A gas turbine stator assembly and a gas turbine

Abstract

The application discloses a gas turbine stator assembly and a gas turbine, and relates to the technical field of gas turbines, aiming to solve the problem of strength reduction caused by deformation of a turbine stator structure due to temperature stress. The gas turbine stator assembly comprises a turbine casing, a plurality of turbine stator vanes and a turbine stator disc. One end of the plurality of turbine stator vanes is fixedly connected to the inside of the turbine casing and extends radially from the inner circumferential surface of the turbine casing. The turbine stator disc is disc-shaped, a through hole is arranged in the center of the disc surface of the turbine stator disc, the through hole is used for penetrating a rotating shaft, the turbine stator vanes are fixedly connected to the outer circumferential surface of the turbine stator disc, and an annular groove is arranged on the disc surface of the turbine stator disc in the direction of a turbine rotor disc of the same stage, and the annular groove is concentrically arranged with the disc surface of the turbine stator disc. The gas turbine comprises the gas turbine stator assembly of the technical scheme. The gas turbine stator assembly provided by the application is used for weakening the influence of deformation of a turbine stator structure due to temperature stress on the strength of the turbine stator structure.

Description

Technical Field

[0001] This invention relates to the field of gas turbine technology, and more particularly to a gas turbine stator assembly and a gas turbine. Background Technology

[0002] New-generation cryogenic liquid rocket staged combustion cycle engines generally employ high-temperature, high-pressure, oxygen-rich gas to drive the turbine. Due to their inherent characteristics, the turbine components are arranged with alternating rotors and stators. The high-temperature, high-pressure, oxygen-rich gas expands and accelerates through the turbine stator, driving the turbine blades to rotate at high speed and output power. In a certain type of high-thrust staged combustion cycle engine, the cryogenic liquid oxygen flowing from the cryogenic oxygen pump acts on the turbine stator structure together with the high-temperature gas. The significant temperature difference between the high-temperature gas and the cryogenic liquid oxygen requires the turbine stator to withstand substantial temperature stress. This leads to deformation of the turbine stator structure, reducing its strength and impairing turbine performance. Summary of the Invention

[0003] The purpose of this invention is to provide a gas turbine stator assembly and a gas turbine, which reduces the deformation of the turbine stator structure caused by temperature stress and avoids the reduction in strength caused by structural deformation.

[0004] To achieve the above objectives, the present invention provides the following technical solution:

[0005] A gas turbine stator assembly, comprising:

[0006] Turbine casing;

[0007] Multiple turbine stator vanes, one end of which is fixedly connected to the inside of the turbine casing and extends radially from the inner circumference of the turbine casing to the turbine casing;

[0008] The turbine stationary disc is disc-shaped. A through hole is opened in the center of the disc surface for the turbine moving disc to pass through. The other end of the turbine stationary blade is fixed to the outer circumference of the turbine stationary disc. An annular groove is opened on the disc surface of the turbine stationary disc facing the same level turbine moving disc. The annular groove is concentric with the disc surface of the turbine stationary disc.

[0009] Optionally, in the above-mentioned gas turbine stator assembly, the radial distance between the groove edge of the annular groove away from the center of the turbine stator disk and the edge of the outer peripheral surface of the turbine stator disk is B, and B is 3mm-5mm larger than the blade thickness at the contact point between the turbine stator blade and the turbine stator disk.

[0010] Optionally, in the above-mentioned gas turbine stator assembly, the axial depth of the annular groove is 1 / 4 to 1 / 3 of the axial length of the turbine stator blade cascade composed of multiple turbine stator blades.

[0011] Optionally, in the above-mentioned gas turbine stator assembly, a cavity is formed around the inner circumference of the turbine stationary disk. The cavity is formed within the turbine stationary disk in a direction radially away from the turbine moving disk along the inner circumference of the turbine stationary disk. The gas turbine stator assembly is also provided with an annular boss. The annular boss is formed on the disk surface of the turbine stationary disk facing the turbine moving disk of the same level. The annular boss is concentrically arranged with the disk surface of the turbine stationary disk. The axial surface of the annular boss is passed through the axial extension surface of the axial surface of the cavity, which is located radially away from the edge of the turbine moving disk.

[0012] Optionally, in the above-mentioned gas turbine stator assembly, the axial height of the annular boss is 1.5mm-4mm.

[0013] Optionally, in the above-mentioned gas turbine stator assembly, the side wall of the annular boss away from the center of the disk is the groove wall of the annular groove near the center of the disk.

[0014] Optionally, in the gas turbine stator assembly described above, the turbine stator blades are electron-beam welded to the turbine casing.

[0015] Optionally, in the above-mentioned gas turbine stator assembly, the thickness of the turbine casing at the electron beam welding point is 1 / 2 to 2 / 3 of the axial length of the turbine stator blade cascade composed of multiple turbine stator blades.

[0016] The present invention also provides a gas turbine, including any of the gas turbine stator assemblies described above.

[0017] Compared to existing technologies that do not have annular grooves, the oxygen pump chamber connected to the inner circumferential surface of the turbine stator disk has a flow path for cryogenic liquid oxygen from the oxygen pump chamber adjacent to the inner circumferential surface and outer wall of the turbine stator disk. The significant temperature difference between the high-temperature gas flow path and the cryogenic liquid oxygen flow path causes the turbine stator disk structure to bear considerable temperature stress. In the gas turbine stator assembly provided by this invention, an annular groove concentric with the flow path of cryogenic liquid oxygen is provided on the disk surface of the turbine stator disk. The recessed groove surface, the outer circumferential surface of the turbine stator disk, and the disk surface of the turbine stator disk form a thin-walled structure. Compared to the overall structure of the turbine stator disk, the thin-walled structure is relatively isolated. In this way, a radially isolated thin-walled structure is actively constructed, allowing the temperature stress borne by the turbine stator disk to be actively released within this radially isolated thin-walled structure, reducing the deformation of the turbine stator structure caused by temperature stress, and avoiding the reduction in strength caused by structural deformation.

[0018] The present invention also provides a gas turbine, including any of the above-mentioned gas turbine stator components and a turbine rotor that cooperates with the gas turbine stator components.

[0019] Compared with the prior art, the beneficial effects of the gas turbine provided by the present invention are the same as those of the gas turbine stator assembly described in the above technical solutions, and will not be repeated here. Attached Figure Description

[0020] The accompanying drawings, which are included to provide a further understanding of the invention and form part of this invention, illustrate exemplary embodiments of the invention and are used to explain the invention, but do not constitute an undue limitation of the invention. In the drawings:

[0021] Figure 1 This is a schematic diagram of the turbine stator assembly and turbine moving disk assembly proposed in an embodiment of the present invention;

[0022] Figure 2 This is a schematic diagram of the structure of a gas turbine stator assembly proposed in an embodiment of the present invention.

[0023] Reference numerals in the attached figures: 1 is the turbine casing, 2 is the turbine stator vane, 3 is the turbine moving disc, 4 is the turbine stationary disc, 41 is the annular groove, 42 is the annular boss, and 5 is the cavity. Detailed Implementation

[0024] To make the technical problems to be solved, the technical solutions, and the beneficial effects of the present invention clearer, the present invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative of the present invention and are not intended to limit the present invention.

[0025] It should be noted that when a component is referred to as being "fixed to" or "set on" another component, it can be directly on or indirectly on that other component. When a component is referred to as being "connected to" another component, it can be directly connected to or indirectly connected to that other component.

[0026] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of this invention, "a plurality of" means two or more, unless otherwise explicitly specified. "Several" means one or more, unless otherwise explicitly specified.

[0027] In the description of this invention, it should be understood that the terms "upper", "lower", "front", "rear", "left", "right", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

[0028] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication of two components or the interaction between two components. Those skilled in the art can understand the specific meaning of the above terms in this invention according to the specific circumstances.

[0029] Please see Figure 1 The gas turbine stator assembly provided in this embodiment of the invention includes: a turbine casing 1, a plurality of turbine stator vanes 2 and a turbine stator disk 4. One end of the plurality of turbine stator vanes 2 is fixedly connected to the inside of the turbine casing 1 and extends radially from the inner circumference of the turbine casing 1. The turbine stator disk 4 is disc-shaped, and a through hole is provided at the center of the disk surface of the turbine stator disk 4 for passing through the turbine moving disk 3. The other end of the turbine stator vanes 2 is fixedly connected to the outer circumference of the turbine stator disk 4. An annular groove 41 is provided on the disk surface of the turbine stator disk 4 facing the same level turbine moving disk 3. The annular groove 41 is concentrically provided with the disk surface of the turbine stator disk 4.

[0030] In specific implementation: A concentric annular groove 41 is provided on the surface of the turbine stationary disk 4 that is in contact with the flow path of cryogenic liquid oxygen. The recessed groove surface of the annular groove 41, the outer peripheral surface of the turbine stationary disk 4, and the disk surface of the turbine stationary disk 4 form a thin-walled structure. Compared with the overall structure of the turbine stationary disk 4, the thin-walled structure is relatively isolated. In this way, a radially isolated thin-walled structure is actively constructed, so that the temperature stress borne by the turbine stationary disk 4 is actively released in this radially isolated thin-walled structure, reducing the deformation of the turbine stationary structure caused by temperature stress and avoiding the reduction in strength caused by structural deformation.

[0031] As one possible implementation, such as Figure 1As shown, the radial distance B between the groove edge of the annular groove 41 away from the center of the turbine stationary disk 4 and the edge of the outer peripheral surface of the turbine stationary disk 4 is 3mm-5mm larger than the blade thickness of the turbine stator 2 at the contact point with the turbine stationary disk 4. To better achieve the effect of actively releasing the temperature stress of the turbine stator structure, the radial distance between the starting position of the annular groove 41 and the outer peripheral surface of the turbine stationary disk 4 is set to a range, which is 3mm-5mm larger than the blade thickness of the turbine stator 2 at the contact point with the turbine stationary disk 4. The purpose of this setting is to control the radial wall thickness of the isolated thin-walled structure formed by the groove surface of the annular groove 41, the outer peripheral surface of the turbine stationary disk 4, and the axial disk surface of the turbine stationary disk 4, so as to better achieve the effect of releasing thermal stress while ensuring the strength of the isolated thin-walled structure. Compared with a turbine stator without the annular groove 41, the isolated thin-walled structure formed by the annular groove 41 has a 20% lower overall structural stress level, improving the reliability of the turbine stator structure.

[0032] As one possible implementation, such as Figure 1 As shown, the axial depth of the annular groove 41 is 1 / 4 to 1 / 3 of the axial length of the turbine stator vane cascade composed of multiple turbine stator vanes 2. It should be noted that when creating the annular groove 41, the distance between the bottom of the annular groove 41 and the cavity 5 needs to be considered, as structural strength must be taken into account. High-pressure liquid oxygen exists in the cavity 5, while the annular groove 41 contains relatively low-pressure, high-temperature combustion gas. If the distance between the bottom of the annular groove 41 and the cavity 5 is too close, the thin wall formed by the close distance will be subjected to a large pressure difference on both sides, leading to instability of the overall structure and easy deformation. If the distance between the bottom of the annular groove 41 and the cavity 5 is too far, that is, the axial depth of the annular groove 41 is too shallow, the isolated thin-walled structure formed by the groove surface of the annular groove 41, the outer peripheral surface of the turbine stationary disk 4, and the axial disk surface of the turbine stationary disk 4 will be too small, and it will not be able to play a good role in actively releasing temperature stress. The beneficial effect of setting the axial depth of the annular groove 41 to 1 / 4-1 / 3 of the axial length of the turbine stator blade cascade composed of multiple turbine stator blades 2 is that the stress level borne by the turbine stationary disk 4 is reduced by 20%, which will not weaken the structural strength, and effectively reduce the impact of the deformation of the turbine stator structure caused by temperature stress on the performance of the turbine stator structure.

[0033] As one possible implementation, a cavity 5 is provided around the inner circumference of the turbine stationary disk 4. The cavity 5 is provided within the turbine stationary disk 4 in a direction radially away from the turbine moving disk 3 along the inner circumference of the turbine stationary disk 4. The gas turbine stator assembly is also provided with an annular boss 42. The annular boss 42 is provided on the disk surface of the turbine stationary disk 4 facing the turbine moving disk 3 of the same level. The annular boss 42 is concentrically arranged with the disk surface of the turbine stationary disk 4. The axial surface of the annular boss 42 is passed through the axial extension surface of the axial surface of the cavity 5, which is radially away from the edge of the turbine moving disk 3. The annular boss 42 is formed on the surface of the turbine stationary disk 4 relative to the turbine moving disk 3 of the same stage, and is concentrically set with the surface of the turbine stationary disk 4. The annular boss 42 is formed around the center of the surface of the turbine stationary disk 4. Since the turbine stationary disk 4 has a disc-shaped structure, on the one hand, in order to make the structure have better strength, the boss is set as an annular boss 42, which is formed around the center of the surface of the turbine stationary disk 4. This maintains the continuity of the annular boss 42 and makes the structure have better strength. On the other hand, the purpose of setting the annular boss 42 is to improve the overall stiffness of the turbine stator structure and reduce the deformation of the turbine stator structure.

[0034] It should be noted that cavity 5 is the oxygen pump chamber, which is filled with low-temperature, high-pressure liquid oxygen. A detachable end-face seal structure is installed between the oxygen pump chamber and the turbine stationary plate 4. During stable engine operation, the high-pressure liquid oxygen in the oxygen pump flows to the end-face seal, where the seal detaches under the action of the high-pressure liquid oxygen. The leaked high-pressure liquid oxygen displaces the high-temperature combustion gas from the turbine chamber, preventing it from flowing back into the oxygen pump chamber and causing cavitation. The low-temperature, high-pressure liquid oxygen displaces the combustion gas and flows between the turbine stator and turbine blades, where it mixes and vaporizes with the high-temperature, high-pressure combustion gas on one side of the turbine stationary plate 4, driving the turbine blades to perform work. The turbine stator, situated between the oxygen pump chamber and the gap between the turbine stationary disk 4 and the turbine moving disk 3, forms a thin-walled structure. One side contacts low-temperature, high-pressure liquid oxygen, while the other side contacts high-temperature, low-pressure combustion gas. This thin-walled structure is subjected to mechanical stress under high pressure, resulting in an axial torsional force. Consequently, the radial surface of the turbine stationary disk 4, near the turbine moving disk 3, is at risk of dynamic-static friction with the moving disk 3. Therefore, an axial annular boss 42 is arranged on the surface of the turbine stationary disk 4 to thicken the easily deformable points of the thin-walled structure, increasing its rigidity and making deformation less likely. Compared to a turbine stator without the axial annular boss 42, the radial deformation under the combined effects of large temperature and pressure differences is reduced by 12.5%, reducing the risk of dynamic-static friction between the turbine stator and the shaft, and improving the safety margin of the radial working clearance between the turbine stator and the turbine moving disk 3.

[0035] It is worth mentioning that the temperature range of the high-temperature low-pressure gas is 400 degrees Celsius to 550 degrees Celsius, while the temperature range of the low-temperature high-pressure liquid oxygen is -160 degrees Celsius to -120 degrees Celsius, and the pressure difference between the two ranges from 6 MPa to 10 MPa.

[0036] Furthermore, the axial height of the annular boss 42 is 1.5mm-4mm. That is, the axial height of the annular boss 42 relative to the surface of the turbine stationary disk 4 is 1.5mm-4mm. As mentioned above, the purpose of setting the annular boss 42 is to increase rigidity and prevent the relatively weak thin-walled structure from deforming due to pressure and temperature differences. Here, it is proposed that the preferred height of the annular boss 42 is 1.5mm-4mm. Within this range, the height of the annular boss 42 can achieve a good effect of increasing rigidity. If it exceeds this range, the effect of increasing rigidity will not be significant, and it will affect the flow of low-temperature oxygen, thus failing to effectively block the high-temperature combustion gas from entering the disk cavity. It will also increase the weight of the structure, causing unnecessary weight burden on the operation of the structure. If the height of the annular boss 42 is lower than this range, the expected effect of increasing rigidity cannot be achieved well.

[0037] In some embodiments, such as Figure 2 As shown, the sidewall of the annular boss 42 away from the center of the disk surface is the groove wall of the annular groove 41 near the center of the disk surface. That is, if the radial surface of the annular groove 41 away from the turbine moving disk 3 is defined as the outer ring surface of the annular groove 41, and the radial surface of the annular groove 41 near the turbine moving disk 3 is defined as the inner ring surface of the annular groove 41, the inner ring surface extends and bends axially to form a sidewall of the annular boss 42. The purpose of this arrangement is to reduce the spatial location of the additional structure on the turbine stationary disk 4 to a smaller range, thereby reducing the impact on the structural strength of the turbine stator compared to setting it in two different locations.

[0038] As one possible implementation, the turbine stator 2 is electron-beam welded to the turbine casing 1. One end of the turbine stator 2 is electron-beam welded to the casing to form a turbine stator assembly. Electron-beam welding uses a high-energy electron beam as a processing heat source, bombarding the guide component and the area to be welded in the casing with a high-energy-density electron beam, causing it to melt rapidly, followed by rapid cooling to achieve the welding purpose. The weld seam after electron-beam welding is not prone to oxidation and has small thermal deformation. During engine operation, high-temperature combustion gases pass through the casing at the weld seam, so the weld seam of the casing is prone to thermal deformation. Due to the excellent performance of electron-beam welding, thermal deformation can be effectively prevented.

[0039] It is worth mentioning that the turbine stator blades 2 can be welded to the turbine casing 1 by other welding methods such as submerged arc welding.

[0040] Furthermore, the thickness of the turbine casing 1 at the electron beam welding point is 1 / 2 to 2 / 3 of the axial length of the turbine stator blade cascade composed of multiple turbine stator blades 2. When using electron beam welding, the structural thickness of the turbine casing 1 at the welding point must be considered. In this embodiment, electron beam welding is performed at the axial length of the turbine stator blade cascade, where the thickness of the turbine casing 1 is 1 / 2 to 2 / 3, to weld the turbine stator blade cascade to the turbine casing 1 into an assembly. The purpose of limiting the structural wall thickness of the turbine casing 1 at the welding point to this range is twofold: firstly, such a wall thickness ensures that the electron beam weld can be fully penetrated, guaranteeing the weld quality, and such an appropriate wall thickness also guarantees the structural strength of the turbine stator; secondly, considering that electron beam welding uses a high-energy electron beam, which releases a large amount of energy, such an appropriate wall thickness can control the welding energy of the electron beam weld, reducing the impact of the heat energy converted at the weld on the deformation of the turbine stator blade cascade structure, thereby preventing the weld heat energy from affecting the turbine starting performance.

[0041] This invention also provides a gas turbine, including the gas turbine stator assembly described in any of the above embodiments and a turbine rotor 3 that cooperates with the gas turbine stator assembly.

[0042] Compared with the prior art, the beneficial effects of the gas turbine provided in the embodiments of the present invention are the same as the beneficial effects of the gas turbine stator assembly provided in the above technical solutions, and will not be repeated here.

[0043] In the description of the above embodiments, specific features, structures, materials, or characteristics may be combined in any suitable manner in one or more embodiments or examples.

[0044] The above description is merely a specific embodiment of the present invention, but the scope of protection of the present invention 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 the present invention should be included within the scope of protection of the present invention. Therefore, the scope of protection of the present invention should be determined by the scope of the claims.

Claims

1. A gas turbine stator assembly characterized by, The gas turbine stator assembly comprises: a turbine casing; a plurality of turbine vanes, one end of which is fixedly connected to the inside of the turbine casing and extends radially from the inner circumferential surface of the turbine casing to the turbine casing; a turbine stator disc, which is disc-shaped, has a through hole in the center of the disc surface for passing through a turbine rotor disc, and has the other end of the turbine vanes fixedly connected to the outer circumferential surface of the turbine stator disc, and has an annular groove concentrically formed in the disc surface facing the turbine rotor disc; the groove of the annular groove away from the center of the disc surface of the turbine stator disc has a radial distance B from the edge of the outer circumferential surface of the turbine stator disc, and the B is greater than the blade thickness of the turbine vanes at the contact with the turbine stator disc by 3-5 mm; and the axial depth of the annular groove is 1 / 4-1 / 3 of the axial length of the turbine stator vane cascade composed of the plurality of turbine vanes.

2. The gas turbine stator assembly of claim 1, wherein, a recess cavity is formed along the circumference of the inner circumferential surface of the turbine stator disc, and the recess cavity is formed in the turbine stator disc in a direction radially away from the turbine rotor disc along the inner circumferential surface of the turbine stator disc; the gas turbine stator assembly is further provided with an annular boss, which is formed on the disc surface of the turbine stator disc facing the turbine rotor disc and is concentrically arranged with the disc surface of the turbine stator disc, and the axial surface of the annular boss is penetrated by the axial extension surface of the axial surface of the recess cavity in a direction radially away from the turbine rotor disc.

3. The gas turbine stator assembly of claim 2, wherein, The axial height of the annular boss is 1.5-4 mm.

4. The gas turbine stator assembly of claim 2, wherein, The side wall surface of the annular boss away from the center of the disc surface is the groove wall surface of the annular groove close to the center of the disc surface.

5. The gas turbine stator assembly of claim 1, wherein, The turbine vanes are electron beam welded to the turbine casing.

6. The gas turbine stator assembly of claim 5, wherein, The thickness of the turbine casing at the electron beam welding position is 1 / 2-2 / 3 of the axial length of the turbine stator vane cascade composed of the plurality of turbine vanes.

7. A gas turbine characterized by, The gas turbine stator assembly comprises any one of claims 1-6.

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