Abstract: In a convergent-divergent flap pair for a turbojet engine nozzle of the variable-geometry convergent-divergent type, the convergent flap and the divergent flap including respective ducts for circulation of cooling air connected to one another to allow cooling of each of the flaps. The duct of the divergent flap includes an impingement cooling cavity defined by two walls of lateral ends provided with air passage orifices through which the impingement cooling cavity opens towards the outside.

Abstract: An apparatus for sealing such a gap may be a plunger seal which may include a flap arm comprising resilient sheets and a wall arm comprising resilient sheets. A proximal end portion of the flap arm may include a closeout seal coupled to the flap. A proximal end portion of the wall arm may include a plunger having a geometry corresponding to that of the closeout seal so the plunger may be matingly received by the closeout seal. When positioned in a gap, the plunger seal may exert a force to urge the flap arm towards the flap and to urge the wall arm and resilient sheet towards the structure to seal the gap.

Abstract: In some embodiments, apparatuses are provided herein useful to sealing a gap between a movable flap and stationary structure, such as a gap between a gas turbine engine nozzle flap and sidewall. An apparatus for sealing such a gap may be a plunger seal which may include a plurality of plunger segments connected together using at least one flexure. When positioned in the gap, the flexures within the plunger segments pivot allowing rotation of each of the plurality of plunger segments about their respective pivot point, such that the plunger assembly seals and contours against the movable surface.

Abstract: Embodiments are directed to boosting aircraft engine performance for takeoff and critical mission segments by reducing airflow used for cooling exhaust gases. The airflow is reduced by stopping an accessory blower or by closing an external air vent Eliminating the cooling airflow to the exhaust has the effect of lowering the backpressure on the engine, which thereby increases maximum engine power.

Abstract: A system and method for vectoring the thrust of a supersonic, air-breathing engine. A thrust vectoring mechanism uses two asymmetrically staggered ramps; one placed at the throat, the other positioned at the exit lip of the nozzle of the engine to re-direct exhaust flow off-axis with the nozzle.

Abstract: One embodiment includes a multistage infrared suppression exhaust system for an aircraft, including: a stage one including a first exhaust conduit to receive a first exhaust air flow at a first temperature-pressure product T1P1, a second exhaust conduit to receive a second exhaust air flow at a second temperature-pressure product T2P2, and a flow integrator mechanically configured to mix the first exhaust air flow with the second exhaust air flow in an integration chamber while preventing back flow into the second exhaust conduit; and a stage two including a stage two cooling airflow to cool the mixed first and second exhaust air flows.

Abstract: A combustor of a liquid rocket engine includes a nozzle unit including a regenerative cooling channel, in which the nozzle unit includes a fuel manifold outer shell, a combustor inner shell, and a combustor outer shell having a downward channel inlet, and the combustor includes a fuel inlet connected to a nozzle neck of the nozzle unit, a fuel manifold formed between the fuel manifold outer shell and the combustor outer shell, and in which fuel introduced from the fuel inlet flows, a downward channel connected in communication with the fuel manifold through the downward channel inlet, and extending in a downward direction from an upper portion of the combustor, a diverting manifold provided at a distal end of the nozzle unit and connected in communication with the downward channel, and an upward channel connected in communication with the diverting manifold and extending in an upward direction of the combustor.

Abstract: A thrust reverser for a gas turbine engine includes a frame, a first reverser door pivotally mounted to the frame, and a second reverser door pivotally mounted to the first reverser door.

Abstract: A dual flow path exhaust assembly for use with a combined turbofan and ramjet engine includes a turbofan engine exhaust duct, a ramjet engine exhaust duct, a combined outlet, and door configured to move between an open position and a closed position to selectively isolate the turbofan engine exhaust duct from the combined outlet.

Abstract: An exhaust nozzle for use with a gas turbine engine includes an outer shroud, an inner plug spaced radially apart from the outer shroud, and at least one support vane that is coupled to the outer shroud. The outer shroud and the inner plug cooperate to provide an exhaust nozzle flow path therebetween. The at least one support vane interconnects the outer shroud and the inner plug to support the inner plug in the exhaust nozzle flow path.

Abstract: An ejector assembly includes a primary nozzle in fluid communication with a primary fluid inlet, a secondary nozzle in fluid communication with a secondary fluid inlet, the primary nozzle being oriented concentrically within the secondary nozzle and the secondary nozzle having a venturi downstream of the primary nozzle, and the primary nozzle having a variable cross-sectional area.

Abstract: Embodiments are directed to boosting aircraft engine performance for takeoff and critical mission segments by reducing airflow used for cooling exhaust gases. The airflow is reduced by stopping an accessory blower or by closing an external air vent. Eliminating the cooling airflow to the exhaust has the effect of lowering the backpressure on the engine, which thereby increases maximum engine power.

Abstract: Exhaust redirecting devices are described that are suitable for use in tilt rotor aircraft. Such devices are constructed of light weight material and permit redirection of exhaust gases from turbojet engines of tilt rotor aircraft as nacelles of the aircraft transition between vertical and horizontal flight. Use of a controller permits coordination between exhaust redirection and nacelle position.

Abstract: A turbine exhaust assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a turbine exhaust case comprised of CMC material and attachable to a turbine case, a tail cone comprised of CMC material that has a leading edge and a trailing edge, and an exhaust mixer comprised of CMC material and coupled to the turbine exhaust case. The exhaust mixer has a plurality of lobes arranged about the tail cone to define an exhaust flow path. A plurality of struts extend from the tail cone to support the exhaust mixer at a location aft of the leading edge of the tail cone. A method of assembling a propulsion system is also disclosed.

Abstract: A louvre offtake arrangement for a gas turbine engine includes a first duct, where a primary flow flows; a second duct, defining an offtake, connected to the first duct at an inlet; and a plurality of louvres arranged at the inlet. At least one louvre protrudes from the second duct into the first duct to divert part of the primary flow towards the second duct.

Abstract: An aircraft tailcone comprising a tailcone fuselage, a turbomachine, for example an APU, housed inside the tailcone, a ram air intake on the tailcone fuselage for the ingestion of ram air towards the interior of the turbomachine compartment, an inlet flap operable reciprocately from an open position in which ingestion of ram air is allowed, and a closed position in which ram air ingestion is prevented. The ram air intake extends annularly along a perimeter of the tailcone fuselage, and the inlet flap is configured such in its closed position that a surface of the inlet flap is substantially flush with the tailcone fuselage.

Abstract: An apparatus comprising a housing, seal, and energy storing device. The housing is coupled to a structure in an exhaust system of an aircraft. The structure is positioned relative to a surface within the exhaust system such that a gap is present between the surface and structure. The seal has an end positioned in contact with the surface to reduce a flow of exhaust through the gap and is coupled to the housing such that the seal extends within the housing. The energy storing device is coupled to the housing and engaged with the seal to allow the seal to translate in a first direction relative to the housing when the gap increases and in a second direction opposite the first direction relative to the housing when the gap decreases such that the seal continues to reduce the flow of exhaust through the gap as the gap changes in size.

Abstract: A oral irrigator device (10) for cleaning teeth, including a system (14) for producing a stream of liquid/air bursts and a nozzle (16) connected to the system, and through which the stream of liquid/air bursts are directed. The nozzle can include a guidance tip (26) including a base portion (30), a tip portion (28), and an orifice (31) through which the stream of liquid/air bursts exits at a flow output pattern, and a flow output pattern selection mechanism (90, 90?) connected to or positioned within the guidance tip. A flow output pattern selection mechanism can be configured to be selectively actuated and to vary, upon the actuation the flow output pattern between a first flow output pattern to at least a second flow output pattern.

Abstract: A gas turbine engine comprises a main gas path having an inner flow boundary wall and an outer flow boundary wall. A turbine exhaust case inner body defines a portion of the inner flow boundary wall of the main gas path. A lobed exhaust mixer defines a portion of the outer flow boundary wall of the main gas path. A stiffener ring is interconnected to at least a number lobes of the lobed exhaust mixer by a plurality of circumferentially spaced-apart struts extending through the main gas path. The stiffener ring is attached to the turbine exhaust case inner body by flexible features, such as circumferentially spaced-apart spring blades.

Abstract: A nozzle for a gas turbine engine. An array of convergent petals hingedly coupled to a nozzle exit of fixed diameter. An array of divergent petals, each divergent petal hingedly coupled to one of the array of convergent petals. A cam surface associated with the array of convergent petals. A pivot point coupled to the array of divergent petals by a first linkage and to a fixed point by a second linkage. A cam follower coupled to the second linkage by a third linkage, the cam follower arranged to abut and travel in contact with the cam surface. An actuator coupled to the cam follower and arranged to translate the cam follower along the cam surface to move the convergent and divergent petals.

Abstract: Components for gas turbine engines are provided. The components include a hybrid skin core cooling cavity defined by a cold wall and a hot wall, wherein the hot wall is exposed to an exterior environment of the component, a hybrid resupply hole formed in the cold wall and fluidly connecting a cold cavity and the hybrid skin core cooling cavity, and a resupply cover located on the cold wall and within the hybrid skin core cooling cavity and positioned relative to the hybrid resupply hole to shield resupply air injected from the cold cavity into the hybrid skin core cooling cavity to minimize losses as the resupply air mixes with air flowing within the hybrid skin core cooling cavity.

Abstract: According to one aspect, an infrared suppression system that utilizes a primary airflow and a secondary air flow for a gas turbine engine is provided. The infrared suppression system includes an exit flap configured to control a flow of exhaust air out of the gas turbine engine and a secondary flow door that is configured to selectively control the secondary air flow into the gas turbine engine. The mixture of the secondary air flow and the primary air flow from the gas turbine engine suppresses an infrared signature produced by the gas turbine engine.

Abstract: An exhaust arrangement is for an aircraft having sensing equipment and a front prop engine. An extension mounts on the underside along an edge of the fuselage to direct exhaust from the engine's exhaust pipes to a point to or past the sensing equipment typically exposed at the mid-section of the aircraft. An extension inlet mounts adjacent the pipe and defines a funnel or conical orifice not directly affixed to the pipe. One or more conductors for the extension then extend from the inlet to convey the exhaust to an outlet of the extension. Preferably, multiple conductors can be used with an expansion joint provided between them. The outlet preferably diverts exhaust away from the tail end of the aircraft. To mount the extension to the aircraft, several arrangements of supports including rods, lugs, angles, and the like hold the extension in vertical, lateral, and axial directions to the aircraft.

Abstract: One embodiment includes a fuel injector. The fuel injector assembly comprises a conduit for conveying fuel from a fuel inlet to a nozzle. The conduit is located in a support, with the conduit, the nozzle, and the support being a single unitary piece. A thermally compliant feature is located at the nozzle which allows the fuel injector to adjust for differential thermal expansion.

Abstract: A strut assembly for a gas turbine engine includes an outer structural case. The outer structural case includes a first mounting pad for mounting a first strut and a second mounting pad for mounting a second strut. The outer structural case further includes a case ligament extending between the first mounting pad and the second mounting pad in a substantially straight direction to reduce an amount of bending stress on the outer structural case.

Abstract: An engine component for a gas turbine engine includes a film-cooled recess comprising a contoured portion defining a step. A hot surface facing hot combustion gas and a cooling surface facing a cooling fluid flow are fluidly coupled by a passage through the engine component. The passage further comprises an inlet in the cooling surface and an outlet in the step. The inlet, passage and outlet are oriented such that the cooling fluid flowing through the passage and exiting the outlet diffuses within the contoured portion prevents premature mixing out with the hot fluid flow.

Abstract: In a gas turbine combustor having an inner and outer liner defining an annular combustion chamber, at least an annular scoop ring provided on each inner and outer combustor liner. The annular scoop ring includes a solid radial inner base provided with bores defined therein and communicating with the combustion chamber to form air dilution inlets. The scoop ring has a radial outer portion in the form of a C-shaped scoop open to receive high velocity annular air flow. The bores of the inlets communicating with the scoop portion to direct the air flow into the combustion chamber whereby the bores of the inlets form jet nozzles to generate air jet penetration and direction within the combustion chamber.

Abstract: A nacelle for a gas turbine engine includes a ring shaped body defining a center axis and having a radially outward surface and a radially inward surface. An aft portion of the radially inward surface includes an axially extending convergent-divergent exit nozzle. An axially extending secondary duct passes through the nacelle in the convergent-divergent exit nozzle. The axially extending secondary duct includes an inlet at a convergent portion of the convergent-divergent exit nozzle and an outlet at a divergent portion of the convergent-divergent exit nozzle.

Abstract: An exhaust system for reducing infrared emissions of a rotary wing aircraft includes a manifold; an opening in the manifold, the opening configured to face upwards and away from the rotary wing aircraft; and a chimney including a wall positioned about the opening, the chimney configured to eject an emission of intermixed secondary air and engine exhaust upwards and away from the rotary wing aircraft.

Abstract: A thrust reverser assembly for a gas turbine engine including a core engine, a nacelle surrounding at least a portion of the core engine defining a bypass duct between the nacelle and the core engine where an outer door movable between a stowed position and a deployed position extends outwards from the nacelle and a blocker door movable between a stowed position and an deployed position extends into an airflow conduit defined by the bypass duct to deflect air outwards.

Abstract: The present disclosure concerns a nacelle for an aircraft turbojet engine that includes a thrust reverser having a cowl displaceable between a deployed position and a stowed position, a fixed structure, and a secondary nozzle. The nozzle includes panels movable in rotation, and hydraulic actuators for sequenced actuation of the cowl and panels. The actuators include a body mounted on the fixed structure, and a rod mounted in the body and is linked to the cowl and panels via a device for transmitting movement of the actuator to the panel. An actuation system having a hydraulic control device controls the displacement of the rod of the actuator. The actuator further includes a device for alternately displacing the rod in translation relative to the body to cause translation of the cowl, or in rotation relative to the body to cause rotation of the nozzle panel.

Abstract: A nozzle assembly for a dual gas turbine engine propulsion system includes a housing mountable proximate to a first bypass passage of a first gas turbine engine and a second bypass passage of a second gas turbine engine, first and second upper doors, and first and second lower doors. Each of the first and second upper doors and the first and second lower doors are pivotally mounted to the housing for movement between a stowed position and a deployed position in which airflow through the first and second bypass passages is redirected relative to respective centerline axes of the first and second gas turbine engines.

Abstract: The present disclosure provides a turbojet engine nacelle having a downstream section including a cowl which is ended with a hinged flap constituting a nozzle. In particular, the hinged flap is hingedly and translatably hinged to the cowl.

Abstract: An exhaust arrangement is for an aircraft having sensing equipment and a front prop engine. An extension mounts on the underside along an edge of the fuselage to direct exhaust from the engine's exhaust pipes to a point to or past the sensing equipment typically exposed at the mid-section of the aircraft. An extension inlet mounts adjacent the pipe and defines a funnel or conical orifice not directly affixed to the pipe. One or more conductors for the extension then extend from the inlet to convey the exhaust to an outlet of the extension. Preferably, multiple conductors can be used with an expansion joint provided between them. The outlet preferably diverts exhaust away from the tail end of the aircraft. To mount the extension to the aircraft, several arrangements of supports including rods, lugs, angles, and the like hold the extension in vertical, lateral, and axial directions to the aircraft.

Abstract: Methods and apparatus for sealing variable area fan nozzles of jet engines are disclosed. An apparatus in accordance with the teachings of this disclosure includes a frame and a seal to be coupled to the frame. The seal to enclose petals of a variable area fan nozzle to substantially prevent airflow between the petals.

Abstract: The exit area of a nozzle assembly is varied by translating a ring assembly located at a rear of the engine nacelle. The ring may be axially translatable along the axis of the engine. As the ring translates, the trailing edge of the ring defines a variable nozzle exit area. Translation of the ring creates an upstream exit at a leading edge of the ring assembly. The upstream exit can be used to bleed or otherwise spill flow excess from the engine bypass duct. As the engine operates in various flight conditions, the ring can be translated to obtain lower fan pressure ratios and thereby increase the efficiency of the engine. Fairings partially enclose actuator components for reduced drag.

Abstract: A gas turbine engine nozzle includes a flap movable relative to a structure. A seal assembly is supported by one of the structure and the flap and includes a seal hinged about an axis. The seal has a sealing profile engaging a seal land of the other of the structure and the flap. A biasing member is configured to urge the hinged seal toward the seal land. A method of sealing a nozzle flap includes supporting a seal relative to a structure along an axis. The seal is urged toward a nozzle flap. The seal rotates about the axis to maintain engagement between the seal and the nozzle flap in response to the urging step.

Abstract: A conduit through which hot combustion gases pass in a gas turbine engine. The conduit includes a wall structure having a central axis and defining an inner volume of the conduit for permitting hot combustion gases to pass through the conduit. The wall structure includes a forward end, an aft end axially spaced from the forward end, the aft end defining a combustion gas outlet for the hot combustion gases passing through the conduit, and a plurality of generally radially outwardly extending protuberances formed in the wall structure. The protuberances each include at least one cooling fluid passage formed therethrough for permitting cooling fluid to enter the inner volume. At least one of the protuberances is shaped so as to cause cooling fluid passing through it to diverge in a circumferential direction as it enters into the inner volume.

Abstract: An exhaust liner assembly including an interface between moveable parts includes a seal liner having a plurality of slots. The size of each slot varies with the circumferential position of the liner assembly to provide a stable cavity pressure within the interface to create a stable barrier about the entire circumference of the exhaust liner assembly.

Abstract: Provided is an electro-hydraulic control system (10) for controlling the movement of a third structure (18) of an engine using a power actuator (26) mountable to the third structure (18), a telescoping fluid porting tube (24) for porting a fluid to the power actuator (26) and a control valve (22) for controlling the flow of movement to effect the movement of the third structure (18). In this way, the third structure (18), such as a nozzle slat of an engine, may be moved via a compact and lightweight electro-hydraulic system.

Abstract: A turbine exhaust case (TEC) of a turbofan aeroengine includes a mixer in an annular wavy configuration having alternate crests and valleys, and a plurality of radial deswirling struts interconnecting an annular hub positioned within the mixer and selected valleys of the mixer. A plurality of secondary trailing edge struts smaller than the deswirling struts in average chord dimensions interconnect the annular hub and other selected valleys free of connection with the deswirling struts.

Abstract: In various embodiments, center body assemblies are described herein. The center body assembly may include an attachment ring, a center body, and a cavity structure. The center body may be configured to be coupled to the attachment ring. The cavity structure may be configured to be coupled to the attachment ring. The cavity structure may be configured to be separated from the center body by a gap. The center body and the cavity structure may be configured to expand independently of each other in response to an increase in temperature. In various embodiments, the cavity structure may comprise a baffle. In various embodiments, the center body may comprise a baffle.

Abstract: An apparatus for flaring an outer tube of a double-walled tube is provided. The apparatus includes a frame, a flaring tool and a clamping arrangement. The frame includes a first portion and a second portion that is laterally disposed in relation to the first portion. The clamping arrangement is disposed on the first portion. The clamping arrangement is configured to receive the double-walled tube therein and position an end of the double-walled tube to face the second portion. The clamping arrangement is operable to axially restrict the double-walled tube. The flaring tool is disposed in an axially movable engagement with respect to the second portion. The flaring tool includes a flaring tip that defines a tapered portion. The tapered portion is configured to face the end of the double-walled tube and flare an end of the outer tube.

Abstract: A variable area mechanism and methods are disclosed. A movable surface comprising a contoured interface surface is operable to rotate about an off-body axis-of-rotation such that the movable surface expands from and retracts into an angle notched surface.

Abstract: A fan housing has a radially outer surface. A pair of radially intermediate surfaces includes a first surface spaced towards a first axial end, which has an inner bore positioned radially outward of an inner bore of a second surface. Vanes extend between the radially outer surface and an outer periphery of the pair of radially intermediate surfaces. A bearing bore extends radially inwardly of the first surface. A mounting lug nub is formed on an outer periphery of the radially outer surface. A first distance is defined between the first axial one end of the fan housing and an opposed axial end. A second distance is defined between the first axial end and a center of the mounting lug. A ratio of the first distance to the second distance is between 2.13 and 2.11. An intermediate part, a ram air fan and a method are also disclosed.

Abstract: A power plant includes a gas turbine unit adapted to feed flue gases into a diverter where they are divided into a recirculated flow that is fed into a mixer together with fresh air to form a mixture that is fed to a gas turbine unit compressor inlet, and a discharged flow, that is fed into a CO2 capture unit. A monitoring system for the mixture oxygen content at the compressor inlet is provided. The monitoring system includes a recirculated flow mass flow rate sensor, a recirculated flow oxygen concentration sensor, mixture mass flow rate sensors, a control unit arranged to process information detected by the recirculated flow mass flow rate sensor, recirculated flow oxygen concentration sensor and mixture mass flow rate sensor, to determine an oxygen concentration upstream of the compressor inlet.

Abstract: A cooling system for a gas turbine engine includes a first structure movable relative to a second structure. The first structure has a cavity. A valve selectively controls fluid flow from a cooling source to the cavity. A valve is configured to move between first and second fluid flow positions in response to movement of the first structure. The first fluid flow position provides a greater amount of cooling fluid from the cooling source to the cavity than in the second fluid flow position.

Abstract: A nozzle for an aircraft powerplant is disclosed which provides two separate flow paths. A flow path is provided in the nozzle for a core flow of the powerplant and another flow path is provided for a bypass flow of the powerplant. The nozzle can have a variety of configurations including, but not limited to, 2D and axisymmetric. Either or both the flow paths can be convergent, divergent, or convergent-divergent, and the flow paths need not be similar between the two. Actuators are provided to manipulate the configuration of the flow paths and the areas of the flow paths. For example, throat and/or exit areas can be manipulated.

Abstract: A system and method of cooling a rocket motor component includes injecting a high pressure liquid coolant through an injector nozzle into a cooling chamber. The cooling chamber having a pressure lower than the high pressure liquid coolant. The liquid coolant flashes into a saturated liquid-vapor coolant mixture in the cooling chamber. The saturated liquid-vapor coolant mixture is at equilibrium at the lower pressure of the cooling chamber. Heat from the rocket motor component to be cooled is absorbed by the coolant. A portion of the liquid portion of the saturated liquid-vapor coolant mixture is converted into gas phase, the converted portion being less than 100% of the coolant. A portion of the coolant is released from the cooling chamber and the coolant in the cooling chamber is dynamically maintained at less than 100% gas phase of the coolant as the thrust and heat generated by the rocket motor varies.

Abstract: A method of managing aircraft exhaust includes providing hot air at a hot air mass flow rate, providing cold air at a cold air mass flow rate, and mixing the hot air and the cold air at a variable hot air mass flow rate to cold air mass flow rate ratio, wherein the variable hot air mass flow rate to cold air mass flow rate ratio is selectively maintained independent of at least one of (1) a variation in the hot air mass flow rate and (2) a variation in a translational speed of the aircraft.