Composite material flat disc flange structure suitable for large-scale solar orientation device of space station

Abstract

The present invention relates to a large-scale, high-rigidity, integrated composite material flat disc flange structure suitable for use as the main structure of a large-scale space station's heliocentric orientation system. This structure belongs to the field of heliocentric orientation system design and manufacturing. The flange structure comprises a cabin interface flange body, various embedded components, precision adjustment gaskets for the TBA slewing mechanism assembly, support components for the drive locking mechanism assembly, precision adjustment gaskets for the toggle locking mechanism assembly, other precision adjustment gaskets, and an L-shaped circumferential reference plate. This structure can meet the 15-year lifespan requirement of the flat disc flange structure under the alternating high and low temperature conditions of a space station's heliocentric orientation system.

Description

Technical Field

[0001] The present invention belongs to the field of solar orientation device design and manufacturing, and relates to a large-size, high-rigidity integrated composite material flat disc flange structure suitable for large-scale solar orientation devices of space stations. Background Art

[0002] The space station's Heliocentric Orientation Mechanism (HOM) is the largest drive mechanism currently used in Chinese spacecraft. Traditionally, the HOM structure of spacecraft is mostly metal. To meet the space station's HOM requirements for lightweight design, the HOM was designed for the first time using composite materials. The cabin interface flange, which provides the primary mounting platform for individual units within the HOM, was designed as a large, high-rigidity, integrated composite flat disc flange to maximize lightweighting and stiffness.

[0003] The main internal components of the space station's solar guidance system (STAR), such as the TBA slewing support assembly, drive locking mechanism assembly, and toggle locking mechanism assembly, are mounted via the cabin interface flange within the STAR main structure. During launch, the cabin interface flange must withstand a load of approximately 460 kg (1400 lb) transmitted from the STAR's top truss end, as well as a load of approximately 613 kg (142 lb) from the STAR itself. While in orbit, the cabin interface flange must withstand the reaction forces transmitted from the TBA slewing support assembly, drive locking mechanism assembly, toggle locking mechanism assembly, and other components during in-orbit operation, while maintaining a 15-year operational lifespan. The stiffness and strength of this cabin interface flange are crucial to the rigidity and strength of the large-scale STAR.

[0004] This flat disc flange structure maximizes the regular layout of the space station's large-scale solar orientation system internal components and integrates the anisotropic properties of carbon fiber composites into an integrated design. Currently, it is the largest composite flat disc flange structure for a solar orientation system in a Chinese spacecraft. This flat disc flange structure offers strong adaptability, high rigidity, and excellent lightweighting.

[0005] During the molding process, the flat disc flange structure adopts special process control to solve the problem of internal defect control during the overall molding and demolding process of large-size, variable-section, and unequal-thickness carbon fiber composite materials, and can meet the 15-year on-orbit long-life requirement of the space station's solar-directed orientation device. Summary of the Invention

[0006] The technical problem solved by the present invention is to overcome the deficiencies of the prior art and propose a flat disc flange structure suitable for a large-scale solar orientation device of a space station.

[0007] In order to cooperate with the realization of the function of the cabin interface flange of the space station's solar orientation device, the solution of the present invention is: a composite material plane disc flange structure suitable for a large-scale solar orientation device of a space station, including a cabin interface flange body and a circular flange edge; the cabin interface flange body and the circular flange edge are an integrally formed carbon fiber composite material structure; the cabin interface flange body is a double-layer structure, and the double-layer structure has embedded parts inside to provide installation interfaces for various components of the solar orientation device; the bottom layer of the double-layer structure forms a plurality of centripetal I-beams along the circumferential direction by setting weight-reducing holes, and the top layer Precision adjustment gaskets are installed corresponding to the vertical beam positions of the centripetal I-beam, which are used to install and adjust the TBA rotation mechanism assembly and the switching locking mechanism assembly in the solar orientation device; the driving locking mechanism assembly support is installed on the top layer according to the position of the driving locking mechanism assembly in the solar orientation device; the weight reduction holes opened on the top layer meet the premise that the assembly installation position and stiffness are met, and the weight reduction holes on the bottom layer constitute through holes; multiple reference plates are installed on the edge of the circular flange to provide a circumferential reference to ensure the coaxial requirements of the internal components of the solar orientation device after installation; the cross beam of the centripetal I-beam is an arc-shaped beam facing the same center of a circle.

[0008] Preferably, the cabin interface flange body and the embedded parts are bonded together by structural adhesive; titanium alloy screws are added between the embedded parts providing installation interfaces for the TBA slewing support mechanism assembly, the drive locking mechanism assembly, and the switch locking mechanism assembly and the cabin interface flange body.

[0009] Preferably, the embedded parts and the precision adjustment gaskets are treated with phosphoric acid anodizing on the bonding surfaces, and the treated aluminum alloy embedded parts and the precision adjustment gaskets are ensured to be bonded within 72 hours.

[0010] Preferably, before bonding the flat disc flange structure to the embedded parts and the precision adjustment gasket, pre-treatment is required to remove the bright resin layer on the surface of the cabin interface flange body that is in contact with the release agent during its own molding process.

[0011] Preferably, the connection area between the outermost circumference of the cabin interface flange body and the flange surface where the resource cabin is installed is designed to be a C-shaped groove structure.

[0012] Preferably, according to the bearing strength requirements of different parts, the double-layer structure adopts a variable cross-section design; the thickness of the cabin interface flange body connected to the resource cabin is greater than the connection area and vertical surface between the C-groove and the resource cabin installation flange surface, i.e., the outermost circumferential thickness of the cabin interface flange body, greater than the I-beam on which the TBA slewing support mechanism assembly and the switching locking mechanism assembly are installed, and greater than the thickness of the area where the central thermal control adapter plate is installed.

[0013] Preferably, when laying layers in the variable cross-section area, the carbon fiber is laid in the thickness change area in an interlaced and decreasing manner, and concentrated mutations are not allowed. For each overall laying cycle, an interlaced and thickened cycle layer is added; wherein, the centripetal I-beam installed by the TBA rotary mechanism assembly and the switching locking mechanism assembly adopts a laying method that combines regional laying with overall laying; the centripetal I-beam installed by the TBA rotary mechanism assembly and the switching locking mechanism assembly is set with a certain proportion of 0° laying along the radial direction of the I-beam in combination with the stiffness requirements; the proportion of 0° laying in the overall laying is not less than 30%; the splicing of the fiber prepreg shall not be set at the corner during laying.

[0014] Preferably, the cabin interface flange body is pre-pressed during the lamination process according to the different thicknesses of the body. The number of pre-pressing times is determined by the thickness of the pre-pressed entity, and pre-pressing is performed at least once every 3 mm, with a pre-pressing temperature of 50-60°C.

[0015] Preferably, the cabin interface flange body needs to be operated in an environment with a temperature of 20±5°C and a humidity of ≤65% during the laying process, and the laying operation time shall not exceed 15 days.

[0016] Preferably, the surface layer of the cabin interface flange body is made of T300 / TDE-85 woven carbon cloth hot-melt prepreg, and the carbon fiber content in the cabin interface flange body is controlled at 60%±3%, which needs to be tested by body sampling.

[0017] Preferably, the carbon fiber material used for the driving locking mechanism assembly support is consistent with the cabin interface flange body.

[0018] Preferably, the cabin interface flange body is formed by using a metal mold, and for TDE-85 resin, the mold is placed in a curing oven at 90±5°C for heat sealing.

[0019] Preferably, during molding, curing is performed at 120±5°C for 2h±10min, and then at 170±5°C for 4h±10min; after curing, the body needs to be sampled and tested to determine whether the degree of curing is ≥90% and the porosity is ≤2%.

[0020] Preferably, sufficient demoulding space is reserved on both sides of each I-beam and the opening side of the C-shaped groove of the cabin interface flange body according to the mold design, and the demoulding angle is 3-5°.

[0021] Preferably, 3-5 mm is reserved between the prepreg edge and the mold edge on both sides of the I-beam and the C-shaped groove of the cabin interface flange body during layering.

[0022] Preferably, when designing a forming mold for the flange surface of the cabin interface flange body for installation at the edge and the resource cabin, a plurality of glue overflow holes need to be reserved along the circumferential direction.

[0023] The cabin interface flange body of the present invention is an integrally molded, integrated high-modulus carbon fiber / epoxy composite material structure, and an adaptive configuration design is performed based on the layout of the key mechanism unit to optimize the force transmission path; through integral molding, the connection links are reduced; the purpose of high rigidity and lightweight is achieved by optimizing the lay-up form; through the mold design, temperature and humidity control, glue overflow groove optimization setting, lay-up time control, pre-pressing, heat sealing and other methods in the overall molding process of the cabin interface flange body, the molding quality of the large-size integrated structure is guaranteed and the internal delamination defects are reduced; through the selection of materials, lay-up method, and thermal matching adaptability design of the connection form, the long life requirement of 15 years of the flat disc flange structure under the high and low temperature alternating conditions (-60 to +80°C) of the large-scale solar orientation device of the space station is met. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] Other features, objects and advantages of the present invention will become more apparent upon reading the detailed description of non-limiting embodiments made with reference to the following drawings:

[0025] Figure 1 This is a schematic diagram of the installation of the cabin interface flange in the main structure of the space station's solar orientation device;

[0026] Figure 2 This is a front view of the interface flange of the main structure cabin of the space station's solar orientation device;

[0027] Figure 3 This is a schematic diagram of the back side of the interface flange of the main structure cabin of the space station's solar orientation device;

[0028] Figure 4 It is a cross-sectional view. DETAILED DESCRIPTION

[0029] The following is combined with Figure 1-4 Further illustrate the present invention. The preferred embodiments of the present invention are only used to help illustrate the present invention and are not used to limit the scope of protection of the present invention. The preferred embodiments of the present invention do not describe all the details in detail, nor do they limit the invention to only the specific implementation methods described. Obviously, many modifications and changes can be made based on the contents of this specification. This specification selects and specifically describes these embodiments in order to better explain the principles and practical applications of the present invention, so that those skilled in the art can make good use of the present invention. The present invention is only limited by the claims and their full scope and equivalents. After reading the contents of the present invention, those skilled in the art can make various changes or modifications to the present invention, and these equivalent changes and modifications also fall within the scope defined by the claims of the present invention.

[0030] To meet the space station's requirements for lightweight, high rigidity, and large dimensions for its large-scale solar-orientation system, and to accommodate the layout of its internal components, an integrated carbon fiber composite flat disc flange structure was employed. Through specialized process control, internal defects during the integral molding and demolding processes for the large, variable-section, and unequal-thickness carbon fiber composites were eliminated, ensuring the 15-year on-orbit lifespan of the space station's solar-orientation system.

[0031] Combined with reference Figure 1 The present invention provides a large-scale, high-rigidity integrated composite flat disc flange structure for the cabin interface flange, located at the bottom of the main structure of the solar orientation device. It also provides an installation platform for other components of the main structure of the solar orientation device and the individual units within the solar orientation device.

[0032] Combined with reference Figure 2 The cabin interface flange provided by the present invention includes a cabin interface flange body 1, various embedded parts 2, a TBA rotary mechanism assembly installation precision adjustment gasket 3, a drive locking mechanism assembly support 4, a switch locking mechanism assembly installation precision adjustment gasket 5, other precision adjustment gaskets 6, and an L-shaped circumferential reference plate 7; wherein the cabin interface flange body 1 and the drive locking mechanism assembly support 4 are carbon fiber composite material structures.

[0033] Combined with reference Figure 2 The cabin interface flange provided by the present invention mainly provides installation interfaces for 8 TBA slewing mechanism assemblies, 2 drive locking mechanism assemblies, 2 switching locking mechanism assemblies, and installation surfaces for other components of the main structure of the solar orientation device. The accuracy of the installation surface is ensured by integral machining of the TBA slewing mechanism assembly installation precision adjustment gasket 3, the drive locking mechanism assembly support 4, the switching locking mechanism assembly installation precision adjustment gasket 5, and other precision adjustment gaskets 6. The above-mentioned precision adjustment gaskets are all made of aluminum alloy and are bonded to the cabin interface flange body 1 with structural adhesive. In order to avoid debonding between the metal parts and the carbon fiber composite material under high and low temperature alternation, and to ensure the 15-year long service life of the TBA slewing support mechanism assembly, the drive locking mechanism assembly, the switching locking mechanism assembly, etc. on orbit, titanium alloy screws are added between the embedded parts of this part of the installation interface and the cabin interface flange body. In order to ensure the reliability of the fixation between the precision adjustment gasket made of aluminum alloy and the cabin interface flange body 1 made of carbon fiber composite material under high and low temperature alternation;

[0034] Combined with reference Figure 2 , the bonding surface of all precision adjustment gaskets needs to be treated with phosphoric acid anodizing, and the treated aluminum alloy gaskets must be bonded within 72 hours;

[0035] Combined with reference Figure 2Before bonding the precision adjustment gasket to the cabin interface flange body 1 using structural adhesive, the bonding surface of the cabin interface flange body 1 must be pre-treated to remove the bright resin layer on the surface of the cabin interface flange body 1 that came into contact with the release agent during its molding process. Then, use a scraper to evenly apply the adhesive to the bonding area, ensuring that the adhesive layer thickness with the precision adjustment gasket is 0.05-0.2mm.

[0036] Combined with reference Figure 2 During the pasting process, the precision adjustment gasket can be tightened by using the screws on the connection holes or by using a special clamp to apply pressure until the adhesive overflows around the precision adjustment gasket. After tightening, clean up the overflowed adhesive to avoid forming excess materials;

[0037] Combined with reference Figure 2 , on the front of the cabin interface flange body 1 (the side where the TBA slewing support mechanism assembly is installed), near the outermost circumferential edge of the cabin interface flange body, an L-shaped circumferential reference piece 7 is set every 45° in the circumferential direction, for a total of 8 pieces. By integrally machining the 8 L-shaped circumferential reference pieces 7, a circumferential reference is provided for the overall machining of the cabin interface flange and the subsequent installation of the TBA slewing support mechanism assembly, etc., to ensure the coaxial requirements of each mechanism inside the large-scale solar orientation device after single-unit installation. The L-shaped circumferential reference piece 7 is made of aluminum alloy material. The L-shaped circumferential reference piece 7 is first glued to the cabin interface flange body 1 using structural adhesive J-133, and then a titanium alloy screw is added for connection;

[0038] Combined with reference Figure 3 The drive locking mechanism assembly support 4 is designed with a high modulus carbon fiber composite material body with a precision adjustment gasket added above it. This ensures the accuracy of the drive locking mechanism assembly mounting surface while meeting the thermal compatibility between the drive locking mechanism assembly support 4 and the cabin interface flange body 1 under the high and low temperature alternation conditions of 15 years on orbit. The high modulus carbon fiber material used in the drive locking mechanism assembly support 4 is consistent with the cabin interface flange body 1;

[0039] Combined with reference Figure 3 The cabin interface flange body 1 is an integrally formed high modulus carbon fiber composite material structure. Through integral molding, the connection links are reduced to achieve the purpose of lightweighting.

[0040] On the front side of the cabin interface flange, the mounting surface of the drive locking mechanism assembly is 47 mm lower than the mounting surface of the TBA slewing support mechanism assembly. In order to reduce the difficulty of integrated molding of the cabin interface flange body 1, this is achieved by rear-connecting the drive locking mechanism assembly support 4;

[0041] Combined with reference Figure 3To accommodate the installation of the solar orientation device on the conical space station resource module, the module interface flange body 1 is designed as an integral flat disc flange structure. Based on the evenly distributed layout of eight TBA slewing mechanism components spaced 45° apart and two switching locking mechanism components spaced 180° apart, a centripetal I-beam is set along the radial direction on the flange surface.

[0042] Combined with reference Figure 3 、 Figure 4 To accommodate the installation of other auxiliary components within the large-scale solar-directed directional control system, such as the thermal control adapter plate, an auxiliary I-beam is added to the cabin interface flange body 1. While ensuring the overall stiffness of the cabin interface flange body 1, the I-beams at different locations adopt different thicknesses to meet the requirements for local installation strength of a single unit. The thermal control adapter plate is installed on the back of the cabin interface flange (facing the resource module); the single-layer thickness of the cabin interface flange body at the resource module connection is 12mm, the thickness of the connection area between the C-groove and the resource module mounting flange surface and the vertical surface (the outermost circumference of the cabin interface flange body) is 9mm, the single-layer thickness of the I-beam for the TBA slewing support mechanism assembly and the switching locking mechanism assembly is 6mm, and the single-layer thickness of the area where the central thermal control adapter plate is installed is 3mm.

[0043] Combined with reference Figure 3 、 Figure 4 The connection area between the outermost circumference of the cabin interface flange body 1 and the flange surface is designed as a C-shaped groove structure. By adjusting the height of the C-shaped groove and the I-beam, the rigidity of the flat disc flange structure in the direction of the mechanism installation surface can be adjusted, thereby ensuring the rigidity of the large-scale solar-direction device during the launch and on-orbit segments.

[0044] Combined with reference Figure 3 The cabin interface flange body 1 adopts a high modulus carbon fiber composite material structure, with a maximum size of 1760mm to match the installation surface of the resource cabin, and a maximum height of 74mm to meet the rigidity requirements of the large-scale solar directional device in the launch and on-orbit segments;

[0045] Combined with reference Figure 3 Figure 4 Based on the mold design, sufficient demolding space must be reserved on both sides of each I-beam and the C-slot opening of the cabin interface flange body 1. A 3-5mm margin is typically reserved between the edges of the prepreg and the mold to facilitate demolding and serve as an overflow channel for the glue. This also reduces damage to the edges of the cabin interface flange body 1 during demolding due to adhesion between the mold and the cabin interface flange body 1 at the edges, and reduces partial layer defects near the demolding opening of the cabin interface flange body 1. The I-beam and C-slot are joined using mold blocks, with a demolding angle of 3-5°.

[0046] Combined with reference Figure 3The carbon fiber layup for the cabin interface flange body 1 utilizes a combination of regional and overall layup, aligned along the long axis of symmetry for the centripetal I-beams used to support the eight TBA slewing and two toggle locking assemblies. To meet stiffness requirements, the I-beams must have a certain proportion of carbon fiber laid along the long sides of the I-beams, with the zero-degree orientation aligned with the long sides. This proportion must account for no less than 30% of the overall layup.

[0047] Combined with reference Figure 3 The surface layer of the cabin interface flange body 1 is made of T300 / TDE-85 carbon fiber woven cloth hot-melt prepreg, which can effectively alleviate the splitting of the surface carbon fiber composite material when the cabin interface flange body 1 is punched;

[0048] Combined with reference Figure 3 When laying out layers in areas with varying cross-sections on the cabin interface flange body 1, high-modulus carbon fiber layers must be laid in areas with varying thicknesses using a staggered thickening method, with no concentrated sudden changes permitted. Typically, one staggered thickening cycle is added to each layup cycle.

[0049] Combined with reference Figure 3 The mold tooling for the cabin interface flange body 1 adopts a combined metal mold of blocks and an integral body, and uses screws distributed on the non-product edge of the mold to pressurize the upper and lower surfaces and vertical surfaces of the I-beam, C-shaped groove, etc. The process is as follows:

[0050] 1) During the layup of the cabin interface flange, first define the radial centripetal direction of the I-beam as 0°, then agree on the layup quadrant according to the quadrant definition of the large-scale solar orientation device;

[0051] 2) For I-beams or C-shaped channels of different sizes, first lay them out on the block mold according to the preset laying order. After each layer is laid out, use a tool to compact it firmly to remove bubbles.

[0052] 3) After the layup is completed, trim the prepreg at the edge of the mold so that there is a 3-5mm gap between the edge and the mold;

[0053] 4) After all the block molds such as I-beams and C-shaped grooves are laid out, vacuum pre-pressing treatment is carried out to ensure the dimensional accuracy before the blocks are molded;

[0054] 5) Laying continuous fibers on the front side of the cabin interface flange body 1 on the integral mold in parallel with the block laying;

[0055] 6) Place the laid-out blocks and fiber plies, starting from the center of the cabin interface flange body 1 to the edge, into the overall mold where the continuous fiber has been laid. Position the block mold and the overall mold and compact them.

[0056] 7) Then, the vertical surface of the C-shaped groove and the back surface of the cabin interface flange body 1 are laid in sequence;

[0057] 8) The outermost layer is covered with T300 woven fabric / TDE-85 hot melt prepreg;

[0058] 9) Use bolt pressure to close the mold and ensure that the front and rear positions of the mold are aligned when closing the mold;

[0059] 10) When closing the mold, place the mold into the curing oven at 900 +10 ℃ for heat sealing;

[0060] 12) After mold closing, cure at 120±5°C for 2h±10min, then at 170±5°C for 4h±10min. After curing, sample the body for testing the degree of cure, which must be ≥90% and the porosity ≤2%.

[0061] 13) After curing is completed, cool naturally and then demould according to the pre-set demoulding angle.

[0062] Combined with reference Figure 3 , the cabin interface flange body 1 requires continuous fiber at the structural corners, and fiber prepreg splicing is not allowed at the corners;

[0063] Combined with reference Figure 3 When designing the molding die for the flange surface where the cabin interface flange body 1 is installed at the edge and the resource cabin, multiple overflow holes need to be reserved along the circumferential direction. The position, number, and size of the overflow holes need to be determined based on the temperature during molding, the viscosity of the resin, the thickness of the composite component in the pressurized area, and the defect level that the composite component in this area needs to achieve. In this case, there are 8 overflow holes of φ16mm, which are used to improve the closed cavity structure of the mold to facilitate the flow of resin and the discharge of air between layers during the mold closing and curing process. This will improve the internal delamination defects of the cabin interface flange body 1 after molding;

[0064] Combined with reference Figure 3 To improve internal delamination defects in the fiber product after molding, the cabin interface flange body 1 requires a pre-pressing step during the layup process, based on the thickness of the different I-beams or C-shaped grooves at different locations of the cabin interface flange body 1, or the flange surface where it is mounted on the resource cabin. The number of pre-pressing steps is determined by the thickness of the pre-pressed body. Generally, pre-pressing is performed at least every 3 mm, at a temperature of 50-60°C.

[0065] Combined with reference Figure 3The cabin interface flange is large and has many layers. To ensure molding quality and reduce delamination defects, the layup process must be carried out in an environment with a temperature of 20±5°C and a humidity of ≤65%. The layup operation time must not exceed 15 days.

[0066] The present invention is applicable to a large-scale, high-rigidity integrated composite material flat disc flange structure for a large-scale solar orientation device of a space station. The main cabin interface flange body 1 of the flat disc flange structure is formed as a whole by metal die-casting after high-modulus carbon fiber composite material plies are laid according to the requirements of large size, high rigidity and lightweight of the large-scale solar orientation device. Its configuration is adaptively designed according to the interface restrictions of the connected resource cabin docking surface and the regular symmetrical layout of the single machine inside the solar orientation device. At the same time, it is designed with unequal wall thickness in different regions in combination with the requirements of load transfer at different positions of the main structure of the solar orientation device. While meeting the requirements of rigidity and strength, the purpose of lightweighting is achieved to the greatest extent. In the design process, the long service life of the main structure of the large-scale solar orientation device of the space station for 15 years in orbit and the matching of internal thermal stress under high and low temperature alternation are fully considered. Necessary control measures are added during the molding process to reduce internal delamination defects and ensure product molding quality.

[0067] Although the present invention has been disclosed above in terms of preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solutions of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solutions of the present invention shall fall within the scope of protection of the technical solutions of the present invention.

Claims

1. A composite material flat disc flange structure suitable for a large-scale solar orientation device in a space station, characterized by: It includes a cabin interface flange body and an annular flange edge; the cabin interface flange body and the annular flange edge are an integrally formed carbon fiber composite material structure; the cabin interface flange body is a double-layer structure, and the double-layer structure has embedded parts inside to provide installation interfaces for various components of the solar orientation device; the bottom layer of the double-layer structure forms a plurality of centripetal I-beams along the circumferential direction by setting weight-reducing holes, and the top layer corresponds to the vertical beam position of the centripetal I-beam and is installed with precision adjustment gaskets for installing and adjusting the TBA rotary mechanism assembly and the switching locking mechanism assembly in the solar orientation device; the driving locking mechanism assembly support is installed on the top layer according to the position of the driving locking mechanism assembly in the solar orientation device; the weight-reducing holes opened on the top layer under the premise of meeting the assembly installation position and stiffness constitute a through hole with the weight-reducing holes on the bottom layer; a plurality of reference plates are installed on the annular flange edge to provide a circumferential reference to ensure the coaxial requirements of various components inside the solar orientation device after installation; the crossbeam of the centripetal I-beam is an arc-shaped beam facing the same center of a circle.

2. The flat disc flange structure according to claim 1, characterized in that: The cabin interface flange body and the embedded parts are bonded together by structural adhesive; titanium alloy screws are added between the embedded parts that provide installation interfaces for the TBA slewing support mechanism assembly, the drive locking mechanism assembly, and the switch locking mechanism assembly and the cabin interface flange body.

3. The flat disc flange structure according to claim 1, characterized in that: The bonding surfaces of embedded parts and precision adjustment gaskets are treated with phosphoric acid anodizing, and the treated aluminum alloy embedded parts and precision adjustment gaskets are guaranteed to be bonded within 72 hours.

4. The flat disc flange structure according to claim 3, characterized in that: Before bonding the flat disc flange structure to the embedded parts and precision adjustment gaskets, pre-treatment is required to remove the bright resin layer on the surface of the cabin interface flange body that has come into contact with the release agent during its own molding process.

5. The flat disc flange structure according to claim 1, characterized in that: The connection area between the outermost circumference of the cabin interface flange body and the flange surface where the resource cabin is installed is designed to be a C-shaped groove structure.

6. The flat disc flange structure according to claim 5, characterized in that: According to the bearing strength requirements of different parts, the double-layer structure adopts a variable cross-section design; the thickness of the cabin interface flange body connected to the resource cabin is greater than the connection area and vertical surface between the C-groove and the resource cabin installation flange surface, that is, the outermost circumferential thickness of the cabin interface flange body, greater than the I-beam where the TBA slewing support mechanism assembly and the switching locking mechanism assembly are installed, and greater than the thickness of the area where the central thermal control adapter plate is installed.

7. The flat disc flange structure according to claim 6, characterized in that: When laying layers in variable cross-section areas, the carbon fiber is laid in the thickness change area in an interspersed and decreasing manner, and concentrated mutations are not allowed. For each overall laying cycle, an interspersed thickening cycle layer is added; among them, the centripetal I-beam installed by the TBA slewing mechanism assembly and the switching locking mechanism assembly adopts a laying method that combines regional laying and overall laying; the centripetal I-beam installed by the TBA slewing mechanism assembly and the switching locking mechanism assembly is set with a certain proportion of 0° laying along the radial direction of the I-beam in combination with the stiffness requirements; the proportion of 0° laying in the overall laying is not less than 30%; the splicing of fiber prepregs shall not be set at the corners during laying.

8. The flat disc flange structure according to claim 7, characterized in that: During the laying process of the cabin interface flange body, a pre-pressing process is added according to the different thicknesses of the body. The number of pre-pressing times is determined by the thickness of the pre-pressed entity, and pre-pressing is performed at least once every 3 mm. The pre-pressing temperature is 50-60°C.

9. The flat disc flange structure according to claim 7, characterized in that: During the laying process, the cabin interface flange body needs to be operated in an environment with a temperature of 20±5°C and a humidity of ≤65%, and the laying operation time shall not exceed 15 days.

10. The flat disc flange structure according to claim 1, characterized in that: The surface layer of the cabin interface flange body is made of T300 / TDE-85 woven carbon cloth hot-melt prepreg, and the carbon fiber content in the cabin interface flange body is controlled at 60%±3%, which needs to be tested by body sampling.

11. The flat disc flange structure according to claim 1, characterized in that: The carbon fiber material used for the drive locking mechanism assembly support is consistent with the cabin interface flange body.

12. The flat disc flange structure according to claim 5, characterized in that: The cabin interface flange body is formed by metal molds. For TDE-85 resin, the mold is placed in a curing oven at 90±5℃ for heat sealing.

13. The flat disc flange structure according to claim 5, characterized in that: During molding, cure at 120±5℃ for 2h±10min, and then cure at 170±5℃ for 4h±10min. After curing, the body needs to be sampled and tested to ensure that the degree of cure is ≥90% and the porosity is ≤2%.

14. The flat disc flange structure according to claim 12, characterized in that: Sufficient demoulding space is reserved on both sides of each I-beam and the opening side of the C-slot of the cabin interface flange body according to the mold design, and the demoulding angle is 3-5°.

15. The flat disc flange structure according to claim 12, characterized in that: When laying the layers, 3-5 mm is reserved on both sides of the I-beam and C-shaped groove of the cabin interface flange body at the edges of the prepreg and the mold.

16. The flat disc flange structure according to claim 12, characterized in that: When designing the flange surface forming mold for the cabin interface flange body at the edge and the resource cabin, multiple glue overflow holes need to be reserved along the circumferential direction.

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