Patent Application
20150345297
This application claims priority to European application 14160866.1 filed Mar. 20, 2014, the contents of which are hereby incorporated in its entirety.
The invention relates to a gas turbine blade comprising an airfoil extending in radial direction from a blade root to a blade tip, defining a span ranging from 0% at the blade root to 100% at the blade tip, and extending in axial direction from a leading edge to a trailing edge, which limit a chord with an axial chord length defined by an axial length of a straight line connecting the leading edge and trailing edge of the airfoil depending on the span. Generally, the gas turbine blade according to the present invention is not restricted to a gas turbine: rotor blades or guide vanes of a turbo-machinery fall legally under the present invention.
The design of rotor blades in a gas turbine engine is of vital importance in terms of efficiency with which the gas flow passing through the gas turbine engine interacts with the blades especially of the at least one turbine of the gas turbine arrangement.
Rotating gas turbine blades must fulfill a multitude of material- and design criteria which consider high mechanical and thermal stresses acting onto the rotating blades during operation. Due to enormous centrifugal forces acting onto rotating blades and an enormous thermal load that must withstand the blades, the main task in the design work of blades is to combine a high degree on stiffness which shall avoid blade vibrations during operation and the possibility of active cooling to enhance load capacity, by providing cooling channels inside the airfoil of a rotating blade. In view of the before requirements an optimized airfoil shape is always sought to improve turbine aerodynamic efficiency.
Rotating blades are arranged in rows which alternate in axial direction with rows of stationary vanes. Every pair a rows including one row of stationary vanes and one row of rotating blades which follows in downstream direction directly forms a so called stage. All stages of the turbine are numbered in sequence beginning with the first stage at the inlet opening of the turbine comprising the first row of stationary vanes followed by the first row of rotating blades.
Normal operation of a gas turbine shows that the stationary vanes, e.g. of the first stage, are excitation sources for vibrations acting onto the following rotating blades in downstream direction in disadvantage manner. It is therefore an object of turbine development to reduce such excitation sources and/or to enhance possibilities of decoupling mechanism to reduce and/or to avoid vibration transmission and excitation onto rotating blades arranged downstream of vanes in the first stage.
An obvious intervention would mean to change the excitation sources itself, but a change of the vanes in the first stage is considered to be expensive and would raise a lot of development work. Proposals to vary the radial length of the blades, i.e. the span of the airfoil which extends from the blade root to the blade tip, would have an impact onto the annulus of the flow path through the turbine which would lead to a major impact on a developments schedule which in view of that is not favorable. Another approach of reducing the tip mass of the rotating blade by reducing the axial chord length of the tip chord, which concerns a straight line connecting the leading edge and trailing edge of the airfoil in the region of the blade tip, resulted in aerodynamic penalty and furthermore a desired frequency shift of the resonant vibration behavior of the rotating blade was not achieved. Finally it was thought about to change the blade material in view of a possible change of Young's modulus, but this idea was dropped because of low cycle fatigue limitations associated with conventionally cast and directionally solidified materials.
All approaches of a desired influence on the vibration behavior of the rotating blades especially arranged within the first stage of a turbine and the turbine aerodynamic efficiency show the complexity of the problem. Major mass redistribution in designing an enhanced shape of the airfoil of a rotating blade is also considered to be difficult because especially rotating blades of the front stages are actively cooled components which are hollow bodies containing a multitude of cooling channel for cooling purpose. The thin metal walls of the rotating blades have to be cooled intensively to fulfill target life. Also the aspect of increasing the shank length of a rotating blade was considered to influence the vibration behavior of the rotating blade itself but was not deemed to be favorable due to the fact that this approach would result in rotor limits at the fire tree region in which cooling air supply via rotor bores is provided so that the rotor outline would also have to be adjusted.
The document U.S. Pat. No. 5,525,038 discloses a rotor blade for a gas turbine engine which is optimized to reduce tip leakage through a tip clearance. The rotor blade provides a significantly bowed surface formed at the tip region extending from the leading edge to the trailing edge of the suction side of the rotor blade. The profile cross-sections along the span of the airfoil of the rotor blade do not vary significantly, at least the axial chord length of the airfoil along the whole span of the rotor blade remains unchanged.
The axial chord length is defined as the length of the projection of the blade, as set in the turbine, onto a line parallel to the turbine axis. This can be seen in for example David Gordon Wilson's “The Design of High-Efficiency Turbomachinery and Gas Turbines”, pp 487-492, published by the MIT Press, Cambridge, Mass., 1984, 5th printing 1991. Particular reference is made to the second figure on page 487.
It is an object of the invention to provide a gas turbine engine rotor blade comprising an airfoil extending in radial direction from a blade root to a blade tip, defining a span ranging from 0% at the blade root to 100% at the blade tip, and extending in axial direction from a leading edge to a trailing edge, which limit a chord with an axial chord length defined by an axial length of a straight line connecting the leading edge and trailing edge of the airfoil depending on the span which provides an enhanced vibration behavior such that resonance excitation does not occur at the rotating blades of the first and following stages.
The object is achieved by the features in the independent claim 1. The invention can be modified advantageously by the features disclosed in the dependent claims as well in the following description especially referring to preferred embodiments.
It has been recognized according to the invention that by increasing the axial chord length at least in a span region from 80% span to 100% span, a significant influence on the resonant vibration behavior of the rotating blade can be exerted without a deterioration of the aerodynamic properties of the airfoil of the rotating blade. The increase of axial chord length is directly combined with an increase of mass in the region of the airfoil tip which influences the mechanical properties, in particular the Eigenfrequencies of the rotating blade.
In a preferred embodiment of the invention the axial chord length of the airfoil of the gas turbine blade increases continuously at least from 70% span to 100% span. advantageously the increase of the axial chord length with increasing span is more or less symmetrical relative to a so called stacking line which is a line on the surface at the pressure side of the airfoil extending from 0% to 100% span at an axial position of 50%±5% of axial chord length.
The inventive gas turbine blade provides in view of its axial chord length a minimum at least in the range between 50%±10% span and 70%±10% span, i.e. the airfoil of the gas turbine blade between 0% span and 50%±10% span is formed with a conventional shape which provides a decreasing axial chord length from 0% span to 50%±10% span. Towards the tip the chord length is increasing again.
An optimized embodiment of an inventive gas turbine blade provides an axial chord length which increases from 50% span to 100% span and provides a minimal axial chord length at 50% span.
The axial increase of the axial chord length in the range between the tailored mid region of the airfoil to the airfoil tip, i.e. 100% span ranges between 5%±5% and 15%±10% of the axial chord length in the tailored mid region of the airfoil.
As a result of the increase of axial chord length along the radial upper part of the airfoil of the turbine blade influence on the eigenfrequency of the turbine blade can be exerted such that the eigenfrequency can be modified in an amount so that resonant excitation can be minimized or even excluded.
To increase the difference between the eigen frequency of the gas turbine blade to the excitation frequency caused by stationary vanes in the first stage even more it is further proposed to bend the leading and trailing edge in the radial upper region of the airfoil additionally. Preferably the bending of the leading and trailing edge depend on a curvature of a stacking line which was already explained before, which is a line on the surface at the pressure side of the airfoil extending from 0% to 100% span at an axial position of 50%±5% of axial chord length. The stacking line is bended in the span region between 50%±10% span and 100% span such that the stacking line encircles at 100% span an angle α with a virtual plane oriented orthogonal to the radial direction and wherein the angle α lies within a plane defined by the stacking line and the radial direction such that for the angle α applies: 12.5°±2.5°≦α≦25°±5°.
For the sake of completeness it should be mentioned that the stacking line can be kept straight between 5%±5% span and 50%±10% span.
Preferably the stacking line provides a curvature within the span region between 50%±10% span and 100% span which is defined by one single radius.
In a further preferred embodiment the rotating blade provides an aspect ratio concerning span to axial chord length at 5%±5% span ranging from 1.6 to 2.1. In case of blades having different span dimensions along the leading and trailing edge the before aspect ratio concerns the span dimension along the trailing edge.
The invention shall subsequently be explained in more detail based on exemplary embodiments in conjunction with the drawings. In the drawings
a, b perspective view on the inventive turbine blade and a top view of vertical stacked airfoil cross sections.
On the right hand side of
Case a) shows a rotor blade commonly used in gas turbines and represents the state of the art. The common rotor blade provides an airfoil 1 which extends radially from a blade root 2 to the blade tip 3. The blade root 2 comprises a shroud 4 and a fire tree shaped blade foot 5 for fixing purpose inside the rotor arrangement. As can be seen from the upper sketch in case a) the commonly known rotor blade provides an airfoil 1 providing a axial chord length 6 which decreases along the whole span from 0% span to 100% span. The rotor blade illustrated in case a) comprises an eigenfrequency which overlaps with the excitation frequency represented by the dashed line box B in the diagram shown in
In case b) an inventive improved rotor blade is illustrated having an airfoil 1 which provides an axial chord length 6 which increases in a span region s from 50% span to 100% span. As can be seen from the side view in the upper sketch in case b) the airfoil 1 has a minimum axial chord length 6 in the range of 50% span. The increase of the axial chord length 6 can also be derived from the front view sketch in the lower part of case b).
The inventive action contributes that the eigenfrequency of the improved airfoil is dropped in comparison to the commonly known blade of case a). Due to the increase of mass in the tip range of the airfoil in case b) the eigenfrequency drops below which means in case of the situation illustrated in the diagram of
Case c) which is illustrated at the right side of
a, b, c show a three side view presentation of an inventive rotor blade as introduced shortly in case c) of
In
As can be seen from the front view of
a shows a perspective view onto the pressure side 11 of an inventive airfoil 1 providing both, an increase of axial chord length 6 in the span range between 50% and 100% span and bending of the leading edge 9 and rear edge 10 within the span region between 50% and 100% span. The bending of the leading 9 and trailing edge 10 depend on the curvature of the stacking line 12 which can be seen in
b shows a vertical projection of different profile cross-sections through the airfoil 1 at different span regions which are indicated in
| Number | Date | Country | Kind |
|---|---|---|---|
| 14160866.1 | Mar 2014 | EP | regional |
