A preparation method of a power cabin applied to a solar aircraft

Abstract

The application provides a power cabin applied to a solar aircraft, and relates to the technical field of aircraft manufacturing, and comprises a power cabin shell, a first connecting flange and a second connecting flange, one end of the power cabin shell is connected with a wing beam of the aircraft through the first connecting flange, and the other end of the power cabin shell is connected with a motor of the aircraft through the second connecting flange; the power cabin shell comprises a composite material layer plate shell and a sandwich rib, the composite material layer plate shell has a central axis, the central axis is a curve, and a foam sandwich rib is arranged on the inner wall of the composite material layer plate shell. Through the power cabin shell structure internally provided with the foam sandwich rib, the foam rib is integrally formed with the composite material layer plate, the power cabin shell is light in weight and high in rigidity, and the requirement of the solar aircraft on a low structure weight coefficient can be met.

Description

Technical Field

[0001] This specification relates to the field of aircraft technology, specifically to a power cabin for solar-powered aircraft and its manufacturing method. Background Technology

[0002] In the field of solar-electric technology, aircraft absorb energy during the day using solar panels, with excess energy stored in energy storage batteries. This stored energy powers the motors during nighttime flight, ensuring long-duration flight across day and night. A solar-electric system mainly includes solar panels that provide electricity; energy storage and power systems such as energy storage batteries, motors, and propellers. These components are primarily mounted on the wings via support structures.

[0003] In recent years, advanced composite materials have been widely used in the structures of solar-powered aircraft. Among them, composite stiffened panel structures have advantages such as integral molding and high load-bearing efficiency, which meet the needs of solar-powered aircraft to reduce structural weight and improve structural efficiency, and can be used as the support structure of their propulsion system.

[0004] For such reinforced closed structural components, the application and uniform distribution of molding pressure are particularly important. Autoclave processes are difficult to apply uniform pressure, and conventional thermal expansion methods are only suitable for parts with smooth internal cavities and no obvious structural features. Therefore, there is an urgent need for a simple and rapid method to prepare reinforced closed structural components. Summary of the Invention

[0005] In view of this, the embodiments of this specification provide a power cabin for use in solar-powered aircraft and a method for its fabrication, so as to achieve the purpose of fabricating a power cabin structure containing foam sandwich ribs.

[0006] The embodiments in this specification provide the following technical solutions:

[0007] A power pod for use in solar-powered aircraft, comprising:

[0008] The power compartment shell, the first connecting flange and the second connecting flange, one end of the power compartment shell is connected to the wing spars of the aircraft through the first connecting flange, and the other end of the power compartment shell is connected to the motor of the aircraft through the second connecting flange;

[0009] The power compartment shell includes a composite material laminate shell and sandwich ribs. The composite material laminate shell has a central axis, which is curved, and the sandwich ribs are set on the inner wall of the composite material laminate shell.

[0010] Furthermore, the composite material laminate shell includes a power compartment shell ply, which is one or more combinations of staggered 0° ply, 90° ply, -45° ply and 45° ply.

[0011] Furthermore, the cross-section of the sandwich rib is one or a combination of semi-circular, circular, elliptical, petal-shaped, square, and shell-shaped.

[0012] A method for fabricating a power compartment for a solar-powered aircraft, comprising the following steps:

[0013] Foam core reinforcement strips are laid inside the rib grooves of the silicone rubber core mold;

[0014] Following the layup sequence of the outer shell prepreg, the outer shell prepreg is laid layer by layer on the surface of the silicone rubber core mold to form a preform of the power compartment shell;

[0015] The preform of the power compartment shell and the silicone rubber core mold are placed together in the metal outer female mold for molding;

[0016] The molding auxiliary materials are wrapped around the outside of the metal outer female mold for molding, and then placed in an oven for heating and curing. After demolding, the cured power compartment blank is obtained.

[0017] After the solidified power compartment blank is trimmed and shaped, the power compartment shell is obtained;

[0018] A set of engine compartment assembly frames is set up according to the relative positions of the wing spars and the motors. Based on the engine compartment assembly frames, the engine compartment shell, the first connecting flange, and the second connecting flange are installed in sequence.

[0019] Furthermore, the preparation method also includes:

[0020] The position and diameter of the motor end of the power compartment shell are determined based on the interface position and diameter between the motor and the power compartment shell.

[0021] The position and diameter of the wing spars end of the power compartment shell are determined based on the interface location and diameter between the wing spars and the power compartment shell.

[0022] The motor end and the wing spars end are respectively used as the two ends of the power compartment shell. Under the premise of not interfering with the wing ribs and skin, the external profile of the power compartment shell is determined by software simulation calculation based on the load conditions of the power compartment shell.

[0023] Furthermore, the preparation method also includes:

[0024] The layup sequence of the prepreg material is determined by using the aerodynamic loads borne by the power compartment and the thrust and torque applied by the power system as load inputs, the allowable material value of the power compartment shell as boundary conditions, and minimizing the structural weight of the power compartment shell as the objective.

[0025] Furthermore, the preparation method also includes:

[0026] Based on the structural characteristics of the power compartment cavity, process clearance, thickness of silicone rubber and hollow metal core, and coefficient of thermal expansion, determine the size, shape, and position of the mold groove of the core mold.

[0027] Using injection molding, a core mold is used to fill the outer wall of a hollow metal core with expanded silicone rubber. After hardening, a silicone rubber core mold with multiple rib grooves is obtained.

[0028] Furthermore, the outer shell prepreg is a fiber-reinforced resin system.

[0029] Furthermore, the foam core reinforcement strips are one or more combinations of PMI foam, PU foam, and Nomax honeycomb.

[0030] Furthermore, the first metal external mold includes a first metal external mold and a second metal external mold. After the preform of the power compartment shell and the silicone rubber core mold are placed in the first metal external mold, the second metal external mold and the first metal external mold are snapped together and fixed.

[0031] Compared with the prior art, the beneficial effects that at least one technical solution adopted in the embodiments of this specification can achieve include at least:

[0032] This invention provides a power cabin shell structure with internal foam core ribs. The foam ribs and composite material layers are integrally formed, resulting in a lightweight and high-rigidity structure that meets the requirements of low structural weight coefficient for solar-powered aircraft. Attached Figure Description

[0033] To more clearly illustrate the technical solutions of the embodiments of this application, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0034] Figure 1 This is a schematic diagram of the overall structure of the power compartment according to an embodiment of the present invention;

[0035] Figure 2 This is a schematic diagram showing the connection between the power compartment, wing spars, and motor according to an embodiment of the present invention;

[0036] Figure 3 This is a schematic diagram of the power compartment shell structure according to an embodiment of the present invention;

[0037] Figure 4 This is an overall structural diagram and a sectional view of the core mold according to an embodiment of the present invention;

[0038] Figure 5 This is an overall structural diagram of the silicone rubber core mold according to an embodiment of the present invention;

[0039] Figure 6 This is a cross-sectional view of the silicone rubber core mold according to an embodiment of the present invention;

[0040] Figure 7 This is a schematic diagram of the structure of the metal outer female mold for molding according to an embodiment of the present invention;

[0041] Figure 8 This is a schematic diagram of the silicone rubber core mold placed inside the metal outer female mold for molding, according to an embodiment of the present invention;

[0042] Figure 9 yes Figure 8 Cross-sectional view of AA;

[0043] Figure 10 This is a schematic diagram of the power compartment being mounted on the power compartment assembly frame according to an embodiment of the present invention;

[0044] Figure 11 This is a schematic diagram of the power compartment shell layering according to an embodiment of the present invention.

[0045] Explanation of reference numerals in the attached drawings: 1. Engine compartment shell; 101. Composite material laminate shell; 102. Foam sandwich ribs; 2. First connecting flange; 3. Second connecting flange; 4. Wing spars; 5. Motor; 6. Silicone rubber core mold; 601. Expanded silicone rubber; 602. Hollow metal inner core; 603. Rib groove; 7. Metal outer female mold for molding; 701. First metal outer female mold; 702. Second metal outer female mold; 8. Engine compartment assembly frame; 9. Core mold; 901. Mold groove. Detailed Implementation

[0046] The embodiments of this application will now be described in detail with reference to the accompanying drawings.

[0047] The following specific examples illustrate the implementation of this application. Those skilled in the art can easily understand other advantages and effects of this application from the content disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of this application, and not all of them. This application can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed based on different viewpoints and applications without departing from the spirit of this application. It should be noted that, in the absence of conflict, the following embodiments and features in the embodiments can be combined with each other. Based on the embodiments in this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0048] It should be noted that various aspects of embodiments within the scope of the appended claims are described below. It will be apparent that the aspects described herein can be embodied in a wide variety of forms, and any particular structure and / or function described herein is merely illustrative. Based on this application, those skilled in the art will understand that one aspect described herein can be implemented independently of any other aspect, and two or more of these aspects can be combined in various ways. For example, any number and aspects set forth herein can be used to implement the device and / or practice the method. Additionally, this device and / or method can be implemented using structures and / or functionalities other than one or more of the aspects set forth herein.

[0049] It should also be noted that the illustrations provided in the following embodiments are only schematic representations of the basic concept of this application. The drawings only show the components related to this application and are not drawn according to the actual number, shape and size of the components in the actual implementation. In the actual implementation, the form, quantity and proportion of each component can be arbitrarily changed, and the layout of the components may also be more complex.

[0050] Furthermore, specific details are provided in the following description to facilitate a thorough understanding of the examples. However, those skilled in the art will understand that the described aspects can be practiced without these specific details.

[0051] The technical solutions provided by the various embodiments of this application are described below with reference to the accompanying drawings.

[0052] like Figure 1 , Figure 2 and Figure 3 As shown, a power pod for use in solar-powered aircraft includes:

[0053] The power cabin shell 1 comprises a first connecting flange 2 and a second connecting flange 3. One end of the power cabin shell 1 is connected to the wing spars of the aircraft via the first connecting flange 2, and the other end of the power cabin shell 1 is connected to the aircraft's motor via the second connecting flange 3. The power cabin shell 1 includes a composite material laminate shell 101 and foam sandwich ribs 102. The composite material laminate shell 101 has a central axis, which is curved. The foam sandwich ribs 102 are disposed on the inner wall of the composite material laminate shell 101. The cross-section of the foam sandwich ribs 102 is one or a combination of several of the following: semi-circular, circular, elliptical, petal-shaped, square, and shell-shaped.

[0054] The composite material laminate shell 101 includes a power compartment shell ply, which is one or more combinations of staggered 0° ply, 90° ply, -45° ply, and 45° ply. In some embodiments, such as Figure 11As shown, the layup sequence of the prepreg, from the outside to the inside, is 45° layup, -45° layup, 90° layup, and 0° layup. In other embodiments, the layup sequence of the prepreg, from the outside to the inside, may also be -45° layup, 45° layup, 0° layup, and 90° layup. The outermost 45° and -45° layups are used to withstand out-of-plane torsional loads, the middle 90° layup is used to withstand lateral loads, and the innermost 0° layup is used to withstand axial loads. The layup sequence for withstanding out-of-plane torsional loads can be changed, as can the layup sequence for withstanding lateral / axial loads.

[0055] A method for fabricating a power compartment for a solar-powered aircraft includes the following steps:

[0056] Step 1: Lay foam core reinforcement strips 102 inside the rib grooves 603 of the silicone rubber core mold 6.

[0057] In some embodiments, the foam core rib 202 is one or more combinations of PMI foam, PU foam, and Nomax honeycomb.

[0058] The second step involves laying the prepreg layers on the surface of the silicone rubber core mold 6 according to the layup sequence of the prepreg, thereby forming the preform of the power compartment shell 1.

[0059] Specifically, the layup sequence of the prepreg is determined by using the aerodynamic loads borne by the power compartment and the thrust and torque applied by the power system as load inputs, the allowable material value of the power compartment shell as boundary conditions, and minimizing the structural weight of the power compartment shell as the objective.

[0060] In some embodiments, the shell prepreg is a fiber-reinforced resin system.

[0061] The third step is to place the preform of the power compartment shell 1 and the silicone rubber core mold 6 into the metal outer female mold 7 for molding.

[0062] Specifically, such as Figure 7 , Figure 8 , Figure 9 As shown, the molding metal external mold 7 includes a first metal external mold 701 and a second metal external mold 702 (a split-type combination). After the preform of the power compartment shell 1 and the silicone rubber core mold 6 are placed in the first metal external mold 701, the second metal external mold 702 is snapped and fixed to the first metal external mold 701. The first metal external mold 701 and the second metal external mold 702 are made of mold steel. After demolding, the outer surface of the blank of the power compartment shell 1 is smooth, and the surface contour accuracy of the power compartment shell 1 is ensured by the molding metal external mold 7.

[0063] Step 4: Wrap the molding auxiliary materials around the outside of the metal outer female mold 7, and place it in an oven to heat and cure. After demolding, the cured power compartment blank is obtained.

[0064] Specifically, the molding accessories include release fabric, breathable felt, and vacuum bags. The release fabric, breathable felt, and vacuum bags are sequentially wrapped around the outside of the metal outer mold 7 used for molding.

[0065] Step 5: After shaping the solidified power compartment blank, the power compartment shell 1 is obtained;

[0066] Step 6: Set up a set of engine compartment assembly frames 8 according to the relative positions of the wing spars 4 and the motor 5. Using the engine compartment assembly frames 8 as a reference, install the engine compartment shell 1, the first connecting flange 2 and the second connecting flange 3 in sequence.

[0067] Before fabricating the power compartment, firstly, the outer contour of the power compartment shell 1 needs to be determined; secondly, a silicone rubber core mold 6 needs to be manufactured for fabricating the power compartment shell 1.

[0068] Specifically, the external outline of the power compartment hull 1 is determined through the following steps:

[0069] The position and diameter of the motor end of the power compartment shell 1 are determined based on the interface position and diameter between the motor 5 and the power compartment shell 1; the position and diameter of the wing spars end of the power compartment shell 1 are determined based on the interface position and diameter between the wing spars 4 and the power compartment shell 1; the motor end and the wing spars end are respectively taken as the two ends of the power compartment shell 1, and the outer contour of the power compartment shell 1 is determined by software simulation based on the load conditions of the power compartment shell 1, provided that there is no interference with the wing ribs and skin.

[0070] Specifically, the location of motor 5 is determined based on the overall input and limitations of the aircraft's propulsion system. The location of wing spars 4 is determined based on the load input (overall load) experienced by the aircraft. Once the positions of the two ends of the power nacelle (motor 5 and wing spars 4) are determined, the diameter of the motor end of the power nacelle shell 1 is determined by the diameter of motor 5, and the diameter of the wing spars end of the power nacelle shell 1 is determined in conjunction with the height of the wing spars. Finally, the shape of the middle section of the power nacelle shell 1 is determined through software simulation calculations based on the aerodynamic loads it experiences and the thrust generated by motor 5. Finite element method software can be used for this simulation.

[0071] Specifically, the silicone rubber core mold 6 is manufactured using the following steps:

[0072] Based on the structural characteristics of the power compartment cavity, process clearances, thickness of the silicone rubber and hollow metal core, and coefficient of thermal expansion, the dimensions and number of mold grooves 901 in the core mold 9 are determined; for example... Figure 4 , Figure 5 , Figure 6As shown, using injection molding, a core mold 9 is used to fill the outer wall of the hollow metal inner core 602 with expanded silicone rubber 601. After hardening, a silicone rubber core mold 6 with multiple rib grooves 603 is obtained. In some embodiments, the silicone rubber core mold 6 is made of two-component precision mold liquid silicone rubber and has a hollow metal inner core 602 embedded inside. This is used to adjust the linear shrinkage rate of the silicone rubber core mold 6 to ≤0.1%, ensuring that the thickness of the composite material layer shell 101 is uniform and the thickness tolerance is ±0.1 mm.

[0073] In some embodiments, the prepreg is cut and spliced ​​according to the fiber direction of the layup sequence to reduce fiber wrinkles on the surface of the composite material laminate shell 101 after molding.

[0074] like Figure 10 As shown, the power compartment assembly frame 8 includes a set of assembly frames for fixing and installing the power compartment shell 1, facilitating the installation of the first connecting flange 2 and the second connecting flange 3 at both ends.

[0075] The power cabin structure of this embodiment was applied to a certain type of solar-powered aircraft and its strength was verified. The results are as follows:

[0076] The composite material laminate shell 101 is made of T700 carbon fiber reinforced epoxy resin. The composite material laminate shell 101 adopts a layup sequence of [45 / -45 / 90 / 0] (45° layup, -45° layup, 90° layup, 0° layup from the inside to the outside) or [0 / 90 / -45 / 45] (0° layup, 90° layup, -45° layup, 45° layup from the inside to the outside). The foam core ribs 102 are made of PMI foam. The first connecting flange 2 and the second connecting flange 3 are made of aluminum alloy.

[0077] Loads (structural loads of 1.0G aerodynamic load, 230 N thrust from motor rotation, and 70 N•m axial torque) were applied to the power cabin structure of this embodiment. Inertial release was used as the boundary condition, and static strength was checked using the Nastran solver. Stress-strain results showed that the maximum tensile strain of the power cabin shell was approximately 475 µε, less than the allowable tensile strain of 5000 µε. The maximum compressive strain was -173 µε, less than the allowable compressive strain of -2000 µε. The maximum shear strain was 657 µε, less than the allowable shear strain of 5000 µε, indicating that the power cabin structure of this embodiment can meet the overall design and operational requirements of solar-powered aircraft.

[0078] Beneficial effects of the embodiments of the present invention:

[0079] The power cabin shell of this invention is an integrally molded structure of foam ribs and composite material layers, which is lightweight and high-rigidity, meeting the requirements of low structural weight coefficient for solar-powered aircraft. The power cabin shell in this invention is a rotating body; the prepreg material is cut according to the fiber direction of the layup design and layered, significantly improving the surface fiber wrinkles of the power cabin after molding. The surface quality and contour accuracy of the power cabin structure in this invention are ensured by the metal outer mold used for molding, eliminating the need for secondary processing and thus guaranteeing the overall rigidity and quality of the structure. The optimized preparation method based on silicone rubber thermal expansion molding in this invention improves upon existing processes. It uses a silicone rubber core mold with an embedded hollow metal core. The structure of the silicone rubber core mold is determined based on the internal cavity structure characteristics of the power cabin shell, process clearance, silicone rubber / hollow metal core thickness, and thermal expansion coefficient distribution. The power cabin shell produced by this method has uniform thickness and minimal structural deformation. Compared to autoclave processes, the preparation method based on silicone rubber thermal expansion molding has lower costs and is more suitable for preparing irregularly shaped power cabin shell structures with internal foam core ribs compared to ordinary thermal expansion methods.

[0080] The various embodiments in this specification are described in a progressive manner. Similar or identical parts between embodiments can be referred to interchangeably. Each embodiment focuses on describing the differences from other embodiments. In particular, the method embodiments described later are relatively simple in description because they correspond to the system; relevant parts can be referred to the descriptions in the system embodiments.

[0081] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A method for manufacturing a power compartment for a solar-powered aircraft, characterized in that, The power pod used in solar-powered aircraft includes: The power cabin shell (1), the first connecting flange (2) and the second connecting flange (3) are connected. One end of the power cabin shell (1) is connected to the wing spars (4) of the aircraft through the first connecting flange (2), and the other end of the power cabin shell (1) is connected to the motor (5) of the aircraft through the second connecting flange (3). The power compartment shell (1) includes a composite material laminate shell (101) and foam sandwich ribs (102). The composite material laminate shell (101) has a central axis, and the central axis is curved. The foam sandwich ribs (102) are arranged on the inner wall of the composite material laminate shell (101). The preparation method includes the following steps: Foam core ribs (102) are laid in the rib groove (603) of the silicone rubber core mold (6). Following the layup sequence of the outer shell prepreg, the outer shell prepreg is applied layer by layer to the surface of the silicone rubber core mold (6) to form a preform of the power compartment shell (1); The preform of the power compartment shell (1) and the silicone rubber core mold (6) are placed together in the metal outer female mold (7) for molding; The molding auxiliary materials are wrapped around the outside of the metal outer female mold (7) for molding, and then placed in an oven for heating and curing. After demolding, the cured power compartment blank is obtained. After the solidified power compartment blank is trimmed and shaped, the power compartment shell (1) is obtained. A set of power compartment assembly frames (8) are set up according to the relative positions of the wing spars (4) and the motor (5). Based on the power compartment assembly frames (8), the power compartment shell (1), the first connecting flange (2), and the second connecting flange (3) are installed in sequence.

2. The method for preparing a power compartment for a solar-powered aircraft according to claim 1, characterized in that, Also includes: The position and diameter of the motor end of the power compartment housing (1) are determined based on the interface position and diameter between the motor (5) and the power compartment housing (1); The position and diameter of the wing spars of the power compartment shell (1) are determined based on the interface position and diameter between the wing spars (4) and the power compartment shell (1); The motor end and the wing spars end are respectively used as the two ends of the power compartment shell (1). Under the premise of not interfering with the wing ribs and skin, the external outline of the power compartment shell (1) is determined by software simulation calculation based on the load conditions of the power compartment shell (1).

3. The method for preparing a power compartment for a solar-powered aircraft according to claim 1, characterized in that, Also includes: The aerodynamic load borne by the power compartment and the thrust and torque applied by the power system are used as load inputs, the allowable material value of the power compartment shell (1) is used as boundary conditions, and the minimum structural weight of the power compartment shell (1) is used as the target to determine the layup sequence of the prepreg material of the outer shell.

4. The method for preparing a power compartment for a solar-powered aircraft according to claim 1, characterized in that, Also includes: Based on the structural characteristics of the power compartment cavity, process clearance, thickness of silicone rubber and hollow metal core and coefficient of thermal expansion, determine the size, shape and position of the mold groove (901) of the core mold (9); Using injection molding, a core mold (9) is used to fill the outer wall of a hollow metal core (602) with expanded silicone rubber (601), and after hardening, a silicone rubber core mold (6) with multiple rib grooves (603) is obtained.

5. The method for preparing a power compartment for a solar-powered aircraft according to claim 1, characterized in that, The prepreg material for the outer shell is a fiber-reinforced resin system.

6. The method for preparing a power compartment for a solar-powered aircraft according to claim 1, characterized in that, The foam sandwich rib (102) is one or more of PMI foam, PU foam, and Nomax honeycomb.

7. The method for preparing a power compartment for a solar-powered aircraft according to claim 1, characterized in that, The metal external mold (7) for molding includes a first metal external mold (701) and a second metal external mold (702). After the preform of the power compartment shell (1) and the silicone rubber core mold (6) are placed in the first metal external mold (701), the second metal external mold (702) and the first metal external mold (701) are snapped together and fixed.

8. The method for preparing a power compartment for a solar-powered aircraft according to claim 1, characterized in that, The composite material laminate shell (101) includes a power compartment shell ply, wherein the power compartment shell ply is one or more combinations of staggered 0° ply, 90° ply, -45° ply, and 45° ply.

9. The method for preparing a power compartment for a solar-powered aircraft according to claim 1, characterized in that, The cross-section of the sandwich rib (102) is one or more of the following: semi-circular, circular, elliptical, petal-shaped, square, and shell-shaped.

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