Abstract: Embodiments of methods and apparatus for providing a sacrificial shield for a fuel injector are provided. According to one example embodiment, a method for shielding a fuel injector associated with a combustion device is provided. The method can include providing at least one shield support device operable to mount to a tip of a fuel injector. The method can also include mounting a plurality of shields to the at least one support device, wherein each of the plurality of shields is spaced apart from each respective adjacent shield by at least one segmentation, and wherein each of the plurality of shields is operable to reduce heat exposure to the tip of the fuel injector.

Abstract: The present invention is directed to a mounting system for a gas turbine transition duct. The mounting system includes a generally C-shaped mounting bracket having a generally radially extending first portion, an arc-shaped second portion, and a generally radially extending third portion. Each of the first portion and third portion includes a spherical bearing and the second portion has a plurality of sets of mounting holes. In another embodiment of the present invention, a gas turbine transition duct is provided having a panel assembly, an aft frame secured to the panel assembly and a mounting system for securing the transition duct to a turbine inlet. The panel assembly includes two formed sheets of metal secured together along axial seams. The panel assembly is secured to an aft frame that is capable of expanding in the circumferential direction due to thermal growth.

Abstract: An aft part of an aircraft, including at least one engine assembly including a turbine engine and a turbine engine mounting pylon. A rigid structure of the mounting pylon includes a longitudinal central box and a connection box carried by the box and projecting from the box. The aircraft structure includes a fuselage lateral extension projecting from it, and the connection box is pressed into contact with the fuselage extension, below it, an attachment mechanism being arranged between these two entities.

Abstract: A system is provided for mounting fuel injectors to a gas turbine engine, including an engine casing having an aperture formed therein and a plurality of fuel injectors, each fuel injector having a respective flange for mounting the fuel injectors to the casing at the aperture so that the fuel injectors extend side-by-side into the engine. The flanges are dismountably sealed to an inner side of the casing. The aperture and the flanges are configured so that, when dismounted, each fuel injector is configured to rotate into an orientation relative to the aperture which allows each of the respective flanges to pass though the aperture and the fuel injector to be withdrawn from the casing. The flanges have respective sealing formations which engage with their neighboring flanges when the fuel injectors are mounted to the casing to close off the aperture and to form seals between the flanges.

Abstract: The invention relates to a propulsion assembly for an aircraft, that comprises: a turbojet engine (1) including a fan casing (3) and an engine casing (5); a pylon for holding said turbojet engine (1); and a device for suspending said turbojet engine under the pylon, that comprises two front suspension connecting rods (9), two rear suspension connecting rods (11), and a thrust acceleration connecting rod (15) extending between the front portion (17) of said engine casing (5) and the pylon in a plane containing the axis (A) of said turbojet engine (1). An additional connecting rod (19) extends between said fan casing (3) and said pylon in a direction substantially parallel to that of the axis (A) of said turbojet engine.

Abstract: A combustor-turbine seal interface is provided for deployment within a gas turbine engine. In one embodiment, the combustor-turbine assembly a combustor, a turbine nozzle downstream of the combustor, and a first compliant dual seal assembly. The first compliant dual seal assembly includes a compliant seal wall sealingly coupled between the combustor and the turbine nozzle, a first compression seal sealingly disposed between the compliant seal wall and the turbine nozzle, and a first bearing seal generally defined by the compliant seal wall and the turbine nozzle. The first bearing seal is sealingly disposed in series with the first compression seal.

Abstract: A device for connecting a thrust reverser front frame to a fan casing includes a crenulated flange secured to the front frame, an annular component to accept this flange secured to the fan casing, and a crenulated annulus of a shape that complements the flange and pivot-mounted on the annular component. The crenulated flange, the annular component and the crenulated annulus are designed in such a way that a rotation of the annulus with respect to the annular component has the effect of locking this flange against this annular component through the collaboration between the respective crenulations of the flange and of the annulus.

Abstract: In order to provide an annular combustion chamber (3) which can be heated without thermal distortion, expands without distortion, and is reliably closed, and for repairs can be opened and permanently closed again and can be reconditioned, the inner and/or outer shell (8, 7) of an annular combustion chamber (3) is produced from a rotationally symmetrical sheet metal. For assembly, the shells are cut open along a split line (16), and the lower half of the inner and of the outer shell (8, 7) is introduced into the lower casing half of a gas turbine. After the insertion of the rotor, the two upper shell halves can likewise be inserted and be connected to the lower halves by welding. In order to improve the service life of the split line weld seam (21), film cooling of the weld seam is routed by an indentation. This may be implemented, for example, by welded-in split line elements (15) which are also suitable for retrofitting.

Abstract: A device for locking an engine on a pylon includes two systems attaching the engine to the pylon, located in attachment regions spaced out in a longitudinal direction of the engine. A second region is adjacent to a center of gravity plane of the engine and closer to a fan than a first region. A first system attaching to the first region is rigid towards torsional, transverse and vertical engine forces. The first region is located at a separation between two compressors of the engine or two turbines, whether the fan is at a rear or front of the engine respectively, the separation being adjacent to a node of a first engine flexural mode. A second system attaching to the second region is rigid towards transverse and vertical forces but less rigid than the first system. The device also includes a third system withstanding thrust forces in the longitudinal direction.

Abstract: A damping device for damping pressure oscillations within a combustion chamber of a turbine is provided. The damping device has a radially extending damping device body. The damping device body has a radially inner circumferential edge portion and a radially outer circumferential edge portion. A through hole is formed in the damping device body and has a tapered shape.

Abstract: In an SEV combustor and a method for reducing emissions in an SEV combustor of a sequential combustion gas turbine, an air/fuel mixture is combusted in a first burner and the hot gases are subsequently introduced into the SEV combustor (1) for further combustion. The SEV combustor (1) includes a chamber having a chamber wall (5) defining a mixing portion (8), for mixing the hot gases with a fuel, and a combustion region (9), at least one inlet (2) for introducing the hot gases into the mixing region (8), at least one inlet (12) for introducing a fuel into the mixing region (8) and at least one inlet (10, 13) for introducing steam into the mixing region.

Abstract: An example engine assembly includes a first attachment structure secured to an engine casing or an engine liner, and a second attachment structure secured to the other of the engine casing or the engine liner. A sliding member is held by the first attachment structure and is slideable relative to the first attachment structure between a first position and a second position. A pin structure moves with the sliding member between the first position and the second position. A link is pivotally connected to the second attachment structure and the pin structure.

Abstract: An aircraft turbomachine with a housing, and first and second fuel injection systems, where each system includes an injector fitted with an attachment plate traversed by three mounting holes positioned according to a first configuration for the first system and according to a second configuration distinct from the first for the second system. Each system also includes a reception device installed on the housing, a sealing joint interposed between the plate and the reception device, and securing screws which attach the plate on to its reception device. Sealing joints are identical for the first and second fuel injection systems.

Abstract: A fixture assembly for securing at least one side rail to a shroud tile of a gas turbine is disclosed. The fixture assembly may include a first support member, a second support member, a first guide block and a second guide block. The first and second guide blocks may be formed from a material that has the same or a similar coefficient of thermal expansion as the material used to form at least one of the shroud tile and the at least one side rail. In addition, the first and second support members may be formed from a material that has a lower coefficient of thermal expansion than the material used to form the first and second guide blocks.

Abstract: A turbine shroud hanger apparatus for a gas turbine engine includes: an arcuate shroud hanger having at least one cooling hole passing therethrough, the cooling hole having an inlet and an outlet; and a filter carried by the shroud hanger positioned upstream of the inlet of the cooling hole, the filter having a plurality of openings formed therethrough which are sized to permit air flow through the cooling hole while preventing the entry of debris particles larger than a preselected size into the cooling hole.

Abstract: A mid-turbine frame is incorporated into a turbine section of a gas turbine engine intermediate a high pressure turbine and a low pressure turbine. The high pressure and low pressure turbines rotate in opposite directions. The mid-turbine frame carries a plurality of vanes to redirect the flow downstream of the high pressure turbine as it approaches the low pressure turbine. In another feature, a power density is defined as the thrust divided by the volume of a turbine section, and the power density is of about 1.5 lbf per in3.

Abstract: An arrangement of components (12, 13, 30) for a wind power plant. A first component (12, 13) has a flange (15, 130) having a flange contact surface (121, 122, 133). A second component (30) has a flange-mounting surface (131, 132) for a flange (15, 130) of the first component (12, 13). Alternatively, the second component (12, 13) has a flange (15, 130) having a flange contact surface (122, 133) and the flange contact surface (122, 133) of the first component (12, 13) and the flange-mounting surface (131, 132) of the second component (30) are arranged opposite each other, or the flange contact surfaces (122, 121) of the components are arranged opposite each other. At least one flange contact surface (121, 122, 133) of a flange (15, 130) and/or the flange-mounting surface (131, 132) of the second component (30) has an outer coating made of a chrome-steel alloy.

Abstract: A near flow path seal for a turbomachine includes a support member having a first end portion that extends to a second end portion through an intermediate portion, and an arm member extending from the first end portion of the support member. The arm member includes a first end that extends to a second end, a first surface extending between the first and second ends, and a second, opposing surface extending between the first and second ends. The second surface includes a plurality of surface features that extend from about the first end toward the second end.

Abstract: A rotating machine has a barrel casing and at least one component carrier disposed therein. The component carrier is aligned and supported within the barrel casing using a plurality of carrier alignment fixtures which have alignment shafts that extend through the barrel casing to the component carrier. Each alignment shaft has an eccentric key pin extending from a distal end thereof and an alignment key rotatably-mounted to each eccentric key pin. The alignment keys mate with corresponding keyway slots defined on the component carrier such that when the carrier alignment fixtures are rotated axially, the corresponding alignment keys bias against the keyway slot and shift the position of the component carrier.

Abstract: A hyperstatic truss including connecting rods, used for suspension of a first ring, forming part of an engine case, inside a second ring concentric to the first ring, the connecting rods being secured at one end to the first ring and at the other end to the second ring. The tensile stiffness of the connecting rods is greater than the compressive stiffness thereof. The truss for example can be used for suspension of a ducted-fan turbine engine with an elongate bypass duct.

Abstract: An exhaust strut is provided and includes a body having an airfoil-shaped cross-section defining a lead edge portion and a trailing edge portion opposite the lead edge portion, the lead edge portion and the trailing edge portion being connected by a pressure side and a suction side opposite the pressure side, at least the lead edge portion and respective sections of the pressure side and the suction side proximate to the lead edge portion being formed of shape memory alloy, and a temperature control system operably disposed at the lead edge portion and the respective sections of the pressure side and the suction side proximate to the lead edge portion to modify a temperature of the shape memory alloy.

Abstract: A turbo machine is provided that includes a turbine nozzle having a wall. The wall provides spaced apart first and second surfaces. A combustor includes a liner having an annular lip that engages one of the first and second surfaces. A floating ring seal is supported by the liner in a slip-fit relationship, for example, using a retainer. The floating ring seal engages the other of the first and second surfaces. The floating ring seal is slidably moveable relative to the liner in a floating direction in response to movement of the liner in a radial direction.

Abstract: A gas turbine system comprising, a diffuser operative to diffuse an airstream output from a compressor, a fuel nozzle operative to receive fuel and emit the fuel in a combustor, and at least one bleed duct operative to direct bleed air from down stream of the combustor to the fuel nozzle.

Abstract: An anti-vibration mount is provided for mounting a first component to a second component. The mount has an elastomeric body which provides a recess into which the first component is received. The mount further has pair of brackets which fit to opposing sides of the elastomeric body sandwiching the first component received in the recess therebetween. At least one of the brackets is arranged to connect the anti-vibration mount and second component together. The mount further has a clamping arrangement which applies clamping pressure across the brackets and thereby compresses the elastomeric body to secure the first component in the recess.

Abstract: A raft assembly for a gas turbine engine is provided. The raft assembly includes a rigid raft formed of a rigid material that has an electrical system and/or a fluid system embedded therein. The raft assembly further includes one or more clamps for mounting tubular members to the raft. The or each clamp has a first clamp block and a second clamp block which, in use, clamp together to grip a tubular member between the blocks. The first block is fixed to the raft. The or each clamp further has a fastener operatively extending between the blocks.

Abstract: The present invention provides a rigid raft formed of rigid composite material. The raft has an electrical system and/or a fluid system embedded therein. The raft further has a tank for containing liquid integrally formed therewith. The tank can be formed of the rigid composite material. The tank can be for a gas turbine engine.

Abstract: Accessories are mounted to a gas turbine engine from a frame attached to a pylon which is used to mount the engine to an airframe. The frame independently mounts the accessories so that engine vibrations are not transmitted to the accessories.

Abstract: The present invention provides a gas turbine engine part which has a primary purpose in the engine which is structural and/or aerodynamic. The part is formed of rigid composite material, and has an electrical system comprising electrical conductors permanently embedded in the composite material. This provides advantages in terms of weight, complexity, and build time.

Abstract: An electrical assembly 600 comprising an electrical raft 200 and an electronic unit 300 is provided to a gas turbine engine 10. The electrical raft 200 has electrical conductors 252 embedded in a rigid material 220, which may be a rigid composite material. The electrical conductors 252 are in electrical contact with the electronic unit 300. When the electronic unit 300 is installed, at least a part 310 of it forms a part of a gas-washed surface of the engine 10. The electronic unit 300 is then easily accessible from the engine 10, and potentially complex and/or heavy access doors/panels may not be required.

Abstract: A combustion chamber in which the injection system is attached so as to prevent any axial motion of the injection system, the combustion chamber includes: a baffle including a tubular portion having a first upstream end surrounded by a first radially projecting collar; and an injection system including a bowl which includes a cylindrical portion inserted into the tubular portion, the cylindrical portion including a second upstream end surrounded by a second radially projecting collar, the second collar axially bearing against the first collar, wherein the combustion chamber includes a clamp which axially clamps the first and the second collars against one another, and a fastener which retains the clamp clamped on either side of the first and second collars.

Abstract: A closing system for a fan cowling on a jet engine of an aircraft that has at least one assembly including a male contrivance integral with the cowling, and a female contrivance integral with the jet engine that has a lower element forming at least one first guide ramp and an upper element forming at least one second guide ramp, with the aforementioned female contrivance being suitable on the one hand for guiding the male contrivance during the opening and/or the closing of the fan cowling, and on the other hand providing a limitation of the radial and circumferential displacements of the male contrivance.

Abstract: A turbine system is disclosed. In one embodiment, the turbine system includes a transition duct. The transition duct includes an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The transition duct further includes an interface member for interfacing with a turbine section. The turbine system further includes a leaf seal contacting the interface member to provide a seal between the interface member and the turbine section.

Abstract: In a gas turbine engine combustor and method for delivering purge gas thereto, a ferrule is generally coupled to the housing and moveable relative thereto. The ferrule has a primary opening through which a combustor component extends into a combustion chamber of the combustor. The ferrule further has a plurality of purge gas openings separate from and in transversely spaced relationship with the primary opening to allow purge gas to flow through the ferrule. In a first position of the ferrule at least one purge gas opening is blocked against the flow of purge gas therethrough and at least one other purge gas opening is unblocked against purge gas flow. In a second position of the ferrule at least one of the blocked purge gas openings of the first ferrule position is unblocked to permit the flow of purge gas therethrough.

Abstract: A system is provided for mounting fuel injectors to a gas turbine engine. The system comprises an engine casing having a plurality of apertures formed therein. The system further comprises a plurality of first and second fuel injectors. Each fuel injector has a respective flange for mounting the fuel injector to the casing at a respective aperture so that the fuel injector extends into the engine. The first fuel injectors have flanges which are dismountably sealed to an inner side of the casing. The second fuel injectors have flanges which are dismountably sealed to an outer side of the casing. The flanges of the first fuel injectors are configured to allow them to pass through the apertures of the second fuel injectors. A first fuel injector can be dismounted from the casing and withdrawn therefrom through the aperture of a dismounted second fuel injector.

Abstract: A support assembly for a turbine system is disclosed. The support assembly includes a transition duct extending between a fuel nozzle and a turbine section. The transition duct has an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis. The support assembly further includes a plurality of protrusions extending from the transition duct. The plurality of protrusions are configured to allow movement of the transition duct about at least one axis.

Abstract: A gas turbine engine including a rotor is disclosed. The rotor includes a stud extending along an axis, rotating elements of a first section, and rotating elements of a second section. The stud includes a first and second external end, the first external end adapted to engage a first pre-load nut or a shaft and the second external end adapted to engage a second pre-load nut or a shaft such that the set of rotating elements are secured. Thus stud includes a first shank connected to the first external end and a second shank connected to the second external end. The first shank is located in the first section and has a first diameter. The second shank is located in the second section and has a second diameter which is greater than the first diameter.

Abstract: An engine mounting apparatus having a yoke carrier block and a main mount block, the yoke carrier block being connectable to the main mount block, the main mount block being connected to the engine by at least one engine attachment link for transferring loads from an engine in a normal load path during normal operation of the engine mounting arrangement, characterized in that the yoke carrier block is connected to a further engine attachment link independently of the main mount block, the further engine attachment link being unloaded during normal operation of the engine mounting arrangement and arranged to transfer loads from the engine in the event of a failure of one or more components in the normal load path.

Abstract: A rotational system for a miniature gas turbine engine includes a rotor shaft rotationally mounted to a forward bearing and an aft bearing. The rotor shaft is located axially with a single threaded rotor nut which provides an axial preload maintained by a fixed compressor wheel and spacers. A preload spring provides dynamic compensation for tolerance variation and undesirable axial movement during rotation of the shaft. A permanent magnet generator is mounted to the rotor shaft behind the forward bearing to generate electrical power for the engine and other accessories. The rotor shaft is inertia welded to a turbine wheel to form a cavity which provides a thermally conductive path away from the aft bearing. A seal ring mounted within a groove within the rotor shaft provides a pneumatic seal to further protect the aft bearing from thermal transfer between the turbine section and the compressor section.

Abstract: A gas turbine engine installation is provided that has a plurality of flexible printed circuit board (FPCB) harnesses to transfer electrical signals, including electrical power, around a gas turbine engine. The plurality of FPCB harnesses is held to the gas turbine engine installation using generally U-shaped clips that have a base from which first and second sidewalls extend. The FPCB harness is held between the tips of base-teeth extending from the base, and side-teeth extending from the first and second sidewalls. In this way, the FPCB harness can be held securely in position by the clip.

Abstract: A shaft cover support for a gas turbine engine is disclosed. The shaft cover support not only provides enhanced support to a shaft cover of the gas turbine engine, but also includes a cooling fluid chamber for passing fluids from a rotor air cooling supply conduit to an inner ring cooling manifold. Furthermore, the shaft cover support may include a cooling shield supply extending from the cooling fluid chamber between the radially outward inlet and the radially inward outlet on the radially extending region and in fluid communication with the cooling fluid chamber for providing cooling fluids to a transition section. The shaft cover support may also provide additional stiffness and reduce interference of the flow from the compressor. In addition, the shaft cover support accommodates a transition section extending between compressor and turbine sections of the gas turbine engine.

Abstract: The present invention relates to an open rotor type aircraft turbine engine (1), comprising a contrarotating propeller receiver (30) and a dual-body gas generator (14) comprising a low-pressure compressor (16) and a high-pressure compressor (18) separated by an intermediate housing (27), said gas generator being arranged upstream from said receiver. According to the invention, the turbine engine further comprises a structural casing (50) for supporting the receiver (30), surrounding the gas generator (14) and having a downstream end (50a) attached to said receiver and an upstream end (50b) attached to said intermediate housing (27). Furthermore, it comprises additional connection means (60) between said structural supporting casing and the gas generator, arranged between the upstream and downstream ends (50b, 50a) of the casing.

Abstract: A shaft cover support for a gas turbine engine is disclosed. The shaft cover support not only provides enhanced support to a shaft cover of the gas turbine engine, but also includes a cooling fluid chamber for passing fluids from a rotor air cooling supply conduit to an inner ring cooling manifold. As such, the shaft cover support accomplishes in a single component what was only partially accomplished in two components in conventional configurations. The shaft cover support may also provide additional stiffness and reduce interference of the flow from the compressor. In addition, the shaft cover support accommodates a transition section extending between compressor and turbine sections of the engine. The shaft cover support has a radially extending region that is offset from the inlet and outlet that enables the shaft cover support to surround the transition, thereby reducing the overall length of this section of the engine.

Abstract: The invention relates to an exhaust system (100) for channeling the streams from a by-pass gas turbine, the system comprising a stream channeling nozzle and an exhaust casing (110) for connecting the channeling nozzle to the outlet of the gas turbine, said channeling nozzle comprising a primary nozzle (120) fastened to the exhaust casing (110) and a secondary nozzle (130) placed around the primary nozzle. The exhaust system further comprises means for fastening the secondary nozzle (130) directly to the exhaust casing (110), said secondary nozzle being supported by the exhaust casing independently of the primary nozzle (120).

Abstract: A gas turbine engine includes first and second stages having a rotational axis. A mid turbine frame is arranged axially between the first and second stages. The mid turbine frame includes a circumferential array of airfoils, and the airfoils each have a curvature provided equidistantly between pressure and suction sides. The airfoils extend from a leading edge to a trailing edge at a midspan plane along the airfoil. An angle is defined between first and second lines respectively tangent to the intersection of the plane and the curvature at airfoil leading and trailing edges. The angle is equal to or greater than about 10°, for example. In one example, an airfoil aspect ratio is less than 1.5.

Abstract: An air cycle machine includes two turbines and a compressor mounted on an integral shaft. The integral shaft includes a plurality of shaft sections that are welded together and machined in a single set-up process into a desired shaft shape to provide highly aligned bearing surfaces. The shaft includes three stops that cooperate with three fasteners to secure the two turbines and the compressor on the shaft.

Abstract: An exemplary turbine engine link assembly includes a link extending longitudinally from a rod end and terminating at a hemispherical end. The rod end is secured to an engine liner or an engine casing. The hemispherical end is biased toward a corresponding hemispherical recess in the engine liner or in the other of the engine liner or the engine casing.

Abstract: In order to reduce a manufacturing cost and a running cost, casings each housing a combustion oscillation detecting device are attached in an alternating manner to cross-flame tubes by which a plurality of combustors arranged annularly in a combustor casing are connected to each other.

Abstract: A bypass turbojet engine including an outer fan duct and an inner fan duct (IFD), between which the secondary flow passes, and a high pressure compressor having an outer casing is disclosed. The IFD is fastened on the outer casing.

Abstract: A transition duct having a thermally free aft frame and being capable of adjusting the natural frequency is disclosed. The aft frame is capable of permitting movement due to thermal gradients with the transition duct. The transition duct utilizes a spring plate located adjacent to an aft mounting bracket, where the spring plate, based on its thickness can either increase or decrease a frequency of the transition duct. Such an arrangement ensures that the transition duct natural frequency does not coincide with or cross other critical engine and/or combustor frequencies.

Abstract: A device is provided for the mutual fixing of two components, which are interconnected via a flanged joint, for preventing a relative movement of the components in relation to each other in the connecting plane of the flanged joint. The device includes a key which engages in both components in a form-fitting manner in the region of the flanged joint. A simplified and flexible application is achieved by the key being guided in one of the components by means of a precision fit, whereas in the other of the components the form fit can be adjusted by adding shims.