Abstract: A rotatable turbine frame includes annular outer and inner bands having a plurality of circumferentially spaced apart struts extending therebetween. An annular drive shaft is fixedly joined to the inner band for transmitting output torque therefrom, and the struts are backwardly radially inclined relative to the direction of rotation of the frame so that gas flow between the struts tends to straighten the inclined struts to effect a compressive load component therein.

Abstract: A hybrid gas turbine engine rotor blade includes an airfoil and a dovetail integrally joined thereto. The airfoil includes an inner portion disposed adjacent to the dovetail to provide a structural transition, and an outer portion disposed thereabove for effecting a pressure rise in airflow channeled thereacross. A nosepiece is disposed adjacent to the leading edge of the airfoil inner portion, and a tailpiece is disposed adjacent to the trailing edge of the airfoil inner portion. The nosepiece, tailpiece, and airfoil inner portion are collectively aerodynamically configured for effecting a pressure rise in the airflow channeled thereacross.

Abstract: A gas turbine engine rotor support system includes outer and inner rotors having respective outer and inner blades interdigitated in respective blade row stages. A stationary rear frame includes a rear support shaft. A rotatable aft frame is disposed forwardly of the rear frame includes an aft support shaft. And, a rotatable forward frame is disposed forwardly of the aft frame and includes a forward support shaft. The forward and aft frames support the outer and inner rotors, and a first bearing is disposed between the aft shaft and the rear shaft for supporting the aft shaft; a second bearing is disposed between the forward shaft and the aft shaft for supporting the forward shaft; and a third bearing is disposed between a core shaft of a core turbine and the forward shaft for supporting the core shaft thereon.

Abstract: A method and assembly are effective for air-starting an aircraft gas turbine engine having a fan powered by a low pressure turbine through a first shaft, and a compressor powered by a high pressure turbine through a second shaft disposed coaxially therewith. A gear train is selectively operatively joined between the first and second shafts by selectively engaging a first clutch for transmitting torque through the first clutch only in one direction from the first shaft to the second shaft when the fan is windmilling for driving the second shaft to allow an air-start of the gas turbine engine during flight. In a preferred embodiment, a second clutch is operatively joined to the gear train and is selectively engageable at speeds of the second shaft below a predetermined release speed, and disengageable at the release speed and above for disconnecting the air-start assembly once the engine has been started.

Abstract: A counterrotating turbine support assembly includes outer and inner rotors having respective outer and inner blades disposed in an outer casing. A mid-frame is disposed forwardly of the rotors and fixedly joined to the outer casing. A rear frame is disposed axially aft of the rotors and is fixedly joined to the outer casing. A forward support shaft is joined at one end to a forward end of the outer rotor and is rotatably supported at an inner end to the mid-frame by a forward bearing. An aft support shaft is joined at one end to an aft end of the outer rotor and is rotatably supported at an inner end to the rear frame by an aft bearing. An intermediate support shaft is fixedly joined at one end to the inner rotor and is rotatably supported at an inner end to the mid-frame by an intermediate bearing.

Abstract: A lubrication system for a gas turbine engine includes first and second coaxially rotor shafts having at least one differential bearing disposed in outer and inner seats thereof. An annular first scoop extends axially from a first portion of the first shaft, and an annular second scoop extends axially from the inner seat of the second shaft. A plurality of first holes extend axially through the first shaft and radially below the first scoop, and a plurality of second holes extend axially through the inner seat of the second shaft and below the second scoop. An annular first shell extends axially between the second scoop and the first shaft from above the first holes. And, oil is injected under the first scoop for flow by centrifugal force from rotation of the first shaft through the first holes and along the first shell into the second scoop for flow through the second holes to lubricate the bearing.

Abstract: A combustion assembly includes a combustor having inner and outer pilot liners, each being in the form of a lobed mixer having outer and inner cold and hot chutes. A plurality of carburetors are joined to a dome disposed at upstream ends of the liners for providing a pilot fuel/air mixture for generating pilot combustion gases in the hot chutes. A plurality of fuel spraybars are disposed downstream from the carburetors and are aligned radially with the cold chutes for selectively injecting main fuel into main airflow for generating a main fuel/air mixture ignitable by the pilot combustion gases. In a preferred and exemplary embodiment of the invention, lean combustion gases are obtained for reducing NO.sub.x emissions in a relatively short residence time.

Abstract: A clearance control assembly in a gas turbine engine includes several stages of cylindrical turbine shroud rings and a one-piece continuous cylindrical housing surrounding the shroud rings in radially outward spaced relation. Each shroud ring is composed of at least a pair of semi-cylindrical segments mounted to the housing. The shroud ring segments are radially positioned by thermal expansion and contraction of the housing in response to temperature-controlled air flow between the housing and shroud segments. Coupling elements joins together opposite edge portions of adjacent shroud ring segments to permit axial movement of the shroud segments toward and away from one another. Lands which protrude outwardly from the housing interior and shroud ring segment exteriors are located between the housing and shroud segments for defining several spiral channels for air flow about the interior of the housing and exterior of the shroud segments.

Abstract: A gas turbine engine incorporates a bearing support assembly which mounts an inner annular bearing arrangement in radially spaced relation inwardly from a stationary outer casing. The bearing arrangement, in turn, mounts a turbine rotor within the outer casing for rotation about a central axis relative to the outer casing. A row of stationary stator vanes are fixedly attached to an extend radially inwardly from the outer casing toward the central axis.

Abstract: A turbine assembly includes a unitary rotor spool having first and second rotor discs. A set of first blades having axial entry dovetails are joined to the first disc, and a set of second blades having axial entry dovetails are joined to the second disc. A unitary annular spacer is disposed axially between the first and second dovetails for preventing the first and second blades from sliding axially toward the spacer. The spacer has an inner diameter greater than an outer diameter of the first rotor disc for allowing the spacer to be assembled over the first rotor disc during assembly.

Abstract: A combustor includes spaced outer and inner liners defining therebetween a combustion zone, and having a plurality of circumferentially spaced tubular domes disposed between forward ends thereof. Each of the domes includes an annular inlet for receiving pilot airflow, a swirler having a tubular ferrule and primary swirl vanes for receiving the pilot airflow, and an outlet disposed in flow communication with the combustion zone. A pilot fuel injector is disposed in each of the domes for injecting pilot fuel into the pilot airflow for generating a pilot fuel/air mixture dischargeable from the dome outlet. A plurality of main fuel spraybars inject main fuel into compressed airflow channeled between the domes and the outer and inner liners for generating a main fuel/air mixture dischargeable concentrically around the pilot fuel/air mixture.

Abstract: An apparatus for controlling clearance between rotor blade tips and shroud segments of a turbine engine casing includes an annular channel extending about the blade tips and having an annular opening located outwardly of and adjacent the blade tips, and separate shroud segments composing the shroud and disposed across the channel opening in circumferential spaced relation. Inlet and outlet tubular members mounted to the casing have inner portions projecting radially inwardly into the annular channel. An annular hollow ring disposed in the annular channel mounts the shroud segments across the channel opening and are mounted to the tubular member inner portions for undergoing movement radially outwardly and inwardly relative to the inner portions in response to thermal expansion and contract of the annular ring. Tubular inlet members introduce a cooling gas into the annular channel from the exterior and tubular outlet members exhausts the cooling gas from the annular channel to the exterior.

Abstract: A mechanical clearance control apparatus is operable for controlling the clearance between the rotor blade tips and the casing shroud of a gas turbine engine. The apparatus includes a shroud segment which defines a circumferential portion of the casing shroud and is disposed in an opening defined by a mounting section of the stationary casing, a channel defined between radially spaced portions of the shroud segment and the stationary casing mounting section, and a wave spring disposed in the channel and preloaded against the shroud segment and mounting section for biasing the shroud segment to move radially inward toward the rotor. Flanges on opposite sides of the shroud segment and on the stationary casing mounting section at opposite sides of the opening cooperate to define inner and outer limits of radial movement of the shroud segment as the shroud segment is moved toward and away from the rotor.

Abstract: A support bearing assembly for supporting a jet aircraft engine high speed rotor shaft where an operating temperature of up to 1500 degrees F. is involved and the bearing unit is exposed to such temperature condition, includes a rolling elements-inner and outer race bearing unit capable of operating at such temperature level without adverse effect and, e.g., these components being comprised of ceramic matrix composites. Since such a bearing unit will have different thermal growth than its metallic support housing, the bearing unit is supported in pairs of opposed cylindrical cages, there being an inner cage pair for holding the inner race and an outer cage pair for holding the outer race. The cages have flange means at one end and circularly, spaced support arms extending from the flanges a distance to cage opposite end locations where the support arms are formed as finger structures which interfit with the race members, the spacing between the support arms defining slots in the cage bodies.

Abstract: A gear-actuated mechanical clearance control apparatus is operable for controlling the clearance between rotor blade tips and the casing shroud of a gas turbine engine. The apparatus includes a pair of annular rings which extend about the casing shroud, a shroud hanger disposed between the casing and an opening in the casing and between the annular rings, and a shroud segment defining a circumferential portion of the casing shroud and disposed in the casing opening. The apparatus also has an actuating mechanism which includes driven gear sectors on the annular rings which face toward one another and a gear disposed between the annular rings and having a peripheral driving gear sector which intermeshes with the driven gear sectors of the annular rings. Upon rotation of the gear, the annular rings undergo opposite circumferential movement causing radial movement of the shroud hanger and the shroud segment therewith to establish the clearance between the shroud segment and rotor blade outer tips.

Abstract: A lightweight rigid turbine engine frame incorporates a one piece annular cast outer casing formed with a circumferential polygonal reinforcing section. The outer casing is separated and spaced from an inner ring by a plurality of radially extending spacer struts. The spacer struts are clamped in compression between the inner ring and outer casing by tensioning members such as bolts. A thermally free floating heat shield is mounted to the bolts in a manner which minimizes thermal stresses by avoiding all undesirable constraint of the heat shield by the engine frame.

Abstract: A jet engine subassembly. The high pressure (HP) turbine includes a ball bearing assembly whose inner race is attached to the HP turbine shaft. The outer race takes radial loads through radially-extending tie rods which pass through the HP turbine nozzle vanes and are connected to the combustor's outer casing. The outer race takes axial loads through a support cone which is attached to the combustor diffurser's radially inner duct through an intermediate forward segment of the combustor inner casing. This eliminates the need for a heavy turbine frame. The ball bearing is the thrust bearing, and does not experience thermal axial growth. This allows the use of unshrouded sloped HP turbine blades.

Abstract: An annular cast inner support member is concentrically mounted within an annual cast outer support member for mounting a plurality of radially extending outlet guide vanes within the flowpath of a gas turbine engine. The outlet guide vanes are resiliently connected to the inner and outer support members to control and limit the compressive loading of the outlet guide vanes caused by thermal stresses during engine operation. A resilient clamped interconnection which is particularly suitable for use with ceramic outlet guide vanes includes a bolted and clamped socket assembly which maintains both radial and circumferential compressive forces on the outlet guide vanes at all times, notwithstanding the differences in thermal growth rates between metal and ceramic materials.

Abstract: A bypass turbofan jet engine subassembly. A row of structural fan stators and a row of structural compressor guide vanes have their outer ends attached to the engine outer duct. The structural stator inner ends are attached by an inner support cone to the outer race of the fan rolling element bearing and the structural guide vane inner ends are attached by another inner support cone to the outer race of the compressor rolling element bearing. The fan and compressor outer races are interconnected by a tie bar. The subassembly eliminates the engine fan and compressor frames.