Gas turbine blade and method for producing such blade

Abstract

A gas turbine blade having a casted metal airfoil, the airfoil has a main wall defining at least one interior cavity, having a first side wall and a second side wall, which are coupled to each other at a leading edge and a trailing edge, extending in a radial direction from a blade root to a blade tip and defining a radial span from 0% at the blade root to 100% at the blade tip. The main airfoil has a radial span dependent chord length defined by a straight line connecting the leading edge and the trailing edge as well as a radial span dependent solidity ratio of metal area to total cross-sectional area. Solidity ratios in a machined zone of the airfoil from 80% to 85% of span are below 35%, in particular all solidity ratios in the zone.

Claims

1. A gas turbine blade, comprising: a casted metal airfoil, said airfoil comprising a main wall defining at least one interior cavity and having a first side wall and a second side wall, which are coupled to each other at a leading edge and a trailing edge, wherein the first and second side walls extend in a radial direction from a blade root to a blade tip and define a radial span from 0% at the blade root to 100% at the blade tip, wherein said airfoil has a radial span dependent chord length defined by a straight line connecting the leading edge and the trailing edge as well as a radial span dependent solidity ratio of metal area to total cross-sectional area, wherein the solidity ratio in a machined zone of the airfoil from 75% to 90% of span is below 35%.

2. The gas turbine blade according to claim 1, wherein the solidity ratio at 80% to 85% of span is below 35%.

3. The gas turbine blade according to claim 2, wherein the airfoil comprises a plurality of solidity ratios in the machined zone, and wherein the plurality of solidity ratios in the machined zone are below 35%.

4. The gas turbine blade according to claim 1, wherein a wall thickness of the main wall extending from an external surface of the main wall to the at least one interior cavity is constant in a zone from 85% to 100% of span.

5. The gas turbine blade according to claim 1, wherein a wall thickness of the main wall extending from an external surface of the main wall to the at least one interior cavity increases by a rate of 1% or greater from 60% to 0% of span.

6. The gas turbine blade according to claim 1, wherein a wall thickness of the main wall at the blade tip extending from an external surface of the main wall to the at least one interior cavity is within a range from 1 to 2 mm.

7. The gas turbine blade according to claim 1, wherein the chord length in a zone from 50% to 70% of span is shorter than the chord length at 100% of span.

8. The gas turbine blade according to claim 7, wherein the airfoil comprises a plurality of chord lengths in the zone, and wherein the plurality of chord lengths in the zone are shorter than the chord length at 100% of span.

9. The gas turbine blade according to claim 1, wherein a trailing edge thickness is thinnest in a zone from 60% to 80% of span.

10. The gas turbine blade according to claim 1, wherein a trailing edge thickness at 100% of span is within a range from 2.5 to 4.0 mm.

11. The gas turbine blade according to claim 1, wherein the machined zone extends along an entire circumference of the airfoil at a given radial height.

12. The gas turbine blade according to claim 1, wherein an external surface of the airfoil is an as-cast region over a partial span starting from the blade root.

13. The gas turbine blade according to claim 12, wherein the partial span is at least in a region from 0% to 5% of span.

14. A method for producing the gas turbine blade according to claim 1, comprising: obtaining the casted airfoil by casting and machining an external surface of said casted airfoil exclusively within a zone from 16% to 100% of span in order to reduce a wall thickness of the main wall and/or a trailing edge thickness in said zone.

15. The method according to claim 14, wherein the machining is done by milling, grinding, EDM or ECM.

16. A gas turbine, comprising: a last turbine stage comprising the gas turbine blade of claim 1.

17. The gas turbine blade according to claim 1, wherein the airfoil comprises a plurality of solidity ratios in the machined zone, and wherein the plurality of solidity ratios in the machined zone are below 35%.

18. The gas turbine blade according to claim 1, wherein the chord length in a zone from 50% to 90% of span is shorter than the chord length at 100% of span.

19. The gas turbine blade according to claim 18, wherein the airfoil comprises a plurality of chord lengths in the zone, and wherein the plurality of chord lengths in the zone are shorter than the chord length at 100% of span.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) FIG. 1 is a perspective view of a gas turbine blade according to an embodiment of the present invention;

(2) FIG. 2 is a front view of the blade;

(3) FIG. 3 is a front view of the blade as FIG. 2 showing machined and as-cast regions

(4) FIG. 4 is a sectional view of the blade along lines IV-IV in FIGS. 1 and 2;

(5) FIG. 5 is a sectional view of the blade along lines V-V in FIGS. 1 and 2;

(6) FIG. 6 is a graph showing the solidity ratio relative to radial span for the blade shown in FIGS. 1 to 4 and for a prior art blade having an as-cast design;

(7) FIG. 7 is a graph showing the ratio of wall thickness/tip wall thickness relative to radial span for the blade shown in FIGS. 1 to 4 and for said prior art blade having the as-cast design;

(8) FIG. 8 is a graph showing the radial span relative to the ratio of the chord length/tip chord length for the blade shown in FIGS. 1 to 4, for a prior art freestanding blade, which is not cored, and for a prior art shrouded blade; and

(9) FIG. 9 is a graph showing the radial span relate to the ratio of tip trailing edge width/trailing edge width for the blade shown in FIGS. 1 to 4 and for said prior art blade having the as-cast design.

DETAILED DESCRIPTION OF INVENTION

(10) FIGS. 1 and 2 show different views of a gas turbine blade 1 according to an embodiment of the present invention. The gas turbine blade 1 comprises a metal airfoil 2 with a main wall having a first side wall 3 and a second side wall 4, which are coupled to each other at a leading edge 5 and a trailing edge 6. The airfoil 2 extends in a radial direction from a blade root 7 to a blade tip 8, defines a radial span s from 0% at the blade root 7 to 100% at the blade tip 8, has a radial span dependent chord length c defined by a straight line connecting the leading edge 5 and the trailing edge 6, and has a radial span dependent solidity ratio r.sub.s of metal area to total cross-sectional area. Moreover, the main wall defines three interior cavities 9, which are separated from each other by partition walls 10 each extending between the first side wall 3 and the second side wall 4.

(11) The gas turbine blade 1 is a casted product, whereas the external surface of the main wall of the casted airfoil 2 is exclusively machined within a zone from 16% to 100% of span s as shown in FIG. 3, advantageously by milling. Thus, the airfoil 2 can be subdivided into an as-cast region 11 extending radially outwards from the blade root 7, a subsequent transition region 12, which may or may not be machined, and a subsequent machined region 13. The machining is done in order to reduce the wall thickness of the main wall as well as the trailing edge thickness in the machined zones or rather in order to achieve the results shown in FIGS. 4 to 9.

(12) FIGS. 4 and 5 show cross sectional views of the airfoil 2 at about 58% of span (FIG. 4) and at 100% of span (FIG. 5). It can be seen by comparison that the wall thickness t at 58% of span is much thicker than at 100% of span. In the present case, the wall thickness at 58% of span is about 4 mm, whereas the wall thickness at 100% of span is about 1 mm.

(13) FIG. 6 shows the solidity ratio r.sub.s relative to radial span s for the blade 1 and for a prior art blade having an as-cast design designated by reference numeral 14. The solidity ratios r.sub.s of the blade 1 are below 35% from 90% to 75% of span s in order to reduce the pull load upon the lower sections, and then revert to conventional levels of 50% to 75% in the lower half of the airfoil 2 that needs thicker walls to bear the pull load exerted by the upper airfoil sections.

(14) FIG. 7 shows the ratio of wall thickness/tip wall thickness relative to radial span s for the blade 1 and for said prior art blade 14 having the as-cast design. The blade 1 has no taper in wall thickness t from 100% to 85% of span, and then tapers greater than 1% in the lower 60% of span. This results in an airfoil that has thin walls at the blade tip 8 and then a higher increase in relative thickness than would be practical with conventional casting processes. It should be noted that both blades 1 and 14 have similar absolute wall thicknesses at 0% of span due to packaging and aerodynamic constraints, but the relative increase in wall thickness is what is critical for mechanical and casting criteria. The wall thickness ratios of blade 1 according to the present invention are generally not possible with conventional casting and are achieved by using adaptive airfoil machining in the upper span regions, i.e. by removing an amount of material in terms of wall thickness reduction that is variable relative to radial span s.

(15) By minimizing pull load in the upper spans thanks to thin walls and low solidity ratios the airfoil 2 can also have reduced chord lengths that are actually lower than the tip chord length until 50% of span. FIG. 8 shows in this context the radial span relative to the ratio of the chord length/tip chord length for the blade 1, for a prior art freestanding blade 14 and for a prior art shrouded blade 16. Since a constant pitch-to-chord ratio of 1:1 is ideal aerodynamically, the ideal chord length should decrease while moving down the airfoil 2. However, this is generally not possible because of the additional metal needed to meet casting requirements and support the pull load of the upper sections of the airfoil 2. The very low solidity ratio r.sub.s of the airfoil 2 from 70% to 100% of span enables shorter chord lengths c from 70% to 50% of span. The prior art free standing blade 15 can achieve lower tip chord multiples in the lower 40% of span only because the airfoil is not cored in this region.

(16) FIG. 9 shows the radial span relative to the ratio of tip trailing edge width/trailing edge width for the blade 1 and for said prior art blade 14 having the as-cast design. The prior art blade 14 having the as-cast design has a continuous increase in trailing edge thickness in accordance with typical taper requirements. The blade 1 has a trailing edge thickness d that is thinnest at about 70% of span as a result of the machining process. This provides further aerodynamic advantage by reducing trailing edge losses. The absolute trailing edge thickness at the blade tip 8 is between 2.5 mm and 3.5 mm.

(17) All of these features combine in an airfoil 2 with an AN.sup.2 greater 7.0 e.sup.7 m.sup.2/min.sup.2.

(18) It should be noted that the described embodiment of a gas turbine blade according to the present invention is not limiting for the invention. Rather, modifications are possible without departing from the scope of protection defined by the accompanying claims.