EP3658751B1 - Blade for a turbine blade - Google Patents

Description

The invention relates to an airfoil for a turbine blade according to the preamble of claim 1, for example in FIG DE 10 2016 123 525 A1 disclosed.

An airfoil corresponding to the preamble of claim 1 has long been known from the extensive prior art. The airfoil and in particular also the entire gas turbine blade are usually produced in an investment casting process, so that cavities are present in the interior of the airfoil. A coolant, mostly cooling air, can flow through these cavities so that the metallic material of the blade and the turbine blade can permanently withstand the high temperatures occurring during operation.

Different cooling concepts, which have long been known, are used for cooling, one of which is referred to as impingement cooling. With this, cooling air jets hit the inner surfaces of the metallic blade wall at an almost perpendicular angle in order to absorb the thermal energy contained therein and then to transport it away with them. The impingement cooling walls required to form the impingement cooling can on the one hand be cast with or on the other hand be provided by installing metal sheet metal inserts. However, due to the manufacturing process, cast impingement cooling systems require a minimum distance between the wall surface to be cooled and the impingement cooling wall having the impingement cooling openings, since the cast cores themselves required a minimum wall thickness for sufficient strength. If the perforated impact cooling wall is mounted as an insert in a blade, additional manufacturing and assembly steps are required, which increase the cost of manufacturing the turbine blade. In addition, on the one hand there can be leaks at the interface between the Impingement cooling insert and cast component and on the other hand signs of wear occur, which can affect the cooling efficiency or the service life.

The object of the invention is therefore to provide a long-lasting airfoil for a turbine blade which enables particularly efficient cooling of the side walls of the airfoil.

This object is achieved according to the invention by an airfoil according to claim 1. Advantageous developments of the device according to the invention are each the subject matter of the dependent claims and the following description.

The present invention proposes that in an airfoil for a turbine blade, comprising a suction-side side wall and a pressure-side side wall, which extend along a profile center line from a common leading edge to a common trailing edge and in a spanwise direction from a root-side end to a head-side end At least partially enclose the cavity, with a first perforated impact cooling wall for impingement cooling of the front edge and at least one further perforated impingement cooling wall for impingement cooling of a section of the suction-side and / or pressure-side side wall, the impingement cooling openings of the first impingement cooling wall and the impingement cooling openings of the at least one being provided along the span inside further baffle cooling wall are fluidically connected in series. In other words, cascaded impingement cooling inside the blade is proposed, with at least one further impingement cooling section, preferably two further impingement cooling sections, per side wall cascading downstream from a first impingement cooling at the leading edge on the suction side and / or pressure side.

The invention is based on the knowledge that series-connected impingement cooling (cascaded impingement cooling) allows the cooling air to be used multiple times and thus to make it more uniform to achieve the temperature distribution along the cross-section. The area of the airfoil which is thermally most heavily loaded, ie the area around the leading edge, is fed with the coolest cooling air and is impingely cooled in a first impingement cooling section. During the first impingement cooling, the cooling air is heated up for the first time and the blade temperature in the vicinity of the leading edge is reduced to a tolerable level. The heated cooling air is then passed into a downstream section of the airfoil and used there again for impingement cooling of the side wall, whereby the temperature of the side wall there is also lowered and the cooling air is heated again. This achieves an efficient use of cooling air, so that - compared with conventional blades - the saved cooling air can be used to increase the efficiency of the gas turbine.

Because the heated cooling air specifically achieves a lower cooling effect in subsequent sections, the thermal displacement across the blade cross-section can be reduced. This can reduce the thermo-mechanical load on the metallic airfoil, which can lead to an increased service life of the airfoil. Due to the fact that the series-connected impingement cooling has low cross-flow components in the spanwise direction, it is comparatively efficient.

An impingement cooling space is also provided between the relevant impingement cooling wall and the inside of the associated side wall, a collecting space being provided downstream of the impinging cooling space in question, which is immediately adjacent to the downstream further impingement cooling wall. The relative terms “upstream” and “downstream” refer to the direction of flow of the cooling air inside the airfoil, unless otherwise stated. The collecting spaces serve as cavities in which the coolant, which has been further heated after impingement cooling, can be collected on the one hand and it can be collected from the other through the impingement cooling openings of the subsequent ones Impact cooling wall can pass through for further impingement cooling. In the event that different flow cross-sections are locally present due to component tolerances along the span, the collecting spaces preferably extend in the spanwise direction over the entire length of the blade. As a result, the pressure in the collecting space can be made uniform.

Furthermore, the collecting space is partially delimited by a projection which is impact-cooled. For this purpose, outlet openings close to the side wall are arranged in a rib, which extends according to a cross-sectional plane from a rib end on the suction side to a rib end on the pressure side. With this configuration, a more uniform temperature of the suction-side and / or pressure-side side wall along the blade profile, that is to say from the front edge in the direction of the rear edge, can be achieved.

More preferably, a supply channel for supplying coolant for cooling the leading edge is provided between the first collecting space and the first impingement cooling space. This supply channel preferably extends over the entire span of the airfoil. It can, more preferably, taper from its foot-side end to the head-side end, so that, provided that the coolant is fed into the supply channel at the foot-side end, it has a larger flow cross-section at the foot-side end than at its head end. This takes into account the fact that, due to the presence of impingement cooling openings in the impingement cooling wall, the amount of coolant present in the supply channel decreases with increasing distance from the end on the foot side. The conical shape of the supply channel therefore leads to an equalization of the flow rate of the coolant along the spreading direction.

According to a further advantageous embodiment, at least one side wall of the blade is preferably on at least one additional baffle cooling wall is provided on both side walls. Thus, the first impingement cooling (the leading edge of the blade) is followed in series by the suction-side impingement cooling and the pressure-side impingement cooling, although the two other impingement cooling systems on both sides of the profile center line are connected in parallel.

It is also of particular advantage if, according to a further preferred embodiment, one of the two further impingement cooling spaces is arranged on the suction side and the other of the two further impingement cooling spaces is arranged on the pressure side and a separate collecting space is connected upstream of each of these two impingement cooling spaces. These can preferably be provided by providing a first separating rib. In this case, the coolant pressures required in the relevant collecting spaces can be set in accordance with the local thermal load on the suction-side and pressure-side side walls so that coolants are used efficiently and locally.

In addition, it is advantageous if a further cavity is provided between two collecting spaces arranged on both sides of the profile center line. This further cavity is preferably separated from the collecting spaces by two second separating ribs. Said cavity can be used on the one hand to reduce the size of the collecting spaces to a desired level if a certain flow velocity is to be achieved in the collecting spaces. On the other hand, the further cavity can also be used to guide a further coolant from a head end to a foot end of the airfoil if this coolant is only to be passed through the airfoil without absorbing thermal energy, if possible.

To avoid leakage of coolant within the blade, it is advantageous if this is monolithic, ie is designed in one piece. Such blades can be produced in particular by means of an additive process. An additive method is understood in particular to be what is known as SLM technology, which is known as "selective laser melting". This technology, also known as 3D printing technology, enables metal components to produce cavities and passage openings that are comparatively small and with exact dimensions compared to turbine blades manufactured using conventional castings.

Even if in the description or in the patent claims some terms are used in the singular or in connection with a numerical word, the scope of the invention for these terms should not be restricted to the singular or the respective numerical word. Furthermore, the words "a" or "an" are not to be understood as numerals, but as indefinite articles.

The above-described properties, features and advantages of the invention and the way in which they are achieved are explained in more detail in connection with the following description of the exemplary embodiments with reference to the following figures.

Here, the figures are shown only schematically, which in particular does not result in any restriction of the implementation of the invention.

Figur 1

eine Turbinenschaufel in einer perspektivischen schematischen Darstellung,

Figur 2

den Querschnitt gemäß der Schnittlinie II-II durch das Schaufelblatt der Turbinenschaufel gemäß Figur 1 als ein erstes Ausführungsbeispiel und

Figur 3

ein zweites Ausführungsbeispiel eines erfindungsgemäßen Schaufelblatts einer Turbinenschaufel.

Show it:

Figure 1

a turbine blade in a perspective schematic representation,

Figure 2

the cross section according to the section line II-II through the blade of the turbine blade according to Figure 1 as a first embodiment and

Figure 3

a second embodiment of an airfoil according to the invention of a turbine blade.

In the following, the technical features which are provided with the same reference symbols are those which have the same technical effect.

A turbine blade 10 relating to the invention is shown in perspective in FIG. The turbine blade 10 is according to FIG Figure 1 designed as a rotor blade. The invention can also be used in a guide vane, not shown, of a guide vane. The turbine blade 10 comprises a blade root 12, which is shaped like a fir tree in cross-section, and a platform 14 arranged thereon. The platform 14 is followed by a blade 16 which is aerodynamically curved. Whether or not the airfoil 16 is covered by a thermal protective layer is irrelevant to the invention. The airfoil 16 comprises a suction side wall 22 and a pressure side wall 24 which, in relation to a hot gas flowing around the airfoil 16, extend from a leading edge 18 to a trailing edge 20. A plurality of openings 28 for blowing out coolant are provided along the rear edge 20 and are separated from one another by webs 30 arranged in between. The blade 16 extends along a spanwise direction from a foot-side end 26 to a head-side end 27. When the turbine blade 10 shown is used in a gas turbine with an axial flow, the spanwise direction coincides with the radial direction of the gas turbine.

Figure 2 shows a sectional view through the airfoil 16 according to the section line II-II as a first embodiment of an airfoil 16 according to the invention, whereas Figure 3 shows a second embodiment for this. In both figures, only the area of the airfoil 16 on the leading edge side with respect to the hot gas of the gas turbine is shown, the rear part and the trailing edge of the airfoil 16 are not visible. The blade 16 and its pressure-side side wall 24 and suction-side side wall 22 extend - as already explained - from the leading edge 18 starting along a profile center line 32 to the rear edge. In the interior of the blade 16, a first perforated, ie provided with impingement cooling openings 42, impingement cooling wall 34 is arranged at a distance from the inner surface of the leading edge 18, so that a first impingement cooling space 36 is formed between them. A supply channel 38 is provided on the side of the first impingement cooling wall 34 opposite the first impingement cooling space 36. This is separated from the remaining cavity of the airfoil 16 by a first rib 40. The first rib 40 extends according to the cross-sectional plane from a rib end 37 on the suction side to a rib end 37 on the pressure side and has outlet openings 39 close to the side wall for the first impingement cooling space 36.

According to the exemplary embodiment shown, four impingement cooling openings 42 lying in the sectional plane are provided in the first impingement cooling wall 34 for surface cooling of the front edge 18 and the directly adjoining suction-side and pressure-side areas of the side walls 22, 24.

Looking along the profile center line 32 in the direction of the rear edge, the first rib 40 is followed by a first collecting space 44 which is separated from a second collecting space 48 by a second rib 46. The latter is also delimited by a third rib 50, so that a third collecting space 52 adjoins it further in the direction of the rear edge. The first collecting space 44 is delimited both on the suction side and on the pressure side by two further impact cooling walls 54. Impingement cooling openings 42 are also arranged in these, so that first further impingement cooling spaces 56 are provided with which corresponding sections of the suction-side and pressure-side side walls 22 and 24 can be impingement-cooled. Second further impingement cooling spaces 59 are separated from the second collecting space 48 by impingement cooling walls 55.

According to the exemplary embodiment shown, all further impingement cooling spaces 56, 59 are laterally delimited by projections 57 extending inward from the side walls 22, 24. Analogously to the first rib 40, the second and third ribs 46, 50 merge at their rib ends 37 into the suction-side and pressure-side side wall 22, 24 and have outlet openings 39 there near the side wall.

Due to the chosen arrangement of collecting spaces 44, 48, 52, impact cooling walls 34, 54, impact cooling spaces 36, 56, outlet openings 39 and projections 57, it is obvious that the coolant is used multiple impact cooling both on the suction side and on the pressure side in several successively connected impingement cooling arrangements reduce the temperatures of the blade walls 22, 24 to a desired level.

It goes without saying that not only the illustrated impingement cooling openings 42, outlet openings 39 are present, but that more of them are distributed along the span in the corresponding walls at the corresponding position, preferably lying in a row.

In detail, a coolant is fed to the supply channel 38 through an opening (not shown) in the turbine blade 10 during operation. There it is distributed over the span of the blade and flows through the individual impingement cooling openings 42 of the first impingement cooling wall 34, forming air jets. The air jets impinge in a known manner on the inner surface of the leading edge and cool it as intended. The coolant then flows through the outlet openings 39 of the first rib 40, after which it strikes the projections 57 with impinging cooling and is deflected by them into the first collecting space 44. From there it flows through the first and second further impact cooling walls 54, 55 for cooling the associated side wall sections. It passes from the first and second impingement cooling spaces 56, 59 through the outlet openings of the ribs 46, 50 into the subsequent collecting spaces 48, 52.

After the coolant has flowed through the cascaded impingement cooling arrangement described above, it reaches the collecting space 52. From there, the coolant can be fed in a known manner and further sections of the airfoil 16 are used for cooling. It is conceivable that on the one hand it is diverted into a kind of meander cooling and then blown out through the rear edge openings 28. It is also possible that the coolant from the interior of the airfoil 16 through film cooling openings (64, Fig. 3 ) is directed to the outside. The combination of both variants can also make technical sense.

Figure 3 shows an alternative embodiment of the turbine blade 10 according to the invention as a second embodiment. In analogy to Figure 2 are in Figure 3 the identical features are provided with the same reference numerals, so that only the differences from the first exemplary embodiment will be discussed below.

Compared to the first exemplary embodiment, separating ribs 58, 60 are provided in the interior of the airfoil 16. A first separating rib 58 extends between the rib 40 and the further rib 46 along the profile center line 32. The separating rib 58 divides the collecting space 44 into two collecting spaces 44a and 44b, of which the former is provided on the suction side and the latter on the pressure side. Two second separating ribs 60 extend along and thus quasi parallel to the profile center line 32 between the rib 46 and the rib 50, but one of them is arranged on the suction side and one on the pressure side.

Just as the first separating rib 58 divides the collecting space 44, the collecting space 48 is off Fig. 2 now divided into two collecting spaces 48a and 48b, however, due to the use of two second separating ribs 60, a further cavity 62 can be provided. The further cavity 62 can be used for different purposes. For example, it is suitable for conveying part of the coolant from the foot-side end 26 of the airfoil 16 to a head-side end 27 of the airfoil 16 without this coming into contact with the comparatively hot side walls 22, 24. In this way, comparatively cool cooling air can be provided at the head-side end 27 of the airfoil, which is particularly advantageous in the case of guide vanes. It is also conceivable that the cavity 62 is hermetically sealed in order to guide the cooling air guided in the sub-collection spaces 48a, 48b closer to the impingement cooling walls 54 and the impingement cooling openings 42 arranged therein.

The features described in the relevant exemplary embodiments and specified in the dependent claims can be combined with one another in any desired manner.

Overall, the invention thus relates to an airfoil 16 for a turbine blade 10, comprising a suction-side side wall 22 and a pressure-side side wall 24, which extend along a profile center line 32 from a common leading edge 18 to a common trailing edge 20 and in a spreading direction from a base end 26 a head end 27 extending at least partially enclosing a cavity, with a first perforated baffle cooling wall 34 provided with openings inside for impingement cooling of the leading edge 18 and at least one further perforated impingement cooling wall 54 for impingement cooling of a section of the suction-side and / or pressure-side vane wall 22, 24 is provided. In order to achieve particularly efficient cooling of the turbine blade, it is proposed that the impingement cooling openings 42 of the first impingement cooling wall 34 and the at least one second impingement cooling wall 54 are fluidically connected in series.

Claims (8)

1. Blade (16) for a turbine blade (10), comprising a suction-side side wall (22) and a pressure-side side wall (24) which enclose a cavity at least partially in a manner which extends along a profile centre line from a common leading edge (18) to a common trailing edge (20) and in a span width direction from a root-side end (26) to a top-side end (27),
   a first impingement cooling wall (34) which is provided with impingement cooling openings (42) for the impingement cooling of the leading edge (18) and at least one further impingement cooling wall (54, 55) which is also provided with impingement cooling openings (42) for the impingement cooling of a section of the suction-side and/or pressure-side side wall (22, 24) being provided along the span width in the interior, the impingement cooling openings (42) of the first impingement cooling wall (34) and the impingement cooling openings of the at least one second impingement cooling wall (54, 55) being connected in series in terms of flow, and an impingement cooling space (36, 56, 59) being provided between the relevant impingement cooling wall (34, 54, 55) and the inner side of the associated side wall (22, 24), and in each case one collecting space (44, 48, 52) being provided downstream of the relevant impingement cooling space (36, 56, 59), which collecting space (44, 48, 52) adjoins the downstream further impingement cooling wall (34, 54, 55) in a directly upstream manner,
   characterized
   in that the collecting space (44, 48) is delimited partially by a projection (57) which is impingement cooled, by outlet openings (39) which are close to the side wall being arranged in a rib (40, 46) which, in accordance with a cross-sectional plane, extends from a suction-side rib end (37) to a pressure-side rib end (37).
2. Blade (16) according to Claim 1,
   a supply duct (38) being provided between the first collecting space (44) and the first impingement cooling space (36).
3. Blade (16) according to Claim 2,
   in each case at least one further impingement cooling wall being provided on at least one side wall (22, 24) of the blade (16), preferably on the two side walls (22, 24).
4. Blade (16) according to Claim 3,
   one of the two further impingement cooling spaces being arranged on the suction side and the other one of the two further impingement cooling spaces being arranged on the pressure side, and a separate collecting space (44a, 44b, 48a, 48b) being connected upstream of each of them.
5. Blade (16) according to Claim 4,
   a further cavity (62) being provided between two collecting spaces which are arranged on both sides of the profile centre line.
6. Blade (16) according to one of the preceding claims which is monolithic.
7. Blade (16) according to Claim 6
   which is produced by means of an additive method.
8. Turbine blade (10) with a blade (16) according to one of the preceding claims.

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