Abstract: A link for supporting a core engine within an outer shell of a gas turbine engine is described. The link, in one form, includes an elongate rod having a first end and a second end, and first and second rod end bearings are secured at the first second ends, respectively. Each end bearing includes a bearing head and a threaded stud extending from the head. The elongate rod includes a first threaded socket located at first rod end and sized to be engaged to the first rod end bearing stud, and a second threaded socket located at second rod end and sized to be engaged to the second rod end bearing stud. The first and second tapered jam nuts are located at the first and second ends of rod. Each tapered jam nut includes a nut end having, for example, a hexagonal head and a tapered portion extending from the nut end. A bore extends through each tapered jam nut, and the bore is sized to have the stud extend therethrough. The bore has internal threads at least at the location of nut end.

Abstract: A gas turbine engine thermal backbone deflection control apparatus to counter the effects of backbone bending due to nacelle aerodynamic, thrust, and maneuver loads by introducing a controlled backbone counter-bending moment by applying an uneven temperature or thermal field around the backbone to induce controlled compression and/or tensioning forces in the backbone or casing between axially spaced apart frames connected by the backbone. One embodiment of the invention provides for flowing thermal control air along an axially and circumferentially sector of the engine casing. The thermal backbone deflection control apparatus is controlled by the engines digital electronic control system using input signals generated by engine thrust load, inlet moment load sensors and/or blade tip clearance sensors.

Abstract: A gas turbine engine backbone deflection control apparatus to counter the effects of backbone bending due to gust, thrust. and maneuver loads by introducing a controlled backbone counter-bending moment by applying controlled tensioning forces only between axially spaced apart frames connected by the backbone. One embodiment of the invention, for a high bypass ratio turbofan engine, provides a hydraulically powered actuator to produce tensioning forces in cables connected between a fan frame and a turbine frame and the actuator is controlled by the engines digital electronic control system using input signals generated by inlet moment load sensors and/or blade tip clearance sensors.

Abstract: A gas turbine engine has a core frame separable from a vane frame for removably attaching the vane frame to a core structure through a mid ring. The mid ring includes an annular forward panel and an annular aft panel, each of which has a radially outer flange and a radially inner flange. The inner flanges are welded to the core frame and the outer flanges are bolted to the vane frame. A plurality of circumferentially spaced stiffeners extend between the forward and aft panels, with each of the stiffeners including a pair of stringers extending diagonally across the stiffeners. Panels may be incorporated with the stringers between the triangular shaped areas adjacent the forward and aft panels of the stiffener. The bolted connection between the mid ring and the vane frame includes a plurality of tongue and groove joints formed in the flanges of the mid ring and vane frame. The tongue and groove joints provide additional radial support for the core engine within the vane frame.

Abstract: In one embodiment, a fan blade includes an airfoil portion, a platform portion, a shank portion, and a generally axial dovetail portion of a predetermined axial dimension for matingly engaging a generally axial dovetail receiving slot of substantially the same predetermined axial dimension in a fan disk member. Aft cleat means, including a portion having substantially the same cross section as the dovetail portion, is provided and is integral with the dovetail portion and projects generally axially therefrom. The aft cleat means strengthens the dovetail portion. Methods are provided for fabricating such a blade and for strengthening existing blades. Other embodiments are disclosed.

Abstract: A sealing arrangement for preventing compressor fan air loss through the relatively wide gaps between adjacent blade platforms required in order to allow for rotation of the blades around their minor root axes in response to bird impact and the like. A thin flexible ribbon-like seal is bonded to one of the two adjacent blade platforms in the fan structure. At the gap, the seal includes a thickened wedge shaped portion with at least one invagination on the inner side facing the rotor axis. At the operating speed of the compressor fan, centrifugal force causes the thickened portion to be forced into the gap between adjacent platforms producing the required seal.

Abstract: A laminated filament composite structure, such as an airfoil for use in an environment in which it is subjected to both foreign object impact and bending is provided with improved leading edge protection. At least one fine wire mesh layer is partially bonded within the composite structure along its neutral bending axis. A portion of the wire mesh layer extends beyond the neutral bending axis and partially around the leading edge where it is bonded to the outer periphery of the primary composite structure. The wire mesh is clad with a metal such as nickel to provide an improved leading edge protective device which is firmly anchored within the composite structure. The neutral bending axis anchoring tends to retain the leading edge protective device intact even after the delamination of the composite structure.