Abstract: A flow control device for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a seal body having a radially inner surface and a radially outer surface and at least one stand-up protruding from the radially outer surface and configured to seal an interrupted surface.

Abstract: A flow control device for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a seal body having a radially inner surface and a radially outer surface and at least one stand-up protruding from the radially outer surface and configured to seal an interrupted surface.

Abstract: An example damper includes a damping member configured to damp a duct wall at an interface between a duct band and the duct wall.

Abstract: A damper for damping vibration of a structural member of a gas turbine engine is disclosed. The damper may include a first metal mesh pad which abuts an outer circumferential surface of the structural member, and a garter spring which abuts the first metal mesh pad. Both the metal mesh pad and the garter spring may completely or partially encircle the structural member. Alternatively, the damper may include a damper cover which encloses the first metal mesh pad and the garter spring and which abuts the outer surface of the structural member. A second metal mesh pad may be inserted between the damper cover and the garter spring. A gas turbine engine which comprises such a damper is also disclosed.

Abstract: An example damper includes a damping member configured to damp a duct wall at an interface between a duct band and the duct wall

Abstract: An inlet case for a small gas turbine engine, where the inlet case is formed as a hybrid piece with a metallic inner shroud cylinder and plastic guide vanes and plastic outer shroud molded around the metallic inner shroud cylinder. The metallic inner shroud cylinder carries an electric generator and provides for a good convective heat transfer surface to carry heat away from the electric generator and into the inlet air flow entering the engine. The plastic guide vanes and outer shroud allow for a light weight inlet case assembly that is also much easier to form the airfoil shapes than a machined metallic inlet case. The plastic guide vanes and outer shroud are injection molded around the metallic inner shroud cylinder to produce a rigid single inlet case. The metallic cylinder includes axial extending slots that receive the lower ends of the guide vanes during the injection molding process.

Abstract: An apparatus and a process for determining the mass required for a wire damper ring to produce the maximum amount of damping in a certain rotary machine. The testing apparatus includes a test piece with an inward facing annular groove and a plurality of expander segments each with the top or outer half of a wire damper ring segment secured to and extending outward to fit within the annular groove. A mandrel with an outer surface having a slight slope is passed through a central opening of the expander segments and displaces the ring segments radially outward and into the annular groove at varying loads. The apparatus is set to vibrating and a piece of the test piece is observed for vibrations. The load applied to the damper ring segments is varied until the vibration is at a minimum. The applied load can now be correlated with the mass required for the damper ring to produce similar damping capability.

Abstract: A low pressure turbine rotor disk for a small twin spool gas turbine engine in which the rotor disk includes a forward side cavity large enough to allow for the bearing assembly that rotatably supports the rotor disk to fit within the cavity in order to shorten the axial distance between the bearings that support the inner rotor shaft on which the turbine rotor disk is secured. Minimizing the bearings spacing allows for a high critical speed for the inner rotor shaft and therefore allows for the small twin spool gas turbine engine to operate at this small scale. The turbine rotor disk also includes a plurality of axial aligned cooling air holes to allow for cooling air from the bearings to flow out from the aft end of the rotor disk. The inner surface of the cavity is an annular surface that forms a seal with knife edges extending outward from the bearing support plate also located within the cavity.

Abstract: A low pressure turbine rotor disk for a small twin spool gas turbine engine in which the rotor disk includes a forward side cavity large enough to allow for the bearing assembly that rotatably supports the rotor disk to fit within the cavity in order to shorten the axial distance between the bearings that support the inner rotor shaft on which the turbine rotor disk is secured. Minimizing the bearings spacing allows for a high critical speed for the inner rotor shaft and therefore allows for the small twin spool gas turbine engine to operate at this small scale. The turbine rotor disk also includes a plurality of axial aligned cooling air holes to allow for cooling air from the bearings to flow out from the aft end of the rotor disk. The inner surface of the cavity is an annular surface that forms a seal with knife edges extending outward from the bearing support plate also located within the cavity.

Abstract: An inner rotor shaft for use in a small twin spool gas turbine engine, the inner rotor shaft having a hollow middle section formed of a smaller diameter hollow section on a compressor end and a larger diameter hollow section on the turbine end of the shaft. Solid shaft end extend from the hollow section to form a forward solid shaft end to secure the fan rotor disk and an aft solid shaft end to secure the turbine rotor disk. A parabolic shaped transition section joins the forward shaft end to the smaller diameter hollow section, and a conical shaped transition section joins the aft shaft end to the larger diameter hollow section. A conical shaped transition piece joins the two hollow sections together to form an inner rotor shaft that can fit within a minimal space between the compressor rotor disk and the annular combustor assembly of the engine.

Abstract: An inner rotor shaft for use in a small twin spool gas turbine engine, the inner rotor shaft having a hollow middle section formed of a smaller diameter hollow section on a compressor end and a larger diameter hollow section on the turbine end of the shaft. Solid shaft end extend from the hollow section to form a forward solid shaft end to secure the fan rotor disk and an aft solid shaft end to secure the turbine rotor disk. A parabolic shaped transition section joins the forward shaft end to the smaller diameter hollow section, and a conical shaped transition section joins the aft shaft end to the larger diameter hollow section. A conical shaped transition piece joins the two hollow sections together to form an inner rotor shaft that can fit within a minimal space between the compressor rotor disk and the annular combustor assembly of the engine.