JP3040660B2 - Gas Turbine Blade Platform Cooling Mechanism - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a cooling mechanism having a thermal stress relaxation structure for a platform in a gas turbine blade.

[0002]

2. Description of the Related Art FIG. 4 is a perspective sectional view of a typical conventional gas turbine hollow rotor blade. In the figure, cooling air flowing from the bottom of the blade root 11 flows in the direction of the arrow to cool the moving blade. That is, the cooling air flowing from the leading edge side 12A flows through the meandering flow path having the fins 13 to cool the blade, flows out from the hole A at the blade top where the chip thinning 14 is provided, and joins the main gas flow. . The cooling air flowing in from the trailing edge side 12B flows in the direction of the arrow through the cooling passage provided with the fins 13 and cools the trailing edge of the blade by the pin fins 15, and then flows out of the blade through the hole or slit B to flow out the main gas. To join. The blades having such a high cooling structure are implanted in the disk 17 with the platforms 16 adjacent to each other in the circumferential direction.

[0003]

However, such a hollow moving blade has an advanced cooling structure for internally cooling a blade root and a blade, but has a shape protruding from a cooling portion. The platform is almost uncooled and does not have a sufficient cooling structure. This requires cooling the hot gas turbine blade platform,
There is a need to reduce the thermal stress generated when cooling this effectively. Incidentally, a temperature difference of 1000 ° C. or more between the gas and the seal air occurs between the gas flow path side of the platform and the rotor side below the platform. SUMMARY OF THE INVENTION The present invention has been made to solve the above-mentioned conventional problems, and has a highly reliable gas turbine operation which can effectively cool a platform surface and reduce thermal stress caused by a temperature difference between an upper surface and a lower surface of the platform. It is intended to provide a wing platform cooling mechanism.

Therefore, the present invention provides a cooling mechanism for a platform of a gas turbine blade, in which a plurality of seal air passage holes penetrating in a radial direction are provided on the platform, and the upper surface ( A shape film outlet is provided in the upper surface of the platform, and a slit extending from the seal air passage hole to the circumferential end face of the platform is provided, and this is used as a means for solving the problem. It is.

[0005]

According to the present invention, a seal air flow passage hole penetrating in the radial direction of the platform is provided, and a shape film outlet is provided on the upper surface of the air flow passage hole. Flows out to the upper surface of the platform through the seal air hole in the direction, but at the time of outflow, it flows so as to crawl while spreading on the upper surface of the platform by the shape film outlet provided on the upper surface of the air hole (the upper surface of the platform). Cool the upper surface. Further, by providing the slit extending from the seal air passage hole to the circumferential end surface of the platform, thermal stress caused by the temperature difference between the upper and lower sides of the platform is reduced by the expansion and contraction of the slit, and the platform is released from tension.

[0006]

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a perspective view of a blade platform according to an embodiment of the present invention; FIG. In the figure, reference numeral 1 denotes a platform on which a plurality of seal air passage holes 2 penetrating in the radial direction are formed. Seal air passage hole 2
, That is, on the upper surface of the platform 1, a shape film outlet 3 is provided. A slit 4 is cut from the seal air passage hole 2 to the circumferential end surface of the platform 1 with respect to the entire thickness of the platform 1. Platform 1
Of the seal air flowing through the lower surface A of the seal air passage hole 2
Flows in the radial direction of the arrow and flows out to the upper surface of the platform 1, and flows in the direction of the thick arrow to cool the upper surface of the platform 1 so as to crawl while spreading over the upper surface of the platform 1 by the shape film outlet 3. The shape film outlet 3 may be oriented in the axial direction as shown in FIG. 1, or may be formed so as to extend in the circumferential direction (from the high pressure ventral side to the low pressure rear side). It may be intermediate. The slit 4 has a temperature difference between the upper and lower surfaces of the platform 1 and the upper and lower surfaces and a temperature difference between the profile portion and the shroud end (the profile portion is cooled and the circumferential end of the shroud is cooled). The thermal stresses caused by the above (absent) are relieved by the expansion and contraction of the slit 4, and the platform 1 is released from tension. FIG. 2 shows a plan view of the slit portion of the platform, and FIG. 3 shows a side view thereof.

[0007]

As described in detail above, according to the present invention, the platform is formed such that the sealing air is blown out from the radial sealing air passage hole through the shape film outlet and crawls on the upper surface of the platform. It spreads and flows, effectively cooling the upper surface of the platform without waste. Further, a slit formed in the circumferential end face of the platform from the seal air passage hole reduces thermal stress caused by a temperature difference between the upper surface and the lower surface of the platform, and releases the platform from tension. Therefore, the effect of further improving the reliability of the rotor blade and opening the way to a higher temperature is great.

[Brief description of the drawings]

FIG. 1 is a perspective view of a platform in a cooling mechanism according to an embodiment of the present invention.

FIG. 2 is a plan view of the platform slit section.

FIG. 3 is a front view of the platform slit section.

FIG. 4 is a perspective view showing a conventional typical gas turbine hollow rotor blade.

[Explanation of symbols]

 DESCRIPTION OF SYMBOLS 1 Platform 2 Seal air channel hole 3 Shape film outlet 4 Slit

Claims (1)

(57) [Claims] In a cooling mechanism for a platform of a gas turbine rotor blade, a plurality of seal air passage holes penetrating in a radial direction are provided on the platform, and an upper surface (platform upper surface) of each of the air passage holes is provided. A cooling mechanism for a gas turbine blade platform, comprising a shape film outlet, and a slit extending from the seal air passage hole to a circumferential end face of the platform.