Rotor blade of a segmented stator axial compressor and method for designing the same

Abstract

This invention discloses a rotor blade of a segmented stacked axial compressor and its design method. The design method includes: determining the axial position of the stacking line; determining the circumferential position of the stacking line: the stacking line is S-shaped in the circumferential direction, and the stacking line includes four control points A, B, C, and D sequentially from the blade root to the blade tip along the height direction; the coordinates of the centroid positions of each basic airfoil section in the first interval from point B to point C conform to a specific formula; and determining the shape of the stacking line based on the determined axial and circumferential positions. This invention, by setting the stacking line of the rotor blade to be S-shaped in the circumferential direction, generates a reverse pressure gradient in the radial direction, causing the gas at both ends to migrate towards the blade center, thereby improving the flow in the end region. This reduces the influence of secondary flow in the blade tip region; and by setting a specific function, the circumferential offset of the stacking line can be controlled by constructing a simple equation, improving the computational efficiency of optimizing the airfoil.

Description

Technical Field

[0001] This invention relates to a rotor blade of a segmented stacked axial flow compressor and its design method. Background Technology

[0002] With the development of technology, modern aero engines face increasingly higher requirements for economy and reliability, posing unprecedented challenges to the performance of various components. The axial compressor is one of the core components of an aero engine, and its performance has a crucial impact on the entire engine. However, as aero engine performance indicators continue to improve, the internal temperature and pressure ratio of the engine gradually increase, leading to increasingly serious losses in the compressor's internal flow system. The severe internal flow losses in multi-stage, high-pressure-ratio axial compressors significantly affect their aerodynamic performance, and consequently, the overall engine performance.

[0003] The main sources of compressor losses are airfoil losses, endwall losses, and leakage losses. The flow in the endwall region of a high-load compressor is very complex, and endwall boundary layer losses and tip leakage losses are the main components of compressor rotor blade losses.

[0004] The compressor is the rotating component inside an aero-engine that experiences the greatest adverse pressure gradient, and achieving efficient airflow pressurization is a major challenge in compressor design. In the corner region formed by the rotor hub and the suction surfaces at the blade roots, separation easily occurs due to the influence of the hub boundary layer and gas undercurrent, resulting in aerodynamic losses and a decrease in compressor efficiency. In the blade tip region, a gap exists between the rotor and the outer casing. Airflow from the rotor blade pressure surface leaks through this gap to the blade suction surface, generating complex secondary flow characteristics such as leakage vortices. Furthermore, the high tangential velocity of the rotor blades generates shock waves at the tips of the compressor's leading-stage rotor blades. These shock waves, leakage vortices, and the outer casing boundary layer interfere with each other, making the flow in the blade tip gap region extremely complex, exhibiting various complex vortex structures. Numerous studies have shown that compressor stall typically occurs in the blade tip and hub regions.

[0005] Therefore, how to effectively control the shock wave and flow field in the blade tip region, thereby reducing the secondary flow in the blade tip region, is a problem that needs to be solved in this field.

[0006] On the other hand, compressor design typically requires multiple rounds of optimization iterations to obtain the ideal compressor aerodynamic blade profile. To save optimization time and avoid over-reliance on designers' experience, compressor aerodynamic optimization methods have gained popularity among designers in recent years. Compressor design optimization methods are interdisciplinary design approaches that combine computational fluid dynamics (CFD) with numerical optimization methods. The basic idea is to use aerodynamic performance as the objective function value, and with the help of CFD three-dimensional numerical simulation tools, to iteratively find the optimal compressor aerodynamic blade profile using optimization algorithms. The parameterized selection method for blade profile design is one of the most critical technologies in compressor aerodynamic optimization. How to extract the key parameters of the compressor's circumferential and radial stacking forms and describe them with appropriate mathematical equations, while taking into account both the characteristics of the compressor blade profile design and the feasibility of the calculation method, is a challenge in the optimization design of compressor blade profile stacking forms. Summary of the Invention

[0007] The technical problem to be solved by the present invention is to overcome the defect of airflow loss in the rotor blades of the segmented stacked axial compressor in the prior art, and at the same time improve the efficiency of blade profile optimization design, so as to provide a rotor blade of the segmented stacked axial compressor and its design method.

[0008] The present invention solves the above-mentioned technical problems through the following technical solution:

[0009] A method for designing rotor blades of a segmented stacked axial compressor, characterized in that the rotor blade includes a blade root and a blade tip, and the rotor blade has a plurality of basic airfoil sections along the height direction from the blade root to the blade tip. The centroids of each basic airfoil section are used as stacking points and superimposed along the height direction of the rotor blade to obtain a stacking line. The design method includes the following steps:

[0010] S1: Determine the axial position of the stacking line;

[0011] S2: Determine the circumferential position of the stacking line: The stacking line is "S"-shaped in the circumferential direction. The stacking line includes four control points A, B, C, and D along the height direction from the leaf root to the leaf tip. Point A is the centroid of the basic airfoil section of the leaf root and is the origin. Point B is located at the position of the maximum circumferential offset of the stacking line in the positive X-axis direction and is a preset point. Point C is located at the position of the maximum circumferential offset of the stacking line in the negative X-axis direction and is a preset point. Point D is the centroid of the basic airfoil section of the leaf tip and is located between points B and C on the X-axis.

[0012] The coordinates (x1, y1) of the centroid positions of each of the basic airfoil sections in the first interval from point B to point C conform to the following formula:

[0013]

[0014] Where, N = (x b -x c ) / 2; α=Pi / (y c –y b Pi = 3.1415926; parameters And M based on the coordinates (x) of point B b y b ) and the coordinates (x) of point C c y c Substituting into the formula yields the result;

[0015] S3: Determine the profile of the stacking line based on the axial and circumferential positions of the centroids of each basic blade section on the stacking line; extend the basic blades along the profile of the stacking line to construct a set of curves to obtain a three-dimensional rotor blade.

[0016] In this technical solution, by setting the stacking line of the rotor blades to an "S" shape in the circumferential direction, a reverse pressure gradient can be generated in the radial direction. This facilitates the migration of low-energy fluid from the blade tip region to the middle of the blade, thereby reducing the influence of secondary flow in the blade tip region, improving the flow conditions at the blade tip, and extending the stable operating range of the compressor rotor blades. Furthermore, the optimal maximum circumferential offset x is determined based on the principle of minimizing losses or maximizing efficiency. b and x c The corresponding height and extension length (defined as positive offset towards the blade head and negative offset towards the blade back) are determined, and the coordinates of the centroid positions of each basic blade section in the first interval conform to a specific function. This allows for the control of the circumferential offset of the centroid positions of each basic blade section in the first interval through a simple equation, improving the computational efficiency of blade optimization. In other words, this significantly improves computational efficiency compared to the existing method of constructing the stacking line through trial and error. It should be noted that the stacking line of the rotor blade is "S"-shaped in the circumferential direction, including cases where the stacking line is approximately "S"-shaped in the circumferential direction.

[0017] Preferably, step S2 includes:

[0018] Extend the first interval downwards at point B along the height direction by a first extension height Δy. b Thus, the first intermediate point E is obtained;

[0019] The coordinates (x2, y2) of the circumferential positions of the centroids of the basic airfoil sections in the second interval from point A to point E conform to the following formula:

[0020]

[0021] Wherein, the parameters h, i, j pass through the coordinates (x, y) of point E. e y e The slope k at point E e and the coordinates of point A (x a y a Substituting into the formula yields the result.

[0022] In this technical solution, the first interval is extended downwards at point B by a first extension height Δy. b This allows the accumulation line to maintain continuous curvature at the turning position, thus achieving a smooth transition. This is beneficial for maintaining high compressor efficiency while keeping the rotor blades at a better strength level. Furthermore, by setting the coordinates of the centroid positions of each basic airfoil section in the second interval to conform to a specific function, the circumferential offset of the centroid positions of each basic airfoil section in the second interval can be controlled by constructing simple equations, thereby improving the calculation efficiency of airfoil optimization.

[0023] Preferably, the first extension height Δy b It is 5% to 20% of the height of the first interval.

[0024] In this technical solution, a first extension height Δy is set. b The range is determined to effectively ensure the continuity of curvature and generate an effective reverse pressure gradient, that is, to avoid the height being too small to ensure the continuity of curvature, while being too large to affect the generation of an effective reverse pressure gradient in the radial direction.

[0025] Preferably, step S2 includes:

[0026] Extend the first interval upwards at point C along the height direction by a second extension height Δy. c Thus, the midpoint F is obtained;

[0027] The coordinates (x3, y3) of the centroid positions of each of the basic airfoil sections in the third interval from point F to point D conform to the following formula:

[0028] x3=k f y3+Q

[0029] Where, k f Let Q be the slope at point F, and let the parameter Q pass through the coordinates (x, y) of point F. f y f The slope k at point F f Substituting into the formula yields the result.

[0030] In this technical solution, the first interval is extended upwards at point C by a first extension height Δy. cThis allows the accumulation line to maintain continuous curvature at the turning position, achieving a smooth transition. This helps maintain high compressor efficiency while ensuring optimal rotor blade strength. Furthermore, by setting the coordinates of the centroid positions of each basic airfoil section in the third interval to conform to a specific function, the circumferential offset of the centroid positions of each basic airfoil section in the second interval can be controlled by constructing simple equations, improving the computational efficiency of airfoil optimization. In addition, by extending the height of point C upwards and combining it with a linear function, the airfoil can form a "wind-blowing" wall effect at the blade tip, preventing high-pressure gas at the blade tip and blade base from migrating to the blade back and reducing the tip secondary flow effect.

[0031] Preferably, the second extension height Δy c It is 5% to 20% of the height of the first interval.

[0032] In this technical solution, a second extension height Δy is set. c The range is determined to effectively ensure the continuity of curvature and generate an effective reverse pressure gradient, that is, to avoid the height being too small to ensure the continuity of curvature, while being too large to affect the generation of an effective reverse pressure gradient in the radial direction.

[0033] Preferably, in step S2, the coordinates (x, y) of point B are... b y b ), y b It is 25% to 35% of the leaf height; and / or,

[0034] In step S2, the coordinates of point C (x) c y c ), y c It accounts for 70% to 85% of the leaf height.

[0035] In this technical solution, the coordinates (x, y) of point B are set. b y b ) in y b The range of values ​​is beneficial for maintaining high compressor efficiency while ensuring that the rotor blades maintain a good strength level.

[0036] By setting the coordinates (x) of point C c y c ) in y c The range of values ​​is beneficial for maintaining high compressor efficiency while ensuring that the rotor blades maintain a good strength level.

[0037] Preferably, in step S1, determining the axial position of the stacking line includes:

[0038] In the axial direction, the centroid position of each of the basic airfoil sections is moved along the height direction to a straight line with the centroid position of the basic airfoil section at the blade root as the reference point; the axial offset of the centroid position of the basic airfoil section at the blade tip is set as the maximum axial offset to determine the centroid position of the basic airfoil section at the blade tip; the axial offset of the centroid positions of the other basic airfoil sections is obtained by linear interpolation.

[0039] In this technical solution, the axial offset of the centroid position of each basic airfoil section is obtained through the above method, thereby improving the calculation efficiency of airfoil optimization.

[0040] Preferably, in step S1, the centroid of the basic airfoil section at the blade tip is positioned further forward than the centroid of the basic airfoil section at the blade root in the axial direction.

[0041] In this technical solution, the centroid of the basic airfoil section at the blade tip is forward in the axial direction than the centroid of the basic airfoil section at the blade root. This forms a forward-swept rotor airfoil, which helps to improve the margin of the rotor blades.

[0042] Preferably, after step S3, the method further includes:

[0043] S4: Repeat steps S1 to S3 multiple times to obtain multiple different rotor blades, obtain the compressor efficiencies corresponding to the multiple different rotor blades, and select the rotor blade corresponding to the highest compressor efficiency as the final rotor blade.

[0044] In this technical solution, in step S4, multiple different maximum axial stacking offsets are substituted into step S1, and multiple different maximum circumferential stacking offsets are substituted into step S2. Steps S1 to S3 are repeated multiple times to obtain multiple different rotor blades. That is, multiple different rotor blades are obtained through an iterative method. These rotor blades correspond to different maximum axial stacking offsets and different maximum circumferential stacking offsets.

[0045] Compressor efficiency is calculated for multiple different rotor blades, and the rotor blade corresponding to the highest compressor efficiency is selected as the final rotor blade to obtain the rotor blade with optimal compressor efficiency, that is, the rotor blade with optimal circumferential and axial stacking.

[0046] A rotor blade for a segmented stacked axial compressor, characterized in that the rotor blade is obtained by the design method of the segmented stacked axial compressor rotor blade described above.

[0047] The positive and progressive effects of this invention are as follows:

[0048] This invention, by setting the stacking line of the rotor blades to an S-shape in the circumferential direction, generates a reverse pressure gradient in the radial direction. This facilitates the migration of low-energy fluid from the blade tip region to the middle of the blade, thereby reducing the influence of secondary flow in the blade tip region, improving the flow conditions at the blade tip, and extending the stable operating range of the compressor rotor blades. Furthermore, the optimal maximum circumferential offset x is determined based on the principle of minimizing losses or maximizing efficiency. b and x c The system sets the corresponding height and extension length, and by setting the coordinates of the centroid positions of each basic airfoil section in the first interval to conform to a specific function, it constructs a simple equation to control the circumferential offset of the centroid positions of each basic airfoil section in the first interval, thereby improving the computational efficiency of airfoil optimization. In other words, this significantly improves computational efficiency compared to the existing method of constructing superimposed lines through trial and error.

[0049] On the other hand, existing technologies often use single-equation forms to describe approximate "S"-shaped curves. However, compressor blade stacking optimization typically involves optimizing multiple variables, such as the maximum circumferential offset and its position, and the root tip tilt angle. In these cases, a single-equation form cannot optimize the key characteristic parameters of compressor circumferential stacking. This invention extracts the most critical design parameters of compressor circumferential stacking—the maximum positive and negative circumferential offsets and their height positions, and the root tip tilt angle—and uses these as input variables. A piecewise function is then used to solve for the offsets at other height positions in a closed loop, ensuring smooth transitions at the connection points of the piecewise function and continuity of at least the first derivative. By ensuring that the key parameters of compressor blade circumferential stacking are fully described, a simple piecewise function is constructed, making the equations easier to solve and improving blade optimization efficiency. In other words, this invention provides a parameterization method that uses the maximum positive and negative circumferential offsets of the compressor circumferential stack, their height positions, and the tilt angle of the root tip position as control parameters. The offsets at other height positions are solved in a closed loop using piecewise function form, which ensures smooth transitions at the connection points of the piecewise function and continuity of at least the first derivative. Attached Figure Description

[0050] Figure 1 This is a flowchart illustrating a preferred embodiment of the design method for rotor blades of a segmented stacked axial compressor according to the present invention.

[0051] Figure 2 This is a structural schematic diagram of the blade root, blade tip, and intermediate basic airfoil section and the center of gravity position of the rotor blade of a segmented stacked axial compressor according to a preferred embodiment of the present invention.

[0052] Figure 3This is a schematic diagram of the axial offset of the stacking line of the rotor blades of a segmented stacked axial compressor according to a preferred embodiment of the present invention.

[0053] Figure 4 This is a schematic diagram of the circumferential offset of the stacking line of the rotor blades of a segmented stacked axial compressor according to a preferred embodiment of the present invention.

[0054] Figure 5 This is a flowchart illustrating a preferred embodiment of the design method for rotor blades of a segmented stacked axial compressor according to the present invention.

[0055] Figure 6 This is a schematic diagram of the final airfoil shape of the rotor blades of a segmented stacked axial compressor according to a preferred embodiment of the present invention.

[0056] Figure 7 This is a schematic diagram of the original blade profile of the rotor blades of a segmented stacked axial compressor according to a preferred embodiment of the present invention. Detailed Implementation

[0057] The present invention will be further illustrated by way of embodiments below, but the present invention is not limited to the scope of the embodiments described herein.

[0058] This embodiment provides a design method for rotor blades of a segmented stacked axial flow compressor.

[0059] The rotor blade includes a blade root and a blade tip. The rotor blade has several basic blade profile sections along the height direction from the blade root to the blade tip. The centroids of each basic blade profile section are used as accumulation points and superimposed along the height direction of the rotor blade to obtain an accumulation line.

[0060] like Figure 1 As shown, the design method for the rotor blades of the segmented stacked axial compressor in this embodiment includes the following steps:

[0061] S1: Determine the axial position of the stacking line;

[0062] S2: Determine the circumferential position of the stacking line: The stacking line is "S"-shaped in the circumferential direction. The stacking line includes four control points A, B, C, and D along the height direction from the leaf root to the leaf tip. Point A is the centroid of the basic airfoil section of the leaf root and is the origin. Point B is located at the position of the maximum circumferential offset of the stacking line in the positive X-axis direction and is a preset point. Point C is located at the position of the maximum circumferential offset of the stacking line in the negative X-axis direction and is a preset point. Point D is the centroid of the basic airfoil section of the leaf tip and is located between points B and C on the X-axis.

[0063] The coordinates (x1, y1) of the centroid positions of each of the basic airfoil sections in the first interval from point B to point C conform to the following formula:

[0064]

[0065] Where, N = (x b -x c ) / 2; α=Pi / (y c –y b Pi = 3.1415926; parameters And M based on the coordinates (x) of point B b y b ) and the coordinates (x) of point C c y c Substituting into the formula yields the result;

[0066] S3: Determine the profile of the stacking line based on the axial and circumferential positions of the centroids of each basic blade section on the stacking line; extend the basic blades along the profile of the stacking line to construct a set of curves to obtain a three-dimensional rotor blade.

[0067] In this way, by setting the stacking line of the rotor blades to be S-shaped in the circumferential direction, a reverse pressure gradient can be generated in the radial direction. This facilitates the migration of low-energy fluid from the blade tip region to the middle of the blade, thereby reducing the influence of secondary flow in the blade tip region, improving the flow conditions at the blade tip, and extending the stable operating range of the compressor rotor blades. Furthermore, the optimal maximum circumferential offset x is determined based on the principle of minimizing losses or maximizing efficiency. b and x c The corresponding height and extension length (defined as positive offset towards the blade head and negative offset towards the blade back) are determined, and the coordinates of the centroid positions of each basic blade section in the first interval conform to a specific function. This allows for the control of the circumferential offset of the centroid positions of each basic blade section in the first interval through a simple equation, improving the computational efficiency of blade optimization. In other words, this significantly improves computational efficiency compared to the existing method of constructing the stacking line through trial and error. It should be noted that the stacking line of the rotor blade is "S"-shaped in the circumferential direction, including cases where the stacking line is approximately "S"-shaped in the circumferential direction.

[0068] Specifically, in step S1, determining the axial position of the stacking line includes:

[0069] The centroid positions of each of the basic airfoil sections are moved along the height direction to a straight line with the centroid position of the basic airfoil section at the blade root as the reference point; the axial offset of the centroid position of the basic airfoil section at the blade tip is set as the maximum axial offset to determine the centroid position of the basic airfoil section at the blade tip; the axial offsets of the centroid positions of the other basic airfoil sections are obtained by linear interpolation.

[0070] Specifically, such as Figure 2 As shown in the figure, the basic leaf-shaped section 111 of the leaf root and its centroid position 1111 are shown, the basic leaf-shaped section 121 of the leaf tip and its centroid position 1211 are shown, and the basic leaf-shaped section 131 of the leaf located in the middle position between the leaf root and the leaf tip and its centroid position 1311 are shown.

[0071] like Figure 3 As shown, in step S1, determining the axial position of the stacking line includes: along the axial direction O, moving the centroid position of each of the basic airfoil sections to a straight line along the height direction, with the centroid position 1111 of the basic airfoil section at the airfoil root as the reference point, that is, placing the centroid position of each of the basic airfoil sections at the initial position. Figure 3 (Represented by dashed lines); Set the axial offset of the centroid position 1211 of the basic airfoil section at the blade tip to the maximum axial offset Ax1, so as to determine the centroid position 1311 of the basic airfoil section at the blade tip, that is, to obtain the final position of the centroid position 1311 of the basic airfoil section at the blade tip. Figure 3 (Represented by solid lines in the text). The axial offset Ax1 of the centroid position 1211 of the basic airfoil section at the blade tip is a preset value, that is, the axial offset Ax1 of the centroid position 1211 of the basic airfoil section at the blade tip is preset as needed.

[0072] The axial offset of the centroid position of the remaining basic blade section is obtained by linear interpolation (linear function). That is, the final position of the centroid of the remaining basic blade section is obtained by linear interpolation. In other words, the axial cumulative offset of the entire blade height is determined by solving the circumferential linear function.

[0073] In this way, the axial offset of the centroid position of each basic airfoil section can be obtained through the above method, thereby improving the calculation efficiency of airfoil optimization.

[0074] Please refer to the following: Figure 3In step S1, the centroid position 1211 of the basic airfoil section at the blade tip is forward of the centroid position 1111 of the basic airfoil section at the blade root in the axial direction. That is, the centroid position 1211 of the basic airfoil section at the blade tip is shifted toward the suction surface of the rotor blade to form a forward-swept rotor airfoil, which helps to improve the margin of the rotor blade; and can play a similar "wind-blocking wall" effect, thereby effectively suppressing blade tip leakage.

[0075] Specifically, such as Figure 4 As shown, in step S2, the stacking line 20 is S-shaped in the circumferential direction. The stacking line 20 includes four control points A, B, C, and D sequentially from the leaf root to the leaf tip along the height direction H. Point A is located at the centroid of the basic leaf shape section at the leaf root, and point A is the origin. That is, in the two-dimensional coordinate system, point A is the origin, and the coordinates of point A (x...) are... a y a (0, 0).

[0076] Point B is located at the position of the maximum circumferential offset of the accumulation line in the positive X-axis direction, and point B is a preset point. Point C is located at the position of the maximum circumferential offset of the accumulation line in the negative X-axis direction, and point C is also a preset point. In other words, this parameterization method, combined with CFD calculations, determines the optimal maximum circumferential offset and the corresponding height of that position based on the principle of minimizing loss or maximizing efficiency; that is, the coordinates of point B (x...). b y b ) and the coordinates (x) of point C c y c The determination is based on the principle of minimizing loss or maximizing efficiency. The accumulation line turns at both points B and C.

[0077] Point D is the centroid of the basic blade section at the blade tip, and point D is located between points B and C on the X-axis.

[0078] It should be noted that the S-shaped accumulation line in the circumferential direction refers to the accumulation line being approximately S-shaped in the circumferential direction. Specifically, points A, B, C, and D are sequentially set along the positive Y-axis (i.e., the height direction of the blade) on the accumulation line. At the same time, point B is located at the position of the maximum circumferential offset of the accumulation line in the positive X-axis direction, and point C is located at the position of the maximum circumferential offset of the accumulation line in the negative X-axis direction. The positions of points A and D on the X-axis are both between the positions of points B and C on the X-axis.

[0079] The coordinates (x1, y1) of the centroid positions of each of the basic airfoil sections in the first interval 21 from point B to point C conform to the following formula:

[0080]

[0081] Where, N = (x b -x c ) / 2; α=Pi / (y c –y b Pi = 3.1415926; parameters And M based on the coordinates (x) of point B b y b ) and the coordinates (x) of point C c y c Substituting into the formula yields the result.

[0082] Preferably, step S2 includes:

[0083] Extend the first interval downwards at point B along the height direction by a first extension height Δy. b Thus, the first intermediate point E is obtained;

[0084] The coordinates (x2, y2) of the circumferential position points of the centroids of each of the basic airfoil sections in the second interval 22 from point A to point E conform to the following formula:

[0085]

[0086] Wherein, the parameters h, i, j pass through the coordinates (x, y) of point E. e y e The slope k at point E e and the coordinates of point A (x a y a Substituting into the formula, we get the slope k at point E. e It can be obtained using the following formula:

[0087]

[0088] Wherein, parameters N, α, This is derived from the formula above.

[0089] Thus, by extending the first interval downwards at point B along the height direction by a first extension height Δy b In other words, the coordinates of the circumferential positions of the centroids of each of the basic airfoil sections between points B and E ensure the curvature continuity of the stacking line at the turning position, thus achieving a smooth transition. This is beneficial for maintaining high compressor efficiency while keeping the rotor blades at a good strength level. Furthermore, by setting the coordinates of the centroids of each basic airfoil section in the second interval to conform to a specific function, the circumferential offset of the centroid position of each basic airfoil section in the second interval can be controlled by constructing a simple equation, thereby improving the calculation efficiency of airfoil optimization. Specifically, the first interval is extended downwards from point B by a first extension height Δy. bThus, the first intermediate point E is obtained. In other words, the coordinates of the circumferential position of the centroid of each of the basic airfoil sections between point B and point E also conform to the formula that the coordinates of the centroid of each of the basic airfoil sections in the first interval 21 from point B to point C conform to.

[0090] Wherein, the first extension height Δy b The height is 5% to 20% of the height of the first interval. Thus, by setting the first extension height Δy... b The range is determined to effectively ensure the continuity of curvature and generate an effective reverse pressure gradient, that is, to avoid the height being too small to ensure the continuity of curvature, while being too large to affect the generation of an effective reverse pressure gradient in the radial direction.

[0091] Preferably, step S2 includes:

[0092] Extend the first interval upwards at point C along the height direction by a second extension height Δy. c Thus, the midpoint F is obtained;

[0093] The coordinates (x3, y3) of the centroid positions of each of the basic airfoil sections in the third interval 23 from point F to point D conform to the following formula:

[0094] x3=k f y3+Q

[0095] Where, k f Let Q be the slope at point F, and let the parameter Q pass through the coordinates (x, y) of point F. f y f The slope k at point F f Substituting into the formula, we get the slope k at point F. f It can be obtained using the following formula:

[0096]

[0097] Wherein, parameters N, α, This is derived from the formula above.

[0098] Thus, by extending the first interval upwards at point C along the height direction by a first extension height Δy cThis allows the accumulation line to maintain continuous curvature at the turning position, achieving a smooth transition. This helps maintain high compressor efficiency while ensuring optimal rotor blade strength. Furthermore, by setting the coordinates of the centroid positions of each basic airfoil section in the third interval to conform to a specific function, the circumferential offset of the centroid positions of each basic airfoil section in the second interval can be controlled by constructing simple equations, improving the computational efficiency of airfoil optimization. In addition, by extending the height of point C upwards and combining it with a linear function, the airfoil can form a "wind-blowing" wall effect at the blade tip, preventing high-pressure gas at the blade tip and blade base from migrating to the blade back and reducing the tip secondary flow effect.

[0099] Wherein, the first interval is extended upward along the height direction at point C by a second extension height Δy. c The coordinates of the circumferential position of the centroid of each of the basic airfoil sections between point C and point F are obtained from the intermediate point F. This means that the coordinates of the centroid of each of the basic airfoil sections in the first interval 21 between point B and point C also conform to the formula.

[0100] Wherein, the second extension height Δy c The height is 5% to 20% of the height of the first interval. Thus, by setting a second extension height Δy... c The range is determined to effectively ensure the continuity of curvature and generate an effective reverse pressure gradient, that is, to avoid the height being too small to ensure the continuity of curvature, while being too large to affect the generation of an effective reverse pressure gradient in the radial direction.

[0101] Preferably, in step S2, the coordinates (x, y) of point B are... b y b ), y b It accounts for 25% to 35% of the leaf height.

[0102] Thus, by setting the coordinates (x) of point B b y b ) in y b The range of values ​​is beneficial for maintaining high compressor efficiency while ensuring that the rotor blades maintain a good strength level.

[0103] Preferably, in step S2, the coordinates (x, y) of point C are... c y c ), y c It accounts for 70% to 85% of the leaf height.

[0104] Thus, by setting the coordinates (x) of point C c y c ) in y c The range of values ​​is beneficial for maintaining high compressor efficiency while ensuring that the rotor blades maintain a good strength level.

[0105] In other words, the coordinates of the circumferential positions of the centroids of the basic blade sections from point A to point E, from point E to point F, and from point F to point D are all governed by different formulas. It is necessary to solve the corresponding formulas piecewise to determine the circumferential stacking offset of the entire blade height.

[0106] Following step S3, the process further includes:

[0107] S4: Repeat steps S1 to S3 multiple times to obtain multiple different rotor blades, obtain the compressor efficiencies corresponding to the multiple different rotor blades, and select the rotor blade corresponding to the highest compressor efficiency as the final rotor blade.

[0108] In step S4, multiple different maximum axial stacking offsets are substituted into step S1, and multiple different maximum circumferential stacking offsets are substituted into step S2. Steps S1 to S3 are repeated multiple times to obtain multiple different rotor blades. That is, multiple different rotor blades are obtained through an iterative method. These rotor blades correspond to different maximum axial stacking offsets and different maximum circumferential stacking offsets.

[0109] Compressor efficiency is calculated for multiple different rotor blades, and the rotor blade corresponding to the highest compressor efficiency is selected as the final rotor blade to obtain the rotor blade with optimal compressor efficiency, that is, the rotor blade with optimal circumferential and axial stacking.

[0110] That is, aerodynamic performance is used as the objective function value, and the compressor efficiency is calculated with the help of mesh generation and CFD three-dimensional numerical simulation tools. The optimal compressor aerodynamic blade profile is then continuously sought through iterative optimization algorithms.

[0111] It should be noted that in other embodiments, step S2, which involves determining the circumferential position of the stacking line, can be completed first, followed by step S1, which involves determining the axial position of the stacking line, or steps S1 and S2 can be completed simultaneously.

[0112] Specifically, such as Figure 5 As shown, Figure 5 This is a flowchart of a method that includes all the steps described above.

[0113] like Figure 6 As shown, this embodiment also provides a rotor blade 10 of a segmented stacked axial compressor, which is obtained by the above-described design method for rotor blades of a segmented stacked axial compressor.

[0114] It should be noted that, Figure 6 The rotor blade 10 with the final airfoil shape is in Figure 7The rotor blades 10' of the original blade shape were obtained through the aforementioned design method for rotor blades of an axial compressor using segmented stacking. This can be derived from... Figure 7 As can be seen, the stacking line 20' of the original blade 10' is a straight line, while the stacking line 20 of the rotor blade 10 obtained in this embodiment is S-shaped in the circumferential direction.

[0115] This embodiment sets the stacking line of the rotor blades to an S-shape in the circumferential direction, which generates a reverse pressure gradient in the radial direction. This facilitates the migration of low-energy fluid from the blade tip region to the middle of the blade, thereby reducing the influence of secondary flow in the blade tip region, improving the flow conditions at the blade tip, and extending the stable operating range of the compressor rotor blades. Furthermore, the optimal maximum circumferential offset x is determined based on the principle of minimizing losses or maximizing efficiency. b and x c The corresponding height and extension length are determined, and the coordinates of the centroid positions of each basic airfoil section in the first interval conform to a specific function. This allows for the control of the circumferential offset of the centroid positions of each basic airfoil section within the first interval through a simple equation, improving the computational efficiency of airfoil optimization. In other words, this significantly improves computational efficiency compared to the trial-and-error method used in existing technologies to construct the stacking curve. On the other hand, existing technologies often use single-equation forms to describe approximate "S"-shaped curves; however, compressor stacking airfoil optimization typically involves optimizing multiple variables, such as the maximum circumferential offset and its position, and the tilt angle at the root tip. In such cases, a single-equation form cannot optimize the key characteristic parameters of the compressor's circumferential stacking. This method extracts the most critical design parameters of the compressor's circumferential stacking: the maximum positive and negative circumferential offsets and their height positions, and the tilt angle at the root tip. These are used as inputs (independent variables). A piecewise function is employed to solve for the offsets at other height positions in a closed-loop manner, ensuring smooth transitions at the connection points of the piecewise function and continuity of at least the first derivative. While ensuring that the key parameters of the compressor airfoil's circumferential stacking are fully described, the simple piecewise function construction makes the equations easy to solve, improving the efficiency of airfoil optimization. In other words, this invention provides a parameterization method that uses the maximum positive and negative circumferential offsets and their height positions, along with the tilt angle at the root tip, as control parameters. A piecewise function is used to solve for the offsets at other height positions in a closed-loop manner, ensuring smooth transitions at the connection points of the piecewise function and continuity of at least the first derivative.

[0116] While specific embodiments of the present invention have been described above, those skilled in the art should understand that these are merely illustrative examples, and the scope of protection of the present invention is defined by the appended claims. Those skilled in the art can make various changes or modifications to these embodiments without departing from the principles and essence of the present invention, but all such changes and modifications fall within the scope of protection of the present invention.

Claims

1. A method of designing a rotor blade of a segmented stator axial flow compressor, characterized in that, The rotor blade includes a blade root and a blade tip. The rotor blade has a plurality of basic airfoil sections along the height direction from the blade root to the blade tip. The centroids of each basic airfoil section are used as accumulation points and superimposed along the height direction of the rotor blade to obtain an accumulation line. The design method includes the following steps: S1: Determine the axial position of the stacking line; S2: Determine the circumferential position of the stacking line: The stacking line is "S"-shaped in the circumferential direction. The stacking line includes four control points A, B, C, and D along the height direction from the leaf root to the leaf tip. Point A is the centroid of the basic airfoil section of the leaf root and is the origin. Point B is located at the position of the maximum circumferential offset of the stacking line in the positive X-axis direction and is a preset point. Point C is located at the position of the maximum circumferential offset of the stacking line in the negative X-axis direction and is a preset point. Point D is the centroid of the basic airfoil section of the leaf tip and is located between points B and C on the X-axis. The coordinates (x1, y1) of the centroid positions of each of the basic airfoil sections in the first interval from point B to point C conform to the following formula: Where, N = (x b -x c ) / 2; α=Pi / (y c –y b Pi = 3.1415926; parameters And M based on the coordinates (x) of point B b y b ) and the coordinates (x) of point C c y c Substituting into the formula yields the result; S3: Determine the profile of the stacking line based on the axial and circumferential positions of the centroids of each basic blade section on the stacking line; extend the basic blades along the profile of the stacking line to construct a set of curves to obtain a three-dimensional rotor blade.

2. The design method for rotor blades of a segmented stacked axial compressor as described in claim 1, characterized in that, Step S2 includes: Extend the first interval downwards at point B along the height direction by a first extension height Δy. b Thus, the first intermediate point E is obtained; The coordinates (x2, y2) of the circumferential position points of the centroids of each of the basic airfoil sections in the second interval from point A to point E conform to the following formula: Wherein, the parameters h, i, j pass through the coordinates (x, y) of point E. e y e The slope k at point E e and the coordinates of point A (x a y a Substituting into the formula yields the result.

3. The design method for rotor blades of a segmented stacked axial compressor as described in claim 2, characterized in that, First extension height Δy b It is 5% to 20% of the height of the first interval.

4. The method of designing a rotor blade of an axial compressor with a segmented stack as recited in claim 1, wherein Step S2 includes: the first interval is extended upward along the height direction at point C by a second extension height Ay c , to obtain an intermediate point F; The coordinates (x3, y3) of the centroid positions of each of the basic airfoil sections in the third interval from point F to point D conform to the following formula: x3 = k f y3 + Q Where, k f Let Q be the slope at point F, and let Q pass through the coordinates (x, y) of point F. f y f The slope k at point F f Substituting into the formula yields the result.

5. The method of designing a rotor blade of an axial compressor with a segmented stack as claimed in claim 4, characterized in that The second extension height Δy c It is 5% to 20% of the height of the first interval.

6. The design method for rotor blades of a segmented stacked axial compressor as described in claim 1, characterized in that, In step S2, the coordinates of point B (x) b y b ), y b It is 25% to 35% of the leaf height; and / or, In step S2, the coordinates of point C (x) c y c ), y c It accounts for 70% to 85% of the leaf height.

7. The method for designing rotor blades of a segmented stacked axial compressor as described in claim 1, characterized in that, In step S1, determining the axial position of the stacking line includes: In the axial direction, the centroid position of each of the basic airfoil sections is moved along the height direction to a straight line with the centroid position of the basic airfoil section at the blade root as the reference point; the axial offset of the centroid position of the basic airfoil section at the blade tip is set as the maximum axial offset to determine the centroid position of the basic airfoil section at the blade tip; the axial offset of the centroid positions of the other basic airfoil sections is obtained by linear interpolation.

8. The method for designing rotor blades of a segmented stacked axial compressor as described in claim 7, characterized in that, In step S1, the centroid of the basic airfoil section at the blade tip is positioned forward of the centroid of the basic airfoil section at the blade root in the axial direction.

9. The method for designing rotor blades of a segmented stacked axial compressor as described in any one of claims 1-8, characterized in that, Following step S3, the following is also included: S4: Repeat steps S1 to S3 multiple times to obtain multiple different rotor blades, obtain the compressor efficiencies corresponding to the multiple different rotor blades, and select the rotor blade corresponding to the highest compressor efficiency as the final rotor blade.

10. A rotor blade of a segmented stator axial flow compressor, characterized in that The rotor blades are obtained by the rotor blade design method of the segmented stacked axial flow compressor as described in any one of claims 1-9.

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