Blade member and rotary machine

The airfoil member of the present invention is provided with an airfoil body, an end wall which is installed at an end part of the airfoil body, a fillet portion which smoothly connects the end part of the airfoil body with the end wall, and a cooling channel which allows a cooling medium to circulate inside the airfoil body and the end wall and in which two main channels are connected so as to bend in a folding manner at a return channel. The return channel is formed so as to run along the fillet portion on a cross section intersecting with a center line of a profile of the airfoil body and also formed in such a manner that the width thereof in the profile thickness direction is greater than the width of the main channel in the profile thickness direction.

TECHNICAL FIELD

The present invention relates to an airfoil member and a rotary machine.

The present application claims the right of priority to Japanese Patent Application No. 2011-096459 filed on Apr. 22, 2011, in Japan, with the content cited herewith.

BACKGROUND ART

As is well known, there is an airfoil member as one of the most important elements in a rotary machine. For example, an airfoil member installed on a stator side performs as a turbine vane for rectifying a working fluid. Further, an airfoil member installed on a rotor side performs as a turbine blade for recovering energy from a working fluid and imparting energy to the working fluid.

Where the above-described airfoil member is used in a high-temperature environment, it is necessary to cool the airfoil member for suppressing oxidation-caused thinning and fatigue of the airfoil member.

For example, a turbine vane described in Patent Document 1 given below is provided with an airfoil body extending in a radial direction of a turbine and an end wall which is formed at the tip end of the airfoil body to extend so as to intersect in the radial direction of the turbine. Inside the airfoil body, a serpentine channel is formed which connects in a meandering manner with a plurality of cooling channels extending in the radial direction of the turbine. Then, cooling air is allowed to circulate into the serpentine channel, thereby cooling the airfoil member.

A turbine blade which has adopted a similar cooling channel is described in Patent Document 2 given below.

PRIOR ART DOCUMENTS

Patent Documents

Patent Document 1: Japanese Published Unexamined Patent Application No. H10-299409

Patent Document 2: Japanese Published Unexamined Patent Application No. 2006-170198

SUMMARY OF THE INVENTION

Problems to be Solved by the Invention

Moreover, in the above-described airfoil member, as shown inFIG. 13andFIG. 14, cooling air c which flows through a serpentine channel74cools an airfoil body70from the inside.

However, in the above-described airfoil member, there is usually formed a fillet portion73which smoothly connects an end part71(base portion) of the airfoil body70with an end wall72(platform) thereof. Therefore, as shown inFIG. 14, around the fillet portion73, a site is developed with a large blade wall thickness which covers from a return part75of the serpentine channel74to an outer surface of the fillet portion73. Thus, there is a possibility that the site having a large blade wall thickness is not sufficiently cooled by the cooling air c.

The present invention has been made in view of the above situation, an object of which is to sufficiently cool a fillet portion.

Means for Solving the Problems

The airfoil member of the present invention is an airfoil member which is provided with an airfoil body, an end wall which is installed at an end portion of the airfoil body in a blade span direction and extends so as to intersect in the blade span direction, a fillet portion which smoothly connects the end portion of the airfoil body with the end wall, and a cooling channel which allows a cooling medium to circulate inside the airfoil body and the end wall and in which two main channels extending along the blade span direction are connected so as to bend in a folding manner at a return channel formed on the end wall side. The return channel is formed so as to run along the fillet portion on a cross section intersecting with a center line of a profile of the airfoil body and also formed in such a manner that the width thereof in a profile thickness direction of the blade is greater than the width of the main channel in the profile thickness direction.

According to the above constitution, since the return channel is formed so as to run along the fillet portion on the cross section intersecting with the center line of the blade profile, from the return channel to the outer surface of the fillet portion has a substantially uniform thickness. It is, thereby, possible to prevent formation of a site which is increased in blade wall thickness and give uniform and sufficient cooling to the fillet portion.

Further, the return channel may be provided on an inner surface of the return channel with a cooling face which is formed along an outer surface of the fillet portion.

Accordingly, the cooling face is provided, by which the fillet portion that is opposite to the cooling face can be cooled more sufficiently.

Further, the return channel may be provided with a projection part which is formed at the center of the airfoil body in the profile thickness direction to guide the flow of the cooling medium to both sides in the profile thickness direction.

Accordingly, the projection part is provided, by which the cooling medium can be guided to both sides in the profile thickness direction. Thereby, it is possible to cool sufficiently the fillet portion positioned on both sides of the return channel in the profile thickness direction.

Further, the return channel may be provided with a cooling hole on a partition wall from an upstream side channel of the main channel which is positioned on the upstream side in the return channel.

Accordingly, compressed air flowing in the proximity of the cooling face of the return channel is exchanged, by which the cooling face is further improved in cooling capability.

Further, the cooling face may be formed in such a manner that a distance of the cooling face from the outer surface of the fillet portion is substantially the same distance from the outer surface of the airfoil body to the inner surface of the main channel.

Accordingly, the distance of the cooling face from the outer surface of the fillet portion is formed so as to be substantially the same distance from the outer surface of the airfoil body to the inner surface of the main channel, thus making it possible to provide uniform cooling between the airfoil body and the fillet portion.

Still further, the cooling face may extend along the center line of the blade profile.

Accordingly, since the cooling face extends along the center line of the blade profile, it is possible to give uniform and sufficient cooling to the fillet portion over a wide range along the center line of the blade profile.

In addition, a rotary machine of the present invention is provided with the above-described airfoil member.

Accordingly, the above-described airfoil member is provided, by which the airfoil member can be enhanced in cooling effect to provide a highly reliable rotary machine.

Effects of the Invention

The airfoil member of the present invention is able to give uniform and sufficient cooling to the fillet portion.

Further, the rotary machine of the present invention can be improved in reliability.

MODE FOR CARRYING OUT THE INVENTION

Hereinafter, a description will be given of embodiments of the present invention with reference to the drawings.

First Embodiment

FIG. 1is a half sectional view which shows a schematic constitution of the gas turbine (rotary machine) GT related to the first embodiment of the present invention. As shown inFIG. 1, the gas turbine GT is provided with a compressor C for generating compressed air (cooling medium) c, a plurality of combustors B for generating combustion gas g by supplying a fuel to the compressed air c supplied from the compressor C, and a turbine (rotary machine) T for obtaining rotary power from the combustion gas g supplied from the combustors B.

In the gas turbine GT, a rotor Rcof the compressor C and a rotor RTof the turbine T are coupled at the shaft ends to extend above a turbine axis P.

In the following description, an extending direction of the rotor RTis referred to as a turbine axial direction, a circumferential direction of the rotor RTis referred to as a turbine circumferential direction, and a radial direction of the rotor RTis referred to as a turbine radial direction.

FIG. 2is an enlarged sectional view which shows a major part of the turbine T, and an enlarged view which shows the major part I inFIG. 1.

As shown inFIG. 2, the turbine T is provided with turbine vanes2and turbine blades (airfoil members)3which are disposed alternately at four stages inside a turbine casing1in the turbine axial direction. The turbine vanes2at each stage constitute an annular turbine vane array, with an interval kept in the turbine circumferential direction, each of which is fixed to the turbine casing1and extends toward the rotor RT.

In a similar manner, the turbine blades3at each stage constitute an annular turbine blade array, with an interval kept in the turbine circumferential direction. They are fixed to rotor disks4A to4D of the rotor RTand also extend toward the turbine casing1.

As shown inFIG. 2, the rotor RTis provided with the rotor disks4A to4D which are stacked in the turbine axial direction and formed in a shaft shape viewed as a whole. First to fourth-stage turbine blades3A to3D are respectively fixed to the outer circumferences by these rotor disks4A to4D. In the following description, where the reference numerals3A to3D,4A to4D, and7A to7D specify individual items, an uppercase letter follows the reference numeral. Where they do not specify individual items, an uppercase letter is omitted.

Of the rotor disks4A to4D, between two disks adjacent in the turbine axial direction, a manifold5is formed which extends in the turbine circumferential direction. A seal disk6, the center axis of which is pointed at the turbine axial direction, is connected to the rotor disk4A on the upstream side of the first-stage rotor disk4A, and the manifold5is formed between the seal disk6and the rotor disk4A.

These manifolds5are continuously connected via a connection hole5adrilled on each of the rotor disks4A to4C and the seal disk6. Compressed air c extracted from the compressor C flows sequentially from the seal disk6side to each of the manifolds5.

Further, the rotor disks4A to4D are respectively formed with radial holes7A to7D for guiding the compressed air c from each of the manifolds5on the upstream side to a cooling channel50(refer toFIG. 3) formed inside each of the turbine blades3A to3D. These radial holes7A to7D are formed in a multiple number respectively on the rotor disks4A to4D, with an interval kept in the turbine circumferential direction.

FIG. 3is a sectional view which is cut along the center line Q of the profile of the turbine blade3(sectional view taken along line II to II inFIG. 4), andFIG. 4is a blade shape sectional view which intersects with the turbine blade3in the blade span direction (sectional view taken along line III to III inFIG. 3).

As shown inFIG. 3, the turbine blade3is constituted continuously with the airfoil body10, the platform (end wall)20and a blade basement30from the other side of the airfoil body10in the blade span direction to one side thereof in the above-described order.

As shown inFIG. 3, the airfoil body10turns the blade span direction to the turbine radial direction and extends, as shown inFIG. 2, from a base end (end part)14formed on the rotor disks4(4A to4D) side (shown inFIG. 3) to a tip end15positioned on the turbine casing1.

Further, as shown inFIG. 4, the airfoil body10turns the profile thickness direction of the blade to the turbine circumferential direction. On the blade-shape cross section intersecting in the blade span direction, a leading edge11is formed in a round shape, while a trailing edge12is formed in a pointed shape. Still further, as shown inFIG. 4, the airfoil body10is formed with a suction side13awhich is formed in a raised shape on a blade surface13on one side in the turbine circumferential direction and a pressure side13bwhich is formed in a recessed shape on the other side in the turbine circumferential direction.

As shown inFIG. 3, the platform20is formed so as to continue to one side in the turbine axial direction with respect to the base end14of the airfoil body10, extending so as to intersect in the blade span direction.

The platform20is smoothly connected to the base end14of the airfoil body10at a fillet portion40. The fillet portion40is formed in a circumferential manner along a contour of the blade-shape cross section of the base end14, and a cross sectional contour which is cut along the blade span direction assumes a quarter-round arc shape (refer toFIG. 6andFIG. 7).

As shown inFIG. 3, the blade basement30is formed so as to continue to one side in the turbine radial direction (close to turbine axis P) with respect to the platform20and formed, for example, in the shape of a Christmas tree or in the shape of a triangle shape. The blade basement30is fitted into a groove of the blade basement (not shown) formed on an outer circumference of the rotor disk4and restrained by an outer circumference part of the rotor disk4(refer toFIG. 2).

The cooling channel50is formed inside the above-constituted turbine blade3.

As shown inFIG. 3andFIG. 4, the cooling channel50is provided with a leading edge-side channel51and serpentine channels52,53which are arranged sequentially from the leading edge11to the trailing edge12.

The leading edge-side channel51extends in the blade span direction from the blade basement30to the tip end15of the airfoil body10on the leading edge11side to which the leading edge-side channel51is closer than the serpentine channels52,53. The leading edge-side channel51is connected at an upstream end thereof to an introduction channel51icommunicatively connected to a radial hole7(refer toFIG. 2). Further, the leading edge-side channel51is communicatively connected to a plurality of cooling holes51h, each of which penetrates through a blade surface13of the leading edge11and a channel wall face of the leading edge-side channel51.

The leading edge-side channel51cools the leading edge11by using the compressed air c which flows from the radial hole7to the tip end15of the airfoil body10and allows the compressed air c to flow through the cooling hole51h, thereby subjecting the leading edge11to shower-head film cooling.

As shown inFIG. 3andFIG. 4, the serpentine channel52is formed in a meandering manner between the leading edge-side channel51and the serpentine channel53so as to be positioned on the center line Q of the blade profile (refer toFIG. 3) and connected with return channels52d,52eat which three main channels52ato52c, each of which extends in the blade span direction, are formed in a U shape.

Each of the three main channels52ato52cextends from the tip end15of the airfoil body10to the base end14of the airfoil body10. Also, these main channels are installed in parallel from the trailing edge12side to the leading edge11side in the above-described order. Then, an outer circumference end of the main channel52ais connected to an outer circumference end of the main channel52bat a return channel52d, and an inner circumference end of the main channel52bis connected to an inner circumference end of the main channel52cat a return channel52e. Further, the main channel52ais connected at the upstream end thereof to the introduction channel52icommunicatively connected to the radial hole7(refer toFIG. 2). Still further, the main channel52cis communicatively connected to a plurality of cooling holes52h, each of which penetrates through the blade surface13and a channel inner wall of the main channel52c.

In the serpentine channel52, the compressed air c flows from the introduction channel52iinto the main channel52a, thereafter, passing through the main channel52a, turning at 180° at the return channel52d, flowing into the main channel52b, passing through the main channel52b, turning at 180° at the return channel52e, and flowing into the main channel52c. In the above-described course, part of the compressed air c which circulates inside the main channel52cflows out obliquely from the cooling hole52has shown inFIG. 4, thereby subjecting the blade surface13to film cooling.

As shown inFIG. 3andFIG. 4, the serpentine channel53is formed in a meandering manner on the trailing edge12side so as to be positioned on the center line Q of the blade profile (refer toFIG. 3), and three main channels53ato53cextending in the blade span direction are connected at the return channels53d,53eformed in a U shape.

Each of the three main channels53ato53cextends from the tip end15of the airfoil body10to the base end14of the airfoil body10and installed in parallel from the leading edge11side to the trailing edge12side in the above-described order. Then, an outer circumference end of the main channel53ais connected to an outer circumference end of the main channel53bat the return channel53d, and an inner circumference end of the main channel53bis connected to an inner circumference end of the main channel53cat the return channel53e. Further, the main channel53ais connected at the upstream end thereof to introduction channels53i,53jcommunicatively connected to the radial hole7. Still further, as shown inFIG. 4, the main channel53cis communicatively connected to a plurality of cooling holes53h, each of which penetrates through the blade surface13and the main channel53c.

In the serpentine channel53, the compressed air c flows from the introduction channels53i,53jinto the main channel53a, thereafter, passing through the main channel53a, turning at 180° at the return channel53d, flowing into the main channel53b, passing through the main channel53b, turning at 180° at the return channel53e, and flowing into the main channel53c. In the above-described course, part of the compressed air c which circulates inside the main channel53cflows out to the surface through the cooling hole53hto conduct film cooling, and the remainder of the compressed air c cools a pin fin at the trailing edge end when flowing out from the trailing edge12.

The compressed air c passes from the introduction channels53i,53jthrough the main channels53a,53b,53cand the return channels53d,53eat the serpentine channel53and is finally discharged into combustion gas. In the course of flowing through the channels by returning to the return channels53d,53e, the compressed air c undergoes pressure loss and thereby gradually decreases in pressure.

In the turbine blade3which has been described as being constituted as above, the above-described return channels52e,53eare formed between the base end14of the airfoil body10and the platform20at the fillet portion40in the turbine radial direction. The return channels52e,53eare formed along the fillet portion40on a cross section intersecting with the center line Q of the blade profile.

The return channels52e,53eare similar in constitution. Thus, in the following description, the return channel53ewill be described with description of the return channel52eomitted.

FIG. 5is an enlarged view which shows a major part inFIG. 3(enlarged view showing a major part IV inFIG. 3),FIG. 6is a sectional view taken along line V to V inFIG. 5, andFIG. 7is a sectional view taken along line VI to VI inFIG. 5.

As shown inFIG. 6andFIG. 7, the return channel53eis formed on a cross section orthogonal to the center line Q of the blade profile (hereinafter, simply referred to as an intersecting cross section of the center line Q of the blade profile) so as to be greater in length L in the profile thickness direction than the main channel53band the main channel53c(L1>L2). Further, as shown inFIG. 6, the return channel53eis formed on the intersecting cross section of the center line Q of the blade profile so as to assume a flat shape in which the length in the blade span direction is smaller than the length L1in the profile thickness direction.

As shown inFIG. 5, the cooling face55is formed along an outer surface40aof the fillet portion40on both sides in the profile thickness direction at the channel inner wall face (inner surface) of the return channel53e.

The cooling face55extends along the center line Q of the blade profile and is also formed, as shown inFIG. 6andFIG. 7, in a circular arc shape along the outer surface40aof the fillet portion40formed in a quarter-round arc shape at the intersecting cross section. More specifically, as shown inFIG. 7, the cooling face55is provided with a cross section contour formed in a quarter-round arc shape along the fillet portion40at the intersecting cross section of the center line Q of the blade profile at a position in the front and back of the partition wall54which separates the main channel53bfrom the main channel53cand, as shown inFIG. 6, also provided with a cross section contour which is formed in a small circular-arc shape along the fillet portion40at a position which includes the partition wall54(intermediate portion53e1).

As shown inFIG. 6andFIG. 7, it is desirable that a channel inner wall60which forms a channel width (L1) of the return channel53eis increased in width substantially up to the proximity of an outer edge40bof the fillet portion40. At the center of the center line Q of the blade profile, the channel width (L1) is increased in width up to the proximity of the outer edge40bof the fillet portion40toward both sides in the profile thickness direction, by which the cooling face55is enlarged along the fillet portion40in the profile thickness direction. The outer edge40bof the fillet portion40refers to a boundary line between the fillet portion40and the surface of the platform20.

Further, where a distance between the outer surface of the turbine blade3and the channel inner wall face of the cooling channel50in a normal line direction of the outer surface of the turbine blade3is defined as a blade wall thickness d, a blade wall thickness d1between the cooling face55and the fillet portion40is formed so as to be substantially the same at respective sites.

Still further, the blade wall thickness d1between the cooling face55and the fillet portion40is formed so as to be substantially equal to a blade wall thickness d2of the main channel53band that of the main channel53c.

In addition, on the blade basement30side of the channel inner wall face of the return channel53e, a projection part56is formed which projects in the normal line direction of the channel inner wall face at the center in the profile thickness direction. As shown inFIG. 6, the projection part56is formed in a trapezoidal shape on each intersecting cross section of the center line Q of the blade profile. As shown inFIG. 5, the projection part56projects in a gradually increasing amount moving from the main channel53bside to the intermediate portion53e1on the return channel53eand, thereafter, projects in a gradually decreasing amount moving from the intermediate portion53e1to the main channel53c.

In addition to guiding the compressed air c as will be described below, the projection part56also adjusts a channel cross section of the return channel53e. In the present embodiment, the projection part56is provided due to the channel width (L1) of the return channel53ebeing, as described above, increased in width in the profile thickness direction, by which the compressed air c flows at the center of the return channel53eand prevents the occurrence of stagnation of flow in the proximity of the channel inner wall60of the intermediate portion53e1, securing uniform flow of the compressed air c of the channel cross section.

Next, a description will be given of actions of the turbine blade3in the above-constituted gas turbine GT.

As described above, when flowing into the serpentine channel53via the introduction channels53i,53j, the compressed air c passes through the main channel53a, turns at 180° at the return channel53d, flows into the main channel53b, passes through the main channel53b, turns at 180° at the return channel53e, and flows into the main channel53c.

Since the cooling face55is formed so as to be substantially uniform in blade wall thickness d1, the compressed air c passing through the return channel53eremoves heat uniformly at sites of the fillet portion40.

That is, as shown inFIG. 5, the compressed air c gives uniform cooling to the sites of the fillet portion40over a direction at which the center line Q of the blade profile extends. Further, on each intersecting cross section of the center line Q of the blade profile (refer toFIG. 6andFIG. 7), the channel width (L1) is increased in width up to the proximity of the outer edge of the fillet portion40, thus making it possible to cool the sites uniformly over the blade span direction of the fillet portion40and in the profile thickness direction thereof.

Further, the compressed air c which has reached the upstream end of the projection part56on the return channel53eis gradually guided to both sides in the profile thickness direction, because the projection part56projects in an increasing amount moving to the intermediate portion53e1. In the return channel53e, the compressed air c guided to both sides in the profile thickness direction carries out cooling mainly via the cooling face55by removing heat from the fillet portion40.

Still further, the projection part56is adjusted for projection amount, thus making it possible to improve uneven flowing of the compressed air c on the channel cross section of the return channel53e.

As so far described, according to the turbine blade3used in the gas turbine GT, the return channel53eis formed so as to run along the fillet portion40on the intersecting cross section of the center line Q of the blade profile. Therefore, the blade wall thickness d is uniform from the return channel53eto the outer surface40aof the fillet portion40. Thereby, it is possible to prevent formation of a site which is increased in blade wall thickness d and to give uniform and sufficient cooling to the fillet portion40. It is, therefore, possible to suppress oxidation-caused thinning and fatigue of the turbine blade3. A similar effect can be obtained also in the return channel52e.

Further, since the cooling face55is provided, it is possible to give sufficient cooling to the fillet portion40that is opposite to the cooling face55.

Further, since the projection part56is provided, the compressed air c is guided to both sides in the profile thickness direction. Thereby, it is possible to give sufficient cooling to the fillet portion40positioned on both sides of the return channel53ein the profile thickness direction.

Further, a distance of the cooling face55to the outer surface40aof the fillet portion40is formed so as to be substantially the same distance from the channel inner wall face of the main channel53bto the outer surface of the airfoil body10and a distance from the channel inner wall face of the main channel53cto the outer surface of the airfoil body10. It is, therefore, possible to carry out uniform cooling between the airfoil body10and the fillet portion40.

Still further, the cooling face55extends along the center line Q of the blade profile. Therefore, the fillet portion40can be uniformly and sufficiently cooled over a wide range along the center line Q of the blade profile.

In addition, the turbine blade3is provided, by which the turbine blade3can be enhanced in cooling effect to improve the reliability.

Second Embodiment

Hereinafter, a description will be given of the second embodiment of the present invention with reference to drawings. In the following description and the drawings used in the description, constituents similar to those which have been already described will be given the same reference numerals, with overlapping description being omitted.

FIG. 8is a sectional view which is cut along the center line of the profile of a turbine vane2of the present embodiment.

In the above-described first embodiment, the present invention is applied to the turbine blade3. However, in the present embodiment, the present invention is applied to the turbine vane2of a turbine T (refer toFIG. 2).

As shown inFIG. 8, an outer shroud (end wall)2bof the turbine vane2is connected to a base end58of an airfoil body2a(outer side in the turbine radial direction, end part) and an inner shroud (end wall)2cof the turbine vane2is connected to a tip end59of the airfoil body2a(inside in the turbine radial direction, end part).

The base end58of the airfoil body2ais smoothly connected to the outer shroud2bat a fillet portion41, and the tip end59of the airfoil body2ais smoothly connected to the inner shroud2cat a fillet portion42.

A serpentine channel (cooling channel)57is formed inside the turbine vane2.

The serpentine channel57is formed in a meandering manner between the leading edge11and the trailing edge12, as shown inFIG. 8, so as to be positioned on the center line Q of the blade profile, as shown inFIG. 4. Five main channels57ato57c,57f,57gextending in the blade span direction are connected at return channels57d(57dA,57dB),57e(57eA,57eB), each of which is formed in a U shape.

Each of the five main channels57ato57c,57f,57gextends from the base end58side of the airfoil body2ato the leading end59side of the airfoil body2a, and they are installed in parallel from the leading edge11side to the trailing edge12side in the above-described order. Then, an inner circumference end of the main channel57ais connected to an inner circumference end of the main channel57bat the return channel57eA, and an outer circumference end of the main channel57bis connected to an outer circumference end of the main channel57cat the return channel57dA. Further, an inner circumference end of the main channel57cis connected to an inner circumference end of the main channel57fat the return channel57eB, and an outer circumference end of the main channel57fis connected to an outer circumference end of the main channel57gat the return channel57dB.

Further, the upstream end of the main channel57ais communicatively connected to a blade ring feed hole70to which compressed air c is supplied. The main channel57gwhich is communicatively connected to a cooling hole53mof the trailing edge12gives convection cooling to an end part of the trailing edge and is discharged, thereafter, into combustion gas.

The above-described return channels57d(57dA,57dB),57e(57eA,57eB) are formed so as to run along the fillet portions41,42on a cross section intersecting with the center line Q of the profile of the airfoil body2a.

More specifically, a cooling face55is formed on each of the return channels57d(57dA,57dB) and57e(57eA,57eB) so as to run along the outer surfaces of the fillet portions41,42. Further, at the center of the return channel57din the profile thickness direction, a projection part57d1is formed which projects in a normal line direction of the channel inner wall face. At the center of the return channel57ein the profile thickness direction, a projection part57e1is formed which projects in the normal line direction of the channel inner wall face.

According to the present embodiment, in addition to obtaining the above-described major effects of the first embodiment, it is possible to sufficiently cool the fillet portions41,42of the turbine vane2.

Action procedures or various shapes and combinations or the like of constituents shown in the above embodiments are only examples and may be changed in various ways based on design requirements and the like, within a scope not departing from the gist of the present invention.

For example, in the above-described first embodiment, the cooling face55has been formed on the intersecting cross section of the center line Q of the blade profile so as to give a circular-arc cross section contour and arranged so as to run along the outer surface of the fillet portion40. However, the cooling face55may be formed so as to give a linear cross section contour which extends obliquely in a tangent line direction of the outer surface of the fillet portion40and arranged so as to run along the outer surface of the fillet portion40. This is also the same in the second embodiment.

Further, in the above-described first embodiment, the present invention has been applied to the return channels52e,53e. However, the present invention may be applied to only one of the return channels52e,53e. Still further, the present invention may be applied to at least one of the return channels the same as the return channels52e,53e. This is also the same in the second embodiment.

Third Embodiment

Hereinafter, a description will be given of the third embodiment of the present invention with reference to drawings. In the following description and the drawings used in the description, constituents similar to those which have been already described will be given the same reference numerals, with overlapping descriptions being omitted.

The present embodiment is an example in which a cooling hole is provided at a return part of the above-described serpentine channel and applicable to both the first embodiment and the second embodiment.FIG. 9is a blade sectional view which shows a turbine blade similar to that used in the first embodiment given inFIG. 3. The example shows a case where the cooling hole is provided on the partition wall of the upstream side at an inlet of each return part in a direction at which compressed air c flows.FIG. 10is an enlarged view which shows a major part ofFIG. 9(an enlarged view which shows the major part VII inFIG. 9).FIG. 11shows a cross section taken along line VIII to VIII when viewed in the leading edge direction around a return channel53egiven inFIG. 10.FIG. 12shows a cross section taken along line IX to IX around the return channel53einFIG. 11. With reference toFIG. 9toFIG. 12, the present embodiment will be described hereinafter by exemplifying the return channel53eof a serpentine channel53.

Each ofFIG. 9andFIG. 10shows an example where a cooling hole53kis installed in the proximity of a base end14of the serpentine channel53, that is, on a partition wall54which separates between main channels53aand53band arranged so as to incline to the bottom of the return channel53e.

Further, as shown inFIG. 11, a channel cross section formed with channel inner walls60on both sides of the return channel53ein the profile thickness direction is increased in channel width (L1) up to the proximity of the outer edge40bof a fillet portion40when viewed from the cross section in the blade span direction, and two cooling holes53kare arranged from the suction side and the pressure side of the main channel53bin such a manner that the compressed air c is blown out toward the channel inner wall60which has been increased in width on both sides thereof. Regarding the cooling hole53k, one end thereof is communicatively connected to an upstream side channel of the main channel53awhich is on the upstream side of the serpentine channel53, while the other end is opened to the main channel53bwhich is on the downstream side.

As shown inFIG. 12, the return channel53eis formed so as to have an enlarged portion61which is expanded to the suction side and the pressure side in the profile thickness direction to a greater extent that the main channels53b,53c. The cooling holes53kinstalled on the suction side and the pressure side are inclined to the center line Q of the blade profile and pointed at a direction at which the compressed air c is in contact with the channel inner wall60.

As described above, the compressed air c flowing through the serpentine channel is decreased in pressure due to pressure loss occurring during which the compressed air c flows through the channel. For example, in the serpentine channel53, the compressed air c which is lower in temperature and which has flowed from introduction channels53i,53jinto the main channel53aflows from the base end side14to the leading end15, turns at 180° to return at the return part53d, flows further down to the base end14and reaches the return part53e, during which the compressed air c is decreased in pressure due to pressure loss inside the channels. That is, there is a certain difference in pressure due to pressure loss between the base end14on the upstream side of the main channel53aand the proximity of the inlet of the return channel53eon the downstream side of the main channel53b. Therefore, due to a difference in pressure between the inlet side of the cooling hole53kand the outlet side thereof, part of the compressed air c flowing through the main channel53aflows into the cooling hole53kand is blown into the return channel53e.

According to the present embodiment, there is a possibility that the compressed air c will stay stagnant at the enlarged portion61which is expanded in the profile thickness direction on the channel cross section of the return channel53e. However, the cooling hole53kof the present embodiment is provided, by which art of the compressed air c blows out obliquely downward toward the channel inner wall60of the return channel53e(radially inward direction). Therefore, the stagnant compressed air which remains at the enlarged portion61is purged to the downstream side and the compressed air flow flowing in the proximity of the cooling face55of the return channel53eis exchanged to improve the cooling capability of the cooling face55.

The cooling hole52kis installed on the partition wall54between the main channels52a,52bin the return channel52eof the serpentine channel52in the same manner, by which a constitution similar to that of the return channel53eis applicable.

Further, the present invention has been applied to the turbine blade3of the above-described turbine T and the turbine vane2of the turbine T. However, the present invention may be applicable to turbine blades and turbine vanes of various types of rotary machines (for example, turbine blades and turbine vanes of the compressor C (refer toFIG. 1)).

INDUSTRIAL APPLICABILITY

The present invention relates to an airfoil member which is provided with an airfoil body, an end wall which is installed at an end part of the airfoil body in a blade span direction and extends so as to intersect in the blade span direction, a fillet portion which smoothly connects the end part of the airfoil body with the end wall, and a cooling channel which allows a cooling medium to circulate inside the airfoil body and the end wall and in which two main channels extending along the blade span direction are connected so as to bend in a folding manner at a return channel formed on the end wall side. The return channel is formed so as to run along the fillet portion on a cross section intersecting with a center line of a profile of the airfoil body and also formed in such a manner that the width thereof in the profile thickness direction is greater than the width of the main channel in the profile thickness direction. The present invention is able to give uniform and sufficient cooling to the fillet portion.

DESCRIPTION OF REFERENCE NUMERALS