Patent Application
20220316353
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-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section.
The compressor or turbine sections may include vanes mounted on vane platforms. Seals may be arranged at leading and trailing edges of such components to reduce cooling flow leakage.
In one exemplary embodiment, a flow path component assembly includes a support structure that is a unitary component having an axial portion and a radial portion and an outer portion arranged between the support structure and a flow path. The outer portion defines a gap between the outer portion and the support structure axial portion. A flange extends into the gap.
In another embodiment according to any of the previous embodiments, the flange extends from the outer portion towards the support structure.
In another embodiment according to any of the previous embodiments, the flange extends from the support structure axial portion towards the outer portion.
In another embodiment according to any of the previous embodiments, a seal is arranged in the gap.
In another embodiment according to any of the previous embodiments, the flange is arranged between the seal and the radial portion.
In another embodiment according to any of the previous embodiments, the seal is arranged between the flange and the radial portion.
In another embodiment according to any of the previous embodiments, the flange is formed in the outer portion and a second flange is formed in the support structure.
In another embodiment according to any of the previous embodiments, the flange is near a leading edge.
In another embodiment according to any of the previous embodiments, the flange is near a trailing edge.
In another embodiment according to any of the previous embodiments, the platform is a vane platform.
In another embodiment according to any of the previous embodiments, the support structure extends between an inner vane platform and an outer vane platform.
In another embodiment according to any of the previous embodiments, the outer portion is formed from a ceramic material.
In another embodiment according to any of the previous embodiments, the support structure is formed from a metallic material.
In another exemplary embodiment, a turbine section for a gas turbine engine includes a plurality of vanes arranged circumferentially about an engine axis. Each vane includes a support structure that is a unitary component having an axial portion and a radial portion, and an outer portion arranged between the support structure and a flow path. The outer portion is formed from a ceramic material. The outer portion defines a gap between the outer portion and the support structure axial portion. A flange extending radially from one of the axial portion and the outer portion. The flange is configured to at least partially block the gap.
In another embodiment according to any of the previous embodiments, a seal is arranged in the gap.
In another embodiment according to any of the previous embodiments, the flange extends from the outer portion towards the support structure.
In another embodiment according to any of the previous embodiments, the flange extends from the support structure towards the outer portion.
In another embodiment according to any of the previous embodiments, the outer portion has an outer axial portion that forms an outer platform of the vane.
In another embodiment according to any of the previous embodiments, the flange is arranged near a leading edge and extends forward of the axial portion and a second flange is arranged near a trailing edge and is forward of an aft end of the axial portion.
In another embodiment according to any of the previous embodiments, the flange is formed in the outer portion and a second flange is formed in the support structure.
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 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 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 over 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), 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 and the low pressure turbine 46 has a pressure ratio that is greater than about five. 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 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. “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. “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 ft/second (350.5 meters/second).
A turbine blade 102 has a radially outer tip 103 that is spaced from a blade outer air seal assembly 104 with a blade outer air seal (“BOAS”) 106. The BOAS 106 may be mounted to an engine case or structure, such as engine static structure 36 via a control ring or support structure 110 and a carrier 112. The engine structure 36 may extend for a full 360° about the engine axis A.
The turbine vane assembly 97 generally comprises a plurality of vane segments 118. In this example, each of the vane segments 118 has an airfoil 116 extending between an inner vane platform 120 and an outer vane platform 122.
Although a vane platform 122 is described, this disclosure may apply to other components, and particularly flow path components. For example, this disclosure may apply to combustor liner panels, shrouds, transition ducts, exhaust nozzle liners, blade outer air seals, or other CMC components. In the illustrated example, the example vane segment 118 is a singlet, meaning the vane segment 118 includes only one airfoil section 116. This disclosure is not limited to singlets, it may be doublets, triplets etc., however. Further, while the example vane segment 118 is in the high pressure turbine section 54, one would understand that this disclosure can be used in other sections of the engine 20 such as the mid-turbine frame 57. Further, although the outer vane platform 122 is generally shown and referenced, this disclosure may apply to the inner vane platform 120.
The outer portion 130 forms the airfoil 116 and inner and outer platforms 120, 122. The outer portion 130 is arranged within the core flowpath C. That is, the outer portion 130 is a portion of the segment 118 that is exposed to the core flowpath C. The outer portion 130 may be formed of a ceramic matrix composite (“CMC”) material. The outer portion 130 may formed of a plurality of CMC laminate sheets. The laminate sheets may be silicon carbide fibers, formed into a braided or woven fabric in each layer. In other examples, the outer portion 130 may be made of a monolithic ceramic. CMC components such as outer portion 130 are formed by laying fiber material, such as laminate sheets or braids, in tooling, injecting a gaseous infiltrant into the tooling, and reacting to form a solid composite component. The component may be further processed by adding additional material to coat the laminate sheets. CMC components may have higher operating temperatures than components formed from other materials.
The outer portion 130 may be spaced from the support structure 132 to form a gap 144 between the outer portion 130 and the support structure 132. The gap 144 may be used for directing cooling flow, for example. In the illustrated example, cooling flow F flows through the axially extending portion 136 of the support 132 to cool the vane assembly 118. In this example, a first seal 150 and a second seal 152 are arranged between the outer portion 130 and the support structure 132. The first seal 150 is arranged near the leading edge 124 and the second seal 152 is arranged near the trailing edge 126. The seals 150, 152 may be L-seals, for example. In this example, the first seal 150 prevents hot gases from the core flowpath C from entering the gap 144, and the second seal 152 prevents cooling flow F from leaking from the vane assembly 118.
The outer platform 122 of the outer portion 130 has a first flange 160 and a second flange 162. The first and second flanges 160, 162 extend radially into the gap 144 between the outer portion 130 and the support structure 132. In this example, the first and second flanges 160, 162 extend radially outward from the outer portion 130 towards the axially extending portion 136 of the support structure 132. A clearance 166 in the axial direction is formed between the flanges 160, 162 and the axially extending portion 136. The clearance 166 may be between about 0.005 and about 0.020 inches (0.127 to 0.508 mm), for example. The flanges 160, 162 partially block the gap 144 and help prevent cooling flow leakage and/or hot gas ingestion by creating a torturous path between the support structure 132 and the outer portion 130. Although a particular arrangement with flanges 160, 162 is shown, the flanges may have a different arrangement, as shown and described below.
The disclosed flange arrangements provide redundant protection against cooling flow leakage or hot gas ingestion. The outer portion 132 may be formed from a ceramic material, which has much higher temperature capabilities than the metallic support structure 132. Thus, cooling flow leakage and/or hot gas ingestion may create unwanted thermal gradients or prematurely wear components. The flanges may help to prevent leakage of cooling air or ingestion of hot gases into the gap 144 between the support structure 132 and the outer portion 130 in the event the seals fail. The flanges form a labyrinth flow discourager by obstructing the cooling air from leaking out by turning the flow and introducing sharp edges and corners. These sharp edges and corners slow the flow of leakage, reducing the amount of cooling air that is leaked through the gap.
In this disclosure, “generally axially” means a direction having a vector component in the axial direction that is greater than a vector component in the circumferential direction, “generally radially” means a direction having a vector component in the radial direction that is greater than a vector component in the axial direction and “generally circumferentially” means a direction having a vector component in the circumferential direction that is greater than a vector component in the axial direction.
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.
Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the true scope and content of this disclosure.
