Abstract: A swirler inlet valve plate for a fuel injector of a gas turbine engine includes a blocker ring and a pivot sleeve mechanically connected to the blocker ring. The blocker ring is formed from at least a portion of a toroid shape. The blocker ring includes a blocking surface perpendicular to an axis of the blocker ring formed from at least a portion of an annulus. The blocker ring also includes a flow surface, an inner circumferential surface of the toroidal shape. The pivot sleeve includes a sleeve portion extending adjacent the blocker ring. The sleeve portion includes a rotation transmission mechanism.

Abstract: A first stage compressor disk of a gas turbine engine includes a body. The body includes a forward end, an aft end, and an outer surface. The body also includes a plurality of forward balancing holes through the outer surface. The forward balancing holes align circumferentially about the body. The body further includes a plurality of aft balancing holes through the outer surface. The aft balancing holes align circumferentially about the body and are located aft of the forward balancing holes. The first stage compressor disk also includes a radial flange at the aft end of the body. The radial flange extends radially outward from the body. The radial flange includes slots for mounting airfoils.

Abstract: A gas turbine engine fuel injector stem includes a single piece of material with a flange, a center body, a gas gallery, and a fluid passage. The flange includes a cylindrical shape, a plurality of mounting holes and a handle hole. The center body extends from a flange cap in the flange axial direction. The gas gallery is adjacent the center body and distal to the flange. The gas gallery includes a ring shape oriented transverse to the flange and includes a gallery opening. The fluid passage extends through the flange, center body, and gas gallery and is in flow communication with the gallery opening.

Abstract: A power system enclosure is provided. The power system enclosure includes a housing accommodating a power source. The power system enclosure further includes an enclosure inlet connected to the housing to allow entry of an intake fluid into the housing. An enclosure exhaust is connected to the housing to route fluid from the housing. Further, a relief damper is disposed in the enclosure exhaust. At least a portion of the relief damper is configured to open when a pressure within the housing exceeds a predetermined threshold value.

Abstract: An apparatus for measuring temperature parameters of a structure is provided. The apparatus comprises: a carrier including at least one receiving portion; and a plurality of temperature sensors disposed within the at least one receiving portion and secured to the receiving portion via a bonding medium, the sensors being distributed along at least one direction of the carrier and configured to measure and record temperature information through microstructural changes corresponding to temperatures experienced by the respective sensors.

Abstract: A method of deploying a modular gas compression plant is disclosed. In an embodiment, the method comprises constructing a building of four substructures including a first lower housing, a first upper housing, a second lower housing, and a second upper housing. The method further includes assembling a turbomachinery equipment into the first lower housing. The method further includes assembling an air inlet duct and an exhaust duct to the first upper housing. The method further includes disassembling the building into the four substructures.

Abstract: A method for manufacturing a stem for a fuel injector of gas turbine engine includes forging a material into a unitary workpiece, which includes forging a top disk having a cylindrical shape, forging a body extending from a base of the top disk, and forging a lower disk connected to the body, distal to the top disk, the lower disk having a cylindrical shape oriented transverse to the top disk. The method further including machining the lower disk to form a gas gallery having a ring shape, and to define a gallery opening through the gas gallery, and machining a first fluid passage through the top disk, the body, and a portion of the gas gallery to the gallery opening, wherein the first fluid passage is in fluid communication with the gallery opening.

Abstract: A pilot liquid tube having a pilot liquid fuel inlet, a pilot liquid fuel conduit, a pilot liquid fuel nozzle, and a shroud configured to shield the pilot liquid fuel nozzle. The pilot liquid tube may be installed and/or removed from an injector of a gas turbine engine while the shroud remains fixed to pilot liquid fuel conduit.

Abstract: An impingement cover (460) for fusing to a gas turbine engine turbine nozzle (451) with a nozzle rail (455) includes a body and a plurality of tabs for forming a first fuse with the nozzle rail (455). The body has a plate like shape and includes a plurality of impingement holes (464), a first edge, a second edge, and a first airfoil cooling hole (470). The plurality of tabs includes a first tab (461) extending from the first edge, and a second tab (462) extending from the second edge. The second tab (462) is located in a different quadrant of the body than the first tab (461).

Abstract: A method and device for modifying a gas turbine engine is disclosed herein. The method for modifying a gas turbine engine includes accessing a first outward end of a first tube installed in the gas turbine engine, the first tube having a tube axis. Then attaching an extraction tool to the first outward end of the first tube, and applying an axial force to the first tube using the extraction tool, the axial force relative to the tube axis, thus breaking a first interference seal between a first interface surface on a first inward end of the first tube and a receiving member of the gas turbine engine. The method continues with removing the first tube from the gas turbine engine.

Abstract: A shield assembly for shielding a housing of an electro-mechanical device from a hot body includes a first sidewall and a second sidewall disposed in a spaced apart relation from each other. A third sidewall extends between the first sidewall and the second sidewall to define a space therebetween. The first, second, and third sidewalls are configured to receive the housing of the electro-mechanical device in the space defined therebetween. The shield assembly further includes a backing plate extending laterally from a bottom end of at least one of the first, second, and third sidewalls. The backing plate is releasably fastened with a bottom end of the housing. The shield assembly further includes a tab member extending away from a top end of at least one of the first, second, and third sidewalls. The tab member is releasably fastened with a bracket associated with the housing of the electro-mechanical device.

Abstract: A liquid gallery for a fuel injector of a gas turbine engine includes a gallery body, a liquid gallery scroll, a plurality of atomizer inlets, and a plurality of atomizer bosses. The liquid gallery scroll extends into the gallery body spanning from a first end to a second end in a circumferential direction. The liquid gallery scroll includes a taper with a cross-sectional area of the liquid gallery scroll reducing from the first end to the second end. Each atomizer inlet is in flow communication with the liquid gallery scroll. Each atomizer boss is aligned with one of the plurality of atomizer inlets.

Abstract: A vane for an Inlet Guide Vane (IGV) of a multistage compressor includes a root portion, a tip portion, and an airfoil extending therebetween. The vane is configured such that a ratio of the maximum thickness of the airfoil to a chord length (Tmax/C) at 50% of the span height (H) taken from the tip portion of the vane is configured to lie in the range of 0.11 to 0.12. In addition to the ratio (Tmax/C) at 50% of the span height (H) being in range of 0.11 to 0.12, a ratio of the maximum thickness to the chord length (Tmax/C) at various points along the airfoil varies with span height (H) of the vane i.e., distance taken from the tip portion of the vane and the local chord length C present at that span height H of the vane taken from the tip portion.

Abstract: A rotor blade for a gas turbine engine has an airfoil, a base integrally joined to the airfoil, and a root integrally joined to the base and mountable in a slot in a rotor hub of the gas turbine engine. The root has a dovetail including at least one contact face that, when mounted, contacts a surface of the slot to retain the rotor blade in the hub. The root includes a neck between the base and the dovetail, and a groove in the neck for redirecting stress in the rotor blade. In certain embodiments, the groove is at a distance from the at least one contact face, has a length less than a length of the dovetail, and/or has an initial non-zero depth at the side of a trailing edge of the airfoil and tapers to a zero depth in the direction of the leading edge of the airfoil.

Abstract: A turbine blade is disclosed. The turbine blade may include a platform, an airfoil extending from one side of the platform, and a neck extending from another side of the platform, wherein the neck includes a forward buttress and an aft buttress. The turbine blade may further include a root extending from the neck, a pocket defined by a plurality of walls and located between the forward buttress and the aft buttress, and a variable radius fillet. The variable radius fillet may be disposed within the pocket and extend between the forward buttress and the aft buttress, wherein a radius of the variable radius fillet increases from the forward buttress to the aft buttress.

Abstract: A gas turbine engine outer premix barrel is disclosed. The outer premix barrel includes a body portion, a barrel portion, and a plurality of vanes. The body portion includes vent air inlets. The barrel portion extends axially aft from the body portion. The plurality of vanes extends axially forward from the body portion. Each vane includes an inward surface located at a radially inner end of the vane, and a plurality of vent passages exiting the vane at the inward surface. Each vent passage is in flow communication with at least one vent air inlet.

Abstract: A damper for a turbine rotor assembly of a gas turbine engine may include a forward plate with a forward face and an aft face, and an aft plate with a forward face and an aft face. The aft face of the forward plate may be connected to the forward face of the aft plate with a longitudinal structure. An area of the aft plate in a plane transverse to the longitudinal structure may be greater than an area of the forward plate in the plane transverse to the longitudinal structure. The damper may also include a pocket on the forward face of the forward plate.

Abstract: An inner bushing assembly (280) to provide a biasing force between a guide vane (260) and an inner ring half (261) of a gas turbine engine compressor (200) is disclosed. The inner bushing assembly (280) includes a first bushing (281), a second bushing (282), and a biasing element (283). The first bushing (281) is configured to be installed about an inner vane shaft (267) of the guide vane (260) adjacent to an airfoil (265) of the guide vane (260). The second bushing (282) is configured to be installed about the inner vane shaft (267) distal to the airfoil (265). The biasing element (283) is configured to be installed about the inner vane shaft (267) between the first bushing (281) and the second bushing (282).

Abstract: An air blocker ring assembly includes a proximal end, a distal end, a blocker ring, and a blocker ring support. The blocker ring has a first protrusion, a second protrusion, and a third protrusion. The blocker ring support has a recess. The first protrusion of the blacker ring is received in the recess. The blocker ring is circumferentially discontinuous about a circumferential slot. At least a portion of the first protrusion is located circumferentially opposite the slot, and the second and third protrusions terminate at the slot.

Abstract: A method for optimizing a rotor assembly for a gas turbine engine is disclosed. The rotor assembly includes rotating parts including a first journal and one or more rotor disks. The method includes determining vector components for a total parallelism vector for each rotating part, determining a total parallelism sum including determining a magnitude of each total parallelism vector and adding the magnitudes into a single sum, and determining a minimum value for the total parallelism sum including selecting the rotor disk build angles and the second journal build angle that result in the smallest value for the total parallelism sum.