Abstract: An exhaust duct of an aircraft engine includes an annular inlet conduit having an inlet central axis, and at least two outlet conduits in flow communication with the inlet conduit. The at least two outlet conduits are located non-parallel to the inlet central axis. Each of the at least two outlet conduits include an outlet port defining a distal end of each of the two outlet conduits. At least one of the outlet ports is non-circular in cross-sectional shape.

Abstract: An exhaust nozzle for a gas turbine engine. The exhaust nozzle comprises a nozzle body having an upstream end axially spaced from a downstream end with respect to an axial centerline of the nozzle body, and a plurality of chevrons circumferentially spaced apart and extending downstream from the downstream end. Each chevron includes an inner wall radially spaced from an outer wall, a root and a tip axially spaced from the root. At least one chevron of the plurality of chevrons includes a first segment extending axially downstream from the root, a second segment extending axially downstream from the first segment, and a third segment extending axially downstream from the second segment to the tip. The inner wall extends along the first segment, the second segment, and third the segment. The first segment is distinguishable from the second segment and the second segment is distinguishable from the third segment.

Abstract: An exterior nozzle member for a turbomachine extending along an axis oriented from upstream to downstream and comprising a primary duct configured to conduct a primary flow, an annular cavity of axis and a turbine, said exterior nozzle member comprising: an acoustic attenuation panel formed by an acoustic attenuation structure on which there are mounted an interior wall and an exterior wall facing the primary duct and the annular cavity, respectively, a nozzle flange connected to the panel and comprising an upstream end configured to be connected to a turbine casing flange and a sealing member connected to the panel and comprising an interior longitudinal branch extending radially inside the nozzle flange in order to thermally protect it from the primary flow, the nozzle flange being made of a first material and the sealing member being made of a second material, the first material having a coefficient of thermal expansion greater than that of the second material.

Abstract: The invention relates to an assembly for an aircraft turbomachine, comprising a central element (1) for the ejection of gas, and a connecting flange (9) interposed between, upstream, a gas outlet (22a) made of metal for a turbomachine, and, at the downstream end, the central element (1). The connecting flange comprises an annular portion (9a) and flexible tabs (11) having axially: a first end (111a) where the tab is connected to the said annular portion, and a second free end (111b), projecting radially inwardly with respect to the first end and to which said tab is attached with the central element (1).

Abstract: An assembly of a pylon and of a nacelle of an aircraft, wherein the nacelle has a structure, a visor, two centering systems and two tensioning systems. Each tensioning system comprises a first fitting, which is attached to the structure and has an engagement wall through which an orifice passes, a second fitting, which is attached to the visor, a threaded rod that is articulated on the second fitting and passes through the orifice, a spring element fitted on the threaded rod against the opposite face of the engagement wall from the second fitting, and an arresting means that blocks the spring element. Each centering system comprises a first fitting, which is attached to the structure and has a bore, and a second fitting, which is attached to the visor and bears a shaft that is accommodated in the bore.

Abstract: An emergency response drone having a multidirectional propulsion system and data capturing equipment and operatively associated computer system for providing information and situational awareness for emergency areas and related targets.

Abstract: A system and method for mitigating thermal loading between engine exhaust structures having different coefficients of thermal expansion. The engine exhaust structure comprises a metallic duct portion, a ceramic duct portion, and a double bipod fitting joining the metallic duct portion to the ceramic duct portion. The double bipod fitting is capable of flexing and taking up the thermal expansion differences between the joined metallic and ceramic ducts across the full temperature spectrum that an engine exhaust structure will experience in service.

Abstract: A combustor assembly for a gas turbine engine includes a dome having a forward surface and an inner surface. The forward surface and the inner surface of the dome at least partially define a slot. The combustor assembly also includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end. The forward end of the liner is positioned within the slot of the dome. The forward end of the liner includes an axial interface surface and a radial interface surface. The axial interface surface defines a radial gap with the inner surface of the dome and the radial interface surface defines an axial gap with the forward surface of the dome. At least one of the radial gap or the axial gap is less than about 0.150 inches during operating conditions of the combustor assembly to prevent an undesirable airflow.

Abstract: A heat exchanger includes a first channel with at least one first wall for porting a first fluid and a second channel with at least one second wall for porting a second fluid. The heat exchanger includes a barrier chamber located between the at least one first wall and the at least one second wall such that a rupture of one of the at least one first wall and the at least one second wall does not result in mixing of the first fluid and the second fluid. The heat exchanger includes a support member that extends between the at least one first wall and the at least one second wall.

Abstract: A waste heat capture system that can be used with at least a gas turbine engine. The system includes: an air scoop connected to a first component, the air scoop configured to direct air from a first duct to an interior compartment of the first component; a second duct along an exterior of the first component; and a thermoelectric material connected to an interior surface of the first component. The interior compartment of the first component is on a first side of the thermoelectric material and the exterior of the first component is on a second side of the thermoelectric material. The first duct is configured to receive air having a first temperature range, and the second duct is configured to receive air having a second temperature range, wherein the second temperature range is an order of magnitude higher than the first temperature range.

Abstract: A heat exchanger is configured to adjust a flow restriction of flow passages through the heat exchanger in response to changes in temperature of elements that define at least a portion of the flow passages. The elements include a first material having a first coefficient of thermal expansion, and a second material having a second coefficient of thermal expansion that is different from the first coefficient of thermal expansion.

Abstract: A flow distributor provided within an exhaust system for a combustion engine configured to generate an exhaust fluid stream. The flow distributor comprising an inlet, and a plate. The plate having at least one perforation defining an outlet, a first peak and a second peak spaced from the first peak.

Abstract: Embodiments are directed toward a monopropellant continuous detonation engine. In some embodiments, the continuous detonation engine includes an engine body, a monopropellant feed assembly, and a detonation initiator. The engine body defines a detonation wave channel. The monopropellant feed assembly delivers monopropellant from a monopropellant storage tank into the detonation wave channel. The detonation initiator initiates continuous detonation of the monopropellant in the detonation wave channel, preferably without a catalyst to promote decomposition of the monopropellant. Accordingly, specific impulse is increased compared to constant-pressure reaction thrusters that catalytically decompose the monopropellant with deflagration combustion.

Abstract: A system (10) includes a modular exhaust collector (30) configured to be arranged in a first orientation (50) or a second orientation (114, 116). The modular exhaust collector (30) is configured to receive an exhaust flow along an inlet axis (54), to direct the exhaust flow along a first direction (36) through an outlet (58) when in the first orientation (50), and to direct the exhaust flow along a second direction (114, 116) through the outlet (58) when in the second orientation (114, 116). The modular exhaust collector (30) includes an exhaust passage (64) to receive the exhaust flow, a plurality of compressor discharge (CD) ports (72), a plurality of flow ports (76), a bottom face (84) opposite the outlet (58) with a first drain (88), and a first side wall (82) with a second drain (88) between the bottom face (84) and the outlet (58).

Abstract: An electromechanical actuation motor includes a main housing, a series of stator windings, and a first circumferential row of fins. The series of stator windings is disposed inside of the main housing. The first circumferential row of fins is connected to and extends radially from an outer surface of the main housing. The main housing and the first circumferential row of fins are formed as a single piece of material via layer-by-layer additive manufacturing. Each fin of the first circumferential row of fins includes a first portion and a second portion. The first portion is integrally formed with and attached to the outer surface of the main housing at a first end of the first portion. The second portion is attached to and integrally formed with a second end of the first portion. The first portion intersects the second portion to form a T-shape.

Abstract: An air inlet device for an air inlet of an aircraft. The air inlet device has air guide arrangements. An air guide arrangement has an inlet plate, an outlet plate, and pivot plates arranged therebetween. The outlet plate and the pivot plates can be rotated about a central axis of the air inlet device by adjustment rings, so that the air guide arrangement is transitioned from a linear state to a curved state, and vice versa. The air inlet device allows for selective switching between a state with low radar signature and a state with high performance of an engine.

Abstract: A center plug for attenuating noise in a gas turbine engine includes an inner skin, the inner skin having a substantially cylindrical shape and extending along an axial centerline; a forward bulkhead disposed proximate a forward end of the inner skin, the forward bulkhead connected to and extending radially outward from the inner skin; an aft bulkhead disposed proximate an aft end of the inner skin, the aft bulkhead connected to and extending radially outward from the inner skin; and an intermediate bulkhead, the intermediate bulkhead positioned intermediate the forward bulkhead and the aft bulkhead, the intermediate bulkhead spaced a distance from the inner skin and extending radially outward from the inner skin, the distance defining a radially inner portion of the intermediate bulkhead and an annular opening.

Abstract: A turbine assembly adapted for use in a gas turbine engine includes a seal segment and a carrier. The seal segment includes a shroud wall that extends circumferentially partway around an axis and a mount post that extends radially outward away from the shroud wall. The carrier is shaped to define a channel that receives at least a portion of the mount post. A mounting insert is used to urge the carrier and mount post toward one another.

Abstract: A nozzle having a forward portion skewed downwards and an aft portion translated downwards provides sufficient clearance between the nozzle and the heat shield structure to prevent contact in the event of large deflections (e.g., as associated with a fan blade Out (FBO) condition). Such large deflections must be accounted for to meet federal aviation regulations and gain airplane CFR 14 Part 25 Certification.

Abstract: A combustor and a gas turbine capable of stably supporting a fuel peg and guiding the flow of air are provided. The combustor may include: a plurality of nozzles configured to eject fuel and air; a flow passage configured to guide a flow of air to be drawn into the nozzles; a plurality of fuel pegs configured to protrude into the flow passage, each of the plurality of fuel pegs including an injection hole through which fuel is discharged; and a peg support configured to be coupled to the fuel pegs to support the fuel pegs.

Abstract: A fan assembly for use in a gas turbine engine of an aircraft includes a fan disk having a number of fan blades and a windage shield coupled to the fan disk to move therewith. The fan assembly supplies air for use in the engine. The windage shield rotates with the fan disk during operation of the gas turbine engine and directs air supplied by the fan blade.

Abstract: A gas turbine engine includes a core engine that has at least a compressor section, a combustor section and a turbine section disposed along a central axis. A fan is coupled to be driven by the turbine section. A fan nozzle is aft of the fan and defines an exit area. The fan nozzle has a body with an airfoil cross-section geometry. The body includes a wall that has a controlled mechanical property distribution that varies in material macro- or micro-structure by location on the wall in accordance with a desired flutter characteristic at the location.

Abstract: A nacelle exhaust nozzle having a deployable noise-reducing component is described. The noise-reducing component includes an annular perforated sleeve coinciding with the inner nacelle loft line and circumscribing a mixed exhaust gas flow exiting the nacelle. The annular perforated sleeve is radially spaced apart from of a displaceable wall of an inflatable envelope that is displaceable between a deployed position, wherein noise-reduction is active, and a retracted position, wherein noise-reduction is inactive. When the inflatable envelope is pressurized, portions of the displaceable wall project through openings in the perforated sleeve and into the exhaust gas flow to form a rough surface at the loft line which causes a reduction in noise level. The portions of the displaceable wall that project through the openings in the perforated sleeve when the inflatable enveloped is pressurized include a plurality of dimples formed on the inner wall and forming the rough surface.

Abstract: A turboprop assembly includes a nacelle with a main nacelle body and a nacelle extension coupled to the main nacelle body. The nacelle extension has a wall that defines an air intake port. The air intake port has a non-circular, non-rectangular, and non-oval shaped perimeter extending in two of three dimensions.

Abstract: A turbofan exhaust nozzle includes: a conical inner shell disposed coaxially inside a surrounding outer shell to define an annular flow duct therebetween terminating in an outlet at a trailing edge of the outer shell and the inner shell terminates in a trailing edge aft of the outlet; and a pylon interrupting circumferentially the duct, wherein at least a portion of the trailing edge of the outer shell is vertically non-coplanar.

Abstract: The invention concerns a combustion chamber (10) comprising a neck (15) downstream of the injection (11) of gases, and downstream of this neck a divergent section (20) whereof the outer face of the wall (30), when in operation, is cooled by a cooling system using a cryogenic product and surrounding this outer face. This divergent section (20), on the inner face (32) of its wall (30), comprises a coating (40) acting as temperature compensator so that the temperature of the inner face (42) of the coating (40) is higher than the condensation temperature of the combustion gases on this inner face (42) under operating conditions, such that no condensation is formed on this inner face (42).

Abstract: A gas turbine engine includes a core defining an engine central longitudinal axis. An inner-fixed structure is radially outward of the core. A core cowl extends from the inner-fixed structure to a trailing edge. A thrust reverser is spaced radially outward of the core cowl to define a fan flow path. A vent has a core cowl inner surface formed as part of the core cowl and a vent inner surface that is spaced radially inward of the core cowl inner surface to define a vent flow path. A reflex member extends from a trailing edge of the core nacelle to impede mixing of the fan flow path and the vent flow path.

Abstract: A turbine exhaust section comprises a turbine exhaust case having radially outer and inner ducts defining therebetween an annular exhaust portion for the hot exhaust gases, and an exhaust mixer projecting axially rearwardly from the turbine exhaust case for mixing the hot exhaust gases with a cooler bypass duct flow. The upstream end of the exhaust mixer surrounds a downstream end of the outer duct and defines therewith an axially extending overlap joint with a radial play between the outer duct and the exhaust mixer. A sliding attachment in the radial direction is provided between the outer duct and the exhaust mixer for accommodating differences in thermal growth during engine operation. The sliding attachment includes a circumferential array of sliding guides extending radially through the axially extending overlap joint. A resilient sealing ring seals the radial play between the outer duct and the exhaust mixer at the overlap joint.

Abstract: A nozzle arrangement is disclosed herein for use with a supersonic jet engine that is configured to produce a plume of exhaust gases. The nozzle arrangement includes, but is not limited to, a nozzle having a trailing edge and a plug body partially positioned within the nozzle. The plug body has an expansion surface and a compression surface downstream of the expansion surface. A protruding portion of the plug body extends downstream of the trailing edge for a length greater than a conventional plug body length. The plug body is configured to shape the exhaust gases to flow substantially parallel to a free stream of air flowing off of the trailing edge of the nozzle and to cause the plume of exhaust gases to isentropically turn the free stream of air to move in a direction parallel to a longitudinal axis of the plug body.

Abstract: A quick disengaging field joint connects a first component of an exhaust system of a gas turbine engine to a second component of the exhaust system. The field joint includes a pair of opposed stepped liners connected via exterior-facing connecting flanges. The field joints can be disassembled entirely from outside the exhaust housing without requiring access to the interior of the exhaust housing.

Abstract: An inlet assembly for a turboprop engine system is provided. The inlet assembly includes an inlet plenum defining first inlet configured to receive air from the atmosphere; an expansion plenum coupled to the inlet plenum and configured to receive the air from the inlet plenum, the inlet plenum and the expansion plenum defining a flow path for the air with a bend of at least 90°; and a second inlet coupled to the expansion plenum and configured to receive the air from the expansion plenum.

Abstract: The construction of an internal/external single expansion ramp nozzle (nxSERN), and method of designing the same, is provided. Initial design parameters for primary stream construction are selected and additional parameters are determined by isentropic relations using the selected design parameters and Prandtl-Meyer function. The nozzle throat input and output angles are determined and used to define an initial portion of the nozzle primary stream lower expansion surface. The nozzle primary stream upper expansion surface and an aft portion of the primary stream lower expansion surface are defined using a method of characteristics. Initial and aft portions of the primary stream lower expansion surface are then connected by a straight line to define the primary stream nozzle.

Abstract: Exhaust centerbody (12?) for a turbine engine, comprising a truncated downstream part (26), which is connected to an upstream part (24) by an annular ridge (22) marking a discontinuity between the outer surfaces of the upstream and downstream parts, the outer surface of the downstream part having a substantially conical general shape, of which the tip (29) is oriented downstream and is positioned in the region of the axis A, the axial half-section of this outer surface defining a line of which the upstream end part is substantially tangential to a straight line (28) passing through the above-mentioned ridge and forming a non-zero angle ? with a tangent (30) to the outer surface of the upstream part, in the region of the ridge, and of which the downstream end part is substantially tangential to a straight line (32) passing through the above-mentioned tip and forming a non-zero angle ? with the axis A.

Abstract: A nozzle arrangement for an engine that is operable in both an air-breathing mode, in which the engine combusts air taken in from atmosphere with hydrogen from a store thereof, and in a rocket mode, in which the engine combusts oxygen from a store thereof with hydrogen from the store thereof. The nozzle arrangement may include a rocket combustion chamber fluidly coupled by a rocket throat to a rocket nozzle. The rocket nozzle includes a first portion adjacent the rocket throat and a second portion remote from the rocket throat and axially moveable relative to the first portion between a rocket position in which they form a substantially contiguous rocket nozzle and an air-breathing position in which they overlap to define an annular throat therebetween. The nozzle arrangement may also include at least one air-breathing combustion chamber arranged to be fluidly coupled to the annular throat when the first and second portions of the nozzle are in the air-breathing position.

Abstract: An exhaust includes a wall that has a first composite material having a first coefficient of thermal expansion and a second composite material having a second coefficient of the thermal expansion that is less than the first coefficient of thermal expansion.

Abstract: A mixer for mixing flows in a turbofan engine is provided. The mixer includes a plurality of chevron lobes, each of the plurality of lobes comprising a crown, a keel, a first trailing edge, a second trailing edge, and a first transverse edge extending between first trailing edge and second trailing edge, said mixer configured to receive two separate incoming exhaust flows and mix the two flows into at least one rotational exhaust flow that is ejected out at least one of said first trailing edge and said second trailing edge.

Abstract: An aircraft gas turbine engine has a row of FLADE fan blades disposed radially outwardly of and drivingly connected to a fan in the engine's fan section. The FLADE fan blades extend across a FLADE duct circumscribing the fan section. A two dimensional air discharge passage is in fluid flow communication with the FLADE duct and with FLADE air upstream and downstream discharge slots in a divergent flap of a two dimensional exhaust nozzle. A valve fully closes the upstream slot when the downstream slot is fully opened and fully opens the upstream slot when the downstream slot is fully closed. The upstream and downstream slots may be located upstream and downstream respectfully of a nozzle discharge area in the nozzle. A sliding deck slides aft or down to open upstream slot and close downstream slot and slides forward or up to close upstream slot and open downstream slot.

Abstract: A turbine exhaust case (TEC) of a turbofan aeroengine includes a mixer for mixing exhaust gases with a bypass air stream, the TEC comprising an annular hub and an annular shroud with the mixer attached to a downstream end of the shroud, the shroud and the mixer surrounding the hub to form an annular exhaust gas duct positioned radially therebetween, a plurality of deswirling struts circumferentially spaced apart with respect to a central axis of the TEC and located within an axial length of the mixer between an upstream end and a downstream end of the mixer, the deswirling struts each having a cambered profile and extending radially across the annular exhaust gas duct and interconnecting the mixer and the hub.

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

Abstract: A gas turbine engine, especially a turbojet or turbofan for an aircraft, includes a combustion section and an exhaust section produces an exhaust gas. The emission products can be reduced by using a gas treatment device with a reactor and an injection device. The reactor produces a nitrogeneous substance such as ammonia. This reducing agent is injected in the gas phase in the exhaust section of the gas turbine engine. By selective non catalytic reaction the Nitrogen oxides of the exhaust gas can be reduced. Furthermore the injection device is arranged such that the temperature window of this reduction process is ranging from 850° C. to 1100° C. The reactor for producing gaseous ammonia can either be a thermal reactor or a high pressure plasma device.

Abstract: Disclosed is a turbofan engine for an aircraft. An inner fan cowl of the turbofan engine bounds an annular cold-stream duct proximate to the engine's central hot-stream generator, and an outer fan cowl bounds the annular cold-stream duct proximate to an outer nacelle cowl. An annular boss is provided along the inner fan cowl, with the annular boss having a rounded cross section that projects with respect to a smooth shape configuration of the inner fan cowl. The annular boss is configured to locally change speed of a cold stream flow in the annular cold-stream duct from a subsonic range to a supersonic range and to originate a first shockwave characteristic in a succession of shockwave characteristics, with the first shockwave characteristic being inclined from its origin toward the rear portion of the turbofan engine.

Abstract: A stage of a space launcher including an engine core, a tank secured above the engine core, and a nozzle including a first deployable nozzle section secured below the engine core, the nozzle further including a second and a third section configured to be situated extending from each other when the nozzle is in a deployed propulsion configuration. The stage further includes a structure for temporarily supporting nozzle sections mounted on the tank, configured to change the nozzle to a configuration for accessing the engine core wherein the nozzle sections are inserted into each other, with the third section bearing against a lower support of the temporary structure.

Abstract: The invention relates to a method for controlling a plurality of actuators (9a, 9b, 9c) for a movable thrust reverser cowl (1), characterized remarkable in that the position deviation between adjacent actuators is measured in real time, and in that the speed profile of the actuator or actuators in question is adjusted by a position deviation exceeding a predetermined threshold.

Abstract: A turbine engine convergent-divergent nozzle including an annular central element and an annular cap arranged coaxially around the central element to co-operate with the central element to define an annular flow channel for a gas stream from the engine. Between a throat section and an ejection section of the nozzle, the central element and the cap present respective internal profiles in longitudinal section that are modeled by curves having respective radii of curvature that are identical in absolute value.

Abstract: An engine apparatus is provided for a flying object, the engine apparatus including a combustion chamber having a throat region and a nozzle region, the nozzle region having a nozzle wall, wherein the nozzle region expands from the throat region towards an exit end relative to a combustion chamber axis, wherein the nozzle region has associated therewith a skirt having a skirt wall, the skirt being positioned downstream relative to the exit end and surrounding the exit end of the nozzle region, and wherein the skirt wall is at an acute angle away from the combustion chamber axis with respect to the nozzle wall, at least at the exit end of the nozzle region.

Abstract: A combined flow straightener and mixer is disclosed as well as a burner for a combustion chamber of a gas turbine having such a mixing device. At least two streamlined bodies are arranged in a structure comprising the side walls of the mixer. The leading edge area of each streamlined body has a profile, which is oriented parallel to a main flow direction prevailing at the leading edge position, and with reference to a central plane of the streamlined bodies, the trailing edges are provided with at least two lobes in opposite transverse directions. The periodic deflections forming the lobes from two adjacent streamlined bodies are out of phase.

Abstract: An engine propulsion system is configured to utilize bursts of media in order to create mechanical energy. The engine propulsion system includes at least one cannon, wherein each cannon is configured to displace the media and further includes a firing pin casing configured to accommodate a firing pin. The firing pin is configured to create the mechanical energy when moved thus allowing the media to exit the cannon.

Abstract: Techniques, devices and systems are disclosed for implementing acoustically triggered propulsion of nano- and micro-scale structures. In one aspect, an ultrasound responsive propulsion device includes a tube that includes one or more layers including an inner layer having an electrostatic surface, and an ultrasound-responsive substance coupled to the inner layer and configured to form gaseous bubbles in a fluid in response to an ultrasound pulse, in which the bubbles exit the tube to propel the tube to move in the fluid.

Abstract: A turbofan engine control system and method includes a core nacelle housing (12), a compressor and a turbine. A turbofan is arranged upstream from the core nacelle and is surrounded by a fan nacelle (34). A bypass flow path (39) is arranged downstream from the turbofan between the core and fan nacelles. The bypass flow path includes a nozzle exit area (40). A controller (50) detects at least one of a take-off condition and a landing condition. The controller changes effectively the nozzle exit area to achieve a thrust vector in response to the take-off and landing conditions.

Abstract: A seal assembly for a turbojet nacelle includes a first end fastened to a first structure, and a second end contacting against a bearing zone of a second structure. The seal assembly further includes a plurality of adjacent blades arranged along the first end and extending longitudinally and perpendicularly thereto. In particular, a portion of the blades has an accordion-shaped structure.