Abstract: A retention system and method for the blades of a rotary machine for preventing forward or aft axial movement of the rotor blades includes a circumferential hub slot formed about a circumference of the machine hub. The rotor blades have machined therein a blade retention slot which is aligned with the circumferential hub slot when the blades are received in correspondingly shaped openings in the hub. At least one ring segment is secured in the blade retention slots and the circumferential hub slot to retain the blades from axial movement. A key assembly is used to secure the ring segments in the aligned slots via a hook portion receiving the ring segments and a threaded portion that is driven radially outwardly by a nut. A cap may be provided to provide a redundant back-up load path for the centrifugal loads on the key. Alternatively, the key assembly may be formed in the blade dovetail.

Abstract: A retention system and method for the blades of a rotary machine for preventing forward or aft axial movement of the rotor blades includes a circumferential hub slot formed about a circumference of the machine hub. The rotor blades have machined therein a blade retention slot which is aligned with the circumferential hub slot when the blades are received in correspondingly shaped openings in the hub. At least one ring segment is secured in the blade retention slots and the circumferential hub slot to retain the blades from axial movement. A key assembly is used to secure the ring segments in the aligned slots via a hook portion receiving the ring segments and a threaded portion that is driven radially outwardly by a nut. A cap may be provided to provide a redundant back-up load path for the centrifugal loads on the key. Alternatively, the key assembly may be formed in the blade dovetail.

Abstract: One form of a retention system includes a pair of aft segments of 180.degree. in length disposed in annular grooves about a compressor spool, the segments having radial projections against which the aft end faces of the base portions of rotor blades abut. The blade base portions are shaped and are received in complementary-shaped slots of the spool. The forward retention system includes a plurality of arcuate segments approximately 30.degree. in length disposed in grooves adjacent the forward end faces of the blades. With the exception of the final pair of blades, the blade base portions thereof are located in the axial complementary slots in pairs, with each forward arcuate segment, except the final segment, being disposed in the groove at a circumferentially spaced location and displaced circumferentially to abut the forward end faces. The final arcuate segment is radially reduced and disposed in the remaining portion of the groove to lie flush with the spool.

Abstract: An apparatus and method for minimizing compressor casing distortion due to bleed air extraction wherein the compressor casing includes an annular outer casing having at least one outlet duct, an annular inner casing disposed co-axially with the outer casing and spaced radially inwardly therefrom to define a plenum, two radial discs extending inwardly from the outer casing defining a plenum, a plurality of small area bleed openings located in the inner casing for imparting a radial velocity to the bleed air greater than the circumferential velocity entering the plenum, an air tube located over each bleed opening, and a baffle plate located on top of the air tubes and communicating with the radial discs for supporting the casings, whereby the bleed air impinges on the outer casing and controls distortion by sufficiently reducing the circumferential temperature gradient forming at the outer casing.

Abstract: The present invention discloses a spring clip made of a directionally solidified nickel-base superalloy for use in a high pressure compressor stator sub-assembly of a gas turbine engine. The spring clip includes a body section and a pair of arms integrally connected to opposing sides of the body section and the spring clip is used as a means for attaching circumferentially segmented flowpath liners to an annular casing. Each arm of the spring clips exerts a spring force against a flowpath liner mount lug causing the mount lug to be seated radially against an inner surface of the casing and axially against an annular seal, wherein the pair of arms of each spring clip engage mount lugs located on circumferentially adjacent ones of the flowpath liners.

Abstract: In a gas turbine engine, a system for maintaining uniform circumferential spacing between adjacent segments of a compressor liner, the liner segments being retained within an outer casing by a flanged connection. The system includes a T-shaped pin retained within the flanged connection radially and circumferentially by engagement with a slot formed in the outer casing, and axially by engagement with a notch formed in the adjacent liner segment, such that the liner segment overlaps the positioning pin. In a preferred embodiment, each liner segment is notched and held in place by a pin which, in turn, is retained within a slot in the compressor casing.

Abstract: A turboshaft engine in which the stator comprises an outer casing surrounding a ring formed by a number of ring elements each having an apertured rib which spaces the element from the casing, the ribs dividing the space between the casing and the ring into separate compartments, and each rib having a sealing element disposed alongside it to cover the apertures and seal one compartment from the next. The casing and the ring are fixed together by virtue of bolts which have their axes inclined to the radial direction so that tightening the bolts causes the casing to press radially on the sealing elements at the same time as pressing them axially against the respective ribs.

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 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 gas turbine engine has a forward rotor, a row of fan blades radially extending from the rotor, and a core turbine engine located rearwardly of the fan blades and rotor and coupled to the rotor for rotatably driving the rotor. An outer annular nacelle surrounds the rotor, fan blades and core engine. An inner annular splitter fairing is disposed rearwardly of the fan blades, surrounds the core engine, and is spaced radially inwardly from the nacelle so as to define therebetween a bypass air flow duct located outwardly from the core engine and rearwardly of the fan blades for producing thrust upon rotation of the rotor and fan blades. A hybrid shape spinner nose is attached about the rotor and projects forwardly therefrom within the nacelle and forwardly of the fan blades.

Abstract: The invention concerns an active clearance control for controlling clearance between a turbine and a casing in a gas turbine aircraft engine.The invention calculates the instantaneous clearance between a turbine casing and a turbine rotor, based on temperature. Two temperatures are involved. First, a steady state temperature (SSTemp) is computed for the rotor and the casing. SSTemp is a predicted, future temperature, which will be attained when the engine reaches steady state operation. Each SSTemp is computed based on presently occurring engine operating conditions, such as selected temperatures, pressures, and rotational speeds.Changes which occur in the SSTemp's indicate the second temperatures, which are the instantaneous temperatures of the casing and rotor. These changes in SSTemp are caused by changes in the present operating conditions, which occur during engine acceleration and deceleration.