Gas turbine

This application claims priority to European application no. 09171142.4, filed 23 Sep. 2009, the entirety of which is incorporated by reference herein.

1. Field of Endeavor

The present invention relates to a gas turbine.

In particular the present invention refers to the sealing of the zone between the guide vane boxes of the high-pressure turbine immediately downstream of the combustion chamber and a fixed frame, such that possible leakages of hot gases flowing in the combustion chamber and/or compressed air used to seal the zone between the combustion chamber and stator airfoil row do not enter the rotor airfoils cooling circuit.

2. Brief Description of the Related Art

In the following reference, will be made to FIG. 1 for describing the relevant parts of the gas turbine; in particular reference will be made to a sequential combustion gas turbine, it is anyhow clear that structures embodying principles of the present invention may be implemented in any gas turbine also not being a sequential combustion gas turbine.

Sequential combustion gas turbines 1 have a compressor (not shown) compressing air and supplying it to first burners (not shown) where fuel is injected and a mixture to be combusted is formed.

Downstream of the first burners a first combustion chamber 2 is provided, where the mixture is combusted to form high pressure hot gases F that are supplied to a high-pressure expansion stage.

The high-pressure expansion stage includes a stator airfoil row 4 separated from the combustion chamber 2 by a first gap 5 , and a rotor airfoil row 6 separated from the stator airfoil row 4 by a second gap 7 ; third gaps 8 are provided between the rotor airfoil row 6 and an annular duct 9 feeding a plurality of side-by-side second burners 10 , wherein further fuel is injected in the hot gases (still rich in air) already partially expanded in the high-pressure expansion stage, such that an ignitable mixture is formed. This ignitable mixture is combusted in a second combustion chamber (not shown) and the hot gases produced are further expanded in a low pressure turbine (not shown).

The stator airfoil row 4 is made of stator airfoils 15 defining between each other guide vanes and having endwalls 16 connected to guide vane boxes 17 .

The guide vane boxes 17 have a box structure and are fed with cooling air A via connections not shown for simplicity.

In particular, the cooling air A comes from the compressor at a temperature of about 450-550° C. and is cooled by an external cooler to a temperature of typically 200-400° C.

Moreover the guide vane boxes 17 are also provided with nozzles 20 that inject the cooling air A into the second gap 7 .

The rotor airfoil row 6 includes a plurality of rotor airfoils 22 having a hollow body provided with an inlet 23 arranged to collect the cooling air A injected from the nozzles 20 .

During operation, the hot gases F formed in the first combustion chamber 2 pass through the stator and rotor airfoil row 4 , 6 such that the rotor airfoil row 6 extracts mechanical power from them.

Moreover, the air A from the guide vane boxes 17 is injected through the nozzles 20 in the second gap 7 towards the rotor airfoil inlets 23 .

As the rotor airfoil row 6 rotate with high speed, it draws the cooling air A injected from the nozzles 20 and makes it to enter the rotor airfoil 22 via the inlets 23 .

The cooling air A entering the rotor airfoils 22 cools the rotor airfoils 22 and is then injected through holes (usually at the leading edge and trailing edge of each rotor airfoil row); the air injected through the leading and trailing edges of the rotor airfoils 22 is indicated by A 2 .

In order to prevent the hot gases F from entering the first gap 5 (the hot gases have a temperature of about 1200-1500° C. and would impair the components close to the first gap 5 ), compressed air (the so-called ‘purge air’) is diverted from the compressor and is injected in the first gap 5 . This air has a temperature of about 450-550° C. and thus is not dangerous for the components close to the gaps 5 .

In addition, in order to prevent the compressed air (purge air) from reaching the rotor airfoil inlet 23 , seals 25 are provided between the stator airfoil endwalls 16 /guide vane boxes 17 and a fixed frame 26 .

Nevertheless, the compressed air diverted from the compressor may leak and pass through the seals 25 and mix with the cooling air A injected in the second gap 7 .

For this reason, the cooling air A flow rate is quite large, such that, in all operating conditions, the air entering the rotor airfoil 22 has a correct temperature to safeguard the rotor airfoil integrity and guarantee their lifetime.

Nevertheless, since the cooling air A flow rate diverted from the compressor into the guide vane boxes is quite large, efficiency of the gas turbine is reduced.

FR 1 351 268 discloses a guide vane box with openings from which compressed air is injected to enter cooling conduits of the rotor airfoils.

GB 2 246 836 discloses a guide vane with first and second passages; from these passages cooling air is injected into cooling passages of the rotor airfoils. The second passages can be blocked off by a Belleville washer.

EP 0 636 765 discloses guide vane boxes with passages from where a flow is injected into rotor airfoil inlets of a rotor cooling circuit.

One of numerous aspects of the present invention includes a gas turbine by which the aforementioned problems of the known art are addressed.

Another aspect of the present invention includes a gas turbine having an increased efficiency when compared with traditional gas turbines.

Further characteristics and advantages of the invention will be more apparent from the description of a preferred but non-exclusive embodiment of the gas turbine according to the invention, illustrated by way of non-limiting example in the accompanying drawings, in which:

FIG. 1 shows a schematic cross section of a portion of a gas turbine according to the prior art;

FIG. 2 shows a schematic cross section of a portion of a gas turbine according to the invention; and

FIGS. 3 and 4 show a particular of guide vane boxes according to two embodiments of the invention.

With reference to the figures, these show a gas turbine 1 having the combustion chamber 2 followed by the stator airfoil row 4 and the rotor airfoil row 6 .

The structure of the gas turbine is the same as that already described; it is thus not described again and with the same reference numbers the same elements are indicated.

In particular, the guide vane boxes 17 are provided with passages 30 connecting a zone 31 upstream of the guide vanes boxes 17 to a zone 32 of the second gap 7 downstream of the guide vanes boxes 17 .

In addition, the mouth 34 of the passages 30 facing the rotor airfoil row 6 is closer to a hot gases path 35 than the nozzles 20 .

The mouth 34 of the passages 30 facing the rotor airfoil row 6 is substantially as close as, or it is closer than, the rotor airfoil inlet 23 to the hot gases path 35 . This permits the flow going out from the mouth 34 not to be drawn from the rotor airfoil row 6 to enter the inlet 23 .

In a first embodiment ( FIG. 4 ), the passages 30 are defined by slots at sidewalls 36 of the guide vane boxes 17 .

In this embodiment, the two contacting sides of two adjacent guide vane boxes may be provided with the slot, such that the passages 34 are defined between two facing slots.

Alternatively only one of the two contacting sidewalls 36 of the adjacent guide vane boxes 17 may be provided with the slot, in which case the passages 30 are defined by the slot of a guide vane box 17 and the flat surface of the adjacent guide vane box 17 .

In a different embodiment ( FIG. 3 ), the passages 30 extend inside of the guide vane boxes 17 and are defined by pipes.

Naturally, in further embodiments the guide vane boxes may be provided with both the slot and the pipes.

In addition, a seal 37 is provided downstream of the mouths 38 of the passages 30 opposite the rotor airfoil row 6 , between the guide vane boxes 17 and the fixed frame 26 . This lets the leakage that may overcome the seals 25 be withheld in a zone separate from the rotor airfoil row 6 .

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

The hot gases pass through the hot gases path 35 and thus they pass through the combustion chamber 2 , the stator airfoil row 4 , and the rotor airfoil 6 .

Through the first gap 5 compressed air (purge air) is supplied in the combustion chamber 2 .

A part of the compressed air (purge air) may leak, overcoming the seals 25 to enter the zone 31 upstream of the guide vane boxes 17 .

Thanks to its high pressure (greater than the pressure inside of the second slot 7 ), the compressed air (purge air) enters the passages 30 through the mouths 38 , passes through the passages 30 and moves out through the mouths 34 entering the second gap 7 in a zone from where it cannot enter the rotor airfoil inlet 23 ; thus the compressed air (purge air) enters the hot gases path 35 .

The additional seal 37 keeps this compressed air (purge air) in a zone adjacent to the mouth 38 of the passage and prevents the high temperature compressed air from being drawn from the high speed rotating rotor airfoil row 6 .

As shown in FIG. 2 , the vane includes a platform 28 having an extension which extends to at least partly close the gap separating the rotor airfoil from the stator airfoil row. The mouths of the passages connecting a zone upstream of the guide vane boxes to a zone of the gap downstream of the guide vane boxes are below the platform extension on a side facing away from a hot gas flow path. The guide vane boxes each include an extension 27 which divides the gap into a radially inner section of the guide vane boxes and a radially outer section side facing the gap. Each of the passage mouths are positioned at a gap in a radially inner section.

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.

1 gas turbine

2 combustion chamber

4 stator airfoil row

5 first gap

6 rotor airfoil row

7 second gap

8 third gap

9 annular gap

10 second burner

15 stator airfoils

16 endwalls of 15

17 guide vane boxes

20 nozzles

22 rotor airfoils

23 rotor airfoil inlet

25 seals

26 fixed frame

30 passages

31 zone upstream of the guide vane boxes

32 zone downstream of the guide vane boxes

34 mouth of 30

35 hot gases path

36 sidewalls of 17

37 seal

38 mouth of 30

A cooling air

A 2 air injected through 22

F hot gases flow

While the invention has been described in detail with reference to exemplary embodiments thereof, it will be apparent to one skilled in the art that various changes can be made, and equivalents employed, without departing from the scope of the invention. The foregoing description of the preferred embodiments of the invention has been presented for purposes of illustration and description. It is not intended to be exhaustive or to limit the invention to the precise form disclosed, and modifications and variations are possible in light of the above teachings or may be acquired from practice of the invention. The embodiments were chosen and described in order to explain the principles of the invention and its practical application to enable one skilled in the art to utilize the invention in various embodiments as are suited to the particular use contemplated. It is intended that the scope of the invention be defined by the claims appended hereto, and their equivalents. The entirety of each of the aforementioned documents is incorporated by reference herein.