JP2012233424A - Variable stator blade mechanism of axial flow type compressor - Google Patents

Abstract

PROBLEM TO BE SOLVED: To reduce pressure losses formed in a blade part without increasing the cost of components, and that, without requiring any special manufacturing process or maintenance.SOLUTION: A variable stator blade mechanism includes blade shafts 32a, 32b which are supported by the outer circumferential surface 2a and inner circumferential surface 33a of an annular flow passage 10 in a rocking manner, and a blade part 31 for variably changing the angle in the annular flow passage with the axis core of the blade shaft as the rocking center. In the blade part, the blade shaft is integrated therewith, and the blade part includes: side faces 31c, 31d which are arranged opposite to each other on the outer circumferential surface or inner circumferential surface of the annular flow passage, and extend to the upstream side and the downstream side of the annular flow passage from the blade shaft; and a positive pressure surface 31a and a negative pressure surface 31b which are continuous to the side faces, and arranged while maintaining the surface side and back side relationship to each other in the thickness direction of the side faces. A raised part 40 to be raised from the positive pressure surface is provided on side edges on the outer circumferential surface side and inner circumferential surface side of the annular flow passage in the positive pressure surface.

Description

  The present invention relates to an axial flow compressor such as a jet engine, and more particularly, to a variable stationary blade mechanism of an axial flow compressor that can vary the angle of a blade according to the rotational speed of the engine.

  Conventionally, a swing mechanism is provided in a stationary blade row of an axial flow compressor, and a variable stationary blade mechanism (VSV =) that varies the angle of the blade portion constituting the stationary blade row according to the rotational speed or output of the engine. Variable Static Vane) is known. In this variable stationary blade mechanism, for example, as shown in Patent Document 1, a blade shaft integrated with a blade portion is supported by a casing so as to be swingable. Further, the blade shaft is provided with a pinion, and an actuator ring having a rack meshing with the pinion is provided so as to be rotatable in the circumferential direction. When the actuator ring is rotated in the circumferential direction, the blade axis swings in the direction around the axis, thereby changing the angle of the blade portion.

  In the above variable stator vane mechanism, the outer peripheral surface of the annular flow channel located on the casing side or the inner peripheral surface of the annular flow channel located on the rotor side, because the blade portion is swung relative to the casing or the rotor. A clearance is formed between the two. Therefore, the air sucked into the annular flow channel leaks from the pressure surface side of the wing portion having a relatively high pressure to the negative pressure surface side of the wing portion having a relatively low pressure through the clearance. Pressure loss occurs.

  Therefore, in the variable stator vane mechanism disclosed in Patent Document 2, a flexible seal member is provided in the clearance formed between the blade portion constituting the stator blade row and the outer peripheral surface and inner peripheral surface of the annular flow path. To reduce the pressure loss.

JP 2010-001821 A JP 2000-345997 A

  However, since the above-described seal member is required to have a temperature change resistance capable of withstanding a violent temperature change and a wear resistance capable of withstanding a swing of the wing part, the seal member having such a temperature change resistance and wear resistance is required. If this is provided, the cost of parts will increase. Further, if the seal member is provided, the manufacturing process becomes complicated accordingly, and if the seal member deteriorates, maintenance must be performed.

  The present invention provides a variable stator vane mechanism for an axial-flow compressor that can reduce pressure loss generated in a blade portion without increasing the cost of parts and without requiring a special manufacturing process or maintenance. The purpose is that.

  In order to solve the above problems, a variable stationary blade mechanism of an axial-flow compressor according to the present invention includes a blade shaft that is swingably supported on one or both of an outer peripheral surface and an inner peripheral surface of an annular flow path. A variable vane mechanism for an axial-flow compressor comprising: a blade section that varies an angle with respect to a flow direction of the intake air sucked into the annular flow path with the axis of the blade shaft as a swing center. The blade portion is integrated with the blade shaft, is disposed facing the outer peripheral surface or the inner peripheral surface of the annular flow path, and has side surfaces extending upstream and downstream of the annular flow channel from the blade shaft; A pressure surface that is continuous with the side surface and is maintained in a side-by-side relationship in the thickness direction of the side surface, and a negative pressure surface that has a lower pressure than the pressure surface due to the flow of intake air sucked into the annular flow path. And either one of the outer peripheral surface side and the inner peripheral surface side edge of the annular flow path at the pressure surface Others The both characterized in that the ridges raised from the pressure surface is provided.

  Further, the variable stator blade mechanism of the axial flow compressor according to the present invention is characterized in that the raised portion starts to rise from the pressure surface on the upstream side of the annular flow path from the blade axis.

  Further, the variable stator blade mechanism of the axial-flow compressor according to the present invention is characterized in that the bulge starts to bulge from the end of the pressure surface located on the upstream side of the annular flow path.

  Further, in the variable stator blade mechanism of the axial flow compressor according to the present invention, the width of one raised portion in the axial direction of the blade shaft is within 40% of the width of the pressure surface in the axial direction of the blade shaft. It is characterized by.

  In the variable stator blade mechanism of the axial compressor of the present invention, the width of one raised portion in the axial direction of the blade axis is 10 to 30% of the width of the pressure surface in the axial direction of the blade axis. It is characterized by that.

  According to the present invention, it is possible to reduce the pressure loss generated in the wing portion without increasing the component cost and without requiring a special manufacturing process or maintenance.

It is a schematic sectional drawing of a jet engine. It is a partial schematic diagram of a compressor. It is an expansion perspective view of the wing | blade part which comprises a stationary blade row | line | column. It is a conceptual diagram of the wing | blade part which comprises a stationary blade row | line | column, (a) is a conceptual diagram which shows the side surface of a wing | blade part, (b) is a figure which shows the pressure surface of a wing | blade part. It is a figure explaining the shape of a protruding part. It is a figure which shows the comparative test data of the blade surface Mach number of the blade part of this embodiment, and the blade surface Mach number of a normal blade part. It is a figure which shows the comparative test data of the pressure loss coefficient of the wing | blade part of this embodiment, and the pressure loss coefficient of a normal wing | blade part.

  Hereinafter, preferred embodiments of the present invention will be described in detail with reference to the accompanying drawings. The dimensions, materials, and other specific numerical values shown in the embodiments are merely examples for facilitating the understanding of the invention, and do not limit the present invention unless otherwise specified. In the present specification and drawings, elements having substantially the same function and configuration are denoted by the same reference numerals, and redundant description is omitted, and elements not directly related to the present invention are not illustrated. To do.

  FIG. 1 is a schematic cross-sectional view of a jet engine. As shown in this figure, a jet engine 1 includes a compressor 3 that compresses air sucked into a casing 2, a combustion chamber 4 that combusts compressed air compressed by the compressor 3, and the combustion chamber 4. And a turbine 5 for converting the jet power of the exhaust jet generated in the combustion process to rotational energy. The rotational energy converted by the turbine 5 is transmitted to the rotor 7 of the compressor 3 through the shaft 6, and the compressor 3 is operated by the rotation of the rotor 7.

  The compressor 3 includes an annular flow path 10 formed at a distance between the casing 2 and the rotor 7. The facing interval between the casing 2 and the rotor 7 is formed so as to gradually decrease from the upstream side of the annular flow path 10 through which air is sucked toward the downstream side connected to the combustion chamber 4. The annular flow path 10 is configured to gradually narrow from the upstream side toward the downstream side. Thereby, the air sucked into the compressor 3 gradually increases in pressure as it is guided from the upstream side to the downstream side of the annular flow path 10.

  FIG. 2 is a partial schematic view of the compressor 3. As shown in this figure, the annular flow path 10 of the compressor 3 is defined by an inner wall surface 2a of the casing 2 and a peripheral surface 7a of the rotor 7, and a plurality of annular flow paths 10 are formed from the upstream side toward the downstream side. The moving blade rows 20 and the plurality of stationary blade rows 30 are alternately arranged.

  Each rotor blade row 20 includes a plurality of blade portions 21 (only one is shown in FIG. 2) protruding from the peripheral surface 7a of the rotor 7 into the annular flow path 10, and these plurality of blade portions 21 are annular flows. They are arranged at equal intervals with respect to the circumferential direction of the path 10 (rotation direction of the rotor 7). Each blade 21 has a shape capable of guiding intake air from the upstream side to the downstream side of the annular flow path 10 by rotating integrally with the rotor 7.

  Each stationary blade row 30 includes a plurality of blade portions 31 (only one is shown in FIG. 2) protruding from the inner wall surface 2 a of the casing 2 into the annular flow path 10. They are arranged at equal intervals in the circumferential direction of the annular flow path 10. In addition, each stationary blade row 30 includes the same number of blade shafts 32 as the blade portions 31 that are swingably supported by the casing 2, and each blade portion 31 is integrated with the blade shaft 32 and has an annular flow. It is held in the road 10. The blade shaft 32 is swung in the direction around the shaft by a swing mechanism (not shown) having a pinion provided outside the casing 2 and a rack meshing with the pinion. The angle of the wing portion 31 varies within the annular flow path 10 with the swinging motion.

  Note that the width of the annular channel 10 in the radial direction (vertical direction in the drawing) is larger on the upstream side than on the downstream side. Therefore, the width (length in the vertical direction in the figure) of the blade portion 31 of the stationary blade row 30 is also larger on the upstream side of the annular flow path 10 than on the downstream side. Therefore, in the stationary blade row 30 positioned on the upstream side of the annular flow path 10, the blade shaft 32 is supported on both the casing 2 side and the rotor 7 side so that the blade portion 31 does not rattle. .

  Specifically, on the upstream side of the annular flow path 10, an annular inner casing 33 (shroud) is fixedly provided between the rotor blade rows 20 like the casing 2. Thus, the blade shaft 32 is supported to be swingable. As a result, in the stationary blade row 30 positioned on the upstream side of the annular flow path 10, the blade portion 31 is supported by the blade shaft 32 in a both-sided manner. Hereinafter, the configuration of the stationary blade row 30 will be described in detail.

  3 is an enlarged perspective view of the wing portion 31 constituting the stationary blade row 30, FIG. 4 is a conceptual diagram of the wing portion 31, FIG. 4 (a) shows a side surface of the wing portion 31, and FIG. ) Shows the pressure surface of the wing part 31. As shown in these drawings, the wing portion 31 is bent at 90 degrees from the pressure surface 31a and the suction surface 31b that are curved in an arc and maintain a front-back relationship with each other, and from the side edges of the pressure surface 31a and the suction surface 31b. Side surfaces 31c and 31d that are continuous with each other. The pressure surface 31a and the suction surface 31b have such a shape that the pressure surface 31a has a relatively higher pressure than the suction surface 31b.

  The side surface 31 c is disposed to face the inner wall surface 2 a of the casing 2 in the annular flow path 10, and the side surface 31 d is disposed to face the outer wall surface 33 a of the inner casing 33 in the annular flow path 10. A blade shaft 32a that is swingably supported by the casing 2 is integrated with the side surface 31c, and a blade shaft 32b that is swingably supported by the inner casing 33 is integrated with the side surface 31d. . At this time, the side surfaces 31c and 31d maintain a dimensional relationship extending to the upstream side and the downstream side of the annular flow path 10 with respect to the blade shafts 32a and 32b. Note that the blade shaft 32 and the blade portion 31 may be configured separately or integrally formed. In any case, in the present embodiment, in the configuration in which the blade shaft 32 and the blade portion 31 are integrated, the blade shaft 32 and the blade portion 31 are integrally formed with the axis of the blade shaft 32 as the center of oscillation. All the configurations in which the blade shaft 32 and the blade portion 31 are in contact with each other or are in a continuous relationship are included.

  As described above, in a state where the blade shaft 32a is integrated with the side surface 31c and the blade shaft 32b is integrated with the side surface 31d, the pressure surface 31a and the negative pressure surface 31b of the blade portion 31 are within the annular flow path 10. It faces the circumferential direction.

  In the following, when the wing part 31 is held in the annular flow path 10, the end of the wing part 31 located on the upstream side of the annular flow path 10 is referred to as a front end part LE, and the downstream side of the annular flow path 10 is Let the edge part of the located wing | blade part 31 be the rear-end part TE. Further, the distance between the front end portion LE and the rear end portion TE is the length L1 of the blade portion 31, the distance between the pressure surface 31a and the suction surface 31b is the thickness L2 of the blade portion 31, and the distance between the side surfaces 31c and 31d. The distance will be described as the width L3 of the wing part 31.

  In the present embodiment, as shown in FIG. 4B, clearance x is formed between the side surface 31 c and the inner wall surface 2 a of the casing 2, and between the side surface 31 d and the outer wall surface 33 a of the inner casing 33. Is done. When the clearance x is formed in this way, in the flow of the intake air sucked into the annular flow path 10, the leakage from the pressure surface 31a having a relatively high pressure to the negative pressure surface 31b having a relatively low pressure. However, in order to reduce this leakage, the raised portion 40 is provided on the pressure surface 31a of the wing portion 31 in the present embodiment.

  The raised portions 40 are provided on both side edges in the width direction of the pressure surface 31a, in other words, on the side edge that continues to the side surface 31c and the side edge that continues to the side surface 31d. Therefore, in the wing part 31, the thickness L2a in the vicinity of both ends in the width direction is larger than the thickness L2b in the vicinity of the center in the width direction. In the present embodiment, the width L4 in the axial direction of the blade shaft 32 in one raised portion 40 is 30% of the width L3 of the pressure surface 31a. However, the relationship between the width L4 of the raised portion 40 and the width L3 of the pressure surface 31a is preferably 10-30% of the width L3, but if it is within 40%, the pressure loss is reduced. A sufficient effect can be obtained.

  Further, the raised portion 40 is provided on the front end portion LE side of the pressure surface 31 a, that is, on the upstream side of the annular flow path 10 with respect to the blade shaft 32. More specifically, the bulging portion 40 starts to bulge from the front end portion LE of the pressure surface 31 a and ends the bulging on the upstream side of the annular flow path 10 from the peripheral surface of the blade shaft 32. That is, the raised portion 40 is provided on the front end portion LE side of the pressure surface 31 a and on both side edges in the width direction of the wing portion 31.

  FIG. 5 is a diagram illustrating the shape of the raised portion 40 in detail. As described above, one bulge 40 is provided in the pressure surface 31a in the vicinity of the side 31c and in the vicinity of the side 31d. Here, the shape of the bulge 40 provided in the vicinity of the side 31c will be described. . As shown in FIG. 5A, in the pressure surface 31a, the end on the side 31c side is L = 0, the end on the side 31d side is L = 100, and the end on the side 31d side from the end on the side 31c side Assume that the distance to the part is indicated by L. FIG. 5B shows a cross section of the pressure surface 31a when L = 0, FIG. 5C shows a cross section of the pressure surface 31a when L = 10, and FIG. FIG. 5E shows a cross section of the pressure surface 31a when L = 30.

  As shown in FIG. 5B, the raised portion 40 is raised at the side edge on the side surface 31c side in a range from the front end portion LE to the peripheral surface of the blade shaft 32a. Then, as shown in FIG. 5C and FIG. 5E, the height of the raised portion 40, that is, the wing portion 31 in the raised portion 40, from the side surface 31 c toward the center in the width direction of the wing portion 31. The thickness L2a of the bulge portion 40 gradually decreases, and the bulge range of the bulge portion 40 that spreads from the front end portion LE toward the rear end portion TE is gradually narrowed. And as shown in FIG.5 (e), the protruding part 40 will lose | disappear completely in the position of L = 30.

  In this way, the bulging portion 40 gradually decreases the bulging height from the side surface 31c toward the width direction center side of the wing portion 31, and the bulging range in the direction from the front end portion LE to the rear end portion TE side. Is gradually narrowing. Here, the raised portion 40 provided on the side surface 31c side has been described, but also in the raised portion 40 provided on the side surface 31d side, from the side surface 31d side to the central side in the width direction of the wing portion 31 as described above. As it goes, the height of the bulge gradually decreases and the range of the bulge gradually decreases.

  FIG. 6 is a diagram showing comparison test data under a predetermined condition between the blade surface Mach number of the blade portion 31 of the present embodiment and the blade surface Mach number of a normal blade portion. However, FIG. 6A shows the distance from the inner peripheral surface (the outer wall surface 33a of the inner casing 33) to the outer peripheral surface (the inner wall surface 2a of the casing 2) of the annular channel 10, as shown in FIG. 4B. Is 100, the blade surface Mach number at a distance = 5 from the inner peripheral surface is shown, and FIG. 6B shows the blade surface Mach number at a distance = 10 from the inner peripheral surface. Yes.

  In addition, the normal wing | blade part which is a comparison test object differs from the wing | blade part 31 of this embodiment only in the point which is not provided with the protruding part 40, and all other shapes, materials, etc. are the same. 6 (a) and 6 (b), the solid line indicates the blade surface Mach number of the blade portion 31 of the present embodiment, and the dotted line indicates the blade surface Mach number of the normal blade portion. Is the blade surface Mach number on the pressure surface side, and SS in the figure is the blade surface Mach number on the suction surface side. When the distance (length) from the front end LE to the rear end TE = L1 is 1, the blade surface Mach number (vertical axis) for each position in the length direction of the blade 31 shown in the horizontal axis in the figure. ) Is as follows.

  In other words, the blade surface Mach number gradually decreases from the front end LE side to the rear end TE side on both the pressure surface and the suction surface. It is shown that it is decelerated and boosted. Here, the blade surface Mach number in the pressure surface of the normal blade portion increases rapidly at the front end portion LE, then gradually decreases toward the rear end portion TE, and rises again in the vicinity of the rear end portion TE. To do. Further, according to the normal blade portion, the difference in blade surface Mach number between the pressure surface and the suction surface, that is, the difference in flow velocity is larger on the front end portion LE side than on the rear end portion TE side. A difference of y1 is generated between the pressure surface and the suction surface on the front end portion LE side with respect to the blade shaft 32 indicated by a one-dot chain line.

  On the other hand, in the wing portion 31 of the present embodiment, the wing surface Mach number is closer to the front end portion LE than the wing shaft 32 due to the influence of the raised portions 40 provided on both side edges in the width direction of the pressure surface 31a. It is higher than the normal wing part 31. That is, in the wing part 31 of the present embodiment, as a result of acceleration of the intake air by the raised part 40, the wing surface Mach number between the pressure surface 31a and the negative pressure surface 31b on the front end LE side with respect to the wing shaft 32. Difference = y2 is smaller than blade surface Mach number difference = y1 in a normal wing part. Thus, if the flow velocity difference between the pressure surface 31a and the negative pressure surface 31b is reduced on the front end portion LE side of the wing portion 31, leakage from the pressure surface 31a to the negative pressure surface 31b generated through the clearance x is reduced. Therefore, the pressure loss caused by the clearance x can be reduced.

  FIG. 7 is a diagram showing comparison test data between the pressure loss coefficient of the wing part 31 of this embodiment and the pressure loss coefficient of a normal wing part. In this figure, the horizontal axis indicates the pressure loss coefficient, and the vertical axis indicates the distance from the inner peripheral surface of the annular flow path 10 (the outer wall surface 33a of the inner casing 33). However, here, the range of the inner 20% of the radial width of the annular channel 10 is shown, and in this embodiment, the range of 2% from the inner peripheral surface of the annular channel 10 is the annular flow. This corresponds to a clearance formed between the inner peripheral surface of the passage 10 (the outer wall surface 33a of the inner casing 33) and the side surface of the wing portion. Moreover, the solid line in the figure indicates the pressure loss coefficient generated by the wing part 31 of the present embodiment, and the dotted line in the figure indicates the pressure loss coefficient generated by the normal wing part.

  As is apparent from this figure, according to the present embodiment, the pressure loss can be reduced within the range where the raised portion 40 is provided, in other words, in the vicinity of the clearance x. In this way, in order to reduce the pressure loss, it is only necessary to provide the raised portion 40 on the blade portion 31 constituting the stationary blade row 30, so there is no need to provide a part such as a clearance seal. Maintenance is not required.

  In the above embodiment, the blade shaft 32 a of the blade portion 31 is supported by the outer peripheral surface (the inner wall surface 2 a of the casing 2) of the annular flow channel 10, and the blade shaft 32 b of the blade portion 31 is supported by the inner periphery of the annular flow channel 10. The case of so-called both-side support supported by the surface (the outer wall surface 33a of the inner casing 33) has been described. However, for example, in FIG. 2, the blade portion 31 may be supported only on the outer peripheral surface side of the annular flow path 10, like the stationary blade row 30 located on the most downstream side of the annular flow path 10. On the contrary, it may be supported only on the inner peripheral surface side of the annular flow path 10.

  Thus, even when the wing part 31 is cantilevered to either the outer peripheral surface or the inner peripheral surface of the annular flow path 10, between the side surface 31c of the wing part 31 and the annular flow path 10, and A clearance x is formed between the side surface 31 d of the wing part 31 and the annular flow path 10. Therefore, even when the wing portion 31 is cantilevered on either the outer peripheral surface or the inner peripheral surface of the annular flow path 10, by providing the raised portion 40 on the pressure surface 31 a of the wing portion 31, the above-described embodiment. Similarly to the above, the pressure loss can be reduced. In any case, the raised portion 40 can be applied to the blade portion 31 in which the blade shaft 32 is supported so as to be swingable on one or both of the outer peripheral surface and the inner peripheral surface of the annular flow path 10.

  Moreover, in the said embodiment, although it was decided that the protruding part 40 was provided in both the outer peripheral surface side edge and the inner peripheral surface side edge of the annular flow path 10 in the pressure surface 31a, Further, it may be provided only on one of the outer peripheral surface side edge and the inner peripheral surface side edge of the annular flow path 10.

  Moreover, in the said embodiment, although the bulging part 40 was provided in the upstream (front end part LE side) of the annular flow path 10 rather than the blade axis | shaft 32 among the pressure surfaces 31a of the wing | blade part 31, a bulging part The position where 40 is provided is not limited to the above embodiment. For example, the raised portion 40 may be provided on the downstream side (rear end TE side) of the annular flow path 10 with respect to the blade shaft 32, or may be continuous from the front end LE to the rear end TE of the wing portion 31. It is good also as providing.

  However, as is clear from FIGS. 6 (a) and 6 (b), the difference in blade surface Mach number (flow velocity) between the pressure surface 31a and the suction surface 31b is larger than that at the front end portion LE than the rear end portion TE side. The side is larger, and more specifically, it is particularly larger on the front end LE side than the blade shaft 32. Thus, when the difference in blade surface Mach number (flow velocity) between the pressure surface 31a and the negative pressure surface 31b increases, the leakage from the clearance x increases accordingly. Therefore, it is desirable that the raised portion 40 starts to rise from the pressure surface 31a at least on the front end LE side of the blade shaft 32.

  Further, on the pressure surface 31a of the wing part 31, the bulge 40 may start bulging from any position between the front end LE and the rear end TE, but the bulge of the bulge 40 from the front end LE. If the air pressure is started, the intake air flow is less likely to be disturbed, and the effect of reducing the pressure loss can be further enhanced.

  Furthermore, in the above-described embodiment, the height of the raised portion 40 and the raised range are gradually reduced from the side edge on the side surface 31c, 31d side of the wing portion 31 toward the center in the width direction of the wing portion 31. However, the raised portion 40 may be provided with the raised height and the raised range uniformly regardless of the position in the width direction of the wing portion 31. However, as in the above-described embodiment, when the bulge height and bulge range are gradually reduced, the flow of intake air is less likely to be disturbed, and the effect of reducing pressure loss can be further enhanced.

  As mentioned above, although preferred embodiment of this invention was described referring an accompanying drawing, it cannot be overemphasized that this invention is not limited to this embodiment. It will be apparent to those skilled in the art that various changes and modifications can be made within the scope of the claims, and these are naturally within the technical scope of the present invention. Is done.

  The present invention relates to an axial flow compressor such as a jet engine, and in particular, can be used for an axial flow compressor in which the angle of a blade portion can be varied according to the rotational speed of the engine.

DESCRIPTION OF SYMBOLS 1 Jet engine 2 Casing 2a Inner wall surface 3 Compressor 7 Rotor 7a Circumferential surface 10 Annular channel 30 Stator blade row 31 Blade part 31a Pressure surface 31b Negative pressure surface 31c, 31d Side surface 32 (32a, 32b) Blade shaft 33 Inner casing 33a Outside Wall 40 ridge

Claims (5)

A blade shaft that is swingably supported on one or both of the outer peripheral surface and the inner peripheral surface of the annular flow path;
A variable vane mechanism for an axial-flow compressor, comprising: a blade portion that varies an angle with respect to a flow direction of intake air sucked into the annular flow path with the axis of the blade shaft as a swing center. There,
The wing part is
The blade shaft is integrated, and is disposed facing the outer peripheral surface or the inner peripheral surface of the annular flow path, and the side surface extends upstream and downstream of the annular flow path from the blade shaft;
A pressure surface that is continuous with the side surface and is maintained in a front-back relationship in the thickness direction of the side surface, and has a lower pressure than the pressure surface due to the flow of intake air sucked into the annular flow path A suction surface, and
Axial-flow compression characterized in that a bulging portion is formed on one or both of the outer circumferential surface side and the inner circumferential surface side edge of the annular flow path on the pressure surface. Variable stator vane mechanism.   The variable vane mechanism for an axial-flow compressor according to claim 1, wherein the raised portion starts to rise from the pressure surface on the upstream side of the annular flow path from the blade axis.   The variable vane mechanism for an axial-flow compressor according to claim 2, wherein the bulge starts to bulge from an end of a pressure surface located on the upstream side of the annular flow path.   The width of one ridge in the axial direction of the blade axis is within 40% of the width of the pressure surface in the axial direction of the blade axis. The variable stator blade mechanism of the axial flow type compressor as described.   5. The axial flow type according to claim 4, wherein the width of the one raised portion in the axial direction of the blade axis is 10 to 30% of the width of the pressure surface in the axial direction of the blade axis. Variable stator vane mechanism of the compressor.

Images