US3353351A - Aerofoil-shaped fluid-cooled blade for a fluid flow machine - Google Patents

Aerofoil-shaped fluid-cooled blade for a fluid flow machine Download PDF

Info

Publication number
US3353351A
US3353351A US508683A US50868365A US3353351A US 3353351 A US3353351 A US 3353351A US 508683 A US508683 A US 508683A US 50868365 A US50868365 A US 50868365A US 3353351 A US3353351 A US 3353351A
Authority
US
United States
Prior art keywords
fluid
blade
wall
sleeve
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.)
Expired - Lifetime
Application number
US508683A
Other languages
English (en)
Inventor
Bill Arthur
Steel Thomas
Poucher Michael
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 PLC
Original Assignee
Rolls Royce PLC
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 PLC filed Critical Rolls Royce PLC
Application granted granted Critical
Publication of US3353351A publication Critical patent/US3353351A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

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
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/16—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades

Definitions

  • ABSTRACT F THE DISCLOSURE A hollow aerofoil-shaped blade for use, for example, as a flame tube-mounted guide vane in a gas turbine engine, the blade having an internal hollow sleeve which is spaced from the blade wall.
  • the sleeve is provided with cooling air outlets for directing some of the cooling fluid supplied to the interior thereof to the space between the sleeve and the blade wall, this fluid subsequently escaping to the exterior of the blade, and is further provided with passages which provide direct communication between the sleeve and the blade wall so that the remainder of the fluid supplied to the sleeve may pass to the exterior of the blade without mixing with fluid in the intervening space.
  • the present invention relates to an aerofoil-shaped blade of type adapted for use in a fluid-flow machine such as a gas turbine engine or a steam turbine.
  • an aerofoil-shaped blade adapted for use in a fluid flow machine, said blade having an external wall, characterised by fluid receiving means surrounded by and spaced from said wall for receiving cooling fluid and directing some of the cooling fluid over the interior of said wall for cooling said wall, said wall having apertures enabling the cooling fluid to escape from the space between said fluid receiving means and said wall, and means defining a passage wherein a portion of the cooling fluid entering the fluid receiving means passes across the said space directly to the exterior of the said blade without mixing with the fluid in the space.
  • the fluid receiving means may comprise a hollow sleeve having apertures for directing said cooling fluid towards a part of said wall which tends to be relatively hot during operation of the engine, and towards the trailing edge of said blade, said wall having apertures in said trailing edge enabling said cooling fluid to escape from the said space.
  • one or more partitions which divide the hollow sleeve into one or more first portions having apertures for directing said cooling fluid into the said space and one or more second portions from which fluid may pass directly from the sleeve to the exterior of the blade through said passage defining means.
  • the said one or more partitions are preferably provided with extensions which project from the sleeve and which are adapted to be disposed in a stream of the fluid for directing fluid from the stream into the sleeve.
  • passage-defining means Preferably there are more than one passage-defining means and means are provided for ensuring a substantially even distribution of air entering said second portion or portions and passing therefrom to the passage-defining means.
  • deflectors each of which projects from the sleeve and which is adapted to be disposed in a stream of the fluid for directing fluid from the stream into a respective second portion of the sleeve.
  • the blade according to the invention may also be provided with spaced-apart flat strips on the leading edge thereof which form with the blade a channel for deflecting a portion of the fluid, which in operation will pass along a path exteriorly of the blade generally from the leading edge to the trailing edge thereof.
  • a fluid-flow machine such as a gas turbine engine, having an inlet at one end for primary combustion air, means for dispersing fuel in said primary combustion air and an outlet at the other end for the combustion gases, and an aerofoil-shaped blade as previously described disposed in said combustion chamber between the inlet and outlet for guiding said combustion gases to said outlet and arranged so that the air from the passage-defining means fOXIILv the secondary combustion air.
  • the invention further provides a fluid-flow machine, such as a gas turbine engine comprising one or more aerofoil-shaped blades as previously described.
  • a fluid-flow machine such as a gas turbine engine comprising one or more aerofoil-shaped blades as previously described.
  • FIGURE 1 is a sectional view of part of a gas turbine engine having aerofoil-shaped blades in accordance with the present invention
  • FIGURE 2 is a perspective cut-away view of the combustion chamber of part of the gas turbine engine of FIGURE 1, and
  • FIGURE 3 is a perspective cut-away view of a blade in accordance with the invention.
  • FIGURE 1 there is shown on one side of the centreline 10 of a gas turbine engine, an axial air compressor 11, a combustion chamber 12 and a power producing turbine 13.
  • a pipe 14 is provided for the supply of fuel to the combustion chamber 12, the fuel being sprayed, in operation against a splash plate 15, and transversely dispersed from a curled end 16 of the splash plate 15 into the air stream from the compressor 11.
  • the air delivered by compressor 11 is divided into a first air stream which enters the combustion chamber 12 at inlet 17 and a second air stream which proceeds in the general direction of arrow 18.
  • the air-fuel mixture is ignited by suitable means (not shown) and the resulting combustion gases expand towards outlet 19 of the combustion chamber 12.
  • T 0 ensure that the temperature of the combustion gases is maintained sufliciently low to avoid damage to the downstream parts of the combustion chamber 12 and the turbine 13, it is preferred to admit cooling or dilution air to the combustion chamber immediately downstream of the zone of mixing of the primary air and fuel. It is also desirable that the blades 20 be cooled against the high temperatures resulting from combustion.
  • the invention enables the blades 20 to be cooled by means of a portion of the second air stream, and accordingly the blades 20 are made hollow and the second air stream is directed into the blades 20 by means of scoops 21, 22 which project thereinto.
  • the main part of the second air stream can leave the interior of each blade 20 through apertures 23, and the remainder after circulating within the blade 20 to cool it leaves through smaller apertures 25 at the trailing edge of blade 20.
  • the apertures 23 are all disposed on one side only of blade 20 but the invention is not limited to this feature, and the apertures may be arranged on both sides if desired.
  • the radially outer 2nd of the leading edge of blade 26 is provided with two apaced-apart strips 26 forming a channel with the leading edge such that a portion of the incoming primary air, which may be a major portion, is deflected radially inwardly transverse to the general flow path of the gases, whereby to promote mixing.
  • the strips 26 in the illustrated embodiment are arranged generally parallel to each other but they may also be disposed to be non-parallel if this is found to provide the desired gas circulation.
  • the combustion chamber 12 is, in this instance, of annular form and is divided into fairly distinct combustion regions by a number of angularly spaced-apart blades 20, the primary and secondary air-streams being divided from the compressor delivery stream alternately angularly around the annular outlet from the compressor 11.
  • Each blade 20 is provided internally with a hollow sleeve 27 the open top portion of one of which can be seen in FIGURE 2 and more detail of which can be seen in FIG. 3.
  • the radially inner portion of the sleeve 27 is sealed to the radially inner wall of combustion chamber 12.
  • the scoop 21 forms an outwardly projecting extension of a partition which extends to the radially inner end of sleeve 27 thus dividing sleeve 27 into tWo chambers 23, 29, and the shape of the scoop 21 is such as to direct the air which strikes it into chamber 2%.
  • Scoop 22 forms an outwardly projecting extension of a partition which extends only part way towards the radially inner end of sleeve 27, so that the air which strikes it can be directed into the radially outer part of compartment 29 to feed the radially outer ones of the apertures 23 through tubes 31, while the air which does not strike it enters the downstream portion of compartment 29 due to the termination of the second air passage and passes in a generally clockwise direction as shown in FIGURE 3 to the radially inner ones of apertures 23 through corresponding tubes 31.
  • the air entering compartment 29 is distributed fairly evenly among the apertures 23.
  • the main portion of the second air stream is directed into chamber 29 by scoop 22, and thus, without mixing with cooling air in the space between the sleeve 27 and the blade 20, passes directly out of sleeve 27 to the combustion chamber 12 through tubes 31, which terminate in the apertures 23 in the exterior wall 20a of blade 20.
  • the tubes 31 also serve to support the sleeve 27 in its spaced-apart relationship to the wall 20a.
  • the second air stream which gathers in compartment 28 of sleeve 27 is used for cooling the exterior wall 26a before passing to the combustion chamber 12.
  • the leading edge of sleeve 27 is formed with apertures 32, so that cooling air can pass from chamber 28 to the space betwen the sleeve 27 and the leading edge of blade 2%), this leading edge tending to be the hottest part of blade 20 during the operation of the engine.
  • the cooling air will circulate from the leading edge sleeve apertures 32 around the sleeve 27, and this air after having helped to cool the wall 20a of the blade 20 by circulating in the space between the sleeve 27 and the wall 20a of the blade 20, and escapes from blade 20 via apertures 25.
  • apertures 25 and 32 have been depicted as slots in the drawings, they could equally well be any other suitable shape, such as circular. It is also to be understood that the invention is not limited to two scoops 21, 22 and two chambers 23, 29 but that any suitable number of scoops and chambers may be used. Further, Where the operational pressure conditions in the combustion chamber permit, it is possible that the leading edge regions of the wall 20a of the blade 20 could be apertures to allow the escape of some cooling air from the space between the sleeve 27 and the leading edge of blade 20, this air then passing as a film rearwardly over the external surface of blade 2% to protect the blade 20 against direct convection from the combustion gases.
  • spaced-apart strips 26 are joined to each other by a web to form a channel which is fixedly disposed slightly upstream of the leading edge of the blade, and the apertures for this film cooling air are formed in the wall 20a of blade 2! on the leading edge of blade 20 so that the air can pass between the leading edge and the channel.
  • An aerofoil-shaped blade adapted for use in a fluid flow machine, said blade having an external wall and fluid receiving means surrounded by and spaced from said wall to define a space for receiving cooling fluid from the fluid receiving means and directing some of the cooling fluid over the interior of said Wall for cooling said wall, said Wall having apertures enabling the cooling fluid to escape from the space between said fluid receiving means and said wall, and means defining a passage wherein a portion of the cooling fluid entering the fluid receiving means passes across the said space directly to the exterior of the said blade without mixing with the fluid in the space.
  • a blade according to claim 1 in which the fluid receiving means comprises a hollow sleeve having apertures for directing said cooling fluid into said space towards a part of said wall which tends to be relatively hot during operation of the engine, and towards the trailing edge of the blade, said wall apertures being in the downstream portion of said blade.
  • a blade according to claim 2 in which there is provided at least one partition which divides the sleeve into a first portion including the apertures in said sleeve for directing said cooling fluid into said space, and at least one second portion from which fluid may pass directly from the sleeve to the exterior of the blade through said passage defining means.
  • a blade according to claim 3 in which said partition is provided with an extension which projects from said sleeve and which is adapted to be disposed in a stream of the fluid for directing fluid from the stream into the sleeve.
  • a blade according to claim 2 in which there is provided a partition dividing the sleeve into tWo portions, said passage defining means being arranged for fluid to escape from the downstream one of the portions to the exterior of the blade, the part of the sleeve around the other portion containing said apertures for directing the cooling fluid into said space.
  • a blade according to claim 3 in which there are more than one passage defining means and means are provided for ensuring a substantially even distribution of air entering said second portion and passing therefrom to each passage defining means.
  • a blade according to claim 6 in which there is at least one deflector which projects from the sleeve and which is adapted to be disposed in the stream of the fluid for directing fluid from the stream into a respective second portion of the sleeve.
  • a blade according to claim 2 in which means are provided for permitting the cooling fluid to circulate in said space from a region adjacent the upstream edge of said blade to the apertures in the downstream edge of the external wall.
  • a blade according to claim 1 in which the passage defining means serves to support the fluid receiving means in fixed spaced relation to said external wall.
  • a blade according to claim 1 in which the upstream edge thereof is provided with spaced apart flat strips which form with the blade 2. channel for deflecting a portion of the fluid, which portion passes along a path exteriorly of the blade generally from the upstream leading edge to the downstream edge thereof.
  • a combustion chamber for a fluid flow machine having an inlet at one end for primary combustion air

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US508683A 1964-12-02 1965-11-19 Aerofoil-shaped fluid-cooled blade for a fluid flow machine Expired - Lifetime US3353351A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB49101/64A GB1034260A (en) 1964-12-02 1964-12-02 Aerofoil-shaped blade for use in a fluid flow machine

Publications (1)

Publication Number Publication Date
US3353351A true US3353351A (en) 1967-11-21

Family

ID=10451151

Family Applications (1)

Application Number Title Priority Date Filing Date
US508683A Expired - Lifetime US3353351A (en) 1964-12-02 1965-11-19 Aerofoil-shaped fluid-cooled blade for a fluid flow machine

Country Status (5)

Country Link
US (1) US3353351A (de)
DE (2) DE1280618B (de)
FR (1) FR1454951A (de)
GB (2) GB1034260A (de)
MY (1) MY6800109A (de)

Cited By (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3412979A (en) * 1965-12-10 1968-11-26 Rolls Royce Aerofoil-shaped blade for use in a fluid flow machine such as a turbine
US4967562A (en) * 1988-12-12 1990-11-06 Sundstrand Corporation Turbine engine with high efficiency fuel atomization
US4967563A (en) * 1988-12-12 1990-11-06 Sundstrand Corporation Turbine engine with high efficiency fuel atomization
US4989404A (en) * 1988-12-12 1991-02-05 Sundstrand Corporation Turbine engine with high efficiency fuel atomization
US5027603A (en) * 1988-12-28 1991-07-02 Sundstrand Corporation Turbine engine with start injector
US5220794A (en) * 1988-12-12 1993-06-22 Sundstrand Corporation Improved fuel injector for a gas turbine engine
US5239818A (en) * 1992-03-30 1993-08-31 General Electric Company Dilution pole combustor and method
US5241818A (en) * 1989-07-13 1993-09-07 Sundstrand Corporation Fuel injector for a gas turbine engine
DE4309131A1 (de) * 1993-03-22 1994-09-29 Abb Management Ag Verfahren und Vorrichtung zur Nachlaufbeeinflussung bei Brennkammereinbauten
EP0848210A2 (de) * 1996-12-13 1998-06-17 Asea Brown Boveri AG Brennkammer mit integrierten Leitschaufeln
EP1270874A1 (de) * 2001-06-18 2003-01-02 Siemens Aktiengesellschaft Gasturbine mit einem Verdichter für Luft
CN104411982A (zh) * 2012-07-06 2015-03-11 斯奈克玛 具有改进叶片轮廓的涡轮机引导叶片
US20160160685A1 (en) * 2009-03-04 2016-06-09 United Technologies Corporation Eliminatin of unfavorable outflow margin
CN104411982B (zh) * 2012-07-06 2016-11-30 斯奈克玛 具有改进叶片轮廓的涡轮机引导叶片
US20170051678A1 (en) * 2015-08-18 2017-02-23 General Electric Company Mixed flow turbocore
US10578028B2 (en) 2015-08-18 2020-03-03 General Electric Company Compressor bleed auxiliary turbine
US20200182068A1 (en) * 2018-12-05 2020-06-11 United Technologies Corporation Axial flow cooling scheme with structural rib for a gas turbine engine
US10822963B2 (en) 2018-12-05 2020-11-03 Raytheon Technologies Corporation Axial flow cooling scheme with castable structural rib for a gas turbine engine

Families Citing this family (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3726604A (en) * 1971-10-13 1973-04-10 Gen Motors Corp Cooled jet flap vane
DE4336143C2 (de) * 1993-10-22 1995-11-16 Erich Wuerzinger Kühlverfahren für Turbomaschinen
US7070386B2 (en) 2004-06-25 2006-07-04 United Technologies Corporation Airfoil insert with castellated end
WO2006053825A1 (de) * 2004-11-16 2006-05-26 Alstom Technology Ltd Gasturbinenanlage und zugehörige brennkammer
US11603766B1 (en) 2022-05-04 2023-03-14 Pratt & Whitney Canada Corp. Turbine stator vanes having inserts and splitter plates

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2847185A (en) * 1953-04-13 1958-08-12 Rolls Royce Hollow blading with means to supply fluid thereinto for turbines or compressors
US3242674A (en) * 1961-05-05 1966-03-29 Lucas Industries Ltd Liquid fuel combustion apparatus

Family Cites Families (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2641040A (en) * 1948-01-02 1953-06-09 Esther C Goddard Means for cooling turbine blades by air
US2858100A (en) * 1952-02-01 1958-10-28 Stalker Dev Company Blade structure for turbines and the like
DE1055884B (de) * 1954-03-02 1959-04-23 Bristol Aero Engines Ltd Flammrohr fuer eine Brennkammer eines Gasturbinenmotors
US2920866A (en) * 1954-12-20 1960-01-12 A V Roe Canada Ltd Hollow air cooled sheet metal turbine blade
FR1177035A (fr) * 1957-05-28 1959-04-20 Snecma Procédé et dispositif de refroidissement d'organes de machines
GB854135A (en) * 1958-03-05 1960-11-16 Rolls Royce Improvements in or relating to combustion equipment
DE1204021B (de) * 1959-04-27 1965-10-28 Rolls Royce Schaufel fuer Axialstroemungsmaschinen, insbesondere Gasturbinen
GB898368A (en) * 1959-06-23 1962-06-06 Rolls Royce Improved combustion chamber
GB938247A (en) * 1962-03-26 1963-10-02 Rolls Royce Gas turbine engine having cooled turbine blading

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2847185A (en) * 1953-04-13 1958-08-12 Rolls Royce Hollow blading with means to supply fluid thereinto for turbines or compressors
US3242674A (en) * 1961-05-05 1966-03-29 Lucas Industries Ltd Liquid fuel combustion apparatus

Cited By (26)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3412979A (en) * 1965-12-10 1968-11-26 Rolls Royce Aerofoil-shaped blade for use in a fluid flow machine such as a turbine
US4967562A (en) * 1988-12-12 1990-11-06 Sundstrand Corporation Turbine engine with high efficiency fuel atomization
US4967563A (en) * 1988-12-12 1990-11-06 Sundstrand Corporation Turbine engine with high efficiency fuel atomization
US4989404A (en) * 1988-12-12 1991-02-05 Sundstrand Corporation Turbine engine with high efficiency fuel atomization
US5220794A (en) * 1988-12-12 1993-06-22 Sundstrand Corporation Improved fuel injector for a gas turbine engine
US5027603A (en) * 1988-12-28 1991-07-02 Sundstrand Corporation Turbine engine with start injector
US5241818A (en) * 1989-07-13 1993-09-07 Sundstrand Corporation Fuel injector for a gas turbine engine
US5239818A (en) * 1992-03-30 1993-08-31 General Electric Company Dilution pole combustor and method
EP0564183A1 (de) * 1992-03-30 1993-10-06 General Electric Company Brennkammer mit Verdünnungsleitschaufeln
EP0564183B1 (de) * 1992-03-30 1997-07-23 General Electric Company Brennkammer mit Verdünnungsleitschaufeln
DE4309131A1 (de) * 1993-03-22 1994-09-29 Abb Management Ag Verfahren und Vorrichtung zur Nachlaufbeeinflussung bei Brennkammereinbauten
US5438821A (en) * 1993-03-22 1995-08-08 Abb Management Ag Method and appliance for influencing the wake of combustion chamber inserts
EP0848210A2 (de) * 1996-12-13 1998-06-17 Asea Brown Boveri AG Brennkammer mit integrierten Leitschaufeln
EP0848210A3 (de) * 1996-12-13 1999-11-17 Asea Brown Boveri AG Brennkammer mit integrierten Leitschaufeln
EP1270874A1 (de) * 2001-06-18 2003-01-02 Siemens Aktiengesellschaft Gasturbine mit einem Verdichter für Luft
US6672070B2 (en) 2001-06-18 2004-01-06 Siemens Aktiengesellschaft Gas turbine with a compressor for air
CN1328492C (zh) * 2001-06-18 2007-07-25 西门子公司 带有空气压缩机的燃气轮机
US9816394B2 (en) * 2009-03-04 2017-11-14 United Technologies Corporation Eliminatin of unfavorable outflow margin
US20160160685A1 (en) * 2009-03-04 2016-06-09 United Technologies Corporation Eliminatin of unfavorable outflow margin
CN104411982B (zh) * 2012-07-06 2016-11-30 斯奈克玛 具有改进叶片轮廓的涡轮机引导叶片
CN104411982A (zh) * 2012-07-06 2015-03-11 斯奈克玛 具有改进叶片轮廓的涡轮机引导叶片
US20170051678A1 (en) * 2015-08-18 2017-02-23 General Electric Company Mixed flow turbocore
US10578028B2 (en) 2015-08-18 2020-03-03 General Electric Company Compressor bleed auxiliary turbine
US10711702B2 (en) * 2015-08-18 2020-07-14 General Electric Company Mixed flow turbocore
US20200182068A1 (en) * 2018-12-05 2020-06-11 United Technologies Corporation Axial flow cooling scheme with structural rib for a gas turbine engine
US10822963B2 (en) 2018-12-05 2020-11-03 Raytheon Technologies Corporation Axial flow cooling scheme with castable structural rib for a gas turbine engine

Also Published As

Publication number Publication date
MY6800109A (en) 1968-12-31
GB1034260A (en) 1966-06-29
DE1476892B2 (de) 1970-11-12
DE1476892A1 (de) 1970-07-16
FR1454951A (fr) 1966-02-11
GB1068280A (en) 1967-05-10
DE1280618B (de) 1968-10-17

Similar Documents

Publication Publication Date Title
US3353351A (en) Aerofoil-shaped fluid-cooled blade for a fluid flow machine
US5288207A (en) Internally cooled turbine airfoil
US2475911A (en) Combustion apparatus
US3800527A (en) Piloted flameholder construction
US4153386A (en) Air cooled turbine vanes
US3374624A (en) Gas turbine engine combustion equipment
US4193260A (en) Combustion apparatus
US4017207A (en) Gas turbine engine
US4573865A (en) Multiple-impingement cooled structure
US3899876A (en) Flame tube for a gas turbine combustion equipment
US3747345A (en) Shortened afterburner construction for turbine engine
US3299632A (en) Combustion chamber for a gas turbine engine
US3793827A (en) Stiffener for combustor liner
US3088281A (en) Combustion chambers for use with swirling combustion supporting medium
US3898797A (en) Cooling arrangements for duct walls
US2657531A (en) Wall cooling arrangement for combustion devices
JPH04227417A (ja) ガスタービンエンジンの燃焼器
US3656297A (en) Combustion chamber air inlet
JPH0259281B2 (de)
US3730644A (en) Gas turbine engine
JPH04227418A (ja) ガスタービンエンジンの燃焼器
US2722803A (en) Cooling means for combustion chamber cross ignition tubes
US9404379B2 (en) Gas turbine shroud assemblies
US3321912A (en) Gas turbine plant
US2560207A (en) Annular combustion chamber with circumferentially spaced double air-swirl burners