Turbine Airfoil The present invention relates to a turbine airfoil which can be used in a gas turbine vane or blade.

The airfoils of gas turbines are typically made of nickel or cobalt based superalloys which show high resistance against the hot and corrosive combustion gases present in gas tur ^¬ bine. However, although such superalloys have considerably high corrosion and oxidation resistance, the high temperatures of the combustion gases in gas turbines require meas ^¬ ures to improve corrosion and/or oxidation resistance fur- ther. Therefore, airfoils of gas turbine blades and vanes are typically at least partially coated with a thermal barrier coating system to prolong the resistance against the hot and corrosive environment. In addition, airfoil bodies are typi ^¬ cally hollow so as to allow a cooling fluid, typically bleed air from the compressor, to flow through the airfoil. Cooling holes present in the walls of the airfoil bodies allow a cer ^¬ tain amount of cooling air to exit the internal passages so as to form a cooling film over the airfoil surface which further protects the superalloy material and the coating applied thereon from the hot and corrosive environment. In particu ^¬ lar, cooling holes are present at the trailing edges of the airfoils as it is shown in US 6,077,036, US 6,126,400,

US 2009/0194356 Al and WO 98/10174, for example. Trailing edge losses are a significant fraction of the over all losses of a turbomachinery blading. In particular, thick trailing edges result in higher losses. For this reason, cooled airfoils with a cutback design at the trailing edge have been developed. This design is realised by taking away material on the pressure side of the airfoil from the trail ^¬ ing edge up to several millimetres towards the leading edge. This measure provides very thin trailing edges which can pro- vide big improvements on the blading efficiency. Airfoils with a cutback designs and thermal barrier coatings are, for example, disclosed in WO 98/10174 Al and EP 1 245 786 A2. However, the beneficial effect on the efficiency can only be achieved if the thickness of the trailing edge is rather small. On the other hand for a blade with thermal barrier coating, the combined thickness of the cast airfoil body wall and the applied thermal barrier coating system exceeds the optimum thickness of the design. In addition, as the flow ve- locity of the gas is the greatest at the trailing edge of the airfoil a thermal barrier coating applied to the trailing edge is prone to high levels of erosion.

It is known to selectively provide a thermal barrier coating system to the airfoil, in particular such that the trailing edge of an airfoil and adjacent regions of an airfoil remain uncoated. Selective coatings are, for example, described in US 6,126,400, US 6,077,036, WO 2005/108746 Al and, with re ^¬ spect to the coating method, in US 2009/0104356 Al .

However, in US 6,077,036 the pressure side of the airfoil is completely uncoated which means that areas which would not suffer from a higher combined thickness of the cast airfoil body and the coating applied thereon remain unprotected against the temperature the hot combustion gas.

WO 2008/043340 Al describes a turbine airfoil with a thermal barrier coating the thickness of which varies over the airfoil surface. However, like in WO 98/10174 the trailing edge is fully coated so that the beneficial effect on blading ef ^¬ ficiency can not be achieved. EP 1 544 414 Al shows a turbine airfoil with a thermal barrier coating the thickness of which varies over the airfoil surface with the trailing edge not being fully coated. In US 6,126,400 the thermal barrier coat- ing only covers about half the airfoil, as seen from the leading edge towards the trailing edge. In US 2009/0104356 Al the method of masking the trailing edge will produce a step in the coating which adversely affects the aerodynamics of the blade. With respect to the mentioned prior art it is an objective of the present invention to provide an improved airfoil and an improved turbine blade or vane.

These objectives are solved by a turbine airfoil as claimed in claim 1 and by a turbine vane or blade as claimed in claim 9. The depending claims contain further developments of the invention .

An inventive turbine airfoil comprises an airfoil body with a leading edge, a trailing edge and an exterior surface. The exterior surface includes a suction side extending from the leading edge to the trailing edge and a pressure side extend ^¬ ing from the leading edge to the trailing edge and being located opposite to the suction side on the airfoil body. The turbine airfoil further comprises a thermal barrier coating system present in a coated surface region, and an uncoated surface region where a thermal barrier coating system is not present. This uncoated surface region extends on the suction side from the trailing edge towards the leading edge to a boundary line located on the suction side between the leading edge and the trailing edge, in particular closer to the trailing edge than to the leading edge. The boundary line may, in particular, mainly extend in radial direction of the airfoil body. The airfoil body comprises a step in the exte- rior surface. This step extends along the boundary line. In particular, the step may be formed such that the surface of the uncoated surface region lies higher than the surface of the typically cast airfoil body in the coated surface region, i.e. when looking along the suction side surface of the un- coated airfoil body from the leading edge towards the trail ^¬ ing edge the step leads to an increased distance from a chord line of the airfoil body profile as compared to a suction side surface without such a step. The height of the step is preferably equal to the thickness of the thermal barrier coating system. "Higher" is meant in a sense that, in relation to a point or a plane located inside of the airfoil, a "higher" exterior surface has a larger distance to the point or plane than a second exterior surface. As a result, the surface which is not higher could be considered as a depression in comparison to the "higher" surface.

The present invention allows to produce very thin trailing edges without thermal barrier coating systems applied thereon and at the same time minimizing or even avoiding a step at the boundary between the coated surface region and the un- coated surface region. This step is minimized or avoided by providing the mentioned step in the surface of the airfoil body. By choosing the height of the step such that it matches the thickness of the thermal barrier coating system to be ap- plied to form the coated surface region the surface of the applied coating in the coated region can be made to match the surface of the uncoated surface region. This allows to pro ^¬ duce a finished surface of the partly coated airfoil which matches the designed definition in the coated surface region as well as in the uncoated surface region. Moreover, since there is no thermal barrier coating at the trailing edge an adverse effect on the airfoil lifetime due to high levels of erosion of the thermal barrier coating at the trailing edge does not occur.

The thermal barrier coating system may, in particular, comprise a thermal barrier coating and a bond coat located be ^¬ tween the thermal barrier coating and the exterior surface of the airfoil body. Typical bond coats are aluminium oxide forming materials, in particular, so called MCrAlY-coatings , where M stands for cobalt and/or nickel, Cr stands for chro ^¬ mium, Al stands for aluminium, and Y stands for yttrium and/or one or more rare earth elements. In case of a coating system including a bond coat the height of the step prefera ^¬ bly corresponds to the combined thickness of the bond coat and the thermal barrier coating.

Furthermore, the inventive turbine airfoil is preferably hol ^¬ low and comprises at least one cooling opening, in particular realised by a cutback design, at the trailing edge. In this way, the trailing edge can be made particularly thin, if the hollow airfoil body comprises a wall the thickness of which is less in the uncoated surface region than in the coated surface region. The thickness of the wall region can, in par ^¬ ticular, decrease over a small transition region on one or both sides of the boundary line. This avoids having a step at the inner surface of the airfoil body at or close to the lo ^¬ cation of the step in the outer surface.

An inventive turbine blade, which in particular is a gas tur ^¬ bine vane or blade, comprises an inventive turbine airfoil. The use of an inventive airfoil allows for producing highly efficient turbomachinery bladings.

Further features, properties and advantages of the present invention will become clear from the following description of an embodiment in conjunction with the accompanying drawings.

Figure 1 schematically shows the structure of the inventive airfoil . Figure 2 shows the trailing edge of the airfoil shown in

Figure 1.

Figure 3 shows a detail of Figure 2. An inventive turbine airfoil may be part of a turbine blade or a turbine vane. Turbine blades are fixed to a rotor of the turbine and rotate together with the rotor. They are adapted for receiving momentum from the flowing combustion gas produced by a combustion system. The turbine vanes are fixed to the turbine casing and form nozzles for guiding on the combustion gases so as to optimize the momentum transfer to the rotor blades. The inventive turbine airfoil can, in general, be used in turbine blades as well as in turbine vanes.

An inventive airfoil 1 is shown in Figure 1. It comprises a cast airfoil body 13, a leading edge 3 at which the flowing combustion gases arrive at the airfoil 1 - the leading edge 3 being the upstream edge - and a trailing edge 5 at which the combustion gases leave the airfoil 1 - the trailing edge 5 being the downstream edge. The exterior surface of the airfoil 1 is formed by a convex suction side 7 and a less con- vex, and typically concave, pressure side 9 which is formed opposite to the suction side 7. Both the suction side 7 and the pressure side 9 extend from the leading edge 3 to the trailing edge 5 and are formed by the exterior surface of the airfoil body wall, i.e. from the surface of the wall showing away from the interior of the airfoil body.

The airfoil body 13 is hollow and comprises, in the present embodiment, a number of interior cavities 11A to HE to allow a cooling fluid, typically bleed air from a compressor of the turbine engine, to flow there through and to cool the airfoil body 13. Moreover, a certain amount of cooling fluid is al ^¬ lowed to leave the internal cavities HA to HE through cool ^¬ ing holes present in the wall of the airfoil body 13 towards its exterior surface so as to form a cooling fluid film over the surface. Note that the cooling holes connecting the inte ^¬ rior cavities HA to 11D with the outside of the airfoil body 13 are not shown in the Figures. The internal cavity HE which is closest to the trailing edge 5 comprises a slit 15 which allows cooling fluid to leave this cavity close to the trailing edge 5. The slit 15 is formed by a cut back in the pressure side 9 of the airfoil 1. This may be done to reduce losses due to a blockage at the trailing edge 5 and, hence, to increase efficiency of the turbomachinery bladings. The loss reducing effect is caused by the decreased thickness of the trailing edge due to the cutback design. In order to reduce the thickness of the trailing edge 5 fur ^¬ ther, the thickness of the wall 17 of the airfoil body 13 is reduced at the suction side 7 of the airfoil in a region ad ^¬ joining the trailing edge 5, as it is best seen in Figure 2. Figure 2 shows the trailing edge 5 of the airfoil 1 and adja- cent airfoil regions. It can be seen that the suction side 7 comprises a thin airfoil region 19 which extends from the trailing edge 5 over a certain length of the airfoil profile towards the leading edge 3. The airfoil body 13 is cast from a high temperature resistive nickel based or cobalt based superalloy and covered with a thermal barrier coating system which reduces corrosion of the airfoil body 13 which would occur due to the hot and corro ^¬ sive combustion gases flowing along the airfoil 1 in opera- tion of a gas turbine. The thermal barrier coating system 21 is best seen in Figure 3 which shows a detail of Figure 2 in the transition region between the regular airfoil body wall 17 and the thin airfoil region 19. The thermal barrier coat ^¬ ing system 21 comprises the actual thermal barrier coating 23, for example zirconium oxide which is at least partially stabilized by yttrium oxide, and a bond coat 25 located be ^¬ tween the surface of the superalloy material the airfoil body 13 is made of and the thermal barrier coating 23. The bond coat is typically an aluminium oxide forming material, in particular an MCrAlY-coating .

A certain minimum wall thickness of the airfoil body wall 17 is necessary for applying such a thermal barrier coating system 21 to the airfoil body 13 so that a coated wall is char- acterized by a minimum thickness. This minimum thickness is, however, thicker than the desired thickness of the thin airfoil region 19. Therefore, no thermal barrier coating system 21 is applied to the thin airfoil region 19 so that the thin airfoil region 19 coincides with an uncoated airfoil region 29 which extends from the trailing edge 5 to a boundary line located between the trailing edge 5 and the leading edge 3, in particular closer to the trailing edge 5 than to the leading edge 3. Typically, the uncoated surface region does not extend over more than 10 to 30 % of the distance between the trailing edge 5 and the leading edge 3. However, the exact distance over which the uncoated surface region 29 extends depends on the actual airfoil design.

The boundary line mainly extends in radial direction of the airfoil body 13, i.e. in a direction from root of the blade or vane towards the tip of the blade or vane. This direction is perpendicular to the plane of the aerodynamic profile shown in the Figures. However, the boundary line does not need to be a straight line but can also be slightly curved so that the distance of the boundary line from the trailing edge 5 varies with the radial position on the suction side sur- face.

According to the embodiment of Figure 2, the uncoated surface region is only present on the suction side 7 and close to the trailing edge 5.

The boundary line is defined by a step 27 in the exterior surface of the cast airfoil body 13. In the present embodi ^¬ ment, the height h of the step 27 corresponds to the thick ^¬ ness of the thermal barrier coating system 21 and is designed such that the surface 33 of the thin airfoil region 19 lies higher than the surface 28 of the airfoil body 13 in the sur ^¬ face region to become coated.

Before the thermal barrier coating system 21 is applied to the surface of the cast airfoil body 13 the suction side 7 is masked between the step 27 and the trailing edge 5 to prevent coating material from adhering to the thin airfoil region 19 which shall become the uncoated airfoil region 29. After the thermal barrier coating system 21 has been applied to the exterior surface of the cast airfoil body 13 and the mask has been removed the surface 31 of the uncoated surface region, the surface of the thermal barrier coating system 21 is smoothly aligned with the surface 33 of the uncoated surface region 29. Hence, no step which could lead to losses is pre ^¬ sent between the coated surface region 30 and the uncoated surface region 29 of the airfoil suction side 7. In addition, as the thin airfoil region 19 between the boundary line and the trailing edge 5 is free from thermal barrier coating not only a very thin trailing edge 5 is achieved but also erosion of the coating due to the high velocities of the combustion gases at the trailing edge 5 are avoided.

To avoid a weak area in the wall 17 of the airfoil body 13 the transition between the regular airfoil body wall 17 and the thin airfoil region 19 is not realised in form of a step but in form of a region in which the thickness of the regular wall 17 gradually decreases from the normal thickness to the thickness of the thin airfoil region 19. In this context, please note that the thickness of the thermal barrier coating system 21 and, hence, the height h of the step 27, is exag ^¬ gerated in the Figures in order to increase its visibility.

The invention has been described with reference to an exem ^¬ plary embodiment of the invention for illustration purposes. However, deviations from the shown embodiment are possible. For example, additional uncoated surface regions may be pre- sent on the suction side and/or the pressure side of the air ^¬ foil. In addition, the thermal barrier coating system may deviate from the thermal barrier coating system used in the described embodiment. Furthermore, although the described air ^¬ foil has five internal cavities for allowing cooling fluid to flow there through the number of internal cavities may be larger or smaller than five.