WO2014130101A2 - Échangeur de chaleur de chambre de combustion de turbine à gaz - Google Patents

Échangeur de chaleur de chambre de combustion de turbine à gaz Download PDF

Info

Publication number
WO2014130101A2
WO2014130101A2 PCT/US2013/068686 US2013068686W WO2014130101A2 WO 2014130101 A2 WO2014130101 A2 WO 2014130101A2 US 2013068686 W US2013068686 W US 2013068686W WO 2014130101 A2 WO2014130101 A2 WO 2014130101A2
Authority
WO
WIPO (PCT)
Prior art keywords
combustor
fuel
heat exchanger
flow
air
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.)
Ceased
Application number
PCT/US2013/068686
Other languages
English (en)
Other versions
WO2014130101A3 (fr
Inventor
Duane A. Smith
William G. Cummings, Iii
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.)
Rolls Royce Corp
Original Assignee
Rolls Royce Corp
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 Rolls Royce Corp filed Critical Rolls Royce Corp
Publication of WO2014130101A2 publication Critical patent/WO2014130101A2/fr
Publication of WO2014130101A3 publication Critical patent/WO2014130101A3/fr
Anticipated expiration legal-status Critical
Ceased legal-status Critical Current

Links

Classifications

    • 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
    • F23R3/005—Combined with pressure or heat exchangers
    • 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
    • F02C7/22—Fuel supply systems
    • F02C7/224—Heating fuel before feeding to the burner
    • 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/30—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply comprising fuel prevapourising devices
    • 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
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03043—Convection cooled combustion chamber walls with means for guiding the cooling air flow
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present disclosure generally relates to gas turbine engine heat exchangers, and more particularly, but not exclusively, to fuel/air heat exchangers.
  • One embodiment of the present disclosure is a unique heat exchanger used within a combustor of a gas turbine engine.
  • Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for exchanging heat between a working fluid in the combustor with a fuel provided to the combustor. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.
  • FIG. 1 depicts one embodiment of a gas turbine engine.
  • FIG. 2 depicts an embodiment of a heat exchanger used in a combustor of the gas turbine engine.
  • FIG. 3 depicts another embodiment of a heat exchanger used in a combustor of the gas turbine engine.
  • a gas turbine engine 50 having a compressor 52, combustor 54, and turbine 56 and can be used in some embodiments as a power source.
  • the gas turbine engine 50 is used as a powerplant for an aircraft.
  • the term "aircraft” includes, but is not limited to, helicopters, airplanes, unmanned space vehicles, fixed wing vehicles, variable wing vehicles, rotary wing vehicles, unmanned combat aerial vehicles, tailless aircraft, hover crafts, and other airborne and/or extraterrestrial (spacecraft) vehicles.
  • present disclosures are contemplated for utilization in other applications that may not be coupled with an aircraft such as, for example, industrial applications, power generation, pumping sets, naval propulsion, weapon systems, security systems, perimeter defense/security systems, and the like known to one of ordinary skill in the art.
  • the gas turbine engine 50 can take a variety of forms in various embodiments. Though depicted as an axial flow single spool engine, in some forms the gas turbine engine 50 can have multiple spools and/or can be a centrifugal or mixed centrifugal/axial flow engine. In some forms the engine 50 can be a turboprop, turbofan, or turboshaft engine. Furthermore, the engine can be an adaptive cycle and/or variable cycle engine. Other variations are also contemplated.
  • the heat exchanger 58 can be used to exchange heat between multiple fluid flow paths at a variety of temperatures, flow rates, pressures, etc.
  • the heat exchanger is a fuel/air heat exchanger, an operation of which will be discussed further below after a description of one form of the combustor 54.
  • the combustor 54 can take any variety of configurations and generally includes an inner combustion portion 60 in which a fuel and working fluid are mixed and combusted, and an exterior portion 64 in which generally no combustion occurs.
  • the combustor is configured to receive working fluid through passage 59 and deliver working fluid to passage 61.
  • the passage 59 can be a compressor passage and the passage 61 can be a turbine passage.
  • the inner combustion portion 60 can take on any variety of configurations, one non- limiting embodiment of which is shown below in FIG. 3.
  • the inner combustion portion 60 can be defined by walls, liners, domes, cans, or combinations thereof.
  • the structures that define the inner combustion portion 60 need not be solid but can be perforated, have slots, holes, etc. for the passage of working fluid such as air.
  • the various openings provided for air entrance to the inner combustion portion 60 can be used to convey working fluid to participate directly in the combustion process, and/or can be used for dilution air, cooling air, etc.
  • the inner combustion portion 60 can be defined by liners offset from each other that are coupled through a combustor dome.
  • fuel injectors or nozzles can protrude through the structure that defines the inner combustion portion 60.
  • the inner combustion portion 60 can include areas that do not locally include a combustion process, but that nevertheless the inner combustion portion 60 is in part defined by structure that generally separates it from the exterior portion 64.
  • some upstream areas of the inner combustion portion 60 that are substantially free from fuel will not include a combustion process, but nevertheless that area will generally be considered part of the inner combustion portion 60.
  • the exterior portion 64 extends between the inner combustion portion 60 and one or more structures that define the exterior portion 64.
  • the exterior portion 64 may not be the same size and shape at all axial/circumferential locations relative to the inner combustion portion 60. In fact, the exterior portion may not entirely surround the inner combustion portion 60.
  • the exterior portion 64 can be defined by various structures of the gas turbine engine.
  • the exterior portion 64 can be defined by a casing, compressor discharge such as through a diffuser, for example, a turbine inlet end, etc.
  • the exterior portion 64 includes a boundary for a flow path for working fluid that is located outside of the inner combustion portion 60 but that nonetheless is a flow path for fluid that is eventually expelled such as through the turbine 56.
  • Fuel can be delivered to the inner combustion portion 60 through a variety of manners including via an injector, nozzle, etc. in any of various states, such as liquid, vapor, mixed, etc.
  • the schematic embodiment disclosed in FIG. 3 depicts a fuel passage 62 which conveys a fuel from a location outside of the combustor 54 to a location into the inner combustion portion 60.
  • the fuel passage 62 traverses the exterior portion 64 and is routed through the heat exchanger 58 prior to being mixed with working fluid and combusted in the inner combustion portion 60.
  • the heat exchanger 58 can include one or more fluid paths that are located in the exterior portion 64, the inner combustion portion 60, intermediate the two portions 60 and 64, or combinations thereof.
  • a fluid flow path for fuel in the heat exchanger 58 is described below.
  • the cooling passage 69 can take a variety of shapes and sizes and can include any number of turns/bends/etc. within and prior to exiting the combustor 54.
  • One non-limiting embodiment of the cooling passage 69 is described below in FIG. 3.
  • FIG. 3 depicts an embodiment of the combustor 54 in which the compressor 52 and turbine 56 are depicted as axial flow turbomachinery components. Other forms and combinations of the compressor 52 and turbine 56 are contemplated herein, whether of the centrifugal or mixed axial-centrifugal types.
  • the combustor 54 of the illustrated embodiment is depicted as a straight flow-through combustor and is of the annular configuration, but other forms of the combustor 54 are also contemplated in other embodiments.
  • a compressor discharge, via a diffuser 66 in the illustrated embodiment provides compressed air to the combustor 54; a turbine inlet, via a turbine inlet guide vane 68 in the illustrated embodiment, receives working fluid from the combustor 54.
  • the combustor 54 depicted in FIG. 3 includes an outer casing 70, outer liner 72, inner liner 74, and inner casing 76.
  • a fuel injector 78 extends into the combustor 54 to deliver fuel to the inner combustion portion 60.
  • the fuel injector 78 delivers fuel to within the inner combustion portion 60 in a liquid form, or combination liquid and vapor.
  • the liquid can be present as a stream, film, droplets, etc.
  • the fuel injector 78 conveys fuel to a vaporizer 80 within which any fuel introduced to the vaporizer in liquid form can turn from the liquid to a vapor.
  • Various configurations of the vaporizer 80 are contemplated beyond the embodiment depicted in FIG. 3.
  • the cooling passage 69 is configured to extend between an area in thermal communication with fuel delivered internal to the combustor 54 to one or more areas outside of the combustor 54.
  • the cooling passage 69 extends from an area in thermal communication with the vaporizer 80, to an inner part of the inner casing 76 before cooling fluid is split to flow aft to the turbine 56 and forward to the compressor 52.
  • the cooling passage 69 can extend along an outer portion of the inner casing 76.
  • the cooling passage 69 can extend across the exterior portion 64 in other locations besides the area depicted in the illustrated embodiment.
  • the cooling passage 69 can extend across the portion 64 on a side opposite the inner combustion portion 60 depicted in the figure.
  • the cooling passage 69 can extend near one of the passages 59, 61. Any variety of other locations, configurations, orientations, etc. of the cooling passage 69 as it exits the combustor 54 are contemplated herein. As seen in the illustrated embodiment, the cooling passage 69 extends across the outer combustion portion 64 as it extends downstream and away from the area that it is in thermal communication with the fuel (in the illustrated embodiment, the vaporizer 80 serves as part of the heat exchange between the fuel and working fluid in the cooling passage 69). In one form the cooling passage 69 is a closed off flow path separate from the outer combustion portion 64.
  • the cooling passage 69 includes an inlet 82 structured to receive working fluid from the passage 59 which is in the form of a compressor discharge through a diffuser in the embodiment of FIG. 3.
  • the inlet 82 can have a variety of shapes and sizes, and in one form can be integrated with one or more components of the combustor.
  • the inlet 82 is offset from a dome 84, but in some forms one or more parts of the inlet 82 can be formed from the dome.
  • the inlet 82 can include an outer lip offset from the dome 84, while the dome 84 itself forms the inner lip. In this manner the inlet 82 can take the form of a scoop.
  • Other configurations are contemplated herein.
  • the inlet 82 can be positioned upstream of, coincident with, or downstream of an area of heat exchange between working fluid that gives up heat and the fuel that absorbs it.
  • One or more cooling passages 69 can be arranged in the combustor(s) 54 of the gas turbine engine 50.
  • cooled cooling air is routed from the cooling passage 69 to both the compressor 52 and the turbine.
  • the cooling passage 69 can route cooled cooling air to either, or both, of the compressor and turbine.
  • the cooling air that is routed within the cooling passage 69 can be pulled from working fluid that would be provided to the inner combustion portion 60 and/or from the working fluid that would be routed to the outer combustion portion 64.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Heat-Exchange Devices With Radiators And Conduit Assemblies (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

La présente invention concerne une turbine à gaz présentant une chambre de combustion dans laquelle un échangeur de chaleur est disposé à l'intérieur de la chambre de combustion. L'échangeur de chaleur peut prendre la forme d'un échangeur de chaleur carburant/air. Dans une forme, l'échangeur de chaleur comprend une voie permettant à l'air de refroidissement d'être transporté jusqu'à un emplacement externe à la chambre de combustion. L'air de refroidissement refroidi transporté à travers la voie peut être obtenu à travers l'action du transfert de chaleur de l'air de refroidissement à un carburant s'écoulant dans l'échangeur de chaleur. L'échangeur de chaleur peut comprendre un vaporiseur de carburant dans une forme.
PCT/US2013/068686 2013-02-23 2013-11-06 Échangeur de chaleur de chambre de combustion de turbine à gaz Ceased WO2014130101A2 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US201361768441P 2013-02-23 2013-02-23
US61/768,441 2013-02-23

Publications (2)

Publication Number Publication Date
WO2014130101A2 true WO2014130101A2 (fr) 2014-08-28
WO2014130101A3 WO2014130101A3 (fr) 2014-11-06

Family

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Family Applications (1)

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PCT/US2013/068686 Ceased WO2014130101A2 (fr) 2013-02-23 2013-11-06 Échangeur de chaleur de chambre de combustion de turbine à gaz

Country Status (2)

Country Link
US (1) US20150000291A1 (fr)
WO (1) WO2014130101A2 (fr)

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Also Published As

Publication number Publication date
US20150000291A1 (en) 2015-01-01
WO2014130101A3 (fr) 2014-11-06

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