Illustrative embodiments pertain to the art of turbomachinery, and specifically to turbine rotor components.
Gas turbine engines are rotary-type combustion turbine engines built around a power core made up of a compressor, combustor and turbine, arranged in flow series with an upstream inlet and downstream exhaust. The compressor compresses air from the inlet, which is mixed with fuel in the combustor and ignited to generate hot combustion gas. The turbine extracts energy from the expanding combustion gas, and drives the compressor via a common shaft. Energy is delivered in the form of rotational energy in the shaft, reactive thrust from the exhaust, or both.
The individual compressor and turbine sections in each spool are subdivided into a number of stages, which are formed of alternating rows of rotor blade and stator vane airfoils. The airfoils are shaped to turn, accelerate and compress the working fluid flow, or to generate lift for conversion to rotational energy in the turbine. Various cooling schemes are employed to ensure part life and durability. Improved cooling schemes may improve part life and provide other advantages.
According to some embodiments, components for gas turbine engines are provided. The components include a component body having a hot side and a cold side, an attachment element extending from the cold side of the component body, a component extension extending from the component body at the attachment element to an end surface, the component extension having a cold-side first portion and a cold-side second portion, wherein the cold-side first portion is a portion of the component extension between the attachment element and the cold-side second portion, and the cold-side second portion is a portion of the component extension extending from the cold-side first portion to the end surface of the component extension, and a cold-side heat transfer augmentation feature on the cold-side second portion of the component extension, wherein the cold-side first portion defines a uniform surface.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the component body is one of a platform of a vane of a gas turbine engine or a blade outer air seal of a gas turbine engine.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the cold-side second portion encompasses between 20% and 80% of the component extension.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the cold-side heat transfer augmentation feature comprises at least one of a pedestal, a trip strip, a chevron trip strip, a spherical bump, a dimple, a bump, and a protrusion.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the cold-side heat transfer augmentation feature comprises a base and a tip.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the component extension is an inner diameter platform aft lip.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the attachment element is a first attachment element and the component comprises a second attachment element and wherein the component extension is a first component extension extending from the component body at the first attachment element, the component further comprising a second component extension extending from the component body at the second attachment element to a second end surface, wherein the second component extension includes a respective cold-side first portion and a respective cold-side second portion, wherein the cold-side second portion of the second component extension includes a cold-side heat transfer augmentation feature.
According to some embodiments, gas turbine engines are provided. The gas turbine engines include a component having a component body having a hot side and a cold side, an attachment element extending from the cold side of the component body, a component extension extending from the component body at the attachment element to an end surface, the component extension having a cold-side first portion and a cold-side second portion, wherein the cold-side first portion is a portion of the component extension between the attachment element and the cold-side second portion, and the cold-side second portion is a portion of the component extension extending from the cold-side first portion to the end surface of the component extension, and a cold-side heat transfer augmentation feature on the cold-side second portion of the component extension, wherein the cold-side first portion defines a uniform surface.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include a turbine section, wherein the component is part of the turbine section and the turbine section defines a hot gaspath therethrough.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include that the component body is a platform of a vane of the gas turbine engine, and wherein the hot side is exposed to the hot gaspath.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include a blade platform located aft of the component, wherein a gap is formed between an end of the blade platform and the component extension.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include that the component is a blade outer air seal of the turbine section.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include that the component is a platform of a blade of the turbine section.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include that the cold-side heat transfer augmentation feature comprises at least one of a pedestal, a trip strip, a chevron trip strip, a spherical bump, a bump, a dimple and a protrusion.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include that the cold-side heat transfer augmentation feature comprises a base and a tip.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include a gas turbine engine part having a part end, wherein the component extension extends adjacent the part end, and wherein the tip defines a portion of the cold-side heat transfer augmentation feature closest to the part end and the base defines a portion of the cold-side heat transfer augmentation feature farthest from the part end.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include that the spacing between the tip and the part end is a minimum gap distance.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include that the component extension is an inner diameter platform aft lip of a vane of the gas turbine engine.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include that the attachment element is a first attachment element and the component comprises a second attachment element and wherein the component extension is a first component extension extending from the component body at the first attachment element, the component further comprising a second component extension extending from the component body at the second attachment element to a second end surface, wherein the second component extension includes a respective cold-side first portion and a respective cold-side second portion, wherein the cold-side second portion of the second component extension includes a cold-side heat transfer augmentation feature.
In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include that the cold-side second portion encompasses between 20% and 80% of the component extension.
The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, the following description and drawings are intended to be illustrative and explanatory in nature and non-limiting.
The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike: The subject matter is particularly pointed out and distinctly claimed at the conclusion of the specification. The foregoing and other features, and advantages of the present disclosure are apparent from the following detailed description taken in conjunction with the accompanying drawings in which like elements may be numbered alike and:
Detailed descriptions of one or more embodiments of the disclosed apparatus and/or methods are presented herein by way of exemplification and not limitation with reference to the Figures.
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 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 exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. An 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 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 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 combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
The engine 20 in one non-limiting example is a high-bypass geared aircraft engine. In a further non-limiting 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. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure 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 (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 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)/(514.7° 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 m/sec).
Although the gas turbine engine 20 is depicted as a turbofan, it should be understood that the concepts described herein are not limited to use with the described configuration, as the teachings may be applied to other types of engines such as, but not limited to, turbojets, turboshafts, and three-spool (plus fan) turbofans wherein an intermediate spool includes an intermediate pressure compressor (“IPC”) between a low pressure compressor (“LPC”) and a high pressure compressor (“HPC”), and an intermediate pressure turbine (“IPT”) between the high pressure turbine (“HPT”) and the low pressure turbine (“LPT”).
Referring now to
A turbine cooling air (TCA) conduit 125 provides cooling air into an outer diameter vane cavity 124 defined in part by an outer platform 119 and the full hoop case 114. The vane 106 is hollow so that air can travel radially into and longitudinally downstream from the outer diameter vane cavity 124, through the vane 106 via one or more vane cavities 122, and into a vane inner diameter cavity 123. The vane inner diameter cavity 123 is defined, in part, by an inner platform 119a. Thereafter air may travel through an orifice 120 in the inner air seal 112 and into a rotor cavity 121. Accordingly, cooling air for at least portions of the vane 106 will flow from a platform region, into the vane, and then out of the vane and into another platform region and/or into a hot gaspath/main gaspath. In some arrangements, the platforms 119, 119a can include ejection holes to enable some or all of the air to be injected into the main gaspath.
It is to be appreciated that the longitudinal orientation of vane 106 is illustrated in a radial direction, but other orientations for vane 106 are within the scope of the disclosure. In such alternate vane orientations, fluid such as cooling air can flow into the vane cavity 122 through an upstream opening illustrated herein as outer diameter cavity 124 and out through a downstream opening in vane cavity 122 illustrated herein as inner diameter cavity 123. A longitudinal span of vane cavity 122 being between such openings.
The vane 106, as shown, includes one or more baffles 126 located within the vane 106. The baffles 126 are positioned within one or more respective baffle cavities 128. The baffle cavities 128 are sub-portions or sub-cavities of the vane cavity 122. In some embodiments, such as shown in
As shown and labeled in
As illustratively shown in
For example, turning to
With reference to
Further, the first component 300 includes component extensions 318, 320 and the second component 302 includes component extensions 322, 324. The component extensions 318, 320, 322, 324 include what is typically called a “lip” located at an extreme end or edge of the component extension. The component extensions 318, 320, 322, 324 are arranged to have a cooling flow (e.g., leakage flow) on the cold side and are exposed to hot flow and/or the main gas path on the hot side. The leakage flow can prevent the hot gas from flowing into areas where it is not desirable to be present. In
With reference to
As noted, the cold side (308, 312) of the component extensions (318, 320, 322, 324) are cooled primarily by leakage flow. Depending on the conditions of this flow, the cooling effectiveness can be quite low, which in turn can cause the component extensions to be hot, particularly along the second portion 320b of the component extension 320. The increased heat at the component extensions can reduce the part life of the components.
Accordingly, embodiments provided herein are directed to improving the effectiveness of the leakage cooling by augmenting a convective heat transfer coefficient in a region near the extreme edge of the component extensions (i.e., along the second portion 320b and/or at the end surface 321 shown in
For example, in accordance with some embodiments of the present disclosure, one or more cold-side heat transfer augmentation feature are formed on the cold-side second portion and/or the end surface of the component extension. The cold-side heat transfer augmentation feature are formed to provide additional wetted area and augment the heat transfer coefficient at or along the surface of the cold-side second portion and/or the end surface of the component extension, reducing metal temperature of the extreme edge of the component extension. The cold-side heat transfer augmentation feature can be a variety of shapes, geometries, or other structures that may range from simple to complex. The shapes can be raised out of the “nominal” surface of the component extension or may be recessed into the “nominal” surface of the component extension. In some embodiments, recessing the cold-side heat transfer augmentation feature can result in reduced part weight and can, in some instances, maintain a desired minimum gap size, thus keeping leakage flow constant.
Turning now to
As shown, the cold side 408 of the component extension 420 is defined by two regions or portions, having a cold-side first portion 420a and a cold-side second portion 420b, similar to that described above with respect to
As illustratively shown in this embodiment, the cold-side heat transfer augmentation feature 428 includes a plurality of bases 427 and tips 429 that define the shape of the cold-side heat transfer augmentation feature 428 of the cold-side second portion 420b. The cold-side first portion 420a is a uniform or smooth surface 425 (i.e., lacking heat transfer augmentation features). In this embodiment, the tips 429 of the cold-side heat transfer augmentation feature 428 are flush or level with the surface 425 of the cold-side first portion 420a and the bases 427 of the cold-side heat transfer augmentation feature 428 are depressed from the level of the surface 425 of the cold-side first portion 420a. The cold-side heat transfer augmentation feature 428 extends along the cold-side second portion 420b from the cold-side first portion 420a to the end surface 421 of the component extension 420.
The cold-side heat transfer augmentation feature 428 of the present disclosure are a set of features on the non-gaspath side (i.e., cold side) of a component extension in a region near the extreme edge or lip (i.e., the end surface) of the component extension (e.g., an inner diameter platform aft-lip of a turbine vane). As noted, the cold-side heat transfer augmentation feature can provide a variety of benefits versus a traditional design which is typically a smooth surface. First, the cold-side heat transfer augmentation feature can be used to promote heat transfer by augmenting the convective heat transfer coefficient and increasing the surface area of the cold side of the component extension. This allows the metal temperatures along the cold-side second portion and/or end surface of the component extension to be decreased, improving part life. Secondly, protruding the cold-side heat transfer augmentation feature can be used to stiffen the structure of the component extension. A structurally stiffened component extension can change the structural characteristics of the region, and can be used to address structural concerns such as high cycle fatigue, low cycle fatigue, and yielding. Finally, selectively removing material to create the cold-side heat transfer augmentation feature can reduce part weight while maintaining cross-sectional areas in the gap between the component extension and another element/part/component of the gas turbine engine. Because leakages are increased with increasing gap size, the thickness of the component extension can be constrained. By incorporating the cold-side heat transfer augmentation feature of the present disclosure, the nominal gap can be maintained (and leakage flow unchanged), while part weight is reduced by the selective removal of material.
In accordance with various embodiments of the present disclosure, the area (or length) of the component extension that is defined as the cold-side second portion, and thus contains the cold-side heat transfer augmentation feature, may be based on a percentage of the total area or length of the component extension. For example, in some embodiments, the cold-side second portion, which includes the cold-side heat transfer augmentation feature, may encompass between 20% and 80% of the cold side of the component extension. As such, the cold-side first portion (i.e., the uniform surface portion) may comprise the remaining 80% to 20% of the cold side of the component extension, respectively.
Turning now to
Although a limited number of examples are schematically shown herein, those of skill in the art will appreciate that the heat transfer augmentation feature of the present disclosure can take any shape, size, geometry, orientation, etc. without departing from the scope of the disclosure. For example, in addition to the embodiments shown herein, the heat transfer augmentation feature can be formed as, without limitation, protrusions, scallops, divoted protrusions (e.g., “golf-ball” like), grids, hatching, wave-like, bumps, etc. Moreover, although the heat transfer augmentation feature is shown on an aft component extension, the heat transfer augmentation features of the present disclosure can be located on forward component extensions.
Turning now to
The cold-side heat transfer augmentation feature 928 is defined by a plurality of features or elements that each extend from a base 927 to a tip 929 thereof. As illustrated, in this embodiment, the base 927 of the cold-side heat transfer augmentation feature 928 is level, flush, or even with the surface 925 of the cold-side first portion 920a. As such, the tip 929 is extended from the base 927 toward the part end 932 of the part 930.
As shown, a minimum gap G1 is formed between the tip 929 (or radial extent) of the cold-side heat transfer augmentation feature 928 and the part end 932. The minimum gap G1 may be a minimum required gap or clearance between the component 900 and the part 930 to prevent contact therebetween (e.g., due to thermal expansion, movement of the relative elements, etc.). Typically, the component extension 920 would be formed of a solid material with all of the spacing between the component extension 920 and the part end 932 being set at the minimum gap G1. However, by using the tip 929 of the cold-side heat transfer augmentation feature 928 to set the minimum gap G1, the rest of the component extension 920, including the cold-side first portion 920a, can be thinner, thus reducing the total part-weight of the component 900, while also maintaining the minimum gap G1. The thinner component extension 920, as compared to typical component extension, creates an increased spacing G2 to be formed between the component extension 920 and the part end 932.
Turning now to
The cold-side heat transfer augmentation feature 1028 is defined by a plurality of features or elements that each extend from a base 1027 to a tip 1029 thereof. As illustrated, in this embodiment, the tip 1029 of the cold-side heat transfer augmentation feature 1028 is level, flush, or even with the surface 1025 of the cold-side first portion 1020a. As such, the base 1029 is recessed from the tip 1027 away the part end 1032 of the part 1030.
As shown, a minimum gap G1 is formed between the tip 1029 (or radial extent) of the cold-side heat transfer augmentation feature 1028 and the part end 1032. The minimum gap G1 may be a minimum required gap or clearance between the component 1000 and the part 1030 to prevent contact therebetween (e.g., due to thermal expansion, movement of the relative elements, etc.). Typically, the component extension 1020 would be formed of a solid material with all of the spacing between the component extension 1020 and the part end 1032 being set at the minimum gap G1. However, by using the tip 1029 of the cold-side heat transfer augmentation feature 1028 to set the minimum gap G1, the rest of the component extension 1020 along the cold-side second portion 1020b, can be thinner, thus reducing the total part-weight of the component 1000, while also maintaining the minimum gap G1. The thinner component extension 1020, as compared to typical component extension, creates an increased spacing G2 to be formed between the component extension 1020 and the part end 1032 at the location of the cold-side second portion 1020b. In this embodiment, the relatively thicker cold-side first portion 1020a (as compared to the embodiment shown in
As used herein, the term “base” of a cold-side heat transfer augmentation feature is a part, surface, or portion of the feature that is farthest from an adjoining part, and thus defines the greatest gap between the component extension and the adjoining part. Further, as used herein, the “tip” of a cold-side heat transfer augmentation feature is the portion, surface, or part of the feature that is closest to an adjoining part, and thus defines the minimum gap between the component extension and the adjoining part. As such, regardless of the mechanism for manufacturing the cold-side heat transfer augmentation feature and/or the shape, geometry, or orientation, the “base” and “tip” are generic terms that have the above definitions.
Advantageously, embodiments provided herein are directed to heat transfer augmentation feature formed in the cold side of component extensions. As such, heat transfer to the cool leakage flow is augmented. Such augmentation can result in cooler metal temperatures, and increased part life. The heat transfer augmentation feature can also impact the structural capability of the component extensions and can be used to tailor modal responses. Finally, in accordance with some embodiments, some arrangements can result in reduced part weight and a subset of these configurations will reduce weight while maintaining the same minimum gap size.
Although the various above embodiments are shown as separate illustrations, those of skill in the art will appreciate that the various features can be combined, mix, and matched to form an airfoil having a desired cooling scheme that is enabled by one or more features described herein. Thus, the above described embodiments are not intended to be distinct arrangements and structures of airfoils and/or core structures, but rather are provided as separate embodiments for clarity and ease of explanation.
As used herein, 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. For example, “about” may include a range of ±8%, or 5%, or 2% of a given value or other percentage change as will be appreciated by those of skill in the art for the particular measurement and/or dimensions referred to herein.
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. It should be appreciated that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” “radial,” “axial,” “circumferential,” and the like are with reference to normal operational attitude and should not be considered otherwise limiting.
While the present disclosure has been described with reference to an illustrative 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.
This invention was made with Government support under contract number W58RGZ-16-C-0046 awarded by the United States Army. The Government has certain rights in the invention.