US20170306764A1 - Airfoil for a turbine engine - Google Patents

Airfoil for a turbine engine Download PDF

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Publication number
US20170306764A1
US20170306764A1 US15/138,624 US201615138624A US2017306764A1 US 20170306764 A1 US20170306764 A1 US 20170306764A1 US 201615138624 A US201615138624 A US 201615138624A US 2017306764 A1 US2017306764 A1 US 2017306764A1
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US
United States
Prior art keywords
protuberance
component
airfoil
film hole
fluid flow
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Abandoned
Application number
US15/138,624
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English (en)
Inventor
Douglas Gerard Konitzer
Scott Ronald Bunker
Robert David Briggs
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
General Electric Co
Original Assignee
General Electric Co
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by General Electric Co filed Critical General Electric Co
Priority to US15/138,624 priority Critical patent/US20170306764A1/en
Assigned to GENERAL ELECTRIC COMPANY reassignment GENERAL ELECTRIC COMPANY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: Briggs, Robert David, BUNKER, RONALD SCOTT, Konitzer, Douglas Gerard
Priority to JP2017078572A priority patent/JP2017198205A/ja
Priority to CA2964139A priority patent/CA2964139A1/en
Priority to EP17167906.1A priority patent/EP3239462A1/en
Priority to CN201710281295.8A priority patent/CN107448300B/zh
Publication of US20170306764A1 publication Critical patent/US20170306764A1/en
Priority to US16/869,831 priority patent/US20200277862A1/en
Priority to US18/337,841 priority patent/US20240159151A1/en
Abandoned legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/002—Cleaning of turbomachines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F05D2220/323—Application in turbines in gas turbines for aircraft propulsion, e.g. jet engines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/185—Two-dimensional patterned serpentine-like
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/23—Three-dimensional prismatic
    • F05D2250/232—Three-dimensional prismatic conical
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • Turbine engines and particularly gas or combustion turbine engines, are rotary engines that extract energy from a flow of combusted gases passing through the engine onto a multitude of rotating turbine blades.
  • Turbine engines for aircraft are often designed to operate at high temperatures to maximize engine efficiency, so cooling of certain engine components can be beneficial or necessary.
  • cooling is accomplished by ducting cooler air from the high and/or low pressure compressors to the engine components that require cooling.
  • Temperatures in the high pressure turbine can be 1000° C. to 2000° C. and the cooling air from the compressor can be around 500° C. to 700° C. While the compressor air is a high temperature, it is cooler relative to the turbine air, and can be used to cool the turbine.
  • Contemporary engine components such as the rotating blades, necessarily account for a portion of the overall engine weight. Decreasing the weight of these engine components is desirable to increase engine efficiency. Decreasing weight of the engine components can be accomplished by utilizing thinner walls for the components, for example. However, thinner walls include a decreased volume through which film holes can extend, which can decrease the effectiveness of the film holes. Thus, it is desirable to utilize thinner walls for the engine components to decrease system weight while providing sufficient length for the film holes to maintain cooling effectiveness.
  • embodiments of the invention relate to a component for a turbine engine, which generates a hot combustion gas flow, and provides a cooling fluid flow defining a cooling fluid flow, includes a wall separating the hot combustion gas flow from the cooling fluid flow.
  • the wall includes a hot surface facing the hot combustion gas flow and a cool surface facing the cooling fluid flow, having a nominal thickness.
  • the component further includes at least one localized, radiused protuberance extending from the cool surface and a film hole extending through the protuberance and the wall.
  • the film hole has a greater length than the nominal thickness of the wall.
  • embodiments of the invention relate to an airfoil for a turbine engine including a wall having a first side adjacent to a first fluid flow and a second side adjacent to a second fluid flow, having a nominal thickness.
  • the component further includes at least one localized, radiused protuberance extending from the wall and a film hole extending through the protuberance and the wall, having a length greater than the nominal thickness of the wall.
  • embodiments of the invention relate to a method of cooling an engine component having a cool surface.
  • the method includes passing a cooling fluid flow along the cool surface and providing at least a portion of the cooling fluid flow through a film hole in a protuberance extending from the cool surface.
  • FIG. 1 is a schematic cross-sectional diagram of a gas turbine engine for an aircraft.
  • FIG. 2 is an isometric view of an airfoil of the gas turbine engine of FIG. 1 .
  • FIG. 3 is a cross-sectional view of the airfoil of FIG. 2 having walls including protuberances with film holes.
  • FIG. 4 is a perspective view of the wall having the protuberance with the film hole extending through the protuberance and the wall.
  • FIG. 5 is a cross-sectional view of the wall of FIG. 4 illustrating a profile of the protuberance and film hole across a direction of a cooling fluid flow.
  • FIG. 6 is a cross-sectional view of the wall of FIG. 4 illustrating a profile of the protuberance and film hole in the direction of the cooling fluid flow.
  • FIG. 7 is a cross-sectional view of another embodiment of the film hole extending in a direction opposite of the cooling fluid flow.
  • FIG. 8 is a cross-sectional view of yet another embodiment the film hole having an inlet on a forward face of the protuberance.
  • FIG. 9 is a cross-sectional view of the protuberance having a recess with the film hole inlet disposed in the recess.
  • FIG. 10 is a perspective view of an elongated protuberance having a film hole offset form the direction of the cooling fluid flow.
  • the described embodiments of the present invention are directed to an engine component for a turbine engine having at least one protuberance disposed on a wall of the engine component with a film hole extending through the protuberance and the wall.
  • the present invention will be described with respect to the turbine for an aircraft gas turbine engine. It will be understood, however, that the invention is not so limited and may have general applicability within an engine, as well as in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications.
  • forward or “upstream” refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component.
  • downstream or “downstream” used in conjunction with “forward” or “upstream” refers to a direction toward the rear or outlet of the engine or being relatively closer to the engine outlet as compared to another component.
  • radial refers to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference.
  • FIG. 1 is a schematic cross-sectional diagram of a gas turbine engine 10 for an aircraft.
  • the engine 10 has a generally longitudinally extending axis or centerline 12 extending forward 14 to aft 16 .
  • the engine 10 includes, in downstream serial flow relationship, a fan section 18 including a fan 20 , a compressor section 22 including a booster or low pressure (LP) compressor 24 and a high pressure (HP) compressor 26 , a combustion section 28 including a combustor 30 , a turbine section 32 including a HP turbine 34 , and a LP turbine 36 , and an exhaust section 38 .
  • LP booster or low pressure
  • HP high pressure
  • the fan section 18 includes a fan casing 40 surrounding the fan 20 .
  • the fan 20 includes a plurality of fan blades 42 disposed radially about the centerline 12 .
  • the HP compressor 26 , the combustor 30 , and the HP turbine 34 form a core 44 of the engine 10 , which generates combustion gases.
  • the core 44 is surrounded by core casing 46 , which can be coupled with the fan casing 40 .
  • a LP shaft or spool 50 which is disposed coaxially about the centerline 12 of the engine 10 within the larger diameter annular HP spool 48 , drivingly connects the LP turbine 36 to the LP compressor 24 and fan 20 .
  • the spools 48 , 50 are rotatable about the engine centerline and couple to a plurality of rotatable elements, which can collectively define a rotor 51 .
  • the LP compressor 24 and the HP compressor 26 respectively include a plurality of compressor stages 52 , 54 , in which a set of compressor blades 56 , 58 rotate relative to a corresponding set of static compressor vanes 60 , 62 (also called a nozzle) to compress or pressurize the stream of fluid passing through the stage.
  • a single compressor stage 52 , 54 multiple compressor blades 56 , 58 can be provided in a ring and can extend radially outwardly relative to the centerline 12 , from a blade platform to a blade tip, while the corresponding static compressor vanes 60 , 62 are positioned upstream of and adjacent to the rotating blades 56 , 58 . It is noted that the number of blades, vanes, and compressor stages shown in FIG. 1 were selected for illustrative purposes only, and that other numbers are possible.
  • the blades 56 , 58 for a stage of the compressor can be mounted to a disk 61 , which is mounted to the corresponding one of the HP and LP spools 48 , 50 , with each stage having its own disk 61 .
  • the vanes 60 , 62 for a stage of the compressor can be mounted to the core casing 46 in a circumferential arrangement.
  • the HP turbine 34 and the LP turbine 36 respectively include a plurality of turbine stages 64 , 66 , in which a set of turbine blades 68 , 70 are rotated relative to a corresponding set of static turbine vanes 72 , 74 (also called a nozzle) to extract energy from the stream of fluid passing through the stage.
  • a single turbine stage 64 , 66 multiple turbine blades 68 , 70 can be provided in a ring and can extend radially outwardly relative to the centerline 12 , from a blade platform to a blade tip, while the corresponding static turbine vanes 72 , 74 are positioned upstream of and adjacent to the rotating blades 68 , 70 .
  • the number of blades, vanes, and turbine stages shown in FIG. 1 were selected for illustrative purposes only, and that other numbers are possible.
  • the blades 68 , 70 for a stage of the turbine can be mounted to a disk 71 , which is mounted to the corresponding one of the HP and LP spools 48 , 50 , with each stage having a dedicated disk 71 .
  • the vanes 72 , 74 for a stage of the compressor can be mounted to the core casing 46 in a circumferential arrangement.
  • stator 63 the stationary portions of the engine 10 , such as the static vanes 60 , 62 , 72 , 74 among the compressor and turbine section 22 , 32 are also referred to individually or collectively as a stator 63 .
  • stator 63 can refer to the combination of non-rotating elements throughout the engine 10 .
  • the airflow exiting the fan section 18 is split such that a portion of the airflow is channeled into the LP compressor 24 , which then supplies pressurized airflow 76 to the HP compressor 26 , which further pressurizes the air.
  • the pressurized airflow 76 from the HP compressor 26 is mixed with fuel in the combustor 30 and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine 34 , which drives the HP compressor 26 .
  • the combustion gases are discharged into the LP turbine 36 , which extracts additional work to drive the LP compressor 24 , and the exhaust gas is ultimately discharged from the engine 10 via the exhaust section 38 .
  • the driving of the LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP compressor 24 .
  • a portion of the pressurized airflow 76 can be drawn from the compressor section 22 as bleed air 77 .
  • the bleed air 77 can be draw from the pressurized airflow 76 and provided to engine components requiring cooling.
  • the temperature of pressurized airflow 76 entering the combustor 30 is significantly increased. As such, cooling provided by the bleed air 77 is necessary for operating of such engine components in the heightened temperature environments.
  • a remaining portion of the airflow 78 bypasses the LP compressor 24 and engine core 44 and exits the engine assembly 10 through a stationary vane row, and more particularly an outlet guide vane assembly 80 , comprising a plurality of airfoil guide vanes 82 , at the fan exhaust side 84 . More specifically, a circumferential row of radially extending airfoil guide vanes 82 are utilized adjacent the fan section 18 to exert some directional control of the airflow 78 .
  • Some of the air supplied by the fan 20 can bypass the engine core 44 and be used for cooling of portions, especially hot portions, of the engine 10 , and/or used to cool or power other aspects of the aircraft.
  • the hot portions of the engine are normally downstream of the combustor 30 , especially the turbine section 32 , with the HP turbine 34 being the hottest portion as it is directly downstream of the combustion section 28 .
  • Other sources of cooling fluid can be, but are not limited to, fluid discharged from the LP compressor 24 or the HP compressor 26 .
  • FIG. 2 is a perspective view of an engine component in the form the turbine blades 68 of the engine 10 of FIG. 1 .
  • the turbine blade 68 is exemplary and that the engine component can include other components requiring cooling.
  • the turbine blade 68 includes a dovetail 90 and an airfoil 92 .
  • the airfoil 92 includes a tip 94 to a root 96 .
  • a span-wise direction can be defined between the tip 94 and the root 96 .
  • the dovetail 90 includes a platform 98 and one or more inlet passages 100 having an outlet 102 .
  • the dovetail 90 and platform 98 can be integral with the airfoil 92 adjoining at the root 96 .
  • the platform 98 helps to radially contain the turbine airflow driven by the airfoil 92 .
  • the dovetail 90 can be configured to mount to a turbine rotor disk 71 ( FIG. 1 ) to rotate the airfoil 92 about the engine centerline 10 .
  • the inlet passages 100 can be fed with a flow of air, such as bypass air 104 .
  • the bypass air 104 is provided to the airfoil 92 at the root 96 exhausting through the outlets 102 .
  • the dovetail 90 is shown in cross-section, such that the inlet passages 100 are housed within the body of the dovetail 90 .
  • the airfoil 92 shown in cross-section, has an interior 110 bounded by an outer wall 112 .
  • the outer wall 112 includes a concave-shaped pressure sidewall 114 and a convex-shaped suction sidewall 116 .
  • a leading edge 118 and a trailing edge 120 are defined at the junction between the pressure and suction sidewalls 114 , 116 , defining a chord-wise distance between the leading and trailing edges 118 , 120 .
  • the airfoil 92 when implemented as a rotating blade as compared to a stationary vane, rotates in a direction such that the pressure sidewall 114 follows the suction sidewall 116 . Thus, as shown in FIG. 3 , the airfoil 92 would rotate upward toward the top of the page.
  • One or more ribs 130 are included in the interior 110 .
  • the ribs 130 can extend between the pressure and suction sidewalls 114 , 116 to define internal chambers 132 .
  • the chambers 132 can be discrete compartments defined within the airfoil 92 .
  • the chambers 132 can be in fluid communication with one another, such as defining a serpentine flow path snaking through the airfoil 92 in the span-wise direction.
  • the ribs 130 and chambers 132 defined by the ribs 130 are exemplary and should not be construed as limiting.
  • the interior 110 or chambers 132 defined therein can also include a plurality of plenums, circuits, micro-circuits, near wall cooling systems, pin banks, or similar structures in non-limiting examples.
  • a protuberance 134 can be disposed on the outer wall 112 .
  • the protuberances 134 are discrete members, defining an increased thickness for the wall 112 .
  • the protuberance can be radiused, rounded, conical, frustoconical, bell-shaped, or non-linear. Additional examples of protuberances can include, but are not limited to, radiused, circular, oval, elliptical, spherical, ellipsoidal, or curvilinear.
  • the protuberances 134 can be integral to the outer wall 112 , or can be mounted thereto. In one non-limiting example, the protuberances 134 can be formed on the airfoil 92 using additive manufacturing. Any number of protuberances 134 can be included on the outer wall 112 or the ribs 130 and can be organized in any manner, such as a linear grouping in the span-wise or chord-wise direction, a pattern, or random placement.
  • the protuberances 134 can be formed on an interior wall of the airfoil 92 or an engine component. In one example, the protuberances 134 can be formed on the rib 130 . In other non-limiting examples, the protuberance 134 can be disposed on walls of cooling structures, such as micro-circuits, cooling mesh, plenums, pin banks, or other component structures requiring cooling.
  • a film hole 136 can extend through the protuberances 134 .
  • the film holes 136 can extend through the protuberances 134 on the outer wall 112 to provide a cooling film along the external surface of the outer wall 112 for cooling the airfoil 92 .
  • the film hole 136 can be a cooling hole such as a cross-over hole.
  • such a cooling hole can provide a flow of cooling fluid among internal cavities or chambers of the engine component, such as adjacent areas channels or a micro circuit.
  • the film hole 136 can be shaped to direct a flow of fluid entering the film hole 132 , passing through the film hole 132 , or exhausting from the film hole 132 .
  • Such shaping can include a converging, diverging, or metering section to direct the flow of fluid, in non-limiting examples.
  • the converging section can increase the flow velocity of the flow of fluid
  • the diverging section can decrease the flow velocity of the flow of fluid
  • the metering section can meter the flow of fluid passing through it.
  • Additional shaping can include an expansion section or a reduction section.
  • the expansion section can include an increasing cross-sectional area to form a diffusion section and the reduction section can include a decreasing cross-sectional area.
  • the shaping of the film hole 136 can include a non-linear film hole 136 .
  • Such a film hole 136 could include curved passages or follow the curvature of the protuberance 134 .
  • the protuberance 134 could be placed on any wall having opposing flows on opposing sides of the wall 118 . Additionally, the film hole 136 can pass through the protuberance to provide a flow between the opposing sides of the wall 118 .
  • the protuberances can be disposed an any wall, such as an internal or external wall, and can include a film hole to provide a flow of fluid through such a wall for providing a cooling film.
  • the airfoil 92 as illustrated is exemplary and non-limiting, and the protuberances 134 can have equal applicability in any other engine component utilizing film holes.
  • FIG. 4 illustrates one protuberance 134 having a film hole 136 extending through the protuberance 134 .
  • the protuberance can be non-rectilinear, including a non-linear surface extending from the first side 144 to the inlet 150 , with the inlet 150 being rounded, transitioning into the film hole 136 .
  • the protuberance is circular.
  • the protuberance can be oval, elliptical, spherical, ellipsoidal, or curvilinear.
  • the protuberance 134 can be symmetrical, being even about an axis parallel to the direction of a cooling fluid flow C.
  • the protuberance 134 is disposed on a wall 140 .
  • the wall 140 can be the outer wall 112 , for example, or any other component wall having a film hole 136 .
  • the wall 140 has a nominal thickness 142 , having a first side 144 and a second side 146 defining a consistent nominal thickness 142 between the sides 144 , 146 .
  • the protuberance 134 is a circular extension, extending from the first side 144 into a first fluid flow 148 .
  • the film hole 136 is disposed in the center of the circular protuberance 134 , having an inlet 150 and an outlet 152 .
  • a passage 154 couples the inlet 150 to the outlet 152 .
  • the nominal thickness 142 can be the thickness for the wall 140 defined as the distance between the first and second side 144 , 146 .
  • a nominal thickness 142 can be determined in many different ways.
  • the nominal thickness 142 for the wall can be a function of the thermal load on the wall 140 , the airfoil 92 , or the engine component.
  • the nominal thickness 142 can be a function of a vibratory force acting on the wall 140 , a pressure differential between on opposing sides of the wall 140 , or the manufacturer required load for the wall 140 during operation. It should be appreciated that the nominal thickness 142 can be determined by multiple methods, such that a minimum operational thickness for the particular wall 140 is determined.
  • the nominal thickness 142 is a minimal operational thickness of the wall 140 , being a function of the thermal load, vibratory force, pressure differentials, load requirements, or other similar method can be respective of minimal operation requirements to maintain safe operation of the engine and individual engine components.
  • the nominal thickness 142 for the wall 140 can reduce engine weight, increasing engine efficiency or performance.
  • the first side 144 can be a cool side or cool surface, adjacent to a first fluid flow 148 , such as a cooling fluid flow C.
  • the second side 146 can be a hot side, or a hot surface, adjacent to a second fluid flow 170 , such as a hot gas flow H.
  • the profile view of the protuberance 134 in FIG. 5 illustrates a height 160 of the protuberance 134 .
  • the height 160 is the maximum distance the protuberance 134 extends from the wall 140 .
  • the height 160 can be determined in multiple different ways.
  • the height 160 can be a function of the nominal thickness 142 . In one example, the height 160 can be at least 50% of the nominal thickness 142 .
  • the height 160 can be equal to the nominal thickness 142 , or greater. In yet another example, the height can be equal to at least 100% of the nominal thickness 142 . It should be appreciated that the height 160 can be anywhere from 5% of the nominal thickness 142 to 200% of the nominal thickness 142 or greater.
  • the height 160 can be a function of the film hole 136 .
  • a length 162 of the film hole 136 can be defined as the distance between the inlet 150 and the outlet 152 .
  • the height 160 can be a function of the length 162 , where a particular film hole 136 can require a particular length 162 to provide an effective flow of fluid.
  • the height 160 can be at least 50% of the length 162 .
  • FIG. 5 illustrates a linear film hole 136 , it should be understood that the film hole 136 need not be linear, and with such a film hole, the length 162 can be measured as the streamline distance between the inlet 150 and the outlet.
  • the film hole 136 as illustrated in FIG. 5 is a perpendicular film hole 136 . In the case of the non-linear film hole, the length 162 will increase. As such, it should be appreciated that in such a case, the height 160 can be at least 30% of the length 162 .
  • the height 160 can be determined as a function of a diameter 164 of the film hole 136 .
  • a particular diameter 164 for a film hole may be required by an engine component, in order to keep structural integrity of the engine component.
  • the diameter 164 can require a particular length 162 for the film hole 136 to maintain an effectiveness, defining a required length-to-diameter ratio (L/D) for the film hole 136 .
  • L/D length-to-diameter ratio
  • the diameter 164 , or the L/D ratio can be used to define the height 160 in order to provide sufficient film hole effectiveness.
  • a side profile view of the protuberance 134 illustrates one orientation of the film hole 136 .
  • the protuberance 134 can be conical, having a portion removed at the inlet 150 of the film hole 136 .
  • the protuberance 135 can have a conic profile, having the inlet 150 disposed on one of the sides of the protuberance.
  • the first cooling flow 148 can be a cooling fluid flow C.
  • a second fluid flow 170 adjacent to the second side 146 of the wall 140 , can be a hot gas flow H.
  • the film hole 136 can be angled in the direction of one of the first and second fluid flows 148 , 170 , or both.
  • the inlet 150 can be positioned upstream of the outlet 152 relative to the cooling fluid flow C.
  • the cooling fluid flow C can be provided through the film hole 136 to the second side 146 as a cooling film 172 to cool the engine component.
  • FIG. 7 illustrates another embodiment of the protuberance 134 having a rounded dimension with a frustoconical shape at the inlet 150 of the film hole 136 .
  • the film hole 136 includes the inlet 150 disposed downstream of the outlet 152 relative to the cooling fluid flow C.
  • Such an orientation can be advantageous for providing an effective film hole length as well as providing multiple directional capabilities for exhausting a fluid from the film hole 136 .
  • FIG. 8 illustrates another embodiment, having a rounded protuberance having the film hole 136 with an inlet 150 offset from the center of the protuberance 134 .
  • the protuberance 134 can be divided into an upstream side 174 and a downstream side 176 .
  • the inlet 150 can be disposed on the upstream side 174 .
  • Such an orientation can be advantageous for determining flow rate entering the film hole 136 .
  • it is contemplated that the inlet 150 can be disposed on the downstream side 176 , or any other position on the protuberance 134 .
  • the position and orientation of the film hole 136 of FIGS. 5-8 is exemplary.
  • the position of the inlet 150 , outlet 152 , and dimension of the passage 154 disposed therebetween can be adapted to control flow rates through the film hole 136 , or adapt the length 162 , diameter 164 , or length-to-diameter ratio for the film hole 136 to provide effective cooling through the film hole 136 .
  • the film hole 136 can be provided with inlet shaping or outlet shaping, to provide a more deterministic flow for a cooling fluid passing through the cooling flow. Such an example would be a diverging outlet which can provide a cooling fluid over a greater cross-sectional area of the engine component.
  • the protuberance 134 can have a height 160 dependent on portions of the engine component, such as the nominal thickness 142 , the length 162 , diameter 164 , or L/D ratio of the film hole 136 .
  • a recess 180 can be formed in the protuberance 134 .
  • the recess 180 can be machined as part of the protuberance 134 , such as during additive manufacturing, or can be removed from the protuberance 134 to form the recess 180 .
  • the recess 180 can be symmetrical, such as a hemispherical shape, while any shape is contemplated. In other non-limiting examples, the recess 180 can be a rectilinear shape, or an arcuate or radiused shape, or any combination thereof.
  • the film hole 136 can be disposed in the recess 180 , having the inlet 150 at least partially formed within the recess 180 .
  • the size or shape of the recess 180 can be used to control the flow rate of a flow of fluid provided to the film hole 136 , or to further reduce component weight in combination with the nominal thickness 142 for the wall 140 .
  • the protuberance 134 can be asymmetrical, having an elongated or offset shape. Such as shape may be desirable to optimize fluid flows within the engine component or for directing a flow toward one or more film holes 136 . Additionally the film hole 136 can be offset from the direction of the first fluid flow 148 along the engine component. The discrete direction of the first fluid flow 148 at the protuberance 134 can be transposed on the protuberance 134 as a transposed axis 190 .
  • the passage 154 shown in a linear example while non-linear film holes are contemplated, can define a passage axis 192 .
  • a film hole angle 194 can be defined between the transposed axis 190 and the passage axis 192 to define the offset relationship of the film hole 136 to the first fluid flow 148 . Additionally, the film hole 136 can be angled relative to a local normal between the first and second sides 144 , 146 .
  • the offset orientation of the film hole 136 or the protuberance 134 can be discrete, relative to an adjacent flow of fluid which can change direction or magnitude at different portions of the engine component.
  • a plurality of protuberances 134 along the engine component can be aligned or patterned, while some of the protuberances 134 or film holes 136 are offset from the direction of the first fluid flow 148 at a portion of the engine component.
  • a method of cooling an engine component can include a cool surface, such as the first side 144 .
  • the method can include passing a cooling fluid flow C along the cool surface 144 and providing at least a portion of the cooling fluid flow C through a film hole, such as the film hole 136 , in a protuberance 134 extending from the cool surface 144 .
  • Providing at least a portion of the cooling fluid flow C can include providing a portion of the cooling fluid flow through a recess 180 in the protuberance 134 prior to providing the cool fluid flow C to the film hole 136 . Additionally, providing a portion of the cooling fluid flow C through the recess 180 can minimize dust accumulation at the film hole 136 or along the cool surface 144 of the engine component.
  • the airfoil 92 or other engine component requiring cooling can utilizing the film hole 136 such as a film hole disposed within the protuberance 134 .
  • the protuberance 134 provides for an increased thickness permitting an increased film hole length 162 to provide effective cooling through the film hole 136 .
  • the use of a protuberance 134 permits the remaining portions of the engine component to have a nominal thickness 142 , which reduces component weight, reducing overall engine weight. A reduced weight provides for better engine efficiency.
  • the protuberances 134 are discrete, having no greater an area than necessary to provide for the casting, drilling, or otherwise forming the film holes 136 through the protuberances 134 in order to have an increased length, diameter, or L/D ratio for the film hole 136 which would otherwise be unachievable within the nominal thickness 142 of the engine component, due to the nominal thickness 142 to manufacturing capabilities of the engine component at the nominal thickness 142 .
  • the protuberances 134 are radiused, reducing drag or resistance caused by the extension of the protuberance 134 into the flow of fluid, such as the cooling fluid flow C, adjacent the protuberance 134 . Further still, the radiused protuberances 134 or recesses 180 therein can provide for reduced dust accumulation, increasing component lifetime or reducing required maintenance.
  • airfoil, engine components, protuberances, or film holes described herein can be formed by additive manufacturing. Such manufacturing can be used to develop the intricate details of the aforementioned, such as specific film hole shaping without the poor yields of such manufacturing as casting, or the imperfections associated with other manufacturing methods such as film hole drilling.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US15/138,624 2016-04-26 2016-04-26 Airfoil for a turbine engine Abandoned US20170306764A1 (en)

Priority Applications (7)

Application Number Priority Date Filing Date Title
US15/138,624 US20170306764A1 (en) 2016-04-26 2016-04-26 Airfoil for a turbine engine
JP2017078572A JP2017198205A (ja) 2016-04-26 2017-04-12 タービンエンジン用のエーロフォイル
CA2964139A CA2964139A1 (en) 2016-04-26 2017-04-13 Airfoil for a turbine engine
EP17167906.1A EP3239462A1 (en) 2016-04-26 2017-04-25 Airfoil for a turbine engine
CN201710281295.8A CN107448300B (zh) 2016-04-26 2017-04-26 用于涡轮发动机的翼型件
US16/869,831 US20200277862A1 (en) 2016-04-26 2020-05-08 Airfoil for a turbine engine
US18/337,841 US20240159151A1 (en) 2016-04-26 2023-06-20 Airfoil for a turbine engine

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US15/138,624 US20170306764A1 (en) 2016-04-26 2016-04-26 Airfoil for a turbine engine

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US16/869,831 Continuation US20200277862A1 (en) 2016-04-26 2020-05-08 Airfoil for a turbine engine

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US15/138,624 Abandoned US20170306764A1 (en) 2016-04-26 2016-04-26 Airfoil for a turbine engine
US16/869,831 Abandoned US20200277862A1 (en) 2016-04-26 2020-05-08 Airfoil for a turbine engine
US18/337,841 Pending US20240159151A1 (en) 2016-04-26 2023-06-20 Airfoil for a turbine engine

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US18/337,841 Pending US20240159151A1 (en) 2016-04-26 2023-06-20 Airfoil for a turbine engine

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US20190186272A1 (en) * 2017-12-18 2019-06-20 General Electric Company Engine component with cooling hole
US20190218940A1 (en) * 2018-01-17 2019-07-18 United Technologies Corporation Dirt separator for internally cooled components
US20200011186A1 (en) * 2016-09-15 2020-01-09 Honeywell International Inc. Gas turbine component with cooling aperture having shaped inlet and method of forming the same
US10626733B2 (en) 2017-10-03 2020-04-21 United Technologies Corporation Airfoil having internal hybrid cooling cavities
US10626734B2 (en) 2017-10-03 2020-04-21 United Technologies Corporation Airfoil having internal hybrid cooling cavities
US10633980B2 (en) 2017-10-03 2020-04-28 United Technologies Coproration Airfoil having internal hybrid cooling cavities
US20210285336A1 (en) * 2020-03-11 2021-09-16 United Technologies Corporation Investment casting core bumper for gas turbine engine article
US11306659B2 (en) * 2019-05-28 2022-04-19 Honeywell International Inc. Plug resistant effusion holes for gas turbine engine
CN114876582A (zh) * 2022-06-28 2022-08-09 西北工业大学 一种涡轮叶片及航空发动机
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US11208900B2 (en) * 2016-09-15 2021-12-28 Honeywell International Inc. Gas turbine component with cooling aperture having shaped inlet and method of forming the same
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US10626733B2 (en) 2017-10-03 2020-04-21 United Technologies Corporation Airfoil having internal hybrid cooling cavities
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EP3597859B1 (en) * 2018-07-13 2023-08-30 Honeywell International Inc. Turbine blade with dust tolerant cooling system
US11306659B2 (en) * 2019-05-28 2022-04-19 Honeywell International Inc. Plug resistant effusion holes for gas turbine engine
US20210285336A1 (en) * 2020-03-11 2021-09-16 United Technologies Corporation Investment casting core bumper for gas turbine engine article
US11242768B2 (en) * 2020-03-11 2022-02-08 Raytheon Technologies Corporation Investment casting core bumper for gas turbine engine article
CN114876582A (zh) * 2022-06-28 2022-08-09 西北工业大学 一种涡轮叶片及航空发动机

Also Published As

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CN107448300B (zh) 2019-12-24
EP3239462A1 (en) 2017-11-01
US20240159151A1 (en) 2024-05-16
US20200277862A1 (en) 2020-09-03
JP2017198205A (ja) 2017-11-02
CN107448300A (zh) 2017-12-08
CA2964139A1 (en) 2017-10-26

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