Liquid rocket engine axial flow turbine air inlet structure and assembling method

Abstract

The invention discloses an axial flow turbine air inlet structure of a liquid rocket engine, and belongs to the technical field of liquid rocket engines, a nozzle cascade ring, a gas collecting ring and an outlet flange are independently designed and manufactured, and in the production process, the turbine gas collecting ring and the nozzle cascade ring are assembled and welded into a first assembly; and then the first assembly and the outlet flange are assembled and welded into the air inlet structure of the axial flow turbine of the liquid rocket engine. According to the method, the machining difficulty can be effectively reduced, the high production quality and the high-precision molded surface of a product are ensured, the product is not prone to deformation after stress is eliminated through heat treatment, and the yield of the axial flow turbine air inlet structure of the liquid rocket engine is effectively improved.

Description

Technical Field

[0001] The invention relates to the technical field of liquid rocket engines, and in particular to an axial flow turbine air intake structure of a liquid rocket engine. Background Art

[0002] As the motive force of the turbopump of a pump-pressure liquid rocket engine, the turbine's main function is to accelerate the expansion of high-temperature and high-pressure combustion gas from the combustion component, generate high-speed airflow to drive the turbine blades to rotate, output shaft work, and provide power for the pump.

[0003] The hypersonic gas flow generated by the large expansion ratio of the impact turbine nozzle cascade can reach 2-3 Mach. Its profile contour and surface finish are key factors affecting the consistency of turbine efficiency. Therefore, the wall surface finish is required to be high enough, and the streamline and velocity distribution of the gas flow entering the gas collection ring and nozzle cascade ring must also be considered to ensure that the gas temperature is uniform and can effectively accelerate and generate enough power to drive the turbine to rotate.

[0004] Turbines usually operate in a complex environment of ultra-high temperature, high pressure and strong vibration, and have high requirements on their structural strength and deformation. Therefore, the structure of the gas collecting ring and nozzle blades needs to have sufficient strength and rigidity to withstand the high temperature and high pressure combustion gas generated by the combustion components to ensure their reliability under extreme working conditions.

[0005] In traditional turbine inlet casing assemblies, the turbine air collecting ring and nozzle blades are usually manufactured separately. The nozzle blade ring is designed as an integrated unit with the outlet flange and combined through a welding process; or the air collecting ring is designed as an integrated unit with the outlet flange and finally welded to the nozzle blade ring. In the existing scheme, the air collecting ring is a thin shell structure, while the flange blank is relatively thick and large. The nozzle blade ring profile has a key impact on the turbine performance. Whether it is precision casting or 3D printing, the structure is relatively complex, easy to deform after heat treatment, many product defects, poor contour accuracy and surface finish, and the profile is not easy to detect, and the performance consistency is poor after batch delivery. In addition, the welding structure of the turbine air collecting ring and the nozzle blade is usually welded by manual welding or automated welding technology to ensure the strength and sealing of the welded joint. In manual welding, welding is difficult and the welding stress is large, so it is easy to cause weld cracks and the nozzle blade profile is also easy to deform.

[0006] In view of this, it is necessary to provide a new technical solution to solve the above problems. Summary of the invention

[0007] In order to solve the above-mentioned technical problems, the present application provides an axial flow turbine air intake structure for a liquid rocket engine, which can effectively reduce the difficulty of processing, ensure high production quality and high-precision surface of the product, and is not easy to deform after heat treatment to eliminate stress, thereby effectively improving the finished product rate of the axial flow turbine air intake structure for the liquid rocket engine.

[0008] A liquid rocket engine axial flow turbine air intake structure, comprising: a turbine collecting ring, a nozzle blade ring and an outlet flange that are independently designed and manufactured; the turbine collecting ring comprises a curved pipe flange with a combustion gas guiding function; the inlet end of the turbine collecting ring is connected to the liquid rocket engine combustion assembly through the curved pipe flange; the inlet end of the nozzle blade ring is connected to the inlet end of the turbine collecting ring; the inlet end of the outlet flange is connected to the outlet end of the nozzle blade ring; the liquid rocket engine axial flow turbine air intake structure is formed by welding the turbine collecting ring, the nozzle blade ring and the outlet flange.

[0009] Preferably, the cross-sectional area of ​​the outlet end of the elbow flange is larger than the cross-sectional area of ​​the inlet end thereof; the inlet end of the elbow flange is smoothly transitionally connected to the outlet end thereof.

[0010] Preferably, the elbow flange has a joint structure for installing a temperature sensor.

[0011] Preferably, the turbine collecting ring and the nozzle blade ring have a positioning mechanism for axial positioning and radial positioning at the connection; the positioning mechanism includes a positioning shaft head arranged on the nozzle blade ring and located at its inlet end and a positioning sleeve arranged on the turbine collecting ring and located at its outlet end; the positioning shaft head can be inserted into the positioning sleeve.

[0012] Preferably, after the positioning shaft head is inserted into the positioning sleeve, the end face on the inlet side is flush with the inner wall on the outlet side of the gas collecting ring; the clearance between the positioning shaft head and the positioning sleeve in the radial direction ranges from 0.02 to 0.05 mm.

[0013] Preferably, the blades in the nozzle blade ring include a straight surface and a curved surface; a semi-Laval nozzle structure cross-sectioned along the axial plane is formed between the straight surface on the blade and the curved surface on the adjacent blade.

[0014] According to another aspect of the present application, a method for assembling an axial flow turbine air intake structure of a liquid rocket engine is provided, which is used for assembling the axial flow turbine air intake structure of the liquid rocket engine, comprising: Independently designed and manufactured turbine air collecting ring, outlet flange and nozzle cascade ring; Inspect the manufactured turbine air collecting ring, outlet flange and nozzle blade ring; Select the turbine collecting ring, outlet flange and nozzle blade ring that have passed the inspection, assemble and weld the turbine collecting ring and nozzle blade ring into a first component; assemble and weld the first component and the outlet flange into an axial turbine intake structure of a liquid rocket engine.

[0015] Preferably, the welding depth range of the turbine air collecting ring and the outlet flange is 4-7 mm; the welding depth range of the turbine air collecting ring and the nozzle blade ring is 4-7 mm; the welding depth range of the nozzle blade ring and the outlet flange is 3-5 mm.

[0016] Preferably, the distance between the welding point between the nozzle blade ring and the outlet flange and the nozzle blade ring blade outlet is in the range of 7-11 mm.

[0017] Preferably, the distance between the welding point between the turbine collecting ring and the nozzle blade ring and the blade profile outlet of the nozzle blade ring is in the range of 7-9 mm.

[0018] Compared with the prior art, this application has at least the following beneficial effects: 1. In the axial flow turbine intake structure of the liquid rocket engine of the present invention, the turbine gas collecting ring, the nozzle blade ring and the outlet flange are all independently designed and manufactured. The separate manufacturing method has a simple structure. Whether it is precision casting or 3D printing technology, it can ensure the high production quality and high-precision surface of the product. It is not easy to deform after heat treatment to eliminate stress. This not only reduces the difficulty of precision casting and the time of 3D printing, but also the abrasive flow process of the nozzle blade channel and the gas collecting ring channel is easy to implement and detect, which can further improve the smoothness of the turbine nozzle blade surface and ensure the consistency of performance.

[0019] 2. The assembly method of the axial-flow turbine intake structure of a liquid rocket engine in the present invention improves welding quality, reduces costs, and reduces deformation and cracks that may occur during welding. It is particularly suitable for mass production and can ensure the consistency of product quality and turbine performance.

[0020] 3. Compared with the traditional welded structure, the assembly method of the axial flow turbine intake structure of the liquid rocket engine in the present invention can achieve precise control of the weld and high-quality welded joints, thereby improving the quality and reliability of welding.

[0021] 4. Compared with the traditional welding structure, the assembly method of the axial flow turbine intake structure of the liquid rocket engine in the present invention can automate the welding process, improve efficiency, reduce manual intervention and production cycle, and further reduce costs; 5. The assembly method of the axial-flow turbine air intake structure of the liquid rocket engine of the present invention can reduce thermal deformation and stress concentration during welding, enhance the stability and durability of the welded joint, and effectively avoid cracks in the thin-walled part of the nozzle blade grid due to excessive welding stress. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] Hereinafter, some specific embodiments of the present invention will be described in detail in an exemplary and non-limiting manner with reference to the accompanying drawings. The same reference numerals in the accompanying drawings indicate the same or similar components or parts. It should be understood by those skilled in the art that these drawings are not necessarily drawn to scale. In the accompanying drawings: Figure 1 It is a three-dimensional structural schematic diagram of the axial flow turbine air intake structure of the liquid rocket engine of the present invention; Figure 2 It is a front structural schematic diagram of the axial flow turbine air intake structure of the liquid rocket engine of the present invention; Figure 3 It is the nozzle cascade ring blade profile diagram; Figure 4 for Figure 2 Sectional view at the AA position; Figure 5 for Figure 4 A partial enlarged view of position B in the middle.

[0023] The above drawings include the following reference numerals: 1. Joint structure; 2. Blades; 3. Throat; 4. Elbow flange; 5. Positioning sleeve; 6. Positioning shaft head; 7. Welding point between turbine collecting ring and outlet flange; 8. Welding point between turbine collecting ring and nozzle blade ring; 9. Welding point between nozzle blade ring and outlet flange; 21. Straight surface; 22. Blade body; 23. Curved surface; 100. Turbine collecting ring; 200. Nozzle blade ring; 300. Outlet flange. DETAILED DESCRIPTION

[0024] In order to make the purpose, technical solution and advantages of the present application clearer, the technical solution of the present application will be clearly and completely described below in combination with the specific embodiments of the present application and the corresponding drawings. Obviously, the described embodiments are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present application.

[0025] like Figure 1-Figure 5 As shown, an axial flow turbine air intake structure of a liquid rocket engine includes: a turbine air collecting ring 100, a nozzle blade ring 200 and an outlet flange 300 which are independently designed and manufactured. The turbine air collecting ring 100 includes a bend flange 4 which has a combustion gas guiding function. The inlet end of the turbine air collecting ring 100 is connected to the combustion assembly of the liquid rocket engine through the bend flange 4. The inlet end of the nozzle blade ring 200 is connected to the inlet end of the turbine air collecting ring 100. The inlet end of the outlet flange 300 is connected to the outlet end of the nozzle blade ring 200. The axial flow turbine air intake structure of the liquid rocket engine is welded by the turbine air collecting ring 100, the nozzle blade ring 200 and the outlet flange 300.

[0026] The cross-sectional area of ​​the outlet end of the elbow flange 4 is larger than the cross-sectional area of ​​the inlet end, and the inlet end of the elbow flange 4 is smoothly transitionally connected to the outlet end. Under this scheme, the cross-sectional area of ​​the elbow in the elbow flange 4 is gradually and evenly expanded, which can not only guide the flow of high-temperature and high-pressure gas, but also help to achieve low loss of high-temperature gas flow, smooth surface, uniform temperature distribution, and high structural strength.

[0027] As another embodiment of the present invention, the elbow flange 4 has a joint structure 1 for installing a temperature sensor. After the temperature sensor is installed in the joint structure 1, it can be used to measure the temperature uniformity of the combustion component.

[0028] The turbine air collecting ring 100 and the nozzle blade ring 200 have a positioning mechanism for axial positioning and radial positioning at the connection, and the positioning mechanism includes a positioning shaft head 6 provided on the nozzle blade ring 200 and located at its inlet end and a positioning sleeve 5 provided on the turbine air collecting ring 100 and located at its outlet end. The positioning shaft head 6 can be inserted into the positioning sleeve 5, thereby realizing the axial positioning and radial positioning of the nozzle blade ring 200 and the turbine air collecting ring 100 at their connection. In addition, the setting of the positioning mechanism can also prevent welding leakage from entering the air collecting ring cavity in the turbine air collecting ring 100, generating unnecessary excess.

[0029] Preferably, after the positioning shaft head 6 is inserted into the positioning sleeve 5, its inlet end face is flush with the inner wall of the gas collecting ring outlet side. The radial clearance between the positioning shaft head 6 and the positioning sleeve 5 is in the range of 0.02-0.05 mm.

[0030] The nozzle cascade ring 200 has a plurality of blades 2, and the blades 2 include a blade body 22, a straight surface 21, and a curved surface 23. The straight surface 21 and the curved surface 23 are respectively arranged on different surfaces of the blade body 22, wherein the curved surface 23 constitutes a blade basin, and in this embodiment, the blade basin is designed using a binary characteristic line method. A semi-Laval nozzle structure cross-sectioned along an axial plane is formed between the straight surface 21 on the blade and the curved surface 23 on the adjacent blade. The throat 3 is the smallest distance between adjacent blades 2.

[0031] It can be understood that the semi-Laval nozzle structure here is equivalent to half a Laval nozzle, which can effectively reduce shock wave losses.

[0032] In addition, in the present embodiment, the throat 8 adopts a smooth transition design, which can significantly improve the aerodynamic performance and efficiency of the supersonic nozzle blade grid. At the same time, under this structure, the area of ​​the throat 8 is easy to detect.

[0033] According to another aspect of the present invention, a method for assembling an axial flow turbine air intake structure of a liquid rocket engine is provided, which is used to assemble the axial flow turbine air intake structure of the liquid rocket engine, and comprises the following steps: Step S1, independently designing and manufacturing a turbine collecting ring 100, an outlet flange 300 and a nozzle blade ring 200.

[0034] The turbine collecting ring 100, the outlet flange 300 and the nozzle blade ring 200 are independently designed and manufactured separately, have a simple structure, are not easily deformed by heat treatment, and reduce the difficulty of precision casting and the time of 3D printing.

[0035] Under this scheme, the abrasive flow process of the nozzle blade ring 200 channel and the turbine air collecting ring 100 channel is easy to implement and detect, which can further improve the smoothness of the turbine nozzle blade surface, ensure the consistency of mass production performance, and make the manufacturing quality easy to ensure and detect. In addition, the turbine air collecting ring 100, the outlet flange 300 and the nozzle blade ring 200 are easy to assemble, the welding structure is simple and reliable, and the welding is not easy to crack.

[0036] It should be noted that the turbine collecting ring 100, the outlet flange 300 and the nozzle blade ring 200 can be processed by precision casting or 3D printing technology. If necessary, the above two solutions can be combined with machining.

[0037] In addition, after the turbine collecting ring 100, the outlet flange 300 and the nozzle blade ring 200 are independently manufactured, surface inspection, polishing and corresponding machining can be performed.

[0038] Step S2 , inspecting the manufactured turbine air collecting ring 100 , outlet flange 300 and nozzle blade ring 200 .

[0039] Step S3, selecting the turbine collecting ring 100, the outlet flange 300 and the nozzle blade ring 200 that have passed the inspection, assembling and welding the turbine collecting ring 100 and the nozzle blade ring 200 into a first assembly; assembling and welding the first assembly and the outlet flange 300 into an axial flow turbine intake structure of a liquid rocket engine.

[0040] Among them, the welding depth range of the turbine air collecting ring and the outlet flange welding point 7 is 4-7mm, the welding depth range of the turbine air collecting ring and the nozzle blade ring welding point 8 is 4-7mm, and the welding depth range of the nozzle blade ring and the outlet flange welding point 9 is 3-5mm, and the angle is 30°-70°. The distance between the nozzle blade ring and the outlet flange welding point 9 and the nozzle blade ring 200 blade outlet is 7-11mm, and the weld is far from the blade trailing edge of the weak part, avoiding the occurrence of welding cracks. The distance range of the turbine air collecting ring and the nozzle blade ring welding point 8 and the nozzle blade ring blade outlet is 7-9mm, and the weld is far from the blade trailing edge of the weak part, avoiding the occurrence of welding cracks.

[0041] In addition, after the first component and the outlet flange 300 are assembled and welded into the axial turbine intake structure of the liquid rocket engine, the turbine intake housing assembly can be finely processed to meet the operating conditions and improve the use effect.

[0042] For ease of description, spatially relative terms such as "above", "above", "on the upper surface of", "above", etc. may be used here to describe the spatial positional relationship between a device or feature and other devices or features as shown in the figure. It should be understood that spatially relative terms are intended to include different orientations of the device in use or operation in addition to the orientation described in the figure. For example, if the device in the accompanying drawings is inverted, the device described as "above other devices or structures" or "above other devices or structures" will be positioned as "below other devices or structures" or "below other devices or structures". Thus, the exemplary term "above" can include both "above" and "below". The device can also be positioned in other different ways (rotated 90 degrees or in other orientations), and the spatially relative descriptions used here are interpreted accordingly.

[0043] It should be noted that the terms used herein are only for describing specific embodiments and are not intended to limit the exemplary embodiments according to the present application. As used herein, unless the context clearly indicates otherwise, the singular form is also intended to include the plural form. In addition, it should be understood that when the terms "comprise" and / or "include" are used in this specification, it indicates the presence of features, steps, operations, devices, components and / or combinations thereof.

[0044] It should be noted that the terms "first", "second", etc. in the specification and claims of the present application and the above-mentioned drawings are used to distinguish similar objects, and are not necessarily used to describe a specific order or sequence. It should be understood that the numbers used in this way can be interchanged where appropriate, so that the embodiments of the present application described herein can be implemented in an order other than those illustrated or described herein.

[0045] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, the present invention may have various modifications and variations. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A liquid rocket engine axial flow turbine intake structure, characterized in that: include: The turbine gas collecting ring, nozzle blade ring and outlet flange are designed and manufactured independently of each other; the turbine gas collecting ring includes a bent pipe flange with a gas guiding function; the inlet end of the turbine gas collecting ring is connected to the liquid rocket engine combustion assembly through the bent pipe flange; the inlet end of the nozzle blade ring is connected to the inlet end of the turbine gas collecting ring; the inlet end of the outlet flange is connected to the outlet end of the nozzle blade ring; the axial flow turbine air intake structure of the liquid rocket engine is welded by the turbine gas collecting ring, the nozzle blade ring and the outlet flange.

2. The liquid rocket engine axial flow turbine air intake structure according to claim 1, characterized in that: The cross-sectional area of ​​the outlet end of the elbow flange is larger than the cross-sectional area of ​​the inlet end thereof; the inlet end of the elbow flange is smoothly transitionally connected with the outlet end thereof.

3. The liquid rocket engine axial flow turbine air intake structure according to claim 2, characterized in that: The elbow flange has a joint structure for installing a temperature sensor.

4. The liquid rocket engine axial flow turbine air intake structure according to claim 1, characterized in that: The turbine collecting ring and the nozzle blade ring have a positioning mechanism for axial positioning and radial positioning at the connection; the positioning mechanism includes a positioning shaft head arranged on the nozzle blade ring and located at its inlet end and a positioning sleeve arranged on the turbine collecting ring and located at its outlet end; the positioning shaft head can be inserted into the positioning sleeve.

5. The liquid rocket engine axial flow turbine air intake structure according to claim 4, characterized in that: After the positioning shaft head is inserted into the positioning sleeve, the end face on the inlet side is flush with the inner wall on the outlet side of the gas collecting ring; the clearance between the positioning shaft head and the positioning sleeve in the radial direction ranges from 0.02 to 0.05 mm.

6. The liquid rocket engine axial flow turbine air intake structure according to claim 1, characterized in that: The blades in the nozzle blade ring include a straight surface and a curved surface; a semi-Laval nozzle structure cross-sectioned along an axial plane is formed between the straight surface on the blade and the curved surface on the adjacent blade.

7. A method for assembling an axial flow turbine intake structure of a liquid rocket engine, characterized in that: Used to assemble the liquid rocket engine axial flow turbine air intake structure according to any one of claims 1 to 6, comprising: Independently designed and manufactured turbine air collecting ring, outlet flange and nozzle cascade ring; Inspect the manufactured turbine air collecting ring, outlet flange and nozzle blade ring; Select the turbine collecting ring, outlet flange and nozzle blade ring that have passed the inspection, assemble and weld the turbine collecting ring and nozzle blade ring into a first component; assemble and weld the first component and the outlet flange into an axial turbine intake structure of a liquid rocket engine.

8. The method for assembling the axial flow turbine air intake structure of a liquid rocket engine according to claim 7, characterized in that: The welding depth range of the turbine air collecting ring and the outlet flange welding place is 4-7mm; the welding depth range of the turbine air collecting ring and the nozzle blade ring welding place is 4-7mm; the welding depth range of the nozzle blade ring and the outlet flange welding place is 3-5mm.

9. The method for assembling the axial flow turbine air intake structure of a liquid rocket engine according to claim 7, characterized in that: The distance between the welding point between the nozzle blade ring and the outlet flange and the nozzle blade ring blade outlet is in the range of 7-11 mm.

10. The method for assembling the axial flow turbine air intake structure of a liquid rocket engine according to claim 7, characterized in that: The distance between the welding point between the turbine collecting ring and the nozzle blade ring and the blade profile outlet of the nozzle blade ring is in the range of 7-9mm.

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