Rotary kinetic fluid motors or pumps – With passage in blade – vane – shaft or rotary distributor...
Reexamination Certificate
1999-06-01
2001-06-05
Look, Edward K. (Department: 3745)
Rotary kinetic fluid motors or pumps
With passage in blade, vane, shaft or rotary distributor...
C415S191000, C416S09700R
Reexamination Certificate
active
06241466
ABSTRACT:
BACKGROUND OF THE INVENTION
The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzle performance and cooling.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases which flow through turbine stages for extracting energy therefrom. In a turbofan engine, a high pressure turbine powers the compressor, and a low pressure turbine powers a fan disposed upstream from the compressor. Each turbine includes a stationary turbine nozzle having vanes mounted between inner and outer bands, followed in turn by a row of rotor blades extending outwardly from a rotor disk.
The high pressure turbine nozzle is disposed at the outlet of the combustor and receives therefrom combustion gases at the hottest temperature, with the temperature decreasing as energy is extracted from the gases in the downstream turbine stages. Both the nozzle vanes and rotor blades have hollow airfoils through which a portion of air bled from the compressor is used for providing cooling thereof. Bleeding cooling air from the compressor necessarily decreases the overall efficiency of the engine, and it is therefore desired to use as little cooling air as possible while adequately cooling the vanes and blades.
The profile or contour of the turbine airfoils is controlled by the specific thermodynamic operating cycle of the engine, and cooperating aerodynamic performance. Each airfoil has a generally concave, pressure side and a generally convex, suction side extending axially between leading and trailing edges and radially between a root and tip. The airfoil increases in thickness just aft of the leading edge and then tapers with a reduced thickness to a thin trailing edge.
Since the trailing edge is thin, it is difficult to cool during operation and typically operates relatively hot which affects the useful life of the airfoil. Trailing edge cooling of the first stage nozzle of the high pressure turbine is particularly critical in view of the hot combustion gases directly received from the combustor.
Trailing edge cooling may be provided in various conventional manners in which the cooling air is channeled inside the airfoil directly behind the trailing edge and is discharged through a row of trailing edge cooling holes thereat. In one design, the trailing edge holes have outlets along the airfoil pressure side which begin at a breakout lip forward of the trailing edge and terminate directly at the trailing edge. Since the thickness of the breakout lip has a practical minimum value to prevent deterioration and oxidation during operation, the breakout distance from the trailing edge to the lip is relatively large.
Accordingly, as the cooling air is discharged through the trailing edge holes, it is heated by the combustion gases which decreases its ability to cool the trailing edge. Furthermore, fluid flow behind the trailing edge locally stagnates in the wake thereof further increasing the difficulty of cooling the trailing edge itself.
Additional considerations in airfoil cooling include the conventionally known backflow margin and blowoff margin. The cooling air is bled from the compressor at a corresponding pressure to ensure a suitable differential pressure between the cooling air inside the airfoil and the pressure of the combustion gases outside the airfoil for driving the cooling air through the airfoil. A suitable backflow margin prevents the reverse flow of combustion gases into cooling air holes in the airfoil. And, a suitable blowoff margin prevents excessive discharge velocities of the cooling air as it exits the cooling holes.
However, in a conventional turbine nozzle design, for example, the cooling air discharged from the trailing edge, pressure side holes has a greater velocity than that of the combustion gases which flow therealong. Accordingly, as the high speed cooling air meets the low speed combustion gases at the hole outlets, mixing losses are created which affects both the overall performance of the engine and affects the ability to cool the airfoil trailing edge.
Accordingly, it is desired to provide an improved turbine airfoil trailing edge configuration for reducing mixing losses and improving trailing edge cooling.
BRIEF SUMMARY OF THE INVENTION
A turbine airfoil includes pressure and suction sides extending between leading and trailing edges and defining an internal cooling air passage. A row of trailing edge holes is disposed in flow communication with the air passage behind the trailing edge. The airfoil is sized in conjunction with an adjacent airfoil for accelerating combustion gases along the pressure side at the trailing edge holes to a velocity at least as high as the velocity of the cooling air discharged from the holes.
REFERENCES:
patent: 4303374 (1981-12-01), Braddy
patent: 4314442 (1982-02-01), Rice
patent: 4353679 (1982-10-01), Hauser
patent: 4545197 (1985-10-01), Rice
patent: 4601638 (1986-07-01), Hill et al.
patent: 5368441 (1994-11-01), Sylvestro et al.
Tung Stephen K.
Weldon Donald C.
General Electric Company
Hess Andrew C.
Look Edward K.
Nguyen Ninh
Young Rodney M.
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