Patent Application
20140030109
This invention relates to a gas-turbine compressor blade, in particular to an aircraft gas turbine or a stationary turbine, having an airfoil fastened to a blade root. The compressor blade has a leading edge or inflow edge, which is also called the blade front edge.
Aircraft gas turbines always face the problem that the compressor blades are subjected to heavy erosion due to sucked-in particles. These are for example grains of sand or the like which impact the compressor blades at high velocity.
As a result of the erosion occurring, it is necessary to replace or repair the compressor blades. Repair is, in the case of blisks in particular, very complex from the engineering viewpoint and also very cost-intensive.
The erosion-resistant materials known from the state of the art are unsuitable as materials for compressor blades. Erosion-resistant coatings too have proven in practice to be unsuitable, as they are only effective for very small particle sizes.
The object underlying the present invention is to provide an aircraft gas-turbine compressor blade which, while being simply designed and easily and cost-effectively producible, has a high erosion resistance.
It is a particular object of the present invention to provide solution to the above problematics by a combination of the features of Claim 1. Further advantageous embodiments of the invention become apparent from the sub-claims.
It is provided in accordance with the invention that at least the area of the leading edge of the compressor blade is made from a material which has a very low modulus of elasticity. A suitable and low modulus of elasticity is in the range of 50 GPa.
The compressor blade in accordance with the invention thus has, due to the low modulus of elasticity, a relatively high elasticity in the area of the leading edge, so that an impact of a foreign object, for example a sand particle or stone, does not lead to damage thanks to the elasticity of the material in the area of the leading edge. In particular, zero or only very low erosion occurs when compared with compressor blades known from the state of the art.
In accordance with the invention, it is particularly advantageous when at least the area of the leading edge is made from a titanium alloy of types Ti-28Nb-1Fe-0.5Si, Ti-36Nb-2Ta-3Zr-0.3O or Ti-24Nb-4Zr-8Sn. By means of an alloy of this type, it is for example possible to reduce erosive wear by about 50% compared with compressor blades made from a conventional alloy, for example Ti64. Compared with compressor blades known from the state of the art made from nickel-based alloys, for example IN718, erosive wear can be reduced by 30%.
It is possible in accordance with the invention to either manufacture the entire compressor blade from the above titanium alloy with the low modulus of elasticity or only to manufacture parts of the compressor blade from this material.
In a particularly favourable embodiment of the invention, it is provided that the leading edge is designed in the form of a leading-edge element which is connected to the airfoil of the compressor blade by means of a welding process. Particularly advantageous is the use of a laser welding process.
The airfoil itself can, in accordance with the invention, be made from one of the conventional titanium alloys, for example Ti64, Ti6246 or Ti6242.
If a separate leading-edge element is provided, the latter extends preferably up to just in front of the blade root, and in the direct transition of the blade root it is not necessary to apply or use the alloy in accordance with the invention.
In the case of separate manufacture of the leading-edge element in accordance with the invention and connection by means of a laser welding process, there is the advantage that due to the thermal influence during the laser welding operation a soft transition is obtained between the low modulus of elasticity of the leading-edge element and the higher modulus of elasticity of the material in the rest of the airfoil. The result is thus a gradual transition of the elastic properties, which acts advantageously on the behaviour of the entire compressor blade component. This effect is based on the fact that during heating of the low-modulus titanium alloy in temperature ranges above around 500° C., the modulus of elasticity rises again to values of conventional titanium alloys, and hence the direct area of the weld is homogeneous.
The compressor blade in accordance with the invention, which can be designed as a single compressor blade, as a blisk blade or as a fan blade, is thus to a high degree resistant against erosion and damage from foreign objects.
The invention thus also relates to a method for the manufacture of a compressor blade in which a leading-edge element is manufactured separately and is connected to the airfoil by means of a welding process, in particular by a laser welding process, where the leading-edge element is made from a low-modulus (in respect of the modulus of elasticity) titanium alloy, for example from Ti-28Nb-1Fe-0.5Si, from Ti-36Nb-2Ta-3Zr-0.3O or from Ti-24Nb-4Zr-8Sn and/or an alloy with a modulus of elasticity of substantially 50 to 70 GPa and where the airfoil is made from an alloy known from the state of the art.
The invention also relates to the use of a low-modulus (of elasticity) titanium alloy, for example Ti-28Nb-1Fe-0.5Si or Ti-36Nb-2Ta-3Zr-0.3O or Ti-24Nb-4Zr-8Sn, in particular for the leading-edge area of a compressor blade.
The present invention is described in the following in light of the accompanying drawing, showing an exemplary embodiment. In the drawing,
The gas-turbine engine 10 in accordance with
The intermediate-pressure compressor 13 and the high-pressure compressor 14 each include several stages, of which each has an arrangement extending in the circumferential direction of fixed and stationary guide vanes 20, generally referred to as stator vanes and projecting radially inwards from the engine casing 21 in an annular flow duct through the compressors 13, 14. The compressors furthermore have an arrangement of compressor rotor blades 22 which project radially outwards from a rotatable drum or disk 26 linked to hubs 27 of the high-pressure turbine 16 or the intermediate-pressure turbine 17, respectively.
The turbine sections 16, 17, 18 have similar stages, including an arrangement of fixed stator vanes 23 projecting radially inwards from the casing 21 into the annular flow duct through the turbines 16, 17, 18, and a subsequent arrangement of turbine blades 24 projecting outwards from a rotatable hub 27. The compressor drum or compressor disk 26 and the blades 22 arranged thereon, as well as the turbine rotor hub 27 and the turbine rotor blades 24 arranged thereon rotate about the engine axis 1 during operation.
| Number | Date | Country | Kind |
|---|---|---|---|
| 10 2012 015 137.3 | Jul 2012 | DE | national |
