Patent ID: 12188371

DETAILED DESCRIPTION

FIG.1schematically illustrates a gas turbine engine20. The gas turbine engine20is disclosed herein as a two-spool turbofan that generally incorporates a fan section22, a compressor section24, a combustor section26and a turbine section28. Alternative engines might include other systems or features. The fan section22drives air along a bypass flow path B in a bypass duct, while the compressor section24drives air along a core flow path C for compression and communication into the combustor section26then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

The exemplary engine20generally includes a low speed spool30and a high speed spool32mounted for rotation about an engine central longitudinal axis A relative to an engine static structure36via several bearing systems38. It should be understood that various bearing systems38at various locations may alternatively or additionally be provided, and the location of bearing systems38may be varied as appropriate to the application.

The low speed spool30generally includes an inner shaft40that interconnects a fan42, a low pressure compressor44and a low pressure turbine46. The inner shaft40is connected to the fan42through a speed change mechanism, which in exemplary gas turbine engine20is illustrated as a geared architecture48to drive the fan42at a lower speed than the low speed spool30. The high speed spool32includes an outer shaft50that interconnects a high pressure compressor52and high pressure turbine54. A combustor56is arranged in exemplary gas turbine20between the high pressure compressor52and the high pressure turbine54. An engine static structure36is arranged generally between the high pressure turbine54and the low pressure turbine46. The engine static structure36further supports bearing systems38in the turbine section28. The inner shaft40and the outer shaft50are concentric and rotate via bearing systems38about the engine central longitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor44then the high pressure compressor52, mixed and burned with fuel in the combustor56, then expanded over the high pressure turbine54and low pressure turbine46. The turbines46,54rotationally drive the respective low speed spool30and high speed spool32in response to the expansion. It will be appreciated that each of the positions of the fan section22, compressor section24, combustor section26, turbine section28, and fan drive gear system48may be varied. For example, gear system48may be located aft of combustor section26or even aft of turbine section28, and fan section22may be positioned forward or aft of the location of gear system48.

In typical gas turbine engines, such as the gas turbine engine20ofFIG.1, the compressor can be regarded as a low-temperature region and includes integrated blade rotors (IBRs) whereas the turbine can be regarded as a high-temperature region and includes rotor discs in which blades can be disconnected from a disc. The IBRs of the compressor are rotational features characterized in that blades are integrally formed with a rotor element. The use of IBRs in turbines or other high-temperature regions as a replacement for rotor discs may improve manufacturability and reliability but has been found to result in high stress levels on the blades. These high stress levels are mainly concentrated at the leading and trailing edges of the blades and are caused by centrifugal forces combined with hot gas exposure.

Therefore, a need exists for an improved IBR design for a turbine-compatible IBR that does not result in high stress levels on the leading and trailing edges of the blades.

Thus, as will be discussed below, an IBR is provided for use in a high-temperature region of a turbine, such as a gas turbine engine. The IBR can be machined and includes a central portion or disc, blades integrally formed with the disc and surfaces between the blades in a circumferential direction. These surfaces are exposed to the hot gas path of the turbine and are characterized as having a cylindrical, curved and/or convex profile to minimize flow separation. The convex profile blends tangentially with cylindrical sections of the disc and, in particular, can be about ⅓ a width of the disc with reference to imaginary lines passing through leading and trailing edge radii.

With reference toFIGS.2and3, an IBR201is provided and includes a disc210, blades220that are integrally formed with the disc210and radially outwardly facing surfaces230of the disc210. The disc210has a generally annular shape, opposite axial sides211,212and a width W in the axial dimension D between the axial sides211,212. The radially outwardly facing surfaces230are provided at a periphery213of the disc210. The blades220are arranged in a circumferential dimension C about the disc210and extend radially outwardly in a radial dimension R. Each blade220can have an airfoil shape with a leading edge221, a trailing edge222opposite the leading edge221, a pressure surface223extending from the leading edge221to the trailing edge222, a suction surface224, which is opposite the pressure surface223and which extends from the leading edge221to the trailing edge222, and a blade tip225. Each radially outwardly facing surface230is disposed adjacent to a pressure surface223of a corresponding blade220and extends in the circumferential dimension C to a suction surface224of a neighboring blade220.

The IBR201can be formed from an initial block of material, such as metallic material or polymeric material for example, which is forged or machined.

Each radially outwardly facing surface230includes leading and trailing wing sections231,232. The leading and trailing wing sections231,232have shared upper surfaces233and shared lower surfaces234. The upper surfaces233of the leading and trailing wing sections231,232cooperatively define a cylindrical plane CP about the periphery213of the disc210. The corresponding blade220for each radially outwardly facing surface230extends radially outwardly from this cylindrical plane CP. The lower surfaces234of the leading and trailing wing sections231,232curvilinearly taper toward the opposite axial sides211,212of the disc210. The leading and trailing wing sections231,232extend axially beyond the leading and trailing edges221,222of the corresponding blade220.

Each radially outwardly facing surface230further includes a primary curved profile235(seeFIG.3) and a secondary curved profile236(seeFIG.2). The primary curved profile235protrudes radially outwardly from the cylindrical plane CP along a chord length L of the corresponding blade220. In accordance with embodiments, a maximum height H that the curved profile235protrudes from the cylindrical plane CP is about ⅓ of the width W of the disc210. The secondary curved profile236is a concave profile that extends in the circumferential dimension C between the pressure surface223of the corresponding blade220and the suction surface224of the neighboring blade220.

With continued reference toFIG.3and with additional reference toFIGS.4A and4B, each radially outwardly facing surface230has a leading edge portion237, a trailing edge portion238and a central portion239which is axially interposed between the leading edge portion237and the trailing edge portion238(the leading edge portion237is shown inFIG.4Aand the trailing edge portion238is shown inFIG.4B). The leading edge portion237, the trailing edge portion238and the central portion239cooperatively form the primary curved profile235.

As shown inFIG.4A, the leading edge portion237corresponds to the lead edge221of the corresponding blade220and blends tangentially with a leading edge fillet2210at a base of the leading edge221of the corresponding blade220. That is, at the base of the leading edge221, the leading edge fillet2210has a curvature401with a uniform or changing radius of curvature from the upper surface233of the leading wing section231and the leading edge portion237is formed to extend tangentially from this curvature401. The trailing edge portion238blends tangentially with a trailing edge fillet2220at a base of the trailing edge222of the corresponding blade220. That is, at the base of the trailing edge222, the trailing edge fillet2220has a curvature402with a uniform or changing radius of curvature from the upper surface233of the trailing wing section232and the leading edge portion238is formed to extend tangentially from this curvature402.

With increasing axial distance from the leading edge221of the corresponding blade220, a curvature of the leading edge portion237(which is initially similar to the curvature401of the leading edge fillet2210allowing for the tangential blending) increases and then reverses direction whereupon the leading edge portion237connects with the central portion239. With increasing axial distance from the trailing edge222of the corresponding blade220, a curvature of the trailing edge portion238(which is initially similar to the curvature402of the trailing edge fillet2220allowing for the tangential blending) increases and then reverses direction whereupon the trailing edge portion238connects with the central portion239.

With continued reference toFIGS.2,3,4A and4B, with reference back toFIG.1and with additional reference toFIG.5, the IBR201can be provided, for example, in the turbine section28of the gas turbine engine20. As shown inFIG.5, the blades220are positioned to aerodynamically interact with the high-temperature fluid flowing through the turbine section28and each radially outwardly facing surface220is thus exposed to a hot gas path.

Technical effects and benefits of the present disclosure are the provision of an IBR for use with a high-temperature region of a turbine. The surfaces of the IBR between the blades, which are exposed to the hot gas path of the turbine, are characterized as having a cylindrical, curved and/or convex profile to minimize flow separation. This leads to eliminations or reductions of high stress levels on the leading and trailing edges of the blades.

The corresponding structures, materials, acts, and equivalents of all means or step plus function elements in the claims below are intended to include any structure, material, or act for performing the function in combination with other claimed elements as specifically claimed. The description of the present disclosure has been presented for purposes of illustration and description, but is not intended to be exhaustive or limited to the technical concepts in the form disclosed. Many modifications and variations will be apparent to those of ordinary skill in the art without departing from the scope and spirit of the disclosure. The embodiments were chosen and described in order to best explain the principles of the disclosure and the practical application, and to enable others of ordinary skill in the art to understand the disclosure for various embodiments with various modifications as are suited to the particular use contemplated.

While the preferred embodiments to the disclosure have been described, it will be understood that those skilled in the art, both now and in the future, may make various improvements and enhancements which fall within the scope of the claims which follow. These claims should be construed to maintain the proper protection for the disclosure first described.