A nozzle for a turbine system includes an airfoil, an inner sidewall, and an outer sidewall. Each of the inner sidewall and outer sidewall includes a peripheral edge defining a pressure side slash face, a suction side slash face, a leading edge face, and a trailing edge face. At least one of the inner sidewall pressure side slash face, the inner sidewall suction side slash face, the outer sidewall pressure side slash face, or the outer sidewall suction side slash face includes a first swept surface extending at a first angle relative to a nominal slash face angle and a second swept surface extending at a second angle relative to the nominal slash face angle. The first and second swept surfaces meet at an arc having a peak that is circumferentially aligned with a stiffening member extending circumferentially on a respective sidewall.

A turbine blade includes: a platform; an airfoil portion extending from the platform in a blade height direction and having a pressure surface and a suction surface extending between a leading edge and a trailing edge; a blade root portion disposed on opposite side of the airfoil portion across the platform in the blade height direction and having a bearing surface; and a shank disposed between the platform and the blade root portion. The shank has: a first lateral surface having a first recess portion and disposed on a pressure surface side and along an extension direction of the blade root portion; and a second lateral surface having a second recess portion and disposed on a suction surface side and along the extension direction of the blade root portion.

Provided is a transition piece including: a first flow passage group formed by arranging a plurality of intra-wall flow passages extending within a plate constituting the transition piece from a side near a gas turbine to a side near a combustor liner; a second flow passage group located on a side near the combustor liner with respect to the first flow passage group; and a plurality of dilution holes that penetrate the plate, and establish communication between a compressed air main flow passage and a combustion gas flow passage, each intra-wall flow passage of the first flow passage group and the second flow passage group having an inlet facing the compressed air main flow passage at an end portion on a side near the gas turbine, and having an outlet facing the combustion gas flow passage at an end portion on a side near the combustor liner, a dilution hole being located nearer to the inlet of an intra-wall flow passage of the second flow passage group than to the outlet of the intra-wall flow passage of the second flow passage group in each of spaces between the intra-wall flow passages adjacent to each other in the second flow passage group.

A suction side fillet portion of a turbine rotor blade includes a central fillet portion located at the center of a length of the suction side fillet portion along an extension direction of the suction side fillet portion, an upstream intermediate fillet portion which is located between the central fillet portion and a front edge that is an upstream end of the suction side fillet portion, and in which a fillet height from an upper surface of a platform portion is higher than a fillet height of the central fillet portion, and a downstream intermediate fillet portion which is located between the central fillet portion and a rear edge that is a downstream end of the suction side fillet portion, and in which a fillet height from the upper surface of the platform portion is higher than the fillet height of the central fillet portion.

A combustor includes a liner having an outlet end to pass combustion gas and a liner flange protruding outward from the outlet end; a transition piece to discharge combustion gas from the liner to a turbine, the transition piece having an inlet end for coupling to the outlet end of the liner and a transition piece flange protruding outward from the inlet end to face the liner flange; and a first elastic support installed on the liner flange to protrude toward the transition piece flange. A force applied from the transition piece elastically deforms an elastic arch of the first elastic support, which includes a movable support that is spaced apart from the liner flange if the force applied from the transition piece does not primarily deform the elastic arch. An auxiliary elastic support installed inside the first elastic support elastically deforms if the force secondarily deforms the elastic arch.

A rotor bow mitigation system for a gas turbine engine includes a control system in communication with an actuator, the control system actuates a power source to repeatedly actuate the actuator for a duration of time to intermittently rotate a rotary component such as a high spool of the gas turbine engine. A method of rotor bow mitigation for a gas turbine engine includes intermittently actuating an actuator of a rotor bow mitigation system in selective mechanical connection with a gear within an accessory gearbox to intermittently rotate the rotary component of the gas turbine engine to mitigate rotor bow conditions.

A process for cold working of the inner surface of a bore in a centrifugal compressor wheel along only a fractional portion of the bore length (i.e., along less than a full axial length of the bore), thereby creating a zone of compressive residual hoop stress in the metal surrounding the bore where the wheel needs the beneficial residual stress. The process purposefully avoids cold working of the bore at locations adjacent to high-stress areas and features of the wheel, where cold working in such locations could negatively impact the wheel's overall life.

Systems and methods are provided for an aircraft propulsor with a fan cowl with one or more tangential restraints. The tangential restraints may be configured to enter respective tangential receiving portions of a fixed portion of the aircraft propulsor. The fan cowl may receive a load that may include a bending load. The tangential restraints may be configured to convert the bending load to membrane tension and thus decrease deflection of the fan cowl. The fan cowl may be a translating sleeve used for a reverse thrust.

To increase the inertia of a sealing lip of a blade for an aircraft turbine engine, and thus improve the service life of such a sealing lip, the sealing lip is conformed so as to have a trough in the outer surface thereof and a corresponding boss in the inner surface thereof, the trough and the boss being defined based on a connection cross section of the sealing lip to a blade body, and being formed at a distance from a free axial end of the sealing lip.

High cycle fatigue (HCF) in a turbine wheel of a sector-divided dual volute turbocharger, particularly a turbocharger where the tongue-to-blade gap is as small as from 1-3% of the wheel diameter, is reduced, by locally increasing the volute cross-sectional area just upstream of the tongues. Thereby, it becomes possible to reduce the force function of the exhaust gas pressure onto the turbine wheel blades. Modifying how the pressure presents itself to the wheel reduces blade excitation and, ultimately, HCF of turbine wheels. In another aspect of the invention, the angle of the tongues are modified to direct the exhaust more directly onto the turbine wheel than conventional tongues. It is surprising that this approach not only accomplishes the desired result, but does this without significant loss of turbine stage efficiency.