EP2813763A1 - Injecteur de carburant et chambre de combustion - Google Patents

Injecteur de carburant et chambre de combustion Download PDF

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Publication number
EP2813763A1
EP2813763A1 EP14171645.6A EP14171645A EP2813763A1 EP 2813763 A1 EP2813763 A1 EP 2813763A1 EP 14171645 A EP14171645 A EP 14171645A EP 2813763 A1 EP2813763 A1 EP 2813763A1
Authority
EP
European Patent Office
Prior art keywords
air
fuel injector
downstream end
splitter
pilot
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.)
Withdrawn
Application number
EP14171645.6A
Other languages
German (de)
English (en)
Inventor
Ian Toon
Robert Dr. Hicks
Michael Dr. Whiteman
Waldemar Lazik
Imon-Kalyan Bagchi
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce Deutschland Ltd and Co KG
Rolls Royce PLC
Original Assignee
Rolls Royce Deutschland Ltd and Co KG
Rolls Royce PLC
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Rolls Royce Deutschland Ltd and Co KG, Rolls Royce PLC filed Critical Rolls Royce Deutschland Ltd and Co KG
Publication of EP2813763A1 publication Critical patent/EP2813763A1/fr
Withdrawn legal-status Critical Current

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Classifications

    • 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/34—Feeding into different combustion zones
    • F23R3/343—Pilot flames, i.e. fuel nozzles or injectors using only a very small proportion of the total fuel to insure continuous combustion
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23D—BURNERS
    • F23D11/00—Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space
    • F23D11/10—Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space the spraying being induced by a gaseous medium, e.g. water vapour
    • F23D11/106—Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space the spraying being induced by a gaseous medium, e.g. water vapour medium and fuel meeting at the burner outlet
    • F23D11/107—Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space the spraying being induced by a gaseous medium, e.g. water vapour medium and fuel meeting at the burner outlet at least one of both being subjected to a swirling motion
    • 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/34—Feeding into different combustion zones
    • F23R3/346—Feeding into different combustion zones for staged combustion

Definitions

  • the present invention relates to a fuel injector and a combustion chamber and in particular to a lean burn fuel injector and a gas turbine engine combustion chamber.
  • each fuel injector is located in a respective one of a plurality of apertures in an upstream end of the combustion chamber.
  • gas turbine engine combustion chamber One type of gas turbine engine combustion chamber is known as a rich burn combustion chamber and another type of gas turbine engine combustion chamber is known as a lean burn combustion chamber.
  • a lean burn type of combustion chamber the fuel and air is mixed such that the fuel to air equivalence ratio is less than one.
  • Lean burn gas turbine engine combustion chambers are being developed which use lean burn technology to reduce the emissions of nitrous oxides (NOx).
  • Lean burn gas turbine engine combustion chambers have lean burn fuel injectors, each of which comprises a pilot fuel injector and a main fuel injector, to enable lean combustion at higher air to fuel ratios than the stoichiometric air to fuel ratio, and to provide high thrusts with low NOx and to supply fuel to the pilot fuel injector at low thrusts to achieve required combustion efficiency, lean blow out margin and altitude relight capability.
  • One type of fuel injector for a lean burn type of combustion chamber comprises a pilot fuel injector and a main fuel injector.
  • the pilot fuel injector is provided between two sets of air swirlers and the main fuel injector is provided between a further two sets of air swirlers.
  • the pilot fuel injector and the main fuel injector are arranged concentrically and the main fuel injector is arranged around the pilot fuel injector.
  • the first two sets of air swirlers provide swirling flows of air which atomise the fuel from the pilot fuel injector and the second two sets of air swirlers provide swirling flows of air which atomise the fuel from the main fuel injector.
  • Each air swirler comprises a plurality of circumferentially spaced radially extending swirl vanes and the swirl vanes extend between concentric members.
  • the four sets of air swirlers are arranged concentrically.
  • This type of fuel injector is also provided with first and second air splitters between the pilot fuel injector and the main fuel injector.
  • the first air splitter has a converging downstream portion and the second air splitter has a diverging downstream portion.
  • the present invention seeks to provide a novel fuel injector for a gas turbine engine combustion chamber which reduces or overcomes the above mentioned problem.
  • the present invention provides a fuel injector comprising a pilot fuel injector and a main fuel injector, the pilot fuel injector comprising at least one pilot air swirler, the main fuel injector comprising a main air blast fuel injector located between an inner main air swirler and an outer main air swirler, a first air splitter located between the at least one pilot air swirler and the inner main air swirler and a second air splitter located between the at least one pilot air swirler and the inner main air swirler, the first air splitter comprising a downstream portion converging to a downstream end, the second air splitter comprising a downstream portion diverging to a downstream end, the downstream end of the second air splitter is downstream of the downstream end of the first air splitter, the downstream end of the second air splitter being downstream of the downstream end of a member defining the outer surface of the outer main air swirler, and the ratio of the distance from the downstream end of the first air splitter to the downstream end of the second air splitter to the diameter of the downstream end of the second air splitter is in
  • the ratio of the distance from the downstream end of the first air splitter to the downstream end of the second air splitter to the diameter of the downstream end of the second air splitter may be in the range of 0.24 to 0.28.
  • the ratio of the distance from the downstream end of the first air splitter to the downstream end of the second air splitter to the diameter of the downstream end of the second air splitter may be in the range of 0.25 to 0.27.
  • the pilot fuel injector may comprise a pilot air blast fuel injector located between an inner pilot air swirler and an outer pilot air swirler.
  • the second air splitter may be located between the first air splitter and the inner main air swirler, an additional air swirler is provided between the first air splitter and the second air splitter to direct air over the second air splitter.
  • the outer main air swirler may comprise a plurality of swirl vanes arranged in an annular duct, the annular duct is defined by a radially inner surface of an outer wall and a radially outer surface of an inner wall.
  • the radially inner surface of the outer wall of the annular duct may converge to a minimum diameter downstream of the swirl vanes, the radial width of the annular duct at the trailing edges of the swirl vanes is in the range of 1.1 to 1.3 times the radial width of the annular duct at the minimum diameter of the radially inner surface of the outer wall of the annular duct.
  • the radially outer surface of the inner wall of the annular duct may converge to a minimum diameter downstream of the swirl vanes, the radial distance of convergence of the radially outer surface of the inner wall is in the range of 0.5 to 1.0 times the radial width of the annular duct at the minimum diameter of the radially inner surface of the outer wall of the annular duct.
  • the axial length of the annular duct from the trailing edges of the swirl vanes to the minimum diameter of the radially inner surface of the outer wall of the annular duct may be in the range of 1.7 to 2.5 times the radial width of the annular duct at the minimum diameter of the radially inner surface of the outer wall of the annular duct.
  • the downstream end of the second air splitter may be downstream of the downstream end of the inner wall of the annular duct.
  • the fuel injector may be provided in a combustion chamber.
  • the combustion chamber may comprise an igniter and the igniter is positioned downstream of the at least one fuel injector.
  • the combustion chamber may be a gas turbine engine combustion chamber.
  • the gas turbine engine may be a turbofan gas turbine engine, a turbo-jet gas turbine engine, a turbo-shaft gas turbine engine or a turbo-prop gas turbine engine.
  • the gas turbine engine may be an aero gas turbine engine, a marine gas turbine engine, an industrial gas turbine engine or an automotive gas turbine engine.
  • the present invention also provides a method of operating a combustion chamber, the combustion chamber comprising an igniter and at least one fuel injector, the igniter being positioned downstream of the at least one fuel injector, the fuel injector comprising a pilot fuel injector and a main fuel injector, the pilot fuel injector comprising at least one pilot air swirler, the main fuel injector comprising a main air blast fuel injector located between an inner main air swirler and an outer main air swirler, a first air splitter located between the at least one pilot air swirler and the inner main air swirler and a second air splitter located between the at least one pilot air swirler and the inner main air swirler, the first air splitter comprising a downstream portion converging to a downstream end, the second air splitter comprising a downstream portion diverging to a downstream end, the downstream end of the second air splitter is downstream of the downstream end of the first air splitter, the downstream end of the second air splitter is downstream of the downstream end of a member defining the outer surface of the outer main air swirler and
  • a turbofan gas turbine engine 10 as shown in figure 1 , comprises in flow series an intake 11, a fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, a combustion chamber 15, a high pressure turbine 16, an intermediate pressure turbine 17, a low pressure turbine 18 and an exhaust 19.
  • the high pressure turbine 16 is arranged to drive the high pressure compressor 14 via a first shaft 26.
  • the intermediate pressure turbine 17 is arranged to drive the intermediate pressure compressor 13 via a second shaft 28 and the low pressure turbine 18 is arranged to drive the fan 12 via a third shaft 30.
  • air flows into the intake 11 and is compressed by the fan 12.
  • a first portion of the air flows through, and is compressed by, the intermediate pressure compressor 13 and the high pressure compressor 14 and is supplied to the combustion chamber 15.
  • Fuel is injected into the combustion chamber 15 and is burnt in the air to produce hot exhaust gases which flow through, and drive, the high pressure turbine 16, the intermediate pressure turbine 17 and the low pressure turbine 18.
  • the hot exhaust gases leaving the low pressure turbine 18 flow through the exhaust 19 to provide propulsive thrust.
  • a second portion of the air bypasses the main engine to provide propulsive thrust.
  • the combustion chamber 15, as shown more clearly in figure 2 is an annular combustion chamber and comprises a radially inner annular wall structure 40, a radially outer annular wall structure 42 and an upstream end wall structure 44.
  • the radially inner annular wall structure 40 comprises a first annular wall 46 and a second annular wall 48.
  • the radially outer annular wall structure 42 comprises a third annular wall 50 and a fourth annular wall 52.
  • the second annular wall 48 is spaced radially from and is arranged radially around the first annular wall 46 and the first annular wall 46 supports the second annular wall 48.
  • the fourth annular wall 52 is spaced radially from and is arranged radially within the third annular wall 50 and the third annular wall 50 supports the fourth annular wall 52.
  • the upstream end of the first annular wall 46 is secured to the upstream end wall structure 44 and the upstream end of the third annular wall 50 is secured to the upstream end wall structure 44.
  • the upstream end wall structure 44 has a plurality of circumferentially spaced apertures 54 and each aperture 54 has a respective one of a plurality of fuel injectors 56 located therein.
  • the fuel injectors 56 are arranged to supply fuel into the annular combustion chamber 15 during operation of the gas turbine engine 10.
  • Each fuel injector 56 comprises a fuel feed arm 58 and a fuel injector head 60, as shown in figure 2 .
  • the fuel feed arm 58 has a first internal passage 62 for the supply of pilot fuel to the fuel injector head 60.
  • the fuel feed arm 58 has a second internal fuel passage 64 for the supply of main fuel to the fuel injector head 60.
  • Each fuel injector head 60 locates coaxially within a respective one of the apertures 54 in the upstream end wall 44 of the combustion chamber 15.
  • the fuel injector head 60 has an axis Y and the fuel feed arm 58 extends generally radially with respect to the axis Y of the fuel injector head 60 and also generally radially with respect to the axis X of the turbofan gas turbine engine 10.
  • the axis Y of each fuel injector head 60 is generally aligned with the axis of the corresponding aperture 54 in the upstream end wall 44 of the combustion chamber 15.
  • a fuel injector 56 according to the present invention is shown more clearly in figure 3 .
  • the fuel injector head 60 comprises a first generally cylindrical member 66, a second generally annular member 68 spaced coaxially around the first member 66 and a third generally annular member 70 spaced coaxially around the second member 68.
  • a plurality of circumferentially spaced swirl vanes 72 extend radially between the first member 66 and the second member 68 to form a first air swirler 71.
  • the second member 68 has a greater axial length than the first member 66 and the first member 66 is positioned at an upstream end 68A of the second member 68 and a generally annular duct 74 is defined between the first member 66 and the second member 68 and the swirl vanes 72 extend radially across the annular duct 74.
  • a generally cylindrical duct 76 is defined radially within the second member 68 at a position downstream of the first member 66.
  • the second member 68 has one or more internal fuel passages 78 which are arranged to receive fuel from the first internal fuel passage 62 in the fuel feed arm 58.
  • the one or more fuel passages 78 are arranged to supply fuel to a fuel swirler 80 which supplies a film of fuel onto a radially inner surface 82 at a downstream end 68B of the second member 68.
  • a plurality of circumferentially spaced swirl vanes 84 extend radially between the second member 68 and the third member 70 to form a second air swirler 83.
  • the second member 68 has a greater axial length than the third member 70 and the third member 70 is positioned at the downstream end 68B of the second member 68 and a generally annular duct 86 is defined between the second member 68 and the third member 70 and the swirl vanes 84 extend across the annular duct 86.
  • the downstream end 70B of the third member 70 is conical and is convergent in a downstream direction.
  • the radially inner surface 82 of the second member 68, the radially outer surface of the second member 68 and the radially inner surface of the third member 70 are all circular in cross-section in a plane perpendicular to the axis Y of the fuel injector head 60 of the fuel injector 56.
  • the downstream end 70B of the third member 70 is downstream of the downstream end 68B of the second member 68 and the downstream end 68B of the second member 68 is downstream of the downstream end 66B of the first member 66.
  • the pilot fuel is supplied by the internal fuel passages 78 and the fuel swirler 80 onto the radially inner surface 82 of the second member 68 and the pilot fuel is atomised by swirling flows of air A and B from the swirl vanes 72 and 84 of the first and second air swirlers 71 and 83 respectively.
  • the first member 66, the second member 68, the third member 70, the first swirler 71 and the second swirler 83 form the pilot fuel injector 59 and in this case the pilot fuel injector 59 is a pilot air blast fuel injector.
  • the first and second air swirlers 71 and 83 are the inner and outer pilot air swirlers respectively for the pilot air blast fuel injector 59.
  • the fuel injector head 60 also comprises a fourth generally annular member 88 spaced coaxially around the third member 70, a fifth member 90 spaced coaxially around the fourth member 88 and a sixth member 92 spaced coaxially around the fifth member 90.
  • a plurality of circumferentially spaced swirl vanes 94 extend radially between the fourth member 88 and the fifth member 90 to form a third air swirler 93.
  • the fifth member 90 has a greater axial length than the fourth member 88 and the fourth member 88 is positioned at the downstream end 90B of the fifth member 90 and a generally annular duct 96 is defined between the fourth member 88 and the fifth member 90 and the swirl vanes 94 extend across the annular duct 96.
  • the fifth member 90 has one or more internal fuel passages 98 which are arranged to receive fuel from the second internal fuel passage 64 in the fuel feed arm 58.
  • the one or more fuel passages 98 are arranged to supply fuel to a fuel swirler 100 which supplies a film of fuel onto the radially inner surface 102 at the downstream end 90B of the fifth member 90.
  • a plurality of circumferentially spaced swirl vanes 104 extend radially between the fifth member 90 and the sixth member 92 to form a fourth air swirler 103.
  • a generally annular duct 106 is defined between the downstream end 90B of the fifth member 90 and the sixth member 92 and the swirl vanes 104 extend across the annular duct 106.
  • the downstream end 88B of the fourth member 88 is conical and is divergent in a downstream direction.
  • the main fuel supplied by the internal fuel passages 98 and fuel swirler 100 onto the radially inner surface 102 of the fifth member 90 is atomised by swirling flows of air C and D from the swirl vanes 94 and 104 of the third and fourth air swirlers 93 and 103 respectively.
  • the fourth member 88, the fifth member 90, the sixth member 92, the third swirler 93 and the fourth swirler 103 form the main fuel injector 61 and in this case the main fuel injector 61 is a main air blast fuel injector 61.
  • the third and fourth air swirlers 93 and 103 are the inner and outer main air swirlers respectively for the main air blast fuel injector 61.
  • the third member 70 and the fourth member 88 are first and second splitters respectively and are used to separate, or split, the flows of fuel and air from the pilot fuel injector 59 from the fuel and air from the main fuel injector 61.
  • the fuel injector head 60 also comprises a plurality of circumferentially spaced swirl vanes 108 which extend radially between the third member 70 and the fourth member 88 to form a fifth air swirler 107.
  • the swirl vanes 108 of the fifth air swirler 107 provide a swirling flow of air E over the radially inner surface of the fourth member 88.
  • the sixth member 92 has a radially inner surface 110, the radially inner surface 110 of the sixth member 92 is generally circular in cross-section in a plane perpendicular to the axis Y of the fuel injector head 60 of the fuel injector 56.
  • the radially inner surface 110 of the downstream end 92B of the sixth member 92 converges to a minimum diameter 114 at a plane arranged perpendicular to the axis Y of the fuel injector head 60 containing the downstream end 90B of the fifth member 90 and then the radially inner surface 110 of the downstream end 92B of the sixth member 92 diverges downstream of the downstream end 90B of the fifth member 90.
  • the radially inner surface 110 of the downstream end 92B of the sixth member 92 downstream of the swirl vanes 104 converges to the minimum diameter 114.
  • the fifth member 90 has a radially outer surface 112, the radially outer surface 112 of the fifth member 90 is generally circular in cross-section in a plane perpendicular to the axis Y of the fuel injector head 60 of the fuel injector 56.
  • the radially inner surface 102 of the fifth member 90 is generally circular in cross-section in a plane perpendicular to the axis Y of the fuel injector head 60 of the fuel injector 56.
  • the radially outer surface 112 of the downstream end 90B of the fifth member 90 converges to a minimum diameter 116 at a plane arranged perpendicular to the axis Y of the fuel injector head 60.
  • the minimum diameter 116 of the radially outer surface 112 of the downstream end 90B of the fifth member 90 is positioned upstream of the minimum diameter 114 of the radially inner surface 110 of the downstream end 92B of the sixth member 92.
  • Figure 4 shows a fuel injector 56 having a half of the fuel injector according to the present invention and a half of the fuel injector not according to the present invention.
  • the upper half of the fuel injector head 60 of the fuel injector 56 is according to the present invention whereas the lower half of the fuel injector head 60 is not according to the present invention.
  • the lower half of the fuel injector head 60 of the fuel injector 56 has the downstream end 70B of the third member 70 positioned too far downstream relative to the downstream end 88B of the fourth member 88 and this results in the pilot fuel Fp from the pilot fuel injector 59 being directed and contained along the axis Y of the fuel injector head 60 and mixing with the main fuel from the main fuel injector 61 downstream of the igniter I location.
  • the upper half of the fuel injector 60 of the fuel injector 56 has the downstream end 70B of the third member 70 positioned so as to produce an "S" shaped flow path for the pilot fuel Fp' supplied from the pilot fuel injector 59 to the main fuel supplied by the main fuel injector 61.
  • the "S" shaped flow path for the pilot fuel Fp' to the main fuel is necessary to minimise emissions and maximise combustion efficiency when the pilot fuel mixes with the main fuel and air flow.
  • the "S" shaped flow path for the pilot fuel Fp' to the main fuel is necessary to achieve relight ignition capability by providing a suitable main fuel and air flow cone angle and entrainment of the pilot fuel into the main fuel and air flow upstream of the igniter location I.
  • the "S" shaped flow path for the pilot fuel Fp' to the main fuel and air flow provides an acceptable temperature at the downstream end 88B of the fourth member 88.
  • Figure 5 shows a fuel injector 56 having a half of the fuel injector according to the present invention and a half of the fuel injector not according to the present invention.
  • the upper half of the fuel injector head 60 of the fuel injector 56 is according to the present invention whereas the lower half of the fuel injector head 60 is not according to the present invention.
  • the lower half of the fuel injector head 60 of the fuel injector 56 has the downstream end 70B of the third member 70 positioned too far upstream relative to the downstream end 88B of the fourth member 88 and this results in the pilot fuel Fp from the pilot fuel injector 59 being directed radially outwards away from the axis Y of the fuel injector head 60 and towards and onto the downstream end 88B of the fourth member 88 producing unacceptable temperature for the downstream end 88B of fourth member 88 due to the impingement of the pilot flame on the downstream end 88 of the fourth member 88.
  • Figure 6 shows that the downstream end 88B of the fourth member 88 of the fuel injector head 60 has a diameter D and the axial distance between the downstream end 70B of the third member 70 and the downstream end 88B of the fourth member 88 of the fuel injector head 60 has a distance L.
  • An acceptable location for the downstream end 70B of the third member 70 and the downstream end 88B of the fourth member 88 is one where the ratio of the axial distance L to the diameter D is in the range of 0.22 to 0.3 to achieve the required "S" shape flow path for the pilot fuel Fp' to the main fuel and air flow in order to achieve the required combustion characteristics discussed previously.
  • the ratio of the distance from the downstream end 70B of the first air splitter 70 to the downstream end 88B of the second air splitter 88 to the diameter of the downstream end 88B of the second air splitter 88 is in the range of 0.22 to 0.30.
  • the ratio of the axial distance L to the diameter D is in the range of 0.24 to 0.28 to achieve the required "S" shape flow path for the pilot fuel Fp' to the main fuel and air flow and more preferably the ratio of the axial distance L to the diameter D is in the range of 0.25 to 0.27 to achieve the required "S" shape flow path for the pilot fuel Fp' to the main fuel and air flow.
  • the advantage of the present invention is that the required aerodynamics are achieved for the fuel injector resulting in minimisation of NOx, maximising combustion efficiency, achieving the required ignition characteristics and lean blow out level and the required temperature for the second splitter.
  • Figure 7 shows the fourth swirler, the main outer air swirler, 103 of the fuel injector 56 in more detail.
  • the radial width L2 of the annular duct 106 at the trailing edges of the swirl vanes 104 is in the range of 1.1 to 1.3 times the radial width T of the annular duct 106 at the minimum diameter 114 of the radially inner surface 110 of the sixth member 92.
  • the axial length L1 of the annular duct 106 from the trailing edges of the swirl vanes 104 to the minimum diameter 114 of the radially inner surface 110 of the sixth member 92 is in the range of 1.7 to 2.5 times the radial width T of the annular duct 106 at the minimum diameter 114 of the radially inner surface 110 of the sixth member 92.
  • the radial distance L3 of convergence of the radially outer surface 112 of the fifth member 90 is in the range of 0.5 to 1.0 times the radial width T of the annular duct 106 at the minimum diameter 114 of the radially inner surface 110 of the sixth member 92.
  • the length L4 of the radially inner surface 110 of the sixth member 90 downstream of the minimum diameter 114 is in the range of 1.0 to 1.2 times the radial width T of the annular duct 106 at the minimum diameter 114 of the radially inner surface 110 of the sixth member 92.
  • the axial length L5 from the position of minimum diameter 116 of the radially outer surface 112 of the fifth member 90 to the position of minimum diameter 114 of the radially inner surface 110 of the sixth member 92 is in the range of 0 to 0.4 times the radial width T of the annular duct 106 at the minimum diameter 114 of the radially inner surface 110 of the sixth member 92.
  • the radius R1 of the radially inner surface 110 of the sixth member 92 at the minimum diameter 114 is in the range of 0.9 to 1.1 times the radial width T of the annular duct 106 at the minimum diameter 114 of the radially inner surface 110 of the sixth member 92.
  • the radius R2 of the radially inner surface 110 of the sixth member 92 at the trailing edges of the swirl vanes 104 is in the range 0.9 to 1.1 times the radial width T of the annular duct 106 at the minimum diameter 114 of the radially inner surface 110 of the sixth member 92.
  • the radius R3 of the radially outer surface 112 of the fifth member 90 at the trailing edges of the swirl vanes 104 is in the range 0.9 to 1.1 times the radial width T of the annular duct 106 at the minimum diameter 114 of the radially inner surface 110 of the sixth member 92.
  • the radially inner surface 110 of the sixth member 92 is defined by a tangent connecting the radius R2 and the radius R1 and the length L4 is the length of a tangent from the radius R2 at the angle ⁇ 1.
  • the angle ⁇ 1 of the divergent portion of the radially inner surface 110 of the sixth member 92 downstream of the position of minimum diameter 114 is in the range of 35° to 45°.
  • the radially outer surface 112 of the fifth member 90 is defined by a tangent connecting the radius R3 and the minimum diameter point 116.
  • the advantage of the arrangement of the arrangement of the fourth swirler, the main outer air swirler, 103 of the fuel injector 56 as described is that the aerodynamic flow of air from the fourth swirler 103 attaches to the radially inner surface 110 of the sixth member 92 both upstream and downstream of the minimum diameter 114 of the radially inner surface 110 and the air from the fourth swirler 103 attaches to the radially outer surface 112 of the fifth member 90.
  • the angle ⁇ 1 of the diverging portion of the radially inner surface 110 of the downstream end 92B of the sixth member 92 is optimised to direct the flow of fuel and air towards the igniter, to enable ignition of the fuel and air.
  • the presence of the diverging portion of the radially inner surface 110 of the downstream end 92 of the sixth member 92 downstream of the minimum diameter 114 of the radially inner surface 110 of the sixth member 92 stabilises the flow cone angle temporarily, which is beneficial in suppressing combustion instabilities and producing consistent ignition.
  • this arrangement prevents flow separation of the air flow from, or significant boundary layer thickness build up on, the radially inner surface 110 of the sixth member 92 and the radially outer surface 112 of the fifth member 90 which impair mixing of the air from the third and fourth swirlers 93 and 103 respectively and the mixing of the fuel into the air from the third and fourth air swirlers 93 and 103.
  • the pilot fuel is supplied by the internal fuel passages 78 and the fuel swirler 80 onto the radially inner surface 82 of the second member 68 and the pilot fuel is atomised by swirling flows of air A and B from the swirl vanes 72 and 84 of the first and second air swirlers 71 and 83 respectively.
  • the first member 66, the second member 68, the third member 70, the first swirler 71 and the second swirler 83 form the pilot fuel injector 59 and in this case the pilot fuel injector 59 is a pilot air blast fuel injector.
  • the main fuel supplied by the internal fuel passages 98 and fuel swirler 100 onto the radially inner surface 102 of the fifth member 90 is atomised by swirling flows of air C and D from the swirl vanes 94 and 104 of the third and fourth air swirlers 93 and 103 respectively.
  • the fourth member 88, the fifth member 90, the sixth member 92, the third swirler 93 and the fourth swirler 103 form the main fuel injector 61 and in this case the main fuel injector 61 is a main air blast fuel injector 61.
  • the pilot fuel and the main fuel are both atomised by high air velocity accelerating the fuel from the respective pre-filming surface, and thus the fuel pressure does not affect the fuel atomisation.
  • An advantage of this type of fuel injector is that pilot fuel only is supplied to the pilot fuel injector 59 for and during a "cold day” take-off, in order to obtain satisfactory combustion efficiency.
  • the pilot fuel injector is provided with an increased number of fuel passages 78, fuel passages 78 with a greater diameter etc. in order to provide a greater flow of fuel to the pilot fuel injector 59 during a "cold day” take off.
  • pilot fuel is supplied to the pilot fuel injector 59 and main fuel is supplied to the main fuel injector 61 for and during a "cold day” take-off, unsatisfactory combustion efficiency, lower than a predetermined level of efficiency, is achieved.
  • Unacceptable combustion efficiency is obtained by supplying pilot fuel and main fuel to the fuel injector because the temperature in the combustion chamber is not hot enough for the main fuel to burn at lean conditions.
  • the modified pilot fuel injector 59 also enables relight ignition to be achieved by supplying pilot fuel only to the pilot fuel injector 59.
  • the present invention has been described with reference to a turbofan gas turbine engine it is equally possible to use the present invention on a combustion chamber of a turbo-jet gas turbine engine, a turbo-shaft gas turbine engine or a turbo-prop gas turbine engine.
  • the present invention has been described with reference to an aero gas turbine engine it is equally possible to use the present invention on a combustion chamber of a marine gas turbine engine, an industrial gas turbine engine or an automotive gas turbine engine.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Fuel-Injection Apparatus (AREA)
EP14171645.6A 2013-06-10 2014-06-09 Injecteur de carburant et chambre de combustion Withdrawn EP2813763A1 (fr)

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CN107559882A (zh) * 2017-07-24 2018-01-09 西北工业大学 一种轴向分级低污染燃烧室
EP3176505B1 (fr) * 2015-11-23 2020-09-30 Rolls-Royce plc Injecteur de carburantet son procédé de fabrication
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