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 leading 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: A gas turbine engine fuel nozzle assemblage (10) comprises a nozzle outer body (25) comprising a premixing section (23) and an interiorly disposed first fuel passageway (22) to supply fuel to the premixing section (23), and more interiorly a central channel (33) comprising an inlet (30) from an air supply, an aperture (19) at an upstream end to receive a lance (18), and a downstream-disposed annular bearing surface (21) adapted to slidingly engage the lance, the fuel nozzle assemblage adapted so that the lance may be alternatably selected as any interchangeable one of a gas lance, an oil lance, and a non-fueling dummy lance. Between the lance and the central channel (33) is formed a cooling fluid passageway (24) that is in fluid communication with the inlet (30) and one or more cooling fluid passageway exits (60). The interchangeability of the lances (18) is effective to provide a desired flexible functionality to said fuel nozzle assemblage.

Abstract: A method and system (14) for monitoring a health of a combustion dynamics sensing system (10) includes monitoring respective dynamic conditions of at least two combustor cans (16) of a can annular combustor (12) of a gas turbine engine with respective dynamic condition sensors (20) associated with each of the cans. The method also includes establishing a baseline relationship between the respective dynamic conditions and then identifying a variance from the baseline relationship indicative of a degraded signal quality provided by a dynamic condition sensor associated with at least one of the cans.

Abstract: A method for monitoring combustion dynamics of a can annular combustor (12) of a gas turbine engine includes monitoring respective dynamic operating conditions of cans (16) of the can annular combustor with respective dynamic operating condition sensors (20) associated with each of the cans. The method also includes grouping the cans into two or more groups according to their respective dynamic operating conditions and identifying a sensor providing an anomalous dynamic operating condition reading for at least one of the cans. The method further includes determining a need to service the identified sensor according to the associated can's group membership.

Abstract: A method and system for controlling combustion in a gas turbine engine (10) includes reducing an overall fuel flow provided to a stage (22) of burners (e.g. 34, 35) of the gas turbine engine until reaching a predetermined dynamic operating condition of the first burner of the stage. The method also includes maintaining, while continuing to reduce the overall fuel flow (e.g. 30), a first portion (e.g. 36), of the overall fuel flow delivered to the first burner at a maintenance level so that the predetermined dynamic operating condition of the first burner is maintained.

Abstract: A transition (200) for a gas turbine engine (100) comprises a transition wall (201) comprising cooling channels (213L, 213R, 223L, 223R) that are adapted to pass a cooling fluid, such as compressed air from a compressor (102) during operation of the gas turbine engine (100) so as to cool combusted hot gases passing through the transition (200). For each cooling channel (213L, 213R, 223L, 223R), respective entry ports (212L, 212L, 222L, 222R) and exit ports (214L, 214R, 224L, 224R) are arranged so as to obtain a performance improvement based upon pressure differentials between the respective entry and exit ports. In various embodiments, a scoop (220) is associated with an entry port (222L, 222R) so as to establish a more elevated pressure differential in the respective cooling channel (223L, 223R). An entry port (330a-h ) may be positioned offset relative to a lower-positioned exit port (340a-h ) so as to so minimize or eliminate intake of heated airflow from a respective nearby exit port (340a-h ).

Abstract: A curved diffuser (210) in a gas turbine engine (201) directs a primary portion of air flow from a compressor (202) through a curved discharge opening (213) into a plenum (220). The curved diffuser (210) also comprises ports (217) through which a secondary portion of air passes into confined space (225) that is defined in part by a pressure boundary element that may be comprised of at least one plate (222) or at least one conduit (306). The at least one plate (222) and the at least one conduit (306) respectively comprise apertures (246, 312) through which pass the secondary portion of air to provide impingement-type cooling to transitions (230, 320). In various embodiments the velocity of the air between adjacent transitions (230, 320) may flow at relatively uniform velocity along the longitudinal distance of the respective transitions (230, 320).

Abstract: A method and system for controlling combustion in a gas turbine combustor (18) includes adjusting an amount of a premix fuel portion (30) of a pilot fuel (28) provided to a premix burner stage (48) of a pilot (46) of the gas turbine combustor. The amount of the premix fuel portion is adjusted from a preset premix fuel portion amount to an adjusted premix fuel portion amount to achieve a desired first operating condition of the combustor. An amount of a diffusion fuel portion (34) of the pilot fuel provided to a diffusion burner stage (50) of the pilot is then adjusted from a preset diffusion fuel portion amount to an adjusted diffusion fuel portion amount to achieve a desired second operating condition of the combustor.

Abstract: Aspects according to the invention relate to a catalytic combustor system for a turbine engine and an associated method. Catalytic combustors are used in connection with turbine engines because they can minimize the formation of oxides of nitrogen during combustion. Despite this emissions advantage, catalytic combustion systems can increase the level of CO in the turbine exhaust. According to aspects of the invention, vortex formation devices includes vortex generators, swirlers and mixers can be placed downstream of each catalytic module surrounding the pilot nozzle so as to form one or more vortices in the otherwise substantially laminar flow exiting the modules. The vortices can create a suction so that a portion of the flow exiting the pilot nozzles is mixed with the flow exiting the catalyst modules. The introduction of the higher temperature pilot flow can accelerate the catalytic reaction time, promoting burnout of the CO formed during combustion.

Abstract: A gas turbine engine (120) includes a cylindrical basket (146) having an axis (14) and a single main burner assembly (12) disposed within the basket. A burner insert (34) is disposed in an annular space between the burner assembly and the basket. The insert includes a face perpendicular to the axis of the basket. A plurality of passageways (114) are formed in the basket, positioned proximate to and downstream of the burner insert for allowing passage of a portion of an oxidizer flow (42) into a combustion chamber (30). A fluid flow path (38), defined between a combustion chamber liner portion (32) of the basket and a casing (40) spaced radially outward from the combustion liner portion, discharges a fluid into a flow reversal region (118) proximate an inlet (20) of the burner assembly. A fuel outlet (44) is disposed in the flow reversal region.

Abstract: Aspects of the invention relate to resonator assemblies for use in non-uniform flow environments. The resonator assemblies include one or more features, such as a box or a scoop, for substantially equalizing the pressure on the resonator. In the box configuration, a box is attached on top of the resonator. The box has a top plate with a plurality of openings and at least one side wall extending from the entire periphery of the top plate. A plenum is defined between the box and the resonator plate. In the scoop configuration, a scoop is attached to the top of the resonator such that the scoop substantially overhangs the resonator. The scoop includes at least one side wall extending substantially perpendicularly therefrom, except for one side without a side wall so as to provide an opening into a space defined between the scoop and the resonator.

Abstract: A combustor (18) for a gas turbine engine (10) includes a combustor burner (34) receiving an oxidizer flow (36). The combustor burner includes an annular vortex generator (38) disposed around a central region (40) of the burner. A fuel outlet (46) is disposed proximate the vortex generator for discharging a combustible fuel (47) into the oxidizer flow. A pilot burner (56) is disposed in the central region of the burner.

Abstract: Aspects of the invention relate to a system for attaching a ceramic liner to a non-ceramic combustor head-end component while accommodating different rates of thermal expansion of these components. One end of the liner can be received within a slot formed by the combustor head-end component. The liner can be held in place by a plurality of pins with each pin passing through a pair of aligned openings in the liner and the head-end component. The end of the liner can be spaced from each of the walls of the slot so as to form a series of gaps in fluid communication. The gaps and the pins can accommodate differential thermal expansion between the liner and combustor head-end component. If desired, at least some of the gaps can be tightly controlled so as to regulate air leakage around the end of the liner.

Abstract: Embodiments of the invention relate to a combustor flow sleeve for a turbine engine. The flow sleeve can be configured to optimize cooling and airflow distribution. The flow sleeve can include first and second sets of openings. A first set of openings can be provided for impingement cooling the areas of the liner that are subjected to high thermal loads. The second set of openings can be provided to more evenly distribute the airflow into the combustor head-end. By focusing the cooling on the areas of need and by making the airflow more uniform, embodiments of the invention can reduce the system pressure drop and enhance the performance and power of the engine.

Abstract: Aspects according to the invention relate to a catalytic combustor system for a turbine engine and an associated method. Catalytic combustors are used in connection with turbine engines because they can minimize the formation of oxides of nitrogen during combustion. Despite this emissions advantage, catalytic combustion systems can increase the level of CO in the turbine exhaust. According to aspects of the invention, vortex formation devices includes vortex generators, swirlers and mixers can be placed downstream of each catalytic module surrounding the pilot nozzle so as to form one or more vortices in the otherwise substantially laminar flow exiting the modules. The vortices can create a suction so that a portion of the flow exiting the pilot nozzle is mixed with the flow exiting thee catalyst modules. The introduction of the higher temperature pilot flow can accelerate the catalytic reaction time, promoting burnout of the CO formed during combustion.

Abstract: Aspects of the invention relate to resonator assemblies for use in non-uniform flow environments. The resonator assemblies include one or more features, such as a box or a scoop, for substantially equalizing the pressure on the resonator. In the box configuration, a box is attached on top of the resonator. The box has a top plate with a plurality of openings and at least one side wall extending from the entire periphery of the top plate. A plenum is defined between the box and the resonator plate. In the scoop configuration, a scoop is attached to the top of the resonator such that the scoop substantially overhangs the resonator. The scoop includes at least one side wall extending substantially perpendicularly therefrom, except for one side without a side wall so as to provide an opening into a space defined between the scoop and the resonator.

Abstract: A gas turbine engine (120) includes a cylindrical basket (146) having an axis (14) and a single main burner assembly (12) disposed within the basket. A burner insert (34) is disposed in an annular space between the burner assembly and the basket. The insert includes a face perpendicular to the axis of the basket. A plurality of passageways (114) are formed in the basket, positioned proximate to and downstream of the burner insert for allowing passage of a portion of an oxidizer flow (42) into a combustion chamber (30). A fluid flow path (38), defined between a combustion chamber liner portion (32) of the basket and a casing (40) spaced radially outward from the combustion liner portion, discharges a fluid into a flow reversal region (118) proximate an inlet (20) of the burner assembly. A fuel outlet (44) is disposed in the flow reversal region.

Abstract: A pre-mixing burner (10) for a gas turbine engine having improved resistance to flashback. Fuel (32) is supplied to a pre-mixing chamber (24) of the burner from a plurality of fuel outlet openings (34) formed in fuel pegs (36) extending into the flow of air (30) passing through the chamber. The fuel outlet openings are formed to direct the fuel in a downstream direction at an angle (A) relative to the direction of the flow of air past the respective fuel peg. This angle imparts a downstream velocity vector (VD) for increasing the net velocity of the air and a normal velocity vector (VN) for directing the fuel away from the wake (44) formed downstream of the fuel peg. Alternate ones of the fuel outlet openings along a single fuel peg may be formed at respective positive (A) and negative (B) angles with respect to a plane (46) extending along the wake in order to minimize the size of the wake.

Abstract: A gas turbine engine (10) includes a plurality of can combustors (19). Each can combustor includes a first stage of burners (46) located at a first radius about the combustor centerline (42) and a second stage of burners (50) located at a second radius greater than the first radius. The second stage of burners may be clocked to an angular position that is not midway between respective neighboring burners of the first stage. Combustion instabilities may be controlled by exploiting variations in combustion parameters created by differential fueling of the two stages.

Abstract: Delivery of a drug is controlled to impart a delay before release after administration by formulating the drug with a disruption agent to provide a core, and coating the core with a regulatory membrane comprising a water-soluble gel-forming polymer and a water-insoluble film-forming polymer.