Structural Composite Airfoil with Directly Attached Front Spar and Related Methods

Abstract

A structural composite airfoil includes a main structural element, an auxiliary structural element defining a trailing edge of the structural composite airfoil, and a leading edge skin panel defining a leading edge of the structural composite airfoil. The main structural element includes an upper skin panel, a lower skin panel, and a front C-channel spar. A first channel of the front C-channel spar faces the leading edge of the structural composite airfoil, and an upper flange of the front C-channel spar forms an acute angle with an elongate span portion of the front C-channel spar. The leading edge skin panel is positioned adjacent to a leading edge region of the main structural element, a first end region of the leading edge skin panel is coupled to the upper flange of the front C-channel spar, and a second end region is coupled to the lower flange of the front C-channel spar.

Description

Technical Field

[0001] This disclosure generally relates to structural composite airfoils and related methods. Background Art

[0002] Aircraft (including fixed-wing aircraft and rotary-wing aircraft) employ various aerodynamic control surfaces, such as ailerons, air brakes, elevators, flaps, rudders, slats, spoilers, etc. By manipulating one or more aerodynamic control surfaces, a pilot can control, for example, the lift generated by the aircraft during takeoff, climb, descent, and landing, as well as the orientation of the aircraft about its pitch, roll, and yaw axes. For example, the trailing edge of the wing of a fixed-wing aircraft typically includes one or more flaps that can move between a retracted position and an extended position. During cruise, the flaps are typically kept in the retracted position. When extended, the flaps increase the camber of the wing. Thus, during takeoff, climb, descent, or landing, the flaps can be partially or fully extended to increase the maximum lift coefficient and effectively reduce the stall speed of the aircraft. The aerodynamic control surfaces are typically airfoils formed of composite materials and are thus referred to herein as structural composite airfoils.

[0003] Structural composite airfoils (such as flaps) have an aerodynamic cross-sectional profile that is typically formed by joining an upper skin adjacent both the leading edge and the trailing edge of the structural composite airfoil to a lower skin. For example, in a conventional configuration of inner and outer flaps, the main structural elements of the flap are defined by joining the upper and lower skins to three spars that extend the width of the flap. The leading edge (which typically includes a bulbous shape) and the trailing edge (which tapers to a thin section) of the structural composite airfoil are typically outside of the main structural elements, thereby forming respective auxiliary structural elements of the flap. Various fasteners and components (e.g., splice straps and / or nut plates) are used to secure the upper and lower skins to the spars and other structures that form the flap. A large number of fasteners can increase cost, manufacturing cycle time, and the weight of the resulting assembly. Accordingly, those skilled in the art continue to work on research and development efforts to improve structural composite airfoils and their manufacture. Summary of the Invention

[0004] Structural composite airfoils and related methods of forming the same as disclosed herein can reduce the number of fasteners, improve the airfoil aerodynamic surface, and / or simplify the manufacturing process of the structural composite airfoil.

[0005] Examples of the structural composite airfoil according to the present disclosure include a main structural element, an auxiliary structural element defining a trailing edge of the structural composite airfoil, and a leading edge skin panel defining a leading edge of the structural composite airfoil. The structural composite airfoil has a leading edge and a trailing edge, and the main structural element extends from a leading edge region to a trailing edge region. The leading edge region of the main structural element is adjacent to the leading edge of the structural composite airfoil.

[0006] The main structural element includes an upper skin panel, a lower skin panel, and a front C-channel spar. An internal volume is defined between the upper skin panel and the lower skin panel. The front C-channel spar includes an upper flange coupled to the upper skin panel and a lower flange coupled to the lower skin panel. A first channel of the front C-channel spar faces the leading edge of the structural composite airfoil, and the upper flange forms an acute angle with the elongated span portion of the front C-channel spar.

[0007] The leading edge skin panel is positioned adjacent to the leading edge region of the main structural element, wherein a first end region of the leading edge skin panel is coupled to the upper flange of the front C-channel spar, wherein a second end region of the leading edge skin panel is coupled to the lower flange of the front C-channel spar, and wherein the leading edge skin panel has a bull nose shape.

[0008] A method of assembling such a structural composite airfoil is also disclosed. In such a method, the upper skin panel is coupled to the upper flange of the front C-channel spar, the lower skin panel is coupled to the lower flange of the front C-channel spar such that an internal volume is defined between the upper skin panel and the lower skin panel, and the leading edge skin panel is coupled to the front C-channel spar. For example, a first end region of the leading edge skin panel is coupled to the upper flange of the front C-channel spar, and a second end region of the leading edge skin panel is coupled to the lower flange of the front C-channel. Description of the Drawings

[0009] Figure 1 is a schematic view of an apparatus according to the present disclosure that may include one or more structural composite airfoils.

[0010] Figure 2 is a schematic side view of an example of a structural composite airfoil according to the present disclosure.

[0011] Figure 3 is a side view of an integral Z-beam formed in the lower skin panel.

[0012] Figure 4 is a side view of an integral Z-beam formed in the upper skin panel.

[0013] Figure 5 is a flow chart representing a disclosed method of forming the disclosed structural composite airfoil. Detailed Description

[0014] ReferenceFigure 1 , the device 12 may include one or more structural composite airfoils 10. The structural composite airfoils 10 may be used in many different industries and applications, such as aerospace, automotive, construction, marine, wind power generation, remote control aircraft, military, entertainment, and / or racing industries. In Figure 1 , an example of the device 12 that may include one or more structural composite airfoils 10 is generally shown in the form of an aircraft 14. The aircraft 14 may take any suitable form, including commercial aircraft, military aircraft, or any other suitable aircraft. Although Figure 1 the aircraft 14 is shown in the form of a fixed-wing aircraft, other types and configurations of aircraft are also within the scope of the aircraft 14 according to the present disclosure, including (but not limited to) rotary-wing aircraft and helicopters.

[0015] The device 12 (e.g., the aircraft 14) may include one or more structural composite airfoils 10. As illustrative non-exclusive examples, the structural composite airfoils 10 may be used in the wing 16 (e.g., the flap 17, which may be an inboard or outboard flap), but other components of the aircraft 14, such as the horizontal stabilizer 18, the vertical stabilizer 20, and other components, additionally or alternatively may include one or more structural composite airfoils 10. Other applications for the structural composite airfoils 10 in the aircraft 14 (or other device 12) may include other wing control surfaces, ailerons, flaperons, air brakes, elevators, slats, spoilers, rudders, canards, and / or winglets. In other industries, examples of the device 12 that includes one or more structural composite airfoils 10 may include or be part of the following: space satellites, transportation vehicles, shipping containers, rapid transit vehicles, vehicle bodies, propeller blades, turbine blades, and / or marine transportation vehicles (e.g., sailboats), etc.

[0016] Figure 2 Illustrative non-exclusive examples of the structural composite airfoil 10 according to the present disclosure are provided. Generally, elements that may be included are shown in solid lines, while optional elements are shown in dashed lines. However, the elements shown in solid lines are not required for all examples, and the elements shown in solid lines may be omitted from a particular example without departing from the scope of the present disclosure.

[0017] The structural composite airfoil 10 has a leading edge 22 and a trailing edge 24, and generally includes a primary structural element 26 and a secondary structural element 28. As used herein, a "primary structural element" is an element or structure that bears flight, ground, or pressurization loads, and whose failure would reduce the structural integrity of the device or assembly of which the structural composite airfoil 10 is a part. As used herein, a "secondary structural element" is an element or structure whose failure does not affect the safety of the device or assembly of which the structural composite airfoil 10 is a part.

[0018] The main structural element 26 extends from the leading edge region 30 to the trailing edge region 32. As Figure 2 shown, the leading edge region 30 is adjacent to the leading edge 22 of the structural composite airfoil 10, but the leading edge region 30 may not actually define the leading edge 22. The leading edge region 30 can be referred to as the region of the main structural element 26 that is closest to the leading edge 22. Similarly, the trailing edge region 32 can be referred to as the region of the main structural element 26 that is closest to the trailing edge 24, but the trailing edge region 32 of the main structural element 26 does not define the trailing edge 24 of the structural composite airfoil 10. As used herein, if a first element or structure is positioned closer to the trailing edge 24 than another element or structure, the first element or structure is said to be "aft" of the other element or structure. Similarly, as used herein, if a first element or structure is positioned closer to the leading edge 22 than another element or structure, the first element or structure is said to be "forward" of the other element or structure.

[0019] The main structural element 26 at least includes an upper skin panel 34, a lower skin panel 36, and a front C-channel spar 38. An internal volume 40 is defined between the upper skin panel 34 and the lower skin panel 36. The front C-channel spar 38 includes an upper flange 42 and a lower flange 44, where the upper flange 42 is coupled to the upper skin panel 34 and the lower flange 44 is coupled to the lower skin panel 36. A first channel 46 of the front C-channel spar 38 faces the leading edge 22 of the structural composite airfoil 10. This arrangement of the front C-channel spar 38 relative to the leading edge 22 can allow for an effective coupling of the leading edge skin panel 54 to the upper skin panel 34 and the lower skin panel 36 (via the front C-channel spar 38) without creating any wobbling in the upper skin panel 34 or the lower skin panel 36, thereby potentially reducing the complexity of manufacturing the upper skin panel 34 and the lower skin panel 36.

[0020] The upper flange 42 forms a first angle 48 with the elongated span portion 50 of the front C-shaped channel spar, and the lower flange 44 forms a second angle 52 with the elongated span portion 50. The first channel 46 is defined by the upper flange 42, the lower flange 44, and the elongated span portion 50. In some examples of the structural composite airfoil 10, the first angle 48 and / or the second angle 52 can be acute angles. Typical conventional airfoil configurations would involve such angles greater than 90 degrees to facilitate removal of parts from the tool and / or the channels of the front spar would be arranged facing the trailing edge of the airfoil. Examples of the presently disclosed structural composite airfoil 10 can advantageously provide a joint between components or elements (e.g., joining the leading edge skin panel 54 and the upper flange 42) without creating flutter in the upper skin panel 34 or the lower skin panel 36 or using splice straps, and / or can reduce the number of parts by reducing or eliminating the number of splice straps, nut plates, and / or other fasteners used to assemble the structural composite airfoil 10. Additionally or alternatively, the upper flange 42 can be angled relative to the elongated span portion 50 to be complementary to the first end region 56 of the leading edge skin panel 54. Similarly, the lower flange 44 can be angled relative to the elongated span portion 50 to be complementary to the second end region 58 of the leading edge skin panel 54.

[0021] The leading edge 22 of the structural composite airfoil 10 is defined by a leading edge skin panel 54, which is typically formed to have a bullnose shape. The leading edge skin panel 54 can be positioned adjacent to the leading edge region 30 of the primary structural element 26, but the leading edge skin panel 54 can be a separate part that is outside or different from the primary structural element 26. In other examples, the leading edge skin panel 54 can be within and / or define the leading edge region 30 of the primary structural element 26, such as in an example where the primary structural element 26 extends to the leading edge 22. The leading edge skin panel 54 is coupled to the upper skin panel 34 and the lower skin panel 36 via the front C-channel spar 38. Specifically, a first end region 56 of the leading edge skin panel 54 is coupled to the upper flange 42 of the front C-channel spar 38, and a second end region 58 of the leading edge skin panel 54 is coupled to the lower flange 44 of the front C-channel spar 38. Since the upper flange 42 of the front C-channel spar 38 is coupled to the leading edge skin panel 54 and the upper skin panel 34, the front C-channel spar 38 effectively couples the leading edge skin panel 54 to the upper skin panel 34. In some examples, the leading edge skin panel 54 does not overlap the upper skin panel 34 on the upper flange 42 (e.g., does not overlap the upper leading edge end 76 of the upper skin panel 34). In a particular example, the upper leading edge end 76 of the upper skin panel 34 can abut the leading edge skin panel 54 (e.g., abut the first end region 56 of the leading edge skin panel 54). In other examples, the upper skin panel 34 can be coupled to the upper flange 42 without contacting the leading edge skin panel 54. Similarly, since the lower flange 44 of the front C-channel spar 38 is coupled to the leading edge skin panel 54 and the lower skin panel 36, the front C-channel spar 38 effectively couples the leading edge skin panel 54 to the lower skin panel 36. In some examples, the leading edge skin panel 54 does not overlap the lower skin panel 36 on the lower flange 44 (e.g., does not overlap the lower leading edge end 78 of the lower skin panel 36). In a particular example, the lower leading edge end 78 of the lower skin panel 36 can abut the leading edge skin panel 54 (e.g., abut the second end region 58). In other examples, the lower skin panel 36 can be coupled to the lower flange 44 without contacting the leading edge skin panel 54.

[0022] The trailing edge 24 of the structural composite airfoil 10 is defined by an auxiliary structural element 28. In various examples of the structural composite airfoil 10, the auxiliary structural element 28 can include a wedge closure, a duckbill closure, a bonded closure, and / or a riveted closure. Examples of suitable trailing edge closures are also disclosed in U.S. Patent No. 10,532,804, titled "Aerodynamic control surface and associated trailing edge close-out method," issued on January 14, 2020, the entire disclosure of which is incorporated herein by reference for all purposes.

[0023] The upper skin panel 34 generally extends from the upper leading edge end 76 to the upper trailing edge end 92. The upper leading edge end 76 corresponds to the end of the upper skin panel 34 closest to the leading edge 22 of the structural composite airfoil 10, and the upper trailing edge end 92 corresponds to the end of the upper skin panel 34 closest to the trailing edge 24 of the structural composite airfoil 10. Similarly, the lower skin panel 36 generally extends from the lower leading edge end 78 to the lower trailing edge end 94. The lower leading edge end 78 corresponds to the end of the lower skin panel 36 closest to the leading edge 22, and the lower trailing edge end 94 corresponds to the end of the lower skin panel 36 closest to the trailing edge 24. As described above, the upper leading edge end 76 and the lower leading edge end 78 can be coupled to the front C-channel spar 38. In some examples, the upper trailing edge end 92 can be coupled to the lower trailing edge end 94. Additionally or alternatively, the upper trailing edge end 92 and / or the lower trailing edge end 94 can form or define the trailing edge 24 of the structural composite airfoil 10.

[0024] The structural composite airfoil 10 can include one or more fasteners that fix various components to each other. For example, a first fastener 80 can couple the leading edge skin panel 54 (e.g., a first end region 56 of the leading edge skin panel 54) to the upper flange 42 of the front C-channel spar 38. In some examples, the first fastener 80 is a plurality of first fasteners 80 spaced along the width of the structural composite airfoil 10 (the width of the airfoil extending into / out of the page) to fix the leading edge skin panel 54 to the front C-channel spar 38 along the first end region 56. The leading edge skin panel 54 can be configured to engage with the upper skin panel 34 via the coupling of both the leading edge skin panel 54 and the upper skin panel 34 to the upper flange 42 without any wobbling being formed in either panel. Additionally or alternatively, since the upper flange 42 can be configured to effectively splice the leading edge skin panel 54 and the upper skin panel 34, the structural composite airfoil 10 can be formed without a separate splice strip connecting the leading edge skin panel 54 and the upper skin panel 34.

[0025] Similarly, a second fastener 82 can couple the leading edge skin panel 54 (e.g., a second end region 58 of the leading edge skin panel 54) to the lower flange 44 of the front C-channel spar 38. In some examples, the second fasteners 82 are a plurality of second fasteners 82 spaced along the width of the structural composite airfoil 10 (the width of the airfoil extending into / out of the page) to secure the leading edge skin panel 54 to the front C-channel spar 38 along the second end region 58. The leading edge skin panel 54 can be configured to mate with the lower skin panel 36 via the coupling of both the leading edge skin panel 54 and the lower skin panel 36 to the lower flange 44 without any rattle being formed in either panel. Additionally or alternatively, since the front flange 44 can be configured to effectively splice the leading edge skin panel 54 and the lower skin panel 36, the structural composite airfoil 10 can be formed without a separate splice strip connecting the leading edge skin panel 54 and the lower skin panel 36. The first fasteners 80 and the second fasteners 82 can be configured such that the leading edge skin panel 54 can be selectively removed from the main structural element 26 by removing the first fasteners 80 and the second fasteners 82.

[0026] A third fastener 84 (or a plurality of third fasteners 84 spaced along the width of the structural composite airfoil 10) can be positioned to couple the upper skin panel 34 to the upper flange 42 of the front C-channel spar 38. The third fasteners 84 generally couple the upper leading edge end 76 of the upper skin panel 34 to the upper flange 42. A fourth fastener 86 (or a plurality of fourth fasteners 86 spaced along the width of the structural composite airfoil 10) can be positioned to couple the lower skin panel 36 to the lower flange 44 of the front C-channel spar 38. The fourth fasteners 86 generally couple the lower leading edge end 78 of the lower skin panel 36 to the lower flange 44. The third fasteners 84 and / or the fourth fasteners 86 can be accessible (e.g., not blind holes) even after the main structural element 26 is assembled. Additionally or alternatively, the third fasteners 84 and / or the fourth fasteners 86 can be fixed permanent fasteners (e.g., hex drive bolts) without nut plates.

[0027] The structural composite airfoil 10 can further include an intermediate C-channel spar 60 and / or a rear C-channel spar 62, one or both of which can form part of the main structural element 26. In Figure 2 the example shown, the main structural element 26 is defined by the front C-channel spar 38, the intermediate C-channel spar 60, the rear C-channel spar 62, and the respective portions of the upper skin panel 34 and the lower skin panel 36 extending between the front C-channel spar 38 and the rear C-channel spar 62. In other examples of the structural composite airfoil 10, the main structural element 26 can be more than Figure 2extends further towards the leading edge 22 as shown. For example, as described above, although the main structural element 26 may extend only between the front C-channel spar 38 and the rear C-channel spar 62, in other examples, the main structural element 26 may optionally extend further forward such that the main structural element 26 may also extend to the leading edge 22 and include the leading edge. Additionally or alternatively, the main member element 26 may be more Figure 2 extends towards the trailing edge 24 as shown. For example, the main structural element 26 may include at least a portion of the structural composite airfoil 10 behind the rear C-channel spar 62.

[0028] In an example including an intermediate C-channel spar 60, the intermediate C-channel spar 60 may include a second channel 64 facing the leading edge 22. The intermediate C-channel spar 60 may be coupled to the upper skin panel 34 and the lower skin panel 36. For example, the intermediate C-channel spar 60 may include an intermediate upper flange 66 coupled to the upper skin panel 34. Additionally or alternatively, the intermediate C-channel spar 60 may include an intermediate lower flange 68 coupled to the lower skin panel 36. The intermediate C-channel spar 60 is positioned behind the front C-channel spar 38.

[0029] In an example including a rear C-channel spar 62, the rear C-channel spar 62 may include a third channel 70 facing the leading edge 22. The rear C-channel spar 62 may be coupled to the upper skin panel 34 and the lower skin panel 36. The rear C-channel spar 62 may include a rear upper flange 72 coupled to the upper skin panel 34. Additionally or alternatively, the rear C-channel spar 62 may include a rear lower flange 74 coupled to the lower skin panel 36. The rear C-channel spar 62 is positioned behind the front C-channel spar 38. In an example of the structural composite airfoil 10 including an intermediate C-channel spar 60 and a rear C-channel spar 62, the rear C-channel spar 62 is positioned behind the intermediate C-channel spar 60.

[0030] A plurality of other fasteners 88 may be utilized to couple the upper skin panel 34 to the intermediate C-channel spar 60 (e.g., the intermediate upper flange 66) and / or the rear C-channel spar 62 (e.g., the rear upper flange 72). Similarly, one or more fasteners 88 may be used to couple the lower skin panel 36 to the intermediate C-channel spar 60 (e.g., the intermediate lower flange 68) and / or the rear C-channel spar 62 (e.g., the rear lower flange 74). Additionally or alternatively, one or more fasteners 88 may be used to couple the upper trailing edge end 92 to the lower trailing edge end 94.

[0031] Each of the upper skin panel 34 and the lower skin panel 36 can be a composite panel formed of multiple (laminated) layers of fiber-reinforced polymer laminated together. For example, the upper skin panel 34 and the lower skin panel 36 can be formed of carbon fiber-reinforced polymer material or glass fiber-reinforced polymer material. In other examples, the upper skin panel 34 and / or the lower skin panel 36 can be a metallic material, a polymer, or other suitable materials.

[0032] In some examples, at least a portion of the upper skin panel 34 can be core-reinforced. As used herein, "core-reinforced" refers to a skin panel having at least a first skin and a low-density core material coupled to the skin. The core-reinforcing material optionally includes a second skin, with the core material sandwiched between the first skin and the second skin to form a sandwich panel. Suitable materials for forming the core-reinforced portion are well known in the art and include honeycomb core materials and metallic core materials, but other core materials are also within the scope of the present disclosure. As an illustrative example, the upper skin panel 34 can include a first upper core-reinforced portion 134, a second upper core-reinforced portion 136, and a third upper core-reinforced portion 138. The first upper core-reinforced portion 134 can be positioned between the front C-channel spar 38 and the intermediate C-channel spar 60, the second upper core-reinforced portion 136 can be positioned between the intermediate C-channel spar 60 and the rear C-channel spar 62, and / or the third upper core-reinforced portion 138 can be positioned between the rear C-channel spar 62 and the upper trailing edge end 92. One or more of the upper core-reinforced portions 134, 136, 138 can be tapered, such as in the regions near the respective sections of the C-channel spars 38, 60, and / or 62. For example, the upper core-reinforced portions 134, 136, and / or 138 can have a height or thickness that extends downward from the upper skin panel 34 toward the lower skin panel 36, where the height or thickness decreases near one or more of the C-channel spars 38, 60, and / or 62, thereby forming a taper. In Figure 2 an example, the thickness of the first upper core-reinforced portion 134 tapers near the front C-channel spar 38 and the intermediate C-channel spar 60, the thickness of the second upper core-reinforced portion 136 tapers near the intermediate C-channel spar 60 and the rear C-channel spar 62, and the thickness of the third upper core-reinforced portion 138 tapers near the rear C-channel spar 62 and the trailing edge 24. In other examples, the height or thickness of one or more of the upper core-reinforced portions 134, 136, and / or 138 can be substantially constant rather than tapering at the locations where the respective upper core-reinforced portions 134, 136, and / or 138 meet the respective C-channel spars 38, 60, 62. In some examples, one or more of the upper core-reinforced portions 134, 136, and / or 138 can be adjacent to the respective C-channel spars 38, 60, and / or 62. Although Figure 2The upper skin panel 34 shown in [Figure] includes three different upper core reinforcement portions 134, 136, 138. However, in other examples, the upper skin panel 34 may be core-reinforced along its entire length, along a greater or lesser portion of its length, and / or may include more or fewer discrete upper core reinforcement sections than Figure 2 shown in [Figure].

[0033] Additionally or alternatively, at least a portion of the lower skin panel 36 may be core-reinforced. As an illustrative example, the lower skin panel 36 includes a first lower core reinforcement portion 140, a second lower core reinforcement portion 142, and a third lower core reinforcement portion 144. The first lower core reinforcement portion 140 may be positioned between the front C-channel spar 38 and the intermediate C-channel spar 60, the second lower core reinforcement portion 142 may be positioned between the intermediate C-channel spar 60 and the rear C-channel spar 62, and / or the third lower core reinforcement portion 144 may be positioned between the rear C-channel spar 62 and the lower trailing edge end 94. One or more of the lower core reinforcement portions 140, 142, 144 may taper, such as in the regions adjacent to the C-channel spars 38, 60, and / or 62 in the respective sections. For example, the lower core reinforcement portions 140, 142, and / or 144 may have a height or thickness that extends upward from the lower skin panel 36 toward the upper skin panel 34, where the height or thickness decreases near one or more of the C-channel spars 38, 60, and / or 62, thereby forming a taper. In Figure 2 the example shown, the thickness of the first lower core reinforcement portion 140 tapers near the front C-channel spar 38 and the intermediate C-channel spar 60, the thickness of the second lower core reinforcement portion 142 tapers near the intermediate C-channel spar 60 and the rear C-channel spar 62, and the thickness of the third lower core reinforcement portion 144 tapers near the rear C-channel spar 62 and the trailing edge 24. In other examples, the height or thickness of one or more of the lower core reinforcement portions 140, 142, and / or 144 may be substantially constant rather than tapering at the locations where the respective lower core reinforcement portions 140, 142, and / or 144 meet the respective C-channel spars 38, 60, 62. In some examples, one or more of the lower core reinforcement portions 140, 142, and / or 144 may abut the respective C-channel spars 38, 60, and / or 62. Although Figure 2 the lower skin panel 36 shown in [Figure] includes three different lower core reinforcement portions 140, 142, 144, in other examples, the lower skin panel 36 may be core-reinforced along its entire length, may be core-reinforced along a greater or lesser portion of its length, and / or may include more or fewer discrete lower core reinforcement portions than Figure 2 shown in [Figure].

[0034] The structural composite airfoil 10 has a length 90, which may also be referred to herein as the chord length 90, and positions along the length 90 may be defined in terms of the percentage of the distance from the leading edge 22 along the length 90. In these respects, the front C-channel spar 38 may be positioned between 0% and 10% of the length 90 away from the leading edge 22. In a particular example, the front C-channel spar 38 is positioned at approximately 5% of the length 90 away from the leading edge 22. In some examples, the front C-channel spar 38 may be positioned as far forward as possible for integration. Additionally or alternatively, the intermediate C-channel spar 60 may be positioned between 20% and 40% of the length 90 away from the leading edge 22, such as at approximately 30% of the length 90 away from the leading edge 22. In some examples, the intermediate C-channel spar 60 may be positioned to balance torsional capabilities within the primary structural element 26 on either side of the intermediate C-channel spar 60. Additionally or alternatively, the rear C-channel spar 62 may be positioned between 40% and 70% of the length 90 away from the leading edge 22, and / or between 50% and 60% of the length 90 from the leading edge 22. In a particular example, the rear C-channel spar 62 may be positioned at approximately 55% of the length 90 away from the leading edge 22. In some examples, the rear C-channel spar 62 may be positioned as far back as possible for integration.

[0035] Some examples of the structural composite airfoil 10 may include an integral Z-spar 100, which may be part of the primary structural element 26, and in some examples, elements behind the integral Z-spar 100 are part of the secondary structural element 28. Thus, positioning the integral Z-spar 100 behind the intermediate C-channel spar 60 and / or the rear C-channel spar 62 (or in place of one or both of these spars) may extend or elongate the length of the primary structural element 26, and / or may increase the percentage of the length 90 of the structural composite airfoil 10 corresponding to the primary structural element 26. In some examples, the integral Z-spar 100 may be formed within the trailing edge region 32 of the primary structural element 26. Figure 3-4 Examples showing such an integral Z-spar 100 Figure 3 Examples showing an integral Z-spar 100 formed in the lower skin panel 36, and Figure 4 Examples showing an integral Z-spar 100 formed in the upper skin panel 34. The integral Z-spar 100 is generally positioned adjacent to the trailing edge 24 of the structural composite airfoil 10, such as by being positioned at least 80% of the length 90 away from the leading edge 22. In some examples, the integral Z-spar 100 may be positioned between 80 - 95% of the length 90 away from the leading edge 22.

[0036] Reference Figure 3, the integral Z - shaped spar 100 may be formed in the lower trailing - edge end 94 of the lower skin panel 36. The integral Z - shaped spar 100 may include a first bend 106, a second bend 108, and a first Z - shaped spar segment 110 extending between the first bend 106 and the second bend 108. In some examples, the first Z - shaped spar segment 110 may be at least substantially perpendicular to the lower skin panel 36 and / or the upper skin panel 34. In some examples, the first Z - shaped spar segment 110 may form an angle greater than 90 degrees and / or greater than 100 degrees with the lower skin panel 36. Additionally or alternatively, the first Z - shaped spar segment 110 may form an angle greater than 90 degrees and / or greater than 100 degrees with the upper skin panel 34. The integral Z - shaped spar 100 may further include a second Z - shaped spar segment 112 extending rearward of the second bend 108. As Figure 3 shown, the second Z - shaped spar segment 112 may be coupled to the upper skin panel 34. In Figure 3 the example shown, the second Z - shaped spar segment 112 is positioned adjacent to the inner surface 114 of the upper skin panel 34. Z - shaped spar fasteners 116 may couple the integral Z - shaped spar 100 to the upper skin panel 34. In some examples, the Z - shaped spar fasteners 116 are recessed into the upper skin panel 34 (e.g., such that the Z - shaped spar fasteners 116 are at least substantially flush or sub - flush with the upper surface 130 of the upper skin panel 34) and extend through the upper skin panel 34 and the second Z - shaped spar segment 112 to couple the integral Z - shaped spar 100 to the upper skin panel 34.

[0037] The integral Z - shaped spar 100 may include a Z - shaped spar joggle 102 in the lower skin panel 36, which may be configured to receive a portion of the trailing - edge closure 104, and the trailing - edge closure may at least partially define the secondary structural element 28 and / or the trailing edge 24 of the structural composite airfoil 10. The Z - shaped spar joggle 102 is actually a small upward offset in the lower skin panel 36 toward the upper skin panel 34 and is generally positioned in front of the first bend 106. As Figure 3 shown, the first cover end region 118 of the trailing - edge closure 104 may be bonded to the lower skin panel 36. Additionally or alternatively, the first cover end region 118 may be riveted or otherwise fastened or coupled to the lower skin panel 36. To produce a smooth surface at the interface and improve aerodynamic performance, as Figure 3As shown, the first cover end region 118 may be slightly recessed into the lower skin panel 36, such as via the Z-beam fitting 102. Depending on the thickness of the first cover end region 118, the Z-beam fitting 102 may be customized to create a larger or smaller recess in the lower skin panel 36 such that the lower panel surface 126 of the lower skin panel 36 is substantially flush with the lower cover surface 128 of the trailing edge closure cover 104 within the first cover end region 118. In other words, the Z-beam fitting 102 may be larger to create a larger recess to receive and engage a given trailing edge closure cover 104 having a thicker first cover end region 118, while the Z-beam fitting 102 may be smaller to create a smaller recess to receive and engage a different given trailing edge closure cover 104 having a thinner first cover end region 118. Any gaps remaining at the interface of the Z-beam fitting 102 and the first cover end region 118 (or elsewhere on the structural composite airfoil 10) may be filled with a sealant, filler material, and / or resin and then smoothed.

[0038] The second cover end region 120 of the trailing edge closure cover 104 may include an integral wedge 122 that may be coupled (e.g., adhesively and / or via one or more fasteners) to the upper skin panel 34, as Figure 3 shown. Alternatively, the integral wedge 122 may be formed integrally with the upper skin panel 34. Still in other examples, the integral wedge 122 may be a separate component that is separate from the trailing edge closure cover 104 and the upper skin panel 34 and may be adhesively or otherwise coupled to the upper skin panel 34 and / or the trailing edge closure cover 104. As an example, the integral wedge 122 may be formed by stacking material layers, molding, and / or machining mating surface profiles to mate with the upper skin panel 34.

[0039] Referring Figure 4 to, the integral Z-beam 100 may be formed in the upper trailing edge end 92 of the upper skin panel 34. In Figure 4 the example shown, the second Z-beam segment 112 is coupled to the lower skin panel 36 and is positioned adjacent to the inner surface 124 of the lower skin panel 36. The Z-beam fasteners 116 couple the integral Z-beam 100 to the lower skin panel 36, where the Z-beam fasteners 116 are recessed into the lower skin panel 36 (e.g., such that the Z-beam fasteners 116 are at least substantially flush or sub-flush with the lower panel surface 126 of the lower skin panel 36) and extend through the lower skin panel 36 and the second Z-beam segment 112 to couple the integral Z-beam 100 to the lower skin panel 36.

[0040] In Figure 4In [description], the integral Z-shaped spar 100 includes a Z-shaped spar fitting 102 in the upper skin panel 34, which is configured to receive a portion of the trailing edge closure 104, and the Z-shaped spar fitting 102 is positioned in front of the first bend 106. The Z-shaped spar fitting 102 is actually a small offset in the upper skin panel 34 towards the lower skin panel 36. In this example, the first lid end region 118 of the trailing edge closure 104 is joined to the upper skin panel 34 rather than the lower skin panel 36. Additionally or alternatively, the first lid end region 118 can be riveted or otherwise fastened or joined to the upper skin panel 34. To create a smooth surface at the interface and improve aerodynamic performance, as Figure 4 shown in [reference], the first lid end region 118 can be slightly recessed into the upper skin panel 34, such as via the Z-shaped spar fitting 102. Depending on the thickness of the first lid end region 118, the Z-shaped spar fitting 102 can be customized to create a larger or smaller recess in the upper skin panel 34, such that the upper panel surface 130 of the upper skin panel 34 is substantially flush with the upper lid surface 132 of the trailing edge closure 104 within the first lid end region 118. In other words, the Z-shaped spar fitting 102 can be larger to create a larger recess to receive and join a given trailing edge closure 104 with a thicker first lid end region 118, while the Z-shaped spar fitting 102 can be smaller to create a smaller recess to receive and join a different given trailing edge closure 104 with a thinner first lid end region 118.

[0041] The second lid end region 120 of the trailing edge closure 104 can include an integral wedge 122, which can be joined (e.g., adhesively and / or via one or more fasteners) to the lower skin panel 36. Alternatively, and as Figure 4 shown in [reference], the integral wedge 122 can be integrally formed with the lower skin panel 36. In still other examples, the integral wedge 122 can be a separate component that is separate from the trailing edge closure 104 and the lower skin panel 36, and can be adhesively joined or otherwise joined to the lower skin panel 36 and / or the trailing edge closure 104. For example, the integral wedge 122 can be formed by stacking material layers, molding, and / or machining mating surface profiles to mate with the lower skin panel 36.

[0042] Figure 5 A flowchart schematically provides an illustrative non-exclusive example of a method 200 according to the present disclosure. In Figure 5 [description], some steps are shown in dashed boxes, indicating that such steps can be optional or can correspond to an optional version of the method according to the present disclosure. That is, not all methods 200 according to the present disclosure are required to include the steps shown in the solid boxes. From the discussion herein, it can be understood that Figure 5The method 200 and steps shown are not restrictive, and other methods and steps are also within the scope of the present disclosure, including methods having a number of steps greater than or less than the number of steps shown.

[0043] Method 200 generally includes coupling an upper skin panel (e.g., upper skin panel 34) to a front C-channel spar (e.g., front C-channel spar 38) at 202, and coupling a lower skin panel (e.g., lower skin panel 36) to the front C-channel spar at 204. Coupling the upper skin panel to the front C-channel spar at 202 generally includes coupling the upper skin panel to the upper flange (e.g., upper flange 42) of the front C-channel spar. Similarly, coupling the lower skin panel to the front C-channel spar at 204 generally includes coupling the lower skin panel to the lower flange (e.g., lower flange 44) of the front C-channel spar. Compared to the prior art, coupling the upper skin panel at 202 and / or coupling the lower skin panel at 204 can be performed using a reduced number of nut plates or other fastening components. Additionally or alternatively, coupling the upper skin panel at 202 and / or coupling the lower skin panel at 204 can be performed without using splice strips. Reducing the number of fasteners or fastening components can reduce the weight of the resulting structural composite airfoil, lower the manufacturing cost, and / or reduce the manufacturing processing time.

[0044] Method 200 further includes coupling a leading edge skin panel (e.g., leading edge skin panel 54) to the front C-channel spar at 206. Coupling the leading edge skin panel at 206 generally includes coupling a first end region (e.g., first end region 56) of the leading edge skin panel to the upper flange of the front C-channel spar and coupling a second end region (e.g., second end region 58) of the leading edge skin panel to the lower flange of the front C-channel spar. The coupling of the leading edge skin panel at 206 can be performed without overlapping the leading edge skin panel on the upper flange of the upper skin panel and the front C-channel spar. Similarly, the coupling of the leading edge skin panel at 206 can be performed without overlapping the leading edge skin panel on the lower flange of the lower skin panel and the front C-channel spar. Coupling the leading edge skin panel at 206 can include coupling the leading edge skin panel without using a splice strip such that the leading edge skin panel can be directly coupled to the front C-channel spar. In some methods 200, coupling the leading edge skin panel at 206 includes adjoining a first end region of the leading edge skin panel to the upper skin panel (e.g., upper leading edge end 76 of upper skin panel 34), which can include forming a lap or splice joint therebetween. Additionally or alternatively, coupling the leading edge skin panel at 206 can include adjoining a second end region of the leading edge skin panel to the lower skin panel (e.g., lower leading edge end 78 of lower skin panel 36) and / or forming a lap or splice joint therebetween.

[0045] In some examples, method 200 includes coupling an upper skin panel to an intermediate C-channel spar (e.g., intermediate C-channel spar 60) at 208, coupling the upper skin panel to a rear C-channel spar (e.g., rear C-channel spar 62) at 210, coupling a lower skin panel to the intermediate C-channel spar at 212, and / or coupling the lower skin panel to the rear C-channel spar at 214. Additionally or alternatively, method 210 may include coupling an auxiliary structural element such as a closure (e.g., auxiliary structural element 28) to the upper skin panel (e.g., upper trailing edge end 92) and / or the lower skin panel (e.g., lower trailing edge end 94) at 216. Additionally or alternatively, method 200 may include forming an integral Z-spar (e.g., integral Z-spar 100) in the lower skin panel or the upper skin panel at 218.

[0046] Illustrative, non-exclusive examples of the inventive subject matter in accordance with the present disclosure are described in the paragraphs listed below:

[0047] A1. A structural composite airfoil (10) having a leading edge (22) and a trailing edge (24), the structural composite airfoil (10) comprising:

[0048] A main structural element (26) extending from a leading edge region (30) to a trailing edge region (32), wherein the leading edge region (30) is adjacent to or defines the leading edge (22) of the structural composite airfoil (10), and wherein the main structural element (26) comprises:

[0049] An upper skin panel (34);

[0050] A lower skin panel (36);

[0051] An internal volume (40) defined between the upper skin panel (34) and the lower skin panel (36); and

[0052] A front C-channel spar (38) including an upper flange (42) coupled to the upper skin panel (34), wherein the front C-channel spar (38) further includes a lower flange (44) coupled to the lower skin panel (36), wherein a first channel (46) of the front C-channel spar (38) faces the leading edge (22) of the structural composite airfoil (10), wherein the upper flange (42) forms a first angle (48) with an elongate span portion (50) of the front C-channel spar (38), wherein the lower flange (44) forms a second angle (52) with the elongate span portion (50), and wherein the first angle (48) is an acute angle;

[0053] An auxiliary structural element (28) defining the trailing edge (24) of the structural composite airfoil (10); and

[0054] The leading edge skin panel (54) defines the leading edge (22) of the structural composite airfoil (10) and is positioned near or within the leading edge region (30) of the primary structural element (26), wherein a first end region (56) of the leading edge skin panel (54) is coupled to the upper flange (42) of the front C-channel spar (38), wherein a second end region (58) of the leading edge skin panel (54) is coupled to the lower flange (44) of the front C-channel spar (38), and wherein the leading edge skin panel (54) has a bull nose shape.

[0055] A1.1. The structural composite airfoil (10) according to paragraph A1, wherein the primary structural element further comprises an intermediate C-channel spar (60) coupled to the upper skin panel (34) and the lower skin panel (36), wherein a second channel (64) of the intermediate C-channel spar (60) faces the leading edge (22) of the structural composite airfoil (10), and wherein the intermediate C-channel spar (60) is positioned behind the front C-channel spar (38).

[0056] A1.2. The structural composite airfoil (10) according to paragraph A1 and / or A1.1, wherein the primary structural element further comprises a rear C-channel spar (62) coupled to the upper skin panel (34) and the lower skin panel (36), wherein a third channel (70) of the rear C-channel spar (62) faces the leading edge (22) of the structural composite airfoil (10), and wherein the rear C-channel spar (62) is positioned behind the intermediate C-channel spar (60).

[0057] A2. The structural composite airfoil (10) according to any one of paragraphs A1 - A1.2, wherein the second angle (52) is an acute angle.

[0058] A3. The structural composite airfoil (10) according to any one of paragraphs A1 - A2, wherein the upper flange (42) is angled relative to the elongated span portion (50) to be complementary to the first end region (56) of the leading edge skin panel (54).

[0059] A4. The structural composite airfoil (10) according to any one of paragraphs A1 - A3, wherein the leading edge skin panel (54) does not overlap with the upper skin panel (34) on the upper flange (42) of the front C-channel spar (38).

[0060] A5. The structural composite airfoil (10) according to any one of paragraphs A1 - A4, wherein the lower flange (44) is angled to be complementary to the second end region (58) of the leading edge skin panel (54).

[0061] A6. The structural composite airfoil (10) according to any one of paragraphs A1 - A5, wherein the leading edge skin panel (54) does not overlap with the lower skin panel (36) on the lower flange (44) of the front C - shaped channel spar (38).

[0062] A7. The structural composite airfoil (10) according to any one of paragraphs A1 - A6, wherein the upper skin panel (34) abuts the leading edge skin panel (54).

[0063] A8. The structural composite airfoil (10) according to any one of paragraphs A1 - A7, wherein the lower skin panel (36) abuts the leading edge skin panel (54).

[0064] A9. The structural composite airfoil (10) according to any one of paragraphs A1 - A8, further comprising a first fastener (80) that couples the leading edge skin panel (54) to the upper flange (42) of the front C - shaped channel spar (38).

[0065] A10. The structural composite airfoil (10) according to any one of paragraphs A1 - A9, further comprising a second fastener (82) that couples the leading edge skin panel (54) to the lower flange (44) of the front C - shaped channel spar (38).

[0066] A11. The structural composite airfoil (10) according to any one of paragraphs A1 - A10, further comprising a third fastener (84) that couples the upper skin panel (34) to the upper flange (42) of the front C - shaped channel spar (38).

[0067] A11.1. The structural composite airfoil (10) according to paragraph A11, wherein the third fastener (84) is not blind, such that when assembling the main structural elements (26), the third fastener is accessible.

[0068] A12. The structural composite airfoil (10) according to any one of paragraphs A1 - A11.1, further comprising a fourth fastener (86) that couples the lower skin panel (36) to the lower flange (44) of the front C - shaped channel spar (38).

[0069] A12.1. The structural composite airfoil (10) according to paragraph A12, wherein the fourth fastener (86) is not blind, such that when assembling the main structural elements (26), the fourth fastener is accessible.

[0070] A13. The structural composite airfoil (10) according to any one of paragraphs A1 - A12.1, wherein the leading edge skin panel (54) is joined to the upper skin panel (34) without any wobbling.

[0071] A14. The structural composite airfoil (10) according to any one of paragraphs A1 - A13, wherein the leading edge skin panel (54) is joined to the lower skin panel (36) without any wobbling.

[0072] A15. The structural composite airfoil (10) according to any one of paragraphs A1 - A14, wherein the leading edge skin panel (54) is joined to the upper skin panel (34) without a splice strip.

[0073] A16. The structural composite airfoil (10) according to any one of paragraphs A1 - A15, wherein the leading edge skin panel (54) is joined to the lower skin panel (36) without a splice strip.

[0074] A17. The structural composite airfoil (10) according to any one of paragraphs A1 - A16, further comprising a plurality of fasteners (84, 86) that couple the front C - channel spar (38) to the upper skin panel (34) and the lower skin panel (36), wherein each fastener (84, 86) of the plurality of fasteners (84, 86) is not a blind hole such that each fastener (84, 86) of the plurality of fasteners (84, 86) is accessible when the front C - channel spar (38) is secured to the upper skin panel (34) and the lower skin panel (36).

[0075] A18. The structural composite airfoil (10) according to any one of paragraphs A1 - A17, wherein the upper skin panel (34) is coupled to the upper flange (42) of the front C - channel spar (38) without a nut plate.

[0076] A19. The structural composite airfoil (10) according to any one of paragraphs A1 - A18, wherein the lower skin panel (36) is coupled to the lower flange (44) of the front C - channel spar (38) without a nut plate.

[0077] A20. The structural composite airfoil (10) according to any one of paragraphs A1 - A19, wherein at least a portion of the upper skin panel (34) is core - reinforced.

[0078] A21. The structural composite airfoil (10) according to any one of paragraphs A1 - A20, wherein at least a portion of the lower skin panel (36) is core - reinforced.

[0079] A22. The structural composite airfoil (10) according to any one of paragraphs A1 - A21, wherein the upper skin panel (34) comprises fiberglass or carbon fiber.

[0080] A23. The structural composite airfoil (10) according to any one of paragraphs A1 - A22, wherein the lower skin panel (36) comprises fiberglass or carbon fiber.

[0081] A24. The structural composite airfoil (10) according to any one of paragraphs A1 - A23, wherein the structural composite airfoil (10) has a length (90), and wherein a position along the length (90) can be defined by a percentage of the distance from the leading edge (22) along the length (90).

[0082] A25. The structural composite airfoil (10) according to paragraph A24, wherein the front C - channel spar (38) is positioned between 0% and 10% of the length (90) away from the leading edge (22).

[0083] A26. The structural composite airfoil (10) according to paragraph A25, wherein the front C - channel spar (38) is positioned at approximately 5% of the length (90) away from the leading edge (22).

[0084] A27. The structural composite airfoil (10) according to any one of paragraphs A24 - A26, wherein the middle C - channel spar (60) is positioned between 20% and 40% of the length (90) away from the leading edge (22).

[0085] A28. The structural composite airfoil (10) according to paragraph A27, wherein the middle C - channel spar (60) is positioned at approximately 30% of the length (90) away from the leading edge (22).

[0086] A29. The structural composite airfoil (10) according to any one of paragraphs A24 - A28, wherein the rear C - channel spar (62) is positioned between 40% and 70% of the length (90) away from the leading edge (22), and / or between 50% and 60% of the length (90) away from the leading edge (22).

[0087] A30. The structural composite airfoil (10) according to paragraph A29, wherein the rear C - channel spar (62) is positioned at approximately 55% of the length (90) away from the leading edge (22).

[0088] A31. The structural composite airfoil (10) according to any one of paragraphs A1 - A30, wherein the structural composite airfoil (10) is a trailing edge flap (17), aileron, flaperon, air brake, elevator, slat, spoiler, canard, rudder, and / or winglet.

[0089] A32. The structural composite airfoil (10) according to any one of paragraphs A1 - A31, wherein the auxiliary structural element (28) includes a wedge closure.

[0090] A33. The structural composite airfoil (10) according to any one of paragraphs A1 - A32, wherein the auxiliary structural element (28) includes a duckbill closure.

[0091] A34. The structural composite airfoil (10) according to any one of paragraphs A1 - A33, wherein the auxiliary structural element (28) includes a combined closure.

[0092] A35. The structural composite airfoil (10) according to any one of paragraphs A1 - A34, wherein the auxiliary structural element (28) includes a riveted closure.

[0093] A36. The structural composite airfoil (10) according to any one of paragraphs A1 - A35, wherein the lower skin panel (36) includes a lower leading edge end (78) and a lower trailing edge end (94), and the lower trailing edge end (94) is opposite to the lower leading edge end (78).

[0094] A37. The structural composite airfoil (10) according to paragraph A36, wherein the lower leading edge end (78) is connected to the front C - shaped channel spar (38).

[0095] A38. The structural composite airfoil (10) according to any one of paragraphs A36 - A37, wherein the lower trailing edge end (94) is connected to the upper trailing edge end (92) of the upper skin panel (34).

[0096] A39. The structural composite airfoil (10) according to any one of paragraphs A36 - A38, wherein the lower trailing edge end (94) forms an integral Z - shaped spar (100).

[0097] A40. The structural composite airfoil (10) according to any one of paragraphs A1 - A39, wherein the main structural element (26) includes an integral Z - shaped spar (100).

[0098] A41. The structural composite airfoil (10) according to paragraph A40, wherein the integral Z-beam (100) is formed by the lower skin panel (36) within the trailing edge region (32) of the main structural element (26).

[0099] A42. The structural composite airfoil (10) according to any one of paragraphs A40 - A41, wherein the integral Z-beam (100) includes a joint configured to receive a portion of the trailing edge closure cover (104).

[0100] A43. The structural composite airfoil (10) according to any one of paragraphs A40 - A42, wherein the integral Z-beam (100) includes a first bend (106), a second bend (108), and a first Z-beam segment (110) extending between the first bend (106) and the second bend (108).

[0101] A44. The structural composite airfoil (10) according to paragraph A43, wherein the first Z-beam segment (110) is substantially perpendicular to the lower skin panel (36) and / or substantially perpendicular to the upper skin panel (34).

[0102] A45. The structural composite airfoil (10) according to paragraph A43 or A44, wherein the integral Z-beam (100) further includes a second Z-beam segment (112) extending rearward of the second bend (108), wherein the second Z-beam segment (112) is coupled to the upper skin panel (34).

[0103] A46. The structural composite airfoil (10) according to paragraph A45, wherein the second Z-beam segment (112) is adjacent to the inner surface (114) of the upper skin panel (34).

[0104] A47. The structural composite airfoil (10) according to paragraph A45 or A46, wherein the second Z-beam segment (112) is coupled to the upper skin panel (34) via a Z-beam fastener (116), wherein the Z-beam fastener (116) is recessed into the upper skin panel (34), and wherein the Z-beam fastener (116) extends through the second Z-beam segment (112).

[0105] A48. The structural composite airfoil (10) according to any one of paragraphs A43 - A47, wherein the joint of the integral Z-beam (100) is in front of the first bend (106).

[0106] A49. The structural composite airfoil (10) according to any one of paragraphs A1 - A48 further includes a trailing edge closure cover (104).

[0107] A50. The structural composite airfoil (10) according to paragraph A50, wherein a first cover end region (118) of the trailing edge closure cover (104) is joined to the lower skin panel (36).

[0108] A51. The structural composite airfoil (10) according to paragraph A49 or A50, wherein a first cover end region (118) of the trailing edge closure cover (104) is recessed into the lower skin panel (36) such that the aerodynamic performance is improved.

[0109] A52. The structural composite airfoil (10) according to any one of paragraphs A49 - A51, wherein a second cover end region (120) of the trailing edge closure cover (104) includes an integral wedge (122) joined to the upper skin panel (34).

[0110] A53. The structural composite airfoil (10) according to any one of paragraphs A1 - A52, wherein the upper skin panel (34) includes an upper leading edge end (76) and an upper trailing edge end (92), wherein the upper trailing edge end (92) is opposite to the upper leading edge end (76).

[0111] A54. The structural composite airfoil (10) according to paragraph A53, wherein the upper leading edge end (76) is joined to the front C - channel spar (38).

[0112] A55. The structural composite airfoil (10) according to any one of paragraphs A53 - A54, wherein the upper trailing edge end (92) is joined to a lower trailing edge end (94) of the lower skin panel (36).

[0113] B1. An aircraft (14) comprising the structural composite airfoil (10) according to any one of paragraphs A1 - A55.

[0114] B2. A trailing edge flap (17) for an aircraft (14) comprising the structural composite airfoil (10) according to any one of paragraphs A1 - A55.

[0115] C1. A method (200) of assembling a structural composite airfoil (10), the method (200) comprising:

[0116] Couple the upper skin panel (34) to the front C-channel spar (38), wherein the structural composite airfoil (10) extends from a leading edge (22) to a trailing edge (24), wherein a first channel (46) of the front C-channel spar (38) faces the leading edge (22) of the structural composite airfoil (10), wherein the front C-channel spar (38) includes an upper flange (42), a lower flange (44), and an elongated span portion (50) extending between the upper flange (42) and the lower flange (44), wherein the step of coupling the upper skin panel (34) to the front C-channel spar (38) includes coupling the upper skin panel (34) to the upper flange (42) of the front C-channel spar (38), and wherein the upper flange (42) forms an acute angle with the elongated span portion (50);

[0117] Couple the lower skin panel (36) to the front C-channel spar (38) such that an internal volume (40) is defined between the upper skin panel (34) and the lower skin panel (36), wherein the upper skin panel (34), the lower skin panel (36), and the front C-channel spar (38) together form at least a part of the main structural element (26) of the structural composite airfoil (10); and

[0118] Couple the leading edge skin panel (54) to the front C-channel spar (38), wherein the leading edge skin panel (54) defines the leading edge (22) of the structural composite airfoil (10), wherein the step of coupling (206) the leading edge skin panel (54) includes coupling a first end region (56) of the leading edge skin panel (54) to the upper flange (42) of the front C-channel spar (38), wherein the step of coupling (206) the leading edge skin panel (54) further includes coupling a second end region (58) of the leading edge skin panel (54) to the lower flange (44) of the front C-channel spar (38), and wherein the leading edge skin panel (54) has a bullnose shape.

[0119] C1.1. The method (200) according to paragraph C1, further comprising coupling the upper skin panel (34) to an intermediate C-channel spar (60), wherein the intermediate C-channel spar (60) is behind the front C-channel spar (38), and wherein a second channel (64) of the intermediate C-channel spar (60) faces the leading edge (22) of the structural composite airfoil (10).

[0120] C1.2. The method (200) according to paragraph C1 or C1.1, further comprising coupling the upper skin panel (34) to a rear C-channel spar (62), and wherein a third channel (70) of the rear C-channel spar (62) faces the leading edge (22) of the structural composite airfoil (10).

[0121] C1.3. The method (200) according to paragraph C1.2, wherein the rear C-channel spar (62) is behind the middle C-channel spar (60).

[0122] C1.4. The method (200) according to any one of paragraphs C1 - C1.3, further comprising coupling (214) the lower skin panel to the rear C-channel spar (62) such that the rear C-channel spar (62) is part of the main structural element (26).

[0123] C1.5. The method (200) according to any one of paragraphs C1 - C1.4, further comprising coupling (212) the lower skin panel (36) to the middle C-channel spar (60) such that the middle C-channel spar (60) is part of the main structural element (26).

[0124] C2. The method (200) according to any one of paragraphs C1 - C1.5, wherein the structural composite airfoil (10) is the structural composite airfoil (10) according to any one of paragraphs A1 - A55.

[0125] C3. The method (200) according to any one of paragraphs C1 - C2, wherein the step of coupling (206) the leading edge skin panel (54) comprises coupling the leading edge skin panel (54) such that the leading edge skin panel (54) does not overlap the upper skin panel (34) on the upper flange (42) of the front C-channel spar (38).

[0126] C4. The method (200) according to any one of paragraphs C1 - C3, wherein the step of coupling (206) the leading edge skin panel (54) comprises coupling the leading edge skin panel (54) such that the leading edge skin panel (54) does not overlap the lower skin panel (36) on the lower flange (44) of the front C-channel spar (38).

[0127] C5. The method (200) according to any one of paragraphs C1 - C4, wherein the step of coupling (206) the leading edge skin panel (54) comprises adjoining a first end region (56) of the leading edge skin panel (54) and the upper skin panel (34).

[0128] C6. The method (200) according to any one of paragraphs C1 - C5, wherein the step of coupling (206) the leading edge skin panel (54) comprises adjoining a second end region (58) of the leading edge skin panel (54) and the lower skin panel (36).

[0129] C7. The method (200) according to any one of paragraphs C1 - C6, wherein the coupling (206) of the leading edge skin panel (54) is performed without using a splice tape.

[0130] C8. The method (200) according to any one of paragraphs C1 - C7, wherein the connection (202) of the upper skin panel (34) to the front C - channel spar (38) is performed without using a nut plate.

[0131] C9. The method (200) according to any one of paragraphs C1 - C8, wherein the connection (204) of the lower skin panel (36) to the front C - channel spar (38) is performed without using a nut plate.

[0132] C10. The method (200) according to any one of paragraphs C1 - C9, further comprising connecting (216) a closure to the upper skin panel (34) and the lower skin panel (36), wherein the closure defines a trailing edge (24) of the structural composite airfoil (10).

[0133] C11. The method (200) according to any one of paragraphs C1 - C10, further comprising forming an integral Z - spar (100) in the lower skin panel (36).

[0134] D1. Use of a structural composite airfoil (10) according to any one of paragraphs A1 - A55 as an inboard flap of an aircraft (14).

[0135] D2. Use of a structural composite airfoil (10) according to any one of paragraphs A1 - A55 as an outboard flap of an aircraft (14).

[0136] As used herein, the terms "selective" and "selectively", when modifying an action, movement, configuration, or other activity of one or more components or features of a device, mean that the particular action, movement, configuration, or other activity is a direct or indirect result of a user's manipulation of the device or of one or more components.

[0137] As used herein, the terms "adapt" and "configured" refer to an element, component, or other subject being designed and / or intended to perform a given function. Thus, the use of the terms "adapt" and "configured" should not be construed to mean that a given element, component, or other subject is merely "able to" perform the given function, but rather that the element, component, and / or other subject has been specifically selected, created, implemented, utilized, programmed, and / or designed for the purpose of performing the function. Within the scope of the present disclosure, an element, component, and / or other recited subject that is recited as being adapted to perform a particular function may alternatively or additionally be described as being configured to perform that function, and vice versa. Similarly, a subject that is recited as being configured to perform a particular function may alternatively or additionally be described as being operable to perform that function.

[0138] As used herein, the phrase "at least one" with respect to a list of one or more entities shall be understood to mean at least one entity selected from any one or more of the entities in the list of entities, but not necessarily including at least one of each entity expressly listed within the list of entities, and not excluding any combination of entities in the list of entities. This definition also allows for the optional presence of entities, whether related or unrelated to those specifically identified, in addition to the entities specifically identified within the list of entities referred to in the phrase "at least one". Thus, by way of non-limiting example, "at least one of A and B" (or equivalently, "at least one of A or B", or equivalently "at least one of A and / or B") can mean, in one embodiment, no at least one of B, optionally including more than one A (and optionally including entities other than B); in another embodiment, no at least one of A, optionally including more than one B (and optionally including entities other than A); in another embodiment, means at least one A, optionally including more than one A, and at least one B, optionally including more than one B (and optionally including other entities). In other words, the phrases "at least one", "one or more", and "and / or" are open-ended expressions that can be both conjunctive and disjunctive in operation. For example, each of the expressions "at least one of A, B, and C", "at least one of A, B, or C", "one or more of A, B, and C", "one or more of A, B, or C", and "A, B, and / or C" can mean A alone, B alone, C alone, A and B together, A and C together, B and C together, or A, B, and C together, and optionally any of the foregoing in combination with at least one other entity.

[0139] Not all devices and methods according to the present disclosure require the various elements of the devices and steps of the methods disclosed herein, and the present disclosure includes all novel and non-obvious combinations and sub-combinations of the various elements and steps disclosed herein. In addition, one or more of the various elements and steps disclosed herein can define an independent inventive subject matter that is independent of and separate from the overall disclosed device or method. Thus, such inventive subject matter does not need to be associated with the specific devices and methods expressly disclosed herein, and such inventive subject matter can find utility in devices and / or methods not expressly disclosed herein.

[0140] As used herein, the phrases "for example", "as an example", and / or the term "example" in short, when used with reference to one or more components, features, details, structures, embodiments, and / or methods in accordance with the present disclosure, are intended to convey that the described components, features, details, structures, embodiments, and / or methods are illustrative, non-exclusive examples of the components, features, details, structures, embodiments, and / or methods in accordance with the present disclosure. Thus, the described components, features, details, structures, embodiments, and / or methods are not intended to be restrictive, required, or exclusive / exhaustive; and other components, features, details, structures, embodiments, and / or methods that are structurally and / or functionally similar and / or equivalent are also within the scope of the present disclosure.

Claims

1. A structural composite airfoil (10) having a leading edge (22) and a trailing edge (24), the structural composite airfoil comprising: A main structural element (26) extending from a leading edge region (30) to a trailing edge region (32), wherein the leading edge region is adjacent to the leading edge of the structural composite airfoil, and wherein the main structural element comprises: an upper skin panel (34); a lower skin panel (36); an internal volume (40) defined between the upper skin panel and the lower skin panel; and a front C-channel spar (38) including an upper flange (42) coupled to the upper skin panel, wherein the front C-channel spar further includes a lower flange (44) coupled to the lower skin panel, wherein a first channel (46) of the front C-channel spar faces the leading edge of the structural composite airfoil, wherein the upper flange forms a first angle (48) with an elongate span portion (50) of the front C-channel spar, wherein the lower flange forms a second angle (52) with the elongate span portion, and wherein the first angle is an acute angle; An auxiliary structural element (28) defining the trailing edge of the structural composite airfoil; A leading edge skin panel (54) defining the leading edge of the structural composite airfoil and positioned adjacent to the leading edge region of the main structural element, wherein a first end region (56) of the leading edge skin panel is coupled to the upper flange of the front C-channel spar, wherein a second end region (58) of the leading edge skin panel is coupled to the lower flange of the front C-channel spar, and wherein the leading edge skin panel has a bullnose shape; and A trailing edge closure (104), wherein: A first closure end region (118) of the trailing edge closure is bonded to the lower skin panel, and the first closure end region of the trailing edge closure is recessed into the lower skin panel; and A second closure end region (120) of the trailing edge closure includes an integral wedge (122) coupled to the upper skin panel.

2. The structural composite airfoil according to claim 1, wherein, The leading edge skin panel does not overlap the upper skin panel on the upper flange of the front C-channel spar, and wherein the leading edge skin panel does not overlap the lower skin panel on the lower flange of the front C-channel spar.

3. The structural composite airfoil according to claim 1, further comprising: A first fastener coupling the leading edge skin panel to the upper flange of the front C-channel spar; A second fastener coupling the leading edge skin panel to the lower flange of the front C-channel spar; A third fastener coupling the upper skin panel to the upper flange of the front C-channel spar, wherein the third fastener is not blind such that the third fastener is accessible when assembling the main structural element; and A fourth fastener coupling the lower skin panel to the lower flange of the front C-channel spar, wherein the fourth fastener is not blind such that the fourth fastener is accessible when assembling the main structural element.

4. The structural composite airfoil according to claim 1, wherein, The leading edge skin panel is joined to the upper skin panel without any wobbling, and wherein the leading edge skin panel is joined to the lower skin panel without any wobbling.

5. The structural composite airfoil according to claim 1, wherein, The structural composite airfoil is a trailing edge flap, aileron, flaperon, air brake, elevator, slat, spoiler, canard, rudder, and / or winglet.

6. The structural composite airfoil according to claim 1, wherein, The auxiliary structural element includes a wedge-shaped closure.

7. The structural composite airfoil according to claim 1, wherein, The lower skin panel includes a lower leading edge end and a lower trailing edge end, wherein the lower trailing edge end is opposite the lower leading edge end, and wherein the lower leading edge end is coupled to the front C-channel spar.

8. The structural composite airfoil according to claim 7, wherein, The lower trailing edge end is coupled to the upper trailing edge end of the upper skin panel.

9. The structural composite airfoil according to claim 1, wherein, The upper skin panel includes an upper leading edge end and an upper trailing edge end, wherein the upper trailing edge end is opposite the upper leading edge end, and wherein the upper leading edge end is coupled to the front C-channel spar.

10. The structural composite airfoil according to claim 9, wherein, The upper trailing edge end is coupled to the integral wedge of the trailing edge closure cover.

11. The structural composite airfoil according to claim 1, wherein, The second angle is an acute angle.

12. The structural composite airfoil according to claim 1, wherein, The front C-channel spar is directly coupled to the upper skin panel and the lower skin panel without splice straps or nut plates.

13. The structural composite airfoil according to claim 1, wherein, The main structural element further includes: An intermediate C-channel spar, coupled to the upper skin panel and the lower skin panel, wherein a second channel of the intermediate C-channel spar faces the leading edge of the structural composite airfoil, and wherein the intermediate C-channel spar is positioned behind the front C-channel spar; and A rear C-channel spar, coupled to the upper skin panel and the lower skin panel, wherein a third channel of the rear C-channel spar faces the leading edge of the structural composite airfoil, and wherein the rear C-channel spar is positioned behind the intermediate C-channel spar.

14. An aircraft, comprising the structural composite airfoil according to claim 1.

15. A trailing edge flap for an aircraft, comprising the structural composite airfoil according to claim 1.

16. A method of assembling a structural composite airfoil (10), the method comprising: Coupling an upper skin panel (34) to a front C-channel spar (38), wherein the structural composite airfoil extends from a leading edge (22) to a trailing edge (24), wherein a first channel (46) of the front C-channel spar faces the leading edge of the structural composite airfoil, wherein the front C-channel spar includes an upper flange (42), a lower flange (44), and an elongated span portion (50) extending between the upper flange and the lower flange, and wherein the step of coupling the upper skin panel to the front C-channel spar includes coupling the upper skin panel to the upper flange of the front C-channel spar, and wherein the upper flange forms an acute angle with the elongated span portion; Coupling a lower skin panel (36) to the front C-channel spar such that an internal volume (40) is defined between the upper skin panel and the lower skin panel, wherein the upper skin panel, the lower skin panel, and the front C-channel spar together form at least a portion of a main structural element (26) of the structural composite airfoil; Couple a leading edge skin panel (54) to the front C-channel spar, wherein the leading edge skin panel defines the leading edge of the structural composite airfoil, and wherein the step of coupling the leading edge skin panel includes coupling a first end region (56) of the leading edge skin panel to the upper flange of the front C-channel spar, and wherein the step of coupling the leading edge skin panel further includes coupling a second end region (58) of the leading edge skin panel to the lower flange of the front C-channel spar, and wherein the leading edge skin panel has a bull nose shape; and Attach a trailing edge closure (104) to the lower skin panel, wherein a first closure end region (118) of the trailing edge closure is bonded to the lower skin panel and is recessed into the lower skin panel; and a second closure end region (120) of the trailing edge closure includes an integral wedge (122) coupled to the upper skin panel.

17. A structural composite airfoil (10) having a leading edge and a trailing edge, the structural composite airfoil comprising: A main structural element (26) extending from a leading edge region (30) to a trailing edge region (32), wherein the leading edge region defines the leading edge of the structural composite airfoil, and wherein the main structural element includes: an upper skin panel (34); a lower skin panel (36); an internal volume (40) defined between the upper skin panel and the lower skin panel; a front C-channel spar (38) including an upper flange (42) coupled to the upper skin panel, and wherein the front C-channel spar further includes a lower flange (44) coupled to the lower skin panel, and wherein a first channel (46) of the front C-channel spar faces the leading edge of the structural composite airfoil, and wherein the upper flange forms a first angle (48) with an elongate span portion (50) of the front C-channel spar, and wherein the lower flange forms a second angle (52) with the elongate span portion, and wherein the first angle is an acute angle; and A leading edge skin panel (54) defining the leading edge of the structural composite airfoil and positioned within the leading edge region of the main structural element, wherein a first end region (56) of the leading edge skin panel is coupled to the upper flange of the front C-channel spar, and wherein a second end region (58) of the leading edge skin panel is coupled to the lower flange of the front C-channel spar, and wherein the leading edge skin panel has a bull nose shape; An auxiliary structural element (28) defining the trailing edge of the structural composite airfoil; and A trailing edge closure (104), wherein: A first closure end region (118) of the trailing edge closure is bonded to the lower skin panel and is recessed into the lower skin panel; and A second closure end region (120) of the trailing edge closure includes an integral wedge (122) coupled to the upper skin panel.

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