Abstract: An active combustion control device provides modulated fuel flow to the combustor of a turbine engine to mitigate combustion instability. The device includes a valve unit and an electronic valve driver controlling translation and rotation components in the valve unit that drive a metering member having a modulating end. The axial position and rotation of the end geometries relative to static flow ports inside the valve unit modulate the fuel flowing from the valve unit to the combustor atomizer. The amplitude and frequency of the modulated fuel flow match that of the instable combustor fuel flow and the phase is offset by 180 degrees so that the resultant fuel flow is at or near a stable state.

Abstract: A combustor includes a cap assembly that extends radially across at least a portion of the combustor, wherein the cap assembly includes an upstream surface axially separated from a downstream surface, and a combustion chamber downstream from the cap assembly. A plurality of tubes extend from the upstream surface through the downstream surface of the cap assembly, and each tube provides fluid communication through the cap assembly to the combustion chamber. The downstream surface is sloped at an angle with respect to the upstream surface.

Abstract: The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of fuel nozzles and a combustion zone downstream of the fuel nozzles. The combustion zone may include a number of steps such that the combustion zone expands in a radial direction downstream of the fuel nozzles.

Abstract: An exemplary combustor includes at least a portion having an inner liner and an outer cover plate, which together form an interposed cooling chamber. A plurality of hollow elements extend from the liner and protrude into the cooling chamber. Each hollow element defines a damping volume connected to an inner volume of the combustion chamber via a calibrated duct. During operation, the hollow elements damp pressure pulsations and, in addition, also transfer heat.

Abstract: A gas turbine engine combustor includes an annulus with one or more circular rows of burners and a means for providing a number of equiangular spaced apart flame temperature nonuniformities around the annulus during engine operation. The number of the flame temperature nonuniformities being equal to a circumferential acoustic mode to be attenuated in the combustor (i.e three, five, or seven). Fuel lines and/or water lines in supply communication with the burners and metering orifices in a portion of the fuel lines and/or the water lines may be used to produce the flame temperature nonuniformities. The annulus of the burners may have an equal number of equiangular spaced apart first and second arcuate segments of the burners and a means for operating the burners in the first segments and operating the burners in the second segments at different first and second flame temperatures respectively.

Abstract: A burner system is provided that includes at least two adjacent burners that are separate from each other. Each of the two burners has at least one combustion chamber and a head end. The head end includes at least a fuel injection and a fuel-air premix. Each burner has a cap with a cap side and cap upper side, wherein at least the cap upper side is arranged ahead of the head end, seen in the direction of flow. In this manner, a burner plenum is formed between the cap upper side and the head end. The at least two burner plenums thus formed have an acoustic connection. A method for damping such a burner system is also provided.

Abstract: A combustor includes a breech end, a plurality of fuel nozzles connected to the breech end and extending downstream from the breech end; and a shroud that circumferentially surrounds the plurality of fuel nozzles inside the combustor. A support extends radially inside at least a portion of the shroud, and a lip seal is between the support and at least one of the plurality of fuel nozzles.

Abstract: A combustor includes an upstream surface that extends radially across at least a portion of the combustor, a downstream surface that extends radially across at least a portion of the combustor and is axially separated from the upstream surface, and a plurality of tubes that extend through the downstream surface. A resonator is upstream from at least one of the tubes, and a fluid passage extends through the resonator and into the least one tube.

Abstract: Problems can arise with regard to noise created by bleed valve arrangements in gas turbine engines. It is known to utilize pressure differential inducing elements such as pepper pots and perforated surfaces in order to attenuate noise from bleed valves. However, these arrangements tend to have a portion which expands such that the exit aperture into a bypass duct wall can significantly affect mechanical strength and operational performance. By providing internally created pepper pots or other pressure differential inducing elements within a path of an arrangement and a constriction to an exit noise attenuation is still achieved but with less detrimental effect with regard to bypass duct wall strength and aerodynamic losses altering operational performance.

Abstract: A method, an apparatus, and a computer program product are provided in connection reducing jet noise from a convergent-divergent nozzle that includes a divergent section through which at least portion of an exhaust gas flows at supersonic speeds. In one example, an apparatus is equipped with one or more injection ports arranged on the divergent section of the convergent-divergent nozzle, one or more delivery pipes coupled to each of the one or more injection ports on the exterior surface side of the divergent section and operable to transport an injection gas from a source to the divergent section through the one or more injection ports, and a controller operable to control at least one parameter associated with introduction of the injection gas. In an aspect, each of the one or more injection ports provides an opening from an interior surface to an exterior surface of the divergent section.

Abstract: A turbomachine combustor includes a combustor cap having a cap surface and a wall that define, at least in part, a resonator volume. A plurality of injection nozzle members extend from the cap surface. Each of the plurality of injection nozzle members include an inner nozzle member and a plurality of outer nozzle members. An adjustable conduit extends through the wall into the resonator volume. The adjustable conduit includes an internal passage having a dimensional parameter. A combustor dynamics mitigation system is operably connected to the combustor cap. The combustor dynamics mitigation system includes a controller configured and disposed to control one a size of the resonator volume and the dimensional parameter of the adjustable conduit to modify combustor dynamics in the combustor.

Abstract: A system for electrically controlling combustion includes a combustion chamber, one or more sensors, an actuator, and a controller. The controller detects dynamic instabilities based upon input regarding conditions in the combustion chamber from the sensors. The actuator electrically modulates combustion, and the controller operates the actuator to counteract the dynamic instabilities.

Abstract: A system includes a compressor that compresses incoming airflow, and a combustor assembly mixing the compressed incoming airflow with fuel and combusting the air and fuel mixture in a combustion zone. The combustor assembly includes a hot side downstream of the combustion zone and a cold side upstream of the combustion zone. The system also includes a turbine receiving products of combustion from the combustor. The combustor assembly includes a resonator positioned in the cold side of the combustor assembly in an annular passage between a flow sleeve and a casing of the combustor assembly.

Abstract: A resonance damper for damping acoustic oscillations within a combustor housing of a gas turbine engine is provided. The resonance damper includes a container, an opening, and a pipe. The container is configured to be attached to an interior wall of the combustor housing and has a cavity. The opening is provided on the container. The pipe is rigidly connected to the opening to define the resonance damper with the cavity.

Abstract: An apparatus for attenuating acoustic oscillations of a gas flow contained in part by a combustor wall of a gas turbine engine combustor, wherein the combustor includes at least one air/fuel mixer, includes at least one resonating tube with a closed end and an open end and a single cavity between the ends. The tube is located on the combustor wall downstream of the air/fuel mixer.

Abstract: A resonance damper for damping acoustic oscillations within a combustor housing of a gas turbine engine is provided. The resonance damper comprises a baffle plate, multiple openings, and multiple tubes. The baffle plate is configured to be attached to an interior wall of the combustor housing. The baffle plate defines a cavity with the interior wall of the combustor housing. The openings are provided on the baffle plate. Each of the tubes is received within each of the openings to define the resonance damper with the cavity.

Abstract: Embodiments of the present application can provide systems and methods for dampening combustor dynamics. According to one embodiment, the system may include a micromixer. The system may also include at least one annular resonator disposed within the micromixer adjacent to a cap face plate or an impingement plate of the micromixer. The at least one annular resonator may include a first side including a number of holes forming a cold side hole pattern, a second side including a number of holes forming a hot side hole pattern, and a cavity substantially defined by the first side and the second side.

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 gas turbine combustor having a Helmholtz resonator in a post-swirler homogenization zone upstream of the combustor zone. Embodiments of the present invention provide Helmholtz resonators that include cavities that occupy a remainder of a post-swirler homogenization zone, which also includes at least one flow-directing structure that passes a fuel/oxidizer mixture from a main swirler assembly into a combustion zone of a combustor. Thus, an open annular region of the combustor is converted into a multi-purpose zone that includes at least one Helmholtz resonator.

Abstract: A fluidic effector provides enhanced plume mixing for an aircraft engine. Air jet injectors are located on both the external and internal cowl surfaces and angled in opposite directions to induce large scale vortices in the exhaust plume. The vortices mix actuation air with the exhaust plume to produce ejector action. The plume mixes out quickly, thereby lowering jet noise and jet exhaust temperature. The injectors have orientations and injection rates that are adjustable to allow variable mixing rates for use at different flight and engine conditions.

Abstract: A fuel nozzle for a turbine engine includes a primary fuel passageway for supplying fuel to a plurality of radially extending fuel injectors arranged around the exterior of the fuel nozzle. A secondary fuel passageway couples an upstream end of the primary fuel passageway to a downstream end of the primary fuel passageway. The secondary fuel passageway acts as a resonator tube to help reduce oscillations in the fuel flowing through the primary fuel passageway.

Abstract: A combustor includes a combustion chamber and an acoustic liner, wherein through-holes through are formed in a wall of the combustion chamber, a plurality of rows of the through-holes is arranged in a lengthwise direction of the combustor, the through-holes of a first through-hole row are shifted with respect to the through-holes of a second through-hole row so that centers of the through-holes of the first and second through-hole rows adjacent each other are staggered in the circumferential direction, cooling grooves are formed so as to avoid the through-holes inside that is sandwiched between the outer peripheral surface and the inner peripheral surface of the combustion chamber, and the cooling grooves include transverse cooling grooves extending in the circumferential direction between rows of the through-holes, and longitudinal cooling grooves extending in the lengthwise direction between the through-holes and connecting the transverse cooling grooves adjacent to each other.

Abstract: A gas turbine combustion chamber is provided. The gas turbine includes a combustion chamber interior and a combustion chamber wall which has a substantially rotationally symmetrical cross-section, wherein on the side of the combustion chamber wall facing away from the combustion chamber interior there is arranged over the entire cross-sectional circumference of the combustion chamber wall a corrugated component which, in combination with the combustion chamber wall, embodies a plurality of separate resonance chambers and wherein openings are incorporated in the combustion chamber wall in such a way that a fluidic connection is established in each case between the combustion chamber interior and one of the resonance chambers and wherein the corrugated component has two locking rings which are connected to the combustion chamber wall in order to seal off the resonance chambers.

Abstract: A turbomachine includes a combustion chamber, and at least one pre-mixer mounted to the combustion chamber. The at least one pre-mixer includes a main body having a first end portion that extends to a second end portion. The first end portion is configured to receive an amount of fuel and an amount of air and the second end portion defines an exit plane from which a fuel-air mixture discharges into the combustion chamber. The turbomachine also includes a combustion dynamics reduction system operatively coupled to the at least one pre-mixer. The combustion dynamics reduction system includes at least one of a boundary layer perturbation mechanism and an acoustic wave introduction system which disrupt a flow pattern of the fuel-air mixture within the at least one pre-mixer.

Abstract: A system for damping combustor nozzle vibrations includes an end cover and a combustion chamber downstream from the end cover. First and second sets of nozzles extend axially between the end cover and the combustion chamber. The second set of nozzles is adjacent to the first set of nozzles. The system includes means for damping vibrations between the nozzles with a gap between the means for damping vibrations. A method for damping combustor nozzle vibrations includes flowing a working fluid through first and second sets of nozzles, wherein the first set of nozzles includes a damping member attached to and circumferentially surrounding at least a portion of the first set of nozzles, and contacting at least one nozzle in the second set of nozzles with the damping member on at least one nozzle in the first set of nozzles.

Abstract: A resonator system for a turbine engine can improve acoustic performance and cooling effectiveness. During engine operation, a combustor liner exhibits alternating hot and cold regions in the circumferential direction corresponding to the non-uniform temperature distribution of the combustion flame. Accordingly, high flow resonators are formed with the liner in substantial alignment with the hot regions of the fluid flow within the liner, and low flow resonators are formed with the liner in substantial alignment with cold regions of the fluid flow within the liner. As a result, appropriate amounts of cooling can be provided to the liner so that cooling air usage is optimized. Alternatively or in addition, the liner can include two or more rows of resonators, which can provide an enhanced acoustic damping response. The resonators in the first row can be aligned with or offset from the resonators in the second row.

Abstract: A combustor for a gas turbine engine and related method is provided in which a plurality of combustor cans are selectively adapted with corresponding resonators. The resonators may, for example, be attached to every can in the consecutive arrangement of combustor cans, every other can, every third can or the like, and may be tuned to the same or first, second, third, etc. frequencies of operation. Such selective tuning is configured to suppress one or more of out-of-phase and in-phase dynamic interaction of streams discharged from adjacent combustor cans by changing the frequencies of pressure oscillation instabilities across the arrangement of consecutive cans.

Abstract: A combustor (100) for a gas turbine engine (30) has a circumferentially extending liner (109) defining at least a portion of an interior combustion chamber (107) and a hot gas path (115). The liner includes a resonator section (112) including at least one resonator (116A, 116B) having a resonator chamber (125A, 125B) formed on an exterior of the liner. A thermal barrier coating (118) is disposed along an inner surface of the liner including an inner surface (130) of the resonator section. The resonator further includes a plurality of apertures (117) and each aperture extends through the liner and the thermal barrier coating at the resonator section and there is fluid flow communication between the combustion chamber and the resonator chamber.

Abstract: A tunable fluid flow control system for controlling combustion dynamics in a combustor is disclosed. The control system includes a fuel supply path having an inlet portion, a first diverted portion, a second diverted portion, and an exit portion. A first diverted portion has a first length and a first diameter. A second diverted portion has a second length and a second diameter and spaced apart from the first diverted portion. The first and second diverted portions converge at a merging location. The exit portion is coupled to the merging location of the first and second diverted portions. At least one flow regulation device is coupled to the inlet portion and configured to divert fuel flow to the first and second diverted portions alternately to generate fuel flow perturbations at the merging location.

Abstract: Certain embodiments of the present disclosure include a combustor resonator having a non-uniform annulus between a combustor assembly and a resonator shell. The combustor resonator may further include resonator necks or passages having non-uniform lengths and geometries.

Abstract: A combustion chamber head of a gas turbine has a substantially annular combustion chamber outer wall 18 as well as a substantially annular combustion chamber inner wall 42 and several burners 6 distributed around the circumference. The combustion chamber head 5 has an inflow-side wall 13 which together with a wall 14 facing the combustion chamber 7 forms a combustion chamber head volume 15. The inflow-side wall 13 is provided with at least one inflow opening 32, the wall 14 facing the combustion chamber 7 is provided with at least one outflow opening 17 for connecting the combustion chamber head volume 15 to the combustion chamber 7, and at least one cooling air duct 29 is provided in the wall 14 facing the combustion chamber 7. A method for cooling and damping of the combustion chamber head is also disclosed.

Abstract: A resonator device for damping a pressure oscillation within a combustion chamber is proposed. The resonator device has a container filled with a gas, an opening in the container and a heating element adapted to generate a flame. The flame is arranged to heat the gas within the container. The resonator device is in a combustion arrangement having a combustion chamber for defining a combustion space for burning fuel. The container is connected to the combustion chamber such that an inside of the container is in communication with the combustion space via the opening. The resonator device has a resonance frequency equal to a pressure oscillation frequency within the combustion chamber under normal load conditions.

Abstract: Certain embodiments include a resonator for the head end of a turbine combustor. The resonator is partially defined by first and second plates disposed about a fuel nozzle in the head end of the turbine combustor. The resonator is further configured to receive an air flow which extends through the resonator and into the fuel nozzle.

Abstract: A gas turbine combustor is provided with a combustor basket where combustion gas flows, the combustion gas being produced by combustion of fuel injected from a nozzle, and a first resonance device and a second resonance device mounted on an outer surface of the combustor basket. The second resonance device is disposed on a downstream side from the first resonance device in a flow of the combustion gas and damps combustion oscillation of a frequency higher than the first resonance device. The first and second resonance devices are acoustic liners each having a housing mounted to the outer surface of the combustor basket. A resonance space surrounded by the housing and the outer surface of the combustor basket communicates with an interior space of the combustor basket via a plurality of acoustic holes formed in the combustor basket.

Abstract: In a gas-turbine combustion burner which is provided, at a distal end thereof, with a fuel spraying hole that sprays fuel into a combustion region formed inside a combustion cylinder of a gas-turbine combustor and in which a fuel flow path that guides the fuel, which is supplied from a fuel source, to the fuel spraying hole is formed in the interior thereof, the impedance of a fuel supply system that guides the fuel from the fuel source to the fuel spraying hole is set so that propagation of pressure fluctuations from the combustion region to the fuel supply system becomes an allowable level or less.

Abstract: A can-annular gas turbine engine including a casing defining an annulus for receiving air flow from a compressor. A combustor basket is supported within the casing, and a swirler assembly is located within the basket and includes an air inlet and an outlet located in a combustion chamber. The swirler assembly is configured to discharge a fuel/air mixture from the outlet downstream into the combustion chamber. An acoustic filter is provided comprising a Herschel-Quincke tube having a first open end at the annulus and a second open end adjacent to the swirler assembly within the combustion chamber. Components of acoustic pressure oscillations travel along two paths of different lengths from one end of the Herschel-Quincke tube to the opposite end of the tube where out of phase components of the pressure oscillations interact to attenuate the pressure oscillations.

Abstract: A noise reduction device, for example for use in a bleed assembly of a gas turbine engine, comprises partitions having apertures which provide contractions and sudden expansions of flow passing through the device, turbulators being provided between the partitions to break up jets issuing from the Apertures. Breaking up the jets before impact with the partitions enables the partitions to be disposed closer together, so enabling adequate noise reduction to be achieved in a device of relatively small thickness.

Abstract: A method is provided for operating a combustion device having at least one premix burner equipped with fuel stagings and is connected to a combustion chamber. The method includes changing, during transient operating states, the fuel staging of the at least one premix burner in accordance with combustion-induced pulsations which occur in the combustion chamber. A pulsation level of the pulsations which occur in the combustion chamber is continuously determined, the fuel staging of the at least one premix burner is changed for lowering the pulsation level if the pulsation level exceeds a predetermined maximum value, and, after a predetermined time interval, by the fuel staging being reset to the undisturbed value which determines the respective operating point if within the predetermined time interval the predetermined maximum value is not exceeded.

Abstract: A dry low NOx (DLN) fuel nozzle of a fuel system is provided and includes one or more downstream orifices formed to define a first fluid path along which fluid is directed to flow, an upstream orifice, located upstream from the one or more downstream orifices, formed to define a second fluid path along which the fluid is directed to flow and connecting passages disposed to fluidly couple the first and second fluid paths of the downstream and upstream orifices, respectively, to one another. At least one of a radial size and an axial location of the upstream orifice is set to cooperatively detune an acoustic impedance of the fuel system.

Abstract: A burner system is provided that includes at least two adjacent burners that are separate from each other. Each of the two burners has at least one combustion chamber and a head end. The head end includes at least a fuel injection and a fuel-air premix. Each burner has a cap with a cap side and cap upper side, wherein at least the cap upper side is arranged ahead of the head end, seen in the direction of flow. In this manner, a burner plenum is formed between the cap upper side and the head end. The at least two burner plenums thus formed have an acoustic connection. A method for damping such a burner system is also provided.

Abstract: An exemplary combustor includes at least a portion having an inner liner and an outer cover plate, which together form an interposed cooling chamber. A plurality of hollow elements extend from the liner and protrude into the cooling chamber. Each hollow element defines a damping volume connected to an inner volume of the combustion chamber via a calibrated duct. During operation, the hollow elements damp pressure pulsations and, in addition, also transfer heat.

Abstract: An intake manifold is provided for inducting combustion air for a compressor of a gas turbine. The intake manifold includes a plurality of walls that delimit against the environment. Devices are provided for controllable change of the mechanical rigidity of the walls to reduce noise which is generated or is emitted in an air intake region during operation of the gas turbine. A gas turbine is also provided and includes an air intake to which air to be compressed is fed via an intake manifold. The intake manifold having a plurality of walls that delimit against the environment. Devices are provided to controllably change a mechanical rigidity of the walls to reduce noise which is generated or emitted in an air intake region during operation of the gas turbine.

Abstract: A gas turbine combustion chamber is provided. The gas turbine includes a combustion chamber interior and a combustion chamber wall which has a substantially rotationally symmetrical cross-section, wherein on the side of the combustion chamber wall facing away from the combustion chamber interior there is arranged over the entire cross-sectional circumference of the combustion chamber wall a corrugated component which, in combination with the combustion chamber wall, embodies a plurality of separate resonance chambers and wherein openings are incorporated in the combustion chamber wall in such a way that a fluidic connection is established in each case between the combustion chamber interior and one of the resonance chambers and wherein the corrugated component has two locking rings which are connected to the combustion chamber wall in order to seal off the resonance chambers.

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: Buzz saw noise propagates through inlet of aviation engine at high power settings. An inlet of aviation engine with blowing slot blowing air grazing over surface of buzz saw lining at speeds higher than inlet mean air flow speed for buzz saw noise control.

Abstract: A combustor is provided and includes a center body having a tubular inner wall, a cartridge disposed within the tubular inner wall to define an outer annular space between an outer surface thereof and the tubular inner wall and an inner annular space having an open end and a closed end opposite the open end and an insert disposed within the outer annular space to envelop the cartridge proximate to the open end, the insert including a body having a tubular forward portion and a partially tubular aft portion, the partially tubular aft portion thereby forming a recessed track having a longitudinal axis substantially aligned with a longitudinal axis of the outer annular space.

Abstract: A Helmholtz resonator is used to detect a presence of a two-phase flow in a conduit. The resonator is sized so that a range of frequencies it is tuned for corresponds to a change in the speed of sound as the fluid changes from all liquid, to liquid plus vapor, and to all vapor. A band-pass filter is used to cancel out any external noise that would pollute the signal from the resonator. The resonator is used to monitor for a two-phase flow occurring within a feed line leading into a turbopump of a rocket engine.

Abstract: A system includes a resonator controller. The resonator controller may adjust a ratio of a first fluid and a second fluid into a resonator, whereby the first fluid is different than the second fluid. Additionally, the resonator controller may control a resonance frequency of the resonator to dampen oscillations based on the ratio.

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 combustor liner may include an annular inner liner and an annular outer liner with a plurality of air holes thereon. The outer liner may be positioned circumferentially around the inner liner such that an annular cooling space is defined between the inner and the outer liner. The combustor liner may also include at least one resonator coupled to the outer liner such that a base of the resonator is separated from the outer liner to form a gap with an external surface of the outer liner. The combustor liner may also include a throat extending from the base of the resonator penetrating the inner liner and the outer liner. The combustor liner may further include a grommet assembly that allows for relative thermal expansion between the inner liner and the outer liner proximate the throat.