US20170248155A1 - Centrifugal compressor diffuser passage boundary layer control - Google Patents
Centrifugal compressor diffuser passage boundary layer control Download PDFInfo
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- US20170248155A1 US20170248155A1 US15/517,262 US201515517262A US2017248155A1 US 20170248155 A1 US20170248155 A1 US 20170248155A1 US 201515517262 A US201515517262 A US 201515517262A US 2017248155 A1 US2017248155 A1 US 2017248155A1
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- diffuser
- bleed
- boundary layer
- flow
- centrifugal compressor
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- 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
- F04D29/42—Casings; Connections of working fluid for radial or helico-centrifugal pumps
- F04D29/44—Fluid-guiding means, e.g. diffusers
- F04D29/441—Fluid-guiding means, e.g. diffusers especially adapted for elastic fluid pumps
- F04D29/444—Bladed diffusers
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- 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
- F01D9/045—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector for radial flow machines or engines
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- 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
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
- F02C3/08—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising at least one radial stage
-
- 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
- F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas- turbine plants for special use
- F02C6/04—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output
- F02C6/06—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas
- F02C6/08—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
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- 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
- F02C7/16—Cooling of plants characterised by cooling medium
- F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D17/00—Radial-flow pumps, e.g. centrifugal pumps; Helico-centrifugal pumps
- F04D17/08—Centrifugal pumps
- F04D17/10—Centrifugal pumps for compressing or evacuating
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
- F04D27/009—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids by bleeding, by passing or recycling fluid
-
- 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
- F04D29/28—Rotors specially for elastic fluids for centrifugal or helico-centrifugal pumps for radial-flow or helico-centrifugal pumps
- F04D29/284—Rotors specially for elastic fluids for centrifugal or helico-centrifugal pumps for radial-flow or helico-centrifugal pumps for compressors
-
- 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
- F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
- F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
- F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
- F04D29/682—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps by fluid extraction
-
- 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
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
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- 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
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- 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
- F05D2240/10—Stators
- F05D2240/12—Fluid guiding means, e.g. vanes
-
- 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
- F05D2250/30—Arrangement of components
- F05D2250/32—Arrangement of components according to their shape
- F05D2250/324—Arrangement of components according to their shape divergent
-
- 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
- F05D2270/10—Purpose of the control system to cope with, or avoid, compressor flow instabilities
- F05D2270/101—Compressor surge or stall
-
- 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
Definitions
- the present invention relates to bleed air from gas turbine engine centrifugal compressors.
- One type of gas turbine engine includes a centrifugal compressor having a rotatable impeller to accelerate and, thereby, increase the kinetic energy of air flowing therethrough.
- a diffuser is generally located immediately downstream of and surrounding the impeller. The diffuser operates to decrease the velocity of the air flow leaving the impeller and transform the energy thereof to an increase in static pressure, thus, pressurizing the air.
- a conventional gas turbine engine typically includes a compressor, combustor, and turbine, both rotating turbine components such as blades, disks and retainers, and stationary turbine components, such as vanes, shrouds, and frames routinely require cooling due to heating thereof by hot combustion gases. Cooling of the turbine, especially the rotating components, is important to the proper function and safe operation of the engine. It is known to bleed cooling air from the centrifugal compressor to help cool the turbine.
- the impact on engine operation may range from relatively benign blade tip distress, resulting in a reduction in engine power and useable blade life, to a rupture of a turbine disk, resulting in an unscheduled engine shutdown.
- turbine cooling air is typically drawn from one or more stages of the compressor and channelled by various means, such as pipes, ducts, and internal passageways to the desired components, such air is not available to be mixed with fuel, ignited in the combustor and undergo work extraction in the primary gas flowpath of the turbine.
- Total cooling flow bled from the compressor is a loss in the engine operating cycle and it is desirable to keep such losses to a minimum.
- Some conventional engines employ clean air bleed systems to cool turbine components in gas turbines using an axi-centrifugal compressor as is done in the General Electric CFE738 engine.
- the turbine cooling supply air exits the centrifugal diffuser through a small gap between the diffuser exit and deswirler inner shroud.
- Other turbine cooling air methods include extracting cooling from the impeller or from a gap between the impeller and the diffuser exit.
- U.S. Pat. No. 5,555,7211 to Bourneuf, et al which issued on Sep. 17, 1996 and is entitled AGas Turbine Engine Cooling Supply Circuit@, discloses using bleed air from an impeller stage of a centrifugal compressor in a turbine cooling supply circuit for a gas turbine.
- U.S. Pat. No. 5,555,721 discloses impeller tip forward bleed flow and impeller tip aft bleed flow for cooling turbine components.
- U.S. Pat. No. 5,555,721 is assigned to the General Electric Co., the same assignee as this patent and is incorporated herein by reference.
- U.S. Pat. No. 8,087,249 to Ottaviano, et al. which issued Jan. 3, 2012, and is entitled ATurbine Cooling Air From A Centrifugal Compressor@ discloses a gas turbine engine turbine cooling system including an impeller and a diffuser directly downstream of the impeller and a bleed for bleeding clean cooling air from downstream of the diffuser.
- U.S. Pat. No. 8,087,249 is assigned to the General Electric Co., the same assignee as this patent and is incorporated herein by reference.
- a diffuser for a centrifugal compressor includes an annular diffuser housing, diffuser vanes axially extending between a forward wall and an aft wall of the diffuser housing, a plurality of diffuser flow passages extending through the housing and spaced about a circumference of the housing.
- the diffuser flow passages are bounded by the diffuser vanes and the forward and aft walls.
- a diffuser boundary layer bleed is provided for bleeding diffuser bleed flow from a diffuser boundary layer in each of the diffuser flow passages.
- the diffuser boundary layer bleed may be configured for bleeding the diffuser bleed flow from the diffuser boundary layer at a position located in a region of flow weakness in each of the diffuser flow passages.
- the diffuser boundary layer bleed may include boundary layer bleed apertures disposed through the forward wall.
- Each of the boundary layer bleed apertures may be a slot including a downstream facing wall angled or canted at an acute cant angle with respect to a downstream diffuser airflow direction in each of the diffuser flow passages respectively.
- the boundary layer bleed apertures may be positioned or located downstream of throat sections of the diffuser flow passages near pressure sides of the diffuser vanes.
- a centrifugal compressor including an annular centrifugal compressor impeller, a diffuser annularly surrounding the impeller, and a plurality of diffuser flow passages extending through a housing of the diffuser and spaced about a circumference of the housing.
- Each of the passages includes a throat section and a diffusing section downstream of the throat section.
- the diffuser flow passages are circumferentially bounded by diffuser vanes extending axially between forward and aft walls of the diffuser and a diffuser boundary layer bleed is provided for bleeding diffuser bleed flow from a diffuser boundary layer in each of the diffuser flow passages.
- the centrifugal compressor may also include a radial clearance between an impeller tip of the impeller and an annular inlet of the diffuser, a means for mixing impeller bleed flow from the radial clearance with diffuser bleed flow from the boundary layer bleed apertures for providing turbine cooling air, and a means for flowing the turbine cooling air to a turbine or a means for flowing impeller bleed flow and the diffuser bleed flow separately to the turbine.
- FIG. 1 is a sectional view illustration of a gas turbine engine with a centrifugal compressor for mixing impeller tip bleed flow and diffuser bleed flow in the compressor section before using the flows for cooling turbine components.
- FIG. 2 is an enlarged sectional view illustration of the centrifugal compressor and a diffuser with diffuser bleed holes illustrated in FIG. 1 .
- FIG. 3 is an aft looking forward perspective view illustration of the diffuser and the diffuser bleed holes through 3 - 3 in FIG. 2 .
- FIG. 4 is an enlarged perspective view illustration of the bleed holes illustrated in FIG. 3 .
- FIG. 5 is a perspective view illustration of a portion of the diffuser and the diffuser bleed holes illustrated in FIG. 2 .
- FIG. 6 is an enlarged sectional view illustration of the centrifugal compressor tip and the diffuser bleed holes illustrated in FIG. 2 .
- FIG. 7 is a sectional view illustration of a gas turbine engine centrifugal compressor with an alternative arrangement for separately flowing impeller tip bleed for cooling turbine components.
- FIG. 8 is a sectional view illustration of the gas turbine engine illustrated in FIG. 7 with an arrangement for separately flowing diffuser bleed flow for cooling turbine components.
- FIG. 9 is an enlarged perspective view illustration of one of the impeller bleed flow ports illustrated in FIG. 7 and as taken through 9 - 9 in FIG. 10 .
- FIG. 10 is a forward looking aft perspective view illustration of an aft casing surrounding the centrifugal compressor and including the impeller and bleed flow ports illustrated in FIGS. 7 and 8 respectively.
- FIG. 11 is cutaway perspective view illustration of impeller bleed flowpaths for one of the impeller bleed flow ports illustrated in FIGS. 7 and 9 .
- FIG. 12 is an enlarged perspective view illustration of one of the diffuser bleed flow ports illustrated in FIG. 8 and as taken through 12 - 12 in FIG. 10 .
- FIG. 13 is cutaway perspective view illustration of a diffuser bleed flowpath through one of the diffuser bleed flow ports illustrated in FIG. 8 and as taken through 12 - 12 in FIG. 10 .
- FIG. 1 Illustrated in FIG. 1 is a gas turbine engine high pressure centrifugal compressor 18 in a high pressure gas generator 10 of a gas turbine engine 8 .
- the high pressure centrifugal compressor 18 is a final compressor stage of a high pressure compressor 4 .
- the high pressure gas generator 10 has a high pressure rotor 12 including, in downstream serial or flow relationship, the high pressure compressor 14 , a combustor 52 , and a high pressure turbine 16 .
- the rotor 12 is rotatably supported about an engine axis 25 by bearings in engine frames not illustrated herein.
- the exemplary embodiment of the high pressure compressor 14 illustrated herein includes a five stage axial compressor 30 followed by the centrifugal compressor 18 having an annular centrifugal compressor impeller 32 .
- Outlet guide vanes 40 are disposed between the five stage axial compressor 30 and the single stage centrifugal compressor 18 .
- Compressor discharge pressure (CDP) air 76 exits the impeller 32 and passes through a diffuser 42 annularly surrounding the impeller 32 and then through a deswirl cascade 44 into a combustion chamber 45 within the combustor 52 .
- the combustion chamber 45 is surrounded by annular radially outer and inner combustor casings 46 , 47 .
- Air 76 is conventionally mixed with fuel provided by a plurality of fuel nozzles 48 and ignited and combusted in an annular combustion zone 50 bounded by annular radially outer and inner combustion liners 72 , 73 .
- the high pressure turbine 16 includes, in downstream serial flow relationship, first and second high pressure turbine stages 55 , 56 having first and second stage disks 60 , 62 .
- a high pressure shaft 64 of the high pressure rotor 12 connects the high pressure turbine 16 in rotational driving engagement to the impeller 32 .
- a first stage nozzle 66 is directly upstream of the first high pressure turbine stage 55 and a second stage nozzle 68 is directly upstream of the second high pressure turbine stage.
- the compressor discharge pressure (CDP) air 76 is discharged from the impeller 32 of the centrifugal compressor 18 , used to combust fuel in the combustor 52 , and to cool components of turbine 16 subjected to the hot combustion gases 54 ; such as, the first stage nozzle 66 , first and second stage shrouds 71 , 69 surrounding the first and second high pressure turbine stages 55 , 56 respectively.
- the high pressure compressor 14 includes a compressor aft casing 110 and a diffuser forward casing 114 as more fully illustrated in FIGS. 1 and 2 .
- the compressor aft casing 110 generally surrounds the axial compressor 30 and the diffuser forward casing 114 generally surrounds the centrifugal compressor 18 and supports the diffuser 42 directly downstream of the centrifugal compressor 18 .
- the compressor discharge pressure (CDP) air 76 is discharged from the impeller 32 of the centrifugal compressor 18 directly into the diffuser 42 .
- the impeller 32 includes a plurality of centrifugal compressor blades 84 radially extending from a rotor disc portion 82 .
- an annular blade tip shroud 90 Opposite and axially forward of the compressor blades 84 is an annular blade tip shroud 90 .
- the shroud 90 is adjacent to blade tips 86 of the compressor blades 84 defining a blade tip clearance 80 therebetween.
- the diffuser 42 includes an annular diffuser housing 20 having a plurality of tangentially disposed diffuser flow passages 22 extending radially therethrough, spaced about a circumference 26 of the housing 20 , and through which diffuser airflow 103 flows in a downstream direction.
- Diffuser vanes 23 axially extend between a forward wall 101 and the aft wall 100 of the diffuser 42 .
- the diffuser vanes 23 circumferentially extend between adjacent ones of the diffuser flow passages 22 .
- the diffuser flow passages 22 are partly defined and circumferentially bounded by the circumferentially spaced apart diffuser vanes 23 .
- Adjacent ones of the passages 22 intersect with each other at radially inner inlet sections 24 of the passages 22 that define a quasi-vaneless annular inlet 27 of the diffuser 42 .
- Each passage 22 further includes a throat section 28 downstream of and integral with the inner inlet section 24 .
- Each passage 22 further includes a diffusing section 99 immediately downstream of the throat section 28 .
- a centrifugal compressor first cooling air source 92 for turbine cooling air 88 is a small predetermined radial clearance (C) located between an impeller tip 36 of the rotating impeller 32 and the annular inlet 27 of the static diffuser 42 .
- Impeller bleed flow 102 from the radial clearance (C) is collected in a radially inner manifold 104 .
- the predetermined. radial clearance (C) is designed to accommodate thermal and mechanical growth of the impeller 32 and is open to or in fluid communication with the radially inner manifold 104 .
- the diffuser airflow 103 on one side of the passage (such as passage 22 ) in multi-passage diffusers (such as the diffuser 42 ) that follow or are downstream of centrifugal impellers (such as the impeller 32 ) is often weak and may be subject to separation. Separation in the passage can generate high losses that lowers engine specific fuel consumption (SFC). This area or region of weak flow 127 is also believed to be a contributor to surge that limits flow range of the compressor.
- SFC engine specific fuel consumption
- a centrifugal compressor stage second cooling air source 94 for turbine cooling air 88 includes a diffuser boundary layer bleed 96 for bleeding diffuser bleed flow 112 from a diffuser boundary layer 113 in each of the diffuser flow passages 22 of the diffuser 42 , illustrated herein as plurality of boundary layer bleed apertures 106 .
- the diffuser boundary layer bleed 96 also referred to as fluidic bleed, helps reduce the weak flow and limit or prevent the unwanted flow separation.
- the diffuser boundary layer bleed 96 bleeds diffuser bleed flow 112 from the diffuser boundary layer 113 into a radially outer manifold 116 .
- the radially inner and outer manifolds 104 , 116 are in fluid communication such that the impeller bleed flow 102 from the radially inner manifold 104 flows into the radially outer manifold 116 .
- the impeller and diffuser bleed flows 102 , 112 are mixed in the radially outer manifold 116 to provide the turbine cooling air 88 which is then ported or otherwise flowed from radially outer manifold 116 through a plurality of circumferentially distributed manifold ports 117 to the high pressure turbine 16 .
- the turbine cooling air 88 may be channelled or flowed therefrom by external piping (not shown) to cool the first and second stage shrouds 71 , 69 (illustrated in FIG. 1 ).
- Substantially axially extending beams or struts 122 separate the radially inner and outer manifolds 104 , 116 and the impeller bleed flow 102 passes between the struts 122 as it flows from the radially inner manifold 104 into the radially outer manifold 116 .
- the fluidic bleed flow illustrated herein as the diffuser boundary layer bleed 96 represents a small amount of flow, less than 1% of the engine core flow. The fluidic bleed is strategically removed near the inception of the weak flow to improve the overall performance of the diffuser.
- the boundary layer bleed apertures 106 may be holes or slots 130 through the forward wall 101 of the diffuser 42 as illustrated herein.
- the boundary layer bleed apertures 106 or slots 130 lead into and are in flow communication with the radially outer manifold 116 .
- the slot 130 is positioned or located downstream of the throat section 28 near a pressure side 126 of the diffuser vane 23 at a position where the flow would begin to show weakness or instability in a diffuser without the diffuser boundary layer bleed 96 . This position is located in what is referred to as a region of flow weakness 127 .
- a slot width W may be sized with manufacturing constraints such as a minimum tool size.
- a slot length L may be selected to enable up to 3% of the engine core flow to be used.
- the slot 130 should ideally be angled such that the diffuser bleed flow 112 exits the slot perpendicular to a forward surface 105 of the forward wall 101 of the diffuser 42 in a radial plane 132 passing through the engine centerline or axis 25 as illustrated in FIG. 5 .
- the slot 130 has radially outer and inner walls 136 , 138 , as illustrated in FIG. 6 , and upstream and downstream facing walls 140 , 142 , as illustrated in FIGS. 4 and 5 respectively, extending through the forward wall 101 .
- the downstream facing wall 142 is designed to scoop boundary layer air 144 in the diffuser boundary layer 113 only.
- the downstream facing wall 142 is angled or canted at an acute cant angle B of less than 90 degrees with respect to the diffuser airflow 103 (parallel to the direction boundary layer air 144 in the downstream direction in the diffuser flow passages 22 of the diffuser 42 .
- an acute cant angle B of 45 degrees is desirable.
- the acute cant angle B is limited by geometry and manufacturing constraints on the outside of the diffuser so that an acute cant angle, for example about 22.5 degrees, is more practical.
- FIGS. 7-13 Illustrated in FIGS. 7-13 is a gas turbine engine with a centrifugal compressor similar to the one illustrated in FIGS. 1-3 but with an alternative arrangement or design for separately gathering and flowing the impeller tip bleed and diffuser bleed flow for cooling turbine components.
- the impeller bleed flow 102 front the radial clearance (C), illustrated in FIG. 9 is flowed into and collected in a radially inner annular manifold 154 illustrated in FIGS. 7 and 9 .
- Inter-manifold apertures 160 are disposed between the inner annular manifold 154 and a plurality of radially outer annular manifolds 156 illustrated in FIGS. 7, 9, and 13 .
- the inter-manifold apertures 160 allow the impeller bleed flow 102 to flow front the inner annular manifold 154 into the outer annular manifolds 156 .
- the impeller bleed flow 102 from the outer annular manifolds 156 is then ported or otherwise flowed through a plurality of circumferentially distributed impeller bleed flow manifold ports 157 , illustrated in FIG. 10 , to the high pressure turbine 16 for turbine cooling.
- the diffuser boundary layer bleed 96 bleeds diffuser bleed flow 112 from the diffuser boundary layer 113 into an annular diffuser bleed manifold 158 from where the diffuser bleed flow 112 is then ported or otherwise flowed through a plurality of circumferentially distributed diffuser bleed manifold ports 159 to the high pressure turbine 16 for turbine cooling.
- FIG. 10 illustrates the relative circumferential and axial locations of the impeller bleed flow manifold ports 157 and the diffuser bleed manifold ports 159 on and through the diffuser forward casing 114 .
Abstract
Description
- This invention was made with government support under government contract No. W911-W6-11-2-0009 by the Department of Defense. The government has certain rights to this invention.
- The present invention relates to bleed air from gas turbine engine centrifugal compressors.
- One type of gas turbine engine includes a centrifugal compressor having a rotatable impeller to accelerate and, thereby, increase the kinetic energy of air flowing therethrough. A diffuser is generally located immediately downstream of and surrounding the impeller. The diffuser operates to decrease the velocity of the air flow leaving the impeller and transform the energy thereof to an increase in static pressure, thus, pressurizing the air.
- A conventional gas turbine engine typically includes a compressor, combustor, and turbine, both rotating turbine components such as blades, disks and retainers, and stationary turbine components, such as vanes, shrouds, and frames routinely require cooling due to heating thereof by hot combustion gases. Cooling of the turbine, especially the rotating components, is important to the proper function and safe operation of the engine. It is known to bleed cooling air from the centrifugal compressor to help cool the turbine.
- Failure to adequately cool a turbine disk and its blading, for example, by providing cooling air deficient in supply pressure, volumetric flow rate or temperature margin, may be detrimental to the life and mechanical integrity of the turbine. Depending on the nature and extent of the cooling deficiency, the impact on engine operation may range from relatively benign blade tip distress, resulting in a reduction in engine power and useable blade life, to a rupture of a turbine disk, resulting in an unscheduled engine shutdown.
- Balanced with the need to adequately cool the turbine is the desire for higher levels of engine operating efficiency which translate into lower fuel consumption and lower operating costs. Since turbine cooling air is typically drawn from one or more stages of the compressor and channelled by various means, such as pipes, ducts, and internal passageways to the desired components, such air is not available to be mixed with fuel, ignited in the combustor and undergo work extraction in the primary gas flowpath of the turbine.
- Total cooling flow bled from the compressor is a loss in the engine operating cycle and it is desirable to keep such losses to a minimum.
- Some conventional engines employ clean air bleed systems to cool turbine components in gas turbines using an axi-centrifugal compressor as is done in the General Electric CFE738 engine. The turbine cooling supply air exits the centrifugal diffuser through a small gap between the diffuser exit and deswirler inner shroud. Other turbine cooling air methods include extracting cooling from the impeller or from a gap between the impeller and the diffuser exit.
- U.S. Pat. No. 5,555,7211 to Bourneuf, et al, which issued on Sep. 17, 1996 and is entitled AGas Turbine Engine Cooling Supply Circuit@, discloses using bleed air from an impeller stage of a centrifugal compressor in a turbine cooling supply circuit for a gas turbine. U.S. Pat. No. 5,555,721 discloses impeller tip forward bleed flow and impeller tip aft bleed flow for cooling turbine components. U.S. Pat. No. 5,555,721 is assigned to the General Electric Co., the same assignee as this patent and is incorporated herein by reference.
- U.S. Pat. No. 8,087,249 to Ottaviano, et al. which issued Jan. 3, 2012, and is entitled ATurbine Cooling Air From A Centrifugal Compressor@ discloses a gas turbine engine turbine cooling system including an impeller and a diffuser directly downstream of the impeller and a bleed for bleeding clean cooling air from downstream of the diffuser. U.S. Pat. No. 8,087,249 is assigned to the General Electric Co., the same assignee as this patent and is incorporated herein by reference.
- Thus, there continues to be a demand for advancements in diffuser design and geometry that improves aerodynamic performance and reduces the overall engine radial envelope.
- A diffuser for a centrifugal compressor includes an annular diffuser housing, diffuser vanes axially extending between a forward wall and an aft wall of the diffuser housing, a plurality of diffuser flow passages extending through the housing and spaced about a circumference of the housing. The diffuser flow passages are bounded by the diffuser vanes and the forward and aft walls. A diffuser boundary layer bleed is provided for bleeding diffuser bleed flow from a diffuser boundary layer in each of the diffuser flow passages.
- The diffuser boundary layer bleed may be configured for bleeding the diffuser bleed flow from the diffuser boundary layer at a position located in a region of flow weakness in each of the diffuser flow passages.
- The diffuser boundary layer bleed may include boundary layer bleed apertures disposed through the forward wall. Each of the boundary layer bleed apertures may be a slot including a downstream facing wall angled or canted at an acute cant angle with respect to a downstream diffuser airflow direction in each of the diffuser flow passages respectively.
- The boundary layer bleed apertures may be positioned or located downstream of throat sections of the diffuser flow passages near pressure sides of the diffuser vanes.
- A centrifugal compressor including an annular centrifugal compressor impeller, a diffuser annularly surrounding the impeller, and a plurality of diffuser flow passages extending through a housing of the diffuser and spaced about a circumference of the housing. Each of the passages includes a throat section and a diffusing section downstream of the throat section. The diffuser flow passages are circumferentially bounded by diffuser vanes extending axially between forward and aft walls of the diffuser and a diffuser boundary layer bleed is provided for bleeding diffuser bleed flow from a diffuser boundary layer in each of the diffuser flow passages.
- The centrifugal compressor may also include a radial clearance between an impeller tip of the impeller and an annular inlet of the diffuser, a means for mixing impeller bleed flow from the radial clearance with diffuser bleed flow from the boundary layer bleed apertures for providing turbine cooling air, and a means for flowing the turbine cooling air to a turbine or a means for flowing impeller bleed flow and the diffuser bleed flow separately to the turbine.
-
FIG. 1 is a sectional view illustration of a gas turbine engine with a centrifugal compressor for mixing impeller tip bleed flow and diffuser bleed flow in the compressor section before using the flows for cooling turbine components. -
FIG. 2 is an enlarged sectional view illustration of the centrifugal compressor and a diffuser with diffuser bleed holes illustrated inFIG. 1 . -
FIG. 3 is an aft looking forward perspective view illustration of the diffuser and the diffuser bleed holes through 3-3 inFIG. 2 . -
FIG. 4 is an enlarged perspective view illustration of the bleed holes illustrated inFIG. 3 . -
FIG. 5 is a perspective view illustration of a portion of the diffuser and the diffuser bleed holes illustrated inFIG. 2 . -
FIG. 6 is an enlarged sectional view illustration of the centrifugal compressor tip and the diffuser bleed holes illustrated inFIG. 2 . -
FIG. 7 is a sectional view illustration of a gas turbine engine centrifugal compressor with an alternative arrangement for separately flowing impeller tip bleed for cooling turbine components. -
FIG. 8 is a sectional view illustration of the gas turbine engine illustrated inFIG. 7 with an arrangement for separately flowing diffuser bleed flow for cooling turbine components. -
FIG. 9 is an enlarged perspective view illustration of one of the impeller bleed flow ports illustrated inFIG. 7 and as taken through 9-9 inFIG. 10 . -
FIG. 10 is a forward looking aft perspective view illustration of an aft casing surrounding the centrifugal compressor and including the impeller and bleed flow ports illustrated inFIGS. 7 and 8 respectively. -
FIG. 11 is cutaway perspective view illustration of impeller bleed flowpaths for one of the impeller bleed flow ports illustrated inFIGS. 7 and 9 . -
FIG. 12 is an enlarged perspective view illustration of one of the diffuser bleed flow ports illustrated inFIG. 8 and as taken through 12-12 inFIG. 10 . -
FIG. 13 is cutaway perspective view illustration of a diffuser bleed flowpath through one of the diffuser bleed flow ports illustrated inFIG. 8 and as taken through 12-12 inFIG. 10 . - Illustrated in
FIG. 1 is a gas turbine engine high pressurecentrifugal compressor 18 in a highpressure gas generator 10 of agas turbine engine 8. The high pressurecentrifugal compressor 18 is a final compressor stage of a high pressure compressor 4. The highpressure gas generator 10 has ahigh pressure rotor 12 including, in downstream serial or flow relationship, the high pressure compressor 14, acombustor 52, and ahigh pressure turbine 16. Therotor 12 is rotatably supported about anengine axis 25 by bearings in engine frames not illustrated herein. - The exemplary embodiment of the high pressure compressor 14 illustrated herein includes a five stage
axial compressor 30 followed by thecentrifugal compressor 18 having an annularcentrifugal compressor impeller 32.Outlet guide vanes 40 are disposed between the five stageaxial compressor 30 and the single stagecentrifugal compressor 18. Compressor discharge pressure (CDP)air 76 exits theimpeller 32 and passes through adiffuser 42 annularly surrounding theimpeller 32 and then through adeswirl cascade 44 into acombustion chamber 45 within thecombustor 52. Thecombustion chamber 45 is surrounded by annular radially outer andinner combustor casings Air 76 is conventionally mixed with fuel provided by a plurality offuel nozzles 48 and ignited and combusted in anannular combustion zone 50 bounded by annular radially outer andinner combustion liners - The combustion produces
hot combustion gases 54 which flow through thehigh pressure turbine 16 causing rotation of thehigh pressure rotor 12 and continue downstream for further work extraction in alow pressure turbine 78 and final exhaust as is conventionally known. In the exemplary embodiment depicted herein, thehigh pressure turbine 16 includes, in downstream serial flow relationship, first and second high pressure turbine stages 55, 56 having first andsecond stage disks high pressure shaft 64 of thehigh pressure rotor 12 connects thehigh pressure turbine 16 in rotational driving engagement to theimpeller 32. Afirst stage nozzle 66 is directly upstream of the first highpressure turbine stage 55 and asecond stage nozzle 68 is directly upstream of the second high pressure turbine stage. - Referring to
FIG. 1 , the compressor discharge pressure (CDP)air 76 is discharged from theimpeller 32 of thecentrifugal compressor 18, used to combust fuel in thecombustor 52, and to cool components ofturbine 16 subjected to thehot combustion gases 54; such as, thefirst stage nozzle 66, first and second stage shrouds 71, 69 surrounding the first and second high pressure turbine stages 55, 56 respectively. The high pressure compressor 14 includes a compressor aftcasing 110 and a diffuser forward casing 114 as more fully illustrated inFIGS. 1 and 2 . The compressor aftcasing 110 generally surrounds theaxial compressor 30 and the diffuser forward casing 114 generally surrounds thecentrifugal compressor 18 and supports thediffuser 42 directly downstream of thecentrifugal compressor 18. The compressor discharge pressure (CDP)air 76 is discharged from theimpeller 32 of thecentrifugal compressor 18 directly into thediffuser 42. - Referring to
FIGS. 2 and 3 , theimpeller 32 includes a plurality ofcentrifugal compressor blades 84 radially extending from arotor disc portion 82. Opposite and axially forward of thecompressor blades 84 is an annularblade tip shroud 90. Theshroud 90 is adjacent toblade tips 86 of thecompressor blades 84 defining ablade tip clearance 80 therebetween. Thediffuser 42 includes anannular diffuser housing 20 having a plurality of tangentially disposeddiffuser flow passages 22 extending radially therethrough, spaced about acircumference 26 of thehousing 20, and through whichdiffuser airflow 103 flows in a downstream direction.Diffuser vanes 23 axially extend between aforward wall 101 and theaft wall 100 of thediffuser 42. - Referring to
FIGS. 2 and 3 , thediffuser vanes 23 circumferentially extend between adjacent ones of thediffuser flow passages 22. Thediffuser flow passages 22 are partly defined and circumferentially bounded by the circumferentially spaced apartdiffuser vanes 23. Adjacent ones of thepassages 22 intersect with each other at radiallyinner inlet sections 24 of thepassages 22 that define a quasi-vanelessannular inlet 27 of thediffuser 42. Eachpassage 22 further includes athroat section 28 downstream of and integral with theinner inlet section 24. Eachpassage 22 further includes a diffusingsection 99 immediately downstream of thethroat section 28. - Referring to
FIGS. 2 and 6 , a centrifugal compressor first coolingair source 92 forturbine cooling air 88 is a small predetermined radial clearance (C) located between animpeller tip 36 of the rotatingimpeller 32 and theannular inlet 27 of thestatic diffuser 42.Impeller bleed flow 102 from the radial clearance (C) is collected in a radiallyinner manifold 104. The predetermined. radial clearance (C) is designed to accommodate thermal and mechanical growth of theimpeller 32 and is open to or in fluid communication with the radiallyinner manifold 104. - Referring to
FIGS. 3-6 , we have found that thediffuser airflow 103 on one side of the passage (such as passage 22) in multi-passage diffusers (such as the diffuser 42) that follow or are downstream of centrifugal impellers (such as the impeller 32) is often weak and may be subject to separation. Separation in the passage can generate high losses that lowers engine specific fuel consumption (SFC). This area or region ofweak flow 127 is also believed to be a contributor to surge that limits flow range of the compressor. - A centrifugal compressor stage second cooling
air source 94 forturbine cooling air 88 includes a diffuser boundary layer bleed 96 for bleeding diffuser bleed flow 112 from adiffuser boundary layer 113 in each of thediffuser flow passages 22 of thediffuser 42, illustrated herein as plurality of boundarylayer bleed apertures 106. The diffuserboundary layer bleed 96, also referred to as fluidic bleed, helps reduce the weak flow and limit or prevent the unwanted flow separation. The diffuser boundary layer bleed 96 bleeds diffuser bleed flow 112 from thediffuser boundary layer 113 into a radiallyouter manifold 116. - The radially inner and
outer manifolds inner manifold 104 flows into the radiallyouter manifold 116. The impeller and diffuser bleed flows 102, 112 are mixed in the radiallyouter manifold 116 to provide theturbine cooling air 88 which is then ported or otherwise flowed from radiallyouter manifold 116 through a plurality of circumferentially distributedmanifold ports 117 to thehigh pressure turbine 16. Theturbine cooling air 88 may be channelled or flowed therefrom by external piping (not shown) to cool the first and second stage shrouds 71, 69 (illustrated inFIG. 1 ). - Substantially axially extending beams or struts 122 separate the radially inner and
outer manifolds struts 122 as it flows from the radiallyinner manifold 104 into the radiallyouter manifold 116. The fluidic bleed flow illustrated herein as the diffuser boundary layer bleed 96 represents a small amount of flow, less than 1% of the engine core flow. The fluidic bleed is strategically removed near the inception of the weak flow to improve the overall performance of the diffuser. - Referring to
FIGS. 3-5 , the boundarylayer bleed apertures 106 may be holes orslots 130 through theforward wall 101 of thediffuser 42 as illustrated herein. The boundarylayer bleed apertures 106 orslots 130 lead into and are in flow communication with the radiallyouter manifold 116. Theslot 130 is positioned or located downstream of thethroat section 28 near apressure side 126 of thediffuser vane 23 at a position where the flow would begin to show weakness or instability in a diffuser without the diffuserboundary layer bleed 96. This position is located in what is referred to as a region offlow weakness 127. A slot width W may be sized with manufacturing constraints such as a minimum tool size. A slot length L may be selected to enable up to 3% of the engine core flow to be used. - The
slot 130 should ideally be angled such that thediffuser bleed flow 112 exits the slot perpendicular to aforward surface 105 of theforward wall 101 of thediffuser 42 in aradial plane 132 passing through the engine centerline oraxis 25 as illustrated inFIG. 5 . However, because of constraints such as the slot extending through or very near abend 134 in theforward wall 101 of thediffuser 42 this angle may be different. Theslot 130 has radially outer andinner walls FIG. 6 , and upstream and downstream facingwalls 140, 142, as illustrated inFIGS. 4 and 5 respectively, extending through theforward wall 101. The downstream facingwall 142 is designed to scoopboundary layer air 144 in thediffuser boundary layer 113 only. Thus, the downstream facingwall 142 is angled or canted at an acute cant angle B of less than 90 degrees with respect to the diffuser airflow 103 (parallel to the directionboundary layer air 144 in the downstream direction in thediffuser flow passages 22 of thediffuser 42. It appears that an acute cant angle B of 45 degrees is desirable. However, the acute cant angle B is limited by geometry and manufacturing constraints on the outside of the diffuser so that an acute cant angle, for example about 22.5 degrees, is more practical. - Illustrated in
FIGS. 7-13 is a gas turbine engine with a centrifugal compressor similar to the one illustrated inFIGS. 1-3 but with an alternative arrangement or design for separately gathering and flowing the impeller tip bleed and diffuser bleed flow for cooling turbine components. Theimpeller bleed flow 102 front the radial clearance (C), illustrated inFIG. 9 , is flowed into and collected in a radially innerannular manifold 154 illustrated inFIGS. 7 and 9 .Inter-manifold apertures 160 are disposed between the innerannular manifold 154 and a plurality of radially outerannular manifolds 156 illustrated inFIGS. 7, 9, and 13 . Theinter-manifold apertures 160 allow theimpeller bleed flow 102 to flow front the innerannular manifold 154 into the outerannular manifolds 156. The impeller bleed flow 102 from the outerannular manifolds 156 is then ported or otherwise flowed through a plurality of circumferentially distributed impeller bleed flowmanifold ports 157, illustrated inFIG. 10 , to thehigh pressure turbine 16 for turbine cooling. - Referring to
FIGS. 8, 10, and 11-13 , the diffuser boundary layer bleed 96 bleeds diffuser bleed flow 112 from thediffuser boundary layer 113 into an annulardiffuser bleed manifold 158 from where thediffuser bleed flow 112 is then ported or otherwise flowed through a plurality of circumferentially distributed diffuser bleedmanifold ports 159 to thehigh pressure turbine 16 for turbine cooling.FIG. 10 illustrates the relative circumferential and axial locations of the impeller bleed flowmanifold ports 157 and the diffuser bleedmanifold ports 159 on and through the diffuser forward casing 114. - While there have been described herein what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein and, it is therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention. Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims.
Claims (24)
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US15/517,262 US20170248155A1 (en) | 2014-10-07 | 2015-08-11 | Centrifugal compressor diffuser passage boundary layer control |
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US201462060991P | 2014-10-07 | 2014-10-07 | |
PCT/US2015/044673 WO2016057112A1 (en) | 2014-10-07 | 2015-08-11 | Centrifugal compressor diffuser passage boundary layer control |
US15/517,262 US20170248155A1 (en) | 2014-10-07 | 2015-08-11 | Centrifugal compressor diffuser passage boundary layer control |
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US15/517,262 Abandoned US20170248155A1 (en) | 2014-10-07 | 2015-08-11 | Centrifugal compressor diffuser passage boundary layer control |
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EP (1) | EP3204616A1 (en) |
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US20170328283A1 (en) * | 2016-05-10 | 2017-11-16 | General Electric Company | Impeller-mounted vortex spoiler |
US20180274376A1 (en) * | 2017-03-27 | 2018-09-27 | General Electric Company | Diffuser-deswirler for a gas turbine engine |
US20180355877A1 (en) * | 2017-06-13 | 2018-12-13 | General Electric Company | Compressor bleed apparatus for a turbine engine |
EP3489527A1 (en) * | 2017-11-28 | 2019-05-29 | Honeywell International Inc. | Compressor with offset diffuser for integral bleed |
US20200325911A1 (en) * | 2019-04-12 | 2020-10-15 | Rolls-Royce Corporation | Deswirler assembly for a centrifugal compressor |
US10830144B2 (en) * | 2016-09-08 | 2020-11-10 | Rolls-Royce North American Technologies Inc. | Gas turbine engine compressor impeller cooling air sinks |
US10876549B2 (en) | 2019-04-05 | 2020-12-29 | Pratt & Whitney Canada Corp. | Tandem stators with flow recirculation conduit |
US11131210B2 (en) | 2019-01-14 | 2021-09-28 | Honeywell International Inc. | Compressor for gas turbine engine with variable vaneless gap |
US11143201B2 (en) | 2019-03-15 | 2021-10-12 | Pratt & Whitney Canada Corp. | Impeller tip cavity |
US11187243B2 (en) | 2015-10-08 | 2021-11-30 | Rolls-Royce Deutschland Ltd & Co Kg | Diffusor for a radial compressor, radial compressor and turbo engine with radial compressor |
US11268536B1 (en) | 2020-09-08 | 2022-03-08 | Pratt & Whitney Canada Corp. | Impeller exducer cavity with flow recirculation |
US11286952B2 (en) | 2020-07-14 | 2022-03-29 | Rolls-Royce Corporation | Diffusion system configured for use with centrifugal compressor |
US11378005B1 (en) | 2020-12-17 | 2022-07-05 | Pratt & Whitney Canada Corp. | Compressor diffuser and diffuser pipes therefor |
US20220268285A1 (en) * | 2021-02-19 | 2022-08-25 | Pratt & Whitney Canada Corp. | Housing for a centrifugal compressor |
US11441516B2 (en) | 2020-07-14 | 2022-09-13 | Rolls-Royce North American Technologies Inc. | Centrifugal compressor assembly for a gas turbine engine with deswirler having sealing features |
US11578654B2 (en) | 2020-07-29 | 2023-02-14 | Rolls-Royce North American Technologies Inc. | Centrifical compressor assembly for a gas turbine engine |
US20230358170A1 (en) * | 2022-05-09 | 2023-11-09 | General Electric Company | Diffuser with passlets |
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US11732731B2 (en) | 2021-10-08 | 2023-08-22 | Honeywell International Inc. | Diffuser and deswirl system with integral tangential onboard injector for engine |
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US11143201B2 (en) | 2019-03-15 | 2021-10-12 | Pratt & Whitney Canada Corp. | Impeller tip cavity |
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US11098730B2 (en) * | 2019-04-12 | 2021-08-24 | Rolls-Royce Corporation | Deswirler assembly for a centrifugal compressor |
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US11815047B2 (en) | 2020-07-14 | 2023-11-14 | Rolls-Royce North American Technologies Inc. | Centrifugal compressor assembly for a gas turbine engine with deswirler having sealing features |
US11441516B2 (en) | 2020-07-14 | 2022-09-13 | Rolls-Royce North American Technologies Inc. | Centrifugal compressor assembly for a gas turbine engine with deswirler having sealing features |
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US11268536B1 (en) | 2020-09-08 | 2022-03-08 | Pratt & Whitney Canada Corp. | Impeller exducer cavity with flow recirculation |
US11378005B1 (en) | 2020-12-17 | 2022-07-05 | Pratt & Whitney Canada Corp. | Compressor diffuser and diffuser pipes therefor |
US20220268285A1 (en) * | 2021-02-19 | 2022-08-25 | Pratt & Whitney Canada Corp. | Housing for a centrifugal compressor |
US11885338B2 (en) * | 2021-02-19 | 2024-01-30 | Pratt & Whitney Canada Corp. | Housing for a centrifugal compressor |
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Also Published As
Publication number | Publication date |
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CN107110180A (en) | 2017-08-29 |
CA2963914A1 (en) | 2016-04-14 |
JP2017530299A (en) | 2017-10-12 |
WO2016057112A1 (en) | 2016-04-14 |
EP3204616A1 (en) | 2017-08-16 |
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