Patent Grant
8157514
The present invention relates to gas turbine engines and, more particularly, to components for gas turbine engines.
A gas turbine engine may be used to power various types of vehicles and systems. One 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, namely, a fan section, a compressor section, a combustor section, a turbine section, and an exhaust section. Other gas turbine engines may not include a fan section, and thereby may include four major sections, namely, a compressor section, a combustor section, a turbine section, and an exhaust section.
The fan section, if applicable, 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 and/or from another source or inlet 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 (if applicable), the compressor section, and the turbine section, typically include a plurality of rotor blades coupled to a rotor disk that is configured to rotate. Such gas turbine engine components may experience stress from operation of the gas turbine engine, such as when portions of the component experience a significantly different range of temperatures from one another.
Accordingly, there is a need for an improved gas turbine engine and/or turbine engine component with a mechanism to help alleviate stress during operation. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.
In accordance with an exemplary embodiment of the present invention, a component for a gas turbine engine having an engine axis is provided. The component comprises a rotor disk and a plurality of airfoils. The rotor disk comprises a web and a rim. The web has a first outer surface at least partially defining a plurality of holes and a plurality of slots. Each of the plurality of slots extends from a corresponding one of the plurality of holes and forms a first angle with the engine axis at the point of intersection with the corresponding one of the plurality of holes. The rim has a second outer surface also at least partially defining the plurality of slots. Each of the plurality of slots forms a second angle with the engine axis at the second outer surface, the second angle being different from the first angle. Each of the plurality of airfoils extends from the second outer surface.
In accordance with another exemplary embodiment of the present invention, a turbine section for a gas turbine engine having an engine axis is provided. The turbine section comprises a rotor disk and a plurality of turbine blades. The rotor disk comprises a web and a rim. The web has a first outer surface at least partially defining a plurality of holes and a plurality of slots. Each of the plurality of slots extends from a corresponding one of the plurality of holes and forms a first angle with the engine axis at the point of intersection with the corresponding one of the plurality of holes. The rim has a second outer surface also at least partially defining the plurality of slots. Each of the plurality of slots forms a second angle with the engine axis at the second outer surface, the second angle being different from the first angle. Each of the plurality of turbine blades extends from the second outer surface.
In accordance with another exemplary embodiment of the present invention, a gas turbine engine is provided. The gas turbine engine has an engine axis, and comprises a compressor, a combustor, and a turbine. The compressor has an inlet and an outlet. The compressor is 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 turbine comprises a rotor disk and a plurality of turbine blades. The rotor disk comprises a web and a rim. The web has a first outer surface at least partially defining a plurality of holes and a plurality of slots. Each of the plurality of slots extends from a corresponding one of the plurality of holes and forms a first angle with the engine axis at the point of intersection with the corresponding one of the plurality of holes. The rim has a second outer surface also at least partially defining the plurality of slots. Each of the plurality of slots forms a second angle with the engine axis at the second outer surface, the second angle being different from the first angle. Each of the plurality of turbine blades extends from the second outer surface.
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 turbine 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 gas turbine jet engine 100 is depicted in
While the gas turbine 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 gas turbine 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 gas turbine 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 gas turbine engine 100 of
The rotor component 200 is depicted in
As depicted in
The web 206 has a first outer surface 210 depicted in
The rim 208 has a second outer surface 216. The second outer surface 216 also at least partially defines the plurality of slots 214, such that each of the plurality of slots 214 forms a second angle B with respect to the engine axis 201 at the second outer surface 216. In a preferred embodiment, the second angle B is different from the first angle. Most preferably, the second angle B is greater than the first angle. For example, in one exemplary embodiment in which the first angle A is equal to zero, the second angle B is equal to fifteen degrees. However, this may vary in other embodiments.
Accordingly, and as depicted in
In a preferred embodiment, the second angle B is at least approximately equal to the angle between a line formed by the tangency points of the airfoil 204 leading and trailing edges at the second outer surface 216 and the engine axis 201 (commonly referenced in the field as the stagger angle), so that each of the slots 214 is at least approximately parallel to the flow path at the rim 208 and the second outer surface 216 thereof. Also in a preferred embodiment, each of the slots 214 is aligned with and parallel to its corresponding hole 212 at the point of intersection of each slot 214 with its corresponding hole 212, such that each of the slots 214 and their corresponding holes 212 are aligned not only with one another but also with the engine axis 201 (and preferably with the first angle A being at least approximately equal to zero, as discussed above).
The angular rotation of the slots 214 and the alignment of the holes 212 and slots 214 with one another and the engine axis 201 provide for improved performance and/or durability of the rotor component 200 and/or for the engine with which the rotor component 200 is utilized. First, the slots 214 provide optimal stress relief from the flow path due to the alignment of the slots 214 with the flow path at the rim 208. Also, the slots 214 provide for optimal durability due to the alignment of the holes 212 with the engine axis 201 and the alignment of the slots 214 with the engine axis 201 at the points in with each of the slots 214 intersects with its corresponding hole 212. Accordingly, these features provide for a reduction in peaking of stresses in edges of each of the holes 212. In addition, this reduction in stress increases the fatigue capability of the rotor component 200, thereby also allowing for the use of an integral dual alloy or cast turbine rotor component 200 to be used if desired.
In the depicted embodiment, each of the plurality of airfoils 204 extends from the second outer surface 216 of the rim 208 in a direction that is generally radially outward from the web 206. In the depicted embodiment, each of the plurality of airfoils 204 extends from a portion of the second outer surface 216 of the rim 208 between two corresponding slots 214 surrounding the portion of the second outer surface 216. Thus, in the depicted embodiment, the second outer surface 216 of the rim 208 alternates between airfoils 204 and slots 214 that extend in generally opposite directions around the perimeter of the rotor disk 202 as shown in
In one preferred embodiment, each of the airfoils 204 comprises a turbine blade, and the rotor component 200 is configured for use in one or more turbines of an engine, such as one or more turbines of the turbine section 108 of the gas turbine engine 100 of
Accordingly, improved rotor components 200 are provided for use in a turbine section, a compressor section, a fan section, and/or another rotor section of a gas turbine engine. The improved rotor components provide for an improved combination of stress relief and durability as a result of the unique angular rotation of the slots 214 and the alignment of the holes 212 and slots 214 with one another and the engine axis 201. Also, improved gas turbine engines 100 are provided with such improved rotor components 200. Accordingly, as noted above, these features provide for a reduction in peaking of stresses in edges of each of the holes 212. In addition, and also as noted above, this reduction in stress increases the fatigue capability of the rotor component 200, thereby also allowing for the use of an integral dual alloy or cast turbine rotor component 200 to be used if desired.
It will be appreciated that the rotor components 200 and engines 100 may differ from those depicted in the Figures and described herein in connection therewith. It will further be appreciated that the rotor components 200 may be implemented in connection with any number of different sections of any number of different types of engines.
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.
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| Number | Date | Country | |
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| 20100239422 A1 | Sep 2010 | US |
