Turbine blade and manufacture thereof

Details Turbine blade and manufacture thereof Turbine blade and manufacture thereof

2000-09-26

2002-03-19

Look, Edward K. (Department: 3745)

Fluid reaction surfaces (i.e., impellers)

Specific blade structure

Laminated, embedded member or encased material

C416S230000, C415S914000

Reexamination Certificate

active

06358013

ABSTRACT:

BACKGROUND OF THE INVENTION
The present invention relates to a gas turbine engine turbine blade having improved gasflow shedding capability.
The present invention also relates to a method of manufacturing said turbine blade.
Present day gas turbine engines operate at extremely high temperatures, eg 1400 C. It follows, that the material from which the turbine blades are manufactured, must be capable of operating in those temperatures for a considerable period of time, in order to ensure commercial viability of the associated engine.
Metals which will perform satisfactorily in such temperatures have been concocted, provided they are of sufficient bulk, as to avoid erosion by the gasflow.
As is well known, the main gasflow surfaces of turbine blades are of aerofoil shape, ie they have a rounded leading edge, suction and pressure surfaces, and terminate in a trailing edge which is thin, relative to the leading portion of the aerofoil. Ideally, the trailing edge should be so thin, that the gasflows from the respective suction and pressure surfaces, on leaving the trailing edge, would flow therefrom in the form of a smooth wake. However, the need to avoid erosion dictates that the trailing edge be rounded, so much so, that the respective gasflows break away from the trailing edge, which reduces the base pressure on the trailing edge extremity, and causes generation of a stream of vortices. This undesirable effect occurs over the full length of the blade trailing edge, and consequently adversely affects the overall operating efficiency of the associated gas turbine engine.
SUMMARY OF THE INVENTION
The present invention seeks to provide an improved gas turbine engine turbine blade.
According to the present invention, a gas turbine engine turbine blade comprises an aerofoil, from the end extremity of the trailing edge of which there projects a plurality of elongate ceramic fibres, in a direction parallel with the mean direction of gasflows which leave said trailing edge during operation of said turbine blade in an associated gas turbine engine, said fibres being arranged in side by side relationship along at least a substantial portion of said trailing edge extremity.
The present invention further provides a method of fixing a plurality of ceramic fibres into the trailing edge portion of a turbine blade so as to protrude therefrom in a direction parallel with the mean direction of gasflows which leave said trailing edge of said turbine blade during operation in a gas turbine engine, comprising the steps of forming a slot in the blade trailing edge extremity, along at least a major portion of the trailing edge length, arranging a plurality of ceramic fibres in side by side relationship, directly or indirectly in said slot, and then squeezing the sides of said slot towards each other, so as to, directly or indirectly, trap and retain said ceramic fibres in the trailing edge portion of said turbine blade.


REFERENCES:
patent: 3779338 (1973-12-01), Hayden et al.
patent: 4789304 (1988-12-01), Gustafson et al.
patent: 4806077 (1989-02-01), Bost
patent: 5401138 (1995-03-01), Mosiewicz
patent: 6139268 (2000-10-01), Murawski et al.
patent: 789883 (1958-06-01), None
patent: 1436724 (1976-05-01), None

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