GB2399602A

GB2399602A - Gas turbine rotor blade

Info

Publication number: GB2399602A
Application number: GB0305956A
Authority: GB
Inventors: Robert Hirst
Original assignee: Alstom SA
Current assignee: Alstom SA
Priority date: 2003-03-15
Filing date: 2003-03-15
Publication date: 2004-09-22
Also published as: GB0305956D0

Abstract

A gas turbine rotor blade comprises a root end to attach to a rotor, and a tip end remote therefrom, the tip end carrying a shroud segment 2 which surrounds the rotor. At least one fin or rib 5 extends radially outwardly from the shroud, the fin having a greater thickness at its trailing edge 6 than its leading edge, relative to the direction of rotation of the rotor. The increased thickness may be achieved by adding material to the upstream end of the fin, and may be over a small part of the fin length. Additionally, the thickness increase may occur over the entire radial height of the fin.

Description

GAS TURBINE ROTOR BLADE

Field of the Invention

This invention relates to a gas turbine rotor blade, to a rotor for a gas turbine, and to a gas turbine engine including such a rotor.

Background to the Invention

In order to reduce the leakage of gases around the blade tips, it is common prac- tice to provide the blade tip with a tip shroud which is a segment of a circumferential ring around the outer ends of the blades. An example of a shrouded rotor is shown in US-A- 5211540.

While this arrangement reduces aerodynamic losses significantly, it is still possible for some leakage to occur around the outside of this shroud. To reduce this leakage fur ther, fins have been provided extending radially outwards of the shroud towards the tur bine casing. US- A-6068443 illustrates such an arrangement. To allow for expansion of the blades from start-up of the engine, a gap may have to be left between the blades and the casing when cold, or a layer of soft metal provided around the inner surface of the casing surrounding the blade tips, so that, on expansion, the fins actually cut into the soft metal to form a sealing groove therein.

Because the tip shrouds need to be formed as separate segments, rather than a continuous ring, to permit assembly and allow for thermal expansion of the blades, there is a possibility of a step being created between one fin and the next as a result of cumula- tive tolerances or manufacturing deviations during the blade casting process. Steps can also result from wear in the interlock area between one shroud and the next, and through "soldiering", i.e. one blade moving against another, for example as a result of thermal ratcheting. Misalignment of the fins can lead to turbulence at the trailing edges, with the possibility of forced induction of hot gases into a cooled shroud cavity.

Summary of the Invention

According to the invention, there is provided a gas turbine rotor blade having a root end for attachment to a rotor and a tip end remote therefore, the tip end carrying a segment of a circumferential shroud which, in use, surrounds the rotor, at least one fin extending radially outwardly of the outer surface of the shroud along the circumferential Specficabon ALSTINTE-P1106 GBA2003-03-13 doc - 2 length of the shroud, the fin having a greater thickness at the trailing edge thereof relative to the direction of rotation of the rotor, in use, then at the leading edge.

Where more than one fin is provided on the shroud, at least the fin furthest up- stream in the direction of gas flow is made thicker at its trailing edge, and preferably each fin has this configuration.

Preferably, the or each fin is made thicker by adding material to the fin on the pressure or upstream face thereof, suitably over a minor part of the circumferential length of the fin.

It has been found that blades in accordance with the invention confer greater effi ciency on the engine and, by reducing potential damage to the shroud, offer longer blade life.

Where the fin tapers radially outwardly from the shroud, the amount of thicken- ing may be uniform over the height of the fin or it too may be tapered so that the greatest thickening occurs at the part of the fin nearest to the shroud.

The shroud segments are preferably designed so as to interlock one with another to provide a substantially continuous circumferential shroud surface under conditions of thermal and centrifugal expansion. Such interlocking designs are well-known in the art, for example from US-A5211540.

The invention also provides a gas turbine rotor comprising a plurality of blades in accordance with the invention, and a gas turbine engine comprising at least one such ro- tor.

Brief Description of the Drawings

In the drawings, which illustrate an exemplarily embodiment of the invention: Figure 1 is a side elevation or view, partly in cross-section of the end portion of a turbine blade and a portion of the turbine casing; and Figure 2 is an end view of the blade shown in Figure 1, with part of an adjacent blade.

Detailed Description of the Illustrated Embodiment

Figures 1 and 2 illustrate one particular form of inter-engaging shroud segment, but it will be appreciated that the invention is not limited to any particular form of shroud

Specification ALSTINTEP1106 GBA-2003-03-13 doc - 3

segment on the blade. The blade 1 rotates within the turbine casing 2, and has a shroud segment 3 provided on its tip. The shroud segment 3 cooperates with the segments on adjacent blades, as shown in Figure 2, to provide a substantially continuous circumferential surface, even under conditions of slight misalignment of the shrouds. Figure 2 shows an extreme example of misalignment to illustrate the effects of the invention.

Each segment 3 is provided with a pair of fins 4 and 5 extending radially outwardly from the segment and along the circumferential length thereof. It will be seen from Figure 2 that, in the event of axial misalignment of the shroud segments 3, the fins 4 and 5 do not align precisely; instead, there is a small step between the adjacent fins, and this can give rise to turbulence at the trailing edge of the fin and to scooping of gases on the pres- sure side, leading in turn to greater leakage of the gases past the fins. This reduces engine efficiency, and can increase wear of the shroud, reducing blade life.

The fin 5 on the pressure side of the rotor is therefore provided with a thickened portion 6 at the trailing edge thereof relative to the direction of rotation. The additional material in the fin 5 is added to the pressure side of the fin so as to deflect the gas flow away from the leading edge of the fin on the adjacent blade, reducing or eliminating the turbulence and scooping effects. The additional thickness of the fin 5 at its trailing edge is preferably equal to the maximum degree of misalignment experienced between two adja- cent shroud segments, in use, while the circumferential length of the fin over which the thickening occurs is such as to ensure smooth deflection of the gases to avoid the genera- tion of turbulence.

While the embodiment shown in the drawings provides thickening only on the pressure-side fin, it will be appreciated that both fins may be thickened in the same man- ner.

Specticaton ALSTINTEP1106 GBA2003-03-13 doc

Claims

Claims 1. A gas turbine rotor blade having a root end for attachment to a

rotor and a tip end remote there from, the tip end carrying a segment of a circumferential shroud which, in use, surrounds the rotor, at least one fin extending radially outwardly of the outer surface of the shroud along the circumferential length of the shroud, the fin having a greater thickness at the trailing edge thereof relative to the direction of rotation of the rotor, in use, than at the leading edge.
2. A rotor blade according to Claim 1 wherein the shroud has more than one fin thereon, the fins being spaced apart in the axial direction of the rotor, and wherein at least the fin furthest upstream in the direction of gas flow has an increased thickness at the trailing edge thereof.
3. A rotor blade according to Claim 1 or 2, wherein the or each fin is made thicker by adding material to the fin on the upstream or pressure face thereof.
4. A rotor blade according to Claim 1, 2 or 3, wherein the increase in thick ness is over a minor part of the circumferential length of the fin.
5. A rotor blade according to any preceding claim, wherein the increase in thickness is uniform over the radial height of the fin.
6. A gas turbine rotor blade substantially as described with reference to, or as shown in, the drawings.
7. A gas turbine rotor comprising a plurality of blades as defined in any pre ceding claim.
8. A gas turbine engine comprising at least one rotor as defined in Claim 7.

Specficabon ALSTINTE-P1106 GBA-2003-03-13 doc