Abstract: A method and control system for controlling compressor output for a gas turbine engine is disclosed. The power output of a gas turbine engine can vary and be below desired output levels due to operating conditions such as ambient temperature and elevation. These operating conditions can lead to lower output of the gas compressor of the turbine engine and lower operating temperatures within or proximate to a turbine of the gas turbine engine and lead to less power output. Additional fuel can be added to increase power to the gas producer shaft and increase turbine temperature of the gas turbine engine. A power transfer device can be used to remove or add power to the gas producer shaft to balance the gas producer mechanical limits and turbine thermal limits at maximum levels and lead to higher power output.

Abstract: A fuel injector for a combustor of a gas turbine engine is disclosed herein. The fuel injector includes a fuel delivery system for receiving and distributing fuel and an injector body. The injector body includes a primary fuel gallery, a pilot fuel gallery, a primary perforated plate, and a pilot distributor plate. The primary fuel gallery is formed as an annular cavity in the injector body and extends around an assembly axis. The primary fuel gallery and the pilot fuel gallery are in flow communication with the fuel delivery system. The primary perforated plate is disposed within the primary fuel gallery and divides the primary fuel gallery. The primary perforated plate having a first perforation to restrict flow. The pilot distributor plate is disposed within the center body assembly and radially inward of the first portion of the primary gallery, adjacent to portions of the tube stem. The pilot distributor plate having a pilot distributor passage to restrict flow.

Abstract: A center body assembly for a lean direct injection fuel injector is disclosed. The center body assembly defines a primary liquid passage, a liquid gallery, a liquid main first passage, a liquid main second passage, and an atomizer assembly. The primary liquid passage supplies liquid fuel to the liquid gallery. The liquid gallery aligns with the primary liquid passage and distributes fuel to the liquid main first passage, and subsequently to the liquid main second passage. The atomizer assembly is in flow communication with the liquid main second passage and provides atomized fuel for combustion.

Abstract: A damped turbine blade assembly for a gas turbine engine is disclosed. The damped turbine blade assembly includes a damper positioned within a first small slot of a first turbine blade and a second large slot of the second turbine blade. A portion of the damper can slidably mate with the second large slot providing a radial and angular connection between the first turbine blade and second turbine blade while allowing movement in a direction tangent to a radial of a center axis of the gas turbine engine. The tangential movement is resisted by friction between the damper contacting the second large slot and provides friction damping against vibrations felt by the turbine blades during operation of the gas turbine engine. The damper can be shaped and/or pre-stressed to control the normal force component of the friction between the damper and the second large slot.

Abstract: This disclosure provides a seal strip assembly for a sealing strip slot to reduce air leakage between gas turbine nozzle segments. The seal strip assembly includes a bimetal element and a backing plate. The bimetal element includes a first layer with a first thermal coefficient of thermal expansion and a second layer with a second coefficient of thermal expansion. The second coefficient of thermal expansion is lower than the first coefficient of thermal expansion allowing the bimetal element to expand with an increase in surrounding temperature.

Abstract: A turbine blade having a base and an airfoil, the base including cooling air inlets and an internal cooling air passageway, and the airfoil including an internal multi-bend heat exchange path beginning at the base and ending at a cooling air outlet at the trailing edge of the airfoil. The airfoil also includes a “skin” that encompasses a tip wall, an inner spar, and a tip flag cooling system.

Abstract: A turbine blade includes a base and an airfoil. The airfoil includes a skin extending from the base and defining a leading edge and a trailing edge opposite the leading edge. The trailing edge includes an inner edge disposed proximate to the base, an outer edge disposed distal the inner edge, and a tuning region edge disposed between the inner edge and the outer edge. The tuning region edge includes an upper transition edge, a middle transition edge, a lower transition edge, an upper tuning edge, and a middle tuning edge. The upper transition edge extends from the outer edge towards the inner edge. The middle transition edge is disposed between the upper transition edge and the inner edge. The lower transition edge is disposed between the middle transition edge and the inner edge. The upper tuning edge is disposed between the upper transition edge and the middle transition edge, being at least partially closer to the leading edge than the middle transition edge.

Abstract: A method for refurbishing an assembly having a first component and a second component is provided. The first component has a fore portion welded to the second component and a flared aft portion in proximity with the second component. The method includes processing the assembly such that the flared aft portion and a part of the fore portion proximal to the flared aft portion of the first component is subject to peening until the flared aft portion is plastically deformed from an initial position to a desired final position with respect to the second component.

Abstract: This disclosure provides a seal assembly to provide a generally smooth transition from a combustion chamber to a turbine within a gas turbine engine. A seal body extends from the combustion chamber towards the turbine and is positioned to seal against unwanted flows between the combustion chamber and a compressed air chamber.

Abstract: A turbine blade having a base and an airfoil, the base including a root end. The airfoil including a skin extending from the base and defining a leading edge, a trailing edge, having a tip end opposite from the root end. The turbine blade further including dividers located within the airfoil. The dividers, leading edge, trailing edge, and skin define serpentine channels within the airfoil. A first multi-bend heat exchange path and a second multi-bend heat exchange path extend through the serpentine channels to cool portion of the turbine blade.

Abstract: A control system for a gas turbine engine is disclosed. In embodiments, control system includes a controller and a high speed recorder. The controller obtains a sensor value from a sensor connected to the gas turbine engine and publishes a tag that includes the type of event, the sensor value, and a timestamp. The high speed recorder checks the tag for an overspeed event. If an overspeed event is detected, the high speed recorder records values provided by the tag.

Abstract: A turbine blade having a base and an airfoil, the base including a root end. The airfoil including a skin extending from the base and defining a leading edge, a trailing edge, having a tip end opposite from the root end. The turbine blade further including dividers located within the airfoil. The dividers, leading edge, trailing edge, and skin define serpentine channels within the airfoil. A first multi-bend heat exchange path and a second multi-bend heat exchange path extend through the serpentine channels to cool portion of the turbine blade.

Abstract: A nozzle segment for a gas turbine engine with turbine airflow passing through the gas turbine engine. The nozzle segment including an upper shroud and an inner hub. The nozzle segment including a first airfoil and second airfoil extending from the upper shroud to the inner hub. The first airfoil and second airfoil including conduits for delivering secondary air to displace a portion of the turbine airflow passing through the gas turbine engine.

Abstract: A system and method for resource estimation of an additive manufacturing device is disclosed herein. The system includes at least one processor and one or more software modules. The processor executes the one or more software modules to determine a number of parts that can be manufactured per manufacturing run, estimate manufacturing time to manufacture a specified number of parts, and determine a total cost to manufacture the number of parts with the additive manufacturing device. The one or more software modules can compare the total cost to the cost of a current manufacturing method.

Abstract: A turbine blade having a base and an airfoil, the base including cooling air inlets and an internal cooling air passageway, and the airfoil including an internal multi-bend heat exchange path beginning at the base and ending at a cooling air outlet at the trailing edge of the airfoil. The airfoil also includes a “skin” that encompasses a tip wall, an inner spar, and a tip flag cooling system.

Abstract: The systems and methods disclosed herein include an assessment system and process for assessing an industrial machine and its various sections, sub-sections, and parts. In embodiments, the assessment system includes an assessment device that includes an assessment overviewer and a part assessor. The assessment overviewer provides an assessor with a selectable industrial machine schematic that illustrates the assessment status of each section of the industrial machine, such as by changing the appearance of each section based its assessment status. The part assessor provides a part assessment interface with engineering instructions and a part grading user interface that provides for a comparison of digital images to complete the part grading.

Abstract: This disclosure provides an air tube for a combustor of a gas turbine engine. The air tube includes an inner tube and an outer tube to deliver discharged compressor air into a combustion chamber of the combustor. The air tube can include struts and fins that can improve the cooling performance of the air tube during operation of the gas turbine engine.

Abstract: A turbine blade having a base and an airfoil, the base including cooling air inlets and an internal cooling air passageway, and the airfoil including an internal multi-bend heat exchange path beginning at the base and ending at a cooling air outlet at the trailing edge of the airfoil. The airfoil also includes a “skin” that encompasses a tip wall, an inner spar, and a tip flag cooling system.

Abstract: This disclosure provides a tube shaping tool for a torque mechanism to modify a tube end. The tube shaping tool can include a cup and an internal assembly. The internal tube assembly can include a stem and an internal body. The stem can be received by the torque mechanism and transmit torque generated from the torque mechanism to the components of the tube shaping tool. The tube shaping tool can roll form the outside diameter of the tube end using standard bearings and a flange bearing attached to the internal body. The flange bearing can also include a flange that can provide facing of the tube end.

Abstract: This disclosure provides a seal strip assembly for a sealing strip slot to reduce air leakage between gas turbine nozzle segments. The seal strip assembly includes a bimetal element and a backing plate. The bimetal element includes a first layer with a first thermal coefficient of thermal expansion and a second layer with a second coefficient of thermal expansion. The second coefficient of thermal expansion is lower than the first coefficient of thermal expansion allowing the bimetal element to expand with an increase in surrounding temperature.