Patent Application
20080273982
The present invention relates to gas turbine engines and, more particularly, to improved gas turbine engine components.
A gas turbine engine may be used to power various types of vehicles and systems. A particular type of gas turbine engine that may be used to power aircraft is a turbofan gas turbine engine. A turbofan gas turbine engine may include, for example, five major sections, a fan section, a compressor section, a combustor section, a turbine section, and an exhaust section.
The fan section is positioned at the front, or “inlet” section of the engine, and includes a fan that induces air from the surrounding environment into the engine, and accelerates a fraction of this air toward the compressor section. The remaining fraction of air induced into the fan section is accelerated into and through a bypass plenum, and out the exhaust section. The compressor section raises the pressure of the air it receives from the fan section to a relatively high level. The compressed air from the compressor section then enters the combustor section, where a ring of fuel nozzles injects a steady stream of fuel. The injected fuel is ignited by a burner, which significantly increases the energy of the compressed air.
The high-energy compressed air from the combustor section then flows into and through the turbine section, causing rotationally mounted turbine blades to rotate and generate energy. Specifically, high-energy compressed air impinges on turbine vanes and turbine blades, causing the turbine to rotate. The air exiting the turbine section is exhausted from the engine via the exhaust section, and the energy remaining in this exhaust air aids the thrust generated by the air flowing through the bypass plenum.
Certain of these gas turbine engine components, such as the fan section, the compressor section, and the turbine section, typically include a plurality of rotor blades coupled to a rotor disk that is configured to rotate. In certain gas turbine engines, the rotor blades are slidably disposed in various slots formed in the rotor disk. While such a configuration is generally effective, it is possible that the blades may experience inadvertent movement in an axial direction during engine operation.
Accordingly, there is a need for an improved turbine engine and/or turbine engine component with a mechanism to prevent, or at least inhibit, axial movement of blades while the turbine engine is operating. The present invention addresses at least this need.
The present invention provides a component for a gas turbine engine. In one embodiment, and by way of example only, the component comprises a rotor disk, a plurality of blade attachments, a plurality of retention flanges, and a plurality of retention tabs. The rotor disk has a plurality of slots formed in an outer surface thereof. Each blade attachment has an attachment region configured to be partially disposed within one of the slots. The plurality of retention flanges extend from the rotor disk outer surface, to thereby at least inhibit axial movement of the blade attachment, when the attachment region is fitted inside its corresponding slot. A retention tab extends from each attachment region and engages one of the retention flanges.
The invention also provides a fan section for a gas turbine engine. In one embodiment, and by way of example only, the fan section comprises an inlet, a rotor disk, a plurality of blade attachments, a plurality of retention flanges, and a plurality of retention tabs. The inlet is adapted to receive air from a surrounding environment. The rotor disk has a plurality of slots formed in an outer surface thereof. Each blade attachment has an attachment region configured to be partially disposed within one of the slots. The plurality of retention flanges extend from the rotor disk outer surface, to thereby at least inhibit axial movement of the blade attachment, when the attachment region is fitted inside its corresponding slot. A retention tab extends from each attachment region and engages one of the retention flanges.
The invention also provides a gas turbine engine. In one embodiment, and by way of example only, the gas turbine engine comprises a compressor, a combustor, a turbine, and a fan section. The compressor has an inlet and an outlet and operable to receive accelerated air through the inlet, compress the accelerated air, and supply the compressed air through the outlet. The combustor is coupled to receive at least a portion of the compressed air from the compressor outlet, and is operable to supply combusted air. The turbine is coupled to receive the combusted air from the combustor and at least a portion of the compressed air from the compressor, and to generate energy therefrom. The fan section comprises an inlet, a plurality of blades, a plurality of retention flanges, and a plurality of retention tabs. The inlet is adapted to receive air from a surrounding environment. The rotor disk has an outer surface forming a plurality of slots. The plurality of blades are attached to the rotor disk. Each blade has an attachment region configured to attach the blade to the rotor disk by fitting within a corresponding slot. The plurality of blades are configured to accelerate a portion of the air, and to supply the accelerated air to the compressor. The plurality of retention flanges extend from the rotor disk outer surface, to thereby at least inhibit axial movement of the blade attachment, when the attachment region is fitted inside its corresponding slot. A retention tab extends from each attachment region and engages one of the retention flanges.
Other independent features and advantages of the preferred airfoil and method will become apparent from the following detailed description, taken in conjunction with the accompanying drawings which illustrate, by way of example, the principles of the invention.
Before proceeding with the detailed description, it is to be appreciated that the described embodiment is not limited to use in conjunction with a particular type of turbine engine or in a particular section or portion of a gas turbine engine. Thus, although the present embodiment is, for convenience of explanation, depicted and described as being implemented in a fan section of a turbofan gas turbine jet engine, it will be appreciated that it can be implemented in various other sections and in various types of engines.
An exemplary embodiment of a turbofan gas turbine jet engine 100 is depicted in
The compressor section 104 includes one or more compressors. In the depicted embodiment, the compressor section 104 includes two compressors, an intermediate pressure compressor 120, and a high pressure compressor 122. However, the number of compressors may vary in other embodiments. The intermediate pressure compressor 120 raises the pressure of the air directed into it from the fan 112, and directs the compressed air into the high pressure compressor 122. The high pressure compressor 122 compresses the air still further, and directs a majority of the high pressure air into the combustion section 106. In addition, a fraction of the compressed air bypasses the combustion section 106 and is used to cool, among other components, turbine blades in the turbine section 108. In the combustion section 106, which includes an annular combustor 124, the high pressure air is mixed with fuel and combusted. The high-temperature combusted air is then directed into the turbine section 108.
The turbine section 108 includes one or more turbines. In the depicted embodiment, the turbine section 108 includes three turbines disposed in axial flow series, a high pressure turbine 126, an intermediate pressure turbine 128, and a low pressure turbine 130. However, it will be appreciated that the number of turbines, and/or the configurations thereof, may vary, as may the number and/or configurations of various other components of the exemplary engine 100. The high-temperature combusted air from the combustion section 106 expands through each turbine, causing it to rotate. The air is then exhausted through a propulsion nozzle 132 disposed in the exhaust section 110, providing addition forward thrust. As the turbines rotate, each drives equipment in the engine 100 via concentrically disposed shafts or spools. Specifically, the high pressure turbine 126 drives the high pressure compressor 122 via a high pressure spool 134, the intermediate pressure turbine 128 drives the intermediate pressure compressor 120 via an intermediate pressure spool 136, and the low pressure turbine 130 drives the fan 112 via a low pressure spool 138. As mentioned above, the engine 100 of
With reference first to
Each slot 210 is configured to slidably receive therein a different, corresponding blade attachment 204, and each slot 210 preferably includes a slot opening 211 for insertion of the corresponding blade attachment 204. The slots 210 are separated by disk posts 215, which are also formed in the outer surface 209, and which extend parallel to the slots 210. The disk posts 215 are also preferably made of titanium, steel, or a nickel-based alloy. However, the disk posts 215 may be made of one or more other materials, and/or may vary in size, number, and/or configuration.
With reference now to
The platform 214 of each blade attachment 204 is coupled to a non-depicted airfoil. The plurality of airfoils are configured to rotate when the attachment region 212 of each blade attachment 204 is fit into its corresponding slot 210 and the engine and the rotor component 200 are operating.
With reference again to
With reference again to
When the rotor component 200 is assembled, the retention flanges 206 and retention tabs 208 preferably engage one another. Thus, when the rotor component 200 rotates, each blade attachment 204 is axially held in place inside its corresponding slot 210. During operation, an axial force further urges the retention flanges 206 and the retention tabs 208 together, to further inhibit blade attachment 204 movement in the axial direction. Preferably, the retention flanges 206 and retention tabs 208 prevent any axial movement of the blade attachment 204.
Accordingly, the retention flanges 206 and the retention tabs 208 of the rotor component 200 can help reduce, and preferably prevent, axial movement of the blade attachments 204 during operation. This can be of particular benefit in designs, such as that depicted in
In addition, the design allows for the rotor component 200 to be assembled and disassembled without removing the engine from its environment. For example, in an aircraft environment, the rotor component 200 can be assembled and disassembled on-wing, thereby making assembly, disassembly, inspection, and/or repair potentially easier and/or less expensive. The rotor component 200 can also be implemented in various different components of various different types of gas turbine engines, for example in a fan section, a compressor section, a turbine section, and/or an axle section, and can also be implemented in any of numerous other different environments.
While the invention has been described with reference to a preferred embodiment, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt to a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this invention, but that the invention will include all embodiments falling within the scope of the appended claims.
This invention was made with Government support under contract no. NAS3-01136 awarded by the National Aeronautics Space Administration (NASA). The Government has certain rights in this invention.
