Patent Application
20140154068
The present disclosure relates generally to airfoil arrays utilized in gas turbine engines and, more particularly, to endwall contouring.
Gas turbine engines typically include a compressor section, a combustor section, and a turbine section, with an annular flow path extending axially through each. Initially, air flows through the compressor where it is compressed or pressurized. The combustor then mixes and ignites the compressed air with fuel, generating hot combustion gases. These hot combustion gases are then directed by the combustor to the turbine where power is extracted from the hot gases by causing blades of the turbine to rotate.
Some sections of the engine include airfoil arrays. Air within the engine moves through fluid flow passages in the arrays. The fluid flow passages are established by adjacent airfoils projecting from laterally extending endwalls. Near the endwalls, the fluid flow is dominated by a flow phenomenon known as a horseshoe vortex, which forms as a result of the endwall boundary layer separating from the endwall as the gas approaches the leading edges of the airfoils. The separated gas reorganizes into the horseshoe vortex. There is a high loss of efficiency associated with the vortex, and this loss is referred to as “secondary” or endwall loss. Accordingly, there exists a need for a way to mitigate or reduce endwall losses.
According to one embodiment of the present disclosure, an airfoil array is disclosed. The airfoil array may comprise an endwall, and a plurality of airfoils radially projecting from the endwall. Each airfoil may have a first side and an opposite second side extending axially in chord between a leading edge and a trailing edge. The airfoils may be circumferentially spaced apart on the endwall thereby defining a plurality of flow passages between adjacent airfoils. The airfoil array may further comprise a convex profiled region extending from the endwall adjacent to the first side of at least one of said plurality of airfoils near the leading edge of the at least one of said plurality of airfoils, and a concave profiled region in the endwall and extending across said at least one of said plurality of flow passages.
In a refinement, a local maximum in radial extent of the convex profiled region may be positioned between the leading edge and mid-chord of the airfoil.
In another refinement, each airfoil may have an axial chord, and a local maximum in radial extent of the convex profiled region may be disposed between about 0% to about 50% of the axial chord.
In another refinement, each flow passage may have a passage width, and a local maximum in radial extent of the convex profiled region may be disposed between about 0% to about 50% of the passage width.
In another refinement, a local minimum in radial extent of the concave profiled region may be positioned between the leading edge and the trailing edge of the airfoil.
In yet another refinement, the concave profiled region may be elongated across the flow passage from the first side of at least one of said plurality of airfoils to the second side of an adjacent airfoil.
In a related refinement, the concave profiled region may be adjacent to the first side of the airfoil near mid-chord.
In another related refinement, the concave profiled region may be adjacent to the second side of the adjacent airfoil along a significant length of the second side.
In another related refinement, the concave profiled region adjacent to the second side of the adjacent airfoil may extend along a majority of the second side of the airfoil.
In yet another related refinement, an extent of the concave profiled region adjacent the first side of the airfoil may be less than an extent of the concave profiled region adjacent the second side of the adjacent airfoil.
In another refinement, each airfoil may have an axial chord and each flow passage may have a passage width. A local minimum in radial extent of the concave profiled region may be disposed between about 0% to about 100% of the axial chord and between about 0% to about 100% of the passage width.
In yet another refinement, the first side may be a pressure side of an airfoil, and the opposite second side may be a suction side of an airfoil.
According to another embodiment, a gas turbine engine is disclosed. The gas turbine engine may comprise a compressor section, a combustor section downstream of the compressor section, and a turbine section downstream of the combustor section. One of the compressor section and the turbine section may have at least one airfoil array including a plurality of airfoils circumferentially spaced apart and projecting radially from an endwall, the airfoils establishing a plurality of flow passages. Each airfoil may have a first side, an opposite second side, a leading edge, and a trailing edge. The endwall may have a convex profiled surface near the leading edge of at least one of said plurality of airfoils, and a concave profiled surface elongated across at least one of said plurality of flow passages from the first side of the at least one of said plurality of airfoils to the second side of an adjacent airfoil.
In a refinement, the convex profiled surface may be adjacent to the first side of each airfoil.
In another refinement, each airfoil may have an axial chord, and the concave profiled surface adjacent to the first side of each airfoil may extend between about 30% to about 80% of the axial chord.
In another refinement, each airfoil may have an axial chord, and the concave profiled surface adjacent to the second side of the adjacent airfoil may extend between about 0% to about 100% of the axial chord.
In another refinement, each airfoil may have an axial chord, and a local minimum in radial extent of the concave profiled surface adjacent to the second side of each airfoil may be positioned between about 0% to about 100% of the axial chord.
In another refinement, each flow passage may have a passage width, and a local minimum in radial extent of the concave profiled surface may be disposed between about 30% to about 80% of the passage width.
In another refinement, a local maximum in radial extent of the convex profiled surface may be disposed between about 0% to about 50% of the passage width, and a local minimum in radial extent of the concave profiled surface may be disposed between about 50% to about 100% of the passage width.
In yet another refinement, the first side may be a pressure side of an airfoil and the opposite second side may be a suction side of an airfoil.
These and other aspects and features of the disclosure will become more readily apparent upon reading the following detailed description when taken in conjunction with the accompanying drawings.
While the present disclosure is susceptible to various modifications and alternative constructions, certain illustrative embodiments thereof, will be shown and described below in detail. It should be understood, however, that there is no intention to be limited to the specific embodiments disclosed, but on the contrary, the intention is to cover all modifications, alternative constructions, and equivalents along within the spirit and scope of the present disclosure.
Referring now to the drawings, and with specific reference to
The turbine section 16 or the compressor section 12 may include at least one airfoil array 20. As shown best in
The endwall 24 may have a plurality of convex profiled regions 36 and a plurality of concave profiled regions 38 configured to direct flow through each of the flow passages 26 within the airfoil array 20. Illustrated in
Each airfoil 22 may have a chord 42, which is defined as a line from the leading edge 32 to the trailing edge 34, and an axial chord 44, which is a projection of the chord 42 onto a plane containing the engine centerline X. Relevant distances may be expressed as a percentage of the length of the axial chord 44, as shown in the percentage scale at the bottom of
From the surface 40 adjacent the convex profiled region 36, the convex region 36 may gradually increase in radial height, or move radially inward toward the annular flow path 18, to a local maximum in radial extent 52. It will be understood that the convex profiled region 36 may extend further than the illustrated contour lines. The convex profiled region 34 may have the local maximum in radial extent 52 positioned between the leading edge 32 and mid-chord of the airfoil 22. The local maximum in radial extent 52 of the convex profiled region 34 may be disposed within an inclusive axial range of about 0% to about 50% of the axial chord and may be disposed within an inclusive lateral range of about 0% to about 50% of the passage width W.
The concave profiled region 38 may be elongated across at least one of said plurality of fluid flow passages 26. Relative to a surface 40 adjacent the concave profiled region 38, the concave region 38 may extend radially outward away from the annular flow path 18. From the surface 40 adjacent the concave profiled region 38, the concave region 38 may gradually decrease in radial height, or move outward radially away from the annular flow path 18, to a local minimum in radial extent 54. The concave region 38 of the endwall 24 may extend across the fluid flow passage 26 from the first side 28 of each airfoil 22 to the second side 30 of the neighboring airfoil 46. It will be understood that the concave profiled region 38 may extend further than the illustrated contour lines.
The concave profiled region 38 may be adjacent to the first side 28 of the airfoil 22 near mid-chord and may be predominantly adjacent to the second side 30 of the neighboring airfoil 46 along a significant length of the second side 30, about a majority of the second side 30 of the airfoil 46. The extent of the concave region 38 adjacent the first side 28 of the airfoil 22 may be less than the extent of the concave region 38 adjacent the second side 30 of the neighboring airfoil 46. For example, adjacent the first side 28 of the airfoil 22, the concave region 38 may extend between an inclusive range of about 30% to about 80% of the axial chord. Adjacent the second side 30 of the neighboring airfoil 46, the concave region 38 may extend between an inclusive range of about 0% to about 100% of the axial chord. The local minimum in radial extent 54 of the concave profiled region 38 may be positioned between the leading edge 32 and the trailing edge 34 of the airfoils 22, 36, or within the inclusive range of about 0% to about 100% of the axial chord. The local minimum in radial extent 54 may also be disposed within an inclusive lateral range of about 0% to about 100% of the passage width, for example, between about 30% to about 80% of the passage width, or for example, between about 50% to about 100% of the passage width.
It will be understood that the endwall 24 contouring described herein may be applied to any type of airfoil array 20 without departing from the scope of the invention. According to an exemplary embodiment, the contoured endwall 24 may be applied to an airfoil array 20 with airfoils 22 having a camber or turning airfoil, as shown in
From the foregoing, it can be seen that the teachings of this disclosure can find industrial application in any number of different situations, including but not limited to, gas turbine engines. Such engines may be used, for example, on aircraft for generating thrust, or in land, marine, or aircraft applications for generating power.
The disclosure described provides a way to mitigate or reduce endwall losses in an airfoil array. By positioning a convex profiled region of the endwall adjacent the first side of the airfoils near the leading edge and a concave profiled region of the endwall elongated across the flow passages, the present invention influences flow through the flow passages, thereby reducing endwall losses due to the horseshoe vortex. Furthermore, the contoured endwall described herein results in an improved aerodynamic performance of the airfoil arrays. Such contouring may minimize aerodynamic losses through blade or vane passages of a gas turbine engine, for example, those of a second stage high pressure turbine vane. In so doing, this may decrease heat, friction and pressure losses, while improving engine efficiency and life of the blade or vane.
While the foregoing detailed description has been given and provided with respect to certain specific embodiments, it is to be understood that the scope of the disclosure should not be limited to such embodiments, but that the same are provided simply for enablement and best mode purposes. The breadth and spirit of the present disclosure is broader than the embodiments specifically disclosed and encompassed within the claims appended hereto.
The present disclosure claims the benefit of priority from U.S. Provisional Patent Application Ser. No. 61/706,983, filed on Sep. 28, 2012, which is incorporated herein by reference.
| Number | Date | Country | |
|---|---|---|---|
| 61706983 | Sep 2012 | US |
