Patent Application
20060280610
This invention relates generally to turbine assemblies, and more particularly, to methods and apparatus for fabricating gas turbine engine rotor blades.
Gas turbine engines typically include a compressor, a combustor, and at least one turbine. The compressor compresses air which is mixed with fuel and channeled to the combustor. The mixture is then ignited for generating hot combustion gases, and the combustion gases are channeled to the turbine which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator.
The turbine includes a rotor assembly and a stator assembly. The rotor assembly includes a plurality of rotor blades extending radially outward from a disk. More specifically, each rotor blade extends radially between a platform adjacent the disk and a blade tip. A combustion gas flowpath through the rotor assembly is bound radially inward by the rotor blade platforms, and radially outward by a plurality of shrouds, wherein each shroud includes at least one seal tooth.
At least one known gas turbine blade is fabricated utilizing a nickel based alloy to produce a turbine blade that has a substantially curved outer shroud. For example, the turbine blade is cast with additional stock that is removed utilizing a simple radial grind operation to produce the finished outer shroud and the seal teeth. More specifically, the nickel based alloy is cast “to size”, i.e. the turbine blade does not require significant machining to produce the finished turbine blade.
However, if the gas turbine blade is fabricated utilizing a different material, for example, a titanium aluminide material, casting a turbine blade that includes a substantially curved shroud and at least one seal tooth, produces a turbine blade that has no “as-cast” surfaces. More specifically, casting a turbine blade utilizing titanium aluminide produces a turbine blade that has a large amount of excess material relative to the final machined configuration, shown in
In one aspect, a method for fabricating a turbine blade for a gas turbine engine is provided. The gas turbine engine includes a turbine including a plurality of turbine blades. The method includes casting at least one turbine blade that includes a blade tip shroud and at least one seal tooth coupled to the blade tip shroud, and removing at least a portion of the turbine blade such that a radially outer surface of the blade tip shroud has at least two substantially flat surfaces.
In another aspect, a turbine rotor blade is provided. The turbine rotor blade includes a blade tip shroud having a radially outer surface including at least two substantially flat surfaces, and at least one seal tooth coupled to said blade tip shroud.
In a further aspect, a gas turbine engine rotor assembly is provided. The rotor assembly includes a low pressure turbine rotor, and a plurality of circumferentially-spaced rotor blades coupled to the low pressure turbine rotor. Each rotor blade includes having a radially outer surface including at least two substantially flat surfaces, and at least one seal tooth coupled to said blade tip shroud.
In operation, air flows through fan assembly 12 and compressed air is supplied to high pressure compressor 14 through booster 22. The highly compressed air is delivered to combustor 16. Hot products of combustion (not shown in
Rotor blades 56 are configured for cooperating with a motive or working fluid, such as air. In the exemplary embodiment, rotor assembly 40 is a turbine, such as low pressure turbine 20 (shown in
Blades 56 rotate about the axial centerline axis up to a specific maximum design rotational speed, and generate centrifugal loads in rotating components. Centrifugal forces generated by rotating blades 56 are carried by portions of rims 50 directly below each rotor blade 56. Rotation of rotor assembly 40 and blades 56 extracts energy from the air which causes turbine 20 to rotate and provide power to drive low pressure compressor 12 (shown in
Rotor blades 56 each include a leading edge 60, a trailing edge 62, and an airfoil 64 extending therebetween. Each airfoil 64 includes a suction side 76 and a circumferentially opposite pressure side 78. Suction and pressure sides 76 and 78, respectively, extend between axially spaced apart leading and trailing edges 60 and 62, respectively and extend in radial span between a rotor blade tip shroud 80 and a rotor blade root 82. A blade chord is measured between rotor blade trailing and leading edges 62 and 60, respectively. In the exemplary embodiment, rotor blades 56 include rotor seal teeth 86 which rotate adjacent to a stator shroud 88 and through a cavity 89 defined by stator shroud 88 and rotor blade tip shroud 80.
In the exemplary embodiment, gas turbine blade 110 is cast utilizing a metallic material. In the exemplary embodiment, gas turbine blade 110 is cast utilizing a titanium aluminide material, for example. In the exemplary embodiment, turbine blade 110 is “overcast”, that is turbine blade 110 includes a portion 116 that must be removed to produce a finished blade 110.
In the exemplary embodiment, after turbine blade 110 is cast, excess portion 116 is removed utilizing an electron discharge machining (EDM) apparatus 122. More specifically, EDM apparatus 122 is configured to remove excess portion 116 such that blade tip shroud 112 and at least one seal tooth 114 are formed, and such that radially outer surface 118 of blade tip shroud 112 has a substantially V-shaped cross-sectional profile 120.
More specifically, blade tip shroud 112 if formed such that blade tip shroud outer surface 118 includes a first non-arcuate portion 130, i.e. a flat, and a second non-arcuate portion 132, or flat, that are coupled together at an apex 134 such that outer surface 118 has a substantially V-shaped cross-sectional profile 120. In the exemplary embodiment, first and second portions 130 and 132 are formed unitarily with turbine blade 110. More specifically, EDM apparatus 122 is utilized to remove portion 116 such that outer surface 118 has a substantially V-shaped cross-sectional profile.
Moreover, each seal tooth 114 includes a first non-arcuate portion 140, or flat, and a second non-arcuate portion 142, or flat, that are coupled together at an apex 144 such each seal tooth 114 has a substantially V-shaped cross-sectional profile 120. In the exemplary embodiment, first and second portions 140 and 142 are formed unitarily with turbine blade 110. More specifically, EDM apparatus 122 is utilized to remove portion 116 such that each seal tooth 114 has a substantially V-shaped cross-sectional profile 120.
The methods described herein facilitate improving the form of the seal teeth and the shroud non-flowpath surface of a gas turbine blade such that the turbine blade can be fabricated utilizing a wire cut EDM apparatus. More specifically, known EDM machines can not produce curved surfaces in the circumferential direction of the turbine blade. Accordingly, the turbine blade described herein does not include the curved form of the shroud and seal teeth of a known turbine blade, rather the arched or curved radially outer surface of the outer shroud is replaced by two flat surfaces, wherein each flat surface spans half the circumferential width of the shroud. Fabricating a turbine blade, that includes an outer shroud having a radially outer surface that is formed by two flat surfaces, facilitates reducing leakage between the seal teeth and the stator shroud. Moreover, although the exemplary embodiment, illustrates an outer shroud and seal teeth each having two substantially flat surfaces, it should be realized that the exemplary turbine blade can include three or more flat surface that define the radially outer surface of the outer shroud and seal teeth.
Accordingly, the methods and apparatus described herein facilitate reducing the amount of machining that is needed to produce LPT Blades from TiAl material, thus reducing cost and cycle time to fabricate a turbine blade. Moreover, the reduced cost and cycle time to fabricate the turbine blade described herein also facilitates reducing the overall weight of the gas turbine engine.
The above-described methods and apparatus are cost-effective and highly reliable. The turbine blade described herein includes a radially outer surface and at least one seal tooth that each include at least two substantially flat surfaces to facilitate reducing the time and cost of fabricating a gas turbine engine.
While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.
