Abstract: A small gas turbine engine with a ceramic turbine to allow for higher turbine inlet temperatures, where a metallic compressor is secured to a ceramic shaft extending from a ceramic turbine to form a single piece ceramic shaft and turbine, where a threaded nut secures a split ring retainer on the compressor end of the ceramic shaft. A hollow thrust runner is compressed between the compressor disk and the turbine disk by the threaded nut to secure rotor together. A centering spring forms a tight fit between the metallic thrust runner and the ceramic shaft on the turbine side.

Abstract: An aero gas turbine engine with a last stage turbine disc of a solid bore design. The solid bore turbine disc includes a center tie bolt that attaches to an upstream face of the solid bore disc. The attachment location is under the bore of the upstream stage. The aft side of the solid disc creates a stub shaft upon which a bearing is directly mounted. No fasteners pass through the solid bore disc. Omission of outboard fasteners allow for ease of assembly, reduced cost, reduced weight, and a reduced part count.

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: Aspects of the invention relate to a system for reducing relative movement between the outer shroud of a compressor diaphragm and the compressor cylinder of a turbine engine, thereby minimizing the wearing of these parts. Embodiments of the invention include a system for applying an axial preload on the outer shroud of a compressor diaphragm, preferably in the same direction as the gas load in the compressor. The system can be installed at or near the horizontal joint such that the axial preload can be applied substantially at the horizontal joint, which is the location of the largest relative movement. In one embodiment, the axial preload system can include a two-piece wedge block and pin. As the pin is driven between the two wedge blocks, the force applied to the pin can be converted to an axial load by the spreading apart of the two wedge block sections.

Abstract: Aspects according to the invention relate to a locking spacer assembly for a slotted turbine engine component. In one embodiment, aspects of the invention can be used in connection with the assembly and disassembly of a disc hosting a row of airfoils as can be found in the compressor or turbine section of a turbine engine. The spacer according to the invention is a multi-part assembly. In one embodiment, the spacer includes at least three sub-components: first and second end supports and a filler. The inner and outer faces of each of these components can have various features to facilitate engagement of the assembly and distribution of centrifugal loads in operation. These parts can be held together by a retainer, which can be received in a cutout provided in the end supports and the filler.