Patent Application
20250198294
Exemplary embodiments of the present disclosure relate generally to gas turbine engines and, in one embodiment, to a fan blade leading edge sheath for a gas turbine engine with a reduced wing thickness.
In a gas turbine engine, air is compressed in a compressor and compressor air is then mixed with fuel and combusted in a combustor to produce a high-temperature and high-pressure working fluid. This working fluid is directed into a turbine in which the working fluid is expanded to generate power. The generated power drives the rotation of a rotor within the turbine through aerodynamic interactions between the working fluid and turbine blades or airfoils. The rotor can be used to drive rotations of a propeller or fan or to produce electricity in a generator.
The air that is compressed in the compressor can be drawn into an inlet of the compressor by the propeller or fan. The propeller or fan includes multiple fan blades, each of which includes a leading edge. Typically that leading edge is protected by a leading edge sheath. This leading edge sheath often exhibits issues that negatively impact its usefulness.
Accordingly, a need exists for an improved leading edge sheath for a fan blade of a gas turbine engine.
According to an aspect of the disclosure, a sheath is provided. The sheath includes a leading edge portion having pressure and suction sides for respective association with pressure and suction sides of an airfoil and first and second wings respectively extending from the pressure and suction sides of the leading edge portion. The first wing includes a first elongate portion and a first trailing edge disposed at an end of and thinned relative to the first elongate portion. The second wing includes a second elongate portion and a second trailing edge disposed at an end of and thinned relative to the second elongate portion.
In accordance with additional or alternative embodiments, an aft edge of the leading edge portion and respective interior surfaces of the first and second wings form a cavity for receiving a leading edge of the airfoil.
In accordance with additional or alternative embodiments, the sheath further includes sheath adhesive by which the sheath is attachable to a leading edge of the airfoil.
In accordance with additional or alternative embodiments, the first and second wings have different chordal lengths.
In accordance with additional or alternative embodiments, respective profiles of each of the first and second trailing edges are curvilinear.
In accordance with additional or alternative embodiments, each of the first and second trailing edges include multiple sections of various thinning slopes.
In accordance with additional or alternative embodiments, each of the first and second trailing edges include surface features to grip an overlying erosion coating.
According to an aspect of the disclosure, an airfoil assembly is provided and includes an airfoil having a leading edge and pressure and suction sides extending from the leading edge and a sheath affixed to the leading edge of the airfoil. The sheath includes a leading edge portion having pressure and suction sides respectively associated with the pressure and suction sides of the airfoil, first and second wings respectively extending from the pressure and suction sides of the leading edge portion and respectively comprising a locally-thinned trailing edge and an erosion coating applied to the pressure and suction sides of the airfoil to overlap with the locally-thinned trailing edge of each of the first and second wings.
In accordance with additional or alternative embodiments, an aft edge of the leading edge portion and respective interior surfaces of the first and second wings form a cavity in which the leading edge of the airfoil is received.
In accordance with additional or alternative embodiments, the airfoil includes metallic materials and the airfoil assembly further includes sheath adhesive to affix the sheath to the leading edge of the airfoil and primer interposed between the erosion coating and the airfoil.
In accordance with additional or alternative embodiments, the erosion coating includes polyurethane.
In accordance with additional or alternative embodiments, the first and second wings have different chordal lengths.
In accordance with additional or alternative embodiments, each of the first and second wings includes an elongate portion and the locally-thinned trailing edge of each of the first and second wings is disposed at an end of and is thinned relative to the corresponding elongate portion.
In accordance with additional or alternative embodiments, respective profiles of the locally-thinned trailing edge of each of the first and second wings are curvilinear.
In accordance with additional or alternative embodiments, the locally-thinned trailing edge of each of the first and second wings includes multiple sections of various thinning slopes.
In accordance with additional or alternative embodiments, the locally-thinned trailing edge of each of the first and second wings includes surface features to grip onto the erosion coating.
According to an aspect of the disclosure, an airfoil assembly method for use with an airfoil having a leading edge and pressure and suction sides extending from the leading edge is provided. The airfoil assembly method includes forming a sheath to include a leading edge portion having pressure and suction sides respectively associated with the pressure and suction sides of the airfoil and first and second wings respectively extending from the pressure and suction sides of the leading edge portion and respectively comprising a locally-thinned trailing edge. The airfoil assembly method further includes adhering the sheath to the leading edge of the airfoil and applying an erosion coating to the pressure and suction sides of the airfoil to overlap with the locally-thinned trailing edge of each of the first and second wings.
In accordance with additional or alternative embodiments, the forming of the sheath includes curvilinearly thinning the locally-thinned trailing edge of each of the first and second wings.
In accordance with additional or alternative embodiments, the applying of the erosion coating is executed such that the erosion coating at least initially overlaps with respective entireties of the locally-thinned trailing edge of each of the first and second wings.
In accordance with additional or alternative embodiments, the airfoil assembly method further includes priming the pressure and suction sides of the airfoil prior to the applying of the erosion coating.
These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings.
The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike:
These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings.
A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the Figures.
The exemplary gas turbine engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged in the gas turbine engine 20 between the high pressure compressor 52 and the high pressure turbine 54. The engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The engine static structure 36 further supports the bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow is compressed by the low pressure compressor 44 and then the high pressure compressor 52, is mixed and burned with fuel in the combustor 56 and is then expanded over the high pressure turbine 54 and the low pressure turbine 46. The high and low pressure turbines 54 and 46 rotationally drive the low speed spool 30 and the high speed spool 32, respectively, in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, geared architecture 48 may be located aft of the combustor section 26 or even aft of the turbine section 28, and the fan section 22 may be positioned forward or aft of the location of geared architecture 48.
With continued reference to
In particular, for certain fan blade assemblies, such as hybrid aluminum fan blade assemblies, the leading edge sheath 202 is provided as a titanium sheath that protects the leading edge of the airfoil section 221. In these or other cases, the leading edge sheath 202 has a U-shaped cross-section that is applied to the leading edge while a protective erosion coating is applied to an aft portion of the leading edge sheath 202 in a manner that leads to the protective erosion coating abutting the airfoil section 221 (i.e., the titanium of the airfoil section 221).
With the construction described above, during operation of the fan 24, precipitate static charge tends to accumulate on each of the multiple fan blades 220 from particles in the air. In addition to offering erosion and corrosion protection, the protective erosion coating effectively dissipates this static charge build-up through titanium components of the airfoil section 221. It has been found, however, that the erosion coating can separate from the trailing edge of the leading edge sheath The separation can cause exposure of underlying primer and possibly progress to a point at which the erosion coating delaminates from the blade. This opens a path for water/electrolyte ingress which increases the risk of corrosion (i.e., galvanic corrosion).
Accordingly, a need exists for an improved leading edge sheath for a fan blade of a gas turbine engine.
Therefore, as will be described below, a leading edge sheath is provided for use with a leading edge of an airfoil section of a fan blade of a gas turbine engine. The leading edge sheath has wings that have reduced thicknesses at trailing edges of the leading edge sheath. An erosion coating is applied to the airfoil section and overlaps with the thinned portions of the leading edge sheath.
With continued reference to
The airfoil 303 includes the leading edge 302, a trailing edge 304, a pressure side 305 extending from the leading edge 302 to the trailing edge 304 and a suction side 306 extending from the leading edge 302 to the trailing edge 304. The leading edge sheath 301 can be affixed to the leading edge 302 of the airfoil 303 and includes a leading edge portion 310 having a pressure side 311 for association with the pressure side 305 of the airfoil 303 and a suction side 312 for association with the suction side 306 of the airfoil 303. The leading edge sheath 301 further includes a first wing 320 and a second wing 330. The first wing 320 extends aft from the pressure side 311 of the leading edge portion 310 and the second wing extends aft from the suction side 312 of the leading edge portion 310. The first wing 320 includes a first elongate portion 321 and a first locally-thinned trailing edge 322. The first locally-thinned trailing edge 322 is disposed at an aft end of the first elongate portion 321. The first elongate portion 321 has a thickness T1 and the first locally-thinned trailing edge 322 has a thickness T2, which is less than T1, so that the first locally-thinned trailing edge 322 is thinned relative to the first elongate portion 321. The second wing 330 includes a second elongate portion 331 and a second locally-thinned trailing edge 332. The second locally-thinned trailing edge 332 is disposed at an aft end of the second elongate portion 331. The second elongate portion 331 has a thickness T3 and the second locally-thinned trailing edge 332 has a thickness T4, which is less than T3, so that the second locally-thinned trailing edge 332 is thinned relative to the second elongate portion 331.
As shown in
With reference to
With continued reference to
In accordance with embodiments, the airfoil 520 includes metallic materials, such as aluminum and/or titanium. In these or other cases, the airfoil assembly 510 further includes sheath adhesive 531 to affix the leading edge sheath 530 to the leading edge of the airfoil 520 and primer 541 interposed between the erosion coating 540 and the airfoil 520.
With the first and second wings including the locally-thinned trailing edge 550 and the erosion coating 540 overlapped with the locally-thinned trailing edge 550, an incidence of erosion coating 540 separation from the leading edge sheath 530 is avoided. In particular, in a case in which the erosion coating 540 is initially provided to overlap with an entirety of the locally-thinned trailing edge 550 as shown in
With reference to
With reference to
In accordance with embodiments, the forming of the sheath of block 801 can include curvilinearly thinning the locally-thinned trailing edge of each of the first and second wings (block 8011) and the applying of the erosion coating of block 804 is executed such that the erosion coating at least initially overlaps with respective entireties of the locally-thinned trailing edge of each of the first and second wings.
Benefits of the features described herein are the provision of a leading edge sheath for a leading edge of an airfoil section of a fan blade of a gas turbine engine with thinned sections at the trailing edges and an erosion coating that overlaps onto the thinned sections to maintain a smooth transition. This avoids the problem of current erosion coatings in that they tend to shrink and pull away from their original position which risks exposing underlying metallic materials and galvanic corrosion. With the overlapped erosion coating, if the erosion coating shrinks and pulls back, the erosion coating still overlaps with the thinned sections of the leading edge sheath and does not expose underlying metallic materials. This reduces the risk of galvanic corrosion.
The term “about” is intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application.
The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof.
While the present disclosure has been described with reference to an exemplary embodiment or embodiments, 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 present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.
