Sandwich flanging structure composite material aircraft cabin door and manufacturing method

Abstract

The invention discloses a sandwich turnup structure composite material aircraft cabin door and a manufacturing method, and relates to the field of aeronautical manufacturing, the sandwich turnup structure composite material aircraft cabin door comprises a cabin door shell made of a composite material, and the cabin door shell comprises an outermost layer skin and a secondary outer layer skin; the side, facing the secondary outer-layer skin, of the edge of the outermost-layer skin is bent to form a turnup, a mounting space is defined by the turnup, a plurality of stringers protruding towards the inner side are integrally formed in the mounting space in the outermost-layer skin, and the extending direction of the stringers is parallel to the length direction of the cabin door; the inner side of the secondary outer layer skin is fixedly connected with at least two composite material cap-shaped beams, and the extension direction of the cap-shaped beams is perpendicular to the extension direction of the stringer; wherein the area, located outside the stringer, between the outermost layer skin and the secondary outer layer skin is filled with a honeycomb core, mounting cavities are formed in the stringer and the cap-shaped beam, and the mounting cavities are filled with foam aluminum alloy core materials. By the adoption of the scheme, the requirements for light weight and large-size bearing capacity of the aircraft cabin door are met.

Description

Technical Field

[0001] This invention relates to the field of aerospace manufacturing technology, and in particular to a sandwich-flanged composite material aircraft door and its manufacturing method. Background Technology

[0002] As the air transport industry continues to develop towards higher efficiency, lower costs, and multi-scenario operations, the requirements for lightweight and high reliability of aircraft structural components are constantly increasing. As a key component in the airframe that simultaneously bears the functions of structural load-bearing, sealing and isolation, and safety protection, the performance of the cabin door directly affects the aircraft's operational safety, fuel economy, and adaptability to different transport scenarios. Especially with the application of cargo aircraft and passenger aircraft converted into cargo aircraft, the increased size and complex stress conditions of cabin doors place more stringent comprehensive requirements on the door structure in terms of weight control, load-bearing capacity, fatigue life, and ease of maintenance.

[0003] In existing aircraft door structures, traditional aluminum alloy doors still account for a large proportion of applications. Although this type of door has a mature design and manufacturing system, its high material density leads to a significant increase in the structural weight ratio in large-size door applications, directly compressing the effective payload space and adversely affecting the overall fuel efficiency of the aircraft. Furthermore, aluminum alloy doors typically rely on multiple machining and riveting assembly processes, resulting in long manufacturing and installation cycles. In scenarios where passenger aircraft are converted into cargo aircraft, this can easily prolong the conversion cycle and reduce the aircraft's ability to quickly enter service. On the other hand, composite material doors have been used in some aircraft models, but existing solutions are mostly concentrated on small-size equipment doors such as APU doors or passenger aisle doors. Their structural design and stress models are mainly designed for relatively uniform and controllable load conditions, significantly limiting their application scenarios.

[0004] In the application of large-size cargo doors in cargo aircraft and passenger-to-freighter converted aircraft, these doors are not only large in size and long in structural span, but also need to withstand complex working conditions such as impact loads, localized concentrated loads, and asymmetric forces generated during cargo loading and unloading. Existing aluminum alloy doors struggle to balance weight and load-bearing capacity, while existing composite material doors generally lack a structural design basis for the aforementioned complex load characteristics, making it difficult to reliably adapt to the usage scenarios of large-size doors while ensuring high load-bearing capacity and structural safety. Therefore, how to simultaneously meet the requirements of lightweight and large-size load-bearing capacity without significantly increasing structural complexity and manufacturing difficulty remains a key technical challenge in the field of aviation door technology that urgently needs to be solved and has long constrained engineering applications. Summary of the Invention

[0005] The purpose of this invention is to provide a sandwich-flanged composite material aircraft door and its manufacturing method, in order to solve the technical problem in the prior art of how to meet the requirements of lightweight and large-size load-bearing capacity without significantly increasing structural complexity and manufacturing difficulty.

[0006] To achieve the above objectives, the technical solution of the present invention is as follows: In a first aspect, a sandwich-flanged composite material aircraft door is provided, comprising: A hatch shell made of composite materials, the hatch shell comprising an outermost skin and a second outermost skin; The edge of the outermost skin is bent toward the side of the second outermost skin to form a flange, and the flange encloses an installation space. Multiple composite material stringers protruding inward are integrally formed in the installation space on the outermost skin, and the extension direction of the stringers is parallel to the length direction of the hatch. At least two composite material cap-shaped beams are fixedly connected to the inner side of the secondary outer skin, and the extension direction of the cap-shaped beams is perpendicular to the extension direction of the stringer. In this area, between the outermost skin and the second outermost skin, and outside the stringer, a honeycomb core is filled. The stringer and the cap beam have mounting cavities inside, which are filled with foamed aluminum alloy core material.

[0007] Furthermore, the thickness of the outermost skin layer is not equal to the thickness of the second outermost skin layer.

[0008] Furthermore, the cavity formed inside the flange is filled with a foamed aluminum alloy core material.

[0009] Furthermore, the foamed aluminum alloy core material filling the flange cavity is a gradient porosity foamed aluminum alloy with isotropic properties, and the porosity of the foamed aluminum alloy core material gradually increases from the region near the inner side of the hatch shell to the outer side of the free end of the flange.

[0010] Furthermore, the edge of the hatch is provided with a composite sealing strip, which includes a titanium alloy skeleton and a sealing material covering the titanium alloy skeleton.

[0011] Furthermore, it also includes an inner panel, which is detachably connected to the hatch housing.

[0012] Furthermore, the honeycomb core and the foamed aluminum alloy core are alternately or interspersed between the outermost skin and the second outermost skin.

[0013] Furthermore, the outermost skin and the second outermost skin are carbon fiber reinforced composite materials, and the honeycomb core is biodegradable polylactic acid.

[0014] Secondly, a method for manufacturing a sandwich-flanged composite aircraft door is provided, for manufacturing a sandwich-flanged composite aircraft door as described above, comprising the following steps: Material is laid on the mold to form the outermost skin, and additional prepreg layers are laid on the outermost skin in areas where flanges are to be formed. Subsequently, material is laid on the outermost skin to integrally form the multiple inwardly protruding composite stringers. The honeycomb core is filled in the area enclosed by the outermost skin and the stringers; Materials are laid on the honeycomb core, the foamed aluminum alloy core, and the stringer to form the secondary outer skin, and at least two composite material cap beams are integrally formed on the inner side of the secondary outer skin. The foamed aluminum alloy core is filled into the internal mounting cavity of the stringer and the cap beams. The laid-out structure is cured to solidify the outermost skin, flange, stringer, second outermost skin, and cap beam into a whole, forming the hatch shell, and the honeycomb core and aluminum foam core are fixed inside the hatch shell.

[0015] Furthermore, the material used to lay the skin, stringers, and cap beams is carbon fiber reinforced resin matrix composite prepreg, and the layup direction of the prepreg is configured to match the main force direction of the aircraft door.

[0016] Furthermore, after the curing step of the laid ply structure, a quality inspection step is also included, which includes: Mechanical properties of the prepreg and honeycomb core material used to manufacture the hatch were sampled and tested. The cured hatch shell is subjected to ultrasonic and X-ray testing to inspect its internal quality; finished product dimensions are also inspected. The external dimensions of the hatch housing are measured to ensure its assembly dimensional accuracy; After the door shell is assembled to the aircraft door frame, an inflation and pressure holding test is conducted to verify the sealing performance.

[0017] The beneficial effects of the sandwich-flanged composite material aircraft door and its manufacturing method provided by this invention are as follows: By designing the hatch shell as a composite sandwich structure consisting of an outermost skin and a second outermost skin, the hatch's weight is effectively reduced while ensuring structural strength. This avoids the problem of excessive weight proportion caused by the high material density of traditional aluminum alloy hatches, thus better meeting the lightweight application requirements of large-size hatches. Simultaneously, the integral molding of the composite material reduces the introduction of metal stiffeners and multi-point riveting connections, which helps to improve the hatch's structural performance without significantly increasing structural complexity and manufacturing difficulty.

[0018] The edge of the outermost skin bends towards the second outermost skin to form a flange, enclosing an installation space and creating a continuous closed edge structure around the hatch, significantly improving the overall stiffness and stability of the hatch edge area. Within this installation space, multiple composite material stringers extending along the hatch length are integrally formed, creating a continuous load-bearing system between the skin and the stringers. This enhances the hatch's in-plane bending and tensile / compressive load-bearing capacity under large span conditions, effectively reducing stress concentration risks and improving the overall load-bearing capacity of large-size hatches. Furthermore, a composite material cap-shaped beam is installed on the inner side of the second outermost skin, its extension direction perpendicular to the stringer extension direction. This creates a crisscrossing load-bearing skeleton structure inside the hatch, which can collaboratively distribute loads from different directions, improving the structural stability of the hatch under asymmetrical stress and localized concentrated load conditions. Furthermore, a honeycomb core is filled between the outermost and second outermost skin layers, and foamed aluminum alloy core material is filled inside the stringers and cap beams. This allows the sandwich structure to provide high specific stiffness support while also having good energy absorption capacity, thus maintaining the advantages of lightweight design while taking into account the load-bearing performance of large-size hatches under complex load environments. Attached Figure Description

[0019] Figure 1 This is a schematic diagram of the structure of the cabin door and the aircraft fuselage in an embodiment of the present invention; Figure 2 This is a schematic diagram of the overall structure of a sandwich-flanged composite material aircraft door according to an embodiment of the present invention; Figure 3 This is an exploded view of a sandwich-flanged composite material aircraft door according to an embodiment of the present invention; Figure 4 This is a schematic diagram of the structure of the girder in an embodiment of the present invention; Figure 5 This is a schematic diagram of the structure of the foamed aluminum alloy core material according to an embodiment of the present invention; Figure 6 This is a cross-sectional view of a sandwich-flanged composite material aircraft door according to an embodiment of the present invention; Figure 7 for Figure 6 Enlarged diagram of point A.

[0020] Reference numerals: 1. Aircraft fuselage; 2. Door shell; 21. Outermost skin; 211. Flanged edge; 212. Stringer; 213. Mounting cavity; 22. Secondary outermost skin; 221. Hat-shaped beam; 23. Honeycomb core; 3. Door rotating hinge end; 4. Door locking mechanism; 5. Composite sealing strip; 6. Inner panel. Detailed Implementation

[0021] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention. Unless otherwise defined, the technical or scientific terms used herein should have the ordinary meaning understood by those skilled in the art. The terms "comprising" and similar expressions used herein mean that the element or object preceding the word covers the element or object listed after the word and its equivalents, but does not exclude other elements or objects.

[0022] The following is in conjunction with the appendix Figure 1 - Appendix Figure 7 The specific embodiments of the present invention will be further described in detail below.

[0023] Reference Figures 1-7 In some embodiments of the present invention, a sandwich-flanged 211 structure composite material aircraft door is provided. The door includes a door shell 2 made of composite material. The door shell 2 has an overall plate-shell structure and is used to install at the door frame of the aircraft fuselage 1 and bear structural loads, airtight loads, and loading / unloading impact loads during flight. The door shell 2 is hinged to the aircraft fuselage 1 via a door rotating hinge end 3. A door locking mechanism 4 is provided at the end of the door shell 2 opposite to the door rotating hinge end 3. The door shell 2 includes at least an outermost skin 21 and a second outermost skin 22, which are spaced apart along the thickness direction to achieve lightweight design while ensuring overall rigidity and load-bearing capacity. In this embodiment, the outermost skin 21 and the second outermost skin 22 are preferably carbon fiber reinforced composite materials (CFRP). The fiber ratio in the main stress direction is increased through the layup design to meet the load-bearing requirements of large-size doors under bending, shearing, and tensile / compressive conditions.

[0024] In some specific embodiments of the present invention, the edge of the outermost skin 21 is bent towards the side of the next outermost skin 22 to form a flange 211. This flange 211 is continuously arranged along the circumference of the hatch and encloses an installation space, thereby forming a closed structural frame at the edge of the hatch. This flange 211 structure improves the bending and torsional resistance of the hatch edge area and provides a spatial basis for the arrangement of internal load-bearing components. In some specific embodiments of the present invention, a cavity is formed inside the flange 211, and the cavity is filled with a foamed aluminum alloy core material. The four edges of the flange adopt a foamed aluminum alloy sandwich design. The isotropic properties of the foamed aluminum alloy improve the strength and stiffness of the flange, effectively suppressing deformation of the hatch edge. Specifically, a foamed metal structure is added to the conventional composite flanged structure, which is an addition to the conventional composite flanged structure. The shear strength of the composite component includes interlaminar and in-plane shear strength, with interlaminar shear strength of 70-120 MPa and in-plane shear strength of 300-400 MPa; shear stiffness (shear modulus) of 40-60 GPa. The shear strength of the foamed aluminum alloy is 5-15 MPa, and the shear stiffness is 1-3 GPa. Typically, the thickness of the composite flange is 3-5 mm, and the thickness of the foamed aluminum alloy is 30-50 mm. Therefore, the equivalent strength and stiffness of the composite flanged structure with added foamed aluminum alloy core are at least doubled and at least 30% higher than those without foamed aluminum alloy core, calculated on an average basis. Compared to honeycomb cores, which have a cross-sectional shear strength of 1-3 MPa and shear stiffness of 0.5-2 GPa, and cross-sectional shear strength of 0.3-1.0 MPa and shear stiffness (shear modulus) of 0.1-0.5 GPa, honeycomb cores have lower cross-sectional shear strength and stiffness than aluminum foam alloys, and their cross-sectional strength and stiffness are also much lower. Therefore, aluminum foam alloys are chosen for the flange filling. Through finite element analysis and topology optimization design, the maximum deformation of a large-size (≥3m×2m) hatch under aerodynamic load of 0.05MPa can be ≤1mm, meeting the aircraft door frame adaptation requirements.

[0025] Preferably, the foamed aluminum alloy core material is a gradient porosity foamed aluminum alloy with isotropic properties, and its porosity gradually increases from the region near the inner side of the hatch shell 2 to the outer side of the free end of the flange 211, thereby ensuring the load-bearing capacity of the root of the flange 211 while improving the energy absorption capacity and structural stability of the edge region.

[0026] Reference Figure 3 and Figure 4In some embodiments of the present invention, multiple inwardly protruding composite material stringers 212 are integrally formed within the installation space of the outermost skin 21. The stringers 212 and the outermost skin 21 are integrally formed through a co-curing process, with their extension direction parallel to the length direction of the door. This integral forming method creates a continuous stress path between the stringers 212 and the skin, avoiding stress concentration problems associated with traditional splicing or riveting structures. This structure is particularly suitable for cargo aircraft and passenger-to-cargo converted aircraft with large doors subjected to localized impact loads during loading and unloading. (Refer to...) Figures 5-7 In this embodiment, the stringer 212 has an installation cavity 213 inside, and the installation cavity 213 is filled with aluminum foam core material. The isotropic properties of aluminum foam alloy are used to improve the compressive and buckling resistance of the stringer 212, while also taking into account the energy absorption effect.

[0027] In some specific embodiments of the present invention, at least two composite material cap beams 221 are fixedly connected to the inner side of the secondary outer skin 22. The cap beams 221 extend along the width direction of the hatch, and their extension direction is perpendicular to the extension direction of the stringer 212. Through the orthogonal arrangement of the stringer 212 and the cap beams 221, a crisscross load-bearing skeleton structure is formed inside the hatch, so that loads from different directions can be distributed and transferred to the entire shell structure. The cap beams 221 are also hollow structures, with mounting cavities 213 formed inside. The mounting cavities 213 are filled with foamed aluminum alloy core material, thereby improving the structural stability of the cap beams 221 without significantly increasing the weight. The cap beams 221 are preferably fixed to the inner side of the secondary outer skin 22 by a secondary adhesive bonding method, making the overall manufacturing process more flexible and easy to adjust under different hatch sizes and load requirements. Preferably, the cap-shaped beam 221 is a variable cross-section cap-shaped CFRP beam. Through biomimetic topology optimization, the non-stressed area material is removed, and the hollow area is filled with gradient porosity foam aluminum alloy. Without reducing the strength, the weight is reduced by 15%-20% compared with the traditional composite door, thus achieving the core requirement of "weight reduction and load increase".

[0028] In some specific embodiments of the present invention, a honeycomb core 23 is filled in the area between the outermost skin 21 and the second outermost skin 22, and outside the stringer 212. The honeycomb core 23 is preferably made of biodegradable polylactic acid, which provides high specific stiffness support, helps to reduce the overall weight of the hatch, and meets environmental protection and recycling requirements.

[0029] In some embodiments, the honeycomb core 23 and the foamed aluminum alloy core are alternately or inlaid between the outermost skin 21 and the second outermost skin 22, so that different areas of the hatch can obtain differentiated structural support according to the stress characteristics, thereby further improving the structural matching degree of large-size hatches under complex load environments.

[0030] In some specific embodiments of the present invention, the thickness of the outermost skin 21 is not equal to the thickness of the next outermost skin 22. Specifically, the thickness of each layer is designed as an unequal-thickness layer based on strength calculations, ranging from 1.2mm to 3.5mm. This thickness gradient design provides higher overall stiffness to the structure near the hatch. Local impact loads are transferred layer by layer through the gradient skin to the main load-bearing cap beam 221, avoiding stress concentration.

[0031] In some specific embodiments of the present invention, the hatch shell 2, through the outermost skin 21, the second outermost skin 22, and the stringers 212, cap beams 221, and sandwich structures disposed therebetween, together form a multi-path load transfer system. The integrally molded composite stringers 212 extending along the length of the hatch mainly bear in-plane tensile, compressive, and bending loads; the composite cap beams 221 extending along the width of the hatch are used to share out-of-plane loads and local concentrated loads. The two work together to disperse the load within the hatch shell 2, avoiding overloading of a single load-bearing component.

[0032] Based on the above structure, the internal mounting cavities 213 of both the stringer 212 and the cap beam 221 are filled with aluminum foam core material. Due to the isotropic mechanical properties of aluminum foam, it can provide balanced support when subjected to multi-directional loads, effectively improving the stability of the stringer 212 and the cap beam 221 under compression, shear, and buckling conditions. Compared to structures relying solely on the honeycomb core 23, this arrangement enables the load-bearing components to possess higher impact resistance and fatigue resistance in complex load environments.

[0033] Meanwhile, honeycomb core 23 is filled in the area between the outermost skin 21 and the second outermost skin 22, outside the stringer 212, providing a lightweight support structure for the hatch in non-load-bearing areas, thereby reducing unnecessary material usage while ensuring overall rigidity. By using a combination of honeycomb core 23 and foamed aluminum alloy core material in different areas, a reasonable balance is achieved between weight control and load-bearing capacity in the hatch structure, avoiding the weight increase problem caused by overall thickening of the layers.

[0034] Furthermore, the outermost skin 21 and the next outermost skin 22 are formed independently and structurally connected to the cap beam 221 via a stringer 212. This ensures the main load-bearing components remain continuous and intact during the forming process, reducing the need for openings or connection breaks along the load-bearing path, thereby lowering the risk of stress concentration and improving the overall structural reliability. This structural design not only meets the load-bearing requirements of large-size hatches but also simplifies the manufacturing process and improves overall manufacturing efficiency.

[0035] In some specific embodiments of the present invention, a composite sealing strip 5 is provided on the edge of the hatch. The composite sealing strip 5 includes a titanium alloy skeleton and a sealing material covering the titanium alloy skeleton. The titanium alloy is Ti-6Al-4V. The titanium alloy skeleton provides structural support, and the sealing material provides airtightness, enabling the hatch to maintain a stable sealing effect under high-altitude flight conditions. In some other embodiments of the present invention, an inner panel 6 is also included, which is detachably connected to the hatch shell 2. The inner panel 6 is preferably a metal or composite material plate, used to withstand direct impacts during cargo loading and unloading, while facilitating later maintenance and replacement, thereby reducing the overall maintenance cost of the hatch.

[0036] In summary, the sandwich-flanged composite aircraft door structure 211 does not achieve weight reduction solely through material replacement. Instead, it achieves overall weight reduction through the synergistic design of composite material properties and structural configuration, ensuring the door's load-bearing capacity meets usage requirements. Specifically, the double-layer composite shell structure formed by the outermost skin 21 and the second outermost skin 22, combined with the rational arrangement of the stringer 212, cap beam 221, and sandwich structure, ensures sufficient strength and stiffness along the main load-bearing paths while reducing material redundancy in non-main load-bearing areas, thus achieving a balance between weight reduction and load-bearing capacity. This structural form is particularly suitable for large-size door applications, increasing effective load-bearing capacity without sacrificing structural safety. Because the door shell 2 is made of carbon fiber reinforced composite material, combined with the double-layer composite layup shell structure, the door as a whole possesses high specific strength and specific stiffness. Based on this, by incorporating a sandwich structure formed by a foamed aluminum alloy core and a honeycomb core 23 inside the door, the door can effectively disperse and absorb energy through a multi-layered structure when subjected to impact loads generated during cargo loading and unloading, reducing the risk of localized stress concentration. Compared to traditional aluminum alloy door structures, this structural configuration improves the door's fatigue resistance under repeated loads, extends its service life, and is suitable for the high-frequency operating conditions of cargo aircraft and passenger-to-freighter converted aircraft.

[0037] Furthermore, the integrated molding structure of the outermost skin 21 and the stringer 212, along with the integral molding of the flange 211, the cap beam 221, and the hatch shell 2, significantly reduces the number of components and connection nodes in the hatch structure. Compared to the multiple processing and assembly steps required for traditional aluminum alloy hatches, such as blanking, stamping, and riveting, this integrated molding structure simplifies the manufacturing process and shortens the processing cycle. The hatch structure differs from the traditional general-purpose composite material hatch design of "composite material layup—honeycomb core 23—composite material layup". By using an integral layup method to form the outermost composite material layup and the stringer 212, and filling the stringer 212 with foamed aluminum alloy core material, the stringer 212 achieves higher strength and stiffness while maintaining continuity; by filling the shell flange 211 with foamed aluminum alloy sandwich material, the structural stability of the hatch edge area is further enhanced. The relatively independent double-layer composite material layup structure avoids the need for openings on the main load-bearing components, which helps improve structural reliability and simplify the processing technology. The application of aluminum foam alloy in stringers 212, cap beams 221 and flanges 211 enables the hatch to have more balanced mechanical properties under multi-directional loads, thereby compensating for the lack of unidirectional force on the honeycomb core 23.

[0038] By incorporating a composite sealing strip 5 made of SMA titanium alloy skeleton and silicone rubber at the edge of the hatch, the hatch can adaptively adjust the sealing pressure under different temperature environments by relying on the temperature self-recovery characteristics of the SMA titanium alloy skeleton. This sealing structure helps maintain stable sealing performance during high-altitude flight and temperature cycling conditions.

[0039] Secondly, in some specific embodiments of the present invention, a method for manufacturing a sandwich-flanged 211 structure composite material aircraft door is provided, for manufacturing a sandwich-flanged 211 structure composite material aircraft door as described above. This manufacturing method is based on a composite material co-curing molding process, and through the coordinated control of material selection, layup direction, and core arrangement, enables the door shell 2 to maintain lightweight while possessing the structural load-bearing capacity required for cargo and passenger-to-cargo conversion applications.

[0040] In this embodiment, the raw materials are prepared first. Based on the hatch load calculation results, carbon fiber reinforced resin matrix composite prepreg is selected as the material for laying out the outermost skin 21, the second outermost skin 22, the stringer 212, and the cap beam 221. The resin content of the prepreg is controlled at 35%–40%. A multi-directional combination method is adopted for the layup direction, wherein the fiber layup ratio along the main stress direction of the hatch is not less than 60%, so as to improve the load-bearing capacity of the hatch under bending and tensile / compressive conditions. The honeycomb core 23 is made of aramid paper, with its honeycomb size controlled at 8–12 mm and its honeycomb wall thickness controlled at 0.03–0.05 mm, so as to achieve structural weight reduction while ensuring support stiffness.

[0041] After the mold is cleaned, the prepreg is laid on the mold to form the outermost skin 21. In the area where the flange 211 is to be formed, the number of prepreg layers is increased to ensure that the flange 211 has sufficient structural thickness and edge stiffness after molding. Subsequently, prepreg is laid on the outermost skin 21 to form at least three inwardly protruding composite stringers through an integral layup, so that the stringers 212 and the outermost skin 21 form a continuous structure during the molding process. For areas with concentrated tensile loads, the local load-bearing capacity is further enhanced by increasing the number of layups. Then, a cavity is formed on the flange 211, and the cavity is filled with a foamed aluminum alloy core material. Preferably, the foamed aluminum alloy core material is a gradient porosity foamed aluminum alloy with isotropic properties, and its porosity gradually increases from the area near the inner side of the hatch shell 2 to the outer side of the free end of the flange 211, thereby ensuring the load-bearing capacity at the root of the flange 211 while improving the energy absorption capacity and structural stability of the edge area.

[0042] After the outermost skin 21 and stringers 212 are laid, the honeycomb core 23 is filled in the area enclosed by the outermost skin 21 and stringers 212. The honeycomb core 23 is positioned and fixed with structural adhesive to ensure that it is densely filled without obvious gaps, so as to reliably bond with the composite material structure during the subsequent curing process.

[0043] Based on the above structure, materials are laid on the honeycomb core 23, the aluminum foam core material, and the stringer 212 to form a secondary outer skin 22. Inside the secondary outer skin 22, at least two composite material cap beams 221 are integrally laid using a conformal laying method, so that the cap beams 221 and the secondary outer skin 22 form a stable load-bearing structure after curing. Simultaneously, the aluminum foam core material is filled into the internal mounting cavities 213 of the stringer 212 and the cap beams 221, providing isotropic support under multi-directional loads and improving the structure's impact and buckling resistance.

[0044] After all the layers and core filling are completed, the laid layer structure is cured to solidify the outermost skin 21, flange 211, stringer 212, second outermost skin 22 and cap beam 221 into a whole hatch shell 2, and fix the honeycomb core 23 and foam aluminum alloy core material inside the hatch shell 2.

[0045] In some embodiments of the present invention, a quality inspection step is included after curing. Specifically, tensile strength testing is performed on each batch of prepreg, with a tensile strength of not less than 3500 MPa and a flexural strength of not less than 2500 MPa; compressive strength testing is performed on the honeycomb core 23, with a compressive strength of not less than 0.3 MPa. Subsequently, ultrasonic testing is performed on the cured door shell 2 to check for delamination or porosity defects, requiring an overall porosity of not more than 2%, and X-ray inspection is performed on the adhesive joints to ensure that there is no debonding. The external dimensions of the door shell are measured using a coordinate measuring machine, with dimensional errors controlled within ±1 mm to ensure its assembly compatibility with the aircraft fuselage 1 door frame. After the door shell 2 is assembled to the aircraft door frame, an inflation and pressure holding test at 0.05 MPa is conducted for 30 minutes without leakage to verify the sealing performance of the door.

[0046] While embodiments of the present invention have been described in detail above, it will be apparent to those skilled in the art that various modifications and variations can be made to these embodiments. However, it should be understood that such modifications and variations fall within the scope and spirit of the present invention. Furthermore, the present invention described herein may have other embodiments and can be implemented or carried out in various ways.

Claims

1. A sandwich-type flanged composite material aircraft door, characterized in that, include: A hatch shell made of composite materials, the hatch shell comprising an outermost skin and a second outermost skin; The edge of the outermost skin is bent toward the side of the second outermost skin to form a flange, and the flange encloses an installation space. Multiple composite material stringers protruding inward are integrally formed in the installation space on the outermost skin, and the extension direction of the stringers is parallel to the length direction of the hatch. At least two composite material cap-shaped beams are fixedly connected to the inner side of the secondary outer skin, and the extension direction of the cap-shaped beams is perpendicular to the extension direction of the stringer. In this area, between the outermost skin and the second outermost skin, and outside the stringer, a honeycomb core is filled. The stringer and the cap beam have mounting cavities inside, which are filled with foamed aluminum alloy core material.

2. The sandwich-flanged composite material aircraft door according to claim 1, characterized in that, The thickness of the outermost skin layer is not equal to the thickness of the second outermost skin layer.

3. The sandwich-flanged composite material aircraft door according to claim 1, characterized in that, The cavity formed inside the flange is filled with a foamed aluminum alloy core material.

4. The sandwich-flanged composite material aircraft door according to claim 3, characterized in that, The foamed aluminum alloy core material filling the flange cavity is a gradient porosity foamed aluminum alloy with isotropic properties, and the porosity of the foamed aluminum alloy core material gradually increases from the region near the inner side of the hatch shell to the outer side of the free end of the flange.

5. The sandwich-flanged composite material aircraft door according to claim 1, characterized in that, The edge of the hatch is provided with a composite sealing strip, which includes a titanium alloy skeleton and a sealing material covering the titanium alloy skeleton.

6. The sandwich-flanged composite material aircraft door according to claim 1, characterized in that, It also includes an inner panel that is detachably connected to the hatch housing.

7. The sandwich-flanged composite material aircraft door according to claim 1, characterized in that, The honeycomb core and the foamed aluminum alloy core are alternately or interspersed between the outermost skin and the second outermost skin.

8. The sandwich-flanged composite material aircraft door according to claim 1, characterized in that, The outermost skin and the second outermost skin are carbon fiber reinforced composite materials, and the honeycomb core is biodegradable polylactic acid.

9. A method for manufacturing a sandwich-flanged composite material aircraft cabin door, used to manufacture a sandwich-flanged composite material aircraft cabin door as described in any one of claims 1-8, characterized in that, Includes the following steps: Material is laid on the mold to form the outermost skin, and additional prepreg layers are laid on the outermost skin in areas where flanges are to be formed. Subsequently, material is laid on the outermost skin to integrally form the multiple inwardly protruding composite stringers. The honeycomb core is filled in the area enclosed by the outermost skin and the stringers; Materials are laid on the honeycomb core, the foamed aluminum alloy core, and the stringer to form the secondary outer skin, and at least two composite material cap beams are integrally formed on the inner side of the secondary outer skin. The foamed aluminum alloy core is filled into the internal mounting cavity of the stringer and the cap beams. The laid-out structure is cured to solidify the outermost skin, flange, stringer, second outermost skin, and cap beam into a whole, forming the hatch shell, and the honeycomb core and aluminum foam core are fixed inside the hatch shell.

10. A method for manufacturing a sandwich-flanged composite material aircraft door according to claim 9, characterized in that, The material used to lay the skin, stringers, and cap beams is carbon fiber reinforced resin matrix composite prepreg, and the layup direction of the prepreg is configured to match the main force direction of the aircraft door.

11. A method for manufacturing a sandwich-flanged composite aircraft door according to claim 9, characterized in that, After the curing step of the laid ply structure, a quality inspection step is also included, which includes: Mechanical properties of the prepreg and honeycomb core material used to manufacture the hatch were sampled and tested. The cured hatch shell is subjected to ultrasonic and X-ray testing to inspect its internal quality; finished product dimensions are also inspected. The external dimensions of the hatch housing are measured to ensure its assembly dimensional accuracy; After the door shell is assembled to the aircraft door frame, an inflation and pressure holding test is conducted to verify the sealing performance.

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