Patent Application
20130058787
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 turbine. However, although such superalloys have considerably high corrosion and oxidation resistance, the high temperatures of the combustion gases in gas turbines require measures to improve corrosion and/or oxidation resistance further. 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 typically 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 certain 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 particular, cooling holes are present at the trailing edges of the airfoils as it is shown in U.S. Pat. No. 6,077,036, U.S. Pat. No. 6,126,400, US 2009/0194356 A1 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 trailing edge up to several millimetres towards the leading edge. This measure provides very thin trailing edges which can provide big improvements on the blading efficiency. An airfoil with a cutback design and a thermal barrier coating is, for example, disclosed in WO 98/10174 A1. 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 velocity 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 U.S. Pat. No. 6,126,400, U.S. Pat. No. 6,077,036 and, with respect to the coating method, in US 2009/0104356 A1.
However, in U.S. Pat. No. 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 A1 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 efficiency can not be achieved. In U.S. Pat. No. 6,126,400 the thermal barrier coating only covers about half the airfoil, as seen from the leading edge towards the trailing edge.
In US 2009/0104356 A1 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 the features of the independent claims. 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 extending 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 airfoil body comprises a step in the exterior 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 cast airfoil body in the coated surface region. 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 uncoated 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 applied 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 produce 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 between 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 chromium, 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 preferably corresponds to the combined thickness of the bond coat and the thermal barrier coating.
Furthermore, the inventive turbine airfoil is preferably hollow 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 particular, 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 location of the step in the outer surface.
An inventive turbine blade, which in particular is a gas turbine 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.
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
The airfoil body 13 is hollow and comprises, in the present embodiment, a number of interior cavities 11A to 11E 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 allowed to leave the internal cavities 11A to 11E through cooling 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 interior cavities 11A to 11D with the outside of the airfoil body 13 are not shown in the Figures. The internal cavity 11E 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 further, the thickness of the wall 17 of the airfoil body 13 is reduced at the suction side 7 of the airfoil in a region adjoining the trailing edge 5, as it is best seen in
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 corrosive combustion gases flowing along the airfoil 1 in operation of a gas turbine. The thermal barrier coating system 21 is best seen in
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 characterized 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.
According to the embodiment of
The boundary line is defined by a step 27 in the exterior surface of the cast airfoil body 13. In the present embodiment, the height h of the step 27 corresponds to the thickness 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 surface 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 present 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 exaggerated in the Figures in order to increase its visibility.
The invention has been described with reference to an exemplary embodiment of the invention for illustration purposes. However, deviations from the shown embodiment are possible. For example, additional uncoated surface regions may be present on the suction side and/or the pressure side of the airfoil. In addition, the thermal barrier coating system may deviate from the thermal barrier coating system used in the described embodiment. Furthermore, although the described airfoil has five internal cavities for allowing cooling fluid to flow there through the number of internal cavities may be larger or smaller than five.
| Number | Date | Country | Kind |
|---|---|---|---|
| 10154125.8 | Feb 2010 | EP | regional |
This application is the US National Stage of International Application No. PCT/EP2011/052169, filed Feb. 15, 2011 and claims the benefit thereof. The International Application claims the benefits of European application No. 10154125.8 EP filed Feb. 19, 2010. All of the applications are incorporated by reference herein in their entirety.
| Filing Document | Filing Date | Country | Kind | 371c Date |
|---|---|---|---|---|
| PCT/EP2011/052169 | 2/15/2011 | WO | 00 | 10/22/2012 |
