US3132484A - Combustion products generator with diverse combustion and diluent air paths - Google Patents

Combustion products generator with diverse combustion and diluent air paths Download PDF

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Publication number
US3132484A
US3132484A US110133A US11013361A US3132484A US 3132484 A US3132484 A US 3132484A US 110133 A US110133 A US 110133A US 11013361 A US11013361 A US 11013361A US 3132484 A US3132484 A US 3132484A
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US
United States
Prior art keywords
combustion
flame tube
annular
air duct
wall
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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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US110133A
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English (en)
Inventor
Lefebvre Arthur Henry
Smith Herbert Frank
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Rolls Royce PLC
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Rolls Royce PLC
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Publication date
Application filed by Rolls Royce PLC filed Critical Rolls Royce PLC
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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B63—SHIPS OR OTHER WATERBORNE VESSELS; RELATED EQUIPMENT
    • B63B—SHIPS OR OTHER WATERBORNE VESSELS; EQUIPMENT FOR SHIPPING 
    • B63B35/00—Vessels or similar floating structures specially adapted for specific purposes and not otherwise provided for
    • B63B35/14—Fishing vessels
    • B63B35/22—Whale catchers; Whale factory vessels
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F02C7/16—Cooling of plants characterised by cooling medium
    • F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
    • 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

Definitions

  • This invention concerns a combustion chamber, e.g. for use in a gas turbine engine.
  • High pressure gas turbine engine combustion chambers have previously been proposed whose flame tubes have been supplied with primary air and with dilution air, such combustion chambers normally also being supplied with secondary air.
  • the primary air is introduced into the flame tubes upstream of the combustion zones therein and provides at least the greater part of the air required to support combustion.
  • the secondary air when used, is introduced into the flame tubes downstream of the combustion zones to assist the combustion gases to undergo reversals of direction and to complete the combustion.
  • the dilution air is introduced into the flame tubes downstream of the secondary air so as to mix with the combustion gases and reduce their temperature to a value acceptable to the turbine of the engine.
  • the dilution air has been supplied to each flame tube from a dilution air duct arranged about the flame tube.
  • the pressure in the dilution air duct has, however, necessarily been higher than that in the flame tube so as to ensure that the dilution air enters the latter. Accordingly there has been a very high compressive buckling load across the outer wall of the flame tube and for this reason high pressure gas turbine engine combustion chambers have normally been made in tube-annular form. That is to say, such combustion chambers have comprised an annular casing containing a number of tubular flame tubes, each said flame tube having a small diameter so as to give the required silliness to withstand the high buckling loads.
  • an annular combustion chamber comprising an annular flame tube, a primary air duct arranged to supply primary air to a combustion zone within the flame tube, fuel supply means for supplying fuel to the said combustion zone, and an annular dilution air duct arranged concentrically within and immediately adjacent to said annular flame tube, said dilution air duct having means for supplying dilution air to the flame tube downstream of the said combustion zone therein.
  • the combustion chamber has an'annular casing within which the annular flame tube is concentrically mounted, a cooling air duct being provided between the said casing and the flame tube, the arrangement being such that no secondary or dilution air is supplied from the cooling air duct to the flame tube.
  • Means are preferably provided for admitting cooling air from said cooling air duct and onto at least part of the internal surface of the outer wall of the annular flame tube.
  • the outer wall of the annular flame tube may comprise a plurality of axially consecutive sections, the downstream portion of each of which is mounted within and spaced from the upstream portion of the adjacent section.
  • the fuel supply means are arranged to cause no obstruction of the dilution air duct.
  • the fuel 3,132,484 Patented May 12, 1964 supply means may comprise fuel injection nozzles and fuel supply piping for supplying fuel to said nozzles, said piping and/or said nozzles being arranged to cross the cooling air duct.
  • the combustion chamber is also preferably provided with fuel ignition means which cross the cooling air duct and which therefore do not obstruct the dilution air duct.
  • the fuel supply means and the primary air duct may be so arranged that the whole mass of the combustion gases undergoes a single reversal of direction.
  • the fuel supply means may comprise fuel injection means arranged centrally of said primary air duct so that primary air is supplied to opposite sides of the fuel injection means, whereby the combustion gases are divided into two portions each of which undergoes single reversal of direction.
  • the invention also comprises a gas turbine engine provided with an annular combustion chamber as set forth above, the dilution air duct preferably being arranged to effect cooling of the root portions of the turbine blades of the engine.
  • FIGURE 1 is a section through part of an annular combustion chamber according to the present invention
  • FIGURE 1A is a partial rear view of the combustion stabilizing member forming part of the chamber of FIG- URE 1;
  • FIGURE 2 is a section through part of a modified embodiment of an annular combustion chamber according to the present invention.
  • a gas turbine engine comprises in flow series a compressor 10, an annular combustion chamber 11, and a turbine 12, the turbine 12 driving the compressor 10 through a shaft 13.
  • the combustion chamber 11 is provided with an outer annular casing 14 within which is concentrically mounted an annular flame tube 15 having outer and inner walls 16, 17 respectively.
  • the outer wall 16 is spaced from the casing 14 to provide an annular cooling air duct 18 whose upstream end receives a supply of cooling air from the compressor 10.
  • the supply of cooling air passing through the cooling air duct 18 prevents overheating of the casing 14, the downstream end of the cooling air duct 18 being arranged to direct the air which has passed therethrough onto the tips of the turbine blades 20 so as to cool the latter.
  • the upstream end of the flame tube 15 communicates, by Way of an annular partition 21 having apertures 22 therein, with an annular primary air duct 23 which is supplied with primary air from the compressor 10.
  • the annular partition is in fact a frusto-conical member converging in the downstream direction of the combustion chamber from the outer wall to the inner wall and acts as a combustion stabilizing means.
  • the unapertured parts of the partition 21 provide sheltered areas within the flame tube 15, these sheltered areas forming part of a combustion zone into which fuel is in jected through fuel injection nozzles 24.
  • the nozzles 24 extend across the cooling air duct 18 and are supplied with fuel through pipes 25 commrmicating with a fuel manifold 26.
  • the primary air is supplied to one side only of the nozzles 24, whereby, as indicated by the arrow 27 the Whole mass of combustion gases undergoes a single reversal of direction.
  • An ignitor plug 28 is provided to effect ignition of the fuel, the ignitor plug being mounted in the casing 14 and crossing the cooling air duct 18 so as to extend into the flame tube 15.
  • the combustion chamber 11 has an inner annular wall 30 which is spaced from and concentrically mounted inwardly of the inner wall 17 of the flame tube 15.
  • walls 17, 30 define between themselves a dilution air duct 31 whose upstream end is provided with dilution air from the compressor 10.
  • the wall 17 of the flame tube has apertures 32 therein through which dilution air from the duct 31 may enter the flame tube downstream of the combustion zone therein.
  • the combustion gases which leave the downstream end of the flame tube 15 and which are directed onto the middle portions of the turbine blades have therefore had their temperature reduced to a value acceptable to the blades 20.
  • the downstream end of the dilution air duct 31 is arranged to direct the air flowing therethrough onto the root portions of the blades 20 to cool the latter.
  • FIGURE 1 the combustion chamber shown therein employs no secondary air.
  • FIGURE 2 The combustion chamber shown in FIGURE 2 is generally similar to that of FIGURE 1, corresponding parts thereof being given the same reference numerals.
  • the primary air duct is bifurcated by a channel section annular wall into branches 36, 37, the fuel injection nozzles 24 being mounted within the channel section wall 35. Accordingly the primary air is supplied to opposite sides of the nozzles 24, whereby, as indicated by the arrows 38, the combustion gases are divided into two portions each of which undergoes a single reversal of direction.
  • the annular flame tube 15 is made up of axially consecutive annular sections 40, 41, 42, the downstream portions of the sections 4-0, 41 being mounted within and spaced from the upstream portions of the sections 41, 42 respectively. Cooling air from the duct 18 may therefore flow onto the internal surface of the outer walls of the sections 41, 42 as indicated by the arrows 43.
  • the air passing through the cooling air duct /18 is employed for cooling purposes only and since it is not employed as secondary or dilution air it does not require to be at a pressure such that it will penetrate into the combustion gas stream in the flame tube 15. Accordingly, the arrangement may be such that there is only a small difference in the pressures prevailing in the cooling air duct 18 and the flame tube 15, whereby the wall 16 of the flame tube 15 is subjected to very little compressive buckling load.
  • Internal wall cooling may be provided by forming the flame tube 15 of axially consecutive sections as shown in FIGURE 2.
  • the temperature distribution along the turbine blades of a gas turbine engine is important, since it affects the stressing of the blade.
  • the root portion of the blade should be at a fairly low temperature whilst the hottest part of the blade should be about a third of the way along the blade from its root. It will be appreciated that this temperature distribution is achieved by the construction shown in the drawings, the position of the dilution air duct 31 being such as to enable part of the air flowing therethrough to be used to cool the root portions of the turbine blades.
  • the shaft 13 has a reduced diameter portion 44 adjacent the combustion chamber 11 so as to make Way for the latter.
  • this reduction in diameter is not at all substantial. This is an advantageous arrangement since the ability to keep the shaft 13 at substantially its maximum diameter substantially throughout its length permits the number of bearings for the shaft 13 to be kept to the minimum.
  • the present invention in contrast to providing the annular dilution air duct externally of the flame tube, provides a flame tube having a smaller radial width for a given flow area and therefore a smaller length if the combustion chamber is designed for a constant length/width ratio.
  • combustion chamber shown in the drawings extends radially outwardly of the compressor 10 of the engine is unimportant in those cases in which the engine is provided with a turbine or with a reheat pipe of large diameter. It is also unimportant in supersonic engines where the air intake may have a larger diameter than that of the compressor.
  • An annular combustion chamber comprising an annular casing, an annular flame tube having concentric radially outer and inner walls, the outer wall being mounted concentrically radially inwardly of and spaced from the annular casing to form a cooling air duct therewith, said flame tube having a combustion zone therein, an annular duct Wall mounted concentrically radially inwardly of and spaced from said inner wall to form a dilution air duct therewith immediately adjacent to said annular flame tube, passages leading from said dilution air duct to the interior of the flame tube downstream only of said combustion zone, an air supply duct supplying air to each of said cooling air duct, flame tube and dilution air duct, fuel supply means for supplying fuel to said combustion zone, a combustion stabilizing member extending across the flame tube upstream of said combustion zone, said member being frusto-conical and converging in the downstream direction of the combustion chamber from said outer wall to said inner wall, and said member defining circumferentially spaced passages adjacent said inner wall

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Life Sciences & Earth Sciences (AREA)
  • Marine Sciences & Fisheries (AREA)
  • Ocean & Marine Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US110133A 1960-05-18 1961-05-15 Combustion products generator with diverse combustion and diluent air paths Expired - Lifetime US3132484A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB17590/60A GB903834A (en) 1960-05-18 1960-05-18 Combustion chamber

Publications (1)

Publication Number Publication Date
US3132484A true US3132484A (en) 1964-05-12

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US110133A Expired - Lifetime US3132484A (en) 1960-05-18 1961-05-15 Combustion products generator with diverse combustion and diluent air paths

Country Status (5)

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US (1) US3132484A (enrdf_load_stackoverflow)
CH (1) CH390622A (enrdf_load_stackoverflow)
DE (2) DE1286332B (enrdf_load_stackoverflow)
GB (1) GB903834A (enrdf_load_stackoverflow)
NL (1) NL264871A (enrdf_load_stackoverflow)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5289687A (en) * 1992-03-30 1994-03-01 General Electric Company One-piece cowl for a double annular combustor

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN116025924B (zh) * 2023-01-13 2024-09-20 南京航空航天大学 一种加力燃烧室外凹腔火焰稳定器后壁面强制冷却装置

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB666062A (en) * 1947-02-28 1952-02-06 Lysholm Alf Gas turbine power plant
US2687010A (en) * 1947-11-03 1954-08-24 Power Jets Res & Dev Ltd Combustion apparatus
GB809514A (en) * 1956-04-03 1959-02-25 Bristol Aero Engines Ltd Improvements in or relating to combustion chambers
US2882681A (en) * 1953-02-24 1959-04-21 Lucas Industries Ltd Air-jacketed annular combustion chambers for jet-propulsion engines, gas turbines or like prime movers
US2884759A (en) * 1956-04-25 1959-05-05 Curtiss Wright Corp Combustion chamber construction
US2907171A (en) * 1954-02-15 1959-10-06 Lysholm Alf Combustion chamber inlet for thermal power plants
US2977760A (en) * 1955-03-16 1961-04-04 Bristol Aero Engines Ltd Annular combustion chambers for use with compressors capable of discharging combustion supporting medium with a rotary swirl through an annular outlet
US2982099A (en) * 1956-10-09 1961-05-02 Rolls Royce Fuel injection arrangement in combustion equipment for gas turbine engines

Family Cites Families (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
NL80817C (enrdf_load_stackoverflow) * 1947-11-03
BE489965A (enrdf_load_stackoverflow) * 1948-09-20
US2664703A (en) * 1949-06-28 1954-01-05 A V Roe Canada Ltd Preheater and vaporizer for gas turbine engines
DE1008963B (de) * 1955-03-28 1957-05-23 Snecma Vorrichtung zum Zuenden und Wiederzuenden von Brennkammern, insbesondere fuer Gasturbinen und Strahltriebwerke
US3088281A (en) * 1956-04-03 1963-05-07 Bristol Siddeley Engines Ltd Combustion chambers for use with swirling combustion supporting medium
GB831656A (en) * 1956-10-09 1960-03-30 Rolls Royce Improvements in or relating to combustion apparatus
DE1102491B (de) * 1957-10-19 1961-03-16 Bristol Siddeley Engines Ltd Verbrennungseinrichtung fuer ein Gasturbinentriebwerk
FR1206830A (fr) * 1958-05-19 1960-02-11 Rolls Royce Perfectionnements aux appareillages de combustion des moteurs à turbine à gaz

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB666062A (en) * 1947-02-28 1952-02-06 Lysholm Alf Gas turbine power plant
US2687010A (en) * 1947-11-03 1954-08-24 Power Jets Res & Dev Ltd Combustion apparatus
US2882681A (en) * 1953-02-24 1959-04-21 Lucas Industries Ltd Air-jacketed annular combustion chambers for jet-propulsion engines, gas turbines or like prime movers
US2907171A (en) * 1954-02-15 1959-10-06 Lysholm Alf Combustion chamber inlet for thermal power plants
US2977760A (en) * 1955-03-16 1961-04-04 Bristol Aero Engines Ltd Annular combustion chambers for use with compressors capable of discharging combustion supporting medium with a rotary swirl through an annular outlet
GB809514A (en) * 1956-04-03 1959-02-25 Bristol Aero Engines Ltd Improvements in or relating to combustion chambers
US2884759A (en) * 1956-04-25 1959-05-05 Curtiss Wright Corp Combustion chamber construction
US2982099A (en) * 1956-10-09 1961-05-02 Rolls Royce Fuel injection arrangement in combustion equipment for gas turbine engines

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5289687A (en) * 1992-03-30 1994-03-01 General Electric Company One-piece cowl for a double annular combustor

Also Published As

Publication number Publication date
CH390622A (de) 1965-04-15
GB903834A (en) 1962-08-22
NL264871A (enrdf_load_stackoverflow) 1964-06-10
DE1286332B (de) 1969-01-02
DE1224989B (de) 1966-09-15

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