Abstract: A seal assembly for a turbine blade tip shroud including a one-piece bent metallic sheet intersegment seal having an axial component and a vertical component adapted to fit in a groove formed in the end walls of each shroud segment, wherein the shroud segment has a platform and radially extending ribs, and the seal fits tightly in the groove providing an axial and a radial component to the seal. A ring seal is also provided for sealing a joint between the shroud and a mounting structure where an annular gap is formed having an "L" shaped cross-section, the ring seal having a "C" shaped configuration and one of the legs of the "C" extending within an axial component of the "L" shaped gap. The leg has a wave pattern with the peaks of the wave in contact with opposite radial walls of the axial component of the gap for creating a gas seal at the joint.

Abstract: A seal assembly for a turbine blade tip shroud including a one-piece bent metallic sheet inter-segment seal having an axial component and a vertical component adapted to fit in a groove formed in the end walls of each shroud segment, wherein the shroud segment has a platform and radially extending ribs, and the seal fits tightly in the groove providing an axial and a radial component to the seal. A ring seal is also provided for sealing a joint between the shroud and a mounting structure where an annular gap is formed having an "L" shaped cross-section, the ring seal having a "C" shaped configuration and one of the legs of the "C" extending within an axial component of the "L" shaped gap. The leg has a wave pattern with the peaks of the wave in contact with opposite radial walls of the axial component of the gap for creating a gas seal at the joint.

Abstract: A cooling system for airfoil vanes in a turbine of a gas turbine engine wherein the airfoil platforms have openings downstream of the insert tube for passing spent platform impingement coolant air into the airfoil cavity downstream of the insert tube to increase the pressure in the aft zone of the airfoil to improve the temperature gradient across the span of the airfoil.

Abstract: A pressurized bearing assembly for turbomachines like gas turbine engines wherein pressurized oil is used to apply an axial force to the outer bearing race of a ball bearing assembly. The force is transmitted through the race to the ball bearings and to the inner race of the bearing assembly, and is further transmitted, via the engine shaft, to the bearings of another, axially spaced apart bearing assembly. The force moves the balls in both assemblies into line-on-line running contact with their respective bearing races, and minimizes vibration in the assemblies.

Abstract: A continuous plasma ignition system includes an exciter 12, an ignitor plug 16 and a cable 18 which connects the excitor 12 to the ignitor plug 16. The ignition system operates at the resonant frequency value f.sub.r of the circuit formed by the exciter, ignitor plug and the cable, such that, a continuously plasma 23 of ionized air is provided across the ignitor gap 22.

Abstract: A fuel nozzle (18) for a gas turbine engine comprises a nozzle stem (55), a nozzle tip assembly (60) and a nozzle sheath (50). A plurality of inlets (82) allow air to flow into the interior (90) of the sheath (50); fuel is flowed into the nozzle (18) through a fuel passage (135) in fluid communication with a fuel manifold. The fuel enters a fuel channel (140) defined by the stem (55) and tip assembly (60), and then passes into a fuel gallery (185) through a plurality of metering holes (190). Fuel swirls out of the tip assembly (60), where it is caught between, and squeezed by, first and second streams of air passing out of radially spaced apart air passages (145) and (220).

Abstract: Adjacent platforms (16) have feather seals (34) in complementary slots (22). Hot grooves (40) carry cooling air across the seal and discharge it into the gap (20) between adjacent platforms. Grooves discharging from abutting surfaces are staggered and have a flow component parallel to the axial gas flow through the turbine.

Abstract: Gas turbine blades 14 are slid axially into disc 12 with a retention tang 22 on each abutting the disc. An axially extending space 28 between blade platforms 26 and the disc receives elongated strip 30. With prebent end 34 resiliently held against the disc the opposite end 36 is bent radially outward against the roots of adjacent blades. A bow 38 biases the blades outwardly, deterring vibration.

Abstract: The present invention discloses a bleed valve for use in gas turbine engines. The valve has a piston, one end of which is positioned inside a housing which is positioned on one side of the by pass fluid flow path. The other end of the piston is positioned in an opening between the primary fluid flow path and the bypass fluid path such that the valve opens and closes in response to a pressure differential between two points in the primary fluid flow path.

Abstract: A bearing support housing for turbomachines like gas turbine engines. The housing is a one piece component having a flexible forward portion and rigid rearward portions, resulting in a housing having a unique combination of properties that allow for normal running of the engine as well as safe operation during unplanned rotor imbalances.

Abstract: The present invention relates to a bearing compartment for a turbomachine such as a gas turbine engine. The compartment includes a bearing compartment, to which is attached, on its front and rear sides, covers that allow for the passage of air through the compartment as well as past the compartment. In particular, the compartment includes a plurality of axially extending and circumferentially spaced apart sealed air cavities surrounding an oil cavity in the compartment; the bearing which supports a shaft extending through the engine is supported by the bearing. The compartment also includes a plurality of axially extending and circumferentially buffer air cavities surrounding the sealed air cavities, thereby providing additional insulate protection to the oil cavity.

Abstract: A split retaining ring 22 has deep internal scallops 26 which clear lugs 36 during installation while the ring fits within lip 40. The ring is opened with ridges 30 clearing lugs 36 while the ring is rotated. Shallow internal scallops rest on and lock with lugs 36 in the final position.

Abstract: The double derivative of the gas generator shaft (16) is sensed and compared to upper and lower limits to determine breaches of these limits within a first predetermined time, in which case a first potential surge condition is declared. The derivative of torque, or the double derivative of shaft speed, of the power shaft (22) is sensed and compared to upper and lower limits to determine breaches of these limits within a second predetermined time, in which case a second potential surge condition is declared. If both a first and second potential surge condition is declared within a third predetermined time, an actual surge condition is declared.

Abstract: Compressor 14 of a gas turbine engine has a final centrifugal stage 18 discharging air into annulus 20. Diffuser 20 are straight and of frustoconical shape receiving air from the annulus. A side cut produces opening 38 in the side of each diffuser. Much of the flow exits through, and is directed by these openings.

Abstract: A bleed valve for a gas turbine engine having a housing of two segments and which form a gas flow path through the compressor, in which a first segment is movable from the second segment thereby creating an opening between the two segments. The moveable segment having one or more arms with rollers attached thereto where the stationary segment has paths on which the rollers travel, as the moveable segment is caused to move away from the stationary segment thereby opening the valve.

Abstract: Fuel flow to the combustor of a gas turbine engine is regulated by controlling the amount of fuel delivered to the primary and secondary fuel orifices of a hybrid fuel nozzle. A one-way valve extending between primary and secondary fuel manifolds is controlled by an electronic sensor that acts upon signals received from sensors at station 4.5 of the engine, and by sensors that measure engine speed. By opening and closing the valve based upon T.sub.4.5 and N.sub.2 speed, optimum engine operating conditions are achieved.

Abstract: An electrical an (10) for the icing-susceptible stator vane stage (16) located in an air inlet section (12) of a turbomachine (not shown). A plurality of stator vanes (20) which are electrically isolated from a plurality of structural shrouds (30, 32) are grouped in one continuous, or a plurality of contiguous arcs about a fanshaft which is generally coincidental with the turbomachine's longitudinal axis (X--X). Low voltage alternating current is provided by an alternator (not shown). This current is directed through a plurality of electrical conductors (26) and is further distributed through each stator vane (20) which is either electrically conductive or employs an electrically conductive medium disposed proximate to an icing susceptible surface of an airfoil contour (21). The electrical current flow, which is generally directed in a radial orientation through each stator vane (20) causes a direct electrically-resistive heating effect to the airfoil contour surfaces (21) of the stator vanes (20).

Abstract: A bleed valve (20) includes a reciprocating sliding element (42) with a circular head (48) joined to an annular wall (50). An annular first chamber (52) defined by the wall (50), the head (48), and the central support (30) has a variable volume dependent upon the relative position of the sliding element (42). The valve (20) includes a housing (22) having a base (26) with first and second annular partitions (62,64) protruding from the base (26). An annular second chamber (66) is defined by the first partition (62), the platform (34) of the central support (30), the second partition (64), and the base (26). An opening (72) in the central support (30) connects the second chamber (66) in fluid communication to the first chamber (52). Control air supplied from an engine source flows through a first bore (78) into the second chamber (66) which is vented to ambient via a second bore (80). The first and second bores (78,80) have a primary and secondary metering orifice (84,86).

Abstract: An annular combustor liner (30) includes louvers (52, 54, 56, 70, 72) for inducing the formation of a single toroidal vortex (68) adjacent a domed upstream portion (34) of the liner (30).

Abstract: A spiral wound containment ring (b 20l ) is received within a radially inward facing channel (38) in an engine case (18).