US3287905A - Combustion chamber for a gas turbine - Google Patents

Combustion chamber for a gas turbine Download PDF

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Publication number
US3287905A
US3287905A US367011A US36701164A US3287905A US 3287905 A US3287905 A US 3287905A US 367011 A US367011 A US 367011A US 36701164 A US36701164 A US 36701164A US 3287905 A US3287905 A US 3287905A
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United States
Prior art keywords
chamber
combustion chamber
air
combustion
gas turbine
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Expired - Lifetime
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US367011A
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English (en)
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Bayard Gaston
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Individual
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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
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/54—Reverse-flow combustion chambers

Definitions

  • the present invention relates to a particular and new embodiment of a combustion chamber of a gas turbine.
  • the combustion chamber is the element of the gas turbine arranged between the compressor and the turbine, inside which is performed the mixture of the fuel fed by the injectors, with the air fed by the compressor and the combustion of said air and fuel mixture.
  • the combustion leads to an important increase in temperature of said air, which increases its energy, the fraction of which may be recovered through expansion is used on the one hand for driving the above-mentioned compressor and, on the other hand, for supplying propelling power in the case of a jet engine or of the mechanical power in the case of a turbine engine.
  • the walls of the combustion chamber define two inner communicating areas: the combustion area which is termed the primary chamber and the dilution area termed secondary chamber.
  • a hot gaseous core with the gases produced through combustion which is essential for the upkeep of the latter combustion.
  • Said core can exist only if there is a local slowing down and a turbulence in the mixture of air and fuel, which mixture should approxirnate stoichiometric proportions.
  • the combustion chamber which forms the object of my invention, is of the annular bent type of which the single figure of the accompanying drawings illustrates in partly sectional perspective view, an embodiment selected by way of example.
  • Said figure shows the preferred shape of the combustion chamber as well as the admission means of the fluxes of primary and secondary air, provided so as to satisfy the general conditions for a good operation as disclosed hereinabove.
  • Said figure also shows the distribution of the fuel injectors and the position of one of the sparking plugs, the other sparking plug occupying, in the example illustrated, a symmetrical position, as well as the particular organization of the ignition area.
  • the combustion chamber which forms the object of the invention is provided with an outer wall 1, approximately cylindrical and incurved in its front part, and with an inner wall 2 also approximately cylindrical and incurved in its front part, said two walls being provided furthermore each, with an arcuate notch in their medial part, said notches being respectivelyy designated by the reference numbers 3 and 4.
  • the rear ends of the walls 1 and 2 are welded over an 3,287,905 Patented Nov. 29, 1966 ICC annulus 5 in the shape of a half tore.
  • the front ends of the walls 1 and 2 open into an annular channel 6 feeding the distributor 7 for the turbine 8.
  • the primary and secondary chambers 9 and 10 are connected through the throttling means 11 formed by the notches 3 and 4.
  • the combustion chamber is arranged inside a casing constituted by the body 12, the support 13 and the ferrule 14 connected together through bolting, so as to form round the combustion chamber an annular channel 15 incurved in the shape of a U, inside which the air fed by the compressor 16 flows.
  • the combustion chamber is held in position in the casing 12-13-14 by abutment screws 17 provided with a stud and passing through the support 13 and engaging the annulus 5 and by abutment screws 18 secured to the wall 19 rigidly secured to the casing.
  • the incurved section of the Wall 1 bears on the screws 18 and furthermore, the front edges of the walls 1 and 2 are inserted in the opening 20 of the distributor 7.
  • the amount of air introduced in the primary chamber or primary flux, required for a good combustion, is adjusted through action on the cross-sectional area provided by the ports 26 and 27 and by the throttled section 11.
  • the secondary flux indicated by the arrow 1 in dotted lines passes through the primary flux of the burnt gases, indicated by the arrow F in dot-and-dash lines and cools said latter flux while it is heated itself and is reflected on the opposite incurved wall 1, so as to form permanently along said wall a layer of gas feeding the roots of the blades of the turbine, the temperature of which is lower than that prevailing in the stream of gas flowing along the wall 2 and feeding the periphery of the turbine.
  • the characteristic features of the compressor allow igniting only a few injectors, to wit: those arranged near the sparking plugs, two in the example illustrated.
  • the ignition area is protected against any objectionable lateral stream of air by partitions 29, 30 arranged longitudinally to either side of each system formed by an injector and the cooperating sparking plug or plugs. Said partitions, welded to the annulus 5, extend over a fraction of the length of the primary chamber 9.
  • the edges of the holes 26 are directed so as to slow down the speed of admission of air.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US367011A 1963-12-09 1964-05-13 Combustion chamber for a gas turbine Expired - Lifetime US3287905A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR4934A FR1388005A (fr) 1963-12-09 1963-12-09 Chambre de combustion de turbine à gaz

Publications (1)

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US3287905A true US3287905A (en) 1966-11-29

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ID=8570510

Family Applications (1)

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US367011A Expired - Lifetime US3287905A (en) 1963-12-09 1964-05-13 Combustion chamber for a gas turbine

Country Status (3)

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US (1) US3287905A (fr)
FR (1) FR1388005A (fr)
GB (1) GB1037923A (fr)

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3902314A (en) * 1973-11-29 1975-09-02 Avco Corp Gas turbine engine frame structure
US4018043A (en) * 1975-09-19 1977-04-19 Avco Corporation Gas turbine engines with toroidal combustors
US4204403A (en) * 1977-05-21 1980-05-27 Rolls-Royce Limited Combustion systems
US4845941A (en) * 1986-11-07 1989-07-11 Paul Marius A Gas turbine engine operating process
WO1990011439A1 (fr) * 1989-03-17 1990-10-04 Sundstrand Corporation Moteur a turbine a gaz compact
US5048298A (en) * 1989-07-10 1991-09-17 Sundstrand Corporation Low cost fuel system for a gas turbine engine
US5058375A (en) * 1988-12-28 1991-10-22 Sundstrand Corporation Gas turbine annular combustor with radial dilution air injection
US5109671A (en) * 1989-12-05 1992-05-05 Allied-Signal Inc. Combustion apparatus and method for a turbine engine
US5140808A (en) * 1989-03-17 1992-08-25 Sundstrand Corporation Gas turbine engine with fuel mainfold system
US5363644A (en) * 1989-12-21 1994-11-15 Sundstrand Corporation Annular combustor
US20070245710A1 (en) * 2006-04-21 2007-10-25 Honeywell International, Inc. Optimized configuration of a reverse flow combustion system for a gas turbine engine
US20120023964A1 (en) * 2010-07-27 2012-02-02 Carsten Ralf Mehring Liquid-fueled premixed reverse-flow annular combustor for a gas turbine engine
WO2024019723A1 (fr) * 2022-07-21 2024-01-25 Raytheon Technologies Corporation Récupération de l'énergie d'électrolyse

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102013222863A1 (de) 2013-11-11 2015-05-13 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenbrennkammer sowie Verfahren zu deren Herstellung

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2296023A (en) * 1941-03-03 1942-09-15 Northrop Aircraft Inc Burner
US2633705A (en) * 1946-03-30 1953-04-07 John Hawkins And Associates Re Segmentally separable gas turbine power plant
GB721899A (en) * 1952-03-06 1955-01-12 Lucas Industries Ltd Liquid fuel combustion apparatus for jet-propulsion engines or gas turbines
US2867267A (en) * 1954-02-23 1959-01-06 Gen Electric Combustion chamber
GB863011A (en) * 1957-12-20 1961-03-15 Constuction De Moteurs D Aviat Burner device for continuous-flow internal combustion engine
US3088278A (en) * 1957-05-01 1963-05-07 Avco Mfg Corp Gas turbine engine

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2296023A (en) * 1941-03-03 1942-09-15 Northrop Aircraft Inc Burner
US2633705A (en) * 1946-03-30 1953-04-07 John Hawkins And Associates Re Segmentally separable gas turbine power plant
GB721899A (en) * 1952-03-06 1955-01-12 Lucas Industries Ltd Liquid fuel combustion apparatus for jet-propulsion engines or gas turbines
US2867267A (en) * 1954-02-23 1959-01-06 Gen Electric Combustion chamber
US3088278A (en) * 1957-05-01 1963-05-07 Avco Mfg Corp Gas turbine engine
GB863011A (en) * 1957-12-20 1961-03-15 Constuction De Moteurs D Aviat Burner device for continuous-flow internal combustion engine

Cited By (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3902314A (en) * 1973-11-29 1975-09-02 Avco Corp Gas turbine engine frame structure
US4018043A (en) * 1975-09-19 1977-04-19 Avco Corporation Gas turbine engines with toroidal combustors
US4204403A (en) * 1977-05-21 1980-05-27 Rolls-Royce Limited Combustion systems
US4845941A (en) * 1986-11-07 1989-07-11 Paul Marius A Gas turbine engine operating process
US5058375A (en) * 1988-12-28 1991-10-22 Sundstrand Corporation Gas turbine annular combustor with radial dilution air injection
US5140808A (en) * 1989-03-17 1992-08-25 Sundstrand Corporation Gas turbine engine with fuel mainfold system
US5033263A (en) * 1989-03-17 1991-07-23 Sundstrand Corporation Compact gas turbine engine
WO1990011439A1 (fr) * 1989-03-17 1990-10-04 Sundstrand Corporation Moteur a turbine a gaz compact
US5048298A (en) * 1989-07-10 1991-09-17 Sundstrand Corporation Low cost fuel system for a gas turbine engine
US5109671A (en) * 1989-12-05 1992-05-05 Allied-Signal Inc. Combustion apparatus and method for a turbine engine
US5363644A (en) * 1989-12-21 1994-11-15 Sundstrand Corporation Annular combustor
US20070245710A1 (en) * 2006-04-21 2007-10-25 Honeywell International, Inc. Optimized configuration of a reverse flow combustion system for a gas turbine engine
US20120023964A1 (en) * 2010-07-27 2012-02-02 Carsten Ralf Mehring Liquid-fueled premixed reverse-flow annular combustor for a gas turbine engine
WO2024019723A1 (fr) * 2022-07-21 2024-01-25 Raytheon Technologies Corporation Récupération de l'énergie d'électrolyse

Also Published As

Publication number Publication date
FR1388005A (fr) 1965-02-05
GB1037923A (en) 1966-08-03

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