Composite curved-swept rotor blade of aero-engine gas compressor and integral disc structure of composite curved-swept rotor blade

Abstract

The invention belongs to the technical field of three-dimensional design of aero-engine gas compressor composite swept-curved rotor blades, and particularly relates to an aero-engine gas compressor composite swept-curved rotor blade and an integral disc structure thereof, and the front edge of the rotor blade is sequentially provided with a blade root sweepback area A1, a sweepforward area A2 and a blade tip sweepforward area A3 from bottom to top; the blade root sweepback area A1, the sweepforward area A2 and the blade tip sweepforward area A3 account for 0-20%, 20-75% and 75-100% of the radial height range of the rotor blade respectively; the blade root sweepback area A1 and the blade tip sweepforward area A3 are provided with dihedral angles; a blade root sweepback area B1, a sweepback degree area B2 and a blade tip straight area B3 are sequentially arranged on the rear edge of the rotor blade from bottom to top; the blade root sweepback area B1, the sweepback passing area B2 and the blade tip straight area B3 account for 0-30%, 30%-80%, 75%-100% and 80%-100% of the radial height range of the rotor blade respectively; the blade root sweepback area B1 and the blade tip straight area B3 are provided with dihedral angles.

Description

Technical Field

[0001] This application belongs to the technical field of three-dimensional design of composite swept and bowed rotor blades for aero-engine compressors, and particularly relates to a composite swept and bowed rotor blade for an aero-engine compressor and its integral disk structure. Background Art

[0002] With the increase in the total pressure ratio and single-stage pressure ratio of the new generation of aero-engine compressors, rotor blades with sweep shapes have begun to be applied to compressors.

[0003] The development trend of compressors is that the number of stages is gradually decreasing, the pressure ratio is gradually increasing, and the compressor load is gradually increasing. Due to the increase in the compressor stage load, the design of rotor blades has faced great challenges. On the premise of ensuring high pressure ratio and high performance, rotor blades require sufficient strength, rigidity, and anti-vibration characteristics. The traditional disk-tenon connection rotor structure is gradually being replaced by the integral bladed disk structure.

[0004] The rotor blades of the integral bladed disk structure are integrally formed on the rotor disk and can be radially divided into the blade tip, blade root, rim, spoke, and hub according to their functions. The rotor blades have a leading edge and a trailing edge in the chord length direction, as Figure 1 shown.

[0005] The flow rate and stable operating margin of the compressor are largely determined by the rotor configuration. The integral bladed disk structure can effectively improve the performance of the compressor compared with the disk-tenon connection rotor structure. However, currently, the integral bladed disk structure adopts a two-dimensional rotor configuration design, and only relying on the sweep shape design of the rotor blades has limited effect on improving the compressor performance, and cannot maximize the potential of the integral bladed disk structure. In addition, in the current compressor design concept, a large stage pressure ratio, flow rate, and wide operating margin cannot be achieved at the same time. Based on this design, the obtained rotor blades are difficult to meet the working requirements of the new generation of compressors for high pressure ratio, large flow rate, and wide margin, and often only consider the aerodynamic performance of the rotor blades, aerodynamic characteristics such as aerodynamic loss at the blade tip, and ignore the mechanical characteristics of the rotor blades, vibration characteristics brought by the sweep shape, and gas-solid coupling characteristics such as order modes.

[0006] In view of the existence of the above technical defects, this application is proposed. Summary of the Invention

[0007] The purpose of this application is to provide a composite swept and bowed rotor blade for an aero-engine compressor and its integral disk structure to overcome or mitigate at least one aspect of the known technical defects.

[0008] The technical solution of this application is as follows:

[0009] On the one hand, a composite swept and bowed rotor blade for an aero-engine compressor is provided, including:

[0010] The leading edge of the rotor blade, from bottom to top, is successively the root sweep region A1, the sweep transition region A2, and the tip sweep region A3;

[0011] The root sweep region A1, the sweep transition region A2, and the tip sweep region A3 account for the radial height ranges of the rotor blade of 0 - 20%, 20% - 75%, and 75% - 100% respectively;

[0012] The root sweep region A1 and the tip sweep region A3 have dihedral angles;

[0013] The trailing edge of the rotor blade, from bottom to top, is successively the root sweep region B1, the sweep transition region B2, and the tip straight region B3;

[0014] The root sweep region B1, the sweep transition region B2, and the tip straight region B3 account for the radial height ranges of the rotor blade of 0 - 30%, 30% - 80%, 75% - 100%, and 80% - 100% respectively;

[0015] The root sweep region B1 and the tip straight region B3 have dihedral angles.

[0016] According to at least one embodiment of the present application, in the above aero - engine compressor compound - swept rotor blade, the sweep angle of the tip sweep region A3 is in the range of 20° to 40°;

[0017] The sweep angle of the root sweep region B1 is in the range of 5° to 30°.

[0018] According to at least one embodiment of the present application, in the above aero - engine compressor compound - swept rotor blade, for the root sweep region A1, the tip sweep region A3, the root sweep region B1, and the tip straight region B3, the dihedral angle changes from top to bottom between - 6° and + 30°, and the middle position of the blade is a non - bent region.

[0019] According to at least one embodiment of the present application, in the above aero - engine compressor compound - swept rotor blade, the sweep angle β and the dihedral angle α of the rotor blade are:

[0020]

[0021] Wherein,

[0022] 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.

[0023] According to at least one embodiment of the present application, in the above aero - engine compressor compound - swept rotor blade,

[0024] γ = arctan(Vy / Vx);

[0025] Among them,

[0026] Vx, Vy, and Vz are the components of the streamline velocity V in the X, Y, and Z directions.

[0027] On the other hand, an integral disk structure for the inlet stage of an aeroengine compressor is provided, and the rotor blades thereon adopt any one of the above-mentioned aeroengine compressor compound bowed and swept rotor blades. Description of the Drawings

[0028] Figure 1 is a schematic diagram of the integral blisk structure;

[0029] Figure 2 is a schematic diagram of the upper sweep distribution of the rotor blade provided by the embodiment of the present application;

[0030] Figure 3 is a distribution curve of the leading edge and trailing edge sweep angles of the rotor blade provided by the embodiment of the present application;

[0031] Figure 4 is a schematic diagram of the dihedral angle provided by the embodiment of the present application;

[0032] Figure 5 is a distribution of adding a dihedral angle feature on the basis of the rotor blade sweep provided by the embodiment of the present application;

[0033] Figure 6 is a schematic diagram of the distribution curve of adding a dihedral angle feature on the basis of the rotor blade sweep provided by the embodiment of the present application;

[0034] Figure 7 is a schematic diagram of the 2D modeling of the rotor blade provided by the embodiment of the present application;

[0035] Figure 8 is a schematic diagram of the 3D modeling of the rotor blade provided by the embodiment of the present application;

[0036] Figure 9 is a schematic diagram of the forward sweep and backward sweep of the rotor blade provided by the embodiment of the present application;

[0037] Figure 10 is a schematic diagram of the meridian angle and local axial tilt angle of the streamline surface under the meridional plane projection view provided by the embodiment of the present application;

[0038] Figure 11 is a schematic diagram of the tangential angle of the rotor blade in the YZ plane provided by the embodiment of the present application.

[0039] 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 Embodiments

[0040] 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 in conjunction with 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 rather than 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 general design.

[0041] In addition, unless otherwise defined, the technical terms or scientific terms used in the description of this application should have the ordinary meanings understood by those of ordinary skill in the field 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 and their equivalents listed after this word, without excluding other related concepts.

[0042] 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, words such as "installation" and "connection" used in the description of this application should be understood in a broad sense. For example, 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.

[0043] 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. The overall blisk structure has good strength and good resistance to vibration and bending. It is suitable for bending design. The dihedral angle can be incorporated into the airfoil design, especially the dihedral feature can be incorporated into the swept airfoil to reduce the aerodynamic loss, secondary flow loss and shock wave in the channel at the blade tip.

[0044] Based on the above embodiments of this application, a compressor composite swept and curved rotor blade and its integral disk structure for an aeroengine are provided. The rotor blade is designed as a combined airfoil structure of "swept shape" and "dihedral angle", and the rotor blade is three-dimensionally designed to fully explore the potential of the overall blisk structure, so as to achieve the best aerodynamic performance, effectively reduce the shock wave Mach number and blade tip loss, improve the endwall performance, reduce the aerodynamic loss of the blade tip clearance, and improve the efficiency of the compressor.

[0045] The swept and curved coupling design of the compressor rotor is a typical three-dimensional configuration design. Separating the leading and trailing edges of the rotor blade can generate a large airflow bend and improve the compressor performance.

[0046] The leading edge of the rotor blade, from bottom to top, is successively the root sweep region A1, the forward sweep transition region A2, and the tip forward sweep region A3, occupying radial height ranges of 0 - 20%, 20% - 75%, and 75% - 100% respectively.

[0047] The trailing edge of the rotor blade, from bottom to top, is successively the root sweep region B1, the sweep transition region B2, and the tip straight region B3, occupying radial height ranges of 0 - 30%, 30% - 80%, 75% - 100%, and 80% - 100% respectively, as Figure 2 shown.

[0048] The swept shape is mainly concentrated at the root and tip of the rotor blade. In particular, the sweep angle of the tip forward sweep region A3 near the tip of the leading edge is in the range of 20 - 40°, and the sweep angle of the root sweep region B1 is in the range of 5° - 30°. The remaining intervals are mainly sweep transition regions.

[0049] The distribution curves of the sweep angles of the leading edge and trailing edge of the rotor blade can be specifically designed as Figure 3 shown, where the forward sweep angle is -, and the backward sweep angle is +.

[0050] On the basis of the swept shape of the rotor blade, a dihedral angle feature is added, as Figure 4 shown. The dihedral angle bending region is mainly concentrated in the root sweep region A1, the tip forward sweep region A3, the root sweep region B1, and the tip straight region B3. That is, the root sweep region A1, the tip forward sweep region A3, the root sweep region B1, and the tip straight region B3 have dihedral angles, and the dihedral angle changes from top to bottom between -6° and +30°. The middle of the blade is a non-bending region, as Figures 5 - 6 shown.

[0051] 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 front 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.

[0052] The 2D shape of the rotor blade is as Figure 7 shown, and the 3D shape is as Figure 8 shown. The forward and backward swept structures of the rotor blade are as Figure 9 shown. In the meridional plane projection view, the meridional angle and the local axial tilt angle of the streamline surface are as Figure 10 shown. In the YZ plane, the tangential angle is as Figure 11 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 stack axis, G is the intersection point of the meridional plane and f2, t1 is the tangent line passing through G and f2, the X-axis is the engine axis direction, the Z-axis is directly above radially, and the XZ plane is the meridional plane of the engine.

[0053]

[0054]

[0055] Among them,

[0056] β is the grazing angle, and α is the dihedral angle;

[0057] is the meridional plane angle, μ is the local axial tilt angle, η is the local tangential tilt angle, and γ is the cylindrical flow angle.

[0058]

[0059] γ = arctan(Vy / Vx);

[0060] Among them,

[0061] 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.

[0062] The above-mentioned compressor composite sweep and bend rotor blade of an aero-engine breaks the traditional two-dimensional design concept, conducts three-dimensional composite modeling and combination design, conducts zonal design on the sweep shape, combines with the dihedral angle design, conducts the design of the "sweep and bend" zones, develops new three-dimensional design technologies, utilizes the integral blisk structure in three-dimensional space, gives full play to the performance of the compressor, can effectively control the airflow bending, improve the performance of the compressor, increase the compressor flow rate and the stable operating range.

[0063] The above-mentioned compressor composite sweep and bend rotor blade of an aero-engine is a rotor blade configuration method that comprehensively considers the coordinated action of the sweep angle and the dihedral angle. Relying on the integral blisk rotor structure, it integrates the three-dimensional configuration design method into the blade configuration design, conducts the sweep shape design for the front and trailing edges, focuses on the dihedral angle and sweep angle design in the area with the largest sweep shape, and considers the comprehensive effects of relevant velocities, tangential tilt angles, stacked axis axial tilt angles, and meridional plane angles on the "dihedral angle and sweep angle", which can provide technical reserves for the next generation of high-performance compressors, and gives the calculation methods for the key parameters "sweep angle" and "dihedral angle" of the three-dimensional configuration, providing a design basis for high-performance high-pressure ratio compressors.

[0064] So far, the technical solutions of the present application have 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 composite swept and bowed rotor blade for an aeroengine, characterized in that, Comprising: The leading edge of the rotor blade from bottom to top is successively the root swept area A1, the forward sweep transition area A2, and the tip forward sweep area A3; The root swept area A1, the forward sweep transition area A2, and the tip forward sweep area A3 account for the radial height range of the rotor blade of 0 - 20%, 20% - 75%, and 75% - 100% respectively; The root swept area A1 and the tip forward sweep area A3 have dihedral angles; The trailing edge of the rotor blade from bottom to top is successively the root swept area B1, the swept transition area B2, and the tip straight area B3; The root swept area B1, the swept transition area B2, and the tip straight area B3 account for the radial height range of the rotor blade of 0 - 30%, 30% - 80%, 75% - 100%, and 80% - 100% respectively; The root swept area B1 and the tip straight area B3 have dihedral angles.

2. The compressor compound bend - swept rotor blade of an aero - engine according to claim 1, characterized in that, The sweep angle of the tip forward sweep area A3 is in the range of 20° - 40°; The sweep angle of the root swept area B1 is in the range of 5° - 30°.

3. The compressor compound bend - swept rotor blade of an aero - engine according to claim 2, characterized in that, For the root swept area A1, the tip forward sweep area A3, the root swept area B1, and the tip straight area B3, the dihedral angle changes from top to bottom between - 6° and + 30°, and the middle position of the blade is a non - bending area.

4. The compressor compound bend - swept rotor blade of an aero - engine according to claim 3, characterized in that, The sweep angle β and the dihedral angle α of the rotor blade are: Wherein, 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.

5. The compressor compound bend - swept rotor blade of an aero - engine according to claim 4, characterized in that, γ = arctan(Vy / Vx); Wherein, Vx, Vy, and Vz are the components of the streamline velocity V in the X, Y, and Z directions.

6. An integral disk structure for the inlet stage of an aeroengine compressor, characterized in that, The rotor blade thereon adopts the compressor compound bend - swept rotor blade of an aero - engine according to claim 5.

Images