Composite fuselage frame and method of manufacturing a frame filler and composite fuselage frame

Abstract

The present application relates to composite fuselage frames and methods of manufacturing frame fillers and composite fuselage frames. A composite fuselage frame for an aircraft includes a web and a shear tie. The web includes a width and a first thickness. The web forms an arc between a first end and a second end relative to a fuselage. The width extends from an outboard side to an inboard side relative to a skin on the fuselage. A shear tie is disposed on the outboard side of the web between a first mousehole and a second mousehole and the shear tie protrudes from the first thickness of the web. The shear tie has a second thickness that extends along the skin between a first stringer and a second stringer. The second thickness of the shear tie is defined by the first thickness and a second polymer matrix composite layer formed from a first material polymer matrix composite ply that provides a frame filler integrated with a second material polymer matrix composite ply of the first thickness.

Description

Technical Field

[0001] This disclosure generally relates to composite fuselage frames for aircraft, and more specifically to composite fuselage frames integrated with frame filler. Various methods for manufacturing laminates for the frame filler and integrating the laminates into the layups of the composite fuselage frame are disclosed. Various additional methods for integrating the layups of the frame filler with the layups of the composite fuselage frame are also disclosed. Background Technology

[0002] Currently, the frame filler is placed manually on the fuselage skin and co-cured with the skin. For example, a total of six fuselage sections require these frame fillers. The current frame filler is co-cured with the fuselage skin laminate, and then the separately manufactured frames are fastened to the skin via stringers on the skin between the frame fillers.

[0003] Therefore, those skilled in the art continue to conduct research and development to reduce the manufacturing time and weight of the fuselage section of an aircraft relative to the skin, frame filler, and fuselage frame that supports the skin. Summary of the Invention

[0004] Examples of composite fuselage frames, methods for manufacturing frame fillers for composite fuselage frames, methods for manufacturing composite fuselage frames for aircraft, and methods for manufacturing fuselage cylindrical sections for aircraft are disclosed. The following is a non-exhaustive list of examples of the subject matter of this disclosure, which may or may not be claimed.

[0005] In one example, the disclosed composite fuselage frame for an aircraft includes an inner chord, a web, and shear-resistant connecting feet. The inner chord is a flange opposite to the shear-resistant connecting feet.

[0006] In another example, the disclosed composite fuselage frame for an aircraft includes a web and shear-resistant connecting feet. The web includes a width and a first predetermined thickness. The web forms an arc relative to the fuselage of the aircraft between a first end and a second end. The width extends from the outer side of the skin relative to the fuselage to the inner side. The shear-resistant connecting feet are disposed on the outer side of the web, between a first rat hole and a second rat hole, and protrude from the first predetermined thickness of the web. The shear-resistant connecting feet have a second predetermined thickness that extends along the skin of the aircraft between a first stringer and a second stringer. The second predetermined thickness of the shear-resistant connecting feet is defined by the first predetermined thickness and a plurality of second polymer matrix composite layers formed of a plurality of first material polymer matrix composite sheets, the plurality of first material polymer matrix composite sheets providing frame filler integrated with the plurality of second material polymer matrix composite sheets of the first predetermined thickness.

[0007] In the example, the disclosed method of manufacturing frame filler for a composite fuselage frame includes: (a) laying a plurality of first material polymer matrix composite sheets on a first laying tool for frame filler to provide a first stack of the first material polymer matrix composite sheets on a first top surface of the first laying tool; and (b) forming the first stack.

[0008] In the example, the disclosed method for manufacturing a composite fuselage frame for an aircraft includes: (a) laying a plurality of first material polymer matrix composite sheets at predetermined locations on a first portion of the top surface of a laying tool for the composite fuselage frame to form a first stack of the first material polymer matrix composite sheets, the predetermined locations being associated with shear-resistant connecting feet of the composite fuselage frame and frame filler; (b) laying a plurality of second material polymer matrix composite sheets on the first material polymer matrix composite sheets and a second portion of the top surface of the laying tool to provide a second stack of the second material polymer matrix composite sheets; and (c) forming the second stack and the first stack.

[0009] In the example, the disclosed method for manufacturing a fuselage section of an aircraft includes: (a) receiving a skin in a manufacturing unit, the skin having at least a first stringer and a second stringer on its inner surface, the skin having the first stringer and the second stringer having been previously manufactured and hardened; (b) mounting the skin onto a manufacturing tool in the manufacturing unit; (c) receiving a composite fuselage frame at the manufacturing unit, the composite fuselage frame having integrated frame filler, the composite fuselage frame having been previously manufactured and hardened; and (d) mounting the composite fuselage frame onto the first stringer and the second stringer on the inner surface of the skin.

[0010] Other examples of the disclosed composite fuselage frame, the method of manufacturing frame filler for the composite fuselage frame, the method of manufacturing a composite fuselage frame for an aircraft, and the method of manufacturing a fuselage section for an aircraft will become apparent from the following detailed description, drawings, and appended claims. Attached Figure Description

[0011] Figure 1 This is a perspective view of an example of a composite fuselage frame with integrated frame filler on shear connection feet;

[0012] Figure 2 This is a perspective view of an example aircraft;

[0013] Figure 3 This is a perspective view of an example of a partially assembled fuselage section;

[0014] Figure 4 This is a cross-sectional side view of another example of a composite fuselage frame;

[0015] Figure 5 yes Figure 4 Another cross-sectional side view of the composite fuselage frame, the other cross-sectional side view showing the polymer matrix composite layer;

[0016] Figure 6 This is an exploded front view showing an example of a polymer matrix composite layer and an outer protective glass fiber layer in another example of a composite fuselage frame;

[0017] Figure 7 It is shown Figure 6 An exploded side view of the polymer matrix composite layer and the outer protective glass fiber layer in the composite fuselage frame;

[0018] Figure 8 yes Figure 4 and Figure 5 A side view of the composite fuselage frame, showing an example of frame filler (indicated by dashed lines) integrated within the composite fuselage frame;

[0019] Figure 9 This is a flowchart illustrating an example of a method for manufacturing frame filler for composite fuselage frames;

[0020] Figure 10 It is shown that it is used for Figure 4 and Figure 5 An exploded front view of an example of a polymer matrix composite layer in an example of a composite fuselage frame filler;

[0021] Figure 11 yes Figure 10 A front view of an example of a frame filler laminate;

[0022] Figure 12 This is a top view of an example of a sheet of composite fabric made of a first material polymer matrix;

[0023] Figure 13 Combination Figure 9 This is a flowchart of another example of a method for manufacturing frame filler for composite fuselage frames;

[0024] Figure 14 This is a rear view of an example of a laying tool, on which a [tool / item] is placed. Figure 8 An example of a laminate of a frame filler (dashed line) and a second material polymer matrix composite layer covered on the laying tool;

[0025] Figure 15 yes Figure 14 An exploded side view of the laying tool, which shows an example of a laminate of frame filler (shown in dashed lines) and a stack of second material polymer matrix composite sheets covering the frame filler;

[0026] Figure 16 This is a top view of an example of a sheet of a second material polymer matrix composite tape;

[0027] Figure 17 Combination Figure 9 and Figure 13 This is a flowchart illustrating yet another example of a method for manufacturing frame filler for composite fuselage frames;

[0028] Figure 18 This is a side view of an example laminate of a composite fuselage frame, showing the integration within the composite fuselage frame. Figure 11 Laminated components of the frame filler in the middle;

[0029] Figure 19 Combination Figure 9 and Figure 13 This is a flowchart illustrating yet another example of a method for manufacturing frame filler for composite fuselage frames;

[0030] Figure 20 yes Figure 14 An exploded side view of the laying tool shows an example of a second material polymer matrix composite sheet covering the laying tool and awaiting placement of the frame filler;

[0031] Figure 21 Combination Figure 9 and Figure 13 This is a flowchart illustrating yet another example of a method for manufacturing frame filler for composite fuselage frames;

[0032] Figure 22 Combination Figure 9 and Figure 13 This is a flowchart of another example of a method for manufacturing frame filler for composite fuselage frames;

[0033] Figure 23 Combination Figure 9 This is a flowchart illustrating yet another example of a method for manufacturing frame filler for composite fuselage frames;

[0034] Figure 24 Combination Figure 9 and Figure 23 This is a flowchart illustrating yet another example of a method for manufacturing frame filler for composite fuselage frames;

[0035] Figure 25 Combination Figure 9 and Figure 23 This is a flowchart illustrating yet another example of a method for manufacturing frame filler for composite fuselage frames;

[0036] Figure 26 Combination Figure 9 and Figure 23This is a flowchart of another example of a method for manufacturing frame filler for composite fuselage frames;

[0037] Figure 27 Combination Figure 9 , Figure 23 and Figure 25 This is a flowchart illustrating yet another example of a method for manufacturing frame filler for composite fuselage frames;

[0038] Figure 28 This is a flowchart illustrating an example of a method for manufacturing a composite fuselage frame for an aircraft;

[0039] Figure 29 Combination Figure 28 This is a flowchart illustrating yet another example of a method for manufacturing a composite fuselage frame for an aircraft;

[0040] Figure 30 Combination Figure 28 This is a flowchart illustrating yet another example of a method for manufacturing a composite fuselage frame for an aircraft;

[0041] Figure 31 Combination Figure 28 This is a flowchart illustrating yet another example of a method for manufacturing a composite fuselage frame for an aircraft;

[0042] Figure 32 Combination Figure 28 This is a flowchart illustrating another example of a method for manufacturing a composite fuselage frame for an aircraft;

[0043] Figure 33 and Figure 34 A flowchart is provided as an example of a method for manufacturing a fuselage section for an aircraft, the method including manufacturing a composite fuselage frame;

[0044] Figure 35 Combination Figure 33 and Figure 34 This is a flowchart of another example of a method for manufacturing a fuselage section for an aircraft, the method including manufacturing a composite fuselage frame;

[0045] Figure 36 It is a block diagram of an aircraft manufacturing and maintenance method, which includes a composite fuselage frame manufactured using one or more examples of the methods disclosed herein, wherein frame filler is integrated in the composite fuselage frame;

[0046] Figure 37 This is a schematic diagram of an aircraft that includes a composite fuselage frame manufactured using one or more examples of the methods disclosed herein, in which frame filler is integrated;

[0047] Figure 38 Combination Figure 28and Figure 30 This is a flowchart illustrating yet another example of a method for manufacturing a composite fuselage frame for an aircraft;

[0048] Figure 39 Is Figure 28 A flowchart illustrating an example of forming the second and first stacks in method 2800;

[0049] Figure 40 This is a side view of an example of a forming tool used for a composite fuselage frame;

[0050] Figure 41 This is a side view of an example compaction tool used for a composite fuselage frame;

[0051] Figure 42 yes Figure 28 A flowchart illustrating another example of the formation of the second and first stacks;

[0052] Figure 43 yes Figure 9 A flowchart illustrating an example of forming the first stack;

[0053] Figure 44 This is a side view of an example of a forming tool used for frame filler;

[0054] Figure 45 This is a side view of an example compaction tool used for frame filler;

[0055] Figure 46 yes Figure 9 A flowchart illustrating another example of forming the first stack;

[0056] Figure 47 yes Figure 13 A flowchart illustrating an example of forming the second stack and the first laminate;

[0057] Figure 48 yes Figure 13 A flowchart illustrating another example of the formation of the second stack and the first laminate;

[0058] Figure 49 yes Figure 23 A flowchart illustrating an example of forming a base and a first laminate;

[0059] Figure 50 yes Figure 23 A flowchart of another example of the formation of the base and the first laminate;

[0060] Figure 51 yes Figure 34 The flowcharts illustrating the formation of the second and first stacks are shown in the figure; and

[0061] Figure 52 yes Figure 34 The flowchart shows another example of the formation of the second and first stacks. Detailed Implementation

[0062] The various examples of composite fuselage frames 100 disclosed herein, methods 900, 1300, 1700, 1900, 2100, 2200, 2300, 2400, 2500, 2600, 2700 for manufacturing frame filler 802 for composite fuselage frames 100, methods 2800, 2900, 3000, 3100, 3200, 3800 for manufacturing composite fuselage frames 100 for aircraft 200, and methods 3300, 3500 for manufacturing fuselage cylindrical sections 310 for aircraft 200 provide techniques for integrating frame filler 802 into shear-resistant connecting feet 114 of composite fuselage frames 100 via two components: clamping arc forming and co-curing. This reduces the total number of manufacturing parts, the total processing time, and the overall weight of the aircraft. The frame filler 802 is laid before forming the composite fuselage frame 100 and integrated into a first sheet form on a tool. The remainder of the composite fuselage frame 100 is laid out normally, and upon curing, the two components become a single continuous part. The frame filler 802 can be formed together with the first layer of the component, minimizing the increase in frame flow time but significantly reducing the current skin co-curing flow time. This also reduces the risk of composite skin construction and reduces weight through appropriately sized frame fillers. For the thousands of frame fillers 802 used per aircraft 200, weight savings are achieved by appropriately reducing the material of each frame filler 802 by two inches. For example, there may be six composite fuselage frame 100 segments that require integrated frame fillers 802. The advantage is shifting time / labor from skin construction to the construction of smaller part features. This also reduces the risk in skin construction, as any defects in the composite fuselage frame 100 will result in much lower scrap / rework costs. This enables improved assembly methods for joining the composite fuselage frame 100 to the fuselage 202 and improved methods for manufacturing the composite fuselage frame 100.

[0063] The technology disclosed herein combines features from two parts, resulting in reduced processing time and weight savings. The weight savings are due to the appropriate dimensional design of the frame filler 802. For example, current skin frame fillers can extend approximately one inch beyond the shear connection feet of the frame in either direction, while the composite fuselage frame 100, in which the frame filler 802 is integrated, will have an appropriate dimension, reducing material by two inches per frame filler for the thousands of frame fillers 802 used in each aircraft 200. For example, the frame filler 802 can be manually laminated prior to the formation of the composite fuselage frame 100. For example, the laminations can be a + / -45 degree oriented BMS8-276 plain weave (PW) material (or a future BMS8-416PW). This material is compatible with the frame resin and can be co-cured with the composite fuselage frame 100. The frame filler 802 can be aligned into the initially formed laminations, for example, via optical laser template projection and vacuum compaction. Once this is complete, the first lamination can be formed using clamping arcs, as is normal for a single-piece frame. The rest of the frame forming process will not be affected, and once cured, these components will form a continuous composite fuselage frame 100, with the frame filler 802 integrated in the composite fuselage frame ready to be installed onto the skin 204 of the fuselage 202.

[0064] The frame filler 802 is integrated with the composite fuselage frame 100 and directly fastened to the skin. The frame filler 802 is co-cured with the composite fuselage frame 100, which is significantly different from any current process. In other words, current aircraft do not use a single-piece frame structure for the fuselage frame, which includes integrated frame filler for attachment to the skin or any form of integrated frame filler within the frame for attachment to the skin. Once the frame filler 802 is laid out as a sheet of the composite fuselage frame 100 and co-cured, the frame filler 802 becomes part of the composite fuselage frame 100. Pre-stacking, + / -45 degree sheet orientation, OLT positioning on the first sheet, and clamping arc formation of the glass fiber / frame filler / and the first sheet are all techniques that can be implemented in any suitable combination. The integrated frame filler design can be used for any other type of composite material or part construction to attach the frame to the fuselage. For example, integrated frame filler designs can be used in composite aircraft to reduce manufacturing risks, save time, weight and materials compared to current processes, and allow for different fuselage skin tooling and manufacturing, which can also lead to time / quality benefits.

[0065] General Reference Figures 1 to 8 and Figure 18 As an example, this disclosure relates to a composite fuselage frame 100 for an aircraft 200. Figure 1 This is a perspective view of an example of a composite fuselage frame 100. Figure 2 This is a perspective view of an example of aircraft 200. Figure 3This is a perspective view of an example of a partially assembled fuselage section 310. Figure 4 This is a cross-sectional side view of another example of the composite fuselage frame 100. Figure 5 yes Figure 4 Another cross-sectional side view of the composite fuselage frame 100 shows the polymer matrix composite layers 502 and 504. Figure 6 This is an exploded front view showing an example of the polymer matrix composite layers 502, 504 and the outer protective glass fiber layer 506 in another example of the composite fuselage frame 100. Figure 7 It is shown Figure 6 An exploded side view of the polymer matrix composite layers 502, 504 and the outer protective glass fiber layer 506 in the composite fuselage frame 100. Figure 8 yes Figure 4 and Figure 5 A side view of the composite fuselage frame 100 shows an example of the frame filler 802 (indicated by dashed lines) integrated within the composite fuselage frame 100. Figure 18 This is a side view of an example of the laminate 1502 of the composite fuselage frame 100, showing its integration within the composite fuselage frame 100. Figure 11 The laminate 1102 of the frame filler 802 in the middle.

[0066] Refer again Figures 1 to 8 and Figure 18 In one or more examples, the composite fuselage frame 100 for the aircraft 200 includes a web 102 and a shear-resistant connecting foot 114. The composite fuselage frame 100 may also include an inner chord 115 located on the side of the web 102 opposite to the shear-resistant connecting foot 114. The web 102 includes a width 103 and a first predetermined thickness 402. The web 102 forms an arc 104 relative to the fuselage 202 of the aircraft 200 between a first end 106 and a second end 108. The width 103 extends from an outer side 404 of the skin 204 on the fuselage 202 to an inner side 406. The shear-resistant connecting foot 114 is disposed on the outer side 404 of the web 102 between a first rat hole 110 and a second rat hole 112, and protrudes from the first predetermined thickness 402 of the web 102. The shear-resistant connecting leg 114 has a second predetermined thickness 408 that extends along the skin 204 of the aircraft 200 between the first stringer 302 and the second stringer 304. The second predetermined thickness 408 of the shear-resistant connecting leg 114 is defined by the first predetermined thickness 402 and a plurality of second polymer matrix composite layers 504, which are formed of a plurality of first material polymer matrix composite sheets 604 that provide frame filler 802 integrated with the plurality of second material polymer matrix composite sheets 602 of the first predetermined thickness 402.

[0067] In another example, the composite fuselage frame 100 also includes a first rat hole 110 and a second rat hole 112. The first rat hole 110 is disposed on the outer side 404 of the web 102 and is sized to fit above a first stringer 302 on the skin 204. The second rat hole 112 is spaced apart from the first rat hole 110 on the outer side 404 of the web 102 and is sized to fit above a second stringer 304 on the skin 204 of the fuselage 202 adjacent to the first stringer 302.

[0068] In yet another example, the composite fuselage frame 100 is configured to provide lateral support 314 to the inner surface 312 of the skin 204 of a fuselage cylindrical section 310 having a plurality of stringers 305. The composite fuselage frame 100 spans the plurality of stringers 305. The composite fuselage frame 100 includes a plurality of shear-resistant connecting feet 114 extending toward the skin 204 between adjacent stringers 305. During the manufacture of the composite fuselage frame 100, each shear-resistant connecting foot 114 has frame filler 802 integrated in each shear-resistant connecting foot, such that the composite fuselage frame 100 includes a plurality of frame filler 802 integrated within the plurality of shear-resistant connecting feet 114.

[0069] In yet another example, the composite fuselage frame 100 further includes a flange 1802 connected to a web 102, which in turn is connected to a shear link 114. The shear link 114 includes frame filler 802 added to the flange 1802. The frame filler 802 is equal to the thickness of the stringer flanges on the first stringers 302 and the second stringers 304. In yet another example of the composite fuselage frame 100, a first predetermined thickness 402 of the web 102 is defined by a first plurality of first polymer matrix composite layers 502 formed by a plurality of second material polymer matrix composite sheets 602. In yet another example of the composite fuselage frame 100, the shear link 114 extends from the web 102 to a distal end 409 of the shear link 114, and the outer surface 410 of the shear link 114 is configured to face the skin 204 and align with a corresponding portion of the skin 204. In another example of the composite fuselage frame 100, a shear connecting foot 114 extends from a first predetermined thickness 402 of the web 102 such that one side of the shear connecting foot 114 is configured to face the skin 204 and fit within the area defined by the first stringer 302, the second stringer 304, and the corresponding portion of the skin 204 between the first stringer 302 and the second stringer 304.

[0070] In another example of the composite fuselage frame 100, the second material polymer matrix composite sheet 602 comprises a thermosetting matrix composite sheet, a thermoplastic matrix composite sheet, or any other suitable unidirectional polymer matrix composite sheet in any suitable combination. In yet another example of the composite fuselage frame 100, the second material polymer matrix composite sheet 602 comprises unidirectional reinforcing fibers and a polymer matrix material. In another example, the unidirectional reinforcing fibers comprise carbon fibers, glass fibers, aramid fibers, natural fibers, polyetheretherketone fibers, or any other suitable reinforcing fibers in any suitable combination. In yet another further example, the polymer matrix material comprises thermosetting resins, thermoplastic resins, epoxy resins, phenolic resins, polyurethane resins, polyimide resins, polyethylene resins, polypropylene resins, polybutylene terephthalate resins, polyamide resins, polyphenylene sulfide resins, polyetherimide resins, polyetherketone ketone resins, polyetheretherketone resins, or any other suitable polymer matrix material in any suitable combination.

[0071] In another example of the composite fuselage frame 100, the first material polymer matrix composite sheet 604 comprises a thermosetting matrix composite sheet, a thermoplastic matrix composite sheet, or any other suitable woven polymer matrix composite sheet in any suitable combination. In yet another example of the composite fuselage frame 100, the first material polymer matrix composite sheet 604 comprises a reinforcing fabric and a polymer matrix material. In yet another example, the reinforcing fabric comprises a plain weave fabric, a 0 / 90-degree woven fabric, a 45-degree woven fabric, or any other suitable woven reinforcing fabric in any suitable combination. In yet another example, the reinforcing fabric comprises woven reinforcing fibers interwoven in at least two directions. In a further example, the woven reinforcing fibers comprise glass fibers, carbon fibers, aramid fibers, natural fibers, polyetheretherketone fibers, or any other suitable reinforcing fibers in any suitable combination. In yet another further example, the polymer matrix material includes thermosetting resins, thermoplastic resins, epoxy resins, phenolic resins, polyurethane resins, polyimide resins, polyethylene resins, polypropylene resins, polybutylene terephthalate resins, polyamide resins, polyphenylene sulfide resins, polyetherimide resins, polyetherketone ketone resins, polyetheretherketone resins, or any other suitable polymer matrix material in any suitable combination.

[0072] In another example of the composite fuselage frame 100, the web 102 and the shear-resistant connecting foot 114 further include an external protective fiberglass layer 506 extending between a first end 106 and a second end 108 of the web 102, and extending along the inner surface 412 and outer surface 410 of the shear-resistant connecting foot 114 between a first rat hole 110 and a second rat hole 112. The external protective fiberglass layer 506 is formed of an external protective fiberglass sheet 606. In yet another example, the composite fuselage frame 100 also includes a single-piece segment 116. In yet another example of the composite fuselage frame 100, multiple composite fuselage frames 100 form a segmented frame 306 for a slice 308 of the fuselage 202 of the aircraft 200.

[0073] General Reference Figures 1 to 27 , Figure 40 , Figure 41 and Figures 43 to 50 As an example, this disclosure relates to methods 900, 1300, 1700, 1900, 2100, 2200, 2300, 2400, 2500, 2600, and 2700 for manufacturing frame filler 802 for composite fuselage frame 100. Figure 1 This is a perspective view of an example of a composite fuselage frame 100. Figure 2 This is a perspective view of an example of aircraft 200. Figure 3 This is a perspective view of an example of a partially assembled fuselage section 310. Figure 4 This is a cross-sectional side view of another example of the composite fuselage frame 100. Figure 5 yes Figure 4 Another cross-sectional side view of the composite fuselage frame 100 shows the polymer matrix composite layers 502 and 504. Figure 6 This is an exploded front view showing an example of the polymer matrix composite layers 502, 504 and the outer protective glass fiber layer 506 in another example of the composite fuselage frame 100. Figure 7 It is shown Figure 6 An exploded side view of the polymer matrix composite layers 502, 504 and the outer protective glass fiber layer 506 in the composite fuselage frame 100. Figure 8 yes Figure 4 and Figure 5 A side view of the composite fuselage frame 100 shows an example of the frame filler 802 (indicated by dashed lines) integrated within the composite fuselage frame 100.

[0074] Figure 9 An example of a method 900 for manufacturing a frame filler 802 for a composite fuselage frame 100 is provided. Figure 10 It is shown that it is used for Figure 4 and Figure 5An exploded front view of an example of polymer matrix composite layers 502, 504 in an example of frame filler 802 of a composite fuselage frame 100. Figure 11 yes Figure 10 A front view of an example of the laminate 1102 of the frame filler 802. Figure 12 This is a top view of an example of a sheet of the first material polymer matrix composite fabric 1202. Figure 13 Combination Figure 9 An example of a method 1300 for manufacturing frame filler 802 for composite fuselage frame 100 is provided. Figure 14 This is a rear view of an example of laying tool 1404. Figure 8 The laminate 1102 of the frame filler 802 (dashed line) is placed on the laying tool 1404, and an example of the second material polymer matrix composite sheet 602 is covered on the laying tool 1404. Figure 15 yes Figure 14 An exploded side view of the laying tool 1404 shows an example of a laminate 1102 (shown in dashed lines) of the frame filler 802 and a second material polymer matrix composite sheet 602 covering the frame filler 802. Figure 16 This is a top view of an example of a sheet of the second material polymer matrix composite tape 1602. Figure 17 Combination Figure 9 and Figure 13 An example of a method 1700 for manufacturing frame filler 802 for composite fuselage frame 100 is provided. Figure 18 This is a side view of an example of the laminate 1502 of the composite fuselage frame 100, showing its integration within the composite fuselage frame 100. Figure 11 The laminate 1102 of the frame filler 802 in the middle.

[0075] Figure 19 Combination Figure 9 and Figure 13 An example of a method 1900 for manufacturing frame filler 802 for composite fuselage frame 100 is provided. Figure 20 yes Figure 14 An exploded side view of the laying tool 1404 shows an example of a second material polymer matrix composite sheet 602 covering the laying tool 1404 and awaiting placement of the frame filler 802. Figure 21 Combination Figure 9 and Figure 13 An example of a method 2100 for manufacturing frame filler 802 for composite fuselage frame 100 is provided. Figure 22 Combination Figure 9 and Figure 13 An example of a method 2200 for manufacturing frame filler 802 for composite fuselage frame 100 is provided. Figure 23 Combination Figure 9An example of a method 2300 for manufacturing frame filler 802 for composite fuselage frame 100 is provided. Figure 24 Combination Figure 9 and Figure 23 An example of a method 2400 for manufacturing frame filler 802 for composite fuselage frame 100 is provided. Figure 25 Combination Figure 9 and Figure 23 An example of a method 2500 for manufacturing frame filler 802 for composite fuselage frame 100 is provided. Figure 26 Combination Figure 9 and Figure 23 An example of a method 2600 for manufacturing frame filler 802 for composite fuselage frame 100 is provided. Figure 27 Combination Figure 9 , Figure 23 and 25 An example of a method 2700 for manufacturing frame filler 802 for composite fuselage frame 100 is provided.

[0076] Figure 40 This is a side view of an example of a forming tool 4004 used for a composite fuselage frame 100. Figure 41 This is a side view of an example of a compaction tool 4104 used for the composite fuselage frame 100. Figure 43 Provided in Figure 9 An example of forming the first stack 1004 in method 900. Figure 44 This is a side view of an example of a forming tool 4404 used for frame filler 802. Figure 45 This is a side view of an example of a compaction tool 4504 used for frame filler 802. Figure 46 Provided in Figure 9 Another example of forming the first stack 1004 in method 900. Figure 47 Provided Figure 13 An example of the formation of the second stack 1506 and the first laminate 1102 in the process. Figure 48 Provided Figure 13 Another example of the formation of the second stack 1506 and the first laminate 1102 in the process. Figure 49 Provided Figure 23 An example of the formation of the base 2002 and the first laminate 1102 in the middle. Figure 50 Provided Figure 23 Another example of the formation of the base 2002 and the first laminate 1102 in the process.

[0077] Refer again Figure 1 , Figures 6 to 12 and Figures 43 to 46In one or more examples, a method 900 for manufacturing frame filler 802 for composite fuselage frame 100 (see...) Figure 9 This includes laying a plurality of first material polymer matrix composite sheets 604 902 on a first laying tool 1002 for a frame filler 802 to provide a first stack 1004 of the first material polymer matrix composite sheets 604 on a first top surface 1006 of the first laying tool 1002. The first stack 1004 is formed at 904.

[0078] In yet another example of method 900, the first material polymer matrix composite sheet 604 comprises a thermosetting matrix composite sheet, a thermoplastic matrix composite sheet, or any other suitable woven polymer matrix composite sheet in any suitable combination. In yet another example of method 900, the first material polymer matrix composite sheet 604 comprises a reinforcing fabric and a polymer matrix material. In another example, the reinforcing fabric comprises a plain weave fabric, a 0 / 90 degree woven fabric, a 45 degree woven fabric, or any other suitable woven reinforcing fabric in any suitable combination. In yet another example, the reinforcing fabric comprises woven reinforcing fibers interwoven in at least two directions. In a further example, the woven reinforcing fibers comprise glass fibers, carbon fibers, aramid fibers, natural fibers, polyetheretherketone fibers, or any other suitable reinforcing fibers. In yet another further example, the polymer matrix material includes thermosetting resins, thermoplastic resins, epoxy resins, phenolic resins, polyurethane resins, polyimide resins, polyethylene resins, polypropylene resins, polybutylene terephthalate resins, polyamide resins, polyphenylene sulfide resins, polyetherimide resins, polyetherketone ketone resins, polyetheretherketone resins, or any other suitable polymer matrix material in any suitable combination. In another example of method 900, the composite fuselage frame 100 also includes a one-piece section 116.

[0079] In yet another example, method 900 further includes cutting 906 a plurality of first material polymer matrix composite sheets 604 from one or more sheets of the first material polymer matrix composite fabric 1202. In this example, method 900 continues from 906 to 902. In another example of method 900, forming 904 the first stack 1004 includes: forming the first stack 1004 4302 onto the top surface 4402 of a forming tool 4404. At 4304, the first stack 1004 is compacted onto the top surface 4502 of a compaction tool 4504. In yet another example of method 900, forming 904 the first stack 1004 includes: forming the first stack 1004 4602 onto the first top surface 1006 of a first laying tool 1002. At 4604, during the formation of 4602, the first stack 1004 is compacted to provide a first laminate 1102 of the frame filler 802.

[0080] Refer again Figures 1 to 11 , Figures 13 to 15 , Figure 40 , Figure 41 and Figures 46 to 48 In one or more examples, method 1300 for manufacturing frame filler 802 for composite fuselage frame 100 (see...) Figure 13 )include Figure 9 Method 900, and from Figure 46 Continuing from 4604 to 1302, a first laminate 1102 of the frame filler 802 is removed from the first laying tool 1002. At 1304, the first laminate 1102 is placed on a first portion 1504 of the second top surface 1402 of the second laying tool 1404 for the composite fuselage frame 100, the first portion 1504 being associated with the frame filler 802 relative to the shear connection foot 114 of the composite fuselage frame 100. At 1306, a plurality of second material polymer matrix composite sheets 602 are laid on the first laminate 1102 and the second portion 1510 of the second top surface 1402 of the second laying tool 1404 to provide a second stack 1506 of the second material polymer matrix composite sheets 602. At 1308, the second stack 1506 and the first laminate 1102 are formed.

[0081] In another example of method 1300, the second material polymer matrix composite sheet 602 comprises a thermosetting matrix composite sheet, a thermoplastic matrix composite sheet, or any other suitable unidirectional polymer matrix composite sheet in any suitable combination. In yet another example of method 1300, the second material polymer matrix composite sheet 602 comprises unidirectional reinforcing fibers and a polymer matrix material. In another example, the unidirectional reinforcing fibers comprise carbon fibers, glass fibers, aramid fibers, natural fibers, polyetheretherketone fibers, or any other suitable reinforcing fibers in any suitable combination. In yet another further example, the polymer matrix material comprises a thermosetting resin, a thermoplastic resin, an epoxy resin, a phenolic resin, a polyurethane resin, a polyimide resin, a polyethylene resin, a polypropylene resin, a polybutylene terephthalate resin, a polyamide resin, a polyphenylene sulfide resin, a polyetherimide resin, a polyetheretherketone resin, a polyetheretherketone resin, or any other suitable polymer matrix material in any suitable combination.

[0082] In yet another example, method 1300 from Figure 46Continuing from 4604 to 1310, a plurality of second material polymer matrix composite sheets 602 are cut from one or more sheets of the second material polymer matrix composite strip 1602. In this example, method 1300 continues from 1310 to 1302. In yet another example of method 1300, forming 1308 the second stack 1506 and the first laminate 1102 includes: forming the second stack 1506 and the first laminate 1102 4702 onto the top surface 4002 of the forming tool 4004. At 4704, the second stack 1506 and the first laminate 1102 are compacted onto the top surface 4102 of the compaction tool 4104. In another example of method 1300, forming 1308 the second stack 1506 and the first laminate 1102 includes: forming 4802 of the second stack 1506 and the first laminate 1102 onto the second top surface 1402 of the second laying tool 1404. At 4804, during the formation of 4802, the second stack 1506 and the first laminate 1102 are compacted to provide the second laminate 1502 of the composite fuselage frame 100.

[0083] Refer again Figures 1 to 10 , Figure 13 , Figure 15 , Figure 17 and Figure 18 In one or more examples, method 1700 for manufacturing frame filler 802 for composite fuselage frame 100 (see...) Figure 17 )include Figure 9 Method 900 and Figure 13 Method 1300, and from Figure 48Continuing from 4804 to 1702, a second laminate 1502 of the composite fuselage frame 100 is cured to harden the composite fuselage frame 100. In this example, after the second laminate 1502 is cured at 1314, the composite fuselage frame 100 includes a web 102 and a flange 1802. The web 102 includes a width 103 and a first predetermined thickness 402. The web 102 forms an arc 104 relative to the fuselage 202 of the aircraft 200 between a first end 106 and a second end 108. The width 103 extends from an outer side 404 of the skin 204 on the fuselage 202 to an inner side 406. The flange 1802 is disposed on the outer side 404 of the web 102 and protrudes from the first predetermined thickness 402 of the web 102. At 1704, method 1700 continues by cutting a first mouse hole 110 from flange 1802 at a first predetermined position. The size of the first mouse hole 110 is set to fit above a first stringer 302 on the skin 204 of the fuselage 202. The first predetermined position is based on a first position of the first stringer 302 on the skin 204. At 1706, a second mouse hole 112, spaced apart from the first mouse hole 110, is cut from flange 1802 at a second predetermined position. The size of the second mouse hole is set to fit above a second stringer 304 adjacent to the first stringer 302 on the skin 204. The second predetermined position is based on a second position of the second stringer 304 on the skin 204.

[0084] In yet another example, the composite fuselage frame 100 is configured to provide lateral support 314 to the inner surface 312 of the skin 204 of a fuselage cylindrical section 310 having a plurality of stringers 305. The composite fuselage frame 100 spans the plurality of stringers 305. The composite fuselage frame 100 includes a plurality of shear-resistant connecting feet 114 extending toward the skin 204 between adjacent stringers 305. During the manufacture of the composite fuselage frame 100, each shear-resistant connecting foot 114 has frame filler 802 integrated in each shear-resistant connecting foot, such that the composite fuselage frame 100 includes a plurality of frame filler 802 integrated within the plurality of shear-resistant connecting feet 114.

[0085] In another example of method 1700, a shear-resistant connecting leg 114 extends from the web 102 between the first rat hole 110 and the second rat hole 112 to the distal end 1804 of the flange 1802. The outer surface 410 of the shear-resistant connecting leg 114 is configured to face the skin 204 and align with a corresponding portion of the skin 204. In yet another example of method 1700, the shear-resistant connecting leg 114 extends from a first predetermined thickness 402 of the web 102 such that one side of the shear-resistant connecting leg 114 is configured to face the skin 204 and fit within the area defined by the first stringer 302, the second stringer 304, and the corresponding portion of the skin 204 between the first stringer 302 and the second stringer 304. In yet another example of method 1700, a plurality of composite fuselage frames 100 form a segmented frame 306 for a section 308 of the fuselage 202 of the aircraft 200.

[0086] In yet another example of method 1700, a portion of flange 1802 between the first rat hole 110 and the second rat hole 112 defines a shear-resistant connecting leg 114 comprising a second predetermined thickness 408. The first predetermined thickness 402 of web 102 is defined by a first plurality of first polymer matrix composite layers 502 formed by a plurality of second material polymer matrix composite sheets 602. The second predetermined thickness 408 of shear-resistant connecting leg 114 is defined by the first predetermined thickness 402 and a second plurality of second polymer matrix composite layers 504 formed by a plurality of first material polymer matrix composite sheets 604, which provide a frame filler 802 integrated with the plurality of second material polymer matrix composite sheets 602 of the first predetermined thickness 402.

[0087] Refer again Figure 1 , Figures 6 to 10 , Figure 13 and Figure 15 In one or more examples, the method of manufacturing the frame filler 802 for the composite fuselage frame 100 includes Figure 9 Method 900 Figure 13 Method 1300, and continuing from 1302, involves integrating the frame filler 802 into the shear connection feet 114 of the composite fuselage frame 100 via a second stack 1506 forming a second material polymer matrix composite sheet 602 via a clamping arc above the first stack 1004 of the first material polymer matrix composite sheet 604. The frame filler 802 provided by the first stack 1004 and the shear connection feet 114 provided by the second stack 1506 are then co-cured to harden the composite fuselage frame 100.

[0088] Refer again Figure 1 , Figures 6 to 9 , Figure 11 , Figure 13 , Figure 14 , Figure 16 , Figure 19 and Figure 20 In one or more examples, method 1900 for manufacturing frame filler 802 for composite fuselage frame 100 (see Figure 19 )include Figure 9 Method 900 Figure 13 Method 1300, and continuing from 1302 to 1902, wherein, before placing the first laminate 1102 (1304) and laying multiple second material polymer matrix composite sheets 602, at least one second material polymer matrix composite sheet 602 is laid on the second top surface 1402 of the second laying tool 1404 to provide the base 2002 of the composite fuselage frame 100 below the first laminate 1102 of the frame filler 802. Method 1900 from Figure 13 1902 continues to 1304. At 1904, after the first laminate 1102 has been placed at 1304, the base 2002 and the first laminate 1102 are formed onto the second top surface 1402 of the second laying tool 1404. At 1906, during the formation of 1904, the base 2002 and the first laminate 1102 are compacted to provide the second laminate 2004 of the intermediate composite fuselage frame 2006. Method 1900 from Figure 13 Method 1906 continues to 1306. In another example, method 1900 starts from... Figure 13 Method 1302 continues to 1908, wherein at least one second material polymer matrix composite layer sheet 602 is cut from one or more sheets of the second material polymer matrix composite strip 1602. In this example, method 1900 continues from 1908 to 1902.

[0089] Refer again Figure 1 , Figure 5 , Figure 6 , Figure 9 , Figures 13 to 15 and Figure 21 In one or more examples, method 2100 for manufacturing frame filler 802 for composite fuselage frame 100 (see...) Figure 21 )include Figure 9 Method 900 Figure 13 Method 1300, and continuing from 1302 to 2102, wherein, prior to placing the first laminate 1304 and laying 1306 a plurality of second material polymer matrix composite sheets 602, an outer protective glass fiber sheet 606 is laid on the second top surface 1402 of the second laying tool 1404 to provide an outer protective glass fiber layer 506 beneath the first laminate 1102 of the frame filler 802. Method 2100 from Figure 132102 continues to 1304 and 1306. At 2104, after the first laminate is placed at 1304 and multiple second material polymer matrix composite sheets 602 are laid at 1306, the outer protective glass fiber sheet 606 is formed together with the second stack 1506 of the second material polymer matrix composite sheets 602 and the first laminate 1102 onto the second top surface 1402 of the second laying tool 1404. At 2106, during the formation of 2104, the outer protective glass fiber sheet 606 is compacted with the second stack 1506 and the first laminate 1102 to provide the second laminate 1502 of the composite fuselage frame 100.

[0090] Refer again Figure 1 , Figure 5 , Figure 6 , Figure 9 , Figure 11 , Figures 13 to 15 and Figure 22 In one or more examples, method 2200 for manufacturing frame filler 802 for composite fuselage frame 100 (see...) Figure 22 )include Figure 9 Method 900 Figure 13 Method 1300 continues from 1302 to 2202, wherein an outer protective glass fiber sheet 606 is laid on a second stack 1506 of the second material polymer matrix composite sheet 602 and a first laminate 1102 of the frame filler 802 currently on the second laying tool 1404 to provide an outer protective glass fiber layer 506 on top of the second stack 1506. At 2204, the outer protective glass fiber sheet 606, together with the second stack 1506 and the first laminate 1102, is formed onto a second top surface 1402 of the second laying tool 1404. At 2206, during the formation of 2204, the outer protective glass fiber sheet 606, together with the second stack 1506 and the first laminate 1102, is compacted to provide a second laminate 1502 of the composite fuselage frame 100.

[0091] Refer again Figure 1 , Figures 6 to 11 , Figure 14 , Figure 16 , Figure 20 , Figure 23 , Figure 40 , Figure 41 , Figure 49 and Figure 50 In one or more examples, method 2300 for manufacturing frame filler 802 for composite fuselage frame 100 (see...) Figure 23 )include Figure 9Method 900, and continuing from 906 to 2302, wherein the first laminate 1102 of the frame filler 802 is removed from the first laying tool 1002. At 2304, at least one second material polymer matrix composite sheet 602 is laid on the second top surface 1402 of the second laying tool 1404 for the composite fuselage frame 100 to provide a base 2002 of the composite fuselage frame 100. At 2306, the first laminate 1102 is placed on the base 2002 at a predetermined position associated with a shear connection foot 114 of the composite fuselage frame 100 and the frame filler 802. At 2308, the base 2002 and the first laminate 1102 are formed.

[0092] In another example, method 2300 further includes cutting 2310 at least one second material polymer matrix composite sheet 602 from one or more sheets of the second material polymer matrix composite strip 1602. In this example, method 2300 continues from 2310 to 2302. In yet another example of method 2300, forming 2308 the base 2002 and the first laminate 1102 includes: forming the base 2002 and the first laminate 1102 4902 onto the top surface 4002 of the forming tool 4004. At 4904, the base 2002 and the first laminate 1102 are compacted onto the top surface 4102 of the compaction tool 4104. In another example of method 2300, forming 2308 the base 2002 and the first laminate 1102 includes: forming 5002 of the base 2002 and the first laminate 1102 onto the second top surface 1402 of the second laying tool 1404. At 5004, during the formation of 5002, the base 2002 and the first laminate 1102 are compacted to provide a second laminate 2004 for the intermediate composite fuselage frame 2006.

[0093] Refer again Figure 1 , Figure 5 , Figure 6 , Figure 9 , Figure 11 , Figure 14 , Figure 20 , Figure 23 and Figure 24 In one or more examples, method 2400 for manufacturing frame filler 802 for composite fuselage frame 100 (see...) Figure 24 )include Figure 9 Method 900 Figure 23 Method 2300 and continuing from 2302 to 2402, wherein an outer protective fiberglass layer 606 is laid on the second top surface 1402 of the second laying tool 1404 to provide an outer protective fiberglass layer 506 below the base 2002 of the composite fuselage frame 100. This method from Figure 232402 continues to 2304 and 2306. At 2404, an outer protective fiberglass sheet 606 is formed together with the base 2002 and the first laminator 1102 onto the second top surface 1402 of the second laying tool 1404. At 2406, during the formation of 2404, the outer protective fiberglass sheet 606 is compacted with the base 2002 and the first laminator 1102 to provide a second laminator 2004 for the intermediate composite fuselage frame 2006.

[0094] Refer again Figure 1 , Figures 6 to 9 , Figure 11 , Figures 14 to 16 , Figure 20 , Figure 23 , Figure 25 and Figure 50 In one or more examples, method 2500 for manufacturing frame filler 802 for composite fuselage frame 100 (see...) Figure 25 )include Figure 9 Method 900 Figure 23 Method 2300, and from Figure 50 Continuing from 5004 to 2502, a plurality of second material polymer matrix composite sheets 602 are laid on a second laminate 2004 of the intermediate composite fuselage frame 2006 currently located on the second layup tool 1404 to provide a second stack 1506 of the second material polymer matrix composite sheets 602 on the second laminate 2004. At 2504, the second stack 1506 is formed onto a third top surface 2008 of the second laminate 2004, in which a first laminate 1102 of the frame filler 802 is integrated. At 2506, during the formation of 2504, the second stack 1506 and the second laminate 2004 of the intermediate composite fuselage frame 2006a are compacted to provide a third laminate 1508 of the composite fuselage frame 100. In another example, method 2500 further includes cutting 2508 a plurality of second material polymer matrix composite sheets 602 from one or more sheets of the second material polymer matrix composite tape 1602. In this example, the method continues from 2508 to 2502.

[0095] Refer again Figure 1 , Figure 5 , Figures 6 to 9 , Figure 11 , Figure 14 , Figure 15 , Figure 20 , Figure 23 and Figure 26 In one or more examples, method 2600 for manufacturing frame filler 802 for composite fuselage frame 100 (see...) Figure 26 )include Figure 9 Method 900 Figure 23 Method 2300 Figure 25 Method 2500, and continuing from 2506 to 2602, wherein an outer protective glass fiber sheet 606 is laid on a third laminate 1508 of the composite fuselage frame 100 currently on a second laying tool 1404 to provide an outer protective glass fiber layer 506 on top of the third laminate 1508. At 2604, the outer protective glass fiber sheet 606 is formed together with a second stack 1506 of a second material polymer matrix composite sheet 602 to a third top surface 2008 of the second laminate 2004 of the intermediate composite fuselage frame 2006, in which a first laminate 1102 of a frame filler 802 is integrated. At 2606, during the formation of 2604, the outer protective glass fiber sheet 606 is compacted together with the second stack 1506 of the second material polymer matrix composite sheet 602 and the second lamination 2004 to provide the third lamination 1508 of the composite fuselage frame 100.

[0096] Refer again Figures 1 to 4 , Figure 9 , Figure 15 , Figure 18 , Figure 23 , Figure 25 and Figure 27 In one or more examples, method 2700 for manufacturing frame filler 802 for composite fuselage frame 100 (see...) Figure 27 )include Figure 9 Method 900 Figure 23 Method 2300 Figure 25Method 2500, and continuing from 2506 to 2702, wherein a third laminate 1508 of the composite fuselage frame 100 is cured to harden the composite fuselage frame 100. In another example of method 2700, after the third laminate 1508 is cured at 2702, the composite fuselage frame 100 includes a web 102 and a flange 1802. The web 102 includes a width 103 and a first predetermined thickness 402. The web 102 forms an arc 104 relative to the fuselage 202 of the aircraft 200 between a first end 106 and a second end 108. The width 103 extends from an outer side 404 of the skin 204 on the fuselage 202 to an inner side 406. The flange 1802 is disposed on the outer side 404 of the web 102 and protrudes from the first predetermined thickness 402 of the web 102. In this example, method 2700 continues from 2702 to 2704, wherein a first mouse hole 110 is cut from flange 1802 at a first predetermined position, and its size is set to fit above a first stringer 302 on the skin 204 of fuselage 202. The first predetermined position is based on a first position of the first stringer 302 on skin 204. At 2706, a second mouse hole 112, spaced apart from the first mouse hole 110, is cut from flange 1802 at a second predetermined position, and the size of the second mouse hole is set to fit above a second stringer 304 adjacent to the first stringer 302 on skin 204. The second predetermined position is based on a second position of the second stringer 304 on skin 204.

[0097] In yet another example, the composite fuselage frame 100 is configured to provide lateral support 314 to the inner surface 312 of the skin 204 of a fuselage cylindrical section 310 having a plurality of stringers 305. The composite fuselage frame 100 spans the plurality of stringers 305. The composite fuselage frame 100 includes a plurality of shear-resistant connecting feet 114 extending toward the skin 204 between adjacent stringers 305. During the manufacture of the composite fuselage frame 100, each shear-resistant connecting foot 114 has frame filler 802 integrated in each shear-resistant connecting foot, such that the composite fuselage frame 100 includes a plurality of frame filler 802 integrated within the plurality of shear-resistant connecting feet 114.

[0098] Refer again Figure 1 , Figures 6 to 10 , Figure 15 and Figure 23 In one or more examples, the method of manufacturing the frame filler 802 for the composite fuselage frame 100 includes Figure 9 Method 900 Figure 23Method 2300, and continuing from 2302, involves integrating the frame filler 802 into the shear connection feet 114 of the composite fuselage frame 100 via a second stack 1506 forming a second material polymer matrix composite sheet 602 via a clamping arc above the first stack 1004 of the first material polymer matrix composite sheet 604. The frame filler 802 provided by the first stack 1004 and the shear connection feet 114 provided by the second stack 1506 are then co-cured to harden the composite fuselage frame 100.

[0099] General Reference Figures 1 to 8 , Figure 10 , Figure 12 , Figures 14 to 16 , Figure 18 , Figures 28 to 32 and Figures 39 to 42 As an example, this disclosure relates to methods 2800, 2900, 3000, 3100, 3200, and 3800 for manufacturing a composite fuselage frame 100 for an aircraft 200. Figure 1 This is a perspective view of an example of a composite fuselage frame 100. Figure 2 This is a perspective view of an example of aircraft 200. Figure 3 This is a perspective view of an example of a partially assembled fuselage section 310. Figure 4 This is a cross-sectional side view of another example of the composite fuselage frame 100. Figure 5 yes Figure 4 Another cross-sectional side view of the composite fuselage frame 100 shows the polymer matrix composite layers 502 and 504. Figure 6 This is an exploded front view showing an example of the polymer matrix composite layers 502, 504 and the outer protective glass fiber layer 506 in another example of the composite fuselage frame 100. Figure 7 It is shown Figure 6 An exploded side view of the polymer matrix composite layers 502, 504 and the outer protective glass fiber layer 506 in the composite fuselage frame 100.

[0100] Figure 8 yes Figure 4 and Figure 5 A side view of the composite fuselage frame 100 shows an example of the frame filler 802 (indicated by dashed lines) integrated within the composite fuselage frame 100. Figure 10 It is shown that it is used for Figure 4 and Figure 5 An exploded front view of an example of polymer matrix composite layers 502, 504 in an example of frame filler 802 of a composite fuselage frame 100. Figure 12 This is a top view of an example of a sheet of the first material polymer matrix composite fabric 1202. Figure 14 This is a rear view of an example of laying tool 1404. Figure 8The laminate 1102 of the frame filler 802 (dashed line) is placed on the laying tool 1404, and an example of the second material polymer matrix composite sheet 602 is covered on the laying tool 1404. Figure 15 yes Figure 14 An exploded side view of the laying tool 1404 shows an example of a laminate 1102 (shown in dashed lines) of the frame filler 802 and a second material polymer matrix composite sheet 602 covering the frame filler 802. Figure 16 This is a top view of an example of a sheet of the second material polymer matrix composite tape 1602. Figure 18 This is a side view of an example of the laminate 1502 of the composite fuselage frame 100, showing its integration within the composite fuselage frame 100. Figure 11 The laminate 1102 of the frame filler 802 in the middle.

[0101] Figure 28 An example of a method 2800 for manufacturing a composite fuselage frame 100 for an aircraft 200 is provided. Figure 29 Combination Figure 28 An example of a method 2900 for manufacturing a composite fuselage frame 100 for an aircraft 200 is provided. Figure 30 Combination Figure 28 An example of a method 3000 for manufacturing a composite fuselage frame 100 for an aircraft 200 is provided. Figure 31 Combination Figure 28 An example of a method 3100 for manufacturing a composite fuselage frame 100 for an aircraft 200 is provided. Figure 32 Combination Figure 28 An example of a method 3200 for manufacturing a composite fuselage frame 100 for an aircraft 200 is provided. Figure 38 Combination Figure 28 and Figure 30 An example of a method 3800 for manufacturing a composite fuselage frame for an aircraft is provided.

[0102] Figure 39 Provided in Figure 28 An example of forming a second stack 1506 and a first stack 1004 in method 2800. Figure 40 This is a side view of an example of a forming tool 4004 used for a composite fuselage frame 100. Figure 41 This is a side view of an example of a compaction tool 4104 used for the composite fuselage frame 100. Figure 42 Provided in Figure 28 Another example of forming a second stack 1506 and a first stack 1004 in method 2800.

[0103] Refer again Figure 1 , Figure 2 , Figure 6 , Figure 8 , Figure 10 , Figure 12 , Figures 14 to 16 , Figure 28 and Figures 39 to 42 In one or more examples, a method 2800 for manufacturing a composite fuselage frame 100 for an aircraft 200 (see...) Figure 28 This includes laying 2802 at predetermined locations multiple first-material polymer matrix composite sheets 604 on a first portion 1504 of the top surface 1402 of a laying tool 1404 for the composite fuselage frame 100 to form a first stack 1004 of the first-material polymer matrix composite sheets 604. The predetermined locations are associated with shear-resistant connecting feet 114 of the composite fuselage frame 100 and frame filler 802. At 2804, multiple second-material polymer matrix composite sheets 602 are laid on the first-material polymer matrix composite sheets 604 and a second portion 1510 of the top surface 1402 of the laying tool 1404 to provide a second stack 1506 of the second-material polymer matrix composite sheets 602. At 2806, the second stack 1506 and the first stack 1004 are formed.

[0104] In another example of method 2800, the first material polymer matrix composite sheet 604 comprises a thermosetting matrix composite sheet, a thermoplastic matrix composite sheet, or any other suitable woven polymer matrix composite sheet in any suitable combination. In yet another example of method 2800, the first material polymer matrix composite sheet 604 comprises a reinforcing fabric and a polymer matrix material. In another example, the reinforcing fabric comprises a plain weave fabric, a 0 / 90 degree woven fabric, a 45 degree woven fabric, or any other suitable reinforcing fabric in any suitable combination. In yet another example, the reinforcing fabric comprises woven reinforcing fibers interwoven in at least two directions. In a further example, the woven reinforcing fibers comprise glass fibers, carbon fibers, aramid fibers, natural fibers, polyetheretherketone fibers, or any other suitable reinforcing fibers in any suitable combination. In yet another further example, the polymer matrix material includes thermosetting resins, thermoplastic resins, epoxy resins, phenolic resins, polyurethane resins, polyimide resins, polyethylene resins, polypropylene resins, polybutylene terephthalate resins, polyamide resins, polyphenylene sulfide resins, polyetherimide resins, polyetherketone ketone resins, polyetheretherketone resins, or any other suitable polymer matrix material in any suitable combination.

[0105] In yet another example of method 2800, the second material polymer matrix composite sheet 602 comprises a thermosetting matrix composite sheet, a thermoplastic matrix composite sheet, or any other suitable unidirectional polymer matrix composite sheet in any suitable combination. In yet another example of method 2800, the second material polymer matrix composite sheet 602 comprises unidirectional reinforcing fibers and a polymer matrix material. In another example, the unidirectional reinforcing fibers comprise carbon fibers, glass fibers, aramid fibers, natural fibers, polyetheretherketone fibers, or any other suitable reinforcing fibers in any suitable combination. In yet another further example, the polymer matrix material is a thermosetting resin, thermoplastic resin, epoxy resin, phenolic resin, polyurethane resin, polyimide resin, polyethylene resin, polypropylene resin, polybutylene terephthalate resin, polyamide resin, polyphenylene sulfide resin, polyetherimide resin, polyetherketone ketone resin, polyetheretherketone resin, or any other suitable polymer matrix material.

[0106] In another example of method 2800, the composite fuselage frame 100 further includes a one-piece segment 116. In yet another example, method 2800 further includes cutting 2808 a plurality of first material polymer matrix composite sheets 604 from one or more sheets of a first material polymer matrix composite fabric 1202. In this example, method 2800 continues from 2808 to 2802. In yet another example, method 2800 continues from 2802 to 2810, wherein a plurality of second material polymer matrix composite sheets 602 are cut from one or more sheets of a second material polymer matrix composite strip 1602. In this example, method 2800 continues from 2810 to 2804. In yet another example of method 2800, forming 2806 the second stack 1506 and the first stack 1004 includes forming 3902 the second stack 1506 and the first stack 1004 onto the top surface 4002 of the forming tool 4004. At 3904, the second stack 1506 and the first stack 1004 are compacted onto the top surface 4102 of the compaction tool 4104. In another example of method 2800, forming the second stack 1506 and the first stack 1004 at 2806 includes forming the second stack 1506 and the first stack 1004 onto the top surface 1402 of the laying tool 1404 during formation 4202. At 4204, the second stack 1506 and the first stack 1004 are compacted to provide a laminate 1502 of the composite fuselage frame 100.

[0107] Refer again Figures 1 to 8 , Figure 10 , Figure 15 , Figure 18 , Figure 28 , Figure 29 and Figure 42In one or more examples, method 2900 for manufacturing composite fuselage frame 100 for aircraft 200 (see...) Figure 29 )include Figure 28 Method 2800 and from Figure 42 Continuing from 4204 to 2902, the laminate 1502 of the composite fuselage frame 100 is cured to harden the composite fuselage frame 100. In another example of method 2900, after the laminate 1502 is cured at 2902, the composite fuselage frame 100 includes a web 102 and a flange 1802. The web 102 includes a width 103 and a first predetermined thickness 402. The web 102 forms an arc 104 relative to the fuselage 202 of the aircraft 200 between a first end 106 and a second end 108. The width 103 extends from an outer side 404 to an inner side 406 relative to the skin 204 on the fuselage 202. The flange 1802 is disposed on the outer side 404 of the web 102 and protrudes from the first predetermined thickness 402 of the web 102. In this example, the method continues from 2902, and at 2904, a first mouse hole 110 is cut from the flange 1802 at a first predetermined position, and the size of the first mouse hole is set to fit above the first stringer 302 on the skin 204. The first predetermined position is based on a first position of the first stringer 302 on the skin 204. At 2906, a second mouse hole 112, spaced apart from the first mouse hole 110, is cut from the flange 1802 at a second predetermined position, and its size is set to fit above the second stringer 304 adjacent to the first stringer 302 on the skin 204. The second predetermined position is based on a second position of the second stringer 304 on the skin 204.

[0108] In yet another example, the composite fuselage frame 100 is configured to provide lateral support 314 to the inner surface 312 of the skin 204 of a fuselage cylindrical section 310 having a plurality of stringers 305. The composite fuselage frame 100 spans the plurality of stringers 305. The composite fuselage frame 100 includes a plurality of shear-resistant connecting feet 114 extending toward the skin 204 between adjacent stringers 305. During the manufacture of the composite fuselage frame 100, each shear-resistant connecting foot 114 has frame filler 802 integrated in each shear-resistant connecting foot, such that the composite fuselage frame 100 includes a plurality of frame filler 802 integrated within the plurality of shear-resistant connecting feet 114.

[0109] In another example, a shear-resistant connecting leg 114 extends from the web 102 between the first rat hole 110 and the second rat hole 112 to the distal end 1804 of the flange 1802, and the outer surface 410 of the shear-resistant connecting leg 114 is configured to face the skin 204 and align with a corresponding portion of the skin 204. In another example, the shear-resistant connecting leg 114 extends from a first predetermined thickness 402 of the web 102 such that one side of the shear-resistant connecting leg 114 is configured to face the skin 204 and fit within the area defined by the first stringer 302, the second stringer 304, and the corresponding portion of the skin 204 between the first stringer 302 and the second stringer 304. In yet another example, a plurality of composite fuselage frames 100 form a segmented frame 306 for a section of the fuselage 202 of the aircraft 200. In yet another further example, a portion of flange 1802 between the first rat hole 110 and the second rat hole 112 defines a shear-resistant connecting leg 114 comprising a second predetermined thickness 408. The first predetermined thickness 402 of web 102 is defined by a first plurality of first polymer matrix composite layers 502 formed by a plurality of second material polymer matrix composite sheets 602. The second predetermined thickness 408 of shear-resistant connecting leg 114 is defined by the first predetermined thickness 402 and a second plurality of second polymer matrix composite layers 504 formed by a plurality of first material polymer matrix composite sheets 604, which provide a frame filler 802 integrated with the plurality of second material polymer matrix composite sheets 602 of the first predetermined thickness 402.

[0110] Refer again Figure 1 , Figures 6 to 8 , Figure 10 , Figure 15 , Figure 28 In one or more examples, the method of manufacturing the frame filler 802 for the composite fuselage frame 100 includes Figure 28 Method 2800, and continuing from 2802, involves integrating the frame filler 802 into the shear connection feet 114 of the composite fuselage frame 100 via a second stack 1506 forming a second material polymer matrix composite sheet 602 via a clamping arc above the first stack 1004 of the first material polymer matrix composite sheet 604. The frame filler 802 provided by the first stack 1004 and the shear connection feet 114 provided by the second stack 1506 are then co-cured to harden the composite fuselage frame 100.

[0111] Refer again Figure 1 , Figure 6 , Figure 10 , Figures 14 to 16 , Figure 20 , Figure 28 and Figure 30In one or more examples, method 3000 for manufacturing frame filler 802 for composite fuselage frame 100 (see...) Figure 30 )include Figure 28 Method 2800. At 3002, before laying 2802 plurality of first material polymer matrix composite sheets 604 and laying 2804 plurality of second material polymer matrix composite sheets 602, at least one second material polymer matrix composite sheet 602 is laid on the top surface 1402 of the laying tool 1404 to provide a base 2002 of the composite fuselage frame 100 beneath the first stack 1004 of the first material polymer matrix composite sheets 604 and the second stack 1506 of the second material polymer matrix composite sheets 602. Method 3000 from Figure 28 Method 3000 continues from 3002 to 2802. From 2802, method 3000 continues to 3004, where, after laying multiple first material polymer matrix composite sheets 604 at 2802, a first stack 1004 of the base 2002 and the first material polymer matrix composite sheets 604 is formed onto the top surface 1402 of the laying tool 1404. At 3006, during the formation of 3004, the first stack 1004 of the base 2002 and the first material polymer matrix composite sheets 604 is compacted to provide a second laminate 2004 of the intermediate composite fuselage frame 2006. Method 3000 continues from Figure 28 Method 3006 continues to 2804. In another example, method 3000 further includes cutting 3008 at least one second material polymer matrix composite layer sheet 602 from one or more sheets of the second material polymer matrix composite tape 1602. In this example, method 3000 continues from 3008 to 3002.

[0112] Refer again Figure 1 , Figure 5 , Figure 6 , Figure 10 , Figure 14 , Figure 15 , Figure 20 , Figure 28 , Figure 30 and Figure 38 In one or more examples, method 3800 for manufacturing frame filler 802 for composite fuselage frame 100 (see...) Figure 38 )include Figure 28 Method 2800 and Figure 30 Method 3000. Method 3800 from Figure 30Continuing from 3002 to 3802, an outer protective glass fiber layer 606 is laid on the top surface 1402 of the laying tool 1404 before laying 2802 plurality of first material polymer matrix composite sheets 604 and 2804 plurality of second material polymer matrix composite sheets 602, to provide an outer protective glass fiber layer 506 beneath the base 2002 of the composite frame 100 and the second stack 1506 of the second material polymer matrix composite sheets 602. Method 3800 from Figure 28 Method 3800 continues from 2802 to 2802. From 2802, method 3800 continues to 3804, where, after laying multiple first material polymer matrix composite sheets 604 at 2802, a first stack 1004 of an outer protective glass fiber sheet 606, a base 2002, and the first material polymer matrix composite sheets 604 is formed onto the top surface 1402 of the laying tool 1404. At 3806, during the formation of 3804, the first stack 1004 of the outer protective glass fiber sheet 606, the base 2002, and the first material polymer matrix composite sheets 604 is compacted to provide a second laminate 2004 for the intermediate composite fuselage frame 2006. Method 3800 continues from Figure 28 3806 continues to 2804.

[0113] Refer again Figure 1 , Figure 5 , Figure 6 , Figure 10 , Figure 14 , Figure 15 , Figure 20 , Figure 28 and Figure 31 In one or more examples, a method 3100 for manufacturing frame filler 802 for composite fuselage frame 100 (see...) Figure 31 )include Figure 28 Method 2800. At 3102, before laying 2802 plurality of first material polymer matrix composite sheets 604 and laying 2804 plurality of second material polymer matrix composite sheets 602, an outer protective glass fiber sheet 606 is laid on the top surface 1402 of the laying tool 1404 to provide an outer protective glass fiber layer 506 beneath the first stack 1004 of the first material polymer matrix composite sheets 604 and the second stack 1506 of the second material polymer matrix composite sheets 602. Method 3100 from Figure 283102 continues to 2802 and 2804. At 3104, after laying multiple first material polymer matrix composite sheets 604 at 2802 and multiple second material polymer matrix composite sheets 602 at 2804, an outer protective glass fiber sheet 606 is formed with a second stack 1506 and a first stack 1004 onto the top surface 1402 of the laying tool 1404. At 3106, during the formation of 3104, the outer protective glass fiber sheet 606 is compacted with the second stack 1506 and the first stack 1004 to provide a laminate 1502 for the composite fuselage frame 100.

[0114] Refer again Figure 1 , Figure 5 , Figure 6 , Figure 10 , Figure 14 , Figure 15 , Figure 28 and Figure 32 In one or more examples, method 3200 for manufacturing frame filler 802 for composite fuselage frame 100 (see...) Figure 32 )include Figure 28 Method 2800 continues from 2804 to 3202, wherein an outer protective glass fiber layer 606 is laid on a second stack 1506 of the second material polymer matrix composite layer 602 and on a first stack 1004 of the first material polymer matrix composite layer 604 currently on the laying tool 1404 to provide an outer protective glass fiber layer 506 on top of the second stack 1506. At 3204, the outer protective glass fiber layer 606 is formed together with the second stack 1506 and the first stack 1004 onto the top surface 1402 of the laying tool 1404. At 3206, during the formation of 3204, the outer protective glass fiber layer 606 is compacted with the second stack 1506 and the first stack 1004 to provide a laminate 1502 of the composite fuselage frame 100.

[0115] General Reference Figures 1 to 8 , Figure 10 , Figure 14 , Figure 15 , Figure 18 , Figures 33 to 35 , Figure 40 , Figure 41 , Figure 51 and Figure 52 As an example, this disclosure relates to methods 3300 and 3500 for manufacturing a fuselage section 310 for an aircraft 200. Figure 1 This is a perspective view of an example of a composite fuselage frame 100. Figure 2 This is a perspective view of an example of aircraft 200. Figure 3This is a perspective view of an example of a partially assembled fuselage section 310. Figure 4 This is a cross-sectional side view of another example of the composite fuselage frame 100. Figure 5 yes Figure 4 Another cross-sectional side view of the composite fuselage frame 100 shows the polymer matrix composite layers 502 and 504. Figure 6 This is an exploded front view showing an example of the polymer matrix composite layers 502, 504 and the outer protective glass fiber layer 506 in another example of the composite fuselage frame 100. Figure 7 It is shown Figure 6 An exploded side view of the polymer matrix composite layers 502, 504 and the outer protective glass fiber layer 506 in the composite fuselage frame 100. Figure 8 yes Figure 4 and Figure 5 A side view of the composite fuselage frame 100 shows an example of the frame filler 802 (indicated by dashed lines) integrated within the composite fuselage frame 100.

[0116] Figure 10 It is shown that it is used for Figure 4 and Figure 5 An exploded front view of an example of polymer matrix composite layers 502, 504 in an example of frame filler 802 of a composite fuselage frame 100. Figure 14 This is a rear view of an example of laying tool 1404. Figure 8 The laminate 1102 of the frame filler 802 (dashed line) is placed on the laying tool 1404, and an example of the second material polymer matrix composite sheet 602 is covered on the laying tool 1404. Figure 15 yes Figure 14 An exploded side view of the laying tool 1404 shows an example of a laminate 1102 (shown in dashed lines) of the frame filler 802 and a second material polymer matrix composite sheet 602 covering the frame filler 802. Figure 18 This is a side view of an example of the laminate 1502 of the composite fuselage frame 100, showing its integration within the composite fuselage frame 100. Figure 11 The laminate 1102 of the frame filler 802 in the middle. Figure 33 and Figure 34 An example of a method 3300 for manufacturing a fuselage section 310 for an aircraft 200 is provided, which includes manufacturing a composite fuselage frame 100 3400. Figure 35 Combination Figure 33 and Figure 34 An example of a method 3500 for manufacturing a fuselage section 310 for an aircraft 200 is provided, which includes manufacturing a composite fuselage frame 100 3400.

[0117] Figure 40This is a side view of an example of a forming tool 4004 used for a composite fuselage frame 100. Figure 41 This is a side view of an example of a compaction tool 4104 used for the composite fuselage frame 100. Figure 51 Provided Figure 34 An example of the formation of the second stack 1506 and the first stack 1004 in the example. Figure 52 Provided Figure 34 Another example of the formation of the second stack 1506 and the first stack 1004 in the process.

[0118] Refer again Figures 1 to 3 , Figures 6 to 8 , Figure 10 , Figure 14 , Figure 15 , Figure 33 , Figure 34 , Figure 40 , Figure 41 , Figure 51 and Figure 52 In one or more examples, a method 3300 for manufacturing the fuselage section 310 for aircraft 200 (see...) Figure 33 The method includes receiving a skin 204 in a manufacturing unit, the skin 204 having at least a first stringer 302 and a second stringer 304 on its inner surface 312. The skin 204 having the first stringer 302 and the second stringer 304 has been pre-manufactured and hardened. At 3304, the skin 204 is mounted on a manufacturing tool in the manufacturing unit. At 3306, a composite fuselage frame 100 in which frame filler 802 is integrated is received in the manufacturing unit. The composite fuselage frame 100 has been previously manufactured and hardened. At 3308, the composite fuselage frame 100 is mounted on the inner surface 312 of the skin 204 above the first stringer 302 and the second stringer 304. The method 3300 also includes manufacturing 3400 (see...). Figure 34A composite fuselage frame 100 includes laying 3402 at predetermined locations a plurality of first material polymer matrix composite sheets 604 on a first portion 1504 of the top surface 1402 of a laying tool 1404 for the composite fuselage frame 100 to form a first stack 1004 of the first material polymer matrix composite sheets 604. The predetermined locations are associated with shear-resistant connecting feet 114 of the composite fuselage frame 100 and frame filler 802. At 3404, a plurality of second material polymer matrix composite sheets 602 are laid on a second portion 1510 of the first material polymer matrix composite sheets 604 and the top surface 1402 of the laying tool 1404 to provide a second stack 1506 of the second material polymer matrix composite sheets 602. At 3406, the second stack 1506 and the first stack 1004 are formed onto the top surface 1402 of the laying tool 1404. At 3408, during the formation of 3406, the second stack 1506 and the first stack 1004 are compacted to provide a laminate 1502 for the composite fuselage frame 100.

[0119] In another example of manufacturing 3400, forming 3406 the second stack 1506 and the first stack 1004 includes: forming 5102 of the second stack 1506 and the first stack 1004 onto the top surface 4002 of the forming tool 4004. At 5104, the second stack 1506 and the first stack 1004 are compacted onto the top surface 4102 of the compaction tool 4104. In yet another example of manufacturing 3400, forming 3406 the second stack 1506 and the first stack 1004 includes: forming 5202 of the second stack 1506 and the first stack 1004 onto the top surface 1402 of the laying tool 1404. At 5204, during the formation of 5202, the second stack 1506 and the first stack 1004 are compacted to provide a laminate 1502 of the composite fuselage frame 100.

[0120] Refer again Figures 1 to 8 , Figure 10 , Figure 15 , Figure 18 and Figures 33 to 35 In one or more examples, a method 3500 for manufacturing the fuselage section 310 for aircraft 200 (see...) Figure 35 )include Figure 33 Method 3300 and Figure 34 Manufacturing of the 3400 composite fuselage frame 100. Method 3500 from Figure 52 5204 continues to 3502, wherein manufacturing the 3400 composite fuselage frame 100 further includes: curing the 3502 composite fuselage frame 100 with a laminate 1502 to harden the composite fuselage frame 100.

[0121] In another example of method 3500, after the laminate 1502 has been cured 3502, the composite fuselage frame 100 includes a web 102 and a flange 1802. The web 102 includes a width 103 and a first predetermined thickness 402. The web forms an arc 104 between a first end 106 and a second end 108 of the fuselage 202 of the aircraft 200. The width 103 extends from an outer side 404 to an inner side 406 of the skin 204 on the fuselage 202. The flange 1802 is disposed on the outer side 404 of the web 102 and protrudes from the first predetermined thickness 402 of the web 102. In this example, manufacturing 3400 of the composite fuselage frame 100 further includes cutting 3504 a first rat hole 110 from the flange 1802 at a first predetermined location, and the size of the first rat hole 110 is set to fit above a first stringer 302 on the skin 204. The first predetermined position is based on a first position of the first stringer 302 on the skin 204. At 3506, a second rat hole 112, spaced apart from the first rat hole 110, is cut from the flange 1802 at a second predetermined position, and its size is set to fit above the second stringer 304 adjacent to the first stringer 302 on the skin 204. The second predetermined position is based on a second position of the second stringer 304 on the skin 204.

[0122] In yet another example, the composite fuselage frame 100 is configured to provide lateral support 314 to the inner surface 312 of the skin 204 of a fuselage cylindrical section 310 having a plurality of stringers 305. The composite fuselage frame 100 spans the plurality of stringers 305. The composite fuselage frame 100 includes a plurality of shear-resistant connecting feet 114 extending toward the skin 204 between adjacent stringers 305. Each shear-resistant connecting foot 114 has frame filler 802 integrated in each shear-resistant connecting foot during the manufacture of the composite fuselage frame 100, such that the composite fuselage frame 100 includes a plurality of frame filler 802 integrated within the plurality of shear-resistant connecting feet 114.

[0123] In another example, a portion of flange 1802 between the first rat hole 110 and the second rat hole 112 defines a shear-resistant connecting leg 114 including a second predetermined thickness 408. The first predetermined thickness 402 of web 102 is defined by a first plurality of first polymer matrix composite layers 502 formed by a plurality of second material polymer matrix composite sheets 602. The second predetermined thickness 408 of shear-resistant connecting leg 114 is defined by the first predetermined thickness 402 and a second plurality of second polymer matrix composite layers 504 formed by a plurality of first material polymer matrix composite sheets 604, which provide a frame filler 802 integrated with the plurality of second material polymer matrix composite sheets 602 of the first predetermined thickness 402.

[0124] In another further example, manufacturing the 3400 composite fuselage frame 100 also includes integrating 3508 frame filler 802 into shear connection feet 114 of the composite fuselage frame 100 via a second stack 1506 of a second material polymer matrix composite sheet 602 formed by clamping arcs over a first stack 1004 of a first material polymer matrix composite sheet 604. At 3510, the frame filler 802 provided by the first stack 1004 and the shear connection feet 114 provided by the second stack 1506 are co-cured to harden the composite fuselage frame 100.

[0125] Examples of composite fuselage frame 100, methods for manufacturing frame filler 802 for composite fuselage frame 100 (900, 1300, 1700, 1900, 2100, 2200, 2300, 2400, 2500, 2600, 2700), methods for manufacturing composite fuselage frame 100 for aircraft 200 (2800, 2900, 3000, 3100, 3200, 3800), and methods for manufacturing fuselage cylindrical section 310 for aircraft 200 (3300, 3500) may be relevant to or used in the context of aircraft manufacturing. While aircraft examples are described, the examples and principles disclosed herein can be applied to other products in the aerospace industry and other industries, such as the automotive industry, space industry, construction industry, and other design and manufacturing industries. Therefore, in addition to aircraft, the examples and principles disclosed herein can be applied to various products used in the manufacture of various types of vehicles and the construction of various types of buildings.

[0126] The foregoing detailed description refers to the accompanying drawings, which illustrate specific examples described in this disclosure. Other examples with different structures and operations do not depart from the scope of this disclosure. In different drawings, the same reference numerals may refer to the same features, elements, or components. Throughout this disclosure, any one of a plurality of items may be referred to individually as an item, and a plurality of items may be referred to collectively as an item and may be represented by the same reference numerals. Furthermore, as used herein, a feature, element, component, or step preceding the word "a" or "an" should be understood to not exclude multiple features, elements, components, or steps unless such exclusion is expressly stated.

[0127] The foregoing provides illustrative, non-exhaustive examples that may, but do not necessarily, claim protection for the subject matter according to this disclosure. Reference to “example” herein means that one or more features, structures, elements, components, characteristics, and / or operational steps described in connection with the example are included in at least one aspect, embodiment, and / or implementation of the subject matter according to this disclosure. Therefore, the phrases “example,” “another example,” “one or more examples,” and similar language throughout this disclosure may, but do not necessarily, refer to the same example. Furthermore, the subject matter characterizing any example may, but does not necessarily include the subject matter characterizing any other example. Moreover, the subject matter characterizing any example may, but does not necessarily, combine with the subject matter characterizing any other example.

[0128] As used herein, a system, apparatus, device, structure, article, element, component, or hardware "constructed" to perform a specified function is indeed capable of performing the specified function without any changes, and not merely has the potential to perform the specified function after further modification. In other words, a system, apparatus, device, structure, article, element, component, or hardware "constructed" to perform a specified function is specifically selected, created, implemented, utilized, programmed, and / or designed for performing the specified function. As used herein, "constructed" means the existing characteristics of the system, apparatus, structure, article, element, component, or hardware that enable the system, apparatus, structure, article, element, component, or hardware to perform the specified function without further modification. For the purposes of this disclosure, a system, apparatus, device, structure, article, element, component, or hardware described as "constructed" to perform a particular function may additionally or alternatively be described as "suitable" and / or "operable to" perform that function.

[0129] Unless otherwise stated, the terms “first,” “second,” “third,” etc., are used merely as labels in this document and are not intended to impose any order, position, or hierarchy on the items referred to by these terms. Furthermore, references to an item such as “second” do not require or exclude the existence of an item such as “first” or a lower-numbered item and / or an item such as “third” or a higher-numbered item.

[0130] As used herein, when used with a list of items, the phrase “at least one of” means that different combinations of one or more of the listed items may be used, and it may be necessary to use only one of each item in the list. For example, “at least one of items A, B, and C” may include, but is not limited to, item A or items A and B. This example may also include items A, B, and C, or items B and C. In other examples, “at least one” may be, for example, but not limited to, two of items A, one of items B, and ten of items C; four items B, seven items C; and other suitable combinations. As used herein, the terms “and / or” and the “ / ” symbol include any and all combinations of one or more associated listed items.

[0131] As used herein, the terms "connection," "link," and similar terms refer to two or more elements that are joined, linked, fastened, attached, connected, communicated, or otherwise associated with each other (e.g., mechanically, electrically, fluidly, optically, electromagnetically). In various examples, the elements may be associated directly or indirectly. As an example, element A may be directly associated with element B. As another example, element A may be indirectly associated with element B, for example, via another element C. It should be understood that not all associations between the various disclosed elements are necessarily represented. Therefore, connectors other than those depicted in the figures may also exist.

[0132] As used herein, the term "approximately" means or indicates a condition that is close to but not exactly close to the stated condition, which still performs the desired function or achieves the desired result. As an example, the term "approximately" means a condition within an acceptable predetermined tolerance or accuracy, such as a condition within 10% of the stated condition. However, the term "approximately" does not exclude a condition that is exactly the stated condition. As used herein, the term "substantially" means a condition that substantially performs the desired function or achieves the desired result.

[0133] The above-mentioned Figures 1 to 8 , Figures 10 to 12 , Figures 14 to 16 , Figure 18 and Figure 20 In this context, "indicated" may refer to a functional element, feature, or component thereof, and does not necessarily imply any specific structure. Therefore, the illustrated structure may be modified, added to, and / or omitted. Furthermore, those skilled in the art will understand that it is not limited to the structures mentioned above. Figures 1 to 8 , Figures 10 to 12 , Figures 14 to 16 , Figure 18 and Figure 20 All elements, features, and / or components described and illustrated herein need to be included in each example, and not all elements, features, and / or components described herein need to be depicted in every illustrative example. Therefore, in Figures 1 to 8 , Figures 10 to 12 , Figures 14 to 16 , Figure 18 and Figure 20 Some of the elements, features, and / or components described and shown herein can be combined in various ways without needing to be included. Figures 1 to 8 , Figures 10 to 12 , Figures 14 to 16 , Figure 18 and Figure 20 Other features described and illustrated in the accompanying drawings and / or disclosures, even if one or more such combinations are not expressly shown herein. Similarly, additional features, not limited to the examples presented, may be combined with some or all of the features shown and described herein. Unless otherwise expressly stated, the features mentioned above are not limited to those in the examples presented herein. Figures 1 to 8 , Figures 10 to 12 , Figures 14 to 16 , Figure 18 and Figure 20 The schematic diagrams depicted are not intended to imply structural limitations regarding the exemplary examples. Rather, while an exemplary structure is shown, it should be understood that this structure can be modified as appropriate. Therefore, modifications, additions, and / or omissions can be made to the illustrated structure. Furthermore, in Figures 1 to 8 , Figures 10 to 12 , Figures 14 to 16 , Figure 18 and Figure 20 In each of these, elements, features, and / or components used for similar or at least substantially similar purposes are designated with the same numerical designations and may be omitted from reference herein. Figures 1 to 8 , Figures 10 to 12 , Figures 14 to 16 , Figure 18 and Figure 20 Each of these elements, features, and / or components is discussed in detail. Similarly, in Figures 1 to 8 , Figures 10 to 12 , Figures 14 to 16 , Figure 18 and Figure 20 Each element, feature, and / or component may not be labeled, but for consistency, the associated reference numerals may be used herein.

[0134] The above-mentioned Figure 9 , Figure 13 , Figure 17 , Figure 19 and Figures 21 to 35 In this document, boxes may represent operations, steps, and / or parts thereof, and the lines connecting the various boxes do not imply any particular order or dependency between the operations or their parts. It should be understood that this does not necessarily represent all dependencies between the various operations disclosed. Figure 9 , Figure 13 , Figure 17 , Figure 19 and Figures 21 to 35 The accompanying disclosures describing the operations of the methods set forth herein should not be construed as requiring a specific order in which the operations are performed. Rather, while an illustrative order is indicated, it should be understood that the order of operations can be modified where appropriate. Therefore, the operations shown can be modified, added to, and / or omitted, and some operations can be performed in a different order or simultaneously. Furthermore, those skilled in the art will understand that not all of the described operations need to be performed.

[0135] Furthermore, references to features, advantages, or similar language used throughout this specification do not imply that all features and advantages that can be implemented using the examples disclosed herein should be or are not present in any single example. Rather, references to features and advantages are understood to mean that a particular feature, advantage, or characteristic described in connection with an example is included in at least one example. Therefore, discussions of features, advantages, and similar language used throughout this disclosure may, but do not necessarily, refer to the same examples.

[0136] It can be like Figure 36 The aircraft manufacturing and maintenance methods shown in 3600 and such Figure 37 Examples of the subject matter disclosed herein are described in the context of the aircraft 3700 shown. In one or more examples, the disclosed composite fuselage frame 100, methods 900, 1300, 1700, 1900, 2100, 2200, 2300, 2400, 2500, 2600, 2700 for manufacturing frame filler 802 for the composite fuselage frame 100 for the aircraft 200, methods 2800, 2900, 3000, 3100, 3200, 3800 for manufacturing the composite fuselage frame 100 for the aircraft 200, and methods 3300, 3500 for manufacturing the fuselage cylindrical section 310 for the aircraft 200 can be used in aircraft manufacturing. During pre-production, maintenance method 3600 may include the specifications and design of the aircraft 3700 (box 3602) and material procurement (box 3604). During production, the manufacturing of components and sub-components of the aircraft 3700 (box 3606) and system integration (box 3608) may occur. Subsequently, the aircraft 3700 may undergo certification and delivery (box 3610) for service (box 3612). During service, the aircraft 3700 may be scheduled for routine maintenance and repair (box 3614). Routine maintenance and repair may include modification, refactoring, refurbishment, etc., of one or more systems of the aircraft 3700.

[0137] Each procedure of maintenance method 3600 may be performed or carried out by a system integrator, a third party, and / or an operator (e.g., a customer). For the purposes of this description, a system integrator may include, but is not limited to, any number of aircraft manufacturers and major system subcontractors; a third party may include, but is not limited to, any number of vendors, subcontractors, and suppliers; and an operator may be an airline, leasing company, military entity, service organization, etc.

[0138] like Figure 37 As shown, an aircraft 3700 produced by maintenance method 3600 may include a fuselage 3702 having multiple advanced systems 3704 and an interior 3706. Examples of advanced systems 3704 include one or more of a propulsion system 3708, an electrical system 3710, a hydraulic system 3712, and an environmental system 3714. Any number of other systems may be included. Although an aerospace example is shown, the principles disclosed herein can be applied to other industries, such as the automotive industry. Therefore, in addition to aircraft 3700, the principles disclosed herein can be applied to other vehicles, such as land vehicles, marine vehicles, space vehicles, etc.

[0139] The disclosed composite fuselage frame 100, methods 900, 1300, 1700, 1900, 2100, 2200, 2300, 2400, 2500, 2600, 2700 for manufacturing frame filler 802 for the composite fuselage frame 100, methods 2800, 2900, 3000, 3100, 3200, 3800 for manufacturing the composite fuselage frame 100 for the aircraft 200, and methods 3300, 3500 for manufacturing the fuselage cylindrical section 310 for the aircraft 200 can be employed during any or more stages of the manufacturing and maintenance method 3600. For example, a component or sub-assembly (box 3606) corresponding to the manufacturing of parts and sub-assemblies can be manufactured or produced in a manner similar to that of a component or sub-assembly (box 3612) produced when the aircraft 3700 enters service. Furthermore, one or more examples of the system, method, or combination thereof may be utilized during the production phase (boxes 3606 and 3608), for example, by substantially accelerating the assembly of aircraft 3700 or reducing the cost of aircraft 3700. Similarly, one or more examples of the system or method implementation, or combinations thereof, may be utilized, for example but not limited to, during the use of aircraft 3700 (box 3612) and / or during maintenance and repair (box 3614).

[0140] The features, advantages, and characteristics described in one example can be combined in any suitable manner in one or more other examples. Those skilled in the art will recognize that the examples described herein can be practiced without the presence of one or more specific features or advantages in a particular example. In other cases, additional features and advantages that may not be present in all examples can be recognized in some examples. Furthermore, various examples of the composite fuselage frame 100, methods 900, 1300, 1700, 1900, 2100, 2200, 2300, 2400, 2500, 2600, 2700 for manufacturing the frame filler 802 for the composite fuselage frame 100, methods 2800, 2900, 3000, 3100, 3200, 3800 for manufacturing the composite fuselage frame 100 for the aircraft 200, and methods 3300, 3500 for manufacturing the fuselage cylindrical section 310 for the aircraft 200, can be modified by those skilled in the art upon reading the specification. This application includes such modifications and is limited only by the scope of the claims.

[0141] This application involves the following terms:

[0142] 1. A method for manufacturing a composite fuselage frame for an aircraft, the method comprising:

[0143] Multiple first material polymer matrix composite sheets are laid at predetermined locations on a first portion of the top surface of a laying tool for the composite fuselage frame to form a first stack of first material polymer matrix composite sheets, the predetermined locations being associated with shear connection feet and frame fillers of the composite fuselage frame.

[0144] Multiple second-material polymer matrix composite sheets are laid on a second portion of the top surface of the first-material polymer matrix composite sheet and the laying tool to provide a second stack of second-material polymer matrix composite sheets; and

[0145] The second stack and the first stack are formed.

[0146] 2. The method according to Clause 1, wherein the first material polymer matrix composite sheet comprises at least one of a thermosetting matrix composite sheet and a thermoplastic matrix composite sheet.

[0147] 3. The method according to Clause 1, wherein the first material polymer matrix composite sheet comprises a reinforcing fabric and a polymer matrix material.

[0148] 4. The method according to Clause 3, wherein the reinforcing fabric comprises at least one of plain weave fabric, 0 / 90 degree woven fabric and 45 degree woven fabric.

[0149] 5. The method according to Clause 3, wherein the reinforcing fabric comprises woven reinforcing fibers interwoven in at least two directions.

[0150] 6. The method according to Clause 5, wherein the braided reinforcing fiber comprises at least one of glass fiber, carbon fiber, aramid fiber, natural fiber and polyetheretherketone fiber.

[0151] 7. The method according to Clause 3, wherein the polymer matrix material comprises at least one of thermosetting resin, thermoplastic resin, epoxy resin, phenolic resin, polyurethane resin, polyimide resin, polyethylene resin, polypropylene resin, polybutylene terephthalate resin, polyamide resin, polyphenylene sulfide resin, polyetherimide resin, polyetherketone ketone resin, and polyetheretherketone resin.

[0152] 8. The method according to Clause 1, wherein the second material polymer matrix composite sheet comprises at least one of a thermosetting matrix composite sheet and a thermoplastic matrix composite sheet.

[0153] 9. The method according to Clause 1, wherein the second material polymer matrix composite sheet comprises unidirectional reinforcing fibers and a polymer matrix material.

[0154] 10. The method according to Clause 9, wherein the unidirectional reinforcing fiber comprises at least one of carbon fiber, glass fiber, aramid fiber, natural fiber and polyetheretherketone fiber.

[0155] 11. The method according to Clause 9, wherein the polymer matrix material comprises at least one of thermosetting resin, thermoplastic resin, epoxy resin, phenolic resin, polyurethane resin, polyimide resin, polyethylene resin, polypropylene resin, polybutylene terephthalate resin, polyamide resin, polyphenylene sulfide resin, polyetherimide resin, polyetherketone ketone resin, and polyetheretherketone resin.

[0156] 12. The method according to Clause 1, wherein the composite fuselage frame further comprises a single-piece section.

[0157] 13. The method according to Clause 1, further comprising:

[0158] The plurality of first material polymer matrix composite layers are cut from one or more sheets of the first material polymer matrix composite fabric.

[0159] 14. The method according to Clause 1, further comprising:

[0160] The plurality of second material polymer matrix composite layers are cut from one or more sheets of the second material polymer matrix composite tape.

[0161] 15. The method according to Clause 1, forming the second stack and the first stack comprises:

[0162] The second stack and the first stack are formed onto the top surface of the forming tool; and

[0163] The second and first stacks are compacted onto the top surface of the compaction tool.

[0164] 16. The method according to Clause 1, forming the second stack and the first stack comprises:

[0165] The second stack and the first stack are formed onto the top surface of the laying tool; and

[0166] During the formation process, the second and first stacks are compacted to provide a laminate for the composite fuselage frame.

[0167] 17. The method according to Clause 16, further comprising:

[0168] The laminate of the composite fuselage frame is cured to harden the composite fuselage frame.

[0169] 18. The method according to Clause 17, wherein, after the laminate is cured, the composite fuselage frame includes a web and a flange, the web including a width and a first predetermined thickness and forming an arc between a first end and a second end relative to the fuselage of the aircraft, the width extending from an outer side to an inner side relative to the skin of the fuselage, the flange disposed on the outer side of the web and projecting from the first predetermined thickness of the web, the method further comprising:

[0170] A first mouse hole is cut from the flange at a first predetermined position, and the size of the first mouse hole is set to fit above a first stringer on the skin, the first predetermined position being based on a first position of the first stringer on the skin; and

[0171] A second mouse hole, spaced apart from the first mouse hole, is cut from the flange at a second predetermined position, and the size of the second mouse hole is set to fit above a second stringer adjacent to the first stringer on the skin, the second predetermined position being based on a second position of the second stringer on the skin.

[0172] 19. The method according to Clause 18, wherein the composite fuselage frame is configured to provide lateral support to the inner surface of the skin of a fuselage cylindrical section having a plurality of stringers, the composite fuselage frame spanning the plurality of stringers, the composite fuselage frame including a plurality of shear connecting feet extending toward the skin between adjacent stringers, each shear connecting foot having frame filler integrated in each shear connecting foot during the manufacture of the composite fuselage frame, such that the composite fuselage frame includes a plurality of frame fillers integrated within the plurality of shear connecting feet.

[0173] 20. The method according to Clause 18, wherein the shear-resistant connecting foot extends from the web between the first and second rat holes to the distal end of the flange, and the outer surface of the shear-resistant connecting foot is configured to face the skin and align with a corresponding portion of the skin.

[0174] 21. The method according to Clause 18, wherein the shear connecting leg extends from the first predetermined thickness of the web such that one side of the shear connecting leg is configured to face the skin and fit within the area defined by the first stringer, the second stringer, and the corresponding portion of the skin between the first stringer and the second stringer.

[0175] 22. The method according to Clause 18, wherein a plurality of composite fuselage frames form a segmented frame for a section of the fuselage of the aircraft.

[0176] 23. The method according to Clause 18, wherein a portion of the flange between the first rat hole and the second rat hole defines the shear-resistant connecting leg, the shear-resistant connecting leg including a second predetermined thickness, wherein the first predetermined thickness of the web is defined by a first plurality of first polymer matrix composite layers formed by the plurality of second material polymer matrix composite sheets, wherein the second predetermined thickness of the shear-resistant connecting leg is defined by the first predetermined thickness and a second plurality of second polymer matrix composite layers formed by the plurality of first material polymer matrix composite sheets, the plurality of first material polymer matrix composite sheets providing the frame filler integrated with the plurality of second material polymer matrix composite sheets of the first predetermined thickness.

[0177] 24. The method according to Clause 1, further comprising:

[0178] The frame filler is integrated into the shear-resistant connecting leg of the composite fuselage frame by forming the second stack of the second material polymer matrix composite sheet via a clamping arc above the first stack of the first material polymer matrix composite sheet; and

[0179] The frame filler provided by the first stack and the shear-resistant connecting feet provided by the second stack are co-cured to harden the composite fuselage frame.

[0180] 25. The method according to Clause 1, further comprising:

[0181] Before laying the plurality of first material polymer matrix composite sheets and laying the plurality of second material polymer matrix composite sheets, at least one second material polymer matrix composite sheet is laid on the top surface of the laying tool to provide the base of the composite fuselage frame under the first stack of the first material polymer matrix composite sheets and the second stack of the second material polymer matrix composite sheets;

[0182] After the plurality of first material polymer matrix composite sheets are laid, the base and the first stack of the first material polymer matrix composite sheets are formed onto the top surface of the laying tool; and

[0183] During the formation process, the first stack of the base and the first material polymer matrix composite sheet is compacted to provide a second laminate of the intermediate composite fuselage frame.

[0184] 26. The method according to Clause 25, further comprising:

[0185] The at least one second material polymer matrix composite layer is cut from one or more sheets of the second material polymer matrix composite tape.

[0186] 27. The method according to Clause 25, further comprising:

[0187] Before laying the plurality of first material polymer matrix composite sheets and the plurality of second material polymer matrix composite sheets, an external protective glass fiber sheet is laid on the top surface of the laying tool to provide an external protective glass fiber layer under the base of the composite frame and the second stack of the second material polymer matrix composite sheets.

[0188] After the plurality of first material polymer matrix composite sheets are laid, the outer protective glass fiber sheet, the base, and the first layer of the first material polymer matrix composite sheets are formed onto the top surface of the laying tool; and

[0189] During the formation process, the first stack of the outer protective glass fiber sheet, the base, and the first material polymer matrix composite sheet is compacted to provide the second laminate of the intermediate composite fuselage frame.

[0190] 28. The method according to Clause 1, further comprising:

[0191] Before laying the plurality of first material polymer matrix composite sheets and the plurality of second material polymer matrix composite sheets, an external protective glass fiber sheet is laid on the top surface of the laying tool to provide an external protective glass fiber layer under the first stack of the first material polymer matrix composite sheets and the second stack of the second material polymer matrix composite sheets.

[0192] After laying the plurality of first material polymer matrix composite sheets and the plurality of second material polymer matrix composite sheets, the outer protective glass fiber sheet is formed together with the second stack and the first stack onto the top surface of the laying tool; and

[0193] During the formation process, the outer protective glass fiber sheet is compacted together with the second and first stacks to provide a laminate for the composite fuselage frame.

[0194] 29. The method according to Clause 1, further comprising:

[0195] An external protective glass fiber layer is laid on top of the second stack of the second material polymer matrix composite layer and the first stack of the first material polymer matrix composite layer to provide an external protective glass fiber layer on top of the second stack.

[0196] The outer protective glass fiber sheet is formed together with the second and first stacks onto the top surface of the laying tool; and

[0197] During the formation process, the outer protective glass fiber sheet is compacted together with the second and first stacks to provide a laminate for the composite fuselage frame.

[0198] 30. A method for manufacturing frame filler for a composite fuselage frame, the method comprising:

[0199] Multiple first material polymer matrix composite sheets are laid on a first laying tool for the frame filler to provide a first stack of first material polymer matrix composite sheets on a first top surface of the first laying tool; and

[0200] The first stack is formed.

[0201] 31. The method according to Clause 30, wherein the first material polymer matrix composite sheet comprises at least one of a thermosetting matrix composite sheet and a thermoplastic matrix composite sheet.

[0202] 32. The method according to Clause 30, wherein the first material polymer matrix composite sheet comprises a reinforcing fabric and a polymer matrix material.

[0203] 33. The method according to Clause 32, wherein the reinforcing fabric comprises at least one of plain weave fabric, 0 / 90 degree woven fabric and 45 degree woven fabric.

[0204] 34. The method according to Clause 32, wherein the reinforcing fabric comprises woven reinforcing fibers interwoven in at least two directions.

[0205] 35. The method according to Clause 34, wherein the braided reinforcing fiber comprises at least one selected from glass fiber, carbon fiber, aramid fiber, natural fiber, and polyetheretherketone fiber.

[0206] 36. The method according to Clause 32, wherein the polymer matrix material comprises at least one of thermosetting resin, thermoplastic resin, epoxy resin, phenolic resin, polyurethane resin, polyimide resin, polyethylene resin, polypropylene resin, polybutylene terephthalate resin, polyamide resin, polyphenylene sulfide resin, polyetherimide resin, polyetherketone ketone resin, and polyetheretherketone resin.

[0207] 37. The method according to Clause 30, wherein the composite fuselage frame further comprises a single-piece section.

[0208] 38. The method according to Clause 30, further comprising:

[0209] The plurality of first material polymer matrix composite layers are cut from one or more sheets of the first material polymer matrix composite fabric.

[0210] 39. The method according to Clause 30, forming the first stack comprises:

[0211] The first stack is formed onto the top surface of the forming tool; and

[0212] The first layer is compacted onto the top surface of the compaction tool.

[0213] 40. The method according to Clause 30, forming the first stack comprises:

[0214] The first layer is formed onto the first top surface of the first laying tool; and

[0215] During the formation process, the first stack is compacted to provide a first laminate of the frame filler.

[0216] 41. The method according to clause 40, further comprising:

[0217] Remove the first laminate of the frame filler from the first laying tool;

[0218] The first laminate is placed at a predetermined position on a first portion of the second top surface of a second laying tool for the composite fuselage frame, the predetermined position being associated with the shear connection foot of the composite fuselage frame and the frame filler.

[0219] Multiple second-material polymer matrix composite sheets are laid on a second portion of the second top surface of the first laminator and the second laying tool to provide a second stack of second-material polymer matrix composite sheets; and

[0220] The second stack and the first laminate are formed.

[0221] 42. The method according to Clause 41, wherein the second material polymer matrix composite sheet comprises at least one of a thermosetting matrix composite sheet and a thermoplastic matrix composite sheet.

[0222] 43. The method according to Clause 41, wherein the second material polymer matrix composite sheet comprises unidirectional reinforcing fibers and a polymer matrix material.

[0223] 44. The method according to Clause 43, wherein the unidirectional reinforcing fiber comprises at least one of carbon fiber, glass fiber, aramid fiber, natural fiber and polyetheretherketone fiber.

[0224] 45. The method according to Clause 43, wherein the polymer matrix material comprises at least one of thermosetting resin, thermoplastic resin, epoxy resin, phenolic resin, polyurethane resin, polyimide resin, polyethylene resin, polypropylene resin, polybutylene terephthalate resin, polyamide resin, polyphenylene sulfide resin, polyetherimide resin, polyetherketone ketone resin, and polyetheretherketone resin.

[0225] 46. ​​The method according to clause 41, further comprising:

[0226] The plurality of second material polymer matrix composite layers are cut from one or more sheets of the second material polymer matrix composite tape.

[0227] 47. The method according to Clause 41, forming the second stack and the first stack comprises:

[0228] The second stack and the first laminate are formed onto the top surface of the forming tool; and

[0229] The second stack and the first lamination are compacted onto the top surface of the compaction tool.

[0230] 48. The method according to Clause 41, forming the second stack and the first stack comprises:

[0231] The second laminate and the first laminator are formed onto the second top surface of the second laying tool; and

[0232] During the formation process, the second stack and the first laminate are compacted to provide the second laminate of the composite fuselage frame.

[0233] 49. The method according to clause 48, further comprising:

[0234] The second laminate of the composite fuselage frame is cured to harden the composite fuselage frame.

[0235] 50. The method according to Clause 49, wherein, after the second laminate has been cured, the composite fuselage frame includes a web and a flange, the web including a width and a first predetermined thickness and forming an arc between a first end and a second end relative to the fuselage of the aircraft, the width extending from an outer side to an inner side relative to the skin of the fuselage, the flange disposed on an outer side of the web and projecting from the first predetermined thickness of the web, the method further comprising:

[0236] A first mouse hole is cut from the flange at a first predetermined position, the size of the first mouse hole being set to fit above a first stringer on the skin of the fuselage, the first predetermined position being based on a first position of the first stringer on the skin; and

[0237] A second mouse hole, spaced apart from the first mouse hole, is cut from the flange at a second predetermined position, and the size of the second mouse hole is set to fit above a second stringer adjacent to the first stringer on the skin, the second predetermined position being based on a second position of the second stringer on the skin.

[0238] 51. The method according to Clause 50, wherein the composite fuselage frame is configured to provide lateral support to the inner surface of the skin of a fuselage cylindrical section having a plurality of stringers, the composite fuselage frame spanning the plurality of stringers, the composite fuselage frame including a plurality of shear connecting feet extending toward the skin between adjacent stringers, each shear connecting foot having frame filler integrated in each shear connecting foot during the manufacture of the composite fuselage frame, such that the composite fuselage frame includes a plurality of frame fillers integrated within the plurality of shear connecting feet.

[0239] 52. The method according to Clause 50, wherein the shear-resistant connecting foot extends from the web between the first rat hole and the second rat hole to the distal end of the flange, and the outer surface of the shear-resistant connecting foot is configured to face the skin and align with a corresponding portion of the skin.

[0240] 53. The method according to Clause 50, wherein the shear connecting leg extends from the first predetermined thickness of the web such that one side of the shear connecting leg is configured to face the skin and fit within the area defined by the first stringer, the second stringer, and the corresponding portion of the skin between the first stringer and the second stringer.

[0241] 54. The method according to Clause 50, wherein a plurality of composite fuselage frames form a segmented frame for a section of the fuselage of the aircraft.

[0242] 55. The method according to Clause 50, wherein a portion of the flange between the first rat hole and the second rat hole defines the shear-resistant connecting leg, the shear-resistant connecting leg including a second predetermined thickness, wherein the first predetermined thickness of the web is defined by a first plurality of first polymer matrix composite layers formed by the plurality of second material polymer matrix composite sheets, wherein the second predetermined thickness of the shear-resistant connecting leg is defined by the first predetermined thickness and a second plurality of second polymer matrix composite layers formed by the plurality of first material polymer matrix composite sheets, the plurality of first material polymer matrix composite sheets providing the frame filler integrated with the plurality of second material polymer matrix composite sheets of the first predetermined thickness.

[0243] 56. The method according to Clause 41, further comprising:

[0244] The frame filler is integrated into the shear-resistant connecting leg of the composite fuselage frame by forming the second stack of the second material polymer matrix composite sheet via a clamping arc above the first stack of the first material polymer matrix composite sheet; and

[0245] The frame filler provided by the first stack and the shear-resistant connecting feet provided by the second stack are co-cured to harden the composite fuselage frame.

[0246] 57. The method according to Clause 41, further comprising:

[0247] Before placing the first laminate and laying the plurality of second material polymer matrix composite sheets, at least one second material polymer matrix composite sheet is laid on the second top surface of the second laying tool to provide the base of the composite fuselage frame under the first laminate of the frame filler;

[0248] After the first laminate is placed, the base and the first laminate are formed onto the second top surface of the second laying tool; and

[0249] During the formation process, the base and the first laminate are compacted to provide a second laminate for the intermediate composite fuselage frame.

[0250] 58. The method according to Clause 57, further comprising:

[0251] The at least one second material polymer matrix composite layer is cut from one or more sheets of the second material polymer matrix composite tape.

[0252] 59. The method according to Clause 41, further comprising:

[0253] Before placing the first laminate and laying the plurality of second material polymer matrix composite sheets, an external protective glass fiber sheet is laid on the second top surface of the second laying tool to provide an external protective glass fiber layer under the first laminate of the frame filler.

[0254] After the first laminate is placed and the plurality of second material polymer matrix composite sheets are laid, the outer protective glass fiber sheet and the second stack of the second material polymer matrix composite sheets and the first laminate are formed together on the second top surface of the second laying tool; and

[0255] During the formation process, the outer protective glass fiber sheet is compacted together with the second laminate and the first laminate to provide the second laminate of the composite fuselage frame.

[0256] 60. The method according to Clause 41, further comprising:

[0257] An external protective glass fiber layer is laid on the second stack of the second material polymer matrix composite layer and the first laminate of the frame filler currently on the second layup tool to provide an external protective glass fiber layer on top of the second stack.

[0258] The outer protective glass fiber sheet is formed together with the second laminate and the first laminator onto the second top surface of the second laying tool; and

[0259] During the formation process, the outer protective glass fiber sheet is compacted together with the second laminate and the first laminate to provide the second laminate of the composite fuselage frame.

[0260] 61. The method according to Clause 40, further comprising:

[0261] Remove the first laminate of the frame filler from the first laying tool;

[0262] At least one second material polymer matrix composite sheet is laid on the second top surface of a second laying tool for the composite fuselage frame to provide the base of the composite fuselage frame;

[0263] The first laminate is placed on the base at a predetermined position, the predetermined position being associated with the shear-resistant connecting foot of the composite fuselage frame and the frame filler; and

[0264] The base and the first laminate are formed.

[0265] 62. The method according to Clause 61, further comprising:

[0266] The at least one second material polymer matrix composite layer is cut from one or more sheets of the second material polymer matrix composite tape.

[0267] 63. The method according to Clause 61, forming the base and the first laminate comprises:

[0268] The base and the first laminate are formed onto the top surface of the forming tool; and

[0269] The base and the first laminate are compacted onto the top surface of the compaction tool.

[0270] 64. The method according to Clause 61, forming the base and the first laminate comprises:

[0271] The base and the first laminate are formed onto the second top surface of the second laying tool; and

[0272] During the formation process, the base and the first laminate are compacted to provide a second laminate for the intermediate composite fuselage frame.

[0273] 65. The method according to Clause 61, further comprising:

[0274] An external protective fiberglass layer is laid on the second top surface of the second laying tool to provide an external protective fiberglass layer below the base of the composite frame.

[0275] The outer protective glass fiber layer is formed together with the base and the first laminate onto the second top surface of the second laying tool; and

[0276] During the formation process, the outer protective glass fiber sheet is compacted together with the base and the first laminate to provide a second laminate for the intermediate composite fuselage frame.

[0277] 66. The method according to Clause 64, further comprising:

[0278] Multiple second material polymer matrix composite sheets are laid on the second laminator of the intermediate composite frame currently located on the second laying tool to provide a second stack of second material polymer matrix composite sheets on the second laminator;

[0279] The second layer is formed onto the third top surface of the second laminate, wherein the first laminate containing the frame filler is integrated into the third top surface of the second laminate; and

[0280] During the formation process, the second lamination of the second stack and the second lamination of the intermediate composite fuselage frame are compacted to provide the third lamination of the composite fuselage frame.

[0281] 67. The method according to clause 66, further comprising:

[0282] The plurality of second material polymer matrix composite layers are cut from one or more sheets of the second material polymer matrix composite tape.

[0283] 68. The method according to Clause 66, further comprising:

[0284] The third laminate of the composite fuselage frame is cured to harden the composite fuselage frame.

[0285] 69. The method according to Clause 68, wherein, after the third laminate is cured, the composite fuselage frame includes a web and a flange, the web including a width and a first predetermined thickness and forming an arc between a first end and a second end relative to the fuselage of the aircraft, the width extending from an outer side to an inner side relative to the skin of the fuselage, the flange disposed on an outer side of the web and projecting from the first predetermined thickness of the web, the method further comprising:

[0286] A first mouse hole is cut from the flange at a first predetermined position, the size of the first mouse hole being set to fit above a first stringer on the skin of the fuselage, the first predetermined position being based on a first position of the first stringer on the skin; and

[0287] A second mouse hole, spaced apart from the first mouse hole, is cut from the flange at a second predetermined position, and the size of the second mouse hole is set to fit above a second stringer adjacent to the first stringer on the skin, the second predetermined position being based on a second position of the second stringer on the skin.

[0288] 70. The method according to Clause 69, wherein the composite fuselage frame is configured to provide lateral support to the inner surface of the skin of a fuselage cylindrical section having a plurality of stringers, the composite fuselage frame spanning the plurality of stringers, the composite fuselage frame including a plurality of shear connecting feet extending toward the skin between adjacent stringers, each shear connecting foot having frame filler integrated in each shear connecting foot during the manufacture of the composite fuselage frame, such that the composite fuselage frame includes a plurality of frame fillers integrated within the plurality of shear connecting feet.

[0289] 71. The method according to Clause 66, further comprising:

[0290] An external protective fiberglass layer is laid on the third laminate of the composite fuselage frame currently on the second laying tool to provide an external protective fiberglass layer on top of the third laminate.

[0291] The outer protective glass fiber layer and the second polymer matrix composite layer are formed together on the third top surface of the second laminate of the intermediate composite frame, wherein the first laminate of the frame filler is integrated into the third top surface of the second laminate; and

[0292] During the formation process, the outer protective glass fiber sheet is compacted together with the second stack and the second lamination of the second material polymer matrix composite sheet to provide the third lamination of the composite fuselage frame.

[0293] 72. The method according to Clause 66, further comprising:

[0294] The frame filler is integrated into the shear-resistant connecting leg of the composite fuselage frame by forming the second stack of the second material polymer matrix composite sheet via a clamping arc above the first stack of the first material polymer matrix composite sheet; and

[0295] The frame filler provided by the first stack and the shear-resistant connecting feet provided by the second stack are co-cured to harden the composite fuselage frame.

[0296] 73. A composite fuselage frame for an aircraft, the composite fuselage frame comprising:

[0297] A web, the web comprising a width and a first predetermined thickness, and forming an arc relative to the fuselage of the aircraft between a first end and a second end, the width extending from the outer side to the inner side of the skin relative to the fuselage; and

[0298] A shear-resistant connecting foot, disposed on the outer side of the web between a first rat hole and a second rat hole, and protruding from a first predetermined thickness of the web, the shear-resistant connecting foot having a second predetermined thickness extending along the skin of the aircraft between a first stringer and a second stringer, and

[0299] The second predetermined thickness of the shear-resistant connecting leg is defined by the first predetermined thickness and a second plurality of second polymer matrix composite layers formed by a plurality of first material polymer matrix composite sheets, the plurality of first material polymer matrix composite sheets providing a frame filler integrated with the plurality of second material polymer matrix composite sheets of the first predetermined thickness.

[0300] 74. The composite fuselage frame according to Clause 73, further comprising:

[0301] The first mouse hole is disposed on the outer side of the web, and the size of the first mouse hole is set to fit above the first stringer on the skin; and

[0302] The second rat hole is spaced apart from the first rat hole on the outer side of the web plate, and the size of the second rat hole is set to fit above the second stringer adjacent to the first stringer on the skin of the fuselage.

[0303] 75. The composite fuselage frame according to Clause 74, wherein the composite fuselage frame is configured to provide lateral support to the inner surface of the skin of a fuselage cylindrical section having a plurality of stringers, the composite fuselage frame spanning the plurality of stringers, the composite fuselage frame including a plurality of shear connecting feet extending toward the skin between adjacent stringers, each shear connecting foot having frame filler integrated in each shear connecting foot during the manufacture of the composite fuselage frame, such that the composite fuselage frame includes a plurality of frame fillers integrated within the plurality of shear connecting feet.

[0304] 76. The composite fuselage frame according to Clause 73, further comprising:

[0305] A flange connected to the web, which in turn is connected to the shear connection foot, wherein the shear connection foot includes frame filler added to the flange, the frame filler being equal to the thickness of the stringer flanges on the first stringer and the second stringer.

[0306] 77. The composite fuselage frame according to Clause 73, wherein the first predetermined thickness of the web is defined by a first plurality of first polymer matrix composite layers formed by the plurality of second material polymer matrix composite sheets.

[0307] 78. The composite fuselage frame according to Clause 73, wherein the shear connecting foot extends from the web to the distal end of the shear connecting foot, and the outer surface of the shear connecting foot is configured to face the skin and align with a corresponding portion of the skin.

[0308] 79. The composite fuselage frame according to Clause 73, wherein the shear connecting leg extends from the first predetermined thickness of the web such that one side of the shear connecting leg is configured to face the skin and fit within the area defined by the first stringer, the second stringer, and the corresponding portion of the skin between the first stringer and the second stringer.

[0309] 80. The composite fuselage frame according to Clause 73, wherein the second material polymer matrix composite sheet comprises at least one of a thermosetting matrix composite sheet and a thermoplastic matrix composite sheet.

[0310] 81. The composite fuselage frame according to Clause 73, wherein the second material polymer matrix composite sheet comprises unidirectional reinforcing fibers and a polymer matrix material.

[0311] 82. The composite fuselage frame according to Clause 81, wherein the unidirectional reinforcing fiber comprises at least one of carbon fiber, glass fiber, aramid fiber, natural fiber and polyetheretherketone fiber.

[0312] 83. The composite fuselage frame according to Clause 81, wherein the polymer matrix material comprises at least one of thermosetting resin, thermoplastic resin, epoxy resin, phenolic resin, polyurethane resin, polyimide resin, polyethylene resin, polypropylene resin, polybutylene terephthalate resin, polyamide resin, polyphenylene sulfide resin, polyetherimide resin, polyetherketone ketone resin, and polyetheretherketone resin.

[0313] 84. The composite fuselage frame according to Clause 73, wherein the first material polymer matrix composite sheet comprises at least one of a thermosetting matrix composite sheet and a thermoplastic matrix composite sheet.

[0314] 85. The composite fuselage frame according to Clause 73, wherein the first material polymer matrix composite sheet comprises reinforcing fabric and polymer matrix material.

[0315] 86. The composite fuselage frame according to Clause 85, wherein the reinforcing fabric comprises at least one of plain weave fabric, 0 / 90 degree woven fabric and 45 degree woven fabric.

[0316] 87. The composite fuselage frame as described in Clause 85, wherein the reinforcing fabric comprises woven reinforcing fibers interwoven in at least two directions.

[0317] 88. The composite fuselage frame according to Clause 87, wherein the woven reinforcing fiber comprises at least one selected from glass fiber, carbon fiber, aramid fiber, natural fiber and polyetheretherketone fiber.

[0318] 89. The composite fuselage frame according to Clause 85, wherein the polymer matrix material comprises at least one of thermosetting resin, thermoplastic resin, epoxy resin, phenolic resin, polyurethane resin, polyimide resin, polyethylene resin, polypropylene resin, polybutylene terephthalate resin, polyamide resin, polyphenylene sulfide resin, polyetherimide resin, polyetherketone ketone resin, and polyetheretherketone resin.

[0319] 90. The composite fuselage frame according to Clause 73, wherein the web and the shear-resistant connecting foot further include an external glass fiber protective layer extending between the first and second ends of the web and extending along the inner and outer surfaces of the shear-resistant connecting foot between the first and second rat holes, the external glass fiber protective layer being formed of the external glass fiber protective layer.

[0320] 91. The composite fuselage frame according to Clause 73, wherein the composite fuselage frame further comprises:

[0321] Single-piece section.

[0322] 92. The composite fuselage frame as described in Clause 73, wherein a plurality of composite fuselage frames form a segmented frame for a section of the fuselage of the aircraft.

[0323] 93. A method for manufacturing a fuselage section for an aircraft, the method comprising:

[0324] A skin is received in a manufacturing unit, the skin having at least a first stringer and a second stringer on its inner surface, the skin having the first stringer and the second stringer having been previously manufactured and hardened;

[0325] The skin is mounted on a manufacturing tool in the manufacturing unit;

[0326] A composite fuselage frame, in which frame filler is integrated, is received at the manufacturing unit; the composite fuselage frame has been previously manufactured.

[0327] The composite fuselage frame is mounted above the first stringer and the second stringer on the inner surface of the skin.

[0328] 94. The method according to Clause 93, wherein manufacturing the composite fuselage frame comprises:

[0329] Multiple first material polymer matrix composite sheets are laid at predetermined locations on a first portion of the top surface of a laying tool for the composite fuselage frame to form a first stack of first material polymer matrix composite sheets, wherein the predetermined locations are associated with shear connection feet of the composite fuselage frame and the frame filler.

[0330] Multiple second-material polymer matrix composite sheets are laid on a second portion of the top surface of the first-material polymer matrix composite sheet and the laying tool to provide a second stack of second-material polymer matrix composite sheets; and

[0331] The second stack and the first stack are formed.

[0332] 95. The method according to Clause 94, forming the second stack and the first stack comprises:

[0333] The second stack and the first stack are formed onto the top surface of the forming tool; and

[0334] The second and first stacks are compacted onto the top surface of the compaction tool.

[0335] 96. The method according to Clause 94, forming the second stack and the first stack comprises:

[0336] The second stack and the first stack are formed onto the top surface of the laying tool; and

[0337] During the formation process, the second and first stacks are compacted to provide a laminate for the composite fuselage frame.

[0338] 97. The method described in Clause 96, further comprising manufacturing the composite fuselage frame:

[0339] The laminate of the composite fuselage frame is cured to harden the composite fuselage frame.

[0340] 98. The method according to Clause 97, wherein, after the laminate is cured, the composite fuselage frame includes a web and a flange, the web including a width and a first predetermined thickness and forming an arc between a first end and a second end relative to the fuselage of the aircraft, the width extending from an outer side to an inner side relative to the skin of the fuselage, the flange disposed on an outer side of the web and projecting from the first predetermined thickness of the web, the manufacture of the composite fuselage frame further comprising:

[0341] A first mouse hole is cut from the flange at a first predetermined position, and the size of the first mouse hole is set to fit above the first stringer on the skin, the first predetermined position being based on a first position of the first stringer on the skin; and

[0342] A second mouse hole, spaced apart from the first mouse hole, is cut from the flange at a second predetermined position, and the size of the second mouse hole is set to fit above the second stringer adjacent to the first stringer on the skin. The second predetermined position is based on the second position of the second stringer on the skin.

[0343] 99. The method according to Clause 98, wherein the composite fuselage frame is configured to provide lateral support to the inner surface of the skin of a fuselage cylindrical section having a plurality of stringers, the composite fuselage frame spanning the plurality of stringers, the composite fuselage frame including a plurality of shear connecting feet extending toward the skin between adjacent stringers, each shear connecting foot having frame filler integrated in each shear connecting foot during the manufacture of the composite fuselage frame, such that the composite fuselage frame includes a plurality of frame fillers integrated within the plurality of shear connecting feet.

[0344] 100. The method according to Clause 98, wherein a portion of the flange between the first rat hole and the second rat hole defines the shear-resistant connecting leg, the shear-resistant connecting leg including a second predetermined thickness, wherein the first predetermined thickness of the web is defined by a first plurality of first polymer matrix composite layers formed by the plurality of second material polymer matrix composite sheets, wherein the second predetermined thickness of the shear-resistant connecting leg is defined by the first predetermined thickness and a second plurality of second polymer matrix composite layers formed by the plurality of first material polymer matrix composite sheets, the plurality of first material polymer matrix composite sheets providing the frame filler integrated with the plurality of second material polymer matrix composite sheets of the first predetermined thickness.

[0345] 101. The method described in Clause 94, further comprising manufacturing the composite fuselage frame:

[0346] The frame filler is integrated into the shear-resistant connecting feet of the composite fuselage frame by forming a second stack of the second material polymer matrix composite sheet via a clamping arc above the first stack of the first material polymer matrix composite sheet; and

[0347] The frame filler provided by the first stack and the shear-resistant connecting feet provided by the second stack are co-cured to harden the composite fuselage frame.

Claims

1. A method for manufacturing a composite fuselage frame for an aircraft, the method comprising: Multiple first material polymer matrix composite sheets are laid at predetermined locations on a first portion of the top surface of a laying tool for the composite fuselage frame to form a first stack of first material polymer matrix composite sheets, the predetermined locations being associated with shear connection feet and frame fillers of the composite fuselage frame. Multiple second-material polymer matrix composite sheets are laid on a second portion of the top surface of the first-material polymer matrix composite sheet and the laying tool to provide a second stack of second-material polymer matrix composite sheets; and The second stack and the first stack are formed.

2. The method according to claim 1, wherein, The first material polymer matrix composite sheet includes at least one of thermosetting matrix composite sheets and thermoplastic matrix composite sheets.

3. The method according to claim 1, wherein, The first material polymer matrix composite sheet includes reinforcing fabric and polymer matrix material.

4. The method according to claim 3, wherein, The reinforcing fabric includes at least one of plain weave fabric, 0 / 90 degree woven fabric, and 45 degree woven fabric.

5. The method according to claim 3, wherein, The reinforcing fabric comprises woven reinforcing fibers interwoven in at least two directions.

6. The method according to claim 5, wherein, The braided reinforcing fiber includes at least one of glass fiber, carbon fiber, aramid fiber, natural fiber, and polyetheretherketone fiber.

7. The method according to claim 3, wherein, The polymer matrix material includes at least one of thermosetting resin, thermoplastic resin, epoxy resin, phenolic resin, polyurethane resin, polyimide resin, polyethylene resin, polypropylene resin, polybutylene terephthalate resin, polyamide resin, polyphenylene sulfide resin, polyetherimide resin, polyetherketone ketone resin, and polyetheretherketone resin.

8. A method for manufacturing frame filler for a composite fuselage frame, the method comprising: Multiple first material polymer matrix composite sheets are laid on a first laying tool for the frame filler to provide a first stack of first material polymer matrix composite sheets on a first top surface of the first laying tool; and The first stack is formed.

9. A composite fuselage frame for an aircraft, the composite fuselage frame comprising: A web, the web including a width and a first predetermined thickness, and forming an arc between a first end and a second end relative to the fuselage of the aircraft, the width extending from the outer side to the inner side of the skin relative to the fuselage; as well as A shear-resistant connecting foot, disposed on the outer side of the web between a first rat hole and a second rat hole, and protruding from a first predetermined thickness of the web, the shear-resistant connecting foot having a second predetermined thickness extending along the skin of the aircraft between a first stringer and a second stringer, and The second predetermined thickness of the shear-resistant connecting leg is defined by the first predetermined thickness and a second plurality of second polymer matrix composite layers formed by a plurality of first material polymer matrix composite sheets, the plurality of first material polymer matrix composite sheets providing a frame filler integrated with the plurality of second material polymer matrix composite sheets of the first predetermined thickness.

10. A method for manufacturing a fuselage section for an aircraft, the method comprising: A skin is received in a manufacturing unit, the skin having at least a first stringer and a second stringer on its inner surface, the skin having the first stringer and the second stringer having been previously manufactured and hardened; The skin is mounted on a manufacturing tool in the manufacturing unit; A composite fuselage frame is received at the manufacturing unit, in which frame filler is integrated, the composite fuselage frame having been previously manufactured; as well as The composite fuselage frame is mounted above the first stringer and the second stringer on the inner surface of the skin.

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