Abstract: The damper arrangement includes a plurality of interconnected volumes and a plurality of necks for connecting the damper to a combustion chamber at a plurality of contact points. The plurality of necks are connected to the plurality of volumes.

Abstract: A system and method for reducing combustion dynamics includes first and second combustors arranged about an axis, and each combustor includes a plurality of tubes that extend axially through at least a portion of the combustor and a combustion chamber downstream from the plurality of tubes. A fuel injector extends through each tube to provide fluid communication into each tube at a fourth axial distance from the combustion chamber. The fourth axial distance in the first combustor is different than the fourth axial distance in the second combustor.

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 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: The invention relating to a A damper for combustion oscillation damping for a gas turbine, includes a resonator cavity with a box or cylinder shape, a neck in flow communication with the resonator cavity and a combustion chamber. The length Lneck of the neck is equal or greater than a maximum distance lmax that a hot gas ingested from the combustion chamber reaches in the neck. The maximum distance lmax is determined according to the following equation: l max = ? t 1 t 2 ? U - p ^ ? ? ? U ? ? ? ? sin ? ( ? ? ? t ) ? ? ? t . With the solution of the present invention, the damper used to damping combustion oscillation may prevent hot gas ingestion during normal operations by uniquely determining the dimensions of the neck connecting the resonator cavity and the combustion chamber.

Abstract: Apparatus and methods for attenuating noise, vibration, and/or acoustic energy in a turbomachine. The apparatus includes a first base plate having first base plate perforations defined therethrough. The apparatus also includes first and second spacers each extending from the first base plate, being spaced apart from each other, and defining a first channel therebetween. The apparatus further includes a first insert disposed in the first channel and having insert perforations defined therethrough, the first base plate, the first and second spacers, and the first insert at least partially defining a first acoustic chamber therebetween, wherein the base plate perforations and the insert perforations are in communication with the first acoustic chamber.

Abstract: A combustion arrangement for a turbine is provided including: a casing; a combustion chamber within the casing, an inner casing volume may be defined as a volume inside the casing but outside the combustion chamber; a partitioning wall partitioning the inner casing volume into first and second volume portions, the partitioning wall may have at least one aperture to allow fluid communication between the first and second volume portions; and a valve may be arranged at the casing to allow an outgoing fluid flow from the inner casing volume to an outside of the casing depending on a valve operating position. The combustion chamber has a combustion entry port for supplying an oxidant into the combustion chamber, where the combustion entry port may be in fluid communication with the first volume portion. The arrangement may be adapted to adjust the valve operating position for damping an oscillation of the arrangement.

Abstract: A Helmholtz damper for a combustor of a gas turbine includes first and second damping volumes. The combustor has a combustion chamber disposed in a housing and closed off by a front plate at which a plurality of burners are exchangeably fastened. The burners are supplied with fuel via fuel lances which extend from outside the housing through a bushing of the housing to the associated burner. The first damping volume has a first end and a second end, the first end of the first damping volume being configured to attach a connecting passage extending to a front panel such that the Helmholtz damper is connectable with the front plate of the combustion chamber in place of one of the burners. The second damping volume has a first end and a second end and is arranged in series with the first damping volume along an axis of the Helmholtz damper with the first end of the second damping volume being detachably connected to the second end of the first damping volume so as to form a combined larger damping volume.

Abstract: A combustor for a gas turbine that includes a front panel, an elongated sleeve with first end and second ends and a burner mounted in the sleeve. The second end of the sleeve seallessly mounted on the front panel. The sleeve and burner are configured to enable slidable mounted of the burner in the sleeve.

Abstract: The damper arrangement include two concentric hollow shapes, each having a wall, wherein the walls form an annular volume therebetween. The damper arrangement further includes one or more necks for connecting to a combustion chamber at corresponding one or more contact points. The one or more necks are connected to the annular volume.

Abstract: The invention relates to a method for damping thermo-acoustic oscillations via a resonator having a resonator volume with thermo-acoustic oscillations induced in a cavity inclined to thermo-acoustic oscillations, and at least some of the thermo-acoustic oscillations being output from the cavity and being input into the resonator volume, with damping oscillations produced in the resonator volume, which are matched to the thermo-acoustic oscillations, with at least some of the damping oscillations being input into the cavity such that the damping oscillations and the thermo-acoustic oscillations overlap in an overlap area, and the oscillations being largely cancelled out. The invention relates to an apparatus for damping thermo-acoustic oscillations in a cavity, comprising a resonator with a resonator volume and a cavity. The invention also relates to a combustion chamber together with a method and an apparatus according to the invention, and to a gas turbine having a combustion chamber.

Abstract: A combustion dynamics control system for an aviation based or land based gas turbine engine employs an acoustic driver that is configured to drive pressure perturbations across a premixed fuel injection orifice to substantially zero in response to a control signal such that fuel flow perturbations across the fuel injection orifice are substantially zero.

Abstract: A bleed assembly for a gas turbine engine is provided. The assembly includes: a duct having an inlet and an outlet; a bleed valve that controls the flow of bleed fluid into the inlet; and a dome-shaped diffuser screen which covers the outlet. The diffuser screen has a plurality of through-holes for passage of the bleed fluid. Each through-hole has one or more nearest-neighbour through-holes at a nearest-neighbour spacing. At the periphery of the diffuser screen, the average nearest-neighbour spacing of the through-holes at a given radial distance from the centre of the diffuser screen increases with increasing radial distance.

Abstract: A combustion chamber (10, 20) for a gas turbine (1) having a housing (12, 23), a combustion zone (21, 22, 47) surrounded by the housing (12, 23), and at least one burner arrangement (11, 25), which has at least one pre-mixing passage (28) opening into the combustion zone (21, 22) for preparation of a fuel/air mixture. To enable suppression of combustion chamber pressure fluctuations, at least one interference body (50) is provided against which a flow can take place and to which a Strouhal number can be assigned, the flow generating fluidically induced sound waves. The interference body (50) is in a passage (48, 59) opening into the combustion zone (21, 22, 47). A fluid (51, 62) flowing through the passage (48, 59) in the direction of the combustion zone generates a sound wave with a frequency f by flowing against the interference body (50), wherein the frequency f substantially corresponds to a dominant frequency of the combustion chamber.

Abstract: A gas turbine including a combustion oscillation suppressing device disposed to a transition piece of a combustor to define a gas space S, and having a plurality of vent holes through which a gas space L and the inside of the transition piece communicate. The combustion oscillation suppressing device is formed of a first member and a second member which define the gas space, the second member having the plurality of vent holes, and a portion of a transition piece wall includes the second member. A distance from the second member to the first member in the radial direction of the transition piece is a height of the gas space, and when the height of the gas space varies along an axial direction of the transition piece, an opening ratio ?p of each of the plurality of vent holes is adjusted depending on the height of the gas space.

Abstract: A noise reducing device in a jet engine that has a cylindrical casing, a cylindrical partition wall that is inserted in the casing while protruding partially from a trailing edge of the casing, and a compressor that compresses air that is taken into the cylindrical partition wall, with the inside of the cylindrical partition wall serving as a duct in which a core stream of high-speed air flows, and the space between the cylindrical partition wall and the casing serving as a duct in which a bypass stream of low-speed air flows, and being connected with a wing of an airplane by a pylon that has a projection portion that extends beyond the casing to the downstream of the core stream and the bypass stream, the noise reducing device includes a nozzle, disposed at the pylon on the downstream of the core stream, that injects a fluid toward a noise generation source that is produced from the mutual approach of an ambient air stream that is produced outside of the bypass stream and the core stream.

Abstract: A combustion chamber (1) for a gas turbine plant, has a combustion chamber wall (10), which is flowed through by combustion gases in the direction of a downstream expansion turbine, the chamber wall (10) has a device (20) for damping thermoacoustic oscillations caused by the combustion gases. At least one resonator tube (22), interacts with the resonator volume (21) and opens out into the combustion chamber (1) with its mouth (M) opposite from the resonator volume (21) in the combustion chamber inner wall (10). At least one feed opening (23, 23?, 23?) for sealing air for sealing the resonator tube mouth (M) is introduced into the combustion chamber (1) from a compressor plenum (2), surrounding the combustion chamber. The at least one first feed opening (23?, 23?) is provided in a region of the combustion chamber wall (10) close to the resonator tube mouth (M) and is aligned such that the sealing air (S) through the feed opening (23?, 23?) flows over the resonator mouth (M).

Abstract: A combustion chamber (1) for a gas turbine plant having a combustion chamber wall (10), through which combustion gases (G) flow in the direction of a downstream gas turbine, wherein the combustion chamber wall (10) has a dampening device (20) for dampening thermoacoustic vibrations caused by the combustion gases (G) and wherein the dampening device (20) comprises at least one Helmholtz resonator. The resonator volume (21) opens into the combustion chamber (1) with the resonator tube opening (M) in the combustion chamber. At least one supply opening (23) with which sealing air (S) for sealing the resonator tube opening (M) is introduced into the combustion chamber (1) via the resonator volume (21) and an at least one resonator tube (22, 22?, 22?) from a compressor plenum (2). The at least one resonator tube (22, 22?, 22?) from a compressor plenum (2).

Abstract: The invention relates to a damper for reducing the pulsations in a combustion chamber of a gas turbine. The damper includes a resonator cavity with an inlet and a neck tube in flow communication with the interior of the combustion chamber and resonator cavity, and a compensation assembly pivotably connected with the neck tube. The compensation assembly is inserted between the resonator cavity and the combustion chamber to permit relative rotation between the combustion chamber and the resonator cavity. With the damper according to the present invention, by way of providing the compensation assembly, it is assured the relative rotation between the combustion chamber and the resonator cavity is compensated, hence operation life is elongated.

Abstract: The invention relates to a damper for reducing pulsations in a gas turbine, which includes an enclosure, a main neck extending from the enclosure, a spacer plate disposed in the enclosure to separate the enclosure into a first cavity and a second cavity and an inner neck with a first end and a second end, extending through the spacer plate to interconnect the first cavity and the second cavity. The first end of the inner neck remains in the first cavity and the second end remains in the second cavity. A flow deflecting member is disposed proximate the second end of the inner neck to deflect a flow passing through the inner neck. With the solution of the present invention, as a damper according to embodiments of the present invention operates, flow field hence damping characteristic in the second cavity constant regardless the adjustment of the spacer plate in the enclosure.

Abstract: An acoustically dampened gas turbine engine having a gas turbine engine combustor with an acoustic damping resonator system is disclosed. The acoustic damping resonator system may be formed from one or more resonators positioned within the gas turbine engine combustor at an outer housing forming a combustor basket and extending circumferentially within the combustor. The resonator may be positioned in a head region of the combustor basket. In one embodiment, the resonator may be positioned in close proximity to an intersection between the outer housing and an upstream wall defining at least a portion of the combustor. The acoustic damping resonator system may mitigate longitudinal mode dynamics thereby increasing an engine operating envelope and decreasing emissions.

Abstract: A turbomachine combustor includes a combustion chamber; a plurality of micro-mixer nozzles mounted to an end cover of the combustion chamber, each including a fuel supply pipe affixed to a nozzle body located within the combustion chamber, wherein fuel from the supply pipe mixes with air in the nozzle body prior to discharge into the combustion chamber; and wherein at least some of the nozzle bodies of the plurality of micro-mixer nozzles have axial length dimensions that differ from axial length dimensions of other of the nozzle bodies.

Abstract: A combustion system for a gas turbine includes a combustion chamber with a wall section separating an outside of the combustion chamber from an inside of the combustion chamber. The wall section has a passage for injecting a combustion medium into the combustion chamber. The combustion system further includes a resonator with a neck section and a resonator chamber, wherein the neck section and the resonator chamber form a resonator volume reducing vibrations within the combustion chamber. The resonator chamber has a first inlet for injecting gaseous medium into the resonator chamber and a second inlet for injecting fuel into the resonator chamber such that a fuel/gas mixture is generated inside the resonator chamber. The neck section is connected from the outside of the combustion chamber to the passage of the wall section such that the combustion medium comprising the fuel/gas mixture is injectable into the combustion chamber.

Abstract: A combustor cap for use with a number fuel nozzles, the combustor cap including a cold side plate, a hot side plate, and a cap cavity extending between the cold side plate and the hot side plate with the number of fuel nozzles extending therethrough. A resonator tube may extend from the cold side plate into the cap cavity.

Abstract: In a silencer, holes are disposed in one row with intervals to each other along one direction, holes forming the one row are disposed in a plurality of rows in a direction perpendicular to the one direction, and positions of the holes in adjacent rows are shifted in the one direction and disposed in a zigzag pattern. A disposition interval of the holes of an outer edge hole row forming one row in at least two sides opposite to each other in the silencing porous plate is set to be larger than a disposition interval of the holes of a row inside the outer edge hole row.

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 combustion device for a gas turbine that includes a portion having first and second walls that defines an inside of the combustion device, a zone between the first and second walls and an outside of the combustion device. A plurality of first passages connect the inside of the combustion device to the zone between the first and second walls and a plurality of second passages connect the zone between the first and second walls to the outside of the combustion device. A plurality of chambers are defined within the zone between the first and second walls, each chamber being connected with one first passage and at least one second passage. Each chamber defining a Helmholtz damper.

Abstract: A combustion arrangement includes a combustion section. Also included is an air discharge section downstream of the combustion section. Further included is a transition region disposed between the combustion section and the air discharge section. Yet further included is a transition piece defining the combustion section and the transition region, wherein the transition piece is configured to carry a combusted gas flow from the combustion section to the air discharge section. Also included is a damping device operatively coupled to the transition piece proximate the air discharge section.

Abstract: A system and method for reducing combustion dynamics includes first and second combustors arranged about an axis, and each combustor includes a plurality of tubes that extend axially through at least a portion of the combustor and a combustion chamber downstream from the plurality of tubes. A fuel injector extends through each tube to provide fluid communication into each tube at a fourth axial distance from the combustion chamber. The fourth axial distance in the first combustor is different than the fourth axial distance in the second combustor.

Abstract: Systems and methods for frequency separation in a gas turbine engine are provided herein. The systems and methods for frequency separation in a gas turbine engine may include determining a hot gas path natural frequency, determining a combustion dynamic frequency, and modifying a compressor discharge temperature to separate the combustion dynamic frequency from the hot gas path natural frequency.

Abstract: The present application provides a fuel delivery system for a combustor with reduced coherence and/or reduced combustion dynamics. The combustor can assembly may include a first manifold for delivering a first flow of fuel to a first set of fuel injectors and a second manifold for delivering a second flow of fuel to a second set of fuel injectors. The first flow of fuel may have a first temperature and the second flow of fuel may have a second temperature. The first temperature may be higher than the second temperature.

Abstract: A system, including a combustor cap assembly, including a first plate configured to divide a combustion chamber from a head end chamber of a combustor, a first sleeve disposed about the first plate, and a first stiffening rib coupled to the combustor cap assembly, wherein the combustor cap assembly is configured to mount in a combustor, and the first stiffening rib is configured to damp vibration caused by combustion dynamics in the combustor.

Abstract: A novel and improved system for cooling and reducing combustion noise in a gas turbine combustor is disclosed. The system includes a flow sleeve for a gas turbine combustor comprising a tubular portion and a conical portion, with a plurality of flow straighteners extending radially inward and between the tubular and conical portions and a plurality of rows of cooling holes extending about the flow sleeve, and an aft ring secured to the outlet end of the conical portion, where the aft ring includes an overlapping piston ring that is able to expand or contract in diameter. A combustion liner extends through the flow sleeve and engages an inlet of a transition duct while the piston ring of the flow sleeve also engages the transition duct to form a seal.

Abstract: The system may include a turbine engine. The turbine engine may include a fuel nozzle. The fuel nozzle may include an air path. The fuel nozzle may also include a fuel path such that the fuel nozzle is in communication with a combustion zone of the turbine engine. Furthermore, the fuel nozzle may include a resonator. The resonator may be disposed in the fuel nozzle directly adjacent to the combustion zone.

Abstract: In a damping device according to the present invention, a damping device 63 is mounted on a bypass pipe 61 that supplies an amount of high-pressure air to a combustor transition piece 33. The damping device 63 includes a fluid introducing unit 71 that forms a fluid introduction space B by covering an outer peripheral portion of the bypass pipe 61, a plurality of acoustic boxes 73a and 73b that forms resonance spaces Da and Db with the base portions connected to the fluid introducing unit 71 and the end portions extending along the outer peripheral portion of the bypass pipe 61 in the circumferential direction, and partition plates 74a and 74b that form resonance ducts Ea and Eb of a predetermined length by partitioning the resonance spaces Da and Db.

Abstract: A combustion arrangement for a turbine is provided including: a casing; a combustion chamber within the casing, an inner casing volume may be defined as a volume inside the casing but outside the combustion chamber; a partitioning wall partitioning the inner casing volume into first and second volume portions, the partitioning wall may have at least one aperture to allow fluid communication between the first and second volume portions; and a valve may be arranged at the casing to allow an outgoing fluid flow from the inner casing volume to an outside of the casing depending on a valve operating position. The combustion chamber has a combustion entry port for supplying an oxidant into the combustion chamber, where the combustion entry port may be in fluid communication with the first volume portion. The arrangement may be adapted to adjust the valve operating position for damping an oscillation of the arrangement.

Abstract: A combustor includes an end cap that extends radially across at least a portion of the combustor. The end cap includes an upstream surface axially separated from a downstream surface. A plurality of tubes extend from the upstream surface through the downstream surface of the end cap to provide fluid communication through the end cap. Each tube in a first set of the plurality of tubes has an inlet proximate to the upstream surface and an outlet downstream from the downstream surface. Each outlet has a first portion that extends a different axial distance from the inlet than a second portion.

Abstract: The present application provides for a combustor for use with a gas turbine engine. The combustor may include a cap member and a number of fuel nozzles extending through the cap member. One or more of the fuel nozzles may be provided in a non-flush position with respect to the cap member.

Abstract: The present application provides a clocked combustor can array for coherence reduction in a gas turbine engine. The clocked combustor can array may include a number of combustor cans positioned in a circumferential array. A first set of the combustor cans may have a first orientation and a second set of the combustor cans may have a second orientation.

Abstract: An airborne mobile platform that has at least one turbofan engine assembly having a fan driven by a core engine, a short nacelle around the fan and a forward portion of the core engine, and a fan exhaust duct through the nacelle. A mixer duct shell is disposed substantially coaxial with and extending forwardly into the fan exhaust duct to provide a mixer duct between the mixer duct shell and a core engine shroud of the core engine. At least a portion of the mixer duct shell has a honeycomb core structure having an inner surface and an outer surface, with an acoustic lining on one of the inner or outer surfaces. The acoustic lining attenuates sound emanating from the turbofan engine assembly.

Abstract: A combustor liner has a plurality of resonators formed thereon. Each resonator has a radially outer wall and at least one side wall. The outer wall can be free of holes. Each resonator has an inner cavity defined between the outer wall, the at least one side wall and the outer peripheral surface of the liner. The at least one side wall of each resonator surrounds a subset of a plurality of holes that extend substantially radially through the liner. A plurality of cooling passages extends generally longitudinally within the liner. Each cooling passage has an inlet in fluid communication with the exterior of the liner and an outlet in fluid communication with the inner cavity of a respective one of the resonators. A coolant, such as compressor air, can enter and flow along the cooling passages to thereby cool the liner and purge the inner cavity of the resonator.

Abstract: An acoustically dampened gas turbine engine having a combustor with an acoustic damping system is disclosed. The acoustic damping system may be formed from an acoustic damping body having at least one orifice configured to receive a combustor nozzle assembly. The acoustic damping body may be positioned in a head region of a combustor basket and may include one or more orifices in the acoustic damping body. The acoustic damping system may mitigate longitudinal mode dynamics thereby increasing an engine operating envelope and decreasing emissions.

Abstract: Combustor. The combustor includes an axially symmetric tube along with means for introducing fuel and air into the tube. A swirler is disposed within the tube to impart rotation in a first direction to the air/fuel mixture. A plurality of holes downstream of the swirler are disposed around the tube and offset at an angle relative to an inward normal to the tube wall. Air is injected through the offset holes to impart rotation to the air/fuel mixture in a second direction opposite to the first direction. A combustion chamber having a diameter larger than that of the tube receives and burns the air/fuel mixture from the tube.

Abstract: A system and method for reducing combustion dynamics includes first and second combustors, and each combustor includes a fuel nozzle and a combustion chamber downstream from the fuel nozzle. Each fuel nozzle includes an axially extending center body, a shroud that circumferentially surrounds at least a portion of the axially extending center body, a plurality of vanes that extend radially between the center body and the shroud, a first fuel port through at least one of the plurality of vanes at a first axial distance from the combustion chamber, the plurality of vanes being located at a second axial distance from the combustion chamber. A second fuel port is provided through the center body at a third axial distance from the combustion chamber. The system further includes structure for producing a combustion instability frequency in the first combustor that is different from the combustion instability frequency in the second combustor.

Abstract: A resonator is provided, having an adaptable resonator frequency for absorbing sound generated by a gas stream of a gas turbine. The resonator includes a neck section, a chamber and a deformable element being deformable under influence of a change of a gas turbine temperature of the gas stream. The shape of the deformable element is predetermined with respect to a respective gas turbine temperature. The neck section and the chamber form a volume of the resonator. The neck section forms a passage coupling the volume with the gas turbine. The deformable element is thermally coupled to the gas turbine in such a way that the shape of the deformable element depends on the respective gas turbine temperature. The deformable element forms a spiral and is installed to the neck section in such a way that an effective diameter of the neck section depends on the gas turbine temperature.

Abstract: A propfan engine has a power-producing core, a plurality of nozzles for exhausting exhaust gas from the core, and a row of propeller blades rotatable about an axis. The blades are positioned downstream of the exhaust nozzles. The exhaust nozzles are rotatable about said axis. The nozzles can be aligned with the gaps between successive propeller blades in order to reduce noise during operation of the engine.

Abstract: A combustion chamber head of a gas turbine has a confinement enclosing a dampening volume (207) and including a combustion chamber-opposite confinement (206) and a combustion chamber-side confinement (210). The combustion chamber-side confinement (210) is provided as perforated wall (210). In the edge area of the combustion chamber-side confinement (210), cooling air can be routed onto the combustion chamber-side confinement (210) via recesses (203) in the confinement (206). This cooling air, which flows along the combustion chamber-side confinement (210), crosses the cooling air flow through the perforated wall (210) in the combustion chamber (101) without mixing with the latter, as both are separated by walls.

Abstract: In one embodiment, a system includes a variable geometry resonator configured to couple to a fluid path upstream from a combustor of a turbine engine. The variable geometry resonator is configured to dampen pressure oscillations in the fluid path and the combustor.

Abstract: A damping device for reducing pressure oscillations in a combustion system includes at least a portion provided with a first, outer wall, a second, inner wall, an intermediate plate interposed between the first wall and the second wall. This intermediate plate forms a spacer grid to define at least one chamber between said first wall and said second wall, first passages connecting each of said at least one chamber to the inner of the combustion system, and second passages connecting each of said at least one chamber to the outer of the combustion system. The second passages open at the same side of said chambers as the first passages, the second passages have a portion extending parallel to the inner wall. This parallel portion of said second passages is equipped with heat transfer enhancing means.

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.