Patent Grant
12000306
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-pressure and temperature exhaust gas flow. The high-pressure and temperature exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section may include low and high pressure compressors, and the turbine section may also include low and high pressure turbines.
Airfoils in the turbine section are typically formed of a superalloy and may include thermal barrier coatings to extend temperature capability and lifetime. Ceramic matrix composite (“CMC”) materials are also being considered for airfoils. Among other attractive properties, CMCs have high temperature resistance. Despite this attribute, however, there are unique challenges to implementing CMCs in airfoils.
A vane arc segment according to an example of the present disclosure includes ceramic matrix composite (CMC) fairing that has an airfoil section that defines suction and pressure side walls, leading and trailing ends, and inner and outer radial ends. A first single-sided platform at the outer radial end projects in a first circumferential direction from the suction side wall. A second single-sided platform at the inner radial end projects in a second, opposite circumferential direction from the pressure side wall. The first single-sided platform is comprised of a fiber layer that extends from the airfoil section at the first radial end and turning into the first single-sided platform. The second single-sided platform is comprised of the fiber layer that extends from the airfoil section at the second radial end and turning into the second single-sided platform.
In a further embodiment of any of the foregoing embodiments, the first single-sided platform includes an edge portion that has a suction side contour that is complementary in shape to the suction side wall of the airfoil section.
In a further embodiment of any of the foregoing embodiments, aft of the edge portion, the first single-sided platform includes a straight portion.
In a further embodiment of any of the foregoing embodiments, the second single-sided platform includes an edge portion that has a pressure side contour that is complementary in shape to the pressure side wall of the airfoil section.
In a further embodiment of any of the foregoing embodiments, forward of the edge portion, the second single-sided platform includes a straight portion.
In a further embodiment of any of the foregoing embodiments, the fiber layer in the airfoil section has a first fiber architecture and the fiber layer in at least one of the first single-sided platform or the second single-sided platform has a second fiber architecture that is different than the first fiber architecture.
In a further embodiment of any of the foregoing embodiments, the CMC fairing is made of a CMC material that has silicon-containing ceramic fiber and a silicon-containing matrix.
A gas turbine engine according to an example of the present disclosure includes a compressor section, a combustor in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor. The turbine section has vane arc segments disposed about a central axis of the gas turbine engine. Each of the vane arc segments has a ceramic matrix composite (CMC) fairing that includes an airfoil section that defines suction and pressure side walls, leading and trailing ends, and inner and outer radial ends. A first single-sided platform at the outer radial end projects in a first circumferential direction from the suction side wall, and a second single-sided platform at the inner radial end projecting in a second, opposite circumferential direction from the pressure side wall. The first single-sided platform is comprised of a fiber layer that extends from the airfoil section at the first radial end and turning into the first single-sided platform, and the second single-sided platform is comprised of the fiber layer that extends from the airfoil section at the second radial end and turning into the second single-sided platform.
A further embodiment of any of the foregoing embodiments includes inner and outer diameter supports supporting the vane arc segments by, respectively, the first single-sided platform and the second single-sided platform.
In a further embodiment of any of the foregoing embodiments, when under aerodynamic loading, the vane arc segments transfer loads to the inner and outer diameter supports via, respectively, the first single-sided platform and the second single-sided platform, and the airfoil section is in compression.
In a further embodiment of any of the foregoing embodiments, the first single-sided platform includes a first platform edge portion that has a suction side contour that is complementary in shape to the suction side wall of the airfoil section.
In a further embodiment of any of the foregoing embodiments, the second single-sided platform includes a second platform edge portion that has a pressure side contour that is complementary in shape to the pressure side wall of the airfoil section.
In a further embodiment of any of the foregoing embodiments, aft of the first platform edge portion, the first single-sided platform includes a first platform straight portion.
In a further embodiment of any of the foregoing embodiments, forward of the second platform edge portion, the second single-sided platform includes a second platform straight portion.
In a further embodiment of any of the foregoing embodiments, the fiber layer in the airfoil section has a first fiber architecture and the fiber layer in at least one of the first single-sided platform or the second single-sided platform has a second fiber architecture that is different than the first fiber architecture.
The present disclosure may include any one or more of the individual features disclosed above and/or below alone or in any combination thereof.
The various features and advantages of the present disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows.
In this disclosure, like reference numerals designate like elements where appropriate and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding elements.
The exemplary 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 first (or low) pressure compressor 44 and a first (or 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 a 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 second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in the exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 may be arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports 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 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded through the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 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, gear system 48 may be located aft of the low pressure compressor, or aft of the combustor section 26 or even aft of turbine section 28, and fan 42 may be positioned forward or aft of the location of gear system 48.
The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), and can be less than or equal to about 18.0, or more narrowly can be less than or equal to 16.0. The geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3. The gear reduction ratio may be less than or equal to 4.0. The low pressure turbine 46 has a pressure ratio that is greater than about five. The low pressure turbine pressure ratio can be less than or equal to 13.0, or more narrowly less than or equal to 12.0. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to an inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1 and less than about 5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (′TSFC)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. The engine parameters described above and those in this paragraph are measured at this condition unless otherwise specified. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45, or more narrowly greater than or equal to 1.25. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]05. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150.0 ft/second (350.5 meters/second), and can be greater than or equal to 1000.0 ft/second (304.8 meters/second).
Vanes in a turbine section of an engine typically include an airfoil section that extends between radially inner and outer platforms that bound the core gas path. In metallic alloy vanes, the airfoil sections are substantially centered on the platforms such that the platforms have near equal overhangs on the pressure side and the suction side of the airfoil section. The side edges of the platforms serve as matefaces and are often used as a sealing interface between vanes, such as with a flat seal in a seal slot. In a ceramic matrix composite (CMC) vane, however, such matefaces and sealing configurations may cause duress from thermal gradients and interlaminar stresses that are not present in metallic vanes. Moreover, turbine vanes require constraints to inhibit motion when loaded by gas path and/or secondary flow forces. Attachment of CMC vanes in an engine and management of stresses, however, is challenging. Attachment features, such as hooks, that are typically used for metallic alloy vanes can result in inefficient loading if employed in CMCs, which may also be sensitive to stress directionality and distress conditions that differ from those of metallic vanes. Additionally, hooks, seal slots, variable thickness walls, gussets, complex-geometry investment casting cores, etc. that may be used in metallic alloy components are generally not acceptable or manufacturable with CMC materials.
As CMC vanes may be single-piece integral structures, there is also considerable difficultly in forming the ceramic fiber layers of the CMC to the desired design shape of the vane. For example, the ceramic fiber layers are first laid up to form the airfoil section. The fabric that overhangs the radial ends of the airfoil section is then draped in opposite directions so as to fan out and form the suction and pressure sides of the platforms. There can be considerable difficulty in bending the fiber plies in opposite directions without forming discontinuities from folds, kinks, or substantial shifting of fibers. To address one or more of the above concerns, the examples set forth herein below disclose CMC vane arc segments that have single-sided platforms.
Each CMC fairing 61 includes an airfoil section 62 that defines suction and pressure side walls 64/66, leading and trailing ends 68a/68b, and outer and inner radial ends 70a/70b. The airfoil section 62 is solid but alternatively may have an internal through-cavity to convey cooling air. At the outer radial end 70a the airfoil section 62 has a first single-sided platform 74 projecting from the suction side wall 64 in a circumferential direction outwardly from the airfoil section 62. At the inner radial end 70b the airfoil section 62 has a second single-sided platform 74 that projects in the opposite circumferential direction outwardly from airfoil section 62.
The platforms 72/74 are single-sided in that they each extend to only one side—the suction side or the pressure side—of the airfoil section 62. As the first single-sided platform 72 projects from the suction side wall 64 the first single-sided platform 72 is a suction single-sided platform. Likewise, as the second single-sided platform 74 projects from the pressure side wall 66, the second single-sided platform 74 is a pressure single-sided platform. At the outer radial end 70a the pressure side wall 66 is absent a platform structure, at least along the profile of the airfoil section 62. Similarly, at the inner radial end 70b the side wall 64 is absent a platform structure, at least along the profile of the airfoil section 62.
Terms such as “inner” and “outer” refer to location with respect to the central engine axis A, i.e., radially inner or radially outer. Moreover, the terminology “first” and “second” as used herein is to differentiate that there are two architecturally distinct structures. It is to be further understood that the terms “first” and “second” are interchangeable in the embodiments herein in that a first component or feature could alternatively be termed as the second component or feature, and vice versa.
The CMC fairings 61 are made of CMC material, shown in partial cutaway at 65. CMC material 65 is comprised of one or more ceramic fiber layers 65a in a ceramic matrix 65b. Example ceramic matrices are silicon-containing ceramic, such as but not limited to, a silicon carbide (SiC) matrix or a silicon nitride (Si3N4) matrix. Example ceramic reinforcement of the CMC are silicon-containing ceramic fibers, such as but not limited to, silicon carbide (SiC) fiber or silicon nitride (Si3N4) fibers. The CMC may be, but is not limited to, a SiC/SiC ceramic matrix composite in which SiC fiber layers are disposed within a SiC matrix. A fiber layer has a fiber architecture, which refers to an ordered arrangement of the fiber tows relative to one another, such as a 2D woven ply or a 3D structure. The CMC fairings 61 are one-piece structures in that the fiber layer or layers are continuous from the platform 72, through the airfoil section 62, and into the platform 74.
Each of the platforms 72/74 includes an edge portion 78 that has a contour that is complementary in shape to the airfoil section 62. The edge portion 78 of the platform 72 has a contour that is complementary in shape to the pressure side wall 66 of the airfoil section 62, i.e., the edge portion 78 fits intimately with the pressure side wall 66. The edge portion 78 of the platform 74 has a contour that is complementary in shape to the suction side wall 64 of the airfoil section 62, i.e., the edge portion 78 fits intimately with the suction side wall 74. Each platform 72/74 also has a straight portion 80. On the platform 72 the straight portion 80 is aft of the edge portion 78, and on the platform 74 the straight portion 80 is forward of the edge portion 78.
When adjacent CMC fairings 61 in the circumferential vane row are brought together, the fairings 61 nest with each other such that the contour on the platform 72 bears against the pressure side wall 66 of the next adjacent fairing 61 in the row, and the contour of the platform 74 bears against the suction side wall 64 of the next adjacent fairing 61 in the other direction in the vane row. Similarly, the straight portions 80 bear against corresponding straight portions 82 of those next fairings 61 in the vane row. The straight portions 80/82 form a platform-to-platform split line fore and aft of the airfoil sections 62.
Referring also to
As the platforms 72/74 are on opposite sides of the airfoil section 62, the line of action between the points or region where the loads are transmitted crosses the airfoil section 62 and represents a cross-corner loading state. A wheelbase, i.e., the distance between the cross-corner points or regions on the platforms 72/74 where the loads are transmitted to the static structures 63a/63b, determines the load-carrying capacity of the CMC fairings 61. In general, increasing the wheelbase (length) corresponds to an increase in load-carrying capacity.
Additionally, as the platforms 72/74 extend in opposite circumferential directions, the cross-corner points or regions on the platforms 72/74 are axially and circumferentially offset from each other. Thus, under aerodynamic loading where the net load acts in a direction from the pressure side wall to the suction side wall, the fairings 61 will tend to rotate but for the constraints by the supports 63a/63b. The resultant transmission of the loads through the fairings 61 places the airfoil section 62 in compression. CMC materials are generally strong in compression loading and weaker in tension loading, which can cause interlaminar stresses. Therefore, compression loading is a favorable loading state for a CMC article such as the CMC fairings 61.
Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments.
The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.
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