Film cooling structure for improving the tip cooling efficiency of a turbine blade and manufacturing method

Abstract

The application discloses a gas film cooling structure and a manufacturing method for improving the tip cooling efficiency of a turbine moving blade. The gas film cooling structure comprises: a plurality of gas film cooling holes, which are arranged on the bottom wall of a groove on the moving blade tip at intervals, the groove is formed by recessing inward from the top wall of the moving blade tip, each gas film cooling hole is across the two sides of a reference flow separation line, the reference flow separation line is the flow separation line of a gas turbine at a rated speed, a tip gap leakage flow generates a pressure side corner vortex and a shaving vortex in the tip gap in opposite rotating directions, the pressure side corner vortex and the shaving vortex form a reverse vortex pair, the line driven by the reverse vortex pair at the bottom wall of the groove is a flow separation line, the tip gap is a radial gap between the moving blade tip and a casing, and the tip gap leakage flow is high-temperature gas passing through the tip gap. The cold gas flow can cool the moving blade tip and reduce the ablation risk of the moving blade tip.

Description

Technical Field

[0001] This invention relates to the field of turbine blade technology, and specifically to a film cooling structure and manufacturing method for improving the cooling efficiency of turbine blade tips under variable speed conditions. Background Technology

[0002] New energy power generation methods such as wind power and photovoltaic power generation are affected by local wind and solar resources, exhibiting characteristics such as instability, large fluctuations, and difficulty in prediction, which seriously impacts the stability of the power grid. Therefore, other power generation methods are needed to match new energy power generation and meet the grid's peak-shaving requirements. Gas turbines, with their compact structure and short start-up and shutdown times, are well-suited to meet the grid's peak-shaving needs. To fully utilize the calorific value of fuel and further improve the power generation efficiency of gas turbines, the compressor pressure ratio and turbine inlet temperature parameters within gas turbines are gradually being increased.

[0003] A turbine stage consists of several stationary blades and rotating blades, where the stationary blades are stationary components and the rotating blades are rotating components. To prevent the rotating blades from contacting and rubbing against the stationary casing during rotation, a radial clearance, called the tip clearance, is provided between the rotating blades and the casing; the portion of the rotating blade adjacent to the tip clearance is called the rotating blade tip. The pressure distribution in the flow field near the pressure side and suction side of the rotating blade tip is different, resulting in a pressure gradient on both sides of the tip. Driven by the pressure gradient at the tip, the high-temperature combustion gas accelerates through the tip clearance without doing work on the rotating blades, thus causing significant leakage losses and reducing the stage efficiency of the turbine.

[0004] To reduce tip clearance leakage flow, decrease leakage losses at the blade tip, and improve turbine stage efficiency, turbine blade designers invented a grooved blade tip design. The grooved blade tip has a simple structure and is easy to manufacture. Compared to a flat blade tip, it only adds a groove in the center of the tip, called the tip groove. The shoulder walls on both sides of the tip groove are called the pressure side shoulder wall and the suction side shoulder wall, respectively, and these two shoulder walls function similarly to sealing teeth. Due to the groove, the tip clearance leakage flow forms a clockwise rotating vortex within the groove, called the pressure side angular vortex; simultaneously, due to the relative motion between the casing and the blade tip, the tip clearance leakage flow forms a counterclockwise rotating vortex within the tip clearance, called the scraping vortex. The pressure side angular vortex and the scraping vortex rotate in opposite directions, forming a counter-rotating vortex pair, which changes the trajectory of the tip clearance leakage flow, ultimately reducing the flow velocity and flow rate, and thus reducing tip clearance leakage losses.

[0005] However, since the scraping vortex and the pressure side vortex rotate in opposite directions and form a reverse vortex pair, the leakage flow in the blade tip gap is driven by this reverse vortex pair to impact the bottom wall of the blade tip groove, thinning the fluid boundary layer at the bottom wall of the blade tip groove, reducing the convective heat transfer efficiency at this point, and forming a high heat transfer coefficient region at the bottom wall of the blade tip groove, making the blade tip easily ablated in this high heat transfer coefficient region.

[0006] In view of the above shortcomings, it is necessary to design a film cooling structure and manufacturing method to improve the cooling efficiency of turbine blade tips. Summary of the Invention

[0007] Therefore, the technical problem to be solved by the present invention is that due to the influence of the reverse vortex pair at the bottom wall of the groove, a high heat transfer coefficient zone is formed at the bottom wall of the moving blade tip, and the moving blade tip is prone to ablation. Thus, an air film cooling structure and manufacturing method for improving the cooling efficiency of turbine moving blade tip are provided.

[0008] To solve the above-mentioned technical problems, the technical solution of the present invention is as follows:

[0009] A film cooling structure for improving the cooling efficiency of turbine blade tips, the film cooling structure comprising:

[0010] Multiple film cooling holes are spaced apart on the bottom wall of a groove on the tip of the moving blade. The groove is formed by indentation from the top wall of the tip of the moving blade. Each film cooling hole spans both sides of a reference flow separation line. The reference flow separation line is the flow separation line of the gas turbine at its rated speed. The pressure side vortex and scraping vortex generated in the tip gap leakage flow in the tip gap form a counter-vortex pair with opposite rotation directions. The counter-vortex pair drives the tip gap leakage flow to form a flow separation line at the bottom wall of the groove. The tip gap is the radial clearance between the tip of the moving blade and the casing. The tip gap leakage flow is the high-temperature gas passing through the tip gap.

[0011] Multiple cold air channels are connected to each of the multiple air film cooling holes.

[0012] Furthermore, the plurality of air film cooling holes are arranged at equal intervals, and the plurality of air film cooling holes are divided into a first air film cooling hole, a last air film cooling hole, and an intermediate air film cooling hole located between the first air film cooling hole and the last air film cooling hole. The first air film cooling hole and the last air film cooling hole are respectively provided corresponding to the leading edge and the trailing edge of the reference flow separation line.

[0013] Furthermore, the cross-section of the air film cooling hole is elliptical, and the center of the ellipse is offset by a preset distance relative to the reference flow separation line towards the suction side shoulder wall of the groove. A perpendicular line is drawn from the center of each ellipse to the tangent line at the corresponding point on the reference flow separation line. The distance between the center of the ellipse and the foot of the perpendicular line is the preset distance. The ratio of the preset distance to the height of the moving blade is 0.0085 to 0.011, and the major axis of the ellipse is located in the direction of the perpendicular line.

[0014] Furthermore, the ratio of the minor axis of the ellipse to the height of the moving blade is 0.0085 to 0.011.

[0015] Furthermore, the distance between the centers of two adjacent ellipses is 6 to 10 times the minor axis of the ellipse.

[0016] Further, the ratio of the major axis to the minor axis of the ellipse is 3.8 to 4.2, or the distance between the positive flow separation line and the negative flow separation line is the separation line spacing, the ratio of the major axis of the ellipse to the separation line spacing is 1.5 to 2, the positive flow separation line is the flow separation line of the gas turbine at 1.2 times the rated speed, and the negative flow separation line is the flow separation line of the gas turbine at 0.8 times the rated speed.

[0017] Furthermore, the projection of each of the cold air channels onto the bottom wall of the groove is perpendicular to the corresponding major axis of the ellipse.

[0018] Furthermore, the included angle between the cold air channel and the bottom wall of the groove is α, where 45°≤α<90°.

[0019] The technical solution of this invention has the following advantages:

[0020] 1. The air film cooling structure for improving the cooling efficiency of turbine blade tips provided by the present invention comprises multiple air film cooling holes spaced apart on the bottom wall of a groove on the blade tip. The groove is formed by the inward indentation of the top wall of the blade tip. Each air film cooling hole spans both sides of the reference flow separation line. Therefore, the cold airflow entering the air film cooling hole from the cold air channel can produce a certain cooling effect on the groove wall, reducing the risk of blade tip ablation.

[0021] 2. The film cooling structure for improving the cooling efficiency of turbine blade tips provided by this invention has an elliptical cross-section for the film cooling holes. The center of the ellipse is offset from the reference flow separation line by a predetermined distance towards the suction side shoulder wall of the groove. A perpendicular line is drawn from the center of each ellipse to the tangent at the corresponding point on the reference flow separation line. The distance between the foot of the perpendicular line and the center of the ellipse is the predetermined distance. The ratio of the predetermined distance to the blade height is 0.0085 to 0.011. The major axis of the ellipse is located in the direction of the perpendicular line. In this way, the film cooling holes can be offset towards the suction side shoulder wall in advance, so that when the speed of the gas turbine is lower than the rated speed, the cold airflow flowing out of the film cooling holes can flow to the suction side shoulder wall, ensuring the cooling effect of the cold airflow on the groove wall.

[0022] 3. The air film cooling structure for improving the cooling efficiency of turbine blade tips provided by the present invention has an elliptical minor axis to blade height ratio of 0.0085 to 0.011, so as to balance the processing difficulty of air film cooling holes and the strength of blade tip wall plate.

[0023] 4. The air film cooling structure for improving the cooling efficiency of turbine blade tips provided by the present invention has an ellipse major axis to ellipse minor axis ratio of 3.8 to 4.2, and the ratio of the preset distance by which the center of the ellipse is offset from the reference flow separation line toward the suction side shoulder wall to the blade height is 0.0085 to 0.011. In this way, the cooling effect of the cold air flowing out of the air film cooling hole on the groove wall is better.

[0024] 5. The air film cooling structure for improving the cooling efficiency of turbine blade tips provided by the present invention has an angle α between the cold air channel and the bottom wall of the groove, where 45°≤α<90°, in order to balance the cooling effect of the cold air flow on the groove wall and the processing and design difficulty of the cold air channel inside the blade.

[0025] A method for manufacturing a film cooling structure to improve the cooling efficiency of turbine blade tips includes the following steps:

[0026] The ratio of the minor axis of the ellipse to the height of the moving blade in the elliptical cross-section of the film cooling hole is determined to be 0.0085 to 0.011; the ratio of the major axis to the minor axis is 3.8 to 4.2; or, the distance between the positive and negative flow separation lines is the separation line spacing, and the ratio of the major axis to the separation line spacing is 1.5 to 2; the pressure side vortex and scraping vortex generated in opposite directions by the tip gap leakage flow in the tip gap form a reverse vortex pair, and the line formed by the reverse vortex pair driving the tip gap leakage flow at the bottom wall of the groove is the flow separation line; the tip gap is the radial clearance between the moving blade tip and the casing; the tip gap leakage flow is the high-pressure combustion gas passing through the tip gap; the positive flow separation line is the flow separation line of the gas turbine at 1.2 times the rated speed; and the negative flow separation line is the flow separation line of the gas turbine at 0.8 times the rated speed.

[0027] Multiple reference points are selected at intervals on the reference flow separation line at the bottom wall of the groove at the tip of the moving blade. The reference flow separation line is the flow separation line of the gas turbine at its rated speed.

[0028] Draw tangents on the reference flow separation line passing through each of the reference points, and draw perpendicular lines from each of the reference points to the corresponding tangents. Offset each of the reference points along the corresponding perpendicular lines towards the suction side shoulder wall of the groove by a preset distance. The ratio of the preset distance to the height of the moving blade is 0.0085 to 0.011.

[0029] Using the reference points after moving the preset distance as the centers of the ellipse and the major axis of the ellipse located on the corresponding vertical line, a plurality of elliptical air film cooling holes are opened on the bottom wall of the groove.

[0030] Furthermore, the distance between two adjacent reference points after moving the preset distance is 6 to 10 times the minor axis of the ellipse.

[0031] Furthermore, it also includes the following steps:

[0032] A cold air passage is created such that the projection of the cold air passage onto the bottom wall of the groove is perpendicular to the major axis of the corresponding ellipse, and the angle between the cold air passage and the bottom wall of the groove is α, where 45°≤α<90°.

[0033] The technical solution of this invention has the following advantages:

[0034] 1. The manufacturing method of the film cooling structure for improving the cooling efficiency of turbine blade tips provided by the present invention determines that the ratio of the minor axis of the ellipse of the film cooling hole with an elliptical cross-section to the blade height is 0.0085 to 0.011; the ratio of the major axis to the minor axis is 3.8 to 4.2, or the distance between the positive and negative flow separation lines is the separation line spacing, and the ratio of the major axis to the separation line spacing is 1.5 to 2; multiple reference points are selected at intervals on the reference flow separation line at the bottom wall of the groove of the blade tip, the reference flow separation line being the flow separation line of the gas turbine at rated speed; and tangents are drawn on the reference flow separation line passing through each reference point. A perpendicular line is drawn from each reference point to the corresponding tangent. Each reference point is then offset by a preset distance along the corresponding perpendicular line towards the suction side shoulder wall of the groove. The ratio of the preset distance to the blade height is 0.0085 to 0.011. Taking each reference point after the preset distance as the center of an ellipse and the major axis of the ellipse as the corresponding perpendicular line, multiple elliptical film cooling holes are opened on the bottom wall of the groove. In this way, when the gas turbine rotates between the rated speed and ±20% of the rated speed, the cold airflow flowing out of the film cooling holes can flow simultaneously towards the pressure side shoulder wall and the suction side shoulder wall of the groove, forming a good cooling effect on the bottom wall of the groove and reducing the risk of easy ablation of the blade tip.

[0035] 2. The manufacturing method of the air film cooling structure for improving the cooling efficiency of turbine blade tips provided by the present invention involves opening a cold air channel, such that the projection of the cold air channel on the bottom wall of the groove is perpendicular to the corresponding major axis of the ellipse, and the angle between the cold air channel and the bottom wall of the groove is α, 45°≤α<90°, so as to take into account both the cooling effect of the cold air flow on the groove wall and the processing and design difficulty of the cold air channel inside the blade. Attached Figure Description

[0036] To more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the specific embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0037] Figure 1 This is a schematic diagram showing the positional relationship between the moving blades and stationary blades and the casing in this invention;

[0038] Figure 2 This is a schematic diagram showing the position of the flow separation line of the gas turbine at different speeds in this invention;

[0039] Figure 3 A schematic diagram showing the position of the air film cooling hole in the groove of the moving blade tip in the invention;

[0040] Figure 4This is a partial enlarged view of the groove on the tip of the moving blade in this invention;

[0041] Figure 5 This is a schematic diagram showing the positional relationship between the film cooling holes and the positive and negative flow separation lines in this invention.

[0042] Figure 6 This is a schematic diagram of the structure of the cooling air passage in the moving blade of the present invention;

[0043] Figure 7a This is a schematic diagram of the airflow in the blade tip gap of a gas turbine at rated speed under the condition that the film cooling is set along the reference flow separation line in this invention.

[0044] Figure 7b This is a schematic diagram of the airflow at 0.8 times the rated speed in a gas turbine under the condition that the film cooling is set along the reference flow separation line in this invention;

[0045] Figure 7c This is a schematic diagram of the airflow at 1.2 times the rated speed in a gas turbine under the condition that the film cooling is set along the reference flow separation line in this invention;

[0046] Figure 8a This is a schematic diagram of the airflow in the blade tip gap of a gas turbine at rated speed under the condition that the film cooling holes are offset towards the suction side shoulder wall in this invention;

[0047] Figure 8b This is a schematic diagram of the airflow in the blade tip gap of a gas turbine at 0.8 times its rated speed under the condition that the film cooling holes are offset towards the suction side shoulder wall in this invention.

[0048] Figure 8c This is a schematic diagram of the airflow in the blade tip gap of a gas turbine at 1.2 times its rated speed under the condition that the film cooling holes are offset towards the suction side shoulder wall.

[0049] Explanation of reference numerals in the attached figures:

[0050] 1. Stationary blade; 2. Moving blade; 3. Casing; 4. Stationary blade end wall; 5. Moving blade end wall; 6. Moving blade tip; 61. Leading edge; 62. Trailing edge; 7. Groove; 71. Pressure side shoulder wall; 72. Suction side shoulder wall; 81. Reference flow separation line; 82. Positive flow separation line; 83. Negative flow separation line; 91. Film cooling hole; 92. Cooling air passage; a. Pressure side angular vortex; b. Scraping vortex; A. Major axis of ellipse; B. Minor axis of ellipse; C. Tip clearance; H. Moving blade height. Detailed Implementation

[0051] The technical solution of the present invention will now be clearly and completely described with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0052] In the description of this invention, it should be noted that the terms "center," "upper," "lower," "left," "right," "vertical," "horizontal," "inner," and "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing the invention and for simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on the invention. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and should not be construed as indicating or implying relative importance.

[0053] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; and they can refer to the internal connection of two components. Those skilled in the art can understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0054] Furthermore, the technical features involved in the different embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

[0055] The present invention will now be described using the first-stage moving blade tip of a GE-E3 turbine as an example. The gas turbine stage consists of a stationary blade 1, a moving blade 2, end walls, and a casing 3. The end walls include a stationary blade end wall 4 and a moving blade end wall 5. The stationary blade 1, casing 3, and stationary blade end wall 4 are stationary components, while the moving blade 2 and moving blade end wall 5 are rotating components. The radially pointed tip of the moving blade 2 is called the moving blade tip 6. A radial clearance exists between the moving blade tip 6 and the casing 3 to prevent friction and contact; this radial clearance is called the tip clearance C. The high-temperature gas passing through the tip clearance is called the tip clearance leakage flow. The ratio of the tip clearance C to the moving blade height H is 0.0085 to 0.011, preferably 0.01.

[0056] The shoulder walls on both sides of the groove 7 (such as Figure 2(As shown) are the pressure-side shoulder wall 71 and the suction-side shoulder wall 72, respectively. The height of the pressure-side shoulder wall 71 and the height of the suction-side shoulder wall 72 are equal, both being 2% of the moving blade height H. The width of the pressure-side shoulder wall 71 and the width of the suction-side shoulder wall 72 are equal, both being 1% of the moving blade height H. Of course, the width and height of the pressure-side shoulder wall 71 and the suction-side shoulder wall 72 can be set to other values ​​according to actual needs, and no specific limitation is made here.

[0057] Example 1

[0058] like Figures 1 to 8c As shown, the air film cooling structure for improving the cooling efficiency of turbine blade tips provided in this embodiment includes multiple air film cooling holes 91 and multiple cold air channels 92, with each cold air channel 92 corresponding to and connected to the air film cooling holes 91.

[0059] Multiple film cooling holes 91 are respectively formed on the bottom wall of the groove 7, which is formed by the inward indentation of the top wall of the automatic blade tip 6. Each film cooling hole 91 spans both sides of the reference flow separation line 81. The reference flow separation line 81 is the flow separation line of the gas turbine at its rated speed. Specifically, the tip clearance leakage flow forms a pressure side vortex a and a scraping vortex b with opposite rotational directions within the tip clearance C. The pressure side vortex a and the scraping vortex b form a reverse vortex pair. The reverse vortex pair drives the tip clearance leakage flow to form the flow separation line at the bottom wall of the groove 7. The tip clearance C is the radial clearance between the moving blade tip 6 and the casing 3. The tip clearance leakage flow is the high-temperature gas passing through the tip clearance C. This reference flow separation line 81 is determined by experimental measurement or numerical calculation. Usually, the reference flow separation line 81 is approximately parallel to the pressure side shoulder wall 71 of the groove 7. Because the position of the flow separation line changes depending on the rotational speed of the gas turbine, for ease of explanation later, the flow separation line at 1.2 times the rated speed is referred to as the positive flow separation line 82, and the flow separation line at 0.8 times the rated speed is referred to as the negative flow separation line 83. The cold airflow entering the film cooling hole 91 through the cold air passage 92 can cool the groove walls on both sides of the flow separation line, producing a certain cooling effect on the groove walls.

[0060] In the first embodiment, the film cooling holes 91 are arranged at equal intervals along the reference flow separation line 81. The plurality of film cooling holes 91 are divided into a first-position film cooling hole, a last-position film cooling hole, and intermediate film cooling holes located between the first-position and last-position film cooling holes. The first-position and last-position film cooling holes are respectively positioned corresponding to the leading and trailing edges of the reference flow separation line 81, such that the first-position and last-position film cooling holes respectively cover the leading and trailing edges of the reference flow separation line 81. Figure 7aAs shown, when the gas turbine speed is at its rated speed, the cold airflow from the film cooling hole 91 can flow towards the pressure side shoulder wall 71 and the suction side shoulder wall 72, resulting in a good cooling effect on the bottom wall of the groove 7; for example... Figure 7b As shown, when the gas turbine speed is lower than the rated speed, the flow separation line will shift relative to the reference flow separation line 81 towards the suction side shoulder wall 72. Consequently, the film cooling hole 91 is located between the pressure side shoulder wall 71 and the flow separation line, and the cold airflow cannot cover the wall surface between the flow separation line and the suction side shoulder wall 72, resulting in poor cooling effect of the cold airflow on the groove wall surface between the flow separation line and the suction side shoulder wall 72. Figure 7c As shown, when the speed of the gas turbine is higher than the rated speed, the flow separation line will shift relative to the reference flow separation line 81 towards the pressure side shoulder wall 71. As a result, the film cooling hole 91 is located between the suction side shoulder wall 72 and the flow separation line. The cold airflow cannot cover the groove wall between the flow separation line and the pressure side shoulder wall 71, resulting in poor cooling effect of the cold airflow on the groove wall between the flow separation line and the pressure side shoulder wall 71.

[0061] As a second implementation method, an improvement upon the first implementation method, it allows the cool airflow to simultaneously flow towards both the pressure-side shoulder wall 71 and the suction-side shoulder wall 72 when the gas turbine is operating between 0.8 and 1.2 times its rated speed, thus achieving a good cooling effect on the groove wall surface. In this implementation method, the design of the film cooling hole 91 is as follows:

[0062] The cross-section of the film cooling hole 91 is elliptical, and there are 8 film cooling holes 91. Of course, the number of film cooling holes 91 can be set according to actual needs, and no specific limitation is made here. Preferably, when the length of the blade tip 6 along the blade tip axial direction (the blade tip axial direction is the same as the length extension direction of the rotating shaft of the gas turbine) is large, the number of film cooling holes 91 is increased accordingly.

[0063] The center of the ellipse is offset relative to the reference flow separation line 81 towards the suction side shoulder wall 72, rather than being located on the reference flow separation line 81. This offset is necessary because, at the same speed variation relative to the rated speed, when the gas turbine speed decreases, the distance the corresponding flow separation line is offset relative to the reference flow separation line 81 towards the suction side shoulder wall 72 is greater than the distance the corresponding flow separation line is offset relative to the reference flow separation line 81 towards the pressure side shoulder wall 71 when the gas turbine speed increases. A perpendicular line is drawn from the center of each ellipse to the tangent line on the reference flow separation line 81 at the corresponding point. The distance between the center of the ellipse and the foot of the perpendicular is a preset distance, and the ratio of this preset distance to the blade height H is 0.0085 to 0.011, preferably 0.01. The major axis A of the ellipse lies on the corresponding perpendicular line. The ratio of the minor axis B of the ellipse to the blade height H is 0.0085 to 0.011, preferably 0.01. If the length of the minor axis B is too small, it will increase the processing difficulty of the film cooling hole 91. If the length of the minor axis B is too large, it will result in a shorter distance between two adjacent film cooling holes 91, thus affecting the structural strength of the wall plate of the blade tip 6. The distance between the centers of the ellipses containing two adjacent film cooling holes 91 is 6 to 10 times the minor axis B of the ellipse. The ratio of the major axis A to the minor axis B of the ellipse is 3.8 to 4.2, preferably 4, to ensure that the cold airflow entering the film cooling hole 91 from the cold air passage 92 can simultaneously flow towards both the pressure side shoulder wall 71 and the suction side shoulder wall 72 when the gas turbine operates at 0.8 to 1.2 times its rated speed. Compared to the first embodiment, this increases the coverage area of ​​the cold airflow on the groove wall, resulting in a better cooling effect on the groove wall, reducing the thermal load on the moving blade tip 6, and thus reducing the risk of the moving blade tip 6 being ablated. Numerical simulation results show that when the ratio of the major axis A to the minor axis B of the film cooling hole 91 is 4, and the preset distance of the ellipse center from the reference flow separation line 81 is 1% of the moving blade height H, the cooling efficiency of the cold airflow on the moving blade tip 6 is the highest. The cooling efficiency η of the groove wall is defined as:

[0064]

[0065] In the formula, T ∞ T represents the turbine's mains temperature (unit: K); W T represents the temperature of the groove wall (unit: K); C The air conditioning temperature is expressed in Kelvin (K).

[0066] In addition, while keeping the location of the ellipse center and the method of determining the minor axis B unchanged, the major axis A of the ellipse can also be determined in other ways. For example, if the aforementioned positive flow separation line 82 and negative flow separation line 83 are determined, the distance between the positive flow separation line 82 and the negative flow separation line 83 can be defined as the separation line spacing. The ratio of the major axis A of the ellipse to the separation line spacing is 1.5 to 2. In this case, the ratio of the major axis A of the ellipse to the minor axis B of the ellipse does not need to strictly adhere to 3.8 to 4.2.

[0067] Whether the ratio of the major axis A to the minor axis B of the ellipse is 3.8 to 4.2, or the ratio of the major axis A to the separation line spacing is 1.5 to 2, the purpose is to ensure that each film cooling hole 91 can simultaneously cover the corresponding positions on the reference flow separation line 81, the positive flow separation line 82, and the negative flow separation line 83. In this way, at 0.8 to 1.2 times the rated speed of the gas turbine, the cold airflow flowing out of the film cooling hole 91 can simultaneously flow towards the suction side shoulder wall 72 and the pressure side shoulder wall 71, so as to achieve better cooling effect on the groove wall, reduce the heat load on the groove wall, and thus reduce the risk of blade tip 6 ablation.

[0068] like Figure 8a As shown, when the gas turbine speed is at its rated speed, the cold airflow from the film cooling hole 91 will flow simultaneously in both directions towards the suction side shoulder wall 72 and the pressure side shoulder wall 71, providing effective cooling protection for the groove wall surface; as Figure 8b As shown, when the gas turbine speed is 0.8 times the rated speed, the negative flow separation line 83 is offset relative to the reference flow separation line 81 towards the suction side shoulder wall 72. However, since the center of the ellipse where the film cooling holes 91 are located has been pre-offset towards the suction side shoulder wall 72, each film cooling hole 91 can still cross the negative flow separation line 83. The cold airflow from the film cooling holes 91 can still flow towards both the suction side shoulder wall 72 and the pressure side shoulder wall 71 simultaneously, thus still providing effective cooling protection for the groove wall surface; as Figure 8c As shown, when the gas turbine speed is 1.2 times the rated speed, the positive flow separation line 82 is offset relative to the reference flow separation line 81 towards the pressure side shoulder wall 71. Each film cooling hole 91 still crosses the positive flow separation line 82, and the cold airflow from the film cooling holes 91 can still flow simultaneously towards both the suction side shoulder wall 72 and the pressure side shoulder wall 71, thus still providing effective cooling protection for the groove wall surface. In summary, the film cooling holes 91 must simultaneously cross the reference flow separation line 81, the positive flow separation line 82, and the negative flow separation line 83 so that the cold airflow from the film cooling holes 91 can simultaneously flow towards both the suction side shoulder wall 72 and the pressure side shoulder wall 71.

[0069] The projection of each cooling air channel 92 onto the bottom wall of the groove 7 is perpendicular to the major axis A of the corresponding film cooling hole 91, further ensuring that the cooling airflow can flow well to both the suction-side shoulder wall 72 and the pressure-side shoulder wall 71 simultaneously. The angle between the cooling air channel 92 and the bottom wall of the groove 7 is α (e.g., ...). Figure 6 As shown in the figure, 45°≤α<90°, so that the cold airflow adheres to the groove wall, thereby improving the cooling protection effect on the groove wall. Numerical simulation results show that when the included angle α is 45°, the average cooling efficiency of the cold airflow on the groove wall is the best. When the included angle α is further reduced, it can no longer provide better cooling protection for the groove wall, and it will also increase the complexity of the cold air channel 92 structure inside the moving blade 2, thereby increasing the design and manufacturing difficulty of the moving blade 2. Specifically, in this invention, the commercial software ANSYS-CFX was used to perform numerical simulation calculations on the cooling efficiency of the groove wall of the first-stage moving blade of the GE-E3 turbine at rated speed. With the position and size of the film cooling hole 91 unchanged, the air blowing ratio N = 1.0, and the only variable being the angle α between the air passage 92 and the bottom wall of the groove 7, and with the angle α set to be 45°, 60° and 90° respectively, the simulation results show that when the angle α is 45°, the average cooling efficiency of the groove wall is 0.39; when the angle α is 60°, the average cooling efficiency of the groove wall is 0.35; and when the angle α is 90°, the average cooling efficiency of the groove wall is 0.31.

[0070] Example 2

[0071] like Figures 1 to 8c As shown, the manufacturing method of the film cooling structure for improving the cooling efficiency of turbine blade tips provided in this embodiment includes the following steps:

[0072] The minor axis B of the ellipse is determined based on the blade height H. The ratio of the minor axis B to the blade height H is 0.0085 to 0.011, preferably 0.01. The ratio of the major axis A to the minor axis B is 3.8 to 4.2, preferably 4. Alternatively, the distance between the positive flow separation line 82 and the negative flow separation line 83 is defined as the separation line spacing. In this case, the ratio of the major axis A to the separation line spacing is 1.5 to 2. In this case, the ratio of the major axis A to the minor axis B does not need to be strictly adhered to. The ratio is 3.8 to 4.2; specifically, the tip clearance leakage flow forms a pressure side vortex a and a scraping vortex b with opposite rotation directions in the tip clearance C. The pressure side vortex a and the scraping vortex b form a reverse vortex pair. The reverse vortex pair drives the tip clearance leakage flow to form a flow separation line at the bottom wall of the groove 7. The positive flow separation line 82 is the flow separation line of the gas turbine at 1.2 times the rated speed, and the negative flow separation line 83 is the flow separation line of the gas turbine at 0.8 times the rated speed.

[0073] A reference flow separation line 81 is defined on the groove 7 of the moving blade tip 6. This reference flow separation line 81 is determined by experimental measurement or numerical calculation. Multiple reference points are selected at equal intervals on the reference flow separation line 81. In this embodiment, the number of reference points is eight. Of course, other values ​​can be set according to actual needs, and no specific limitation is made here. The multiple reference points include a first reference point, a last reference point, and an intermediate reference point located between the first reference point and the last reference point. Preferably, the first reference point and the last reference point are respectively set to the leading edge of the reference flow separation line 81 (here, the leading edge is the part of the reference flow separation line 81 that is closer to the leading edge 61 of the moving blade tip 6) and the trailing edge of the reference flow separation line 81 (here, the trailing edge is the part of the reference flow separation line 81 that is closer to the trailing edge 62 of the moving blade tip 6). The reference flow separation line 81 is the flow separation line of the gas turbine at the rated speed.

[0074] Tangents to the reference flow separation line 81 are drawn through each reference point, and perpendicular lines are drawn from the corresponding tangents at each reference point. Each reference point is then offset by a predetermined distance along its corresponding perpendicular line towards the suction-side shoulder wall 72. The ratio of this predetermined distance to the blade height H is 0.0085 to 0.011, preferably 0.01. The reason the center of the ellipse is offset relative to the reference flow separation line 81 towards the suction-side shoulder wall 72, rather than located on the reference flow separation line 81, is that at the same speed variation relative to the rated speed, when the gas turbine speed decreases, the corresponding flow separation line relative to the reference flow separation line 81... The distance by which the quasi-flow separation line 81 is offset toward the suction-side shoulder wall 72 is greater than the distance by which the flow separation line corresponding to the gas turbine speed increases is offset toward the pressure-side shoulder wall 71 relative to the reference flow separation line 81. Numerical simulation results show that when the ratio of the major axis A to the minor axis B of the ellipse of the film cooling hole 91 is 4, and the preset distance of the ellipse center from the reference flow separation line 81 is 1% of the blade height H, the cooling efficiency of the cold airflow flowing out of the film cooling hole 91 on the groove wall is the highest. The distance between two adjacent reference points after offset is 6 to 10 times the minor axis B of the ellipse.

[0075] Using each offset reference point as the center of the ellipse and making the major axis A of the ellipse lie on the corresponding vertical line, multiple elliptical air film cooling holes 91 are opened on the bottom wall of the groove 7.

[0076] A cooling air passage 92 is created. First, ensure that the projection of the cooling air passage 92 onto the bottom wall of the groove 7 is perpendicular to the major axis A of the ellipse. This further ensures that the cooling airflow can flow effectively to both the suction-side shoulder wall 72 and the pressure-side shoulder wall 71 simultaneously. Then, set the angle between the cooling air passage 92 and the bottom wall of the groove 7 to α, where 45° ≤ α < 90°. This allows the cooling airflow to adhere to the groove wall, thereby improving the cooling and protection effect of the cooling airflow on the groove wall and ultimately improving the cooling and protection effect on the blade tip 6. Numerical simulation results show that when α is 45°, the average cooling efficiency of the cooling airflow on the groove wall is the best. When the angle α is further reduced, it is no longer possible to provide better cooling and protection for the groove wall, and it also increases the complexity of the cooling air passage 92 structure inside the blade 2, thus increasing the design and manufacturing difficulty of the blade 2. Specifically, in this invention, the commercial software ANSYS-CFX was used to perform numerical simulation calculations on the cooling efficiency of the groove wall of the first-stage moving blade of the GE-E3 turbine at rated speed. With the position and size of the film cooling hole 91 unchanged and the air blowing ratio N = 1.0, the only variable was the angle α between the air passage 92 and the bottom wall of the groove 7. The simulation results showed that when α is 45°, the average cooling efficiency of the groove wall is 0.39; when α is 60°, the average cooling efficiency of the groove wall is 0.35; and when α is 90°, the average cooling efficiency of the groove wall is 0.31. In addition, the projection of the air passage 92 on the bottom wall of the groove 7 is perpendicular to the major axis A of the ellipse of the corresponding film cooling hole 91.

[0077] Obviously, the above embodiments are merely illustrative examples for clear explanation and are not intended to limit the implementation. Those skilled in the art will recognize that other variations or modifications can be made based on the above description. It is neither necessary nor possible to exhaustively list all possible implementations here. However, obvious variations or modifications derived therefrom are still within the scope of protection of this invention.

Claims

1. A film cooling structure for improving the cooling efficiency of turbine blade tips, characterized in that, The air film cooling structure includes: Multiple film cooling holes (91) are spaced apart on the bottom wall of the groove (7) on the blade tip (6). The groove (7) is formed by recessing inward from the top wall of the blade tip (6). Each film cooling hole (91) crosses the reference flow separation line (81). The reference flow separation line (81) is the flow separation line of the gas turbine at the rated speed. The blade tip gap leakage flow generates pressure side vortex (a) and scraping vortex (b) with opposite rotation directions in the blade tip gap (C). The pressure side vortex (a) and the scraping vortex (b) form a reverse vortex pair. The reverse vortex pair drives the blade tip gap leakage flow to form a flow separation line at the bottom wall of the groove (7). The blade tip gap (C) is the radial gap between the blade tip (6) and the casing (3). The blade tip gap leakage flow is high-temperature gas passing through the blade tip gap (C). Multiple cold air channels (92) are connected one-to-one with multiple air film cooling holes (91); The cross-section of the air film cooling hole (91) is elliptical. The center of the ellipse is offset by a preset distance relative to the reference flow separation line (81) toward the suction side shoulder wall (72) of the groove (7). A perpendicular line is drawn from the center of each ellipse to the tangent line passing through the corresponding point on the reference flow separation line (81). The distance between the center of the ellipse and the foot of the perpendicular line is the preset distance. The ratio of the preset distance to the blade height (H) is 0.0085 to 0.

011. The major axis (A) of the ellipse is located in the direction of the perpendicular line. The ratio of the minor axis (B) of the ellipse to the height (H) of the moving blade is 0.0085 to 0.011; The distance between the centers of two adjacent ellipses is 6 to 10 times the minor axis (B) of the ellipse; The ratio of the major axis (A) to the minor axis (B) of the ellipse is 3.8 to 4.2, or the distance between the positive flow separation line (82) and the negative flow separation line (83) is the separation line spacing, the ratio of the major axis (A) to the separation line spacing is 1.5 to 2, the positive flow separation line (82) is the flow separation line of the gas turbine at 1.2 times the rated speed, and the negative flow separation line (83) is the flow separation line of the gas turbine at 0.8 times the rated speed; The projection of each of the cold air channels (92) onto the bottom wall of the groove (7) is perpendicular to the corresponding major axis (A) of the ellipse; The included angle between the cold air channel (92) and the bottom wall of the groove (7) is α, where 45°≤α<90°.

2. The film cooling structure for improving the cooling efficiency of turbine blade tips according to claim 1, characterized in that, The plurality of air film cooling holes (91) are arranged at equal intervals. The plurality of air film cooling holes (91) are divided into a first air film cooling hole, a last air film cooling hole and an intermediate air film cooling hole located between the first air film cooling hole and the last air film cooling hole. The first air film cooling hole and the last air film cooling hole are respectively provided corresponding to the leading edge and the trailing edge of the reference flow separation line (81).

3. A method for manufacturing a film cooling structure to improve the cooling efficiency of turbine blade tips, characterized in that, Includes the following steps: The ratio of the minor axis (B) of the elliptical film cooling hole (91) with an elliptical cross-section to the blade height (H) is determined to be 0.0085 to 0.011; the ratio of the major axis (A) of the ellipse to the minor axis (B) is 3.8 to 4.2, or the distance between the positive flow separation line (82) and the negative flow separation line (83) is the separation line spacing, and the ratio of the major axis (A) of the ellipse to the separation line spacing is 1.5 to 2; the leakage flow in the blade tip gap generates pressure side angle vortices (a) and scraping vortices (b) with opposite rotation directions in the blade tip gap (C), the pressure side angle... The vortex (a) and the scraping vortex (b) form a counter-vortex pair, and the counter-vortex pair drives the tip gap leakage flow to form a flow separation line at the bottom wall of the groove (7). The tip gap (C) is the radial gap between the moving blade tip (6) and the casing (3). The tip gap leakage flow is the high-temperature combustion gas passing through the tip gap (C). The positive flow separation line (82) is the flow separation line of the gas turbine at 1.2 times the rated speed, and the negative flow separation line (83) is the flow separation line of the gas turbine at 0.8 times the rated speed. Multiple reference points are selected at intervals on the reference flow separation line (81), which is the flow separation line of the gas turbine at its rated speed; Make a tangent line on the reference flow separation line (81) passing through each of the reference points, make a perpendicular line from each of the reference points to the corresponding tangent line, and shift each of the reference points along the corresponding perpendicular line towards the suction side shoulder wall (72) of the groove (7) by a preset distance. The ratio of the preset distance to the size of the moving blade height (H) is 0.0085 to 0.

011. Using the reference points after moving the preset distance as the center of the ellipse and the major axis (A) of the ellipse located on the corresponding vertical line, a plurality of elliptical air film cooling holes (91) are opened on the bottom wall of the groove (7).

4. The method for manufacturing a film cooling structure for improving the cooling efficiency of turbine blade tips according to claim 3, characterized in that, The distance between two adjacent reference points after moving the preset distance is 6 to 10 times the minor axis (B) of the ellipse.

5. The method for manufacturing a film cooling structure for improving the cooling efficiency of turbine blade tips according to claim 3 or 4, characterized in that, It also includes the following steps: A cold air passage (92) is opened, such that the projection of the cold air passage (92) on the bottom wall of the groove (7) is perpendicular to the corresponding major axis (A) of the ellipse, and the angle between the cold air passage (92) and the bottom wall of the groove (7) is α, 45°≤α<90°.

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