EP0554325B1 - Gasturbinenbrennkammer und deren arbeitsweise - Google Patents

Gasturbinenbrennkammer und deren arbeitsweise Download PDF

Info

Publication number
EP0554325B1
EP0554325B1 EP91918884A EP91918884A EP0554325B1 EP 0554325 B1 EP0554325 B1 EP 0554325B1 EP 91918884 A EP91918884 A EP 91918884A EP 91918884 A EP91918884 A EP 91918884A EP 0554325 B1 EP0554325 B1 EP 0554325B1
Authority
EP
European Patent Office
Prior art keywords
fuel
combustion chamber
mixing zone
air
air mixing
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
Application number
EP91918884A
Other languages
English (en)
French (fr)
Other versions
EP0554325A1 (de
Inventor
Ian James Toon
Stephen James O'dell
Jeffrey Douglas Willis
Paul Royston Currin
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 PLC
Original Assignee
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
Priority claimed from GB909023004A external-priority patent/GB9023004D0/en
Priority claimed from GB919118945A external-priority patent/GB9118945D0/en
Application filed by Rolls Royce PLC filed Critical Rolls Royce PLC
Publication of EP0554325A1 publication Critical patent/EP0554325A1/de
Application granted granted Critical
Publication of EP0554325B1 publication Critical patent/EP0554325B1/de
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

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
    • F23R3/346—Feeding into different combustion zones for staged combustion
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23C—METHODS OR APPARATUS FOR COMBUSTION USING FLUID FUEL OR SOLID FUEL SUSPENDED IN  A CARRIER GAS OR AIR 
    • F23C6/00—Combustion apparatus characterised by the combination of two or more combustion chambers or combustion zones, e.g. for staged combustion
    • F23C6/04—Combustion apparatus characterised by the combination of two or more combustion chambers or combustion zones, e.g. for staged combustion in series connection
    • F23C6/045—Combustion apparatus characterised by the combination of two or more combustion chambers or combustion zones, e.g. for staged combustion in series connection with staged combustion in a single enclosure
    • F23C6/047—Combustion apparatus characterised by the combination of two or more combustion chambers or combustion zones, e.g. for staged combustion in series connection with staged combustion in a single enclosure with fuel supply in stages
    • 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
    • F23R3/12—Air inlet arrangements for primary air inducing a vortex
    • F23R3/14—Air inlet arrangements for primary air inducing a vortex by using swirl vanes

Definitions

  • the present invention relates to a gas turbine combustion chamber as disclosed e.g. in FR-A- 2 085 942, and to a method of operating a gas turbine engine combustion chamber.
  • staged combustion is required in order to minimise the quantity of the oxides of nitrogen (NOx) produced.
  • NOx oxides of nitrogen
  • the fundamental way to reduce emissions of nitrogen oxides is to reduce the combustion reaction temperature, and this requires premixing of the fuel and all the combustion air before combustion takes place.
  • the oxides of nitrogen (NOx) are commonly reduced by a method which uses two stages of fuel injection.
  • Our UK patent no. 1489339 discloses two stages of fuel injection to reduce NOx.
  • staged combustion both stages of combustion seek to provide lean combustion and hence the low combustion temperatures required to minimise NOx.
  • lean combustion means combustion of fuel in air where the fuel to air ratio is low i.e. less than the stoichiometric ratio.
  • the present invention seeks to provide a novel gas turbine combustion chamber, and a novel method of operating a gas turbine engine combustion chamber.
  • the present invention provides a gas turbine engine combustion chamber comprising first air intake means, primary fuel injector means and a first fuel and air mixing zone, the first fuel and air mixing zone being defined by at least one annular wall and an upstream wall connected to the upstream end of the annular wall, the upstream wall having at least one aperture, the first air intake means comprising at least one first radial flow swirler and at least one second radial flow swirler, each first radial flow swirler being arranged to supply air into the first fuel and air mixing zone through said aperture, each second radial flow swirler being arranged to supply air into the first fuel and air mixing zone through said aperture, each first radial flow swirler being positioned axially downstream of the respective second radial flow swirler with respect to the axis of the combustion chamber, each first radial flow swirler being arranged to swirl air in the opposite direction to the respective second radial flow swirler, the primary fuel injector means being arranged to supply fuel into at least one of the passages defined between the vanes of each of the first
  • each pilot fuel injector is aligned with a respective one of the apertures to supply fuel into the first fuel and air mixing zone.
  • the primary fuel injector means is arranged to supply fuel into all the passages defined between the vanes of the first radial flow swirler.
  • the primary fuel injector means is arranged to supply fuel into all the passages defined between the vanes of the second radial flow swirler.
  • the primary fuel injector means is arranged to supply fuel into the radially outer region of the passages between the vanes.
  • the primary fuel injector means may comprise a hollow cylindrical member arranged to extend axially with respect to the combustion chamber, the cylindrical member has a plurality of apertures spaced apart axially along the cylindrical member to inject fuel into the passages.
  • the apertures may be arranged to direct the fuel radially inwardly.
  • the primary fuel injector means may be arranged to inject gas fuel or evaporated liquid fuel.
  • the pilot fuel injector may be arranged to inject gas fuel, or liquid fuel.
  • the combustion chamber may be tubular and has a single aperture in its upstream wall.
  • the combustion chamber may further comprise secondary air intake means, secondary fuel injector means and a secondary fuel and air mixing zone, the secondary fuel and air mixing zone is annular and surrounds the first fuel and air mixing zone, the secondary fuel and air mixing zone being defined at its radially outer extremity by a second annular wall, the secondary fuel injector means being arranged to supply fuel into the upstream end of the secondary fuel and air mixing zone, the secondary fuel and air mixing zone being in fluid flow communication at its downstream end with the interior of the combustion chamber downstream of the first fuel and air mixing zone.
  • the secondary air intake may be downstream of the first air intake means.
  • the secondary fuel and air mixing zone may be defined at its radially inner extremity by a third annular wall.
  • the annular wall may have a first portion defining the first fuel and air mixing zone, a second portion of increased diameter downstream of the first portion and a third frusto conical portion interconnecting the first and second portions.
  • the third conical portion may have a plurality of equicircumferentially spaced apertures arranged to direct the secondary fuel and mixture from the secondary fuel and air mixing zone as a plurality of jets in a downstream direction towards the centre line of the combustion chamber.
  • the apertures may be slots.
  • the downstream end of the second annular wall may be secured to the third conical portion of the annular wall.
  • Cooling air may be supplied to an annular chamber defined between the annular wall and the third annular wall.
  • the secondary fuel injector means may comprise a plurality of equi-circumferentially spaced fuel injectors.
  • the secondary fuel injector means may be arranged to inject gas fuel or evaporated liquid fuel.
  • the downstream end of the first portion of the annular wall reduces in diameter to a throat.
  • the combustion chamber may comprise tertiary air intake means, tertiary fuel injector means and a tertiary fuel and air mixing zone, the tertiary fuel and air mixing zone is annular and surrounds the secondary combustion zone, the tertiary fuel and air mixing zone is defined at its radially outer extremity by a fourth annular wall, the tertiary fuel injector means is arranged to supply fuel into the upstream end of the tertiary fuel and air mixing zone, the tertiary fuel and air mixing zone is in fluid flow communication at its downstream end with a tertiary combustion zone in the interior of the combustion chamber downstream of the secondary combustion zone.
  • the annular wall may have a fourth portion of larger diameter than the second portion downstream of the second portion and defining the tertiary combustion zone, a fifth frusto conical portion interconnecting the second and fourth portions.
  • the downstream end of the second portion of the annular wall may reduce in diameter to a throat.
  • the tertiary air intake may be downstream of the second air intake means.
  • the tertiary fuel and air mixing zone may be defined at its radially inner extremity by a fifth annular wall.
  • the fifth conical portion may have a plurality of equi-circumferentially spaced apertures arranged to direct the tertiary fuel and air mixture from the tertiary fuel and air mixing zone as a plurality of jets in a downstream direction towards the centreline of the combustion chamber.
  • the apertures may be slots.
  • the downstream end of the fourth annular wall may be secured to the fifth conical portion of the annular wall.
  • the tertiary fuel injector means may comprise a plurality of equi-circumferentially spaced fuel injectors.
  • the tertiary fuel injectors means may be arranged to inject gas fuel or evaporated liquid fuel.
  • Fuel may only be supplied from the pilot fuel injector into the first fuel and air mixing zone from the start of operation of the gas turbine engine until a predetermined output power level is obtained, fuel is supplied from the primary fuel injector means into at least one of the passages defined between the vanes of the first radial flow swirler and into at least one of the passages defined between the vanes of the second radial flow swirler to flow into the first fuel and air mixing zone for output power levels greater than the predetermined level, and simultaneously fuel is supplied from the secondary fuel injector means into the secondary fuel and air mixing zone to flow into the interior of the combustion chamber downstream of the first fuel and air mixing zone.
  • Fuel may be supplied from the pilot fuel injector only into the first fuel and air mixing zone from the start of operation of the gas turbine engine until a predetermined output power level is obtained, supplying fuel from the primary fuel injector means into at least one of the passages defined between the vanes of the first radial flow swirler and into at least one of the passages defined between the vanes of the second radial flow swirler to flow into the first fuel and air mixing zone for output power levels greater than a predetermined level, and simultaneously supplying fuel into the secondary fuel and air mixing zone to flow into the secondary combustion zone in the interior of the combustion chamber downstream of the first fuel and air mixing zone, supplying fuel into the tertiary fuel and air mixing zone to flow into the tertiary combustion zone in the interior of the combustion chamber downstream of the secondary combustion zone for output power levels greater than a second predetermined level and for ambient air temperatures greater than a predetermined temperature.
  • the predetermined output power level may be 35 to 40% power.
  • Figure 1 is a view of a gas turbine engine having a combustion chamber assembly and fuel injector according to the present invention.
  • Figure 2 is an enlarged longitudinal cross-sectional view through the combustion chamber shown in Figure 1.
  • Figure 3 is a further enlarged longitudinal cross-sectional view through the upstream end of the combustion chamber assembly shown in Figure 2.
  • Figure 4 is a cross-section in the direction of arrows G-G in figure 3, and
  • Figure 5 is a cross-sectional view in the direction of arrows H-H in Figure 3.
  • Figure 6 is a graph of percentage base load fuel flow versus percentage load for the combustion chamber shown in Figure 3.
  • Figure 7 is an enlarged longitudinal cross-sectional view through the upstream end of an alternative combustion chamber assembly according to the present invention.
  • Figure 8 is an enlarged longitudinal cross-sectional view through the upstream end of a further combustion chamber assembly according to the present invention.
  • An industrial gas turbine engine 10 shown in Figure 1, comprises in axial flow series an inlet 12, a compressor section 14, a combustion chamber assembly 16, a turbine section 18, a power turbine section 20 and an exhaust 22.
  • the turbine section 18 is arranged to drive the compressor section 14 via one or more shafts (not shown).
  • the power turbine section 20 is arranged to drive an electrical generator 26, via a shaft 24.
  • the power turbine section 20 may be arranged to provide drive for other purposes.
  • the operation of the gas turbine engine 10 is quite conventional, and will not be discussed further.
  • the combustion chamber assembly 16 is shown more clearly in Figures 2 to 5.
  • a plurality of compressor outlet guide vanes 28 are provided at the axially downstream end of the compressor section 14, to which is secured at their radially inner ends an inner annular wall 30 which defines the inner surface of an annular chamber 34.
  • a diffuser is defined between an annular wall 32 and the upstream portion of the inner annular wall 30. The downstream end of the inner annular wall 30 is secured to the radially inner ends of a row of nozzle guide vanes 90 which direct hot gases from the combustion chamber assembly 16 into the turbine section 18.
  • the combustion chamber assembly 16 comprises a plurality of equally circumferentially spaced tubular combustion chambers 36.
  • the axes of the tubular combustion chambers 36 are arranged to extend in generally radial directions.
  • the inlets of the tubular combustion chambers 36 are at their radially outermost ends and their outlets are at their radially innermost ends.
  • Each of the tubular combustion chambers 36 comprises an upstream wall 44 secured to the upstream end of an annular wall 37.
  • a first, upstream, portion 38 of the annular wall 37 defines a first fuel and air mixing zone 64, and a second, downstream portion 42 of the annular wall is interconnected with the first portion 38 by a third portion 40.
  • the second portion 42 of the annular wall has a greater diameter than the first portion 38, and the third portion 40 is frusto conical.
  • a plurality of equally circumferentially spaced transition ducts 46 are provided, and each of the transition ducts 46 has a circular cross-section at its upstream end.
  • the upstream end of each of the transition ducts 46 is located coaxially around the downstream end of a corresponding one of the tubular combustion chambers 36, and each of the transition ducts 46 connects and seals with an angular section of the nozzle guide vanes 90.
  • a plurality of cylindrical casings 48 are provided, and each cylindrical casing 48 is located coaxially around a respective one of the tubular combustion chambers 36. Each cylindrical casing 48 is secured to a respective boss 52 on an annular engine casing 50. A number of chambers 54 are formed between each tubular combustion chamber 36 and its respective cylindrical casing 48.
  • each transition duct 46 has a bracket 56 which extends radially, with respect to the upstream end of the transition duct, and the engine casing 50 has a plurality of pairs of brackets 58.
  • Each bracket 56 is pivotally secured to a respective one of the pairs of brackets 58 by a pin 60, to provide a pivot mounting which is described more fully in our copending UK patent application no. 9019089.3 filed 1 September 1990.
  • the upstream wall 44 of each of the tubular combustion chambers 36 has an aperture 62 to allow the supply of air and fuel into the first fuel and air mixing zone 64.
  • a plurality of first radial flow swirlers are provided and each first radial flow swirler is arranged coaxially with the aperture 62 in the upstream wall 44 of the respective tubular combustion chamber 36.
  • a plurality of second radial flow swirlers are provided and each second radial flow swirler is arranged coaxially with the aperture 62 in the upstream wall 44 of the respective tubular combustion chamber 36.
  • the first radial flow swirlers are positioned axially downstream, with respect to the axis of the tubular combustion chamber, of the second radial flow swirlers.
  • Each first radial flow swirler comprises a first side plate 66, a second side plate 68 and a plurality of first vanes 70.
  • the first side plate 66 has a central aperture arranged coaxially with the aperture 62 in the upstream wall 44, and the plate 66 is secured to the upstream wall 44.
  • the first vanes 70 extend axially between and are secured to the first and second side plates 66 and 68 respectively.
  • a number of passages 72 are formed between the first vanes 70 for the flow of air.
  • Each second radial flow swirler comprises a plurality of second vanes 74 and a third side plate 76.
  • the second vanes 74 extend axially between the second side plate 68 and the third side plate 76.
  • the second side plate 68 has a central aperture arranged coaxially with the aperture 62 in the upstream wall 44, and has a shaped annular lip 78 which extends in an axially downstream direction into the aperture 62.
  • a number of passages 80 are formed between the second vanes 74 for the flow of air.
  • the first and second vanes 70,74 of the first and second radial flow swirlers are arranged to swirl air in opposite directions, as seen from Figures 4 and 5.
  • a first annular air intake 82 is defined axially between the radially outer end of each first side plate 66 and a closure plate 84 at the outer end of each cylindrical casing 48.
  • a plurality of pilot fuel injectors 86 are provided, and each pilot fuel injector 86 is arranged coaxially with the aperture 62 of one of the tubular combustion chambers 36 to supply fuel through the aperture 62 into the first fuel and air mixing zone 64.
  • a plurality of primary fuel injectors 88 are provided for each of the tubular combustion chambers 36.
  • Each of the primary fuel injectors 88 comprises a hollow cylindrical member which extends axially with respect to the tubular combustion chamber 36.
  • Each of the hollow cylindrical members passes axially through the third side plate 76 and the second side plate 68 and locates in a blind hole in the first side plate 66.
  • Each of the hollow cylindrical members is arranged to pass axially through one of the passages 80 between the second vanes 74 and through one of the passages 72 between the first vanes 70.
  • the hollow cylindrical members are positioned towards the radially outer region of the passages 72,80, and have axially spaced apertures 90 to inject fuel into the first radial flow swirler assembly and axially spaced apertures 92 to inject fuel into the second radial flow swirler assembly.
  • the apertures 90 and 92 are arranged to direct the fuel radially inwardly.
  • a second annular fuel and air mixing zone 94 surrounds the first fuel and air mixing zone 64 of each tubular combustion chamber 36.
  • Each second annular fuel and air mixing zone 94 is defined between a second annular wall 96 and a third annular wall 98.
  • the second annular wall 96 defines the radially outer extremity of the second fuel and air mixing zone 94 and the third annular wall 98 defines the radially inner extremity of the second fuel and air mixing zone 94.
  • the axially upstream end 100 of each third annular wall 98 is secured to the first side plate 66 of the first radial flow swirler of the respective tubular combustion chamber 36.
  • a second annular air intake 102 is defined axially between the upstream end of each second annular wall 96 and the upstream end 100 of the respective third annular wall 98 to supply air into the second annular fuel and air mixing zones 94.
  • a plurality of secondary fuel injectors 104 are provided for each of the tubular combustion chambers 36.
  • Each of the secondary fuel injectors 104 comprises a hollow cylindrical member which extends axially with respect to the tubular combustion chamber 36.
  • Each of the hollow cylindrical members passes axially through the upstream end 100 of the third annular wall 98 to supply fuel into the second fuel and air mixing zone 94.
  • Each second and third annular wall 96,98 is arranged coaxially around the first portion 38 of the respective annular wall.
  • the second and third annular walls 96 and 98 are secured to the respective third frusto conical portion 40, and each frusto conical portion 40 is provided with a plurality of circumferentially spaced apertures 106 which are arranged to direct fuel and air into a second combustion zone 112 in the tubular combustion chambers 36, in a downstream direction towards the axis of the tubular combustion chamber 36.
  • the apertures 106 may be circular or slots.
  • Each first side plate 66 is provided with a plurality of apertures 108 to supply cooling air into an annular space 110 between the upstream portion 38 of the annular wall and the third annular wall 98 for cooling of the annular wall.
  • the annular wall may be formed from a laminated structure comprising spaced perforated inner and outer sheets which give transpiration cooling of the annular wall.
  • primary air A flows through the first air intake 82 and through the first and second radial flow swirlers.
  • the lips 78 direct the primary air into the first fuel and air mixing zone or primary combustion zone 64.
  • the flows of air from the first and second radial flow swirlers are in opposite directions and this produces opposed flow vortices B and C.
  • a shear layer D is formed between the vortices B and C which improves mixing turbulence.
  • the pilot injectors 86 only are used at low power settings, that is less than about 40% power. They inject gas or pre-evaporated liquid fuel at a narrow angle only into the primary air which has passed through the second radial flow swirlers to create a locally fuel rich mixture on the axes of the tubular combustion chambers 36. Diffusion causes the fuel to mix with the primary air in the vortex B. Vortex C remains an air only region. Thus a locally fuel rich mixture is created on the combustion chamber 36 centreline which sustains combustion in the primary combustion zone 64.
  • the primary fuel injectors 88 are not used, during low power operation, and thus only primary air exits from the downstream end of the passages 72 and 80 formed between the respective vanes 70 and 74 of the first and second swirler assemblies.
  • the pilot injectors 86 are not used, and all the fuel supplied into the combustion chamber 36 is supplied from the primary and secondary injectors 88 and 104 respectively.
  • the primary fuel injectors 88 inject gas, or pre-evaporated liquid fuel, into the passages 72 and 80 formed between the respective vanes 70 and 74 of the first and second swirler assemblies.
  • the secondary fuel injectors 104 inject gas, or pre-evaporated liquid fuel, into the second fuel and air mixing zone 94 to mix with secondary air entering the second fuel and air mixing zone 94 through the second annular intake 102.
  • the first and second radial flow swirler assemblies direct the fuel and air mixture towards the centreline of the tubular combustion chamber 36 before it is turned so that it flows parallel to the centreline of the combustion chamber 36.
  • the fuel is entrained into both vortex B and vortex C which have opposite swirl, and the shear layer D between the two vortices improves the mixing turbulence.
  • Secondary air E flows through the second air intake 102 into the secondary air and fuel mixing zone 94.
  • the secondary air and fuel is mixed as it flows axially downstream through the second fuel and air mixing zone 94.
  • the resulting fuel and air mixture formed in the secondary air and fuel mixing zone 94 is injected through the apertures 106 into the second downstream portion 42 of the tubular combustion chamber 36 where secondary combustion occurs in the second combustion zone 112.
  • the fuel and air mixture injected from the second fuel and air mixing zone 94 is in the form of discrete jets F which are directed in a downstream direction towards the centreline of the tubular combustion chamber 36. This ensures good penetration of the secondary fuel and air mixture into the gases from the primary combustion zone 64 and hence good mixing. Interaction of the secondary fuel and air mixture jets F with cooling air flowing over the inner surfaces of the downstream portion 42 of the combustion chamber 36 is minimised because of this angling of the jets F towards the centreline of the combustion chamber.
  • the graph in Figure 6 illustrates how the fuel flow to the pilot, primary and secondary injectors 86,88 and 104 respectively varies with the power, or load, setting of the gas turbine engine.
  • pilot injectors 86 are supplied with fuel at power settings below 35% power. At power settings above 35% fuel is supplied simultaneously to the primary and secondary injectors 88 and 104, and the supply of fuel to the pilot injectors 86 is terminated. At a power, or load setting of 35%, 83% of the fuel supplied to each combustion chamber is supplied to the primary injectors 88 and the remaining 17% of the fuel is supplied to the secondary injectors 104. As the power, or load, setting is increased the total quantity of fuel supplied to each combustor increases and the total quantity of fuel supplied to the primary injectors and secondary injectors increases.
  • the percentage of the total fuel supplied to the combustion chamber, which is supplied to the primary injectors 88 decreases gradually from 83% at 35% power setting to approximately 50% at 100% power setting.
  • the percentage of the total fuel supplied to the combustion chamber which is supplied to the secondary injectors 104 increases gradually from 17% at 35% power setting to approximately 50% at 100% power setting.
  • the percentage of fuel supplied to the primary injectors 88 preferably decreases gradually from 78% at 40% power setting to 50% at 100% power setting whilst the percentage of fuel supplied to the secondary injectors 104 increases from 22% at 40% power setting to 50% at 100% power setting.
  • the first fuel and air mixing zone 64 is supplied with fuel so that it has a constant maximum temperature of 1800°K (1527°C) to prevent disassociation of nitrogen at higher temperatures, and hence prevent the formation of NOx.
  • the second combustion zone 112 is supplied with fuel so that it also has a constant maximum temperature of 1800°K (1527°C),and has a minimum temperature of 1500°K (1227°C) to prevent the build up of carbon monoxide etc.
  • the mimimum temperature is 1550°K.
  • the temperature of the flame in the first fuel and air mixing zone 64 remains substantially constant, or within a predetermined range of temperatures, so that the emissions of NOx remains low.
  • the margin between the required flame temperature and the temperature at which the flame is extinguished varies. In some circumstances the flame may be extinguished in the first fuel and air mixing zone. In order to provide an adequate margin between the flame temperature and the temperature at which the flame is extinguished, a greater proportion of fuel could be supplied to the first fuel and air mixing zone 64.
  • this solution is not desirable because the flame temperature is increased and thus the emissions of NOx is increased.
  • An alternative combustion chamber assembly 136 shown in Figure 7, is substantially the same as that shown in Figures 2 to 5 and the same reference numerals have been used to designate like parts.
  • the combustion chamber assembly 136 differs from that shown in Figures 2 to 5 in that the downstream end of the first portion 38 of the annular wall 37 has a frusto conical portion 120 which reduces in diameter to a throat 122.
  • the third frustoconical portion 40 interconnects the first portion 38 and the second portion 42, and the second portion 42 still has a greater diameter than the first portion 38.
  • the reduction in diameter at the downstream end of the first portion 38, in combination with a constant temperature in the first fuel and air mixing zone or combustion zone 64 allows a suitable margin between the flame temperature and the temperature at which the flame is extinguished to be maintained with variations of power setting between 35% and 100% power to prevent the flame in the first fuel and air mixing zone 64 being extinguished.
  • combustion chambers shown in Figures 2 to 5 and in Figure 7 are suitable for operation across the full power range for ambient air temperatures in the range of -30°C to +30°C or higher.
  • a further combustion chamber assembly 236, shown in figure 8, is similar to that shown in figure 7 and the same reference numerals have been used to designate like parts.
  • the combustion chamber assembly 236 differs from that shown in Figure 7 in that each of the tubular combustion chambers 236 also comprises a fourth portion 130 positioned downstream of and interconnected to, the second portion 42 by a fifth portion 132.
  • the fourth portion 130 of the annular wall has a greater diameter than the second portion 40, and the fifth portion 132 is frustoconical.
  • the downstream end of the second portion 42 of the annular wall 37 has a frustoconical portion 134 which reduces in diameter to a throat 136.
  • a third annular fuel and air mixing zone 138 surrounds the second combustion zone 112 of each tubular combustion chamber 236.
  • Each third annular fuel and air mixing zone 138 is defined between a fourth annular wall 140 and a fifth annular wall 142.
  • the fourth annular wall 140 defines the radially outer extremity of the third fuel and air mixing zone 138 and the fifth annular wall 142 defines the radially inner extremity of the third fuel and air mixing zone 138.
  • a third annular air intake 144 is defined between the upstream ends of the fourth and fifth annular walls 140 and 142 respectively to supply air into the third annular fuel and air mixing zones 138.
  • a plurality of tertiary fuel injectors 146 are provided for each of the tubular combustion chambers 236.
  • Each fourth and fifth annular wall 140,142 is arranged coaxially around the second portion 42 of the respective annular wall.
  • the fourth and fifth annular walls 140 and 142 are secured to the respective fifth frustoconical portion 132, and each frustoconical portion 132 is provided with a plurality of circumferentially spaced apertures 148 which are arranged to direct fuel and air into a tertiary combustion zone 150, in the tubular combustion chambers 236, in a downstream direction towards the axis of the tubular combustion chambers 236.
  • the apertures 148 may be circular or slots.
  • primary air A flows through the first air intake 82 and through the first and second radial flow swirlers.
  • the lip 78 directs the primary air into the first fuel and air mixing zone, or primary combustion zone, 64.
  • the flows of air from the first and second radial flow swirlers are in opposite directions to improve mixing turbulence.
  • the pilot injectors 86 only are used at low power settings, that is less than about 40% power. They inject the gas or pre-evaporated liquid fuel at a narrow angle only into the primary air which has passed through the second radial flow swirlers to create a locally fuel rich mixture on the axes of the tubular combustion chambers 236. Diffusion causes the fuel to mix with the primary air in the vortex B. Vortex C remains an air only region. Thus a locally rich mixture is created on the combustion chamber 236 centreline which sustains combustion in the primary combustion zone 64.
  • the primary fuel injectors 88 are not used, during low power operation and thus only primary air exits from the downstream end of the passages 72 and 80 formed between the respective vanes 70 and 74 of the first and second swirler assemblies.
  • the pilot injectors 86 are not used, and all the fuel supplied into the combustion chamber 236 is supplied from the primary and secondary injectors 88 and 104 respectively or from the primary, secondary and tertiary injectors 88,104 and 146 respectively.
  • the primary fuel injectors 88 inject gas, or pre-evaporated liquid fuel, into the passages 72 and 80 formed between the respective vanes 70 and 74 of the first and second swirler assemblies.
  • the secondary fuel injectors 104 inject gas, or pre-evaporated liquid fuel, into the second fuel and air mixing zone 94 to mix with the secondary air entering the second fuel and air mixing zone 94 through the second annular intake 102.
  • the first and second radial flow swirlers direct the fuel and air mixture towards the centreline of the tubular combustion chambers 236 before it is turned so that it flows parallel to the centreline of the combustion chamber 236.
  • the fuel is entrained into both vortex B and vortex C which have opposite swirl, and the shear layer D improves mixing turbulence.
  • Secondary air E flows through the second air intake 102 into the secondary air and fuel mixing zone 94.
  • the secondary air and fuel is mixed as it flows axially downstream through the second fuel and air mixing zone 94.
  • the resulting fuel and air mixture formed in the secondary air and fuel mixing zone 94 is injected through the apertures 106 into the second portion 42 of the tubular combustion chamber 236 where secondary combustion occurs in the second combustion zone 112.
  • the reduction in diameter at the downstream end of the first portion 38, provided by the frustoconical portion 120 and the throat 122 allows a suitable margin between the flame temperature in the primary combustion zone 64 and the temperature at which the flame is extinguished with variations in power setting to prevent the flame in the primary combustion zone 64 being extinguished. This enhances the recirculation of hot combustion products into the primary combustion zone 64 to reignite the fuel and air mixture.
  • the reduction in diameter at the downstream end of the second portion 40, provided by the frustoconical portion 134 and the throat 136 allows a suitable margin between the flame temperature in the secondary combustion zone 112 and the temperature at which the flame is extinguished with variations in power setting to prevent the flame in the secondary combustion zone 112 being extinguished. This enhances the recirculation of hot combustion products into the secondary combustion zone 112 to reignite the fuel and air mixture, by producing recirculation zones J.
  • the primary and secondary fuel injectors 88 and 104 respectively supply fuel into the primary and secondary combustion zones 64 and 112 respectively for power settings between 40% and 100% power.
  • the tertiary fuel injectors 146 do not supply fuel into the tertiary combustion zone 150 at low ambient air temperatures at any power setting.
  • the amount of fuel supplied to the primary injectors 88 is increased to maintain the temperature in the primary combustion zone 64 at 1800°K. This is important to ensure optimum combustion for NOx reduction, and to maintain a high enough temperature in the secondary combustion zone 112 for combustion to continue in the secondary combustion zone 112.
  • the primary and secondary fuel injectors 88 and 104 respectively supply fuel into the primary and secondary combustion zones 64 and 112 respectively for lower power settings between 40% and a predetermined power setting.
  • the primary, secondary and tertiary fuel injectors 88,104 and 146 respectively supply fuel into the primary, secondary and tertiary combustion zones 64,112 and 150 respectively.
  • the minimum power setting at which the primary, secondary and tertiary fuel injectors 88,104 and 146 respectively supply fuel into the primary, secondary and tertiary combustion zones 64,112 and 150 respectively increases from the predetermined power setting at high ambient air temperature operation.
  • the tertiary fuel injectors 146 are not supplied with fuel at any power setting.
  • the temperature in the first fuel and air mixing zone 64 is maintained at about 1800°K, and the temperature in the second combustion zone 112 is maintained at about 1740°K and the temperature in the tertiary combustion zone 150 is varied between 1550°K and 1800°K.
  • the tertiary fuel injectors 146 do not supply fuel to the tertiary combustion zone 150 and the amount of fuel supplied by the secondary fuel injectors 104 into the secondary combustion zone 112 is increased to increase its temperature to 1850°K.
  • the system then acts as a two staged combustor.
  • the combination of the secondary fuel and air mixing zone 94 and secondary combustion zone 112 together with the tertiary fuel and air mixing zone 138 and tertiary combustion zone 150 allows reduced emissions of NOx to be achieved at all power settings between 40% and 100% power over a wide range of pressure ratios and velocity profiles without the need for variable geometry air intakes for the combustion chambers 236.
  • the industrial gas turbine engine will be provided with a control system which controls the fuel supplied to the pilot, primary and secondary injectors in accordance with the power demanded for the combustion chambers shown in Figures 2 to 5 and 7.
  • the industrial gas turbine engine will be provided with a control system which controls the fuel supplied to the pilot, primary, secondary and tertiary injectors in accordance with the power demanded and the ambient air temperature for the combustion chamber shown in Figure 8.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)

Claims (36)

  1. Brennkammer (36) für ein Gasturbinentriebwerk mit einem ersten Lufteinlaß (82), mit einer Primärbrennstoff-Einspritzeinrichtung (88) und einer ersten Brennstoff/Luft-Mischzone (64), wobei die erste Brennstoff/Luft-Mischzone (64) durch wenigstens eine Ringwand (37) und eine stromaufwärtige Wand (44) definiert ist, die mit dem stromaufwärtigen Ende der Ringwand (37) verbunden ist, wobei die stromaufwärtige Wand (44) wenigstens eine Öffnung (62) aufweist, und der erste Lufteinlaß wenigstens eine erste radiale Strömungsverwirbelungseinrichtung (66, 68, 70) und wenigstens eine zweite radiale Strömungsverwirbelungseinrichtung (68, 74, 76) aufweist, und jede erste radiale Strömungsverwirbelungseinrichtung (66, 68, 70) Luft in die erste Brennstoff/Luft-Mischzone (64) über die Öffnung (62) zuführt und jede zweite radiale Strömungsverwirbelungseinrichtung (68, 74, 76) Luft in die erste Brennstoff/Luft-Mischzone (64) über die Öffnung (62) einleitet, und wobei die erste radiale Strömungsverwirbelungseinrichtung (66, 68, 70) axial stromab der zweiten radialen Strömungsverwirbelungseinrichtung (68, 74, 76) in bezug auf die Achse der Brennkammer (36) angeordnet ist und jede erste radiale Strömungsverwirbelungseinrichtung (66, 68, 70) die Luft im Gegensinn zu der von der zweiten radialen Strömungsverwirbelungseinrichtung (68, 74, 76) verwirbelten Luft richtet, und wobei die Primärbrennstoff-Einspritzeinrichtung (88) Brennstoff in wenigstens einen der Kanäle (80) einleitet, der zwischen den Leitschaufeln (74) jeder zweiten radialen Strömungsverwirbelungseinrichtung (68, 74, 76) definiert ist,
    dadurch gekennzeichnet, daß die Primärbrennstoff-Einspritzeinrichtung (88) Brennstoff in wenigstens einen der Kanäle (72) eintreten läßt, die zwischen den Leitschaufeln (70) jeder ersten radialen Strömungsverwirbelungseinrichtung (66, 68, 70) definiert sind.
  2. Brennkammer nach Anspruch 1, welche wenigstens einen Pilotbrennstoff-Injektor (86) aufweist und jeder Pilotbrennstoff-Injektor (86) auf jeweils eine der Öffnungen (62) ausgerichtet ist, um Brennstoff durch die Öffnung (62) in die erste Brennstoff/Luft-Mischzone (64) einzuspritzen.
  3. Brennkammer nach Anspruch 1, bei welcher die Primärbrennstoff-Einspritzeinrichtung (88) Brennstoff in sämtliche Kanäle (72) einleitet, die zwischen den Leitschaufeln (70) der ersten radialen Strömungsverwirbelungseinrichtung (66, 68, 70) ausgebildet sind.
  4. Brennkammer nach Anspruch 1, bei welcher die Primärbrennstoff-Einspritzeinrichtung (88) Brennstoff in sämtliche Kanäle (80) einführt, die zwischen den Leitschaufeln (74) der zweiten radialen Strömungsverwirbelungseinrichtung (68, 74, 76) definiert sind.
  5. Brennkammer nach Anspruch 1, bei welcher die Primärbrennstoff-Einspritzeinrichtung (88) Brennstoff in den radial äußeren Bereich der Kanäle (72, 80) zwischen den Leitschaufeln (70, 74) einspritzt.
  6. Brennkammer nach Anspruch 1, bei welcher die Primärbrennstoff-Einspritzeinrichtung (88) ein hohlzylindrisches Glied aufweist, welches sich axial bezüglich der Brennkammer (36) erstreckt, wobei das zylindrische Glied mehrere Öffnungen (90) aufweist, die im axialen Abstand längs des zylindrischen Gliedes angeordnet sind, um Brennstoff in die Kanäle (72, 80) einzuspritzen.
  7. Brennkammer nach Anspruch 6, bei welcher die Öffnungen (90) den Brennstoff radial nach innen richten.
  8. Brennkammer nach Anspruch 1, bei welcher die Primärbrennstoff-Einspritzeinrichtung (88) gasförmigen Brennstoff oder verdampften flüssigen Brennstoff einbläst.
  9. Brennkammer nach Anspruch 2, bei welcher der Pilotbrennstoff-Injektor (86) gasförmigen oder flüssigen Brennstoff einspritzt.
  10. Brennkammer nach Anspruch 1, bei welcher die Brennkammer (36) rohrförmig ausgebildet ist und eine einzige Öffnung (62) in ihrer stromaufwärtigen Wand (44) aufweist.
  11. Brennkammer nach Anspruch 2, welche außerdem einen Sekundär-Lufteinlaß (102), eine Sekundärbrennstoff-Einspritzeinrichtung (104) und eine sekundäre Brennstoff/Luft-Mischzone (94) aufweist, wobei die sekundäre Brennstoff/Luft-Mischzone (94) ringförmig ausgebildet ist und die erste Brennstoff/LuftMischzone (64) umschließt, und die sekundäre Brennstoff/LuftMischzone (94) an ihrem radial äußeren Ende durch eine zweite Ringwand (96) definiert ist, und wobei die sekundäre BrennstoffEinspritzeinrichtung (104) Brennstoff in das stromaufwärtige Ende der sekundären Brennstoff/Luft-Mischzone (94) einleitet, und wobei die sekundäre Brennstoff/Luft-Mischzone (94) an ihrem stromabwärtigen Ende mit einer Sekundärverbrennungszone (112) im Inneren der Brennkammer (36) stromab der ersten Brennstoff/ Luft-Mischzone (64) in Strömungsverbindung steht.
  12. Brennkammer nach Anspruch 11, bei welcher die Ringwand (37) einen ersten, die erste Brennstoff/Luft-Mischzone (64) definierenden Abschnitt (38), einen zweiten, die Sekundärverbrennungszone (112) definierenden Abschnitt (42) mit einem gegenüber dem ersten Abschnitt (38) vergrößerten Durchmesser und einen dritten kegelstumpfförmigen Abschnitt (40) aufweist, der den ersten Abschnitt (38) mit dem zweiten Abschnitt (42) verbindet.
  13. Brennkammer nach Anspruch 12, bei welcher sich das stromabwärtige Ende (120) des ersten Abschnitts (38) der Ringwand (37) im Durchmesser zu einer Einschnürung (122) vermindert.
  14. Brennkammer nach Anspruch 11, bei welcher der Sekundär-Lufteinlaß (102) stromab der ersten Lufteinlaßeinrichtung (82) angeordnet ist.
  15. Brennkammer nach Anspruch 11, bei welcher die Sekundär-Brennstoff/Luft-Mischzone (94) an ihrem radial inneren Ende durch eine dritte Ringwand (98) definiert ist.
  16. Brennkammer nach Anspruch 12, bei welcher der dritte konische Abschnitt (40) mehrere in Umfangsrichtung im gleichen Abstand angeordnete Öffnungen (106) aufweist, die das SekundärBrennstoff/Luft-Gemisch aus der Sekundär-Brennstoff/Luft-Mischzone (94) in Gestalt mehrerer Strahlen austreten lassen, die stromab in Richtung auf die Mittellinie der Brennkammer (36) verlaufen.
  17. Brennkammer nach Anspruch 16, bei welcher die Öffnungen (106) Schlitze sind.
  18. Brennkammer nach Anspruch 12, bei welcher das stromabwärtige Ende der zweiten Ringwand (96) an dem dritten konischen Abschnitt (40) der Ringwand (37) festgelegt ist.
  19. Brennkammer nach Anspruch 15, bei welcher Kühlluft einer Ringkammer (110) zugeführt wird, die zwischen der Ringwand (37) und der dritten Ringwand (98) ausgebildet ist.
  20. Brennkammer nach Anspruch 11, bei welcher die Sekundärbrennstoff-Einspritzeinrichtung (104) mehrere in Umfangsrichtung im gleichen Winkelabstand zueinander angeordnete Injektoren aufweist.
  21. Brennkammer nach Anspruch 11, bei welcher die Sekundärbrennstoff-Einspritzeinrichtung (104) gasförmigen Brennstoff oder verdampften flüssigen Brennstoff zuführt.
  22. Brennkammer nach Anspruch 11, welche weiter einen Tertiär-Lufteinlaß (144), eine tertiäre Brennstoff-Einspritzeinrichtung (146) und eine tertiäre Brennstoff/Luft-Mischzone (138) aufweist, wobei die tertiäre Brennstoff/Luft-Mischzone (138) ringförmig ausgebildet ist und die Sekundärverbrennungszone (112) umschließt, und wobei die tertiäre Brennstoff/LuftMischzone (138) an ihrem radial äußeren Ende durch eine vierte Ringwand (140) begrenzt ist und die tertiäre Brennstoff-Einspritzeinrichtung (146) den Brennstoff in das stromaufwärtige Ende der tertiären Brennstoff/Luft-Mischzone (138) einspritzt, und wobei die tertiäre Brennstoff/Luft-Mischzone (138) an ihrem stromabwärtigen Ende mit einer tertiären Verbrennungszone (150) im Inneren der Brennkammer (136) stromab der Sekundärverbrennungszone (112) in Strömungsverbindung steht.
  23. Brennkammer nach Anspruch 22, bei welcher die Ringwand (37) stromab des zweiten Abschnitts (42) einen vierten Abschnitt (130) mit einem Durchmesser aufweist, der größer ist als der Durchmesser des zweiten Abschnitts, wodurch die tertiäre Verbrennungszone (150) definiert wird, und wobei ein fünfter hohlkegelstumpfförmiger Abschnitt (132) den zweiten Abschnitt (42) mit dem vierten Abschnitt (130) verbindet.
  24. Brennkammer nach Anspruch 22, bei welcher das stromabwärtige Ende des zweiten Abschnitts (42) der Ringwand sich im Durchmesser zu einer Einschnürung (136) verjüngt.
  25. Brennkammer nach Anspruch 22, bei welcher der tertiäre Lufteinlaß (144) stromab des zweiten Lufteinlasses (102) liegt.
  26. Brennkammer nach Anspruch 22, bei welcher die tertiäre Brennstoff/Luft-Mischzone (138) an ihrem radial inneren Ende durch eine fünfte Ringwand (142) definiert ist.
  27. Brennkammer nach Anspruch 23, bei welcher der fünfte konische Abschnitt (132) mehrere in Umfangsrichtung im gleichen Winkelabstand zueinander angeordnete Öffnung (148) aufweist, die das tertiäre Brennstoff/Luft-Gemisch aus der tertiären Brennstoff/Luft-Mischzone (138) in Gestalt mehrerer Strahlen stromab nach der Mittellinie der Brennkammer (236) richten.
  28. Brennkammer nach Anspruch 27, bei welcher die Öffnungen (148) als Schlitze ausgebildet sind.
  29. Brennkammer nach Anspruch 23, bei welcher das stromabwärtige Ende der vierten Ringwand (140) an dem fünften konischen Abschnitt (132) der Ringwand (37) festgelegt ist.
  30. Brennkammer nach Anspruch 22, bei welcher die tertiäre Brennstoff-Einspritzeinrichtung (146) mehrere in Umfangsrichtung im gleichen Winkelabstand zueinander angeordnete Injektoren aufweist.
  31. Brennkammer nach Anspruch 22, bei welcher die tertiäre Brennstoff-Einspritzeinrichtung (146) gasförmigen Brennstoff oder verdampften flüssigen Brennstoff einspritzt.
  32. Verfahren zum Betrieb einer Brennkammer eines Gasturbinentriebwerks gemäß Anspruch 11 mit den folgenden Schritten: vom Beginn des Anlaufs des Gasturbinentriebwerks bis zum Erreichen eines vorbestimmten Ausgangsleistungspegels wird Brennstoff von dem Pilot-Brennstoffinjektor (86) nur in die erste Brennstoff/Luft-Mischzone (64) eingespritzt; für eine Ausgangsleistung, die über dem vorbestimmten Pegel liegt, wird Brennstoff von der Primärbrennstoff-Einspritzeinrichtung (88) in wenigstens einen der Kanäle (72), die zwischen den Leitschaufeln (40) der ersten radialen Strömungsverwirbelungseinrichtung (66, 68, 70) gebildet sind, und in wenigstens einen der Kanäle (80) eingespritzt, die zwischen den Leitschaufeln (74) der zweiten radialen Strömungsverwirbelungseinrichtung (68, 74, 76 gebildet sind, damit dieser Brennstoff in die erste Brennstoff/Luft-Mischzone (64) gelangt, wobei gleichzeitig Brennstoff von der Sekundärbrennstoff-Einspritzeinrichtung (104) in die Sekundär-Brennstoff/Luft-Mischzone (94) eingespritzt wird, damit dieser in die Sekundärverbrennungszone (112) im Inneren der Brennkammer (36) stromab der ersten Brennstoff/ Luft-Mischzone (64) gelangt.
  33. Verfahren nach Anspruch 32, bei welchem der vorbestimmte Ausgangsleistungspegel 35 bis 40 % beträgt.
  34. Verfahren nach Anspruch 33, bei welchem der Anteil des Brennstoffs, der von der Primärbrennstoff-Einspritzeinrichtung zugeführt wird, zwischen 75 und 50 % der Gesamtbrennstoffmenge liegt, die der Brennkammer bei einer Ausgangsleistung zwischen 40 % bis 100 % zugeführt wird.
  35. Verfahren zum Betrieb einer Brennkammer eines Gasturbinentriebwerks gemäß Anspruch 22 mit den folgenden Schritten: von Beginn des Betriebs des Gasturbinentriebwerks bis zum Erreichen eines vorbestimmten Ausgangsleistungspegels wird Brennstoff vom Pilotbrennstoff-Injektor (86) nur in die erste Brennstoff/Luft-Mischzone (64) eingespritzt; für Ausgangsleistungspegel über dem vorbestimmten Leistungspegel wird Brennstoff von der Primärbrennstoff-Einspritzeinrichtung (88) in wenigstens einen der Kanäle (72) zwischen den Leitschaufeln (70) der ersten radialen Strömungsverwirbelungseinrichtung (66, 68, 70) und in wenigstens einen der Kanäle (80) eingespritzt, die zwischen den Leitschaufeln (74) der zweiten radialen Strömungsverwirbelungseinrichtung (68, 74, 76) gebildet sind, damit dieser Brennstoff in die erste Brennstoff/Luft-Mischzone (64) gelangen kann, und gleichzeitig wird Brennstoff der Sekundär-Brennstoff/Luft-Mischzone (94) zugeführt, damit dieser in die Sekundärverbrennungszone (112) im Inneren der Brennkammer (236) stromab der ersten Brennstoff/Luft-Mischzone (64) strömen kann; für Ausgangsleistungspegel, die größer sind als ein zweiter vorbestimmter Ausgangspegel und bei Umgebungslufttemperaturen, die über einer vorbestimmten Temperatur liegen, wird der Brennstoff in die tertäre Brennstoff/Luft-Mischzone (138) eingespritzt, damit er in die tertiäre Verbrennungszone (150) im Inneren der Brennkammer (236) stromab der Sekundärverbrennungszone (112) strömen kann.
  36. Verfahren nach Anspruch 35, bei welchem der vorbestimmte Ausgangsleistungspegel 35 % bis 40 % Leistung beträgt.
EP91918884A 1990-10-23 1991-09-26 Gasturbinenbrennkammer und deren arbeitsweise Expired - Lifetime EP0554325B1 (de)

Applications Claiming Priority (5)

Application Number Priority Date Filing Date Title
GB9023004 1990-10-23
GB909023004A GB9023004D0 (en) 1990-10-23 1990-10-23 A gas turbine engine combustion chamber and a method of operating a gas turbine engine combustion chamber
GB9118945 1991-09-05
GB919118945A GB9118945D0 (en) 1991-09-05 1991-09-05 A gas turbine engine combustion chamber and a method of operating a gas turbine engine combustion chamber
PCT/GB1991/001658 WO1992007221A1 (en) 1990-10-23 1991-09-26 Gasturbine combustion chamber and method of operation thereof

Publications (2)

Publication Number Publication Date
EP0554325A1 EP0554325A1 (de) 1993-08-11
EP0554325B1 true EP0554325B1 (de) 1995-07-26

Family

ID=26297840

Family Applications (1)

Application Number Title Priority Date Filing Date
EP91918884A Expired - Lifetime EP0554325B1 (de) 1990-10-23 1991-09-26 Gasturbinenbrennkammer und deren arbeitsweise

Country Status (4)

Country Link
EP (1) EP0554325B1 (de)
JP (1) JP3077939B2 (de)
DE (1) DE69111614T2 (de)
WO (1) WO1992007221A1 (de)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7003960B2 (en) 2000-10-05 2006-02-28 Alstom Technology Ltd Method and appliance for supplying fuel to a premixing burner
US9879635B2 (en) 2007-11-12 2018-01-30 GETAS GESELLSCHAFT FüR THERMODYNAMISCHE ANTRIEBSSYSTEME MBH Axial piston engine and method for operating an axial piston engine

Families Citing this family (37)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB9122965D0 (en) * 1991-10-29 1991-12-18 Rolls Royce Plc Turbine engine control system
GB2278431A (en) * 1993-05-24 1994-11-30 Rolls Royce Plc A gas turbine engine combustion chamber
US5394688A (en) * 1993-10-27 1995-03-07 Westinghouse Electric Corporation Gas turbine combustor swirl vane arrangement
GB2284884B (en) * 1993-12-16 1997-12-10 Rolls Royce Plc A gas turbine engine combustion chamber
GB9410233D0 (en) 1994-05-21 1994-07-06 Rolls Royce Plc A gas turbine engine combustion chamber
ES2101663T3 (es) * 1994-07-13 2001-12-16 Volvo Aero Corp Camara de combustion de bajas emisiones para motores de turbina de gas.
US5813232A (en) * 1995-06-05 1998-09-29 Allison Engine Company, Inc. Dry low emission combustor for gas turbine engines
DE19532264C2 (de) * 1995-09-01 2001-09-06 Mtu Aero Engines Gmbh Einrichtung zur Aufbereitung eines Gemisches aus Brennstoff und Luft an Brennkammern für Gasturbinentriebwerke
GB9520002D0 (en) * 1995-09-30 1995-12-06 Rolls Royce Plc Turbine engine control system
GB9611235D0 (en) * 1996-05-30 1996-07-31 Rolls Royce Plc A gas turbine engine combustion chamber and a method of operation thereof
GB9809371D0 (en) 1998-05-02 1998-07-01 Rolls Royce Plc A combustion chamber and a method of operation thereof
GB9813972D0 (en) 1998-06-30 1998-08-26 Rolls Royce Plc A combustion chamber
GB9818160D0 (en) 1998-08-21 1998-10-14 Rolls Royce Plc A combustion chamber
GB9911871D0 (en) 1999-05-22 1999-07-21 Rolls Royce Plc A gas turbine engine and a method of controlling a gas turbine engine
GB9911867D0 (en) 1999-05-22 1999-07-21 Rolls Royce Plc A combustion chamber assembly and a method of operating a combustion chamber assembly
GB9915770D0 (en) 1999-07-07 1999-09-08 Rolls Royce Plc A combustion chamber
GB0019533D0 (en) 2000-08-10 2000-09-27 Rolls Royce Plc A combustion chamber
GB0111788D0 (en) 2001-05-15 2001-07-04 Rolls Royce Plc A combustion chamber
GB0515034D0 (en) 2005-07-21 2005-08-31 Rolls Royce Plc Method and system for operating a multi-stage combustor
US7631500B2 (en) * 2006-09-29 2009-12-15 General Electric Company Methods and apparatus to facilitate decreasing combustor acoustics
US7886539B2 (en) * 2007-09-14 2011-02-15 Siemens Energy, Inc. Multi-stage axial combustion system
JP5172468B2 (ja) * 2008-05-23 2013-03-27 川崎重工業株式会社 燃焼装置および燃焼装置の制御方法
US8215116B2 (en) * 2008-10-02 2012-07-10 General Electric Company System and method for air-fuel mixing in gas turbines
AU2009352303B2 (en) * 2009-09-13 2015-07-30 Lean Flame Inc. Method of fuel staging in combustion apparatus
WO2011109514A1 (en) 2010-03-02 2011-09-09 Icr Turbine Engine Corporatin Dispatchable power from a renewable energy facility
US8984895B2 (en) 2010-07-09 2015-03-24 Icr Turbine Engine Corporation Metallic ceramic spool for a gas turbine engine
US9051873B2 (en) 2011-05-20 2015-06-09 Icr Turbine Engine Corporation Ceramic-to-metal turbine shaft attachment
US10094288B2 (en) 2012-07-24 2018-10-09 Icr Turbine Engine Corporation Ceramic-to-metal turbine volute attachment for a gas turbine engine
JP5460846B2 (ja) * 2012-12-26 2014-04-02 川崎重工業株式会社 燃焼装置および燃焼装置の制御方法
US9534790B2 (en) * 2013-01-07 2017-01-03 General Electric Company Fuel injector for supplying fuel to a combustor
JP6318443B2 (ja) * 2013-01-22 2018-05-09 三菱日立パワーシステムズ株式会社 燃焼器、及び回転機械
GB201312974D0 (en) * 2013-07-19 2013-09-04 Siemens Ag Turbine engine control system
US20150159877A1 (en) * 2013-12-06 2015-06-11 General Electric Company Late lean injection manifold mixing system
EP3236157A1 (de) 2016-04-22 2017-10-25 Siemens Aktiengesellschaft Wirbler zum mischen von kraftstoff mit luft in einem verbrennungsmotor
JP6906381B2 (ja) * 2017-07-03 2021-07-21 株式会社東芝 燃焼装置およびガスタービン
RU193249U1 (ru) * 2019-03-04 2019-10-21 Публичное Акционерное Общество "Одк-Сатурн" Газосборник камеры сгорания газотурбинного двигателя
CN114526497B (zh) * 2022-01-07 2023-02-07 清华大学 双缩口组合旋流式中心分级高温升燃烧室

Family Cites Families (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE1039785B (de) * 1957-10-12 1958-09-25 Maschf Augsburg Nuernberg Ag Brennkammer mit hoher Waermebelastung, insbesondere fuer Verbrennung heizwertarmer, gasfoermiger Brennstoffe in Gasturbinenanlagen
US3605405A (en) * 1970-04-09 1971-09-20 Gen Electric Carbon elimination and cooling improvement to scroll type combustors
US3722216A (en) * 1971-01-04 1973-03-27 Gen Electric Annular slot combustor
US3811278A (en) * 1973-02-01 1974-05-21 Gen Electric Fuel injection apparatus
DE3819898A1 (de) * 1988-06-11 1989-12-14 Daimler Benz Ag Brennkammer fuer eine thermische stroemungsmaschine
US4910957A (en) * 1988-07-13 1990-03-27 Prutech Ii Staged lean premix low nox hot wall gas turbine combustor with improved turndown capability

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7003960B2 (en) 2000-10-05 2006-02-28 Alstom Technology Ltd Method and appliance for supplying fuel to a premixing burner
US9879635B2 (en) 2007-11-12 2018-01-30 GETAS GESELLSCHAFT FüR THERMODYNAMISCHE ANTRIEBSSYSTEME MBH Axial piston engine and method for operating an axial piston engine

Also Published As

Publication number Publication date
JPH06502240A (ja) 1994-03-10
JP3077939B2 (ja) 2000-08-21
WO1992007221A1 (en) 1992-04-30
EP0554325A1 (de) 1993-08-11
DE69111614D1 (de) 1995-08-31
DE69111614T2 (de) 1995-12-21

Similar Documents

Publication Publication Date Title
EP0554325B1 (de) Gasturbinenbrennkammer und deren arbeitsweise
US5319935A (en) Staged gas turbine combustion chamber with counter swirling arrays of radial vanes having interjacent fuel injection
US6240732B1 (en) Fluid manifold
EP1431543B1 (de) Injektor
US5628192A (en) Gas turbine engine combustion chamber
EP0982545B1 (de) Brennkammer und Verfahren zum Betrieb
JP5364275B2 (ja) 燃焼システムにおけるＮＯｘエミッションを低減するのを可能にするための方法及びシステム
US6993916B2 (en) Burner tube and method for mixing air and gas in a gas turbine engine
US5865024A (en) Dual fuel mixer for gas turbine combustor
EP0700499B1 (de) Verbrennungskammer eines gasturbinenmotors
US5511375A (en) Dual fuel mixer for gas turbine combustor
US6826913B2 (en) Airflow modulation technique for low emissions combustors
JP5642357B2 (ja) タービンエンジン用予混合装置
US20090056336A1 (en) Gas turbine premixer with radially staged flow passages and method for mixing air and gas in a gas turbine
US6253538B1 (en) Variable premix-lean burn combustor
US8015814B2 (en) Turbine engine having folded annular jet combustor
EP2041494B1 (de) Vormischungsinjektoren für einen gasturbinenmotor
EP1106928A1 (de) Brennstoffsystemanordnung und Verfahren zur gestuften Brennstoffzufuhr in zugleich mit flüssig- und gasförmigen Brennstoffen getriebenen Gasturbinen
EP0710347A1 (de) Zweibrennstoff-einspritzdüse
EP1297283A1 (de) Brennstoffeinspritzdüse für emissionsarme gasturbinenvormischbrennkammer
US4179881A (en) Premix combustor assembly
EP1243854B1 (de) Brennstoffeinspritzventil

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

17P Request for examination filed

Effective date: 19930312

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): DE FR GB

17Q First examination report despatched

Effective date: 19941215

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): DE FR GB

ET Fr: translation filed
REF Corresponds to:

Ref document number: 69111614

Country of ref document: DE

Date of ref document: 19950831

PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

26N No opposition filed
REG Reference to a national code

Ref country code: GB

Ref legal event code: IF02

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: FR

Payment date: 20050815

Year of fee payment: 15

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: GB

Payment date: 20050817

Year of fee payment: 15

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: DE

Payment date: 20050818

Year of fee payment: 15

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: DE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20070403

GBPC Gb: european patent ceased through non-payment of renewal fee

Effective date: 20060926

REG Reference to a national code

Ref country code: FR

Ref legal event code: ST

Effective date: 20070531

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: GB

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20060926

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: FR

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20061002