EP0203431A1 - Prallkühlung für einen Turbineneinlasskanal - Google Patents

Prallkühlung für einen Turbineneinlasskanal Download PDF

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Publication number
EP0203431A1
EP0203431A1 EP86106295A EP86106295A EP0203431A1 EP 0203431 A1 EP0203431 A1 EP 0203431A1 EP 86106295 A EP86106295 A EP 86106295A EP 86106295 A EP86106295 A EP 86106295A EP 0203431 A1 EP0203431 A1 EP 0203431A1
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EP
European Patent Office
Prior art keywords
impingement
transition duct
air
cooling
combustor
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Granted
Application number
EP86106295A
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English (en)
French (fr)
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EP0203431B2 (de
EP0203431B1 (de
Inventor
Lewis Berkley Davis, Jr.
Walter Walls Goodwin
Charles Even Steber
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General Electric Co
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General Electric Co
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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • 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
    • 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
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling
    • 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
    • F23R2900/03044—Impingement cooled combustion chamber walls or subassemblies

Definitions

  • the present invention relates to gas turbine engines and, more particularly, to apparatus for cooling a transition duct employed to conduct hot gasses from a combustor to a turbine stage of an advanced heavy duty gas turbine engine.
  • a large heavy duty gas turbine engine conventionally employs a plurality of cylindrical combustor stages operated in parallel to produce hot energetic gas for introduction into the first turbine stage of the engine.
  • the first turbine stage preferably receives the hot gas in the shape of an annulus.
  • a transition duct is disposed between each of the combustor stages and the first turbine stage to change the gas flow field exiting each combustor from a generally cylindrical shape to one which forms part of an annulus. The gas flow from all of the transition ducts thus produces the desired annular flow.
  • thermodynamic efficiency of which a heat engine is capable depends on the maximum temperature of its working fluid which, in the case of a gas turbine, is the hot gas exiting the combustor stages.
  • the maximum feasible temperature of the hot gas is limited by the operating temperature limit of the metal parts in contact with this hot gas, and on the ability to cool these parts below the hot gas temperature.
  • the task of cooling the transition duct of an advanced heavy duty gas turbine engine, which is the one addressed by the present invention, is difficult because currently known cooling methods are either inadequate, or carry unacceptable penalties.
  • the entire external surface of the transition duct is exposed to relatively cool air discharged from the compressor, which supplies the total air flow for the gas turbine.
  • the flow of air over the exterior of the transition duct to the combustor causes passive cooling.
  • Some portions of the exterior of the transition duct are relatively well cooled by passive cooling, but others are poorly cooled thereby.
  • the portions of the exterior of the transition duct that are most poorly cooled are generally in structurally weaker areas, which are also areas more highly heated by the hot gas therewithin.
  • the maximum combustor exit temperature must be limited by the maximum allowed metal,temperature of the most poorly cooled areas of of the transition duct.
  • Another cooling technique which has found use in cooling the exterior of the transition duct employs an impingement plate, baffle or sleeve disposed a short distance away from the transition duct outer surface.
  • the impingement sleeve contains an array of holes through which compressor discharge air passes to generate an array of air jets which impinge on and cool the outer surface of the transition duct.
  • U.S. Patent No. 3,652,181 discloses such an impingement cooling technique for a transition duct in which the impingement sleeve surrounds only a portion of the transition duct. After impacting the surface to be cooled, the spent impingement air flows in the space between the transition duct outer surface and the impingement sleeve, towards holes in the transition duct. The air passing through these holes mixes with, and reduces the hot gas temperature just ahead of, the root area of the turbine blades and thus helps reduce the metal temperature of this portion of the turbine blades. Depending upon the heat transfer rate from the hot gas and the maximum allowed metal temperature, this method can use less cooling air than film cooling to maintain acceptable metal temperatures, and can be used in combination with film cooling to further reduce metal temperature. However, even the combination of impingement and film cooling for a transition duct would require more cooling air than is available in an advanced heavy duty gas turbine.
  • impingement cooling of a combustion component can either consume a portion of the air flow allocated to the combustion process, or be performed in series with the combustor such that the air used to cool a combustion component is subsequently used in the combustion process. It is the series mode of cooling a transition duct which is addressed by the present invention.
  • the pressure drop of an impingement cooling system essentially is generated by two components. First, there is a pressure drop needed to accelerate the air through the impingement sleeve holes to create the jets which impinge on the surface to be cooled. The second is more subtle, and is largely ignored in other known impingement cooling applications.
  • the spent impingement air If the spent impingement air is to be used in the combustor, it must be collected and brought to the combustor. The collection naturally takes place between the impingement sleeve and the external surface of the transition ducts, and it will be seen that, as one moves towards the combustor, the air flow velocity must steadily increase as more air is collected. The second component of pressure drop occurs due to the requirement to reaccelerate each additional quantity of spent impingement air to the velocity of that air already moving towards the combustor.
  • the local magnitude of the heat transfer in an impingement cooling system is determined by a number of variables.
  • these variables include the cooling air properties, the local distance between the impingement sleeve and the transition duct surface, the hole size, spacing and array pattern, the impingement air jet velocity, and the velocity of air flowing perpendicular to the air jet such as, for example, air resulting from the collection of spent impingement air.
  • An air jet formed by an opening in an impingement plate must traverse the space separating the impingement plate from the surface to be cooled, and must impact the surface to be cooled with sufficient velocity and in sufficient volume to effect the desired cooling.
  • the analysis of such jet impingement is relatively simple when only a single jet is involved. However, when an array of jets is used, the impingement air flowing away after impingement from one jet, captured between the surface being cooled and the impingement plate, tends to produce a crossflow of air which interferes with the cooling action of other jets, particularly those downstream in the direction along which the impingement air must flow to exit the constraining space.
  • a crossflow of air passing through the space between an aperture and the surface to be cooled may prevent the aperture-produced air jet from reaching the surface to be cooled, or may reduce the effectiveness of any portion of the air jet which may reach the surface to be cooled.
  • the actual cooling effects of an array of jets is difficult to predict, and so may only be derived empirically.
  • the spacing between these larger holes is preferably varied relative to the spacing of the smaller holes, to establish a desired impingement cooling intensity as required by the transition duct design.
  • the present invention provides impingement cooling for a transition duct in an advanced heavy duty gas turbine engine.
  • the transition duct is cooled by impingement jets formed by apertures in a sleeve spaced a distance from the surface to be cooled.
  • the sleeve is configured so as to duct spent impingement air towards the combustor, where it can be subsequently used for mixing with, and combustion of, the fuel, or for cooling of the combustor.
  • the distance between the impingement sleeve and the transition duct surface is varied to control the velocity of air crossflow from spent impingement air in order to minimize the pressure loss due to crossflow.
  • the cross-sectional areas of the apertures are varied to project impingement jets over the various distances and crossflow velocities. Generally, larger aperture areas are used with larger distances.
  • the distance between the impingement sleeve and the transition duct increases systematically towards the combustor as the quantity of spent impingement air increases to a maximum value at the intersection of the combustor and the transition duct.
  • the combination of variations in distance, aperture size, and inter-aperture spacing is utilized to vary the impingement cooling intensity to compensate for the variable internal heat load and also to produce the desired temperature distribution over the surface of the transition duct according to design requirements.
  • the aforementioned variations are optimized to minimize the air flow pressure drop ahead of the combustion system which achieving the required cooling intensity according to design requirements.
  • an impingement cooling apparatus for cooling a surface, the surface being disposed in a compressed air environment, comprising an impingement plate spaced a distance from the surface, a plurality of apertures in the impingement plate, means for producing a pressure drop across the impingement plate whereby each of the apertures produces an impingement jet directed toward the surface, the apertures having an area, the apertures being spaced apart by a spacing and at least one of the distance, the area and the spacing being varied over the impingement plate to control a cooling in the surface.
  • an impingement cooling apparatus for cooling a surface of a transition duct disposed between a combustor and a turbine stage of a gas turbine engine, the transition duct being disposed in a compressed air plenum, comprising an impingement sleeve surrounding the transition duct and spaced a distance therefrom to form a flow volume therebetween, a plurality of apertures in the impingement sleeve, each of the plurality of apertures having an area, adjacent ones of the apertures being separated by a spacing, a closed end at a turbine end of the flow volume, an exit at a combustor end of the flow volume, a flow sleeve surrounding the combustor, a flared entry portion at an end of the flow sleeve adjacent overlapping the exit and forming an aerodynamic converging shape therebetween, a flow of air through the aerodynamic converging shape flowing toward the combustor being effective to reduce a pressure at the exit below a pressure
  • an impingement cooling apparatus for cooling an enclosed surface formed by a wall affixed to a transition duct, the transition duct being enclosed by an impingement sleeve, the wall passing through an opening in the impingement sleeve, the transition duct and the impingement sleeve being disposed in a plenum effective to contain a pressurized air environment, comprising an impingement insert within the wall having a planar bottom spaced a distance from the enclosed surface, a plurality of apertures in the planar bottom, means for producing a pressure drop across the impingement insert whereby each of the apertures produces an impingement jet directed toward the enclosed surface, the apertures having an area, the apertures being spaced apart by a spacing, the enclosed surface including at least one film cooling aperture through the transition duct for exhausting spent impingement cooling air from between the impingement insert and the enclosed surface and the area and the spacing being varied over the planar bottom to tailor a cooling in the surface.
  • Gas turbine engine 10 includes a plurality of combustors 12, only one of which is shown, uniformly disposed with respect to a longitudinal axis thereof. In one type of gas turbine engine 10, ten combustors 12 are employed. Fuel and primary combustion air are injected into combustor 12 through a fuel nozzle 14. The fuel and air, ignited by a spark plug 16, burn within combustor 12. The hot products of combustion and heated excess air pass through a transition duct 18 to the inlet end of a turbine stage 20.
  • Combustor 12 and transition duct 18 are contained within a plenum 22 to which a supply of compressed air is fed from a compressor outlet 24 of gas turbine engine 10.
  • Compressed air from compressor outlet 24 flows along the surface of combustor 12 where it is admitted to the interior of combustor 12 through conventional apertures (not shown) in the surface thereof.
  • the air thus admitted to the interior of combustor 12 enters into the combustion reaction downstream of fuel nozzle 14 or may be directed as a cooling film along the inner surface of combustor 12.
  • Some compressed air may also be employed for diluting the hot gas to control and profile the temperature of the effluent of combustor 12.
  • a flow sleeve 26 may be provided surrounding combustor 12 for improving the flow of air along the walls thereof.
  • transition duct 18 The outside surface of transition duct 18 is convectively cooled by compressed air flowing from the compressor outlet 24 toward combustor 12.
  • a radially inner surface 28 of transition duct 18 is disposed in the direct flow of compressed air as it changes direction after exiting compressor outlet 24.
  • a portion 30 of radially inner surface 28 nearer a combustor end 32 of transition duct 18 is more than adequately cooled.
  • a portion 34 of radially inner surface 28 nearer a turbine end 36 is cooled less strongly.
  • a radially outer surface 38 of transition duct 18 is protected from the direct flow of compressed air from compressor outlet 24.
  • a portion 40 of radially outer surface 38 nearer combustor end 32 is cooled by compressed air flowing about the circumference of transition duct 18 on its way to combustor 12. Such cooling is substantially less effective than that experienced by radially inner surface 28.
  • a portion 42 of radially outer surface 38 nearer turbine end 36 is most poorly cooled since very little compressed air circulates therepast. Thus, the cooling effectiveness on transition duct 18 tends to decrease from combustor end 32 to turbine end 36.
  • the cooling problem on portion 42 is additionally complicated by the fact that the hot gas flowing within transition duct 18 is strongly turned in this region. Thus, highly effective convective heat transfer from the hot gas operates on portion 42.
  • portion 42 becomes the hottest part of transition duct 18 and provides the effective limit on the temperature of the hot gas which can be admitted thereto from combustor 12.
  • the resulting unequal temperatures on transition duct 18 may set up troublesome thermal expansion patterns and possibly cause premature failure of transition duct 18.
  • portions 34 and 42 near turbine end 36 of transition duct 18 are less robust than are portions 30 and 40 near combustor end 32, and are thus less capable of withstanding higher temperatures. At least part of this reduction in robustness ensues from the connection of an aft support 44 to portion 42.
  • the temperatures of portions 30 and 40 should be approximately equal and may be permitted to rise substantially higher than the temperatures of portions 34 and 42.
  • the temperatures of portions 34 and 42 should be approximately equal.
  • FIG. 2 there is shown a plate 46 whose surface is to be cooled by impingement cooling.
  • An impingement plate 48 spaced from the surface of plate 46, is pierced by a plurality of holes 50, 52 and 54.
  • a closed end 56 bridges plate 46 and impingement plate 48 forms a chamber 58.
  • fAn exit 60 in chamber 58 provides the only opening through which all air injected through holes 50, 52 and 54 must exit.
  • impingement plate 48 a pressure drop across impingement plate 48 is effective to produce air jets flowing through holes 50, 52 and 54.
  • Hole 50 being closest to closed end 56, forms an impingement jet which impinges on plate 46.
  • the air from hole 50 After impinging on plate 46, the air from hole 50 must flow toward exit 60 as indicated by an air flow arrow 62.
  • Air in the impingement jet formed by hole 52 whose flow is indicated by an air flow arrow 64, must penetrate the crossflow created by the air injected by hole 50. Assuming that the volumes of air injected into chamber 58 by holes 50 and 52 are equal, then the volume of air formed in the combined air flows from holes 50 and 52 is twice the volume from hole 50 alone.
  • the combined air flow downstream of hole 52 has twice the volume and twice the velocity of the crossflow air in air flow arrow 62 arriving at hole 52.
  • This combined volume forms the crossflow through which hole 54 must project its jet upon plate 46.
  • the total air passing downstream of hole 54 has thrice the velocity of that upstream of hole 52.
  • the embodiment of the invention shown in Fig. 3A permits tailoring the cooling to produce a desired temperature pattern on transition duct 18.
  • An impingement sleeve 66 surrounding, and spaced from, transition duct 18 forms a flow volume 68 therebetween which is substantially sealed at turbine end 36 and is open at combustor end 32 thereof.
  • Impingement sleeve 66 is pierced by a plurality of apertures 70 for training a plurality of impingement jets which impinge upon transition duct 18.
  • the spent impingement air since the spent impingement air must all flow toward an exit 72 at combustor end 32, its massflow must increase systematically toward exit 72.
  • the overall pressure drop across the impingement sleeve or the difference between the pressure in plenum 22 (the compressor discharge pressure) and that at exit 72 of flow volume 68. For example, it may be desirable to limit this pressure drop to less than two percent of the compressor discharge pressure.
  • the overall pressure drop through impingement sleeve 66 results from the accumulation of the pressure drop across apertures 70 and the pressure required to accelerate the spent impingement air up to the crossflow velocity in flow volume 68.
  • the velocity of a gas flowing in an enclosed channel varies inversely as the cross-sectional area of the channel.
  • the height of flow volume 68 increases from turbine end 36 to combustor end 32. This tends to reduce the air flow velocity near exit 72 compared to the velocity the air would attain if the smaller height of flow volume 68 were continued throughout its length. This permits taking advantage of a small height of flow volume 68 near turbine end 36 where the crossflow mass flow rate is small, while still limiting the velocity of the cross flow nearer exit 72.
  • the apertures 70 in the first band of apertures about impingement sleeve 66 adjacent turbine end 36 are shown much more closely spaced than are those in the last band of apertures 70 adjacent exit 72. Also, the spacing between the first two bands of apertures at turbine end 36 is much smaller than the spacing between the last two bands of apertures adjacent exit 72. Systematic variation in hole-to-hole and band-to-band spacing is seen at intermediate points.
  • apertures 70' in flow sleeve 26 permit that portion of the combustor air flow which does not pass through impingement sleeve 66 to combine with the impingement air flow spent prior to commencing combustion.
  • the number, size and distribution of apertures 70' are selected to permit the desired airflow, and create the required overall pressure drop for the impingement sleeve.
  • a seal 73 between flow sleeve 26 and impingement sleeve 66 permits considerable misalignment therebetween while preventing air flow from entering at their junction. Such entry would imbalance the air flow split between them.
  • FIG. 3B An alternate embodiment of the invention shown in Fig. 3B is quite similar to that shown in Fig. 3A. The principal difference is in the configuration of flow sleeve 26 and the junction between exit end 32 of impingement sleeve 66 and flared entry portion 74 of slow sleeve 26. An enlarged view of this junction is shown in Fig. 4, in which exit 72 is surrounded by a flared entry portion 74 of flow sleeve 26, creating an annular flow passage 78. Annular flow passage 78 takes the place of apertures 70' (Fig. 3A) having an area calculated to permit the required air flow to pass while creating the required overall pressure drop for impingement sleeve 66.
  • This embodiment requires precise control of the size of annular flow passage 78 in order to achieve consistent flow split and pressure drop performance among ten or more combustors operating in parallel, as is the case in a conventional or advanced heavy duty gas turbine engine.
  • aft support 44 includes a generally circular wall 80 welded at substantially its entire perimeter to transition duct 18 and extending through a circular opening 82 in impingement sleeve 66, thus forming a blind cup-shaped volume 84 which is open to plenum 22 at its upper end but which is substantially closed at its lower end.
  • a complete disclosure of the structure and function of aft support 44 is contained in U.S. Patent No. 4,422,288 whose disclosure is incorporated herein by reference.
  • transition duct 18 is curved outward toward cup-shaped volume 84 in this cross section.
  • the following disclosed technique for providing cooling to the portion of transition duct 18 which is enclosed within circular wall 80 provides an excellent example of the power and flexibility for tailoring the impingement cooling of a surface over which differences in heat load, distance and air cross-flow volume are all encountered.
  • An impingement insert 86 having an upward-directed wall 90 and a planar bottom 92 is tightly fitted into cup-shaped volume 84 with planar bottom 92 spaced from the surface of transition duct 18.
  • Upward-directed wall 90 preferably includes a flange 94 at its upper extremity for attachment to the inner surface of circular wall 80.
  • Flange 94 is preferably attached to circular wall 80 using, for example, welding.
  • An annular space 96 between upward-directed wall 90 and circular wall 80 permits insert 86 and wall 90 to reach the same temperature before they are joined at flange 94 thus minfmizing the thermal stress at this joint.
  • a plurality of apertures 98 in planar bottom 92 permit the pressurized air in plenum 22 to form impingement jets for cooling an enclosed surface 100 of transition duct 18 within circular wall 80.
  • film cooling apertures 102 are disposed in two staggered rows lU4 and 106 located near the upstream edge of planar bottom 92 with respect to the gas flow within transition duct 18. As best illustrated in Fig. 6. film cooling apertures 102 are inclined in the direction of gas flow thereby encouraging film cooling of the inner surface of transition duct 13 by the air passing therethrough. Such film cooling strongly modifies the local heat load downstream of film cooling apertures 102.
  • Apertures 98 are arranged in nine rows 108-124, each aligned transverse to the gas-flow path.
  • the three apertures 98 closest to the center of each of rows 114, 116 and 118 are of relatively small diameter. This smallness is in response to two factors, 1) this region of enclosed surface 100 is strongly film cooled by film cooling apertures 102, and 2) planar bottom 92 and enclosed surface 100 are spaced relatively close together, as seen in the cross section through row 116 in Fig. 5.
  • the outer three apertures 98 in rows 114, 116 and 118 become progressively larger in response to the increasing distance over which the impingement jets must be projected (see Fig. 5).
  • Rows 108 and 124 contain apertures 98 of intermediate size and closest spacing. This is in response to the combination of the shorter distance between planar bottom 92 and enclosed surface 100 in these locations (see Fig. 6) as well as the fact that there are no upstream impingement jets to produce a crossflow to interfere with the projection of cooling air upon enclosed surface 100.
  • Row 110 and 122 contain apertures 98 of larger size and wider spacing to compensate for the presence of crossflow from upstream impingement jets as well as the increasing distance (see Fig. 6).
  • the present invention is capable of tailoring the cooling provided by impingement jet cooling over an area where the three variables of heat load, distance and air crossflow are present in independent fields over the areas of interest.
  • air crossflow velocity is controlled by purposely increasing the distance between transition duct 18 and impingement sleeve 66 and compensating for the increased distance by increasing the diameters of apertures 70.
  • the spacing of the larger-diameter apertures 70 is increased to control the air mass flow density.
  • the distance is generally fixed by the design of transition duct 18.
  • the varying distances are accommodated by suitably controlling the diameter and spacing of apertures 98. Additionally, the problem of disposing of the spent impingement air is solved by employing the spent impingement air for film cooling and by further modifying the diameter and spacing of apertures 98 to compensate for the resulting variation in the heat load over enclosed surface 100.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Transition And Organic Metals Composition Catalysts For Addition Polymerization (AREA)
  • Devices That Are Associated With Refrigeration Equipment (AREA)
  • Heat Treatments In General, Especially Conveying And Cooling (AREA)
EP19860106295 1985-05-14 1986-05-07 Prallkühlung für einen Turbineneinlasskanal Expired - Lifetime EP0203431B2 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US73401885A 1985-05-14 1985-05-14
US734018 1985-05-14

Publications (3)

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EP0203431A1 true EP0203431A1 (de) 1986-12-03
EP0203431B1 EP0203431B1 (de) 1990-11-22
EP0203431B2 EP0203431B2 (de) 1996-05-22

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EP19860106295 Expired - Lifetime EP0203431B2 (de) 1985-05-14 1986-05-07 Prallkühlung für einen Turbineneinlasskanal

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EP (1) EP0203431B2 (de)
JP (1) JPS629157A (de)
AU (1) AU593551B2 (de)
CA (1) CA1263243A (de)
DE (1) DE3675690D1 (de)
NO (1) NO162887C (de)

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EP0239020A2 (de) * 1986-03-20 1987-09-30 Hitachi, Ltd. Gasturbinenbrennkammer
EP0273126A1 (de) * 1986-11-25 1988-07-06 General Electric Company Gasturbinenbrennkammer
EP0599055A1 (de) * 1992-11-27 1994-06-01 Asea Brown Boveri Ag Gasturbinenbrennkammer
GB2285503A (en) * 1993-12-22 1995-07-12 Snecma Combustion chamber having a multi-perforated wall
DE19720786A1 (de) * 1997-05-17 1998-11-19 Abb Research Ltd Brennkammer
GB2328011A (en) * 1997-08-05 1999-02-10 Europ Gas Turbines Ltd Combustor for gas or liquid fuelled turbine
EP1146289A1 (de) * 2000-04-13 2001-10-17 Mitsubishi Heavy Industries, Ltd. Kühlstruktur für das Endstück einer Gasturbinenbrennkammer
GB2372093A (en) * 2000-12-22 2002-08-14 Alstom Power Nv Arrangement for cooling a component
EP1650503A1 (de) * 2004-10-25 2006-04-26 Siemens Aktiengesellschaft Verfahren zur Kühlung eines Hitzeschildelements und Hitzeschildelement
EP1850070A2 (de) * 2006-04-24 2007-10-31 General Electric Company Verfahren und System zur Verringerung von Druckverlusten in Gasturbinentriebwerken
WO2009103636A1 (de) * 2008-02-20 2009-08-27 Alstom Technology Ltd. Thermische maschine
EP2738469A1 (de) * 2012-11-30 2014-06-04 Alstom Technology Ltd Gasturbinenteil mit wandnaher Kühlanordnung
EP3048250A1 (de) * 2015-01-20 2016-07-27 United Technologies Corporation Kühlsystem mit überkühlter luft mit ringförmigem mischkanal
US10415478B2 (en) 2015-01-20 2019-09-17 United Technologies Corporation Air mixing systems having mixing chambers for gas turbine engines

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US5157188A (en) * 1985-03-19 1992-10-20 Phillips Petroleum Company Methane conversion
CA1309873C (en) * 1987-04-01 1992-11-10 Graham P. Butt Gas turbine combustor transition duct forced convection cooling
GB2221979B (en) * 1988-08-17 1992-03-25 Rolls Royce Plc A combustion chamber for a gas turbine engine
US6494044B1 (en) * 1999-11-19 2002-12-17 General Electric Company Aerodynamic devices for enhancing sidepanel cooling on an impingement cooled transition duct and related method
US7010921B2 (en) * 2004-06-01 2006-03-14 General Electric Company Method and apparatus for cooling combustor liner and transition piece of a gas turbine
US7878002B2 (en) * 2007-04-17 2011-02-01 General Electric Company Methods and systems to facilitate reducing combustor pressure drops
US8474266B2 (en) * 2009-07-24 2013-07-02 General Electric Company System and method for a gas turbine combustor having a bleed duct from a diffuser to a fuel nozzle
US8516822B2 (en) * 2010-03-02 2013-08-27 General Electric Company Angled vanes in combustor flow sleeve
US8252251B2 (en) * 2010-03-30 2012-08-28 General Electric Company Fluid cooled reformer and method for cooling a reformer
US8359867B2 (en) * 2010-04-08 2013-01-29 General Electric Company Combustor having a flow sleeve
US8276391B2 (en) * 2010-04-19 2012-10-02 General Electric Company Combustor liner cooling at transition duct interface and related method
US9506359B2 (en) * 2012-04-03 2016-11-29 General Electric Company Transition nozzle combustion system
US9447971B2 (en) * 2012-05-02 2016-09-20 General Electric Company Acoustic resonator located at flow sleeve of gas turbine combustor
WO2020092896A1 (en) * 2018-11-02 2020-05-07 Chromalloy Gas Turbine Llc System and method for providing compressed air to a gas turbine combustor

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GB849255A (en) * 1956-11-01 1960-09-21 Josef Cermak Method of and arrangements for cooling the walls of combustion spaces and other spaces subject to high thermal stresses
US3384346A (en) * 1966-02-01 1968-05-21 Rolls Royce Aerofoil shaped blade for a fluid flow machine such as a gas turbine engine
US3652181A (en) * 1970-11-23 1972-03-28 Carl F Wilhelm Jr Cooling sleeve for gas turbine combustor transition member
US3806276A (en) * 1972-08-30 1974-04-23 Gen Motors Corp Cooled turbine blade
FR2221020A5 (de) * 1973-03-09 1974-10-04 Gen Electric
FR2311176A1 (fr) * 1975-05-16 1976-12-10 Bbc Brown Boveri & Cie Ailette refroidie de turbine
US4339925A (en) * 1978-08-03 1982-07-20 Bbc Brown, Boveri & Company Limited Method and apparatus for cooling hot gas casings
GB2104965A (en) * 1981-08-31 1983-03-16 Gen Electric Multiple-impingement cooled structure
GB2112869A (en) * 1981-12-31 1983-07-27 Westinghouse Electric Corp Cooled airfoil

Cited By (30)

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EP0239020A3 (en) * 1986-03-20 1989-01-18 Hitachi, Ltd. Gas turbine combustion apparatus
US4872312A (en) * 1986-03-20 1989-10-10 Hitachi, Ltd. Gas turbine combustion apparatus
EP0239020A2 (de) * 1986-03-20 1987-09-30 Hitachi, Ltd. Gasturbinenbrennkammer
EP0273126A1 (de) * 1986-11-25 1988-07-06 General Electric Company Gasturbinenbrennkammer
EP0599055A1 (de) * 1992-11-27 1994-06-01 Asea Brown Boveri Ag Gasturbinenbrennkammer
DE4239856A1 (de) * 1992-11-27 1994-06-01 Asea Brown Boveri Gasturbinenbrennkammer
US5388412A (en) * 1992-11-27 1995-02-14 Asea Brown Boveri Ltd. Gas turbine combustion chamber with impingement cooling tubes
GB2285503A (en) * 1993-12-22 1995-07-12 Snecma Combustion chamber having a multi-perforated wall
GB2285503B (en) * 1993-12-22 1998-02-18 Snecma Combustion chamber having a multi-perforated wall
US6106278A (en) * 1997-05-17 2000-08-22 Abb Research Ltd. Combustion chamber
DE19720786A1 (de) * 1997-05-17 1998-11-19 Abb Research Ltd Brennkammer
US6134877A (en) * 1997-08-05 2000-10-24 European Gas Turbines Limited Combustor for gas-or liquid-fuelled turbine
EP0896193A3 (de) * 1997-08-05 2000-07-26 European Gas Turbines Limited Gasturbinenbrennkammer
EP0896193A2 (de) * 1997-08-05 1999-02-10 European Gas Turbines Limited Gasturbinenbrennkammer
GB2328011A (en) * 1997-08-05 1999-02-10 Europ Gas Turbines Ltd Combustor for gas or liquid fuelled turbine
EP1146289A1 (de) * 2000-04-13 2001-10-17 Mitsubishi Heavy Industries, Ltd. Kühlstruktur für das Endstück einer Gasturbinenbrennkammer
US6553766B2 (en) * 2000-04-13 2003-04-29 Mitsubishi Heavy Industries, Ltd. Cooling structure of a combustor tail tube
GB2372093A (en) * 2000-12-22 2002-08-14 Alstom Power Nv Arrangement for cooling a component
US6615588B2 (en) 2000-12-22 2003-09-09 Alstom (Switzerland) Ltd Arrangement for using a plate shaped element with through-openings for cooling a component
GB2372093B (en) * 2000-12-22 2005-06-15 Alstom Power Nv Arrangement for cooling a component
EP1650503A1 (de) * 2004-10-25 2006-04-26 Siemens Aktiengesellschaft Verfahren zur Kühlung eines Hitzeschildelements und Hitzeschildelement
EP1850070A2 (de) * 2006-04-24 2007-10-31 General Electric Company Verfahren und System zur Verringerung von Druckverlusten in Gasturbinentriebwerken
EP1850070A3 (de) * 2006-04-24 2014-08-06 General Electric Company Verfahren und System zur Verringerung von Druckverlusten in Gasturbinentriebwerken
WO2009103636A1 (de) * 2008-02-20 2009-08-27 Alstom Technology Ltd. Thermische maschine
US8272220B2 (en) 2008-02-20 2012-09-25 Alstom Technology Ltd Impingement cooling plate for a hot gas duct of a thermal machine
EP2738469A1 (de) * 2012-11-30 2014-06-04 Alstom Technology Ltd Gasturbinenteil mit wandnaher Kühlanordnung
US9945561B2 (en) 2012-11-30 2018-04-17 Ansaldo Energia Ip Uk Limited Gas turbine part comprising a near wall cooling arrangement
EP3048250A1 (de) * 2015-01-20 2016-07-27 United Technologies Corporation Kühlsystem mit überkühlter luft mit ringförmigem mischkanal
US10100738B2 (en) 2015-01-20 2018-10-16 United Technologies Corporation Overcooled air cooling system with annular mixing passage
US10415478B2 (en) 2015-01-20 2019-09-17 United Technologies Corporation Air mixing systems having mixing chambers for gas turbine engines

Also Published As

Publication number Publication date
JPS629157A (ja) 1987-01-17
NO162887B (no) 1989-11-20
JPH0524337B2 (de) 1993-04-07
NO861900L (no) 1986-11-17
CA1263243A (en) 1989-11-28
EP0203431B2 (de) 1996-05-22
DE3675690D1 (de) 1991-01-03
AU5735386A (en) 1986-11-20
AU593551B2 (en) 1990-02-15
EP0203431B1 (de) 1990-11-22
NO162887C (no) 1990-02-28

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