Patent Application
20190338649
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.
Airfoils may incorporate various cooling cavities located adjacent external sidewalls. Such cooling cavities are subject to both hot material walls (exterior or external) and cold material walls (interior or internal). Although such cavities are designed for cooling portions of airfoil bodies, improved cooling designs may be desirable.
According to some embodiments, airfoils for gas turbine engines are provided. The airfoils include an airfoil body extending between a leading edge and a trailing edge in an axial direction, between a pressure side and a suction side in a circumferential direction, and between a root and a tip in a radial direction, a first transitioning leading edge cavity located adjacent one of the pressure side and the suction side proximate the root of the airfoil body and transitioning axially toward the leading edge as the first transitioning leading edge cavity extends radially toward the tip, and a second transitioning leading edge cavity adjacent the other of the pressure side and the suction side and adjacent the leading edge proximate the root of the airfoil body and transitioning axially toward the trailing edge as the second transitioning leading edge cavity extends radially toward the tip. A portion of the second transitioning leading edge cavity shields a portion of the first transitioning leading edge cavity proximate the root of the airfoil body.
In addition to one or more of the features described above, or as an alternative, further embodiments of the airfoils may include that the second transitioning leading edge cavity comprise an impingement portion proximate the root.
In addition to one or more of the features described above, or as an alternative, further embodiments of the airfoils may include that the impingement portion of the second transitioning leading edge cavity shields the first transitioning leading edge cavity.
In addition to one or more of the features described above, or as an alternative, further embodiments of the airfoils may include that the second transitioning leading edge cavity is located aft of the first transitioning leading edge cavity proximate the tip.
In addition to one or more of the features described above, or as an alternative, further embodiments of the airfoils may include that the second transitioning leading edge cavity spans the airfoil body between the pressure side and the suction side proximate the tip.
In addition to one or more of the features described above, or as an alternative, further embodiments of the airfoils may include that the first transitioning leading edge cavity forms a film cooling cavity along the leading edge at the tip of the airfoil body.
In addition to one or more of the features described above, or as an alternative, further embodiments of the airfoils may include that the airfoil body has a first thickness along the leading edge proximate the root and a second thickness along the leading edge proximate the tip, wherein the first thickness is different from the second thickness.
In addition to one or more of the features described above, or as an alternative, further embodiments of the airfoils may include that the first thickness is less than the second thickness.
In addition to one or more of the features described above, or as an alternative, further embodiments of the airfoils may include that the first thickness is between 0.020″ and 0.045″, and the second thickness is between 0.045″ and 0.070″.
In addition to one or more of the features described above, or as an alternative, further embodiments of the airfoils may include at least one main body cavity located aft of the first transitioning leading edge cavity and the second transitioning leading edge cavity.
According to some embodiments, core assemblies for forming airfoils of gas turbine engines are provided. The core assemblies include a first transitioning leading edge cavity core positioned to form a portion of one of a pressure side and a suction side of a formed airfoil body proximate a root of the formed airfoil body, the first transitioning leading edge cavity core transitions axially forward as the first transitioning leading edge cavity extends radially toward a tip of the formed airfoil body to define a portion of a leading edge of the formed airfoil body at the tip, and a second transitioning leading edge cavity core positioned adjacent the first transitioning leading edge cavity core when arranged to form the airfoil, wherein the second transitioning leading edge cavity core is positioned to form a portion of the other of the pressure side and the suction side proximate the root of the formed airfoil body and transitions axially aftward of the first transitioning leading edge cavity core as the second transitioning leading edge cavity core extends radially toward the tip of the formed airfoil body.
In addition to one or more of the features described above, or as an alternative, further embodiments of the core assemblies may include that the second transitioning leading edge cavity core comprise an impingement cavity core adjacent the leading edge of the formed airfoil body and proximate the root.
In addition to one or more of the features described above, or as an alternative, further embodiments of the core assemblies may include that the impingement cavity core of the second transitioning leading edge cavity core is arranged to shield the first transitioning leading edge cavity.
In addition to one or more of the features described above, or as an alternative, further embodiments of the core assemblies may include that the second transitioning leading edge cavity core is located aft of the first transitioning leading edge cavity core proximate the tip of the formed airfoil body.
In addition to one or more of the features described above, or as an alternative, further embodiments of the core assemblies may include that the second transitioning leading edge cavity core spans the formed airfoil body between the pressure side and the suction side proximate the tip of the formed airfoil body.
In addition to one or more of the features described above, or as an alternative, further embodiments of the core assemblies may include that the first transitioning leading edge cavity core is arranged to form a film cooling cavity along the leading edge at the tip of the formed airfoil body.
In addition to one or more of the features described above, or as an alternative, further embodiments of the core assemblies may include at least one main body cavity core located aft of the first transitioning leading edge cavity core and the second transitioning leading edge cavity core.
According to some embodiments, gas turbine engines are provided. The gas turbine engines include a turbine section having a plurality of airfoils. At least one airfoil includes an airfoil body extending between a leading edge and a trailing edge in an axial direction, between a pressure side and a suction side in a circumferential direction, and between a root and a tip in a radial direction, a first transitioning leading edge cavity located adjacent one of the pressure side and the suction side proximate the root of the airfoil body and transitioning axially toward the leading edge as the first transitioning leading edge cavity extends radially toward the tip, and a second transitioning leading edge cavity adjacent the other of the pressure side and the suction side and adjacent the leading edge proximate the root of the airfoil body and transitioning axially toward the trailing edge as the second transitioning leading edge cavity extends radially toward the tip. A portion of the second transitioning leading edge cavity shields a portion of the first transitioning leading edge cavity proximate the root of the airfoil body.
In addition to one or more of the features described above, or as an alternative, further embodiments of the gas turbine engines may include that the second transitioning leading edge cavity comprise an impingement portion proximate the root.
In addition to one or more of the features described above, or as an alternative, further embodiments of the gas turbine engines may include that the impingement portion of the second transitioning leading edge cavity shields the first transitioning leading edge cavity.
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 can be 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 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. 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 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”).
The turbine 200 is housed within a case 212, which may have multiple parts (e.g., turbine case, diffuser case, etc.). In various locations, components, such as seals, may be positioned between the airfoils 201, 202 and the case 212. For example, as shown in
As shown and labeled in
Typically, airfoil cooling includes impingement cavities for cooling various hot surfaces of the airfoils. For example, it may be desirable to position a leading edge impingement cavity immediately adjacent to the external leading edge of the airfoil (e.g., left side edge of the airfoils 201, 202). The leading edge impingement cavity is typically supplied cooling airflow from impingement apertures which serve as conduits for cooling air that originates within the leading edge cooling cavities of the airfoil. Once in the leading edge impingement cavity, the cooling air flow is expelled through an array of shower head holes, thus providing increased convective cooling and a protective film to mitigate the locally high external heat flux along the leading edge airfoil surface.
Traditionally, investment casting manufacturing processes utilize hard tooling “core dies” to create both external airfoil and internal cooling geometries. In order to fabricate internal cooling geometries, it is required that the definition of the features be created in the same relative orientation (approximately parallel) to the “pull” direction of the core die tooling. As a result, the orientation and location of any internal cooling features is limited by virtue of core tooling/core die manufacturing processes used for investment casting of turbine airfoils. Further, various cooling feature may require drilling through the external walls or surfaces of the airfoil to fluidly connect to internal cavities thereof (e.g., to form film cooling holes). The orientation of the local internal rib geometry and positioning of the impingement cooling apertures is necessary to ensure optimal internal convective heat transfer characteristics are achieved to mitigate high external heat flux regions.
For example, turning now to
As shown in
Air that impinges into the leading edge cavity 320 (or other forward and side cooling cavities 320, 322, 324) may be expunged onto a hot external surface of the airfoil 300 through one or more film cooling holes 336. During manufacturing of the airfoil 300, the film cooling holes 336 may be drilled into or through the external surfaces of the airfoil body 302. With reference to
The skin core cavities described above may be very efficient at cooling the hot wall of the airfoil, but this efficiency may degrade as the hot wall thickness increases. Accordingly, to maintain improved cooling, thin airfoil exterior walls may be preferable. However, other considerations may require increased thickness external walls of the airfoil. For example, one region of an airfoil that may require an increased external wall thickness is the leading edge of the airfoil where the part must be designed to withstand foreign object damage “FOD” (e.g., debris passing through the hot gas path and contacting and/or impacting the leading edge of the airfoil). To take advantage of skin core cavity cooling and also being able to withstand FOD, embodiments of present disclosure are directed to airfoils and cores for making the same that incorporate a modified cooling scheme that has a transition from a skin core cavity to an impingement cavity configuration. This transition can be employed, in some embodiments, toward an outer diameter or outer span of the airfoil. Further, the impingement cavity configuration may incorporate film cooling at the outer spans. Accordingly, a more robust airfoil design can be achieved as compared to just impingement cooling or just skin core cooling.
Turning now to
The airfoil 400, as shown, is arranged as a blade having an airfoil body 402 that extends from a platform 404. The airfoil body 402 attaches to or is connected to the platform 404 at a root 406 (i.e., inner diameter) and extends radially outward to a tip 408 (i.e., outer diameter). The platform 404 may be integrally formed with or attached to an attachment element 410 and/or the airfoil body 402, the attachment element 410 being configured to attach to or engage with a rotor disc for installation of the airfoil 400 to the rotor disc. The airfoil body 402 extends in an axial direction A from a leading edge 412 to a trailing edge 414, and in a radial direction R from the root 406 to the tip 408. In the circumferential direction C, the airfoil body 402 extends between a pressure side 416 and a suction side 418.
The airfoil body 402 defines a number of internal cooling cavities. For example, as shown in
As noted, the first transitioning leading edge cavity 422 transitions from being proximate the pressure side 416 to being proximate the leading edge 412. The second transitioning leading edge cavity 424 transitions from being proximate the leading edge 412 and the suction side 418 to being proximate both the pressure and suction sides 416, 418. Proximate the root 406, as shown in cross-section in
The first transitioning leading edge cavity 422 is located aft of the impingement portion 424b of the second transitioning leading edge cavity 424 at the root 406. Accordingly, the amount of heat pickup within the first transitioning leading edge cavity 422 at the root 406 will be reduced, thus keeping the temperature of the air within the first transitioning leading edge cavity 422 relatively cool as compared to the air within the second transitioning leading edge cavity 424 at the root 406.
As the first and second transitioning leading edge cavities 422, 424 extend radially outward toward the tip 408, the geometries of the first and second transitioning leading edge cavities 422, 424 change. For example, as shown in
Proximate the tip 408 of the airfoil body 402, as shown in
In some embodiments, one or both of the transitioning leading edge cavities (or portions thereof) can include one or more heat transfer augmentation features. Heat transfer augmentation features can include, but are not limited to, turbulators, trip strips (including, but not limited to normal, skewed, segmented skewed, chevron, segmented chevron, W-shaped, and discrete W's), pin fins, hemispherical bumps and/or dimples, as well as non-hemispherical shaped bumps and/or dimples, etc.
Accordingly, in accordance with some embodiments of the present disclosure, a cooling passage starts as a pressure side skin core on the inner diameter of the part and is used to efficiently cool the pressure side inner diameter. There is little risk of impact damage at these spans and the heat load is generally controlled due to concern regarding a combination of high stress and temperature in the same region. The skin core is then brought forward to the leading edge to act as a film cooling cavity for the outer diameter. At the outer diameter, where the part is more likely to have a higher heat load and has an elevated risk of impact damage, an impingement scheme with cooling air is employed. This type of configuration will be balanced to provide an optimal balance of damage tolerance and cooling effectiveness.
Additionally, embodiments provided herein may enable improved robustness while provide the cooling described herein (e.g., shifting of cooling air from the leading edge aftward and relatively cooler air forward to the leading edge). For example, turning to
As shown, the airfoil 530 has an airfoil body 532 defining a first transitioning leading edge cavity 534 and a second transitioning leading edge cavity 536. The first transitioning leading edge cavity 534 is proximate to a pressure side 538 at the root of the airfoil body 532 (as shown in
As shown in
Turning now to
The first transitioning leading edge cavity core 654 is arranged at the pressure side of the formed airfoil and is arranged to form a cavity that is substantially protected from the thermal pick up that occurs at the leading edge of the formed airfoil, as shown and described above. The first transitioning leading edge cavity core 654 then transitions forward to form a film cooling scheme at the tip of the formed airfoil. The second transitioning leading edge cavity core 656 is arranged forward of the first transitioning leading edge cavity core 654 at the root of the formed airfoil and includes an impingement cavity core 658. The second transitioning leading edge cavity core 656 will transition aftward of the first transitioning leading edge cavity core 654 proximate the tip of the formed airfoil. The second transitioning leading edge cavity core 656 can include one or more core elements to join the impingement cavity core 658 to the rest of the second transitioning leading edge cavity core 656 to form one or more impingement holes therebetween in a formed airfoil, as shown and described above. Further, the first transitioning leading edge cavity core 654 can include one or more core elements to form film cooling holes in an airfoil body of a formed airfoil, as will be appreciated by those of skill in the art (or film cooling holes may be drilled or otherwise formed post-airfoil body formation).
Advantageously, embodiments described herein can incorporate skin cavity/core (e.g., thin wall) cooling at various locations but may also include improved FOD protection where needed. Accordingly, embodiments provided herein can enable improved part life and thrust specific fuel consumption.
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.
