Abstract: A turbomachine including an annular combustion chamber, a sectorized turbine nozzle arranged at an outlet from the chamber, and a sealing mechanism interposed axially between the chamber and the nozzle, the sealing mechanism including an annular gasket that is axially resilient. The gasket includes a first axial bearing mechanism for bearing against a downstream end of the chamber and a downstream annular lip that is sectorized, each sector of the downstream lip being in alignment with a sector of the nozzle and including a second axial bearing mechanism for bearing against an upstream end of the nozzle sector.

Abstract: A combustor liner for a combustor of a turbine engine includes a single-walled generally cylindrical liner that extends from a head end to an aft end. A venturi is formed on the combustor liner at a location between the head and aft ends. The venturi is formed by first and second straight portions of the combustor liner that angle inward to form a reduced diameter portion. The combustor liner may also include an air deflector to deflect air down onto the exterior of the reduced diameter portion of the single-walled combustor liner. The combustor liner may also include dilution holes, and turbulator rings.

Abstract: Fuel and air are injected in a first poloidal flow in a first poloidal direction within a first annular zone of an annular combustor. A first combustion gas from the at least partial combustion of the fuel and air is discharged into an annular transition zone of the annular combustor and transformed to a second combustion gas therein within an at least partial second poloidal flow followed by an at least partial third poloidal flow in the annular transition zone, wherein the direction of the second poloidal flow is opposite to that of the first and third poloidal flows. The second combustion gas is discharged into a second annular zone of the annular combustor, and then transformed to a third combustion gas therein before being discharged therefrom, responsive to which a back pressure is generated in the annular combustor.

Abstract: A shell for a combustor liner includes a cold side, a hot side, a row of cooling holes and a jet wall. The jet wall projects from the hot side for creating a wall shear jet of increased velocity cooling flow in a tangential direction away from the row of cooling holes and along an adjacent heat shield cold side wall.

Abstract: A combustor liner is provided. The combustor liner may include an upstream portion and a downstream end portion. The upstream portion may have a radius and a length along a generally longitudinal axis. The downstream end portion may have a radius and a length along the generally longitudinal axis. The downstream end portion may define a plurality of channels. Each of the plurality of channels may extend helically through the length of the downstream end portion. Each of the plurality of channels may be configured to flow an air flow therethrough, cooling the downstream end portion.

Abstract: Certain embodiments include a first individual sector configured to fit together with a plurality of individual sectors to form a combustor cap assembly of a turbine combustor, wherein the first individual sector is configured to fixedly attach to a first fuel nozzle of a plurality of fuel nozzles, the first individual sector comprises a first substantially enclosed cavity configured to surround the first fuel nozzle, and the first substantially enclosed cavity is configured to receive a cooling air flow.

Abstract: A system for supplying a working fluid to a combustor includes a fuel nozzle, a combustion chamber downstream from the fuel nozzle, and a flow sleeve that circumferentially surrounds the combustion chamber. A plurality of fuel injectors are circumferentially arranged around the flow sleeve to provide fluid communication through the flow sleeve to the combustion chamber. A distribution manifold circumferentially surrounds the plurality of fuel injectors, and a fluid passage through the flow sleeve and into the distribution manifold provides fluid communication through the flow sleeve to the plurality of fuel injectors.

Abstract: Present embodiments are directed toward reducing low flow or no flow situations in a head end of a turbine combustor. Embodiments include a system having a turbine combustor. The turbine combustor includes a fuel nozzle having an inner shell and an outer shell, and a feed cap disposed about the fuel nozzle and having an outer wall and a back plate. The back plate joins respective upstream ends of the outer shell of the fuel nozzle and the outer wall of the feed cap. The turbine combustor is configured to flow a first pressurized air via an air path extending along the outer wall of the feed cap, the back plate of the feed cap, and into the fuel nozzle.

Abstract: A method and apparatus are described that include a transition piece including a cooling sleeve. The cooling sleeve includes a first end and an opposite second end, the cooling sleeve is coupled to an inner wall of the transition piece, such that an annular passage is defined between the inner wall and the cooling sleeve. The first end defines an annular inlet and second end defines an annular outlet.

Abstract: A combustion apparatus in a gas turbine engine comprises a combustor shell for receiving air, a fuel injection system associated with the combustor shell, a first fuel supply structure, and a shield structure. The fuel supply structure is in fluid communication with a source of fuel for delivering fuel from the source of fuel to the fuel injection system and comprises a first fuel supply elements including a first section extending along a first path having a component in an axial direction and a second section extending from the first section along a second path having a component in a circumferential direction. The shield structure is associated with at least a portion of the second section of the first fuel supply element.

Abstract: A flow sleeve for use in a turbine engine including a compressor section, a combustion section, and a turbine section downstream from the combustion section. The flow sleeve includes a sleeve wall defining a pre-mixing passage of the combustion section and has a forward end proximate to a cover plate of the combustion section and an opposed aft end. The sleeve wall includes a first wall section extending axially from the sleeve wall forward end and a second wall section extending from the first wall section between the forward and aft ends of the sleeve wall toward the aft end of the sleeve wall. A cavity is formed between the first wall section and the second wall section.

Abstract: A turbine combustor liner assembly includes a combustor liner having upstream and downstream ends; a transition duct attached to the downstream end of the combustor liner; a first flow sleeve surrounding the combustor liner, with a first radial flow passage therebetween; and a first annular inlet at an aft end of the flow sleeve, the inlet provided with a plurality of circumferentially spaced, angled flow vanes arranged to swirl air entering the first radial flow passage via the annular inlet.

Abstract: A fuel injector assembly for use in a turbine engine having a combustion section and a turbine section downstream from the combustion section. The fuel injector assembly includes a flow sleeve defining a pre-mixing passage of the combustion section and including a sleeve wall having a forward end proximate to a cover plate of the combustion section and an opposed aft end. An annular cavity in fluid communication with a source of fuel is formed in the sleeve wall adjacent the aft end. A fuel dispensing structure is associated with the cavity and includes at least one fuel distribution aperture for distributing fuel from the cavity to the pre-mixing passage.

Abstract: A turbine engine having a plenum for passing fluids from an outlet of a compressor to an inlet of a combustor that may increase the efficiency of the turbine engine. The turbine engine may include a combustor, a compressor positioned upstream of the combustor, a transition channel extending from the compressor to the combustor, and a shell extending between the compressor and a combustor portal and positioned around the at least one transition channel. The turbine engine may also include an axial diffusor in the shell near the at least one transition channel, wherein the axial diffusor may include a fluid flow recess in a trailing edge of the axial diffusor. The turbine engine may also include a wave protrusion extending from a surface positioned radially inward of the axial diffusor. The fluid flow recess and the wave protrusion may reduce fluid flow loss within the shell.

Abstract: An annular combustion chamber for a gas turbine engine includes radially inner and outer walls connected together by a chamber end wall including openings, each of which receives a fuel injection system. Heat protection deflectors are fastened to the chamber end wall. Holes are formed through the chamber end wall to pass cooling air onto an upstream face of each deflector. The inner or outer edge of a deflector presents a sealing rim engaging the respective inner or outer wall of the chamber.

Abstract: A combustor includes first and second annular casings with a joint between the first and second annular casings. A flow sleeve surrounds a combustion chamber to define an annular passage, and an annular shield inside the first annular casing extends downstream from the joint and prevents a working fluid flowing through the annular passage from contacting at least a portion of the first annular casing. A method for reducing thermal stresses in a combustor includes flowing a working fluid from a compressor through an annular passage between a combustion chamber and a flow sleeve inside the combustor, shielding the working fluid flowing through the annular passage from contact with at least a portion of a joint between first and second annular casings, and shielding the working fluid flowing through the annular passage from contact with at least a portion of the first annular casing downstream from the joint.

Abstract: A gas turbine system comprising, a diffuser operative to diffuse an airstream output from a compressor, a fuel nozzle operative to receive fuel and emit the fuel in a combustor, and at least one bleed duct operative to direct bleed air from down stream of the combustor to the fuel nozzle.

Abstract: A can annular combustion arrangement (12) for a gas turbine engine (10), including: a combustor can (14) having an inlet (46) at a head end (62); an annular chamber (30) surrounding the combustor can (14) for delivering a flow of compressed air from a plenum (18) to the inlet (46); a fuel injector (49) configured to inject a flow of fuel into the flow of compressed air; and a flow tripping device (60) surrounding the combustor can (14) and disposed in the annular chamber (30) downstream of the fuel injector (49).

Abstract: The present invention relates to combustors, and provides a combustion liner that may at least partially contain a combustion process. The combustion liner may include a support structure that supports a plurality of panels. The panels may be in the form of thermal barrier panels, e.g., that are attached to the support structure and/or other parts of the machine.

Abstract: Disclosed is a combustor including a baffle plate having at least one through baffle hole and at least one fuel nozzle extending through the at least one through baffle hole. At least one diluent shroud is affixed to the at least one baffle plate and is configured to guide a diluent flow toward a mixing chamber of the at least one fuel nozzle. Further disclosed is a method for introducing a diluent flow into a mixing chamber of a fuel nozzle including urging the diluent flow from a plenum through a baffle plate gap between a baffle plate and an outer surface of the fuel nozzle. The diluent flow is directed via at least one diluent shroud extending from the baffle plate toward a plurality of air swirler holes extending through a fuel nozzle tip. The diluent flow is flowed through the plurality of air swirler holes into the mixing chamber.

Abstract: A combustor within which a combustion zone is defined is provided and includes an annular liner having a first mixing hole defined therein at a first axial position, a flow sleeve, having a second mixing hole defined therein at a second axial position, the flow sleeve surrounding the liner to form a first flow space at an exterior of the liner, a port, coupled to the flow sleeve at the second axial position, which is configured to remove air from the first flow space via the second mixing hole, and a shield, having a third mixing hole defined therein at the second axial position, the shield being disposed to shield the liner and to form a second flow space within the liner, which is communicable with the combustion zone via the third mixing hole and with the first flow space via the first mixing hole.

Abstract: A combustor for an industrial turbine includes a single transition piece transitioning directly from a combustor head-end to a turbine inlet. The transition piece includes an inner surface and an outer surface. The inner surface bounds an interior space for combusted gas flow from the combustor head-end to the turbine inlet. The outer surface at least partially defines an area for compressor discharge air flow. The transition piece includes a plurality of apertures configured to allow compressor discharge air flow into the interior space. Each of the plurality of apertures extends from an entry portion on the outer surface to an exit portion on the inner surface.

Abstract: A compressor diffuser for a gas turbine includes: a compressor diffuser having an upstream end and a downstream end, the downstream end defined by a peripheral annular edge, the annular edge formed with a plurality of substantially axially-oriented slots extending from an opening at the downstream edge in an upstream direction.

Abstract: A thermal machine, especially a gas turbine, includes an annular combustor which is outwardly delimited by an outer shell and an inner shell (33) and through which a hot gas axially flows. The outer shell and inner shell (33) are each provided with a concentric cooling shroud (31) which is attached at a distance on their outer side, forming a cooling passage (32) through which cooling passage (32) cooling air flows in a direction which is opposite to the hot gas flow. The cooling of the combustor is improved by at least one of the cooling shrouds (31), on the side on which the cooling air enters the cooling passage (32), having an outwardly curved, rounded inlet edge (37) for improving the inflow conditions.

Abstract: A combustor includes a first flow path and a second flow path. The first flow path places a diffuser in indirect fluid communication with a combustor cap assembly by way of an intervening lower combustor annular passageway. The second flow path places the diffuser in direct fluid communication with the combustor cap assembly.

Abstract: A combustor assembly in a gas turbine engine. The combustor assembly includes a combustor device coupled to a main engine casing, a first fuel injection system, a transition duct, and an intermediate duct. The combustor device includes a flow sleeve for receiving pressurized air and a liner disposed radially inwardly from the flow sleeve. The first fuel injection system provides fuel that is ignited with the pressurized air creating first working gases. The intermediate duct is disposed between the liner and the transition duct and defines a path for the first working gases to flow from the liner to the transition duct. An intermediate duct inlet portion is associated with a liner outlet and allows movement between the intermediate duct and the liner. An intermediate duct outlet portion is associated with a transition duct inlet section and allows movement between the intermediate duct and the transition duct.

Abstract: A combustor is provided and includes a liner through which fluid fed from at least two injection points flows from a head end to an interior of a transition piece, a first one of the at least two injection points being axially proximate to a fluid impenetrable coupling between the liner and the transition piece and defining apertures disposed in fluid communication with a first passage leading to the head end, and a second one of the at least two injection points being disposed axially between the apertures and the head end and upstream from the apertures relative to a direction of fluid flow through the first passage, the second one of the at least two injection points being formed of openings disposed in fluid communication with a second passage leading to the head end.

Abstract: A combustor assembly that includes a casing that includes a plenum, a combustor liner positioned within the plenum and defining a combustion chamber there within, and a transition piece that includes a forward portion that is coupled to the combustion liner and an aft portion that extends from the forward portion. The aft portion includes a transition piece frame. The transition piece frame includes an upstream surface, a downstream surface, and a radially inner surface extending between the upstream and downstream surfaces. The radially inner surface includes a leading edge that extends between the radially inner surface and the upstream surface. A plurality of first cooling passages extend from the upstream surface to the downstream surface. Each of the first cooling passages is oriented obliquely with respect to the leading edge from the upstream surface towards the downstream surface.

Abstract: A system, in one embodiment, includes a turbine engine. The turbine engine includes a combustor that includes a hollow annular wall having a combustor liner. The turbine engine also includes first flow path in a first direction through the hollow annular wall. The turbine engine further includes a second flow path in a second direction that is opposite the first direction through the hollow annular wall. The second flow path may include one or more film holes configured to supply a cooling film to a downstream end portion of the combustor liner.

Abstract: A combustor assembly for a turbine engine includes a combustor liner, a flow sleeve and a baffle ring. The flow sleeve surrounds the combustor liner. An annulus is formed between the flow sleeve and the combustor liner. A plurality of row of cooling holes are formed in the flow sleeve. The baffle ring radially surrounds the combustor liner and is located in the annulus.

Abstract: A combustor assembly for a turbomachine includes combustor housing and a flow sleeve arranged between the combustor housing and a combustor liner. The flow sleeve defines a first annular fluid passage, and a second annular fluid passage. A quaternary cap is mounted to the combustor housing. The quaternary cap includes a first fluid plenum fluidly connected to the first annular fluid passage, a second fluid plenum fluidly connected to the second annular fluid passage, and a plurality of vanes fluidly connected to each of the first and second fluid plenums. Each of the plurality of vanes includes a body portion having a first fluid channel coupled to the first fluid plenum, and a second fluid channel coupled to the second fluid plenum. The first fluid channel extends completely through the body portion and the second fluid channel extends partially into the body portion.

Abstract: A resilient annular seal structure is disposed radially between an aft end portion of a combustor liner and a forward end portion of a transition piece, the resilient annular seal structure configured to form a first annular cavity radially between the forward end portion of the transition piece and the aft end portion of said combustor. At least one transfer tube radially extends from the second flow sleeve through the second flow annulus to the transition piece, and is arranged to supply compressor discharge cooling air radially from an area outside the first and second substantially axially extending flow annuli directly to the resilient annular seal structure and to the aft end of the combustor liner.

Abstract: A thermal machine including a wall defining a hot gas duct for transferring a hot gas stream and a cooling jacket disposed at a distance from the wall on an outside of the hot gas duct so as to define a cooling duct with an inlet and an outlet. The cooling duct is configured to conduct a cooling medium along an external face of the wall from the inlet to an outlet in a direction counter to a flow of hot gas in the hot gas duct. An impingement cooling plate is disposed at the inlet of the cooling duct and includes cooling baffle holes configured such that cooling medium entering the cooling duct through the cooling baffle holes flows in a direction perpendicular to the wall. The impingement cooling plate is positioned such that an inflow-side edge sealingly abuts the wall of the hot gas duct so as to reduce a transverse flow of the cooling medium through the cooling duct.

Abstract: A turbomachine includes a compressor, a combustor operatively connected to the compressor, an end cover mounted to the combustor, and an injection nozzle assembly operatively connected to the combustor. The injection nozzle assembly includes a first end portion that extends to a second end portion, and a plurality of tube elements provided at the second end portion. Each of the plurality of tube elements defining a fluid passage includes a body having a first end section that extends to a second end section. The second end section projects beyond the second end portion of the injection nozzle assembly.

Abstract: A combustor assembly that includes a casing that includes a plenum, a combustor liner positioned within the plenum and defining a combustion chamber there within, and a transition piece that includes a forward portion that is coupled to the combustion liner and an aft portion that extends from the forward portion. The aft portion includes a transition piece frame. The transition piece frame includes an upstream surface, a downstream surface, and a radially inner surface extending between the upstream and downstream surfaces. The radially inner surface includes a leading edge that extends between the radially inner surface and the upstream surface. A plurality of first cooling passages extend from the upstream surface to the downstream surface. Each of the first cooling passages is oriented obliquely with respect to the leading edge from the upstream surface towards the downstream surface.

Abstract: A method and apparatus for cooling a combustor liner in a combustor are disclosed. In one embodiment, a combustor is disclosed. The combustor includes a transition piece, and an impingement sleeve at least partially surrounding the transition piece and at least partially defining a generally annular flow path therebetween. The combustor further includes an injection sleeve mounted to one of the transition piece or the impingement sleeve and positioned radially outward of the impingement sleeve, the injection sleeve at least partially defining a flow channel configured to flow working fluid to the flow path. In another embodiment, a method for cooling a combustor liner in a combustor is disclosed. The method includes flowing a working fluid through a flow channel at least partially defined by an injection sleeve, and exhausting the working fluid from the flow channel into a flow path adjacent the combustor liner.

Abstract: A system includes an air flow conditioner configured to mount in an air chamber separated from a combustion chamber of a turbine combustor. The air flow conditioner comprises a perforated annular wall configured to direct an air flow in both an axial direction and a radial direction relative to an axis of the turbine combustor. In addition, the air flow conditioner is configured to uniformly supply the air flow into air inlets of one or more fuel nozzles.

Abstract: A combustor for a combustion turbine engine, the combustor that includes: a chamber defined by an outer wall and forming a channel between windows defined through the outer wall toward a forward end of the chamber and at least one fuel injector positioned toward an aft end of the chamber; a screen; and a standoff comprising a raised area on an outer surface of the outer wall near the periphery of the windows; wherein the screen extends over the windows and is supported by the standoff in a raised position in relation to the outer surface of the outer wall and the windows.

Abstract: In a gas turbine combustor and a gas turbine, an air passage (54) that supplies combustion high-pressure air, a pilot nozzle (44), a main fuel nozzle (45), and a top hat nozzle (47) that supply fuel, and openings (64) that supply film air (cooling air) are provided with respect to a combustor inner cylinder (42), and contraction members (71, 72, and 73) are arranged along a circumferential direction on an inner wall surface in a downstream of a flow direction of combustion gas in the inner cylinder (42). The contraction members (71, 72, and 73) are provided in a predetermined area in the circumferential direction excluding penetrating portions (74, 75, and 76), which do not disturb the flow of cooling air, thereby enabling to suppress generation of carbon monoxide and the like and to suppress the occurrence of unstable combustion.

Abstract: A combustor assembly for a turbine engine includes a cap assembly which gradually increases the volume of compressed air provided to the combustor assembly, and which slows the compressed air in a controlled manner before delivering the compressed air into a combustor area. The cap assembly includes inner and outer sleeves which are mounted in a concentric fashion. The annular space formed between the inner and outer sleeves gradually increases from an aft end to a forward end. A stepped portion is formed on the inner sleeve of the cap assembly, and this stepped portion is joined to a corresponding combustor liner.

Abstract: A combustor assembly for a turbine engine includes a combustor liner, a flow sleeve and a plenum ring. The flow sleeve surrounds the combustor liner to form an annulus radially between the combustor liner and the flow sleeve. The flow sleeve has a plurality of rows of cooling holes. The plenum ring radially surrounds the combustor liner in the annulus. The plenum ring has a plurality of bypass tubes for directing axial air flow and radial flow chambers for directing radial air flow.

Abstract: A combustor for a combustion turbine engine that includes: a chamber defined by an outer wall and forming a channel between windows defined through the outer wall toward a forward end of the chamber and at least one fuel injector positioned toward an aft end of the chamber; and a multilayer screen filter comprising at least two layers of screen over at least a portion of the windows and at least one layer of screen over the remaining portion of the windows. The windows include a forward end and a forward portion, and an aft end and an aft portion. The multilayer screen filter is positioned over the windows such that, in operation, a supply of compressed air entering the chamber through the windows passes through at least one layer of screen. The multilayer screen filter is configured such that the aft portion of the windows include at least two layers of screen, and the forward portion of the windows includes one less layer of screen than the aft portion of the windows.

Abstract: A gas turbine transition duct apparatus is provided comprising first and second turbine transition ducts and a strip seal. The strip seal may comprise a sealing element and a spring structure.

Abstract: A gas turbine engine includes an exhaust liner having cold and hot sheets radially spaced from one another and interconnected by a band arranged in a cavity between the hot and cold sheets. In one example, the band is Z-shaped to permit thermal growth of the hot sheets relative to the cold sheets in the axial and radial directions. The hot sheets include axially adjacent portions that provide slots that are in fluid communication with the space. Cooling fluid is provided to the cavity through impingement jets in the cold sheet. The slots are arranged radially between the hot sheet portions.

Abstract: A combustion duct for a gas turbine engine has a combustion liner to receive products of combustion, and deliver the products of combustion downstream toward a turbine rotor. An outer housing is positioned radially outwardly of the combustion liner. A flow sleeve is positioned radially intermediate the outer housing and the combustion liner. A chamber radially outwardly of the flow sleeve receives cooling air. A plurality of holes through the flow sleeve deliver cooling air from the chamber against an outer periphery of the combustion liner. A plurality of tabs are associated with at least some of the holes in the flow sleeve. The tabs are positioned to extend radially inwardly on a downstream side of the holes.

Abstract: A gas turbine combustor includes a combustion chamber defined by a combustor liner, the combustor liner having an upstream end cover supporting one or more nozzles arranged to supply fuel to the combustion chamber where the fuel mixes with air supplied from a compressor. A transition duct is connected between an aft end of the combustion chamber liner and a first stage turbine casing, the transition duct supplying gaseous products of combustion to the first stage turbine nozzle. One or more additional fuel injection nozzles are arranged at an aft end of the transition duct for introducing additional fuel into the transition duct upstream of the first stage turbine nozzle.

Abstract: A combustor for a turbine engine includes an outer liner having a row of circumferentially distributed outer combustion air holes and an inner liner circumscribed by the outer liner and having a row of circumferentially distributed inner combustion air holes. The inner and outer liners each include at least a major air hole having a first hole size, an intermediate air hole having a second hole size, and a minor air hole having a third hole size. The first, second, and third hole sizes are all different from each other.

Abstract: Transition duct assemblies and gas turbine engine systems involving such assemblies are provided. In this regard, a representative a transition duct assembly for a gas turbine engine includes: an impingement sheet having cooling holes formed therethrough, an inlet end and a non-flanged outlet end, the impingement sheet being operative to be positioned about an exterior of a transition duct such that cooling air is directed to flow about the transition duct; the non-flanged outlet end of the impingement sheet being operative to attach the impingement sheet to the transition duct such that the inlet end is positioned adjacent to an intake end of the transition duct and the outlet end is positioned adjacent to an exhaust end of the transition duct.

Abstract: A combustor assembly for a turbine includes a combustor and a combustor liner; a first flow sleeve surrounding the combustor liner forming a first substantially axially-extending flow annulus radially therebetween. The first flow sleeve has a first plurality of apertures formed about a circumference thereof for directing compressor discharge air as cooling air radially into said first flow annulus. A transition piece is connected to the combustor liner, the transition piece adapted to carry hot combustion gases to the turbine, and a second flow sleeve surrounds the transition piece forming a second substantially axially-extending flow annulus radially therebetween. The second flow sleeve has a second plurality of apertures for directing compressor discharge air as cooling air radially into the second flow annulus, the first substantially axially-extending flow annulus connecting with the second substantially axially-extending flow annulus.

Abstract: A combustor assembly for a turbine engine includes a fuel nozzle located at an upstream end of the combustor assembly. The fuel nozzle includes both a fuel delivery passageway and a purge air passageway. The purge air passageway conveys a flow of purge air to cool the fuel nozzle. The combustor assembly also includes a combustion product return line that conveys a flow of combustion products from a position downstream of a combustion zone of the combustor back to the fuel nozzle. The combustion products are mixed with purge air, and the mixture is delivered into a combustion zone located just downstream of the fuel nozzle. The addition of the combustion products into the purge air reduces the amount of oxygen in mixed flow, which helps to reduce the generation of undesirable nitrogen oxide combustion byproducts. Also, the greater heat of the combustion products helps to increase the temperature of the mixture, which helps to maintain the stability of a pilot flame fed by the mixture.