Classifications

• • 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
  • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
• • 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
  • F05D2250/00—Geometry
  • F05D2250/10—Two-dimensional
  • F05D2250/18—Two-dimensional patterned
  • F05D2250/183—Two-dimensional patterned zigzag
• • 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
  • F05D2260/00—Function
  • F05D2260/20—Heat transfer, e.g. cooling
  • F05D2260/221—Improvement of heat transfer
  • F05D2260/2212—Improvement of heat transfer by creating turbulence
• • 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
• • 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
  • F05D2270/00—Control
  • F05D2270/01—Purpose of the control system
  • F05D2270/17—Purpose of the control system to control boundary layer
• • 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

• This disclosure relates to a gas turbine engine. More particularly, the disclosure relates to a serpentine cooling passage that may be incorporated into a gas turbine engine component, such as an airfoil.
• Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.
• Both the compressor and turbine sections may include alternating series of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine.
• turbine blades rotate and extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine.
• the turbine vanes which generally do not rotate, guide the airflow and prepare it for the next set of blades.
• cooling passages having turns that provide a serpentine shape, which create undesired pressure losses.
• Some of the cooling passages may include portions having turbulence promoters that enhance the cooling effects of the cooling flow through the cooling passage.
• a gas turbine engine component includes a structure having a cooling passage providing upstream and downstream portions separated from one another by an inner wall and fluidly connected by a bend.
• the downstream portion includes an outer wall opposite the inner wall to provide a downstream region extending between the inner and outer walls.
• a turbulence promoter extends from the outer wall adjacent to the bend in the downstream portion. The turbulence promoter is absent from a stagnation region adjoining the inner wall adjacent to the bend in the downstream portion.
• the downstream region includes a downstream width defined by the inner and outer walls.
• the downstream width corresponds to the sum of a stagnation width and a vena contracta width.
• the stagnation region provides the stagnation width and the vena contracta region provides the turbulence promoter.
• the bend is greater than 90°.
• the bend is between 135° and
• the bend is 180°.
• the turbulence promoter is provided by chevron- shaped trip strips.
• the chevron-shape is provided by multiple legs meeting at an apex.
• the apex points in the direction of incoming flow through the passage.
• the turbulence promoter corresponds to a first turbulence promoter, and includes a second turbulence promoter provided in the downstream region that extends from the inner wall to the outer wall downstream from the first turbulence promoter.
• the turbulence promoter corresponds to multiple turbulators comprising pins.
• the gas turbine engine component includes an airfoil that includes pressure and suction walls spaced apart from one another and are joined at leading and trailing edges.
• the airfoil includes the cooling passage arranged between the pressure and suction walls.
• the cooling passage extends in the radial direction from a root supporting the airfoil toward a tip of the airfoil.
• a gas turbine engine in another exemplary embodiment, includes a compressor section and a turbine section.
• a combustor section is arranged between the compressor and turbine sections.
• One of the compressor and turbine sections includes an airfoil.
• the airfoil includes pressure and suction walls spaced apart from one another and are joined at leading and trailing edges extending in a radial direction.
• the airfoil has a cooling passage arranged between the pressure and suction walls that extend toward a tip of the airfoil.
• the cooling passage provides upstream and downstream portions separated from one another by an inner wall and fluidly connected by a bend.
• the downstream portion includes an outer wall opposite the inner wall to provide a downstream region that extends between the inner and outer walls.
• a turbulence promoter extends from the outer wall adjacent to the bend in the downstream portion, but the turbulence promoter is absent from a stagnation region adjoining the inner wall adjacent to the bend in the downstream portion.
• the airfoil provides a turbine blade.
• the downstream region includes a downstream width defined by the inner and outer walls.
• the downstream width corresponds to the sum of a stagnation width and a vena contracta width.
• the stagnation region provides the stagnation width and the vena contracta region provides the turbulence promoter.
• the bend is greater than 90°.
• the bend is 180°.
• the turbulence promoter is provided by chevron- shaped trip strips.
• the turbulence promoter corresponds to multiple turbulators comprising pins.
• Figure 1 schematically illustrates a gas turbine engine embodiment.
• Figure 2A is a perspective view of the airfoil having the disclosed cooling passage.
• Figure 2B is a plan view of the airfoil illustrating directional references.
• Figure 3 is a schematic cross-sectional view of fluid flow through a cooling passage having a bend, which creates a stagnation region.
• Figure 4 is one example cross-sectional view of the cooling passage illustrated in Figure 3 with turbulence promoters configured to minimize the stagnation region.
• Figure 5 is another example cross-sectional view of the cooling passage illustrated in Figure 3 with turbulence promoters configured to minimize the stagnation region.
• FIG. 1 schematically illustrates an example gas turbine engine 20 that includes a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
• Alternative engines might include an augmenter section (not shown) among other systems or features.
• the fan section 22 drives air along a bypass flow path B while the compressor section 24 draws air in along a core flow path C where air is compressed and communicated to a combustor section 26.
• air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section 28 where energy is extracted and utilized to drive the fan section 22 and the compressor section 24.
• turbofan gas turbine engine depicts a turbofan gas turbine engine
• the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.
• the example 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.
• the low speed spool 30 generally includes an inner shaft 40 that connects a fan 42 and a low pressure (or first) compressor section 44 to a low pressure (or first) turbine section 46.
• the inner shaft 40 drives the fan 42 through a speed change device, such 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 (or second) compressor section 52 and a high pressure (or second) turbine section 54.
• the inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the engine central longitudinal axis X.
• a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54.
• the high pressure turbine 54 includes at least two stages to provide a double stage high pressure turbine 54.
• the high pressure turbine 54 includes only a single stage.
• a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure” compressor or turbine.
• the example low pressure turbine 46 has a pressure ratio that is greater than about five (5).
• the pressure ratio of the example low pressure turbine 46 is measured prior to an inlet of the low pressure turbine 46 as related to the pressure measured at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
• 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 as well as setting airflow entering the low pressure turbine 46.
• the core airflow C is compressed by the low pressure compressor 44 then by the high pressure compressor 52 mixed with fuel and ignited in the combustor 56 to produce high speed exhaust gases that are then expanded through the high pressure turbine 54 and low pressure turbine 46.
• the mid-turbine frame 57 includes vanes 59, which are in the core airflow path and function as an inlet guide vane for the low pressure turbine 46. Utilizing the vane 59 of the mid- turbine frame 57 as the inlet guide vane for low pressure turbine 46 decreases the length of the low pressure turbine 46 without increasing the axial length of the mid-turbine frame 57. Reducing or eliminating the number of vanes in the low pressure turbine 46 shortens the axial length of the turbine section 28. Thus, the compactness of the gas turbine engine 20 is increased and a higher power density may be achieved.
• the disclosed gas turbine engine 20 in one example is a high-bypass geared aircraft engine.
• the gas turbine engine 20 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10).
• the example geared architecture 48 is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3.
• the gas turbine engine 20 includes a bypass ratio greater than about ten (10: 1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor 44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.
• 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.
• the flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (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 1.50. In another non- limiting embodiment the low fan pressure ratio 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)] ° '5 .
• the "Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second.
• the disclosed serpentine cooling passage may be used in various gas turbine engine components. For exemplary purposes, a turbine blade 64 is described. It should be understood that the cooling passage may also be used in vanes, blade outer air seals, and turbine platforms, for example.
• each turbine blade 64 is mounted to the rotor disk.
• the turbine blade 64 includes a platform 76, which provides the inner flow path, supported by the root 74.
• An airfoil 78 extends in a radial direction R from the platform 76 to a tip 80.
• the turbine blades may be integrally formed with the rotor such that the roots are eliminated.
• the platform is provided by the outer diameter of the rotor.
• the airfoil 78 provides leading and trailing edges 82, 84.
• the tip 80 is arranged adjacent to a blade outer air seal (not shown).
• the airfoil 78 of Figure 2B somewhat schematically illustrates exterior airfoil surface extending in a chord-wise direction C from a leading edge 82 to a trailing edge 84.
• the airfoil 78 is provided between pressure (typically concave) and suction (typically convex) wall 86, 88 in an airfoil thickness direction T, which is generally perpendicular to the chord-wise direction C.
• Multiple turbine blades 64 are arranged circumferentially in a circumferential direction A.
• the airfoil 78 extends from the platform 76 in the radial direction R, or spanwise, to the tip 80.
• the airfoil 78 includes a cooling passage 90 provided between the pressure and suction walls 86, 88.
• the exterior airfoil surface may include multiple film cooling holes (not shown) in fluid communication with the cooling passage 90.
• the cooling passage 90 includes an upstream portion 92 and a downstream portion 94 fluidly connected by a bend 96.
• a common inner wall 98 separates the upstream and downstream portions 92, 94.
• the upstream portion 92 includes an upstream region 100 providing an upstream width (indicated by arrow) through which cooling flow F passes before reaching the bend 96.
• flow F through the bend 96 creates a stagnation region 104 within a downstream region 102 of the downstream portion 94 as the flow F is forced to turn.
• the stagnation region 104 is where the velocity is closest to zero. Typically, the sharper the bend 96, the larger the stagnation region 104 having a stagnation width (indicated by arrow).
• the flow F through the upstream portion 92 has a fairly uniform velocity. As the flow F enters the bend 96, the inside fluid stream lines accelerate faster, but cannot abruptly change directions. As a result, the flow F accelerates and onto wall 108 and creates a stagnation region at the inner wall 98 near the bend 96 in the downstream portion 94.
• the downstream flow is not required to turn as sharply and therefore does not stagnate to the same extent.
• the downstream region 102 has a downstream width (indicated by arrow) provided between the inner and outer walls 98, 108.
• the downstream width corresponds to the sum of the width of the stagnation region 104 and the outer flow region, or vena contracta 106.
• the vena contracta is the narrowest flow width where the fluid velocity of the flow F is at its maximum.
• a coefficient of contraction corresponds to the ratio of the stagnation width 104 relative to the vena contracta 106.
• the separation region 104 extends for about x/Dh of -1.6, wherein "x" is axial distance in flow direction and "Dh" is hydraulic diameter of the passage. This may vary depending on geometry and flow velocity.
• trip strips 112 are used in both the upstream and downstream portions 92, 94 of the passage 190.
• the trip strips 112 are arranged in rows 110 in a repeating chevron shape.
• the trip strips 112 are distributed uniformly throughout the passage 90 so that the chevrons extend across the entire width of the upstream and downstream portions for the length of the passage.
• the bend 196 is greater than 90° and, for example, between 135° and 225°. In the example shown, the bend 196 is 180°.
• the trip strips 112 are provided in the upstream and downstream portions 192, 194 of the passage 190.
• the trip strips 112 include angled legs joined to one another at an apex 116.
• the apex 116 is positioned pointing in the direction of incoming flow F.
• the legs 114 of adjoining chevrons may also be joined.
• trip strips 112 are omitted near the inner wall 198 and adjacent to the bend 196 to provide an open area 118.
• downstream trip strips in the downstream portion extend a greater distance across the downstream width than the trip strips adjacent to the bend.
• the flow F at the downstream portion 194 increases the pressure near the outer wall 108, which forces more air to the inner wall, reducing the contraction region and the extent of the separation bubble 104.
• Turbulators 126 are provided in the upstream and downstream portions 292, 294 of the passage 290.
• the turbulators 126 are provided by pins of any suitable cross-sectional shape and may be arranged in rows 210.
• Some of the turbulators 126 are omitted near the inner wall 298 and adjacent to the bend 296 to provide an open area 218.
• the turbulators 126 near the outer wall 308 adjacent to the bend 296 remain.
• the flow F at the downstream portion 294 increases the pressure near the outer wall 208, which reduces the width of the stagnation region (indicated by the dashed lines).

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• Engineering & Computer Science (AREA)
• Mechanical Engineering (AREA)
• General Engineering & Computer Science (AREA)
• Turbine Rotor Nozzle Sealing (AREA)

Applications Claiming Priority (2)

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• 2012
  • 2012-09-14 US US13/618,178 patent/US20140219813A1/en not_active Abandoned
• 2013
  • 2013-09-05 EP EP13836460.9A patent/EP2895694A4/de not_active Withdrawn
  • 2013-09-05 WO PCT/US2013/058282 patent/WO2014042955A1/en active Application Filing

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