阅读说明：本技术 具有直接联接的前翼梁的结构复合翼型件及相关方法 (Structural composite airfoil with directly coupled front spar and related method ) 是由 布赖恩·格鲁纳 彼得·舒普 于 2021-05-20 设计创作，主要内容包括：一种结构复合翼型件,包括主要结构元件、限定结构复合翼型件的后缘的辅助结构元件以及限定结构复合翼型件的前缘的前缘蒙皮面板。主要结构元件包括上蒙皮面板、下蒙皮面板和前C形通道翼梁。前C形通道翼梁的第一通道面向结构复合翼型件的前缘,并且前C形通道翼梁的上凸缘与前C形通道翼梁的细长跨度部形成锐角。前缘蒙皮面板与主要结构元件的前缘区域相邻定位,前缘蒙皮面板的第一端部区域联接至前C形通道翼梁的上凸缘,并且第二端部区域联接至前C形通道翼梁的下凸缘。(A structural composite airfoil includes a primary structural element, a secondary 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 primary structural elements include an upper skin panel, a lower skin panel, and a forward C-channel spar. The first channel of the front C-channel spar faces the leading edge of the structural composite airfoil, and the upper flange of the front C-channel spar forms an acute angle with the elongated span of the front C-channel spar. A leading edge skin panel is positioned adjacent a leading edge region of the primary structural element, a first end region of the leading edge skin panel being coupled to the upper flange of the forward C-channel spar, and a second end region being coupled to the lower flange of the forward C-channel spar.)

1. A structural composite airfoil having a leading edge and a trailing edge, comprising:

a primary structural element extending from a leading edge region to a trailing edge region, wherein the leading edge region is adjacent to the leading edge of the structural composite airfoil, wherein the primary structural element comprises: an upper skin panel; a lower skin panel; an interior volume defined between the upper skin panel and the lower skin panel; and a forward C-channel spar comprising an upper flange coupled to the upper skin panel, wherein the forward C-channel spar further comprises a lower flange coupled to the lower skin panel, wherein a first channel of the forward C-channel spar faces the leading edge of the structural composite airfoil, wherein the upper flange forms a first angle with an elongated span of the forward C-channel spar, wherein the lower flange forms a second angle with the elongated span, and wherein the first angle is an acute angle;

an auxiliary structural element defining the trailing edge of the structural composite airfoil; and

a leading edge skin panel defining the leading edge of the structural composite airfoil and positioned adjacent to the leading edge region of the primary structural element, wherein a first end region of the leading edge skin panel is coupled to the upper flange of the forward C-channel spar, wherein a second end region of the leading edge skin panel is coupled to the lower flange of the forward C-channel spar, and wherein the leading edge skin panel has a bull-nose shape.

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

3. The structural composite airfoil of claim 1, further comprising:

a first fastener coupling the leading edge skin panel to the upper flange of the forward C-channel spar;

a second fastener coupling the leading edge skin panel to the lower flange of the forward C-channel spar;

a third fastener coupling the upper skin panel to the upper flange of the forward C-channel spar, wherein the third fastener is not blind, such that the third fastener is accessible when the primary structural element is assembled; and

a fourth fastener coupling the lower skin panel to the lower flange of the forward C-channel spar, wherein the fourth fastener is not blind, such that the fourth fastener is accessible when the primary structural element is assembled.

4. The structural composite airfoil of claim 1, wherein the leading edge skin panel is engaged with the upper skin panel without any sloshing, and wherein the leading edge skin panel is engaged with the lower skin panel without any sloshing.

5. The structural composite airfoil of claim 1, wherein the structural composite airfoil is a trailing edge flap, aileron, flaperon, airbrake, elevator, slat, spoiler, canard, rudder, and/or winglet.

6. The structural composite airfoil of claim 1, wherein the auxiliary structural element comprises a wedge enclosure.

7. The structural composite airfoil of 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, wherein the lower leading edge end is coupled to the forward C-channel spar.

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

9. The structural composite airfoil of claim 1, further comprising a trailing edge closure cap, wherein a first cap end region of the trailing edge closure cap is bonded to the lower skin panel.

10. The structural composite airfoil of claim 9, wherein the first cap end region of the trailing edge closure cap is recessed into the lower skin panel.

11. The structural composite airfoil of claim 9, wherein the second cap end region of the trailing edge closure cap includes an integral wedge coupled to the upper skin panel.

12. The structural composite airfoil of 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, wherein the upper leading edge end is coupled to the forward C-channel spar.

13. The structural composite airfoil of claim 12, wherein the upper trailing edge end is coupled to an integral wedge of a trailing edge closure cap.

14. The structural composite airfoil of claim 1, wherein the second angle is an acute angle.

15. The structural composite airfoil of claim 1, wherein the forward C-channel spar is coupled directly to the upper and lower skin panels without a splice or nut plate.

16. The structural composite airfoil of claim 1, wherein the primary structural element further comprises:

an intermediate C-channel spar coupled to the upper and lower skin panels, 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 aft of the forward C-channel spar; and

an aft C-channel spar coupled to the upper and lower skin panels, wherein a third channel of the aft C-channel spar faces the leading edge of the structural composite airfoil, and wherein the aft C-channel spar is positioned aft of the intermediate C-channel spar.

17. An aircraft comprising the structural composite airfoil of claim 1.

18. A trailing edge flap for an aircraft comprising the structural composite airfoil of claim 1.

19. A method of assembling a structural composite airfoil, the method comprising:

coupling an upper skin panel to a forward C-channel spar, wherein the structural composite airfoil extends from a leading edge to a trailing edge, wherein a first channel of the forward C-channel spar faces the leading edge of the structural composite airfoil, wherein the forward C-channel spar includes an upper flange, a lower flange, and an elongated span extending between the upper flange and the lower flange, wherein coupling the upper skin panel to the forward C-channel spar includes coupling the upper skin panel to the upper flange of the forward C-channel spar, and wherein the upper flange forms an acute angle with the elongated span;

coupling a lower skin panel to the forward C-channel spar such that an interior volume is defined between the upper skin panel and the lower skin panel, wherein the upper skin panel, the lower skin panel, and the forward C-channel spar together form at least a portion of a primary structural element of the structural composite airfoil; and

coupling a leading edge skin panel to the forward C-channel spar, wherein the leading edge skin panel defines the leading edge of the structural composite airfoil, wherein coupling the leading edge skin panel comprises coupling a first end region of the leading edge skin panel to the upper flange of the forward C-channel spar, wherein coupling the leading edge skin panel further comprises coupling a second end region of the leading edge skin panel to the lower flange of the forward C-channel spar, wherein the leading edge skin panel has a bull-nose shape.

20. A structural composite airfoil having a leading edge and a trailing edge, comprising:

a primary structural element extending from a leading edge region to a trailing edge region, wherein the leading edge region defines the leading edge of the structural composite airfoil, wherein the primary structural element comprises: an upper skin panel; a lower skin panel; an interior volume defined between the upper skin panel and the lower skin panel; a forward C-channel spar including an upper flange coupled to the upper skin panel, wherein the forward C-channel spar further includes a lower flange coupled to the lower skin panel, wherein a first channel of the forward C-channel spar faces the leading edge of the structural composite airfoil, wherein the upper flange forms a first angle with an elongated span of the forward C-channel spar, wherein the lower flange forms a second angle with the elongated span, and wherein the first angle is an acute angle; and a leading edge skin panel defining the leading edge of the structural composite airfoil and positioned within the leading edge region of the primary structural element, wherein a first end region of the leading edge skin panel is coupled to the upper flange of the forward C-channel spar, wherein a second end region of the leading edge skin panel is coupled to the lower flange of the forward C-channel spar, and wherein the leading edge skin panel has a bull-nose shape; and

an auxiliary structural element defining the trailing edge of the structural composite airfoil.

Technical Field

The present disclosure relates generally to structural composite airfoils and related methods.

Background

Aircraft, including fixed wing aircraft and rotary wing aircraft, employ various aerodynamic control surfaces such as ailerons, airbrakes, elevators, flaps, rudders, slats, spoilers, and the like. By manipulating one or more aerodynamic control surfaces, a pilot can control the lift generated by the aircraft, such as 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 a wing of a fixed wing aircraft typically includes one or more flaps that are movable between a retracted position and an extended position. During cruising, the flaps are normally held in a retracted position. When extended, the flap increases camber of the wing. Thus, during take-off, climb, descent or landing, the flaps may 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 from composite materials and are therefore referred to herein as structural composite airfoils.

Structural composite airfoils (such as flaps) have an aerodynamic cross-sectional profile that is typically formed by coupling an upper skin adjacent to both the leading and trailing edges of the structural composite airfoil to a lower skin. For example, in conventional configurations of inboard and outboard flaps, the primary structural elements of the flap are defined by coupling upper and lower skins to three spars that extend the width of the flap. The leading edge (typically comprising a bull-nose shape) and the trailing edge (tapering to a thin cross-section) of the structural composite airfoil are typically outboard of the primary structural elements, forming respective secondary 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 spar and other structures forming the flap. The large number of fasteners can increase the cost, manufacturing cycle time, and weight of the resulting assembly. Accordingly, those skilled in the art continue to strive for research and development efforts to improve structural composite airfoils and their manufacture.

Disclosure of Invention

The structural composite airfoils and associated methods of forming the same as disclosed herein may reduce the number of fasteners, improve the airfoil aerodynamic surfaces, and/or simplify the manufacturing process of the structural composite airfoils.

Examples of structural composite airfoils according to the present disclosure include a primary structural element, a secondary 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 primary structural elements extend from the leading edge region to the trailing edge region. The leading edge region of the primary structural element is adjacent to the leading edge of the structural composite airfoil.

The primary structural elements include an upper skin panel, a lower skin panel, and a forward C-channel spar. An interior volume is defined between the upper skin panel and the lower skin panel. The forward C-channel spar includes an upper flange coupled to the upper skin panel and a lower flange coupled to the lower skin panel. The 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 of the front C-channel spar.

A leading edge skin panel is positioned adjacent to a leading edge region of the primary structural element, wherein a first end region of the leading edge skin panel is coupled to the upper flange of the forward C-channel spar, wherein a second end region of the leading edge skin panel is coupled to the lower flange of the forward C-channel spar, and wherein the leading edge skin panel has a bull-nose shape.

Methods of assembling such structural composite airfoils are also disclosed. In such a method, an upper skin panel is coupled to an upper flange of the forward C-channel spar, a lower skin panel is coupled to a lower flange of the forward C-channel spar such that an interior volume is defined between the upper and lower skin panels, and the forward edge skin panel is coupled to the forward C-channel spar. For example, a first end region of the leading edge skin panel is coupled to an upper flange of the forward C-channel spar, and a second end region of the leading edge skin panel is coupled to a lower flange of the forward C-channel.

Drawings

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

FIG. 2 is a schematic side view of an example of a structural composite airfoil according to the present disclosure.

Figure 3 is a side view of an integral Z-shaped spar formed in the lower skin panel.

Figure 4 is a side view of an integral Z-shaped spar formed in the upper skin panel.

FIG. 5 is a flow chart representing the disclosed method of forming the disclosed structural composite airfoil.

Detailed Description

Referring to FIG. 1, one or more structural composite airfoils 10 may be included in an apparatus 12. The structural composite airfoil 10 may be used in many different industries and applications, such as the aerospace, automotive, construction, marine, wind power, remote control, military, entertainment, and/or racing industries. In FIG. 1, an example of a device 12 that may include one or more structural composite airfoils 10 is shown generally in the form of an aircraft 14. The aircraft 14 may take any suitable form, including a commercial aircraft, a military aircraft, or any other suitable aircraft. Although FIG. 1 illustrates the aircraft 14 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) rotorcraft and helicopters.

The apparatus 12 (e.g., aircraft 14) may include one or more structural composite airfoils 10. As an illustrative, non-exclusive example, the structural composite airfoil 10 may be used in a wing 16 (e.g., a flap 17, which may be an inboard or outboard flap), but other components of the aircraft 14, such as horizontal stabilizers 18, vertical stabilizers 20, and other components, may additionally or alternatively include one or more structural composite airfoils 10. Other applications for the structural composite airfoil 10 in the aircraft 14 (or other equipment 12) may include other wing control surfaces, ailerons, flaperons, airbrakes, elevators, slats, spoilers, rudders, canards, and/or winglets. In other industries, examples of equipment 12 including one or more structural composite airfoils 10 may include or be part of: space satellites, vehicles, shipping containers, fast vehicles, vehicle bodies, propeller blades, turbine blades, and/or marine vehicles (e.g., sailboats), among others.

FIG. 2 provides an illustrative, non-exclusive example of a structural composite airfoil 10 according to the present disclosure. In general, 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 necessary 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.

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 is subject to flight, ground, or pressurization loads, and failure thereof will reduce the structural integrity of the device or component of which the structural composite airfoil 10 is a part. As used herein, an "auxiliary structural element" is an element or structure whose failure does not affect the safety of the equipment or component of which the structural composite airfoil 10 is a part.

The primary structural elements 26 extend from a leading edge region 30 to a trailing edge region 32. As shown in FIG. 2, 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 may be referred to as the region of the primary structural element 26 closest to the leading edge 22. Similarly, the trailing edge region 32 may be referred to as the region of the primary structural element 26 closest to the trailing edge 24, but the trailing edge region 32 of the primary structural element 26 does not define the trailing edge 24 of the structural composite airfoil 10. As used herein, a first element or structure is said to be "behind" another element or structure if the first element or structure is located closer to the trailing edge 24 than the other element or structure. Similarly, as used herein, a first element or structure is said to be "forward" of another element or structure if the first element or structure is positioned closer to the leading edge 22 than the other element or structure.

The primary structural elements 26 include at least an upper skin panel 34, a lower skin panel 36, and a forward C-channel spar 38. An interior volume 40 is defined between the upper skin panel 34 and the lower skin panel 36. The forward C-channel spar 38 includes an upper flange 42 and a lower flange 44, with the upper flange 42 coupled to the upper skin panel 34 and the lower flange 44 coupled to the lower skin panel 36. The first channel 46 of the forward C-shaped channel spar 38 faces the leading edge 22 of the structural composite airfoil 10. This arrangement of the forward C-channel spar 38 relative to the leading edge 22 may allow for efficient coupling of the leading edge skin panel 54 to the upper and lower skin panels 34, 36 (via the forward C-channel spar 38) without creating any slosh in the upper and lower skin panels 34, 36, potentially reducing the complexity of manufacturing the upper and lower skin panels 34, 36.

The upper flange 42 forms a first angle 48 with an elongated span 50 of the forward C-shaped channel spar, and the lower flange 44 forms a second angle 52 with the elongated span 50, the first channel 46 being defined by the upper flange 42, the lower flange 44, and the elongated span 50. In some examples of the structural composite airfoil 10, the first angle 48 and/or the second angle 52 may be acute. Typical conventional airfoil configurations will involve such angles of greater than 90 degrees to facilitate removal of the part from the tool, and/or the channel of the front airfoil will be disposed facing the trailing edge of the airfoil. The presently disclosed examples of the structural composite airfoil 10 may advantageously provide for joining between components or elements (e.g., joining the leading edge skin panel 54 and the upper flange 42) without creating sloshing in the upper skin panel 34 or the lower skin panel 36 or utilizing a splice strap, and/or may 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 may be angled relative to the elongate span 50 so as to be complementary to the first end region 56 of the leading edge skin panel 54. Similarly, the lower flange 44 may be angled relative to the elongate span 50 so as to be complementary to the second end region 58 of the leading edge skin panel 54.

The leading edge 22 of the structural composite airfoil 10 is defined by a leading edge skin panel 54 that is generally shaped to have a bull nose (bullnose) shape. The leading edge skin panel 54 may be located adjacent the leading edge region 30 of the primary structural element 26, but the leading edge skin panel 54 may be a separate part that is external to or distinct from the primary structural element 26. In other examples, the leading edge skin panel 54 may be within the leading edge region 30 of the primary structural element 26 and/or define the leading edge region 30 of the primary structural element 26, such as in examples where the primary structural element 26 extends to the leading edge 22. The leading edge skin panel 54 is coupled to the upper and lower skin panels 34, 36 via the leading C-channel spar 38. Specifically, a first end region 56 of leading edge skin panel 54 is coupled to upper flange 42 of forward C-channel spar 38, and a second end region 58 of leading edge skin panel 54 is coupled to lower flange 44 of forward C-channel spar 38. Because the upper flange 42 of the forward C-channel spar 38 is coupled to the leading edge skin panel 54 and the upper skin panel 34, the forward 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 may 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 may be coupled to the upper flange 42 without contacting the leading edge skin panel 54. Similarly, because the lower flange 44 of the forward C-channel spar 38 is coupled to the leading edge skin panel 54 and the lower skin panel 36, the forward 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 may abut the leading edge skin panel 54 (e.g., abut the second end region 58). In other examples, the lower skin panel 36 may be coupled to the lower flange 44 without contacting the leading edge skin panel 54.

The trailing edge 24 of the structural composite airfoil 10 is defined by a secondary structural element 28. In various examples of the structural composite airfoil 10, the secondary structural elements 28 may include wedge closures, duckbill closures, bond closures, and/or rivet closures. Examples of suitable trailing edge closures are also disclosed in U.S. patent No.10,532,804 entitled "Aerodynamic control surface and associated trailing edge closure method" (issued on 14.1.2020), the entire disclosure of which is incorporated herein by reference for all purposes.

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 and lower leading edge ends 76, 78 may be coupled to the forward C-channel spar 38. In some examples, upper trailing edge end 92 may be coupled to lower trailing edge end 94. Additionally or alternatively, the upper trailing edge end 92 and/or the lower trailing edge end 94 may form or define the trailing edge 24 of the structural composite airfoil 10.

The structural composite airfoil 10 may include one or more fasteners that secure the various components to one another. For example, the first fastener 80 may couple the leading edge skin panel 54 (e.g., the first end region 56 of the leading edge skin panel 54) to the upper flange 42 of the forward C-channel spar 38. In some examples, the first fasteners 80 are a plurality of first fasteners 80 spaced along the width of the structural composite airfoil 10 (the width of the airfoil extending in/out of the page) to secure the leading edge skin panel 54 to the forward C-channel spar 38 along the first end region 56. The leading edge skin panel 54 may 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 sloshing being formed in either panel. Additionally or alternatively, because the upper flange 42 may be configured to effectively splice the leading edge skin panel 54 and the upper skin panel 34, the structural composite airfoil 10 may be formed without a separate splice strap connecting the leading edge skin panel 54 and the upper skin panel 34.

Similarly, second fasteners 82 may couple leading edge skin panel 54 (e.g., second end region 58 of leading edge skin panel 54) to lower flange 44 of leading C-channel spar 38. In some examples, the second fastener 82 is a plurality of second fasteners 82 spaced along the width of the structural composite airfoil 10 (the width of the airfoil extending in/out of the page) to secure the leading edge skin panel 54 to the leading C-channel spar 38 along the second end region 58. The leading edge skin panel 54 may be configured to bond 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 sloshing being formed in either panel. Additionally or alternatively, because the forward flange 44 may be configured to effectively splice the leading edge skin panel 54 and the lower skin panel 36, the structural composite airfoil 10 may be formed without a separate splice strap connecting the leading edge skin panel 54 and the lower skin panel 36. The first and second fasteners 80, 82 may be configured such that the leading edge skin panel 54 may be selectively removed from the primary structural element 26 by removing the first and second fasteners 80, 82.

A third fastener 84 (or a plurality of third fasteners 84 spaced along the width of the structural composite airfoil 10) may be positioned to couple the upper skin panel 34 to the upper flange 42 of the forward 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. The fourth fastener 86 (or a plurality of fourth fasteners 86 spaced along the width of the structural composite airfoil 10) may be positioned to couple the lower skin panel 36 to the lower flange 44 of the forward C-channel spar 38. The fourth fastener 86 generally couples the lower leading edge end 78 of the lower skin panel 36 to the lower flange 44. Third fastener 84 and/or fourth fastener 86 may be accessible (e.g., not blind) even after assembly of primary structural element 26. Additionally or alternatively, the third fastener 84 and/or the fourth fastener 86 may be fixed permanent fasteners (e.g., hex drive bolts) without a nut plate.

The structural composite airfoil 10 may further include an intermediate C-channel spar 60 and/or an aft C-channel spar 62, either or both of which may form a portion of the primary structural element 26. In the example shown in fig. 2, the primary structural elements 26 are defined by respective portions of the forward C-channel spar 38, the intermediate C-channel spar 60, the aft C-channel spar 62, and the upper and lower skin panels 34, 36 extending between the forward C-channel spar 38 and the aft C-channel spar 62. In other examples of the structural composite airfoil 10, the primary structural element 26 may extend more toward the leading edge 22 than shown in FIG. 2. For example, as described above, while the primary structural elements 26 may extend only between the forward and aft C-channel spars 38, 62, in other examples, the primary structural elements 26 may optionally extend further forward such that the primary structural elements 26 may also extend to and include the leading edge 22. Additionally or alternatively, the primary member element 26 may extend more toward the trailing edge 24 than shown in FIG. 2. For example, the primary structural element 26 may comprise at least a portion of the structural composite airfoil 10 aft of the aft C-channel spar 62.

In examples 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 and lower skin panels 34, 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. An intermediate C-channel spar 60 is positioned aft of the forward C-channel spar 38.

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

The upper skin panel 34 may be coupled to the intermediate C-channel spar 60 (e.g., the intermediate upper flange 66) and/or the aft C-channel spar 62 (e.g., the aft upper flange 72) with a plurality of other fasteners 88. 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 aft C-channel spar 62 (e.g., the aft lower flange 74). Additionally or alternatively, one or more fasteners 88 may be used to couple upper trailing edge end 92 to lower trailing edge end 94.

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

In some examples, at least a portion of the upper skin panel 34 may be core-stiffened. As used herein, "core-stiffened" refers to a skin panel having at least a first skin and a low density core material coupled to the skin. The core reinforcement material optionally includes a second skin, the core material being sandwiched between the first skin and the second skin to form a sandwich panel. Suitable materials for forming the core reinforcement portion are well known in the art and include honeycomb core materials and metal core materials, although other core materials are also within the scope of the present disclosure. As an illustrative example, the upper skin panel 34 may include a first upper core stiffened portion 134, a second upper core stiffened portion 136, and a third upper core stiffened portion 138. A first upper core reinforcement portion 134 may be positioned between the forward C-channel spar 38 and the intermediate C-channel spar 60, a second upper core reinforcement portion 136 may be positioned between the intermediate C-channel spar 60 and the aft C-channel spar 62, and/or a third upper core reinforcement portion 138 may be positioned between the aft C-channel spar 62 and the upper trailing edge end 92. One or more of the upper core reinforcement portions 134, 136, 138 may be tapered, such as in a region proximate to a respective section of the C-channel spar 38, 60, and/or 62. For example, the upper core reinforcement portions 134, 136, and/or 138 may have a height or thickness extending downward from the upper skin panel 34 toward the lower skin panel 36, wherein the height or thickness decreases proximate to one or more of the C-channel spars 38, 60, and/or 62, thereby forming a taper. In the example of fig. 2, the thickness of the first upper core reinforcement portion 134 tapers near the front C-channel spar 38 and the intermediate C-channel spar 60, the thickness of the second upper core reinforcement 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 reinforcement 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 upper core reinforcement portions 134, 136, and/or 138 may be substantially constant rather than tapering where the respective upper core reinforcement portion 134, 136, and/or 138 meets the respective C-channel spar 38, 60, 62. In some examples, one or more of the upper core reinforcement portions 134, 136, and/or 138 may abut the respective C-channel spar 38, 60, and/or 62. Although the upper skin panel 34 shown in fig. 2 includes three different upper core stiffened portions 134, 136, 138, in other examples, the upper skin panel 34 may be core stiffened along its entire length, along a greater or lesser portion of its length, and/or may include more or fewer separate upper core stiffened sections than shown in fig. 2.

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. A first lower core reinforcement portion 140 may be positioned between the front C-channel spar 38 and the intermediate C-channel spar 60, a 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 a 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 be tapered, such as in a region of the respective section proximate the C-channel spars 38, 60, and/or 62. For example, the lower core reinforcement portions 140, 142, and/or 144 may have a height or thickness extending upwardly from the lower skin panel 36 toward the upper skin panel 34, wherein the height or thickness decreases proximate to one or more of the C-channel spars 38, 60, and/or 62, thereby forming a taper. In the example of fig. 2, 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 lower core reinforcement portions 140, 142, and/or 144 may be substantially constant rather than tapering where the respective lower core reinforcement portion 140, 142, and/or 144 meets the respective C-channel spar 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 spar 38, 60, and/or 62. Although the lower skin panel 36 shown in fig. 2 includes three different lower core stiffened portions 140, 142, 144, in other examples, the lower skin panel 36 may be core stiffened along its entire length, may be core stiffened along a greater or lesser portion of its length, and/or may include more or less separate lower core stiffened portions than shown in fig. 2.

The structural composite airfoil 10 has a length 90, which may also be referred to herein as a chord length 90, and the location along the length 90 may be defined in terms of a percentage of the distance from the leading edge 22 along the length 90. In these aspects, the forward 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 forward C-channel spar 38 is positioned approximately 5% of the length 90 away from the leading edge 22. In some examples, the forward C-channel spar 38 may be positioned as 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 about 30% of the length 90 away from the leading edge 22. In some examples, the intermediate C-channel spar 60 may be positioned for balancing torsional capability within the primary structural elements 26 on either side of the intermediate C-channel spar 60. Additionally or alternatively, the aft 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 away from the leading edge 22. In particular examples, the aft C-channel spar 62 may be positioned approximately 55% of the length 90 away from the leading edge 22. In some examples, the aft C-channel spar 62 may be positioned as aft as possible for integration.

Some examples of the structural composite airfoil 10 may include an integral Z-shaped spar 100, which may be part of the primary structural element 26, and in some examples, the element aft of the integral Z-shaped spar 100 is part of the secondary structural element 28. Accordingly, positioning the integral Z-shaped spar 100 behind (or in place of one or both of) the intermediate C-channel spar 60 and/or the aft C-channel spar 62 may extend or extend 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-shaped spar 100 may be formed within the trailing edge region 32 of the primary structural element 26. Fig. 3-4 illustrate an example of such an integral Z-shaped spar 100, fig. 3 illustrates an example of an integral Z-shaped spar 100 formed in the lower skin panel 36, and fig. 4 illustrates an example of an integral Z-shaped spar 100 formed in the upper skin panel 34. The integral Z-shaped spar 100 is generally positioned adjacent the trailing edge 24 of the structural composite airfoil 10, such as by being positioned at least 80% away from the length 90 of the leading edge 22. In some examples, the integral Z-shaped spar 100 may be positioned between 80-95% of the length 90 away from the leading edge 22.

Referring to FIG. 3, an 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 with the lower skin panel 36 that is greater than 90 degrees and/or greater than 100 degrees. Additionally or alternatively, the first Z-shaped spar segment 110 may form an angle with the upper skin panel 34 of greater than 90 degrees and/or greater than 100 degrees. The integral Z-shaped spar 100 may further include a second Z-shaped spar section 112 extending aft of the second bend 108. As shown in fig. 3, the second Z-shaped spar section 112 may be coupled to the upper skin panel 34. In the example shown in fig. 3, the second Z-shaped spar segment 112 is located adjacent to an interior 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, Z-shaped spar fasteners 116 are recessed into upper skin panel 34 (e.g., such that Z-shaped spar fasteners 116 are at least substantially flush or sub flush with upper panel surface 130 of upper skin panel 34) and extend through upper skin panel 34 and second Z-shaped spar segments 112 to couple integral Z-shaped spar 100 to upper skin panel 34.

The integral Z-spar 100 may include a Z-spar joint (joggle)102 in the lower skin panel 36, which may be configured to receive a portion of a trailing edge closure cap 104, which may at least partially define the trailing edge 24 and/or the auxiliary structural elements 28 of the structural composite airfoil 10. The Z-shaped spar joint 102 is in fact a small offset in the lower skin panel 36 upwards towards the upper skin panel 34 and is positioned generally forward of the first bend 106. As shown in fig. 3, the first cover end region 118 of the trailing edge closure cover 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 create a smooth surface at the interface and improve aerodynamic performance, as shown in fig. 3, the first cap end region 118 may be slightly recessed into the lower skin panel 36, such as via the Z-spar joint 102. Depending on the thickness of the first cap end region 118, the Z-shaped spar joint 102 may be tailored 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 cap surface 128 of the trailing edge closure cap 104 within the first cap end region 118. In other words, Z-shaped spar joint 102 may be larger to create a larger recess to receive and engage a given trailing edge closure cap 104 having a thicker first cap end region 118, while Z-shaped spar joint 102 may be smaller to create a smaller recess to receive and engage a different given trailing edge closure cap 104 having a thinner first cap end region 118. Any gaps remaining at the interface of the Z-spar joint 102 and the first cap 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.

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., bonded and/or coupled via one or more fasteners) to the upper skin panel 34, as shown in fig. 3. Alternatively, the integral wedges 122 may be integrally formed with the upper skin panel 34. In still other examples, integral wedge 122 may be a separate component from trailing edge closure cap 104 and from upper skin panel 34, and may be bonded or otherwise coupled to upper skin panel 34 and/or trailing edge closure cap 104. By way of example, the integral wedges 122 may be formed by stacking layers of material, molding, and/or by machining mating surface profiles to mate with the upper skin panel 34.

Referring to FIG. 4, an integral Z-shaped spar 100 may be formed in the upper trailing edge end 92 of the upper skin panel 34. In the example shown in fig. 4, the second Z-shaped spar segment 112 is coupled to the lower skin panel 36 and is positioned adjacent to an inner surface 124 of the lower skin panel 36. Z-shaped spar fasteners 116 couple the integral Z-shaped spar 100 to the lower skin panel 36, with the Z-shaped spar fasteners 116 being recessed into the lower skin panel 36 (e.g., such that the Z-shaped spar fasteners 116 are at least substantially flush or sub-flush with a lower panel surface 126 of the lower skin panel 36) and extending through the lower skin panel 36 and the second Z-shaped spar segment 112 to couple the integral Z-shaped spar 100 to the lower skin panel 36.

In fig. 4, integral Z-spar 100 includes a Z-spar joint 102 in upper skin panel 34 that is configured to receive a portion of trailing edge closure cap 104, where Z-spar joint 102 is positioned forward of first bend 106. The Z-shaped spar joint 102 is actually a small offset in the upper skin panel 34 toward the lower skin panel 36. In this example, the first cover end region 118 of the trailing edge closure cover 104 is bonded to the upper skin panel 34 rather than the lower skin panel 36. Additionally or alternatively, the first cover end region 118 may be riveted or otherwise fastened or coupled to the upper skin panel 34. To create a smooth surface at the interface and improve aerodynamic performance, as shown in fig. 4, the first cap end region 118 may be slightly recessed into the upper skin panel 34, such as via a Z-shaped spar joint 102. Depending on the thickness of the first cap end region 118, the Z-shaped spar joint 102 may be tailored 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 cap surface 132 of the trailing edge closure cap 104 within the first cap end region 118. In other words, Z-spar joint 102 may be larger to create a larger recess to receive and engage a given trailing edge closure cap 104 having a thicker first cap end region 118, while Z-spar joint 102 may be smaller to create a smaller recess to receive and engage a different given trailing edge closure cap 104 having a thinner first cap end region 118.

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., bonded and/or coupled via one or more fasteners) to the lower skin panel 36. Alternatively, and as shown in fig. 4, the integral wedges 122 may be integrally formed with the lower skin panel 36. In still other examples, the integral wedge 122 may be a separate component from the trailing edge closure cap 104 and from the lower skin panel 36, and may be bonded or otherwise coupled to the lower skin panel 36 and/or the trailing edge closure cap 104. For example, the integral wedges 122 may be formed by stacking layers of material, molding, and/or by machining mating surface profiles to mate with the lower skin panel 36.

Fig. 5 schematically provides a flow chart representing an illustrative, non-exclusive example of a method 200 according to the present disclosure. In fig. 5, some steps are shown in dashed boxes, indicating that such steps may be optional or may correspond to an alternative 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 box. It will be understood from the discussion herein that the method 200 and steps illustrated in fig. 5 are not limiting, and that other methods and steps are within the scope of the present disclosure, including methods having a greater or lesser number of steps than illustrated.

The method 200 generally includes coupling an upper skin panel (e.g., upper skin panel 34) to a forward C-channel spar (e.g., forward C-channel spar 38) at 202, and coupling a lower skin panel (e.g., lower skin panel 36) to the forward C-channel spar at 204. Coupling the upper skin panel to the forward C-channel spar at 202 generally includes coupling the upper skin panel to an upper flange (e.g., upper flange 42) of the forward C-channel spar. Similarly, coupling the lower skin panel to the forward C-channel spar at 204 generally includes coupling the lower skin panel to a lower flange (e.g., lower flange 44) of the forward C-channel spar. Coupling the upper skin panel at 202 and/or the lower skin panel at 204 may be performed with a reduced number of nut plates or other fastening components as compared to conventional techniques. Additionally or alternatively, coupling the upper skin panel at 202 and/or coupling the lower skin panel at 204 may be performed without the use of a splice strap. Reducing the number of fasteners or fastening components may reduce the weight of the resulting structural composite airfoil, reduce manufacturing costs, and/or reduce manufacturing tooling time.

Method 200 also includes coupling a leading edge skin panel (e.g., leading edge skin panel 54) to the leading 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 an upper flange of the leading C-channel spar and coupling a second end region (e.g., second end region 58) of the leading edge skin panel to a lower flange of the leading C-channel spar. The coupling of the leading edge skin panels at 206 may be performed without overlapping the upper skin panels and the leading edge skin panels on the upper flanges of the forward C-channel spars. Similarly, the coupling of the leading edge skin panels at 206 may be performed without overlapping the lower skin panels and the leading edge skin panels on the lower flanges of the forward C-channel spars. Coupling the leading edge skin panel at 206 may include coupling the leading edge skin panel without the use of a splice strap such that the leading edge skin panel may be coupled directly to the forward C-channel spar. In some methods 200, coupling the leading edge skin panel at 206 includes abutting a first end region of the leading edge skin panel with an upper skin panel (e.g., the upper leading edge end 76 of the upper skin panel 34), which may include forming a lap joint or splice therebetween. Additionally or alternatively, coupling the leading edge skin panel at 206 may include abutting and/or forming a lap joint or splice between the second end region of the leading edge skin panel and the lower skin panel (e.g., the lower leading edge end 78 of the lower skin panel 36).

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 an upper skin panel to an aft C-channel spar (e.g., aft C-channel spar 62) at 210, coupling a lower skin panel to an intermediate C-channel spar at 212, and/or coupling a lower skin panel to an aft C-channel spar at 214. Additionally or alternatively, the method 210 may include coupling an auxiliary structural element (e.g., auxiliary structural element 28), such as a closure, 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, the method 200 may include forming an integral Z-shaped spar (e.g., the integral Z-shaped spar 100) in the lower skin panel or the upper skin panel at 218.

Illustrative, non-exclusive examples of the inventive subject matter according to this disclosure are described in the paragraphs listed below:

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

a primary 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 a leading edge (22) of the structural composite airfoil (10) or defines the leading edge (22) of the structural composite airfoil (10), wherein the primary structural element (26) comprises:

an upper skin panel (34);

a lower skin panel (36);

an interior volume (40) defined between the upper skin panel (34) and the lower skin panel (36); and

a forward C-channel spar (38) comprising an upper flange (42) coupled to the upper skin panel (34), wherein the forward C-channel spar (38) further comprises a lower flange (44) coupled to the lower skin panel (36), wherein a first channel (46) of the forward 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 elongated span (50) of the forward C-channel spar (38), wherein the lower flange (44) forms a second angle (52) with the elongated span (50), and wherein the first angle (48) is an acute angle;

an auxiliary structural element (28) defining a trailing edge (24) of the structural composite airfoil (10); and

a leading edge skin panel (54) defining a leading edge (22) of the structural composite airfoil (10) and positioned adjacent to or within a 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 forward 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 forward C-channel spar (38), and wherein the leading edge skin panel (54) has a bull-nose shape.

A1.1. The structural composite airfoil (10) according to paragraph a1, wherein the primary structural element further includes an intermediate C-channel spar (60) coupled to the upper and lower skin panels (34, 36), wherein the second channel (64) of the intermediate C-channel spar (60) faces the leading edge (22) of the structural composite airfoil (10), wherein the intermediate C-channel spar (60) is positioned aft of the forward C-channel spar (38).

A1.2. The structural composite airfoil (10) according to paragraph A1 and/or a1.1, wherein the primary structural element further includes an aft C-channel spar (62) coupled to the upper skin panel (34) and the lower skin panel (36), wherein the third channel (70) of the aft C-channel spar (62) faces the leading edge (22) of the structural composite airfoil (10), and wherein the aft C-channel spar (62) is positioned aft of the intermediate C-channel spar (60).

A2. The structural composite airfoil (10) according to any of paragraphs A1-a1.2, wherein the second angle (52) is an acute angle.

A3. The structural composite airfoil (10) according to any of paragraphs a1-a2, wherein the upper flange (42) is angled relative to the elongate span (50) to complement the first end region (56) of the leading edge skin panel (54).

A4. The structural composite airfoil (10) according to any of paragraphs a1-A3, wherein the leading edge skin panel (54) does not overlap the upper skin panel (34) on the upper flange (42) of the forward C-channel spar (38).

A5. The structural composite airfoil (10) according to any of paragraphs a1-a4, wherein the lower flange (44) is angled to complement the second end region (58) of the leading edge skin panel (54).

A6. The structural composite airfoil (10) according to any of paragraphs a1-a5, wherein the leading edge skin panel (54) does not overlap the lower skin panel (36) on the lower flange (44) of the forward C-channel spar (38).

A7. The structural composite airfoil (10) according to any of paragraphs a1-a6, wherein the upper skin panel (34) abuts the leading edge skin panel (54).

A8. The structural composite airfoil (10) according to any of paragraphs a1-a7, wherein the lower skin panel (36) abuts the leading edge skin panel (54).

A9. The structural composite airfoil (10) according to any of paragraphs a1-A8, further comprising a first fastener (80) coupling the leading edge skin panel (54) to the upper flange (42) of the leading C-channel spar (38).

A10. The structural composite airfoil (10) according to any of paragraphs a1-a9, further comprising a second fastener (82) coupling the leading edge skin panel (54) to the lower flange (44) of the leading C-channel spar (38).

A11. The structural composite airfoil (10) according to any of paragraphs a1-a10, further comprising a third fastener (84) coupling the upper skin panel (34) to the upper flange (42) of the forward C-channel spar (38).

A11.1. The structural composite airfoil (10) according to paragraph a11, wherein the third fastener (84) is not blind (blind) such that the third fastener is accessible when the primary structural element (26) is assembled.

A12. The structural composite airfoil (10) according to any of paragraphs A1-a11.1, further comprising a fourth fastener (86) coupling the lower skin panel (36) to the lower flange (44) of the forward C-channel spar (38).

A12.1. The structural composite airfoil (10) according to paragraph a12, wherein the fourth fastener (86) is not blind, such that the fourth fastener is accessible when the primary structural element (26) is assembled.

A13. The structural composite airfoil (10) according to any of paragraphs A1-a12.1, wherein the leading edge skin panel (54) is joined with the upper skin panel (34) without any sloshing.

A14. The structural composite airfoil (10) according to any of paragraphs a1-a13, wherein the leading edge skin panel (54) engages the lower skin panel (36) without any sloshing.

A15. The structural composite airfoil (10) according to any of paragraphs a1-a14, wherein the leading edge skin panel (54) is joined with the upper skin panel (34) without a splice strap.

A16. The structural composite airfoil (10) according to any of paragraphs a1-a15, wherein the leading edge skin panel (54) is joined with the lower skin panel (36) without a splice strap.

A17. The structural composite airfoil (10) according to any of paragraphs a1-a16, further comprising a plurality of fasteners (84, 86) coupling the forward C-channel spar (38) to the upper and lower skin panels (34, 36), wherein each fastener (84, 86) of the plurality of fasteners (84, 86) is not blind, such that each fastener (84, 86) of the plurality of fasteners (84, 86) is accessible when the forward C-channel spar (38) is secured to the upper and lower skin panels (34, 36).

A18. The structural composite airfoil (10) according to any of paragraphs a1-a17, wherein the upper skin panel (34) is coupled to the upper flange (42) of the forward C-channel spar (38) without a nut plate.

A19. The structural composite airfoil (10) according to any of paragraphs a1-a18, wherein the lower skin panel (36) is coupled to the lower flange (44) of the forward C-channel spar (38) without a nut plate.

A20. The structural composite airfoil (10) according to any of paragraphs a1-a19, wherein at least a portion of the upper skin panel (34) is core reinforced.

A21. The structural composite airfoil (10) according to any of paragraphs a1-a20, wherein at least a portion of the lower skin panel (36) is core reinforced.

A22. The structural composite airfoil (10) according to any of paragraphs a1-a21, wherein the upper skin panel (34) includes glass or carbon fibers.

A23. The structural composite airfoil (10) according to any of paragraphs a1-a22, wherein the lower skin panel (36) includes glass or carbon fibers.

A24. The structural composite airfoil (10) according to any of paragraphs a1-a23, wherein the structural composite airfoil (10) has a length (90), and wherein the position along the length (90) may be defined by a percentage of the distance along the length (90) from the leading edge (22).

A25. The structural composite airfoil (10) according to paragraph a24, wherein the forward C-channel spar (38) is positioned between 0% and 10% of a length (90) away from the leading edge (22).

A26. The structural composite airfoil (10) according to paragraph a25, wherein the forward C-channel spar (38) is positioned about 5% of the length (90) away from the leading edge (22).

A27. The structural composite airfoil (10) according to any of paragraphs a24-a26, wherein the intermediate C-channel spar (60) is positioned between 20% and 40% of a length (90) away from the leading edge (22).

A28. The structural composite airfoil (10) according to paragraph a27, wherein the intermediate C-channel spar (60) is positioned about 30% of the length (90) away from the leading edge (22).

A29. The structural composite airfoil (10) according to any of paragraphs a24-a28, wherein the aft C-channel spar (62) is positioned between 40% and 70% of a length (90) away from the leading edge (22), and/or between 50% and 60% of the length (90) away from the leading edge (22).

A30. The structural composite airfoil (10) according to paragraph a29, wherein the aft C-channel spar (62) is positioned about 55% of the length (90) away from the leading edge (22).

A31. The structural composite airfoil (10) according to any of paragraphs a1-a30, wherein the structural composite airfoil (10) is a trailing edge flap (17), a flap, a flaperon, an airbrake, an elevator, a slat, a spoiler, a canard, a rudder, and/or a winglet (winglet).

A32. The structural composite airfoil (10) according to any of paragraphs a1-a31, wherein the secondary structural element (28) comprises a wedge seal.

A33. The structural composite airfoil (10) according to any of paragraphs a1-a32, wherein the secondary structural element (28) comprises a duckbill closure.

A34. The structural composite airfoil (10) according to any of paragraphs a1-a33, wherein the auxiliary structural element (28) includes a bonded closure.

A35. The structural composite airfoil (10) according to any of paragraphs a1-a34, wherein the auxiliary structural element (28) comprises a riveted closure.

A36. The structural composite airfoil (10) according to any of paragraphs a1-a35, wherein the lower skin panel (36) includes a lower leading edge end (78) and a lower trailing edge end (94), wherein the lower trailing edge end (94) is opposite the lower leading edge end (78).

A37. The structural composite airfoil (10) according to paragraph a36, wherein the lower leading edge end (78) is coupled to the forward C-channel spar (38).

A38. The structural composite airfoil (10) according to any of paragraphs a36-a37, wherein the lower trailing edge end (94) is coupled to the upper trailing edge end (92) of the upper skin panel (34).

A39. The structural composite airfoil (10) according to any of paragraphs a36-a38, wherein the lower trailing edge end (94) forms an integral Z-shaped spar (100).

A40. The structural composite airfoil (10) according to any of paragraphs a1-a39, wherein the primary structural element (26) includes an integral Z-shaped spar (100).

A41. The structural composite airfoil (10) according to paragraph a40, wherein the integral Z-shaped spar (100) is formed from the lower skin panel (36) within the trailing edge region (32) of the primary structural element (26).

A42. The structural composite airfoil (10) according to any of paragraphs a40-a41, wherein the integral Z-shaped spar (100) includes a joint configured to receive a portion of a trailing edge closure cap (104).

A43. The structural composite airfoil (10) according to any of paragraphs a40-a42, wherein the integral Z-shaped spar (100) includes a first bend (106), a second bend (108), and a first Z-shaped spar section (110) extending between the first bend (106) and the second bend (108).

A44. The structural composite airfoil (10) according to paragraph a43, wherein the first Z-shaped spar segment (110) is substantially perpendicular to the lower skin panel (36) and/or substantially perpendicular to the upper skin panel (34).

A45. The structural composite airfoil (10) according to paragraph a43 or a44, wherein the integral Z-shaped spar (100) further includes a second Z-shaped spar segment (112) extending aft of the second bend (108), wherein the second Z-shaped spar segment (112) is coupled to the upper skin panel (34).

A46. The structural composite airfoil (10) according to paragraph a45, wherein the second Z-shaped spar segment (112) is adjacent to the inner surface (114) of the upper skin panel (34).

A47. The structural composite airfoil (10) according to paragraph a45 or a46, wherein the second Z-shaped spar segment (112) is coupled to the upper skin panel (34) via Z-shaped spar fasteners (116), wherein the Z-shaped spar fasteners (116) are recessed into the upper skin panel (34), and wherein the Z-shaped spar fasteners (116) extend through the second Z-shaped spar segment (112).

A48. The structural composite airfoil (10) according to any of paragraphs a43-a47, wherein a joint of the integral Z-shaped spar (100) is forward of the first bend (106).

A49. The structural composite airfoil (10) according to any of paragraphs a1-a48, further comprising a trailing edge closure cap (104).

A50. The structural composite airfoil (10) according to paragraph a50, wherein the first cover end region (118) of the trailing edge closure cover (104) is bonded to the lower skin panel (36).

A51. The structural composite airfoil (10) according to paragraph a49 or a50, wherein the first cover end region (118) of the trailing edge closure cover (104) is recessed into the lower skin panel (36) such that aerodynamic performance is improved.

A52. The structural composite airfoil (10) according to any of paragraphs a49-a51, wherein the second cover end region (120) of the trailing edge closure cover (104) includes an integral wedge (122) coupled to the upper skin panel (34).

A53. The structural composite airfoil (10) according to any 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 the upper leading edge end (76).

A54. The structural composite airfoil (10) according to paragraph a53, wherein the upper leading edge end (76) is coupled to the forward C-channel spar (38).

A55. The structural composite airfoil (10) according to any of paragraphs a53-a54, wherein the upper trailing edge end (92) is coupled to the lower trailing edge end (94) of the lower skin panel (36).

B1. An aircraft (14) comprising the structural composite airfoil (10) according to any of paragraphs a1-a 55.

B2. A trailing edge flap (17) for an aircraft (14), comprising a structural composite airfoil (10) according to any of paragraphs A1-A55.

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

coupling (202) an upper skin panel (34) to a forward 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 forward C-channel spar (38) faces the leading edge (22) of the structural composite airfoil (10), wherein the forward C-channel spar (38) includes an upper flange (42), a lower flange (44), and an elongated span (50) extending between the upper flange (42) and the lower flange (44), wherein the step of coupling (202) the upper skin panel (34) to the forward C-channel spar (38) includes coupling the upper skin panel (34) to the upper flange (42) of the forward C-channel spar (38), and wherein the upper flange (42) forms an acute angle with the elongated span (50);

coupling (204) the lower skin panel (36) to the forward C-channel spar (38) such that an interior 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 forward C-channel spar (38) together form at least a portion of a primary structural element (26) of the structural composite airfoil (10); and

coupling (206) a leading edge skin panel (54) to the leading C-channel spar (38), wherein the leading edge skin panel (54) defines a 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 an upper flange (42) of the leading 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 a lower flange (44) of the leading C-channel spar (38), wherein the leading edge skin panel (54) has a bull-nose shape.

C1.1. The method (200) according to paragraph C1, further comprising coupling (208) the upper skin panel (34) to an intermediate C-channel spar (60), wherein the intermediate C-channel spar (60) is aft of the forward C-channel spar (38), and wherein the second channel (64) of the intermediate C-channel spar (60) faces the leading edge (22) of the structural composite airfoil (10).

C1.2. The method (200) according to paragraph C1 or C1.1, further comprising coupling (210) the upper skin panel (34) to the aft C-channel spar (62), and wherein the third channel (70) of the aft C-channel spar (62) faces the leading edge (22) of the structural composite airfoil (10).

C1.3. The method (200) according to paragraph C1.2, wherein the rear C-channel spar (62) is rearward of the intermediate C-channel spar (60).

C1.4. The method (200) of any of paragraphs C1-C1.3, further comprising coupling (214) the lower skin panel to the aft C-channel spar (62) such that the aft C-channel spar (62) is part of the primary structural element (26).

C1.5. The method (200) of any of paragraphs C1-C1.4, further comprising coupling (212) the lower skin panel (36) to the intermediate C-channel spar (60) such that the intermediate C-channel spar (60) is part of the primary structural element (26).

C2. The method (200) of any of paragraphs C1-C1.5, wherein the structural composite airfoil (10) is the structural composite airfoil (10) of any of paragraphs A1-A55.

C3. The method (200) according to any of paragraphs C1-C2, wherein the step of coupling (206) the leading edge skin panel (54) includes 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 leading C-channel spar (38).

C4. The method (200) according to any of paragraphs C1-C3, wherein the step of coupling (206) the leading edge skin panel (54) includes 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 leading C-channel spar (38).

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

C6. The method (200) according to any of paragraphs C1-C5, wherein the step of coupling (206) the leading edge skin panel (54) includes abutting the second end region (58) of the leading edge skin panel (54) and the lower skin panel (36).

C7. The method (200) of any of paragraphs C1-C6, wherein the joining (206) of the leading edge skin panel (54) is performed without the use of a splicing tape.

C8. The method (200) of any of paragraphs C1-C7, wherein the coupling (202) of the upper skin panel (34) to the forward C-channel spar (38) is performed without the use of a nut plate.

C9. The method (200) of any of paragraphs C1-C8, wherein the coupling (204) of the lower skin panel (36) to the forward C-channel spar (38) is performed without the use of a nut plate.

C10. The method (200) according to any of paragraphs C1-C9, further comprising coupling (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).

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

D1. Use of a structural composite airfoil (10) according to any of paragraphs a1-a55 as an inboard flap of an aircraft (14).

D2. Use of a structural composite airfoil (10) according to any of paragraphs a1-a55 as an outboard flap of an aircraft (14).

As used herein, the terms "selective" and "selectively" when modifying an action, motion, configuration, or other activity of one or more components or features of a device mean that the particular action, motion, configuration, or other activity is a direct or indirect result of a user manipulating an aspect or one or more components of the device.

As used herein, the terms "adapted" and "configured" mean that an element, component, or other subject matter is designed and/or intended to perform a given function. Thus, the use of the terms "adapted" and "configured" should not be construed to mean that a given element, component, or other subject matter is merely "capable" of performing a given function, but rather that an element, component, and/or other subject matter is specifically selected, created, implemented, utilized, programmed, and/or designed for the purpose of performing the function. It is within the scope of the present disclosure that elements, components, and/or other recited subject matter recited as being suitable for performing a particular function may additionally or alternatively be described as being configured to perform that function, and vice versa. Similarly, subject matter recited as being configured to perform a particular function may additionally or alternatively be described as being operable to perform that function.

As used herein, the phrase "at least one" in reference to a list of one or more entities should be understood to mean at least one entity selected from any one or more entities in the list of entities, but does not necessarily include at least one of each of the entities explicitly listed within the list of entities, and does not exclude any combination of entities in the list of entities. This definition also allows for the selective presence of entities, whether related or unrelated to those specifically identified entities, in addition to the specifically identified entities within the list of entities referred to by the phrase "at least one". Thus, as a 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 refer, in one embodiment, to the absence of at least one of B, optionally including more than one a (and optionally including entities other than B); in another embodiment, at least one of a is absent, optionally including more than one B (and optionally including entities other than a); in another embodiment, at least one, optionally more than one, a, and at least one, optionally more than one, B (and optionally other entities) are referred to. In other words, the phrases "at least one," "one or more," and/or "are open-ended expressions that, in operation, may be both conjunctive and disjunctive. 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" may 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 in combination with at least one other entity, of the foregoing.

Not all devices and methods in accordance with the present disclosure require the various elements and method steps of the devices disclosed herein, and the present disclosure includes all novel and nonobvious combinations and subcombinations of the various elements and steps disclosed herein. Further, one or more of the various elements and steps disclosed herein may define independent inventive subject matter that is independent of and separate from the entirety of the disclosed apparatus or method. Thus, such inventive subject matter need not be associated with the particular apparatus and methods explicitly disclosed herein, and such inventive subject matter may find utility in apparatus and/or methods not explicitly disclosed herein.

As used herein, the phrase "for example," the phrase "as an example" and/or simply the term "example," when used in reference to one or more components, features, details, structures, embodiments, and/or methods according to the present disclosure, is intended to convey that the described components, features, details, structures, embodiments, and/or methods are illustrative, non-exclusive examples of components, features, details, structures, embodiments, and/or methods according to the present disclosure. Thus, the described components, features, details, structures, embodiments, and/or methods are not intended to be limiting, required, or exclusive/exhaustive; and other components, features, details, structures, embodiments, and/or methods that include structurally and/or functionally similar and/or equivalent components, features, details, structures, embodiments, and/or methods are also within the scope of the present disclosure.