METHOD FOR MANUFACTURING A REAR SECTION OF AN AIRCRAFT AND AIRCRAFT REAR SECTION

Abstract

A method for manufacturing a composite assembly of an empennage and rear-fuselage having a continuous skin solution. The method obtains parts of the sub-structure. For each part, it is obtained a plurality of stringers performs and frames preforms by composite tooling. The frames are transferred to curing frames molds and a sub-structure skin is obtained. Furthermore, the method includes integrating the parts over an integration tool having cavities for locating the curing frames molds and the stringer performs. Furthermore, the method includes co-curing the integration tool in one shot on an autoclave, demolding the sub-structure skin sections and disassembling the curing frame molds to obtain the composite assembly of the rear section.

Claims

1. A method for manufacturing a composite assembly of a rear section of an aircraft having a continuous skin solution, the method comprising: obtaining parts of a sub-structure by: obtaining a plurality of stringers preforms formed by composite tooling; obtaining a plurality of frames preforms formed by composite tooling; transferring the plurality of frames to curing frames molds; and obtaining a sub-structure skin section formed by composite tooling; positioning the parts over an integration tool comprising cavities configured to locate the curing frames molds and the stringers preforms by: positioning the curing frames molds in the cavities; locating the stringer preforms in the cavities; and laying-up the sub-structure skin sections; co-curing the integration tool with the integrated parts in a single operation on an autoclave; demolding the sub-structure skin sections to obtain a continuous skin; disassembling the curing frame molds to obtain co-cured parts; and joining the co-cured parts to form the composite assembly of the rear section.

2. The method of claim 1, wherein obtaining a plurality of stringers preforms comprises: laying up the stringers preforms in a flat plate; and hot-forming the stringers preforms with a membrane in a stringers male tool or female tool.

3. The method of claim 1, wherein obtaining a plurality of frames preforms comprises: laying up the frames preforms in a flat plate; and hot-forming the frames preforms with a membrane in a male tool.

4. The method of claim 1, further comprising obtaining a plurality of joining parts for each part by: laying up the joining parts in a flat plate; and hot-forming the joining parts with a membrane in a male or female tool.

5. The method of claim 1, wherein laying-up the sub-structure skin sections further comprises: laying-up the sub-structure skin sections over the integration tool; and adding internal molds inside stringers preforms profile.

6. The method of claim 1, wherein laying-up the sub-structure skin sections further comprises: laying-up the sub-structure skin sections over a dedicated tool; pre-curing the sub-structure skin sections; and transferring the pre-cured layup of the sub-structure skin sections from the dedicated tool to the integration tool.

7. The method of claim 1, wherein integrating the parts in an integration tool, comprises integrating the two or more parts in a male integration tool.

8. The method of claim 1, wherein integrating the two or more parts in an integration tool, comprises integrating the two or more parts in a female integration tool.

9. The method of claim 1, wherein joining the two or more co-cured parts to obtain the composite assembly of the rear section, comprises joining the two or more co-cured parts by mechanical means or by adhesive bonding means.

Description

SUMMARY OF THE DRAWINGS

[0026] For a better understanding the above explanation and for the sole purpose of providing an example, some non-limiting drawings are included that schematically depict a practical embodiment.

[0027] FIGS. 1A and 1B shows a conventional T-shape vertical tail plane and interface fittings comprising longitudinal load fittings and transversal load fittings.

[0028] FIG. 2 shows examples of conventional vertical tail plane lugs and lateral attachments.

[0029] FIG. 3 shows conventional fairings to cover interface fittings as shown in FIGS. 1A, 1B and 2.

[0030] FIGS. 4A and 4B shows a first way to join two sub-structure halves according to the present disclosure.

[0031] FIGS. 5A and 5B shows a second way to join two sub-structure halves according to the present disclosure.

[0032] FIGS. 6A, 6B and 6C show a third way to join two sub-structure halves according to the present disclosure.

[0033] FIGS. 7Ai, 7Aii, 7Bi and 7Bii show a fourth way to join two sub-structure halves according to the present disclosure.

[0034] FIGS. 8A and 8B show a fifth way to join two sub-structure halves according to the present disclosure.

[0035] FIG. 9 shows an example of an integration tool according to the present disclosure.

DETAILED DESCRIPTION OF THE INVENTION

[0036] FIG. 1A shows a conventional T-shape vertical tail plane (100). The vertical tail plane (100) comprises a rudder (120) and elevators (130) and interface fittings (110). FIG. 1B shows the interface fittings divided by longitudinal load fittings (111) and transversal load fittings (112) of a conventional vertical tail plane of an aircraft.

[0037] FIG. 2 shows examples of conventional vertical tail plane lugs and lateral attachments (200).

[0038] FIG. 3 shows conventional fairings (300) to cover interface fittings as the ones shown in FIGS. 1A, 1B and 2.

[0039] FIGS. 4A and 4B show a first way to join two sub-structures halves according to the present disclosure in order to easier the sub-structure manufacturing. FIG. 4A shows a first skin half (410) and a second skin half (420). The first and the second skin halves (410) and (420) can be manufactured in one single shot including the stringers (415) and the joining parts (425). Frames (440) and spars (430) are joined to the two half skins (410) and (420) through the joining parts (425) and joining means comprised in the frames (440) and spars (430) as shown in FIG. 4B. In some examples, the joining means comprise clips.

[0040] FIGS. 5A and 5B show a second way to join two sub-structure halves according to the present disclosure in order to easier the sub-structure manufacturing of an empennage of an aircraft. FIG. 5A shows a first skin half (510) and a second skin half (520) both manufactured in one shot and including stringers (515). Frames (540) and spars (530) are joined to the two half skins (510) and (520) through the joining means which are lately added as shown in FIG. 5B.

[0041] FIGS. 6A, 6B and 6C show a third way to join two sub-structure halves according to the present disclosure in order to easier the sub-structure manufacturing of an empennage of an aircraft. A first sub-structure half comprises a first skin half (610), a spar half (630) and a frame half (640), as well as stringers (625) and (optionally) joining parts (615). A second sub-structure half comprises a second skin half (620), a second spar half (635) and a second frame half (645), as well as stringers and (optionally) joining parts.

[0042] The frames halves (640) and (645) and spar halves (630) and (635) can be integrated or joined to each half skin (610) and (620), respectively by mechanical discrete means. The frames webs joint (650) is located at a symmetrical plane. The frames webs are joined in double shear manner with additional splices or single shear established directly on webs. The half skins (610) and (620) could be also manufactured in one shot including stringers (615) and half frames and spars which can be integrated through co-curing, co-bonding or equivalent methods, e.g. welding in thermoplastic as shown in FIG. 6C. Hence, the skin half-shells (610) and (620) are joined at the lower position as well as frames (640), (645) by mechanical means.

[0043] FIGS. 7Ai, 7Aii, 7Bi and 7Bii show a fourth way to join two sub-structure halves according to the present disclosure in order to easier the sub-structure manufacturing of an empennage of an aircraft. FIGS. 7Ai, ii shows a first skin half (710) manufactured on one shot with stringers (715), spars (730) and frames (740). A second skin half (720) includes stringers (715) and (optionally) joining parts (725) to close the ensemble by joining means that can be discrete or continuous joining means. FIGS. 7Bi, ii shows the second skin half (720) without joining parts (725).

[0044] FIGS. 8A and 8B show an isometric view of another option for the fourth way to join the two halves shown in FIGS. 7A and 7B. FIG. 8A shows the first skin half (710) manufactured on one shot including stringers (715) and two external frames (742) and (744) and the second skin half (720 including stringers (715) and the remaining intermediate frames (746). Shells are assembled directly to the frame feet or joining parts as shown in FIG. 8B.

[0045] All previously showed processes permit to include frame shear ties in the manufacturing process of half skins by continuous surface contact means as resin or bonding interfaces performed by co-curing, co-bonding or secondary co-bonding or equivalent methods. It makes easier assemble the frames. The rig must include, in this case, stringers and shear tie allocations. Alternatively integrated frames are possible, in which the frame feet (external flanges) directly joins the skin.

[0046] FIG. 9 shows an integration tool (900) according to the present disclosure. The integrated tool is used in an example of a manufacturing process according to the present disclosure that is being described below:

[0047] The manufacturing process disclosed herein permits obtaining a composite assembly of an empennage of an aircraft having a continuous skin solution based on the integration of two or more halves of sub-structure. The manufacturing process uses pre-impregnated Carbon fiber reinforced polymer (CFRP) composite.

[0048] For each half of the sub-structure, the proposed manufacturing process comprises a plurality of steps. In particular, the process comprises a step for obtaining stringer preforms. In this regard, omega stringers or T-stringers can be layed-up in a flat plate and formed in a subsequent hot-forming process with membrane in a tool that forces the composite to get the desire form. The tooling performed to obtain the stringer preforms can be a male concept or a female concept. This process is also applicable to joining parts as frames and spars shear-ties. In some examples, in order to perform the perform stringers lay-up, a forming process with a press tool can be performed. Roll forming and pultrusion processes can be used for manufacturing the stringers preforms.

[0049] Furthermore, the proposed manufacturing process comprises a step for obtaining frames preforms. Frames preforms are laid-up in a flat plate and formed in a subsequent hot-forming process with membrane in a male tool. After that, frame preforms are transferred to each corresponding curing mold. In some examples, stamping, braiding, roll forming and pultrusion processes could be used in order to obtain the frames preforms. Furthermore, frames could be cured separately and integrated to the skin afterwards through mechanical or chemical (co-bonded, bonding) systems.

[0050] In order to obtain the skin of the sub-structure, CFRP composite layers are laminated over a 3D shaped tool by means of a fibre placement machine. In a first scenario, the skin could be layed-up directly over an integration tool. In this regard, internal molds can be added inside the (omega) stringers profile to allow a smooth surface for the direct laying of the skin over the integration tool. In a second scenario the skin could be laid-up directly over a dedicated tool. In this respect, a transfer operation of the skin to the integration tool is required.

[0051] Integration of stringers and frames is performed in an integration male tool (900 with cavities for frames mold location (920 and cavities for stringers preforms positioning (910 as shown in FIG. 9. Frame preforms are transferred to curing mould and positioned in the integration tool (900 as shown in FIG. 9 following the location of the stringers in their corresponding cavities (910. In a first scenario, the skin is lay-up over the integration tool or alternatively, it can be transferred from its dedicated laying tool. In some examples, external caul-plates (if needed) and a vacuum bag can be located in the integration tool. In some examples, a female concept for integration tool is also applicable.

[0052] Furthermore, composite curing is performed. The integration tool with composite laminate inside is co-cured on an auto-clave in one-shot process. Latterly, the demolding operation is performed out of the autoclave. The skin is demolded in a vertical way with frame molds attached. The frame molds will be disassembled in a longitudinal way. Same process could be applicable to CFRP composite dry-fibers materials. Resign transfer molding, RTM or vacuum infusion process could be applied with the similar integration tooling concept.

[0053] Even though reference has been made to a specific embodiment of the invention, it is obvious for a person skilled in the art that the composite assembly described herein is susceptible to numerous variations and modifications, and that all the details mentioned can be substituted for other technically equivalent ones without departing from the scope of protection defined by the attached claims.

[0054] While at least one exemplary embodiment of the present invention(s) is disclosed herein, it should be understood that modifications, substitutions and alternatives may be apparent to one of ordinary skill in the art and can be made without departing from the scope of this disclosure. This disclosure is intended to cover any adaptations or variations of the exemplary embodiment(s). In addition, in this disclosure, the terms comprise or comprising do not exclude other elements or steps, the terms a or one do not exclude a plural number, and the term or means either or both. Furthermore, characteristics or steps which have been described may also be used in combination with other characteristics or steps and in any order unless the disclosure or context suggests otherwise. This disclosure hereby incorporates by reference the complete disclosure of any patent or application from which it claims benefit or priority.