EP1320661A1

EP1320661A1 - Gas turbine blade

Info

Publication number: EP1320661A1
Application number: EP01980405A
Authority: EP
Inventors: Peter Tiemann
Original assignee: Siemens AG
Current assignee: Siemens AG
Priority date: 2000-09-26
Filing date: 2001-09-18
Publication date: 2003-06-25
Anticipated expiration: 2021-09-18
Also published as: JP4669202B2; WO2002027146A1; US6874988B2; EP1191189A1; EP1320661B1; JP2004510091A; DE50113551D1; US20040022630A1

Abstract

The invention relates to a gas turbine blade (1) having a combined convective cooling effected by a meandering cooling channel (21) and having an impact cooling effected via an impact cooling insert (37). The impact cooling insert (37) is arranged inside a first partial section (23) of the meandering cooling channel (21) extending along the front edge (8) of the blade pan. The impact cooling insert (37) tapers along this first partial section (23).

Description

description

Gas turbine blade

The invention relates to a gas turbine blade with an airfoil leading edge and an airfoil trailing edge and with an internal cooling structure, comprising a meandering cooling channel with partial sections directed along the airfoil axis for guiding a cooling fluid from the airfoil leading edge to the airfoil trailing edge.

No. 5,468,125 discloses a hollow gas turbine blade which can be cooled by cooling air. The cooling air is blown into cooling chambers of the hollow gas turbine blade, which run parallel to the blade axis, where it continuously cools the hot surface of the gas turbine blade from the inside. The incoming cooling air, which has not yet been heated, is first led past the leading edge of the gas turbine blade, which is exposed to particularly high temperatures and must therefore be cooled particularly efficiently. after the

Cooling air has also cooled the other areas of the blade through the blade, it leaves the blade at the trailing edge of the blade via bores.

The object of the invention is to provide a gas turbine blade that uses a cooling fluid to cool the gas turbine blade in a particularly efficient manner.

According to the invention, this object is achieved by specifying a gas turbine blade directed along an airfoil with an airfoil leading edge and an airfoil trailing edge and with an internal cooling structure, comprising an aander-shaped cooling duct with partial sections directed along the airfoil axis for guiding a cooling fluid from the airfoil leading edge to the airfoil trailing edge, the rear edge of the airfoil Sections runs along the front edge of the airfoil and an entry area for the Has cooling fluid and an outlet area for the cooling fluid, the first section having an impingement cooling insert which, with its insert front directed towards the front airfoil, runs parallel to the front edge of the airfoil, the impingement cooling insert tapering towards the outlet area.

The invention proceeds from the recognition that the airfoil leading edge can always be cooled sufficiently efficiently in a conventional internal cooling a Gasturbinenschau- fei by a meandering cooling channel is not, as the thermally particularly highly loaded outside ^• the airfoil leading edge has a relatively small surface area on the inside faces the front edge of the airfoil. Purely convective cooling using a

Cooling fluid flow in the meandering channel partial area on the leading edge of the airfoil can under certain circumstances be insufficient to sufficiently lower the temperature of the leading edge of the airfoil. On the other hand, the invention is based on the observation that cooling by means of an impact cooling insert is precisely the case with

The front edge of the airfoil allows greater heat dissipation due to the higher cooling capacity of the impingement cooling, but the cooling of the airfoil as a whole by the impingement cooling insert is comparatively inefficient, since the cooling fluid absorbs less heat overall. At a. gas turbine blade with meandering cooled by cooling air, for example, the cooling fluid emerging from the trailing edge after passing through the meandering channel is warmer than that which also emerges from impingement cooling from a trailing blade edge _. Cooling fluid.

The invention now for the first time combines impingement cooling with meandering channel cooling in such a way that the advantages of these two methods are exploited without being equally exposed to the disadvantages of the respective methods. This is achieved in that the airfoil pre <-υ ω w) P ¹ P »cπ on CD cn o Cπ tr PJ P- d 0 03 P- CQ St.

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everywhere is supplied with impingement cooling air. This means a particularly homogeneous cooling.

c) The impingement cooling insert is preferably surrounded by air guiding ribs directed transversely to the blade axis, which guide cooling fluid emerging from the impingement cooling insert around the impingement cooling insert in the direction of the rear edge of the airfoil. By means of such air guide ribs, the cooling fluid, after it has impinged on the airfoil wall, is guided along the outer wall of the impingement cooling insert away from the airfoil leading edge and then enters the free part of the first section. The free part of the first section is the part in which the impact cooling insert is not arranged. The air-guiding ribs are further preferably directed in relation to a plane oriented perpendicular to the blade axis in such a way that they additionally direct the cooling fluid in one direction from the inlet region to the outlet region. The cooling fluid entering the free part of the first section therefore already has a flow component in the direction of the main flow in this first section. The flow guidance by means of the air guide ribs thus enables a flow of the cooling fluid through the gas turbine blade that is as vortex-free as possible and therefore particularly favorable in terms of pressure loss.

d) The gas turbine blade is preferably designed as a guide blade which is designed with an inner ring. When the guide vane is used in a gas turbine, the inner ring serves to seal a hot gas duct of the gas turbine from a rotor of the gas turbine. An inner ring cooler leads from the impact cooling insert to the inner ring. While in conventional cooling of the gas turbine blade solely by convective cooling of a cooling fluid flowing in a meandering channel, the efficiency of cooling an inner ring of a gas turbine et 10 d φ

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An internal cooling system of the gas turbine blade 1 is explained in more detail below:

A meandering cooling duct 21 leads through the interior of the gas turbine blade 1. The meandering cooling duct 21 is made up of sections 23, 25, 27 directed along the blade axis 3. These sections 23, 25, 27 are separated from one another by ribs 31. The first subsection 23 runs along the front edge 8 of the airfoil. In the meandering cooling duct 21, the inside of the

Blade area 7 arranged turbulators 29, which provide for the generation of turbulence in a cooling fluid flowing through the meandering cooling channel 21, which in turn results in improved heat transfer to the cooling fluid. The first partial section 23 is open to the fastening area 5 and there has an entry area 33 for cooling fluid. The end of the first section 23 adjoining the inner ring 9 forms an outlet region 35 for cooling fluid from the first section 23, which then enters the second section 25. An impact cooling insert 37 is arranged in the first section 23. This impingement cooling insert 37 tapers conically from the inlet region 33 to the outlet region 35, so that three successive intersection surfaces F1, F2, F3 along the blade axis become smaller compared to one another along this direction. The impact cooling insert 37 is oriented so that it runs parallel to the front edge 8 of the airfoil with its insert front. It extends over the entire length of the leading edge 8 of the airfoil. The tapering of the impingement cooling insert 37 frees the first partial section 23 more and more in one direction from the inlet area to the outlet area. By means of a linearly oblique course of an insert rear side 41 of the impingement cooling insert 37 opposite the insert front side 39, the first partial section 23 is thus bisected, so to speak, in half into a half occupied by the impingement cooling insert 37 and a half free from the impingement cooling insert 37. The impingement cooling insert ^'37 has uniformly distributed impingement cooling holes 43. Air guiding ribs 51 surrounding the impingement cooling insert 37 are arranged on the inside of the airfoil area 7. These air guide ribs 51 extend transversely to the blade axis 3. At the same time, they are inclined with respect to a plane oriented perpendicular to the blade axis 3. The air guide ribs 51 each end before they enter the free part of the first section 23.

7 film cooling openings 53 are provided in the area of the airfoil trailing edge 10 in the airfoil area.

The impingement cooling insert 37 opens out in the area of the inner ring 9 at an inner ring cooling duct 55.

When the gas turbine guide vane 1 is used, it is arranged in a gas turbine (not shown) and hot gas flows around it. The high thermal load requires cooling by means of a cooling fluid 61, which is the

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Claims

claims

1. A gas turbine blade (1) directed along a blade axis (3) with a blade edge (8) and a blade edge (10) and with an internal cooling structure, comprising a meandering cooling duct (21) with sections (23) directed along the blade axis (3) 25, 27) for guiding a cooling fluid (61) from the airfoil leading edge (8) to the airfoil trailing edge (10), a first of the partial sections (23, 25, 27) along the

Blade front edge (8) extends and has an inlet region (33) for the cooling fluid (61) and an outlet region (35) for the cooling fluid (61), characterized in that the first section (23) has an impact cooling insert (37) with it the insert front face (39) directed towards the front airfoil (8) runs parallel to the front airfoil edge (8), the impingement cooling insert (37) tapering toward the outlet region (35).

2. Gas turbine blade (1) according to claim 1, wherein the impact cooling insert (37) covers the entire inlet area (33).

3. Gas turbine blade (1) according to claim 1, in which the impingement cooling insert (37) tapers in its cross-sectional area (F1, F2, F3) in proportion to a quantity of cooling fluid emerging from the impingement cooling insert (37) measured along the blade axis (3).

4. The gas turbine blade (1) according to claim 1, wherein the impact cooling insert (37) is surrounded by air guide ribs (51) directed transversely to the blade axis (3), the cooling fluid (61) emerging from the impact cooling insert (37) around the impact cooling insert (37) around towards the trailing edge of the airfoil (10).

5. The gas turbine blade (1) according to claim 4, in which the air guide ribs (51) are directed relative to a plane oriented perpendicularly to the blade axis (3) in such a way that they additionally move the cooling fluid (61) in a direction from the inlet area (33) to Guide the exit area (35).

6. The gas turbine blade (1) according to claim 1, which is designed as a guide blade with an inner ring (9), the inner ring (9) sealing a hot gas duct of the gas turbine against a rotor of the gas turbine when the guide blade is installed in a gas turbine, and wherein an inner ring cooling duct ( 55) leads from the impact cooling insert (37) to the inner ring (9).