A launch vehicle interstage and method of determining parameters

Abstract

The application provides a launch vehicle interstage section and a parameter determination method, and relates to the application field of launch vehicle interstage section structures.The launch vehicle interstage section comprises: a flange of a turn-up edge arranged at a first end of a cabin section; a flange of a mouth groove arranged at a second end of the cabin section; the flange of the turn-up edge is connected with the first end of the cabin section through a first transition section; the flange of the mouth groove is connected with the second end of the cabin section through a second transition section; and the cabin section, the flange of the turn-up edge, the flange of the mouth groove, the first transition section and the second transition section are integrally formed.The scheme of the application realizes effective transmission and bearing of the load of the launch vehicle interstage section structure, and simultaneously reduces the weight of the launch vehicle interstage section structure to the maximum extent.

Description

Technical Field

[0001] This invention relates to the field of rocket stage interstage structure applications, and in particular to a method for determining the parameters of a launch vehicle stage interstage. Background Technology

[0002] The interstage section is one of the main load-bearing structures of a launch vehicle, responsible for docking with adjacent upper and lower stages, bearing various external loads during rocket transport and flight, and also housing corresponding instruments and equipment. Metal interstage sections have been successfully applied in various types of launch vehicles, with high levels of structural design methods and reliability. Composite materials, due to their lower structural mass, are of great significance for reducing rocket weight, and the application research of carbon fiber composite interstage sections for launch vehicles has also received considerable attention. Because composite materials are anisotropic, their load-bearing and transfer methods and molding processes differ significantly from those of metal structures, thus requiring targeted structural design and molding processes. Summary of the Invention

[0003] This invention provides a method for determining the interstage section and parameters of a launch vehicle, which solves the problem of excessive weight of the interstage section structure in the prior art.

[0004] To solve the above-mentioned technical problems, the technical solution of the present invention is as follows:

[0005] This invention provides a launch vehicle stage intersection, comprising:

[0006] hull section;

[0007] A flange installed at the first end of the cabin section;

[0008] A slotted flange is installed at the second end of the cabin section;

[0009] The flanged flange is connected to the first end of the cabin section via a first transition section;

[0010] The slotted flange is connected to the second end of the cabin section via a second transition section;

[0011] The cabin section, flange, grooved flange, first transition section, and second transition section are integrally formed.

[0012] Optionally, the connection between the cabin section and the flanged flange forms a first preset angle, the first transition section is disposed at the first preset angle, and the first transition section has a first preset thickness value.

[0013] Optionally, the connection between the cabin section and the slot flange forms a second preset angle, the second transition section is disposed at the second preset angle, and the second transition section has a second preset thickness value.

[0014] Optionally, the launch vehicle stage interstage further includes:

[0015] Openings provided on the cabin section;

[0016] A reinforcement zone is provided at a predetermined position of the opening;

[0017] The reinforcement area has a third preset thickness value.

[0018] Optionally, a metal structural member is provided on the flange.

[0019] Optionally, the slotted flange is provided with a plurality of first grooves;

[0020] A through hole is provided in the first groove.

[0021] The present invention also provides a method for determining parameters of a launch vehicle interstage section, wherein the interstage section structure is any one of the above-described interstage section structures, and the determination method includes:

[0022] Obtain the cross-sectional load value of the flange;

[0023] Obtain the cross-sectional load value of the grooved flange;

[0024] Obtain the thickness parameters of the cabin section;

[0025] Based on the thickness parameters of the cabin section and the cross-sectional load value of the flange, the thickness parameters of the first transition section are determined.

[0026] Based on the thickness parameters of the cabin section and the cross-sectional load value of the slot flange, the thickness parameters of the second transition section are determined.

[0027] Optionally, it also includes:

[0028] The buckling value of the cabin structure is determined based on the cross-sectional load values ​​of the flanged flange, the cross-sectional load values ​​of the slotted flange, and the thickness parameters of the cabin section.

[0029] Optionally, determining the thickness parameters of the first transition section based on the thickness parameters of the cabin section and the cross-sectional load value of the flange includes:

[0030] Based on the cross-sectional load value of the flange and the thickness parameters of the cabin section, the first preset included angle is determined;

[0031] The thickness parameters of the first transition section are determined by the first preset angle.

[0032] Optionally, determining the thickness parameters of the second transition section based on the thickness parameters of the cabin section and the cross-sectional load value of the slotted flange includes:

[0033] The second preset included angle is determined based on the cross-sectional load value of the grooved flange and the thickness parameters of the cabin section;

[0034] The thickness parameters of the second transition section are determined by the second preset angle.

[0035] The above-described solution of the present invention has at least the following beneficial effects:

[0036] The interstage section of a launch vehicle according to the present invention includes: a cabin section; a flanged flange disposed at a first end of the cabin section; a slotted flange disposed at a second end of the cabin section; the flanged flange is connected to the first end of the cabin section via a first transition section; the slotted flange is connected to the second end of the cabin section via a second transition section; the cabin section, the flanged flange, the slotted flange, the first transition section, and the second transition section are integrally formed. This achieves effective transfer and bearing of the structural load of the interstage section of the launch vehicle, while minimizing the weight of the interstage section structure. Attached Figure Description

[0037] Figure 1 This is a perspective view of the interstage section of the launch vehicle according to the present invention;

[0038] Figure 2 yes Figure 1 Enlarged view at point C;

[0039] Figure 3 yes Figure 1 An enlarged view at point A;

[0040] Figure 4 yes Figure 1 An enlarged view at point B;

[0041] Figure 5 This is an enlarged schematic diagram of the opening reinforcement hole in the interstage section of the launch vehicle according to the present invention;

[0042] Figure 6 This is a flowchart of the method for determining the parameters of the interstage section of a launch vehicle according to the present invention;

[0043] Explanation of reference numerals in the attached figures:

[0044] 1. Cabin section; 2. Groove flange; 3. Opening; 4. Flanged flange; 5. Reinforcement area; 6. Metal structural component; 7. First groove; 8. Through hole; 9. Screw; 10. Support plate nut; 11. Longitudinal rib; 12. Transition section. Detailed Implementation

[0045] Exemplary embodiments of the present disclosure will now be described in more detail with reference to the accompanying drawings. While exemplary embodiments of the present disclosure are shown in the drawings, it should be understood that the present disclosure may be implemented in various forms and should not be limited to the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the disclosure to those skilled in the art.

[0046] like Figure 1 As shown, an embodiment of the present invention proposes a launch vehicle stage interstage section, comprising:

[0047] Section 1 of the cabin;

[0048] The flange 4 is installed at the first end of the cabin section 1;

[0049] The slotted flange 2 is installed at the second end of the cabin section 1;

[0050] The flange 4 is connected to the first end of the cabin section 1 via a first transition section;

[0051] The slotted flange 2 is connected to the second end of the cabin section 1 via a second transition section;

[0052] The cabin section 1, flange 4, groove flange 2, first transition section and second transition section are integrally formed.

[0053] In this embodiment, by setting grooved flanges 2 and flanged flanges 4 at both ends of the cabin section 1, the rocket stage inter-section structure is formed by integral molding. Furthermore, a structural design using a mesh-skin composite material structure is adopted. Under the premise of process feasibility, this achieves effective transfer and bearing of structural loads while minimizing the weight of the structure. In this embodiment, the first transition section is located near the flanged flange, and the second transition section is located near the grooved flange.

[0054] Specifically, section 1 of the cabin is made of composite material. Due to the load-bearing requirements of the connection at both ends, the connection areas at both ends (the locations of the flanged flange or the slotted flange) need to be structurally thickened.

[0055] With the connection areas at both ends thickened, the cylindrical section of the cabin is relatively thin. Therefore, the connection areas at both ends will form a region of stiffness discontinuity and stress concentration with the cylindrical section of cabin 1. Under axial compressive load, the connection areas at both ends will experience local radial deformation, causing the overall structure to buckle under a lower axial compressive load.

[0056] Between the thickened areas at both ends and the cylindrical structure of section 1, a transition section 12 with a certain thickness is provided. The axial length of this transition section is not less than 20 times the thickness gradient, where the thickness gradient is the thickness difference between the thickened areas at both ends and other unthickened areas.

[0057] The longitudinal reinforcement 11 is circumferentially dispersed in the thickness transition section 12 to gradually reduce its height. Along the axial direction, this creates a sloping downward trend in the height of the longitudinal reinforcement 11, with a slope angle not exceeding 30°. Figure 2 As shown, the transition section is thicker closer to the flange.

[0058] After the longitudinal reinforcement 11 is broken up, it is inserted into the interior of the slope transition layer, so that the ends of the longitudinal reinforcement 11 and the transition section 12 form a strong connection under the curing process.

[0059] The proportion of circumferential ply in the thickness transition section 12 is increased to improve the region's resistance to radial deformation. The proportion of circumferential ply thickness is not less than 40% of the total thickness.

[0060] Secondly, due to the local radial deformation caused by the axial compressive load, the two ends of the cabin section 1 are subjected to different loads. In the part with smaller axial tensile load, the flanged flange connection structure is selected; in the part with larger axial tensile load, the slotted flange connection structure is selected.

[0061] In an optional embodiment of the present invention, the connection between the cabin section 1 and the flange 4 forms a first preset angle, the first transition section is disposed at the first preset angle, and the first transition section has a first preset thickness value.

[0062] In this embodiment, the cabin section 1 and the flange 5 are formed using a continuous layering method, thereby improving the load transfer capacity of the flange structure, such as... Figure 3 As shown, the flange 4 structure and the cabin section 1 adopt a continuous layering molding method, thereby improving the structural load transfer capacity.

[0063] In this embodiment, a flange 4 is used to connect the parts with smaller axial tensile loads. Therefore, a certain angle is formed between the flange 4 structure and the first end of the cabin section 1. The thickness of this angle is determined by the thickness of the cabin section 1 cylinder and the structure formed by the flange 4. The continuous layering molding method is still used.

[0064] In an optional embodiment of the present invention, the connection between the cabin section 1 and the slot flange 2 forms a second preset angle, the second transition section is disposed at the second preset angle, and the second transition section has a second preset thickness value, such as... Figure 4 As shown.

[0065] In this embodiment, the grooved flange 2 is a preformed composite material solid structure. The solid structure ply body and the cylinder body of the cabin section 1 are formed continuously. The grooved flange 2 and the end face area are plyed with circumferential ply to improve the tensile strength of the grooved flange 2 area. Continuous ply wrapping structures are interspersed between the circumferential ply, so that the circumferential ply and the wrapping structure are formed as one, effectively dispersing and transferring the axial load of the opening area to the non-opening area.

[0066] A metal gasket structure is added to area 2 of the grooved flange. With the help of the rigidity of the metal gasket structure, the concentrated load at the local point of action is distributed to the end face of the entire opening area, reducing the local concentrated force and improving the overall load-bearing capacity of the structure.

[0067] In an optional embodiment of the present invention, the interstage section of the launch vehicle further includes:

[0068] Opening 3 is provided on the cabin section 1;

[0069] A reinforcement area 5 is provided at a preset position of the opening 3;

[0070] The reinforcing area 5 has a third preset thickness value.

[0071] In this embodiment, the opening location (opening 3) of the cabin section 1 needs local structural reinforcement to eliminate the impact of the pre-buckling at the local opening location on the overall resistance to axial compression instability, such as... Figure 5 As shown.

[0072] Opening 3 is located in the center of the area surrounded by the mesh reinforcement.

[0073] Local reinforcement of the opening is achieved by increasing the thickness of the composite material skin surrounding the corresponding mesh reinforcement area. Additionally, a reinforcement zone 5 of a certain thickness is provided around the opening 3, and the thickness of the reinforcement zone 5 is also achieved through a layering process, thus ensuring that the opening 3 can withstand a certain range of loads.

[0074] The reinforcing layer is inserted into the mesh reinforcement layer around the perimeter of the enclosure and co-cured to improve the structural connection strength and overall stiffness of the area.

[0075] The thickness of the local opening reinforcement is iteratively confirmed through finite element buckling calculation to ensure that the overall instability load meets the design requirements.

[0076] In an optional embodiment of the present invention, a metal structural member 6 is disposed on the flange 4.

[0077] In this embodiment, a metal connection structure 6 is provided at the flange position of the flange. The metal connection 6 adopts a double-hole parallel mode. The screw 9 connects the support plate nut 10 to the cabin section 1 and the rocket launch section, thereby improving the torsional resistance of the connection hole and reducing the compressive stress of the metal parts on the composite material structure.

[0078] In an optional embodiment of the present invention, a plurality of first grooves 7 are provided on the grooved flange 2;

[0079] A through hole 8 is provided in the first groove 7.

[0080] In this embodiment, the cabin section 1 is connected to the rocket launch section by means of screws through the through holes 8 provided in the multiple first grooves 7.

[0081] like Figure 6 As shown, the present invention also provides a method for determining the parameters of a launch vehicle interstage section, wherein the interstage section structure is as described in any of the above claims, and the determination method includes:

[0082] Step 11, obtain the cross-sectional load value of the flange;

[0083] Step 12, obtain the cross-sectional load value of the grooved flange;

[0084] Step 13: Obtain the thickness parameters of the cabin section;

[0085] Step 14: Determine the thickness parameters of the first transition section based on the thickness parameters of the cabin section and the cross-sectional load value of the flange.

[0086] Step 15: Determine the thickness parameters of the second transition section based on the thickness parameters of the cabin section and the cross-sectional load value of the slot flange.

[0087] In an optional embodiment of the present invention, step 13 may include:

[0088] Step 131: Determine the buckling value of the cabin section structure based on the cross-sectional load values ​​of the flanged flange, the cross-sectional load values ​​of the slotted flange, and the thickness parameters of the cabin section.

[0089] In this embodiment, local radial deformation occurs under the action of the cross-sectional load values ​​of the flange and the grooved flange, resulting in overall buckling of the entire structure under a lower axial compressive load, thereby obtaining the buckling value of the cabin section structure.

[0090] In an optional embodiment of the present invention, step 14 may include:

[0091] Step 141: Determine the first preset included angle based on the cross-sectional load value of the flange and the thickness parameters of the cabin section;

[0092] Step 142: Determine the thickness parameter of the first transition section using the first preset angle.

[0093] In this embodiment, based on the cross-sectional load value of the flange and the cabin section forming a certain curvature, the curvature and the thickness between the cabin section form a first preset angle. At this angle, by adopting a continuous lay-up forming method, the first transition section has a certain thickness parameter, thereby obtaining the thickness parameter of the first transition section.

[0094] In an optional embodiment of the present invention, step 15 may include:

[0095] Step 151: Determine the second preset included angle based on the cross-sectional load value of the slot method and the thickness parameters of the cabin section;

[0096] Step 152: Determine the thickness parameters of the second transition section using the second preset angle.

[0097] In this embodiment, based on the cross-sectional load value of the slotted flange and the cabin section forming a certain curvature, the curvature and the thickness between the cabin section form a second preset angle. At this angle, by adopting a continuous lay-up forming method, the second transition section has a certain thickness, thereby obtaining the thickness parameter of the second transition section.

[0098] This paper proposes a method for determining the interstage section and parameters of a launch vehicle, involving a cabin section; a flanged flange at the first end of the cabin section; a slotted flange at the second end of the cabin section; the flanged flange is connected to the first end of the cabin section via a first transition section; the slotted flange is connected to the second end of the cabin section via a second transition section; the cabin section, flanged flange, slotted flange, first transition section, and second transition section are integrally formed. Under the premise of technological feasibility, this method achieves effective load transfer and bearing of the interstage section structure of the launch vehicle, while minimizing the weight of the interstage section structure.

[0099] It should be noted that the above-described embodiments are all applicable to the embodiments of this method and can achieve the same technical effect.

[0100] The above description represents the preferred embodiments of the present invention. It should be noted that those skilled in the art can make various improvements and modifications without departing from the principles of the present invention, and these improvements and modifications should also be considered within the scope of protection of the present invention.

Claims

1. A stage section of a launch vehicle, characterized in that, include: Cabin section (1); The flange (4) is installed at the first end of the cabin section (1); The slotted flange (2) is installed at the second end of the cabin section (1); The flange (4) is connected to the first end of the cabin section (1) through a first transition section; The slotted flange (2) is connected to the second end of the cabin section (1) via a second transition section; The cabin section (1), flange (4), groove flange (2), first transition section and second transition section are integrally formed; The connection between the cabin section (1) and the flange (4) forms a first preset angle, the first transition section is set at the first preset angle, and the first transition section has a first preset thickness value; The connection between the cabin section (1) and the slot flange (2) forms a second preset angle, the second transition section is set at the second preset angle, and the second transition section has a second preset thickness value; Also includes: An opening (3) is provided on the cabin section (1); A reinforcement area (5) is provided at a preset position of the opening (3); The reinforcing area (5) has a third preset thickness value; Metal structural component (6) installed on the flange (4); The grooved flange (2) is provided with a plurality of first grooves (7); The first groove (7) is provided with a through hole (8).

2. A method for determining parameters of the interstage section of a launch vehicle, characterized in that, The rocket stage interstage structure is the rocket stage interstage structure as described in claim 1, and the determination method includes: Obtain the cross-sectional load value of the flange; Obtain the cross-sectional load value of the grooved flange; Obtain the thickness parameters of the cabin section; Based on the thickness parameters of the cabin section and the cross-sectional load value of the flange, the thickness parameters of the first transition section are determined. Based on the thickness parameters of the cabin section and the cross-sectional load value of the slot flange, the thickness parameters of the second transition section are determined.

3. The method for determining the parameters of the interstage section of a launch vehicle according to claim 2, characterized in that, Also includes: The buckling value of the cabin section structure is determined based on the cross-sectional load values ​​of the flanged flange, the cross-sectional load values ​​of the slotted flange, and the thickness parameters of the cabin section.

4. The method for determining the parameters of the interstage section of a launch vehicle according to claim 2, characterized in that, The determination of the thickness parameters of the first transition section based on the thickness parameters of the cabin section and the cross-sectional load value of the flange includes: Based on the cross-sectional load value of the flange and the thickness parameters of the cabin section, the first preset included angle is determined; The thickness parameters of the first transition section are determined by the first preset angle.

5. The method for determining the parameters of the interstage section of a launch vehicle according to claim 2, characterized in that, The determination of the thickness parameters of the second transition section based on the thickness parameters of the cabin section and the cross-sectional load value of the slotted flange includes: The second preset included angle is determined based on the cross-sectional load value of the grooved flange and the thickness parameters of the cabin section; The thickness parameters of the second transition section are determined by the second preset angle.

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