Patent Application
20070243061
The present invention will be more fully described by way of example with reference to the accompanying drawings in which:
A turbofan gas turbine engine 10, as shown in
The high-pressure turbine 24 of the gas turbine engine 10 is shown more clearly in
The platforms 42 of the turbine rotor blades 32 and the platforms 48 of the turbine stator vanes 36 have upstream and downstream portions 42A, 42B, 48A and 48B, which extend axially towards each other. Thus the turbine rotor blades 32 in a stage of turbine rotor blades 32 have upstream portions 42A of the platforms 42, which extend in an upstream direction towards the downstream portions 48B of the platforms 48 of the stage of turbine stator vanes 36 immediately upstream of the stage of turbine rotor blades 32. The stage of turbine stator vanes 36 immediately upstream of the stage of turbine rotor blades 32 has downstream portions 48B of the platforms 48, which extends in a downstream direction towards the upstream portions 42A of the platforms 42 of the turbine rotor blades 32 and the downstream portions 48B of the platforms 48 of the turbine stator vanes 36 are arranged radially around the upstream portions 42A of the platforms 42 of the turbine rotor blades 32.
Similarly the turbine rotor blades 32 in the stage of turbine rotor blades 32 have downstream portions 42B of the platforms 42, which extend in a downstream direction towards upstream portions 48A of the platforms 48 of the stage of turbine stator vanes 36 immediately downstream of a stage of turbine rotor blades 32. The stage of turbine stator vanes 36 immediately downstream of the stage of turbine rotor blades 32 has upstream portions 48A of the platforms 48, which extend in an upstream direction towards the downstream portions 42B of the platforms 42 of the turbine rotor blades 32 and the downstream portions 42B of the platforms 42 of the turbine rotor blades 32 are arranged radially around the upstream portions 48A of the platforms 48 of the turbine stator vanes 36.
A clearance, or seal, 43 is formed between the upstream portions 42A of the platforms 42 of the stage of turbine rotor blades 32 and the downstream portions 48B of the platforms 48 of the upstream stage of turbine stator vanes 36 and a clearance, or seal, 45 is formed between the downstream portions 42B of the platforms 42 of the stage of turbine rotor blades 32 and the upstream portions 48A of the platforms 48 of the downstream stage of turbine stator vanes 36.
These clearances, or seals, 43 and 45 control the amount of cooling air A flowing from within the interior of the high-pressure turbine 24 into the flow path B through the turbine 24 and control the flow of hot gases C from the turbine flow path into the interior of the high-pressure turbine 24. The platforms 42, 48 of the turbine rotor blades 32 and turbine stator vanes 36 overlap to provide a smooth flow line for the inner boundary of the flow path through the high-pressure turbine 24.
The downstream portions 48B of the platforms 48 of the turbine stator vanes 36 of the upstream stage of turbine stator vanes 36 comprises a shape memory alloy member or a bimetallic member. The downstream portions 42B of the platforms 42 of the turbine rotor blades 32 comprises a shape memory alloy member or a bimetallic member.
The shape memory alloy members, or the bimetallic members, of the downstream portions 42B and 48B of the platforms 42 and 48 of the turbine rotor blades 32 and turbine stator vanes 36 respectively are arranged such that above a predetermined temperature, for example in the range 800° C. to 1000° C. the shape memory alloy members or bimetallic members change shape, bend radially outwardly, to increase the clearances 45 and 43 respectively in order to allow a greater flow of cooling air A through the clearances 45 and 43 into the flow path B through the high-pressure turbine 24, as shown by the dashed lines in
The shape memory alloy members, or the bimetallic members, of the downstream portions 42B and 48B of the platforms 42 and 48 of the turbine rotor blades 32 and turbine stator vanes 36 respectively are arranged such that below the predetermined temperature, for example in the range 800° C. to 1000° C. the shape memory alloy members or bimetallic members change shape, bend radially inwardly, back to their original positions to decrease the clearances 45 and 43 respectively in order to allow a lesser flow of cooling air A through the clearances 45 and 43 into the flow path B through the high-pressure turbine 24 and to prevent the flow C of hot gases from the flow path B to the interior of the high-pressure turbine 24 as shown by the full lines in
There are many metals and/or alloys, which have non-linear thermal coefficients of expansion in this temperature region. A bimetallic member would use metals and/or alloys chosen to give a large mismatch in thermal coefficients of expansion in this temperature range to give maximum movement of the bimetallic member, but a small mismatch in thermal coefficients of expansion at temperatures lower than this temperature range to minimise movements and reduce the possibility of contact between the radially adjacent portions of the platforms. Thus the bimetallic member 60 comprises two metals/alloy members 62, 64 with different thermal coefficients of expansion. The metal member 62 with the lower thermal coefficient of expansion is arranged radially further from the axis of the high-pressure turbine 24, radially nearer the flow path B through the high-pressure turbine 24, than the metal member 64 with the higher thermal coefficient of expansion as shown in
The shape memory alloy member for example may comprise a nickel-titanium-palladium shape memory alloy, an iron-nickel-cobalt-titanium shape memory alloy, an iron-manganese-silicon shape memory alloy or an iron-manganese-carbon shape memory alloy.
The shape memory alloy member may be pre-stressed. The shape memory alloy members or bimetallic members of the portions 42B and 48B of the platforms 42 and 48 of the turbine rotor blades 32 and turbine stator vanes 36 may be continuous annular members or part annular members.
In some instances the turbine rotor blades 32 may have shrouds at their radially outer ends to define a portion of the outer boundary of the flow path through the high turbine.
Although the present invention has been described with reference to a high-pressure turbine it may also be used in an intermediate pressure turbine or a low-pressure turbine. Although the present invention has been described with reference to turbine blades and turbine vanes, it may be applicable to compressor blades and compressor vanes.
| Number | Date | Country | Kind |
|---|---|---|---|
| 0607560.0 | Apr 2006 | GB | national |
