Turbine blade and gas turbine

Abstract

The disclosure provides a turbine blade and a gas turbine, and relates to the technical field of aero-engines. The turbine blade comprises: an inner wall forming an impingement cavity and having a plurality of impingement holes in communication with the impingement cavity, cooling airflow from the impingement cavity forming impingement jets via the plurality of impingement holes; an outer wall sleeved outside the inner wall and spaced from the inner wall to form an impingement flow channel; and a flow guide mechanism connected between the inner wall and the outer wall and configured to divide the impingement flow channel into a plurality of sub-flow channels for guiding wall surface jets formed after impingement jets impinge on the outer wall; wherein the outer wall forms a plurality of film holes downstream of the sub-flow channels, and the flow guide mechanism is further configured to inhibit interference of upstream wall surface jets flowing along the sub-flow channels on downstream impingement jets in the sub-flow channels.

Description

Technical Field

[0001] At least one embodiment of this disclosure relates to the field of aero-engine technology, and more particularly to a turbine blade and a gas turbine. Background Technology

[0002] One of the key ways to improve the efficiency of gas turbine cycles is to continuously increase the turbine inlet temperature, which requires hot-end components such as turbine blades to have more effective cooling capabilities to withstand the resulting extreme high-temperature environment.

[0003] In related technologies, turbine blades with double walls achieve higher overall cooling efficiency than traditional single-wall or simple composite cooling by constructing a fine cooling flow path between the inner and outer double walls. However, under non-uniform heat loads, the wall jet formed by the cooling airflow after impact tends to generate strong crossflow downstream, interfering with the cooling effect of subsequent impact jets and limiting further improvements in cooling uniformity and efficiency. Summary of the Invention

[0004] In view of this, the present disclosure provides a turbine blade and a gas turbine that can improve cooling uniformity and overall cooling efficiency.

[0005] As a first aspect of the present disclosure, a turbine blade is provided, comprising: an inner wall forming an impact cavity and having a plurality of impact holes communicating with the impact cavity, wherein cooling airflow from the impact cavity forms an impact jet through the plurality of impact holes; an outer wall sleeved outside the inner wall and spaced apart from the inner wall to form an impact channel; and a flow guiding mechanism connected between the inner wall and the outer wall and configured to divide the impact channel into a plurality of sub-channels for guiding the flow of a wall jet formed after the impact jet impacts the outer wall; wherein the outer wall forms a plurality of film perforations downstream of the sub-channels, and the flow guiding mechanism is further configured to suppress interference of the upstream wall jet flowing along the sub-channel to the downstream impact jet within the sub-channels.

[0006] According to an embodiment of this disclosure, the flow guiding mechanism includes: a plurality of parallel first ribs adapted to divide at least a portion of the impact flow channel into a plurality of sub-flow channels and guide the wall jet along the sub-flow channels to the air film orifice; a plurality of flow guiding components, respectively spaced apart from the first ribs in the plurality of sub-flow channels, the flow guiding components including: a plurality of second ribs, the second ribs being disposed between two adjacent impact holes, the second ribs being adapted to guide the wall jet to flow downstream from the gap between the second ribs and the first ribs, so as to prevent the wall jet located on one side of the second ribs from directly impacting the impact jet located on the other side of the second ribs.

[0007] According to an embodiment of this disclosure, the flow guiding component further includes: a plurality of connecting ribs disposed between two adjacent second ribs, and the two adjacent connecting ribs are connected to both ends of the extension direction of a second rib, forming a wave-shaped flow guiding component with the second rib.

[0008] According to an embodiment of the present disclosure, the first rib extends along the inner wall in a direction orthogonal to the blade height direction, the second rib extends in the blade height direction, and the plurality of air film holes are disposed at one end of the plurality of sub-channels near the trailing edge.

[0009] According to an embodiment of the present disclosure, the first rib extends in the blade height direction, the second rib extends along the inner wall in a direction orthogonal to the blade height direction, and the plurality of air film holes are disposed at opposite ends of the plurality of sub-channels.

[0010] According to an embodiment of the present disclosure, the axes of two impact holes located on both sides of a second rib are configured to be offset from each other in the extending direction of the second rib; the distance between the axes of the two impact holes located on both sides of a second rib in the extending direction of the second rib is configured to be 0.1-5 times the diameter of the impact hole.

[0011] According to an embodiment of this disclosure, in an orthographic projection parallel to the extending direction of the first rib, the projected area of ​​the second rib is 10%-90% of the projected area of ​​the sub-channel.

[0012] According to an embodiment of this disclosure, the aforementioned flow guiding mechanism is disposed near the leading edge of the turbine blade.

[0013] According to embodiments of this disclosure, the distance between the aforementioned air film hole and the nearest impact hole is configured to be 1-10 times the diameter of the impact hole; and / or, the angle between the axis of the aforementioned air film hole and the outer wall is configured to be 25°-90°; and / or, the impact distance of the aforementioned impact hole is configured to be 0.5-6 times the diameter of the impact hole; and / or, the spacing between two adjacent impact holes within the aforementioned sub-channel is configured to be 2-10 times the diameter of the impact hole; and / or, the spacing between two adjacent first ribs is configured to be less than 10 times the diameter of the impact hole.

[0014] As a second aspect of the present disclosure, a gas turbine is provided, comprising: a compressor adapted to compress inhaled external air into compressed air; a combustion chamber adapted to mix and burn the compressed air with fuel to generate gas; and a turbine connected to the combustion chamber, the turbine having any of the aforementioned turbine blades and adapted to rotate under the action of the gas.

[0015] In this embodiment, by setting a flow guiding mechanism in the impact channel between the inner and outer walls, the wide channel space is divided into multiple restricted sub-channels. After the cooling airflow forms an impact jet through the impact hole and impacts the outer wall, the wall jet can only flow along the corresponding sub-channel, avoiding the disorderly diffusion of the wall jet in the impact channel to form a large-scale crossflow. At the same time, the physical structure of the flow guiding mechanism can block the upstream wall jet from interfering with the downstream impact jet in the sub-channel, cutting off the interaction path between the upstream wall jet and the downstream impact jet. The upstream wall jet flows downstream along the predetermined sub-channel, and the lateral expansion (i.e., diffusion along the inner surface of the outer wall) is suppressed, thereby reducing the impact, entrainment, or momentum interference of the upstream wall jet on the newly emerging impact jet ejected from the adjacent downstream impact hole. This ensures that the impact jet ejected from each impact hole can maintain high momentum and directionality, directly and effectively impacting the preset target area on the inner surface of the outer wall, thus enhancing the impact heat transfer effect. Attached Figure Description

[0016] The above and other objects, features and advantages of this disclosure will become clearer from the following description of embodiments with reference to the accompanying drawings, in which:

[0017] Figure 1 A perspective view of a turbine blade according to an embodiment of the present disclosure is schematically shown;

[0018] Figure 2 A partial perspective view of a turbine blade according to an embodiment of the present disclosure is schematically shown;

[0019] Figure 3 A partial side view of a turbine blade according to a first embodiment of the present disclosure is schematically shown;

[0020] Figure 4 A partial side view of a turbine blade according to a second embodiment of the present disclosure is schematically shown;

[0021] Figure 5 A partial side view of a turbine blade according to a third embodiment of the present disclosure is schematically shown;

[0022] Figure 6 A partial side view of a turbine blade according to a fourth embodiment of the present disclosure is schematically shown; and

[0023] Figure 7 A partial cross-sectional view of a turbine blade according to this disclosure is schematically shown.

[0024] The annotations in the attached figures are explained as follows:

[0025] 1. Inner wall; 11. Impact chamber; 12. Impact hole;

[0026] 2. Outer wall; 21. Air film pores;

[0027] 3. Flow guiding mechanism; 31. First rib; 32. Flow guiding assembly; 321. Second rib; 322. Connecting rib;

[0028] 4. Impact channel; 41. Sub-channel;

[0029] 5. Tail margin;

[0030] 6. Leading edge. Detailed Implementation

[0031] To make the objectives, technical solutions, and advantages of this disclosure clearer, the following detailed description is provided in conjunction with specific embodiments and the accompanying drawings.

[0032] The terminology used herein is for the purpose of describing particular embodiments only and is not intended to limit this disclosure. The terms “comprising,” “including,” etc., as used herein indicate the presence of the stated features, steps, operations, and / or components, but do not exclude the presence or addition of one or more other features, steps, operations, or components.

[0033] All terms used herein, including technical and scientific terms, have the meanings commonly understood by those skilled in the art, unless otherwise defined. It should be noted that the terms used herein are to be interpreted in a manner consistent with the context of this specification, and not in an idealized or overly rigid way.

[0034] When using expressions such as "at least one of A, B, and C," the meaning should generally be interpreted according to the understanding of someone skilled in the art. For example, "a system having at least one of A, B, and C" should include, but is not limited to, systems having A alone, having B alone, having C alone, having A and B, having A and C, having B and C, and / or having A, B, and C. Similarly, when using expressions such as "at least one of A, B, or C," the meaning should generally be interpreted according to the understanding of someone skilled in the art. For example, "a system having at least one of A, B, or C" should include, but is not limited to, systems having A alone, having B alone, having C alone, having A and B, having A and C, having B and C, and / or having A, B, and C.

[0035] It should also be noted that the directional terms mentioned in the embodiments, such as "up," "down," "front," "back," "left," and "right," are only for reference to the directions in the accompanying drawings and are not intended to limit the scope of protection of this disclosure. Throughout the drawings, the same elements are represented by the same or similar reference numerals. Conventional structures or constructions will be omitted where they may cause confusion in understanding this disclosure.

[0036] In the process of developing this disclosure, it was discovered that a typical turbine blade with a double-walled structure mainly consists of three parts: an inner wall, internal supports (such as column ribs or partition ribs), and an outer wall. The inner wall typically has impact holes to form an impact jet that cools the outer wall. The column ribs or partition ribs connect the inner and outer walls, providing support and enhancing heat transfer. Cooled gas impacts the outer wall from the impact chamber formed on the inner wall through the impact holes, then flows out through film cooling holes, forming a heat-insulating film on the outer surface of the blade.

[0037] However, the heat load varies significantly at different locations on the turbine blades, and the impact cavity formed on the inner wall makes it difficult to adjust the cooling arrangement. The full film cooling scheme also results in a certain waste of cooling air, which is not conducive to the refined cooling design of the turbine blades.

[0038] In addition, the non-uniform air film pore arrangement can cause cold air to accumulate in the impact chamber, and the resulting crossflow will have an adverse effect on impact cooling.

[0039] Faced with the increasing thermal load on turbine blades and the stringent requirements of limited cooling air volume, the development of a more efficient non-full film covered double-wall structure is of great significance for turbine blade cooling design.

[0040] Figure 1 A perspective view of a turbine blade according to an embodiment of the present disclosure is shown schematically.

[0041] As one aspect of this disclosure, a turbine blade is provided. (Refer to...) Figure 1 As shown, the turbine blade includes an inner wall 1, an outer wall 2, and a flow guiding mechanism 3. The inner wall 1 forms an impact chamber 11 and has multiple impact holes 12 communicating with the impact chamber 11. Cooling airflow from the impact chamber 11 forms an impact jet through the multiple impact holes 12. The outer wall 2 is fitted outside the inner wall 1 and spaced apart from the inner wall 1 to form an impact channel 4. The flow guiding mechanism 3 is connected between the inner wall 1 and the outer wall 2 and is configured to divide the impact channel 4 into multiple sub-channels 41 that guide the flow of the wall jet formed after the impact jet impacts the outer wall 2. Among them, the outer wall 2 forms multiple film perforations 21 downstream of the sub-channels 41, and the flow guiding mechanism 3 is also configured to suppress the interference of the upstream wall jet flowing along the sub-channels 41 on the downstream impact jet.

[0042] In this embodiment, by providing a flow guiding mechanism 3 within the impact channel 4 between the inner wall 1 and the outer wall 2, the wide channel space is divided into multiple restricted sub-channels 41. After the cooling airflow forms an impact jet through the impact hole 12 and impacts the outer wall 2, the wall jet can only flow along the corresponding sub-channel 41, avoiding the disordered diffusion of the wall jet within the impact channel 4 to form a large-scale crossflow (i.e., disordered diffusion along the inner surface of the outer wall 2). At the same time, the solid structure of the flow guiding mechanism 3 can block the interference of the upstream wall jet in the sub-channel 41. The impact jet cuts off the interaction path between the upstream wall jet and the downstream impact jet. The upstream wall jet flows downstream along the predetermined sub-channel 41, and its lateral expansion is suppressed. This reduces the impact, entrainment, or momentum interference of the upstream wall jet on the newly emerging impact jet ejected from the adjacent downstream impact hole 12. It ensures that the impact jet ejected from each impact hole 12 can maintain high momentum and directionality, and directly and effectively impact the preset target area on the inner surface of the outer wall 2, thereby enhancing the impact heat transfer effect.

[0043] After the cooling airflow completes the impact heat exchange in each sub-channel 41, it can flow smoothly down the sub-channel 41 to the downstream film cooling hole 21 and be discharged, forming a heat-insulating film on the outer surface of the turbine blade. The setting of the flow guiding mechanism 3 also makes the flow rate and flow state of the cold air in each sub-channel 41 stable. Combined with the film cooling hole 21 that is specifically opened in the downstream of the sub-channel 41 on the outer wall 2, the cooling airflow can be finely distributed to adapt to the heat load differences at different positions of the turbine blade.

[0044] In addition, the flow guiding mechanism 3 is connected between the inner wall 1 and the outer wall 2. Its own structural strength can provide support for the inner wall 1 and the outer wall 2, enhance the overall rigidity of the double-wall structure, and withstand the aerodynamic load and thermal stress during the operation of the turbine blade.

[0045] According to embodiments of this disclosure, the air film pores 21 may be arranged in a local area of ​​the outer wall 2.

[0046] As an example, sub-channel 41 can be configured along the blade chord direction ( Figure 1 Extending in the Y direction (as shown). Multiple film cooling holes 21 are formed at one end of the outer wall 2 near the trailing edge 5. The airflow formed by the cooling working fluid in the sub-channel 41 is driven by the chordal static pressure gradient and the kinetic energy of the wall jet, flows toward the trailing edge 5, and exits from the film cooling holes 21.

[0047] In this embodiment, the sub-channel 41 extending along the blade chord direction can follow the natural tendency of the cooling medium inside the turbine blade to flow towards the trailing edge 5 driven by the chordal static pressure gradient. Driven by the kinetic energy of the wall jet, the cooling medium can flow smoothly to the trailing edge 5, reducing local resistance losses during the flow process and improving the delivery efficiency of the cooling medium. This arrangement concentrates the film cooling holes 21 at the end of the outer wall 2 near the trailing edge 5, forming a continuous and complete heat-insulating film in the high-heat-load region of the trailing edge 5. This achieves directional distribution of the cooling medium to the high-temperature region. Simultaneously, the unidirectional flow path avoids the cooling airflow from turning back or flowing back within the channel, effectively suppressing the interference of crossflow on the upstream impinging jet and ensuring the synergistic effect of impinging cooling and film cooling.

[0048] In some alternative embodiments, the sub-channel 41 can be configured along the blade height direction ( Figure 1 (Extended in the X direction as shown). The airflow formed by the cooling medium flows from the impact point towards the blade root and blade tip. The middle part along the blade height direction is the upstream side, and the two sides near the blade root and blade tip are the downstream sides. Multiple film cooling holes 21 are formed on the outer wall 2 at both ends of the sub-channel 41 (i.e., near the blade root and blade tip). When the wall jet flows to both ends of the sub-channel 41, the cooling medium can be discharged through the multiple film cooling holes 21 to form a local film cooling protection in the blade root or blade tip area.

[0049] In this embodiment, the sub-channel 41 extends parallel to the blade height direction, allowing the cooling medium to be diverted from the middle of the blade to both ends (blade root and blade tip). The film cooling holes 21, correspondingly located at both ends of the sub-channel 41, can directly form local film cooling protection in the blade root and blade tip regions, enhancing cooling in these areas. This bidirectional flow pattern shortens the flow distance of the cooling medium from the impact point to the film cooling holes 21, reducing the temperature rise of the airflow within the channel and maintaining a lower temperature for the ejected film cooling medium. This improves the heat insulation effect of the film cooling. Simultaneously, the symmetrical flow structure helps to balance the distribution of cool air along the blade height direction, preventing localized accumulation or insufficient cool air.

[0050] In some alternative embodiments, the extension direction of the sub-channel 41 near the middle along the blade height direction can be approximately parallel to the blade chord direction, and the extension direction of the sub-channels 41 on both sides of the middle sub-channel 41 can also be approximately parallel to the blade height direction. Multiple film cooling holes 21 are formed near the blade root, blade tip, and trailing edge 5. In this embodiment, the combined arrangement of the middle sub-channel 41 extending along the blade chord direction and the two side sub-channels 41 extending along the blade height direction combines the advantages of both chord and height extension, achieving a refined zoned design of the turbine blade cooling channels. The middle sub-channel 41 guides the cooling airflow towards the trailing edge 5, while the two side sub-channels 41 direct the airflow to the blade root and blade tip. Combined with the film cooling holes 21 formed at the blade root, blade tip, and trailing edge 5, film cooling coverage can be simultaneously formed in multiple critical heat load areas of the blade, adapting to the complex non-uniform heat load distribution characteristics of the blade. By using sub-channels 41 with different extension directions to divert and control the cooling airflow, the flow space of the impact channel 4 is further divided, which structurally limits the diffusion range of the wall jet to the greatest extent and weakens the mutual influence of crossflow between different areas. At the same time, the cooling medium can be delivered in multiple parallel paths according to the channel partitions. While meeting the enhanced cooling needs of multiple areas, the overall distribution ratio of cold air volume is optimized, and the cooling efficiency and cold air utilization rate are improved.

[0051] In some illustrative embodiments, the extension direction of the sub-channel 41 may form an angle with the blade height direction or the blade chord direction.

[0052] According to the turbine blades of this disclosure, the flow of the cooling working fluid in the impact channel 4 is adjusted by setting the flow guiding mechanism 3, which suppresses the adverse effects of crossflow and improves the cooling effect. In conjunction with the arrangement of the non-fully covered air film holes 21, the goal of efficient utilization of cold air is achieved.

[0053] According to an embodiment of this disclosure, the outer side of the outer wall 2 is the gas contact side.

[0054] Figure 2 A partial perspective view of a turbine blade according to an embodiment of the present disclosure is schematically shown. Figure 3 A partial side view of a turbine blade according to a first embodiment of the present disclosure is schematically shown. Figure 4 A partial side view of a turbine blade according to a second embodiment of the present disclosure is shown schematically.

[0055] According to embodiments of this disclosure, see Figures 1 to 4As shown, the flow guiding mechanism 3 includes a plurality of parallel first ribs 31 and a plurality of flow guiding components 32. The plurality of first ribs 31 are adapted to divide at least a portion of the impact channel 4 into a plurality of sub-channels 41 and guide the wall jet along the sub-channels 41 toward the film gas aperture 21. The plurality of flow guiding components 32 are respectively spaced apart from the first ribs 31 in the plurality of sub-channels 41, and each flow guiding component 32 includes a plurality of second ribs 321. The second ribs 321 are disposed between two adjacent impact apertures 12, and are adapted to guide the wall jet downstream from the gap between the second ribs 321 and the first ribs 31 (e.g., along...). Figure 3 (As indicated by the arrow) flow to prevent the wall jet located on one side of the second rib 321 from directly impacting the impact jet located on the other side of the second rib 321.

[0056] The second fin 321 can protect the impact jet and weaken the adverse effects of the upstream crossflow, increase the heat transfer area, and enhance heat transfer by utilizing the wall jet generated by the impact.

[0057] As an example, the number of impact holes in each sub-channel 41 can be greater than 4. If the number of impact holes 12 in the sub-channel 41 exceeds 6, air film holes 21 can be arranged at one end of the sub-channel 41 or at both ends of the sub-channel 41.

[0058] As an example, multiple parallel first ribs 31 can extend along the blade height direction, thereby dividing the impact channel 4 into multiple sub-channels 41 extending along the blade height direction. Within each sub-channel 41, multiple second ribs 321 can be arranged along the blade chord direction, and multiple film-forming holes 21 are disposed at opposite ends of the multiple sub-channels 41. The extending direction of the second ribs 321 can be perpendicular to or at an acute angle to the first ribs 31.

[0059] In this implementation, the first rib 31, extending along the blade height, constrains the mainstream cooling airflow along a bidirectional path from the impact point to the blade root and tip, achieving comprehensive cooling across the entire span of the blade. The second ribs 321, arranged along the blade chord direction, effectively block the lateral diffusion of the wall jet in the chord direction by being perpendicular or inclined to the mainstream direction, and guide the airflow to both sides of the sub-channel 41. The second ribs 321, together with the first rib 31, form a finer flow channel within the sub-channel 41, forcing the wall jet to flow stably from the gaps between the second ribs 321 to the blade root and tip, while preventing direct impact interference from the upstream wall jet on other impact jets within the same sub-channel 41. After reaching both ends of the sub-channel 41, the airflow is discharged through pre-set film cooling holes 21. This configuration allows the cooling medium obtained from the impact point in the middle of the sub-channel 41 to be efficiently distributed and transported to the blade root and blade tip, two areas with high heat load due to the secondary flow effect of the end wall. It also forms an effective protective air film in these areas, optimizes the spatial distribution of the cooling airflow, enhances the cooling effect in the end region, and improves the uniformity of the overall temperature field of the blade.

[0060] Or, such as Figure 2 As shown, multiple parallel first ribs 31 can extend along the blade chord direction orthogonal to the blade height direction of the turbine blade, thereby dividing the impact flow channel 4 into multiple sub-flow channels 41 extending along the blade chord direction. Multiple film cooling holes 21 are disposed at one end of the multiple sub-flow channels 41 near the trailing edge 5. In each chordal sub-flow channel 41, multiple second ribs 321 can be arranged along the blade height direction, and the extension direction of the second ribs 321 can be perpendicular to or at an acute angle to the first ribs 31.

[0061] In this implementation, the first rib 31, extending along the blade chord direction, forms a clearly defined chordal cooling channel through physical separation, constraining the mainstream direction of the wall jet to a path from the leading edge 6 to the trailing edge 5, thereby directly utilizing the static pressure gradient along the blade chord direction as the flow driving force. The second rib 321, extending in the blade height direction, effectively blocks the lateral diffusion of the wall jet in the blade height direction through its arrangement perpendicular or approximately perpendicular to the mainstream direction, and forces the airflow forward through the gap defined by the second rib 321 and the first rib 31. This prevents the upstream wall jet from directly impacting and interfering with the newly generated downstream impinging jet. The wave-shaped guide assembly 32, through the continuous structure provided by the connecting rib 322, makes the guiding surface for the airflow present a continuous but non-straight shape in the blade height direction. This helps to further suppress lateral flow and vortex generation within the sub-channel 41, promoting stable convergence of the airflow towards the trailing edge 5.

[0062] The cooling airflow, guided and enhanced in an orderly manner, is concentrated and delivered to the trailing edge 5 region at the end of the sub-channel 41, and discharged in a concentrated manner through the pre-set air film hole 21 there. This forms an effective air film coverage on the outer surface of the trailing edge 5 of the blade, achieving efficient and concentrated cooling of areas that are difficult to cool, and improving the utilization efficiency of the cooling working fluid and the cooling reliability of the trailing edge 5.

[0063] Or, such as Figure 4 As shown, along the leaf height direction, the first rib 31 located in the middle of the leaf extends along the leaf chord direction, and the first rib 31 near the leaf root or leaf tip extends along the leaf height direction. One end of the first rib 31 near the leaf root or leaf tip can be connected to the first rib 31 located in the middle of the leaf, or form a gap with the first rib 31 located in the middle of the leaf.

[0064] The heights of the first rib 31 and the second rib 321 can be set to be equal to the height of the impact channel 4 to ensure complete separation and guidance of the impact channel 4.

[0065] In this embodiment, the first rib 31 forms multiple sub-channels 41 by physically separating the impact channel 4, constraining the wall jet formed after the impact jet hits the outer wall 2, which might otherwise diffuse randomly in all directions, within the sub-channels 41 in a specific direction, and guiding the wall jet to flow directionally along the sub-channels 41 to multiple film gas holes 21. The second rib 321, disposed between adjacent impact holes 12, forms a narrower gap channel between the first ribs 31. The second rib 321 can block and redirect the lateral flow component in the sub-channel 41, forcing the wall jet to flow downstream mainly from the gap between the second rib 321 and the first rib 31. This blocks the wall jet located on one side of the second rib 321 from laterally impacting the newly formed impact jet column on the other side of the second rib 321, avoiding mutual interference and momentum cancellation between the impact jets generated at different times and the wall jets. This ensures the cooling efficiency of each impact jet on the outer wall 2 and makes the cooling airflow in the sub-channel 41 more orderly and stable, ultimately improving the utilization rate of the cooling medium and the temperature uniformity of the blade wall.

[0066] As an example, the second rib 321 can be located at the centerline of adjacent impact holes 12 within a sub-channel 41, or the arrangement of the second rib 321 between adjacent impact holes 12 can be adjusted according to the impact jet parameters to achieve the best cooling effect.

[0067] In the orthographic projection along the Z direction, the projection shape of the first rib 31 can be wavy or straight. The Z direction is perpendicular to the blade height direction and the blade chord direction.

[0068] The cross-sectional shape of the second rib 321 can be rectangular, semi-circular, triangular, etc.

[0069] According to the turbine blades of this disclosure, the crossflow formed by the upstream impact jet is blocked by the second rib 321 which is separated from the first rib 31, thus protecting the downstream impact jet. After the impact jet reaches the outer wall 2, it is restricted by the second rib 321 on both sides of the impact hole 12. The cold air flows out from the gap formed between the second rib 321 and the first rib 31 on both sides, and moves downstream along the first rib 31, scouring the surfaces of the first rib 31 and the second rib 321, thereby improving the heat exchange efficiency.

[0070] Figure 5 A partial side view of a turbine blade according to a third embodiment of the present disclosure is shown schematically.

[0071] According to embodiments of this disclosure, see Figure 5 As shown, the flow guiding assembly 32 also includes a plurality of connecting ribs 322. The plurality of connecting ribs 322 are disposed between two adjacent second ribs 321, and two adjacent connecting ribs 322 are connected to the two ends of the extension direction of a second rib 321, forming a wave-shaped flow guiding assembly 32 with the second rib 321.

[0072] In some illustrative embodiments, when the first rib 31 extends along the blade height direction to form the blade height sub-channel 41, the wave-shaped flow guide assembly 32 is disposed in the sub-channel 41, the extension direction of the second rib 321 can be parallel to the blade chord direction, and the connecting rib 322 connects the end of the second rib 321 along the blade height direction.

[0073] Alternatively, when the first rib 31 extends along the blade chord direction to form a chordal sub-channel 41, the corrugated guide assembly 32 is disposed within the sub-channel 41, and the extension direction of the second rib 321 can be parallel to the blade height direction. The connecting rib 322 connects to the end of the second rib 321 along the blade chord direction. It should be understood that the extension of the first rib 31 along the blade height direction or the blade chord direction does not mean that the first rib 31 is a rectangular rib. The shape of the first rib 31 can match the shape of the impact channel 4 formed between the inner wall 1 and the outer wall 2.

[0074] In this implementation, the connecting rib 322 establishes a physical connection between the ends of adjacent second ribs 321, integrating the discretely distributed second ribs 321 into a structurally coherent wave-shaped guide wall. This enhances the structural integrity of the guide assembly 32 within the sub-channel 41 and its continuous guiding effect on the flow field. The wave-shaped profile allows the wall jet to be smoothly guided downstream by the connecting rib 322 after flowing through the gaps between the second ribs 321, avoiding flow separation or stagnant vortices behind the ends of the second ribs 321. By forming a continuous but non-linear flow boundary, the lateral diffusion and secondary flow of the cooling airflow within the sub-channel 41 can be more effectively suppressed, ensuring that the wall jet flows more concentratedly and stably along the axial direction of the sub-channel 41 towards the target film orifice 21 region.

[0075] Meanwhile, the structurally coherent guide wall helps to establish a more uniform longitudinal static pressure gradient within the sub-channel 41, further reducing flow losses and optimizing the overall flow efficiency from the impact zone to the outlet of the film gas hole 21.

[0076] According to an embodiment of this disclosure, by providing a second rib 321, the impact jet can be surrounded by the first rib 31 and the second rib 321 on three sides. After the impact jet reaches the outer wall 2, it is restricted by the first rib 31 and the second rib 321. The cooling medium flows out from one side of the notch formed by the first rib 31 and the second rib 321 and moves downstream along the first rib 31.

[0077] The second rib 321 further increases the heat transfer area inside the turbine blades. As the cooling medium after the upstream impact flows along the second rib 321, surface heat transfer is also improved. The first rib 31 serves to increase the heat transfer area and guide the impact flow.

[0078] Figure 6 A partial side view of a turbine blade according to a fourth embodiment of the present disclosure is shown schematically.

[0079] According to embodiments of this disclosure, see Figure 6 As shown, the axes of the two impact holes 12 located on either side of a second rib 321 are configured to be offset from each other in the extending direction of the second rib 321. The distance D1 between the axes of the two impact holes 12 located on either side of a second rib 321 in the extending direction of the second rib 321 is configured to be 0.1-5 times the diameter D of the impact hole 12. That is, 0.1D≤D1≤5D.

[0080] For example, the distance D1 between the axes of the two impact holes 12 located on both sides of a second rib 321 in the extending direction of the second rib 321 can be configured as any value among 0.1D, 0.5D, D, 1.5D, 2D, 2.5D, 3D, 3.5D, 4D, 4.5D and 5D.

[0081] In this embodiment, the axes of adjacent impact holes 12 are staggered in a direction parallel to the second rib 321, and the stagger distance D1 is controlled within the range of 0.1 to 5 times the diameter of the impact hole 12, resulting in the following synergistic technical effects: One of the core functions of the second rib 321 is to block the direct interference of the jets from its two side walls. The staggered design first causes the impact points of the impact jets generated by these two holes to be offset from each other on the inner surface of the outer wall 2 along the extension direction of the second rib 321. This causes the main diffusion directions of the two wall jets generated immediately after the impact to be spatially separated at the root. Subsequently, the second rib 321 itself further physically blocks the lateral mixing of these two jets.

[0082] By precisely controlling the misalignment distance, the isolation effect can be optimized: when D1 is too small (e.g., less than 0.1 times the aperture), the misalignment is insufficient, the jet impact points are too close, and some jets or wall jets may still cross the top of the second rib 321 and mix; when D1 is too large (e.g., greater than 5 times the aperture), it may exceed the hydrodynamic influence range that the second rib 321 can effectively cover, and will reduce the coverage density of the impact cooling. By controlling D1 within the above range, the isolation effect of the second rib 321 on the flow fields on both sides can be maximized while ensuring the coverage of the impact cooling. This more thoroughly prevents the wall jet on one side from directly impacting and weakening the newly formed impact jet column on the other side, ensuring that the cooling efficiency of each impact jet is fully utilized, and ultimately promoting a more stable and orderly longitudinal flow within the sub-channel 41.

[0083] According to an embodiment of the present disclosure, in an orthographic projection parallel to the extension direction of the first rib 31, the projected area of ​​the second rib 321 is 10%-90% of the projected area of ​​the sub-channel 41.

[0084] As an example, in an orthographic projection parallel to the extension direction of the first rib 31, the projected area of ​​the second rib 321 is any value among 10%, 20%, 30%, 40%, 50%, 60%, 70%, 80%, and 90% of the projected area of ​​the sub-channel 41.

[0085] In this implementation, by ensuring that the projected area of ​​the second rib 321 accounts for no less than 10% of the projected area of ​​the sub-channel 41, sufficient physical presence of the second rib 321 is ensured to effectively block and redirect the wall jet, preventing disordered lateral diffusion of the wall jet within the sub-channel 41 and guaranteeing isolation from adjacent impact jets. By ensuring that the proportion is no higher than 90%, excessive blockage of the flow channel by the second rib 321 is avoided, which would lead to a sharp increase in flow resistance. This proportion range allows the cooling airflow to maintain a sufficiently unobstructed path to the target film cooling hole 21 region while being adequately guided and disturbed to enhance heat transfer, thereby achieving an optimal balance between improving wall cooling uniformity and maintaining system flow efficiency. The specific value of the proportion can be adjusted according to the heat load intensity of the area where the sub-channel 41 is located. A higher proportion can be used in high heat load areas to enhance flow organization and heat transfer, while a lower proportion can be used in low heat load areas to reduce flow pressure loss.

[0086] According to an embodiment of this disclosure, the flow guiding mechanism 3 is disposed near the leading edge 6 of the turbine blade.

[0087] The leading edge 6 is the region that bears the direct impact of the incoming gas flow and the highest stagnation temperature, exhibiting the highest heat flux density and the most stringent requirements for cooling efficiency. A flow guiding mechanism 3 is installed at this location to immediately guide and regulate the cooling jet ejected from the impact chamber 11 through the impact hole 12 after completing its impact cooling of the outer wall 2. Through the combined action of the main flow channel formed by the first rib 31 and the wave-shaped flow guiding wall formed by the second rib 321 and connecting rib 322, the cooling airflow within the complex curvature region of the leading edge 6 is effectively organized, preventing flow separation and stagnation. This ensures that the wall jet is efficiently guided and covers the predetermined area of ​​the outer surface of the leading edge 6, providing a stable, uniform, and sufficient cooling air source for the dense film cooling holes 21 opened on the outer wall 2, guaranteeing the integrity of film cooling. Simultaneously, the flow guiding mechanism 3 fully utilizes its isolation and protection function for the impact jet in this critical region of the leading edge 6, avoiding mutual interference between jets that may be caused by high-dynamic incoming pressure fluctuations, thereby ensuring the reliability of the basic impact cooling effect.

[0088] According to embodiments of this disclosure, the distance D2 between the air film orifice 21 and the nearest impact orifice 12 is configured to be 1-10 times the diameter D of the impact orifice 12. That is, D≤D2≤10D.

[0089] For example, the distance D2 between the air film hole 21 and the nearest impact hole 12 can be configured as any value among D, 2D, 3D, 4D, 5D, 6D, 7D, 8D, 9D, and 10D.

[0090] In this implementation, when the distance is not less than one times the diameter of the impact hole 12, it ensures that the impact jet is not directly affected by the suction interference from the gas film hole 21 during the initial development stage of impacting the outer wall 2 and forming the wall jet. This ensures that the local convective heat transfer effect of the impact jet on the inner surface of the outer wall 2 can be fully utilized. When the distance is not greater than ten times the diameter of the impact hole 12, it avoids excessive attenuation of the wall jet due to friction before reaching the gas film hole 21. This ensures that the cooling airflow with sufficient momentum and energy can be effectively discharged through the gas film hole 21 and form a highly adhesive and fully covered gas film layer on the outer surface of the blade. This distance range, combined with the guiding effect of the flow guiding mechanism 3 (first rib 31 and second rib 321), optimizes the entire flow path length from the impact hole 12 to the gas film hole 21. While ensuring the core impact cooling efficiency, it provides a stable and sufficient air source for downstream gas film cooling, improving the overall cooling efficiency of the turbine blade under limited cooling medium conditions.

[0091] Figure 7 A partial cross-sectional view of a turbine blade according to this disclosure is schematically shown.

[0092] According to embodiments of this disclosure, referring to Figure 7As shown, the angle α between the axis of the air film orifice 21 and the outer wall 2 is configured to range from 25° to 90°. That is, 25° ≤ α ≤ 90°. As an example, the angle α between the axis of the air film orifice 21 and the outer wall 2 can be configured to any value among 25°, 30°, 35°, 40°, 45°, 50°, 55°, 60°, 65°, 70°, 75°, 80°, 85°, and 90°.

[0093] In this implementation, by setting the lower limit of the included angle α to 25°, it is ensured that the cooling medium ejected from the orifice has a sufficient velocity component parallel to the surface of the blade outer wall 2. This parallel velocity component is the hydrodynamic basis for forming a stable, wall-attached gas film layer, effectively isolating the high-temperature combustion gas from the wall surface. At the same time, this angle range avoids the problems of excessive jet adhesion, insufficient lateral diffusion, and easy stripping by the main combustion gas flow caused by an angle that is too small (such as close to 0°). By setting the upper limit of the included angle to 90°, which allows vertical injection, design options are provided for areas where higher penetration depth is required to counteract strong lateral pressure gradients or strong secondary flows. The specific selection of the included angle α is matched with the incoming flow direction provided by the flow guide mechanism 3, the local external combustion gas flow direction, and the pressure field, so that the discharged cooling jet can integrate into the mainstream with optimal momentum and direction, achieving effective coverage while minimizing aerodynamic losses and cooling medium consumption.

[0094] The angle formed between the axis of the air film hole 21 and the first rib 31 is in the range of 0 to 90°. The angle between the axis of the air film hole 21 and the first rib 31 can be configured to any value among 0°, 10°, 20°, 30°, 40°, 50°, 60°, 70°, 80° and 90°.

[0095] According to embodiments of this disclosure, the impact distance H of the impact hole 12 is configured to be 0.5-6 times the diameter D of the impact hole 12. That is, 0.5D≤H≤6D.

[0096] As an example, the impact distance H of the impact hole 12 can be configured to any value among 0.5D, D, 2D, 3D, 4D, 5D, and 6D.

[0097] The cross-sectional shape of the impact hole 12 can be set to circular or elliptical according to different heat load requirements.

[0098] In this implementation, when the impact distance is not less than 0.5 times the orifice diameter, sufficient space is provided for the jet to complete the transition from in-orifice flow to free jet, allowing it to fully develop and form a stable jet core region before impacting the surface. This ensures that the jet has sufficient momentum and concentration to disrupt the wall thermal boundary layer and generate a stagnation region with a high heat transfer coefficient. When the impact distance is not greater than 6 times the orifice diameter, excessive radial diffusion and kinetic energy attenuation of the jet before reaching the surface are limited, ensuring that the jet maintains a high velocity and concentrated impact force upon impact, thus maintaining a strong local convective heat transfer effect.

[0099] According to embodiments of this disclosure, the spacing D3 between two adjacent impact holes 12 within a sub-channel 41 is configured to be 2-10 times the diameter D of the impact hole 12. That is, 2D≤D3≤10D.

[0100] As an example, the spacing D3 between two adjacent impact holes 12 within a sub-channel 41 can be configured as any value among 2D, 3D, 4D, 5D, 6D, 7D, 8D, 9D, and 10D.

[0101] In this implementation, by ensuring that the spacing D3 between adjacent impact holes 12 within the same sub-channel 41 is not less than twice the orifice diameter, sufficient radial development space is provided for each impact jet. This effectively prevents excessive mixing or mutual interference between two adjacent jets before impacting the surface, ensuring that each jet can independently form its core high heat transfer rate stagnation region, thus guaranteeing the strength of the basic impact cooling. By ensuring that the spacing D3 is not greater than 10 times the orifice diameter, excessively large, unimpacted cooling weak zones between adjacent jet action areas are avoided, ensuring that the wall jets can merge in time to form a continuous and uniform wall-attached flow, providing uniform inflow conditions for subsequent extended convection cooling.

[0102] According to an embodiment of this disclosure, the spacing D4 between two adjacent first ribs 31 is configured to be less than 10 times the diameter D of the impact hole 12 in a direction orthogonal to the extending direction of the first rib 31. That is, D4 ≤ 10D.

[0103] As an example, the spacing D4 between two adjacent first ribs 31 can be configured as any value among D, 2D, 3D, 4D, 5D, 6D, 7D, 8D, 9D, and 10D in a direction orthogonal to the extension direction of the first rib 31.

[0104] The spacing between two adjacent first ribs 31 can be adjusted according to the local heat load in a direction orthogonal to the extension direction of the first rib 31.

[0105] In this implementation, the narrower width of the sub-channel 41 allows the first rib 31 to more effectively restrict the diffusion of the wall jet across the cross section, forcing the cooling airflow to flow in the designed direction within a more defined channel, enhancing the organization of the airflow, which is conducive to forming a stable and concentrated wall jet, and ensuring that it can be effectively delivered to the pre-positioned film orifice 21.

[0106] As a second aspect of this disclosure, a gas turbine is provided, comprising a compressor, a combustion chamber, and a turbine. The compressor is adapted to compress inhaled outside air into compressed air. The combustion chamber is adapted to mix and burn the compressed air with fuel to produce gas. The turbine is connected to the combustion chamber and has any of the turbine blades described above, adapted to rotate under the action of the gas.

[0107] The flow guiding mechanism 3 can be arranged on the suction surface of the turbine blade near the leading edge 6. Alternatively, the flow guiding mechanism 3 can be arranged on the pressure surface. When the flow guiding mechanism 3 is arranged on the pressure surface, the flow direction of the cooling medium in the impact channel can be either acute or obtuse with the mainstream direction.

[0108] It should be noted that the arrangement of the flow guiding mechanism 3 is not limited to turbine blades, but can also be used for turbine moving blades, end walls, and flame tube cooling.

[0109] The embodiments of this disclosure have been described above. However, these embodiments are for illustrative purposes only and are not intended to limit the scope of this disclosure. Although various embodiments have been described above, this does not mean that the measures in the various embodiments cannot be used advantageously in combination. The scope of this disclosure is defined by the appended claims and their equivalents. Various substitutions and modifications can be made by those skilled in the art without departing from the scope of this disclosure, and all such substitutions and modifications should fall within the scope of this disclosure.

Claims

1. A turbine blade, characterized in that, include: The inner wall forms an impact cavity and has multiple impact holes communicating with the impact cavity. Cooling airflow from the impact cavity forms an impact jet through the multiple impact holes. The outer wall is fitted outside the inner wall and spaced apart from the inner wall to form an impact flow channel; A flow guiding mechanism is connected between the inner wall and the outer wall and is configured to divide the impact flow channel into multiple sub-channels that guide the wall jet flow formed after the impact jet hits the outer wall; The outer wall forms multiple air film holes downstream of the sub-channel, and the flow guiding mechanism is also configured to suppress the interference of the upstream wall jet flowing along the sub-channel on the downstream impact jet within the sub-channel.

2. The turbine blade according to claim 1, characterized in that, The flow guiding mechanism includes: Multiple parallel first ribs are adapted to divide at least a portion of the impact channel into multiple sub-channels and guide the wall jet along the sub-channels to the film gas pore; Multiple flow guiding components are respectively disposed at intervals from the first rib in multiple sub-channels, the flow guiding components including: Multiple second ribs are provided between two adjacent impact holes. The second ribs are adapted to guide the wall jet to flow downstream from the gap between the second rib and the first rib, so as to prevent the wall jet located on one side of the second rib from directly impacting the impact jet located on the other side of the second rib.

3. The turbine blade according to claim 2, characterized in that, The flow guiding component also includes: Multiple connecting ribs are disposed between two adjacent second ribs, and two adjacent connecting ribs are connected to the two ends of the extension direction of a second rib, forming a wave-shaped flow guide assembly with the second rib.

4. The turbine blade according to claim 2 or 3, characterized in that, The first rib extends along the inner wall in a direction orthogonal to the blade height direction, the second rib extends in the blade height direction, and the plurality of air film holes are disposed at one end of the plurality of sub-channels near the trailing edge.

5. The turbine blade according to claim 2 or 3, characterized in that, The first rib extends in the blade height direction, the second rib extends along the inner wall in a direction orthogonal to the blade height direction, and the plurality of air film holes are disposed at opposite ends of the plurality of sub-channels.

6. The turbine blade according to claim 2 or 3, characterized in that, The axes of the two impact holes located on either side of a second rib are configured to be offset from each other in the extending direction of the second rib; The distance between the axes of the two impact holes located on either side of a second rib in the extending direction of the second rib is configured to be 0.1-5 times the diameter of the impact hole.

7. The turbine blade according to claim 2, characterized in that, In a projection parallel to the extension direction of the first rib, the projected area of ​​the second rib is 10%-90% of the projected area of ​​the sub-channel.

8. The turbine blade according to any one of claims 1-3, characterized in that, The flow guiding mechanism is located near the leading edge of the turbine blade.

9. The turbine blade according to claim 2, characterized in that, The distance between the air film pore and the nearest impact pore is configured to be 1-10 times the diameter of the impact pore; And / or, the angle between the axis of the air film pore and the outer wall is configured to be in the range of 25°-90°; And / or, the impact distance of the impact hole is configured to be 0.5-6 times the diameter of the impact hole; And / or, the spacing between two adjacent impact holes within one of the sub-channels is configured to be 2-10 times the impact hole diameter; And / or, the spacing between two adjacent first ribs is configured to be less than 10 times the diameter of the impact hole.

10. A gas turbine, characterized in that, include: An air compressor is used to compress inhaled outside air into compressed air. A combustion chamber is used to mix and burn the compressed air with fuel to produce combustion gas; A turbine, in communication with the combustion chamber, the turbine having turbine blades as described in any one of claims 1-9, adapted to rotate under the action of the combustion gas.

Images