Compressor rotor blade with variable blade tip bending angle and integral blade disc structure of compressor rotor blade

Abstract

The invention belongs to the technical field of aero-engine compressor design, and particularly relates to a compressor rotor blade with a variable blade tip bend angle and a blisk structure of the compressor rotor blade, and the front edge of the rotor blade is sequentially provided with a straight transition area A1, a sweepforward area A2 and a sweepforward area A3 from top to bottom; the straight transition area A1, the sweepforward area A2 and the sweepforward area A3 respectively account for 0-50%, 50-80% and 80-100% of the radial height of the rotor blade; a sweepback area B1, a turning transition area B2 and a sweepforward area B3 are sequentially arranged on the front edge of the rotor blade from top to bottom. The sweepback area B1, the turning transition area B2 and the sweepforward area B3 respectively account for 0-70%, 70-80% and 80-100% of the radial height of the rotor blade; the dihedral angle is set to be gradually reduced from + 4 degrees to 0 degree within the range of 0-10% of the radial height of the rotor blade, and the dihedral angle is set to be gradually increased from 0 degree to + 10 degrees within the range of 70-100% of the radial height.

Description

Technical Field

[0001] This application belongs to the technical field of aero-engine compressor design, and particularly relates to a compressor rotor blade with a variable tip bend angle and an integral blisk structure thereof. Background Art

[0002] The pressure ratio and flow rate of the next-generation aero-engine compressor are gradually increasing. Having fewer stages, a high pressure ratio, and high efficiency has become an inevitable trend for future compressors.

[0003] The rear-stage rotor in the compressor bears almost the highest temperature load and aerodynamic load, and often cooperates with the outlet-stage stator design to improve the performance and margin of the engine. The configuration of the last rotor with the traditional disk-finger connection structure is not conducive to vibration reduction and large aerodynamic loads, and moreover, it greatly limits the configuration of the rotor blade, making it difficult to meet the usage requirements of the rear-stage rotor in the next-generation engine.

[0004] The rotor blade of the integral blisk structure is integrally formed on the rotor disk, and can be radially divided into a blade tip, a blade root, a rim, a spoke, and a hub according to functions. The rotor blade has a leading edge and a trailing edge in the chord length direction. The integral blisk structure can reduce the vibration of the blade tip and the blisk, and improve the performance of the rear-stage rotor. However, the configuration of the rotor blade usually adopts two-dimensional design, which cannot fully exert the potential of the integral blisk structure, and the end-wall performance is average, and it cannot well meet the usage requirements of high-performance compressors.

[0005] In view of the existence of the above technical defects, this application is proposed. Summary of the Invention

[0006] The purpose of this application is to provide a compressor rotor blade with a variable tip bend angle and an integral blisk structure thereof, so as to overcome or mitigate at least one aspect of the known technical defects.

[0007] The technical solution of this application is as follows:

[0008] On the one hand, a compressor rotor blade with a variable tip bend angle is provided, including:

[0009] The leading edge of the rotor blade is successively a straight transition region A1, a forward-swept region A2, and a forward-swept region A3 from top to bottom;

[0010] The ranges of the straight transition region A1, the forward-swept region A2, and the forward-swept region A3 respectively accounting for the radial height of the rotor blade are 0 - 50%, 50% - 80%, and 80% - 100%;

[0011] The leading edge of the rotor blade is successively a backward-swept region B1, a transition region B2, and a forward-swept region B3 from top to bottom;

[0012] The swept region B1, the transition region B2, and the forward-swept region B3 respectively occupy ranges of 0 - 70%, 70% - 80%, and 80% - 100% of the radial height of the rotor blade;

[0013] Within the range of 0 - 10% of the radial height of the rotor blade, the dihedral angle is gradually reduced from +4° to 0°, and within the range of 70% - 100% of the radial height, the dihedral angle is gradually increased from 0° to +10°.

[0014] According to at least one embodiment of the present application, in the compressor rotor blade with the tip bend angle variation described above, the tip sweep angles of the leading edge and trailing edge of the rotor blade are in the range of -20° to -40°, and the sweep angles at the root and mid-span positions of the trailing edge are controlled at +10°.

[0015] According to at least one embodiment of the present application, in the compressor rotor blade with the tip bend angle variation described above, the sweep angle β and the dihedral angle α of the rotor blade are:

[0016]

[0017]

[0018]

[0019] γ = arctan(Vy / Vx);

[0020] Wherein,

[0021] is the meridian plane angle, μ is the local axial tilt angle, η is the local tangential tilt angle, and γ is the cylindrical flow angle;

[0022] Vx, Vy, and Vz are the components of the streamline velocity V in the X, Y, and Z directions.

[0023] According to at least one embodiment of the present application, in the compressor rotor blade with the tip bend angle variation described above, the tip region of the rotor blade has a variable tangential angle, and the tangential angle range is between 0° and 60°.

[0024] According to at least one embodiment of the present application, in the compressor rotor blade with the tip bend angle variation described above, the tangential angle of the rotor blade gradually increases from the leading edge to the trailing edge direction.

[0025] On the other hand, a rear-stage integral disk structure in an aeroengine compressor is provided, on which the rotor blades adopt any one of the compressor rotor blades with the tip bend angle variation described above.

[0026] A compressor rotor blade with a variable tip bend angle and its integral blisk structure are provided. Relying on the integral blisk structure, a three-dimensional configuration design is carried out. Based on the integral blisk, a design of sweep and lean partitioning is carried out. At the tip and root positions, a three-dimensional composite modeling design is carried out. A dihedral angle design is carried out in the forward sweep area of the front trailing edge. When applied to the mid- and rear-stage rotors, the endwall performance of the compressor can be improved, the airflow bending can be effectively controlled, and the performance of the compressor can be enhanced. It provides a technical reserve for the next generation of high-performance compressors, and provides a calculation method for the key parameters of the three-dimensional configuration, namely the sweep angle and the dihedral angle design, which can provide a design basis for high-performance high-pressure ratio compressors. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] Figure 1 FIG. is a schematic diagram of the upper sweep distribution of the rotor blade provided by the embodiment of the present application;

[0028] Figure 2 FIG. is a curve of the upper sweep and its dihedral angle distribution of the rotor blade provided by the embodiment of the present application;

[0029] Figure 3 FIG. is a schematic diagram of the two-dimensional modeling of the rotor blade provided by the embodiment of the present application;

[0030] Figure 4 FIG. is a schematic diagram of the three-dimensional modeling of the rotor blade provided by the embodiment of the present application;

[0031] Figure 5 FIG. is a schematic diagram of the meridian angle and the local axial tilt angle of the streamline surface under the meridional projection view of the rotor blade provided by the embodiment of the present application;

[0032] Figure 6 FIG. is a schematic diagram of the tangential angle design on the rotor blade provided by the embodiment of the present application;

[0033] Figure 7 FIG. is Figure 6 a partial cross-sectional view in the J1-J1 direction of

[0034] Figure 8 FIG. is Figure 6 a partial cross-sectional view in the J2-J2 direction of

[0035] Figure 9 FIG. is Figure 6 a partial cross-sectional view in the J3-J3 direction of

[0036] For better illustration of this embodiment, some contents of the drawings will be omitted, enlarged or reduced, and are only used for exemplary illustration and should not be construed as a limitation to the present application. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0037] To make the technical solutions and their advantages of this application clearer, the following will further describe the technical solutions of this application clearly and completely with reference to the accompanying drawings. It can be understood that the specific embodiments described herein are only partial embodiments of this application, which are only used to explain this application and do not limit this application. It should be noted that for the convenience of description, only the parts related to this application are shown in the drawings, and other related parts can refer to the usual design.

[0038] In addition, unless otherwise defined, the technical terms or scientific terms used in the description of this application should have the ordinary meaning understood by those of ordinary skill in the art to which this application belongs. The "including" used in the description of this application means that the concept appearing before this word covers the concepts listed after this word and their equivalents, without excluding other related concepts.

[0039] In addition, the words indicating directions used in the description of this application are only used to indicate relative directions or positional relationships. When the absolute position of the object being described changes, its relative positional relationship may also change accordingly. It should also be noted that unless otherwise clearly specified and limited, the words such as "installation" and "connection" used in the description of this application should be understood in a broad sense. For example, the connection can be a fixed connection or a detachable connection; it can be a mechanical connection or an electrical connection; it can be directly connected or indirectly connected through an intermediate medium. Those skilled in the art can understand their specific meanings in this application according to the specific situation.

[0040] The compressor rotor blade is composed of stacked airfoils. The airfoil is composed of a pressure surface and a suction surface facing each other front and back. The chord length extends from the leading edge to the trailing edge, and the radial height extends from the blade root to the blade tip. Based on the integral blisk, the middle and rear stage rotor blades are designed. The combined airfoil structure of "swept shape" and "dihedral angle" can effectively reduce shock waves and tip losses, improve the end wall performance of the compressor, increase the flow rate and efficiency of the compressor, fully exploit the potential of the integral blisk structure, and meet the usage requirements of high-performance compressors. Based on this, the embodiments of this application provide a compressor rotor blade with a variable tip bend angle and its integral blisk structure.

[0041] The swept coupling design of the compressor rotor is a typical three-dimensional configuration design. Separating the leading and trailing edge sweeps of the rotor blade can generate a larger airflow bend compared to the two-dimensional configuration design of the rotor blade, improving the compressor performance.

[0042] The leading edge of the rotor blade is successively a straight transition zone A1, a forward sweep zone A2, and a forward sweep zone A3 from top to bottom, respectively occupying the radial height ranges of 0 - 50%, 50% - 80%, and 80% - 100%.

[0043] The leading edge of the rotor blade is successively the swept-back area B1, the transition area B2, and the forward-swept area B3 from top to bottom, and the ranges of the radial height they occupy are 0 - 70%, 70% - 80%, and 80% - 100% respectively.

[0044] The swept shape distribution of the rotor blade is as Figure 1 shown. The swept shape is mainly concentrated in the mid-span and tip regions of the leading edge of the rotor blade, and the tip region of the trailing edge. The forward-swept angle is -, and the swept-back angle is +. The swept angles of the leading edge and the tip of the trailing edge of the rotor blade are mainly controlled within -20° to -40°, and the swept angles at the root and mid-span positions of the trailing edge are controlled at +10°. The root and mid-span positions of the leading edge are straight transition areas, as Figure 2 shown.

[0045] The dihedral angle can be used to reduce the tip loss, aerodynamic loss, secondary flow loss, and shock wave in the channel. On the basis of the swept shape, adding the dihedral angle feature can further significantly improve the aerodynamic performance of the rotor blade.

[0046] The dihedral angle bending area can be mainly designed and concentrated within the swept-back area B1 and the forward-swept area B3. The dihedral angle is between 0° and +10°. The mid-span position is a non-bending area. The dihedral angle gradually decreases from +4° to 0° within the radial height range of 0 - 10%, and gradually increases from 0° to +10° within the radial height range of 70% - 100%, as Figure 2 shown.

[0047] The swept shape and the dihedral angle of the rotor blade have a great influence on the performance and stability of the compressor. The compound swept shape of the leading and trailing edges is determined according to the performance and stability of the compressor. The leading-edge aerodynamic dihedral angle is defined with respect to the streamline surface. The 2D modeling of the rotor blade is as Figure 3 shown, and the 3D modeling is as Figure 4 shown. Under the meridional projection view, the meridional angle and the local axial inclination angle of the streamline surface are as Figure 5 shown. f1 is the tip, f2 is the blade shape, f3 is the root, LE is the leading edge, TR is the trailing edge, AK is the streamline surface, CE is the blade stacking axis, the G point is the intersection point of the meridional plane and f2, t1 is the tangent line passing through the G point and f2, the X-axis is the engine axis direction, the Z-axis is directly above the radial direction, and the XZ plane is the meridional plane of the engine.

[0048]

[0049]

[0050] Among them,

[0051] β is the swept angle, and α is the dihedral angle;

[0052] is the meridian plane angle, μ is the local axial tilt angle, η is the local tangential tilt angle, and γ is the cylindrical airflow angle.

[0053]

[0054] γ = arctan(Vy / Vx);

[0055] Wherein,

[0056] Vx, Vy, and Vz are the components of the streamline velocity V in the X, Y, and Z directions, and the initial velocity V0 corresponds to the leading edge position.

[0057] The compressor rotor blade with variable tip bend angle disclosed in the above embodiments relies on the integral blisk structure for three-dimensional configuration design. Based on the integral blisk, the design of bend and sweep zoning is carried out. At the tip and root positions, three-dimensional composite modeling design is carried out. At the forward sweep area of the front trailing edge, dihedral angle design is carried out. It is applied in the middle and rear stages of the rotor, which can improve the endwall performance of the compressor, effectively control the airflow bending, improve the performance of the compressor, provide technical reserves for the next generation of high-performance compressors, and provide a calculation method for the key parameters of the three-dimensional configuration, namely the sweep angle and the dihedral angle design, which can provide a design basis for high-performance high-pressure ratio compressors.

[0058] In order to further improve the aerodynamic / endwall performance in the middle and rear areas of the compressor, a variable tangential angle can be added to the tip region of the rotor blade, and the tip is bent circumferentially. The range of the bent tangential angle is between 0° and 60°, as Figures 6 - 9 shown, η-11 is the tangential angle of the non-bent area near the leading edge, η-21 is the tangential angle of the circumferential bend near the leading edge, η-12 is the tangential angle of the non-bent area near the blade middle, η-22 is the tangential angle of the circumferential bend near the blade middle, η-13 is the tangential angle of the non-bent area near the trailing edge, η-23 is the tangential angle of the circumferential bend near the trailing edge. The tangential angle can be specifically limited within the range of 10° to 25°, and gradually increases from the leading edge to the trailing edge direction.

[0059] The compressor rotor blade with variable tip bend angle disclosed in the above embodiments considers the influence of the tangential angle change on the configuration and aerodynamics of the rotor blade, and carries out the variable "tangential angle" design within the entire chord length range, forming a configuration design of the sweep + dihedral angle + variable tangential angle of the rotor blade, which can fully explore the potential of the integral blisk structure to achieve the best aerodynamic performance, reduce shock waves and tip losses, and improve the endwall performance and tip clearance loss.

[0060] So far, the technical solution of the present application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the protection scope of the present application is obviously not limited to these specific embodiments. Without departing from the principle of the present application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will fall within the protection scope of the present application.

Claims

1. A compressor rotor blade with a variable tip bend angle, characterized in that, Comprising: The leading edge of the rotor blade, from top to bottom, is successively a straight transition zone A1, a swept-forward zone A2, and a swept-forward zone A3; The ranges of the straight transition zone A1, the swept-forward zone A2, and the swept-forward zone A3 respectively accounting for the radial height of the rotor blade are 0 - 50%, 50% - 80%, and 80% - 100%; The leading edge of the rotor blade, from top to bottom, is successively a swept-back zone B1, a transition zone B2, and a swept-forward zone B3; The ranges of the swept-back zone B1, the transition zone B2, and the swept-forward zone B3 respectively accounting for the radial height of the rotor blade are 0 - 70%, 70% - 80%, and 80% - 100%; Within the range of 0 - 10% of the radial height of the rotor blade, the dihedral angle gradually decreases from +4° to 0°, and within the range of 70% - 100% of the radial height, the dihedral angle gradually increases from 0° to +10°.

2. The compressor rotor blade with variable tip bend angle according to claim 1, wherein The tip sweep angles of the leading edge and trailing edge of the rotor blade are in the range of -20° to -40°, and the sweep angles at the root and mid-span positions of the trailing edge are controlled at +10°.

3. The compressor rotor blade with variable tip bend angle according to claim 2, wherein The sweep angle β and dihedral angle α of the rotor blade are: γ = arctan(Vy / Vx); Wherein, where θ is the meridian plane angle, μ is the local axial tilt angle, η is the local tangential tilt angle, and γ is the cylindrical air flow angle; Vx, Vy, and Vz are the components of the streamline velocity V in the X, Y, and Z directions.

4. The compressor rotor blade with variable tip bend angle according to claim 3, wherein The tip region of the rotor blade has a variable tangential angle, and the tangential angle range is between 0° and 60°.

5. The compressor rotor blade with variable tip bend angle according to claim 4, wherein The tangential angle of the rotor blade gradually increases from the leading edge to the trailing edge direction.

6. An overall disk structure for the middle and rear stages of a compressor of an aeroengine, characterized in that, The upper rotor blade adopts the compressor rotor blade with variable tip bend angle according to claim 5.

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