JPH0337302A - gas turbine equipment - Google Patents

Abstract

(57) [Summary] This bulletin contains application data before electronic filing, so abstract data is not recorded.

Description

[Detailed Description of the Invention] [Industrial Application Field] The present invention relates to an axial flow turbine having stator blades and rotor blades;
The present invention relates to a gas turbine including a compressor, a combustor, a precooler for precooling air from the compressor, and more specifically, to a cooling structure for first stage stationary blades of the axial flow turbine.

[Prior Art] In gas turbines, efforts are being made to increase the temperature of the turbine inlet gas, for example, to 1400° C., in order to improve the thermal efficiency characteristics or output characteristics.

When the inlet gas temperature of the turbine increases in this way, it is necessary to lower the temperature of the blades to the permissible temperature of the material of the blades. To this end, efforts are being made to improve the cooling structure and increase the amount of cooling air.

However, these cooling methods have limitations.

In particular, if the amount of cooling air is increased too much, the performance will drop significantly and there will be no advantage of increasing the inlet temperature. Therefore, as a way to substantially lower the surface temperature without consuming too much cooling air, a cooling method that cools the compressor discharge air once and guides it inside the blade has been proposed. Transaction, 106.4 (198
4 years) p756-764 (ASMH Transacti
on, Vol. 106. &4, 1984. p756-7
64). The third section schematically shows the outline.
To explain with a diagram, the gas turbine device has a compressor C1
It consists of a combustor n, a turbine T, and a precooler A. The air extracted from the outlet of the compressor C or from the plenum chamber of the combustor B is compressed by the compressor C and therefore has a high temperature and high pressure.

This high-temperature, high-pressure air is once cooled down in the precooler A, and then sent into the cavity in the first stage stator vane located downstream of the combustor of the turbine T, which operates in particularly high-temperature gas. The air flows out into the mainstream gas through the holes, cooling the first stage stationary blades and rotating rotor blades.

[Problems to be Solved by the Invention] However, in the case of such a cooling system, the air with a total pressure of P2 extracted from the compressor C has a total pressure of P2 due to the pressure loss that occurs when passing through the piping and the precooler. Total pressure P lower than
,become. Although this pressure loss depends on the size of the piping and precooler, the total pressure loss ratio ((p, -p, )/P, ) is about 5 mm. On the other hand, most of the air discharged from the compressor C passes through the combustor B and flows into the turbine T as high-temperature, high-pressure gas at a total pressure P4. Pressure loss also occurs in combustor B, but its total pressure loss rate (p2-p*)/P is about 2.5 to 5z. Therefore, the pressure P3 of the cooling air supplied to the inside of the first stage stator vane is smaller than the total pressure P4 of the two ports of the first stage stator vane, and if a crack occurs in the first stage stator vane after long-term use, the pressure will pass through this crack. There is a possibility that the gas with a high temperature, that is, the mainstream gas from the combustor B, flows back into the interior of the stator blade. When reverse flow occurs, the temperature of the stator blades increases and their lifespan is extremely shortened, resulting in a decrease in the reliability of the gas turbine.

As a solution to the above problem, it may be possible to provide a separate compressor and use a separate compressor to increase the pressure that has fallen through piping, precoolers, etc.; however, if a separate compressor is provided, Another problem arises in that the power loss driving the compressor increases, eliminating any performance advantage, and also complicating the turbine system.

SUMMARY OF THE INVENTION Accordingly, an object of the present invention is to provide a gas turbine stator blade cooling system that uses pre-cooled cooling air and prevents mainstream gas from flowing back into the first stage stator blades.

[Means for Solving the Problems] In order to solve the above problems, the present invention divides the inside of a first stage stator vane into at least two chambers, and a compressor is installed in the chamber located near the leading edge of the stator vane. The configuration is such that the air discharged from the engine is guided, and the air that has passed through the precooler is introduced into the chamber located closer to the trailing edge.

[Operations/Effects] In the present invention, the first stage stator vane is cooled by the above-mentioned respective air introduced into each chamber inside the first stage stator vane. Incidentally, the surface pressure distribution of the first stage stationary blade is highest at the leading edge and becomes lower toward the trailing edge. The pressure of the compressor discharge air directly introduced into the chamber near the leading edge of the stationary blade is higher than the surface pressure of the leading edge, and the pressure of the air introduced into the chamber near the trailing edge is higher than the pressure of the air introduced into the chamber near the trailing edge. Although the pressure is lower than the trailing edge surface pressure, even if a crack occurs in the first stage stator blade, the mainstream gas will not flow back into the stator blade through this crack. do not have.

Therefore, it is also possible to adopt a cooling method in which film cooling holes are provided at the leading edge, and to strengthen cooling by lowering the temperature of the cooling air using a precooler at the center or trailing edge of the blade.

[Example] An example of the present invention will be described below with reference to FIGS. 1 and 2.

FIG. 1 is a sectional view showing the vicinity of the first stage stationary blade of a gas turbine. Although not shown, a compressor, a combustor, and a precooler are provided as in FIG. 3.

The first stage stationary blade 2 is fixed by a retainer ring 3 fitted to the turbine casing 5 and a support ring 4 located on the rotating shaft side. , combustor transine 1
Combustion gas flows in from the engine piece 1, is accelerated by changing its outflow direction by the first stage stator blade 2, and hits the rotor blade 6 to generate rotational force. Two chambers 20 and 21 are formed inside the stationary blade 2, and a core plug 811o is inserted into each of these chambers.

The air from the compressor passes through the precooler and is cooled by the precooler,
It passes through the opening 22 of the turbine casing 5, the hole 23 of the retainer ring 3 and the annular air chamber 24.

Then, it passes through the impingement force bar 9 provided for cooling the end wall 25 of the stator blade 2 and is guided into the core plug 10, and is guided into the core plug 10 through a number of small holes (simplified illustration). After cooling the inner wall of the chamber 20 of the stator blade 2 by blowing air from the stator blade 2 (not shown), the air is blown out through a number of film holes provided in the stator blade 2 (holes through which air flows out to form an air film on the stator blade surface). ) 14 and the outflow hole 15 leading to the trailing edge of the stationary blade into the mainstream gas. On the other hand, the compressor discharge air that does not pass through the precooler is guided through the combustor chamber 11 to the support ring 4, passes through the impingement force bar 7 for cooling the end wall 26, and is introduced into the core plug holder 8. After cooling the inner wall of the chamber 21 of the stationary blade 2 by blowing air from a large number of small holes (not shown) provided in the core plug 8, the large number of film holes 1 provided in the chamber 20 of the stationary blade 2 are cooled.
3 to flow out into the mainstream gas.

The true pressure distribution of the first stage stationary blade 2 is 21st! As shown in I, it is highest on the leading edge side 30 and becomes lower toward the trailing edge side 31. In view of this, in this embodiment, as cooling air, compressor discharge air with high pressure almost equal to the compressor discharge pressure is sent to the chamber 20 near the leading edge side, and via a precooler to the chamber 21 near the trailing edge side. The pressure is reduced by
As for the pressure, air whose pressure has been reduced by the pressure loss in the precooler and its piping system is introduced. As shown in FIG. 2, the compressor discharge air pressure P2 supplied to the chamber 20 is higher than the surface pressure of any part of the first stage stationary blade 2, so air flows out from the film hole 13.

Furthermore, even if a crack occurs in the stator blade 2 at the chamber 20, the mainstream gas will not flow back into the stator blade.On the other hand, the surface pressure of the stator blade surface against the chamber 21 is lower than that at the leading edge. The pressure P of the air introduced into the chamber 21 from the precooler is higher than the surface pressure;
The air introduced into the film hole 14 and the trailing edge outflow hole 1
Furthermore, even if a crack occurs in the stator blade 2 in the chamber 21, the mainstream gas will not flow back into the stator blade.

In the above embodiment, the chambers were provided in the stator blade 2 at the front and rear positions, but there are chambers near the leading edge of the stator blade (first chamber), a chamber toward the trailing edge (third chamber), and a chamber near the leading edge of the stator blade (first chamber). Three chambers may be provided, such as a chamber between them (the second chamber). In this case, the compressor discharge air is introduced into the first and second chambers, and the air from the precooler is introduced into the third chamber. Compressor discharge air in 1 chamber, 2nd chamber, 3rd chamber
A similar effect can be achieved by introducing air from a precooler into the chamber. &! A G chamber may also be provided.

It can be done according to this. The present invention can also be applied to a case where a film cooling hole is provided in the leading edge of the stator blade 2, and compressor discharge air is introduced into a chamber on the leading edge side of the blade, and air from a precooler is introduced into a chamber on the trailing edge side.

[9! As described above, according to the present invention, high-pressure compressor discharge air is introduced into the chamber near the leading edge of the first-stage stator vane, which is placed in the highest temperature atmosphere, so that the air in front of the vane is The trailing edge side, which has a large heat load and is difficult to cool due to the structure of the blade, is cooled by introducing low-temperature air that has passed through a precooler into a chamber near the trailing edge side, which has a good cooling effect. Even if a crack occurs in the blade, high temperature mainstream gas is prevented from flowing back into the stator blade, resulting in a highly reliable gas turbine.

[Brief explanation of drawings]

FIG. 1 shows a first example of a gas turbine showing an embodiment of the present invention.
FIG. 2 is a cross-sectional view of the vicinity of a stage stator vane, FIG. 2 is a diagram showing a cross section of a first stage stator vane and pressure distribution on the blade surface, and FIG. 3 is a schematic overall view showing a gas turbine device having a precooler. 1... Combustor transition piece 2... 1st stage stator vane 3... Retainer ring 4...
・・Support ring 5・Turbine casing 6・
・・Rotating blade 7・・Inbingement capa 8
...Core plug 9...Impingement cover 10...Core plug 11...Combustor chamber 12...
・Stator blade cross section 13... Film hole 14... Film hole 15... Trailing edge outflow hole 20.21... Chamber
30... Pre-record 31... Trailing edge diagram Cooling air from precooler l: Combustor transition piece 2: 1st stage stationary blade 3: Retainer ring 4: Support ring 5, Turbine casing 6: 1st stage rotor blade 7.9: Impingement force bar 8, IQ: Core plug diagram 8.10: Core plug 12: Stator blade cross section 13, 14: Film hole 15: Trailing edge outflow hole diagram

Claims (1)

[Claims] 1. In a gas turbine device consisting of a compressor, a combustor, an axial flow turbine that operates with mainstream gas from the combustor, and a precooler that precools a part of the air compressed by the compressor,
The space between the leading edge and the trailing edge of the first stage stator blade of the axial flow turbine is divided into at least two chambers, and compressor discharge air is directly guided as cooling air into the chamber located closer to the leading edge of the stator blade. A gas turbine stator blade cooling device, characterized in that the compressor discharge air is passed through a precooler and cooled air is guided as cooling air into the chamber located near the trailing edge of the stator blade.