JP7041033B2 - Axial flow compressor - Google Patents

Description

The present invention relates to an axial flow compressor, and more particularly to an axial flow compressor with an bleed structure used in an aircraft gas turbine engine or the like.

Axial flow compressors used in gas turbine engines for aircraft are designed to accommodate the large inflow air volume during rated operation (high power operation) such as during cruising, which occupies most of the operating time. For this reason, under small flow rate operating conditions during non-rated operation such as idling and taxiing, the inflow conditions are different from those rated, the blade train does not operate stably, turning stall occurs, and pressure efficiency decreases. Invites, and the total pressure loss becomes large.

In view of this, it is known that a swirling stall is suppressed by providing an air extraction hole in a part of the fluid passage of the compressor and extracting a part of the compressed air (for example, Patent Document 1).

Japanese Unexamined Patent Publication No. 59-168296

However, the conventional one cannot sufficiently avoid the occurrence of turning stall, and the effect of suppressing the decrease in the pressure efficiency of the compressor is small.

As a result of diligent research by the present inventor, the action of suppressing the decrease in pressure efficiency of the axial flow compressor by bleeding is related to the position of the bleeding hole with respect to the cord position and the span position of the stationary wing portion, and the bleeding hole is defined as the stationary wing portion. It has been found that the pressure loss of the axial flow compressor is satisfactorily reduced by providing it at a specific cord position and span position.

The problem to be solved by the present invention is to provide an air extraction hole at an appropriate position for reducing the pressure loss of the axial flow compressor, and to effectively reduce the pressure loss of the axial flow compressor by bleeding air.

The axial flow compressor according to one embodiment of the present invention includes a cylindrical casing (14), a rotor (20) rotatably provided in the casing (14), and a central axis of the rotor (20). A plurality of rotor blades (39) provided on the outer peripheral surface (20B) of the rotor (20) with a predetermined pitch around (X), and after the rotor blades (39) in the axial direction of the rotor (20). An axial flow compressor (36) having a plurality of blades (41) provided on the inner peripheral surface (14A) of the casing (14) so as to be adjacent to the side of the rotor (20). The bleed air passage (72) including the bleed air opening (70) opened in the outer peripheral surface (20B) is provided, and the bleed air opening (70) is the free end edge of the stationary blade (41) facing the rotor (20). (41A) is open toward the free end edge at a position in front of the position 1/2 of the front end (41B) of the free end edge (41A) in the axial direction.

According to this configuration, the peeling generated in the downstream direction of the stationary blade (41) by the bleed air is suppressed, the turning stall is less likely to occur, and the pressure loss of the axial flow compressor (36) is effectively reduced.

In the axial flow compressor, preferably, the bleed air opening (70) is in the range of 10% to 20% from the front end (41B) in the axial direction of the free end edge (41A) of the stationary blade (41). Located inside.

According to this configuration, the pressure loss of the axial flow compressor (36) is significantly reduced by the bleed air, especially in the small inflow air amount.

In the axial flow compressor, preferably, a plurality of the bleed air openings (70) are provided around the central axis (X) of the rotor (20) at equal pitches.

According to this configuration, prevention of air vortices from diffusing in the downstream direction of the vane (41) near the leading edge of the vane (41) is prevented all around the central axis (X) of the rotor (20). It is done uniformly over.

In the axial flow compressor, preferably, each bleed passage (72) extends backward from the bleed opening (70) at a predetermined angle (θ) with respect to the outer peripheral surface of the rotor (20). ..

According to this configuration, compressed air containing a vortex easily flows from each bleed air opening 70 to the bleed air passage (72) during unrated operation, and the pressure loss of the axial flow compressor (36) during unrated operation is effectively reduced. Is done.

In the axial flow compressor, the predetermined angle (θ) is preferably in the range of 20 to 40 degrees.

According to this configuration, compressed air containing a vortex easily flows from each bleed air opening (70) to the bleed air passage (72) during non-rated operation, and the pressure loss of the axial flow compressor (36) during unrated operation is reduced. Is done effectively.

According to the axial flow compressor according to the present invention, the pressure loss of the axial flow compressor due to bleed air can be effectively reduced.

Sectional drawing which shows the outline of the gas turbine engine for the aircraft which uses the axial flow compressor of this embodiment. Sectional drawing which shows the main part of the axial flow compressor of this embodiment graphically. Inflow air volume-Graph showing pressure drop coefficient characteristics Graph showing the relationship between total pressure and span length

Hereinafter, embodiments of the axial flow compressor according to the present invention will be described with reference to FIGS. 1 to 5.

First, an outline of a gas turbine engine (turbofan engine) for an aircraft in which the axial flow compressor of the present embodiment is used will be described with reference to FIG. 1.

The gas turbine engine 10 has a substantially cylindrical outer casing 12 and an inner casing 14 arranged concentrically with each other. The inner casing 14 rotatably supports the low-pressure system rotary shaft (rotor) 20 by the front first bearing 16 and the rear first bearing 18 inside. The low-pressure system rotary shaft 20 rotatably supports the high-pressure system rotary shaft 26 by the hollow shaft by the front second bearing 22 and the rear second bearing 24 on the outer periphery. The low-pressure system rotary shaft 20 and the high-pressure system rotary shaft 26 are concentrically arranged, and their central axes are indicated by reference numerals X.

The low-pressure rotating shaft 20 includes a substantially conical tip portion 20A protruding forward from the inner casing 14. A front fan 28 including a plurality of fan blades 29 spaced in the circumferential direction is provided on the outer periphery of the tip portion 20A. On the downstream side of the front fan 28, a plurality of stator vanes 30 including an outer end joined to the outer casing 12 and an outer end joined to the inner casing 14 are provided at predetermined intervals in the circumferential direction. On the downstream side of the stator vane 30, a bypass duct 32 having an annular cross section formed between the outer casing 12 and the inner casing 14 and a circle formed concentrically with the inner casing 14 (concentric with the central axis X). An air compression duct (annular fluid passage) 34 having an annular cross section is provided in parallel.

An axial compressor 36 is provided at the inlet of the air compression duct 34. In the axial flow compressor 36, the front and rear two rows of rotor blade rows 38 provided on the outer periphery of the low pressure system rotary shaft 20 and the front and rear two rows of stationary blade rows 40 provided on the inner casing 14 are adjacent to each other in the axial direction. And have them alternately.

The rotor blade row 38 has a plurality of rotor blades 39 (FIG.) formed by a cantilever extending radially outward from the outer peripheral surface 20B of the low-pressure system rotary shaft 20 at a predetermined pitch around the central axis X of the low-pressure system rotary shaft 20. 2) is included. The stationary blade row 40 has a predetermined pitch around the central axis X of the low pressure system rotating shaft 20 so as to be adjacent to the rear side of the rotor blade row 38 in the axial direction of the low pressure system rotating shaft 20. It includes a plurality of stationary blades 41 (see FIG. 2) with cantilever beams extending inward in the radial direction from 14A (see FIG. 2).

A centrifugal compressor 42 is provided at the outlet of the air compression duct 34. The centrifugal compressor 42 has an impeller 44 provided on the outer periphery of the high-pressure system rotary shaft 26. At the outlet of the air compression duct 34, a stationary blade row 46 located on the upstream side of the impeller 44 is provided. A diffuser 50 fixed to the inner casing 14 is provided at the outlet of the centrifugal compressor 42.

A combustion chamber member 54 is provided on the downstream side of the diffuser 50 to define a backflow combustion chamber 52 to which compressed air is supplied from the diffuser 50. The inner casing 14 is provided with a plurality of fuel injection nozzles 56 for injecting fuel into the backflow combustion chamber 52. The backflow combustion chamber 52 produces high-pressure combustion gas by burning a mixture of fuel and air. A nozzle guide vane row 58 is provided at the outlet of the backflow combustion chamber 52.

A high-pressure turbine 60 and a low-pressure turbine 62 for injecting the combustion gas generated in the backflow combustion chamber 52 are provided on the downstream side of the backflow combustion chamber 52. The high-pressure turbine 60 includes a high-pressure turbine wheel 64 fixed to the outer periphery of the high-pressure system rotary shaft 26. The low-pressure turbine 62 is located on the downstream side of the high-pressure turbine 60, and has a plurality of nozzle guide vane rows 66 fixed to the inner casing 14 and a plurality of low-pressure turbine wheels 68 provided on the outer periphery of the low-pressure system rotary shaft 20. Have alternating directions.

When starting the gas turbine engine 10, the high-pressure system rotary shaft 26 is rotationally driven by a starter motor (not shown). When the high-pressure system rotary shaft 26 is rotationally driven, the air compressed by the centrifugal compressor 42 is supplied to the backflow combustion chamber 52, and combustion gas is generated by combustion of the air-fuel mixture in the backflow combustion chamber 52. .. The combustion gas is sprayed onto the high-pressure turbine wheel 64 and the low-pressure turbine wheel 68 to rotate these turbine wheels 64 and 68.

As a result, the low-pressure system rotary shaft 20 and the high-pressure system rotary shaft 26 rotate, the front fan 19 rotates, the axial flow compressor 36 and the centrifugal compressor 42 are operated, and compressed air is supplied to the backflow combustion chamber 52. .. As a result, the gas turbine engine 10 continues to operate even after the starter motor is stopped.

During the operation of the gas turbine engine 10, a part of the air sucked by the front fan 28 passes through the bypass duct 32 and is ejected rearward, and generates a main thrust particularly during low-speed flight. The rest of the air sucked by the front fan 28 is supplied to the backflow combustion chamber 52 and burns as an air-fuel mixture with the fuel, and the combustion gas contributes to the rotational drive of the low-pressure system rotary shaft 20 and the high-pressure system rotary shaft 26, and then rearward. It spouts out and generates thrust.

Next, the bleed air structure provided in the axial flow compressor 36 will be described with reference to FIG.

The low-pressure rotary shaft 20 forming the rotor has a low-pressure bleed passage 72 including a circular bleed opening 70 opened toward the air compression duct 34 (compressed air flow path) on the outer peripheral surface 20B of the low-pressure rotary shaft 20. A plurality of them are formed at equal pitches around the central axis X of the system rotation shaft 20.

A part of the compressed air generated by the axial-flow compressor 36 flows from each bleed air opening 70 to the bleed air passage 72, and is drawn out of the axial-flow compressor 36.

Each bleed opening 70 has a cord position (a position parallel to the axial direction) of the stationary blade 41 from the front end 41B of the free end edge 41A in the axial direction of the free end edge 41A facing the low pressure system rotation axis 20 of the stationary blade 41. It opens toward the free end edge 41A of the stationary blade 41 at a position in front of the 1/2 position.

Each bleed opening 70 is preferably located within a range of 10% to 20% from the front end 41B in the axial direction of the free end edge 41A of the stationary blade 41. Assuming that the cord length in the axial direction of the free end edge 41A is LC, each bleed opening 70 is on the front side of 0.5LC, that is, on the front side of 1 / 2LC, preferably the front end of the free end edge 41A of the stationary blade 41 in the axial direction. It is located at the code position within the range of 41B to 0.1LC to 0.2LC.

FIG. 3 is a graph showing the relationship between the inflow air amount and the pressure loss coefficient characteristic when the bleed air opening 70 is not provided and when the bleed air opening 70 is provided at different cord positions.

In FIG. 3, the characteristic line A shows that the bleed air opening 70 is located within the range of 0.1LC to 0.2LC from the front end 41B, and the characteristic line B shows the bleed air opening 70 in the range of 0.4LC to 0.5LC from the front end 41B. When located inside, the characteristic line C is the inflow air amount-pressure loss when the bleed air opening 70 is located within the range of 0.8 LC from the front end 41B, and the characteristic line D is the case where the bleed air opening 70 is not provided. Shows the coefficient characteristics.

From this graph, if the bleed air opening 70 is located on the front side of the front end 41B to the front side of 1 / 2LC, preferably within the range of 0.1LC to 0.2LC from the front end 41B, the pressure loss coefficient characteristic is small, that is, It can be seen that the pressure loss is small.

FIG. 4 is a graph showing the relationship between the span position and the total pressure when the bleed air opening 70 is not provided and when the bleed air opening 70 is provided at different cord positions. In this graph, the position of the outer peripheral surface 20B of the low pressure system rotary shaft 20 (outermost end position) is set to the span length 0, and the position of the inner peripheral surface 14A of the inner casing 14 (outermost end position) is set to the span length 1.

From this graph, it can be seen that when the bleed air opening 70 is provided, the reduction of the total pressure on the span length 0 side, that is, the reduction of the pressure loss becomes good.

In the present embodiment, the bleed air opening 70 is located in the cord position within the range of 0.1LC to 0.2LC from the front end 41B to the front side of 1 / 2LC, preferably from the front end 41B to the stationary blade 41 from the span length 0 side. By opening toward the free end edge 41A of the above, compressed air containing a vortex generated in the axial flow compressor 36 flows from each bleed air opening 70 to the bleed air passage 72, and is efficiently bleeded out of the axial flow compressor 36. , The peeling that occurs in the downstream direction of the stationary blade 41 is suppressed. As a result, in the axial flow compressor 36 of the present embodiment, reduction of pressure loss is effectively suppressed, and high air compression efficiency can be obtained.

Since a plurality of bleeding openings 70 are provided around the central axis X of the low pressure system rotating shaft 20 at equal pitches, bleeding is performed at each position around the central axis X of the low pressure system rotating shaft 20 in the downstream direction of the stationary blade 41. Suppressing the peeling that occurs in the low pressure system rotating shaft 20 is uniformly performed over the entire circumference around the central axis X. As a result, the pressure loss of the axial flow compressor 36 is effectively reduced.

Each bleed passage 72 has a predetermined angle θ, preferably an angle in the range of 20 to 40 degrees, with respect to the outer peripheral surface 20B of the low pressure system rotary shaft 20 behind the bleed opening 70, that is, on the airflow advancing side of the axial flow compressor 36. It is tilted and extended with θ.

Since each bleed air passage 72 is inclined in this way, compressed air easily flows from each bleed air opening 70 to the bleed air passage 72, and the pressure loss of the axial flow compressor 36 is effectively reduced.

Although the present invention has been described above with respect to its preferred embodiments, the present invention is not limited to such embodiments and can be appropriately modified without departing from the spirit of the present invention.

For example, the bleed air openings 70 do not necessarily have to be provided in a plurality, and may be a single number. The bleed air opening 70 may be an ellipse, a rectangle, an oval extending along the root of the stationary blade 41, or the like, in addition to the circular shape.

In addition, not all of the components shown in the above embodiments are indispensable, and they can be appropriately selected as long as they do not deviate from the gist of the present invention.

10: Gas turbine engine 12: Outer casing 14: Inner casing 14A: Inner peripheral surface 16: Front first bearing 18: Rear first bearing 19: Front fan 20: Low pressure system rotary shaft (rotor)
20A: Tip portion 20B: Outer peripheral surface 22: Front second bearing 24: Rear second bearing 26: High-pressure system rotating shaft 28: Front fan 29: Fan blade 30: Stator vane 32: Bypass duct 34: Air compression duct 36 : Axial flow compressor 38: Turbine row 39: Turbine 40: Static blade row 41: Static wing 41A: Free end edge 41B: Front end 42: Centrifugal compressor 44: Impeller 46: Static blade row 50: Diffuser 52: Backflow Combustion chamber 54: Combustion chamber member 56: Fuel injection nozzle 58: Nozzle guide vane row 60: High pressure turbine 62: Low pressure turbine 64: High pressure turbine wheel 66: Nozzle guide vane row 68: Low pressure turbine wheel 70: Extraction opening 72: Extraction passage

Claims (3)

With a cylindrical casing,
A rotor rotatably provided in the casing,
A plurality of blades provided on the outer peripheral surface of the rotor with a predetermined pitch around the central axis of the rotor, and
An axial flow compressor having a plurality of stationary blades provided on the inner peripheral surface of the casing so as to be adjacent to the rear side of the moving blade in the axial direction of the rotor.
It has an bleed air passage including an bleed air opening opened on the outer peripheral surface of the rotor.
The bleed air opening is only within a range of 10% to 20% of the span length of the free end edge from the front end of the free end edge in the axial direction of the free end edge of the stationary blade facing the rotor. Axial flow compressor that opens toward the edge. The axial flow compressor according to claim 1, wherein a plurality of the bleed air openings are provided around the central axis of the rotor at equal pitches. The axial flow compressor according to claim 1 or 2 , wherein each bleed passage extends from the bleed opening so as to incline rearward at a predetermined angle with respect to the outer peripheral surface of the rotor.

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