BACKGROUND OF THE INVENTION
This application relates to an integrally bladed rotor, such as utilized in gas turbine engines, wherein an outer rim has a discontinuity.
Gas turbine engines typically include a plurality of sections mounted in series. A fan section may deliver air to a compressor section. The compressor section may include high and low compression stages, and delivers compressed air to a combustion section. The air is mixed with fuel in the combustion section and burned. Products of this combustion are passed downstream over turbine rotors.
The compressor section includes a plurality of rotors having a plurality of circumferentially spaced blades. Recently, these rotors and blades have been formed as an integral component, called an “integrally bladed rotor.”
In one known integrally bladed rotor, blades extend from an outer rim. The outer rim in integrally bladed rotors is subject to a number of stresses, and in particular, hoop stresses. The hoop stresses can cause the life of the integrally bladed rotor to be reduced due to thermal fatigue.
SUMMARY OF THE INVENTION
In the disclosed embodiment of this invention, discontinuities are formed in the outer rim of an integrally bladed rotor. In the disclosed embodiment, the discontinuity extends through the entire axial and radial width of the outer rim.
These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 schematically shows a gas turbine engine.
FIG. 2 shows an integrally bladed rotor according to an embodiment of the present invention.
FIG. 3 shows a detail of the inventive integrally bladed rotor.
FIG. 4 is a perspective view of the FIG. 3 integrally bladed rotor.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
FIG. 1 shows a gas turbine engine 10. As known, a fan section 14 moves air and rotates about an axial center line 12. A compressor section 16, a combustion section 18, and a turbine section 20 are also centered on the axial center line 12. FIG. 1 is a highly schematic view; however, it does show the main components of the gas turbine engine. Further, while a particular type of gas turbine engine is illustrated in FIG. 1, it should be understood that the present invention extends to other types of gas turbine engines.
FIG. 2 shows an integrally bladed rotor 80, such as may be utilized for the high stage compression section. The integrally bladed rotor 80 includes an outer rim 82, a plurality of circumferentially distributed blades 84, a central hub 48, and a plurality of channels 86. The channels 86 extend through the axial width of the rotor 80. Channels 86 and discontinuities 88, 90 and 92 (see FIGS. 3 and 4) address the hoop stresses discussed earlier.
FIG. 3 shows integrally bladed rotor 80. In integrally bladed rotor 80, a discontinuity 88, 90, 92 is formed through a radial extent of the outer rim 82. As shown, a central enlarged, seal holding portion 90 is formed between two smaller slots 88 and 92. As can be appreciated, the radially inner slot 92 extends to the channel 86.
As shown in FIG. 4, the outer slot 88 extends across the axial width of the rotor 80. Seals 96 may be inserted in the enlarged portion 90 of the discontinuity. The seal 96 is shown as a wire seal, however, other seals, such as brush seals or W seals, may be utilized. The seals prevent recirculation of gases from the radially outer face of the outer rim 82 into the channels 86.
Although embodiments of this invention have been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.
Patent Application
20110182745
