Abstract: A liner assembly for a combustor of a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a grommet with a multiple of grommet cooling passages.

Abstract: A gas turbine engine assembly includes a support component, an engine component, and a seal. The annular support component is formed to include a notch. The engine component is mounted in spaced-apart relation to the support component so that a gap is formed between the support component and the engine component. The seal is located in the notch and is adapted to close the gap.

Abstract: Aspects of the disclosure are directed to a cooling design feature for inclusion in a liner of an aircraft, comprising: a plurality of angled holes, and at least one through hole separating all combinations of any two of the angled holes, wherein the at least one through hole is oriented at an angle that is substantially perpendicular to a surface of the liner, and wherein each of the plurality of angled holes are non-parallel to the at least one through hole.

Abstract: A gas turbine cooling system efficiently cools a first-stage turbine wheel and attachment portions of first-stage turbine blades. The gas turbine cooling system includes a group of impingement cooling holes provided in a partition wall as a stationary component that separates an exit space of a compressor and a wheel space located upstream of a turbine wheel for ejecting compressed air in the exit space toward the turbine wheel and attachment portions of the turbine blades.

Abstract: A gas turbine engine. The engine includes a first compressor coupled to a first turbine by a first shaft, the first turbine having first and second turbine stages. A first combustor is provided downstream of the first compressor and upstream of the first stage of the first turbine. A second combustor is provided downstream of the first stage of the first turbine, and upstream of the second stage of the first turbine. A further turbine is provided downstream of the first turbine, and is coupled to a further compressor by a further shaft.

Abstract: Gas turbine unit (GTV) provides compressed air and steam methane-hydrogen mixture to a combustion chamber to enrich combustion products and cooling by evaporation or superheating of water steam. The temperature of heat exchange processes of the gas turbine unit is increased by additional fuel combustion in the steam-methane-hydrogen mixture postcombustion flow extracted at the output from the additional free work gas turbine, and before supply of steam-methane-hydrogen mixture to the combustion chamber it is previously cooled to the temperature of 200+240° C. with simultaneous differential condensation of water steam. The condensate is processed for preparation of methane steam-gas mixture and low pressure water steam which is passed through the additional free work gas turbine.

Abstract: A hybrid diffuser case for a gas turbine engine includes an outer diffuser case wall that extends downstream from an annular outer pre-diffuser case shroud. An inner diffuser case wall extends downstream from an annular inner pre-diffuser case shroud. A multiple of struts extend between the annular outer pre-diffuser case shroud and the annular inner pre-diffuser case shroud.

Abstract: A combustor for use in a gas turbine engine includes a chute and a combustion liner defining a combustion chamber and a chute-receiving aperture that extends through the combustion liner. The chute extends through the chute-receiving aperture of the combustion liner and defines a passageway sized to convey air from an environment outside the combustion chamber through the combustion liner into the combustion chamber.

Abstract: A disc turbine engine includes a multi disc engine in which each disc engine includes a turbine blade, and low-pressure compressor blade, a high-pressure compressor blade and a bearing that runs the disc engine freely around a shaft. Each disc engine has its own cooling system, the compressor's blades act as cooling fins for the turbine blade, and air bleeding from the high-pressure compressor to the lower pressure compressor through a hollow turbine blade. Cooling the nozzle is by attaching the nozzle to the guide fan and by air bleeding through the hollow body. There is no stator in between the disc engine and no large shaft is required because the power produced by each turbine blade is consumed by its own compressor. The weight and cost of this engine will be less than other engines at the same thrust output.

Abstract: A method for operating a closed loop regenerative thermodynamic power generation cycle system is presented. The method includes supplying a high-temperature working fluid stream at a first pressure P1 to an expander, and extracting a partially expanded high temperature working fluid stream from the expander at a second pressure P2. Each of the first pressure P1 and the second pressure P2, are higher than a critical pressure of the working fluid; and the second pressure P2 is lower than P1. The method further includes regeneratively supplying the extracted high temperature working fluid stream at the second pressure P2 to a low temperature working fluid stream at the first pressure P1. A closed loop regenerative thermodynamic power generation cycle system is also presented.

Abstract: A cooled gas turbine engine component comprises a wall having first and second surfaces. The second surface has a plurality of recesses and each recess has a planar upstream end surface arranged so that it hangs over the upstream end of the recess. Each recess has a depth equal to the required depth plus the thickness of the thermal barrier coating to be deposited. The wall has a plurality of angled effusion cooling apertures extending from the first surface towards the second surface. Each effusion cooling aperture has an inlet in the first surface and an outlet in the end surface of a corresponding one of the recesses in the second surface. Each recess has smoothly curved transitions from the end surface and side surfaces to the second surface. Blocking of the effusion cooling apertures by thermal barrier coating is reduced.

Abstract: A method of controlling a supply of a fuel to a combustor of a gas turbine having a compressor upstream of the combustor is provided. The method includes: supplying the fuel to the combustor; obtaining a property value of at least one physical property (PT8, PT7, Tinlet, THBOV) of air used for burning the fuel in the combustor; estimating a heat input (HIengmodel) of the fuel supplied to the combustor based on the property value; measuring a Caloric Value (LCVmea) of the fuel upstream of the combustor; adjusting the estimated heat input (HIengmodel) based on the measured Caloric Value (LCVmea); and controlling a fuel valve regulating the supply of the fuel to the combustor based on the adjusted estimated heat input (HIexpected) and a demanded heat input (FFDEM).

Abstract: A shroud retention system may generally include a shroud hanger and first and second hooked components extending from an outer hanger wall of the hanger. The first hooked component may include a first wall extending from the outer hanger wall and a first rail extending from the first wall. The second hooked component may include a second wall extending from the outer hanger wall and a second rail extending from the second wall. In addition, the system may also include a shroud segment having a shroud wall configured to be positioned radially between the outer hanger wall and the first and second rails. Moreover, the system may include a retention spring positioned within a radial space defined between the outer hanger wall and the first and second rails that is configured to apply a radial spring force against the shroud segment.

Abstract: A gas turbine engine comprises a main fan that delivers air into a bypass duct and into a core engine. A heat exchanger is positioned within the bypass duct and receives a fluid to be cooled from a component associated with the gas turbine engine. A heat exchanger fan is positioned to draw air across the heat exchanger and a control for the heat exchanger fan. The control is programmed to stop operation of the fan during certain conditions, and to drive the heat exchanger fan under other conditions. A method of forming a heat exchanger is also disclosed.

Abstract: Several embodiments include a system and method for providing thrust recovery in an aircraft engine. The system and method enables a wider range of safe operation for aircraft where the aircraft engine is rapidly reactivated from a substantially deactivated state. The method thereby reduces noise output and fuel usage during descent and shortens runway lengths and occupancy time required for landing. Thrust recovery is provided via the use of stored bleed air being re-injected back into the aircraft engine. An onboard air storage tank and a system of valves facilitating this method are disclosed.

Abstract: A liner portion has an exhaust liner portion with an inner face extending to an outward extending flange. Cooling grooves formed in the flange have a radially outward inlet point and a radially inner outlet point. A flow area at the inlet point is smaller than the flow area at the outlet point.

Abstract: A nozzle includes a nozzle body defining a longitudinal axis. The nozzle body includes an air passage having a radial swirler and a converging conical cross-section. A fuel circuit is radially outboard from the air passage with respect to the longitudinal axis. The fuel circuit extends from a fuel circuit inlet to a fuel circuit annular outlet. The fuel circuit includes a plurality of helical passages to mitigate gravitational effects at low fuel flow rates. Each helical passage of the fuel circuit opens tangentially with respect to the fuel circuit annular outlet into an outlet of the air passage.

Abstract: A combustor can assembly includes a plurality of combustor cans spaced circumferentially about a gas turbine engine. Each combustor can is coupled in flow communication with at least one fuel manifold via a respective can fuel line. The combustor can assembly also includes a first interconnecting fuel line that includes a first end and a second end. The first end is coupled in flow communication with the can fuel line of a first combustor can, and the second end is coupled in flow communication with the can fuel line of a second combustor can that is not circumferentially adjacent to the first combustor can. The combustor can assembly further includes a first control device operatively coupled to the can fuel line of the first combustor can. The first control device is operable to change a dynamic operational characteristic of the first and second combustor cans independently of other combustor cans.

Abstract: In an embodiment, a method includes flowing an exhaust gas from a turbine of a gas turbine system to an exhaust gas compressor of the gas turbine system via an exhaust recirculation path; evaluating moist flow parameters of the exhaust gas within an inlet section of the exhaust gas compressor using a controller comprising non-transitory media programmed with instructions and one or more processors configured to execute the instructions; and modulating cooling of the exhaust gas within the exhaust recirculation path, heating of the exhaust gas within the inlet section of the exhaust gas compressor, or both, based on the evaluation.

Abstract: A control valve for a gas turbine engine air starter. The control valve includes a spherical or ball valve element rotatable between a closed position and an open position to connect an internal flow passage to the air starter. A series of camming elements interconnecting a piston and a valve stem for the spherical valve element convert linear displacement of the piston into rotational movement of the valve for rapid opening.