EP3163028A1 - Compressor apparatus - Google Patents

Compressor apparatus Download PDF

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Publication number
EP3163028A1
EP3163028A1 EP16195207.2A EP16195207A EP3163028A1 EP 3163028 A1 EP3163028 A1 EP 3163028A1 EP 16195207 A EP16195207 A EP 16195207A EP 3163028 A1 EP3163028 A1 EP 3163028A1
Authority
EP
European Patent Office
Prior art keywords
splitter
dimension
blades
stator
chord
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.)
Withdrawn
Application number
EP16195207.2A
Other languages
German (de)
English (en)
French (fr)
Inventor
Anthony Louis DIPIETRO JR.
Gregory John Kajfasz
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
Publication of EP3163028A1 publication Critical patent/EP3163028A1/en
Withdrawn legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/541—Specially adapted for elastic fluid pumps
    • F04D29/542—Bladed diffusers
    • F04D29/544—Blade shapes
    • 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
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F04D29/324—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
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/146—Shape, i.e. outer, aerodynamic form of blades with tandem configuration, split blades or slotted 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
    • F01D9/00—Stators
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D19/00—Axial-flow pumps
    • 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
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/142—Shape, i.e. outer, aerodynamic form of the blades of successive rotor or stator blade-rows
    • F01D5/143—Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour

Definitions

  • This invention relates generally to turbomachinery compressors and more particularly relates to rotor and stator airfoils of such compressors.
  • a gas turbine engine includes, in serial flow communication, a compressor, a combustor, and turbine.
  • the turbine is mechanically coupled to the compressor and the three components define a turbomachinery core.
  • the core is operable in a known manner to generate a flow of hot, pressurized combustion gases to operate the engine as well as perform useful work such as providing propulsive thrust or mechanical work.
  • One common type of compressor is an axial-flow compressor with multiple stages each including a rotating disk with a row of axial-flow airfoils, referred to as rotor blades.
  • this type of compressor also includes stationary airfoils alternating with the rotor airfoils, referred to as stator vanes.
  • the stator vanes are typically bounded at their inner and outer ends by arcuate endwall structures (e.g. a hub or a case).
  • thermodynamic cycle efficiency it is generally desirable to incorporate a compressor having the highest possible pressure ratio (that is, the ratio of inlet pressure to outlet pressure). It is also desirable to include the fewest number of compressor stages. However, there are well-known inter-related aerodynamic limits to the maximum pressure ratio and mass flow possible through a given compressor stage.
  • stator vane row including stator vane airfoils and splitter airfoils.
  • FIG. 1 illustrates a portion of a gas turbine engine, generally designated 10.
  • the engine 10 has a longitudinal centerline axis 11 and includes, in axial flow sequence, a fan 12, a low-pressure compressor or "booster” 14, and a high-pressure compressor ("HPC") 16.
  • HPC high-pressure compressor
  • the fan 12 and booster 14 are driven by a low-pressure turbine ("LPT") which is not illustrated in FIG. 1 , via an inner shaft 18.
  • LPT low-pressure turbine
  • the rotating fan 12 operates to generate a pressurized fan flow of air, some of which enters the booster 14 and the HPC 16, and some of which is discharged through a bypass duct 20.
  • the rotating booster 14 supercharges the flow into the HPC 16.
  • HPC 16 is driven by a high-pressure turbine ("HPT") which is not illustrated in FIG.1 , via an outer shaft 22.
  • HPC operates to generate a core flow which passes through a core of the engine 10.
  • turbofan engine While the illustrated example is a high-bypass turbofan engine, the principles of the present invention are equally applicable to other types of engines such as low-bypass turbofans, turbojets, and turboshafts.
  • forward or “front” refer to a location relatively upstream in an air flow passing through or around a component
  • aft or “rear” refer to a location relatively downstream in an air flow passing through or around a component. The direction of this flow is shown by the arrow “F” in FIG. 1 .
  • the HPC 16 is configured for axial fluid flow, that is, fluid flow generally parallel to the centerline axis 11. This is in contrast to a centrifugal compressor or mixed-flow compressor.
  • the HPC 16 includes a number of stages, each of which includes a rotor comprising a row of airfoils or blades mounted to a rotating disk, and row of stationary airfoils or vanes. The vanes serve to turn the airflow exiting an upstream row of blades before it enters the downstream row of blades.
  • FIGS. 2-6 illustrate a portion of a rotor 38 constructed according to a first exemplary embodiment of the present invention and suitable for inclusion in the HPC 16.
  • the rotor 38 may be incorporated into one or more of the stages in the aft half of the HPC 16, particularly the last or aft-most stage.
  • the rotor 38 includes a disk 40 with a web 42 and a rim 44. It will be understood that the complete disk 40 is an annular structure mounted for rotation about the centerline axis 11.
  • the rim 44 has a forward end 46 and an aft end 48.
  • An annular flowpath surface 50 extends between the forward and aft ends 46, 48.
  • Each compressor blade extends from a root 54 at the flowpath surface 50 to a tip 56, and includes a concave pressure side 58 joined to a convex suction side 60 at a leading edge 62 and a trailing edge 64.
  • each compressor blade 52 has a span (or span dimension) "S1" defined as the radial distance from the root 54 to the tip 56, and a chord (or chord dimension) "C1" defined as the length of an imaginary straight line connecting the leading edge 62 and the trailing edge 64.
  • its chord C1 may be different at different locations along the span S1.
  • the relevant measurement is the chord C1 at the root 54.
  • the flowpath surface 50 is not a body of revolution. Rather, the flowpath surface 50 has a non-axisymmetric surface profile. As an example of a non-axisymmetric surface profile, it may be contoured with a concave curve or "scallop" 66 between each adjacent pair of compressor blades 52.
  • the dashed lines in FIG. 4 illustrate a hypothetical cylindrical surface with a radius passing through the roots 54 of the compressor blades 52. It can be seen that the flowpath surface curvature has its maximum radius (or minimum radial depth of the scallop 66) at the compressor blade roots 54, and has its minimum radius (or maximum radial depth "d" of the scallop 66) at a position approximately midway between adjacent compressor blades 52.
  • this scalloped configuration is effective to reduce the magnitude of mechanical and thermal hoop stress concentration at the airfoil hub intersections on the rim 44 along the flowpath surface 50. This contributes to the goal of achieving acceptably-long component life of the disk 40.
  • An aerodynamically adverse side effect of scalloping the flowpath 50 is to increase the rotor passage flow area between adjacent compressor blades 52. This increase in rotor passage through flow area increases the aerodynamic loading level and in turn tends to cause undesirable flow separation on the suction side 60 of the compressor blade 52, at the inboard portion near the root 54, and at an aft location, for example approximately 75% of the chord distance C1 from the leading edge 62.
  • An array of splitter blades 152 extend from the flowpath surface 50.
  • One splitter blade 152 is disposed between each pair of compressor blades 52. In the circumferential direction, the splitter blades 152 may be located halfway or circumferentially biased between two adjacent compressor blades 52, or circumferentially aligned with the deepest portion d of the scallop 66. Stated another way, the compressor blades 52 and splitter blades 152 alternate around the periphery of the flowpath surface 50.
  • Each splitter blade 152 extends from a root 154 at the flowpath surface 50 to a tip 156, and includes a concave pressure side 158 joined to a convex suction side 160 at a leading edge 162 and a trailing edge 164. As best seen in FIG.
  • each splitter blade 152 has a span (or span dimension) "S2" defined as the radial distance from the root 154 to the tip 156, and a chord (or chord dimension) "C2" defined as the length of an imaginary straight line connecting the leading edge 162 and the trailing edge 164.
  • S2 span (or span dimension)
  • C2 chord (or chord dimension)
  • its chord C2 may be different at different locations along the span S2.
  • the relevant measurement is the chord C2 at the root 154.
  • the splitter blades 152 function to locally increase the hub solidity of the rotor 38 and thereby prevent the above-mentioned flow separation from the compressor blades 52.
  • a similar effect could be obtained by simply increasing the number of compressor blades 152, and therefore reducing the blade-to-blade spacing. This, however, has the undesirable side effect of increasing aerodynamic surface area frictional losses which would manifest as reduced aerodynamic efficiency and increased rotor weight. Therefore, the dimensions of the splitter blades 152 and their position may be selected to prevent flow separation while minimizing their surface area.
  • the splitter blades 152 are positioned so that their trailing edges 164 are at approximately the same axial position as the trailing edges of the compressor blades 52, relative to the rim 44. This can be seen in FIG. 3 .
  • the span S2 and/or the chord C2 of the splitter blades 152 may be some fraction less than unity of the corresponding span S1 and chord C1 of the compressor blades 52. These may be referred to as "part-span” and/or “part-chord” splitter blades.
  • the span S2 may be equal to or less than the span S1.
  • the span S2 is 50% or less of the span S 1. More preferably for the least frictional losses, the span S2 is 30% or less of the span S1.
  • the chord C2 may be equal to or less than the chord C1.
  • the chord C2 is 50% or less of the chord C1.
  • the disk 40, compressor blades 52, and splitter blades 152 may be constructed from any material capable of withstanding the anticipated stresses and environmental conditions in operation.
  • suitable alloys include iron, nickel, and titanium alloys.
  • FIGS. 2-6 the disk 40, compressor blades 52, and splitter blades 152 are depicted as an integral, unitary, or monolithic whole. This type of structure may be referred to as a "bladed disk” or "blisk".
  • the principles of the present invention are equally applicable to a rotor built up from separate components (not shown).
  • FIGS. 7-11 illustrate a portion of a rotor 238 constructed according to a second exemplary embodiment of the present invention and suitable for inclusion in the HPC 16.
  • the rotor 238 may be incorporated into one or more of the stages in the aft half of the HPC 16, particularly the last or aft-most stage.
  • the rotor 238 includes a disk 240 with a web 242 and a rim 244. It will be understood that the complete disk 240 is an annular structure mounted for rotation about the centerline axis 11.
  • the rim 244 has a forward end 246 and an aft end 248.
  • An annular flowpath surface 250 extends between the forward and aft ends 246, 248.
  • Each compressor blade 252 extends from a root 254 at the flowpath surface 250 to a tip 256, and includes a concave pressure side 258 joined to a convex suction side 260 at a leading edge 262 and a trailing edge 264.
  • each compressor blade 252 has a span (or span dimension) "S3" defined as the radial distance from the root 254 to the tip 256, and a chord (or chord dimension) "C3" defined as the length of an imaginary straight line connecting the leading edge 262 and the trailing edge 264.
  • its chord C3 may be different at different locations along the span S3.
  • the relevant measurement is the chord C3 at the root 254.
  • the compressor blades 252 are uniformly spaced apart around the periphery of the flowpath surface 250.
  • a nondimensional parameter called “solidity” is defined as c/s, where "c” is equal to the blade chord as described above.
  • the compressor blades 252 may have a spacing which is significantly greater than a spacing that would be expected in the prior art, resulting in a blade solidity significantly less than would be expected in the prior art.
  • the flowpath surface 250 is depicted as a body of revolution (i.e. axisymmetric).
  • the flowpath surface 250 may have a non-axisymmetric surface profile as described above for the flowpath surface 50.
  • the reduced blade solidity will have the effect of reducing weight, improving rotor performance, and simplify manufacturing by minimizing the total number of compressor airfoils used in a given rotor stage.
  • An aerodynamically adverse side effect of reduced blade solidity is to increase the rotor passage flow area between adjacent compressor blades 252. This increase in rotor passage through flow area increases the aerodynamic loading level and in turn tends to cause undesirable flow separation on the suction side 260 of the compressor blade 252, at the inboard portion near the root 254, and at an aft location, for example approximately 75% of the chord distance C3 from the leading edge 262, also referred to as "hub flow separation".
  • the compressor blade spacing may be intentionally selected to produce a solidity low enough to result in hub flow separation under expected operating conditions.
  • An array of splitter blades 352 extend from the flowpath surface 250.
  • One splitter blade 352 is disposed between each pair of compressor blades 252. In the circumferential direction, the splitter blades 352 may be located halfway or circumferentially biased between two adjacent compressor blades 252. Stated another way, the compressor blades 252 and splitter blades 352 alternate around the periphery of the flowpath surface 250.
  • Each splitter blade 352 extends from a root 354 at the flowpath surface 250 to a tip 356, and includes a concave pressure side 358 joined to a convex suction side 360 at a leading edge 362 and a trailing edge 364. As best seen in FIG.
  • each splitter blade 352 has a span (or span dimension) "S4" defined as the radial distance from the root 354 to the tip 356, and a chord (or chord dimension) "C4" defined as the length of an imaginary straight line connecting the leading edge 362 and the trailing edge 364.
  • S4 span (or span dimension)
  • C4 chord (or chord dimension)
  • its chord C4 may be different at different locations along the span S4.
  • the relevant measurement is the chord C4 at the root 354.
  • the splitter blades 352 function to locally increase the hub solidity of the rotor 238 and thereby prevent the above-mentioned flow separation from the compressor blades 252.
  • a similar effect could be obtained by simply increasing the number of compressor blades 252, and therefore reducing the blade-to-blade spacing. This, however, has the undesirable side effect of increasing aerodynamic surface area frictional losses which would manifest as reduced aerodynamic efficiency and increased rotor weight. Therefore, the dimensions of the splitter blades 352 and their position may be selected to prevent flow separation while minimizing their surface area.
  • the splitter blades 352 are positioned so that their trailing edges 364 are at approximately the same axial position as the trailing edges 264 of the compressor blades 252, relative to the rim 244. This can be seen in FIG. 8 .
  • the span S4 and/or the chord C4 of the splitter blades 352 may be some fraction less than unity of the corresponding span S3 and chord C3 of the compressor blades 252. These may be referred to as "part-span” and/or "part-chord” splitter blades.
  • the span S4 may be equal to or less than the span S3.
  • the span S4 is 50% or less of the span S3.
  • the span S4 is 30% or less of the span S3.
  • the chord C4 may be equal to or less than the chord C3.
  • the chord C4 is 50% or less of the chord C3.
  • the disk 240, compressor blades 252, and splitter blades 352 using the same materials and structural configuration (e.g. monolithic or separable) as the disk 40, compressor blades 52, and splitter blades 152 described above.
  • the bypass duct 20 includes an array of airflow-shaped fan outlet guide vanes (“OGVs") 400 bounded at inboard and outboard ends by a core cowl 402 and a fan cowl 404, respectively.
  • the booster 14 includes several rows of airflow-shaped booster guide vanes 406 bounded at inboard and outboard ends, respectively by an inner band 408 and a casing 410.
  • the HPC 16 includes several rows of airflow-shaped compressor stator vanes 452 bounded by an inner band 444 and a casing 470, respectively.
  • the OGVs 400, booster guide vanes 406, and compressor stator vanes 452 may all be considered to be “stator airfoils”.
  • FIGS. 12 and 13 illustrate a portion of one row of the compressor stator vanes 452.
  • the illustrated stator vanes 452 may be incorporated into one or more of the stages of the HPC 16.
  • the stator vanes 452 constitute "stator airfoils" which are conceptually representative of any stator structure, and the principles described herein could optionally be incorporated into the outlet guide vanes 400 and/or the booster guide vanes 406.
  • the inner band 444 defines an annular inner flowpath surface 450 extending between forward and aft ends 446, 448.
  • the casing 470 defines an annular outer flowpath surface 472 extending between forward and aft ends 474, 476.
  • the stator vanes 452 extend between the inner and outer flowpath surfaces 450, 472.
  • Each stator vane 452 extends from a root 454 at the inner flowpath surface 450 to a tip 456 at the outer flowpath surface 472, and includes a concave pressure side 458 joined to a convex suction side 460 at a leading edge 462 and a trailing edge 464.
  • each stator vane 452 has a span (or span dimension) "S5" defined as the radial distance from the root 454 to the tip 456, and a chord (or chord dimension) "C5" defined as the length of an imaginary straight line connecting the leading edge 462 and the trailing edge 464.
  • its chord C5 may be different at different locations along the span S5.
  • the relevant measurement would be the chord C5 at the root 454 or tip 456.
  • the stator vanes 452 are uniformly spaced apart around the periphery of the inner flowpath surface 450.
  • the stator vanes 452 have a mean circumferential spacing "s", defined as described above (see FIG. 13 ).
  • a nondimensional parameter called “solidity” is defined as c/s, where "c” is equal to the vane chord as described above.
  • the stator vanes 452 may have a spacing which is significantly greater than a spacing that would be expected in the prior art, resulting in a vane solidity significantly less than would be expected in the prior art.
  • the inner and outer flowpath surfaces 450, 472 are depicted as bodies of revolution (i.e. axisymmetric structures).
  • either or both of the inner or outer flowpath surfaces 450, 472 may have a non-axisymmetric surface profile as described above for the flowpath surface 50.
  • the reduced vane solidity will have the effect of reducing weight, improving stator performance, and simplify manufacturing by minimizing the total number of airfoils used in a given stator stage.
  • An aerodynamically adverse side effect of reduced stator solidity is to increase the rotor passage flow area between adjacent stator vanes 452. This increase in stator passage through flow area increases the aerodynamic loading level and in turn tends to cause undesirable flow separation on the suction side 460 of the stator vane 452, at the inboard portion near the root 454, and at an aft location, for example approximately 75% of the chord distance C5 from the leading edge 462, also referred to as "hub flow separation".
  • stator vane spacing may be intentionally selected to produce a solidity low enough to result in endwall separation under expected operating conditions.
  • one or both of the inner and outer flowpath surfaces 450, 472 may be provided with an array of splitter vanes.
  • an array of splitter vanes 552 extend radially inward from the outer flowpath surface 472.
  • One splitter vane 552 is disposed between each pair of stator vanes 452. In the circumferential direction, the splitter vanes 552 may be located halfway or circumferentially biased between two adjacent stator vanes 452. Stated another way, the stator vanes 452 and splitter vanes 552 alternate around the periphery of the outer flowpath surface 472.
  • Each splitter vane 552 extends from a root 554 at the outer flowpath surface 472 to a tip 556, and includes a concave pressure side 558 joined to a convex suction side 560 at a leading edge 562 and a trailing edge 564.
  • each splitter vane 552 has a span (or span dimension) "S6" defined as the radial distance from the root 554 to the tip 556, and a chord (or chord dimension) "C6” defined as the length of an imaginary straight line connecting the leading edge 562 and the trailing edge 564.
  • its chord C6 may be different at different locations along the span S6.
  • the relevant measurement is the chord C6 at the root 554.
  • the splitter vanes 552 function to locally increase the hub solidity of the stator and thereby prevent the above-mentioned flow separation from the stator vanes 452.
  • a similar effect could be obtained by simply increasing the number of stator vanes 452, and therefore reducing the vane-to-vane spacing. This, however, has the undesirable side effect of increasing aerodynamic surface area frictional losses which would manifest as reduced aerodynamic efficiency and increased stator weight. Therefore, the dimensions of the splitter vanes 552 and their position may be selected to prevent flow separation while minimizing their surface area.
  • the splitter vanes 552 are positioned so that their trailing edges 564 are at approximately the same axial position as the trailing edges 464 of the stator vanes 452, relative to the outer flowpath surface 472. This can be seen in FIG.
  • the span S6 and/or the chord C6 of the splitter vanes 552 may be some fraction less than unity of the corresponding span S5 and chord C5 of the stator vanes 452. These may be referred to as "part-span” and/or "part-chord” splitter vanes.
  • the span S6 may be equal to or less than the span S5.
  • the span S6 is 50% or less of the span S5.
  • the span S6 is 30% or less of the span S5.
  • the chord C6 may be equal to or less than the chord C5.
  • the chord C6 is 50% or less of the chord C5.
  • FIG. 13 illustrates an array of splitter vanes 652 extending radially outward from the inner flowpath surface 450.
  • One splitter vane 652 is disposed between each pair of stator vanes 652.
  • the splitter vanes 652 may be identical to the splitter vanes 552 described above, in terms of their shape, circumferential position relative to the stator vanes 452, and their span and chord dimensions.
  • splitter vanes may optionally be incorporated at the inner flowpath surface, or the outer flowpath surface 472, or both.
  • the compressor apparatus described herein with splitter blades and/or splitter vanes increases the endwall solidity level locally, reduces the endwall aerodynamic loading level locally, and suppresses the tendency of the airfoil portion adjacent the endwall to want to separate in the presence of the non-axisymmetric contoured endwall flowpath surface, or with a reduced airfoil count on an axisymmetric flowpath.
  • the use of a partial-span and/or partial-chord splitter blade or vane is effective to keep the solidity levels of the middle and upper sections of the airfoil unchanged from a nominal value, and therefore to maintain middle and upper airfoil section performance.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Geometry (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP16195207.2A 2015-10-26 2016-10-24 Compressor apparatus Withdrawn EP3163028A1 (en)

Applications Claiming Priority (1)

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US14/923,005 US20170114796A1 (en) 2015-10-26 2015-10-26 Compressor incorporating splitters

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EP3163028A1 true EP3163028A1 (en) 2017-05-03

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US (1) US20170114796A1 (ja)
EP (1) EP3163028A1 (ja)
JP (1) JP2017082784A (ja)
CN (1) CN107035435B (ja)
BR (1) BR102016024900A2 (ja)
CA (1) CA2945103A1 (ja)

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EP3358138B1 (en) * 2017-02-07 2021-06-23 Doosan Heavy Industries & Construction Co., Ltd. Pre-swirler for gas turbine
US11480063B1 (en) * 2021-09-27 2022-10-25 General Electric Company Gas turbine engine with inlet pre-swirl features
US11959393B2 (en) 2021-02-02 2024-04-16 General Electric Company Turbine engine with reduced cross flow airfoils

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US10473118B2 (en) * 2014-08-29 2019-11-12 Siemens Aktiengesellschaft Controlled convergence compressor flowpath for a gas turbine engine
CN107687447A (zh) * 2017-09-29 2018-02-13 广东威灵电机制造有限公司 扩压器
TWI678471B (zh) * 2018-08-02 2019-12-01 宏碁股份有限公司 散熱風扇
DE102019213932B4 (de) 2019-09-12 2022-05-19 Deutsches Zentrum für Luft- und Raumfahrt e.V. Verdichterstufe mit variabler Statorschaufelneigung
US11401824B2 (en) * 2019-10-15 2022-08-02 General Electric Company Gas turbine engine outlet guide vane assembly
US11242770B2 (en) 2020-04-02 2022-02-08 General Electric Company Turbine center frame and method

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US20170114796A1 (en) 2017-04-27
CN107035435A (zh) 2017-08-11
CA2945103A1 (en) 2017-04-26

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