Turbomachinery component

A turbomachinery blade for a gas turbine engine is provided and includes an airfoil extending between a leading edge and a trailing edge. In one embodiment the turbomachinery blade is a compressor blade. The blade can include a platform attached to the airfoil on one side, the other side being attached to a stalk having a lower attachment portion useful for being received in a compressor disk. The blade includes an undercut beneath a portion of the airfoil, preferably beneath the leading edge and/or trailing edge of the airfoil. In one form the undercut is located in a corner of the platform and extends partially along two sides of the platform.

TECHNICAL FIELD

The present invention relates to rotating gas turbine engine components, and more particularly, but not exclusively, to reducing vibratory stresses in rotating compressor blades of gas turbine engines.

BACKGROUND

Improving the ability of gas turbine engine rotating components to withstand stresses, such as vibratory stresses for example, remains an area of interest. Some existing systems, however, have various shortcomings relative to certain applications. Accordingly, there remains a need for further contributions in this area of technology.

SUMMARY

One embodiment of the present invention is a unique turbomachinery blade. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for reducing stresses in a turbomachinery blade. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

DETAILED DESCRIPTION OF THE ILLUSTRATIVE EMBODIMENTS

Referring toFIG. 1, there is illustrated one embodiment of a gas turbine engine20which includes a fan section21, a compressor section22, a combustor section23, and a turbine section24that are integrated together to produce an aircraft flight propulsion engine. This type of gas turbine engine is generally referred to as a turbo-fan. Other types of gas turbine engines are also contemplated, such as, but not limited to, turboprops, turbojets, and turboshafts. It is important to realize that there are a multitude of ways in which the gas turbine engine components can be linked together. The gas turbine engine can have any number of spools. One form of a gas turbine engine includes a compressor, a combustor, and a turbine that have been integrated together to produce an aircraft flight propulsion engine. As used herein, the term “aircraft” includes, but is not limited to, helicopters, airplanes, fixed wing vehicles, variable wing vehicles, rotary wing vehicles, unmanned combat aerial vehicles, tailless aircraft, hover crafts, and other vehicles. Further, the present inventions are contemplated for utilization in other applications that may not be coupled with an aircraft such as, for example, industrial applications, power generation, pumping sets, naval propulsion, weapon systems, security systems, perimeter defense/security systems, and the like known to one of ordinary skill in the art.

The compressor section22includes a rotor25having a plurality of compressor blades26coupled thereto. The rotor25is affixed to a shaft27that is rotatable within the gas turbine engine20. A plurality of compressor vanes28are positioned within the compressor section22to direct the fluid flow relative to blades26. Turbine section24includes a plurality of turbine blades30that are coupled to a rotor disk31. The embodiment of the turbine section24depicted inFIG. 1includes a relatively low pressure turbine and a relatively high pressure turbine. The rotor disk31is affixed to the shaft27, which is rotatable within the gas turbine engine20. Energy extracted in the turbine section24from the hot gas exiting the combustor section23is transmitted through shaft27to drive the compressor section22. Further, a plurality of turbine vanes32are positioned within the turbine section24to direct the hot gaseous flow stream exiting the combustor section23.

The turbine section24provides power to a fan shaft33, which drives the fan section21. The fan section21includes a fan34having a plurality of fan blades35. Air enters the gas turbine engine20in the direction of arrows A and passes through the fan section21into the compressor section22and a bypass duct36. Further details related to the principles and components of a conventional gas turbine engine will not be described herein as they are believed known to one of ordinary skill in the art.

Referring toFIG. 2and with continuing reference toFIG. 1, as previously set forth, the compressor stage22may include a rotor or compressor wheel assembly116. A cross sectional view of a portion of the compressor wheel assembly116positioned in compressor housing124is set forth inFIG. 2. The compressor wheel assembly116preferably comprises a compressor wheel126and one or more compressor blades128. The orientation of the compressor blades128and the compressor wheel126is such that air flows in a generally axially aft direction as indicated by the arrow ofFIG. 2labeled “AIR FLOW”. The compressor wheel126is generally normal to air flow and extends circumferentially about a center axis of the gas turbine engine. The compressor blade128depicted inFIG. 2can be a compressor blade from any location within the compressor section22. To set forth just one non-limiting example, the compressor blade128depicted inFIG. 2can be coupled to a fourth rotor in a multi-rotor compressor.

The compressor wheel126includes a blade retaining slot130disposed therein. In the illustrative embodiment, the blade retaining slot130preferably has a dovetail shape. Other slot configurations and/or shapes are contemplated as within the scope of the present application. As discussed further below, an attachment portion of the compressor blades128fits within and engages the blade retaining slot130, the compressor blades128extending circumferentially around a center axis of the gas turbine engine20. Although not illustrated, in some forms the compressor wheel126and blades128can be formed as a unitary whole.

Referring collectively toFIGS. 2 and 3, each compressor blade128includes an airfoil section132and a stalk138. The stalk138includes a root section134with an attachment portion141. The stalk138also includes a platform portion136that provides a surface for the smooth passage of airflow thereover. The root section134is mountable within the blade retaining slot130and may be inserted therein through a loading slot (not shown). The root section134has an attachment portion141that fits within and engages the blade retaining slot130of the compressor wheel126as illustrated. An upper portion of the stalk138defines the platform136of the compressor blade128. The platform136extends between first end139and opposite second end140. In one form when completely assembled, adjacent compressor blades128are preferably positioned so that platforms136of adjacent compressor blades128abut one another.

The airfoil section132of each compressor blade128includes a leading edge142and a trailing edge144. The airfoil section132includes a number of characteristics such as, but not limited to, sweep, camber and twist, to set forth just a few non-limiting examples. In one form the airfoil section132can be highly swept. In any event, various embodiments of the airfoil section132can have a variety of different characteristics.

The stalk138includes a stalk leading edge section146and a stalk trailing edge section148. An upper portion of the stalk leading edge section146and the stalk trailing edge section148define a portion of the platform136and extends beyond the root134to the first and second opposite ends139and140, respectively. In some forms the stalk leading edge section146is positioned below a portion of the leading edge142of the airfoil section132and the stalk trailing edge section148is positioned below a portion of the trailing edge144of the airfoil section132.

As illustrated inFIG. 3, the compressor blade128is shown having one embodiment of an undercut150that can be used in some applications to mitigate the effects of vibrations such as the stresses that accompany vibrations. Though the undercut150is shown relative to a compressor blade in the illustrative embodiment, it can be used in other types of gas turbine engine blades. In the illustrative embodiment the undercut150is positioned beneath the trailing edge144of the airfoil section132which in some applications is a critical area of stress. In other embodiments the undercut150can alternatively and/or additionally be positioned beneath the leading edge142which can also be a critical area of stress. The undercut150serves to vary the stiffness of the structure. In one form the stiffness of the structure away from the undercut150drives the load path away from that area. When the undercut area is placed under a trailing edge or leading edge of the blade128, the stiffness is driven away thus driving the load path away from that area. The reduction in load across that area reduces vibratory stress for a given vibration. The undercut150can be created by removing some amount of material from the blade128after it is formed, and/or forming the blade128at the same time as at least some portion of the undercut150. To set forth just a few non-limiting examples, the undercut150can be formed by milling away select portions of the stalk and/or platform. The blade128having the undercut150can also be cast, forged, or assembled from separate pieces (airfoil, platform, stalk, root section). The undercut can be formed in the platform136, the stalk138, or both.

The illustrated shape of undercut150is exemplary and other shapes are contemplated as within the scope of the application. In one embodiment the undercut150begins at first end139and extends only a portion of the way toward opposite second end140. However, it is also contemplated as within the scope of the application that the undercut might not include either of ends139and140, but instead only span some portion of the length between the two ends. The depth, width and thickness of the undercut150may be tailored as desired to achieve a desired property, such as a high cycle fatigue design requirements for a respective gas turbine engine. In some embodiments the undercut150can be disposed equally on either side of the leading edge142and/or trailing edge144. In some forms the undercut150can be positioned unequally on either side of the leading edge142and/or trailing edge144. The undercut150can also extend along the blade128to any given location along either or both sides of the platform136. In some cases such location can be referred to as a chord location. Various other shapes and combinations are contemplated.

As illustrated inFIG. 3, the undercut150may include an upper surface152, a side surface154, and a back surface156. The height, width and depth of the undercut150defines the position of the upper surface152, the side surface154, and the back surface156. In the illustrative embodiment the upper surface152includes a portion of the lower surface of platform136. The side surface154can be positioned within the stalk138a predetermined distance from a trailing outside edge of the platform136. The back surface156may be positioned at a predetermined depth within the stalk138from an end139of the platform136. Though the top surface152and back surface156are shown having relatively flat shapes, other embodiments can have a variety of other shapes. In addition, though the side surface154is shown having a curvilinear shape, in other embodiments the side surface154can have other shapes. Not all embodiments need have a well defined upper surface152, side surface154, or back surface156. In some forms the undercut150can take other forms such as a scoop or scallop. In short, the undercut150can have a variety of shapes, forms, and sizes.

As illustrated inFIG. 2, having a complete platform136may be useful in some embodiments because of the need to create a fluid tight seal, or relatively fluid tight seal, between a lower surface158of the platform136and the compressor wheel126. As previously mentioned, during assembly, a plurality of compressor blades128are preferably positioned in the compressor wheel126such that adjacent compressor blades128will be positioned so that the platforms136of adjacent compressor blades128abut one another at respective ends139and140.

In one aspect of the present application the vibration mitigating undercut can be formed on an underside surface of the platform beneath the leading edge and/or trailing edge of the airfoil. The stiffness of the stalk away from the undercut drives the load path created during operation of the gas turbine engine away from the leading and/or trailing edge of the airfoil. The reduction in load across the critical areas reduces the vibratory stress in the critical feature for a given vibration. The depth, width and thickness of the undercut can be tailored to achieve high cycle fatigue design requirements of gas turbine engines utilizing the compressor blade.

In one embodiment of the application there is a compressor blade for a gas turbine engine. The blade includes an airfoil extending between a leading edge and a trailing edge. The blade further includes a stalk having a lower attachment portion and an upper portion defining a platform. The platform has a first side and a second side. A portion of the first side of the platform is connected to the airfoil. The blade further includes at least one undercut in the stalk beneath a portion of the airfoil.

In one refinement of the application the undercut in the stalk is located beneath at least a portion of the trailing edge of the airfoil.

In another refinement of the application the undercut in the stalk is located beneath at least a portion of the leading edge of the airfoil.

In another refinement of the application the airfoil is highly swept.

In another refinement of the application the platform extends between a first end and a second end. The undercut is in the platform, and the undercut begins at the first end and extends only a portion of the way toward the second end.

In another refinement of the application the undercut is in the platform and the undercut is located beneath at least a portion of the leading edge of the airfoil. The platform includes a second undercut located beneath at least a portion of the trailing edge of the airfoil.

In another refinement of the application the attachment portion is dovetail shaped.

In another embodiment of the application there is a compressor blade for a gas turbine engine. The blade includes a stalk. A lower portion of the stalk defines an attachment section. An upper portion of the stalk defines a platform. The platform has an upper surface and a lower surface. The upper and lower surfaces extend between a first outside edge and a second outside edge. An airfoil is attached to the upper surface of the platform. The airfoil has a leading edge positioned at about the first outside edge. The airfoil also has a trailing edge positioned at about the second outside edge. The blade further includes an undercut in the lower surface of the platform. At least a portion of the undercut is positioned beneath at least one of the leading edge or the trailing edge of the airfoil.

In one refinement the undercut is positioned beneath at least a portion of the trailing edge of the airfoil.

In another refinement the undercut is positioned beneath at least a portion of the leading edge of the airfoil.

In another refinement there is a second undercut in the bottom surface of the platform. The second undercut is positioned beneath at least a portion of the leading edge of the airfoil.

In another refinement the attachment section is dovetail shaped.

In another refinement the airfoil is highly swept.

In another refinement the undercut in the platform begins at the first outside edge and extends only a portion of the way toward the second outside edge.

In another embodiment of the application there is a compressor stage of a gas turbine engine. The compressor stage includes a compressor wheel having a plurality of blade retaining slots. The compressor stage further includes a plurality of compressor blades. Each blade is positioned in one of the blade retaining slots. Each compressor blade includes an airfoil having a leading edge and a trailing edge. Each compressor blade further includes a stalk defining a platform having an upper side and a lower side. The airfoil is connected to the upper side of the platform. The stalk includes an attachment portion mountable within the respective blade retaining slot. The stalk further includes means for driving the load pathway away from at least a portion of the airfoil for loads generated by rotation of the compressor wheel.

In one refinement the means for driving the load pathway away from at least a portion of the airfoil comprises at least one undercut located in the stalk.

In another refinement the undercut is positioned beneath at least a portion of the trailing edge of the airfoil.

In another refinement the undercut is positioned beneath at least a portion of the leading edge of the airfoil.

In another refinement the airfoil is highly swept.

In another refinement there is a second undercut located beneath the platform at a leading edge of the airfoil.

One aspect of the present application provides a compressor blade for a gas turbine engine, comprising an airfoil extending between a leading edge and a trailing edge and operable to affect a change in total pressure between an upstream side of the airfoil and a downstream side of the airfoil, a stalk having a lower attachment portion and an upper portion defining a platform, the platform having a first side and a second side, a portion of the first side of the platform being coupled to the airfoil, and at least one undercut in the stalk beneath a portion of the airfoil.

Another aspect of the present application provides an apparatus comprising a rotatable blade of a gas turbine engine including a stalk having a lower portion defining an attachment section and an upper portion of the stalk defining a platform, the platform having an upper surface and a lower surface an airfoil extending from the upper surface of the platform, and an undercut in the lower surface of the platform and partially extending along one side of the platform, at least a portion of the undercut positioned beneath at least one of the leading edge or the trailing edge of the airfoil.

A further aspect of the present application provides a compressor stage of a gas turbine engine, comprising a compressor wheel having a plurality of blade retaining slots, a plurality of compressor blades, each blade being positioned in one of the blade retaining slots, the plurality of compressor blades comprising an airfoil having a leading edge and a trailing edge, a stalk defining a platform having an upper side and a lower side, the airfoil being connected to the upper side of the platform, wherein the stalk includes an attachment portion mountable within the respective blade retaining slot, the stalk further including means for driving the load pathway away from at least a portion of the airfoil.