Abstract: An improved brush seal for reducing swirl, recirculation and turbulence of an approaching gas stream to reduce irregular bristle wear and operating costs. The brush real may include a baffle of the honeycomb type or circumferentially stacked plates type, positive or negative steps in the flow path approaching a seal stage, and recirculation breakers.

Abstract: A split ring seal having an outer split ring and an inner split ring with the inner and outer split rings having overlapping portions. The split ring seal is located between a seating seal and a carrier and separates a high pressure cavity from a low pressure cavity. Movement of the split ring seal is restricted by sidewalls of the carrier. A diagonal edge on the outer ring reacts against centrifugal loading from the inner ring so that said outer ring is urged axially against a sidewall of the carrier to form a positive seal for blocking any leakage path between the high and low pressure cavities.

Abstract: A turbine frame includes first and second coaxially disposed rings having a plurality of circumferentially spaced apart struts extending therebetween. A plurality of clevises join respective first ends of the struts to the first ring for removably joining the struts thereto. Each of the clevises includes a base removably fixedly joined to the first ring, and a pair of legs extending away from the base and spaced apart to define a U-shaped clevis slot receiving the strut first end. The first strut end is removably fixedly joined to the clevis legs by a pair of expansion bolts. The clevis base includes a central aperture aligned with a first port in the first ring for providing access therethrough. And, the strut first end abuts the first ring for carrying compressive loads therebetween and providing a seal therewith.

Abstract: The present invention provides an axial blade to disk retaining assembly having a plurality of circumferentially disposed lugs arranged in a ring around an axial facing surface of the blade and disk assembly and a plurality of circumferentially disposed hooks arranged in a ring around the axial facing surface and in circumferential alternating positions with the hooks and radially spaced apart from and open towards the lugs. A split retaining ring is provided that is operable to be placed between said hooks and lugs to axially secure the blades on the disk. In the preferred embodiment each hook is disposed on a disk post between the blades and each lug is disposed on a blade root.

Abstract: A tube fitting for use with a flammable fluid is disposed in an auto-ignition zone of a gas turbine engine. The tube fitting has a pair of separable casings defining a passage to accommodate flow therethrough of the flammable fluid, and the casings have a pair of seals associated in sealing engagement with each other to seal about the passage. A restriction is arranged in one of the casings for limiting at least to a preselected flow rate any leakage flow of the flammable fluid from the passage past the seals upon an interruption in the sealing engagement thereof in the event of the occurrence of the separation of the casing.

Abstract: A sealing device for compressor axial dovetail rotor blades for a gas turbine engine comprising a sealing member having a finger portion extending from a main body portion which engages a disk hook located on each disk post and forms a first point of contact about which the sealing member rotates. The sealing member further comprises an extended skirt portion which engages the blade platform of the rotor blades forming a second point of contact. The main body has a center of gravity positioned such that during operation of the engine the main body pivots due to centrifugal force about the disk hook such that the skirt portion engages the blade platform. A split ring is located between the disk posts and a tail portion extending from the lower portion of the sealing member and engages the disk post and tail portion during operation of the engine thereby forming a third point of contact.

Abstract: A turbine engine of the type having a compressor section, combustor section and high-pressure turbine section in which compressor air is conveyed internally of the compressor and turbine sections to cool the turbine interstage volume, a first stage turbine disk includes a shaft supporting a second stage turbine disk, a turbine interstage seal is secured by boltless engagement to the disks and is prestressed to resist flexure, and the first stage disk carries a forward seal attached by boltless connections. The interstage seal includes a bore having a rearwardly-extending arm engaging the aft shaft in a bayonet connection to prevent deflection of the bore and includes passages to allow air flow around the bore. The forward seal includes radially-extending vanes to convey cooling air along the disk, and is secured by inward and outward bayonet connections, the inward connection including locking pins which are arranged to perform a balancing function.

Abstract: A slip joint connects coaxially oriented tubular members in a manner to allow axial slippage while maintaining concentricity without generating large axial loads. The joint comprises a first tubular member having a smaller outside diameter than an inside diameter of a second member whereby the first member is slidable into the second member. A pair of spaced lands circumscribe the first member and define a groove therebetween. The slip joint has a seal ring position in the groove. The lands have a plurality of circumferentially spaced nodes extending radially outward so as to contact the inner diameter of the second member.

Abstract: The present invention provides a phase independent damper and damper assembly capable of damping rotor blade vibrations in the axial, circumferential, and radial directions. One embodiment, particularly useful in an aircraft gas turbine engine compressor rotor, provides a damper and damper assembly including a generally axially extending blade damper operable to fictionally engage a generally axially extending circumferentially facing surface of a disk and axially extending angled surface under the platform of the blade. The invention provides an additional advantage of allowing simple and easy modification of existing engines to incorporate the damper of the present invention.

Abstract: A seal assembly for use in a compressor stage of a turbine engine having a rotor disk which is attached to a plurality of rotor blades. Each rotor blade has an axially oriented dovetail attachment which connects to the rotor disk. Each rotor blade has an aftward side which is in contact with a high pressure region and a forward side which is in contact with a low pressure region. A plurality of seal segments form an annular ring located radially inward from the plurality of rotor blades. Each seal segment has an offset center of gravity which causes each seal segment to be securely attached to the rotor disk as centrifugal forces act upon each seal segment during rotation.

Abstract: A method and system for diverting leakage air back into the flow path of a turbine engine. A stator vane assembly is connected to a shroud assembly at the radially inner end of the stator vane assembly, the shroud assembly is provided with a scoop which is placed in the path of leakage air traversing in a forward direction from the high pressure static side of the stator vane to the low static pressure side of the stator vane. The leakage path is located between the stator vane assembly and a rotating member. The scoop intercepts the leakage air and re-directs the leakage air into an airflow path of the turbine engine with an aftward component of velocity.

Abstract: A fan blade damper system includes a fan disk integrally connected to a disk post. A retainer is fastened to the disk post and is molded to retain a damper block. The top portion of the damper block is shaped to make frictional contact with the underside of the platform of a fan blade, the underside of the platform being the region of the fan blade which experiences the highest platform vibratory amplitude when the fan blade is rotated. The damper block is thus mounted in a statically determinant fashion. The rotation of the fan blade produces centrifugal forces which produce reactionary forces in the damper block and retainer such that a maximum damping force amplitude is reacted to the fan disk.

Abstract: A rotor assembly for supporting rotor blades includes a plurality of axially adjoining discrete disks each having a rim, web, and hub. The rims include axial dovetail grooves for collectively supporting a respective blade dovetail therein. Each of the disks includes at least one axially extending annular arm joined integrally with the disk and fixedly joined to an adjacent one of the arms. The disks are fixedly joined together solely by the blade dovetails in the rim dovetail grooves and by the arms so that upon a crack failure of one of the disks, centrifugal load from the failed disk is transferred to an adjacent disk.

Abstract: A process for identificaiton, evaluation and removal of microshrinkage of investment cast superalloy parts and parts produced by the process. Parts free of deleterious microshrinkage capable of longer life at existing stress levels or use at higher alternating stresses are produced by first subjecting the parts to hot isostatic pressing to eliminate nonsurface connected subsurface microshrinkage. Next, in order to expose near-surface microshrinkage microscopically connected to the surface, the parts are immersed in an acid solution for a time sufficient to uniformly remove at least about 0.005 inches of the original as cast surface. The exposed microshrinkage is evaluated for acceptability or removal by standard nondestructive techniques, such as liquid penetration evaluation.

Abstract: An air channeling assembly includes an outer casing having a series of spaced apart air ports spaced radially outwardly from an inner casing to define a plenum therebetween. A series of baffles are disposed in the plenum and spaced radially inwardly from the outer casing. Each of the baffles includes an intermediate opening joined in flow communication with a respective one of the air ports, and first and second ends extending circumferentially oppositely away from the intermediate opening. The baffle provides a flow shield between the inner and outer casing around each of the air ports and channels air through the intermediate opening and respective air ports.

Abstract: A rotating labyrinth seal especially useful for effecting sealing between two plenums in aircraft gas turbine engines comprising a base and a plurality of radially-directed seal teeth rings extending circumferentially around the outer peripheral surface of the base. Each of the seal teeth rings has a body portion and a tip portion, with the body portion being substantially thicker than the tip portion. Seal teeth rings of the invention exhibit improved resistance to fatigue crack propagation compared to conventional seal teeth rings.

Abstract: A gas turbine engine having an outer annular turbine rotor, which is counterrotatable relative to an inner annular turbine rotor, includes attachment structure for attaching turbine blades in circumferential rows to the inner rotor. The attachment structure includes first and second male members of hook-like configurations spaced axially along and circumferentially extending about the rotor, and first and second grooves defined by flanges attached to the rotor and axially spaced along and circumferentially extending about the rotor. The grooves open in a common axial direction for receiving in mated relation respectively and concurrently the hook-like configured first and second male members. The attachment structure also includes several different locking arrangements for locking the male members in mated relation with the grooves to thereby restrain the turbine blade against radial and axial movements relative to the rotor.