Fluid reaction surfaces (i.e. – impellers) – With heating – cooling or thermal insulation means – Changing state mass within or fluid flow through working...
Reexamination Certificate
2001-12-12
2003-09-02
Look, Edward K. (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
With heating, cooling or thermal insulation means
Changing state mass within or fluid flow through working...
C164S369000, C029S889721
Reexamination Certificate
active
06612811
ABSTRACT:
BACKGROUND OF THE INVENTION
The present invention relates generally to a turbine nozzle for a gas turbine engine and, in particular, to an airfoil utilized therein having at least one of an inner cooling slot and an outer cooling slot at the trailing edge thereof configured to have a variable fillet between a recessed wall and a side wall so as to reduce stress on the airfoil.
It will be appreciated that a nozzle segment for the high pressure turbine of a gas turbine engine typically includes a pair of hollow airfoils with integral inner and outer flowpath bands. These pieces are cast separately, partially machined, brazed together, and subsequently finish machined to form the nozzle segment. The hollow airfoil is fed internally with cooling air which then flows through trailing edge slots that exit the aft cavity of the airfoil and discharges through openings in the trailing edge of the airfoil. This cooling air then performs convection cooling as it passes along the trailing edge slot within the airfoil. When such air discharges to the flowpath through the openings in the airfoil trailing edge, it provides film cooling for the airfoil trailing edge.
Turbine airfoils with trailing edge cooling slots inherently have a step between the slot and the rib between the slots. It has been found that the step in the cooling slot closest to the nozzle bands at the inner and outer airfoil/flowpath intersection causes a large stress concentration with high thermal stresses present, which can then result in trailing edge axial cracks. The cracks ultimately propagate through the airfoil section and lead to premature failure of the turbine nozzles. The cooling slot itself cannot be removed since overheating of the trailing edge of the airfoil would result.
Moreover, the step is difficult to grind smooth because of its proximity to the airfoil/band junction.
It will be understood that the hollow airfoil cavities and trailing edge cooling slots are formed during a casting process by ceramic core which is produced separately and combined with a wax pattern prior to casting. On previous designs, corner fillets for the trailing edge slot are created by the ceramic core and minimized in order to reduce slot blockage and maintain cooling flow area. During manufacturing, however, the ceramic core is subjected to auto-finishing to remove unwanted core material around the core die splitline. It has been found that this process often removes some, if not all, of the external corner fillet on the core and results in a sharp internal corner in the finished casting. This corner acts as a stress concentration and can initiate cracking of the airfoil trailing edge.
It will be recognized that an attempt to address a similar problem for a turbine blade in a gas turbine engine is disclosed in U.S. Pat. No. 6,062,817, entitled “Apparatus and Methods For Cooling Slot Step Elimination,” which is also owned by the assignee of the present invention. A turbine blade is disclosed therein where at least a portion of a step between an airfoil trailing edge slot and a platform is eliminated. An airfoil core utilized to cast the turbine blade includes a tab for forming a continuous and smooth contour from a first trailing edge slot recessed wall to a juncture of the airfoil. In this way, stress concentration is reduced, thereby improving the longevity and performance of the turbine blade.
Thus, in light of the foregoing, it would be desirable for an improved airfoil design to be developed for use with a turbine nozzle which reduces stress concentrations at the steps of the cooling slots located adjacent the inner and outer nozzle bands without adversely affecting the cooling flow from such slots. It would also be desirable to modify the core utilized so as to eliminate the opportunity for additional stress concentrations created by the auto-finishing manufacturing process.
BRIEF SUMMARY OF THE INVENTION
In a first exemplary embodiment of the invention, an airfoil for a turbine nozzle assembly of a gas turbine engine is disclosed as including an outer side wall, an inner side wall, a leading edge extending from the outer side wall to the inner side wall, a trailing edge extending from the outer side wall to the inner side wall, a concave surface extending from the leading edge to the trailing edge on a pressure side of the airfoil, a convex surface extending from the leading edge to the trailing edge on a suction side of the airfoil, an outer cooling slot, an inner cooling slot, and at least one middle cooling slot formed in the concave side of the airfoil adjacent the trailing edge. Each of the cooling slots also includes a recessed wall, an inner slot side wall, an outer slot side wall, an inner corner fillet located between the inner slot side wall and the recessed wall, and an outer corner fillet located between the outer slot side wall and the recessed wall, wherein one of the inner and outer corner fillets of at least one of the inner and outer cooling slots forms a variable contour from an opening in the concave surface to an exit plane of the trailing edge cooling slots. More specifically, the corner fillet forming the variable contour is radiused in a first plane substantially perpendicular to the slot exit plane from the opening to the exit plane. The airfoil also includes a junction between the corner fillet forming the variable contour and an end portion of the airfoil, wherein the junction is radiused in a second plane substantially perpendicular to the slot exit plane from the opening to the exit plane.
In a second exemplary embodiment of the invention, an airfoil core for a turbine airfoil is disclosed as including a wedge channel for forming a hollow portion of an airfoil and a plurality of fingers extending from the wedge channel, wherein at least one of the fingers located at an end is configured to have a distal portion with a predetermined radius from a first side wall to a second side wall. The distal portion of the finger is radiused in a first plane substantially perpendicular to an axis through the finger and radiused in a second plane substantially parallel to the axis through the finger.
In a third exemplary embodiment of the invention, a method of fabricating an airfoil of a turbine nozzle is disclosed as including the steps of inserting a mold within a die and injecting a slurry into the die. An airfoil is formed that includes an outer side wall, an inner side wall, a leading edge extending from the outer side wall to the inner side wall, a trailing edge extending from the outer side wall to the inner side wall, a concave surface extending from the leading edge to the trailing edge on a pressure side of the airfoil, a convex surface extending from the leading edge to the trailing edge on a suction side of the airfoil, and a plurality of cooling slots formed in the concave side of the airfoil adjacent the trailing edge, each of the cooling slots further including a recessed wall and a pair of slot side walls, and a variable contour for a corner fillet between the recessed wall and one of the slot side walls of a cooling slot adjacent at least one of the inner and outer side walls of the airfoil from an opening in the concave surface to an exit plane of the trailing edge cooling slots. In this way, the corner fillet is formed with a radius in a first plane substantially perpendicular to the slot exit plane that gradually increases from a minimum radius at the opening to a maximum radius at the slot exit plane. The method also includes the step of forming a junction between the corner fillet and an end portion of the airfoil, wherein the junction is radiused in a second plane substantially perpendicular to the slot exit plane from the opening to the exit plane.
REFERENCES:
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patent: 6183192 (2001-02-01), Tressler et al.
Heffron Todd S.
Morgan Clive A.
Andes William Scott
Davidson James P.
General Electric Company
Look Edward K.
White Dwayne J.
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