High-pressure turbine moving blade suction surface echelon cooling structure and design method thereof

Abstract

The invention discloses a high-pressure turbine moving blade suction surface echelon cooling structure and a design method thereof, and belongs to the technical field of aero-engine turbine blade cooling. According to the structure, the suction surface of a movable blade is divided into an upstream gill area, a midstream main body area and a downstream area in the fuel gas flowing direction, a blade root area, a blade middle area and a blade tip area are divided in the unfolding direction, the midstream blade middle area is in uniform cooling layout, and the blade root side, the blade tip side, the gill area and the downstream area are in echelon layout with the hole diameter decreasing gradually from the end to the middle. And partition differentiation air film cooling is achieved. The design method comprises the steps of working condition and flowing heat load acquisition, region division, reference cooling layout and end region echelon cooling layout determination, parameter checking, structure output and the like. Compared with a traditional uniform hole distribution scheme, the cooling efficiency, the air film covering uniformity and the cold air utilization efficiency can be remarkably improved by distributing cooling air in an echelon mode and optimizing matching of air film outflow and local flowing, and the method is suitable for thermal protection of the suction surface of the high-pressure turbine moving blade of the engine.

Description

Technical Field

[0001] This invention belongs to the field of aero-engine turbine blade cooling technology, and relates to the design of air film cooling layout and flow heat transfer control of high-pressure turbine moving blades. Specifically, it relates to a novel stepped cooling structure and its design method for the suction surface of high-pressure turbine moving blades, which is used to improve the cooling efficiency and temperature distribution uniformity of the suction surface of the moving blades. Background Technology

[0002] Film cooling (FSC) of turbine blades is one of the core methods for thermal protection of high-temperature components in aero-engines. It aims to block the heating of the turbine blade wall by spraying cooling gas through film cooling holes on the blade surface, serving two main purposes: first, to remove some heat through the cooling gas; and second, to isolate the turbine blade wall from the high-temperature combustion gases, thereby reducing localized temperatures, protecting the blade surface, and preventing material creep or thermal damage. Currently, turbine blade FSC often employs a uniform arrangement, neglecting the temperature distribution in different regions and the entrainment and detachment effects of vortex structures on the outflow. With the continuous increase in turbine inlet temperature, this uniform arrangement is highly susceptible to localized cooling failure, leading to overheating and subsequent thermal ablation, severely impacting engine operational safety.

[0003] Specifically, unlike guide vanes, the flow and heat transfer environment of high-pressure turbine blades in aero-engines, especially the suction surface, is far more complex. The suction surface undergoes strong acceleration and deceleration processes, with localized adverse pressure gradients and significant curvature, making it prone to boundary layer thickening, separation, and reattachment. Simultaneously, the high turbulence, unsteady impacts, and circumferential inhomogeneity introduced by the upstream guide vane wake cause significant spatial and temporal variations in the wall heat flux density and aerodynamic load on the suction surface. Inside the rotor channel, the surface flow becomes even more complex due to the combined effects of secondary flow and rotation.

[0004] For the suction surface of the moving blade, its flow characteristics are as follows: (1) Overall, the flow is constricted along the flow direction, unlike the guide vane, and the spanwise distribution along the blade is no longer symmetrical. (2) In the blade mid-area, the flow is relatively stable, but the coverage area is much smaller than the corresponding area of ​​the turbine guide vane. (3) In the blade root area, under the action of strong channel vortices and centrifugal and Coriolis forces pointing towards the blade tip, the flow shifts significantly towards the blade mid-area, resulting in a large area of ​​uncovered air film. Compared with the guide vane, the uncovered air film area is close to the blade mid-area, which can easily lead to overheating and inducing thermal ablation in this area. (4) In the blade tip area, the leakage flow from the pressure surface points towards the blade root, while the centrifugal and Coriolis forces point towards the blade tip. After the superposition of the two, the flow shifts slightly towards the blade mid-area, and the overall uncovered air film area is small, even smaller than the uncovered area of ​​the guide vane in this part. (5) In the gill area, the curvature changes drastically, and the flow undergoes a stepwise change from pressure to pressure, the air film's ability to adhere to the wall weakens, and it is very easy to detach from the wall, inducing high-temperature failure.

[0005] Under the aforementioned complex flow conditions, the film gas outflow from the suction surface of high-pressure turbine blades becomes extremely complex. In summary: in the mid-blade region, the film gas is relatively stable, but its coverage area is relatively small; in the blade root region, the film gas shifts significantly, resulting in a large uncovered area; in the blade tip region, the film gas shifts slightly, resulting in a smaller uncovered area; and in the gill region, the film gas easily detaches from the blade wall. Although some studies have begun to incorporate factors such as endwall secondary flow intensity, leakage flow distribution, and unsteady loads into the film gas design, proposing some improved perforation strategies for local areas, these efforts are generally localized and empirical adjustments. A systematic design approach for the entire suction surface of high-pressure turbine blades has not yet been developed, making it difficult to achieve a reasonable allocation of cooling resources and an overall improvement in cooling performance under the constraint of limited cooling gas flow.

[0006] In summary, existing film cooling designs for the suction surface of high-pressure turbine blades are still insufficient in fully considering the complex three-dimensional flow structure, heat load distribution, and the effects of rotational additional forces. This makes it difficult to achieve differentiated matching of cooling needs in different regions and overall optimization of coolant utilization efficiency. Therefore, designing a film cooling layout that can optimize the flow characteristics of different regions of the suction surface to achieve uniform and effective thermal protection is a pressing technical problem that needs to be solved in this field. Summary of the Invention

[0007] (a) Purpose of the invention The purpose of this invention is to propose a tiered cooling structure and design method for the suction surface of a high-pressure turbine blade. Its essential feature is the tiered arrangement of the film cooling holes, specifically: in the blade mid-section, a uniform arrangement is used; in the blade root section, a tiered cooling method is adopted, with the hole diameter distributed in a tiered manner, i.e., larger holes are used near the root, decreasing in diameter towards the mid-section along the spanwise direction, so that the film cooling outflow can effectively resist the entrainment and rotational effects of channel vortices; in the blade tip section, a tiered cooling method is adopted, with the hole diameter distributed in a tiered manner, i.e., larger holes are used near the tip .... To resist the effects of leakage flow blow-off and rotation, a centralized tiered cooling method is adopted in the downstream area of ​​the blade root, with at least two rows of film cooling holes arranged locally, and the hole diameters are distributed in a tiered manner to further suppress the blow-off effect of channel vortices and rotation on the film cooling outflow. In the downstream area of ​​the blade tip, a centralized tiered cooling method is adopted, with at least two rows of film cooling holes arranged locally, and the hole diameters are distributed in a tiered manner to further suppress the blow-off effect of leakage flow and rotation on the film cooling outflow. Compared with the blade root area, the number of centralized cooling film cooling holes in the blade tip area is less, and the hole diameter variation along the span is smaller. In the gill area, a small hole arrangement is used upstream and a large hole arrangement downstream to enhance the film cooling adhesion ability. Overall, the film cooling arrangement is adjusted according to requirements to achieve a reasonable and effective distribution of cool air and achieve the most ideal cooling effect, possessing the advantages of strong functionality and clear application.

[0008] (II) Technical Solution To achieve the objective of this invention and solve its technical problems, the present invention adopts the following technical solution: The first objective of this invention is to provide a tiered cooling structure for the suction surface of a high-pressure turbine blade, used for differentiated film cooling of the suction surface of the high-pressure turbine blade in different zones. This improves the temperature distribution of the suction surface and suppresses the entrainment and blow-off of the film by the secondary flow, especially under conditions of high turbine inlet temperature, particularly above the temperature resistance limit of the blade substrate material. The suction surface of the blade body is sequentially divided into an upstream gill zone, a midstream main body zone, and a downstream zone along the gas flow direction; wherein, The upstream gill region is the area on the suction surface of the moving blade that is close to the trailing edge of the guide vane and has significant changes in curvature and pressure gradient. It adopts a stepped-increase air film pore layout along the flow direction, with upstream small pore rows and downstream large pore rows arranged in the flow direction. Each of the upstream small pore rows and the downstream large pore rows has one or more rows. The air film pores are evenly distributed along the span, and the pore diameter of the air film pores in the downstream large pore row is larger than that of the upstream small pore air film pores, forming the reference pore diameter of the entire air film pore area. This allows the cold air ejected from the upstream and downstream rows to form a cold air film attached to the suction surface in the upstream gill region. The midstream main region is divided in the spanwise direction into the midstream blade root region near the rim, the midstream blade tip region near the tip, and the midstream blade middle region between the two. The midstream blade middle region adopts a uniform cooling layout, with its film cooling pores arranged in a single or multiple rows. The diameter of each row of film cooling pores is basically the same and corresponds to the reference diameter, and each film cooling pore is uniformly distributed along the spanwise direction. The midstream blade root region and the midstream blade tip region both adopt a stepped cooling layout. In the spanwise direction, there are continuous rows of film cooling pore arrays corresponding to each row of film cooling pores in the midstream blade middle region. The diameter of each film cooling pore gradually decreases from the side near the rim or tip towards the midstream blade middle region in the spanwise direction, forming a stepped distribution of pore diameter transitioning from the end region to the blade middle region. The downstream region is divided in the spanwise direction into a downstream leaf root region near the rim and a downstream leaf tip region near the tip. Both the downstream leaf root and downstream leaf tip regions adopt a tiered cooling layout, with at least two arrays of exhaust film holes arranged in the flowwise direction. The diameter of each exhaust film hole gradually decreases from the side near the rim or tip towards the middle region in the spanwise direction. Moreover, each exhaust film hole is located in the downstream areas where the air film is sparse or not covered at the upper and lower ends of the midstream main region in the spanwise direction. This makes the exhaust film holes in the upstream gill region, the midstream main region, and the downstream region together form a tiered cooling structure in the spanwise and flowwise directions, with the diameter gradually decreasing from the end region to the middle region, which strengthens the weak cooling areas on the leaf root and leaf tip sides.

[0009] The second objective of this invention is to provide a design method for the above-mentioned high-pressure turbine blade suction surface tiered cooling structure, wherein the method, when implemented, includes at least the following steps: SS1. Obtaining Operating Conditions and Flow Heat Loads: Based on the geometric parameters of the target high-pressure turbine stage and representative design operating conditions, information on the wall temperature, local heat flux density, static pressure, and secondary flow intensity distribution in the end region on the suction surface of the moving blades is obtained through CFD numerical calculations and / or experimental tests. SS2. Suction surface region division: Based on the static pressure gradient, curvature change and secondary flow intensity distribution at various points on the suction surface of the moving blade, it is divided into the upstream gill region, the midstream main body region and the downstream region along the gas flow direction. The midstream main body region is further divided into the midstream blade root region, the midstream blade middle region and the midstream blade tip region in the spanwise direction. The downstream region is further divided into the downstream blade root region and the downstream blade tip region in the spanwise direction. SS3. Determination of the reference cooling layout in the midstream main area: Determine the number of rows, spanwise position and reference aperture of single or multiple rows of uniform film cooling holes in the midstream blade mid-section along the gas flow direction, so as to form a uniform cooling film under the premise of meeting the overall cooling gas flow constraint. Based on this, determine the arrangement position of film cooling holes in the midstream blade root area and midstream blade tip area along the spanwise direction, corresponding to each row in the midstream blade mid-section and arranged in a continuous row. SS4. Determination of the tiered cooling layout in the terminal and gill regions: Based on the secondary flow and leakage flow intensity and the easily thinned areas of the gas film in the midstream leaf root region, midstream leaf tip region, downstream leaf root region, and downstream leaf tip region, the aperture distribution and number of rows of each exhaust film pores are determined in a progressively decreasing manner from the end to the middle along the longitudinal direction. In the upstream gill region, the positions and aperture differences between the upstream small pore rows and the downstream large pore rows are determined along the gas flow direction, forming a tiered cooling layout that connects with the midstream main region. SS5. Parameter Verification and Structural Output: Under the constraints of total cooling air flow, turbine stage aerodynamic losses and blade structural strength, the aperture, row spacing and aperture spacing of the film cooling holes in each region are iteratively verified and optimized until a high-pressure turbine blade suction surface tiered cooling structure that meets the design requirements is formed.

[0010] (III) Technical Effects Compared with the prior art, the high-pressure turbine blade suction surface tiered cooling structure and its design method of the present invention have the following beneficial and significant technical effects: (1) Unlike the conventional arrangement from leaf root to leaf tip, the tiered film cooling layout of this invention features uniform distribution in the leaf mid-area, tiered distribution in the leaf root and leaf tip areas, and concentrated distribution in the leaf root downstream and leaf tip downstream areas, highlighting its structural characteristics. Specifically, this invention divides the suction surface of the moving blade along the flow direction into an upstream gill area, a midstream main area, and a downstream area, and further distinguishes the leaf root area, leaf mid-area, and leaf tip area along the span. Combining the static pressure gradient, curvature changes, and secondary flow and leakage flow intensity of each area, it implements differentiated pore layout and tiered pore size distribution, enabling targeted enhanced cooling of the gill area, tip area, and downstream weak area. This significantly improves the continuity and stability of the film cooling coverage and significantly reduces the maximum metal temperature and temperature gradient in local high heat load areas.

[0011] (2) The tiered film cooling layout of the present invention can be added to the blade surface at appropriate positions from the leaf root to the leaf tip as needed. Furthermore, the present invention constructs a uniformly cooled reference film cooling perforation array in the midstream leaf region, and adopts a tiered layout in the midstream leaf root region, midstream leaf tip region, downstream leaf root region, and downstream leaf tip region, with the perforation diameter decreasing progressively from the tip to the middle along the spanwise direction. In the gill region, a combination of upstream small perforation arrays and downstream large perforation arrays is used to achieve a cooling distribution strategy that is moderate in the middle and enhanced at the ends. Under the premise that the total cooling airflow is basically limited, the cooling air utilization efficiency is effectively improved, alleviating the problem of insufficient cooling in the end region and excessive cooling in some areas in the traditional uniform perforation scheme.

[0012] (3) The high-pressure turbine blade suction surface tiered cooling structure of the present invention is specifically designed for the weak air film coverage area of ​​the turbine blade caused by secondary flow and rotation in the channel, so as to achieve targeted and efficient cooling. This layout is implemented locally on the blade surface, and the cooling effect is well improved. Attached Figure Description

[0013] Figure 1 The diagram shows a tiered film cooling layout for the suction surface of the turbine blades. Figure 2 The image shown is a top view of the uniform cooling layout in the middle section of the midstream blades; Figure 3 The image shown is a top view of the tiered cooling layout in the midstream leaf root region; Figure 4 The image shown is a top view of the tiered cooling layout in the midstream blade tip area; Figure 5 The image shown is a top view of the concentrated tiered cooling layout in the downstream blade root area. Figure 6 The image shown is a top view of the concentrated tiered cooling layout in the downstream blade tip area; Figure 7 The image shown is a top view of the cooling layout of the small pores upstream and the large pores downstream of the gill area. Figure 8 The image shown is a partial cross-sectional view of the tiered cooling layout. Figure 9 The image shown is a partial top view of the tiered cooling layout; Figure 10 The diagram shows the design method of the stepped cooling structure for the suction surface of a high-pressure turbine blade.

[0014] Explanation of reference numerals in the attached diagram: 1-Suction surface of turbine blade, 2-Midstream blade middle area, 3-Midstream blade root area, 4-Midstream blade tip area, 5-Downstream blade root area, 6-Downstream blade tip area, 7-Upstream gill area, 8-Air film pore, 9-Air film pore inlet, 10-Air film pore outlet, 11-Flow direction, 12-Spread direction, 13-Main stream, 14-Jet. Detailed Implementation

[0015] The purpose of this invention is to propose a tiered cooling structure for the suction surface of a high-pressure turbine blade and its design method. To make the objectives, technical solutions, and advantages of this invention clearer, the technical solutions in the embodiments of this invention will be described in more detail below with reference to the accompanying drawings. The described embodiments are some, but not all, embodiments of this invention, and are exemplary and intended to explain the invention, not to limit it. Based on the embodiments of this invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this invention.

[0016] Example 1: Cascaded Cooling Structure for Suction Surface of High-Pressure Turbine Blades As a specific example, Figure 1 The diagram shows the film cooling layout obtained by the present invention based on the stepped cooling design method for the suction surface of a high-pressure turbine blade. This layout is used for differentiated film cooling of the suction surface of the high-pressure turbine blade, aiming to improve the temperature distribution of the suction surface and suppress the entrainment and blow-off of the film cooling by secondary flow, especially under conditions of high turbine inlet temperature, particularly exceeding the temperature resistance limit of the blade substrate material. Compared to conventional film cooling, which features a uniform perforation arrangement along the spanwise direction from blade root to blade tip, the present invention divides the suction surface of the blade body into an upstream gill zone, a midstream main body zone, and a downstream zone along the gas flow direction. The midstream blade middle zone 2 uses a uniform perforation layout, the midstream blade root zone 3 uses a stepped perforation layout, the midstream blade tip zone 4 uses a stepped perforation layout, the downstream blade root zone 5 uses a concentrated stepped perforation layout, the downstream blade tip zone 6 uses a concentrated stepped perforation layout, and the upstream gill zone uses a layout with varying perforation sizes. More specifically: Figure 7 The image shown is a top view of the cooling layout of the small pores upstream and the large pores downstream of the gill area. y - z (Surface). In this embodiment of the invention, the upstream gill region 7 is the area on the suction surface of the moving blade near the trailing edge of the guide vane, where the curvature and pressure gradient change significantly. It employs a progressively increasing air film pore layout along the flow direction, with upstream small pore rows and downstream large pore rows arranged in the flow direction. Each of the upstream small pore rows and downstream large pore rows has one or more rows, and the air film pores are evenly distributed along the spanwise direction. The pore diameter of the downstream large pore row is larger than that of the upstream small pore row, forming the reference pore diameter for the entire air film pore area. This allows the cold air ejected from the upstream and downstream rows to form a cold air film adhering to the suction surface in the upstream gill region. Furthermore, in the upstream gill region 7, the air film pores preferably adopt a cylindrical structure, and the pore diameter of the air film pores in the upstream small pore row along the flow direction is 0.3 mm. D ~0.6 D The downstream large-aperture gas film pore diameter is D And as a reference aperture, in which DThe numerical range is 0.4~0.8mm. The upstream small-hole row is preferably set to one row, with 10~30 air film holes; the downstream large-hole row is preferably set to one row, with 10~30 air film holes. The upstream and downstream hole rows can be arranged in a straight line or in a staggered arrangement. The hole spacing is... P , P In 2 D ~5 D Between. The spacing between the holes is... Q , Q In 2 D ~3 D Between them, a pilot air film layer with good adhesion is formed by the upstream small-pore row, and a high-momentum enhanced air film layer is formed by the downstream large-pore row. This improves the adhesion of the air film to the suction surface in the upstream gill area under the action of unfavorable pressure gradient and strong curvature while ensuring sufficient cooling flow.

[0017] Figure 2 The image shown is a top view of the uniform cooling layout in the middle section of the midstream blades. y - z (Surface). In this embodiment of the invention, the midstream blade middle region adopts a uniform cooling layout, and the film cooling pores are cylindrical with a diameter of [missing information]. D And consistent with the reference aperture, D The diameter is between 0.4 and 0.8 mm. The film cooling pores in the mid-section of the blade can be arranged in a single row or multiple rows, with 5 to 12 pores per row. The pores in the same row are equidistant in the spanwise direction, forming a uniform cooling film with consistent pore size and uniform spacing in the mid-section of the blade, providing the area and flow rate basis for the tiered enhanced cooling in the end zone.

[0018] Figure 3 The image shown is a top view of the tiered cooling layout in the midstream leaf root region. y - z (Surface). In this embodiment of the invention, in the tiered cooling layout of the midstream leaf root region, the film cooling pores are cylindrical structures with their diameters distributed in a tiered manner along the spanwise direction. Near the strong channel vortex position on the root rim side, the film cooling pore diameter is... D 1, D 1 is distributed at 1.5 D ~3 D Between; while near the mid-leaf region, the air film pore size is D 2, D 2< D 1, D 2 distributed in 1 D ~1.5 D Between them. The tiered cooling layout in the blade root region and the uniform cooling layout in the blade mid-section are connected in rows along the span, and the number of holes in each row is the same. The number of tiered cooling holes in each row in the blade root region is between 4 and 6. The high momentum jet with a larger diameter at the root resists the entrainment effect of the secondary flow in the channel and smoothly transitions to the blade mid-section.

[0019] Figure 4 The image shown is a top view of the tiered cooling layout in the midstream blade tip region. y - z (Surface). In this embodiment of the invention, in the tiered cooling layout of the midstream blade tip region, the film cooling pores are cylindrical structures with their diameters distributed in a tiered manner along the spanwise direction. Near the location of strong leakage flow at the tip, the film cooling pore diameter is... D 3, D 3 is distributed in 1.2 D ~2 D Between. And near the mid-leaf region, the air film pore size is... D 4, D 4< D 3, D 4 distributed in 1 D ~1.2 D Between them. The tiered cooling layout in the blade tip region and the uniform cooling layout in the blade mid-section are connected in rows along the span, and the number of holes in each row is the same. The number of tiered cooling holes in each row in the blade tip region is between 3 and 4. The larger diameter jet at the tip enhances the ability to resist leakage flow in the tip gap, while also taking into account the cooling balance between the blade tip region and the blade mid-section.

[0020] Figure 5 The image shown is a top view of the concentrated tiered cooling layout in the downstream blade root region. y - z (Surface). In this embodiment of the invention, in the tiered cooling layout of the downstream blade root region, the film cooling pores are cylindrical structures, concentrated in the area not covered by the film cooling at the blade root, and their diameters are distributed in a tiered manner along the span, specifically decreasing in diameter from near the rim side towards the central region. Near the strong channel vortex position at the root, the film cooling pore diameter is... D 1, D 1 is distributed at 1.5 D ~3 D Between; decreasing along the spanning aperture until D The centralized tiered cooling system is arranged in two rows along the flow direction, upstream and downstream, with 4-6 holes in each row. The upstream and downstream hole rows can be arranged in a straight line or in a staggered pattern. The hole spacing is... P , P In 2 D ~6 D Between. The spacing between the holes is... Q , Q In 2 D ~6 D Between these, a locally concentrated enhanced cooling air film with a stepped change in pore size is formed in the area downstream of the leaf root where the air film is easily sparse or uncovered, thereby further suppressing the blowing effect of channel vortex and rotational additional force on the air film.

[0021] Figure 6 The image shown is a top view of the concentrated tiered cooling layout in the downstream blade tip area. y -z (Surface). In this embodiment of the invention, in the tiered cooling layout of the downstream blade tip region, the film cooling pores are cylindrical structures, concentrated in the area not covered by the film cooling at the blade tip, and their diameters are distributed in a tiered manner along the span, specifically decreasing in diameter from near the blade tip towards the middle region. Near the location of strong leakage flow at the tip, the film cooling pore diameter is... D 3, D 3 is distributed in 1.2 D ~2 D Between; decreasing along the spanning aperture until D The centralized tiered cooling system is arranged in two rows along the flow direction, upstream and downstream, with 3-4 holes in each row. The upstream and downstream hole rows can be arranged in a straight line or a staggered arrangement. The hole spacing is... P , P In 2 D ~4 D Between. The spacing between the holes is... Q , Q In 2 D ~4 D Between these, a moderately concentrated and appropriately intense tiered cooling zone is formed in the area downstream of the blade tip where the air film is easily sparse or uncovered, thereby suppressing leakage flow purging the air film while also controlling flow loss.

[0022] Figure 8 Partial cross-sectional view of the tiered film cooling layout ( x - z (Cross section), the angle between the centerline of the air film pores in each region and the horizontal direction (i.e., the cross section perpendicular to the spanwise direction relative to the shape of the suction surface) is... α , α Between 30° and 60°. Figure 9 Partial top view of the tiered film cooling layout ( x - z (Cross section), along the span, the angle between the centerline of the film pore and the flow direction is... β , β It can be 0°, or it can be adjusted in the opposite direction according to the direction of the secondary flow and the rotational additional force. The adjustment range is between 0° and 75°, so that the wall projection direction of the air film jet forms an angle with the transverse velocity component of the local secondary flow, so as to simultaneously take into account the air film adhesion and anti-entrainment ability in different areas, and improve the air film coverage effectiveness of the tiered cooling structure in three-dimensional complex flow.

[0023] In this embodiment of the invention, the film cooling layout obtained based on the turbine blade suction surface tiered cooling design method can have simple cylindrical holes or complex irregularly shaped holes in each region. Preferably, the film cooling holes in each region at least partially adopt an irregularly shaped hole structure with an expanding outlet section or a laterally expanded outlet profile. The inlet is approximately cylindrical, and the outlet is set as a locally expanded section or has an outwardly expanding profile extending along the span. By reducing the cold air jet velocity and increasing the jet cross-sectional area, the momentum ratio between the jet and the mainstream is reduced, improving the adhesion performance and lateral expansion capability of the film cooling system. At the same time, the turbulence of the jet is reduced, making the cold air film more uniformly cover the suction surface, improving the film cooling efficiency, especially under high blowing ratio conditions, it can effectively suppress the jet lifting and blowing-off phenomena.

[0024] It should be noted that the high-pressure turbine blade suction surface tiered cooling structure of this embodiment implements a differentiated cooling strategy based on the flow characteristics and heat load distribution of different regions of the suction surface: In the end regions such as the blade root and blade tip, where the secondary flow and leakage flow intensity are high, a tiered perforation pattern with gradually decreasing aperture from the tip to the middle is used along the spanwise direction to obtain a larger aperture and higher momentum film jet at the tip, thereby enhancing the local film's anti-entrainment and anti-blow-off capabilities; in the midstream mainstream region, a uniform perforation pattern with the same aperture as the reference is used to form a stable and continuous reference cooling film, providing an area and flow rate reference for overall cooling; in the upstream gill region, a double-layer film layout with upstream small orifice leader and downstream large orifice reinforcement is used to cope with the adverse effects of reverse pressure gradient and strong curvature. Through the above multi-region synergistic effect, while ensuring that the total cooling gas consumption is controllable, the uniformity of the overall film coverage and cooling effectiveness of the suction surface are significantly improved, achieving comprehensive optimization between cooling performance and flow loss.

[0025] Example 2: Design Method for Cascade Cooling Structure of Suction Surface of High-Pressure Turbine Blades Based on Embodiment 1 above, Embodiment 2 further provides a design method for the aforementioned high-pressure turbine blade suction surface cascade cooling structure, such as... Figure 10 As shown, its implementation includes the following steps: SS1. Obtaining Operating Conditions and Flowing Heat Loads: Based on the geometric parameters of the target high-pressure turbine stage and representative design conditions, CFD numerical calculations and / or experimental tests are used to obtain information on the wall temperature, local heat flux density, static pressure, and secondary flow intensity distribution in the end region on the suction surface of the moving blades.

[0026] As a preferred method, CFD numerical calculations employ the Reynolds-averaged Navier-Stokes equations combined with the SST k-ω turbulence model to solve the flow field of the target high-pressure turbine stage under design conditions. The computational domain must include the entire flow path of the moving blades and the tip clearance region. The wall boundary conditions are set as a no-slip adiabatic wall or a wall with constant heat flux density. The wall temperature distribution of the suction surface of the moving blades is obtained by solving the conjugate heat transfer equations. An anisotropic heat conduction model is used in the solid domain, and the interface between the fluid domain and the solid domain is coupled using temperature and heat flux continuity conditions. The intensity of secondary flow in the end region is quantitatively characterized by calculating the channel vorticity, horseshoe vorticity, and tip leakage vorticity. When the vorticity value is greater than the preset level of the mainstream vorticity average (e.g., twice), it is determined to be a strong secondary flow region, providing basic data for subsequent region division and tiered cooling layout.

[0027] SS2. Suction surface area division: Based on the static pressure gradient, curvature change, and secondary flow intensity distribution at various points on the suction surface of the blade, it is divided into the upstream gill region, the midstream main body region, and the downstream region along the gas flow direction. The midstream main body region is further divided into the midstream blade root region, the midstream blade middle region, and the midstream blade tip region in the spanwise direction. The downstream region is divided into the downstream blade root region and the downstream blade tip region in the spanwise direction. As a preferred approach, when dividing the suction surface region of the blade, a comprehensive multi-indicator criterion is adopted, mainly based on the sign and rate of change of the static pressure gradient, the change of wall curvature, and the intensity of secondary flow in the end region: the region where the static pressure gradient changes significantly from the compressive to the adverse pressure and the curvature changes drastically is identified as the upstream gill region; the middle region where the static pressure gradient is relatively stable, the curvature changes are small, and the secondary flow intensity is moderate is classified as the midstream main region; the downstream region near the blade root end wall and the blade tip gap region where the secondary flow intensity and leakage flow intensity are significantly higher than those in the middle region is classified as the downstream region, and further subdivided into the midstream blade root region, the midstream blade middle region, the midstream blade tip region, the downstream blade root region, and the downstream blade tip region based on the distribution of the spanwise pressure field and the secondary flow intensity.

[0028] SS3. Determination of the baseline cooling layout for the midstream main area: In the midstream blade mid-section, the number of rows, spanwise positions, and reference apertures of single or multiple rows of uniform film cooling holes are determined along the gas flow direction. This ensures the formation of a uniform cooling film while meeting the overall cooling gas flow constraints. Based on this, the locations of consecutive rows of film cooling holes corresponding to each row in the midstream blade root and tip regions are determined along the spanwise direction. When determining the reference cooling layout in the midstream blade mid-section, the reference film cooling hole apertures are first determined based on the maximum wall heat flux density and allowable wall temperature in the midstream blade mid-section, combined with the total available cooling gas mass flow rate and channel pressure ratio. DThe number of air film pores is determined; then, based on the temperature gradient distribution along the spanwise and flowwise directions in the midstream blade middle region, the spanwise position and pore spacing of each air film pore are determined to ensure that a continuous and uniform cooling air film is formed in the midstream blade middle region. The number and position information of the air film pores in the midstream blade middle region are used as the benchmark constraint for the tiered cooling layout of the midstream blade root region, the midstream blade tip region, and the downstream regions.

[0029] SS4. Determination of the tiered cooling layout for the terminal and gill regions: Based on the secondary flow and leakage flow intensity and the gas film thinning region in the midstream leaf root region, midstream leaf tip region, and downstream leaf root region and downstream leaf tip region, the aperture distribution and number of each exhaust film pores are determined in a progressively decreasing manner from the end to the middle along the longitudinal direction. In the upstream gill region, the position and aperture difference relationship between the upstream small pore row and the downstream large pore row are determined along the gas flow direction, forming a tiered cooling layout that connects with the midstream main region.

[0030] When determining the gradation distribution of air film pore size in the midstream and downstream blade root and tip regions, the peak local heat flux density, the intensity of secondary or leakage flow in the end region, and the spatial range of areas where the air film is easily thinned or not covered are comprehensively considered. A pore size distribution strategy that gradually decreases from the end to the middle is adopted: pores with a larger diameter than the reference pore size are set near the rim or blade tip. D The air film pores are designed to resist strong entrainment or purging effects, and the pore size is gradually reduced in stages as it transitions towards the central region until it approaches the reference pore size. D The number and distribution of holes in each stage are determined according to the air film coverage requirements to ensure that the distribution of hole diameters in each stage can enhance local cooling in the end area and achieve a smooth connection with the uniform cooling air film in the middle area of ​​the midstream blade.

[0031] Furthermore, when determining the location and pore size difference between the upstream small pore rows and the downstream large pore rows in the upstream gill region, the upstream small pore rows were positioned in the initial region before and after the transition from the hydrostatic gradient to the adverse hydrostatic gradient, and in the initial region of wall curvature change. The downstream large pore rows were positioned in the downstream region where the adverse hydrostatic gradient intensifies and the separation risk is higher. The pore size of the upstream small pores was determined as the reference pore size. D 0.3 to 0.6 times the diameter of the pilot gas film layer with good adhesion is used to first form a pilot gas film layer to stabilize the near-wall flow, and the downstream large pore diameter is determined as the reference pore diameter. D It is used to generate a reinforced air film layer with higher momentum and thickness on the basis of the pilot air film, thereby jointly improving the air film adhesion ability and cooling stability of the gill area under adverse pressure gradient and strong curvature.

[0032] SS5. Parameter Verification and Structure Output: Under the constraints of total cooling air flow, turbine stage aerodynamic losses, and blade structural strength, the aperture, row spacing, and aperture spacing of the air film vents in each region are iteratively checked and optimized until a high-pressure turbine blade suction surface tiered cooling structure that meets the design requirements is formed.

[0033] When iteratively verifying and optimizing the pore diameter, row spacing, and pore spacing of the film cooling vents in each region, a comprehensive evaluation index system is constructed. The main constraint indicators are the overall highest wall temperature of the moving blade suction surface, the temperature gradient of the blade root, blade tip, and gill area, and the increase in the total pressure loss of the turbine stage. The optimization objectives are the average film cooling efficiency, cold air utilization rate, and end-area film coverage rate. The pore diameter gradient distribution, row spacing, and pore spacing are adjusted in combination using parameter scanning or optimization algorithms. When each constraint indicator meets the preset engineering limit and the optimization objective reaches or approaches the expected threshold, the final design parameters are output. Otherwise, the adjustment continues until the comprehensive evaluation requirements are met. Furthermore, before forming the high-pressure turbine blade suction surface tiered cooling structure that meets the design requirements, the structural integrity and manufacturability of the determined film cooling hole layout are checked. This includes: checking the influence of hole diameter and hole spacing in each region on the blade body strength and stiffness to ensure that the blade safety factor meets the design specifications under centrifugal load, thermal load, and vibration load; determining the upper and lower limits of the minimum hole diameter, minimum hole spacing, and hole axis inclination angle based on the processing capability; if the film cooling hole parameters in some regions exceed the processing range, the hole diameter, number of holes, or position are locally modified without destroying the overall tiered cooling concept and regional allocation principle, thereby ensuring that the tiered cooling structure meets both aerodynamic and thermal protection performance and is engineering feasible.

[0034] Based on the above steps, to verify the effectiveness and engineering applicability of the design method in this embodiment, numerical comparative analysis was conducted under representative high-temperature and high-load conditions, using a traditional blade suction surface cooling scheme with uniformly distributed perforations along the spanwise direction and a tiered cooling structure scheme obtained according to the design method of this embodiment. The comparison results show that in the high-heat-load end regions such as the blade root, blade tip, and gill regions, the tiered cooling structure designed in this embodiment can significantly reduce the maximum wall temperature of the blade suction surface and significantly mitigate the wall temperature gradient. The end region air film coverage and air film continuity are superior to the uniformly distributed perforation scheme. Simultaneously, it maintains a cooling margin and temperature level comparable to the traditional scheme in the middle region of the mid-spindle, thus avoiding the problem of excessive cooling in the middle region due to enhanced end region cooling. Furthermore, through comprehensive optimization of the perforation distribution, perforation row spacing, and perforation spacing, this embodiment controls the increase in total turbine stage pressure loss within an acceptable engineering range, achieving a comprehensive balance between cooling performance, aerodynamic losses, and structural integrity. Therefore, the design method provided in this embodiment can significantly improve the overall thermal protection performance of the suction surface of the high-pressure turbine blade and the local end region under the constraint of limited cold air flow, and provides a feasible design approach for the safe and reliable operation of high thrust-to-weight ratio aero engines at higher turbine inlet temperatures.

[0035] The objectives of this invention have been fully and effectively achieved through the above embodiments. Those skilled in the art will understand that this invention includes, but is not limited to, the contents described in the accompanying drawings and the specific embodiments described above. Although the invention has been described with reference to what is currently considered the most practical and preferred embodiments, it should be understood that the invention is not limited to the disclosed embodiments, and any modifications that do not depart from the functional and structural principles of the invention will be included within the scope of the claims.

Claims

1. A tiered cooling structure for the suction surface of a high-pressure turbine blade, characterized in that, The suction surface of the moving blade is divided into the upstream gill region, the midstream main body region, and the downstream region along the gas flow direction, wherein: The upstream gill region is the area on the suction surface of the moving blade that is close to the trailing edge of the guide vane and has significant changes in curvature and pressure gradient. It adopts a stepped-increase air film pore layout along the flow direction, with upstream small pore rows and downstream large pore rows arranged in the flow direction. Each of the upstream small pore rows and the downstream large pore rows has one or more rows. The air film pores are evenly distributed along the span, and the pore diameter of the air film pores in the downstream large pore row is larger than the pore diameter of the upstream small pore air film pores, forming the reference pore diameter of the entire air film pore. The midstream main region is divided in the spanwise direction into the midstream blade root region near the rim, the midstream blade tip region near the tip, and the midstream blade middle region between the two. The midstream blade middle region adopts a uniform cooling layout, with its film cooling pores arranged in a single or multiple rows. The diameter of each row of film cooling pores is basically the same and corresponds to the reference pore diameter, and each film cooling pore is uniformly distributed along the spanwise direction. Both the midstream blade root region and the midstream blade tip region adopt a stepped cooling layout, with film cooling pore arrays arranged in the spanwise direction, corresponding to each row of film cooling pores in the midstream blade middle region and arranged in continuous rows. The diameter of each film cooling pore gradually decreases in the spanwise direction from the side near the rim or tip towards the midstream blade middle region. The downstream region is divided into a downstream blade root region near the rim and a downstream blade tip region near the tip in the spanwise direction. Both the downstream blade root region and the downstream blade tip region adopt a stepped cooling layout. In the flow direction, each region has at least two arrays of exhaust film holes, an upstream exhaust hole and a downstream exhaust hole. The diameter of each exhaust film hole gradually decreases from the side near the rim or blade tip towards the middle region in the spanwise direction. Furthermore, each exhaust film hole is located in the downstream areas where the air film is easily sparse or not covered at the upper and lower ends of the midstream main region in the spanwise direction.

2. The high-pressure turbine blade suction surface cascade cooling structure according to claim 1, characterized in that, The pore size of the air film pores in the upstream gill region's upstream small pore row is 0.3~0.6 mm. D The downstream large-aperture gas film pore diameter is D And as a reference aperture, in which D The numerical range is 0.4~0.8mm; the number of air film pores in the spanwise direction of both the upstream small pore row and the downstream large pore row is 10~30, and the pore spacing of each air film pore is... P 2~5 D The spacing between upstream hole rows Q 2~3 D Furthermore, the upstream and downstream orifices are arranged in a straight or staggered manner according to local geometry and flow conditions. The upstream small orifices form a pilot air film layer with good adhesion, while the downstream large orifices form a high momentum enhanced air film layer.

3. The high-pressure turbine blade suction surface cascade cooling structure according to claim 1 or 2, characterized in that, The midstream blade region adopts a uniform cooling layout, and the film cooling pores are cylindrical with a diameter of [missing information]. D The film cooling pores are arranged in a single or multiple rows along the flow direction, with 5 to 12 pores per row. The pores in the same row are equidistant in the spanwise direction, forming a uniform cooling film with consistent pore size and uniform spacing in the middle region of the midstream blade.

4. The high-pressure turbine blade suction surface cascade cooling structure according to claim 3, characterized in that, In the stepped cooling layout of the midstream leaf root region, the diameter of the film cooling pores decreases in a stepped manner along the spanwise direction, with the diameter of the film cooling pores near the rim side decreasing progressively. D 1 is 1.5~3 D The pore size of the air film pores near the middle section of the mid-upper leaf D 2 represents 1~1.5 D and D 2< D 1. Each row of air film pores has 4 to 6 pores, and they are arranged in a continuous row with the corresponding air film pores in the middle zone of the midstream leaf in the span direction.

5. The high-pressure turbine blade suction surface cascade cooling structure according to claim 4, characterized in that, In the stepped cooling layout of the midstream blade tip region, the diameter of the film cooling pores decreases in a stepped manner along the span, with the diameter of the film cooling pores near the blade tip decreasing in that direction. D 3 is 1.2~2 D The pore size of the air film pores near the middle section of the mid-upper leaf D 4 represents 1~1.2 D and D 4 less than D 3. Each row of air film pores has 3 to 4 pores, and they are arranged in a continuous row with the corresponding air film pores in the middle zone of the mid-stream leaf in the span direction.

6. The high-pressure turbine blade suction surface cascade cooling structure according to claim 5, characterized in that, In the tiered cooling layout of the downstream blade root region, the diameter of each exhaust film vent gradually decreases in the spanwise direction from the rim side towards the central region. The diameter of the exhaust film vent near the rim side is [missing information]. D 1 and D 1 in 1.5~3 D Within the range, the aperture diameter decreases gradually along the spanwise direction until the reference aperture diameter is reached. D Each venting membrane has 4 to 6 vent holes, and the spacing between the vent holes in the spanning direction is the same. P 2~6 D The spacing between adjacent air film pore rows in the flow direction Q 2~6 D The upstream and downstream rows are arranged in a straight or staggered manner.

7. The high-pressure turbine blade suction surface cascade cooling structure according to claim 6, characterized in that, In the tiered cooling layout of the downstream blade tip region, the diameter of each exhaust film vent gradually decreases in the spanwise direction from the blade tip side towards the central region. The diameter of the exhaust film vent near the blade tip side is [missing information]. D 3 and D 3 in 1.2~2 D Within the range, the aperture diameter decreases gradually along the spanwise direction until the reference aperture diameter is reached. D Each venting membrane has 3 to 4 vent holes, and the spacing between the vent holes in the span direction is the same. P 2~4 D The spacing between adjacent air film pore rows in the flow direction Q 2~4 D The upstream and downstream rows are arranged in a straight or staggered manner.

8. The high-pressure turbine blade suction surface cascade cooling structure according to claim 1, characterized in that, The centerline of the air film pore in each region has a jet angle α relative to the shape of the suction surface on a cross section perpendicular to the span, with α selected in the range of 30° to 60°. In the projection plane parallel to the suction surface, there is a flow direction deflection angle β, with β being 0° or reversed in the range of 0° to 75° according to the combined direction of the secondary flow and the rotational additional force in the local channel.

9. A design method for a tiered cooling structure for the suction surface of a high-pressure turbine blade as described in any one of claims 1 to 8, characterized in that, It should include at least the following steps: SS1. Based on the geometric parameters of the target high-pressure turbine stage and representative design conditions, obtain information on wall temperature, local heat flux density, static pressure, and secondary flow intensity distribution in the end region through CFD numerical calculations and / or experimental tests. SS2. Based on the static pressure gradient, curvature change and secondary flow intensity distribution at various points on the suction surface of the moving blade, it is divided into the upstream gill region, the midstream main body region and the downstream region along the gas flow direction. The midstream main body region is further divided into the midstream blade root region, the midstream blade middle region and the midstream blade tip region in the spanwise direction. The downstream region is further divided into the downstream blade root region and the downstream blade tip region in the spanwise direction. SS3. Determine the number of rows, spanwise position and reference aperture of a single or multiple rows of uniform film gas holes in the midstream blade mid-section along the gas flow direction, so as to form a uniform cooling film gas under the premise of satisfying the overall cold gas flow constraint. Based on this, determine the arrangement position of film gas holes in the midstream blade root region and midstream blade tip region along the spanwise direction, corresponding to each row in the midstream blade mid-section and arranged in a continuous row. SS4. Based on the secondary flow and leakage flow intensity and the gas film thinning region in the midstream leaf root region, midstream leaf tip region, and downstream leaf root region and downstream leaf tip region, the aperture distribution and number of rows of each exhaust film pores are determined in a progressively decreasing manner from the end to the middle along the longitudinal direction. In the upstream gill region along the gas flow direction, the position and aperture difference relationship of the upstream small hole row and the downstream large hole row are determined to form a stepped cooling layout that connects with the midstream main region. SS5. Under the constraints of total cooling air flow, turbine stage aerodynamic losses and blade structural strength, the diameter of the air film vents, the row spacing and the vent spacing of each region are iteratively checked and optimized until a high-pressure turbine blade suction surface tiered cooling structure that meets the design requirements is formed.

10. The method according to claim 9, characterized in that, In step SS1, the CFD numerical calculation uses the Reynolds-averaged NS equations combined with the SST k-ω turbulence model to solve the flow field of the target high-pressure turbine stage under design conditions. The wall temperature distribution of the blade suction surface is obtained by solving the conjugate heat transfer equations. The solid domain adopts an anisotropic heat conduction model, and the interface between the fluid domain and the solid domain is coupled by temperature and heat flow continuity conditions. The intensity of secondary flow in the end region is quantitatively characterized by calculating the channel vortex, horseshoe vortex, and tip leakage vortex. When the vortex value is greater than the preset level of the average value of the mainstream vortex, it is determined to be a strong secondary flow region.

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