US20150323185A1 - Turbine engine and method of assembling thereof - Google Patents

Turbine engine and method of assembling thereof Download PDF

Info

Publication number
US20150323185A1
US20150323185A1 US14/570,838 US201414570838A US2015323185A1 US 20150323185 A1 US20150323185 A1 US 20150323185A1 US 201414570838 A US201414570838 A US 201414570838A US 2015323185 A1 US2015323185 A1 US 2015323185A1
Authority
US
United States
Prior art keywords
combustor
flow
turbine
compressor
accordance
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Abandoned
Application number
US14/570,838
Other languages
English (en)
Inventor
Peter Daniel Silkowski
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
General Electric Co
Original Assignee
General Electric Co
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by General Electric Co filed Critical General Electric Co
Priority to US14/570,838 priority Critical patent/US20150323185A1/en
Assigned to GENERAL ELECTRIC COMPANY reassignment GENERAL ELECTRIC COMPANY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: SILKOWSKI, PETER DANIEL
Publication of US20150323185A1 publication Critical patent/US20150323185A1/en
Priority to EP15196094.5A priority patent/EP3034796A1/fr
Priority to BR102015029799A priority patent/BR102015029799A2/pt
Priority to CA2913910A priority patent/CA2913910A1/fr
Priority to JP2015239000A priority patent/JP2016114055A/ja
Priority to CN201510931236.1A priority patent/CN105697147A/zh
Abandoned legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/14—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/10—Air inlet arrangements for primary air
    • F23R3/12—Air inlet arrangements for primary air inducing a vortex
    • F23R3/14—Air inlet arrangements for primary air inducing a vortex by using swirl vanes
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23P—METAL-WORKING NOT OTHERWISE PROVIDED FOR; COMBINED OPERATIONS; UNIVERSAL MACHINE TOOLS
    • B23P15/00—Making specific metal objects by operations not covered by a single other subclass or a group in this subclass
    • B23P15/04—Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from several pieces
    • 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
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • 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
    • 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
    • F04D27/02—Surge control
    • F04D27/0246—Surge control by varying geometry within the pumps, e.g. by adjusting vanes
    • 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
    • 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
    • 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
    • F04D29/56—Fluid-guiding means, e.g. diffusers adjustable
    • F04D29/563—Fluid-guiding means, e.g. diffusers adjustable specially adapted for elastic fluid pumps
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/002—Wall structures
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/10—Air inlet arrangements for primary air
    • F23R3/12—Air inlet arrangements for primary air inducing a vortex
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/26—Controlling the air flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/34—Feeding into different combustion zones
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/44—Combustion chambers comprising a single tubular flame tube within a tubular casing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
    • 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
    • F05D2240/35—Combustors or associated equipment
    • 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
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49316—Impeller making
    • Y10T29/4932—Turbomachine making
    • Y10T29/49323—Assembling fluid flow directing devices, e.g., stators, diaphragms, nozzles

Definitions

  • the present disclosure relates generally to turbine engines and, more specifically, to an improved turbine engine component architecture utilizing swirling combustion, such as those found in ultra compact combustors.
  • Rotary machines such as gas turbines, are often used to generate power with electric generators.
  • Gas turbines for example, have a gas path that typically includes, in serial-flow relationship, an air intake, a compressor, a combustor, a turbine, and a gas outlet.
  • Compressor and turbine sections include at least one row of circumferentially-spaced rotating buckets or blades coupled within a housing.
  • At least some known turbine engines are used in cogeneration facilities and power plants. Engines used in such applications may have high specific work and power per unit mass flow requirements.
  • a first set of guide vanes is coupled between an outlet of the compressor and an inlet of the combustor.
  • the first set of guide vanes facilitates reducing swirl (i.e., removing bulk swirl) of a flow of air discharged from the compressor such that the flow of air is channeled in a substantially axial direction towards the combustor.
  • a second set of guide vanes is coupled between an outlet of the combustor and an inlet of the turbine.
  • the second set of guide vanes facilitates increasing swirl (i.e., reintroducing bulk swirl) of a flow of combustion gas discharged from the combustor such that flow angle requirements for the inlet of the turbine are satisfied.
  • redirecting the flows of air and combustion gas with the first and second sets of guide vanes increases operating inefficiencies of the gas turbine.
  • including additional components, such as the first and second sets of guide vanes generally adds weight, cost, and complexity to the gas turbine.
  • a turbine engine in one aspect, includes a compressor configured to discharge a flow of air at a first flow angle and a combustor coupled downstream from the compressor.
  • the combustor is configured to combust the flow of air with fuel to form a flow of combustion gas discharged from the combustor at a second flow angle.
  • the turbine engine also includes a turbine coupled downstream from the combustor.
  • the turbine includes an inlet configured to receive the flow of combustion gas having a flow angle within a predetermined range, wherein the combustor is oriented such that the second flow angle is within the predetermined range.
  • a method of assembling a turbine engine includes coupling a combustor downstream from a compressor configured to discharge a flow of air at a first flow angle.
  • the combustor is configured to combust the flow of air with fuel to form a flow of combustion gas configured to discharge from the combustor at a second flow angle.
  • the method also includes coupling a turbine downstream from the combustor, the turbine comprising an inlet configured to receive the flow of combustion gas from the combustor having a flow angle within a predetermined range, and orienting the combustor such that the second flow angle is within the predetermined range.
  • FIG. 1 is a schematic illustration of an exemplary gas turbine engine
  • FIG. 2 is an enlarged schematic illustration of an exemplary combustion assembly that may be used with the gas turbine engine shown in FIG. 1 ;
  • FIG. 3 is a schematic illustration of an exemplary flow path of the combustion assembly along line 3 - 3 shown in FIG. 2 ;
  • FIG. 4 is an enlarged schematic illustration of an alternative combustion assembly that may be used with the gas turbine engine shown in FIG. 1 ;
  • FIG. 5 is a schematic illustration of an exemplary flow path of the combustion assembly along line 5 - 5 shown in FIG. 4 ;
  • FIG. 6 is a schematic illustration of the flow path shown in FIG. 3 along line 6 - 6 shown in FIG. 3 .
  • Approximating language may be applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about”, “approximately”, and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value.
  • range limitations may be combined and/or interchanged. Such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise.
  • Embodiments of the present disclosure relate to turbine engines and methods of assembling thereof. More specifically, the turbine engines described herein include a combustor that operates with bulk swirl flow of combustion gas channeled towards an inlet of a turbine.
  • the turbine generally includes at least one row of rotating buckets, or rotor blades, coupled to a hub in a fixed orientation and the flow of combustion gas channeled towards the turbine needs to have a flow angle within a predetermined range to ensure the bulk swirl requirements at the inlet of the turbine are satisfied.
  • an orientation of the bulk swirl combustor is selected to ensure the flow of combustion gas discharged towards the turbine has a flow angle within the predetermined range.
  • Leveraging bulk swirl combustors to satisfy flow angle requirements of the turbine enables removal of one or both of guide vanes and turbine nozzles positioned at opposing ends of the combustor. Moreover, removing turbine nozzles from the turbine engine reduces the complexity of the turbine engine by eliminating component cooling requirements with compressor bleed air. As such, a length of the turbine is decreased, a weight of the turbine engine is reduced, and turbine efficiency is increased.
  • the terms “axial” and “axially” refer to directions and orientations that extend substantially parallel to a centerline of the turbine engine.
  • the terms “radial” and “radially” refer to directions and orientations that extend substantially perpendicular to the centerline of the turbine engine.
  • the terms “circumferential” and “circumferentially” refer to directions and orientations that extend arcuately about the centerline of the turbine engine.
  • fluid as used herein includes any medium or material that flows, including, but not limited to, air, gas, liquid and steam.
  • FIG. 1 is a schematic illustration of a gas turbine engine 10 including a low pressure compressor 12 , a high pressure compressor 14 , and a combustor 16 coupled downstream from high pressure compressor 14 .
  • Gas turbine engine 10 also includes a high pressure turbine 18 coupled downstream from combustor 16 , a low pressure turbine 20 coupled downstream from high pressure turbine 18 , and a power turbine 22 coupled downstream from low pressure turbine 20 .
  • the compressed air is discharged towards combustor 16 and mixed with fuel to form a flow of combustion gas discharged towards turbines 18 and 20 .
  • the flow of combustion gas drives turbines 18 and 20 about a centerline 24 of gas turbine engine 10 .
  • FIG. 2 is an enlarged schematic illustration of an exemplary combustion assembly 100 that may be used with gas turbine engine 10 (shown in FIG. 1 ), and FIG. 3 is a schematic illustration of an exemplary flow path 102 of combustion assembly 100 along line 3 - 3 (shown in FIG. 2 ).
  • compressor 14 discharges a flow of air 104 towards combustor 16
  • combustor 16 combusts the flow of air 104 with fuel (not shown) to form a flow of combustion gas 106 .
  • the flow of combustion gas 106 is discharged from combustor 16 towards turbine 18 .
  • compressor 14 includes a row 108 of rotor blades 110 defining an outlet 112 of compressor 14 .
  • Rotor blades 110 are in a first fixed orientation relative to centerline 24 such that the flow of air 104 (each shown in FIG. 3 ) is discharged from compressor 14 at a first flow angle.
  • flow angle is defined as a ratio of circumferential velocity to axial velocity of a flow of fluid.
  • Combustor 16 is coupled downstream from compressor 14 and configured to combust the flow of air 104 with fuel to form the flow of combustion gas 106 discharged from combustor 16 .
  • combustor 16 is a bulk swirl combustor in a second fixed orientation relative to centerline 24 such that the flow of combustion gas 106 is discharged from combustor 16 at a second flow angle.
  • combustor 16 includes a corrugated combustor dome 114 and a plurality of combustion devices 116 coupled to corrugated combustor dome 114 . As such, combustion devices 116 are in the second fixed orientation causing the flow of combustion gas 106 to be discharged at the second flow angle.
  • the second fixed orientation is selected such that the second flow angle is within a predetermined range to satisfy the flow angle requirements of inlet 122 , as will be described in more detail below. More specifically, the orientation of combustor 16 is selected as a function of a velocity gradient between the flow of air entering combustor 16 and the flow of combustion gas discharged from combustor 16 . The velocity gradient is induced by the addition of heat during the combustion process.
  • combustor 16 is any bulk swirl combustor that enables gas turbine engine 10 to function as described herein.
  • combustor 16 includes a trapped vortex cavity configuration.
  • Turbine 18 is coupled downstream from combustor 16 and includes a row 118 of rotor blades 120 in a third fixed orientation defining an inlet 122 of turbine 18 .
  • Combustion gas 106 channeled towards turbine 18 needs to have a flow angle within a predetermined range to satisfy the flow angle requirements of inlet 122 , and the predetermined range is based on an orientation of rotor blades 120 .
  • combustor 16 is in a second fixed orientation such that the flow of air 104 enters combustor 16 at the first flow angle, and the flow of combustion gas 106 is discharged from combustor 16 at the second flow angle.
  • the orientation of combustor 16 is selected based on a difference between the first flow angle and the predetermined range, and an expected difference between flow angles of fluid entering and being discharged from combustor 16 .
  • the second fixed orientation of combustor 16 is selected to ensure the second flow angle is within the predetermined range for satisfying the flow angle requirements of inlet 122 .
  • FIG. 4 is an enlarged schematic illustration of an alternative combustion assembly 124 that may be used with gas turbine engine 10 (shown in FIG. 1 ), and FIG. 5 is a schematic illustration of an exemplary flow path 126 of combustion assembly 124 along line 5 - 5 (shown in FIG. 4 ).
  • combustion assembly 124 includes a variable guide vane assembly 128 positioned between a last stage rotor (not shown) of compressor 14 and combustor 16 .
  • Variable guide vane assembly 128 includes a plurality of variable guide vanes 130 .
  • Variable guide vanes 130 are also selectively actuated based on the operating condition of gas turbine engine 10 , as will be described in more detail below.
  • compressor 14 discharges a flow of air 104 through variable guide vane assembly 128 and towards combustor 16 , and combustor 16 combusts the flow of air 104 with fuel (not shown) to form a flow of combustion gas 106 .
  • the flow of combustion gas 106 is discharged from combustor 16 towards turbine 18 .
  • Variable guide vane assembly 128 is positioned between compressor 14 and combustor 16 to ensure the flow angle of combustion gas 106 channeled towards turbine 18 is within the predetermined range for inlet 122 across a wide range of operating conditions for gas turbine engine 10 .
  • variable guide vane assembly 128 is selectively operable to modify the flow angle of the flow of air 104 discharged from compressor 14 and channeled towards combustor 16 .
  • variable guide vanes 130 are selectively actuated to rotate about an axis (not shown) extending substantially radially from centerline 24 (shown in FIG. 4 ).
  • variable guide vanes 130 actuate into a position such that the flow angle of the flow of air 104 discharged from combustor 16 is within the predetermined range for inlet 122 .
  • variable guide vanes 130 actuate into a position that modifies the flow angle of the flow of air 104 before entering combustor 16 .
  • the orientation of variable guide vanes 130 is selected such that the second flow angle of the flow of combustion gas 106 discharged from combustor 16 is within the predetermined range for inlet 122 .
  • FIG. 6 is a schematic illustration of flow path 102 along line 6 - 6 (shown in FIG. 3 ).
  • flow path 102 extends between compressor 14 and turbine 18 in a substantially axial direction along centerline 24 .
  • Flow path 102 is defined by an outer flow duct 132 and an inner flow duct 134 .
  • Flow path 102 also includes a variable cross-sectional area defined between outer and inner flow ducts 132 and 134 along the substantially axial direction.
  • the outer and inner flow ducts 132 and 134 are tailored such that a mean radial location 135 extending therebetween, and a cross-sectional area defined therebetween, facilitates ensuring the flow angle of combustion gas 106 (shown in FIG.
  • flow path 102 includes a first section 136 extending between compressor 14 and combustor 16 , and a second section 138 extending between combustor 16 and turbine 18 .
  • the cross-sectional area of first section 136 progressively increases in size from compressor 14 to combustor 16
  • the cross-sectional area of second section 138 progressively decreases in size from combustor 16 to turbine 18 .
  • Progressively increasing the cross-sectional area of first section 136 decelerates the flow of air 104 (shown in FIG.
  • the turbine engine and methods described herein relate to leveraging bulk swirl combustors to enable a component architecture of the turbine engine to be modified.
  • an orientation of a bulk swirl combustor is selected to ensure a flow of combustion gas discharged towards a turbine has a flow angle within a predetermined range. Leveraging bulk swirl combustors to satisfy flow angle requirements of the turbine enables removal of one or both of guide vanes and turbine nozzles positioned at opposing ends of the combustor.
  • variable guide vanes or a variable-area turbine flow path is implemented to ensure flow angle requirements for the turbine are satisfied at a wide range of operational modes, for example.
  • An exemplary technical effect of the turbine engine and methods described herein includes at least one of: (a) removing redundant components from the turbine engine; (b) reducing a weight and length of the turbine engine; and (c) increasing an operational efficiency of the turbine engine.
  • Exemplary embodiments of the gas turbine engine are described above in detail.
  • the assembly is not limited to the specific embodiments described herein, but rather, components of systems and/or steps of the methods may be utilized independently and separately from other components and/or steps described herein.
  • the configuration of components described herein may also be used in combination with other processes, and is not limited to practice with only gas turbine engines and related methods as described herein. Rather, the exemplary embodiment can be implemented and utilized in connection with many applications where leveraging bulk swirl combustion is desirable.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Physics & Mathematics (AREA)
  • Geometry (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US14/570,838 2014-05-07 2014-12-15 Turbine engine and method of assembling thereof Abandoned US20150323185A1 (en)

Priority Applications (6)

Application Number Priority Date Filing Date Title
US14/570,838 US20150323185A1 (en) 2014-05-07 2014-12-15 Turbine engine and method of assembling thereof
EP15196094.5A EP3034796A1 (fr) 2014-05-07 2015-11-24 Moteur à turbine et son procédé d'assemblage
BR102015029799A BR102015029799A2 (pt) 2014-12-15 2015-11-27 motor de turbina
CA2913910A CA2913910A1 (fr) 2014-05-07 2015-12-03 Moteur a turbine et methode d'assemblage connexe
JP2015239000A JP2016114055A (ja) 2014-05-07 2015-12-08 タービンエンジン及びその組立方法
CN201510931236.1A CN105697147A (zh) 2014-05-07 2015-12-15 涡轮发动机和其组装的方法

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US201461989855P 2014-05-07 2014-05-07
US14/570,838 US20150323185A1 (en) 2014-05-07 2014-12-15 Turbine engine and method of assembling thereof

Publications (1)

Publication Number Publication Date
US20150323185A1 true US20150323185A1 (en) 2015-11-12

Family

ID=54367508

Family Applications (4)

Application Number Title Priority Date Filing Date
US14/570,838 Abandoned US20150323185A1 (en) 2014-05-07 2014-12-15 Turbine engine and method of assembling thereof
US14/706,679 Active 2035-11-21 US10082076B2 (en) 2014-05-07 2015-05-07 Ultra compact combustor having reduced air flow turns
US16/133,362 Active 2035-06-21 US11053844B2 (en) 2014-05-07 2018-09-17 Ultra compact combustor
US17/332,504 Pending US20210285370A1 (en) 2014-05-07 2021-05-27 Ultra compact combustor

Family Applications After (3)

Application Number Title Priority Date Filing Date
US14/706,679 Active 2035-11-21 US10082076B2 (en) 2014-05-07 2015-05-07 Ultra compact combustor having reduced air flow turns
US16/133,362 Active 2035-06-21 US11053844B2 (en) 2014-05-07 2018-09-17 Ultra compact combustor
US17/332,504 Pending US20210285370A1 (en) 2014-05-07 2021-05-27 Ultra compact combustor

Country Status (5)

Country Link
US (4) US20150323185A1 (fr)
EP (1) EP3034796A1 (fr)
JP (1) JP2016114055A (fr)
CN (1) CN105697147A (fr)
CA (1) CA2913910A1 (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20190271268A1 (en) * 2018-03-01 2019-09-05 General Electric Company Turbine Engine With Rotating Detonation Combustion System
US10527287B2 (en) * 2016-04-13 2020-01-07 Safran Helicopter Engines Injectors for gas turbine combustion chamber
US10969107B2 (en) 2017-09-15 2021-04-06 General Electric Company Turbine engine assembly including a rotating detonation combustor
US11434831B2 (en) * 2018-05-23 2022-09-06 General Electric Company Gas turbine combustor having a plurality of angled vanes circumferentially spaced within the combustor
US11536456B2 (en) * 2017-10-24 2022-12-27 General Electric Company Fuel and air injection handling system for a combustor of a rotating detonation engine

Families Citing this family (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10704787B2 (en) * 2016-03-30 2020-07-07 General Electric Company Closed trapped vortex cavity pilot for a gas turbine engine augmentor
CN106091312B (zh) * 2016-06-20 2021-10-08 珠海格力电器股份有限公司 空调室内机
US10823418B2 (en) 2017-03-02 2020-11-03 General Electric Company Gas turbine engine combustor comprising air inlet tubes arranged around the combustor
US20190017441A1 (en) * 2017-07-17 2019-01-17 General Electric Company Gas turbine engine combustor
US11073286B2 (en) * 2017-09-20 2021-07-27 General Electric Company Trapped vortex combustor and method for operating the same
US10823422B2 (en) * 2017-10-17 2020-11-03 General Electric Company Tangential bulk swirl air in a trapped vortex combustor for a gas turbine engine
CN110529250B (zh) * 2018-05-23 2022-08-02 通用电气公司 燃气涡轮发动机
ES2739725B2 (es) * 2018-08-01 2022-06-30 De Juan Miguel Angel Holgado Combustor para turbina y turbina obtenida
GB201820206D0 (en) * 2018-12-12 2019-01-23 Rolls Royce Plc A fuel spray nozzle
WO2020245638A1 (fr) * 2019-06-06 2020-12-10 Miguel Angel Holgado De Juan Brûleur pour turbine et turbine obtenue
US11885497B2 (en) * 2019-07-19 2024-01-30 Pratt & Whitney Canada Corp. Fuel nozzle with slot for cooling
JP2021127755A (ja) * 2020-02-17 2021-09-02 三菱重工業株式会社 ２軸式ガスタービン
AT523924B1 (de) * 2020-10-01 2022-01-15 Comb Bay One E U Verfahren und Vorrichtung zur rekursiven sequentiellen Verbrennung

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20100037618A1 (en) * 2008-08-12 2010-02-18 Richard Charron Transition with a linear flow path for use in a gas turbine engine
US7721547B2 (en) * 2005-06-27 2010-05-25 Siemens Energy, Inc. Combustion transition duct providing stage 1 tangential turning for turbine engines
US20130086908A1 (en) * 2010-06-15 2013-04-11 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber arrangement of axial type of construction
US20150369135A1 (en) * 2012-04-17 2015-12-24 Walter R. Laster Device for improved air and fuel distribution to a combustor

Family Cites Families (42)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2930192A (en) * 1953-12-07 1960-03-29 Gen Electric Reverse vortex combustion chamber
US3722216A (en) * 1971-01-04 1973-03-27 Gen Electric Annular slot combustor
US3955361A (en) * 1971-12-15 1976-05-11 Phillips Petroleum Company Gas turbine combustor with controlled fuel mixing
GB1424197A (en) * 1972-06-09 1976-02-11 Lucas Industries Ltd Combustion chambers for gas turbine engines
FR2402068A1 (fr) * 1977-09-02 1979-03-30 Snecma Chambre de combustion anti-pollution
US4445338A (en) 1981-10-23 1984-05-01 The United States Of America As Represented By The Secretary Of The Navy Swirler assembly for a vorbix augmentor
US4606190A (en) * 1982-07-22 1986-08-19 United Technologies Corporation Variable area inlet guide vanes
DE68906594T2 (de) 1988-04-25 1993-08-26 United Technologies Corp Staubabscheider fuer eine luftgekuehlte schaufel.
US5207064A (en) 1990-11-21 1993-05-04 General Electric Company Staged, mixed combustor assembly having low emissions
US5165226A (en) * 1991-08-09 1992-11-24 Pratt & Whitney Canada, Inc. Single vortex combustor arrangement
US5289685A (en) * 1992-11-16 1994-03-01 General Electric Company Fuel supply system for a gas turbine engine
US5638682A (en) 1994-09-23 1997-06-17 General Electric Company Air fuel mixer for gas turbine combustor having slots at downstream end of mixing duct
US5613363A (en) 1994-09-26 1997-03-25 General Electric Company Air fuel mixer for gas turbine combustor
US5857339A (en) 1995-05-23 1999-01-12 The United States Of America As Represented By The Secretary Of The Air Force Combustor flame stabilizing structure
US5791148A (en) 1995-06-07 1998-08-11 General Electric Company Liner of a gas turbine engine combustor having trapped vortex cavity
US5619855A (en) 1995-06-07 1997-04-15 General Electric Company High inlet mach combustor for gas turbine engine
DE19541303A1 (de) 1995-11-06 1997-05-28 Siemens Ag Gasturbine
US5918458A (en) 1997-02-14 1999-07-06 General Electric Company System and method of providing clean filtered cooling air to a hot portion of a gas turbine engine
GB2343486B (en) 1998-06-19 2000-09-20 Rolls Royce Plc Improvemnts in or relating to cooling systems for gas turbine engine airfoil
US6295801B1 (en) 1998-12-18 2001-10-02 General Electric Company Fuel injector bar for gas turbine engine combustor having trapped vortex cavity
US6286298B1 (en) 1998-12-18 2001-09-11 General Electric Company Apparatus and method for rich-quench-lean (RQL) concept in a gas turbine engine combustor having trapped vortex cavity
US6286317B1 (en) 1998-12-18 2001-09-11 General Electric Company Cooling nugget for a liner of a gas turbine engine combustor having trapped vortex cavity
US6286300B1 (en) 2000-01-27 2001-09-11 Honeywell International Inc. Combustor with fuel preparation chambers
US8272219B1 (en) 2000-11-03 2012-09-25 General Electric Company Gas turbine engine combustor having trapped dual vortex cavity
US6796130B2 (en) 2002-11-07 2004-09-28 Siemens Westinghouse Power Corporation Integrated combustor and nozzle for a gas turbine combustion system
GB2398863B (en) 2003-01-31 2007-10-17 Alstom Combustion Chamber
GB2426555A (en) * 2005-05-28 2006-11-29 Siemens Ind Turbomachinery Ltd Turbocharger air intake
FR2888631B1 (fr) * 2005-07-18 2010-12-10 Snecma Turbomachine a distribution angulaire de l'air
US7588415B2 (en) * 2005-07-20 2009-09-15 United Technologies Corporation Synch ring variable vane synchronizing mechanism for inner diameter vane shroud
GB2429515B (en) * 2005-08-11 2008-06-25 Rolls Royce Plc Cooling method and apparatus
WO2007102807A1 (fr) 2006-03-06 2007-09-13 United Technologies Corporation Chambre de combustion annulaire à flux en biais pour moteur à turbine
EP2076714B1 (fr) 2006-10-18 2016-08-31 Lean Flame, INC. Dispositif de libération/conversion d'énergie
US7762058B2 (en) 2007-04-17 2010-07-27 Pratt & Whitney Rocketdyne, Inc. Ultra-compact, high performance aerovortical rocket thruster
US9822649B2 (en) 2008-11-12 2017-11-21 General Electric Company Integrated combustor and stage 1 nozzle in a gas turbine and method
BR112012005612A2 (pt) 2009-09-13 2016-06-21 Lean Flame Inc pré-misturador de entrada para aparelho de combustão
GB0916432D0 (en) 2009-09-21 2009-10-28 Rolls Royce Plc Separator device
US8739550B2 (en) 2009-09-30 2014-06-03 Precision Combustion, Inc. Two stage combustor with reformer
GB0920094D0 (en) 2009-11-17 2009-12-30 Alstom Technology Ltd Reheat combustor for a gas turbine engine
EP2362148A1 (fr) 2010-02-23 2011-08-31 Siemens Aktiengesellschaft Injecteur de carburant et assemblage de tourbillonnement avec mélangeur à lobes
DE102010009043B4 (de) 2010-02-23 2013-11-07 Iav Gmbh Ingenieurgesellschaft Auto Und Verkehr Statischer Mischer für eine Abgasanlage einer Brennkraftmaschine
RU2550370C2 (ru) 2011-05-11 2015-05-10 Альстом Текнолоджи Лтд Центробежная форсунка с выступающими частями
US9291063B2 (en) 2012-02-29 2016-03-22 Siemens Energy, Inc. Mid-section of a can-annular gas turbine engine with an improved rotation of air flow from the compressor to the turbine

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7721547B2 (en) * 2005-06-27 2010-05-25 Siemens Energy, Inc. Combustion transition duct providing stage 1 tangential turning for turbine engines
US20100037618A1 (en) * 2008-08-12 2010-02-18 Richard Charron Transition with a linear flow path for use in a gas turbine engine
US20130086908A1 (en) * 2010-06-15 2013-04-11 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber arrangement of axial type of construction
US20150369135A1 (en) * 2012-04-17 2015-12-24 Walter R. Laster Device for improved air and fuel distribution to a combustor

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10527287B2 (en) * 2016-04-13 2020-01-07 Safran Helicopter Engines Injectors for gas turbine combustion chamber
US10969107B2 (en) 2017-09-15 2021-04-06 General Electric Company Turbine engine assembly including a rotating detonation combustor
US11536456B2 (en) * 2017-10-24 2022-12-27 General Electric Company Fuel and air injection handling system for a combustor of a rotating detonation engine
US20190271268A1 (en) * 2018-03-01 2019-09-05 General Electric Company Turbine Engine With Rotating Detonation Combustion System
US11434831B2 (en) * 2018-05-23 2022-09-06 General Electric Company Gas turbine combustor having a plurality of angled vanes circumferentially spaced within the combustor
US11840967B2 (en) * 2018-05-23 2023-12-12 General Electric Company Gas turbine engine

Also Published As

Publication number Publication date
US20190226398A1 (en) 2019-07-25
CN105697147A (zh) 2016-06-22
CA2913910A1 (fr) 2016-06-15
JP2016114055A (ja) 2016-06-23
US20150323184A1 (en) 2015-11-12
EP3034796A1 (fr) 2016-06-22
US20210285370A1 (en) 2021-09-16
US11053844B2 (en) 2021-07-06
US10082076B2 (en) 2018-09-25

Similar Documents

Publication Publication Date Title
US20150323185A1 (en) Turbine engine and method of assembling thereof
US10794396B2 (en) Inlet pre-swirl gas turbine engine
US20180363676A1 (en) Inlet pre-swirl gas turbine engine
US11391211B2 (en) Waste heat recovery system
US10458247B2 (en) Stator of an aircraft turbine engine
CN107178425B (zh) 一种燃气涡轮发动机和其中的通气孔组件
US10077660B2 (en) Turbine engine assembly and method of manufacturing
US10196982B2 (en) Gas turbine engine having a flow control surface with a cooling conduit
US10823422B2 (en) Tangential bulk swirl air in a trapped vortex combustor for a gas turbine engine
CN107956598B (zh) 燃气涡轮发动机
US10718340B2 (en) Gas turbine manufacturing method
US20170159449A1 (en) Gas turbine engine with fillet film holes
US10513944B2 (en) Manifold for use in a clearance control system and method of manufacturing
EP3239511A1 (fr) Ensemble de turboréacteur et procédé d'assemblage
US20170016620A1 (en) Combustor assembly for use in a gas turbine engine and method of assembling
US10641490B2 (en) Combustor for use in a turbine engine
US11274563B2 (en) Turbine rear frame for a turbine engine
US20180216576A1 (en) Supersonic turbofan engine
EP3392468B1 (fr) Diffuseur d'échappement avec couronnes d'aubes directrices variables pour un moteur à turbine à gaz
US11840967B2 (en) Gas turbine engine
EP3182009A1 (fr) Chemise de combustion destinée à être utilisée dans un dispositif combustor et procédé de fabrication
US10774661B2 (en) Shroud for a turbine engine
US11401835B2 (en) Turbine center frame
EP3112634A1 (fr) Alternative d'architecture-design de moteur distribué avancé
BR102015029799A2 (pt) motor de turbina

Legal Events

Date Code Title Description
AS Assignment

Owner name: GENERAL ELECTRIC COMPANY, NEW YORK

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:SILKOWSKI, PETER DANIEL;REEL/FRAME:034510/0293

Effective date: 20141215

STCB Information on status: application discontinuation

Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION