Gas turbine

A gas turbine is provided including a compressor, a combustion chamber, a guide vane row and a rotor airfoil row. The guide vane row includes a plurality of guide vane airfoils having a blade and an inner platform. The ratio between the pitch and the leading edge diameter of the guide vane airfoils is between 6.3-7.6 and the ratio between the platform length and the leading edge diameter of the guide vane airfoils is between 4.0-5.5.

FIELD OF INVENTION

The present invention relates to gas turbines.

BACKGROUND

Gas turbines are known to comprise a compressor, a combustion chamber and a turbine.

Different gas turbines comprise a compressor, a first combustion chamber and a high pressure turbine; thus these gas turbines comprise a second combustion chamber and a low pressure turbine.

In the following particular reference will be made to high pressure turbines, it is anyhow clear that the present invention may be implemented in any kind of turbine, also not being the high pressure turbine or a turbine stage facing the combustion chamber.

Turbines have at least a guide vane row and a rotor blade row.

Each guide vane row is made of stator airfoils having an inner and an outer platform facing respective inner and outer walls of the combustion chamber; moreover the inner and outer platforms are separated from the inner and outer combustion chamber walls by an inner and an outer gap.

During operation the hot gases generated in the combustion chamber from the combustion of a fuel with the compressed air coming from the compressor, pass through the turbine to deliver mechanical power to the rotor.

As known in the art, when hot gases impinge on an obstacle, a high static pressure zone is generated.

Thus, as during operation the hot gases passing through the turbine impinge on the guide vane airfoils, in the zone upstream of the guide vane row a high static pressure zone is generated.

In particular the high static pressure is not uniform, but has peaks in correspondence with the leading edges of the guide vane airfoils.

This effect is particularly relevant in the first guide vane row after the combustion chamber.

In addition, the hot gases path (i.e. the duct wherein the hot gases generated in the combustion chamber pass through) has a first constricting cross section zone followed by a second expanding cross section zone followed by a third constricting cross section zone.

In the second expanding cross section zone a transition between the combustion chamber and the platforms of the guide vane airfoils is provided.

It is clear that this expanding portion makes the hot gases static pressure in the transition zone between the combustion chamber and the guide vane platforms (i.e. in the zone upstream of the leading edges of the guide vane blades) to further increase.

Such high static pressure causes the risk that hot gases enter the gaps, and damage the components nearby (so-called “gas ingestion”).

Because of the particular shape of the hot gases path, this risk is mainly relevant at the inner gap.

SUMMARY

The disclosure is directed to a gas turbine including at least a combustion chamber, a guide vane row and a rotor airfoil row. The guide vane row includes a plurality of guide vane airfoils including a blade and an inner platform. A ratio between a pitch and a leading edge diameter of the guide vane airfoils is between 6.3-7.6 and a ratio between a platform length and the leading edge diameter of the guide vane airfoils is between 4.0-5.5. The platform length is defined by the axial distance between a leading edge of a guide vane blade and an inner guide vane platform inlet measured at half height of the guide vane blade.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Introduction to the Embodiments

The technical aim of the present invention is therefore to provide a gas turbine by which the said problems of the known art are eliminated or sensibly reduced.

Within the scope of this technical aim, an object of the invention is to provide a gas turbine by which the risk of gas ingestion caused by the high static pressure upstream of the guide vane airfoil leading edges, in particular in the inner gap between the combustion chamber and the guide vane row, is very low.

The reliability of the gas turbine is thereby increased with respect to traditional gas turbines.

The technical aim, together with these and further objects, are attained according to the invention by providing a gas turbine in accordance with the accompanying claims.

DETAILED DESCRIPTION

With reference to the figures, shown are a portion of a gas turbine that comprises a compressor (not shown), a combustion chamber2(partially shown) and a turbine stage immediately downstream of the combustion chamber2that comprises a guide vane row3.

The combustion chamber2has an annular shape and is defined by an inner wall4and an outer wall5.

The guide vane row3comprises a plurality of guide vane airfoils each having a blade7, an inner platform8and an outer platform9; the inner platforms8of the adjacent guide vane airfoils in combination with the outer platforms9of the adjacent guide vane airfoils define an annular hot gases path.

Between the combustion chamber inner wall4and the guide vane inner platform8there is provided an inner gap11; correspondingly between the combustion chamber outer wall5and the guide vane outer platform9there is provided an outer gap12.

Downstream of the guide vane row3a rotor airfoil row is provided; the rotor airfoil row is not shown.

FIG. 1shows pitch P, being the circumferential distance between the leading edges15of two adjacent guide vane blades7and the leading edge diameter D, being the diameter of the guide vane blade7at the leading edge15; these parameters are measured at half height of the guide vane blade7.

Moreover,FIG. 2shows the platform length L at the inner diameter, being the axial distance measured at half height of the guide vane blade7between the leading edge15of a guide vane blade7and the guide vane inner platform inlet16.

Advantageously, the ratio between the pitch P and the leading edge diameter D of the guide vane airfoils is between 6.3-7.6, preferably between 6.7-7.1 and more preferably 6.8-7.0.

Moreover, the ratio between the platform length L and the leading edge diameter D of the guide vane airfoils is between 4.0-5.5, preferably between 4.5-5.0 and more preferably 4.6-4.8.

In addition, the area of the gas path at least in the zone of the first guide vane row3continuously decreases.

FIG. 2shows a plane17defining the cross section of the hot gases path at the platform inlet16and a plane18defining the cross section of the hot gases path at the leading edges15of the guide vane blades7.

Advantageously, the annulus constriction in the zone of the first guide vane row3, defined by the ratio between the hot gases path area at the cross section defined by the plane17and the hot gases path area at the cross section defined by the plane18, is comprised between 1.0-1.5, preferably 1.1-1.4 and more preferably 1.2-1.3.

Advantageously this annulus constriction provides a hot gases path cross section that is continuously decreasing, thereby avoiding expanding zones wherein the static pressure of the hot gases increases.

Moreover, the inner gap11and the outer gap12are aligned with each other with respect to a plane20perpendicular to the gas turbine axis21.

The operation of the gas turbine of the invention is apparent from that described and illustrated and is substantially the following.

A fuel/compressed air mixture is combusted in the combustion chamber2forming hot gases that flow through the hot gases path and, in particular, pass through the guide vane row3.

In a zone22of the hot gases path upstream of the guide vane airfoils, the static pressure of the hot gases that impinge on the guide vane blades7increases.

Nevertheless as the gap11is far away from the leading edges15of the guide vane blades7, the high static pressure does not cause (or causes in a very limited amount) the hot gases to enter into the inner gap11.

In addition, only a low amount of hot gases enters into the outer gap12because of the shape of the outer platform and because of the distance between the leading edges15of the guide vane blades7and the outer gap12.

Moreover, the fact that the hot gases path cross section continuously decreases, in particular in the zone upstream of the guide vane row3, helps to reduce the hot gases static pressure upstream of the guide vane row3and, in addition, to increase the stability of the hot gases flow and to counteract the flow separation.

In this respectFIG. 3shows the profile of the hot gases path in the zone between the end of the combustion chamber2and the guide vane row3for an embodiment of the gas turbine according to the invention and according to the prior art.

In particular, inFIG. 3the continuous line indicates the profile of the hot gases path of the embodiment of the invention, and the dashed line the profile of the hot gases path of an embodiment of the prior art; moreover inFIG. 3also the positions of the gap11in the embodiment of the invention and prior art are indicated.

FIG. 3clearly shows that in the embodiment of the invention the gap11is located in a constricting cross section zone of the hot gases path, whereas according to the prior art the gap11is located in an expanding cross section zone of the hot gases path.

The gas turbine conceived in this manner is susceptible to numerous modifications and variants, all falling within the scope of the inventive concept; moreover all details can be replaced by technically equivalent elements.

In practice the materials used and the dimensions can be chosen at will according to requirements and to the state of the art.

REFERENCE NUMBERS

2combustion chamber3guide vane row4inner wall of the combustion chamber5outer wall of the combustion chamber7blade of the guide vane airfoil8inner platform of the guide vane airfoil9outer platform of the guide vane airfoil11inner gap between4and812outer gap between5and915leading edge of the guide vane blade16platform inlet17hot gases path cross section at the platform inlet1618hot gases path cross section at the leading edges1520plane perpendicular to the gas turbine axis2121gas turbine axis22hot gases path zone upstream of the guide vane row3P pitchD leading edge diameter of the guide vane bladeL platform length