US2510645A - Air nozzle and porting for combustion chamber liners - Google Patents

Air nozzle and porting for combustion chamber liners Download PDF

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Publication number
US2510645A
US2510645A US705866A US70586646A US2510645A US 2510645 A US2510645 A US 2510645A US 705866 A US705866 A US 705866A US 70586646 A US70586646 A US 70586646A US 2510645 A US2510645 A US 2510645A
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liner
air
nozzle
combustion
ports
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US705866A
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English (en)
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Kenton D Mcmahan
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General Electric Co
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General Electric Co
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Priority to FR962862D priority Critical patent/FR962862A/fr
Application filed by General Electric Co filed Critical General Electric Co
Priority to US705866A priority patent/US2510645A/en
Priority to GB28500/47A priority patent/GB635946A/en
Priority to CH265334D priority patent/CH265334A/de
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    • 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
    • F23R3/045—Air inlet arrangements using pipes
    • 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
    • F23R3/06—Arrangement of apertures along the flame tube
    • 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
    • F23R3/60—Support structures; Attaching or mounting means

Definitions

  • This invention relates to ombustion chambers or combustors for generating hot gases under pressure, as for use in thermal powerplants such as gas turbines. More specifically, the invention relates to an improvement in the Nerad type combustor, described generally in patent application Serial No. 501,106, filed September 3, 1943, in the name of Anthony J. Nerad, now abandoned, also a continuation-in-part Serial No. 750,- 015, filed May 23, 1947.
  • the Nerad combustor is so designed that high velocity jets of combustion supporting fluid are directed in an exactly radial direction into the combustion space from inlet ports in a substantially cylindrical liner, which defines the combustion space.
  • a characteristic flow path ' is obtained which includes a tore or smoke ring vortex at the closed end of the liner, the establishment and maintenance of which is essential to eflicient operation of the combustor.
  • a most important factor affecting the tore is the uniformity with which air is supplied to the ports in the liner.
  • the afore-mentioned Nerad application describes various ways to secure the required uniformity of air supply.
  • an object of the invention is to provide an improved pressurized combustor of the Nerad type which is substantially insensitive to variations in the direction from which the combustion air approaches the air inlet ports, and which gives more stable combustion characteristics with uniform temperature distribution and a lower over-all pressure drop through the combustor.
  • Another object is to provide means for insuring a satisfactory flow of air to the inlet openings in the end dome and the liner of the combustor.
  • a further object is to provide an improved liner for a Nerad combustor which will operate .efficiently regardless of variations in the angle at which the combustion air approaches the inlet ports of the liner.
  • my combustor assembly comprises a thin-walled cylindrical member I open at either end. At one end the cylinder l is provided with a flange secured by suitable threaded fastenings 2 to a corresponding flange of an inlet elbow or air adapter 3. While the inlet elbow 3 may be formed as a casting, it is shown in Fig. 1 as fabricated from thin sheet metal provided with a boss 4 defining a threaded opening coaxial with the cylindrical casing l. The elbow 3 constitutes an inlet conduit through which air under pressure enters from a suitable source, such as a compressor (not shown), as indicated by the arrow 5.
  • a suitable source such as a compressor (not shown), as indicated by the arrow 5.
  • a spray nozzle For introducing fluid fuel into the combustion space a spray nozzle is provided, as indicated at 6.
  • This nozzle may be of any suitable type. Where the fuel used is a liquid such as kerosene or gasoline, the so-called duplex nozzle may be used, as described in patent application Serial No. 622,604, filed October 16, 1945, in the names of Charles D. Fulton and David C. Ipsen. Such nozzles require two supply conduits I and 8. It should be understood that many other types of nozzle may also be used; and the specific details of the spray nozzle are not material to an understanding of the present invention.
  • the nozzle tip 9 projects into the inlet elbow 3 coaxial with the casing I, and has a portion threadedly engaging the central. opening in boss 4.
  • the nozzle 6 discharges the fluid fuel in the form of a hollow cone, indicated diagrammatically at 30 in Fig. 1.
  • an inner liner indicated generally at Hi This comprises a generally cylindrical but preferably somewhat tapered casing having its smaller end adjacent the inlet elbow 3.
  • the construction of the inner liner may be seen by referring to Fig. 2 in conjunction with Fig. 1. It is formed by punching a plurality of longitudinal rows of combustion air inlet ports ll, the precise proportions and arrangement of which is specifled more particularly in the above-mentioned Nerad application. Between adjacent rows of combustion air ports II are longitudinal rows of nozzles l2, formed by striking the metal of the liner outwardly as shown in Fig. 1 so as to define cooling air inlet nozzles arranged to admit air in the manner indicated by the arrows l3 in Fig. 1.
  • This cooling air l3 forms a thin continuous film of pure comparatively cool air flowing over the inner surface of the liner it so as to keep hot products of combustion and incompletely burned fuel particles from contact with the relatively cool metal of the liner.
  • This arrangement prevents deposition of carbon particles as described in the above-mentioned Nerad application.
  • a row of spaced struck-out portions or "dimples" ll which serve a purpose noted hereinafter.
  • Parallel to the discharge edge of the liner and spaced intermediate the dimples I4 and the last circumferential row of combustion air inlets II is a row of plain circular ports IS, the purpose of which will also be noted hereinafter.
  • the liner formed as shown in Fig. 2, is rolled upto define a somewhat tapering cylinder as in Fig. 1, the side edges It being secured together by suitable means, as for instance seam-welding.
  • each of the combustion air inlet ports ii is provided with a short nozzle ll projecting radially outward from the liner.
  • the arrangement of these nozzles is indicated in Fig. 1. and a single nozzle is shown to an enlarged scale in Fig. 3.
  • the nozzles may advantageously be secured to the liner by welding to a struck-out flange 22 provided around the edge of each port ll.
  • Nerad combustion chamber in its very simplest form, is somewhat sensitive to changes in the direction from which the air supply approaches the inner liner. For most efiieient and successful'operation of this type of combusto it is essential that the combustion air entering the combustion space defined within the liner l through the ports ll shall form discrete jets with their axes approaching very closely to the radial direction.
  • the air inlet ports I I produce jets having their axes substantially in a plane normal to the longitudinal axis of the liner ll, the axes of the jets from any given circumferential row of ports I i meeting at the axis of the liner.
  • an important object of the invention is to provide means for stabilizing the direction of the combustion air jets entering the combustion space within the liner Ill through the ports H.
  • this may be achieved by providing, the nozzles ll extending a straight cylindrical discharge portion and a well-rounded entrance portion. It is desirable that the discharge edge is of the nozzle lie in a plane exactly normal to the axis I! of the nozzle. Likewise, the axis iii of the nozzle must be normal to the axis of the liner, as indicated in Fig. 1.
  • the radius of curvature of the nozzle inlet may advantageously be of the order of one quarter the inside diameter d of the nozzle. It will be apparent that this nozzle has a contracting inlet portion, and a non-expanding discharge portion. This is important, since it is desired to produce a high velocity jet of good penetrating power.
  • the axial length of the air nozzles II decreases progressively from the closed or fuel nozzle end of the liner to the open discharge end.
  • This arrangement is paraciaeec ticularly advantageous when the liner is ,some-' what tapered and is contained within a straight cylindrical outer housing, as in Fig. 1. It will be apparent that the tapered inner liner ill forms an annular air supply space with the cylindrical outer housing I which decreases progressively in' cross-sectional area along the length of the liner.
  • the nozzles ll may all be of the same axial length, if desired, and if the increase in size and weight is not objectionable.
  • the inlet or fuel nozzle end of the liner I is closed by a hemispherical end dome 25.
  • This dome structure has a central opening surrounded by an axially extending flange 26 adapted to slide snugly over the cylindrical nozzle tip 9.
  • the dome is provided with one or more circumferential rows of nozzles formed by struck-out tongue portions 21. The function of these nozzles is to form a thin film of cooling and insulating air flowing over the inner surface of the dome 25 so as to prevent carbon deposition thereon.
  • the nozzles 21 perform a somewhat similar function for the end dome 25 as is performed for the liner l0 by the cooling air nozzles l2.
  • the cooling and insulating air from the nozzles 21 flows radially inward in the manner of the arrows 28, across the exposed end surfaces of the nozzle tip 9, after which it reverses its direction and flows radially outward, as indicated by the arrow 29.
  • the member 25 is referred to herein as a dome, I intend this term to include generally closures for the fuel nozzle end of the liner of shapes other than the hemispheral configuration shown in Fig. 1, for instance substantially flat disc members, as disclosed in the above-mentioned Nerad application.
  • is provided with a number of spaced air inlet ports 33. These are of such a size, location, and number that the air supplied through the inlet elbow or air adapter 3 will flow uniformly into the space or plenum chamber 33 defined between the end dome 25 and the shroud 38, with substantially no difference in pressure between the air in the elbow 3 and that in the plenum chamber 3i. Thus with substantially no pressure drop across the air inlet openings 33 in the shroud 3
  • is provided with a central opening having a plurality of circumferentially spaced bosses 35 welded around the outside thereof. These bosses are adapted to be secured. by suitable threaded fastenings 36 to the ring 4 provided on the outer wall of elbow 3. It will be seen that the shroud 3
  • the plain circular ports l5 at the discharge end of liner III are trimmer holes, provided for the purpose of facilitating final balancing" of the combustioncharacteristics of the apparatus. It will be observed that the combustion inlet nozzles I1 and the cooling air nozzles l2 are arranged in an entirely uniform symmetrical manner. It may sometimes be found, when the combustor is built and tested, that the temperature distribution in the hot gases discharged from the liner is not exactly uniform. This may be corrected by suitable selection of the size. number, and location of the trimmer holes. since these holes have no associated nozzles corresponding to the nozzles ll, they may be readily punched in any size and arrangement found necessary to achieve completely uniform temperature distribution. The precise arrangement required can be determined only by analysis and testing of a model. 4
  • the discharge end of liner I0 is supported in the following manner.
  • the downstream end of the outer casing l is provided with a flange adapted to be secured by suitable threaded fastenings 31 to a cooperating flange at the upstream end of a discharge elbow indicated generally at 38.
  • an annular member having an outer circumferential portion 39 disposed between the cooperating flanges.
  • the inner circumferential portion of this annular member is spun or otherwise shaped to form an annular convolution 40 and a cylindrical portion 4
  • the dimples l4 formed in the discharge edge of liner H] are of such a radial height that the outside diameter of the discharge edge, measured over the dimples, is substantially the same or very slightly less than the inner diameter of the cylinder 4
  • the discharge end of liner I is thus supported in a longitudinally slidable manner within the ring 4i. Since the nozzle end of the liner is held rigidly by the threaded fastenings 36, the discharge edge of the liner in slides axially to a limited extent within the supporting ring 4!, as the length ,of the liner-changes as a result of differential thermal expansion between the hot liner and the comparatively cool outer casing I.
  • the annular convolution provided in the support member permits differential thermal expansion between the inner cylindrical portion 4
  • also serves to support the inner liner 42 of the discharge elbow assembly 38, in a manner which will be obvious from a consideration of Fig. 1.
  • the outer casing 43 of the elbow assembly may be provided with a flange at the downstream end secured by threaded fastenings 44 or other suitable means to a cooperating flange of a conduit 45, which may for instance be the inlet to a gas turbine.
  • the downstream end of the inner liner 42 is provided with a circumferential row of dimples 46 which may be similar in structure and are identical in function to the struck-out portions l4 provided in the discharge edge of the liner l0.
  • supports the downstream end of the inner liner 42 and the upstream end of the liner a in the manner described above in connection with the discharge end of the liner l0 and the upstream end of liner 42.
  • The. supporting ring members for the respective inner liners are provided with a circumferential row of openings 41 through which cooling air flows from the plenum chamber 24.
  • This coolant flowing through the annular space defined between the respective members 42, 43 and 45, 45a serves effectively to cool the inner liners and to reduce the transmission of heat from the inner liners to the outer casings. It will be obvious that the combustion and cooling air flowing through the plenum chamber 24 likewise reduces the transmission of heat to the outer combustor housing I.
  • the dimples l4 cooperate with the supporting ring 4
  • some of the cooling air indicated by the arrows 49 flows through the spaces defined between the dimples 46 in the manner of the arrows 50 to form a cooling and insulating film on the inner surface of liner 45a.
  • the air inlet elbow 3 may be fabricated from sheet metal or cast of aluminum, magnesium, or other suitable metal. Ordinarily it is necessary that the inner combustor liner Ill and the end dome 25 as well as the conduit liners 42 and 45a be made of expensive temperature-resistlng alloys such as various stainless steels. I have found however that the marked improvement in uniformity of temperature and stability of gas flow achieved with my invention makes it possible to fabricate the liner III of low temperature metals, such as ordinary mild steel.
  • the outer housings I, 43 and 45 may be made of ordinary inexpensive sheet metal stock.
  • the present invention has been described as an improvement in the combustor invented by Mr. A. J. Nerad, it will be obvious to those skilled in the art that it may also find application in other combustion devices operating at high pressures and with substantial pressure differentials across the liner defining the combustion space, the combustion supporting fluid being admitted with high spouting velocity through ports or nozzles the size and location of which needs be very carefully chosen in order to obtain a desired flow path within the reaction space.
  • the invention makes possible the accurate establishment of a jet having a precisely determined direction, and being insensitive to irregularities or unsyinmetricity of the flow of fluid approaching the liner.
  • a fluid fuel combustor of the type having an elongated liner of circular cross-section tapering from a minor diameter at one end to a major diameter at the other end and defining a combustion space within and having a plurality of circumferentially spaced longitudinal rows of ports for admitting combustion air in discrete jets, a cylindrical outer housing enclosing the liner and spaced therefrom to define an air supply chamber decreasing progressively in crosssectional area from the smaller end of the liner towards the larger end, an end dome member closing the smaller end of the liner and having ports for admitting cooling air to the combustion space adjacent the closed end thereof, and means for admitting fluid fuel to the combustion space, the combination of non-expanding nozzle means associated with each of the combustion air inlet ports for stabilizing the direction of jet discharge into the liner, each of said nozzles having its axis normal to the axis of the liner and projecting from the inner surface of the liner radially outward into the air supply chamber, each of the nozzles in the circum
  • a fluid fuel combustor of the type having an elongated liner of circular cross-section defining a combustion space Within and having a plurality of circumferentially spaced longitudinal rows of ports for admitting combustion air in discrete jets, an outer housing enclosing the liner and spaced therefrom to define an air supply chamber, an end dome member closing the inlet end of the liner and having ports for admitting cooling air to the combustion space, and means for supplying fuel to the combustion space adjacent the closed end thereof
  • a fluid fuel combustor of the type having a cylindrical outer housing spaced radially from an inner elongated liner of circular crosssection tapering from a minor diameter at one end to a major diameter at the other end and defining a combustion space within, said liner having a plurality of circ-umferentially spaced longitudinal rows of ports adapted to admit combustion air in discrete radial jets, and an end dome member closing the small end of the liner, the larger end being open for the dischar of hot products of combustion, the combination of non-expanding nozzle means associated with each of said ports for stabilizing the direction of jet discharge into the liner, each nozzle having a rounded contracting entrance and being arranged with its axis normal to the axis of the liner and extending radially outward from the inner surface of the liner, the nozzles adjacent the closed end of the liner having an axial length of the order of five-eighths the inner diameter of the nozzle, the remaining nozzles be- 10 ing of
  • a fluid fuel combustor of the type having an elongated liner of circular cross-section tapering from a minor diameter at one end to a major diameter at the other end and-defining a combustion space within, said liner having a plurality of circumferentially spaced longitudinal rows of ports adapted to admit combustion air in discrete radial jets, and-an end dome member closing the small endof the liner, the larger end being open for the discharge of hot products of combustion, the combination of non-expanding nozzle means associated with each of said ports for stabilizing the direction of jet discharge into the liner, each nozzle being arranged with it axis normal to the axis of the liner and extending radially outward from the inner surface of the liner, the nozzles adjacent the closed end of the liner having rounded contracting entrances and an axial length of the order of five-eighths the inner diameter of the nozzle.
  • a fluid fuel combustor of the type having a substantially cylindrical liner with a plurality of circumferentially spaced longitudinal rows of ports for admitting combustion air in radial discrete jets
  • a fluid fuel combustor of the type having a liner defining a combustion space and having a plurality of spaced openings adapted to admit combustion air in strong discrete jets the direction of which is important to establishment of a desired flow path inside the liner, the combination of separate non-expanding nozzle means associated with each opening and extending from the inner surface of the liner radially outward i'or stabilizing the direction of the jet discharged into the liner, each of the nozzles having a well rounded contracting entrance portion and a cylindrical discharge portion and an overall axial length of the order of five-eighths the inner diameter of the nozzle.
  • a fluid fuel combustor having an elongated liner of substantially circular cross-section defining a combustion space within, and an end dome member closing one end of the liner and having ports for the admission of air to the combustion space
  • the combination of walls defining an outer casing enclosing the liner and end dome and spaced from each to define therewith an air supply chamber, said walls also forming an inlet opening for receiving air under pressure
  • the combination of shroud means spaced from and enclosing said end dome and having ports adapted to supply air uniformly to the end dome ports at substantially the pressure ofthe air supplied to the inlet of said casing whereby the direction of the jets from the end dome ports is substantially unaffected by unsymmetricity or changes in the direction of approach of the air entering through said inlet opening.
  • a combustion device of the type having an elongated liner of circular cross-section tapering from a minor diameter at one end to a major diameter at the other end and defining a combustion space within, said liner having a plurality of circumferentiaily spaced longitudinal rows of ports adapted to admit combustion air in discrete radial Jets, an end dome member closing the small end of the liner and having louvers for admitting jets of air, the larger end being open for the discharge of hot products of combustion, the combination of non-expanding nozzle means associated with said ports for stabilizing the direction of jet discharge into the liner, each nozzle having a well-rounded contracting entrance and a non-expanding discharge portion and extending radially outward from the inner surface of the liner, the nozzles adjacent the closed end of the liner having an overall axial length of the order of five-eighths the inner diameter of the nozzle, the remaining nozzles being of progressively shorter axial length toward the discharge end of the liner, shroud means

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Gas Burners (AREA)
US705866A 1946-10-26 1946-10-26 Air nozzle and porting for combustion chamber liners Expired - Lifetime US2510645A (en)

Priority Applications (4)

Application Number Priority Date Filing Date Title
FR962862D FR962862A (en, 2012) 1946-10-26
US705866A US2510645A (en) 1946-10-26 1946-10-26 Air nozzle and porting for combustion chamber liners
GB28500/47A GB635946A (en) 1946-10-26 1947-10-24 Improvements in and relating to combustion chambers
CH265334D CH265334A (de) 1946-10-26 1948-02-28 Mit flüssigem Brennstoff gespiesener Brenner.

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US705866A US2510645A (en) 1946-10-26 1946-10-26 Air nozzle and porting for combustion chamber liners

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US2510645A true US2510645A (en) 1950-06-06

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CH (1) CH265334A (en, 2012)
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GB (1) GB635946A (en, 2012)

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US4704869A (en) * 1983-06-08 1987-11-10 Hitachi, Ltd. Gas turbine combustor
US4875339A (en) * 1987-11-27 1989-10-24 General Electric Company Combustion chamber liner insert
US4887432A (en) * 1988-10-07 1989-12-19 Westinghouse Electric Corp. Gas turbine combustion chamber with air scoops
US20020184889A1 (en) * 2001-06-06 2002-12-12 Snecma Moteurs Fastening a CMC combustion chamber in a turbomachine using the dilution holes
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
US20030000223A1 (en) * 2001-06-06 2003-01-02 Snecma Moteurs Mounting for a CMC combustion chamber of a turbomachine by means of flexible connecting sleeves
US6536201B2 (en) * 2000-12-11 2003-03-25 Pratt & Whitney Canada Corp. Combustor turbine successive dual cooling
US20030129555A1 (en) * 2001-12-25 2003-07-10 Yuji Mukai Burner for hydrogen generation system and hydrogen generation system having the same
US20080092547A1 (en) * 2006-09-21 2008-04-24 Lockyer John F Combustor assembly for gas turbine engine
EP2019264A1 (fr) * 2007-07-26 2009-01-28 Snecma Chambre de combustion d'une turbomachine
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FR2959795A1 (fr) * 2010-05-05 2011-11-11 Snecma Virole de chambre de combustion a contact glissant avec le carter de chambre
US20120189424A1 (en) * 2011-01-24 2012-07-26 Propheter-Hinckley Tracy A Mateface cooling feather seal assembly
US20170176006A1 (en) * 2015-12-16 2017-06-22 Rolls-Royce Deutschland Ltd & Co Kg Wall of a structural component, in particular of a gas turbine combustion chamber wall, to be cooled by means of cooling air
US20190249874A1 (en) * 2018-02-14 2019-08-15 General Electric Company Liner of a Gas Turbine Engine Combustor
US11009232B2 (en) * 2016-09-05 2021-05-18 Ansaldo Energia Switzerland AG Combustor device for a gas turbine engine and gas turbine engine incorporating said combustor device

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US2625792A (en) * 1947-09-10 1953-01-20 Rolls Royce Flame tube having telescoping walls with fluted ends to admit air
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US2575923A (en) * 1948-12-29 1951-11-20 Gen Electric Method and apparatus for pumping volatile liquids
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US2709338A (en) * 1953-01-16 1955-05-31 Rolls Royce Double-walled ducting for conveying hot gas with means to interconnect the walls
US2795108A (en) * 1953-10-07 1957-06-11 Westinghouse Electric Corp Combustion apparatus
DE1021646B (de) * 1953-12-07 1957-12-27 Gen Elek C Company Brennkammer
US2930192A (en) * 1953-12-07 1960-03-29 Gen Electric Reverse vortex combustion chamber
US2836379A (en) * 1954-05-18 1958-05-27 Gen Dyanmics Corp Ramjet wing system for jet propelled aircraft
US2884049A (en) * 1955-01-17 1959-04-28 Martin E Barzelay Spray drying apparatus
DE1043719B (de) * 1955-09-15 1958-11-13 Gen Electric Abschlusshaube fuer das Flammrohr einer Gasturbinen-Brennkammer
DE1044523B (de) * 1955-09-15 1958-11-20 Gen Electric Abschlusshaube fuer das Flammrohr einer Gasturbinen-Brennkammer
US3342403A (en) * 1964-06-22 1967-09-19 Power Jets Res & Dev Ltd Machine having a rotor supported between end-plates
US3186697A (en) * 1964-12-23 1965-06-01 Mid Continent Metal Products C Gas-fired heater
US3371482A (en) * 1965-06-14 1968-03-05 Snecma Jet propulsion casings having fuel drainage means
FR2194881A1 (en, 2012) * 1972-08-02 1974-03-01 Gen Electric
US3866417A (en) * 1973-02-09 1975-02-18 Gen Electric Gas turbine engine augmenter liner coolant flow control system
FR2333126A1 (fr) * 1975-11-29 1977-06-24 Rolls Royce Dispositif de refrigeration pour chambre de combustion de moteur a turbine a gaz
US4301657A (en) * 1978-05-04 1981-11-24 Caterpillar Tractor Co. Gas turbine combustion chamber
US4286943A (en) * 1979-08-21 1981-09-01 Joseph J. Petlak Air heater
US4704869A (en) * 1983-06-08 1987-11-10 Hitachi, Ltd. Gas turbine combustor
US4875339A (en) * 1987-11-27 1989-10-24 General Electric Company Combustion chamber liner insert
US4887432A (en) * 1988-10-07 1989-12-19 Westinghouse Electric Corp. Gas turbine combustion chamber with air scoops
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
US6536201B2 (en) * 2000-12-11 2003-03-25 Pratt & Whitney Canada Corp. Combustor turbine successive dual cooling
US20020184889A1 (en) * 2001-06-06 2002-12-12 Snecma Moteurs Fastening a CMC combustion chamber in a turbomachine using the dilution holes
US20030000223A1 (en) * 2001-06-06 2003-01-02 Snecma Moteurs Mounting for a CMC combustion chamber of a turbomachine by means of flexible connecting sleeves
US6668559B2 (en) * 2001-06-06 2003-12-30 Snecma Moteurs Fastening a CMC combustion chamber in a turbomachine using the dilution holes
US6823676B2 (en) * 2001-06-06 2004-11-30 Snecma Moteurs Mounting for a CMC combustion chamber of a turbomachine by means of flexible connecting sleeves
US20030129555A1 (en) * 2001-12-25 2003-07-10 Yuji Mukai Burner for hydrogen generation system and hydrogen generation system having the same
US20080092547A1 (en) * 2006-09-21 2008-04-24 Lockyer John F Combustor assembly for gas turbine engine
US7975487B2 (en) * 2006-09-21 2011-07-12 Solar Turbines Inc. Combustor assembly for gas turbine engine
FR2919380A1 (fr) * 2007-07-26 2009-01-30 Snecma Sa Chambre de combustion d'une turbomachine.
US20100043449A1 (en) * 2007-07-26 2010-02-25 Snecma Device for attaching a combustion chamber
US8028530B2 (en) 2007-07-26 2011-10-04 Snecma Device for attaching a combustion chamber
EP2019264A1 (fr) * 2007-07-26 2009-01-28 Snecma Chambre de combustion d'une turbomachine
US8490400B2 (en) 2008-09-15 2013-07-23 Siemens Energy, Inc. Combustor assembly comprising a combustor device, a transition duct and a flow conditioner
US20100064693A1 (en) * 2008-09-15 2010-03-18 Koenig Michael H Combustor assembly comprising a combustor device, a transition duct and a flow conditioner
FR2959795A1 (fr) * 2010-05-05 2011-11-11 Snecma Virole de chambre de combustion a contact glissant avec le carter de chambre
US20120189424A1 (en) * 2011-01-24 2012-07-26 Propheter-Hinckley Tracy A Mateface cooling feather seal assembly
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US20170176006A1 (en) * 2015-12-16 2017-06-22 Rolls-Royce Deutschland Ltd & Co Kg Wall of a structural component, in particular of a gas turbine combustion chamber wall, to be cooled by means of cooling air
US10429069B2 (en) * 2015-12-16 2019-10-01 Rolls-Royce Deutschland Ltd & Co Kg Wall of a structural component, in particular of gas turbine combustion chamber wall, to be cooled by means of cooling air
US11009232B2 (en) * 2016-09-05 2021-05-18 Ansaldo Energia Switzerland AG Combustor device for a gas turbine engine and gas turbine engine incorporating said combustor device
US20190249874A1 (en) * 2018-02-14 2019-08-15 General Electric Company Liner of a Gas Turbine Engine Combustor
US10890327B2 (en) * 2018-02-14 2021-01-12 General Electric Company Liner of a gas turbine engine combustor including dilution holes with airflow features

Also Published As

Publication number Publication date
GB635946A (en) 1950-04-19
FR962862A (en, 2012) 1950-06-22
CH265334A (de) 1949-11-30

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