Vibration control structure for rotor blades

Abstract

To provide a vibration control structure for rotor blades with excellent workability, which reduces the number of components and man-hours for assembly, and does not need a large space, in a rotary machine such as a gas turbine engine.SOLUTION: A vibration control structure for a plurality of rotor blades (10) planted on an outer periphery of a disk (5a) includes: a pedestal (17) located on an outer peripheral surface (50) of the disk (5a); and an engaging protrusion (19) in which an attachment part (14) attached to the disk (5a) in the rotor blades (10) engages with an engagement groove (18) provided inside the disk (5a). Between the outer peripheral surface (50) and the pedestal (17), a damper (21) is provided so as to span the plurality of rotor blades (10).SELECTED DRAWING: Figure 2

Description

[Technical field]

[0001] The present disclosure relates to a vibration damping structure for suppressing vibration of rotor blades in a rotary machine such as a gas turbine engine. [Background technology]

[0002] For example, in the compressor of a gas turbine engine, multiple rotor blades are attached to the outer periphery of the disk, and there is a concern that the rotor blades may resonate depending on the operating conditions. As a countermeasure, a vibration control part called a damper is usually installed on each rotor blade, but this increases the number of parts and the assembly man-hours. Therefore, in the following Patent Document 1, a wire is stretched and inserted through multiple rotor blades to suppress the vibration of the rotor blades. [Prior art documents] [Patent documents]

[0003] [Patent Document 1] Japanese Patent Application Publication No. 10-325302 Summary of the Invention [Problem to be solved by the invention]

[0004] However, the technology of Patent Document 1 is difficult to employ in small gas turbine engines because the wires are passed through the rotor blades, and because a large space is required to pass the wires through.

[0005] An object of the present disclosure is to provide a vibration damping structure for a rotor blade which reduces the number of parts and the number of assembly steps, does not require a large space, and is easy to work with. [Means for solving the problem]

[0006] In order to achieve the above object, the vibration damping structure of a rotor blade according to the present disclosure comprises: A vibration damping structure having a plurality of rotor blades planted on the outer periphery of a disk, an attachment portion of the rotor blade attached to the disk has a seat located on an outer circumferential surface of the disk and an engagement protrusion that engages with an engagement groove provided inside the disk, A damper spanning the plurality of rotor blades is interposed between the outer circumferential surface and the base.

[0007] According to this configuration, the damper is interposed between the outer peripheral surface of the disk and the base, so that the occurrence of resonance of the rotor blades can be suppressed. Also, since the damper spans multiple rotor blades, the number of dampers is suppressed, and as a result, the number of parts and the number of assembly steps are reduced. Furthermore, since the damper is only interposed between the outer peripheral surface of the disk and the base, the workability is good. Effect of the Invention

[0008] According to the vibration damping structure for rotor blades disclosed herein, a damper spanning multiple rotor blades is interposed between the outer peripheral surface of the disk and the base of the rotor blades, thereby reducing the number of parts and assembly labor, and providing a vibration damping structure for rotor blades that does not require a large space and is easy to work with. [Brief description of the drawings]

[0009] [Figure 1] 1 is a partially cutaway side view showing a gas turbine engine equipped with a vibration damping structure for a rotor blade according to the present disclosure; [Diagram 2] FIG. 2 is a partial enlarged cross-sectional view of a rotor blade according to an embodiment of the present disclosure. [Diagram 3] FIG. 2 is a plan view showing a main portion of a vibration damping structure for a rotor blade according to the present disclosure. [Figure 4] FIG. 2 is a perspective view showing a main portion of a vibration damping structure for a rotor blade according to the present disclosure. [Diagram 5] FIG. 2A is a perspective view showing a method of incorporating a damper according to the present disclosure into a rotor blade, and FIG. 2B is an enlarged perspective view showing a main portion surrounded by a rectangular dashed line in FIG. [Figure 6] FIG. 13 is a perspective view showing a modified example of the damper according to the present disclosure. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0010] Hereinafter, an embodiment of the present disclosure will be described with reference to the drawings, but the present disclosure is not limited to this embodiment.

[0011] FIG. 1 shows a gas turbine engine (hereinafter, sometimes simply referred to as a "gas turbine") GT using the vibration damping structure of the rotor blade of the present disclosure. This gas turbine GT includes a compressor 1, a combustor 2, and a turbine 3 as main components. Intake air IA taken into the gas turbine GT from an intake passage 11 is compressed by the compressor 1 to become compressed air CA, which is sent to the combustor 2, where it is mixed with fuel gas F supplied to the combustor 2 and combusted as an air-fuel mixture. High-temperature combustion gas CG generated in the combustor 2 is sent to the turbine 3 to rotate the turbine 3, and is then discharged to the outside from an exhaust passage 30 as exhaust gas EG. The rotational power of the turbine 3 is output via a rotating shaft 5. A load such as a generator is connected to the rotating shaft 5 of the gas turbine GT via a reduction gear (not shown).

[0012] The compressor 1 includes a disk 5a constituting a part of the rotating shaft 5 of the gas turbine GT, and a compressor casing 7. Multiple stages of compressor stator vanes 8 are attached at predetermined intervals to the inner periphery of the compressor casing 7. Meanwhile, multiple stages of compressor rotor blades 10 are provided on the outer periphery of the disk 5a so as to be located downstream of the compressor stator vanes 8 of each stage. The combination of the compressor stator vanes 8 and the compressor rotor blades 10 compresses air IA sucked from an intake passage 11.

[0013] FIG. 2 shows a rotor blade 10 according to an embodiment of the present disclosure. A plurality of rotor blades 10 are planted on the disk 5a at equal intervals in the circumferential direction. The rotor blade 10 has a blade portion 12 and an attachment portion 14 on the base side. The attachment portion 14 attached to the disk 5a has a base 17 located on the outer peripheral surface 50 of the disk 5a and an engagement protrusion 19 that engages with an engagement groove 18 provided inside the disk 5a. The base 17 of the rotor blade 10 is slidably inserted into the engagement groove 18 in the circumferential direction to attach the rotor blade 10 to the disk 5a. In detail, the attachment portion 14 of the rotor blade 10 in FIG. 2 is inserted into the disk 5a from an insertion recess 20 shown in FIG. 3 formed in a part of the outer peripheral surface 50 of the disk 5a and fed forward in the circumferential direction, thereby planting a plurality of rotor blades 10 on the outer periphery of the disk 5a.

[0014] The gap between the outer peripheral surface 50 of the disk 5a and the base 17 is larger on the low-pressure side in the disk axial direction AX than on the high-pressure side. By increasing the gap GA1 between the outer peripheral surface 50 of the disk 5a and the base 17 on the low-pressure side in the disk axial direction AX, which is the intake side of the compressed air CA, the ease of assembling the rotor blade 10 and the disk 5a is improved. By decreasing the gap GA2 between the outer peripheral surface 50 of the disk 5a and the base 17 on the high-pressure side, which is the exhaust side of the compressed air CA, the leakage flow LA from the rear, which is the high-pressure side, can be prevented.

[0015] A damper 21 is interposed in a gap GA1 between the seat 17 on the low pressure side and the outer peripheral surface 50 of the disk 5a. As shown in FIG. 4, the damper 21 is a plate material that straddles the seats 17 of the multiple rotor blades 10 arranged in the circumferential direction of the disk 5a. In this embodiment, the damper 21 is an annular plate material that straddles all the rotor blades 10. The material of the damper 21 is, for example, stainless steel or Inconel. The plate thickness is set to be as thin as about 0.2 to 0.4 mm, for example, 0.3 mm, thereby reducing the mass of the damper 21 and suppressing the outward displacement of the damper 21 due to centrifugal force. In addition, on the low pressure side of the disk axial direction AX where the gap between the outer peripheral surface 50 of the disk 5a and the seat 17 is large, resonance of the rotor blades 10 is likely to occur, but the damper 21 also has the effect of suppressing this resonance.

[0016] Fig. 5 shows a method of assembling the damper 21 to the rotor blade 10. In Fig. 5(a), as described above, when the rotor blade 10 is to be installed on the disk 5a, the damper 21 is inserted from the circumferential direction between the outer peripheral surface 50 of the disk 5a and the base 17 with only one or two rotor blades 10 remaining. By giving the damper 21, which is a plate material, a slight bend with a large radius of curvature in advance, it can be smoothly assembled into the rotor blade 10, and the workability is improved. Both ends 21a, 21b of the damper 21 are set to abut against each other with almost no circumferential gap as shown in Fig. 5(b).

[0017] 6, the plate material 21 may be divided into a plurality of pieces in the circumferential direction, for example, into two pieces. Even in this case, each plate material 21 spans a plurality of rotor blades 10, and therefore the number of plate materials 21 is less than the number of rotor blades 10. The damper 21 may be divided into three or more pieces. Also, instead of the plate material 21, a small diameter wire may be used.

[0018] The vibration damping structure for the rotor blade 10 shown in Fig. 2 of this embodiment is a vibration damping structure for multiple rotor blades 10 planted on the outer periphery of a disk 5a, and an attachment portion 14 attached to the disk 5a of the rotor blade 10 has a base 17 located on the outer periphery 50 of the disk 5a and an engagement protrusion 19 that engages with an engagement groove 18 provided inside the disk 5a, and a damper 21 that spans the multiple rotor blades 10 is interposed between the outer periphery 50 and the base 17. According to this configuration, the occurrence of resonance in the rotor blade 10 can be suppressed by a small number of dampers 21 interposed between the outer periphery 50 of the disk 5a and the base 17. In addition, leakage flow LA from the high pressure side to the low pressure side through gaps GA1 and GA2 between the base 17 of the rotor blade 10 and the outer periphery 50 of the disk 5a is also suppressed.

[0019] In the vibration damping structure for the rotor blades 10 according to the second aspect of the present embodiment, the damper 21 is annular and spans all of the rotor blades 10. According to this configuration, only a single damper 21 is required, so that the number of parts and the number of assembly steps can be further reduced.

[0020] In the vibration damping structure of the rotor blade 10 according to the third aspect of this embodiment, the damper 21 is made of a plurality of plate members divided in the circumferential direction of the disk 5a. This configuration can reduce the number of parts and the number of assembly steps compared to the case where a damper 21 is assembled for each rotor blade. In addition, since the damper 21 is a plate member divided in the circumferential direction, it has a certain width along the disk axial direction AX, and therefore the circumferential position is determined by the plate members adjacent in the circumferential direction butting against each other.

[0021] In the vibration control structure for the rotor blade 10 according to the fourth aspect of the present embodiment, the damper 21 is disposed on the low pressure side of the rotor blade 10 in the disk axial direction AX. This configuration can effectively suppress resonance of the rotor blade 10 that occurs on the low pressure side of the disk 5a due to a large gap GA1 between the rotor blade 10 and the disk 5a.

[0022] In the vibration damping structure for the rotor blade 10 according to the fifth aspect of the present embodiment, the rotor blade 10 is applied to the rotor blade of a compressor in a gas turbine engine GT. With this configuration, resonance that is likely to occur in the rotor blade 10 of the compressor 1 can be effectively suppressed.

[0023] The vibration damping structure of the present disclosure is not limited to the compressor 1 of a gas turbine engine GT, but can also be applied to the turbine 3. In addition, it can also be applied to a rotary machine such as a blower having rotor blades 10.

[0024] The present disclosure is not limited to the above-described embodiments, and various additions, modifications, and deletions are possible without departing from the scope of the present disclosure. Therefore, such additions, modifications, and deletions are also included in the scope of the present disclosure. [Explanation of symbols]

[0025] 1: Compressor 2: Combustor 3: Turbine 5: Rotation axis 5a:Disc 7: Casing 8: Static blade 10: Moving blade 11: Intake passage 12: Wings 14: Mounting part 17: Pedestal 18: Engagement groove 19: Engagement protrusion 20: Insertion recess 21: Damper 30: Exhaust passage 50: Outer disc surface AX: Disk axial direction CA: Compressed air CG: Combustion gas EG: Exhaust gas F: Fuel gas IA: Intake air LA: Leakage flow GA1: (Low pressure side) Gap GA2: (High pressure side) Gap GT: Gas turbine

Claims

1. A vibration damping structure having a plurality of rotor blades planted on the outer periphery of a disk, an attachment portion of the rotor blade attached to the disk has a seat located on an outer circumferential surface of the disk and an engagement protrusion that engages with an engagement groove provided inside the disk, a damper extending across the plurality of rotor blades is interposed between the outer circumferential surface and the base;

2. 2. The vibration damping structure for rotor blades according to claim 1, wherein the damper is annular and extends across all of the rotor blades.

3. 2. The vibration damping structure for a rotor blade according to claim 1, wherein the damper is made of a plurality of plate members divided in the circumferential direction of the disk.

4. 3. The vibration damping structure for a rotor blade according to claim 1, wherein the damper is disposed on a low pressure side of the rotor blade in the disk axial direction.

5. 3. The vibration damping structure for a compressor blade of a gas turbine engine according to claim 1 or 2, wherein the rotor blade is a rotor blade of a compressor of a gas turbine engine.

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