Fluid reaction surfaces (i.e. – impellers) – With heating – cooling or thermal insulation means – Changing state mass within or fluid flow through working...
Reexamination Certificate
2000-01-11
2001-07-24
Verdier, Christopher (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
With heating, cooling or thermal insulation means
Changing state mass within or fluid flow through working...
C416S189000, C416S192000, C415S115000
Reexamination Certificate
active
06264428
ABSTRACT:
FIELD OF THE INVENTION
This invention relates to a cooled aerofoil for a gas turbine engine.
BACKGROUND OF THE INVENTION
Gas turbine engines of the axial flow type conventionally include turbines that are made up of axially alternate annular arrays of radially extending stator aerofoil vanes and rotary aerofoil blades. The demands of modern gas turbine engines dictate that the gases that flow through, and thereby drive, the turbine are at extremely high temperature. As the gases flow through the turbine, their temperature progressively falls as they drive the turbine. However, notwithstanding this, the gas temperatures in the higher pressure regions of the turbine are so high that some form of aerofoil cooling is required.
Conventionally, turbine aerofoils, both blades and vanes, are cooled internally with air that has been tapped from the gas turbine engine's compressor. Using engine compressor air in this manner does, however, carry a penalty in terms of the overall operating efficiency of the engine. Thus generally speaking, the larger the percentage of air taken from the compressor, the greater the adverse effect there is upon overall engine operating efficiency.
Many efforts have been made in the past to make efficient use of compressor-derived air in the cooling of aerofoils. These efforts have centered mainly around the design of aerofoils provided with internal passages for the flow of cooling air. Typically such passages are in a serpentine configuration to provide convection cooling and some of the air from the passages is exhausted through small holes that provide communication between the passages and the aerofoil external surface. As the air is exhausted from the holes, it forms a film that provides additional aerofoil cooling.
Aerofoils cooled in this manner are often complex internally and hence difficult and expensive to manufacture. Moreover, they may not be as effective as is desirable in providing overall aerofoil cooling in view of the air pressure losses that are associated with flowing the cooling air through the many turns in the small diameter passages within the aerofoil.
It has been suggested in FR2,569,225 to provide a hollow aerofoil in which the walls of the aerofoil are provided with radially extending passages. The passages are in communication with the hollow aerofoil interior and also with the external surface of the aerofoil. Cooling air is supplied to the central aerofoil chamber from where it flows into the radially extending passages. From the radially extending passages, it flows on to the aerofoil exterior surface to provide film cooling thereof.
Although such aerofoils are cooled effectively, the pursuit of greater engine efficiency makes yet more effective cooling a highly desirable objective. It is an object of the present invention to provide such an aerofoil.
SUMMARY OF THE INVENTION
According to the present invention a cooled aerofoil for a gas turbine engine is hollow having a wall defining a central, lengthways extending plenum, said wall being configured to define a leading edge, trailing edge, suction flank and pressure flank of said aerofoil, said suction and pressure flanks interconnecting said leading and trailing edges and having a plurality of passages extending lengthways therethrough, means being provided to supply cooling air to each of said suction flank passages at one extent of said aerofoil, said suction flank passages being configured to subsequently direct that cooling air into said central plenum at the opposite extent of said aerofoil to thereby provide convection cooling of said suction flank, said pressure flank being provided with a first array of apertures interconnecting said pressure flank passages with said central plenum to facilitate the flow of at least some of said cooling air from said central plenum into said pressure flank passages and a second array of apertures interconnecting said pressure flank passages with the outer surface of said pressure flank to facilitate the exhaustion of said cooling air from said pressure flank passages to provide film cooling of said pressure flank outer surface.
Means may be provided to direct some of the cooling air from said central plenum to the region of said aerofoil trailing edge to provide cooling thereof.
Said means to direct some of said cooling air from said central plenum to the region of said aerofoil trailing edge may comprise a secondary plenum positioned alongside said central plenum, said secondary plenum being interconnected with said central plenum to provide flow communication therebetween, and is also in communication with the external surface of said aerofoil adjacent said trailing edge to facilitate the exhaustion of said cooling air from said secondary plenum.
Said pressure flank passages are preferably configured to define internal walls upon which said cooling air from said central plenum impinges having flowed through said first array of apertures.
Said pressure flank passages may be of elongate cross-sectional shape.
Said aerofoil may be in the form of a rotor blade, having a root portion at one extent thereof for attachment to a rotor disc of a gas turbine engine and a tip at the other end thereof.
Preferably said suction flank passages each extend from the region of said root portion to the region of said tip portion, said means to supply cooling air to said suction flank passages being located in said root portion region so that said cooling air is exhausted from said suction flank passages into said central plenum in the region of said tip portion.
Preferably the portion of said suction flank between said suction flank passages and said central plenum is so configured and arranged as to carry the majority of centrifugal loads operationally imposed upon said aerofoil.
Said aerofoil is preferably provided with at least one lengthways extending passage adjacent said leading edge thereof that is in communication with said cooling air supply means in said root portion so as to be provided with a supply of cooling air independent of that supplied to said central plenum.
Cooling passages may be provided in said leading edge region to exhaust cooling air from said leading edge passage on to said aerofoil external surface to provide film cooling thereof.
Said aerofoil may be provided with a shroud at said tip.
Said suction flank passages may each be of generally serpentine configuration.
REFERENCES:
patent: 3427001 (1969-02-01), Malley et al.
patent: 3533711 (1970-10-01), Kercher
patent: 3533712 (1970-10-01), Kercher
patent: 3628885 (1971-12-01), Sidenstick et al.
patent: 4474532 (1984-10-01), Pazder
patent: 4992206 (1991-02-01), Ohtomo et al.
patent: 5215431 (1993-06-01), Derrien
patent: 5342172 (1994-08-01), Coudray et al.
patent: 5387086 (1995-02-01), Frey et al.
patent: 5562409 (1996-10-01), Livsey et al.
patent: 5700131 (1997-12-01), Hall et al.
patent: 5702232 (1997-12-01), Moore
patent: 5813836 (1998-09-01), Starkweather
patent: 0562944 A (1993-09-01), None
patent: 1467197 SP (1977-03-01), None
patent: 2319567 A (1998-05-01), None
patent: 60-135606 (1985-07-01), None
patent: 60-198305 (1985-10-01), None
Dailey Geoffrey M
Evans Peter A
McCall Richard A B
Manelli Denison & Selter PLLC
Rolls-Royce plc
Taltavull W. Warren
Verdier Christopher
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