Abstract: A gas turbine engine has: an engine shaft rotatable about a rotation axis; at least two bearings spaced apart from one another along the rotation axis, the at least two bearings having inner races for rotation with the engine shaft, outer races, and rolling elements disposed radially between the inner races and the outer races; a bearing support extending axially along the rotation axis from one of the at least two bearings to the other, the bearing support secured to both of the outer races of the at least two bearings, the bearing support secured to a structural case of the gas turbine engine via frangible tabs.

Abstract: Methods and systems of operating a gas turbine engine in a low-power condition are provided. In one embodiment, the method includes supplying fuel to a combustor by supplying fuel to a first fuel manifolds and a second fuel manifold of the gas turbine engine. The method also includes, while supplying fuel to the combustor by supplying fuel to the first fuel manifold: stopping supplying fuel to the second fuel manifold; and using a pump to drive gas into the second fuel manifold to flush fuel in the second fuel manifold into the combustor and hinder coking in the second fuel manifold and associated fuel nozzles.

Abstract: A system is described that includes a controllable component of an engine configured to regulate fuel flow to the engine, a digital control unit configured to control the engine by at least communicating with the controllable component of the engine, and a protection component configured to disable communication between the digital control unit and the controllable component of the engine. The system further includes an analog control unit configured to control the engine by at least communicating with the controllable component of the engine in response to the protection component disabling communication between the digital control unit and the controllable component of the engine.

Abstract: A method of calibrating a gas turbine engine having a propulsive fan, and an engine core, the method including: measuring a total thrust generated by the engine; measuring the thrust generated by the engine core; measuring first and second engine performance parameters; based on the total thrust and engine core thrust, determining a thrust generated by the propulsive fan; providing a first power setting parameter associating the fan thrust with the first engine performance parameter; and providing a second power setting parameter associating the engine core thrust with the second engine performance parameter.

Abstract: A method and system for operating a gas turbine engine coupled to an aircraft propeller are described herein. The method comprises detecting a command for unfeathering the propeller, inhibiting shaft shear detection logic in response to detecting the command for unfeathering the propeller, detecting completion of the unfeathering of the propeller, and enabling the shaft shear detection logic in response to detecting the completion of the propeller unfeathering.

Abstract: The present disclosure concerns control a hybrid electric gas turbine system (300) for an aircraft. The system comprises an electric generator (308) and a gas turbine (309) to form a generator system, an electric motor (303) and a fan (302) to form a propulsor (301), a controller (306) and an electric storage unit (307). After receiving a command for a change in demand for thrust, the controller (306) determines an operational profile that minimises a function comprising a measure of fuel supplied to the gas turbine (309), a transfer of electric power from or to the electric storage unit (307) and a difference between measures of current and demanded thrust over a time period. The controller then operates the electric motor (303), gas turbine (309) and electric storage unit (307) according to the determined operational profile over the time period.

Abstract: An apparatus for providing foreign object debris protection an air intake of an aircraft engine. The apparatus includes a frame and a plurality of cross-members. The cross-members are positioned in the frame to define a plurality of screen openings. At least one of the cross-members has an aerodynamically efficient cross section.

Abstract: A gas turbine engine includes: an engine case; an engine rotation shaft rotatably supported by the engine case via a bearing member; a first flange extending from the engine case to define a first annular contact surface facing in an axial direction; and a bearing case having an inner end supporting the bearing member and a second flange defining a second annular contact surface that coaxially contacts the first annular contact surface. The second flange is joined to the first flange so as to be movable radially relative to the first flange when applied with a load greater than or equal to a predetermined value in a direction to move the second flange radially relative to the first flange. An outer circumferential edge of one of the first and second flanges is provided with an opposing piece opposing an outer circumferential edge of the other via a radial gap.

Abstract: There is disclosed a bearing assembly for a gas turbine engine. The bearing assembly has a central axis and includes at least two bearings being axially spaced from one another relative to the central axis and configured to rotatably support a rotating component of the gas turbine engine. A bearing housing includes bearing supports supporting the at least two bearings. The bearing housing is securable to a casing of the gas turbine engine via hairpin structures being axially spaced from one another.

Abstract: A method of operating a multi-angle, multi-wave array may comprise, emitting a first light at a blue wavelength, emitting a second light at an infrared wavelength, emitting a third light at an ultraviolet wavelength, and detecting a scattered light from each of the first light, the second light, and the third light at a plurality of light sensing devices wherein the detection of scattered light is determinative between categories of foreign object debris including solid objects and particulates including silicate sand, water vapor, dust, volcanic ash, sea-salt aerosol, and smoke.

Abstract: A combustor section of a gas turbine engine includes a combustor having a combustor inlet, and a combustor bypass passage having a passage inlet located upstream of the combustor inlet. The combustor bypass passage is configured to divert a selected bypass airflow around the combustor. A combustor bypass valve is located at the combustor bypass passage to control the selected bypass airflow along the combustor bypass passage. A method of operating a gas turbine engine, includes urging a core airflow from a compressor section toward a combustor section, flowing a first portion of the core airflow into the combustor section via a combustor inlet, and flowing a second portion of the core airflow into a combustor bypass passage via a combustor bypass valve, thereby bypassing the combustor with the second portion of the core airflow.

Abstract: There is disclosed a gas turbine engine including a bearing housing having at least two bearings axially spaced from one another relative to a central axis. The bearing housing has a case between the at least two bearings having a joint configured for relative axial movement between the bearing supports. A method of operating a bearing assembly including the bearing housing is also disclosed.

Abstract: A detonation transfer assembly is disclosed. A detonation transfer assembly may comprise an external casing comprising an input end and an output end axially opposite the input end, an explosive column spanning axially inside the external casing, a primary explosive disposed within the explosive column, and a secondary explosive disposed within the explosive column axially between the primary explosive and the output end. The primary explosive and/or the secondary explosive may comprise a thermally insensitive initiation material that resists at least one of detonation or thermal degradation in response to temperature increase rate of 3.3° C. per hour over at least twenty hours.

Abstract: A support assembly for a load-bearing unit, a gas turbine engine including the support assembly, and a method of operation of the support assembly are provided. The support assembly includes a support element, a damper, and a variable stiffness member. The support element supports the load-bearing unit. The damper supports the support element and is configured to provide dampening of the load-bearing unit. The variable stiffness member is positioned between the damper and the load-bearing unit. The variable stiffness member is configured to provide a serial dampening of the load-bearing unit with the damper. The variable stiffness member includes a shape memory alloy.

Abstract: A propulsion system according to an exemplary aspect of the present disclosure includes, among other things, a pressurant selectively released from a pressure tank to drive a pump to sustain propellant flow for main combustion.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan and a braking system. The braking system is configured to selectively engage the fan during ground windmilling to apply a first level of braking to slow rotation of the fan. Further, when the rotation of the fan sufficiently slows, the braking system is further configured to apply a second level of braking more restrictive than the first level of braking.

Abstract: A double frangible bearing support structure supports a low pressure rotor of an aircraft engine. The support structure has a first bearing assembly including a first bearing supported by a first bearing support adapted to buckle or frange when subject to a predetermined critical load resulting from an abnormal rotor imbalance. The support structure has a second bearing assembly comprising a second bearing having rolling elements disposed between inner and outer races. The outer race is connected to a second bearing support by means of frangible bolts adapted to fail when subject to a predetermined critical load resulting from radial displacements and loads of the low pressure rotor following decoupling/franging at the first bearing support.

Abstract: A bearing arrangement rotatably supports a shaft of an aircraft engine. The bearing arrangement comprises a bearing having rolling elements disposed between inner and outer races. The inner race is affixed to the shaft. A decoupler normally structurally couples the outer race of the bearing to a stator structure of the engine. The decoupler is configured to release the bearing from the stator structure when subject to a predetermined critical load. A bumper is mounted to the stator structure and encircles the bearing. The bumper has a radially inwardly facing surface disposed in close proximity to a radially outer surface of the outer race of the bearing and defines therewith a radial gap to accommodate and constrain an orbiting motion of the rotor about the central axis of the engine after decoupling at the bearing.

Abstract: In one aspect, a ram air turbine stall recovery system includes a ram air turbine, a hydraulic pump powered by the ram air turbine with the hydraulic pump having a pump inlet and a pump outlet, an electric generator powered by the ram air turbine, a bypass line that connects the pump outlet to the pump inlet, a bypass valve located on the bypass line between the pump outlet and the pump inlet; and a controller responsive to an input signal that indicates a potential stall condition of the ram air turbine. The controller causes the bypass valve to open to allow a hydraulic fluid to flow through the bypass line in order to reduce a pressure within the hydraulic pump.

Abstract: A method of adjusting a setpoint value of at least one parameter that has an influence on thrust of a gas turbine engine propelling an aircraft during a stage of flight of the aircraft, the method including: obtaining a current value of at least one operating variable of the engine; extracting from a pre-established table a decrement value for the at least one parameter associated with the current value of the at least one operating variable of the engine; and adjusting the setpoint value of the at least one parameter by applying thereto the decrement value extracted from the table.

Abstract: The present application provides an inlet bleed heat system for supplying a flow of bleed air to a flow of incoming air into a compressor of a gas turbine engine. The inlet bleed heat system may include an air knife and a silencer panel. The air knife may include a compressor bleed air port in communication with the flow of bleed air and a discharge gap to discharge the flow of bleed air into the flow of incoming air. The air knife may and the silencer panel may form an integrated air knife/silencer panel.

Abstract: A gas turbine engine having at least one spool assembly, the at least one spool assembly including a fan rotor, a compressor disposed downstream of the fan rotor, a turbine and a shaft connecting the fan rotor, compressor and turbine, a joint affixed to an upstream end of the shaft, and including a first link connecting the fan rotor to the shaft and a second link connecting the compressor to the shaft, the second link being less rigid then the first link.

Abstract: In a method for validating speed of a fan of an electronic device, different speeds of the fan are set for validation. The method controls the fan to operate under each of the set speeds in turn, and controls a camera device to continuously capture images of the fan with a shutter speed of the camera device corresponding to the set speed. The images are transmitted to a comparison unit to be analyzed for sameness and clarity. The fan is determined to work abnormally if not all the images are sameness or if at least one image is unclear, and the fan is determined to work normally if all the images are sameness or clear. A validation report is generated to indicate whether the fan is normal or abnormal according to the determination.

Abstract: This invention relates to the operation of gas turbine engines, and in particular to determining deterioration of components during operation. In a specific embodiment, the invention is concerned with determining the actions to be taken when a foreign body impact has been detected, for example on a fan blade. Accordingly, the invention provides a method to take one or more FOD detection apparatus, analyse the likely (probabilistic) outcome, and provide a system to determine subsequent action that assures safety whilst minimising operational disruption. This invention provides a method and apparatus to identify FOD or bird impact to gas turbine fan blades, assessing the damage that may have occurred whist still in flight and determining post impact actions, including replacement parts.

Abstract: The presently disclosed device provides a method and means for ensuring that containers of all types and sizes are vented or purged to atmospheric or environmental conditions upon the interior or exterior of the container reaching a critical temperature pressure, or humidity. Specifically, the presently disclosed invention integrates shape memory polymer (SMP) based, thermally activated fasteners into the venting systems. The result is a venting system that increases munitions safety without compromising the venting effectiveness, structural integrity, or the required bullet/fragment impact resistance requirements of the system.

Abstract: A method and a system for detecting the ingestion of an object by an aircraft turbine engine during a mission is disclosed. The method includes: acquiring during the mission digital images of the operating fan of the turbine engine, these images being acquired at an acquisition frequency proportional to the speed of rotation of the fan and to the number of blades of the fan; identifying the different phases of the mission of the aircraft for each phase of the mission; comparing the images of the fan with at least one reference image corresponding to sound operation of the fan; and, if necessary, identifying each abnormal image of the fan which differs from the corresponding reference image, the identification of an abnormal image of the fan corresponding to the crossing of an alert level for detecting the ingestion of an object by the turbine engine.

Abstract: An insensitive munitions casing, such as a motor case, for housing a propellant charge is disclosed. The casing is formed by winding, wrapping or rolling a fiber, tape or sheet of material such that the casing has a plurality of reinforcing layers. At least some of the reinforcing layers have discontinuous interfaces. Other reinforcing layers may overlap the discontinuous interfaces. The reinforcing layers are coated with adhesive bonding material for adhering the reinforcing layer to itself and bonding the discontinuous interfaces. The adhesive bonding material has a bond-strength breakdown temperature which is below the auto-ignition temperature of the propellant charge. Once the bond strength breakdown temperature of the adhesive is reached, the bond fails at the discontinuous interface, thereby enabling release of propellant charge pressurization.

Abstract: A nozzle is provided and includes an outer annulus defined by an exterior wall and an interior wall and including air inlets through which air flows to a fuel mixing zone and a combustion zone, an inner annulus disposed within the interior wall and including a fuel volume into which fuel is fed to a distal end thereof, which is adjacent to and isolated from the combustion zone and an airflow line, disposed between the fuel volume and the interior wall, through which air flows to the combustion zone with the airflow line and the combustion zone isolated from the fuel volume, and a fuse configured to melt during a flameholding incident and to form a breach through which fuel flows from the fuel volume, bypassing the fuel mixing zone, to a fuel burning zone downstream from the fuel mixing zone.

Abstract: A rotor decoupler system for a gas turbine engine is disclosed, comprising a load reduction system disposed between a bearing system and the bearing support. The load reduction system comprises a fuse and a damper such that the damper transfers at least a part of the radial load from the rotor to the bearing support when a fuse fails. In one exemplary embodiment the damper comprises a wire mesh, disposed around the outer race of the bearing. In another exemplary embodiment the damper comprises a rope damper assembly. In yet another exemplary embodiment, a load reduction system is disposed between a support structure and a frame, comprising an extensional fuse and a damper disposed between the support structure and the frame such that the damper transfers at least a part of the unbalance load to the frame when an extensional fuse fails.

Abstract: There is disclosed a solid fuel rocket motor including a center-perforated solid fuel grain. The solid fuel grain may be adapted to burn outwardly from a center surface facing the center perforation when ignited by a primary igniter during a normal mode of operation. The solid fuel grain may be further adapted to burn longitudinally from an end face when ignited by a secondary igniter during a safety mode of operation.

Abstract: A device for, in a gas turbine engine, braking a turbine including a rotor driving a shaft capable of rotating with respect to a stator in the event of said shaft breaking is disclosed. The device includes a first braking member provided with at least one abrasive element and a second braking member including a ring-shaped element (120A) made of a material capable of being eroded by the abrasive element. One of the two braking members being secured to the rotor and the other of the two braking members being secured to the stator. The braking members come into contact with one another through axial displacement of the rotor once the shaft has broken. The abrasive element of the first braking member eroding the ring-shaped element of the second braking member.

Abstract: The invention relates to a turbomachine thermomechanical part forming a body of revolution about a longitudinal axis, and including at least one abradable ring for a labyrinth seal. In characteristic manner, the abradable ring is made up of angular sectors that present different stiffnesses between adjacent pairs of sectors. The invention is applicable to a compressor, a turbine, or to a rotor and stator assembly.

Abstract: The invention relates to a method and a system for detecting the ingestion of an object by an aircraft turbine engine during a mission. The method includes the steps of acquiring during the mission digital images of the operating fan of the turbine engine, these images being acquired at an acquisition frequency proportional to the speed of rotation of the fan and to the number of blades of the fan (E10), identifying the different phases of the mission of the aircraft (E40), for each phase of the mission, comparing the images of the fan with at least one reference image corresponding to sound operation of the fan (E60), and, if necessary, identifying each abnormal image of the fan which differs from the corresponding reference image, the identification of an abnormal image of the fan corresponding to the crossing of an alert level for detecting the ingestion of an object by the turbine engine (E90).

Abstract: A method for assembling a gas turbine engine includes coupling a low-pressure turbine to a core turbine engine, coupling a counter-rotating fan assembly including a forward fan assembly and an axially aft fan assembly to the low-pressure turbine such that the forward fan assembly rotates in a first direction and the aft fan assembly rotates in an opposite second direction, and coupling a booster compressor directly to the low-pressure turbine such that the booster compressor rotates in the first direction.

Abstract: A turbine for a sprinkler is disclosed for self-governing its rotational velocity. As a rate of fluid through the sprinkler increases, particularly when air is used to flush the sprinkler system, a portion of the turbine shifts outwardly so as to decrease alignment of vanes located thereon with directed water streams for controlling the rotation of the turbine.

Abstract: An emergency fuel shut-off system for a gas turbine engine comprises a fuel shut-off device to be actuated for terminating a fuel supply to the engine in response to a mechanical actuation force and an actuator adapted for creating the mechanical actuation force when a rotor shaft of the engine fails. A fuse apparatus is provided to link the actuator with the shut-off device in order to transmit the mechanical actuation force. The fuse apparatus disables transmission of the mechanical actuation force when the transmitted mechanical actuation force exceeds a pre-selected level.

Abstract: A casing for a gas turbine includes a construction providing improved structural efficiency. Improved load paths and means for transmitting loads in the engine case are disclosed.

Abstract: A rotor decoupler system for a gas turbine engine is disclosed, comprising a load reduction system disposed between a bearing system and the bearing support. The load reduction system comprises a fuse and a damper such that the damper transfers at least a part of the radial load from the rotor to the bearing support when a fuse fails. In one exemplary embodiment the damper comprises a wire mesh, disposed around the outer race of the bearing. In another exemplary embodiment the damper comprises a rope damper assembly. In yet another exemplary embodiment, a load reduction system is disposed between a support structure and a frame, comprising an extensional fuse and a damper disposed between the support structure and the frame such that the damper transfers at least a part of the unbalance load to the frame when an extensional fuse fails.

Abstract: A turbine for a sprinkler is disclosed for self-governing its rotational velocity. As a rate of fluid through the sprinkler increases, particularly when air is used to flush the sprinkler system, a portion of the turbine shifts outwardly so as to decrease alignment of vanes located thereon with directed water streams for controlling the rotation of the turbine.

Abstract: A method for surface treating a titanium gas turbine engine component. The method includes providing a gas turbine engine component having a titanium-containing surface. The component is heated to a temperature sufficient to diffuse carbon into the titanium and below 1000° F. The surface is contacted with a carbon-containing gas to diffuse carbon into the surface to form carbides. Thereafter, the carbide-containing surface is coated with a lubricant comprising a binder and a friction modifier. The binder preferably including titanium oxide and the friction modifier preferably including tungsten disulfide. The coefficient of friction between the surface and another titanium-containing surface is less than about 0.6 in high altitude atmospheres.

Abstract: An arrangement of bearing supports for a rotating shaft mounted on bearings connected to an engine stator structure by bearing supports of which at least one is made up of N>1 elements effectively acting in parallel to simultaneously connect the bearing to the structure, of which N?1 of the N elements form this connection by fusible links, is made up of a first bearing supported by a first bearing support and a second bearing supported by a second bearing support, that of the two bearing supports which makes up the N>1 elements being the second bearing support.

Abstract: The invention relates to a method of detecting damage in a bearing supporting at least one rotary shaft of an engine in rotation, the method consisting in: defining a measurement period corresponding to a range of speeds of rotation of the shaft during renewable activity at low speeds of operation of the engine; acquiring over the entire measurement period a vibratory signal representative of acceleration of components of the engine; sampling the vibratory signal as a function of the speed of rotation of the shaft during the measurement period; transforming the sampled vibratory signal into a frequency signal to obtain frequency spectrum lines as a function of the speed of rotation of the shaft; calculating the mean of the amplitudes of the spectrum lines; determining amplitude peaks around multiples of the theoretical frequency of a damaged roller; calculating the ratio between each amplitude peak and the amplitude level determined for a sound bearing; and comparing the ratio that is obtained with at least o

Abstract: A fluid jet apparatus includes a housing defining a propellant cavity, a fluid cavity, and a passage through the housing in fluid communication with the fluid cavity. The apparatus further includes a membrane separating the propellant cavity and the fluid cavity, a propellant disposed in the propellant cavity, and a fluid retained in the fluid cavity. The propellant urges the fluid from the fluid cavity through the passage upon initiation of the propellant. A munition system includes a munition and at least one fluid jet apparatus adapted to vent the munition. A method includes initiating a propellant to produce a rapidly expanding gas and urging a fluid through a passageway with the gas to produce a fluid jet.

Abstract: A device for venting a container housing an energetic material includes an insulating portion, a charge holder disposed in the insulating portion, and an explosive cutting charge disposed in the charge holder. The device further includes a thermally activated initiation device, and a transfer line coupling the thermally activated initiation device and the explosive cutting charge. An apparatus includes a container, an energetic material disposed within the container, and a device, disposed within the container, for venting the container. The device for venting the container includes an insulating portion, a charge holder disposed in the insulating portion, and an explosive cutting charge disposed in the charge holder. The device for venting the container further includes a thermally activated initiation device, and a transfer line coupling the thermally activated initiation device and the explosive cutting charge.

Abstract: The turbomachine of the invention extends longitudinally along an axis, and includes a rotor attached to drive shaft, arranged to rotate around an axis, supported by at least a first bearing, mounted on the fixed structure of the turbomachine by a bearing support element. The turbomachine is characterised by the fact that it includes a stop ring, mounted on the fixed structure of the turbomachine, to cooperate with the support element of the first bearing and, in the event of displacement of the rotor in relation to the fixed structure, to perform a function of axial retention of the rotor in an even manner, with no angle effect between the axis of the turbomachine and the axis of the drive shaft.

Abstract: A gas turbine engine afterburner igniter is provided with primary fuel duct means for injecting a jet of fuel into a gas stream directed into a combustion chamber. A spring is positioned within the bore of said fuel duct. The spring being moveable relative to said duct such that during operation said spring moves downstream though the bore due to the passage of fuel therein and returns to it's original position thereafter. During operation the spring abrades the internal surface of the bore thus removing carbon deposits.

Abstract: A method for fabrication of a locomotive diesel engine turbocharger turbine stage so as to avoid harmonic vibration in the turbine blades when the engine is operated at pre-selected throttle settings. Upon determining throttle settings for the engine and modeling of the turbocharger therefor, at least one natural vibration frequency of the turbine blades and a turbine nozzle vane excitation frequency as a function of turbocharger rotation speed are determined. Then the data obtained is analyzed to ascertain whether at least one data coincidence, that is whether a rotation speed correlative of a throttle setting has a coincidence over a predetermined range of both a natural frequency and said turbine vane nozzle excitation. If not, the model is suitable for fabrication of the turbocharger; if so, the model is modified by changing blade stiffness and/or vane numbering until no data coincidences are present.

Abstract: A device is provided which can serve one or more of the multiple functions of impeding rotation of a shaft de-coupled turbine rotor, selectively impeding rotation of a first spool of the engine, for example to permit a second spool to generate power for use in ground operation, and facilitating reduced aircraft ground speed.

Abstract: The invention relates to a turbomachine comprising a rotor with a drive shaft centered on the axis of the turbomachine by a first bearing and a second bearing which bearings are supported respectively by a first-bearing-support piece and a second-bearing-support piece which are secured to one another and connected to the turbomachine fixed structure by a decoupling device. This turbomachine comprises means designed to collaborate with at least one element of the turbomachine fixed structure in order to perform a dual function, that of preventing the bearing supports, from rotating and that of radially retaining the drive shaft in the event of the bearings becoming decoupled.

Abstract: A mechanical-electronic shut-off system detects a shaft failure and initiates the shut-off of the fuel supply. It features, on the free end of a reference shaft (2) connected the to energy-consuming end of the respective shaft (1), an axially moveable signal trip element (13, 14) held under pre-load (12) whose locking arrangement (17,18) is released via a radial driver arrangement (17, 19) by rotary movement in the event of a shaft failure. The resultant relative rotation of the shaft (1) enables the signal trip element to move towards a sensor (21) or a switching element. An electric signal so produced instantly interrupts the further supply of fuel by means of an electronic control and avoids or controls a dangerous overspeed condition of the failed shaft.