US2742762A - Combustion chamber for axial flow gas turbines - Google Patents

Combustion chamber for axial flow gas turbines Download PDF

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Publication number
US2742762A
US2742762A US229071A US22907151A US2742762A US 2742762 A US2742762 A US 2742762A US 229071 A US229071 A US 229071A US 22907151 A US22907151 A US 22907151A US 2742762 A US2742762 A US 2742762A
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US
United States
Prior art keywords
combustion chamber
carbon
air
flow
combustion
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
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US229071A
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English (en)
Inventor
Malcolm S Kuhring
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National Research Council of Canada
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National Research Council of Canada
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Filing date
Publication date
Priority to BE514534D priority Critical patent/BE514534A/xx
Priority to NL79375D priority patent/NL79375C/xx
Application filed by National Research Council of Canada filed Critical National Research Council of Canada
Priority to US229071A priority patent/US2742762A/en
Priority to GB12152/52A priority patent/GB731054A/en
Priority to DEN5542A priority patent/DE935287C/de
Priority to FR1065482D priority patent/FR1065482A/fr
Priority to CH317633D priority patent/CH317633A/fr
Application granted granted Critical
Publication of US2742762A publication Critical patent/US2742762A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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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
    • 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
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/10—Air inlet arrangements for primary air

Definitions

  • This invention relates to combustion chambers for gas turbines and the like and more particularly it relates to a means for overcoming the problems created by the formation of carbon in the primary combustion zone, which carbon formation under normal circumstances tends to block the primary air inlets, necessitating dismantling and cleaning of these ports to restore efficient operation of the combustion chamber.
  • Combustion chambers for gas turbines and the like are generally constructed from very thin sheet metal in order to overcome a situation known as thermal shock which arises due to the material being exposed to extreme ranges of temperature.
  • main uses in which gas turbines have heretofore been employed is in jet propulsion units for aircraft, it is desirable in most instances to keep the overall weight of all component parts of gas turbine combustion chambers as low as is possible, consistent with the parts having the required strength to fulfill the function that they are to perform.
  • Gas turbine combustion chambers which are in use in aircraft jet propulsion units may be divided into two general types.
  • the combustion chamber is generally cylindrical in form and has an inner wall or flame tube defining the combustion chamber proper which is closed at the forward end, except for the primary air holes which are described below, and an outer wall surrounding the inner wall through which air is supplied.
  • Air is supplied to the inside of the combustion chamber proper through holes in the inner wall which are usually divided into two groups, the first of which is a group of holes in the forward end ofthe casing through which primary air is admitted, and the second group being rather larger holes spaced around theinner wall somewhat downstream of the forward end of the combustion chamber, for the admission of secondary combustion air.
  • the fuel is normally supplied to this type. of combustion chamber by nozzle means extending into the interior of the combustion chamber proper, centrally at the forward end thereof and facing axially to the rear.
  • the second general type of combustion chamber is annular in form and differs from the first type above mentioned principally in this respect only.
  • the essential parts and their position correspond to those detailed above for the cylindrical type of combustion chamber.
  • the maximum temperature of a surface upon which carbon mayform in an oxidizing atmosphere without being immediately burnt off is from about 660 to 750 C., depending to some extent upon the nature of the surface and the form of carbon which is deposited, it will be appreciated that whereas substantially no can bon will tend to build up on the walls of the combustion chamber in the secondary combustion zone because of the high temperatures existing there, the temperature of the wall of the combustion chamber in the region of the primary air inlets is such that carbon may form and build up' thereon, and it has been found in practice that considerable build up of carbon occurs around and about the inside end of the primary air inlet ports, which build up, after a number of hours of continuous operation of the combustion chamber, seriously restricts the flow of primary air into the combustion chamber and accordingly materially reduces the efficiency of the combustion chamber as a whole.
  • my invention consists in forming the primary air inlets themselves in such a manner that the air stream flowing through them will completely fill the space provided at the extreme inner edge of the inlet port.
  • This may be done in a number of ways without materially increasing the Weight of the combustion chamber or increasing its susceptibility to thermal shock but, generally speaking, all methods of doing this involve the provision of a somewhat elongated port which is sufliciently long that the air stream flowing through it under conditions of operation will fill the passage at its discharge end and said, that the ports may not be elongated by thickening an appreciable area of the wall'because of considerations of weight and thermal shock.
  • Fig. 1 is a fragmentary cross section of a typical combustion chamber of the cylindrical type
  • Fig. 2 is an enlarged fragmentary cross sectional view of the end of the combustion chamber showing the type of holes normally employed as primary'air inlet ports;
  • FIG. 3 is a greatly enlarged fragmentary view of one of the air inlet ports shown in Figure 2, illustrating the flow of air therethrough;
  • Fig. 4 is a greatly enlarged fragmentary cross section of one of the primary inlet holes illustrating the progressive build up of carbon which occurs during operation;
  • Fig. 5A is a fragmentary cross section of a primary air inlet port according to the invention showing the manner in which the carbon builds up in stages during operation without interfering with the flow of air therethrough;
  • Fig. SB is a similar fragmentary cross section illustrating an alternative form of the invention.
  • Fig. 5C is a similar fragmentary cross section illus- (rating a still further embodiment of the invention.
  • Fig. 5D is a fragmentary cross section of a still further embodiment of the invention illustrating the desired air flow characteristics according to the invention
  • Fig. SE is a fragmentary cross section of a still further embodiment of the invention.
  • Fig. 6 is a front end elevation of an example of a primary air inlet port; according to the invention with typical dimensions;
  • Fig. 7 is a longitudinal section of the air inlet port shown in Figure 6, taken along the line 7--7.
  • a combustion chamber for internal cornbustion gas turbine engines consists of an outer shell 7, an inner shell or flame tube 6, some form of fuel supply 2 and spray nozzle 1.
  • the primary air passes through a plurality of holes 4 in the end or head of the inner shell or flame tube. These holes are intentionally small to permit the flow only of sufficient air to support combustion in this area of the flame tube. This air flow is shown by the dotted arrows and is known as primary air.
  • the flame tube 6, including the head normally operates at somewhat high temperatures.
  • the closed head or end 9 is not nearly as hot as the open end and fuel droplets or vapor impinging on the surface of the head tend to form carbon which, in a period of time, tends to close or partially close the primary air holes 4 blocking the flow of primary air. This may limit the period which may be declared as the service life of the cornbustion chambers as the balance of air flow through the various portions of the combustion chamber are vitally important.
  • the primary holes 4 are usually punched or drilled and appear as shown in Figure 2.
  • Air flow through a simple hole of this nature in a thin plate produces what is known as a vena contracta or contraction of flow.
  • Increasing the thickness of the plate and, consequently the length of the passage alleviates this condition somewhat but not entirely and in addition increases the weight and imposes greater thermal loads on the material.
  • the vena contracta induces an annular zone of turbulent back flow on the downstream side of the plate, as shown by the arrows in Figure 3.
  • This air flow causes fuel in one form or another to impinge on the surface of the plate in and around the holes and carbon is deposited progressively, as shown at a, b and c in Figure 4.
  • the air flow is completely shut off in this area and not only has the efficiency of the combustion chamber been impaired progressively, wasting fuel, rendering the engine harder to start and raising turbine temperatures, but it becomes necessary to dismantle the combustion chambers and remove the carbon.
  • the plate is effectively thickened locally, and the passage is improved aerodynamically so that while the original flow rate will be maintained, with the improved flow pattern and elimination of local eddies and fuel impingement at the edges of the primary air holes, carbon build up does not block the holes, but'takes place substantially as shown at a b and c in Figures 5A, 5B and 5C.
  • the thickening of the plate locally may be accomplished by an actual integral thickening of the plate as at 10, shown in Figure 5B, or by the fitting of a separate piece or liner IL in the form of a hollow rivet 12, preferably of thin metal, to the drilled or stamped hole, as shown in Figure 513 or by the insertion of a machined or formed hollow metal liner or passage as shown in Figure 5C.
  • the resulting passage should be longer than the original thickness of the plate in which it is fitted, and it may be of any length desired, although there is no advantage in increasing its length beyond three to four times its internal diameter.
  • the critical feature is that under conditions of operation there must be full flow at the extreme inside end of the passage, as shown at 13 in Figure 5D.
  • the insert must be fixed rigidly to the plate so that it will not become detached either by the flow of gases, from vibration, dirierential thermal expansion or other cause. For this reason, it may be necessary for the insert to project slightly downstream from the plate. This length should be kept to a minimum to assist in'preventing the insert from being heated by the direct action of the burning gases in the combustion chamber. It has been shown that if the insert does project materially inside the combustion chamber, carbon tends to build up on the outer surface of the insert and, while it does not tend to block the hole, this is not desirable. In Figures 5D and C are shown alternative methods by which the hollow rivet 12 may be attached without projecting downstream beyond the surface of the plate.
  • the upstream end of the insert should be in the form of what is known as a bellmouth. However, any similar form, such as an enlargement of the entry, countersinking, etc. will improve the air flow pattern.
  • the internal diameter may be reduced slightly if the same rate of mass air fiow is desired or the number of holes may be reduced to maintain the same rate of mass flow, due to the higher coeflicient of discharge of such an orifice compared to that of a straight punched or drilled hole.
  • the following example illustrates one form of the in vention and provides dimensional data in respect to a type of primary air inlet-port according to the invention which was used in an experimental combustion chamber.
  • EXAMPLE 1 The end of a flame tube of a combustion chamber was punched with a number of holes in the normal way, as illustrated in Figure 2. In a number of the holes thus punched were secured the inserts shown in Figures 6 and'7. These inserts were secured on the upstream face of the end of the flame tube, and the actualdimensions of the air inlet port insert were as follows:
  • my invention provides a very simple means of controlling the build up of carbon around the primary air inlet passageways of a combustion chamber in such a way that the deposition of carbon does not have an adverse effect upon the etficiency of the primary air inlet ports.
  • a combustion chamber formed from thin sheet material and having a closed inlet end and an open discharge end; fuel inlet means adjacent said closed end; and a plurality of generally tubular inserts extending through said closed end in axial alignment with said combustion chamber and surrounding said fuel inlet means, said inserts each having a length which is substantially greater than the thickness of the material from which said end is formed, not extending appreciably into said combustion chamber, and being formed to provide an inlet passage for primary air which is arranged to project a stream of air axially into said combustion chamber which is of coincident cross section with the innermost extremity of said passage whereby build-up of carbon deposit around said'innermost extremity is confined to an annular space surrounding said airstream, and growth of said deposit does not tend to restrict the flow of said airstream.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Pre-Mixing And Non-Premixing Gas Burner (AREA)
US229071A 1951-05-31 1951-05-31 Combustion chamber for axial flow gas turbines Expired - Lifetime US2742762A (en)

Priority Applications (7)

Application Number Priority Date Filing Date Title
BE514534D BE514534A (en(2012)) 1951-05-31
NL79375D NL79375C (en(2012)) 1951-05-31
US229071A US2742762A (en) 1951-05-31 1951-05-31 Combustion chamber for axial flow gas turbines
GB12152/52A GB731054A (en) 1951-05-31 1952-05-13 Gas turbine combustion chamber
DEN5542A DE935287C (de) 1951-05-31 1952-05-20 Verbrennungskammer von Gasturbinen mit primaerer und sekundaerer Verbrennungszone
FR1065482D FR1065482A (fr) 1951-05-31 1952-05-26 Chambre de combustion pour turbines à gaz et engins analogues
CH317633D CH317633A (fr) 1951-05-31 1952-05-29 Dispositif de combustion d'installation à turbine à gaz

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US229071A US2742762A (en) 1951-05-31 1951-05-31 Combustion chamber for axial flow gas turbines

Publications (1)

Publication Number Publication Date
US2742762A true US2742762A (en) 1956-04-24

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Application Number Title Priority Date Filing Date
US229071A Expired - Lifetime US2742762A (en) 1951-05-31 1951-05-31 Combustion chamber for axial flow gas turbines

Country Status (7)

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US (1) US2742762A (en(2012))
BE (1) BE514534A (en(2012))
CH (1) CH317633A (en(2012))
DE (1) DE935287C (en(2012))
FR (1) FR1065482A (en(2012))
GB (1) GB731054A (en(2012))
NL (1) NL79375C (en(2012))

Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2890569A (en) * 1954-05-03 1959-06-16 Phillips Petroleum Co Removal of carbon deposits in jet engines
US2902823A (en) * 1956-11-21 1959-09-08 Clarence E Wagner Design for a stainless steel or aluminum gas generator wall spraying system for combustion chamber
US3210935A (en) * 1959-09-23 1965-10-12 Lyman C Fisher Jetevator for missile control
US3874169A (en) * 1971-01-14 1975-04-01 Stal Laval Turbin Ab Combustion chamber for gas turbines
US3981142A (en) * 1974-04-01 1976-09-21 General Motors Corporation Ceramic combustion liner
WO1980002451A1 (en) * 1979-05-08 1980-11-13 R Babington Improvements in liquid fuel burners
US5050385A (en) * 1982-10-06 1991-09-24 Hitachi, Ltd. Inner cylinder for a gas turbine combustor reinforced by built up welding
US5209067A (en) * 1990-10-17 1993-05-11 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Gas turbine combustion chamber wall structure for minimizing cooling film disturbances
US5323605A (en) * 1990-10-01 1994-06-28 General Electric Company Double dome arched combustor
US5331815A (en) * 1992-03-23 1994-07-26 General Electric Company Impact resistant combustor
US5755093A (en) * 1995-05-01 1998-05-26 United Technologies Corporation Forced air cooled gas turbine exhaust liner
US20030182943A1 (en) * 2002-04-02 2003-10-02 Miklos Gerendas Combustion chamber of gas turbine with starter film cooling
US20050241316A1 (en) * 2004-04-28 2005-11-03 Honeywell International Inc. Uniform effusion cooling method for a can combustion chamber
JP2008169840A (ja) * 2007-01-09 2008-07-24 General Electric Co <Ge> シンブル、スリーブ並びに燃焼器アセンブリの冷却法
US20100037622A1 (en) * 2008-08-18 2010-02-18 General Electric Company Contoured Impingement Sleeve Holes
US20100122537A1 (en) * 2008-11-20 2010-05-20 Honeywell International Inc. Combustors with inserts between dual wall liners
US20130078582A1 (en) * 2011-09-27 2013-03-28 Rolls-Royce Plc Method of operating a combustion chamber
JP2014077630A (ja) * 2012-10-10 2014-05-01 General Electric Co <Ge> 流体を分離するためのシステム及び方法
US20150285498A1 (en) * 2014-04-02 2015-10-08 United Technologies Corporation Grommet assembly and method of design
US20160047549A1 (en) * 2014-08-15 2016-02-18 Rolls-Royce Corporation Ceramic matrix composite components with inserts
US20160201908A1 (en) * 2013-08-30 2016-07-14 United Technologies Corporation Vena contracta swirling dilution passages for gas turbine engine combustor

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR1421372A (fr) * 1964-09-16 1965-12-17 Perfectionnement apporté aux chambres de combustion

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2227666A (en) * 1936-12-10 1941-01-07 Bbc Brown Boveri & Cie Starting up system for heat producing and consuming plants
US2471101A (en) * 1945-03-31 1949-05-24 Charles E Feinberg Secondary combustion air distribution control for bridge wall furnaces
US2510645A (en) * 1946-10-26 1950-06-06 Gen Electric Air nozzle and porting for combustion chamber liners
GB650462A (en) * 1947-11-03 1951-02-28 Power Jets Res & Dev Ltd Improvements in or relating to combustion apparatus
GB650528A (en) * 1947-11-03 1951-02-28 Power Jets Res & Dev Ltd Improvements in or relating to combustion apparatus
US2547619A (en) * 1948-11-27 1951-04-03 Gen Electric Combustor with sectional housing and liner
US2601390A (en) * 1946-11-07 1952-06-24 Westinghouse Electric Corp Combustion chamber for gas turbines with circumferentially arranged pulverized solidfuel and air nozzles
US2631429A (en) * 1948-06-08 1953-03-17 Jr Harold M Jacklin Cooling arrangement for radial flow gas turbines having coaxial combustors

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE709065C (de) * 1935-07-17 1941-08-07 Rene Alexandre Arthur Couzinet Speiseeinrichtung fuer Gasturbinen
GB603485A (en) * 1946-02-23 1948-06-16 Armstrong Siddeley Motors Ltd Liquid fuel combustion chamber
DE804982C (de) * 1947-06-23 1951-05-04 Armstrong Siddeley Motors Ltd Brennkammer fuer Brennkraftturbeinen

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2227666A (en) * 1936-12-10 1941-01-07 Bbc Brown Boveri & Cie Starting up system for heat producing and consuming plants
US2471101A (en) * 1945-03-31 1949-05-24 Charles E Feinberg Secondary combustion air distribution control for bridge wall furnaces
US2510645A (en) * 1946-10-26 1950-06-06 Gen Electric Air nozzle and porting for combustion chamber liners
US2601390A (en) * 1946-11-07 1952-06-24 Westinghouse Electric Corp Combustion chamber for gas turbines with circumferentially arranged pulverized solidfuel and air nozzles
GB650462A (en) * 1947-11-03 1951-02-28 Power Jets Res & Dev Ltd Improvements in or relating to combustion apparatus
GB650528A (en) * 1947-11-03 1951-02-28 Power Jets Res & Dev Ltd Improvements in or relating to combustion apparatus
US2631429A (en) * 1948-06-08 1953-03-17 Jr Harold M Jacklin Cooling arrangement for radial flow gas turbines having coaxial combustors
US2547619A (en) * 1948-11-27 1951-04-03 Gen Electric Combustor with sectional housing and liner

Cited By (27)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2890569A (en) * 1954-05-03 1959-06-16 Phillips Petroleum Co Removal of carbon deposits in jet engines
US2902823A (en) * 1956-11-21 1959-09-08 Clarence E Wagner Design for a stainless steel or aluminum gas generator wall spraying system for combustion chamber
US3210935A (en) * 1959-09-23 1965-10-12 Lyman C Fisher Jetevator for missile control
US3874169A (en) * 1971-01-14 1975-04-01 Stal Laval Turbin Ab Combustion chamber for gas turbines
US3981142A (en) * 1974-04-01 1976-09-21 General Motors Corporation Ceramic combustion liner
US4298338A (en) * 1976-12-30 1981-11-03 Owens-Illinois, Inc. Liquid fuel burners
WO1980002451A1 (en) * 1979-05-08 1980-11-13 R Babington Improvements in liquid fuel burners
US5050385A (en) * 1982-10-06 1991-09-24 Hitachi, Ltd. Inner cylinder for a gas turbine combustor reinforced by built up welding
US5323605A (en) * 1990-10-01 1994-06-28 General Electric Company Double dome arched combustor
US5209067A (en) * 1990-10-17 1993-05-11 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Gas turbine combustion chamber wall structure for minimizing cooling film disturbances
US5331815A (en) * 1992-03-23 1994-07-26 General Electric Company Impact resistant combustor
US5755093A (en) * 1995-05-01 1998-05-26 United Technologies Corporation Forced air cooled gas turbine exhaust liner
US20030182943A1 (en) * 2002-04-02 2003-10-02 Miklos Gerendas Combustion chamber of gas turbine with starter film cooling
US7124588B2 (en) * 2002-04-02 2006-10-24 Rolls-Royce Deutschland Ltd & Co Kg Combustion chamber of gas turbine with starter film cooling
US20050241316A1 (en) * 2004-04-28 2005-11-03 Honeywell International Inc. Uniform effusion cooling method for a can combustion chamber
JP2008169840A (ja) * 2007-01-09 2008-07-24 General Electric Co <Ge> シンブル、スリーブ並びに燃焼器アセンブリの冷却法
US8281600B2 (en) * 2007-01-09 2012-10-09 General Electric Company Thimble, sleeve, and method for cooling a combustor assembly
US20100251723A1 (en) * 2007-01-09 2010-10-07 Wei Chen Thimble, sleeve, and method for cooling a combustor assembly
US20100037622A1 (en) * 2008-08-18 2010-02-18 General Electric Company Contoured Impingement Sleeve Holes
US8161752B2 (en) * 2008-11-20 2012-04-24 Honeywell International Inc. Combustors with inserts between dual wall liners
US20100122537A1 (en) * 2008-11-20 2010-05-20 Honeywell International Inc. Combustors with inserts between dual wall liners
US20130078582A1 (en) * 2011-09-27 2013-03-28 Rolls-Royce Plc Method of operating a combustion chamber
JP2014077630A (ja) * 2012-10-10 2014-05-01 General Electric Co <Ge> 流体を分離するためのシステム及び方法
US20160201908A1 (en) * 2013-08-30 2016-07-14 United Technologies Corporation Vena contracta swirling dilution passages for gas turbine engine combustor
US20150285498A1 (en) * 2014-04-02 2015-10-08 United Technologies Corporation Grommet assembly and method of design
US10443848B2 (en) * 2014-04-02 2019-10-15 United Technologies Corporation Grommet assembly and method of design
US20160047549A1 (en) * 2014-08-15 2016-02-18 Rolls-Royce Corporation Ceramic matrix composite components with inserts

Also Published As

Publication number Publication date
NL79375C (en(2012))
GB731054A (en) 1955-06-01
DE935287C (de) 1955-11-17
CH317633A (fr) 1956-11-30
BE514534A (en(2012))
FR1065482A (fr) 1954-05-26

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