Cooling structure for turbine stator blade of gas turbine
By designing a multi-channel structure and cooling mode in the turbine stator blades of the gas turbine, the problem of the limited number of times the cold air flow direction can be changed was solved, the cooling efficiency and heat transfer capacity were improved, and uniform cooling of the blades and stable operation at high temperatures were achieved.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2026-02-02
- Publication Date
- 2026-03-27
AI Technical Summary
In the existing internal cooling structure of gas turbine stator blades, the number of times the cold air flow direction changes is limited, resulting in low heat transfer efficiency.
A cooling structure for the turbine stator blades of a gas turbine is designed, which adopts a multi-layer channel structure within the blade body, including a first channel, a second channel, and a third channel, each with a straight and staggered rib structure. The channels are divided into multiple sub-channels by partition ribs. The design combines the impact cooling and convection cooling modes of the sleeve to optimize the flow path of the cold air.
It enhances the turbulence of cold air inside the blade, improves cooling efficiency, ensures uniform cooling and heat transfer in all areas of the blade, and extends the service life of the blade.
Smart Images
Figure CN121738702A_ABST
Abstract
Description
TECHNICAL FIELD
[0001] The present application relates to the field of gas turbine technology, and particularly relates to a gas turbine turbine vane cooling structure. BACKGROUND
[0002] As a core power equipment in the field of high-end equipment manufacturing, the working efficiency of the gas turbine is directly related to the operating temperature, and increasing the working temperature has become the core direction of the development of gas turbine technology. The turbine blade is a key component that withstands high-temperature working medium impact in the turbine structure of the gas turbine, and its cooling efficiency directly determines the overall working capacity and operation reliability of the gas turbine. Therefore, the blade cooling technology is one of the core research topics in the field of gas turbine.
[0003] Cooling technology mainly includes two directions of external cooling and internal cooling. The external cooling method mainly includes film cooling, thermal barrier coating, porous material, etc. The internal cooling method mainly includes disturbance convection cooling, jet impact cooling, etc.
[0004] The internal cooling mainly adopts the staggered rib structure, and the staggered rib structure is mostly integral or only partitioned in the axial direction. The number of changes in the cold gas flow direction is limited, the turbulence degree is insufficient, and the heat transfer efficiency is low. SUMMARY
[0005] The purpose of the present application is to solve the problem that the internal cooling structure mainly adopts the staggered rib structure, the number of changes in the cold gas flow direction is limited, and the heat transfer efficiency is low. Further, a gas turbine turbine vane cooling structure is provided.
[0006] The technical scheme of the present application is: a gas turbine turbine vane cooling structure, comprising: a blade body;
[0007] The blade body has a first channel, a second channel and a third channel which are sequentially communicated from the top to the tail, and the first channel and the third channel are respectively provided with in-line rib structures, and the second channel is provided with a staggered rib structure;
[0008] The in-line rib structure in the first channel is located in the leading edge region of the blade body, the staggered rib structure in the second channel is located in the middle region of the blade body, and the in-line rib structure in the third channel is located in the trailing edge region of the blade body;
[0009] A plurality of first partition ribs which are spaced apart and arranged along the air flow direction are arranged in the second channel, and the second channel is divided into a plurality of second sub-channels by the plurality of first partition ribs.
[0010] Further, the second channel comprises a front channel and a rear channel, the front channel is located near one side of the leading edge region of the blade body, and the rear channel is located near one side of the trailing edge region of the blade body, and the number of second sub-channels in the rear channel is greater than that in the front channel.
[0011] Further, the staggered rib structure in the second channel is formed by a plurality of first inclined ribs arranged in a first direction and a plurality of second inclined ribs arranged in a second direction, and the first direction and the second direction are perpendicular.
[0012] Further, a plurality of the first inclined ribs are located on the suction surface, a plurality of the second inclined ribs are located on the pressure surface, or a plurality of the first inclined ribs are located on the pressure surface, and a plurality of the second inclined ribs are located on the suction surface.
[0013] Further, it further comprises a sleeve mounted on the front end of the blade body, and the middle of the sleeve has a cavity, and the cavity is communicated with the first channel through the impact hole.
[0014] Further, the leading edge region of the blade body is provided with a boss for fixing the sleeve.
[0015] Further, the in-line rib structure in the first channel is distributed on the outside of the sleeve.
[0016] Further, a plurality of second partition ribs are arranged in the first channel, and the first channel is divided into a plurality of first sub-channels by the plurality of second partition ribs, and the number of the first sub-channels is less than that of the second sub-channels.
[0017] Further, a plurality of the second partition ribs are distributed on the outside of the sleeve.
[0018] Further, a plurality of third partition ribs are arranged in the third channel, and the third channel is divided into a plurality of third sub-channels by the plurality of third partition ribs, and the number of the third sub-channels is greater than that of the second sub-channels.
[0019] Compared with the prior art, the present application has the following effects:
[0020] 1、The gas turbine turbine vane cooling structure provided by the present application adopts in-line rib structure in the leading edge region and the trailing edge region of the blade, which can ensure the smoothness of the cold air flow, reduce the flow resistance, meet the basic cooling demand of the region, adopt staggered rib structure in the middle region, cooperate with a plurality of first partition ribs to form a plurality of second sub-channels, increase the number of changes of the cold air flow direction in the blade, enhance the turbulence degree of the cold air in the blade, and further enhance the heat transfer capacity of the cold air and increase the cooling efficiency. BRIEF DESCRIPTION OF DRAWINGS
[0021] Figure 1 is the overall structure schematic diagram of the present application;
[0022] Figure 2 is Figure 1 the longitudinal section view after the sleeve is hidden in the present application.
[0023] In the figure: 1, blade body; 2, first channel; 3, second channel; 4, third channel; 5, first partition rib; 6, front channel; 7, rear channel; 8, first inclined rib; 9, second inclined rib; 10, sleeve; 11, cavity; 12, impact hole; 13, boss; 14, second partition rib; 15, third partition rib. DETAILED DESCRIPTION
[0024] Specific implementation method one: combined Figure 1 , Figure 2 This embodiment is described, and the embodiment includes a blade body 1, the blade body 1 has a first channel 2, a second channel 3 and a third channel 4 which are sequentially communicated from the top to the tail, the first channel 2 and the third channel 4 are respectively provided with in-line rib structures, and the second channel 3 is provided with staggered rib structures, the in-line rib structure in the first channel 2 is located in the leading edge region of the blade body 1, the staggered rib structure in the second channel 3 is located in the middle region of the blade body 1, and the in-line rib structure in the third channel 4 is located in the trailing edge region of the blade body 1, the second channel 3 is provided with a plurality of first partition ribs 5 which are spaced apart and arranged along the air flow direction, the second channel 3 is divided into a plurality of second sub-channels through the plurality of first partition ribs 5, the spacing direction of the plurality of first partition ribs 5 is perpendicular to the cold air flow direction, and the arrangement direction of the plurality of first partition ribs 5 is parallel to the cold air flow direction.
[0025] The gas turbine turbine vane cooling structure of the embodiment adopts the in-line rib structure in the leading edge region and the trailing edge region, can ensure the smoothness of the cold air flow, reduce the flow resistance, meet the basic cooling requirement of the region at the same time, adopts the staggered rib structure in the middle region, cooperates with the plurality of first partition ribs 5 to form the plurality of second sub-channels, can increase the number of changes of the cold air flow direction in the blade, enhances the turbulence degree of the cold air in the blade, and further enhances the heat transfer capacity of the cold air and increases the cooling efficiency.
[0026] Specific implementation method two: combined Figure 1 , Figure 2The difference between the embodiment and the first specific embodiment is that the second channel 3 includes a front channel 6 located near the leading edge region of the blade body 1 and a rear channel 7 located near the trailing edge region of the blade body 1, and the number of second sub-channels in the rear channel 7 is greater than that in the front channel 6. As the cold gas flows from the front end to the rear end, its temperature rises after absorbing heat, and the cooling capacity gradually decreases. The rear channel 7 is provided with more sub-channels, which can further refine the cold gas flow path, increase the heat transfer area, compensate for the attenuation of the cooling capacity of the cold gas, and ensure the overall cooling effect of the second channel 3 is balanced. The other components and connection relationships are the same as those in the first specific embodiment.
[0027] The third specific embodiment is a combination of the first specific embodiment and the second specific embodiment. Figure 1 Figure 2 The difference between the embodiment and the first specific embodiment is that the second channel 3 includes a front channel 6 located near the leading edge region of the blade body 1 and a rear channel 7 located near the trailing edge region of the blade body 1, and the number of second sub-channels in the rear channel 7 is greater than that in the front channel 6. As the cold gas flows from the front end to the rear end, its temperature rises after absorbing heat, and the cooling capacity gradually decreases. The rear channel 7 is provided with more sub-channels, which can further refine the cold gas flow path, increase the heat transfer area, compensate for the attenuation of the cooling capacity of the cold gas, and ensure the overall cooling effect of the second channel 3 is balanced. The other components and connection relationships are the same as those in the first specific embodiment.
[0028] The fourth specific embodiment is a combination of the first specific embodiment and the third specific embodiment. Figure 1 Figure 2 The difference between the embodiment and the third specific embodiment is that the first inclined ribs 8 are located on the suction surface, and the second inclined ribs 9 are located on the pressure surface, or the first inclined ribs 8 are located on the pressure surface, and the second inclined ribs 9 are located on the suction surface. The inclined ribs are arranged on the suction surface and the pressure surface of the blade, respectively, to ensure that the cold gas can be disturbed in the entire area of the channel cross section, avoid the problem of local smooth flow and weak heat transfer caused by arranging inclined ribs on a single wall, realize overall heat transfer enhancement in the channel, and improve the uniformity of the blade wall temperature distribution. The other components and connection relationships are the same as those in the third specific embodiment.
[0029] The fifth specific embodiment is a combination of the first specific embodiment and the fourth specific embodiment. Figure 1 Figure 2 The difference between the embodiment and the specific embodiment one is that the sleeve 10 is installed on the front end of the blade body 1, the middle of the sleeve 10 has a cavity 11, the cavity 11 is communicated with the first channel 2 through the impact hole 12, the cold air in the cavity 11 of the sleeve 10 impacts the inside of the first channel 2 at high speed through the impact hole 12, a jet impact cooling effect is formed, a large amount of heat in the leading edge region can be quickly taken away, the impact + convection composite cooling mode is formed by cooperating with the convection cooling of the in-line rib in the first channel 2, the cooling efficiency of the leading edge region is greatly improved, and the service life of the blade is prolonged. The other components and connection relationships are the same as those in the specific embodiment one.
[0030] Specific embodiment six: combination Figure 1 , Figure 2 The difference between the embodiment and the specific embodiment five is that the blade body 1 is provided with a boss 13 for fixing the sleeve 10, the boss 13 is in contact with the outer side of the sleeve 10, the boss 13 provides stable mounting support for the sleeve 10, avoids loosening or displacement of the sleeve 10 under the harsh working conditions of high speed, high temperature and high pressure of the gas turbine, ensures the communication precision of the impact hole 12 and the first channel 2, guarantees the stability and reliability of the impact cooling effect, and improves the mechanical strength of the overall structure. The other components and connection relationships are the same as those in the specific embodiment five.
[0031] Specific embodiment seven: combination Figure 1 , Figure 2 The difference between the embodiment and the specific embodiment five is that the in-line rib structure in the first channel 2 is distributed on the outer side of the sleeve 10, the in-line rib structure not only avoids interference with the sleeve 10 and the impact hole 12, fully utilizes the internal space of the first channel 2, but also forms a synergistic effect with the impact cooling of the sleeve 10, the cold air sprayed out of the impact hole 12 forms an orderly convection flow under the guidance of the in-line rib after impacting the channel wall, reduces the flow resistance, ensures that the cooling airflow covers the entire first channel 2, and improves the uniformity and efficiency of the cooling of the leading edge region. The other components and connection relationships are the same as those in the specific embodiment five.
[0032] Specific embodiment eight: combination Figure 1 , Figure 2This embodiment differs from specific embodiment seven in that it has multiple second partition ribs 14 spaced apart and arranged along the airflow direction within the first channel 2. These second partition ribs 14 divide the first channel 2 into multiple first sub-channels, with fewer first sub-channels than second sub-channels. The spacing direction of the second partition ribs 14 is perpendicular to the airflow direction, while the arrangement direction of the second partition ribs 14 themselves is parallel to the airflow direction. The first channel 2 corresponds to the leading edge region, whose cooling demand is lower than that of the middle region. Therefore, fewer first sub-channels are used. This subdivision ensures uniform distribution of cold air while preventing excessive flow resistance due to too many first sub-channels. This difference in the number of first and second sub-channels allows for precise matching of cooling intensity in different regions, optimizing overall flow resistance and improving the operating efficiency of the gas turbine while meeting cooling requirements. Other components and connections are the same as in specific embodiment seven.
[0033] Specific Implementation Method Nine: Combining Figure 1 , Figure 2 This embodiment differs from Specific Embodiment Eight in that multiple second partition ribs 14 are distributed on the outer side of the sleeve 10. The placement of these second partition ribs on the outer side of the sleeve 10 matches the arrangement of the straight rib structure, avoiding interference with the jet from the impact hole 12 and ensuring that the impact cooling effect is not affected. Simultaneously, the second partition ribs 14, together with the straight ribs, guide the cold air to form an orderly flow within the first sub-channel, further improving the efficiency of convection cooling and ensuring temperature uniformity in the leading edge region. Other components and connections are the same as in Specific Embodiment Eight.
[0034] Specific Implementation Method Ten: Combining Figure 1 , Figure 2 This embodiment differs from Specific Embodiment 1 in that it has multiple third partition ribs 15 spaced apart and arranged along the airflow direction within the third channel 4. These third partition ribs 15 divide the third channel 4 into multiple third sub-channels, with more third sub-channels than second sub-channels. The spacing direction of the multiple third partition ribs 15 is perpendicular to the airflow direction, while the arrangement direction of the multiple third partition ribs 15 themselves is parallel to the airflow direction. Since the trailing edge region is relatively narrow, and the air temperature has increased after passing through the first two channels, resulting in a decrease in cooling capacity, the presence of more third sub-channels maximizes the heat transfer area, refines the airflow path, and ensures sufficient contact between the airflow and the trailing edge region wall, compensating for the reduction in cooling capacity. This also adapts to the spatial structure of the trailing edge region, preventing localized overheating and ensuring the overall uniformity of the blade's cooling effect, further enhancing the blade's high-temperature resistance and operational reliability. Other components and connections are the same as in any of Specific Embodiments 1 to 9.
[0035] The working principle of this implementation method is as follows:
[0036] The cold air first enters the cavity 11 inside the sleeve 10 from the top of the blade, then passes through the impact hole 12, and enters the first channel 2, the second channel 3 and the third channel 4 in sequence to cool each area of the blade. Finally, it flows out from the tail end of the blade. When it enters a different channel, the cold air is divided into multiple parts, which is the same as the number of corresponding sub-channels.
[0037] The scope of this invention is not limited to the above-described embodiments; a combination of one or more specific embodiments can also achieve the purpose of the invention.
Claims
1. A gas turbine turbine vane cooling structure, characterized by, Comprising: a blade body (1); the blade body (1) has a first channel (2), a second channel (3) and a third channel (4) sequentially communicated from the top to the tail, the first channel (2) and the third channel (4) are respectively provided with in-line rib structures, and the second channel (3) is provided with staggered rib structures; the in-line rib structure in the first channel (2) is located in the leading edge region of the blade body (1), the staggered rib structure in the second channel (3) is located in the middle region of the blade body (1), and the in-line rib structure in the third channel (4) is located in the trailing edge region of the blade body (1); a plurality of first partition ribs (5) are arranged in the second channel (3) and spaced apart along the air flow direction, and the second channel (3) is divided into a plurality of second sub-channels by the plurality of first partition ribs (5).
2. A gas turbine turbine vane cooling structure according to claim 1, wherein The second channel (3) includes a front channel (6) and a rear channel (7), the front channel (6) is located near the leading edge region of the blade body (1), the rear channel (7) is located near the trailing edge region of the blade body (1), and the number of second sub-channels in the rear channel (7) is greater than that in the front channel (6).
3. The gas turbine turbine vane cooling structure according to claim 1, wherein The staggered rib structure in the second channel (3) is formed by a plurality of first inclined ribs (8) arranged in a first direction and a plurality of second inclined ribs (9) arranged in a second direction, and the first direction and the second direction are perpendicular.
4. A gas turbine turbine vane cooling structure according to claim 3, wherein A plurality of first inclined ribs (8) are located on the suction surface, a plurality of second inclined ribs (9) are located on the pressure surface, or a plurality of first inclined ribs (8) are located on the pressure surface, and a plurality of second inclined ribs (9) are located on the suction surface.
5. A gas turbine turbine vane cooling structure as claimed in claim 1, characterized by Further comprising: a sleeve (10) mounted on the front end of the blade body (1), the middle of the sleeve (10) has a cavity (11), and the cavity (11) is communicated with the first channel (2) through an impact hole (12).
6. A gas turbine turbine vane cooling structure according to claim 5, wherein The leading edge region of the blade body (1) is provided with a boss (13) for fixing the sleeve (10).
7. A gas turbine turbine vane cooling structure according to claim 5 wherein, The in-line rib structure in the first channel (2) is distributed on the outside of the sleeve (10).
8. A gas turbine turbine vane cooling structure according to claim 7, wherein A plurality of second partition ribs (14) are arranged in the first channel (2) and spaced apart along the air flow direction, and the first channel (2) is divided into a plurality of first sub-channels by the plurality of second partition ribs (14), and the number of first sub-channels is less than that of second sub-channels.
9. A gas turbine turbine vane cooling structure according to claim 8, wherein A plurality of second partition ribs (14) are distributed on the outside of the sleeve (10).
10. A gas turbine turbine vane cooling structure according to any one of claims 1-9, characterized in that, A plurality of third partition ribs (15) are arranged in the third channel (4) and spaced apart along the air flow direction, and the third channel (4) is divided into a plurality of third sub-channels by the plurality of third partition ribs (15), and the number of third sub-channels is greater than that of second sub-channels.