A design method for high-pressure turbine groove tip with shoulder wall and transverse rib joint shaping and blade

Abstract

The present invention provides a joint shaping design method of the shoulder wall and transverse rib of the tip of a high-pressure turbine groove and a blade. The method firstly translates the pressure side shoulder wall in the range of 10% to 80% of the axial chord length inwardly for a set distance for the groove tip, and tilts the shoulder wall to the pressure side at a certain angle. The pressure side shoulder wall shaping can increase the tip leakage resistance and reduce the tip leakage loss. Next, four air film holes are arranged along the middle arc line at the bottom of the tip groove, and the center position of the holes is determined by the arc length ratio of the middle arc line occupied, which can facilitate the fine layout of the air film holes. Then, a transverse rib structure is arranged between adjacent air film holes, and the control parameters are the width of the transverse rib, the angle between the transverse rib and the middle arc line, and the position of the transverse rib. The transverse rib structure divides the tip groove into a front cavity and a rear cavity, which can regulate the transport and mixing mechanism of the air film jet, improve the air film cooling effect of the front part of the tip groove, and improve the cooling performance of the turbine tip while reducing the tip leakage loss.

Description

Technical Field

[0001] The invention belongs to the field of aero-engine turbine design, and in particular relates to a high-pressure turbine groove tip design method and a blade with shoulder wall and transverse rib combined molding. Background Art

[0002] The high-pressure turbine is a key hot-end component of an aircraft engine, and its aero-thermal performance and reliability are crucial to the high performance and long service life of the aircraft engine. In particular, in recent years, with the continuous increase in the temperature before the turbine inlet and the stage load, the high-pressure turbine tip has become a high-pressure area of ​​thermal failure, and a new flow control and cooling layout method is urgently needed.

[0003] The high-pressure turbine blades are high-speed rotating parts. In order to prevent collision and friction with the casing, there must be a gap between the blade tip and the casing. Due to the existence of the circumferential pressure gradient, a part of the mainstream fluid will flow from the pressure side through the tip gap area to the suction side, forming a tip leakage flow and other related secondary vortex systems. Since the tip leakage flow cannot do work, it will cause part of the mainstream energy loss; at the same time, these large-scale secondary vortex systems will undergo strong shear and mixing with the mainstream fluid, resulting in corresponding mixing losses. Studies have shown that an increase of 1% in the blade tip clearance can lead to a 0.8% to 1.8% decrease in vortex magnitude efficiency.

[0004] In addition to causing high aerodynamic losses, the tip clearance also poses severe challenges to the cooling layout and heat transfer design of the blade tip. On the one hand, the leakage flow in the turbine tip area is characterized by fast flow velocity, thin boundary layer, and large heat transfer coefficient, especially for the flow reattachment area on the blade top surface, the leakage vortex impact suction surface area, and the trailing edge area of ​​the blade. The convective heat transfer coefficients of these areas with the high-temperature mainstream are extremely high, and high-temperature creep and ablation are prone to occur. On the other hand, the existence of a complex secondary vortex system at the blade tip will affect the cold air transport and wall adhesion in the blade tip area, making the cooling of the blade tip more difficult.

[0005] In order to improve the aero-thermal performance of turbine blade tips and enhance the cooling effect of blade tips, domestic and foreign scholars and technicians have successively proposed blade tip shaping methods including grooves and winglets. Among them, groove blade tip shaping can effectively reduce blade tip leakage loss and heat transfer coefficient, and its configuration is easy to process, so it has been widely used. Although shaping methods such as blade tip winglets can help reduce blade tip leakage loss, they have the disadvantages of difficult cooling and processing, so they are less used. Summary of the invention

[0006] In order to solve the problems existing in the prior art, the present invention provides a blade tip shaping design method for reducing the leakage loss and blade tip heat load of a high-pressure turbine moving blade. Based on the groove blade tip in actual application, the shoulder wall on the pressure surface side of the blade tip is pushed inward and tilted to suppress the leakage flow, and at the same time, transverse ribs are arranged in the front part of the groove to regulate the mixing and covering effects of the cooling airflow.

[0007] In order to achieve the above object, the technical solution adopted by the present invention is: a design method for a high-pressure turbine groove blade tip with shoulder wall and transverse rib combined molding, comprising the following steps:

[0008] S1, for the tip area of ​​the high-pressure turbine moving blade, the inner cavity profile of the groove is formed by uniformly offsetting the blade profile inward by a distance t, and a chamfer is used near the trailing edge of the blade for smooth connection; at the same time, the inner cavity profile of the groove is extended downward by a distance h, and a blade tip groove with uniform shoulder wall width and groove depth is formed by Boolean cutting;

[0009] S2, the pressure side shoulder wall is translated into the groove chamber by a set distance, and the shoulder wall is tilted toward the pressure surface of the blade by a set angle to form the pressure side shoulder wall shape. The main control parameters include the translation distance w S and the tilt angle θ;

[0010] S3, the pressure side shoulder wall shaping area and the adjacent unshaped areas at both ends are smoothly connected using spline surfaces;

[0011] S4, four air film holes are arranged along the middle arc line of the bottom section of the blade tip groove, and the control parameters include the aperture D and the center position x1-x4; the center position is determined by the arc length ratio of the middle arc line;

[0012] S5, arrange the transverse rib shape in the front area of ​​the blade tip groove, and its control parameters include the transverse rib shape width w R , the relative position d between the transverse rib shape and the leading edge of the blade and the inclination angle α of the transverse rib structure; the transverse rib structure divides the tip groove chamber into a front cavity and a rear cavity;

[0013] S6, using global optimization algorithm and CFD simulation, taking isentropic efficiency and maximum tip temperature as optimization targets or constraints, finely optimize the geometric parameters of the pressure side shoulder wall shape and the groove cross rib shape, as well as the layout parameters of the film holes at the bottom of the groove, and obtain the layout parameters;

[0014] S7. According to the layout parameters, the CFD simulation technology is used to perform a gas-thermal coupling analysis on the turbine groove blade tip with the proposed shoulder wall and transverse rib joint modeling to obtain the optimal geometric parameters corresponding to different operating conditions.

[0015] The groove depth h is 2.5% of the relative blade height, and the shoulder wall thickness t is 1.7% of the relative blade height.

[0016] The pressure side shoulder wall shaping range is controlled between 10% and 80% of the axial chord length.

[0017] The distance w that the pressure side shoulder wall moves inward S , which is between 50% and 150% of the local shoulder thickness, and the inclination angle θ of the pressure side shoulder is between 45° and 75°; the translation distance of the top profile of the pressure side shoulder is determined by w S -hcos(θ) calculation.

[0018] Firstly, the film holes are evenly arranged on the bottom surface of the blade tip groove along the mid-arc line. Then, the position parameters of the film holes are fine-tuned using a global optimization algorithm to adapt to the complex flow structure in the blade tip groove.

[0019] The included angle α between the groove transverse rib shape and the perpendicular line of the center arc line ranges from -30° to 0°.

[0020] The optimal geometric parameter combination of the pressure side shoulder wall shape and the groove cross rib shape is obtained through a global optimization algorithm based on the Kriging proxy model.

[0021] A turbine blade, the blade tip of the turbine blade is obtained based on the high-pressure turbine groove blade tip design method of the shoulder wall and transverse rib combined molding described in the present invention.

[0022] Compared with the prior art, the present invention has at least the following beneficial effects: the present invention pushes the pressure side shoulder wall inward and tilts it, increases the turning angle of the tip leakage flow entering the tip gap, thereby increasing the leakage resistance to reduce the tip leakage loss. At the same time, four air film holes are arranged along the middle arc line at the bottom of the tip groove, and the position of the center of the circle is determined by the arc length ratio of the middle arc line, which can facilitate the designer to perform parametric layout and fine adjustment of the air film holes; in addition, a transverse rib is arranged between the first and second air film holes in the front of the groove, which is conducive to enhancing the air film cooling effect in the front of the groove; through the joint shaping of the pressure side shoulder wall and the groove transverse rib, the cooling performance of the turbine tip can be improved while reducing the tip leakage loss. BRIEF DESCRIPTION OF THE DRAWINGS

[0023] Figure 1 The present invention provides a high-pressure turbine rotor blade structure that adopts a pressure-side shoulder wall and a groove transverse rib combined with a mold;

[0024] Figure 2 It is a schematic diagram of the turbine groove blade tip structure using joint shaping;

[0025] Figure 3 A cross-sectional view of the blade tip pressure side shoulder wall;

[0026] Figure 4 This is a schematic diagram of the layout of the air film holes at the bottom of the blade tip groove;

[0027] Figure 5 This is a schematic diagram of the blade tip groove transverse rib modeling;

[0028] Figure 6 The streamline distribution of the blade tip area after the blade tip joint shaping provided by the present invention is adopted;

[0029] Explanation of the symbols in the accompanying drawings: In the figure, 1—the tip of the high-pressure turbine moving blade, 2—the trailing edge of the blade, 3—the leading edge of the blade, 4—the pressure surface of the blade, 11—the tip groove, 12—the pressure side shoulder wall shape, 13—the cross rib shape, 14—the air film hole at the bottom of the groove, 121—the starting position of the pressure side shoulder wall shape, 122—the ending position of the pressure side shoulder wall shape, 123—the front turning point of the pressure side shoulder wall shape, 124—the rear turning point of the pressure side shoulder wall shape, 5—the suction surface of the blade, 6—the casing, θ—the inclination angle of the pressure side shoulder wall shape, w S —Translation distance of pressure side shoulder wall shaping, h—Deepness of blade tip groove, c—Size of blade tip clearance, t—Thickness of groove shoulder wall, 141—First air film hole, 142—Second air film hole, 143—Third air film hole, 144—Fourth air film hole, d—Relative distance from the intersection of the front side of the transverse rib shaping and the median arc line to the leading edge of the blade, α—Angle between the transverse rib shaping and the perpendicular line of the median arc line, w R —width of the transverse rib shape, 15—mid-arc line of the cross section where the groove bottom is located, 71—platform vortex of the pressure side shoulder wall shape, 72—pressure side swirling vortex, 73—scraping vortex, 74—leakage vortex, 75—channel vortex. DETAILED DESCRIPTION

[0030] The following will take a typical first-stage rotor blade of a high-pressure turbine as an example to introduce in detail the specific implementation measures of the blade tip shaping in the present invention:

[0031] Figure 1 The geometric structure of the first-stage moving blade of a high-pressure turbine formed by adopting the blade tip shaping method provided by the present invention is given, and its blade profile adopts the GE-E3 first-stage moving blade of a high-pressure turbine provided in a NASA public report. In order to better explain the influence mechanism and effect of blade tip shaping, the internal chamber structure of the blade and the air film holes on the blade surface are reasonably simplified, and only the geometric shape and cooling layout of the blade tip area are explained.

[0032] Figure 2 A schematic diagram of the structure of a high-pressure turbine blade tip 1 after combined modeling is given. The groove blade tip mainly includes four main structures, namely, a blade tip groove 11, a pressure side shoulder wall modeling 12, a transverse rib modeling 13, and a groove bottom air film hole 14. For the commonly used turbine groove blade tip, the present invention proposes a combined modeling method of the blade tip pressure side shoulder wall and the groove transverse rib and a parameterized layout method of the groove air film hole;

[0033] First, combine Figure 3 The cross-sectional view of the blade tip area is given to explain the modeling method and modeling parameters of the blade tip groove. For the blade tip groove modeling, the blade top profile is first offset inward by a set distance to obtain the basic contour line of the inner cavity of the blade tip groove. In the area near the trailing edge 2, a rounding is used for smooth transition. The thickness t of the shoulder wall of the groove is determined by the offset distance. After determining the thickness of the shoulder wall of the groove, the groove structure of the blade tip can be obtained by giving the blade tip groove depth h and using the Boolean excision method. Figure 3 The tip groove depth h is taken as 2.5% of the relative blade height, and the shoulder wall thickness t is taken as 1.7% of the relative blade height.

[0034] Further, combined with Figure 3 The cross-sectional view of the blade tip area is given to explain the shaping method and shaping parameters of the pressure side shoulder wall; for the pressure side shoulder wall shaping 12, it is first necessary to determine the range of the shoulder wall shaping, and comprehensively consider the application effect and processing difficulty of the pressure side shoulder wall shaping, and control the pressure side shoulder wall shaping range between 10% and 80% of the axial chord length. The overall shaping range of the pressure side shoulder wall is determined by the starting position 121 of the pressure side shoulder wall shaping and the ending position 122 of the pressure side shoulder wall shaping, and the range of the pressure side shoulder wall uniformly pushed inward is further determined by the front turning point 123 of the pressure side shoulder wall shaping and the rear turning point 124 of the pressure side shoulder wall shaping. S It represents the translation distance of the pressure side shoulder wall modeling. The recommended value is between 50% and 150% of the shoulder wall thickness. Figure 3 In the example given, the distance that the pressure side shoulder wall is translated inward is approximately equal to the thickness of the shoulder wall. Then, the top line of the shoulder wall is translated outward for a certain distance to make the shoulder wall inclined. The inclination angle is represented by θ, and its recommended value range is 45° to 75°. Figure 3 In the example given, the pressure side shoulder wall has an inclination angle of about 60°. Therefore, the translation distance of the pressure side shoulder wall top profile can be calculated by w S -hcos(θ) is calculated to ensure the required shoulder wall inclination angle. In the specific implementation process, the pressure side profile of the blade tip area of ​​the original design is first extracted and interpolated with a cubic spline curve, and the profile points within the range of the front turning point 123 of the pressure side shoulder wall shape and the rear turning point 124 of the pressure side shoulder wall shape are translated inward by w S. Then, a spline curve is used to smoothly connect the transition area between the starting position 121 of the shape and the front turning point 123 of the pressure side shoulder wall shape and between the rear turning point 124 of the pressure side shoulder wall shape and the end position 122 of the pressure side shoulder wall shape, so that the bottom profile of the pressure side shoulder wall shape can be obtained. In a similar manner, according to the inclination angle θ of the pressure side shoulder wall and the groove depth h, the top profile of the pressure side shoulder wall shape can be obtained. The above method can ensure that the blade tip groove depth h and the blade tip clearance c remain unchanged while obtaining the pressure side shoulder wall shape 12.

[0035] Further, combined with Figure 4 The layout method and parameters of the film holes of the blade tip groove are described. Along the middle arc line 15 of the groove contour, four film holes 14 are arranged at the bottom of the blade tip groove, namely the first film hole 141, the second film hole 142, the third film hole 143 and the fourth film hole 144. The main control parameters of the film holes are the aperture D and the relative positions of the centers of the circles x1, x2, x3 and x4. Specifically, x1 represents the distance from the center of the first film hole 141 to the leading edge point of the blade, and x2, x3 and x4 represent the distances between the centers of two adjacent film holes, respectively. In order to facilitate the parametric layout of the film holes, x1-x4 are all normalized by the total arc length of the middle arc line. For example, x2=0.1 means that the distance between the centers of the second film hole 142 and the first film hole 141 on the middle arc line is 10% of the total arc length of the middle arc line. In order to make the cooling coverage of the blade tip groove more sufficient and uniform, the film holes are first evenly arranged in the entire groove, and then the position parameters of the film holes are fine-tuned to adapt to the complex flow structure in the blade tip groove. At the same time, the diameter D of the film holes is adjusted within a certain range, so that the cooling air can be dynamically adjusted according to the cooling needs of different positions. For turbine blade tips with different geometries and working conditions, the optimal film hole position and diameter parameters can be obtained by fine optimization within the recommended range.

[0036] Combination Figure 4 and Figure 5 The cross-rib shaping method and shaping parameters of the blade tip groove are described. The cross-sectional shape and position of the groove cross-rib shaping 13 are mainly determined by the width w of the cross-rib shaping. R , the angle α between the groove transverse rib and the perpendicular line of the middle arc line 15, and the distance d from the intersection of the groove transverse rib and the middle arc line 15 to the leading edge 3 of the blade. R The value of is recommended to be equal to the shoulder wall thickness t. On the one hand, it can avoid the processing difficulties caused by too thin transverse ribs, and on the other hand, it can avoid the cooling performance degradation caused by too thick transverse ribs. The angle α is negative on the left side of the perpendicular line of the middle arc line 15 and positive on the right side. Figure 5The angle shown in is a negative value. Compared with the positive angle, the negative angle helps to guide the leakage flow to the pressure side, which is beneficial to improve the gas-thermal performance of the blade tip. Therefore, the recommended value range of the angle α is -30° to 0°. The cross rib is arranged between the first film hole and the second film hole, which can divide the blade tip groove into two chambers, the front and the middle and rear, which helps to improve the cooling effect of the film cooling air on the front area of ​​the blade tip groove. Therefore, the position parameter d of the blade tip groove cross rib is given after the film hole cooling layout is completed, and fine-tuned to achieve a better cooling effect. According to the above three parameters, the cross-sectional contour line of the groove cross rib shape can be obtained. Then the contour line of the groove cross rib is stretched upward from the bottom of the groove to be flush with the top surface of the shoulder wall, and smoothly connected to the shoulder walls on both sides of the blade pressure surface 4 and the blade suction surface 5, so as to obtain the cross rib shape of the blade tip groove.

[0037] Furthermore, the geometric parameters of the high-pressure turbine groove tip with shoulder wall and transverse rib joint shaping are finely optimized by combining the global optimization algorithm based on the Kriging model and CFD simulation technology. During optimization, the isentropic efficiency and the maximum tip temperature can be used as optimization targets or constraints according to the needs of the designers. The groove tip structure in the attached figure of the present invention is a specific example obtained through fine optimization. Specifically, the position parameters of the pressure side shoulder wall shaping optimized in this example are 10% to 60% of the axial chord length, and the translation distance w S is 0.7 mm, the inclination angle θ is 60°, and the width w of the groove cross rib shape optimized in this example R The angle α between the groove transverse rib shape and the perpendicular line of the center arc line 15 is 0°, and the distance d from the intersection of the groove transverse rib shape and the center arc line 15 to the blade leading edge 3 is 19.5% of the total length of the center arc line 15.

[0038] Finally, the effect of the blade tip shaping method provided by the present invention is described. Figure 6The flow and vortex structure near the blade tip using the combined shaping of the shoulder wall and the transverse rib is given. A small platform vortex 71 is formed on the shaping platform of the pressure side shoulder wall due to the flow turning; the leakage flow entering the groove from the pressure surface 4 of the blade and the leading edge 3 of the blade forms a pressure side swirl vortex 72 in the tip groove 11 near the pressure side shoulder wall; near the upper casing 6, a scraping vortex 73 is formed due to the relative movement between the upper casing and the moving blade; the leakage flow flows outward from the shoulder wall of the suction surface 5 of the blade to form a leakage vortex 74; the circumferential pressure gradient in the turbine blade grid leads to the formation of a channel vortex 75. Compared with the original groove blade tip, the combined shaping of the blade tip of the present invention improves the efficiency of the turbine stage and reduces the metal temperature in the turbine blade tip area. The reason is that the pressure-side shoulder wall shape increases the turning angle of the tip leakage flow, thereby increasing the leakage resistance and reducing the tip leakage loss. In addition, the groove cross rib shape hinders the front airflow from flowing backward, changing the migration and mixing mechanism of the cold air inside the groove, making it difficult for the cold air in the front chamber of the groove to expand backward, thereby enhancing the film cooling effect in front of the tip groove. A vortex structure is formed behind the cross rib, so that the cold air of the film hole behind the rib can cover a larger area. In summary, the combined shape of the shoulder wall and the cross rib can reduce the tip leakage loss while improving the cooling effect of the tip area.

Claims

1. A design method for a high-pressure turbine groove blade tip with shoulder wall and transverse rib joint molding, characterized in that: The following steps are involved: S1, for the tip area of ​​the high-pressure turbine rotor blade, the inner cavity profile of the groove is formed by uniformly offsetting the blade profile inward by a distance t, and a chamfer is used in the near-trailing edge area for smooth connection; at the same time, the inner cavity profile of the groove is extended downward by a distance h, and a blade tip groove (11) with a uniform shoulder wall width and groove depth is formed by Boolean cutting; S2, the pressure side shoulder wall is translated into the groove chamber by a set distance, and the shoulder wall is tilted toward the pressure surface of the blade by a set angle to form a pressure side shoulder wall shape (12). Its main control parameters include the translation distance w S and the tilt angle θ; S3, the pressure side shoulder wall shaping area and the adjacent unshaped areas at both ends are smoothly connected using spline surfaces; S4, arranging four air film holes (14) along the middle arc line of the bottom cross section of the blade tip groove, wherein the control parameters thereof include the aperture D and the center position x1-x4; the center position thereof is determined by the arc length ratio of the middle arc line occupied; S5, arranging a transverse rib shape (13) in the front area of ​​the blade tip groove (11), and its control parameters include the width w of the transverse rib shape R , the relative position d between the transverse rib shape and the leading edge (3) of the blade, and the inclination angle α of the transverse rib structure; the transverse rib structure divides the blade tip groove chamber into a front cavity and a rear cavity; S6, using global optimization algorithm and CFD simulation, taking isentropic efficiency and maximum tip temperature as optimization targets or constraints, finely optimize the geometric parameters of the pressure side shoulder wall shape and the groove cross rib shape, as well as the layout parameters of the film holes at the bottom of the groove, and obtain the layout parameters; S7. According to the layout parameters, the CFD simulation technology is used to perform a gas-thermal coupling analysis on the turbine groove blade tip with the proposed shoulder wall and transverse rib joint modeling to obtain the optimal geometric parameters corresponding to different operating conditions.

2. The design method for high-pressure turbine groove blade tip with shoulder wall and transverse rib joint molding according to claim 1 is characterized in that: The groove depth h is 2.5% of the relative blade height, and the shoulder wall thickness t is 1.7% of the relative blade height.

3. The design method for high-pressure turbine groove blade tip with shoulder wall and transverse rib joint molding according to claim 1 is characterized in that: The pressure side shoulder wall shaping range is controlled between 10% and 80% of the axial chord length.

4. The design method for high-pressure turbine groove blade tip with shoulder wall and transverse rib joint molding according to claim 1 is characterized in that: The distance w that the pressure side shoulder wall moves inward S , which is between 50% and 150% of the local shoulder thickness, and the inclination angle θ of the pressure side shoulder is between 45° and 75°; the translation distance of the top profile of the pressure side shoulder is determined by w S -hcos(θ) calculation.

5. The design method of high-pressure turbine groove blade tip with shoulder wall and transverse rib joint molding according to claim 1 is characterized in that: Firstly, the film holes are evenly arranged on the bottom surface of the blade tip groove along the mid-arc line. Then, the position parameters of the film holes are fine-tuned using a global optimization algorithm to adapt to the complex flow structure in the blade tip groove.

6. The design method for high-pressure turbine groove blade tip with shoulder wall and transverse rib joint molding according to claim 1 is characterized in that: The included angle α between the groove transverse rib shape and the perpendicular line of the center arc line ranges from -30° to 0°.

7. The design method for high-pressure turbine groove blade tip with shoulder wall and transverse rib joint molding according to claim 1 is characterized in that: The optimal geometric parameter combination of the pressure side shoulder wall shape and the groove cross rib shape is obtained through a global optimization algorithm based on the Kriging proxy model.

8. A turbine blade, characterized in that: The tip of the turbine blade is obtained based on the high-pressure turbine groove tip design method of shoulder wall and transverse rib combined molding as described in any one of claims 1 to 7.

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