US3735589A - Walls of combustion chambers - Google Patents

Walls of combustion chambers Download PDF

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Publication number
US3735589A
US3735589A US00148715A US3735589DA US3735589A US 3735589 A US3735589 A US 3735589A US 00148715 A US00148715 A US 00148715A US 3735589D A US3735589D A US 3735589DA US 3735589 A US3735589 A US 3735589A
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United States
Prior art keywords
upstream
combustion chamber
holes
downstream
wall
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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US00148715A
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English (en)
Inventor
J Caruel
J Castellant
S Coutor
A Lacroix
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Safran Aircraft Engines SAS
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SNECMA SAS
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Publication date
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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
    • F23R3/08—Arrangement of apertures along the flame tube between annular flame tube sections, e.g. flame tubes with telescopic sections
    • 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
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • ABSTRACT A combustion chamber mounted in a flow of air at high pressure and the wall of which is provided with dilution holes located in the path of a cooling air film derived from the said flow through holes provided at a substantially truncated junction connecting an upstream wall element with a downstream wall element offset relative to the upstream element towards theoutside of the chamber, wherein the upstream wall element is extended rearwardly of this junction by means of a tongue substantially parallel to the rear element, and wherein this junction has an inclination between 15 and 25, preferably about 20, the holes are arranged in zig-zag very near the upstream element so as to provide in the downstream direction a divergent portion of unperforated wall of substantial length, and the dilution holes are surrounded by flanges projecting into the combustion chamber.
  • the invention relates to dilution type combustion chambers having walls cooled by air films, more especially of the type used in jet engines for aircraft.
  • a combustion chamber is constructed to ensure combustion of fuel in a flow. of air under high pressure. It is mounted within a wall called a flame tube arranged longitudinally in the flow of air.
  • This wall is provided with lateral openings through which dilution air, derived from the high pressure flow, enters transversely into the chamber, and with and arrangement for cooling comprising means for forming an air film on the inner surface of the walls in order to protect them from the direct action of the flame.
  • This method of cooling is known in the art under the name of film cooling.
  • an efficient cooling may be achieved by providing the truncated joint with an angle of inclination between about 15 and 25, and preferably 20, by arranging the holes in staggered relation adjacent the upstream wall element so as to leave downstream a divergent nonperforated wall portion of substantial length, and by surrounding the dilution holes with flanges projecting into the combustion chamber.
  • the flanges of the dilution holes preferably extend into the combustion chamber substantially to the level of the said tongue i.e., the distance of the ofiset of the downstream wall section from the upstream wall section.
  • the dilution holes are not circular but oblong, with their long axes in the longitudinal direction of the combustion chamber, that is to say in the direction of the flow of the air film.
  • FIG. 1 shows a combustion chamber in longitudinal cross-section
  • FIG. 2 shows on a larger scale the part of FIG. 1 within the dot-dash lines
  • FIG. 3 is a development along the arrow III in'FIG.
  • FIG. 4 is a partial cross-section on an enlarged scale along the line IV-IV in FIG. 1 and FIG. 5 is a view in the direction of the arrow V in FIG. 4.
  • the combustion chamber shown is annular; its wall comprises, as known in the art, an annular base 1, an annular outer part 2, and an annular inner part 3.
  • the base 1 carries a plurality of burners d in crown arrangement and the rear edges of the outer and irmer parts 2 and 3 form an annular outlet orifice 5.
  • the outer part 2 is provided with two rows of openings 7, h.
  • the inner part 3 has also two rows of openings 9, 10.
  • Illa and llllb are provided for mounting the combustion chamber in a jet engine, (not shown), between the compressor and the turbine (not shown) of this engine.
  • Air at high pressure compressed by the compressor flows in a flow F at the outside of the ring formed by the combustion chamber and in a flow F inside this ring. A part of this air also enters the combustion chamber through holes such as 7, 8, 9, m and 12, surrounding the burnem 4 and between them.
  • Fuel is injected into the burners 4 and lit, by means not shown, so that it burns in the primary air entering through holes such as 7 and 9 and its combustion is furthered and continues in the secondary and dilution air which enters through the openings and 110. Cooling air enters at 12.
  • the hot gases escape at high speed through the opening 5 for driving the turbine, and the combustion chamber is at low pressure relative to the external alr flows F, F
  • the outer part 2 of the wall of the combustion chamber is made of three metal sleeves, the diameters of which increase from upstream to downstream, and the inner part 3 is made of three sleeves with stepped diameters decreasing from upstream to downstream.
  • Each sleeve has at the beginning a connecting part which is welded to The wall element located immediately upstream slightly in front of the rear edge thereof.
  • FIG. 2 shows on a larger scale the connection of the center sleeve 13 of the outer part 2 with the sleeve 14 located upsneam therefrom.
  • the sleeve 13 of generally cylindrical shape is connected upstream by a truncated section 13a, with an apex angle of about 20, to cylindrical portion 13b of smaller diameter which is located on the sleeve 14 and fixed thereto by welded joints (not shown).
  • the truncated portion 13a is perforated by a crown of holes 15 which, as shown in FIG. 3, are arranged in zig-zag or staggered relation in two rows quite near the cylindrical portion 13b, so that the truncated part 13a is provided with perforations over the major part of its length.
  • the sleeve M extends towards the rear forming a cylindrical tongue or skirt 140, the rear edge 14b of which is located slightly beyond the truncated portion 13a.
  • the dilution holes 8 provided in the sleeve 13 are arranged in a ring at a certain distance downstream of the rear edge Mb of the tongue 14a.
  • the ring of apertures 8 comprises a plurality of large apertures 16, alternating with smaller apertures 17, wherein the apertures are regularly spaced apart. They are not circular but oblong, with their large dimension pointing in the longitudinal direction of the combustion chamber.
  • Each opening 16 or 17 has a flange 16a or 17a, bent back towards the interior of the combustion chamber. These flanges 17a extend substantially to the level of the tongue 14a and the flanges 16a extend even further into the chamber.
  • the upper sleeve 14 and the lower sleeve 18 are connected respectively to the base 1 and to the center sleeve 13, as already described with reference to FIG. 2. It may be seen from FIG. 1 that the truncated parts 14b and 1814 are used for these connections, each perforated by a ring of holes in staggered arrangement 19 and 20 respectively, and also showing the tongues 14a and 130.
  • the inner part 3 of the wall is formed by three sleeves 13', 14, 18' connected to each other and to the base 1 in a similar manner, wherein the elements playing the same role as in the outer part 1 are marked with the same reference numerals, with the addition of an apostrophe.
  • the ring of dilution holes provided in the sleeve 13' has holes corresponding to the holes 16 and 17 in FIGS. 4 and 5, equipped with flanges, such as 16a, projecting into the combustion chamber at least to the level of the tongue l4'a.
  • the openings 7 and 9, provided respectively in the sleeves 14 and 14, are circular and equipped with flanges 7a and 9a projecting into the interior of the combustion chamber.
  • the lower sleeves l8 and 18' are not provided with dilution holes and are respectively welded to annular elements 21 and 21' defining the outlet orifice 5.
  • the part of the combustion chamber of which the construction is most critical is clearly that comprising the sleeves l3 and 13'. It is this part which is exposed to the most intensive thermal radiation and the dilution air must be introduced into this part in large amounts, transversely of the longitudinal direction of the flame, in order to generate the necessary turbulence, but without inhibiting the flow of hot gases towards the outlet orifice 5 and without impairing the cooling of the walls.
  • the air enters the holes 15 at a lesser speed than if the truncated part 13a were more inclined.
  • the air leaving the divergent channel 22 flows through the annular channel 23 with parallel walls, protected against the turbulence occurring in the combustion chamber, and leaves this channel 23 along the arrow F forming a homogenous film without turbulence flowing at low speed over the inner surface of the sleeve 13.
  • the air leaving the holes 15 has impinged on the tongue 14a, which contributes to its cooling.
  • the good cooling of the part of the sleeve 13 surrounding the holes 16, 17 and downstream thereof may be explained by the co-operation of two elements, namely the flanges 16a, 17a, the height of which prevents the air film from being swept along by the jets of dilution air entering through the holes, and the homogenous and comparatively slow flow of this air film which makes it possible to pass around the flanges 16a, 17a and to reform downstream thereof.
  • the oblong shape of the dilution holes makes it possible to provide the air with the required flow crosssection by reducing the transverse dimensions of the holes. In this manner, the distance between the parts of the cooling air film flowing on either side of the holes is reduced, which facilitates the restoration of the film downstream of the holes. Moreover, the jets of dilution air which enter the combustion chamber present less of a risk to form a barrier in front of the outlet 5.
  • the alternation of openings such as 16 with smaller openings such as 17 favors the dilution and the restoration of the cooling air film downstream of the openings.
  • the combustion chamber according to the invention may be a flame tube in the form of a bag open at its downstream end and provided with a burner forming part of an assembly of flame tubes arranged around the axis of a jet engine or gas turbine.
  • the dilution and film cooling device according to the invention may be used only in the part of the combustion chamber exposed to the worse heat effects, whilst the other parts can be cooled by other means.
  • a combustion chamber operable in a flow of high pressure air having an outer wall extending generally longitudinally in the flow of air and comprising an upstream wall section, a downstream wall section offset radially outward of the chamber relative to said upstream section, and an outwardly flaring wall section connecting said upstream and downstream sections and having an inclination angle in the range of about 15 to 25 relative to said upstream section, said outwardly flaring section comprising an upstream portion penetrated by a multiplicity of apertures arranged in staggered relation and an imperforate downstream portion of substantial length; said upstream wall section including an extension projecting substantially parallel to the downstream wall section and extending longitudinally 6 wherein the length of said flanges at least substantially equals the extent of the radial offset of said downstream wall section.

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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)
  • Measuring Oxygen Concentration In Cells (AREA)
US00148715A 1970-06-02 1971-06-01 Walls of combustion chambers Expired - Lifetime US3735589A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR7020243A FR2093115A5 (pt) 1970-06-02 1970-06-02

Publications (1)

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US3735589A true US3735589A (en) 1973-05-29

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US00148715A Expired - Lifetime US3735589A (en) 1970-06-02 1971-06-01 Walls of combustion chambers

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US (1) US3735589A (pt)
DE (1) DE2126648C3 (pt)
FR (1) FR2093115A5 (pt)
GB (1) GB1323007A (pt)

Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3899881A (en) * 1974-02-04 1975-08-19 Gen Motors Corp Combustion apparatus with secondary air to vaporization chamber and concurrent variance of secondary air and dilution air in a reverse sense
DE3007209A1 (de) * 1979-03-01 1980-09-11 Snecma Vorrichtung zum kuehlen der wand einer brennkammer
US4233123A (en) * 1978-12-18 1980-11-11 General Motors Corporation Method for making an air cooled combustor
US6101814A (en) * 1999-04-15 2000-08-15 United Technologies Corporation Low emissions can combustor with dilution hole arrangement for a turbine engine
JP2001147017A (ja) * 1999-11-01 2001-05-29 General Electric Co <Ge> オフセット希釈燃焼器ライナ
US6606861B2 (en) 2001-02-26 2003-08-19 United Technologies Corporation Low emissions combustor for a gas turbine engine
US20090293490A1 (en) * 2008-05-28 2009-12-03 Rolls-Royce Plc Combustor wall with improved cooling
US20100218503A1 (en) * 2009-02-27 2010-09-02 Honeywell International Inc. Plunged hole arrangement for annular rich-quench-lean gas turbine combustors
US20100218504A1 (en) * 2009-02-27 2010-09-02 Honeywell International Inc. Annular rich-quench-lean gas turbine combustors with plunged holes
US20170059159A1 (en) * 2015-08-25 2017-03-02 Rolls-Royce Corporation Cmc combustor shell with integral chutes
US10670270B2 (en) 2016-02-01 2020-06-02 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber with wall contouring
WO2021118567A1 (en) * 2019-12-12 2021-06-17 Siemens Energy Global GmbH & Co. KG Combustor liner in gas turbine engine

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CA980584A (en) * 1972-11-10 1975-12-30 Edward E. Ekstedt Double walled impingement cooled combustor
FR2312654A1 (fr) * 1975-05-28 1976-12-24 Snecma Perfectionnements aux chambres de combustion pour moteurs a turbine a gaz
US5481867A (en) * 1988-05-31 1996-01-09 United Technologies Corporation Combustor
GB2441342B (en) * 2006-09-01 2009-03-18 Rolls Royce Plc Wall elements with apertures for gas turbine engine components

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2447482A (en) * 1945-04-25 1948-08-24 Westinghouse Electric Corp Turbine apparatus
US2657531A (en) * 1948-01-22 1953-11-03 Gen Electric Wall cooling arrangement for combustion devices
GB700017A (en) * 1951-03-05 1953-11-25 Lucas Industries Ltd Combustion chambers for prime movers
US2867267A (en) * 1954-02-23 1959-01-06 Gen Electric Combustion chamber
US3046742A (en) * 1959-01-05 1962-07-31 Gen Motors Corp Combustion apparatus
US3447318A (en) * 1966-12-08 1969-06-03 Snecma Wall element for combustion chambers
US3490230A (en) * 1968-03-22 1970-01-20 Us Navy Combustion air control shutter

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2447482A (en) * 1945-04-25 1948-08-24 Westinghouse Electric Corp Turbine apparatus
US2657531A (en) * 1948-01-22 1953-11-03 Gen Electric Wall cooling arrangement for combustion devices
GB700017A (en) * 1951-03-05 1953-11-25 Lucas Industries Ltd Combustion chambers for prime movers
US2867267A (en) * 1954-02-23 1959-01-06 Gen Electric Combustion chamber
US3046742A (en) * 1959-01-05 1962-07-31 Gen Motors Corp Combustion apparatus
US3447318A (en) * 1966-12-08 1969-06-03 Snecma Wall element for combustion chambers
US3490230A (en) * 1968-03-22 1970-01-20 Us Navy Combustion air control shutter

Cited By (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3899881A (en) * 1974-02-04 1975-08-19 Gen Motors Corp Combustion apparatus with secondary air to vaporization chamber and concurrent variance of secondary air and dilution air in a reverse sense
US4233123A (en) * 1978-12-18 1980-11-11 General Motors Corporation Method for making an air cooled combustor
DE3007209A1 (de) * 1979-03-01 1980-09-11 Snecma Vorrichtung zum kuehlen der wand einer brennkammer
US6101814A (en) * 1999-04-15 2000-08-15 United Technologies Corporation Low emissions can combustor with dilution hole arrangement for a turbine engine
JP2001147017A (ja) * 1999-11-01 2001-05-29 General Electric Co <Ge> オフセット希釈燃焼器ライナ
US6606861B2 (en) 2001-02-26 2003-08-19 United Technologies Corporation Low emissions combustor for a gas turbine engine
US20090293490A1 (en) * 2008-05-28 2009-12-03 Rolls-Royce Plc Combustor wall with improved cooling
US20100218503A1 (en) * 2009-02-27 2010-09-02 Honeywell International Inc. Plunged hole arrangement for annular rich-quench-lean gas turbine combustors
US20100218504A1 (en) * 2009-02-27 2010-09-02 Honeywell International Inc. Annular rich-quench-lean gas turbine combustors with plunged holes
US8141365B2 (en) 2009-02-27 2012-03-27 Honeywell International Inc. Plunged hole arrangement for annular rich-quench-lean gas turbine combustors
US8171740B2 (en) * 2009-02-27 2012-05-08 Honeywell International Inc. Annular rich-quench-lean gas turbine combustors with plunged holes
US20170059159A1 (en) * 2015-08-25 2017-03-02 Rolls-Royce Corporation Cmc combustor shell with integral chutes
US11796174B2 (en) 2015-08-25 2023-10-24 Rolls-Royce Corporation CMC combustor shell with integral chutes
US10670270B2 (en) 2016-02-01 2020-06-02 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber with wall contouring
WO2021118567A1 (en) * 2019-12-12 2021-06-17 Siemens Energy Global GmbH & Co. KG Combustor liner in gas turbine engine

Also Published As

Publication number Publication date
FR2093115A5 (pt) 1972-01-28
GB1323007A (en) 1973-07-11
DE2126648A1 (de) 1971-12-09
DE2126648B2 (de) 1978-05-18
DE2126648C3 (de) 1978-12-21

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