Abstract: A method for inspecting a gas turbine engine using a maintenance tool includes assembling a plurality of rail segments of a rail system of the maintenance tool within a core air flowpath of the gas turbine engine. The gas turbine engine includes a compressor section, a combustion section, and a turbine section together defining at least in part the core air flowpath. The method also includes moving a maintenance head of the maintenance tool along the plurality of rail segments of the rail system to perform maintenance operations within the core air flowpath of the gas turbine engine.

Abstract: A dual wall liner for a gas turbine engine may comprise a shell having a first side and a second side, a panel contacting the shell, the panel at least partially defining a hot gas path through which a hot gas flows, wherein the first side of the shell faces the panel, wherein the shell includes a thermal barrier coating (TBC) disposed on the first side of the shell. The TBC may thermally protect the shell from heat from a hot gas path.

Abstract: A maintenance tool for gas turbine engine includes a rail system having a plurality of rail segments insertable through one or more inspection holes of the gas turbine engine for assembly within a core air flowpath of the gas turbine engine. The maintenance tool additionally includes a maintenance head movable along the plurality of rail segments of the rail system for performing maintenance operations within the core air flowpath.

Abstract: A booster splitter for a gas turbine engine and a method of additively manufacturing the booster splitter are provided. The booster splitter includes an annular outer wall defining an internal fluid passageway in fluid communication with a fluid supply and terminating in discharge ports that eject a flow of fluid into the compressor section of the gas turbine engine. The internal fluid passageway may also be in fluid communication with heating plenums of a first plurality of airfoils for heating those airfoils.

Abstract: A combustion device burns fuel ammonia in a combustor using combustion air, and includes a catalyst reduction unit which is configured to reduce nitrogen oxides in a combustion exhaust gas supplied from the combustor, in which at least a part of the fuel ammonia is supplied to the catalyst reduction unit as a reducing agent for the nitrogen oxides in the combustion exhaust gas.

Abstract: A system is provided for multi-engine coordination of gas turbine engine motoring in an aircraft. The system includes a controller operable to determine a motoring mode as a selection between a single engine dry motoring mode and a multi-engine dry motoring mode based on at least one temperature of a plurality of gas turbine engines and initiate dry motoring based on the motoring mode.

Abstract: A core duct assembly for a gas turbine engine, the core duct assembly including: a core duct including an outer and an inner wall, the outer wall having an interior surface; a gas flow path member extending across the gas flow path at least partly between the inner and outer walls, the rotor blade having a radial span extending from a blade platform to a blade tip, wherein an upstream wall axis is defined as an axis tangential to a point on a first portion of the interior surface of the outer wall of the core duct extending downstream from the gas flow path member, the upstream wall axis lying in a longitudinal plane of the gas turbine engine containing the rotational axis of the engine, and wherein the upstream wall axis intersects the rotor blade at a point spaced radially inward from the blade tip of the rotor blade.

Abstract: A method for controlling thrust from a turbojet that is fuel flow rate regulated by a high limit value for providing protection against surging of a compressor of the turbojet is provided. The method includes: obtaining a first thrust value corresponding to a first operating point of the compressor on the high limit value, the high limit value taking account of an underestimate of the fuel flow rate; controlling the turbojet to reach the first thrust value; monitoring the turbojet to detect underspeed of the compressor; and where applicable: obtaining a second thrust value corresponding to a second operating point that guarantees a predetermined margin relative to the high limit value so as to obtain protection against underspeed of the turbojet; and controlling the turbojet to reach the second value.

Abstract: A fuel assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a fuel valve that meters flow of fluid between a combustor and a fuel supply, and a controller in communication with the fuel assembly. The controller is programmed to receive data corresponding to a present rotational speed of a gas turbine engine component and data corresponding to at least one present flight condition, and is programmed to cause a flow rate from the fuel valve to change in response to determining that a rate of change in an acceleration rate relating to the present rotational speed meets at least one predetermined threshold. At least one predetermined threshold relates to an engine light-off event of the combustor and is based upon the at least one present flight condition.

Abstract: The regenerative cooling system (100) is provided for a regenerative heat engine (1) and comprises a cooling chamber (79) which surrounds a gas expander (78), leaving open a gas circulation space (80) between said chamber (79) and said expander (78), a working gas (81) expelled from the gas expander (78) circulating in said space (80) before returning to a regenerative heat exchanger (5) where it is cooled, a large portion of the heat of said gas (81) being reintroduced into the thermodynamic cycle of the regenerative heat engine (1).

Abstract: A gas turbine engine generates noise during use, and one particularly important flight condition for noise generation is take-off. A gas turbine engine that has high efficiency provides low noise, in particular from the fan and the turbine that drives the fan. Values are defined for a noise parameter NP that results in a gas turbine engine having reduced combined fan and turbine noise.

Abstract: A method of controlling plume exhaust heat and/or noise radiation from a turbofan engine assembly having a short nacelle. A mixer duct shell is supported such that a downstream edge of the short nacelle overlays an upstream portion of the mixer duct shell. A first portion of fan exhaust may be routed through the mixer duct shell between its inner surface and an outer surface of a core engine shroud. A second portion of fan exhaust may be routed over an outer surface of the mixer duct shell. At least one of the inner surface and an outer surface of the mixer duct shell may have an acoustic lining including a honeycomb core structure.

Abstract: A fuel nozzle for a gas turbine engine is generally provided. The fuel nozzle includes an outer sleeve extended circumferentially around a fuel nozzle centerline and extended along a longitudinal direction substantially co-directional to the fuel nozzle centerline. The outer sleeve defines a plurality of first radially oriented air inlet ports through the outer sleeve in circumferential arrangement relative to the fuel nozzle centerline. The fuel nozzle further includes a centerbody positioned radially inward of the outer sleeve. The centerbody is extended along the longitudinal direction substantially co-directional to the fuel nozzle centerline and wherein the centerbody is concentric to the fuel nozzle centerline and the outer sleeve. The centerbody defines a plurality of second radially oriented air inlet ports through the centerbody in circumferential arrangement relative to the fuel nozzle centerline.

Abstract: A combustor (3) includes a plurality of first burners extending along an axis (P) and arranged at intervals in the circumferential direction, a base plate (31) having support openings configured to support the plurality of first burners and a plurality of through-holes (34) through which base plate air flows toward a downstream side, a plurality of extension pipes supported by the base plate (31) to correspond to the first burners and configured to guide a pre-mixed gas ejected from the first burners toward the downstream side, and a base plate air guide section (35) configured to change a direction of at least some of the base plate air ejected from the through-holes (34) at a downstream side of the base plate (31).

Abstract: A gas turbine engine system for detecting control sensor override includes a plurality of temperature sensors coupled to the gas turbine engine system. The temperature sensors are configured to generate a plurality of signals representative of exhaust gas temperatures of the gas turbine engine. The system includes an on-site monitoring system coupled in communication to the plurality of temperature sensors. The on-site monitoring system has a processor programmed to continuously receive the plurality of signals from the temperature sensors. In addition, the processor is programmed to analyze the plurality of signals to verify the accuracy of the exhaust gas temperatures associated with the plurality of signals, and to detect a jumpered temperature sensor of the plurality of temperature sensors.

Abstract: A combustor system for a GT system may include: a plurality of burners, each burner including an inflow region for receiving a combustion air flow and a mixing zone disposed downstream of the inflow region for receiving the air flow and a fuel flow; a combustion chamber disposed downstream of the mixing zone; a fuel flow valve system disposed to control the fuel flow to each of the plurality of burners; a combustion sensor configured to determine a combustion parameter; and an exhaust sensor configured to determine an exhaust parameter. A control system may be connected to the combustion sensor, the exhaust sensor and fuel flow valve system. The control system, in response to the gas turbine system operating at a low partial load, redistributes the fuel flow to at least one burner of the plurality of burners as a function of a predetermined emission limit.

Abstract: An exemplary tiltrotor aircraft with a hybrid drive system includes a first propulsion system having a first engine and a first supplemental driver operably coupled to a first proprotor that is operable between a helicopter mode and an airplane mode and a second propulsion system having a second engine and a second supplemental driver operably coupled to a second proprotor that is operable between a helicopter mode and an airplane mode.

Abstract: A fuel injection system may include a fuel spray nozzle including a nozzle head and a nozzle stem. The nozzle head may include an air channel and a swirler in fluid communication with the air channel. The nozzle stem may extend into a compressor discharge pressure cavity and convey fuel to the air channel. The air channel may combine air from the swirler with the fuel from the nozzle stem and convey a mixture of fuel and air to a combustor. The fuel injection system may further include a scoop. The scoop may be coupled to the fuel spray nozzle. The scoop may receive air that flows into the compressor discharge pressure cavity from a diffusor. An outer surface of the fuel spray nozzle and an inner surface of the scoop define a duct in fluid communication with the swirler. The swirler may receive air from the duct.

Abstract: A gas turbine system, and a control apparatus and method thereof, can control the amount of compressed air supplied to each of a combustor and an anti-icing unit in order to improve the droop control performance of the gas turbine system. The control apparatus of the gas turbine system may include a sensing unit to measure a rotor speed of the turbine; and a compressed air distribution unit to adjust a distribution ratio of the compressed air supplied to the combustor to the compressed air supplied to the anti-icing unit, based on the measured rotor speed. The control apparatus can increase the amount of compressed air supplied to the combustor even though the revolution number of the turbine connected with a system frequency may be significantly lowered.

Abstract: A gas turbine engine combustor includes a liner defining at least in part a combustion chamber, a first side exposed to the combustion chamber, a second side opposite the first side, and a dilution hole extending from the second side to the first side. The liner includes an airflow feature on the first side of the liner adjacent to the dilution hole and extending into the combustion chamber to increase a cooling of the liner.