Abstract: An inlet for a turbofan engine, including an inlet wall surrounding an inlet flow path. The inlet wall extends axially from an upstream end to a downstream end adjacent the fan. The inlet wall has a shape defining a plurality of teeth circumferentially spaced around the inlet. The teeth extend axially and project radially inwardly toward the central longitudinal axis. A central portion of the inlet flow path has a cross-sectional dimension measured diametrically between opposed teeth, the cross-sectional dimension varying along the axial direction. The central portion defines a geometric throat at a minimum value of the cross-sectional dimension. The inlet wall is shaped so that the geometric throat is axially spaced from the upstream end and the downstream end. A gas turbine engine and a method of shielding tips of fan blades from impact by an object having a predetermined minimum dimension are also discussed.

Abstract: A gas turbine engine nacelle comprising an intake liner. The liner includes a plurality of cells. Each cell includes an open radially inner end in fluid communication with an interior side of the nacelle, and an open radially outer end in fluid communication with an exterior side of the nacelle. Each open end of each cell defines a respective cross sectional area. The intake liner further comprises radially inner and outer facing sheets overlying a respective radially inner and outer open ends of the respective cell. Each facing sheet defines at least one aperture overlying at least one cell, an overlying portion of the respective aperture having a smaller cross sectional area than the respective open end of the respective cell.

Abstract: A device for the extraction of bleed air from flowing air at or in an aircraft engine is provided. The device includes a means for specific adjustment of an inlet cross section of an opening at or in the area of a wall of the aircraft engine and a flow guide means for a boundary layer flow.

Abstract: A ram air turbine movable between a stowed position and a deployed position is provided including a rotatable hub assembly. A hub face is arranged at a first end of the rotatable hub assembly. The hub face has a non-planar configuration such that a limited portion of a surface area of the hub face is arranged within a plane oriented perpendicular to a direction of travel of an airflow upstream from the hub face when the ram air turbine is in a deployed position.

Abstract: A gas turbine engine includes a nacelle defining a centerline axis and an annular splitter radially inward from the nacelle. A spinner is radially inward of the nacelle forward of a compressor section. A fan blade extends from a fan blade platform. A distance X is the axial distance from a first point to a second point, wherein the first point is defined on a leading edge of the annular splitter and the second point is defined on a leading edge of the fan blade where the fan blade meets the fan blade platform. A distance H is the radial distance from the first point to the second point. The relative position of the first point and the second point is governed by the ratio of X H ? 1.5 for reducing foreign object debris (FOD) intake into the compressor section.

Abstract: A nacelle structure (10) surrounding a fan, the structure comprising an air inlet (100) having an inside surface (100A) defining an air feed channel; and at least one ring (22) arranged in the air inlet and selectively movable between a deployed position in which the ring is moved radially inwards from the inside surface to provide an obstacle to shock waves coming from the fan, and a retracted position in which the ring is moved radially outwards from the deployed position so as to become flush with the inside surface and thus reconstitute the air feed channel.

Abstract: An aircraft turbine engine includes a fan duct having a wall; an air passage arranged in the wall and including an air inlet opening at the wall, the air passage being designed to receive part of the flow of air from the fan duct across the opening; an air intake housing located above the air inlet opening and fixed to the wall, the housing being arranged to intercept part of the air flow in the fan duct, and successively including, in the air flow direction, an upstream wall then a downstream wall, an upstream opening and a downstream opening arranged on the upstream wall and the downstream wall, respectively, the inside of the housing being in fluidic connection to the air inlet opening, and a movable sealing means between an “open” position; and a “closed” position relative to the downstream opening and the air inlet opening.

Abstract: A ducted fan UAV that can be collapsed into a stowed configuration and then deployed for flight by, for example, inflating the duct to a deployed configuration. The UAV includes a plurality of rotor blades, a plurality of struts and a plurality of control vanes each being pivotally mounted to a center body by a hinge so that the rotor blades, the struts and the control vanes can be folded into the stowed configuration to be substantially parallel to the center body and be unfolded into the deployed configuration to be substantially perpendicular to the center body. The UAV also includes a pressurization system providing a pressurant to a chamber within the duct so as to inflate the duct and cause the struts, the rotor blades and the control vanes to move from the stowed configuration to the deployed configuration.

Abstract: In an intake structure of an aircraft that takes in air from a front side of an airframe, a bump is provided on a surface of the airframe of the aircraft at a portion immediately before an intake. The bump is formed so as to smoothly protrude from the front side of the airframe while having a width that is wider than a width of the intake.

Abstract: An engine inlet for efficient operation at both design Mach number and off-design Mach numbers with an inlet having a caret configuration with rotatably extending effective leading edges on the inlet from a retracted position aligned with a nominal Mach number shock wave to an extended position aligned with an off-design Mach number shock wave.

Abstract: A blower assembly includes a housing having a scrolled wall, a motor having an output shaft extending within the housing, an impeller positioned within the housing and mounted to the output shaft for creating an airflow along an airflow path within the housing, an airflow outlet, and a scroll cut-off. The scroll cut-off is a membrane positioned between the scrolled wall and the air flow outlet, and forms a substantially continuous surface with the scrolled wall and with the airflow outlet, along the airflow path. In a first position, the membrane is substantially taut and in a second position, the membrane is relaxed. Alternatively, the membrane is inflatable via a valve in fluid communication with the membrane.

Abstract: An apparatus is formed of a cowl and a flow inlet formed on an interior surface of the cowl. The flow inlet has a supersonic compression section attached to a subsonic diffusion section at a throat. The supersonic compression section includes an at least partially elliptical compression ramp which extends along an approximately 180 degree arc along the interior surface. The flow inlet may form part of an aircraft. A method of air flow using the flow inlet is also disclosed.

Abstract: A silencer apparatus for a gas turbine inlet system ducting is disclosed. The inlet ducting contains a main silencer and a pre-silencer, and Inlet Bleed Heat (“IBH”) located between the silencers. The pre-silencer decreases the noise level from the turbine compressor and makes the air flow/temperature profiles more uniform. The main silencer reduces noise from the IBH and the remaining noise from the compressor to an appropriate level. The main silencer is comprised of a first plurality of sound-absorbing splitters disposed along the gas flow direction in the gas turbine inlet ducting. The pre-silencer is comprised of a second plurality of sound-absorbing splitters disposed along the gas flow direction in the gas turbine inlet ducting, but staggered with respect to the first plurality of splitters to thereby block a direct line of travel for noise acoustical waves from the compressor travelling opposite the gas flow direction in the inlet ducting.

Abstract: A nacelle for a turbofan gas turbine engine is provided. An inner wall of the nacelle defines an air intake which directs air into the fan section of the engine. The intake has, in flow series, an intake lip, a throat and a diffuser. The diffuser has, in flow series, first and second flow conditioning sections over both of which the flow cross-sectional area of the diffuser increases with increasing downstream distance from the throat. In addition, over the second section the nacelle inner wall lies substantially on a surface of an oblique circular cone having an apex which is offset from the centerline of the engine.

Abstract: An engine inlet for efficient operation in both subsonic and supersonic flight wherein the inlet has a caret configuration is rotatable about an off-body axis for compression ramp angle and capture area variation and a diffuser is engaged to the inlet in a scrubbing relationship to maintain a seal upon rotation of the inlet.

Abstract: Methods are provided forming a metallic nacelle inlet structure for a turbine engine aircraft propulsion system with an axis. The nacelle inlet structure includes an inlet lip portion and an outer barrel portion. One of the method may include steps of: forming the inlet lip portion; forming the outer barrel portion; arranging the inlet lip portion axially adjacent the outer barrel portion; and friction stir welding the inlet lip portion to the outer barrel portion to provide a circumferentially extending weld joint between the inlet lip portion and the outer barrel portion.

Abstract: An apparatus comprises a leading edge of an inlet. The leading edge of the inlet is positioned relative to a direction of air flow such that a total pressure of air along the leading edge of the inlet is equalized within selected tolerances.

Abstract: A swept gradient air boundary layer diverter for an aircraft. The aircraft includes a fuselage and an air inlet for an engine of the aircraft, where the air inlet includes a cowl at a leading edge of the inlet. The diverter includes a V-shaped ramp portion formed in the fuselage in an area proximate to and in front of the cowl where the ramp portion extends downward away from an outer surface of the fuselage towards an inside of the aircraft. The diverter also includes a V-shaped trough portion formed into the fuselage and being positioned adjacent to and integral with the ramp portion between the ramp portion and the air inlet. Air flowing over the fuselage towards the cowl is expanded and compressed by the ramp portion and the trough portion so as to create pressure gradients that generate vortices to redirect boundary layer airflow around the air inlet.

Abstract: An aircraft gas-turbine engine has an engine cowling-nacelle which surrounds a core engine in a tubular way, characterized in that at least one inflow-side part of the engine cowling-nacelle is telescopically movable against the flow direction.

Abstract: A battery enclosure is provided that externally shorts the enclosed battery, or batteries, immediately prior to the battery being punctured. As a result, the battery's state-of-charge is decreased, leading to a less severe reaction when the battery case is punctured.

Abstract: Device for exhaust air cooling of aircraft air-conditioning systems comprising a bypass channel that bypasses an aircraft air-conditioning system arranged downstream of a ram air channel with a ram air inlet channel portion, and merges into an outlet channel, arranged downstream of the aircraft air-conditioning system, before a discharge opening for exhaust air, a common inlet being provided for the ram air channel and the bypass channel.

Abstract: A nacelle for a gas turbine engine system includes a housing that has a pressure side and a suction side. A passage extends between the pressure side and the suction side for permitting airflow from the pressure side to the suction side. The passage includes a first curved section that has an inlet at the pressure side. A linear section is connected with the first curved section. A second curved section is connected with the linear section and has an outlet at the suction side. The outlet has a cross sectional area smaller than a cross sectional area of the linear section.

Abstract: An aircraft nacelle comprising an air intake configured to channel an air flow in the direction of an engine assembly which comprises a fan with an axis of rotation, said air intake comprising, according to a flow direction of the air flow, a lip with a leading edge and an internal conduit which extends said lip towards the fan, wherein the air intake comprises a plurality of protrusions distributed over the circumference of the air intake and following each protrusion, a concave form configured in order to achieve compression of the air flow.

Abstract: An engine inlet assembly includes an inlet duct having a first inlet duct leg and a second inlet duct leg, the first inlet duct leg and the second inlet duct leg extending toward a common inlet duct outlet located at an engine inlet. The first inlet duct leg and the second inlet duct leg are fixedly secured to an airframe of an aircraft at an air inlet. The duct outlet has a floating interface with the engine inlet.

Abstract: A flight vehicle includes a fuselage and a gas turbine engine. The gas turbine engine is coupled to the fuselage to provide thrust when air surrounding the flight vehicle is admitted to the gas turbine engine and combusted with fuel. The flight vehicle further includes a generator coupled to the gas turbine engine to provide power to equipment included in the vehicle.

Abstract: An engine system including an air intake duct positioned upstream of an engine cylinder may include an ice ingestion feature for retaining condensation. An ice ingestion feature may include indents formed in the bottom of an air intake duct wall. As such, the ice ingestion feature may include compartments and/or grooves of varying depths, widths, and/or angles such that the ice retention rate may be based on the surface area of the compartment or grooves.

Abstract: A supersonic inlet includes a relaxed isentropic compression surface to improve net propulsive force by shaping the compression surface of the inlet to defocus the resulting shocklets away from the cowl lip. Relaxed isentropic compression shaping of the inlet compression surface functions to reduce the cowl lip surface angle, thereby improving inlet drag characteristics and interference drag characteristics. Supersonic inlets in accordance with the invention also demonstrate reductions in peak sonic boom overpressure while maintaining overall engine performance.

Abstract: A combined engine includes a turbopump including a pump injecting hydrogen into a heater arranged in an outer casing downstream from a central body, and a subsonic turbine driving the pump, which turbine receives partially-expanded hydrogen collected at an outlet from the heater to apply the hydrogen to a supersonic turbine to operate the engine as a turbojet. The hydrogen from the supersonic turbine is collected in tubes inside the central body to be sent to a combustion chamber defined downstream from the central body, while the hydrogen that is partially expanded in the subsonic turbine is sent directly to the combustion chamber via injectors to operate the engine as a ramjet.

Abstract: An air intake system suitable for a supersonic vehicle is disclosed. The system includes a channel comprising an inlet and a side wall and a plenum coupled to the side wall. The plenum is configured to accept a flow of coolant. In certain embodiments, the coolant is the waste coolant from an on-board electronics cooling system. The system also includes a porous region in the side wall configured to allow a flow of bleed air from the channel through the porous region of the side wall into the plenum so as to aid the transition to supersonic flow. In certain embodiments, the flow of the bleed air is reduced at supersonic speeds by pressurization of the plenum with the coolant.

Abstract: An inlet particle separator system for a vehicle engine includes a hub section, a shroud section, a splitter, and a hub suction flow passage. The shroud section surrounds at least a portion of the hub section and is spaced apart therefrom to define a main flow passageway that has an air inlet. The splitter is disposed downstream of the air inlet and extends into the passageway to divide the main flow passageway into a scavenge flow path and an engine flow path. The hub suction flow passage has a hub suction inlet port and a hub suction outlet port. The hub suction inlet port extends through the hub section and is in fluid communication with the air inlet. The hub suction outlet port extends through the splitter and is in fluid communication with the scavenge flow path.

Abstract: An example inlet duct screen assembly includes a grid portion including a plurality of apertures and a frame portion circumscribing the grid portion. The grid portion and the frame portion are formed from a single sheet of material.

Abstract: A supersonic inlet includes a relaxed isentropic compression surface to improve net propulsive force by shaping the compression surface of the inlet to defocus the resulting shocklets away from the cowl lip. Relaxed isentropic compression shaping of the inlet compression surface functions to reduce the cowl lip surface angle, thereby improving inlet drag characteristics and interference drag characteristics. Supersonic inlets in accordance with the invention also demonstrate reductions in peak sonic boom overpressure while maintaining overall engine performance.

Abstract: An air inlet (10) for an engine (20) of an aircraft (50), the air inlet (10) having a duct (2) feeding outside air to said engine (20) together with a filter system combining at least one filter (3) and at least one bypass device (4). Said filter (3) and said bypass device (4) are fastened on a common support (1) situated at one end of said duct (2) and on a flank of a fuselage (51) of said aircraft (50). Said bypass device (4) comprises a cover (5) and an actuator (6) controlling opening of said cover (5). Thus, when said filter (3) no longer allows outside air to pass therethrough, said cover (5) is opened, and outside air passes via the bypass device (4) in order to feed said engine (20) with outside air.

Abstract: A rotating inlet cowl for a turbine engine includes a rotation axis. The rotating inlet cowl includes a forward cone defining a forward end of the inlet cowl. The forward end is configured to be eccentric relative to the rotation axis of the inlet cowl. Furthermore, the forward cone is truncated by a truncation surface defining the forward end of the inlet cowl.

Abstract: A supersonic fan has at least two fan blades extending radially from a rim surface of a hub and is circumscribed by a duct. An inboard leading edge having an inboard sweep angle rotates at approximately Mach 1 or greater. A low-pressure zone generates propulsion. An inboard flow regime may migrate outboard from the rim surface and contaminate the low-pressure zone, reducing propulsion. An outboard leading edge having an outboard sweep angle extends radially from the inboard leading edge to form an apex. The inboard and outboard sweep angles are each at least approximately 30 degrees and in the opposite direction. An inboard vortex forming near to and as a result of the apex trails circumferentially across the fan blades and is positioned to substantially confine the inboard flow regime to be inboard of the apex, thereby preserving the low-pressure zone and increasing propulsion for the supersonic fan.

Abstract: An aircraft nacelle with a connecting device between a first conduit of an air intake and a second conduit of a motor, both conduits being arranged end-to-end, the connecting device including an annular flange connected to the air intake, an annular flange connected to the motor pressed against annular flange of the air intake, the flanges extending in a plane perpendicular to the longitudinal axis of the nacelle; bolt holes formed in annular flanges and disposed plumb to one another; and connecting elements located in bolt holes, for connecting the annular flanges, wherein at least one flange includes at least one opening interposed between a bolt hole and junction zone of the flange and/or between a bolt hole and outer edge of the flange.

Abstract: A method of securing two components to each other includes locating a mounting strip at a first mounting flange of a first component. The mounting strip is slidably affixed to the first component and includes a plurality of strip openings and a tightening ramp located at each strip opening of the plurality of strip openings. A second component is located such that a second mounting flange of the second component abuts the first mounting flange. A plurality of fasteners are located at the second mounting flange. A fastener of the plurality of fasteners extends through each strip opening of the plurality of strip openings. The mounting strip is slid in a direction to increase tension on the plurality of fasteners via an increasing height of the tightening ramp under a fastener head of the fastener, thereby securing the first mounting flange to the second mounting flange.

Abstract: A turbofan engine which has a composite fan case surrounding a fan with a plurality of fan blades is disclosed. The composite fan case includes a containment zone having an inner fabric layer composed of resin-impregnated fibers substantially uni-axially oriented in a common angular direction corresponding to a blade release angle of the fan blades. The fan case also includes a composite outer shell and an energy absorbing core disposed radially between the inner fabric layer and the composite outer shell. The energy absorbing core includes non resin impregnated multidirectional fibers.

Abstract: A nacelle inlet is provided for a turbo-fan engine for an aircraft. The nacelle inlet includes a lip skin having first and second pad-ups, an inner barrel including a forward flange and a bulkhead. The bulkhead, the lip skin and the inner barrel are attached together at mating surfaces. The bulkhead and the lip skin are attached together at mating surfaces. Each mating surface defines a surface geometry which approximates a developable surface having parallel ruling lines. The developable surfaces are formed by machining.

Abstract: A supersonic inlet having a cowl lip configured to capture the conic shock and exhibit a zero or substantially zero cowl angle is disclosed herein. The inlet may be configured to employ a relaxed isentropic compression surface and an internal bypass. The nacelle bypass may prevent flow distortions, introduced by the capture of the conic shock, from reaching the turbomachinery, thereby allowing the cowl angle to be reduced to zero or substantially zero. Such a cowl angle may reduce the inlet's contribution to the overall sonic boom signature for a supersonic aircraft while allowing for an increase in engine pressure recovery and a subsequent improvement in generated thrust by the engine.

Abstract: A variable area engine inlet of a helicopter is provided and includes first inlet portions disposed to face one another in opposite directions and at a distance from one another and second inlet portions extending between the first inlet portions and being disposed to face one another in opposite directions and at a distance from one another. The first and second inlet portions define a capture area and at least one or both of the first and second inlet portions include a movable portion disposed to occupy and move between first and second positions. The first position is associated with a non-constricted condition of the capture area and the second position is associated with a constricted condition of the capture area.

Abstract: A gas turbine engine has an inlet duct formed to have a shape with a first ellipse in one half and a second ellipse in a second half. The second half has an upstream most end which is smaller than the first ellipse. The inlet duct has a surface defining the second ellipse which curves away from the first ellipse, such that the second ellipse is larger at an intermediate location. The second ellipse is even larger at a downstream end of the inlet duct leading into a fan.

Abstract: An inlet door system includes a duct, an inlet door, and a plurality of openings. The duct is configured to extend from the auxiliary power unit (APU) to an intake opening formed in an outer surface of an aircraft. The duct includes an inlet port, an outlet port, and a duct sidewall extending between the inlet port and the outlet port. The inlet door includes an inner surface, an outer surface, and an outer peripheral edge between the inner and outer surfaces. The door is rotationally coupled to the duct, and is configured to selectively rotate between a closed position and a plurality of open positions. The openings extend through the inlet door between the inner surface and outer surface, and each opening is disposed adjacent to the outer peripheral edge.

Abstract: A method and apparatus for supplying air to a precooler. Air flow is created through a fan duct in an engine system. A first portion of the air flow is directed into a first inlet of an inlet system to feed a first half of the precooler. A second portion of the air flow is directed through the fan duct into a second inlet of the inlet system to feed a second half of the precooler.

Abstract: An apparatus comprises a leading edge of an inlet. The leading edge of the inlet is positioned relative to a direction of air flow such that a total pressure of air along the leading edge of the inlet is equalized within selected tolerances.

Abstract: An inner fixed structure of a thrust reverser includes an inner skin, an outer skin, a cellular core disposed between the inner skin and the outer skin, and at least one crack and delamination stopper extending radially between the inner skin and the outer skin.

Abstract: An aircraft nacelle including a pipe (32), a peripheral wall (34), a lip (36), and a front frame (38) connecting the peripheral wall (34) and the pipe (32) and forming with the lip (36) a space (50) in which the hot air that is provided for frost treatment can circulate, with the pipe (32) including a coating (44) for the composite material acoustic treatment, is characterized in that the nacelle includes at least one element (52) made of a heat-conducting material inserted between the lip (36) and the pipe (32) ensuring the continuity of the aerodynamic surfaces of the lip (36) and the pipe (32) and the propagation of heat from the space (50) toward the rear of the nacelle, whereby the at least one element (52) includes a coating (54) for acoustic treatment that is made of a heat-resistant material.

Abstract: An acoustic treatment panel that is connected to an air intake of an aircraft nacelle, includes at least one acoustically resistive structure (30) and a reflective layer (32), between which are located—in a direction that is essentially perpendicular to the longitudinal direction of the nacelle—bands (34) of alveolar cells spaced in such a way as to allow the passage of hot air provided for a frost treatment, characterized in that it includes pipes (36) for hot air that are each delimited by at least one partition that extends from the acoustically resistive layer (30) up to the reflective layer (32) in such a way as to insulate the bands (34) of cells in the longitudinal direction.

Abstract: One example of an engine mounted inlet plenum for a rotorcraft includes an inlet flange, an outlet flange and a plenum duct that connects the inlet flange and the outlet flange. The inlet flange, which defines an inlet to flow air into a barrier filtration system, is configured to be substantially aligned with a closed cowl door assembly of the rotorcraft. In some implementations, the plenum duct is configured to extend to an outer mold line (OML) of the rotorcraft such that the inlet flange is aligned with the OML. The outlet flange defines an outlet to flow air filtered by the barrier filtration system into an engine of the rotorcraft. The plenum duct allows air to flow from the inlet to the outlet.

Abstract: A fan blade for a gas turbine engine includes an airfoil that includes leading and trailing edges joined by pressure and suction sides to provide an exterior airfoil surface that extends in a radial direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord. A circumferential coordinate is scaled by the local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading and trailing edges at the span location.