Abstract: A link for supporting a core engine within an outer shell of a gas turbine engine is described. The link, in one form, includes an elongate rod having a first end and a second end, and first and second rod end bearings are secured at the first second ends, respectively. Each end bearing includes a bearing head and a threaded stud extending from the head. The elongate rod includes a first threaded socket located at first rod end and sized to be engaged to the first rod end bearing stud, and a second threaded socket located at second rod end and sized to be engaged to the second rod end bearing stud. The first and second tapered jam nuts are located at the first and second ends of rod. Each tapered jam nut includes a nut end having, for example, a hexagonal head and a tapered portion extending from the nut end. A bore extends through each tapered jam nut, and the bore is sized to have the stud extend therethrough. The bore has internal threads at least at the location of nut end.

Abstract: A turbojet engine includes a compressor and turbine, with a combustor disposed therebetween for generating exhaust gas. An afterburner includes an annular casing and liner spaced radially inwardly therefrom to define a cooling duct therebetween. A recoup duct is disposed in flow communication with the compressor for receiving a portion of compressed air therefrom as recoup air. A plenum is disposed at an upstream end of the afterburner casing, and includes an inlet disposed in flow communication with the recoup duct for receiving the recoup air, and an outlet disposed in flow communication with the afterburner cooling duct for discharging the recoup air therein.

Abstract: An exhaust nozzle includes a plurality of circumferentially spaced apart exhaust flaps, and a plurality of scissor linkages joined to respective ones thereof. Each linkage includes a frame pivotally joined to a respective one of the flaps. A carriage is slidably mounted to the frame. A pair of circumferentially oppositely extending side links are pivotally joined at one end to the carriage, and pivotally joined at opposite ends to corresponding side links to form an interconnected ring thereof. A pair of circumferentially oppositely extending equalizer links are pivotally joined at one end to the frame, and pivotally joined at opposite ends to corresponding ones of the side links to maintain coordinated angular position of the side links pivotable on the carriage, and in turn maintain circumferential spacing between adjacent ones of the flaps.

Abstract: The present invention provides coolable exhaust nozzle flap and seal assemblies having liners that are spaced inward and apart from their respective flap and seal cool walls, to form a cooling passage therebetween that converges in the axial downstream flow direction. One embodiment provides single wall multi-hole film cooled nozzle flap and seal liners particularly for use individually or in combination on convergent and divergent hollow flaps and seals for variable exhaust nozzles. Flap and seal coolable assemblies may be unitary structures that are integrally fabricated or formed such as by casting.

Abstract: A gas turbine engine fuel injection assembly includes a combustor case surrounding a combustor having a dome. An arcuate manifold is disposed inside the combustor case adjacent to the dome, and includes a plurality of fuel injectors disposed in flow communication therewith for receiving fuel for injection into the combustor. An inlet stem is joined in flow communication with the manifold for channeling fuel thereto, and includes an inner fitting disposed in a mounting port in the combustor case. A mounting adaptor is joined to the combustor case at the mounting port, and circumferentially engages the inner fitting to restrain torsional movement thereof. An outer fitting extends in the adaptor and engages the inner fitting in flow communication therewith for supplying fuel thereto.

Abstract: A stiffened composite is provided exhibiting multi-directional reinforcement with continuous long fibers. The composite has a substrate and serpentine stiffeners adhered thereto. The stiffeners merge toward, join with and diverge from adjacent stiffeners to provide a stiffener matrix on the substrate. The serpentine stiffener permit the matrix to exhibit multi-directional reinforcement with the continuous long fibers within the serpentine ribs. The composite structures are useful as aircraft bypass ducts, and as other composite products.

Abstract: A design system for designing a tube for a gas turbine engine is described. The system including, in one embodiment, a computer workstation having a plurality of executable modules stored therein. The modules include a tube design module for routing a tube between tube fittings, a tube editing module for editing tube assembly geometry, and a tube detailing module for preparing detail drawings.

Abstract: A process for periodically calibrating an algorithm in a control unit of a system for balancing loads on a thrust bearing of a gas turbine engine rotor is disclosed involving the steps of initializing a calibration of the control unit algorithm, causing the engine to attain a crossover condition, measuring the residual load on the rotor thrust bearing, calculating a residual load on the rotor thrust bearing by means of the control unit algorithm, comparing the measured residual load and the calculated residual load on the rotor thrust bearing to determine a difference therebetween, and modifying the control unit algorithm to compensate for the difference between the measured and calculated residual loads on the rotor thrust bearing.

Abstract: A rotor blade for a rotor of a gas turbine engine having an axis of rotation including a root portion, a platform portion connected to the root portion and having a damper pocket formed therein, an airfoil portion connected to the platform portion, a generally bar-shaped damping member loosely arranged in the damper pocket having at least one scrubbing surface, and at least one leg extending from the bar-shaped damping member for retaining the bar-shaped damping member in the damper pocket. The bar-shaped damping member is slidably displaceable and rotatable within the damper pocket during rotation of the rotor. The damper pocket in the platform portion has a rear surface with an upper portion and a lower portion at an angle to the upper portion, a pair of spaced side surfaces, and a pair of spaced lower surfaces extending from the rear surface lower portion which are substantially coplanar.

Abstract: A low profile joint is provided for a strut tube extending radially through a strut extending between an outer casing and an inner hub in a gas turbine engine frame. The strut tube has a longitudinal axis and a closed distal end. A sideseat is spaced from the distal end and is disposed substantially perpendicularly to the longitudinal axis to define a flow orifice. A secondary tube has a ballnose at a distal end thereof disposed in abutting contact with the sideseat for channeling fluid therebetween. A fastener joins together the strut and secondary tubes in compression between the ballnose and sideseat to maintain sealed contact therebetween for channeling fluid between the strut and secondary tubes.

Abstract: A turbine blade includes an airfoil having a root, tip cap, and laterally opposite pressure and suction sides extending between leading and trailing edges. A cooling circuit is disposed inside the airfoil and extends from the tip cap to the root for circulating a coolant. The airfoil also includes a pair of squealer tips extending radially upwardly from the tip cap along the pressure and suction sides which are spaced apart from the leading to trailing edges to define an upwardly open tip cavity. A thermal conductor is fixedly joined to the squealer tips and extends radially inwardly to the tip cap for conducting heat thereto for removal by the coolant. The conductor preferably surrounds the tip cavity inboard of the squealer tips and is coextensive therewith for being protected thereby.

Abstract: A turbulator configuration formed on an inner surface of at least one wall of a cooling passage for a rotor blade in a gas turbine engine is disclosed, where the cooling passage has an upstream end, a downstream end, and a longitudinal axis therethrough. The turbulator configuration includes a first set of longitudinally spaced turbulators having a specified height and length for tripping a thermal boundary layer along the cooling passage wall to enhance heat transfer between the coolant and the wall, as well as a second set of longitudinally spaced turbulators having a specified height and length for generating turbulence in a core flow region of the cooling passage to reduce the thickness of the thermal boundary layer downstream therefrom.

Abstract: A turbine blade includes a dovetail, shank, platform, and airfoil. A cooling circuit extends radially therethrough for circulating a coolant. A thermal conductor is disposed on a lower surface of the platform for conducting heat from the platform to the shank for removal by the coolant in the cooling circuit.

Abstract: A turbine rotor includes a rotor disk having a plurality of dovetail slots defining therebetween respective disk posts. Each post has a top land and a plurality of tangs spaced radially therebelow inside adjacent ones of the slots. A plurality of turbine blades each having a dovetail are mounted in respective ones of the dovetail slots for radially retaining the blades. A platform is joined integrally with the dovetail and extends circumferentially to adjoin adjacent ones of the platforms to define a radially inner flowpath. An airfoil extends integrally from the platform for extracting energy from combustion gases flowable thereover. A thermal barrier coating is fixedly bonded to each of the disk post top lands for providing thermal insulation between the platforms and the disk posts for reducing disk post temperature.

Abstract: A system for controlling thrust forces on a thrust bearing in a rotating structure of a gas turbine engine at designated operating points including a device for providing thrust load compensation to the thrust bearing, a control for operating the thrust load compensation device, and a sensor for detecting rotational cage speed of the thrust bearing. The sensor provides a signal to the control when the rotational cage speed of the thrust bearing drops below a specified ratio of the rotational speed for the rotating structure, the signal being indicative of an incipient skid condition for the thrust bearing. The control then causes the thrust load compensation device to provide an additional predetermined load on the thrust bearing when it receives the signal from the sensor so that a resultant load thereon is within a specified load range which extends the life of the thrust bearing.

Abstract: A turbine blade includes an airfoil having an internal cooling circuit therein. The airfoil extends from a root to a tip cap, and includes laterally opposite pressure and suction sides extending between a leading edge and an opposite trailing edge over which is flowable a combustion gas. A pair of squealer tips extend radially upwardly from the tip cap along the pressure and suction sides, and are spaced apart between the leading and trailing edges to define an upwardly open tip cavity. At least one of the squealer tips includes a slot extending radially inwardly to the tip cap, with the slot also extending along the squealer tip between the leading and trailing edges. A plurality of spaced apart supply holes extend radially through the tip cap in the slot in flow communication with the cooling circuit for channeling the coolant into the slot for cooling the squealer tip.

Abstract: A gas turbine engine fan blade includes a main body having a radially spaced apart root and tip, and axially spaced apart leading and trailing edges. A front shell is fixedly joined to a portion of the body leading edge to collectively define with the body an airfoil having first and second opposite sides extending between the root, tip, shell, and trailing edge for pressurizing air flowing thereover. The shell is hollow to offset toward the trailing edge a collective center of gravity of the airfoil to reduce steady stress at the airfoil leading edge.

Abstract: A fan blade includes a spar having first and second opposite sides. The spar sides include a fatigue-life enhanced surface layer, with the spar first side having a cavity bounded in part by the surface layer. A skin closes the cavity and is bonded to the spar for maintaining the enhanced surface layer inside the cavity.

Abstract: A washer faced spring assembly is disclosed for a motion absorbing and motion damping application which includes a first washer having a circular or polygonal hole therethrough centered about a central axis, a second washer disposed opposite to the first washer also having a circular or polygonal hole therethrough centered about a central axis disposed generally coincident with the hole of the first washer, and a spring connected between the first washer and the second washer, the spring including a plurality of fingers providing planar or full surface contact between the first and second washers. The washer faced spring assembly may also include a base portion utilized in conjunction with the plurality of fingers, where the base portion has one of several alternative shapes, such as circular or polygonal.

Abstract: A cooling passageway construction particularly adapted for cooling the interior of a heated turbine engine assembly and/or airfoil blade includes an undulating partition which subdivides the passageway into first and second subpassages. The partition defines localized minimum throat areas within each subpassage to produce localized shear jet flow between the partition and the walls of the passageway. Bleed holes may be formed through the partition to generate additional turbulence within the coolant flow and further enhance heat transfer.