Patent Application
20190316472
This disclosure relates to gas turbine engines and particularly to internally cooled airfoils of rotor blades and stator vanes.
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 combustor 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 section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
As is well known, the aircraft engine industry is experiencing a significant effort to improve the gas turbine engine's performance while simultaneously decreasing its weight. The ultimate goal has been to attain the optimum thrust-to-weight ratio. One of the primary areas of focus to achieve this goal is the “hot section” of the engine since it is well known that engine's thrust/weight ratio is significantly improved by increasing the temperature of the turbine gases. However, turbine gas temperature is limited by the metal temperature constraints of the engine's components. Significant effort has been made to achieve higher turbine operating temperatures by incorporating technological advances in the internal cooling of the turbine blades.
Serpentine core cooling passages have been used to cool turbine blades. The serpentine cooling passage is arranged between the leading and trailing edge core cooling passages in a chord-wise direction. One typical serpentine configuration provides “up” passages arranged near the leading and trailing edges fluidly joined by a “down” passage. This type of cooling configuration may have inadequacies in some applications. To this end, a double wall cooling configuration has been used to improve turbine blade cooling.
In a double wall blade configuration, thin hybrid skin core cavity passages extend radially and are provided in a thickness direction between the core cooling passages and each of the pressure and suction side exterior airfoil surfaces to separate “hot” and “cold” walls. Double wall cooling has been used as a technology to improve the cooling effectiveness of a turbine blades, vanes, blade out air seals, combustor panels, or any other hot section component. Often, core support features are used to form resupply holes and interconnect a main body core to a hybrid skin core. The main body core creates the core passages, and the hybrid skin core creates the skin passages.
Some turbine stator vanes incorporate an internal radially extending baffle, which is supported with respect to the outer airfoil wall by interiorly extending radial ribs. The baffle is typically constructed from a thin sheet of metal and is arranged centrally within the airfoil. Cooling fluid is supplied to the baffle through, for example, an outer platform. Holes in the baffle provide impingement cooling from the baffle onto an inner surface of the outer airfoil, or “hot”, wall.
With traditional double wall configurations, a cooling benefit is derived from passing coolant air from the internal radial flow and/or serpentine cavities through the “cold” wall via cooling holes (aka resupply holes) and impinging the flow on the “hot” wall. The cooling effectiveness of airfoils having a double wall arrangement is affected by a variety of factors, including cooling hole geometry provided in the “cold” wall. The relative thinness of the inner, “cold” wall can limit the cooling hole geometry that can be used, which may limit the cooling effectiveness of the airfoil.
In one exemplary embodiment, an airfoil includes pressure and suction side walls that extend in a chord-wise direction between leading and trailing edges. The pressure and suction side walls extend in a radial direction to provide an exterior airfoil surface. A core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a tip. An outer cooling passage is arranged in one of the pressure and suction side walls to form a hot side wall and a cold side wall. The hot side wall defines a portion of the exterior airfoil surface and the cold side wall defines a portion of the core cooling passage. The core cooling passage, the outer cooling passage and a cooling hole fluidly interconnects the core and outer cooling passages. The cold side wall has a first thickness and a protrusion that extends from the cold side wall beyond the first thickness to a second thickness. The cooling hole is arranged at least partially in the protrusion.
In a further embodiment of the above, the first thickness of the cold side wall circumscribes the protrusion.
In a further embodiment of any of the above, the cooling hole includes a diffuser.
In a further embodiment of any of the above, the cooling hole includes an inlet at the core cooling passage and an outlet at the outer cooling passage. The inlet is arranged at the protrusion.
In a further embodiment of any of the above, the cooling hole includes an inlet at the core cooling passage and an outlet at the outer cooling passage. The outlet is arranged at the protrusion.
In a further embodiment of any of the above, the protrusion is arranged in one of the core and outer cooling passages. One of the core and outer cooling passages is configured to receive a fluid moving in a flow direction. The protrusion has leading and trailing surfaces with respect to the flow direction.
In a further embodiment of any of the above, the protrusion is provided in both of the core and outer cooling passages.
In a further embodiment of any of the above, the leading surface has a corner and/or a smooth contour.
In a further embodiment of any of the above, the trailing surface has a corner and/or a smooth contour.
In a further embodiment of any of the above, the airfoil is a turbine blade. The hot and cold side walls are integral with one another.
In a further embodiment of any of the above, the airfoil is a stator vane. The cold side wall is provided by a baffle supported by the hot side wall. The baffle is discrete from the hot side wall.
In a further embodiment of any of the above, ribs and/or pin fins support the baffle relative to the hot side wall.
In another exemplary embodiment, a gas turbine engine include a combustor section arranged fluidly between compressor and turbine sections. An airfoil is arranged in the turbine section. The airfoil includes pressure and suction side walls that extend in a chord-wise direction between leading and trailing edges. The pressure and suction side walls extend in a radial direction to provide an exterior airfoil surface. A core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a tip. An outer cooling passage is arranged in one of the pressure and suction side walls to form a hot side wall and a cold side wall. The hot side wall defines a portion of the exterior airfoil surface and the cold side wall defines a portion of the core cooling passage. The core cooling passage, outer cooling passage and a cooling hole fluidly interconnects the core and outer cooling passages. The cold side wall has a first thickness and a protrusion that extends from the cold side wall beyond the first thickness to a second thickness. The cooling hole is arranged at least partially in the protrusion.
In a further embodiment of the above, the first thickness of the cold side wall circumscribes the protrusion.
In a further embodiment of any of the above, the cooling hole includes a diffuser.
In a further embodiment of any of the above, the cooling hole includes an inlet at the core cooling passage and an outlet at the outer cooling passage. One of the inlet and the outlet is arranged at the protrusion.
In a further embodiment of any of the above, the protrusion is arranged in one of the core and outer cooling passages. One of the core and outer cooling passages is configured to receive a fluid moving in a flow direction. The protrusion has leading and trailing surfaces with respect to the flow direction. The leading surface has a corner and/or a smooth contour.
In a further embodiment of any of the above, the protrusion is arranged in one of the core and outer cooling passages. One of the core and outer cooling passages is configured to receive a fluid moving in a flow direction. The protrusion has leading and trailing surfaces with respect to the flow direction. The trailing surface has a corner and/or a smooth contour.
In a further embodiment of any of the above, the airfoil is a stator vane. The cold side wall is provided by a baffle supported by the hot side wall. The baffle is discrete from the hot side wall. The hot side wall includes ribs and/or pin fins that extend inwardly to support the baffle.
In another exemplary embodiment, a casting mold for an airfoil includes a shell and a core assembly together defining pressure and suction side walls that extend in a chord-wise direction between leading and trailing edges. The pressure and suction side walls extend in a radial direction to provide an exterior airfoil surface. A core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a tip. An outer cooling passage is arranged in one of the pressure and suction side walls to form a hot side wall and a cold side wall. The hot side wall defines a portion of the exterior airfoil surface and the cold side wall defines a portion of the core cooling passage. The core cooling passage and the outer cooling passage and a cooling hole fluidly interconnects the core and outer cooling passages. The cold side wall has a first thickness and a protrusion that extends from the cold side wall beyond the first thickness to a second thickness. The cooling hole is arranged at least partially in the protrusion.
The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:
The embodiments, examples and alternatives of the preceding paragraphs, the claims, or the following description and drawings, including any of their various aspects or respective individual features, may be taken independently or in any combination. Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible.
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 X 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 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 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 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 is 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 X 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 geared architecture 48 may be varied. For example, geared architecture 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 geared architecture 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 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).
The disclosed cooling configuration is particularly beneficial for turbine blades and vanes of a gas turbine engine where internal cooling of the airfoil is desired, although the disclosed arrangement may also be used in the compressor section. For exemplary purposes, a double wall stator vane 63 is described in connection with
Referring to
The section includes a fixed stage 60 that provides a circumferential array of vanes 63 arranged axially adjacent to a rotating stage 62 with a circumferential array of blades 69. In the example, the vanes 63 include an outer diameter portion 64 having hooks 65 that support the array of vanes 63 with respect to a case structure. An airfoil 68 extends radially from the outer platform 64 to an inner diameter portion or platform 66. It should be understood that the disclosed vane arrangement could be used for vane structures cantilevered at the inner diameter portion of the airfoil.
Referring to
Referring to
One or more ribs (or pin fins) 90 are provided between and connect the pressure and suction sides 78, 80. In one example, the one or more ribs 90 extend radially. The ribs 90 separate a trailing edge cooling cavity 88 from the perimeter cavity 86. In one example, holes may be provided in the ribs 90 to provide cooling fluid from the perimeter cavity 86 into the trailing edge cooling cavity 88. Fluid exits the trailing edge 76 as is known. In another example, the one or more ribs extend axially or extend with an angle.
An impingement cooling arrangement 92 is provided to cool the leading edge 74. In the example, a portion of the baffle 72 includes impingement cooling holes 94 that provide impingement cooling fluid to an interior or backside of the exterior wall 82 at the leading edge 74.
In one example, the baffle 72 is provided by sheet steel, for example, a single sheet, and includes an outer contour generally free of protrusions. The outer contour is provided by plastic deformation. In another example, the baffle may be cast. The cooling holes, such as the impingement cooling holes 94, are provided in the baffle 72 using at least one of drilling, laser drilling, electro discharge machining, or casting.
The exterior wall 82 includes an interior surface 98 from which multiple pin fins extend interiorly to a terminal end. Alternatively, the pin fins may extend outward from the baffle to a terminal end. The terminal end supports the baffle 72, which is a separate, discrete structure from the exterior wall 82 in the example. In one example, the pin fins 96 are arranged in rows and radially spaced from one another, as best shown in
The blade 69 also includes a double wall arrangement. Referring to
The airfoil 178 of
In the example, the airfoil 178 includes a serpentine cooling passage 190 provided between the pressure and suction walls 186, 188. The serpentine cooling passage 190 provides a core cooling passage. The disclosed skin core and cooling hole arrangement may be used with other cooling passage configurations, including non-serpentine cooling passage arrangements. As will be appreciated from the disclosure below, it should be understood that the central core passages from which resupply flow is bled might consist of a single radial core passage, and or multiple radial central core passages. Additionally one or more radial flow central core cooling passages may also be combined with a central core passage serpentine consisting of two or more continuous central cooling passages from which resupply flow may also be supplied.
Referring to
Referring to
As shown in
Referring to
Referring to
The protrusion 1220 may be provided on either side of the cold side wall 1202, that is, in the outer cooling passage 1212 (
The protrusion 1220 have leading and trailing surfaces 1222, 1224 with respect to the flow direction that are designed to achieve a desired flow characteristic. For example, the leading surface 1222 may have a sharp corner to generate turbulent flow for enhanced cooling and/or block debris for cleaner airflow. The leading surface 1222 may have a smooth contour to encourage laminar flow over the cooling hole 1210. The trailing surface 1224 may have a sharp corner to generate a recirculation flow or tumble downstream from the cooling hole 1210, or a smooth contour.
The cooling holes 210 may have various shapes. One or more cooling holes 210 may be fluidly connected to each discrete skin passage 198, or outer cooling passage. The exit 214 may provide a diffuser (right cooling hole 210 in
Different diffuser geometries 218, 318, 418 are shown in
Referring to
The exit hole geometry as illustrated in
Supply holes 210 with a diffuser 518, 618, 718 are shown in
The exit hole geometries, as illustrated in
Using techniques typically used in external film cooling, one may orient the core support resupply cooling or cooling hole features in the streamwise direction of the cooling air flow in the hybrid skin core cavity cooling passage. By improving the relative alignment of the two separate flow streams the momentum mixing associated with the differences in the inertial Coriolis and buoyancy forces between the two separate flow streams will be significantly reduced. In so doing the high pressure losses typically observed between the two independent flow streams emanating from a cooling hole 210 oriented normal to the downstream flow field within the hybrid skin core cavity cooling passage 198 can be significantly reduced and the resulting mixing length can be dissipated quickly along the streamwise direction of cooling flow within the hybrid skin core cavity cooling passage. These techniques may also be used to manufacture the disclosed vane if an integral baffle is desired.
Additive manufacturing and Fugitive Core casting processes enable design flexibility in gas turbine manufacturing. One of the design spaces that additive opens up is in the design of ceramic cores used in the investment casting process. Traditional ceramic cores are made with a core die, which has a finite number of “pull planes.” These pull planes restrict the design of ceramic cores to prevent features from overhanging in the direction that the die is pulled when the cores are removed. Additive manufacturing and Fugitive Core processes can remove those manufacturing restrictions, as dies are no longer required to create the ceramic cores of the internal cooling passages and internal convective cooling features, such as trip strips, pedestals, ribs, cooling holes, etc.
An additive manufacturing process may be used to produce an airfoil for the disclosed blade and vane. Alternatively, a core for casting may be constructed using additive manufacturing and/or Fugitive Core manufacturing may be used to provide the correspondingly shaped cooling hole geometries when casting the airfoil. These advanced manufacturing techniques enable unique core features to be integrally formed and fabricated as part of the entire ceramic core body and then later cast using conventional loss wax investment casting processes. Alternatively powdered metals suitable for aerospace airfoil applications may be used to fabricate airfoil cooling configurations and complex cooling configurations directly. The machine deposits multiple layers of powdered metal onto one another. The layers are joined to one another with reference to CAD data, which relates to a particular cross-section of the airfoil. In one example, the powdered metal may be melted using a direct metal laser sintering process or an electron-beam melting process. With the layers built upon one another and joined to one another cross-section by cross-section, an airfoil with the above-described geometries may be produced. The airfoil may be post-processed to provide desired structural characteristics. For example, the airfoil may be heated to reconfigure the joined layers into a single crystalline structure.
A casting mold is shown in
It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements would benefit herefrom. Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the embodiments of the present invention. Additionally it is important to note that any complex multi-facetted cooling hole geometries that bridge centrally located core cooling passages and peripherally located cooling passages can be created at any number of radial, circumferential, and/or tangential locations within an internal cooling configuration. The quantity, size, orientation, and location will be dictated by the necessity to increase the local thermal cooling effectiveness and achieve the necessary thermal performance required to mitigate hot section part cooling airflow requirements, as well as, meet part and module level durability life, stage efficiency, module, and overall engine cycle performance and mission weight fuel burn requirements.
Although the different examples have specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples.
Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.
