JP5047078B2 - gas turbine - Google Patents

Description

  The present invention relates to a turbine blade and a gas turbine particularly suitable for use in a turbine rotor blade of a gas turbine.

In general, a moving blade of a gas turbine is exposed to a high-temperature gas, and is therefore formed of a material having heat resistance to ensure the strength of the moving blade. Furthermore, convection cooling with a turbulator provided with a turbulator in the cooling chamber (cavity) formed in the blade and a shower head cooling hole for blowing cooling air to the outside are provided at the leading edge of the blade with high heat load. Cooling by various methods such as film cooling is performed (for example, see Patent Document 1).
JP-T 8-5055917

  However, due to the increase in the temperature of the gas flowing around the moving blades as the efficiency of gas turbines increases in recent years, it is expected that the conventional cooling method for the leading edge of the moving blades will be insufficient. There is a need for improved capabilities.

  For example, the area where the shower head cooling hole is provided at the leading edge of the moving blade is widened, and the shower head cooling hole is formed up to the hub side at the leading edge of the moving blade, i.e., at the radially inner end. A method for improvement is conceivable.

In particular, the fillet portion (connecting portion) between the leading edge and the hub has a high thermal load and is a region where the thermal barrier coating (TBC) formed on the surface of the moving blade is likely to be peeled off. It is desired to keep the metal temperature of the rotor blade (the temperature of the metal part) within a predetermined range by convection cooling with cooling air flowing through the holes.
In other words, in the region where the TBC is peeled off, the high temperature gas is in direct contact with the metal portion of the blade, and the metal temperature is likely to rise, and the blade strength is likely to be lowered. In order to suppress the decrease in the blade strength, it is desired to form a shower head cooling hole in the vicinity of the fillet portion described above.

  However, since a plurality of through holes are formed in the region where the shower head cooling hole is formed, the strength of the moving blade may be reduced as compared with a region where the shower head cooling hole is not formed. In particular, the vicinity of the hub side end portion (fillet portion) at the leading edge needs to ensure the strength against vibration of the moving blade, so the shower head cooling hole is formed near the hub side end portion at the leading edge. There was a problem that it was difficult.

  On the other hand, a so-called high load blade, that is, a moving blade with a large turning angle, has a problem that it is difficult to form a shower head cooling hole on the leading edge because thermal stress tends to concentrate on the leading edge.

  Specifically, a moving blade with a large turning angle has a length of a pressure surface (abdominal side) curved in a concave shape and a length of a suction surface (back side surface) curved in a convex shape in the blade cross-sectional view. The difference between is large. When such a moving blade is heated by high-temperature gas, the difference in the amount of thermal elongation between the pressure surface and the suction surface becomes larger than that of a moving blade having a small turning angle. Therefore, distortion due to the difference in the amount of thermal expansion, that is, thermal stress tends to concentrate on the region on the pressure surface side at the leading edge.

  When a through hole such as a shower head cooling hole is formed in a region where stress is concentrated in this way, defects such as cracks are likely to occur, and it is difficult to form the shower head cooling hole on the front edge.

The present invention has been made to solve the above-described problem, and an object of the present invention is to provide a gas turbine including a turbine blade that can improve the reliability of the blade that contacts the high-temperature gas.

In order to achieve the above object, the present invention provides the following means.
The gas turbine of the present invention includes a pressure surface curved in a concave shape, a suction pressure surface curved in a convex shape, and an airfoil portion having a front edge and a rear edge to which the pressure surface and an end of the suction surface are connected, A plurality of plate portions extending between the pressure surface and the suction surface to divide the inside of the airfoil portion and to form a plurality of cooling passages through which a cooling fluid for cooling the airfoil portion flows; Among the plurality of plate portions, at least the plate portion closest to the front edge is tapered from the suction surface side toward the pressure surface side, and the thickness thereof decreases in a tapered shape. The connecting portion with the positive pressure surface is provided with a turbine blade having lower rigidity than the connecting portion with the negative pressure surface.

According to the present invention, the pressure surface restraint by the plate portion closest to the front edge is weaker than the suction surface restraint, so that the pressure surface is easily deformed. Therefore, when the turbine blades are heated by high-temperature gas, etc., even if there is a difference between the amount of thermal elongation on the pressure surface and the amount of heat elongation on the suction surface, the pressure surface is easily deformed, that is, the stress at the leading edge. Concentration is eased.
Further, by reducing the thickness of the plate portion in a tapered shape from the suction surface toward the pressure surface, the connection portion of the plate portion with the pressure surface is surely lower in rigidity than the connection portion with the suction surface.
And these can improve the reliability of the turbine blade in contact with the high-temperature gas.

In the above invention, the distance from the front edge on the pressure surface to the connecting portion with the plate portion closest to the front edge, and the plate portion closest to the front edge from the front edge on the suction surface It is desirable that the distance to the connection part is substantially equal.

  According to the present invention, when the turbine blade is heated, the amount of thermal elongation from the leading edge to the connection portion on the pressure surface and the amount of thermal elongation from the leading edge to the connection portion on the suction surface become substantially equal. Stress concentration at the edge is relaxed.

According to the gas turbine of the present invention, restraint of the pressure side by the plate portion of the turbine blade weaker than constraint on the suction surface, since the positive pressure surface is easily deformed That elongation, since the stress concentration at the leading edge is reduced, high temperature There is an effect that the reliability of the blade in contact with the gas can be improved.

According to the gas turbine of the present invention, when the turbine blade is heated, and the thermal extension amount from the leading edge to the connecting portion of the pressure side, it becomes substantially equal to the thermal extension amount from the leading edge to the connecting portion of the suction surface Therefore, there is an effect that it is possible to improve the reliability of the blade in contact with the high temperature gas.

A turbine rotor blade of a gas turbine according to an embodiment of the present invention will be described with reference to FIGS. 1 and 2.
FIG. 1 is a cross-sectional view for explaining the blade shape and the like of the turbine rotor blade according to the present embodiment. FIG. 2 is a perspective view illustrating a schematic configuration of the turbine rotor blade of FIG.
In the present embodiment, the turbine blade according to the invention of the present application is described as being applied to a first stage blade in a turbine section of a gas turbine. However, the present invention is not limited to a single stage blade and is applied to other blades. There is no particular limitation.
Furthermore, a well-known structure can be used as a gas turbine and a turbine part, and it does not specifically limit.

  As shown in FIG. 1 and FIG. 2, the turbine rotor blade (turbine blade) 1 is provided with a blade attachment portion 2, a platform 3, and an airfoil portion 4.

As shown in FIG. 2, the blade attachment portion 2 extends radially inward (lower side in FIG. 2) and is detachably fitted to a rotor (not shown) of the turbine. It is a part used when attaching the moving blade 1 to a rotor.
In addition, as a shape of the wing | blade attachment part 2, well-known shapes, such as the shape which provided the so-called Christmas tree shape in the edge part of the shank 21, can be employ | adopted and it does not specifically limit.

  As shown in FIG. 2, the shank 21 is a member formed in a substantially prismatic shape extending from the platform 3 in the radial direction, that is, toward the rotor. In other words, it is a member arranged between the platform 3 and the rotor.

  As shown in FIG. 2, the platform 3 is a substantially plate-like member that is disposed at the radially outer end of the shank 21.

As shown in FIG. 2, the airfoil 4 is a member that extends from the outer peripheral surface of the platform 3 (the upper surface in FIG. 2) toward the radially outer side (the upper side in FIG. 2).
In the present embodiment, description will be made by applying to a case where the turbine rotor blade 1 is a so-called high-load blade, that is, a rotor blade having a large turning angle.
Here, as a moving blade having a large turning angle, for example, a moving blade having a turning angle of about 110 ° or more can be exemplified.

  As shown in FIG. 1, the airfoil portion 4 includes a positive pressure surface 41 that is a concave curved surface, a negative pressure surface 42 that is a convex curved surface, and an upstream side of the high-temperature fluid flowing through the turbine portion (left side in FIG. 1). A front edge 43 that is an end of the rear edge 44 and a rear edge 44 that is an end on the downstream side (the right side in FIG. 1) are provided.

  Inside the airfoil portion 4, a plurality of ribs (plate portions) 45 extending between the pressure surface 41 and the suction surface 42 to divide the inside of the airfoil portion 4, and cooling for cooling the airfoil portion 4. And a cooling passage 46 through which air flows.

As shown in FIG. 1, the rib 45 is a substantially plate-like member that extends in a direction intersecting the positive pressure surface 41 and the negative pressure surface 42, and a plurality of ribs 45 are provided from the front edge 43 toward the rear edge 44 at intervals. It is a member.
Among the plurality of ribs 45, the first rib 45 </ b> A disposed on the most front edge 43 side is a substantially plate shape formed such that the plate thickness decreases in a tapered shape from the negative pressure surface 42 toward the positive pressure surface 41. It is a member.

  Furthermore, the distance L1 from the connection portion 47A between the first rib 45A and the positive pressure surface 41 to the front edge 43, the distance L2 from the connection portion 47B between the first rib 45A and the negative pressure surface 42 to the front edge 43, Are approximately equal in length.

  The cooling passage 46 is a space formed inside the airfoil portion 4 and is a space partitioned by the ribs 45. Cooling air (cooling fluid) that cools the pressure surface 41, the suction surface 42, the front edge 43, the rear edge 44, and the like of the airfoil 4 is supplied to the inside of the cooling passage 46.

Next, relaxation of stress concentration at the leading edge 43 of the turbine rotor blade 1 having the above-described configuration will be described.
While the gas turbine having the turbine blade 1 is in operation, the high-temperature gas flows around the turbine blade 1 and the temperature of the turbine blade 1 rises.

Then, the pressure surface 41 and the suction surface 42 of the turbine rotor blade 1 are each increased in length by thermal expansion, as shown in FIG. The amount of thermal expansion of the pressure surface 41 and the suction surface 42 is affected by the length of the pressure surface 41 and the suction surface 42 because the constituent materials are the same, that is, the linear expansion coefficient is the same.
In the case of the turbine rotor blade 1 having a large turning angle as in the present embodiment, the suction surface 42 is longer than the pressure surface 41 and there is a difference in the length thereof, so that the heat of the suction surface 42 is higher than the pressure surface 41. The elongation amount is long and a difference occurs in the thermal elongation amount.

The first rib 45A is the width dimension of the connecting portion 47A of the positive pressure surface 41 side is thin rather than the connecting portion 47B of the suction surface 42 side, the rigidity of the connection portion 47A side of the connecting portion 47B side It is low Kuna' than rigidity. Therefore, the first rib 45 </ b> A restrains the thermal expansion of the suction surface 42 and allows the thermal expansion of the pressure surface 41 in comparison.
Therefore, the difference in thermal elongation between the positive pressure surface 41 and the negative pressure surface 42 in the vicinity of the front edge 43 where the positive pressure surface 41 and the negative pressure surface 42 are joined is reduced.

  Further, since the distance L1 from the connecting portion 47A to the front edge 43 and the distance L2 from the connecting portion 47B to the front edge 43 are substantially equal to each other, before the positive pressure surface 41 and the negative pressure surface 42 are joined together. The difference in thermal elongation between the pressure surface 41 and the suction surface 42 in the vicinity of the edge 43 is further reduced.

  As described above, since the difference in thermal elongation near the front edge 43 is reduced, stress concentration caused by the difference in thermal elongation is reduced.

According to the above configuration, by reducing the thickness of the first rib 45A from the negative pressure surface 42 side in a tapered shape toward the positive pressure surface 41 side, restraining of the pressure surface 41 by the first rib 45A is suction surface 42 Therefore, the positive pressure surface 41 is easily deformed. Therefore, when the turbine rotor blade 1 is heated by a high-temperature gas or the like, even if there is a difference between the thermal elongation amount at the pressure surface 41 and the thermal elongation amount at the suction surface 42, the pressure surface 41 is heated at the suction surface 42 . Since the deformation easily follows the elongation, the stress concentration at the leading edge 43 is alleviated.

As a result, since the structural strength at the base of the leading edge 43 can be afforded, a through hole for supplying cooling air from the cooling passage 46 inside the airfoil portion 4, a so-called shower head cooling hole is provided at the root of the leading edge 43 . Can be formed. Thereby, the cooling performance of the turbine rotor blade 1 can be improved, and in particular, the reliability in strength against a high-temperature working fluid can be improved.

  Furthermore, since the distance L1 from the connecting portion 47A to the front edge 43 and the distance L2 from the connecting portion 47B to the front edge 43 are substantially equal in length, when the turbine rotor blade 1 is heated, the pressure surface 41, the amount of thermal elongation from the leading edge 43 to the connecting portion 47A and the amount of thermal elongation from the leading edge 43 to the connecting portion 47B of the suction surface 42 are substantially equal, so that the stress concentration on the pressure surface 41 side of the leading edge 43 Is alleviated.

It is a sectional view explaining the airfoil etc. of the turbine rotor blade concerning one embodiment of the present invention. It is a perspective view explaining schematic structure of the turbine rotor blade of FIG.

Explanation of symbols

1 Turbine blade (turbine blade)
4 Airfoil part 41 Positive pressure surface 42 Negative pressure surface 43 Front edge 44 Rear edge 45 Rib (plate part)
46 Cooling passage 47A, 47B Connection L1, L2 distance

Claims (2)

An airfoil having a pressure surface curved in a concave shape, a suction surface curved in a convex shape, and a leading edge and a trailing edge to which the pressure surface and an end of the suction surface are connected;
A plurality of plate portions extending between the pressure surface and the suction surface to divide the inside of the airfoil portion and form a plurality of cooling passages through which a cooling fluid for cooling the airfoil portion flows;
Is provided,
Of the plurality of plate portions, at least the plate portion closest to the front edge is tapered from the suction surface side toward the pressure surface side, and the thickness thereof is tapered.
Connecting portions between the pressure surface of the plate portion, the gas turbine rigidity than connection portion between the negative pressure surface, characterized in that the lower plate turbine blades is provided. A distance from the front edge of the positive pressure surface to the connecting portion with the plate portion closest to the front edge ;
The gas turbine according to claim 1, wherein a distance from the front edge on the suction surface to a connection portion with the plate portion closest to the front edge is substantially equal.

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