Gas turbine blade

Abstract

The invention refers to a gas turbine blade including 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. The axial chord length increases at least from 80% span to 100% span.

Claims

1. A gas turbine 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, wherein the axial chord length increases at least from 80% span to 100% span and the axial chord length provides a minimum at least in the range between 50%±10% span and 70%±10% span.

2. The gas turbine blade according to claim 1, wherein the axial chord length increases at least from 70% to 100% span.

3. The gas turbine blade according to claim 1, wherein the axial chord length increases from 50% span to 100% span and provides a minimum at 50% span.

4. The gas turbine blade according to claim 1, wherein the leading edge and the tailing edge separate a suction and a pressure surface of the airfoil, both surfaces extending radially between the blade root and the blade tip and axially between the leading and trailing edge and being mutually opposed surfaces of the airfoil along a circumferential direction which is orthogonal to the axial and radial direction, and that the leading and trailing edge are bent within at least one span region.

5. The gas turbine blade according to claim 4, wherein the leading and the trailing edge are bent in a circumferential direction towards the suction surface side of the airfoil.

6. The gas turbine blade according to claim 4, wherein the at least one span region is between 50%±10% span and 100% span.

7. The gas turbine blade according to claim 4, wherein bending of the leading and trailing edge depend on a curvature of a 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, and said stacking line is bent in the span region between 50%±10% span and 100% span such that the stacking line encircles at 100% span at an angle α with a virtual plane oriented orthogonal to the radial direction, wherein the angle α is in a plane defined by the stacking line and the radial direction, for the angle applies:
(12.5°±2.5°)≦α≦(25°±5°).

8. The gas turbine blade according to claim 7, wherein the stacking line is straight between 0% span and 50%±10% span.

9. The gas turbine blade according to claim 7, wherein stacking line provides a curvature within the span direction, and the curvature is defined by one single radius.

10. The gas turbine blade according to claim 1, wherein the blade is an activity-cooled rotating turbine blade having cooling channels inside the airfoil.

11. A gas turbine blade according to claim 1, wherein the blade provides an aspect ratio span/axial chord length at 5%±5% span ranging from 1.6 to 2.1.

12. The gas turbine blade according to claim 1, wherein the blade is suitable for use as a rotor blade or a guide blade vane of a turbo-machinery.

Description

DESCRIPTION OF THE DRAWINGS

(1) The invention shall subsequently be explained in more detail based on exemplary embodiments in conjunction with the drawings. In the drawings

(2) FIG. 1 shows on the left hand side a diagram which illustrates resonance frequency behavior, e.g. of vanes and blades in the front stage of a gas turbine,

(3) FIG. 2 three side view presentation of an enhanced embodiment of the inventive turbine blade and

(4) FIG. 3a, b perspective view on the inventive turbine blade and a top view of vertical stacked airfoil cross sections.

DETAILED DESCRIPTION

(5) FIG. 1 shows on the left hand side a diagram which illustrates resonance frequency behavior of vanes and blades in the first stage of a gas turbine. Along the abscissa of the diagram values are indicated representing the engine speed. Along the ordinate of the diagram vibrating frequency are indicated. The dashed line box B indicates the source of excitation depending on the engine speed, in which resonance excitation of the blades of the gas turbine can occur.

(6) On the right hand side of FIG. 1 three different embodiments a), b) c) of rotor blades of a gas turbine are illustrates. The upper view in each case shows a side view of a rotor blade and the corresponding lower view shows the blade in a perspective front view.

(7) 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 FIG. 1 left hand side. This leads to a reduced life time due to a high amount of vibrational impact.

(8) 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).

(9) 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 FIG. 1 left hand side there is nearly no overlap between the resonance frequency of the blade of case b) and the excitation frequency range indicated by the dashed line box B. Therefore the improved blade illustrated in case b) provides a significant enhanced vibrational behavior which is clearly robust against vibrational excitation. This leads to an effective enhancement of the aerodynamic behavior and prolongs lifetime of the blade clearly.

(10) Case c) which is illustrated at the right side of FIG. 1 shows a rotor blade which provides an axial chord length increase as explained in case b), which can be derived from the upper view in case c) but additionally provides a bending of the airfoil 1 in circumferential direction towards the suction side 7 of the airfoil 1. Bending of the airfoil 1 is limited in a span region preferably between 50% span and 100% span which can be derived from the lower sketch of case c). The additional bending of the airfoil 1 as described before and as will be discussed in more detail below leads to an enhanced frequency behavior of the rotor blade which is illustrated in the diagram of FIG. 1 left hand side. The eigenfrequency of a rotor blade as disclosed in case c) provides a significant lower eigenfrequency which is clearly below the airfoil illustrated in case b). This leads to a significant frequency separation relative to the excitation frequency characterized by the dashed line box B of FIG. 1.

(11) FIG. 2 a, b, c show a three side view presentation of an inventive rotor blade as introduced shortly in case c) of FIG. 1. The FIG. 2a shows a front view, FIG. 2b shows the side view and FIG. 2c shows the rear view of an inventively formed rotor blade.

(12) In FIG. 2b it is assumed that the flow direction 8 of the gas flow in a turbine is directed from the left hand side to the right hand side, so that the left edge of the illustration represents the leading edge 9 and the right edge represents the rear edge 10 of the airfoil 1. The suction side 7 of the airfoil 1 in FIG. 2b faces towards the observer. The blade has an radially extension which is called span s which extends from 0% span at the blade root (not shown) to 100% span which corresponds to the blade tip 3. The axial chord length 6 varies along the whole span s but increases inventively from a mid range span preferably from 50% span to 100% span. The increase of axial chord length 6 leads automatically to an increase of mass in the blade tip region which influences the resonance frequency of the rotor blade significantly. The amount of increase of the axial chord length 6 from the mid-range span region to 100% span is about 5%±5% to 15%±10% related to the axial chord length 6 of 50% span of the airfoil 1. This increase is illustrated in FIG. 2b by the vertical dashed lines.

(13) As can be seen from the front view of FIG. 2a the leading edge 9 is bended as well the rear edge 10 which cannot be seen on the front view in a span range between 50% span and 100% span. The bending is oriented towards the suction side 7 of the airfoil 1 of the rotor blade. Bending of the leading edge 9 as well of the rear edge 10 is defined by a curvature of a so called stacking line which is a line on the surface at the pressure side 11 of the airfoil 1 extending from 0% to 100% span at an axial position of 50.+−0.5% of axial chord length 6. The curvature of the stacking line within the span region between 50% and 100% span is defined by one single radius preferably which can be seen more clearly in FIG. 3a.

(14) FIG. 3a 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 FIG. 3a which is the line on the surface of the pressure side 11 extending from 0% to 100% span at an axial position of 50%±5% of axial chord length 6. The stacking line 12 is almost straight between 0% span and 50%±10% span and is bended in the span region between 50%±10% span and 100% span such that the stacking line 12 encircles at 100% span an angle α with the virtual plane 13 orientated 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 the angle α is between 12.5°±2.5° and 25°±5°. The curvature of the stacking line within the upper span region is defined by on single radius preferably. In other preferred embodiments the stacking line additionally can provide at least one straight section along the upper span region.

(15) FIG. 3b shows a vertical projection of different profile cross-sections through the airfoil 1 at different span regions which are indicated in FIG. 3a by roman numerals I to VIII. The profile cross section I corresponds to the profile cross-section at 0% span and the profile cross section VIII corresponds to the profile cross-section at 100% span. The vertical projection in radial direction shows a significant geometrical offset of the profile cross section within the span region 50% span to 100% i.e. the profile cross sections V to VIII. The geometrical offset is caused both by an offset in circumferential direction towards the suction side 7 of the airfoil 1 and further by an increase of axial chord length 6 from 50% span to 100% span.