JPH05156901A - Gas turbine cooling stationary blade - Google Patents

Abstract

PURPOSE:To reduce a metal temperature of a trailing edge which is subjected to a high heat load and a thermal stress so as to increase reliability and service life of a gas turbine cooling stationary blade by increasing an effect of cooling the trailing edge. CONSTITUTION:A cavity in a gas turbine stationary blade 1 is divided into a trailing edge A and other front sections by a bulk head 9, the front sections are cooled by blowing cooling air which is supplied to core plugs 2, 3 inserted into the front sections from many small holes of the core plugs, 2, 3. The air is blown from blow off holes 5, 6 provided on the front surface of the blade 1 and the outside of the blade 1 is film cooled. A slit section 8 at a rear end of the core plug 3 is fitted closely in a slit of bulk head 9, and the cooling air in the core plugs 2, 3 is introduced into the trailing edge A from the slit section 8 and flown out therefrom to the outside of the blade 1. The trailing edge A is therefore, cooled, independent of the front sections of the blade 1.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a cooling structure for a stationary blade of a turbine section of a gas turbine, which supplies cooling air for increasing a cooling effect to a trailing edge portion of a blade having a high heat load, and a small amount of cooling air. The present invention relates to a cooling structure for a gas turbine vane, which improves cooling of the vane and improves efficiency of the entire gas turbine.

[0002]

2. Description of the Related Art A cross section of a conventional gas turbine cooling vane is shown in FIG. 2 (a), and a vertical cross section taken along a curved surface passing through its center is shown in FIG. 2 (b). In this gas turbine cooling vane (vane), the cooling air introduced into the core plugs 2 and 3 provided in the vane 1 having the partition wall 4 having the communication hole 10 is innumerably opened in the core plug. It is blown as a high-speed jet toward the inner surface of the blade from a small hole (not shown), the blade wall is cooled by the impinging jet, and then some air flows out from the film holes 5 and 6 formed in the blade 1. To cool the outer surface of the blade, and some of the air merges after passing through the gap between the core plug and the inner surface of the blade to cool the air for cooling the trailing edge of the blade.
As a result, it passes between the pin fins 7 at the trailing edge of the blade and flows out of the blade from the trailing edge of the blade.

[0003]

Air blown out from the small holes innumerable in the core plug exchanges heat with the amount of heat flowing from the outside of the blade to cool the blade, and the air itself is heated by heat exchange. To rise. The rate of temperature rise depends on the amount of heat from the outside of the blade and the air blowing speed. In the conventional stationary vane cooling structure, the air after the temperature rise flows into the trailing edge of the vane. On the other hand, generally, the distribution of the heat load outside the blade is C in FIG.
It is like a line, and as can be seen from this figure, the heat load on the trailing edge of the blade is high. However, in the conventional vane cooling structure,
There is a problem in that the blade trailing edge portion having a high heat load is cooled by high temperature air.

The metal temperature distribution on the outer surface of the stationary blade and the cooling air temperature distribution by the conventional cooling structure are as shown by lines A and B in FIG. 3, respectively. In this figure, the numbers 1 to 8 on the horizontal axis are shown in FIG.
The positions indicated by the numbers 1 to 3 in (a) are shown.
Indicates the inside of the core plug. As shown in FIG. 3, the temperature of the air blown from the inside of the core plug rises after colliding with the inner surface of the blade, and further rises before the inflow of the trailing edge of the blade. Also, when accelerating from the entrance of the trailing edge of the wing toward the exit,
Heat flows in and the temperature of the cooling air further rises. The metal temperature on the outer surface of the blade also rises from the inlet (point) of the trailing edge of the blade toward the trailing edge (point) of the blade, and becomes maximum at the trailing edge of the blade. As a result, the blade trailing edge is subjected to material deterioration due to high temperature oxidation and thermal stress (compressive stress) due to high blade metal temperature, and depending on the operating state of the gas turbine, cracking, crack development, and crack development due to repeated thermal fatigue, Furthermore, there is a fear of destruction.

As described above, in the above prior art, since the cooling air temperature is high at the trailing edge of the blade even though the heat load outside the blade is high, the metal temperature at the trailing edge of the blade rises, causing high temperature oxidation and heat. There is a problem of crack generation due to large stress.

It is an object of the present invention to provide a gas turbine cooling stationary blade which improves the cooling performance of the trailing edge portion of the blade and lowers the metal temperature of the trailing edge portion of the blade, thereby improving the reliability of the gas turbine. ..

[0007]

SUMMARY OF THE INVENTION A gas turbine cooling vane according to the present invention has the construction described in each of the claims.

[0008]

According to the present invention, at the trailing edge of the gas turbine stationary blade,
It is possible to supply cooling air having the same temperature as the cooling air of the blades other than the blade trailing edge. As a result, as shown by the following formula, it is possible to lower the blade trailing edge metal temperature by ΔT _m as compared with the conventional case. _{T m = T g -η (T} g -T c) T m '= T g -η (T g -T c') ΔT m = T m -T m '= η (T c -T c') = ηΔT _c where T _{m is a} conventional blade trailing edge metal temperature, T _m 'is a blade trailing edge metal temperature according to the present invention, T _{g is a} high temperature gas temperature outside the blade,
T _c : Conventional blade trailing edge cooling air temperature, T _c ′: Blade trailing edge cooling air temperature according to the present invention, η: Cooling efficiency (constant). That is, in the present invention, the temperature of the blade trailing edge cooling air is lower than that of the conventional one by ΔT _c , and the blade trailing edge metal temperature can be lowered by ΔT _m accordingly, whereby the blade trailing edge portion is cooled. It is possible to suppress the high temperature oxidation and the thermal stress level of, and improve the stability and reliability of the gas turbine stationary blade.

[0009]

1 is a cross-sectional view of a gas turbine stationary blade according to an embodiment of the present invention, and FIG. 1 is a vertical cross-section taken along a curved surface passing through the center thereof.
Shown in (a) and (b). The same parts as those in FIG. 2 are denoted by the same reference numerals.

In FIG. 1, air pressurized by a gas turbine compressor or air further cooled by an external cooling air cooling device is supplied as cooling air 11. The cooling air 11 is branched and supplied to a front core plug 2 and a rear core plug 3 that cool the front side of the stationary blade 1. Innumerable small holes (not shown) are formed in these core plugs, and the cooling air blown out from the core plugs through these innumerable small holes (not shown) collides with the inner surface of the blade 1 and the collision jet flows. To cool the wings. At this time, this cooling air absorbs heat and its temperature rises. The cooling air after the collision merges in the gap between the inner surface of the blade 1 and the core plug, and is jetted downstream from the film holes 5 and 6 provided in the blade 1 to the outside of the blade,
A low temperature air film is formed between the high temperature gas outside the blade and the outer surface of the blade to enhance the cooling effect of the blade.

The rear edge of the rear core plug 3 has a rear edge of a protruding core plug having a narrow slit formed therein.
The core plug trailing edge fins 8 are closely fitted to the slits provided in the partition wall 9 that divides the inside of the blade 1 into a blade trailing edge region and other regions. As a result, the cooling air 13 for the trailing edge portion of the blade is not the air after being blown out from a large number of small holes (not shown) of the core plug and colliding with the inner surface of the blade to raise the temperature, but the rear core plug. The low-temperature cooling air 11 inside 3 is guided to the blade trailing edge portion directly from the slits of the core plug trailing edge fins 8 and flows out from the blade trailing edge end to the outside of the blade. This makes it possible to keep the metal temperature at the trailing edge of the blade low. By adjusting the slit width above, it is also possible to adjust the flow rate of the cooling air passing through it.
Therefore, when the cooling effect of the trailing edge of the blade is excessive, the flow rate of cooling air passing through the slit can be reduced.

FIG. 5 shows another embodiment. In this embodiment, the partition wall 9 for partitioning the cavity in the blade 1 into the blade trailing edge region and the other region is the same as in the above embodiment, but the core plug has the slit-shaped core plug trailing edge fins. 8 is not provided, instead, core plugs 2 and 3 in the blade are provided.
The lower end wall is formed by projecting an opening 20 below the lower end of the core plug 2 and guiding a part of the cooling air 11 inside the core plugs 2 and 3 into a passage 18 provided in the lower end wall 17 of the blade 1. After cooling 17, the communication hole 17 '
Lead to the trailing edge of the wing. Further, the upper end wall 16 of the blade 1 is also provided with a communication hole 16 ', and a part of the cooling air 11 before entering the core plug flows directly from the upper side to the blade trailing edge portion through the communication hole 16'. Is done.
These air cools the trailing edge of the blade and flows out from the trailing edge of the blade. This structure enhances cooling of the trailing edge of the blade. In this case, since the cooling air flows into the blade trailing edge portion from the upper side and the lower side, it tends to be difficult for the cooling air to flow on the blade radial center line O. To prevent this, the blade trailing edge portion is formed. By increasing the number of rows of pin fins 7 as cooling elements to increase the air resistance on the upper side and the lower side, it is preferable that the air also flows well on the blade radial center line O.

In each of the above-described embodiments, as is clear from the structure, the cooling air blown out from the numerous small holes provided on the surface of the core plug and colliding with the inner surface of the blade to raise the temperature is introduced into the trailing edge of the blade. The temperature of the cooling air for cooling the trailing edge of the blade introduced into the trailing edge of the blade without mixing with the cooling air is the same as the temperature of the cooling air in the core plug.

FIG. 6 shows the blade metal temperature distribution and the cooling air temperature distribution according to the embodiment of the present invention in comparison with the conventional example. The numbers 1 to 8 on the horizontal axis represent the positions shown by the numbers 1 to 8 in FIGS. 1A, 2A, and 5A, respectively. Is the core plug 3
Represents the interior. According to the present invention, the air temperature T _c ′ for cooling the trailing edge of the blade is the same as the air temperature inside the core plug (point), which is about 150 times the temperature T _{c of} the air for cooling the trailing edge of the blade. ~ 200 ° C lower. As a result, the metal temperature at the trailing edge of the blade is lowered by the following ΔT _m . ΔT _m = η (T _c −T _{c ′} ) η: Cooling efficiency Therefore, the high temperature oxidation and thermal stress at the trailing edge of the blade are alleviated, and the blade life can be extended.

[0015]

According to the present invention, the metal temperature of the blade trailing edge portion of the gas turbine stationary blade can be made lower than that of the conventional one, so that the thermal stress of the blade and the level of high temperature oxidation can be reduced. This has the effects of improving the reliability of the gas turbine stationary blade and extending its life.

[Brief description of drawings]

FIG. 1 is a cross-sectional view and a vertical cross-sectional view of a gas turbine cooling vane according to an embodiment of the present invention,

FIG. 2 is a transverse sectional view and a longitudinal sectional view of a conventional gas turbine cooling vane;

FIG. 3 is a diagram showing distributions of surface metal temperature and cooling air temperature of a conventional gas turbine cooling vane;

FIG. 4 is a diagram showing a distribution of an external heat load of a gas turbine stationary blade,

FIG. 5 is a horizontal sectional view and a vertical sectional view of a gas turbine cooling vane according to another embodiment of the present invention,

FIG. 6 is a diagram showing the distribution of the metal temperature and the cooling air temperature of the gas turbine cooling vane of the present invention in comparison with those in the prior art.

[Explanation of symbols]

1 ... Gas turbine stationary blade 2, 3 ... Core plug 4 ... Partition wall 5, 6 ... Film outlet 7 ... Pin fin 8 ... Core plug trailing edge fin 9 ... Partition wall 10 ... Communication hole 11 ... Cooling air 12 ... Between blade inner surface and core plug Gap 16 ... Upper end wall 16 '... Communication hole 17 ... Lower end wall 17' ... Communication hole 18 ... Air passage 20 ... Opening part

Claims (3)

[Claims] 1. A gas turbine cooling vane having a cavity inside for introducing cooling air to the cavity, wherein the cavity in the vane is divided into a region at the trailing edge of the vane and a region other than that. A partition wall is provided, and the cooling air introduced into the core plug inserted therein is blown out from a large number of small holes of the core plug into the region other than the region of the trailing edge of the blade to supply the cooling air. Together with the other, on the other hand, in the region of the blade trailing edge portion, cooling air is supplied via a blade trailing edge portion cooling air supply passage that is independent of the cooling air supplied into the region other than the region of the blade trailing edge portion. A gas turbine cooling vane characterized by being configured to supply. 2. The air supply passage for cooling the blade trailing edge portion,
The rear edge end portion of the core plug is formed as a protruding slit portion, and the protruding slit portion is closely fitted to the slit provided in the partition wall and opened in the region of the blade trailing edge portion, thereby forming a structure. The gas turbine cooling stator vane according to claim 1. 3. The air supply passage for cooling the blade trailing edge portion,
The lower portion of the core plug is formed as a projecting opening,
The opening is opened in the air passage formed in the lower end wall of the vane, and the air passage in the lower end wall is communicated with the trailing edge region of the vane. The gas turbine cooling stationary blade according to claim 1, wherein the gas turbine cooling stationary blade is constituted by providing an air passage extending from an upper end wall of the blade to an area of the blade trailing edge portion.