Abstract: A spacer for a diaphragm assembly coupling is disclosed. The spacer includes a base, a spacing portion, and a spacing body edge. The base includes a base body including a base edge. The spacing portion includes a spacing body and a spacing flange. The spacing body extends from the base and includes an outer diameter that is smaller than that of the base. The spacing flange extends outward from the spacing body and is spaced apart from the base. The spacing body edge is located at an intersection of the spacing body and the base body. A reference line extending from the spacing body edge to the base edge forms an angle from 10 to 30 degrees with the spacer axis.

Abstract: A turbine assembly of an engine is disclosed. The turbine assembly includes a nozzle ring for receiving exhaust gases from a combustion chamber having a plurality of guide vanes for guiding the exhaust gases, a stator ring disposed downstream to the nozzle ring, and a rotor disposed circumferentially within the stator ring. The stator ring includes a plurality of stator vanes adapted to receive the exhaust gases through the guide vanes of the nozzle ring. The stator ring is axially movable between a first position and a second position along at least one cross-key pin disposed between the stator ring and a turbine casing. Each of the plurality of stator vanes moves into a throat plane of each pair of the plurality of guide vanes in the first position of the stator ring, and moves out of the throat plane in the second position of the stator ring.

Abstract: An insert tube of a turbine vane is disclosed. The insert tube includes a pressure side wall, a suction side wall opposite and spaced apart from the pressure side wall, a leading edge, and a trailing edge opposite the leading edge. The insert tube includes a plurality of cooling channels spaced along the pressure side wall. The insert tube includes an indented portion located between the trailing edge and the pressure side wall.

Abstract: A method and device for retrofitting a gas turbine engine for improved hot day performance are disclosed. The method can include removing a first selected stator bladerow from the plurality of compressor stages, the first selected stator bladerow having a first inlet swirl angle and including a first plurality of fixed stator vanes. Each stator vane of the first plurality of fixed stator vanes can have a first stator vane angle. The method can also include providing a first improved stator bladerow to replace the first selected stator bladerow. The first improved stator bladerow can have a second plurality of fixed stator vanes, each having a second stator vane angle smaller than the first stator vane angle. The method can also include replacing the first selected stator bladerow with the first improved stator bladerow to produce an increased pressure ratio and flow rate compared to the first selected stator bladerow.

Abstract: A modal analysis is performed on a blade and disk assembly of a turbomachine, including creating a finite element model of the blade and disk assembly made up of a plurality of nodes corresponding to points on the blade and disk assembly and characterized by a simplifying assumption of a rigidly attached connection between the blade and disk. The modal analysis also includes selecting a target vibration frequency at which a target amplitude of vibration has been one of measured or estimated for a target node during operation, and deriving modal velocity vectors from the target amplitude at the target node.

Abstract: A method for manufacturing a metallic component is disclosed. The method includes forming a metallic component with a support structure using an additive manufacturing process. The support structure includes support walls arranged to form flow passages with a predetermined cross-sectional area. The method also includes placing the metallic component with the support structure into a chamber and sealing the chamber. The method further includes introducing a fuel mixture into the chamber after sealing the chamber. The method still further includes igniting the fuel mixture in the chamber to remove one or more of the support walls of the support structure from the metallic component.

Abstract: A grommet for a combustor of a gas turbine engine is disclosed. The grommet includes a lower platform and a raised platform. The lower platform includes a plurality of top slots located in the top surface and a plurality of bottom slots located in the bottom surface. Each top slot and bottom slot may extend radially along the lower platform. The plurality of top slots may be spaced equidistantly from one another along the circumference of the lower platform. The plurality of bottom slots may be spaced equidistantly from one another along the circumference of the lower platform.

Abstract: A fuel injector for a combustion engine includes an injector head including a nozzle, a premixer, and a distributor structured to distribute a plurality of different fuels to different sets of fueling orifices in the premixer. A pilot assembly of the fuel injector is coupled to the premixer and includes a first fueling passage for a first fuel and a second fueling passage for a second fuel. Multiple sets of fueling orifices are positioned within the fuel injector, the fueling orifice sets being selectively connectable to a plurality of different fuel supplies, and both located and sized so as to accommodate a wide range of flow rates to enable a combustion engine coupled with the fuel injector to operate on fuels having a range of Wobbe indices and compositions.

Abstract: A fuel injector for a combustion engine includes an injector head including a nozzle, a premixer, and a distributor structured to distribute a plurality of different fuels to different sets of fueling orifices in the premixer. A pilot assembly of the fuel injector is coupled to the premixer and includes a first fueling passage for a first fuel and a second fueling passage for a second fuel. Multiple sets of fueling orifices are positioned within the fuel injector, the fueling orifice sets being selectively connectable to a plurality of different fuel supplies, and both located and sized so as to accommodate a wide range of flow rates to enable a combustion engine coupled with the fuel injector to operate on fuels having a range of Wobbe numbers and compositions.

Abstract: A fuel injector for a combustion engine includes an injector head including a nozzle, a premixer, and a distributor structured to distribute a plurality of different fuels to different sets of fueling orifices in the premixer. A pilot assembly of the fuel injector is coupled to the premixer and includes a first fueling passage for a first fuel and a second fueling passage for a second fuel. Multiple sets of fueling orifices are positioned within the fuel injector, the fueling orifice sets being selectively connectable to a plurality of different fuel supplies, and both located and sized so as to accommodate a wide range of flow rates to enable a combustion engine coupled with the fuel injector to operate on fuels having a range of Wobbe numbers and compositions.

Abstract: A turbine blade for a gas turbine engine may include at least two wrapped, serpentine-shaped internal cooling paths. A first one of the serpentine-shaped internal cooling paths may include a first passage that extends radially along a leading edge of the turbine blade from adjacent a root end of the turbine blade to adjacent a tip end of the turbine blade. The first passage may be configured to provide fresh cooling fluid to the leading edge. A second passage downstream of the first passage may be configured to discharge spent cooling fluid from the first passage of the first one of the serpentine-shaped internal cooling paths across a plurality of flow disrupters positioned along an upper span of a trailing edge of the turbine blade before exiting from the trailing edge of the turbine blade. A second one of the serpentine-shaped internal cooling paths may be configured to supply fresh cooling fluid to a lower span of the trailing edge of the turbine blade.

Abstract: A turbine blade for a gas turbine engine may include an outer, peripheral wall extending along a suction side and a pressure side of the blade from a leading edge to a trailing edge of the blade and from a root end to a tip end of the blade, a substantially vertical bifurcating internal wall extending between the leading edge and the trailing edge of the blade in between the suction side and the pressure side of the blade, a plurality of pressure side cooling fluid passages defined between the peripheral wall on the pressure side of the blade and the bifurcating internal wall and extending at least part way from the root end to the tip end of the blade, and a plurality of suction side cooling fluid passages defined between the peripheral wall on the suction side of the blade and the bifurcating internal wall and extending at least part way from the tip end to the root end of the blade.

Abstract: A fuel supply system for a combustion chamber having at least two combustion zones in which each combustion zone is arranged at a height different from the other includes a pair of fuel injectors that are configured to supply fuel to respective ones of the combustion zones. The pair of fuel injectors are disposed at different heights corresponding to the different heights of the combustion zones, each of the heights being taken in reference with one of a lower one of the fuel injectors and a horizontal midplane of the combustion chamber. The system also includes a flow control device in fluid communication with at least the lower one of the fuel injectors. The flow control device is configured to selectively regulate a supply of fuel to the lower fuel injector based, at least in part, on an amount of fuel supplied to an upper one of the fuel injectors.

Abstract: A system may be configured to use heat energy to produce power and potable water. The system may include an organic rankine cycle (ORC) subsystem configured to receive heat energy from one or more sources and convert that heat energy into usable power. The system may also include an air gap membrane distillation (AGMD) subsystem configured to receive heat energy from the ORC subsystem and use the heat energy to convert impure water into potable water.

Abstract: A turbine stage for a gas turbine engine is disclosed. The turbine stage includes a turbine disk and a turbine diaphragm, forming a cavity there between. The turbine stage also includes an exit flow discourager comprised of at least two teeth. The teeth may be located radially apart from each other, each tooth including a length extending in the axial direction and a width extending in the radial direction. A channel is formed between the teeth and an axially adjacent surface. A recirculation region may be formed in between each pair of teeth.

Abstract: An airfoil for a turbomachine such as a gas turbine engine includes a pressure side and a suction side, a flow-distributing forward wall and an inner cooling wall. A switchbacked passage extends through the airfoil, and the flow-distributing forward wall has a plurality of jetting orifices formed therein and connecting the switchbacked passage to a forward cavity. The jetting orifices are oriented to produce wall jets of cooling fluid directed in an upstream direction toward a back side of the leading edge. A second passage extends between the forward cavity and outlets of the airfoil, so as to convey cooling fluid in a downstream direction toward the outlets.

Abstract: Exhaust ducting for a gas turbine engine includes a lower plenum, an upper plenum, an exhaust fume isolator, and an inner duct seal positioned between the upper plenum and the exhaust fume isolator. The lower plenum is in flow communication with an exhaust collector of the gas turbine engine. The upper plenum is adjacent and in flow communication with the lower plenum. The exhaust fume isolator includes an isolation chamber and mixed exhaust ducting. The isolation chamber is adjacent the upper plenum and in flow communication with the enclosure. The mixed exhaust ducting is in flow communication with the isolation chamber and extends from the isolation chamber to the enclosure exhaust.

Abstract: An exhaust system for a gas turbine engine is provided. The exhaust system includes an exhaust diffuser having a diffuser axis. The exhaust diffuser includes an outer diffuser wall. The exhaust diffuser also includes an inner diffuser wall. The exhaust diffuser further includes a diffuser inlet formed by the first and second upstream ends of the outer diffuser wall and the inner diffuser wall. The exhaust diffuser includes a diffuser exit. The exhaust system also includes an exhaust collector coupled with the exhaust diffuser and adapted to collect and redirect exhaust gases in a radial direction. The exhaust collector includes a circumferential wall. The exhaust collector also includes a forward wall. The exhaust collector further includes an aft wall, wherein the aft wall is angled with respect to the diffuser axis. The exhaust collector includes a splitter disposed between the forward wall and the aft wall of the exhaust collector.

Abstract: An airfoil for a turbomachine such as a gas turbine engine includes a switchbacked passage for conveying cooling fluid, and flow-distribution passages in a forward wall so as to direct cooling fluid from the internal passage for back side impingement upon a leading edge of the airfoil. An inner cooling wall forms a passage for discharging the spent cooling air from a trailing edge of the airfoil after impingement, and cools one of a pressure side and a suction side of the airfoil by way of conduction. The switchbacked passage may have a serpentine form.

Abstract: A system and method for detecting the rotation breakaway of a spacer from a compressor rotor disk are disclosed herein. The method includes detecting a disk sensed feature and a spacer sensed feature on the compressor rotor disk and the spacer respectively. The method also includes comparing the timing between the two to a predetermined threshold to determine whether the relative position of the spacer to the compressor rotor disk exceeds a predetermined amount. The predetermined amount may be selected to determine whether an imbalance, rubbing, or binding can occur in the gas turbine engine or to determine whether anti-rotation features have been broken.