Patent Application
20250067187
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 gas turbine engine according to an example of the present disclosure includes a ceramic matrix composite (CMC) vane arc segment that includes first and second platforms and an airfoil section that extends radially therebetween. The airfoil section has an internal through-cavity. The first and second platforms have, respectively, platform inlet and outlet ports connected to the internal through-cavity for conveying cooling air. First and second metallic vane supports are radially mounted between the first and second metallic vane supports. The second metallic vane support has a plenum and a plenum inlet port connected with the platform outlet port for receiving the cooling air from the internal through-cavity. A seal is located radially between the second platform and the second metallic vane support. The seal circumscribes the platform outlet port to limit leakage of the cooling air between the second platform and the second metallic vane support.
In a further embodiment of any of the foregoing embodiments, the seal is a rope seal.
In a further embodiment of any of the foregoing embodiments, the rope seal includes ends that are secured together such that the rope seal is a continuous ring.
In a further embodiment of any of the foregoing embodiments, the seal is a tadpole seal having a bulb section and a tail section extending from the bulb section.
A further embodiment of any of the foregoing embodiments includes at least one fastener securing the tail section to the second metallic vane support.
In a further embodiment of any of the foregoing embodiments, the at least one fastener comprises a pin that extends through a hole in the tail section.
In a further embodiment of any of the foregoing embodiments, the at least one fastener comprises a retainer clip clamping onto the tail section.
In a further embodiment of any of the foregoing embodiments, the second metallic support is a tangential onboard injector (TOBI).
In a further embodiment of any of the foregoing embodiments, the TOBI includes fore and aft annular walls and an outer diameter annular wall. The fore and aft annular walls and the outer diameter annular wall define the plenum therebetween. The TOBI has an axially-oriented nozzle for discharging the cooling air from the plenum in an aft direction.
In a further embodiment of any of the foregoing embodiments, the second metallic vane support includes a seal groove in which the seal is at least partially disposed.
In a further embodiment of any of the foregoing embodiments, the second metallic vane support includes first and second eyelets that each open on one end thereof to a side face of the second metallic vane support and on an opposed end thereof to the seal groove. The seal is routed through the first and second eyelets.
In a further embodiment of any of the foregoing embodiments, the vane arc segment includes a spar extending through the internal through-cavity and protruding from the second platform. The seal circumscribes the spar.
A gas turbine engine according to an example of the present disclosure includes a ceramic matrix composite (CMC) vane arc segment that has first and second platforms and an airfoil section that extends radially therebetween. The airfoil section has an internal through-cavity receiving pressurized cooling air. The first and second platforms have, respectively, platform inlet and outlet ports connected to the internal through-cavity conveying the pressurized cooling air. A tangential onboard injector (TOBI) radially supports the CMC vane arc segment. The TOBI has a plenum and a plenum inlet port connected with the platform outlet port. The plenum receives the pressurized cooling air from the internal through-cavity via the plenum inlet port. A seal is located radially between the second platform and the TOBI. The seal circumscribes the platform outlet port and limiting leakage of the pressurized cooling air between the second platform and the TOBI. The pressurized cooling air loads the vane arc segment against the seal causing the seal to compress.
In a further embodiment of any of the foregoing embodiments, the seal is selected from the group consisting of a rope seal and a tadpole seal.
In a further embodiment of any of the foregoing embodiments, the TOBI includes fore and aft annular walls and an outer diameter annular wall. The fore and aft annular walls and the outer diameter annular wall define the planum therebetween. The TOBI has an axially-oriented nozzle for discharging the cooling air from the plenum in an aft direction.
In a further embodiment of any of the foregoing embodiments, the TOBI includes a seal groove in which the seal is at least partially disposed, first and second eyelets that each open on one end thereof to a side face of the TOBI and on an opposed end thereof to the seal groove, and the seal is routed through the first and second eyelets.
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. Terms such as “first” and “second” used herein are to differentiate that there are two architecturally distinct components or features. Furthermore, the terms “first” and “second” are interchangeable in that a first component or feature could alternatively be termed as the second component or feature, and vice versa.
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)]0.5. 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).
The vane arc segments 62 are formed of a ceramic matrix composite (“CMC”). A CMC is formed of ceramic fiber tows that are disposed in a ceramic matrix. As an example, the CMC may be, but is not limited to, a SiC/SiC composite in which SiC fiber tows are disposed within a SiC matrix. The fiber tows are arranged in a fiber architecture, which refers to an ordered arrangement of the tows relative to one another, such as a braided or woven architecture.
The vane arc segments 62 are radially supported through the platform 68 by one or more first metallic supports 74. For example, the support or supports 74 may be an engine case or an intermediate structure, such as a spar or carrier, that attaches to an engine case. The vane arc segments 62 may be supported by a flange or flanges, or other features designed for CMC vane attachment. The vane arc segments 62 are supported at the inner diameter by a second metallic support 76. In the example shown, the second metallic support 76 is a tangential onboard injector (TOBI), although it is to be understood that the examples herein are also applicable at the first metallic support 74 and to supports other than TOBIs. The metallic supports 74/76 are formed of an alloy, such as but not limited to, nickel- or cobalt-based alloys.
As shown in
The vane arc segments 62 are situated on the outer diameter annular wall 80 such that the TOBI 76 radially supports each of the vane arc segments 62. In this regard, at least a portion of the surface of the outer diameter annular wall 78c is in interfacial contact with the inner diameter surface of the inner platform 70. Such interfacial contact may facilitate load distribution and secondary sealing between the TOBI 76 and the vane arc segments 62.
The outer diameter annular wall 78c includes plenum inlet ports 86 that are connected, respectively, with the platform outlet ports 72b to receive the cooling air flow from each of the vane arc segments 62 into the plenum 80. For example, the ports 72b/86 are approximately equal in area and are radially aligned with each other. The cross-sectional flow area of the ports 72b/86 may be selected to meter the cooling air flow for a desired downstream cooling effect, to provide pressures that reduce leakage, and/or to reduce pressure-driven stresses in the TOBI 76.
In general, a TOBI is a structure in a gas turbine engine at an inner diameter location of the turbine vanes that receives cooling air from the vanes and redirects the cooling air through nozzles in an axially aft direction to cool downstream components, such as a portion of a turbine disc. The cooling air that is provided to the vane arc segments 62, and subsequently to the TOBI 76, is generally at a higher pressure than the pressure in the core gaspath C (e.g., at cruise). As a result, the interfaces where the TOBI 76 and the vane arc segments 62 meet are potentially subject to leakage of cooling air into the core gaspath C. In addition to the pressure differential, a further challenge to limiting such leakage is that the CMC material of the vane arc segments 62 and the alloy of the TOBI differ in coefficient of thermal expansion, thereby causing changes in the gap sizes that provide potential leak paths over the relevant operating temperature range. Such thermal differences may also challenge retention and proper positioning of a seal in a space that could vary in size over the temperatures range.
In these regards, a seal 88 is provided in the interface radially between the outer diameter annular wall 80 and the platform 70 for limiting leakage of the cooling air. For instance, the seal 88 is a ring that circumscribes the ports 72b/86 to seal the interface. In the illustrated example, the seal 88 is a rope seal. A rope seal is formed of high-temperature resistance fibers, such as ceramic fibers. The fibers may be braided, knitted, or woven. Example ceramic fibers include, but are not limited to, oxide fibers. For instance, the ceramic fibers are NEXTEL fibers, which are composed of Al2O3, SiO2, and B2O3. Optionally, a rope seal may include a sheath surrounding a fiber core. The sheath can be an overbraid or foil that surrounds a fiber core.
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In a further example depicted in
In another example shown in
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.
