Metal founding – Process – Shaping liquid metal against a forming surface
Reexamination Certificate
1999-03-22
2002-01-22
Elve, M. Alexandria (Department: 1725)
Metal founding
Process
Shaping liquid metal against a forming surface
C164S122100
Reexamination Certificate
active
06340047
ABSTRACT:
BACKGROUND OF THE INVENTION
The present invention relates generally to gas turbine engines, and, more specifically, to casting of turbine airfoils therein.
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 downstream through multiple turbine stages that extract energy therefrom. Since the turbine stages are heated by the hot combustion gases, they are typically internally cooled by using a portion of the pressurized air bled from the compressor.
A typical turbine stage includes an annular turbine stator or nozzle having a plurality of circumferentially spaced apart nozzle vanes extending radially between outer and inner bands. Disposed downstream from the nozzle is a row of circumferentially spaced apart turbine rotor blades extending radially outwardly from a supporting rotor disk.
The vanes and blades define airfoils having respective aerodynamic geometries for maximizing efficiency of energy extraction from the combustion gases. A typical airfoil includes a generally concave, pressure side and an opposite, generally convex, suction side extending axially between leading and trailing edges, and radially between a root and a tip.
In a nozzle vane, the airfoil extends radially between the outer and inner bands and is typically formed in a one-piece casting. In a rotor blade, the airfoil tip is spaced from a surrounding turbine shroud, with the root of the airfoil being integrally formed with a dovetail which mounts the blade in a complementary dovetail slot formed in the perimeter of the rotor disk.
Since turbine blades rotate during operation they are subject to considerable centrifugal force and corresponding stress, with the force increasing the complexity of cooling the blade. A typical blade includes an internal cooling circuit formed by multiple, radially extending flow passages or channels through which the cooling air is channeled. The blade airfoil is initially internally cooled by the air which is then discharged through various holes extending though the walls of the airfoil.
Due to the aerodynamic profile of the airfoil, the heat transfer coefficient between the hot combustion gases and the airfoil varies over the pressure and suction sides between the leading and trailing edges and between the root to tip. Accordingly, the internal cooling circuit varies in complexity for best utilizing the limited cooling air to cool the different portions of the airfoil differently in response to the varying heat influx from the combustion gases. Many compromises must be made in defining the internal cooling circuit due to the aerodynamic limitations of channeling the cooling air therethrough, and while balancing the centrifugal and thermal stress experienced by the blade during operation.
A high pressure turbine rotor blade typically includes a dedicated cooling passage or channel behind its leading edge, a dedicated cooling passage behind its trailing edge, and a multi-pass serpentine cooling passage disposed axially therebetween and extending radially between the root and tip of the blade airfoil. The flow passages typically also include turbulators in the form of small ribs extending from the inside surface of the airfoil which trip a portion of the cooling air as it flows radially through the cooling passages for enhancing cooling air heat transfer. The airfoil typically includes several radial rows of film cooling holes extending through the walls thereof for discharging the internal cooling air in corresponding films along the outer surface of the airfoil for providing film cooling thereof.
In order to precisely form the external and internal features of the airfoil, turbine rotor blades are typically cast using high-strength superalloys. In the lost wax method of casting, a ceramic casting core is initially molded to precisely define the internal cooling circuit, including any turbulators or other features desired. The core is then surrounded by wax to define the desired metal portions of the blade, and the wax is then surrounded by a ceramic outer shell.
The wax is removed, and molten metal is injected into the space previously occupied by the wax. The metal solidifies, the shell is removed, and the core is leached away leaving behind the cast blade, including its airfoil and dovetail having the desired precise configurations thereof, both externally and internally. The various holes in the airfoil, such as the film cooling holes, may then be suitably drilled therein.
Some turbine blades, such as stage two blades, have relatively long airfoils which require relatively long casting cores. Since the typical casting core includes multiple legs for matching the multiple internal flow channels of the airfoil, the legs are slender and subject to movement and breakage during the casting process. Misaligned core legs correspondingly change the dimensions of the resulting airfoil, and can lead to out-of-specification locally thick or thin regions for which the airfoil may be rejected. And, core breakage during the casting process also may result in rejection of the cast blade.
As a solution to this problem, it is known to provide one or more core ties between adjacent legs to fixedly join together the legs for reducing undesirable movement therebetween during the casting process and reducing the likelihood of core breakage. However, the ties necessarily define a corresponding tie hole in the intermediate airfoil rib through which a portion of the cooling air being channeled through the flow channels is short circuited. Cooling air short circuits in the complex internal flow channels reduce the cooling efficiency of the available air and correspondingly adversely affect the useful life of the blade during operation.
Accordingly, it is desired to provide an improved method of casting turbine airfoils which reduces the adverse effects of core ties used in the casting thereof.
BRIEF SUMMARY OF THE INVENTION
A gas turbine engine airfoil is cast around a core having a plurality of legs to form matching flow channels in the airfoil. The legs have a tie extending therebetween to maintain alignment. And, the tie is relocated along the core span to reduce differential static pressure of the cooling air across the resulting tie hole formed by the core tie.
REFERENCES:
patent: 3533712 (1970-10-01), Kercher
patent: 4148350 (1979-04-01), Rossman
patent: 4474532 (1984-10-01), Pazder
patent: 4627480 (1986-12-01), Lee
patent: 5243759 (1993-09-01), Brown et al.
patent: 5337805 (1994-08-01), Green et al.
patent: 5358029 (1994-10-01), Baveja et al.
patent: 5505250 (1996-04-01), Jago
patent: 6186741 (2001-02-01), Webb et al.
Andes William Scott
Elve M. Alexandria
General Electric Company
Hess Andrew C.
Tran Len
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