AXIAL COMPRESSOR

Abstract

An axial compressor includes: a cylindrical casing; a rotor rotatably disposed in the casing; rotor blades arranged on an outer circumferential surface of the rotor around a central axis of the rotor; stationary blades arranged on an inner circumferential surface of the casing at a position adjacent to and behind the rotor blades in an axial direction, each stationary blade having a tip opposing the rotor and having a chord length defined between a leading edge and a trailing edge of the tip; and at least one bleed passage having a bleed opening that opens out in the outer circumferential surface of the rotor. The bleed opening opens at a position located ahead of a position axially spaced rearward from the leading edges of the tips of the stationary blades by one half of the chord length in such a manner that the bleed opening faces toward the tips.

Description

TECHNICAL FIELD

The present invention relates to an axial compressor, and particularly to an axial compressor equipped with a bleed structure used in gas turbine engines for aircraft or the like.

BACKGROUND ART

Axial compressors used in gas turbine engines for aircraft are designed to operate properly with a large inflow air volume at rated operation (during large-output operation), such as when cruising, which occupies a large part of the operation time. At non-rated operation, such as when idling or taxiing, the inflow air volume is small and the inflow condition differs from the inflow condition at the rated operation. Therefore, at non-rated operation, the vane cascade may not operate stably, rotating stall may occur, the pressure efficiency may decrease, and/or the total pressure loss may become large.

To address such problems, it is known to provide a bleed hole in a part of a fluid passage of a compressor to suppress rotating stall by bleeding part of the compressed air (JP 559-168296A, for example).

However, in the prior art arrangement, the occurrence of rotating stall cannot be avoided sufficiently, and the effect of suppressing the decrease in the pressure efficiency of the compressor is small.

As a result of earnest research made by the present inventor, it has been found that the effect of suppressing the decrease in the pressure efficiency of the axial compressor by air bleed depends on the position of the bleed hole with respect to the chord position of the stationary blade section, and the pressure loss in the axial compressor is favorably reduced by providing the bleed hole at a particular chord position (range).

SUMMARY OF THE INVENTION

A primary object of the present invention is to provide an axial compressor in which a bleed hole is provided at an appropriate position such that the pressure loss in the axial compressor is effectively reduced by air bleed.

Means to Accomplish the Task

One embodiment of the present invention provides an axial compressor (36), comprising: a cylindrical casing (14); a rotor (20) rotatably disposed in the casing; multiple rotor blades (39) arranged on an outer circumferential surface (20B) of the rotor (20) at a predetermined pitch around a central axis (X) of the rotor; multiple stationary blades (41) arranged on an inner circumferential surface (14A) of the casing (14) at a position adjacent to and behind the rotor blades (39) in an axial direction of the rotor, each stationary blade having a tip (41A) opposing the rotor and having a chord length (LC) defined between a leading edge (41B) and a trailing edge (41C) of the tip; and at least one bleed passage (72) having a bleed opening (70) that opens out in the outer circumferential surface (20B) of the rotor (20), wherein the bleed opening (70) opens at a position located ahead of a position axially spaced rearward from the leading edges (41B) of the tips (41A) of the stationary blades (41) by one half of the chord length in such a manner that the bleed opening faces toward the tips.

According to this arrangement, the air bleed suppresses laminar flow separation produced in the downstream direction of the stationary blades, whereby rotating stall becomes less likely to occur and the pressure loss in the axial compressor is effectively reduced.

Preferably, the bleed opening (70) is located axially rearward of the leading edges (41B) of the tips (41A) of the stationary blades (41) by 10-20% of the chord length.

According to this arrangement, the pressure loss in the axial compressor air bleed is reduced remarkably by air bleed particularly when the inflow air volume is small.

Preferably, the at least one bleed passage (72) comprises multiple bleed passages arranged around the central axis (X) of the rotor (20) at a regular pitch.

According to this arrangement, spreading of air vortices produced in the vicinity of the front edges of the stationary blades to the downstream side of the stationary blades can be prevented evenly all around the central axis of the rotor.

Preferably, the at least one bleed passage (72) extends from the bleed opening (70) obliquely rearward at a predetermined angle (θ) relative to the axial direction of the rotor (20).

According to this arrangement, at non-rated operation, the compressed air containing vortices can flow from the bleed opening to the bleed passage easily and smoothly, whereby the pressure loss in the axial compressor is effectively reduced at non-rated operation.

Preferably, the predetermined angle (θ) is in a range from 20 to 40 degrees.

According to this arrangement, at non-rated operation, the compressed air containing vortices can flow from the bleed opening to the bleed passage easily and smoothly, whereby the pressure loss in the axial compressor is effectively reduced at non-rated operation.

Thus, the axial compressor of the present invention can reduce the pressure loss in the axial compressor effectively by air bleed.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a sectional view showing an overall structure of a gas turbine engine for aircraft including an axial compressor according to one embodiment of the present invention;

FIG. 2 is a sectional view schematically showing a part of the axial compressor;

FIG. 3 is a graph showing inflow air volume-pressure loss coefficient characteristics; and

FIG. 4 is a graph showing a relationship between total pressure and a span length.

DESCRIPTION OF THE PREFERRED EMBODIMENT(S)

In the following, an embodiment of an axial compressor according to the present invention will be described with reference to FIGS. 1 to 5.

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

As shown in FIG. 1, a gas turbine engine 10 includes a substantially cylindrical outer casing 12 and an inner casing 14 that are arranged coaxially. The inner casing 14 rotatably supports a low pressure rotary shaft (rotor) 20 therein via a front first bearing 16 and a rear first bearing 18. A tubular high pressure rotary shaft 26 is arranged so as to be rotatable around an outer circumference of an axially intermediate portion of the low pressure rotary shaft 20. The front portion of the high pressure rotary shaft 26 is supported by the inner casing 14 via a front second bearing 22 while the rear portion of the same is supported by the low pressure rotary shaft 20 via a rear second bearing 24. The low pressure rotary shaft 20 and the high pressure rotary shaft 26 are arranged coaxially, and the central axis thereof is denoted by a reference sign “X.”

The low pressure rotary shaft 20 includes a substantially conical tip portion 20A that protrudes more forward than the inner casing 14. An outer circumference of the tip portion 20A is provided with a front fan 28 including multiple fan blades 29, which are made of titanium alloy or the like and arranged to be spaced apart from one another in the circumferential direction. Multiple stator vanes 30, each having an outer end joined to the outer casing 12 and an inner end joined to the inner casing 14, are arranged on a downstream side of the front fan 28 so as to be spaced apart from one another at a predetermined interval in the circumferential direction. On a downstream side of the stator vanes 30, a bypass duct 32 defined between the outer casing 12 and the inner casing 14 to have an annular cross-sectional shape and an air compression duct (annular fluid passage) 34 defined coaxially (to be coaxial with the central axis X) in the inner casing 14 to have an annular cross-sectional shape are provided in parallel with each other.

An axial compressor 36 is provided in an inlet of the air compression duct 34. The axial compressor 36 includes two (front and rear) rotor blade tows 38 provided on an outer circumference of the low pressure rotary shaft 20 and two (front and rear) stationary blade rows 40 provided in the inner casing 14, such that the rotor blade rows 38 and the stationary blade rows 40 are arranged adjacent to each other and alternate in the axial direction.

Each of the rotor blade rows 38 includes multiple rotor blades 39 (see FIG. 2) extending radially outward from an outer circumferential surface 20B of the low pressure rotary shaft 20 in a cantilever fashion and arranged around the central axis X of the low pressure rotary shaft 20 at a predetermined pitch. Each of the stationary blade rows 40 includes multiple stationary blades 41 (see FIG. 2) extending radially inward from an inner circumferential surface 14A of the inner casing 14 (see FIG. 2) in a cantilever fashion and arranged around the central axis X of the low pressure rotary shaft 20 at a predetermined pitch at a position adjacent to and behind the corresponding rotor blade row 38 in the axial direction of the low pressure rotary shaft 20.

A centrifugal compressor 42 is provided in an outlet of the air compression duct 34. The centrifugal compressor 42 includes impellers 44 provided on an outer circumference of the high pressure rotary shaft 26. A stationary blade row 46 is provided in the outlet of the air compression duct 34 on an upstream side of the impellers 44. Further, a diffuser 50 is provided at an outlet of the centrifugal compressor 42, wherein the diffuser is fixed to the inner casing 14.

On a downstream side of the diffuser 50, a combustion chamber member 54 is provided to define a reverse-flow combustion chamber 52 to which compressed air is supplied from the diffuser 50. The inner casing 14 is provided with multiple fuel injection nozzles 56 for injecting fuel into the reverse-flow combustion chamber 52. The reverse-flow combustion chamber 52 produces high-pressure combustion gas by combusting air-fuel mixture therein. A nozzle guide vane row 58 is provided in an outlet of the reverse-flow combustion chamber 52.

On a downstream side of the reverse-flow combustion chamber 52, a high pressure turbine 60 and a low pressure turbine 62 are provided such that the combustion gas produced in the reverse-flow combustion chamber 52 is blown thereto. The high pressure turbine 60 includes a high pressure turbine wheel 64 fixed to an outer circumference of the high pressure rotary shaft 26. The low pressure turbine 62 is provided on a downstream side of the high pressure turbine 60 and includes multiple nozzle guide vane rows 66 fixed to the inner casing 14 and multiple low pressure turbine wheels 68 provided on an outer circumference of the low pressure rotary shaft 20 arranged in an axially alternating manner.

At the start of the gas turbine engine 10, a starter motor (not shown in the drawings) drives the high pressure rotary shaft 26 to rotate. Once the high pressure rotary shaft 26 starts rotating, the air compressed by the centrifugal compressor 42 is supplied to the reverse-flow combustion chamber 52, and air-fuel mixture combustion takes place in the reverse-flow combustion chamber 52 to produce combustion gas. The combustion gas is blown to the high pressure turbine wheel 64 and the low pressure turbine wheels 68 to rotate the turbine wheels 64, 68.

Thereby, the low pressure rotary shaft 20 and the high pressure rotary shaft 26 rotate, which causes the front fan 28 to rotate and brings the axial compressor 36 and the centrifugal compressor 42 into operation, whereby the compressed air is supplied to the reverse-flow combustion chamber 52. Therefore, the gas turbine engine 10 continues to operate after the starter motor is stopped.

During the operation of the gas turbine engine 10, part of the air suctioned by the front fan 28 passes through the bypass duct 32 and is blown out rearward, and generates the main thrust particularly in a low-speed flight. The remaining part of the air suctioned by the front fan 28 is supplied to the reverse-flow combustion chamber 52 and mixed with the fuel and combusted, and the combustion gas is used to drive the low pressure rotary shaft 20 and the high pressure rotary shaft 26 to rotate before being blown out rearward to generate thrust.

Next, an air bleed structure provided in the axial compressor 36 will be described with reference to FIG. 2.

The low pressure rotary shaft 20 constituting the rotor is formed with multiple bleed passages 72 arranged around the central axis X of the low pressure rotary shaft 20 at a regular pitch, wherein each bleed passage 72 includes a circular bleed opening 70 that opens out in the outer circumferential surface 20B of the low pressure rotary shaft 20 toward the air compression duct 34 (compressed air passage).

Part of the compressed air produced by the axial compressor 36 flows from the bleed openings 70 to the bleed passages 72 to be bled to the outside of the axial compressor 36.

As shown in FIG. 2, each stationary blade 41 has a tip (free end edge) 41A opposing the low pressure rotary shaft (rotor) 20, and the tip 41A of each stationary blade 41 has a leading edge (front end) 41B and a trailing edge (rear end) 41C such that a chord length LC of the stationary blade 41 is defined as a length between the leading edge 41B and the trailing edge 41C.

With respect to the cord position (position in a direction parallel to the chord of each stationary blade 41 or the axial direction), each bleed opening 70 opens at a position ahead of a position axially spaced rearward from the leading edges 41B of the tips 41A of the stationary blades 41 by one half of the chord length LC (0.5LC) in such a manner that the bleed opening 70 faces toward the tips 41A of the stationary blades 41. It is to be noted that the position of the bleed opening 70 may be measured as the position of the center of the bleed opening 70.

Preferably, in the embodiment illustrated in FIG. 2, each bleed opening 70 is located in a range from a position axially spaced rearward from the leading edges 41B of the tips 41A of the stationary blades 41 by 10% of the chord length LC to a position axially spaced rearward from the leading edges 41B by 20% of the chord length LC. Namely, the bleed opening 70 is preferably located in a chord position range of 0.1LC-0.2LC with respect to the leading edges 41B.

FIG. 3 is a graph showing a relationship between the inflow air volume and the pressure loss coefficient for a case where no bleed openings 70 are provided and cases where the bleed openings 70 are provided at respective different chord positions.

In FIG. 3, a characteristic curve A represents the inflow air volume-pressure loss coefficient characteristics in a case where the bleed openings 70 are located in a range of 0.1LC-0.2LC axially rearward of the leading edges 41B, a characteristic curve B represents the inflow air volume-pressure loss coefficient characteristics in a case where the bleed openings 70 are located in a range of 0.4LC-0.5LC axially rearward of the leading edges 41B, a characteristic curve C represents the inflow air volume-pressure loss coefficient characteristics in a case where the bleed openings 70 are located in a range of 0.8LC-0.9LC axially rearward of the leading edges 41B, and a characteristic curve D represents the inflow air volume-pressure loss coefficient characteristics in a case where no bleed openings 70 are provided.

As will be appreciated from this graph, when the bleed openings 70 are located ahead of the position axially spaced rearward from the leading edges 41B by one half of the chord length LC (0.5LC), preferably in a range of 0.1LC-0.2LC axially rearward of the leading edges 41B, the pressure loss coefficient is small, namely, the pressure loss is small.

FIG. 4 is a graph showing a relationship between the span position and the total pressure for a case where no bleed openings 70 are provided and cases where the bleed openings 70 are provided at respective different chord positions. In this graph, the position of the outer circumferential surface 20B of the low pressure rotary shaft 20 (innermost end position) is represented by a span length of 0, and the position of the inner circumferential surface 14A of the inner casing 14 (outermost end position) is represented by a span length of 1.

As will be appreciated from this graph, provision of the bleed openings 70 improves (or suppresses) the decrease in the total pressure or the pressure loss on the side of the span length of 0.

In the present embodiment, because the bleed openings 70 open ahead of a position axially spaced rearward from the leading edges 41B by 0.5 LC, and are preferably located in the chord position range of 0.1LC-0.2LC axially rearward of the leading edges 41B, such that the bleed openings 70 face toward the tips 41A of the stationary blades 41 from the side of the span length of 0, the compressed air produced in the axial compressor 36 and containing vortices flows from each bleed opening 70 to the corresponding bleed passage 72 to be bled efficiently to the outside of the axial compressor 36, whereby the laminar flow separation that occurs in the downstream direction of the stationary blades 41 is suppressed. Thereby, in the axial compressor 36 of the present embodiment, the pressure loss is effectively suppressed so that a high air compression efficiency is achieved.

Since the bleed openings 70 are provided at multiple positions around the central axis X of the low pressure rotary shaft 20 at a regular pitch, air bleed takes place at these positions around the central axis X of the low pressure rotary shaft 20, so that the suppression of the laminar flow separation produced in the downstream direction of the stationary blades 41 can be achieved evenly all around the central axis X of the low pressure rotary shaft 20. Thereby, the pressure loss in the axial compressor 36 can be reduced effectively.

Each bleed passage 72 extends from the corresponding bleed opening 70 rearward or in the direction of airflow in the axial compressor 36 and obliquely at a predetermined angle θ relative to the outer circumferential surface 20B (or axial direction) of the low pressure rotary shaft 20. Preferably, the angle θ is in a range from 20 to 40 degrees.

Because each bleed passage 72 is inclined as described above, the compressed air can flow from each bleed opening 70 to the corresponding bleed passage 72 easily and smoothly, whereby the pressure loss in the axial compressor 36 is reduced effectively.

In the foregoing, the present invention has been described in terms of the preferred embodiments thereof, but the present invention is not limited to the foregoing embodiments and various alterations and modifications may be made as appropriate.

For instance, it is not necessarily indispensable to provide multiple bleed openings 70 (or multiple bleed passages 72), and it may be possible to provide only a single bleed opening. The bleed openings 70 may not be circular, and may be of any other shape such as an ellipse, a rectangle, or an oblong circle that extends along the tips of the stationary blades 41.

Also, not all of the structural elements shown in the above embodiment(s) are necessarily indispensable and they may be selectively used as appropriate without departing from the scope of the present invention.

Claims

1. An axial compressor, comprising:

a cylindrical casing;

a rotor rotatably disposed in the casing;

multiple rotor blades arranged on an outer circumferential surface of the rotor at a predetermined pitch around a central axis of the rotor;

multiple stationary blades arranged on an inner circumferential surface of the casing at a position adjacent to and behind the rotor blades in an axial direction of the rotor, each stationary blade having a tip opposing the rotor and having a chord length defined between a leading edge and a trailing edge of the tip; and

at least one bleed passage having a bleed opening that opens out in the outer circumferential surface of the rotor,

wherein the bleed opening opens at a position located ahead of a position axially spaced rearward from the leading edges of the tips of the stationary blades by one half of the chord length in such a manner that the bleed opening faces toward the tips.

2. The axial compressor according to claim 1, wherein the bleed opening is located axially rearward of the leading edges of the tips of the stationary blades by 10-20% of the chord length.

3. The axial compressor according to claim 1, wherein the at least one bleed passage comprises multiple bleed passages arranged around the central axis of the rotor at a regular pitch.

4. The axial compressor according to claim 1, wherein the at least one bleed passage extends from the bleed opening obliquely rearward at a predetermined angle relative to the axial direction of the rotor.

5. The axial compressor according to claim 4, wherein the predetermined angle is in a range from 20 to 40 degrees.