Abstract: A device for improving the effectiveness of the acoustic treatment(s) in an aircraft propulsion unit that includes a pipe (36) into which a gas stream flows, whereby the pipe (36) is delimited by two approximately concentric inside walls (40) and outside walls (42), wherein the device includes at least one partition (38, 38?) that extends over at least one portion of the length of the pipe (36) between the inside and outside walls (40, 42) and in that the profile of at least one wall (44, 46) of the at least one partition (38, 38?) is defined so as to increase the number of reflections of an acoustic beam on at least one of the inside and/or outside walls (40, 42) of the pipe and/or on at least one of the walls (44, 46) of the at least one partition (38, 38?).

Abstract: Embodiments of the present invention provide resonators (260, 460) that have lateral walls (268, 270) disposed at non-square angles relative to the liner's longitudinal (and flow-based) axis (219) such that a film cooling of substantial portions of an intervening strip (244, 444) is provided from apertures (226A, 226B, 426) in a resonator box (262, 462) adjacent and upstream from the intervening strip (244, 444). This film cooling also cools weld seams (280) along the lateral walls (268, 270) of the resonator boxes (262, 462). In various embodiments the lateral wall angles are such that film cooling may be provided to include the most of the downstream portions of the intervening strips (244, 444). These downstream portions are closer to the combustion heat source and therefore expected to be in greater need of cooling.

Abstract: A gas turbine engine having a ramburner is disclosed. The ramburner is disposed downstream of a gas turbine engine combustor and receives an engine exhaust flow from the gas turbine engine combustor. The ramburner also accepts a bypass air. Fuel is injected into the ramburner and a combustion reaction is auto-initiated based upon local gas temperatures. No mechanical flame holders need be used. A slidable valve may be used to vary the amount of bypass air into the ramburner. A movable cowl and a plug nozzle form an exit flow path of the gas turbine engine. The movable cowl can be positioned to vary a throat area and exit area of the gas turbine engine based upon the operation of the ramburner, which may be influenced by the amount of bypass air entering the ramburner.

Abstract: An end cap for a fuel injector of a turbine engine is disclosed. The end cap includes an annular first surface including a plurality of perforations. The annular first surface is exposed to a combustion chamber of the turbine engine. The end cap is coupled to an end face of the fuel injector to define an enclosed cavity. The enclosed cavity and the plurality of perforations form an array of Helmholtz resonators.

Abstract: A combustion device includes a plurality of mixing devices into which a fluid containing oxygen and a fuel are introduced and mixed to form a mixture. The combustion device also includes a combustion chamber in which the mixture formed in the mixing devices is burnt. Each mixing device has a conical body with a lance projecting into the conical body, where the fuel is injectable into the conical body. Tips of the lances of different mixing devices have different distances from corresponding open ends of the respective conical bodies.

Abstract: A resonance chamber (42) has an outer wall (32) with coolant inlet holes (34A-C), an inner wall (36) with acoustic holes (38), and side walls (40A-C) between the inner and outer walls. A depression (33A-C) in the outer wall has a bottom portion (50) that is close to the inner wall compared to peaks (37A-C) of the outer wall. The coolant inlet holes may be positioned along the bottom portion of the depression and along a bottom portion of the side walls to direct coolant flows (44, 51) toward impingement locations (43) on the inner wall that are out of alignment with the acoustic holes. This improves impingement cooling efficiency. The peaks (37A-C) of the outer wall provide volume in the resonance chamber for a target resonance.

Abstract: A combustor assembly is provided comprising a combustor casing; a liner coupled to the combustor casing; a burner assembly coupled to the combustor casing; and at least one resonator assembly. The resonator assembly comprises a resonator outer plate having at least one opening, a resonator side wall coupled to the resonator outer plate, and a resonator inner plate defined by a portion of the liner. The resonator inner plate is coupled to the resonator side wall and has at least one slot formed therein having an aspect ratio of at least 4:1.

Abstract: In a bypass turbomachine, at least one air takeoff orifice is provided in the primary channel, said orifice leading to an air takeoff pipe housed inside an air inlet sleeve of the nacelle. The air takeoff pipe opens out in the vicinity of the pylon into two air diffusion pipes each secured to a respective maintenance cover, each air diffusion pipe opening out into an air injection pipe secured to a thrust reverser cover and itself opening out to the outside of the nacelle via the trailing edge thereof, each air injection pipe being suitable for uncoupling from the corresponding air diffusion pipe when the corresponding thrust reverser cover slides downstream.

Abstract: Device for damping oscillations of a combustion chamber that includes at least one resonator connected to a pre-chamber in a vibration-damping manner. The pre-chamber is connected to a combustion chamber in a vibration-damping manner via at least one passage channel. This abstract is not intended to define the invention disclosed in the specification, nor intended to limit the scope of the invention in any way.

Abstract: A combustion control system for a turbine engine is disclosed. The combustion control system includes a fuel injector having a main fuel supply and pilot fuel supply coupled to a combustor of the turbine engine. The combustion control system also includes a sensor coupled to a transfer tube. The transfer tube is fluidly coupled to the combustor, and the sensor is configured to detect a pressure pulse in the combustor. A semi-infinite coil is also coupled to the transfer tube. The combustion control system also includes a controller electrically connected to the sensor. The controller is configured to compare an amplitude of the pressure pulse within a frequency range to a threshold amplitude, and adjust the pilot fuel supply in response to the comparison.

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: 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: A gas turbine engine test cell has a turbine detuner capable of recovering kinetic energy from exhaust gases emitted by a gas turbine engine while also detuning the exhaust flow to reduce unwanted infrasound. The gas turbine engine test cell includes a test cell building, a thrust frame for mounting the gas turbine engine, and the turbine detuner disposed downstream of the thrust frame for extracting energy from the exhaust gases of the gas turbine engine when in operation. The turbine detuner has an inlet for receiving the exhaust gases, a kinetic energy recovery mechanism (e.g. stator and rotor) for converting the kinetic energy of the exhaust gases into rotary power, and an outlet through which de-energized exhaust gases are emitted after being de-energized by the kinetic energy recovery mechanism. By eliminating the augmentor, the test cell is more compact. The turbine detuner not only reduces infrasound but also recovers otherwise wasted energy.

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 gas-turbine combustion chamber wall for a lean-burning gas-turbine combustion chamber with a combustion chamber casing 2, 3 and several combustion chamber segments arranged in the combustion chamber casing 2, 3 and forming a combustion chamber wall 10. Each of the combustion chamber segments is supplied with cooling air with film cooling via axial and/or radial cooling holes 16, 17, with the cooling holes 16, 17 being axially spaced from each other and with dampening openings 19a being provided in this area for the introduction of cooling air.

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: 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 combustion chamber for a gas turbine engine comprising at least one Helmholtz resonator having a resonator cavity and a resonator neck in flow communication with the interior of the combustion chamber. The cavity extends around at least part of the neck and is spaced apart therefrom to define a cooling chamber therebetween. The cooling chamber allows the neck and the cavity to be cooled with a flow of cooling air. The sunken neck allows the resonator to be located within enclosures with minimal volume without compromising on the damping properties afforded.

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 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: The present application provides a swirler for a turbine combustor. The swirler may include an outer wall, a passage defined by the outer wall, and an end wall. The end wall may include a number of apertures therethrough.

Abstract: In order to realize a stable decrease in NOx, a gas turbine combustor is supplied which can reduce combustion vibration. A combustor (3) is provided with a first box body (30), which is installed outside an object body (20) such as a combustor basket (6), a transition piece (7) or a bypass duct (11) so as to form a first internal space (31) having a predetermined capacity; and a first throat (32) having a predetermined length which has one end (32a) open to a side wall (20a) of the object body (20) and has the other end (32b) open to a first internal space (31); wherein, a first resistive element (33) having a multiple number of through-holes is inserted and engaged to one end (32a).

Abstract: A combustor arrangement for a gas turbine engine (31) has a split line (42) and a plurality of burners (20,37) arranged in an annular ring (40). The burners (46) of the combustor on either side of the split line (42) have a separation distance of at least two times the average separation distance between burners (48) distant from the split line (42).

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 gas turbine combustion system has a combustion system wall delimiting a flow path for hot and pressurized combustion gas and at least one resonator with a resonator volume. The resonator volume is delimited by resonator walls, where one of the resonator walls is located adjacent to or is formed by the combustion system wall. The resonator has at least one cooling fluid supply opening which is open towards a cooling fluid source. It further has a neck opening which is open towards the flow path and which is implemented in the form of a neck slot.

Abstract: One or more Helmholtz-type resonators (270) is/are provided at the junction (260) of a combustor (220) and a combustion chamber (240) of a gas turbine engine (100). In one embodiment, adjacent Helmholtz-type resonators (290, 291, 292), which may be separated by respective baffles (285), have different volumes that help provide for damping different undesired combustion-generated acoustic pressure waves. In some embodiments, a structural member (435) may be provided between adjacent Helmholtz-type resonators (425, 426, 427, 428) at the junction. At least one of the plurality of Helmholtz-type resonators comprises one or more inlet openings (480), and one or more exit openings (482). Embodiments (370, 425-429) are described in which Helmholtz-type resonators provided at the junction are enlarged in size using various approaches.

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: An object of the present invention is to provide a gas turbine that improves appearance of a combustion oscillation suppressing device disposed to a transition piece of a combustor. To achieve the object, a combustion oscillation suppressing device 34 is disposed to a transition piece 32 of a combustor to define a gas space S, and has a plurality of vent holes 343 through which a gas space L and the inside of the transition piece 32 communicate with each other. The combustion oscillation suppressing device 34 has: a first member 341 that has a substantially U-shaped cross-section formed by press-working, and is disposed on the outer circumference of the transition piece 32 with a substantially U-shaped cross-section opening thereof facing the transition piece 32; and a second member 342 that defines the gas space S between the first member 341 and the second member 342 by blocking the substantially U-shaped cross-section opening of the first member 341, and has the vent holes 343.

Abstract: A system is provided for mixing gas flows in a gas turbine engine. The system includes an annular wall structure which internally delimits a channel for a first gas flow. The wall structure includes a plurality of first wall portions. At least one of the first wall portions is movably arranged in a radial direction between a first position and a second position for mixing an external gas flow with the first gas flow.

Abstract: A modification method for reducing emissions from an annular shaped combustor of a gas turbine plant, having uniformly spaced circumferentially mounted premix burners (20), includes the steps of: removing at least one burner (20), thereby disrupting the spatial uniformity of the remaining the burners (20); and modifying the combustor air distribution system so as to compensate for the increased burner pressure drop of the remaining burners, thus enabling the modified combustor to operate at a load equivalent to the unmodified combustor. Emission reduction is enabled by the increase in the gas velocity of the burner for a given load further enabled by the flame stabilizing effect of disrupting the spatial uniformity.

Abstract: A fuel manifold assembly for a gas turbine engine comprises an annular fuel manifold and a plurality of fuel nozzles circumferentially distributed about the fuel manifold. The fuel manifold has at least one fuel conveying passage in fluid flow communication with the plurality of fuel nozzles and defines at least one location susceptible to overheating between two of the plurality of fuel nozzles. A slot extends through the fuel manifold in the susceptible location to reduce heat transfer in the fuel manifold while maintaining the fuel manifold assembly dynamically balanced.

Abstract: Methods and devices for increasing the sound attenuation properties of a noise suppression device are provided. Various embodiments include sound-attenuating cavities with improved sound capturing openings. In some embodiments, the sound capturing openings include holes surrounded by projections. In other embodiments, the sound capturing openings include angled slats. Embodiments also include sound capturing devices that include the improved sound capturing openings, such as silencers, vent hoods, barrier walls, air exchanges, and air intake and exhaust ducts.

Abstract: Methods and devices for increasing the sound attenuation properties of a noise suppression device are provided. In some embodiments, a sound attenuating conduit has an improved shape that provides increased sound attenuation properties. In other embodiments, the sound attenuating conduits are stacked together, and shaped to increase the air flow cross-section. Still other embodiments provide machinery that includes the improved sound attenuating conduits.

Abstract: A method for operating a turbofan engine assembly is provided. The turbofan engine assembly includes a core cowl which circumscribes the core gas turbine engine, a nacelle positioned radially outward from the core cowl, a fan nozzle duct having an area defined between the core cowl and a portion of the nacelle, a first cowl and a second cowl that define a portion of the nacelle wherein the second cowl is repositionable with respect to the first cowl, and a thrust reverser assembly wherein the second cowl surrounds the thrust reverser assembly. The method includes varying an operating speed of the fan assembly from a first operating speed to a second operating speed. The method further includes selectively positioning the second cowl between a first operational position and a second operational position to vary the area of the fan nozzle duct to facilitate improving engine efficiency at the second operating speed.

Abstract: A device for improving the effectiveness of the acoustic treatment(s) in an aircraft propulsion unit that includes a pipe (36) into which a gas stream flows, whereby the pipe (36) is delimited by two approximately concentric inside walls (40) and outside walls (42), wherein the device includes at least one partition (38, 38?) that extends over at least one portion of the length of the pipe (36) between the inside and outside walls (40, 42) and in that the profile of at least one wall (44, 46) of the at least one partition (38, 38?) is defined so as to increase the number of reflections of an acoustic beam on at least one of the inside and/or outside walls (40, 42) of the pipe and/or on at least one of the walls (44, 46) of the at least one partition (38, 38?).

Abstract: A system for damping thermo-acoustic instability in a combustor device for a gas turbine, the combustor device including at least one combustion chamber, in particular of an annular type, and at least one burner associated to the combustion chamber and mounted in a position corresponding to a front portion set upstream of the combustion chamber; the damping system including at least one Helmholtz resonator including a casing defining inside it a pre-set volume and a neck for hydraulic connection between the pre-set volume and the combustion chamber, the neck being connected to one side of the combustion chamber at a distance from the front upstream portion thereof provided with the at least one burner. The casing of the resonator includes structure which varies the pre-set volume within a pre-set range and structure which delivers a cooling fluid.

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 combustor includes at least one combustor liner defining a combustion chamber capable of directing combustion products toward a turbine. At least one combustor sleeve is located outside of the combustion chamber and is capable of reducing a magnitude of acoustic waves in the combustion chamber. The at least one combustor liner and the at least one combustor sleeve define at least one flow channel therebetween. Further, a combustor includes at least one combustor liner defining a combustion chamber capable of directing combustion products toward turbomachinery. At least one combustor sleeve disposed outside of the combustion chamber and is capable of controlling distribution of fluid flow in the combustor to modify a uniformity of the fluid flow to the combustion chamber.

Abstract: A high-bypass gas turbine engine includes a variable area fan nozzle with an acoustic system having an acoustic impedance.

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 combustor arrangement for a gas turbine engine (31) has a split line (42) and a plurality of burners (20,37) arranged in an annular ring (40). The burners (46) of the combustor on either side of the split line (42) have a separation distance of at least two times the average separation distance between burners (48) distant from the split line (42).

Abstract: A combustion device, such as an afterburner 4 for a gas turbine engine, comprises primary and secondary fuel delivery means 12, 20. Pressure fluctuations in the afterburner 4 are damped by reducing the supply of fuel through the primary fuel delivery means 12 during pressure peaks. This is achieved by means of a spill valve 18, controlled by a controller 22 in response to pressure in the afterburner 4 as sensed by a transducer 24. The spill valve 18 diverts a proportion of the fuel to the secondary fuel delivery means 20. The secondary fuel delivery means 20 is positioned upstream of the primary fuel delivery means 12 by a distance X, such that the secondary fuel delivery means serves to supply additional fuel to the combustion zone 26 during pressure troughs.

Abstract: A multi-layered honeycomb structure adapted to reduce and/or eliminate thermal deformation is disclosed herein. In some embodiments, walls of the honeycomb structure comprise a first layer, optionally, a second layer, and a core layer adjacent to the first layer or between the first and second layers. The first and second layers may be compositions of Inconel or other high strength materials. The core layer may be copper or another thermally conductive material. The core layer is adapted to transmit heat between a first region of a structure and a second region. In this manner, heat can be transferred from the heated region to an unheated region, thereby reducing the temperature difference between the regions and thus the amount of thermal deformation.

Abstract: A resonator for damping pressure waves or instabilities in a system supported by acoustic energy wherein the resonator adopts the use of counter-bored openings on the downstream side of the resonator within the flow path of the system. The resonator includes a first member and a second member where the second member includes a plurality of openings therethrough with each hole having a counter-bore on its upstream side. The first member has a size substantially smaller than the diameter of said flow path and a first plurality of openings therethrough. The openings are in fluid communication with the flow path. The second member has a size generally equal to the first member and a second plurality of openings therethrough, which openings are in fluid communication with the flow path.

Abstract: A combustor wall of a gas turbine engine is provided with an acoustic damper component. The component has a first metering passage, a first damping chamber, a first damping passage, a second damping chamber and a second damping passage. Air flows through the damper to be ejected into the combustion chamber from the second damping passage at a selected velocity and volumetric flow. The flow being sufficient to damp instabilities from the combustion process.

Abstract: An apparatus and method for suppressing dynamic instability in a bleed duct of a gas turbine engine includes a fan bypass duct configured to permit a flow of air through the gas turbine engine. The fan bypass duct defines a fan duct surface, and the bleed duct has an inlet in fluid communication with the fan bypass duct and a flow control valve having an opened position and a closed position. A flow diverter is positioned on the fan duct surface proximate the inlet of the bleed duct and diverts the flow of air from the inlet when the flow control valve is closed, while permitting a portion of the flow of air to enter the inlet when the flow control valve is opened.

Abstract: A method for operating a furnace with a multi-burner system for generating a hot gas is provided. The furnace has a combustion chamber with a plurality of burners, each having a fuel feed. The method includes regulating a first fuel feed to at least one of the plurality of burners as a function of pressure pulsations that occur in the combustion chamber so as to achieve a steady operation of the furnace. The multi-burner system may be a gas turbine, preferably of a power plant.

Abstract: In a bypass turbomachine, at least one air takeoff orifice is provided in the primary channel, said orifice leading to an air takeoff pipe housed inside an air inlet sleeve of the nacelle. The air takeoff pipe opens out in the vicinity of the pylon into two air diffusion pipes each secured to a respective maintenance cover, each air diffusion pipe opening out into an air injection pipe secured to a thrust reverser cover and itself opening out to the outside of the nacelle via the trailing edge thereof, each air injection pipe being suitable for uncoupling from the corresponding air diffusion pipe when the corresponding thrust reverser cover slides downstream.

Abstract: A gas turbine power generator plant, intended to reduce its noise by making small the intake and exhaust outlets of the cooing air channel of a case, comprises an engine core in which a turbine, a compressor and a generator are installed on the same axis, a combustor for burning air for combustion compressed by the compressor and supplying the air to the turbine, a radiator for cooling a coolant or a lubricant, a cooling fan for ventilating the radiator with cooling air, an electric power converter for converting electric power generated by the generator, and the case for housing these constituent elements. And, a combustion air channel passing the compressor, the combustor and the turbine and a cooling air channel passing the radiator, the cooling fan and the electric power converter are formed as mutually independent channels from intake to exhaust.

Abstract: An exhaust flow nozzle for use with a jet engine. The exhaust flow nozzle includes a beveled downstream edge portion that helps to direct noise generated by the jet engine upwardly away from a ground surface during takeoff and landing operations. One specific embodiment makes use of a linearly moveable flow altering component that is positioned outside the final flow nozzle. When the flow altering component is in its retracted position it has no effect on the exhaust flow leaving the flow nozzle. Alternative embodiments involve the use of varying edge shapes for the final flow nozzle.