Abstract: A centrifugal impeller having an inlet and a plurality of outlets, the outlets including at least one outlet in a forward direction of the impeller and at least one outlet in a rearward direction of the impeller. A flow path distance of the outlet in the forward direction can be greater than a flow path distance of the outlet in the rearward direction in order to provide a greater pressure in the outlet having the greater flow path distance. In another embodiment, the flow path volume in the forward direction can be greater than the flow path volume in the rearward direction in order to provide a greater flow volume in the forward direction. In another embodiment, the number of flow paths in the forward direction can be greater than the number of flow paths in the rearward direction in order to provide a greater flow volume in the forward direction.

Abstract: A turbomachine with a row of blades in which the blades are structured such that the blades can untwist at higher rotational speeds in order to provide optimal aerodynamics at different rotational speeds. Each blade is formed with a thin slot between the pressure side wall and the suction side wall so that stresses from centrifugal forces occurring during high rotations will not pass from adjacent walls and allow for the blade to untwist more than a solid blade. Each blade is formed from a fiber laminated composite material with the fibers oriented at different directions to form different stiffness in each direction so that a desired untwist can occur based upon a given blade mass and rotational speed. A thin sheet of a slippery material can be placed within the thin slot and between the adjacent airfoil walls to allow slippage between the adjacent walls and to promote damping for the blade.

Abstract: A thin compressor blade used in a small aircraft gas turbine engine, in which the blade includes a main blade portion formed from a fiber reinforced resinous composite material to provide a lightweight and strong blade, and in which the leading edge or the trailing edge of the blade is formed from a foam metal reinforced composite material that defines a sharp edge for improved performance of the thin blade. The fibers of the main body extend into the edge region in order that the resin used to bind the foam metal will also bind with the fibers of the laminates to secure the edge to the main blade body. The blade tip can also be formed from the porous foam metal and resin material to make a strong and lightweight rotor blade.

Abstract: A compressor blade for a gas turbine engine that includes a composite leading edge insert attached to a titanium blade by a row of fastener pins along an aft edge of the insert to prevent unzipping. Each insert is also secured by an electrochemically disbondable composition for easy removable of a damaged insert from an integrally bladed disk. One or more of the fastener pins provides for an electrical conductive path that bypasses the disbondable composition during a lightening strike so that the insert does not release. Each replaceable insert includes a root portion that fits within a slot of the rotor disk, with each insert being secured against radial displacement by two shear pins secured within two shear pin slots formed in both the insert root portion and the rotor disk slot. The shear pin slots are curved and follow the curvature of the airfoil at the junction to the rotor disk.

Abstract: A turbine blade and platform assembly in which the platform is a separate piece that is secured to the blade through shear pins that extend along slots formed between opposed abutting parts of the blade and the platform. Because the platform is detached from the blade, the blade can be made of a single crystal superalloy with a reduction in casting defects. The platform has one or more airfoil shaped openings or slots on the outer surface in which a blade is inserted. The shear pins are inserted to secure the platform to the blade. In another embodiment, the blade root is inserted into an opening formed in the rotor disk, and shear pins are inserted into slots formed in the root and the disk opening in order to secure the blade to the rotor disk. With this embodiment, the blade can be made of a ceramic material because only compressive forces are formed to hold the blade in place.

Abstract: A rotor disc and rotor blade assembly in which the rotor blades include curved shear pins to secure the blade to the rotor disc. The blade includes an airfoil portion with an airfoil outline shape at the junction of the airfoil with the platform, and the root extends below the platform and has the general shape of the airfoil at the junction. An opening in the rotor disc also has the shape of the airfoil outline such that the root will fit snuggly within the opening in the rotor disc. Curved slots in both the opening and the root wall are formed on both sides of the opening and the blade root. Shear pins having a rectangular cross sectional shape fit within the slots to secure the blade to the disc. Because the root curvature follows the airfoil shape, the shear pins are located below the airfoil walls and absorb the shear stress caused by bending forces acting on the airfoil from the fluid reaction.

Abstract: A stator vane assembly for a gas turbine engine, the stator vane including an airfoil portion formed from a single crystal material and two platforms attached to the ends of the airfoil by shear pins that fit within slots formed between the platform and airfoil. The shear pin slots extend along the contour of the airfoil where the lowest stress levels are located. The platforms include openings having the same cross sectional shape as the curved airfoil. Because of the shear pin retainers between the airfoil and the platforms, the airfoil can be formed from a single crystal material while the platforms are un-coupled from the airfoil and can be made from a different material.

Abstract: A turbo machine airfoil, such as a blade or a vane in a compressor, in which the airfoil root portion includes screw threads having a dove tail cross sectional shape, and the rotor disk includes holes having similar screw threads such that a blade or vane can be easily replaced by unscrewing the old blade or vane and screwing in a new blade or vane. A locking pin biased by a spring is located in the root of the airfoil to lock the airfoil in the threaded hole to prevent the airfoil from loosening due to vibrations and to lock the airfoil in place at a desired angle of attack for maximum efficiency of the turbo machine. The removable cover plate is located on the casing and over the rotor disk such that the blade can be removed through the opening without the need to disassemble the turbo machine.

Abstract: A fan type compressor includes a rotor disk with a plurality of root insertion openings to receive a blade. Each insertion opening includes two shear pin slots on the sides, and the blade includes a root portion with shear pin slots formed on the pressure side and the suction side of the root. The blade is inserted into the insertion opening such that the slots are aligned, and a shear pin is inserted into both slots to secure the blade within the rotor disk. The shear pin slots are formed below the airfoil curvature and follow the airfoil curvature at the root portion. The shear pins are flexible enough to bend during insertion into the slots, and are strong enough to prevent from shearing off during rotor disk operation. The blade attachment structure provides for reduction of the dead weight from the attachment structure in order to simulate an IBR while providing for easy removal and replacement of a damaged blade.

Abstract: A gas turbine engine used in an aircraft includes a fan to provide a bypass flow through the engine. The fan delivers some airflow to the compressor for compression, and a fuel is mixed with the compressed air and delivered to a turbine to drive the engine and compressor and fan. The fan also delivers airflow to a central bypass passage located inward of the compressor. Guide vanes both upstream and downstream of the combustor each with cooling air passages therein, pass the bypass airflow from the inner bypass passage into an outer bypass passage located outward of the combustor and the turbine. The bypass flow through the guide vanes also cool the guide vanes and pass over and around the combustor to cool the combustor. The bypass airflow exits the engine at a location outward of the hot gas flow from the turbine.

Abstract: A rotor shaft that drives a load and is supported by a plurality of bearings includes a first bearing to support a distal end of the shaft, a second bearing to support a distal end of the load, and a third bearing to support both proximal ends of the shaft and the load in order to dynamically uncouple the load from the rotor shaft and to eliminate a fourth bearing of the prior art. The invention is also used in a small gas turbine engine in which the turbine shaft is connected to an electric generator, and the third or middle bearing supports the proximal ends of the generator and the rotor shaft to dynamically uncouple the turbine shaft from the generator rotor.

Abstract: A composite turbine blade and fan blade assembly for use in a turbo machine such as a gas turbine engine. The turbine blade assembly includes a plurality of metallic turbine blades secured to a rotor disk and forming an outer shroud assembly. A fiber reinforced composite fan blade assembly includes an annular ring with a plurality of fan blades extending there from. The annular ring has a conical shape to allow for easy installation over the outer shroud assembly, and is secured to the outer shroud assembly to force the fan blade assembly to rotate with the turbine blades, yet allow for a slight radial movement between the fan blades and the turbine blades. The turbine blades are formed of a high temperature super alloy to allow for use in a gas turbine engine, and the fan blade assembly is made from a fiber reinforced composite to allow for higher rotational speeds and smaller diameters for smaller gas turbine engines without exceeding allowable stress limits.

Abstract: A compressor includes a stator having a plurality of vane sectors each with an outer shroud having a forward and a rearward guide vane extending there form, the vane sector being mounted in an annular slot of the casing, the guide vanes forming a gap between the tips and an inner shroud, and a spring member located on the outer shroud in the rearward portion to bias the rearward shroud in a radial direction, where a space is formed in the slot to allow for the outer shroud to move in the radial direction when the rearward guide vane tip makes contact with the inner shroud due to thermal growth of the stator components. This arrangement allows for the rearward guide vane tip to maintain contact with the inner shroud without inducing high compressive stresses in the stator components due to the thermal growth.

Abstract: A locking spacer assembly for filling a void between adjacent components in a turbine engine. In at least one embodiment, the locking spacer assembly may be configured to be inserted between adjacent turbine blades in a disc groove in a turbine blade stage assembly. The locking spacer assembly may be formed from fore and aft end supports and a locking device having fore and aft angled surfaces for urging the fore and aft end supports into lateral recesses in the disc groove. The locking spacer assembly may also include a mid spacer positioned between the fore and aft end supports. A retainer may be attached to the locking device to draw the fore and aft angled surfaces of the locking device into contact with the fore and aft support surfaces and urge the first and second end supports into lateral recesses in the disc groove.

Abstract: The present invention is a shroud tip clearance control ring for use in a gas turbine engine in which a rotary blade and a shroud ring form a gap that forms a blade outer air seal, and where the shroud ring is formed of two shroud segment halves, each shroud segment including a pin located near an end of the segment, the pins being slidably within slots formed on the casing, and a drive member moves along a direction in which the pins are moved along the slots to change a radius of the shroud ring in a way such that the radius remains substantially the same along a 360 degree angle of the shroud ring. The drive member includes two positions, one that places the shroud ring in a radially inward position, and another position that places the shroud ring in a radially outward position.

Abstract: A blade or a vane preferably used in a gas turbine engine, the blade or vane being formed of a metallic core or insert with a ceramic airfoil surface bonded to the metallic insert. The insert includes a series of concave and convex portions arranged along the outer surface, in which a plurality of hoop fibers are secured within a resin matrix to firmly secure the metallic insert to the ceramic airfoil body. Axial loads between the insert and the ceramic body are transferred to the hoop fibers. The hoop fibers are strong enough to resist extending in length that would permit the hoop fibers from sliding over the widest portion of the insert.

Abstract: A bearing damper includes a coiled wire wrapped around projections formed on an outer race of the bearing, and supported by a casing having projections extending toward the coiled wire. Projections on the outer race are aligned with recesses on the casing, while projections on the casing are aligned with recesses in the outer race so that the coiled wire can flex between the projections that support the wire and provide a damping function. The coiled wire is formed of a plurality of strands. Flexing of the wire causes adjacent strands to rub and dissipate energy from the vibrations. The tension in the coiled wire can be adjusted in order to vary the damping capability of the bearing damper.

Abstract: A compressor blade that has a blade root, an airfoil having a first end, and a second end opposite the first end, the second end having at least one edge, and the airfoil is made of a first material having a first modulus of elasticity. A blade platform connects the blade root to the first end of the airfoil, and a flexible seal is connected to the airfoil adjacent the second end, and the seal is made of a second material having a modulus of elasticity that is substantially less than the first modulus of elasticity.

Abstract: An apparatus for actuating variable stage vanes is provided having a plurality of pivot arms, a synchronizing ring, and apparatus for pivotly attaching the pivot arms to the synchronizing ring. Each pivot arm includes a first end for fixed attachment with one of the vanes. The synchronizing ring includes a first flange, a second flange, a web extending between the flanges, and a plurality of openings disposed in the web. The apparatus for pivotly attaching the pivot arms to the synchronizing ring are disposed within the openings.

Abstract: The inner shroud of the vane of a vane assembly for the compressor section of a gas turbine engine is fabricate from a composite material preferably imide resin reinforced with carbon fibers and is formed into an open ended box-like member segment supporting no less than two vanes or airfoils. The segments are interlocked by a necked down edge adapted to slip into the open end of the adjacent inner shroud of the stator vane that serve to seal the gas path and provide damping. Rub strips are integrally formed on the bottom surface of the inner shroud.