The present invention relates to turbine airfoils, i.e. to rotating blades or vanes, in particular for heavy-duty industrial gas turbines and cooling methods therefore, as well as to turbines equipped with such airfoils.
In the field of heavy duty industrial gas turbine engines it is important to ensure that the component exposed to the hot gas flow, in particular downstream of the combustion chamber, is at a temperature level which does not harm the corresponding components. Therefore rotating or stationary gas turbine airfoils, typically made of or at least based on metal, have to be cooled internally. To this end they comprise cooling channels provided in the air foils which are supplied with cooling air typically discharged from the compressor end exit. On the one hand cooling is effected by circulation of this cooling air in these internals channels, on the other hand by bores provided in the wall structure of the air foil leading to a blowing-out of the cooling medium and a film cooling at the location of the exits of the cooling hole and downstream thereof.
In particular the air foil trailing edge is required to be maintained at a low metal temperature.
In an embodiment, the present invention provides a turbine blade or vane including at least one internal radial channel for the circulation of cooling medium bordered on a pressure side by a pressure side wall and on a suction side by a suction side wall joined at a upstream side at a leading edge and at and downstream side at the trailing edge. At least one exit hole extends through at least one of the pressure side wall or the suction side wall for blowing out of cooling medium from the internal radial channel to a medium surrounding the blade or vane. At least one trailing edge exit hole along the trailing edge has a surfacial exit opening disposed at the pressure side of the trailing edge.
Exemplary embodiments of the present invention are described in the following with reference to the drawings, in which:
In an embodiment, the present invention provides an improved cooling scheme for rotating airfoils or stationary airfoils of heavy-duty industrial gas turbines in particular. In particular an improved scheme for film cooling in the trailing edge region of such airfoils shall be provided.
Specifically, the proposed turbine blade or vane comprises at least one internal radial channel, typically if there are several, these are separated from each other by radially extending separation walls, for the circulation of cooling medium. These cooling medium channels are bordered on a pressure side of the airfoil by a pressure side wall and on a suction side of the airfoil by a suction side wall, respectively joined at an upstream side at a radially extending leading edge of the blade/vane and at a downstream side at a radially extending trailing edge of the blade/vane, wherein the turbine blade or vane typically comprises at least one exit hole (so-called film cooling holes) through at least one of pressure side wall or suction side wall or the tip of the blade for the blowing out of cooling medium from the internal channel to the medium surrounding the blade or vane, i.e. to the surrounding hot gas machine airflow.
In accordance with an embodiment of the present invention, this structure is characterised in that along the trailing edge there is at least one trailing edge exit hole the surfacial exit opening of which is located at the pressure side of the trailing edge.
According to a preferred embodiment of the turbine blade or vane the trailing edge exit hole blows out cooling air to the medium surrounding the blade or vane under an angle α with respect to the pressure side wall surface at the blowing-out point, which preferably is in the range of 5-45°, more preferably in the range of 5-30°. In other words the cooling airflow takes place not parallel to the hot gas stream but is somewhat directed into the hot gas stream at the point of exit of the hole.
According to yet another preferred embodiment, the trailing edge side of the surfacial exit opening of the trailing edge exit hole is located close to the trailing edge. This means that it is preferably located not more than 50 mm, more preferably not more than 30 mm, most preferably not more than 10 mm upstream of the trailing edge along the pressure side wall surface. It is however normally not located at the trailing edge so the exit opening it is not along the line of the trailing edge or touching the line of the trailing edge.
According to another preferred embodiment, along the trailing edge and along a radial direction at least two, preferably at least four, trailing edge exit holes are located supplied via individual cooling medium bores connecting the trailing edge exit holes to the internal radial channel. Typically the holes are distributed equally along the trailing edge the distance being the pitch of the row of holes. This pitch, expressed as the ratio of the distance P of the centres of adjacent holes to the diameter d of the holes along the edge is typically in the range of P/d=2-8 for a typical blade in the field of heavy duty industrial gas turbines.
Preferably and according to yet another preferred embodiment, at least one of the bores and/or trailing edge exit holes is inclined with respect to an axial direction of the machine. This can be with a positive or negative angle β which is preferably in the range of 0-50°, more preferably in the range of 10-40°. Preferentially all the bores and/or trailing edge exit holes are inclined with the same angle, preferably with a positive angle β being defined as radially outwards in a downstream direction.
Preferentially the trailing edge exit hole comprises a bore connecting the internal radial channel with the medium surrounding the blade or vane so basically penetrating the wall structure of the blade, and the bore comprises, on its side connecting to the internal radial channel, a circular cylindrical section, and on its side to the surface of the blade or vane a widening section conically widening towards the surface of the blade or vane, wherein preferably the ratio of the length of the circular cylindrical section to the total length of the circular cylindrical section and the widening section is in the range of 0.2-0.7, preferably in the range of 0.2-0.5.
The widening can be in a fully circular manner, i.e. in the sense that the diameter of the circular cross section is gradually increasing towards the surfacial exit hole. On the other hand, and this is preferred, the conical widening can be such that in a direction perpendicular to the plane of the surface of the pressure side wall, the diameter stays constant, while it increases in a direction parallel to the plane of the surface of the pressure side wall. Like that the cross-section becomes increasingly oval or racetrack shaped with an increasing ratio of the long axis to the short axis along and towards the exit of the hole. This fan like widening leads to a particularly good and efficient spreading of the cooling air over the surface of the blade.
Yet another preferred embodiment is characterised in that the trailing edge exit hole comprises a bore connecting the internal radial channel with the medium surrounding the blade or vane, and in that the ratio of the length L of the bore to the diameter d of the bore is in the range of L/d=5-50, preferably in the range of L/d=20-40.
Such a blade typically comprises at least one radial leading-edge cooling passage located closest to the leading edge, at least one intermediate cooling passage as well as at least one trailing edge cooling passage located closest to the trailing edge, and the trailing edge exit hole is supplied by the trailing edge cooling passage, which preferably itself is supplied by cooling medium flow in a radially outward direction by a meander or serpentine type cooling medium circulation within the blade through the further cooling passages.
Preferentially, and according to another preferred embodiment of the invention, at the trailing edge the pressure sidewall of the blade/vane comprises a step recessed towards the suction side. This step can be a casted slot. In this case, at least one trailing edge exit hole can for example at least partly open towards the surrounding medium in the region of this step, wherein preferably at least part of, preferably at least the totality of the surfacial opening of the trailing edge exit hole is located in a radially extending leading-edge surface of the step. This leading-edge surface of the step is particularly preferably at an angle in the range of 60-120°, more preferably in the range of 75-105° with respect to a radially extending bottom surface of the step, wherein most preferably the leading-edge surface is oriented essentially perpendicularly to the hot gas flow on the pressure side and the bottom surface essentially parallel to the hot gas flow on the pressure side.
The trailing edge exit hole can be supplied by cooling medium via a bore which fully opens in a radially extending leading-edge surface of the step and which is distanced from the bottom surface of the step, expressed as a function of the length T of the step along the gas flow direction and the diameter d of the bore, and as a function of the depth t of the step essentially perpendicularly to the gas flow direction and the diameter d of the bore in the range of T/d=8-12, preferably T/d=9-11 or around T/d=10, and in the range of t/d=1.0-1.8 preferably t/d=1.3-1.7, or around t/d=1.5.
The cross-section of the bore, in particular at the point of exit, be it in such a step or just on the pressure side of the blade/vane, can be circular, oval, elliptical or racetrack shaped, preferably in the latter cases with the long axis aligned along a radial direction.
The trailing edge exit hole can, in the alternative, be supplied by cooling medium via a bore which only partly opens in a radially extending leading-edge surface of the step and which at least partly, preferably over the full length, channels through the bottom surface of the step forming scarfed holes.
Typically such a blade/vane is at least partly based on metal and/or ceramics, coated or uncoated, and it is a rotating or stationary turbine aerofoil.
Furthermore the present invention pertains to a turbine, preferably a gas turbine with a turbine blade as outlined and defined above.
As pointed out above, in an embodiment, the present invention provides a design of film cooling holes which are aligned with the pressure side of the trailing edge, and which can significantly reduce the metal temperatures of the airfoil, thereby extending the component lifetime. In the following several different concepts to implement this general scheme will be shown and discussed.
Within this hollow profile there are located separating walls 10, extending radially between the foot of the blade and the tip of the blade, separating the above-mentioned individual cooling passages from each other. Typically the separating walls 10 extend between the two side walls on the pressure side 8 and the suction side 9, respectively, and can either be, as illustrated in
As illustrated schematically in
A second cooling airflow is fed into channel 2, the more leading edge oriented of the two intermediate cooling passages, at the foot of the blade and also travels radially outwards through channel 2 as illustrated schematically by arrow 14. In this case, as there are no film cooling holes, in the tip portion of the blade 5 there is a passage between the intermediate cooling passage 2 and the intermediate cooling passage 3, so at the tip portion there is one or a series of holes in the separating wall 10 separating these two channels 2, 3, such that the cooling air passes, as illustrated schematically by arrow 16, to the intermediate cooling passage 3 at the trailing edge side and then travels radially inwards towards the axis of the machine as illustrated schematically by arrow 15. At the foot of the blade 5, this cooling airflow stream is again redirected through a hole or a series of holes in the separating walls 10 between channels 3 and 4 and enters the trailing edge cooling passage at the trailing edge side on the foot thereof. It then again travels upwards in a radial direction towards the tip of the blade and cools the walls bordering the trailing edge cooling passage 4 from the interior side, as this is illustrated by arrows 17 and 18.
Correspondingly therefore, the cooling medium follows, in a meander or serpentine type fashion, the arrows 14-18 through the channels 2-4.
According to an embodiment of the present invention, the cooling airflow 18 travelling in the trailing edge cooling passage 4 at least partly exits in the region of the trailing edge 7 via one or a series of trailing edge exit holes 22, so via a trailing edge coolant ejection 21.
In accordance with this embodiment as illustrated in
The bore 44 is thereby arranged at an angle α with respect to the pressure side wall surface plane at the trailing edge, as schematically illustrated by line 19 in
This pressure side bleed ejection of the coolant flow enables the air foil to operate at a higher inlet hot gas temperature, while maintaining the same (or lower) cooling air consumption relative current operating hot gas temperature.
To summarize in
In
Furthermore the holes are distanced in a radial direction by a pitch P. In addition to that the channels and also the exit holes 22 are not aligned along the axis but are inclined, in a direction radially downstream outwards as illustrated in
Furthermore the actual exit holes 22 are specifically structured in a widening manner as will be illustrated in more detail by using the illustration of
As one can see in
The widening can be realized, as this is specifically illustrated in Figures a and b, by only widening in a direction essentially radial with respect to the machine, so the widening is only visible in the illustration b), while in the illustration a) there is no widening within the section 27. However there can also be widening, in the sense of a full tubular widening along both directions.
The widening as illustrated in
A different embodiment of the invention is shown in
On the other hand, and arranged essentially perpendicularly to this surface 45, there is the bottom surface 35 of the step 34, which is arranged essentially parallel to the chord line of the blade and which, in this case, approximately half the full width of the blade in this very terminal section at the trailing edge 7.
According to this embodiment, the bore 44 of the trailing edge coolant ejection 21 terminates in the above-mentioned leading edge side surface 45 and thus enters the step 34. As illustrated in
In the alternative, and as illustrated in
Yet another embodiment with such a step 34 is illustrated in
While the invention has been particularly shown and described with reference to preferred embodiments thereof, it will be understood by those skilled in the art that various changes in form and details may be made therein without departing from the spirit and scope of the invention.
This application claims priority to International Patent Application No. PCT/EP2011/053831, filed Mar. 15, 2011, and U.S. Provisional Patent Application No. 61/315,470, filed Mar. 19, 2010, which are hereby incorporated by reference herein in their entirety.
Number | Date | Country | |
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61315470 | Mar 2010 | US |
Number | Date | Country | |
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Parent | PCT/EP2011/053831 | Mar 2011 | US |
Child | 13622055 | US |