Abstract: A burner including a gas injector for injecting fuel into the burner for a gas turbine engine is provided. The burner is provided with a quarl, which surrounds a combustion room, at least a first channel emerging into the combustion room for providing the combustion room with air mixed with fuel, a swirler for mixing the air and the fuel located at the inlet of the channel, at least one tube for the provision of fuel at an inlet of at least one channel, wherein the tube is provided with a plurality of diffuser holes distributed along the tube acting as gas injectors for effectively distributing fuel in a flow of air passing through the channel.

Abstract: A gas turbine includes a compressor and at least one combustor downstream from the compressor. The combustor includes a burner having an inner shroud extending axially along at least a portion of the burner, an outer shroud radially separated from the inner shroud and extending axially along at least a portion of the burner, and a plurality of stator vanes extending radially between the inner shroud and the outer shroud. The stator vanes have an inner end proximate the inner shroud and an outer end proximate the outer shroud. The burner further includes a vortex tip at one of either the inner end or the outer end of the stator vanes. The vortex tip provides a gap between the inner end and the inner shroud or the outer end and the outer shroud, and the vortex tip includes a plurality of fuel ports. The gas turbine further includes a turbine downstream from the combustor.

Abstract: A fuel nozzle 110 having a plurality of swirl vane 130 on an outer peripheral surface thereof is installed within a burner tube 120, with a clearance 121 being provided. Each swirl vane 130 progressively curves from an upstream side toward a downstream side (inclines along a circumferential direction) in order to swirl compressed air A flowing through an air passage 111 to form a swirl air flow a. Here, curvature of each swirl vane 130 is greater on its outer peripheral side than on its inner peripheral side. By suppressing occurrence of an air streamline heading from the inner peripheral side toward the outer peripheral side, therefore, flow velocity on the inner peripheral side and flow velocity on the outer peripheral side become equal, thus preventing flashback on the inner peripheral side.

Abstract: A mixer assembly for a gas turbine engine, including: a primary fuel injector having a central axis for injecting a primary fuel spray into a primary air stream, wherein fuel is provided at a desired rate and droplets of the fuel spray are within a desired size range; a secondary fuel injector positioned radially outwardly of the primary fuel injector for injecting a secondary fuel spray into a secondary air stream spaced radially outwardly of and surrounding the primary air stream; and, a primary air jet positioned between the primary fuel injector and the secondary fuel injector to direct a portion of an incoming air stream between the primary air stream and the secondary air stream.

Abstract: A fuel spraying apparatus of a gas turbine engine includes a first fuel spraying section configured to spray fuel for diffusion combustion and a second fuel spraying section configured to spray fuel for premixed combustion. The second fuel spraying section is disposed so as to surround the first fuel spraying section. The apparatus further includes an air curtain generating section disposed between the first fuel spraying section and the second fuel spraying section. The air curtain generating section is configured to generate an air curtain which defines an outer edge of a diffusion combustion area in a combustion chamber.

Abstract: A gas turbine engine pilot assembly includes a swirler having high and low pressure sides. A hood at least partially encloses the swirler on the high pressure side. The hood is secured over the swirler, in one example. The hood includes an aperture creating a tortuous path from the high pressure side to the low pressure side through the swirler. The hood reduces the differential pressure across the swirler by reducing the velocity and pressure of the air before entering the swirler. In one example, the hood includes first and second spaced apart walls interconnected by a perimeter wall. The walls form a generally annular structure, in one example. At least one of the walls includes an array of apertures communicating with a cavity interiorly arranged within the walls upstream from the swirler. Air from the high pressure side flows through the apertures and is slowed before passing through the swirler and into a combustion chamber.

Abstract: A system includes a turbine combustor fuel nozzle. The turbine combustor fuel nozzle includes a swirl vane. The turbine combustor fuel nozzle also includes an injection hole configured to inject fluid in a downstream region of the swirl vane. The injection of fluid in a downstream region of the swirl vane may be in response to detection of a condition indicative of a flame inside the turbine combustor fuel nozzle.

Abstract: A fuel nozzle guide plate has mirror symmetry about a first plane perpendicular to the fuel nozzle guide plate and mirror symmetry about a second plane parallel to the fuel nozzle guide plate. The fuel nozzle guide plate is attached to an air swirler, which is mounted in an opening of a bulkhead for mixing air and fuel at an upstream side of a combustor.

Abstract: A fuel nozzle for a gas turbine engine includes a nozzle body having a longitudinal axis. An elongated annular air passage is defined within the nozzle body. A plurality of circumferentially spaced apart axially extending swirl vanes is disposed within the annular air passage. Each swirl vane has an axially swept cross-sectional profile along the radial extent thereof.

Abstract: A fuel nozzle assembly for a gas turbine engine includes a flange and a pre-mix tube. The flange includes a first end that is configured to couple to an end cover of the combustor, and a second end that is opposite said first end. The pre-mix tube is coupled at a first end to the flange second end. The flange and the pre-mix tube are fabricated to operate at a natural frequency that is different from an operating frequency of the gas turbine engine.

Abstract: This invention provides a gas turbine combustor, which can enable only a swirler and a heat shield to be readily taken out from the combustor upon the exchange the swirler and the heat shield. The gas turbine combustor (1) includes: a combustion cylinder (8) constituting a combustion chamber (11); a fuel injection unit (12) adapted for supplying fuel (F) to a head portion of the combustion cylinder (8); a support member (27) configured for allowing the fuel injection unit (12) to be supported by the combustion cylinder (8); and a heat shield (28) adapted for heat-insulating the support member (27) from combustion gas in the combustion chamber (11), wherein the fuel injection unit (12) includes a fuel injection valve (13) adapted for injecting the fuel (F), and a swirler (14) adapted for supplying compressed air (CA) to the fuel (F) injected from the fuel injection valve (13) while swirling the compressed air (CA).

Abstract: An injection device for injecting a mixture of air and fuel into a combustion chamber of a turbomachine is disclosed. The injection device includes two fuel circuits and a Venturi which improves the carburation.

Abstract: A fuel nozzle assembly includes a swirler assembly having an inlet end, an outlet end, a shroud inner surface and a hub outer surface. The inner surface has a first diameter adjacent to the inlet end and a second diameter adjacent to the outlet end defining a differential diameter ratio. A plurality of vanes are coupled to the swirler assembly and extend between the shroud inner surface and the hub outer surface. The vanes have a pair of opposing sidewalls joined at a leading edge and at an axially-spaced trailing edge, and a first height adjacent to the leading edge and a second height adjacent to the trailing edge. The first height and the second height define a differential height ratio, wherein the differential diameter ratio, the differential height ratio, or both are configured to provide convergent flow through the fuel nozzle.

Abstract: A fuel nozzle for a turbine engine has a central body member with a pilot, a surrounding barrel housing, a mixing duct and an air inlet duct. The fuel nozzle additionally has a main fuel injection device located between the air inlet duct and the mixing duct. The main fuel injection device is configured to introduce a flow of fuel into the barrel member to create a fuel/air mixture which is then premixed with a swirler. The fuel/air mixture then further mixes in the mixing duct and exits the nozzle into a combustor for combustion. The geometry of the fuel nozzle ensures that pressure waves from the combustor do not create a time varying fuel to air equivalence ratio in the flow through the nozzle that achieves a resonance with the pressure waves.

Abstract: A fuel injector assembly includes a fuel injector that provides fuel to be mixed with air for ignition in a combustor. The fuel injector assembly includes a fuel shroud and a plurality of spanning members. The spanning members are spaced circumferentially on the fuel shroud between the fuel shroud and a base plate assembly, which supports the fuel injector assembly within the combustor. The spanning members space the fuel shroud from the base plate assembly such that an annular gap is formed therebetween. The gap permits a flow of purge air therethrough to be mixed with compressed air entering the fuel shroud and with fuel provided by the fuel injector. The spanning members comprise elongate members having upstream ends that are circumferentially offset from opposed downstream ends so that the spanning members provide tangential velocity to the purge air flowing through the annular gap.

Abstract: Provided is a gas turbine combustor capable of reducing the size of a low-temperature air layer of pilot air formed between a pilot flame and a premixed flame and of improving the flame stability of the premixed flame. A gas turbine combustor, which is provided with a pilot burner that is provided at the center portion of a combustor main body formed in a cylindrical shape to form a pilot flame, and a plurality of main burners arranged so as to surround the outer periphery of the pilot burner to form a premixed flame, includes, as the ignition improving part, a channel blocking member that reduces the size of the low-temperature air layer of the pilot air formed between the pilot flame and the premixed flame.

Abstract: A gas turbine combustor is provided which can accomplish stable high-load combustion, high combustion efficiency, a low concentration of CO, and NOx in a small-sized combustion chamber. The gas turbine combustor includes an inner liner 12 and an outer liner 14 being concentric and cylindrical and an end liner 16 closing between upstream ends thereof and forms a hollow cylindrical combustion chamber 18 therein. The gas turbine combustor further includes a swirling air flow forming device 22 introducing combustion air 7a into the vicinity of the end liner in the combustion chamber 18 from the outside and forming a swirling air flow, a fuel ejector 24 ejecting fuel 8 in the swirling direction to form a premixed swirling flow, and an igniter 26 igniting the premixed swirling flow to form a tubular flame surface 11.

Abstract: A nozzle includes an inlet at an upstream end of the nozzle, a discharge outlet at a downstream end of the nozzle, and a fluid delivery passage extending between the inlet and the discharge outlet. Exterior and interior walls of the nozzle have downstream tip ends that are relatively longitudinally movable at one or more interfaces. An internal insulating gap is interposed between the interior and exterior walls to insulate fuel from ambient temperature conditions exterior to the nozzle. One or more flexible seal structures internal to the nozzle isolate a portion of the insulating gap from any ambient fluid entering into the gap through the one or more interfaces while providing relative movement between interior and exterior walls of the nozzle.

Abstract: A gas turbine combustor is provided. The gas turbine combustor has a pilot nozzle having a pilot fuel injection port at a downstream end, in terms of the direction flow of a fuel, and lies on the axial centreline of the gas turbine combustor. It also has a main combustion zone, downstream of the pilot nozzle, and a pilot cone projecting from the vicinity of the pilot fuel injection port and having a diverged end adjacent to the main combustion zone, wherein a half angle of the pilot cone is increasing in the downstream direction.

Abstract: A method includes receiving fuel and air into a cup of a turbine fuel nozzle. The method also includes mixing the fuel and air at least partially within the cup. In addition, the method includes directing a fuel air mixture toward a turbine combustor. The method also includes shielding an inner wall of the cup with a blanket of a protective fluid flow to reduce the possibility of flame holding along the inner wall. The protective fluid flow excludes a combustible mixture of fuel and air.

Abstract: A combustor for a gas-turbine engine including a burner head, a combustion chamber disposed downstream of the burner head, a swirler for creating a swirling flow of air in the combustion chamber, and a fuel nozzle disposed in the burner head. The fuel nozzle is disposed giving rise to a first angle of exit of the fuel from a downstream face of the burner head of >±0° with respect to a longitudinal axis of the combustor, this first angle lying in a first plane passing through the longitudinal axis. The fuel also exits at a second angle from the downstream face of >±0° with respect to the first plane, the second angle lying in a second plane orthogonal to the first plane.

Abstract: A fuel nozzle auxiliary vane comprising a vane mountable base comprising a fuel inlet, wherein the vane mountable base is configured to mount to a surface of a main vane disposed in an airflow path of a fuel nozzle. The fuel nozzle auxiliary vane also includes a body extending from the vane mountable base, wherein the body comprises a fuel passage that turns from the fuel inlet to a fuel outlet, and the fuel outlet has a fuel outlet direction generally crosswise to a fuel inlet direction through the fuel inlet.

Abstract: A gas turbine engine swirler/nozzle apparatus has a swirler having a central axis and a nozzle. The nozzle has an outlet end with a plurality of outlets about the central axis and having an asymmetry about the central axis. The apparatus may be formed as a reengineering of a baseline apparatus having a symmetric nozzle.

Abstract: A combustor for mixing a flow of compressed air from a compressor and a flow of fuel from a fuel source. The combustor may include a first swirler for mixing the flow of compressed air and the flow of fuel into a first fuel-air flow, a second flow source for providing a second flow downstream of the first swirler, and a second swirler for mixing the first fuel-air flow and the second flow.

Abstract: The burner (1) of a gas turbine includes a duct (2) which houses four vortex generators (3) and a lance (7) that carries one or more nozzles (8) for injecting a fuel within the duct (2). The lance (7) extends from one of the vortex generators (3).

Abstract: A flameless burner for a gas turbine engine includes a burner body having a longitudinal axis, an upstream section and a downstream section. The upstream section of the burner body defines a primary swirl generating chamber having air swirlers associated therewith. The primary swirl generating chamber is adapted and configured to receive compressor discharge air through the air swirlers. The strong swirl of the compressor discharge air forms a recirculation zone that entrains combustion product gases toward the burner body. Fuel injectors are operatively connected to the downstream section of the burner body for issuing fuel into the recirculated combustion product gases.

Abstract: A premixing device includes an air inlet, at least one fuel inlet slot having a wall profile configured to form a fuel boundary layer along a portion of a wall of the premixing device, a mixing chamber, and at least one diverging fuel injection slot jet disposed inside the at least one fuel inlet slot, the slot jet being configured to create a flow separation region in a diverging portion thereof to generate mixing turbulence at an outlet of the slot jet to aerodynamically enhance a mixing of the fuel from the boundary layer with compressed air without causing a boundary layer flow separation and a flame holding in the mixing chamber. Low-emission combustors, gas turbine combustors, methods for premixing a fuel and an oxidizer in a combustion system, a gas turbine, and a gas to liquid system using the premixing device are also disclosed.

Abstract: A fuel nozzle tip for use with a combustor and a method of assembling the fuel nozzle tip are provided. The fuel nozzle tip includes a fuel tube and an air collar coupled to the fuel tube. The fuel tube includes a first plurality of circumferentially-spaced fuel openings and a second plurality of circumferentially-spaced fuel openings. The fuel tube is configured to channel fuel into a mixing zone defined within the combustor. The air collar includes a plurality of circumferentially-spaced air openings configured to discharge air into the mixing zone. At least one of the plurality of air openings is oriented to facilitate generating a swirl number of greater than 0.6 within the mixing zone.

Abstract: A fuel nozzle for a turbine engine is configured to deliver a large volume of a fuel which has a relatively low amount of energy per unit volume. The fuel nozzle includes a fuel swirler plate having fuel delivery apertures which are angled with respect to the flat surfaces of the swirler plate. A nozzle cap covers the end of the fuel nozzle to create a swirl chamber at the outlet end. The nozzle cap may include a plurality of air inlet apertures to allow the air to enter the swirl chamber.

Abstract: A gas-turbine burner having a plurality of main swirl generators, each having an inlet flow opening formed by the main swirl generator edge, is provided. In order to achieve a uniform flow of combustion air through the main swirl generator, the gas-turbine burner has an inlet-flow guide means with a flow guide surface which runs from one of the inlet-flow openings to an adjacent inlet-flow opening, to which the main swirl generator edges which form the inlet-flow openings are connected, and the flow guide surface widens from there radially upwards. The main swirl generators are central-symmetrically arranged around a pilot burner and the flow guide surface runs radially outside the main swirl generators.

Abstract: A burner for use in a gas turbine engine includes a burner tube having an inlet end and an outlet end; a plurality of air passages extending axially in the burner tube configured to convey air flows from the inlet end to the outlet end; a plurality of fuel passages extending axially along the burner tube and spaced around the plurality of air passage configured to convey fuel from the inlet end to the outlet end; and a radial air swirler provided at the outlet end configured to direct the air flows radially toward the outlet end and impart swirl to the air flows. The radial air swirler includes a plurality of vanes to direct and swirl the air flows and an end plate. The end plate includes a plurality of fuel injection holes to inject the fuel radially into the swirling air flows. A method of mixing air and fuel in a burner of a gas turbine is also provided. The burner includes a burner tube including an inlet end, an outlet end, a plurality of axial air passages, and a plurality of axial fuel passages.

Abstract: A combustion burner 10A according to one embodiment of the present invention includes: a fuel nozzle 110; a burner tube 120 forming the air passage 111 between the burner tube 120 and the fuel nozzle 110; swirler vanes (swirler vanes) 130 arranged in a plurality of positions in the circumferential direction on the external circumferential surface of the fuel nozzle 110, each extending along the axial direction of the fuel nozzle 110, and gradually curving from upstream toward downstream; and a liquid fuel injecting hole 133A from which a liquid fuel is injected to a surface of each of the swirler vanes 130. The combustion burner 10A also includes multi-purpose injecting holes 11-1 to 11-3 as a cooling unit that cools a part of a vane pressure surface 132a of the swirler vane 130 on which the liquid fuel LF hits.

Abstract: A premixed combustion burner for a gas turbine is provided that can efficiently premix fuel and air to produce fuel gas having a uniform concentration, while reliably achieving prevention of flash back by making the flow rate of fuel gas substantially uniform. The premixed combustion burner for a gas turbine has: a fuel nozzle; a burner cylinder arranged so as to surround the fuel nozzle and form an air passageway between itself and the fuel nozzle; and swirler vanes that are arranged along an axial direction of the fuel nozzle in a plurality of positions around the circumferential direction of an outer circumference surface of the fuel nozzle and that gradually curve from an upstream side to a downstream side to spin the air traveling within the air passageway from the upstream side to the downstream side. A cutaway section is provided in a rear edge section on an inner circumference side of the swirler vane.

Abstract: A system that includes a turbine fuel nozzle comprising an air-fuel premixer. The air-fuel premixed includes a swirl vane configured to swirl fuel and air in a downstream direction, wherein the swirl vane comprises an internal coolant path from a downstream end portion in an upstream direction through a substantial length of the swirl vane.

Abstract: A mixer assembly for use in a combustion chamber of a gas turbine engine includes a pilot mixer, a main mixer and a centerbody positioned between the pilot mixer and the main mixer. The pilot mixer includes an annular pilot housing having a hollow interior and a pilot fuel nozzle mounted in the housing and adapted for dispensing droplets of fuel to the hollow interior of the pilot housing. The main mixer includes: a main housing surrounding the pilot housing and defining an annular cavity; a plurality of fuel injection ports for introducing fuel into the annular cavity; and, a swirler arrangement including at least one swirler positioned upstream from the fuel injection ports, wherein each swirler of the arrangement has a plurality of vanes for swirling air traveling through such swirler to mix air and the droplets of fuel dispensed by the fuel injection ports.

Abstract: An embodiment of the present invention provides a method and system of operating a combustion system that has Lean direct injection (LDI) functionality. The method and system provides a passive cooling system for an injector of the LDI system. The method and system may also provide a means to direct the flow of the fluid exiting the injector of the LDI system.

Abstract: A gas turbine engine has a combustor module including an annular combustor having a liner assembly that defines an annular combustion chamber. The liner assembly includes an inner liner and an outer liner that circumscribes the inner liner. The inner liner has a plurality of first combustion air admission holes passing therethrough and the second liner has a plurality of second combustion air admission holes passing therethrough. Each of the first and the second combustion air admission holes has a swirler disposed therein.

Abstract: An apparatus and method fabricating a deflector-flare cone for a combustor is provided. The combustor includes an air swirler annular about a centerline axis of the combustor wherein the swirler includes an annular exit downstream of the swirler. The deflector-flare cone includes a single annular body including an engagement end configured to support the deflector-flare cone, an annular divergent portion extending downstream from the engagement end. The annular divergent portion includes a radially outer annular deflector portion and a radially inner annular flare cone portion that are separated by an annular gap extending between the deflector portion and the flare cone portion. The deflector-flare cone includes a plurality of cooling passages extending through the single annular body of the deflector-flare cone. The plurality of cooling passages are spaced circumferentially about the centerline axis and are configured to be coupled in flow communication with a cooling fluid source.

Abstract: A swirl generator, a method for preventing flashback in a burner with a swirl generator and a burner are provided. The swirl generator includes a central fuel distributor element, an outer wall enclosing the central fuel distributor element and bounding an axial flow channel for combustion air, swirl vanes extending in a radial direction to the outer wall and giving the flowing combustion air a tangential flow component, and a separating wall enclosing the central fuel distributor element and being positioned radially within the outer wall. The separating wall divides the flow channel into a radially inner channel segment and a radially outer channel segment. The radially inner channel segment allows the combustion air to pass without giving it a tangential flow component or while giving it a tangential flow component counter to the orientation of the tangential flow component in the radially outer channel segment.

Abstract: A gas turbine burner, comprising at least one swirler, the swirler having at least one air inlet opening, at least one air outlet opening positioned downstream to the air inlet opening and at least one swirler air passage extending from the at least one air inlet opening to the at least one air outlet opening which is delimited by swirler air passage walls; a fuel injection system which comprises fuel injection openings arranged in at least one swirler air passage wall so as to inject fuel into the swirler air passage; and an air injection system which comprises air injection openings arranged in at least one swirler air passage wall and positioned downstream of the fuel injection openings for injecting air into the swirler air passage.

Abstract: A mixer assembly and a method for forming a fuel-air mixture is combinable with a burner system of a heat engine, especially a gas turbine plant.

Abstract: A lean direct injection fuel nozzle for a gas turbine is disclosed which includes a radially outer main fuel delivery system including a main inner air swirler defined in part by a main inner air passage having a radially inner wall with a diverging downstream surface, an intermediate air swirler radially inward of the main inner air swirler for providing a cooling air flow along the downstream surface of the radially inner wall of the main inner air passage, and a radially inner pilot fuel delivery system radially inward of the intermediate air swirler.

Abstract: A fuel nozzle for a gas turbine engine is disclosed which includes a nozzle body having a longitudinal axis, an elongated annular air passage defined within the nozzle body, and a plurality of circumferentially spaced apart axially extending swirl vanes disposed within the annular air passage, wherein each swirl vane has multiple joined leads and a variable thickness along the axial extent thereof.

Abstract: The combustion chamber includes a chamber endwall perforated by at least one passage hole. An injection system is mounted so that it can slide diametrically with respect to the passage hole. A deflector is mounted on the chamber endwall. The deflector includes a fitting ring which includes first tenons. A sleeve which is coaxial to the fitting ring includes second tenons. The first and the second tenons allow the passage of the first tenons between the second tenons and then the engagement of the first tenons behind the second tenons. The injection system is slidably mounted in the sleeve. Antirotation means are provided to prevent a rotation of the injection system with respect to the sleeve.

Abstract: A fuel nozzle manifold comprising a flange, a stem and a swirler is provided. The flange has a first fluid inlet fluidly connected to a radially extending first flow passage, the stem includes at least a first axially extending and only partially circumferentially extending flow channel, and the swirler has at least a first radially extending premix passage. The flange and the stem can comprise a homogeneous component, or two separate components fluidly connecting the first axially extending flow channel to the first flow passage, to form a fluid connection between the flange and the stem, the swirler comprises another component fitted together with the flange and stem component and fluidly connecting the first premix passage and the first flow channel.

Abstract: An injector includes a surface and an injector hole formed in the surface. The injector also includes a groove formed in the surface, the groove surrounding the injector hole.

Abstract: A system for suppressing combustion instability in a turbomachine includes at least one combustion chamber operatively connected to the turbomachine, and at least one pre-mixer mounted to the at least one combustion chamber. The at least one pre-mixer is configured to receive an amount of fuel and an amount of air that is combined and discharged into the at least one combustion chamber. In addition, the turbomachine includes a combustion instability suppression system operatively associated with the at least one pre-mixer. The combustion instability suppression system is configured to create a combustion asymmetry. The combustion asymmetry facilitates combustion instability suppression in the turbomachine.

Abstract: In one embodiment, a system includes a fuel nozzle that includes a fuel injector configured to output a fuel flow and a premixer tube disposed about the fuel flow output from the fuel injector. The premixer tube includes a perforated portion and a non-perforated portion, and the non-perforated portion is downstream of the perforated portion.

Abstract: A combustor includes a burner tube that receives fuel for combustion from a fuel delivery system. An axial swirler is installed in the burner tube. The burner tube is flared downstream of the swirler. The structure provides for improved combustion stability while extending lean blowout and low frequency dynamics margins, which in turn serve to further reduce NOx emissions.

Abstract: A mixer assembly for use in a combustion chamber of a gas turbine engine includes a pilot mixer, a main mixer, and a fuel manifold. The pilot mixer includes: an annular pilot housing having a hollow interior, a primary fuel injector mounted in the pilot housing, a plurality of axial swirlers positioned upstream from the primary fuel injector, and a plurality of secondary fuel injection ports for introducing fuel into the hollow interior of the pilot housing. The main mixer includes: a main housing surrounding the pilot housing and defining an annular cavity, a plurality of fuel injection ports, and at least one swirler positioned upstream from the plurality of fuel injection ports. The fuel manifold is in flow communication with the plurality of secondary fuel injection ports in the pilot mixer and the plurality of fuel injection ports in the main mixer.