A method for organizing flow in boundary layer of high-load turbine end zone

Abstract

The present invention relates to a flow organization method for a boundary layer at a high-load turbine end region, comprising the following steps: SS1, determining the setting position of a boundary layer distortion correction blade; SS2, determining the distortion angle of the boundary layer distortion correction blade; SS3, determining the three-dimensional blade row design of the boundary layer distortion correction blade. The present invention eliminates the negative angle of attack caused by the distortion of the end wall boundary layer by using an airflow correction blade near the proximal region of the rotor-static interface, so that the velocity direction of the proximal region fluid entering the downstream blade row channel is basically consistent with the mainstream velocity direction, thereby achieving the control of the distortion of the boundary layer at the end wall of the high-load turbine, optimizing the flow field structure of the inlet end region, reducing the secondary flow loss, and improving the aerodynamic performance of the high-load turbine.

Description

Technical Field

[0001] The present invention belongs to the technical field of impeller machinery, and relates to a flow organization method for a boundary layer in a high-load turbine end region, and specifically to a method for controlling the distortion of a boundary layer in a high-load turbine end region by using airflow correction blades at a rotor-static interface. Background Art

[0002] In high bypass ratio aviation turbofan engines, the relatively low inlet gas temperature and pressure limit the output power of a single-stage low-pressure turbine. At the same time, due to the limitations of fan speed and high bypass ratio, the number of low-pressure turbine stages is usually large, and it is one of the largest components inside the turbofan engine, and its weight can account for 20% to 30% of the weight of the entire engine. On the other hand, the low-pressure turbine has a significant impact on the performance of high bypass ratio turbofan engines. Taking the engine fuel consumption rate as an example, in the cruising state, if the low-pressure turbine efficiency increases by 1%, the engine fuel consumption rate will decrease by 0.7% to 0.9%, becoming one of the components that affects the performance of the entire engine the most. With the continuous pursuit of high bypass ratio turbofan engine performance, high-load low-pressure turbine design has become a trend. Reducing the number of low-pressure turbine blades, or even the number of stages, through high-load blade design is an effective way to reduce the weight of the low-pressure turbine and improve engine performance, and has received widespread attention from researchers. In recent years, high lift, low aspect ratio, and large turning angle blades have become a new trend in low-pressure turbine design. However, the increase in blade load will inevitably increase the lateral pressure difference in the airflow channel and the flow-direction adverse pressure gradient in the blade boundary layer, which can easily lead to boundary layer separation, aggravate blade profile losses and secondary flow losses in the end area. Therefore, studying the complex flow in the end area of ​​the high-load low-pressure turbine and improving the aerodynamic performance of the low-pressure turbine are of great significance in the refined design of the low-pressure turbine of a high-performance, high-bypass ratio turbofan engine.

[0003] At present, we have a comprehensive understanding of the formation and development mechanism of secondary flow in the end area of ​​high-load turbine blades under steady conditions. However, under steady flow conditions, the complex geometric and aerodynamic environment such as unsteady effects and potential effects inside the real engine is not taken into account, resulting in the research results under this condition being unable to accurately reflect the complex evolution process of secondary flow in the end area of ​​high-load low-pressure turbines. Therefore, more and more researchers have begun to study the secondary flow in the end area of ​​high-load low-pressure turbines under turbine stages or unsteady conditions. Among them, the distortion of the end wall boundary layer is one of the flow characteristics caused by the unsteady interference of the rotor and stator blade rows.

[0004] Taking the migration of airflow from the moving blade row to the stationary blade row as an example, the boundary layer distortion at the inlet end area mainly occurs when the airflow inside the turbine stage flows from the stationary blade row to the moving blade row, or migrates from the moving blade row to the stationary blade row. Due to the relative motion between the turbine stages, the boundary layer is distorted under the action of the wall shear force. Figure 1As shown in the figure, the cause of the distortion of the boundary layer is analyzed through the velocity triangle: the rotor speed u, the absolute velocity c of the upstream outlet airflow and its relative velocity w form a velocity triangle. Since the outlet absolute velocity c2 of the mainstream is relatively large, under the action of the upstream rotating end wall, its airflow enters the stator channel at a velocity w2 close to zero angle of attack. However, for the low-energy fluid in the end zone boundary layer, since the outlet absolute velocity c2' is relatively small, under the action of the upstream rotating end wall, this part of the fluid enters the downstream stator channel at a velocity w2' with a relatively large negative angle of attack, causing a large deviation between the low-energy fluid in the inlet end wall boundary layer and the mainstream velocity direction, resulting in the distortion of the stator inlet end wall boundary layer morphology, affecting the development of the secondary flow in the end zone of the downstream blade channel.

[0005] The study found that the distortion of the boundary layer of the end wall increases the negative angle of attack of the fluid in the proximal area, which increases the thickness of the boundary layer in the proximal area of ​​the leading edge of the blade, resulting in a large flow loss in the inlet end area. As the speed of the upstream rotating end wall increases, the flow loss gradually increases, deteriorating the aerodynamic performance of the high-load turbine. Therefore, how to control the boundary layer distortion of the end area of ​​the high-load turbine, organize and optimize the end area flow field structure in the cascade channel, thereby reducing the secondary flow loss and improving the aerodynamic performance of the turbine, is an important issue that needs to be solved urgently. Summary of the invention

[0006] In order to reduce the flow loss, the present invention proposes a flow organization method for the boundary layer of the end area of ​​a high-load turbine. By designing an airflow correction blade near the proximal end area of ​​the rotor-static interface, the negative angle of attack caused by the distortion of the end wall boundary layer is eliminated, thereby achieving the control of the distortion of the end wall boundary layer of the high-load turbine, optimizing the end area flow field structure, and reducing secondary flow losses.

[0007] The technical solution adopted by the present invention to solve its technical problems is:

[0008] A flow organization method for a high-load turbine end region boundary layer of the present invention comprises the following steps:

[0009] SS1, determine the setting position of the boundary layer distortion correction blade:

[0010] A boundary layer distortion correction blade is arranged between the upstream rotor blade and the downstream stator blade, and the distance between the leading edge of the boundary layer distortion correction blade and the rotor-stator interface and the distance between the trailing edge of the boundary layer distortion correction blade and the leading edge of the downstream stator blade are determined.

[0011] SS2, determine the twist angle of the boundary layer twist correction blade:

[0012] The rim blade inlet structural angle α1, the hub blade inlet structural angle α2, the rim blade outlet structural angle, and the hub blade outlet structural angle are defined, wherein the rim and hub inlet structural angles are determined by the actual outlet angle of the low-energy fluid in the upstream rotor end area, and according to the end wall boundary layer distortion law, the hub blade inlet structural angle α2 gradually decreases with the increase of the spanwise height, and the rim blade inlet structural angle α1 is smaller than the hub blade inlet structural angle α2; the rim and hub outlet structural angles are determined by the angle at which the end area fluid enters the downstream stator blade channel, and in order to ensure that the outlet airflow angle of the fluid in the trailing edge end area of ​​the boundary layer distortion correction blade is consistent, the rim blade outlet structural angle and the hub blade outlet structural angle are the same.

[0013] SS3, 3D Cascade Design for Boundary Layer Distortion Correction Blades:

[0014] SS31, determine the actual blade size: follow the above steps to complete the design of the rim blade and hub blade, and enlarge or reduce the blade size proportionally according to the actual turbine size.

[0015] SS32, determine the actual size of blades with different cross-sections: Since the real aircraft engine is an annular structure, the flow area of ​​the hub and the rim are different, and the pitch t is also different. In order to ensure that the boundary layer distortion correction blade has uniform load distribution along the span direction, the hub blade pitch and the rim blade pitch need to be determined according to the hub radius and the rim radius. At the same time, the pitch at the corresponding position is determined according to the different cross-sectional radius of the blade. The calculation formula is as follows:

[0016]

[0017] Where R is the radius of the blade section (including the rim blade section and the hub blade section), and n is the number of blades, whose value is determined based on the radius of the reference blade section and the pitch.

[0018] Then, according to the ratio of the pitches between different spanwise heights, the scaling ratio of the blade profile is determined, and the scaling calculation formula for the actual height section blade profile size is as follows:

[0019]

[0020] Where t1 is the reference blade pitch, t2 is the calculated blade pitch, x1 is the reference blade size, and x2 is the calculated blade size.

[0021] SS33, 3D stacking: After determining the position and size of the blade sections at each height, perform 3D spanwise stacking of the 2D blade profile.

[0022] Preferably, the distance l1 between the leading edge of the boundary layer distortion correction blade and the rotor-static interface is 3 mm; in order to ensure that the trailing wake of the boundary layer distortion correction blade 4 is mixed before reaching the leading edge of the downstream stator blade 6, the distance l2 between the trailing edge of the boundary layer distortion correction blade and the leading edge of the downstream stator blade is 1Cx, where Cx is the axial chord length of the boundary layer distortion correction blade.

[0023] Preferably, the rotation direction of the upstream rotor is from the pressure side to the suction side.

[0024] Preferably, the flow direction of the end zone fluid after correction by the boundary layer distortion correction blades changes as follows: the velocity direction of the upstream rotor inlet is c1', and affected by the rotating end wall, the velocity direction of the fluid in the proximal zone of the upstream rotor outlet is along an oblique upward direction, and enters the boundary layer distortion correction blade channel at a speed w2' with a near-zero angle of attack, and finally leaves the boundary layer distortion correction blade channel at a speed direction of c3', and enters the downstream stator 5 blade channel at a small negative angle of attack at this speed direction. At this time, the velocity direction is basically consistent with the velocity direction of the mainstream incoming flow.

[0025] Preferably, the hub blade inlet structural angle is 35°, the rim blade inlet structural angle is 5°, and the rim blade outlet structural angle and the hub blade outlet structural angle are both -60°.

[0026] Preferably, the cross-sectional radius R of the reference blade profile is 200 mm, the pitch t is 66.8 mm, the number of blades n is calculated to be 18, the hub radius is 150 mm through the number of blades, and the pitch of the hub blade profile is 52.36 mm.

[0027] Preferably, the hub blade profile is reduced by 0.7838 times according to the reference blade profile size.

[0028] Preferably, the stacking schemes include the following three: leading edge alignment, center alignment and trailing edge alignment.

[0029] Preferably, three-dimensional stacking is performed according to the trailing edge alignment, and all blade section models are stacked along the axial direction of the blade row; in this process, it is ensured that the trailing edge of each section is aligned with the reference plane to form a continuous three-dimensional blade row structure.

[0030] Preferably, h is the stacking height, which is determined by the boundary layer thickness in the end area, that is, the boundary layer distortion height range. The boundary layer thickness is 0.05H, wherein H is the rotor blade height, which is 100 mm. Therefore, the stacking height h is 5 mm.

[0031] The advantages of the present invention compared with the prior art are:

[0032] (1) The present invention proposes a flow organization method for the boundary layer of the end area of ​​a high-load turbine, that is, an airflow correction blade is used near the proximal area of ​​the rotor-static interface to eliminate the negative angle of attack caused by the distortion of the end wall boundary layer, so that the velocity direction of the fluid in the proximal area entering the downstream blade channel is basically consistent with the mainstream velocity direction, thereby achieving the control of the distortion of the boundary layer of the end wall of the high-load turbine, optimizing the flow field structure of the inlet end area, reducing the secondary flow loss, and improving the aerodynamic performance of the high-load turbine.

[0033] (2) Due to the effect of fluid viscosity, the inlet flow forms a typical velocity gradient along the span direction, resulting in different degrees of distortion of the low-energy fluid in the proximal region along the span direction. In view of this phenomenon, based on the end region flow organization method, the present invention further designs an end wall boundary layer distortion correction blade. On the one hand, the inlet structural angle of the blade gradually decreases along the span direction, so that the velocity direction of low-energy fluids with different degrees of distortion entering the blade channel is close to the optimal angle of attack; on the other hand, the outlet structural angle of the blade remains consistent along the span direction to ensure that the velocity direction of the low-energy fluid in the end region entering the downstream blade channel is consistent with the mainstream velocity direction.

[0034] (3) The present invention adopts a trailing edge alignment stacking scheme, which significantly reduces potential airflow disturbances in the blade channel flow field by ensuring that the trailing edges of all blade sections are precisely aligned in the same vertical plane. This precise alignment not only helps to maintain the continuity and uniformity of the flow field, but also reduces the flow separation and secondary flow losses caused by the misalignment of the trailing edges of the blades; it can more effectively control the flow of the end zone boundary layer and reduce performance losses caused by boundary layer distortion. Therefore, the control method proposed in the present invention not only improves the aerodynamic performance of the turbine, but also helps to improve the performance of the entire engine, including increasing thrust, reducing fuel consumption and extending the service life of the engine; reducing stress concentration and potential fatigue damage problems caused by structural inconsistencies. This consistency is crucial to ensuring the stability and reliability of the turbine in long-term operation.

[0035] (4) Through the above design and optimization, the present invention not only solves the problem of boundary layer distortion in the end zone, but also improves the overall performance of the turbine, including reducing flow losses and improving aerodynamic efficiency. These advantages are of great significance for the refined design of the low-pressure turbine of a high-performance high-bypass ratio turbofan engine, and help improve the overall performance and efficiency of the engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0036] Figure 1 It is a schematic diagram of velocity triangle analysis under rotation-stationary interference;

[0037] Figure 2 A flow chart of a high-load turbine end zone boundary layer flow organization method of the present invention;

[0038] Figure 3This is a schematic diagram of the boundary layer distortion correction principle of the present invention;

[0039] Figure 4 It is a schematic diagram of the boundary layer distortion correction result of the present invention;

[0040] Figure 5 The boundary layer distortion correction blade design diagram of the present invention;

[0041] Figure 6 It is a schematic diagram of the boundary layer distortion correction blade cascade of the present invention.

[0042] Reference numerals:

[0043] Upstream rotor 1, upstream rotor blades 2, rotor-stator interface 3, boundary layer distortion correction blades 4, downstream stator 5, downstream stator blades 6. DETAILED DESCRIPTION

[0044] In order to make the purpose, technical scheme and advantages of the implementation of the present invention clearer, the technical scheme in the embodiment of the present invention will be described in more detail below in conjunction with the drawings in the embodiment of the present invention. In the drawings, the same or similar reference numerals throughout represent the same or similar elements or elements with the same or similar functions. The described embodiments are part of the embodiments of the present invention, not all of the embodiments. The embodiments described below with reference to the drawings are exemplary and are intended to be used to explain the present invention, and should not be construed as limitations on the present invention. Based on the embodiments in the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention. The structure and technical scheme of the present invention are further described in detail below in conjunction with the drawings, and an embodiment of the present invention is given.

[0045] A flow organization method for a high-load turbine end region boundary layer of the present invention comprises the following steps:

[0046] SS1, determine the setting position of the boundary layer distortion correction blade:

[0047] A boundary layer distortion correction blade 4 is arranged between the upstream rotor blade 2 and the downstream stator blade 6 to determine the distance between the leading edge of the boundary layer distortion correction blade 4 and the rotor-stator interface 3 and the distance between the trailing edge of the boundary layer distortion correction blade 4 and the leading edge of the downstream stator blade 6.

[0048] SS2, determine the twist angle of the boundary layer twist correction blade:

[0049] The rim blade inlet structural angle α1, the hub blade inlet structural angle α2, the rim blade outlet structural angle, and the hub blade outlet structural angle are defined, wherein the rim and hub inlet structural angles are determined by the actual outlet angle of the low-energy fluid in the upstream rotor end area, and according to the end wall boundary layer distortion law, the hub blade inlet structural angle α2 gradually decreases with the increase of the spanwise height, and the rim blade inlet structural angle α1 is smaller than the hub blade inlet structural angle α2; the rim and hub outlet structural angles are determined by the angle at which the end area fluid enters the downstream stator blade channel, and in order to ensure that the outlet airflow angle of the fluid in the trailing edge end area of ​​the boundary layer distortion correction blade is consistent, the rim blade outlet structural angle and the hub blade outlet structural angle are the same.

[0050] SS3, 3D Cascade Design for Boundary Layer Distortion Correction Blades:

[0051] SS31, determine the actual blade size: follow the above steps to complete the design of the rim blade and hub blade, and enlarge or reduce the blade size proportionally according to the actual turbine size.

[0052] SS32, determine the actual size of blades with different cross-sections: Since the real aircraft engine is an annular structure, the flow area of ​​the hub and the rim are different, and the pitch t is also different. In order to ensure that the boundary layer distortion correction blade has uniform load distribution along the span direction, the hub blade pitch and the rim blade pitch need to be determined according to the hub radius and the rim radius. At the same time, the pitch at the corresponding position is determined according to the different cross-sectional radius of the blade. The calculation formula is as follows:

[0053]

[0054] Where R is the radius of the blade section (including the rim blade section and the hub blade section), and n is the number of blades, whose value is determined based on the radius of the reference blade section and the pitch.

[0055] Then, according to the ratio of the pitches between different spanwise heights, the scaling ratio of the blade profile is determined, and the scaling calculation formula for the actual height section blade profile size is as follows:

[0056]

[0057] Where t1 is the reference blade pitch, t2 is the calculated blade pitch, x1 is the reference blade size, and x2 is the calculated blade size.

[0058] SS33, 3D stacking: After determining the position and size of the blade sections at each height, perform 3D spanwise stacking of the 2D blade profile.

[0059] Further, when determining the setting position of the boundary layer distortion correction blade in step SS1: the principle of the boundary layer flow organization method in the high-load turbine end area is as follows: Figure 2As shown, in order to control the distortion of the boundary layer in the end area, a boundary layer distortion correction blade 4 is arranged between the upstream rotor blade 2 and the downstream stator blade 6. The distance between the leading edge of the boundary layer distortion correction blade 4 and the rotor-static interface 3 is l1. In order to ensure the distortion control effect, l1 is 3 mm in the present invention. The distance between the trailing edge of the boundary layer distortion correction blade 4 and the leading edge of the downstream stator blade 6 is l2. In order to ensure that the trailing wake of the boundary layer distortion correction blade 4 is mixed before reaching the leading edge of the downstream stator blade 6, l2 is 1Cx in the present invention, where Cx is the axial chord length of the boundary layer distortion correction blade. The rotation direction of the upstream rotor 1 is from the pressure surface to the suction surface.

[0060] from Figure 3 It can be seen that the flow direction of the end zone fluid changes after correction by the boundary layer distortion correction blades. The velocity direction of the upstream rotor inlet is c1'. Affected by the rotating end wall, the velocity direction of the fluid in the proximal zone of the upstream rotor outlet is along the oblique upward direction, and enters the boundary layer distortion correction blade channel at a speed w2' with a near-zero angle of attack, and finally leaves the boundary layer distortion correction blade channel at a speed direction of c3', and enters the downstream stator 5 blade channel at a small negative angle of attack at this speed direction. At this time, the velocity direction is basically consistent with the velocity direction of the mainstream flow.

[0061] The boundary layer distortion correction results are as follows Figure 4 As shown in the figure, from the perspective of the trailing edge of the blade, the upstream end wall rotates clockwise. Affected by the rotation effect, the low-energy fluid in the boundary layer of the rear end wall of the cascade has a large deviation from the velocity direction of the mainstream, and the velocity profile along the span direction evolves into a two-dimensional twisted tilt. Subsequently, the distorted end zone fluid enters the boundary layer distortion correction cascade channel, and recovers to a gradient-changing one-dimensional linear boundary layer near the trailing edge exit of the boundary layer distortion correction blade, and enters the downstream stator cascade channel with this velocity distribution.

[0062] When determining the twist angle of the boundary layer twist correction blade in the further step SS2: Due to the effect of fluid viscosity, the inlet flow forms a typical velocity gradient along the span direction, resulting in different degrees of twisting of the low-energy fluid in the proximal region along the span direction, that is, the fluid in the end region along the span direction has different inlet angles, and the inlet angle gradually decreases with the increase of the span height, such as Figure 3 In order to allow the proximal fluid to enter the boundary layer correction blade channel at a more appropriate angle of attack, the present invention designs a boundary layer distortion correction blade, and its blade design is as follows Figure 5As shown, the inlet structural angle refers to the angle formed by the fluid entering the blade and the leading edge of the blade, while the outlet structural angle refers to the angle formed by the fluid leaving the blade and the trailing edge of the blade. α1 is the inlet structural angle of the rim blade, and α2 is the inlet structural angle of the hub blade. Since the end zone fluid inlet angle gradually decreases with the increase of spanwise height, the blade inlet structural angle should also gradually decrease with the increase of spanwise height. Therefore, the inlet structural angle of the rim blade is smaller than the inlet structural angle of the hub blade. In the present invention, the inlet structural angle of the hub blade is 35°, the inlet structural angle of the rim blade is 5°, and the outlet structural angle is -60°.

[0063] At the same time, the present invention achieves the purpose of changing the inlet structural angle while keeping the outlet structural angle unchanged by controlling the bending angle of the blade. The present invention achieves the purpose of gradually reducing the inlet structural angle with the increase of the spanwise height by controlling the bending angle of the blade, that is, the inlet structural angle of the hub blade is greater than the inlet structural angle of the rim blade. Such a design makes the velocity direction of low-energy fluids with different degrees of distortion closer to the optimal angle of attack when entering the blade channel, which helps to improve the aerodynamic performance of the fluid. At the same time, keeping the outlet structural angle unchanged along the spanwise height ensures that the velocity direction of the low-energy fluid in the end area is consistent with the mainstream velocity direction when leaving the blade channel and entering the downstream blade, thereby optimizing the structure of the entire flow field and reducing flow losses.

[0064] In the further step SS3, the three-dimensional blade cascade design of the boundary layer distortion correction blade is determined: the specific design steps of the boundary layer distortion correction blade are as follows, and the three-dimensional blade cascade design result is as follows: Figure 6 As shown:

[0065] 1) Boundary layer distortion correction blade profile design: Complete the design of the rim blade profile and hub blade profile according to the above steps, and enlarge or reduce the designed blade profile size proportionally according to the actual turbine size. In addition, the following design procedures should be followed during the blade profile design process:

[0066] First, analyze the flow characteristics of the upstream rotor end area: It is necessary to conduct an in-depth analysis of the flow characteristics of the upstream rotor end area, including the velocity, pressure distribution and streamline shape of the fluid. Obtain detailed flow data of the low-energy fluid in the rotor end area through computational fluid dynamics (CFD) simulation or experimental measurement. Secondly, determine the outlet angle: The outlet angle refers to the angle between the direction of the fluid flowing out of the rotor blade and the axial direction of the rotor blade. It is a key parameter affecting the subsequent boundary layer flow. According to the analysis results, determine the actual outlet angle when the low-energy fluid is discharged from the rotor blade, and adjust the inlet structural angle of the corresponding blade section accordingly to ensure that the fluid can smoothly enter the boundary layer distortion correction blade. Finally, design the rim blade and hub blade: According to the determined outlet angle, design the rim blade and hub blade to adapt to the flow characteristics of the end area. After completing the preliminary design, evaluate the performance of the blade through CFD simulation or other aerodynamic analysis tools. According to the evaluation results, make necessary adjustments to the blade to optimize its aerodynamic performance. By designing the rim blade profile and hub blade profile adapted to the flow characteristics of the end area, the boundary layer flow is effectively controlled, the secondary flow loss is reduced, and the overall aerodynamic efficiency of the turbine is improved.

[0067] 2) Determine the actual size of blades with different cross-sections: Since the real aircraft engine is an annular structure, the flow area of ​​the hub and the rim are different, and the pitch t is also different. In order to ensure that the height load of the boundary layer distortion correction blade is evenly distributed along the span direction, the hub blade pitch and the rim blade pitch need to be determined according to the hub radius and the rim radius. At the same time, the pitch of the corresponding position is determined according to the different cross-sectional radius of the blade. The calculation formula is as follows:

[0068]

[0069] Where R is the radius of the blade section (including the rim blade section and the hub blade section), and n is the number of blades, whose value is determined based on the radius of the reference blade section and the pitch.

[0070] In the present invention, the cross-sectional radius R of the reference blade profile is 200 mm, the pitch t is 66.8 mm, and the number of blades n is calculated to be 18. The hub radius is 150 mm from the number of blades, so the pitch of the hub blade profile is 52.36 mm.

[0071] Then, according to the ratio of the pitches between different spanwise heights, the scaling ratio of the blade profile is determined, and the scaling calculation formula for the actual height section blade profile size is as follows:

[0072]

[0073] Where t1 is the reference blade pitch, t2 is the calculated blade pitch, x1 is the reference blade size, and x2 is the calculated blade size.

[0074] In the present invention, the hub airfoil should be reduced by 0.7838 times according to the reference airfoil size.

[0075] 3) Three-dimensional stacking: After determining the position and size of the blade section at each height, perform three-dimensional spanwise stacking on the two-dimensional blade profile.

[0076] There are three types of stacking schemes: leading edge alignment, center alignment, and trailing edge alignment. In the leading edge alignment scheme, the leading edges of blade sections at all heights are aligned on the same straight line, which ensures the high consistency of the leading edge of the blade and helps maintain the continuity of the flow field. In the center alignment scheme, the center points of blade sections at all heights are aligned on the same straight line, which ensures the symmetry of the center of the blade and helps balance the force on the blade. In the trailing edge alignment scheme, the trailing edges of blade sections at all heights are aligned on the same straight line, which ensures the high consistency of the trailing edge of the blade and helps reduce the impact of the wake.

[0077] The boundary layer distortion correction blade designed in the present invention has the characteristics of different leading edge structural angles and the same trailing edge structural angles along the span direction. Therefore, the trailing edge alignment stacking scheme is selected for three-dimensional stacking, and all blade section models are stacked along the axial direction of the blade cascade. In this process, it is ensured that the trailing edge of each section is aligned with the reference plane to form a continuous three-dimensional blade cascade structure. The stacking result is shown in Figure 6 As shown, h is the stacking height, which is determined by the boundary layer thickness in the end zone, that is, the boundary layer distortion height range. In the present invention, the boundary layer thickness is 0.05H, wherein H is the rotor blade height, which is 100 mm in the present invention. Therefore, the stacking height h is 5 mm.

[0078] The advantages of the present invention compared with the prior art are:

[0079] (1) The present invention proposes a flow organization method for the boundary layer of the end area of ​​a high-load turbine, that is, an airflow correction blade is used near the proximal area of ​​the rotor-static interface to eliminate the negative angle of attack caused by the distortion of the end wall boundary layer, so that the velocity direction of the fluid in the proximal area entering the downstream blade channel is basically consistent with the mainstream velocity direction, thereby achieving the control of the distortion of the boundary layer of the end wall of the high-load turbine, optimizing the flow field structure of the inlet end area, reducing the secondary flow loss, and improving the aerodynamic performance of the high-load turbine.

[0080] (2) Due to the effect of fluid viscosity, the inlet flow forms a typical velocity gradient along the span direction, resulting in different degrees of distortion of the low-energy fluid in the proximal region along the span direction. In view of this phenomenon, based on the end region flow organization method, the present invention further designs an end wall boundary layer distortion correction blade. On the one hand, the inlet structural angle of the blade gradually decreases along the span direction, so that the velocity direction of low-energy fluids with different degrees of distortion entering the blade channel is close to the optimal angle of attack; on the other hand, the outlet structural angle of the blade remains consistent along the span direction to ensure that the velocity direction of the low-energy fluid in the end region entering the downstream blade channel is consistent with the mainstream velocity direction.

[0081] (3) The present invention adopts a trailing edge alignment stacking scheme, which significantly reduces potential airflow disturbances in the blade channel flow field by ensuring that the trailing edges of all blade sections are precisely aligned in the same vertical plane. This precise alignment not only helps to maintain the continuity and uniformity of the flow field, but also reduces the flow separation and secondary flow losses caused by the misalignment of the trailing edges of the blades; it can more effectively control the flow of the end zone boundary layer and reduce performance losses caused by boundary layer distortion. Therefore, the control method proposed in the present invention not only improves the aerodynamic performance of the turbine, but also helps to improve the performance of the entire engine, including increasing thrust, reducing fuel consumption and extending the service life of the engine; reducing stress concentration and potential fatigue damage problems caused by structural inconsistencies. This consistency is crucial to ensuring the stability and reliability of the turbine in long-term operation.

[0082] (4) Through the above design and optimization, the present invention not only solves the problem of boundary layer distortion in the end zone, but also improves the overall performance of the turbine, including reducing flow losses and improving aerodynamic efficiency. These advantages are of great significance for the refined design of the low-pressure turbine of a high-performance high-bypass ratio turbofan engine, and help improve the overall performance and efficiency of the engine.

[0083] Through the above embodiments, the purpose of the present invention is fully and effectively achieved. Those skilled in the art can understand that the present invention includes but is not limited to the contents described in the drawings and the above specific embodiments. Although the present invention has been described with respect to the most practical and preferred embodiments currently considered, it should be understood that the present invention is not limited to the disclosed embodiments, and any modification that does not deviate from the functional and structural principles of the present invention will be included in the scope of the claims.

Claims

1. A flow organization method for a boundary layer at a high-load turbine end region, characterized in that: The following steps are involved: SS1, determine the setting position of the boundary layer distortion correction blade: A boundary layer distortion correction blade (4) is arranged between an upstream rotor blade (2) and a downstream stator blade (6), and a distance between a leading edge of the boundary layer distortion correction blade (4) and a rotor-stator interface (3) and a distance between a trailing edge of the boundary layer distortion correction blade (4) and a leading edge of a downstream stator blade (6) are determined; SS2, determine the twist angle of the boundary layer twist correction blade: The rim blade inlet structural angle α1, the hub blade inlet structural angle α2, the rim blade outlet structural angle, and the hub blade outlet structural angle are defined, wherein the rim and hub inlet structural angles are determined by the actual outlet angle of the low-energy fluid in the upstream rotor end area, and according to the end wall boundary layer distortion law, the hub blade inlet structural angle α2 gradually decreases with the increase of the spanwise height, and the rim blade inlet structural angle α1 is smaller than the hub blade inlet structural angle α2; the rim and hub outlet structural angles are determined by the angle at which the end area fluid enters the downstream stator blade channel, and in order to ensure that the outlet airflow angle of the trailing edge end area fluid of the boundary layer distortion correction blade (4) is consistent, the rim blade outlet structural angle and the hub blade outlet structural angle are the same; SS3, 3D Cascade Design for Boundary Layer Distortion Correction Blades: SS31, determine the actual blade size: follow the above steps to complete the design of the rim blade and hub blade, and enlarge or reduce the blade size proportionally according to the actual turbine size; SS32, determine the actual size of blades with different cross-sections: Since the real aircraft engine is an annular structure, the flow area of ​​the hub and the rim are different, and the grid pitch t is also different. In order to ensure that the boundary layer distortion correction blade (4) has a uniform load distribution along the span direction, the hub blade pitch and the rim blade pitch need to be determined according to the hub radius and the rim radius. At the same time, the grid pitch at the corresponding position is determined according to the different cross-sectional radius of the blade. The calculation formula is as follows: Where R is the radius of the blade section, including the rim blade section and the hub blade section, and n is the number of blades, whose value is determined according to the radius of the reference blade section and the pitch; Then, according to the ratio of the pitches between different spanwise heights, the scaling ratio of the blade profile is determined, and the scaling calculation formula for the actual height section blade profile size is as follows: Where, t1 is the reference blade pitch, t2 is the calculated blade pitch, x1 is the reference blade size, and x2 is the calculated blade size; SS33, 3D stacking: After determining the position and size of the blade sections at each height, perform 3D spanwise stacking of the 2D blade profile.

2. A flow organization method for a high-load turbine end zone boundary layer according to claim 1, characterized in that: The distance l1 between the leading edge of the boundary layer distortion correction blade and the rotor-static interface is 3 mm; in order to ensure that the trailing wake of the boundary layer distortion correction blade 4 is mixed before reaching the leading edge of the downstream stator blade 6, the distance l2 between the trailing edge of the boundary layer distortion correction blade and the leading edge of the downstream stator blade is 1Cx, where Cx is the axial chord length of the boundary layer distortion correction blade.

3. A flow organization method for a high-load turbine end zone boundary layer according to claim 2, characterized in that: The rotation direction of the upstream rotor (1) is from the pressure side to the suction side.

4. The flow organization method of the boundary layer of the high-load turbine end region according to claim 1 is characterized in that: After the boundary layer distortion correction blade is corrected, the flow direction of the end zone fluid changes as follows: the velocity direction of the upstream rotor inlet is c1'. Affected by the rotating end wall, the velocity direction of the fluid in the proximal zone of the upstream rotor outlet is along the upward direction, and enters the boundary layer distortion correction blade channel at a speed w2' with a near-zero angle of attack, and finally leaves the boundary layer distortion correction blade channel at a speed direction of c3', and enters the downstream stator (5) blade channel at a small negative angle of attack at this speed direction. At this time, the velocity direction is basically consistent with the velocity direction of the mainstream inflow.

5. The flow organization method of the boundary layer of the high-load turbine end region according to claim 1 is characterized in that: The hub blade inlet structural angle is 35°, the rim blade inlet structural angle is 5°, and the rim blade outlet structural angle and the hub blade outlet structural angle are both -60°.

6. The flow organization method of the boundary layer of the high-load turbine end region according to claim 1 is characterized in that: The cross-sectional radius R of the reference blade profile is 200 mm, the pitch t is 66.8 mm, and the number of blades n is calculated to be 18. The hub radius is 150 mm through the number of blades, and the pitch of the hub blade profile is 52.36 mm.

7. The flow organization method of the boundary layer of the high-load turbine end region according to claim 1 is characterized in that: The hub blade profile is reduced by 0.7838 times according to the reference blade size.

8. The flow organization method of the boundary layer of the high-load turbine end region according to claim 1 is characterized in that: There are three stacking schemes: leading edge alignment, center alignment, and trailing edge alignment.

9. A flow organization method for a boundary layer at a high-load turbine end region according to claim 8, characterized in that: Perform three-dimensional stacking according to the trailing edge alignment, and stack all blade section models along the axial direction of the blade cascade; In this process, it is ensured that the trailing edge of each section is aligned with the reference plane to form a continuous three-dimensional cascade structure.

10. A flow organization method for a boundary layer at a high-load turbine end region according to claim 9, characterized in that: h is the stacking height, which is determined by the boundary layer thickness in the end area, that is, the boundary layer distortion height range. The boundary layer thickness is 0.05H, where H is the rotor blade height, which is 100mm. Therefore, the stacking height h is 5mm.

Images