Thermoplastic aerostructure with localized ply isolation and method for forming aerostructure

Abstract

To provide an efficient process for connecting structural composite components of a structure without causing damage to the components of the structure.SOLUTION: There is provided a composite wingbox structure 10 formed of reinforced thermoplastic. The composite structure 10 includes: carbon fiber reinforcement; and a plurality of insulation elements 60 to localize heat formed during a process of manufacturing the structure 10. The process of manufacturing the wingbox 10 includes the steps of: interleaving a series of insulations elements 60 within a plurality of thin films 50 and consolidating the insulation elements 60 and the thin films 50 to form a laminate 40. The laminate 40 is then aligned with a support structure 15 such that the insulation elements 60 overlie the support structure 15. The laminate 40 is then fused to the support structure 15 using a non-contact heating process, such as inductive welding.SELECTED DRAWING: Figure 1

Description

[Technical Field]

[0001] This application claims priority to U.S. Provisional Patent Application No. 62 / 666,193, filed May 3, 2018, the disclosure of which is incorporated herein by reference in its entirety.

[0002] The present invention relates to the field of composite materials. In particular, the present application relates to structures formed from a plurality of components made of composite materials. The present invention is particularly intended for application in the field of aircraft structures formed from a plurality of composite elements. [Background technology]

[0003] Composite materials have been used in a wide variety of applications where the benefits of lightweight, high-strength materials outweigh material cost. For example, historically, aircraft structures have been formed from lightweight metals, such as aluminum and, more recently, titanium. However, modern aircraft are constructed largely from composite materials. A commonly used material in the aerospace industry is carbon fiber reinforced thermosetting plastic. Such materials allow for the formation of complex structures that, once hardened, permanently retain their shape. However, this advantage limits the ability to fuse the formed structure with separate structures. Instead, the separate elements are connected using separate connectors, e.g., fasteners. While structural elements formed from reinforced thermoplastic materials can be connected without separate fasteners, the process of melting the elements requires heating the components above temperatures that could cause damage to the structure. Therefore, an efficient process is needed to connect composite components of a structure without damaging the structural components. Summary of the Invention [Means for solving the problem]

[0004] In view of the above, according to one aspect, the present invention provides a laminate structure configured to localize heat generated during an induction welding process. In particular, the laminate is configured to localize heat to specific regions of one or more specific plies of a multi-layer carbon fiber reinforced thermoplastic composite laminate. According to one aspect, the heat is localized to specific regions of specific plies by implementing multiple layers of insulating elements configured and embedded in a pattern within the layers of the laminate. Optionally, the laminate can be implemented in an aircraft structure, such as a wing box.

[0005] According to another aspect, the present invention provides a composite aircraft structure having a plurality of longitudinally elongated spars and ribs interconnected to the plurality of spars to form a grid, each of the spars and ribs having an upper element with a layer of unidirectional or woven carbon fiber reinforced thermoplastic material forming a weld zone. A composite skin is connected to the grid of spars and ribs. The laminate includes at least five plies of carbon fiber reinforced thermoplastic material. In a first ply, the carbon fibers are oriented in a first direction. In a second ply, the carbon fibers are oriented in the first direction, and the first ply overlies and is directly connected to the second ply. A third ply has carbon fibers oriented in a second direction transverse to the first direction, and a fourth ply has carbon fibers oriented in the second direction, and the third ply overlies and is directly connected to the fourth ply. A fifth ply has carbon fibers oriented in a direction transverse to the carbon fibers of the upper elements of the spars and ribs. The laminate also includes a plurality of elongated insulating elements formed of an electrically insulating material. Each of the insulating elements has a length and a width, with at least some of the insulating elements having a length that substantially exceeds the width. The insulating elements form a grid configured substantially similar to the grid formed by the ribs and spars, such that the insulating elements overlie the upper elements of the ribs and spars. The grid of insulating elements is disposed between the second and third layers. Optionally, each insulating element comprises glass fibers embedded in a matrix of thermoplastic material. The insulating elements may also include a plurality of elongated spar insulators overlying the spars and a plurality of elongated rib insulators overlying the ribs, the rib insulators being shorter than the spar insulators. The rib insulators may extend between the spar insulators without substantially overlying the spar insulators. The spar insulators may also optionally be spaced apart from one another.

[0006] Optionally, the composite skin may include an additional layer of carbon fiber reinforced thermoplastic material, the grid of insulating elements may form a first grid, and the plurality of insulating elements may form a second grid configured substantially similarly to the first grid of insulating elements, the second grid being disposed between the additional layers of carbon fiber reinforced thermoplastic material such that the second grid is aligned with the first grid.

[0007] The aircraft structure may have a second composite skin connected to a grid of the ribs and spars, and each spar and rib may include a lower element having a layer of unidirectional or woven carbon fiber reinforced thermoplastic material forming a lower weld zone, and the second skin may be connected to the lower element of the ribs and spars.

[0008] According to yet another aspect, the present invention provides an aircraft structure in which a plurality of insulating elements form a grid, the insulating elements being configurable to define a plurality of openings.

[0009] According to another aspect, the present invention provides an aircraft structure formed of a thermoplastic material comprising a polyaryletherketone class semi-crystalline thermoplastic material.

[0010] According to another aspect, the present invention provides an aircraft structure formed from a thermoplastic material having a melting point above 500 degrees Fahrenheit.

[0011] According to yet another aspect, an aircraft structure is provided, in which a laminate includes an insulating layer and a structural layer, the insulating layer configured such that when the bottom layer is heated above its melting point by an induction welding head, a thermoplastic material of the structural layer remains below its melting point.

[0012] According to yet another aspect, an aircraft structure is provided, in which a laminate includes insulating layers and structural layers, the insulating element configured such that a bottom layer of the structural layers has a melting point, and the insulating element is configured to prevent the structural layers other than the bottom layer from heating above the melting point when the bottom layer is exposed to an electromagnetic field sufficient to heat a portion of a bottom of the laminate above the melting point.

[0013] Similarly, the present invention also provides an aircraft structure having a laminate formed of a plurality of structural layers and a plurality of insulating elements embedded within a bottom layer, the thermoplastic material of the bottom layer having a melting point, and the insulating elements configured to prevent melting of the thermoplastic material of the structural layers when the bottom layer is welded to the separate aircraft structure by induction welding.

[0014] According to yet another aspect, the present invention provides a composite aircraft structure formed of a plurality of longitudinally elongated spars, a plurality of elongated ribs connected to the plurality of spars to form a grid, and a composite skin. The composite skin includes a plurality of structural layers, a plurality of insulating elements, and a welded layer. The structural layers are each formed of a unidirectional or woven carbon fiber reinforced thermoplastic material. The insulating elements are formed of an electrically insulating material and form a plurality of insulating grids, which are aligned with one another and disposed between the structural layers. The welded layers are also formed of a unidirectional or woven carbon fiber reinforced thermoplastic material. The aircraft structure also includes a plurality of connecting elements, each having a layer of unidirectional or woven carbon fiber reinforced thermoplastic material. Each connecting element is connected to one of the spars or one of the ribs. The carbon fibers of the welded layers are oriented in a first direction, and the carbon fibers of the connecting elements are oriented in a second direction transverse to the first direction. The welded layers overlie and are connected to the connecting elements.

[0015] The present invention further provides a composite laminate for use in aircraft structures, the laminate including a plurality of structural layers, a plurality of insulating elements, and a welded layer. The structural layers are each formed of a unidirectional or woven carbon fiber-reinforced thermoplastic thin film, with the fiber orientation of at least a portion of the structural layers transverse to the fiber orientation of adjacent structural layers. The insulating elements are formed of an electrically insulating material and form a plurality of insulating grids. The insulating grids are aligned with one another and embedded between adjacent structural layers whose carbon fiber orientations are transverse to one another. The welded layer can also be formed of a unidirectional or woven carbon fiber-reinforced thermoplastic thin film.

[0016] The present invention further provides a method for forming a composite wing box. The method includes providing a plurality of spars, connecting a plurality of ribs to the spars, and forming a composite skin. The composite skin includes a plurality of structural layers, a plurality of insulating layers, and a welded layer. Each structural layer is formed of a unidirectional or woven carbon fiber reinforced thermoplastic material. Each insulating layer is formed of a plurality of spaced apart electrical insulating elements. Each insulating element has a length and a width, the length being substantially greater than the width. The insulating layers are aligned with one another and disposed between the structural layers. The insulating layers are configured such that the insulating elements overlie the spar and ribs when the composite skin overlies the spar and ribs. The welded layer is formed of a unidirectional or woven carbon fiber reinforced thermoplastic material. The method includes placing the composite skin over the ribs and spars such that the layers of the composite skin overlie the ribs and spars. The composite skin is then welded to the ribs and spars. The welding process includes bringing an induction welding head over the composite laminate adjacent the rib and spar, where the induction welding head applies an electromagnetic field through the composite laminate to heat the weld layer above the melting point of the thermoplastic material of the weld layer, and where the insulating layer prevents the structural layer from heating above the melting point of the thermoplastic material.

[0017] Optionally, the method can include connecting a plurality of connecting elements to the ribs and the spars. Each connecting element can include a layer of unidirectional or woven carbon fiber reinforced thermoplastic material, and the welding step can include welding the welded layer to the connecting element. Also, the carbon fibers in the welded layer can be oriented in a first direction, and the carbon fibers in the connecting elements can be oriented in a second direction. Connecting the plurality of connectors can include connecting the connectors such that the second direction is transverse to the first direction.

[0018] According to yet another aspect, the present invention provides a method including providing a composite laminate having a plurality of insulating layers, the providing the composite laminate including arranging the plurality of insulating elements into a grid.

[0019] According to yet another aspect, the present invention provides a method including providing a composite laminate having a plurality of insulating layers, the method including configuring the plurality of insulating elements such that a plurality of openings are formed between the insulating elements.

[0020] According to yet another aspect, the present invention provides a method comprising forming the composite skin, the method comprising providing a plurality of pairs of the structural layers, wherein each structural layer of a pair is adjacent to and directly connected to the other structural layer of the pair, and the two structural layers of the pair are oriented such that the fibers of both layers are substantially parallel. Adjacent pairs of structural layers are oriented such that the fibers of one pair are transverse to the fibers of the adjacent pair. Optionally, the method includes providing the insulating layer by interposing an insulating layer between the one pair and the adjacent pair. An insulating layer can also be interposed between each adjacent pair of structural layers.

[0021] According to another aspect, a method is provided for forming a composite laminate for an aircraft structure. The method includes providing a plurality of structural layers, each formed of a unidirectional or woven carbon fiber reinforced thermoplastic film; and providing a plurality of insulating elements formed of an electrically insulating material, each insulating element having a length and a width, the length being substantially greater than the width. The method further includes fabricating a plurality of insulating layers, each layer formed in a pattern in which the insulating elements are spaced apart by configuring the insulating elements. The insulating layers are positioned between the structural layers by aligning the pattern with each insulating layer. The structural layers are configured so that the bottom layer forms a weld layer. The method includes integrating the structural layers and the insulating layers by applying heat and pressure to form a composite laminate in which the insulating layers are embedded within the structural layers. Optionally, the configuring step includes configuring the insulating elements so that the insulating elements extend between the spaced apart insulating elements. The configuring step can also include configuring the insulating elements so that there is a gap between each individual insulating element.

[0022] According to yet another aspect, a method is provided that includes placing the consolidated laminate over a tie layer of unidirectional carbon fiber reinforced thermoplastic material such that the fibers of the weld layer are oriented transverse to the fibers of the tie layer; and inducing an electromagnetic field that penetrates the thickness of the laminate and the tie layer to heat the weld layer and the tie layer sufficiently to weld the weld layer and the tie layer together.

[0023] According to another aspect, a method is provided that includes inducing the electromagnetic field by moving an induction coil over the stack along a pattern of the insulating layers in the stack.

[0024] The present invention further provides a method comprising inducing the electromagnetic field by concentrating the field within a width of a plurality of insulating elements, optionally wherein the interconnect layer is rigidly coupled to the aircraft structure such that the inducing comprises welding the laminate to the interconnect layer, thereby fixedly coupling the laminate to the aircraft structure.

[0025] According to yet another aspect, a method is provided in which the aircraft structure has a plurality of structural elements forming a pattern, the method comprising aligning the pattern of the structural elements in the laminate with the pattern of the insulating layer. The insulating layer can be formed of insulating elements arranged into a grid. Optionally, arranging the insulating elements in a pattern can include configuring the insulating elements such that a plurality of openings are formed between the insulating elements.

[0026] While methods and apparatus are described herein with reference to several embodiments and exemplary drawings, those skilled in the art will understand that the inventive methods and apparatus for sorting items using dynamically reconfigurable sorting arrays are not limited to the described embodiments or drawings. It should be understood that the drawings and their detailed description are not intended to limit the embodiments to the particular forms disclosed. Rather, the intent is to cover all modifications, equivalents, and alternatives falling within the scope of the method and apparatus for sorting items using one or more dynamically reconfigurable sorting arrays as defined by the appended claims. Any headings used herein are merely organizational and are not intended to limit the scope of the description or the claims. As used herein, the words "can" and "may" are used in a permissive sense (i.e., in the sense of possibility) rather than a mandatory sense (i.e., in the sense of must). As used herein, the words "include," "including," and "includes" mean "including, but not limited to." [Brief explanation of the drawings]

[0027] The foregoing summary, as well as the following detailed description of the preferred embodiments of the present invention, are best understood when read in conjunction with the accompanying drawings. [Figure 1] FIG. 1 is an exploded perspective view of a wing box implementing aspects of the present invention. [Figure 2] FIG. 2 is a diagrammatic representation of a system forming the wingbox of FIG. [Figure 3] FIG. 3 is an enlarged perspective view of a portion of the wing box of FIG. [Figure 4] FIG. 4 is a schematic side view of the wing box portion illustrated in FIG. DETAILED DESCRIPTION OF THE INVENTION

[0028] Referring now generally to the drawings, and particularly to FIG. 1 , a composite aircraft structure is generally designated 10. In the present example, the aircraft structure 10 is a wing box having an upper skin 40 and a lower skin, both of which are supported by a support structure 15. The upper and / or lower skins are formed of a carbon fiber reinforced thermoplastic material. The composite skin includes insulating elements that allow the skin to be efficiently welded to the support structure without heating a significant portion of the skin, which could be damaged if heated. In particular, the skin 40 is formed of a laminate configured to localize the heating action produced by an induction welding head and limit heating of the skin, such that only the portion of the skin within the weld zone is heated to the welding temperature.

[0029] Details of the aircraft structure will now be described in greater detail with reference to Figure 1. In Figure 1, the aircraft structure implements the laminate within a wing box structure. However, it should be understood that this is merely an exemplary structure. The laminate can be implemented within a variety of composite structures. Additionally, the laminate has particular application in aircraft structures, but is not limited to use in aircraft structures.

[0030] The aircraft structure 10 of FIG. 1 is a wing box, including a support structure 15 that is an inner layer of the wing box and a cover that overlies the support structure. In the present example, the cover 40 is referred to as a skin and is shown in FIG. 3. For clarity, the skin is shown in FIG. 1 with the structural layers separated from the insulating layers, and only one structural layer (designated 50) is shown. FIG. 1 also illustrates the assembly with one skin overlying the top of the support structure. However, it should be understood that the assembly 10 can also include a lower skin that covers the bottom of the support structure. The lower skin can be formed from a laminate similar to the laminate of the upper skin 40. The top and / or bottom skins can also wrap around the leading and / or trailing edges of the support structure to form a closed assembly.

[0031] The support structure 15 can have any number of configurations depending on the intended use of the structure. In the present example, the support structure 15 is formed of a plurality of elongated spars and ribs. A forward spar 20 extends the entire length of the wing box. A first end of the forward spar is connected to a first end rib 30, and a second end of the forward spar is connected to a second end rib 30, such that the forward spar spans the entire distance between the two end ribs. The end ribs 30 extend the entire width of the wing box and are oriented transversely to the forward spar 20. The forward spar 20 is an elongated beam having a central web of height and thickness, the height being substantially greater than the thickness. The web of the forward spar can be implemented in any variety of beam configurations, such as a box beam, channel beam, or I-beam. In the present example, the spar can be a channel beam with the web extending between an upper surface 22 and a lower surface 24. As will be described below, the upper and lower surfaces 22, 24 may form connecting surfaces that connect the skin 40 to the spar.

[0032] An aft spar 25, spaced apart from the forward spar 20, is also connected to the two end ribs 30 so that the forward spar spans the entire length between the two end ribs. The aft spar 25 can be configured substantially similar to the forward spar. In the present example, the aft spar 25 is substantially similar to the forward spar, except that the aft spar is shorter than the forward spar. Thus, the interconnected spars 20, 25 and end ribs 30 generally form the sides of a box-shaped support structure 15. The support structure 15 can also include multiple additional elements that stiffen and / or strengthen the structure. For example, the support structure 15 can include additional spars that extend the entire length of the structure. In the present example, the support structure 15 includes multiple intermediate ribs 35.

[0033] In this embodiment, the ribs 30, 35 are tapered in height to form a tapered wing box. In particular, the height of the ribs 30, 35 adjacent the forward spar 20 is greater than the height of the ribs 30, 35 adjacent the aft spar 25. The intermediate ribs 35 are spaced apart from the end ribs 30 and from each other. The ribs 30, 35 may be configured as elongated beams similar to the spars 20, 25 described above. In the present example, the ribs 30, 35 are configured similar to channel beams, with a central web extending between transverse upper and lower legs. Also, similar to the spars, the upper and lower legs of the ribs 30, 35 may form surfaces that contact the skin 40 and provide connection points between the skin and the support structure 15. Specifically, the end ribs 30 may include upper and lower flanges 32, 34 that form contact surfaces, and the intermediate ribs 35 may include upper and lower flanges 37, 38 that form contact surfaces.

[0034] Referring now to Figures 1 and 3-4, the laminate forming the skin 40 will be described in greater detail. The laminate includes multiple structural layers 50, 52, 54, 56, and 58 with multiple insulating elements 60 embedded between one or more of the structural layers. Each structural layer has reinforcing elements embedded in a matrix material. Depending on the application, the reinforcing elements may be any type of reinforcing material. By way of example, the reinforcing elements may be elongated strands or fibers of glass or carbon. For example, exemplary carbon fibers are continuous, high-strength, high-strain-rate PAN-based fibers with tows of 3,000 to 12,000 fibers. These reinforcing fibers can be surface-treated and sized to improve interlaminar shear properties with the matrix material. However, it should be understood that these materials are intended as exemplary. Other materials may be utilized depending on the intended application of the laminate.

[0035] The reinforcing elements are embedded in a matrix material, e.g., a polymer. Depending on the application, any type of polymer can be used for the matrix material, including amorphous, crystalline, and semi-crystalline polymers. In the present example, the matrix material is a thermoplastic material, e.g., a thermoplastic elastomer. More specifically, the thermoplastic material is a semi-crystalline thermoplastic material. In particular, the thermoplastic material may be a thermoplastic polymer of the polyaryletherketone (PAEK) family, including, but not limited to, polyetheretherketone (PEEK) and polyetherketoneketone (PEKK).

[0036] As mentioned above, the structural layers 50, 52, 54, 56, 58 may be composite materials, which may be carbon fiber reinforced thermoplastic materials. In particular, the thin films may be thermoplastic prepregs, which are thin films in which reinforcement materials have been pre-impregnated with resin. For example, the prepregs may be thermoplastic prepregs manufactured by coating reinforcement fibers with a thermoplastic matrix. Such prepreg thin films can be reheated and reformed by heating the thin film above the melting point of the thermoplastic matrix. Some exemplary prepreg materials that can be used to form the structural elements 25, 26 include, but are not limited to, CETEX, such as those sold under the product names TC1200, TC1225, and TC1320, manufactured by TenCate Advanced Composites USA (Morgan Hill, California, USA). TC1200 is a carbon fiber reinforced semi-crystalline PEEK composite material with a glass transition temperature (T g )143°C / 289°F, melting point (T m ) 343°C / 649°F. TC1225 is a carbon fiber reinforced semi-crystalline PAEK composite with T g 147℃ / 297℃, T m TC1320 is a carbon fiber reinforced semi-crystalline PEKK composite with a T g150℃ / 318℃, T m is 337°C / 639°F.

[0037] Referring again to FIG. 1 , the stack 40 has multiple structural layers configured to carry structural loads. Multiple insulating layers 60 are embedded within the structural layers. While the insulating layers can be configured to carry structural loads, the insulating layers in the current example are configured to insulate the stack from heating without carrying significant structural loads (if any). To illustrate the details of the insulating layers 60, the insulating layers 60 are shown in FIG. 1 separate from the structural layers of the stack, as well as the single structural layer 50 of the stack. In the illustrated embodiment, FIG. 1 illustrates multiple aligned, overlapping insulating layers 60, but it should be understood that the number and location of these insulating layers will vary depending on the application.

[0038] The insulating layer 60 is embedded in a matrix material, preferably a thermoplastic material that can be heat-fused with the other layers of the laminate. By heat-melting the matrix material of the insulating layer, the insulating layer is formed integrally with the laminate. In particular, in the present example, the insulating layer is formed of a composite material having a matrix substantially similar to the matrix material of the structural layer. For example, the resistive element is carbon fiber and the matrix material is a thermoplastic material, such as a semi-crystalline thermoplastic material of the polyaryletherketone (PAEK) family.

[0039] The insulating layer 60 provides electrical insulation between portions of the structural layers when an electromagnetic field is applied to the stack adjacent to the insulating layer 60. The insulating layers 60 can comprise any type of electrically insulating material. Preferably, these insulating layers 60 comprise a thermoplastic material. In the present example, the insulating layer comprises one or more thin films of a glass-reinforced thermoplastic material.

[0040] The insulating layer 60 is configured to insulate specific zones or regions of the laminate from the heating process. It should be understood that while the insulating layer 60 limits heating of the structural layers, the insulating layer does not limit heating by providing significant thermal insulation. Instead, the insulating layer selectively isolates various regions from being heated by induction heating by providing electrical insulation.

[0041] For example, during the process of welding the skin 40 to the support structure 15, the insulating layer limits the heat generated during the welding process so that only selected plies or plies of the laminate are heated, and only certain areas of the selected plies or plies are heated. To this end, the insulating layer 60 is configured to form a pattern corresponding to the portions of the laminate intended to be heated. Specifically, the insulating layer 60 is configured to form a pattern corresponding to the portions of the laminate intended to be welded. In particular, the pattern can correspond to the areas of the laminate that will contact the portions to be welded. For example, the insulating layer 60 is formed to correspond to the pattern formed by the top surfaces (e.g., top surfaces 22, 27, 32, and 37) of the spars 20, 25 and the ribs 30, 35.

[0042] Each insulating layer 60 can be formed from a single piece of insulating material configured in a desired pattern. In some cases, the pattern can be solid, with the insulating material simply forming blocks, stripes, or similar patterns. However, in the present example, the insulating layer 60 forms a pattern with a plurality of openings so that the area over which the insulating layer 60 overlies the stack is less than half the area of the stack. In some cases, the insulator forms a pattern with sufficient open area so that the area over which the insulator overlies the stack is less than 40% of the total area of the stack. Furthermore, in some applications, the insulating layer forms a pattern with sufficient open area so that the area over which the insulator overlies the stack is less than 30% of the total area of the stack. In some applications, it may be desirable to configure an insulating layer with sufficient open area so that the area over which the insulator overlies the stack is less than 20% of the total area of the stack. Furthermore, the insulating layer in the exemplary structure above forms a pattern that covers at least about 5% of the total area of the stack.

[0043] Referring again to FIG. 1 , in the present example, the insulation layer 60 forms a grid pattern of intersecting stripes of insulating material. Specifically, the insulation layer 60 is formed of a plurality of elongated insulation elements having a length and a width, the length being substantially greater than the width. An exemplary material for forming the insulation elements is fiberglass tape embedded in a thermoplastic material. A first group of insulation elements forms spar insulation elements 70 that extend at least substantially the entire length of the spars 20, 25. As shown in FIG. 1 , the spar insulation elements 70 are laid so that a first spar insulation element overlies the upper surface 22 of the forward spar 20 and a second spar insulation element 70 overlies the upper surface 27 of the aft spar 25. A pair of end rib insulation elements 80 are laid adjacent the ends of the aft spar so that the end rib insulation elements overlie the upper surfaces 32 of the end ribs. Similarly, a plurality of insulation elements 82 extend between the spar insulation elements 72 and overlie the intermediate rib 35. It will be seen that the ribs 30, 35 extend aft beyond the aft spar 25. Thus, the intermediate rib insulation elements can extend beyond the spar insulation elements 70 and overlie the aft spar 25. The rib insulation elements 82 and the spar insulation elements 70 overlap to form an insulation element having a double thickness. However, in the present example, the insulation elements of the insulation layer do not substantially overlap, so the entire insulation layer is a single layer thick. Therefore, to prevent significant overlap of insulation elements, the intermediate rib insulation elements are formed in two segments. A first segment 82 extends beyond the intermediate spar and spans the entire space between the spar insulation elements 70, while a second segment 84 extends beyond the aft end of the intermediate rib 35 such that the second segment extends away from the spar insulation elements 70 overlying the aft spar.

[0044] From the above, it can be seen that the insulating layer 60 forms a pattern with a plurality of open areas. In particular, the insulating layer forms a pattern that forms a grid connecting the laminate 40 with the tops of the ribs and spars.

[0045] In some applications, it may be desirable to vary the pattern or size of the various insulating layers so that they insulate various plies of the stack. However, in the present example, each of the insulating layers 60 is substantially similar. Specifically, each insulating layer 60 has a similar size and pattern. Also, as shown in FIG. 1, each of the insulating layers 60 of the stack is aligned so that the insulating elements forming the pattern of each layer overlie corresponding elements in the other insulating layers.

[0046] As described above, the structural layers 50, 52, 54, 56, and 58 and the insulating layer 60 can each comprise a layer of reinforced thermoplastic composite material. Therefore, the structural layers and the insulating layers can be joined together by fusing the layers together. Specifically, the layers can be consolidated by applying sufficient heat above the melting point of the thermoplastic matrix and applying sufficient pressure to fuse the layers together. In this manner, the insulating layer 60 is embedded within the laminate between the structural layers.

[0047] The number of structural plies and their orientation within the laminate will vary depending on the application. Additionally, the orientation of the plies and the placement of the insulation layers can be varied depending on the desired thermal insulation. Figures 3 and 4 show an exemplary layup of the laminate 40. However, it should be understood that the number of plies, the orientation of the plies, and the number and location of the insulation layers 60 are merely exemplary. The present invention is not limited to the layup shown in this exemplary laminate.

[0048] The laminate includes ten structural layers, designated 50a, b, 52a, b, 54a, b, 56a, b, and 58a, b. As noted above, each structural layer may be a thin film or layer of composite material, such as unidirectional or woven carbon fiber-reinforced thermoplastic tape. The fiber orientation of each structural layer may vary to provide strength in multiple directions. However, in the present example, the layers may be laid in pairs, with the fiber orientations of both layers in the pair being parallel. For example, the first layer of the laminate, designated 50a, has a 90° fiber orientation, and the second layer of the laminate, designated 50b, also has a 90° fiber orientation. These first two layers, 50a, 50b, form an adjacent pair and are bonded together. The third and fourth structural layers comprise a second pair of layers, 52a, 52b, with each layer of this second pair having a 0° fiber orientation. The fifth and sixth structural layers comprise a third pair of plies 54a, 54b. Each ply of this third pair has a 45° fiber orientation. The seventh and eighth structural layers comprise a fourth pair of plies 56a, 56b. The laminate is formed symmetrically about its midline, which in this example is the third pair of plies. Therefore, the fourth pair of structural plies 56a, 56b can have a fiber orientation similar to that of the second pair of structural plies (e.g., 0°). Similarly, the ninth and tenth plies form a fifth pair, each ply of this fifth pair can have a 90° fiber orientation. Note that the plies do not have to be laid in pairs. For example, the plies are laid in pairs as shown in FIG. 4. However, in FIG. 3 the top and bottom layers 52, 58 are laid as a single layer, while the six plies between them are laid in pairs.

[0049] The insulating layers 60 are selectively inserted between the structural layers of the laminate. While an insulating layer can be inserted between each layer of the laminate, in the present example, the insulating layers are inserted between fewer than all structural layers. Specifically, the insulating layers are inserted between layers having a selected characteristic. For example, in the present example, the selected characteristic is a change in fiber direction. The insulating layers 60 are inserted between adjacent layers having different fiber angles. Therefore, in the above layup, the insulating layers are interposed between pairs of structural layers, but not between pairs of layers. More specifically, with reference to FIGS. 3-4, a first insulating layer 60a is disposed between layer 50 (FIG. 3) or 50b (FIG. 4) and layer 52a because the interface between layer 50 / 50b and layer 52a is an interface where the fibers of the adjacent layers are transverse to each other (i.e., the fiber angle changes from 90° to 0°). Similarly, second insulating layer 60b is disposed between layers 52b and 54a where the fiber angle changes from 0° to 45°. Third insulating layer 60c is disposed between the third and fourth pair where the fiber angle changes from 45° to 0° (i.e., at the interface of layers 54b and 56a). Fourth insulating layer 60d is disposed between the fourth and fifth pair where the fiber angle changes from 0° to 90° (i.e., at the interface of layers 56b and 58a, or layer 58 in FIG. 3).

[0050] It should be noted that the thicknesses of the layers in the figures are not to scale and, in some cases, have been exaggerated for illustrative purposes only. For example, in FIG. 2, the structural layers 50, 52, 54, 56, and 58 and the insulating layer 60 are shown with gaps between adjacent layers. However, it should be understood that the layers of the stack are integral layers, with different layers fused together. Furthermore, as described above, each insulating layer 60 is described as a single layer of insulating material formed from one or more insulating elements. However, in certain applications, it may be desirable to add one or more of the insulating layers so that the insulating layer is formed from multiple layers of insulating material. Also, as shown in FIG. 2, the insulating layer 60 need not be a continuous layer extending the entire length and width of the stack, as shown in FIG. 1. Instead, the insulating layer 60 may cover only a small portion of the area of the structural layer.

[0051] The laminate 40 can be formed using a variety of processes. Methods for forming the laminate 40 from multiple reinforced thermoplastic layers are described in detail below, as well as methods for using the laminate to form aircraft structures.

[0052] Multiple layers of carbon fiber-reinforced thermoplastic tape are laid one on top of the other to form structural plies. The fiber orientation of the plies can be varied, and insulating plies can be positioned at the interfaces of plies with transverse fiber orientations. For example, the layer can be formed with ten structural plies and four insulating layers oriented at 90°, 90°, Ins, 0°, 0°, Ins, 45°, 45°, Ins, 0°, 0°, Ins, 90°, 90° (where "Ins" refers to the insulating layers). The insulating layers can be formed with one or more insulating elements forming a pattern. The insulating layers can be aligned through the thickness of the stack so that the pattern of each insulating layer overlaps the pattern of the other insulating layers in the stack. In this exemplary stack, the carbon fiber layers of the structural layers are formed with PEEK / AS4 carbon fiber-reinforced unidirectional tape, and the insulating layers are formed with PEEK / S2 glass fiber-reinforced thermoplastic unidirectional tape.

[0053] The structural and insulating layers are consolidated to form a laminate by heating the assembled layers under pressure. For example, the assembly is heated to a temperature above its melting point. In the current example, the assembled layers are heated to about 725°C under a pressure of about 30 psi. After the assembled layers reach 725°C, the pressure is increased to about 100 psi and the assembly is held at the elevated temperature for an extended period of time, for example, about 30 minutes. The pressure is then removed and the consolidated laminate is allowed to cool to ambient temperature.

[0054] The laminate formed as described above can be joined with one or more separate components to form a structure. One exemplary structure is an aircraft structure. More specifically, the laminate is configured to allow welding of the laminate to a separate structure. Furthermore, the laminate is configured to isolate heat generated during the welding process from one or more selected plies of the laminate. The heat generated can also be isolated from selected regions of the selected plies. The process of welding the laminate to a separate component is described in more detail below.

[0055] As described above, the laminate is formed with one or more insulating layers. The insulating layers can be formed based on a pattern. In particular, the pattern can be similar to the contact areas of the separate elements or assemblies to be welded to the laminate. For example, referring to FIG. 1, the insulating layer can form a grid similar to the grid formed by the ribs 30, 35 and spars 20, 25 of the support structure 15. To this end, the method includes aligning the pattern of the insulating layer with the grid formed by the ribs and spars. Referring to FIG. 2, the top surface of the support structure has one or more connecting elements for welding the support structure to the laminate. In particular, the ribs 30, 35 and spars 20, 25 can be formed such that the top surfaces 22, 27, 32, 37 form connecting elements. Alternatively, the connecting elements can be separate elements that can be fixedly connected to the ribs and spars.

[0056] In the illustrated embodiment, the connecting element is a ply of carbon fiber reinforced thermoplastic material similar to the material forming the structural plies of the laminate 40. In FIG. 2, the connecting element is shown at 39 and is illustrated as a laminate fixedly connected to the top of the flange 37 of the intermediate rib 35. However, it should be understood that the connecting element may be an integral part of the underlying structure. For example, the rib 35 may be formed such that the flange 37 is formed of carbon fiber reinforced thermoplastic material that acts as the connecting element. Alternatively, the connecting element may be a separate element that is mechanically secured, for example, by clips or fasteners.

[0057] The fiber direction of the carbon fibers in the connecting element 39 is oriented transverse to the fiber direction of the bottom ply 58b of the stack. The stack 40 is placed over the structure so that the bottom layer 58b contacts the connecting element 39, with the insulating layers 60a, b, c, d aligned with the connecting element. An induction welding head 100 is then operatively positioned relative to the stack to induce an electromagnetic field that penetrates the thickness of the stack. Specifically, the induction welding head 100 moves over the top surface of the stack. The welding head need not contact the top surface; however, the welding head is sufficiently close to the top surface of the stack to induce an electromagnetic field that penetrates the thickness of the stack with sufficient strength to weld the bottom layer 58a and the connecting element 39.

[0058] The welding head 100 heats the bottom layer and the connecting elements through electromagnetic induction, which generates eddy currents. The welding head includes an electronic oscillator that sends a high-frequency alternating current through an electromagnet. A high-velocity alternating magnetic field then passes through the laminate and connecting elements, generating eddy currents that in turn heat the bottom layer and connecting elements.

[0059] In the present example, the welding head induces an electromagnetic field through the laminate and the connecting element such that adjacent layers having transverse carbon fibers heat in response to the electromagnetic field. However, layers separated by the insulating layer 60 do not heat above their melting points in response to the electromagnetic field. In particular, in the present example, the electromagnetic field generated by the welding head is sufficient to heat the bottom layer 58a and the connecting element 39 above the melting point of the thermoplastic matrix material of the bottom layer and the connecting element. When the bottom layer and the connecting element are raised above the melting point under pressure, the bottom layer fuses with the connecting element, welding the two elements together. Also, as noted above, the insulating layer 60 insulates adjacent layers from the induced heat generated by the welding head. Notably, it should be noted that a portion of the electromagnetic field extends beyond the edges of the insulating element of the insulating layer. Therefore, the electromagnetic field may induce some heating in the structural ply. However, the induction heating is substantially less than the heating induced between non-insulating layers (i.e., 58b and 39) and substantially below the melting point of the structural layer. The insulating layer thereby limits heating of the structural layer below its melting point while allowing heating of the bottom layer and the connecting elements above its melting point. Thus, while the insulating layer does not necessarily prevent all heating of a layer adjacent to it, for purposes of this method, the insulating layer prevents heating of that layer if it limits heating of that layer substantially below its melting point.

[0060] 2, the welding head 100 is shown as inducing an electromagnetic field that penetrates the laminate to weld the laminate to the rib 35. The method includes controlling the movement of the welding head so that the welding head is brought over the laminate to weld multiple points of the laminate to the support structure 15. In particular, the method can include controlling the position of the welding head to follow the pattern formed by the insulating layer 60, such that the welding head welds the laminate to the support structure along areas that correspond to the pattern formed by the insulating layer 60.

[0061] In the above description, the laminate has been described as a flat panel laminate. However, it should be understood that the present invention is not limited to flat panel structures. For example, the laminate may be used in a variety of structures and in a variety of applications, and may have particular application in aerospace to provide various components, including, but not limited to, fuselages, nacelles, and airfoils, such as wings, elevators, etc. The laminate 40 described above may be formed into curved structures and may be connected to separate elements or structures.

[0062] It will be understood by those skilled in the art that the above-described embodiments can be changed or modified without departing from the spirit of the present invention. Therefore, it should be understood that the present invention is not limited to the particular embodiments described herein, but is intended to include all modifications that fall within the spirit of the present invention as defined in the appended claims. [Explanation of symbols]

[0063] 1...Translation 2...Translation 10...Composite aircraft structures, aircraft structures, assemblies 15...Support structure 20...Front spar 22...Top surface 24...bottom surface 25...Rear spar 27...Top surface 30...First end rib, second end rib 32...Top surface, upper flange 34...Lower flange 35...Intermediate rib 37...Top surface, upper flange 38...Lower flange 39...Connected elements 40...Upper skin, cover, laminate 50...Structural layer 50a...Structural layer, first layer 50b...Structural layer, second layer 52...Structural Layer 52a...Structural layer, third structural layer 52b...Structural layer, fourth structural layer 54...Structural Layer 54a...Structural layer, fifth structural layer 54b...Structural layer, 6th structural layer 56...Structural Layer 56a...Structural layer, 7th structural layer 56b...Structural layer, 8th structural layer 58...Structural Layer 58a...Structural layer, bottom layer 58b...Structural layer, bottom layer 60...Multiple insulating elements 60a...first insulating layer 60b...second insulating layer 60c...Third insulating layer 60d...fourth insulating layer 70...Spar insulation element 70...Second spar insulating element 80...End rib insulating element 82...insulating element, first segment 84...Second segment 100...induction welding head, welding head

Claims

1. 1. A composite aircraft structure comprising: a plurality of longitudinally elongated spars, each having an upper element with a layer of unidirectional or woven carbon fiber reinforced thermoplastic material forming a weld zone; a plurality of elongated ribs interconnected with said plurality of spars to form a grid, each rib having an upper element with a layer of unidirectional or woven carbon fiber reinforced thermoplastic material forming a weld zone; a composite skin connected to the grid of spars and ribs, a first ply of unidirectional carbon fiber reinforced thermoplastic material, the carbon fibers being oriented in a first direction; a second ply of unidirectional carbon fiber reinforced thermoplastic material, the carbon fibers being oriented in the first direction, the first ply overlying and directly connected to the second ply; and a third ply of unidirectional carbon fiber reinforced thermoplastic material, the carbon fibers oriented in a second direction transverse to the first direction; and a fourth ply of unidirectional carbon fiber reinforced thermoplastic material, the carbon fibers being oriented in the second direction, the third ply overlying and directly connected to the fourth ply; and a fifth ply of unidirectional carbon fiber reinforced thermoplastic material, the carbon fibers oriented transverse to the carbon fibers of the upper element of the spar and rib; and a plurality of elongated insulating elements formed of an electrically insulating material, each insulating element having a length and a width, at least some of the insulating elements having a length that substantially exceeds the width, the insulating elements forming a grid configured substantially similar to the grid formed by the ribs and spars such that the insulating elements overlie the upper elements of the ribs and spars, the grid of the insulating elements being disposed between the second layer and the third layer; a composite skin having A composite aircraft structure having:

2. 10. The aircraft structure of claim 1, wherein the first, second, third, fourth, and fifth plies are unidirectional carbon fiber reinforced thermoplastic material.

3. 3. An aircraft structure according to claim 1 or 2, wherein each insulating element comprises glass fibres embedded in a matrix of thermoplastic material.

4. An aircraft structure according to any one of claims 1 to 3, wherein the insulating element comprises a plurality of elongated spar insulation pieces overlying the spar.

5. 5. The aircraft structure of claim 4, wherein the insulation element comprises a plurality of elongated rib insulation members overlying the ribs, the rib insulation members having a length less than a length of the spar insulation members.

6. 6. An aircraft structure according to claim 5, wherein said rib insulation extends between said spar insulation members without substantially overlapping said spar insulation members.

7. 7. An aircraft structure according to claim 5 or 6, wherein the spar insulation is spaced apart from one another.

8. 8. The aircraft structure according to claim 1, wherein the composite skin comprises an additional layer of carbon fiber reinforced thermoplastic material, the grid of insulating elements being a first grid, and the plurality of insulating elements forming a second grid configured substantially similarly to the first grid of insulating elements, the second grid being disposed between the additional layer of carbon fiber reinforced thermoplastic material such that the second grid is aligned with the first grid.

9. 9. An aircraft structure according to any one of claims 1 to 8, comprising a second composite skin connected to a grid of said ribs and spars.

10. 10. An aircraft structure according to claim 9, wherein each spar and rib has a lower element having a layer of unidirectional carbon fiber reinforced thermoplastic material forming a lower weld zone.

11. 11. An aircraft structure according to claim 10, wherein the second skin is connected to the lower element of the rib and spar.

12. 12. An aircraft structure according to any one of claims 1 to 11, wherein the grid formed by the insulating elements has a plurality of openings.

13. 13. The aircraft structure according to claim 1, wherein the thermoplastic material of the first to fifth plies comprises a semi-crystalline thermoplastic material of the polyaryletherketone family.

14. 14. An aircraft structure according to any one of claims 1 to 13, wherein the thermoplastic material of the carbon fibre reinforced plies has a melting point above 500 degrees Fahrenheit.

15. 15. The aircraft structure according to claim 1, wherein the insulating element is configured such that when the fifth ply is heated by an induction welding head to a temperature above its melting point, the thermoplastic material of the first to fourth plies remains below its melting point.

16. 15. The aircraft structure according to any one of claims 1 to 14, wherein the thermoplastic material of the fifth layer has a melting point, and the insulating element is configured to prevent heating of the second and third layers above the melting point when the fifth laminate is exposed to an electromagnetic field sufficient to heat a portion of the fifth laminate above the melting point.

17. 15. The aircraft structure according to claim 1, wherein the thermoplastic material of the fifth layer has a melting point, and the insulating element is configured to prevent the thermoplastic material of the first to fourth layers from melting when the fifth layer is welded to the rib and spar upper elements by induction welding.

18. 1. A composite aircraft structure comprising: a plurality of longitudinally elongated spars; a plurality of elongated ribs connected to the plurality of spars to form a grid; A composite skin, a plurality of structural layers, each of which is formed from a carbon fiber reinforced thermoplastic thin film; a plurality of insulating elements formed of an electrically insulating material, the insulating elements forming a plurality of insulating grids, the insulating grids being aligned with one another, the insulating grids being disposed between the structural layers; and A welded layer formed from a carbon fiber reinforced thermoplastic thin film a composite skin having a plurality of connecting elements having a layer of unidirectional carbon fiber reinforced thermoplastic material, each connecting element being connected to one of the spars or the ribs; and the carbon fibers of the welding layer are oriented in a first direction, and the carbon fibers of the connecting element are oriented in a second direction transverse to the first direction; The welding layer overlies and is connected to the connecting element. Composite aircraft structure.

19. 20. The composite aircraft structure of claim 18, wherein the weld layer and the connecting element comprise a unidirectional carbon fiber reinforced thermoplastic material.

20. 20. An aircraft structure according to claim 18 or 19, wherein each insulating element comprises glass fibres embedded in a matrix of thermoplastic material.

21. An aircraft structure according to any one of claims 18 to 20, wherein the insulating element comprises a plurality of elongated spar insulation pieces overlying the spar.

22. 22. The aircraft structure of claim 21, wherein the insulation element comprises a plurality of elongated rib insulation members overlying the ribs, the rib insulation members having a length less than a length of the spar insulation members.

23. 23. The aircraft structure of claim 22, wherein the rib insulation extends between the spar insulation without substantially overlapping the spar insulation.

24. 24. An aircraft structure according to claim 22 or 23, wherein the spar insulation is spaced apart from one another.

25. 25. An aircraft structure according to any one of claims 18 to 24, comprising a second composite skin connected to a grid of said ribs and spars.

26. 26. An aircraft structure according to claim 25, wherein a plurality of said connectors connect said second skin with said ribs and spars.

27. An aircraft structure according to any one of claims 18 to 26, wherein the grid formed of the insulating elements has a plurality of openings.

28. 28. An aircraft structure according to any one of claims 18 to 27, wherein the thermoplastic material of the structural layer comprises a semi-crystalline thermoplastic material of the polyaryletherketone family.

29. 29. An aircraft structure according to any one of claims 18 to 28, wherein the thermoplastic material of the structural layer has a melting point above 500 degrees Fahrenheit.

30. 30. The aircraft structure according to any one of claims 18 to 29, wherein the insulating element is configured to keep the thermoplastic material of the structural layer below its melting point when the weld layer is heated above its melting point by an induction welding head.

31. 30. The aircraft structure according to any one of claims 18 to 29, wherein the thermoplastic material of the welding layer has a melting point, and the insulating element is configured to prevent the thermoplastic material of the structural layer from melting when the welding layer is welded to the connecting element by induction welding.

32. 1. A composite laminate for use in an aircraft structure, comprising: a plurality of structural layers, each of which is formed from a unidirectional carbon fiber reinforced thermoplastic thin film, the fiber orientation of at least a portion of the structural layers being transverse to the fiber orientation of an adjacent structural layer; a plurality of insulating elements formed of an electrically insulating material, the insulating elements forming a plurality of insulating grids, the insulating grids being aligned with one another and embedded between adjacent structural layers within the structural layer with the fiber directions of the carbon fiber layers transverse to one another; A welded layer formed from a carbon fiber reinforced thermoplastic thin film A composite laminate having:

33. 33. The composite laminate of claim 32, wherein the carbon fiber reinforced thermoplastic film comprises a unidirectional carbon fiber reinforced thermoplastic film.

34. 34. A composite laminate according to claim 32 or 33, wherein each insulating element comprises glass fibres embedded in a matrix of thermoplastic material.

35. 35. A composite laminate according to any one of claims 32 to 34, wherein the insulating element comprises a plurality of elongated first insulating bodies having a width and a length, the length being substantially greater than the width.

36. 33. The composite laminate of claim 32, wherein the insulating element comprises a plurality of elongated second insulators, the second insulators having a length that is less than a length of the first insulators.

37. 37. The composite laminate of claim 36, wherein the second insulation extends between the first insulation without substantially overlapping the first insulation.

38. 38. A composite laminate according to claim 36 or 37, wherein the first insulators are spaced apart from one another.

39. 40. A composite laminate according to any one of claims 32 to 39, wherein the grid formed of the insulating elements has a plurality of openings.

40. 40. The composite laminate of any one of claims 32 to 39, wherein the thermoplastic material of the structural layer comprises a polyaryletherketone class semi-crystalline thermoplastic material.

41. 41. The composite laminate of any one of claims 32 to 40, wherein the thermoplastic material of the structural layer has a melting point above 500 degrees Fahrenheit.

42. 42. A composite laminate according to any one of claims 32 to 41, wherein the insulating element is configured to keep the thermoplastic material of the structural layer below its melting point when the weld layer is heated above its melting point by an induction welding head.

43. 43. A composite laminate according to any one of claims 32 to 42, wherein the thermoplastic material of the welded layer has a melting point, and the insulating element is configured to prevent the thermoplastic material of the structural layer from melting when the welded layer is welded to a separate thermoplastic element.

44. 44. A composite laminate as claimed in any one of claims 32 to 43, wherein the structural layer comprises a plurality of pairs of unidirectional carbon reinforced thermoplastic plies, each of said pairs comprising adjacent overlapping plies in which the carbon fibre fibre orientations are substantially parallel.

45. 45. The composite laminate of claim 44, wherein one of said grids of insulating elements is disposed between adjacent pairs of structural layers.

46. 1. A method of forming a composite wing box, comprising: providing a plurality of spars; connecting a plurality of ribs to the plurality of spars; forming a composite skin, the composite skin comprising: a plurality of structural layers, each of which is formed from a carbon fiber reinforced thermoplastic thin film; a plurality of insulating layers, each formed of a plurality of spaced apart electrically insulating elements, each insulating element having a length and a width, the length being substantially greater than the width, the insulating layers being aligned with one another and positioned between the structural layers, the insulating layers being configured such that the insulating elements overlie the spars and ribs when the composite skin overlies the spars and ribs; A welded layer formed from a carbon fiber reinforced thermoplastic thin film forming a composite skin having placing the composite skin over the ribs and spars such that the layer of the composite skin overlies the ribs and spars; welding the composite skin to the ribs and spars, the welding step including moving an induction welding head over the composite laminate adjacent the ribs and spars, the induction welding head applying an electromagnetic field that penetrates the composite laminate to heat the weld layer above a melting point of the thermoplastic material of the weld layer, and the insulating layer preventing the structural layer from heating above the melting point of the thermoplastic material; A method having the following.

47. 47. The method of claim 46, including connecting a plurality of connecting elements to the rib and the spar, each connecting element comprising a layer of unidirectional carbon fiber reinforced thermoplastic material, and wherein the welding step includes welding the weld layer to the connecting element.

48. 48. The method of claim 47, wherein the carbon fibers of the welded layer are oriented in a first direction and the carbon fibers of the connecting element are oriented in a second direction, and wherein connecting the plurality of connectors comprises connecting the connectors such that the second direction is transverse to the first direction.

49. 49. The method of any one of claims 46 to 48, wherein providing the composite laminate having a plurality of insulating layers comprises arranging the plurality of insulating elements into a grid.

50. 50. The method of any one of claims 46 to 49, wherein providing the composite laminate having a plurality of insulating layers comprises configuring the plurality of insulating elements such that a plurality of openings are formed between the insulating elements.

51. 51. The method of any one of claims 46-50, wherein forming the composite skin comprises providing a plurality of pairs of structural layers, each layer of a structural layer pair adjacent to and directly connected to the other structural layer of the pair, the two layers of the pair oriented so that the fibers of both layers are substantially parallel, and adjacent pairs of structural layers oriented so that the fibers of one pair are transverse to the fibers of the adjacent pair.

52. 52. The method of claim 51, wherein the step of providing an insulating layer comprises interposing an insulating layer between the one pair and the adjacent pair.

53. 53. The method of claim 52, wherein the method of providing an insulating layer comprises interposing an insulating layer between each adjacent pair of structural layers.

54. 54. The method of any one of claims 46 to 53, wherein the composite skin has an upper surface and a lower surface, the weld layer forming the lower surface, and moving the induction head comprises moving the induction head above the upper surface.

55. 1. A method of forming a composite laminate for an aircraft structure, comprising: providing a plurality of structural layers, each of the plurality of structural layers being formed from a carbon fiber reinforced thermoplastic thin film; providing a plurality of insulating elements formed of an electrically insulating material, each insulating element having a length and a width, the length being substantially greater than the width; fabricating a plurality of insulating layers, each layer formed in a pattern such that the insulating elements are spaced apart by the step of configuring the insulating elements; disposing the insulating layers between the structural layers, the disposing step including aligning the pattern with each insulating layer; The structural layers are configured such that the bottom layer forms a weld layer. disposing the insulating layer between the structural layers; combining the structural layers and the insulating layers by applying heat and pressure to the structural layers and the insulating layers to form a composite laminate in which the insulating layers are embedded within the structural layers; A method having the following.

56. 56. The method of claim 55, wherein the configuring step comprises configuring a plurality of the insulating elements such that the insulating elements extend between the plurality of spaced apart insulating elements.

57. 57. A method according to claim 55 or 56, wherein the configuring step comprises configuring the insulating elements so that there is a gap between each individual insulating element.

58. 58. The method according to any one of claims 55 to 57, placing the consolidated laminate on a tie layer of unidirectional carbon fiber reinforced thermoplastic material such that the fibers of the weld layer are oriented transverse to the fibers of the tie layer; inducing an electromagnetic field that penetrates the thickness of the laminate and the tie layer to heat the weld layer and the tie layer sufficiently to weld the weld layer and the tie layer together; The method according to claim 1, wherein

59. 60. The method of claim 58, wherein inducing the electromagnetic field comprises moving an induction coil over the stack along the pattern of the insulating layer.

60. 60. The method of claim 58 or 59, wherein inducing the electromagnetic field comprises concentrating the electromagnetic field within the width of the insulating element.

61. 61. The method of claims 58-60, wherein the bonding layer is rigidly coupled to the aircraft structure such that the step of inducing an electromagnetic field welds the laminate to the bonding layer, thereby fixedly coupling the laminate to the aircraft structure.

62. 62. The method of claim 61, wherein the aircraft structure has a plurality of structural elements forming a pattern, the method including aligning the pattern of the structural elements with the pattern of the insulating layer.

63. 63. The method of any one of claims 55 to 62, wherein arranging the insulating elements in a pattern comprises arranging the insulating elements into a grid.

64. 64. The method of any one of claims 55 to 63, wherein configuring the insulating elements in a pattern comprises configuring the insulating elements such that a plurality of openings are formed between the insulating elements.

65. 65. The method of any one of claims 55 to 64, wherein providing the plurality of structural layers comprises providing a plurality of pairs of structural layers made of unidirectional carbon fiber reinforced material, each layer of a pair adjacent to and directly connected to the other structural layer of the pair, the two layers of the pair being oriented so that the fibers of both layers are substantially parallel, and adjacent pairs of structural layers being oriented so that the fibers of one pair are transverse to the fibers of the adjacent pair.

66. 66. The method of claim 65, wherein the step of providing an insulating layer comprises interposing an insulating layer between the one pair and the adjacent pair.

67. 67. The method of claim 66, wherein the method of providing an insulating layer comprises interposing an insulating layer between each adjacent pair of structural layers.

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