another adhesive (in a “pre-preg” form). In other embodiments, the epoxy or other adhesive can be disposed on the first sheet, the second sheet, and/or the core (e.g., using a wetting-out process), prior to consolidating the elements of the assembly  156 . The core  152  can include a foam core, a honeycomb core or another suitable core, made from fiberglass, aluminum, and/or other suitable low cost, lightweight materials. 
         [0021]    In  FIG. 3D , a vacuum bag  170  has been placed over the assembly  156 , and has been sealed to the mold element  140  using a sealant tape  173  or other suitable sealant. Suitable breathers and release films (not shown in  FIG. 3D ) are used to control the flow of epoxy and/or other resins, adhesives, or matrix materials under vacuum. The vacuum bag  170  is attached to a vacuum source  172  via a vacuum line  171 . The manufacturing process can then include drawing a vacuum on the volume interior to the vacuum bag  170 , thereby compressing and consolidating the elements of the assembly  156 . The assembly  156  can, while under vacuum, be exposed to an elevated temperature (e.g., in a suitable oven or autoclave) to cure the epoxy and/or other adhesive materials that form the bonds between the elements of the assembly  156 . 
         [0022]    In a typical arrangement, the entire assembly  156  is bonded and co-cured simultaneously. In particular, the first and second sheets  151   a ,  151   b , the intermediate core  152 , the solar cells  110 , and the electrical connectors  114  (along with other circuit elements, where appropriate) are cured simultaneously. The protective cover  153  can be applied afterwards, or can be co-cured along with the foregoing elements. 
         [0023]      FIGS. 4A-4C  illustrate representative cross-sectional views of completed assemblies  156 . For purposes of illustration, only portions of the assemblies are shown in  FIGS. 4A-4C , and several of the dimensions shown in  FIGS. 4A-4C  have been exaggerated.  FIG. 4A  illustrates a portion of an assembly  156  having a single solar cell  110 . As shown in  FIG. 4A , the pressure (indicated by arrows P) exerted by the vacuum (and/or other suitable arrangement), on the assembly  156  during the manufacturing process by can force the solar cell  110  into the underlying portions of the assembly  156 . In a particular embodiment, the solar cell  110  can accordingly form and simultaneously fill a cell recess  157  that is indented into the first sheet  151   a  and, optionally, the core  152 . For example, when the core  152  is a foam core, it can readily indent under vacuum, and when the core  152  is a honeycomb or more rigid core, the indentation may be limited to the first sheet  151   a.    
         [0024]    During the curing process, the first adhesive  154 , the second adhesive  155  and/or the epoxy and/or other adhesive carried by the first sheet  151   a , consolidate and bond the solar cell  110  firmly to the assembly  156 . In particular, one or more of the foregoing adhesives can bond the second surface  113  of the solar cell  110  to the bottom  159  of the cell recess  157 . One or more of the foregoing adhesives can bond the edge or side surfaces  116  of the solar cell  110  to corresponding walls  161  of the cell recess  157  via strong cohesive and adhesive bonds. Accordingly, the solar cell  110  becomes highly integrated with the assembly  156  that carries it. 
         [0025]    It is expected that the foregoing arrangement will increase the overall strength of the assembly  156 . In particular, the solar cells  110  can carry loads that would normally be carried by the aircraft structure. Such loads include, but are not limited to, compression loads. For example, when the assembly  156  forms part of a wing upper surface, the upper surface can be placed in compression as the wing bends upwardly. Because the solar cells  110  are intimately integrated with the assembly  156  (e.g., via direct contact between the edge/side surfaces  116  of the solar cells  110 , and the adjacent walls  161  of the cell recess  157 ), they can carry the foregoing compression loads. Solar cells that are merely surface mounted on a wing (as is the case with some conventional arrangements) cannot suitably carry such compression loads. 
         [0026]    Because assemblies  156  in accordance with the present technology take advantage of the load-bearing capability of the solar cells  110 , other elements of the aircraft that would otherwise be required to provide this load capability can be reduced or eliminated, thus reducing the overall aircraft weight. For example, in an embodiment for which the assembly  156  forms a portion of a wing surface that is supported internally by ribs or other structural elements, the number of ribs or other structural elements can be reduced to produce a corresponding weight reduction. Reducing the aircraft weight can reduce cost, and/or can increase the endurance and/or the payload capacity of the aircraft. 
         [0027]    In particular embodiments, the solar cells can be bare, monocrystalline solar cells, e.g., cells without encapsulants or other packaging. An advantage of monocrystalline cells (as opposed to thin film cells) is that monocrystalline cells have a high energy concentration and can accordingly produce a greater amount of energy per square meter than can thin film cells. A drawback with monocrystalline cells is that they can be very brittle and are also typically very flat. However, by integrating monocrystalline solar cells with the underlying structure in the manner described above, the brittle characteristics of the solar cell  110  can be mitigated and the support provided by the underlying and integrally connected substrate  150  can allow the solar cells  110  to bend without breaking. In particular, assemblies  156  made in accordance with the techniques described herein can be bent into a curved shape having a 1.5 foot radius, without breaking the solar cells  110  carried by the substrate  156 . Even if a solar cell  110  does break under load, the presence of the underlying structural elements can be sufficient to preserve at least some of the electrical continuity provided by the solar cells  110  and the contacts  114 , so that the solar cells  110  maintain at least some functionality. 
         [0028]      FIG. 4B  is a cross-sectional view of a representative assembly  156  illustrating two solar cells  110  connected by an electrical connector  114 . In one aspect of this embodiment, each solar cell  110  forms a corresponding cell recess  157  as it is compacted during the manufacturing process. An inter-cell recess  158  is formed between neighboring cell recesses  157 . An inter-cell recess bottom  162  is defined at least in part by the protective cover  153 . If the protective cover  153  is not included in the assembly  156 , the inter-cell recess bottom  162  can be defined at least in part by the electrical connector  114 . In a particular embodiment, the inter-cell recess  158  is present only directly above the electrical connector  114 . At positions away from the electrical connector  114  (e.g., into and out of the plane of  FIG. 4B ), the first sheet  151   a  and the core  152  can project upwardly between the solar cells  110  so that the outwardly facing surface  160  of the overall assembly  156  remains generally flat. 
         [0029]      FIG. 4C  illustrates an assembly  156  having surface contours in accordance with still further embodiments of the present technology. For purposes of comparison, the inter-cell recess bottom  162  shown in  FIG. 4B  is also shown in  FIG. 4C , along with other potential contours for the inter-cell recess bottom, illustrated as first-third recess bottoms  162   a - 162   c . Each of these inter-cell recess bottom contours  162   a - 162   c  can result from placing the solar cells  110  closer to each other (thus reducing the depth of the corresponding inter-cell recess  158 ), up to the limit indicated by the third recess bottom contour  162   c , which is generally flat. In still further embodiments, the assembly  156  can include an additional filler material  164 , which is placed within the inter-cell recess  158  to reduce the aerodynamic discontinuity presented by the recess, and thereby improve aerodynamic performance. However, in other embodiments, the roughness provided the inter-cell recesses  158  (which may not be visible to the naked eye) can provide aerodynamic advantages, e.g., by delaying flow separation over the wing surface. Accordingly, the size of the inter-cell recesses  158 , and whether and to what extent the inter-cell recesses  158  are filled, can be selected based at least in part on the Reynold&#39;s number of local flow. 
         [0030]      FIGS. 1A-4C  and the associated discussion above describe in general terms techniques and corresponding components used to form load-bearing aircraft structures that include solar cells. The following examples provide additional details of particular embodiments for representative components and techniques. 
       Solar Cells 
       [0031]    Representative solar cells in accordance with embodiments of the present technology include polycrystalline silicon cells, monocrystalline silicon cells, gallium arsenide cells, and/or other suitable solar cell constructions. In particular aspects of these embodiments, the solar cells are bare, e.g., without plastic or other mounting substrates and hardware to reduce weight and to improve the integration of the solar cells with the underlying aircraft structure. In a representative embodiment, the solar cell  110  can weigh approximately 500 grams per square meter and the underlying structure can weigh approximately 150 grams per square meter, producing a total assembly weight of approximately 650 grams per square meter. This is expected to be a significant weight reduction compared to existing solar aircraft structures which are typically at least 50% heavier. Other conventional structures, which may include solar cells sandwiched between two sheets of Mylar®, may be relatively light, but are expected to be significantly less durable than structures formed in accordance with embodiments of the present technology. 
       Sheets 
       [0032]    The first and second sheets can include unwoven or woven fabric, for example, a plain weave fiberglass fabric having a weight of approximately 0.7 ounces per square yard to approximately 5.6 ounces per square yard. In other embodiments, the sheets can include a plain weave carbon fiber having a weight in the same or a similar range. In still further embodiments, the sheets can have other constructions. For example, the sheets can be formed from Kevlar® or other high strength lightweight materials. In addition to these criteria, the sheets can be selected for low cost, tolerance for environmental factors, ease of manufacturability, and formability, among others. In any of these embodiments, as discussed above, the sheets can be pre-impregnated with epoxy or another suitable adhesive, or the adhesive can be applied to the sheets at the time the assembly is manufactured. Suitable epoxies include a 105 Epoxy System available from West Systems of Bay City, Mich., having a cure time of from about 6 hours to about 24 hours based on the selection of the associated hardener. Another representative epoxy includes the MGS L285 Epoxy System, available from MGS Kunstharz Probukte Gmbh of Stuttgart, Germany. Suitable cure temperatures for these materials are within the ranges typically provided by the manufacturer. 
       Core 
       [0033]    Suitable cores for placement between the sheets described above can include foam cores, e.g., Rohacell Foam Core available from Evonik Industries of Essen, Germany, or Airex Foam Core available from Airex AG of Sins, Switzerland. Other suitable cores include Nomex honeycomb core and aluminum honeycomb core, both available from HexCel of Stamford, Conn. 
       Protective Cover 
       [0034]    Suitable protective covers include Coverite Microlite materials, available from Coverite of Champaign, Ill., and So-Lite available from Mountain Models of Appleton, Wis. 
         [0035]    During a typical bonding process, the epoxy bonds quickly to the back surface  113  of the solar cell  110 , thus preventing the epoxy from flowing, seeping, or leaking around or beyond the sides  116  of the solar cell where it might interfere with the ability of the solar cell  110  to receive solar radiation. Because the typical manufacturing process includes placing the assembly under vacuum during curing, bubbles tend not to form in the assembly  156 . This can be a significant advantage over other bonding techniques because such bubbles, if present during manufacture, can expand and burst when the aircraft reaches high altitude, thus damaging or destroying the surrounding structure. 
         [0036]    One feature of at least some of the foregoing embodiments described above with reference to  FIGS. 1A-4C  is that the aircraft can include solar cells that are intimately integrated with the aircraft structure. In particular, the solar cells can be bonded to the aircraft structure at the same time the aircraft structure itself is formed and cured. This arrangement can increase the strength of the mechanical and/or chemical bonds between the solar cells and the underlying structure, thus allowing the solar cells to form load-bearing portions of the overall structure. As discussed above, this in turn can reduce the requirements for other load-bearing structures of the aircraft which in turn can reduce the overall weight of the aircraft. 
         [0037]    From the foregoing, it will be appreciated that specific embodiments of the present technology have been described herein for purposes of illustration, but that various modifications may be made without deviating from the technology. For example, the protective cover material, the foam core, the first and second sheets, and/or other elements of the overall structure can have compositions different than those expressly disclosed herein. In particular embodiments, the solar cells can be carried by other structures, in addition to or in lieu of the wings and stabilizers described above. Such structures can include fuselage or empennage surfaces, and/or other surfaces that are sufficiently exposed to solar indication. The aircraft can be ground-launched, dropped from other aircraft, and/or deployed in other manners. Particular embodiments of solar cells integrated with corresponding structure were described in the context of high altitude, long endurance aircraft. Other embodiments of the presently-disclosed technology can be implemented in varying manners on other aircraft and/or in the context of other structures, e.g., blimps, spacecraft, and/or other applications where low weight and low cost can provide significant advantages. 
         [0038]    Certain aspects of the technology described in the context of particular embodiments may be combined or eliminated in other embodiments. For example, the protective coating can be eliminated in at least some embodiments for which such a coating is unnecessary, and/or the benefit of such a coating does not outweigh the cost and/or the weight of the coating. Further, while advantages associated with certain embodiments of the present technology have been described in the context of those embodiments, other embodiments may also exhibit such advantages, and not all embodiments need necessarily exhibit such advantages to fall within the scope of the present technology. Accordingly, the present disclosure and associated technology can encompass other embodiments not expressly described or shown herein. The following examples provide additional embodiments of the present technology.