|| List of recent Gas Turbine-related patents
| Aircraft engine systems and methods for operating same|
A device and a method for determining a residual life expectancy of a rotor of a gas turbine. The method includes receiving at a computer operating conditions of the gas turbine, receiving a gas turbine rotor inspection result, updating, based on the operating conditions of the gas turbine and the gas turbine rotor inspection result, a database for a fleet corresponding to the gas turbine, and calculating the residual life expectancy of the rotor of the gas turbine..
| System and method for 0n-wing engine trim verification|
Devices and methods relating to gas turbine engines and engine temperature trim verification are disclosed. An exemplary method comprises acquiring signals representing a plurality of engine parameters measured while the engine is operating and determining a recommended trim thermocouple resistance based at least partly on the measured parameters.
| Systems and methods for use in adapting the operation of a gas turbine|
A control system for use in adapting the operation of a gas turbine is provided. The control system is configured to be coupled to at least one component of the gas turbine.
| Solid oxide fuel cell system|
A solid oxide fuel cell system (10) comprises a solid oxide fuel cell stack (12) and a gas turbine engine (14). The solid oxide fuel cell stack (12) comprises a plurality of solid oxide fuel cells (16).
| Edge seal for gas turbine engine ceramic matrix composite component|
A gas turbine engine ceramic matrix composite (cmc) component includes first and second outer layers of plies, and an intermediate layer of plies between the first and second outer layers of plies. The intermediate layer of plies is offset relative to the first and second outer layers of plies.
| Split coating mask system for gas turbine engine component|
A mask assembly for a gas turbine engine component includes a second mask that at least partially overlaps a first mask to fit a platform of a gas turbine engine component.. .
| Blade leading edge tip rib|
A rotor blade for a gas turbine engine includes a leading edge tip rib projecting outwardly from an airfoil of the blade at a tip region thereof. The tip rib continuously surrounds a leading edge of the airfoil and extends rearwardly from the leading edge along respective pressure and suction side surfaces to thereby alter the blade tip leakage vortex structure and strength, resulting in a stage efficiency benefit..
| Gas turbine engine blade and disk|
In some examples, an apparatus may include a gas turbine engine component comprising a blade, a platform forming an airflow surface from which the blade extends on one side, and a motion weld receiving surface disposed on a non-airflow side of the platform, and a disk having a motion weld disk portion to receive the motion weld receiving surface of the gas turbine engine component and form a motion weld coupling when the disk and gas turbine engine component are coupled, wherein, prior to coupling, the gas turbine engine component further includes a blade stalk disposed on the non-airflow side of the platform and in which is formed the motion weld receiving surface.. .
| Gas turbine engine component|
Embodiments are disclosed of a turbomachinery component having an outer part that is captured between a cap and a base of the component. In one form, the outer part is placed in compression between the cap and the base.
| Exhaust section for bypass gas turbine engines|
A turbine exhaust section comprises a turbine exhaust case having radially outer and inner ducts defining therebetween an annular exhaust portion for the hot exhaust gases, and an exhaust mixer projecting axially rearwardly from the turbine exhaust case for mixing the hot exhaust gases with a cooler bypass duct flow. The upstream end of the exhaust mixer surrounds a downstream end of the outer duct and defines therewith an axially extending overlap joint with a radial play between the outer duct and the exhaust mixer.
| Flight gas turbine with a first rotatable shaft|
The present invention describes an aircraft gas turbine with a first rotatable shaft and a second shaft arranged coaxially thereto and which at least in an area close to the shaft end is non-rotatably connected to the first shaft. Recesses are provided in shaft areas at a distance from the connecting area between the shafts, said recesses overlapping one another at least in some areas when a defined twist of the first shaft relative to the second shaft is exceeded.
| Flow diverter to redirect secondary flow|
An assembly for a gas turbine engine includes a seal and a flow diverter. The flow diverter is disposed adjacent the seal to direct a secondary gas flow that passes across the seal away from a rotor cavity such that the secondary gas flow travels back toward a main gas flow path of the gas turbine engine..
| Gas turbine engine systems involving tip fans|
Gas turbine engine systems involving tip fans are provided. In this regard, a representative gas turbine engine system includes: a multi-stage fan having a first rotatable set of blades and a second counter-rotatable set of blades, the second rotatable set of blades defining an inner fan and a tip fan and being located downstream of the first set of rotatable blades; and an epicyclic differential gear assembly operative to receive a torque input and differentially apply the torque input to the first set of blades and the second set of blades..
| Active turbine or compressor tip clearance control|
A gas turbine engine includes an annular plenum defined with an outer skin and a perforated inner skin for receiving selective air flow to impinge a support case which supports shrouds of the rotor assemblies of the engine therein for active tip clearance control of the rotor assemblies. In one embodiment a bobbin-type transfer tube for supplying cooling air into the plenum, is provided between an outer case of the engine an the plenum such that the thermally induced relative movement of the outer case and the plenum is permitted..
| Methods and apparatus for measuring axial shaft displacement within gas turbine engines|
In some embodiments, an apparatus includes a target member, a sensor and a mounting assembly. The target member is coupled to a shaft that is disposed within an engine housing and that rotates about an axis.
| Axial oil scoop for a gas turbine engine|
An axial oil scoop for a gas turbine engine includes an outer surface with a concave profile, the outer surface includes a multiple of radial holes.. .
| Air particle separator|
A gas turbine engine is disclosed having a particle separator and an ejector disposed downstream of the particle separator and structured to entrain a dirty flow from the separator. A container of working fluid can be placed in flow communication with the ejector to provide an ejector flow to entrain the dirty flow from the particle separator.
| Method for providing a frequency response for a combined cycle power plant|
The disclosure refers to a method for providing a frequency response for a combined cycle power plant connected to an electric grid. The combined cycle power plant includes a gas turbine engine and a steam turbine engine.
| Gas turbine engine system and supersonic exhaust nozzle|
One embodiment of the present invention is a unique gas turbine engine system. Another embodiment is a unique exhaust nozzle system for a gas turbine engine.
| Gas turbine engine buffer system|
A gas turbine engine includes a fan, a compressor section, and a turbine section configured to drive the compressor section and the fan. A buffer system is configured to communicate a buffer supply air to a portion of the gas turbine engine.
| Combined cycle power plant|
A combined cycle engine is used to provide power to a vehicle. In one form the combined cycle engine includes two engines coupled through a gearbox.
| Gas turbine variable focus laser ignition|
A laser ignition system for a gas turbine engine includes a combustion chamber. The system comprises a laser source for generating a continuous laser beam during an ignition process of the combustion chamber; and a dynamic laser focus apparatus positioned outside of the combustion chamber and focusing the laser beam into a continuously varying focal point to generate a laser kernel moving within a spray of air/fuel mixture injected into the combustion chamber..
| Gas turbine engine and method for operating a gas turbine engine|
One embodiment of the present disclosure is a unique method for operating a gas turbine engine during flight operation of the gas turbine engine in an aircraft. Another embodiment of the present disclosure is a unique gas turbine engine.
| Method for adjusting a natural gas temperature for a fuel supply line of a gas turbine engine|
The method for adjusting a natural gas temperature for a fuel supply line of a gas turbine engine includes measuring by infrared analysis the natural gas percentage content of methane (ch4), ethane (c2h6), propane (c3h8), butane (c4h10), carbon dioxide (co2), calculating the nitrogen (n2) percentage content as the complement to 100 of the measured percentage content of methane (ch4), ethane (c2h6), propane (c3h8), butane (c4h10), carbon dioxide (co2), calculating an index indicative of the natural gas energy content and adjusting the natural gas temperature on the basis of the index.. .
| Systems and methods to control combustion dynamic frequencies|
Systems and methods for frequency separation in a gas turbine engine are provided herein. The systems and methods for frequency separation in a gas turbine engine may include determining a hot gas path natural frequency, determining a combustion dynamic frequency, and modifying a compressor discharge temperature to separate the combustion dynamic frequency from the hot gas path natural frequency..
| Flow conditioner in a combustor of a gas turbine engine|
A combustor in a gas turbine includes a liner having an interior volume defining a main combustion zone, a fuel injection system for delivering fuel into the main combustion zone, and a flow sleeve that defines, with the liner, a passageway for air to flow on its way to be mixed with fuel from the fuel injection system, wherein the mixture is burned in the main combustion zone to create hot combustion gases. The combustor further includes a flow conditioner including at least one panel having a configuration such that air is able to pass through the panel(s) on its way to the passageway, wherein at least a substantial portion of the air that enters the passageway for being burned in the main combustion zone passes through the panel(s)..
| Thermally compliant dual wall liner for a gas turbine engine|
A stiffener for a liner assembly of a gas turbine engine includes a second resilient member arranged on a first resilient member at substantially ninety (90) degree orientation.. .
| Rich burn, quick mix, lean burn combustor|
A combustor for a gas turbine includes a fuel nozzle having a central swirler that circumferentially surrounds a downstream end of the fuel nozzle. A primary combustion zone is defined within the central swirler.
| Variable area fan nozzle with wall thickness distribution|
A gas turbine engine includes a core engine that has at least a compressor section, a combustor section and a turbine section disposed along a central axis. A fan is coupled to be driven by the turbine section.
| Laser-ignition combustor for gas turbine engine|
The combustor has a laser ignitor mounted to the casing, remotely from the liner of the combustion chamber. The laser ignitor has an igniter beam path for igniting the fuel and air mixture in the combustion chamber, the igniter beam path extending at least partially across the air plenum surrounding the liner and into the combustion chamber through a corresponding beam path aperture provided in the liner..
| Gas turbine engine variable geometry flow component|
A variable geometry mechanism suitable for use in a gas turbine engine is disclosed in which movable vane segments which are coupled to a rotatable ring, or rings, are used to change an aerodynamic property of a working fluid flowing through the gas turbine engine. The movable vane segments can be rotated through the ring, or rings, between a first position associated with the first vane and a second position associated with a second vane to place the movable vane segments in proximity to one or the other of the first and second vanes of the gas turbine engine.
| Gas turbine inlet filter with replaceable media cartridges|
A filter arrangement for a system within which fluid is filtered and an associated method of providing the arrangement. A frame of the arrangement includes an outer periphery shape that is complementary to a shape for location of the filter arrangement therein and a plurality of filter cartridge seat slots arranged in v-shaped pairs.
|Determining the deterioration of a gas turbine engine in use|
This invention provides a method and apparatus to identify fod or bird impact to gas turbine fan blades, assessing the damage that may have occurred whist still in flight and determining post impact actions, including replacement parts.. .
|Ceramic powders and methods therefor|
A ceramic powder and method of forming the ceramic powder capable of being used in coatings to allow components to survive in high temperatures environments, such as the hostile thermal environment of a gas turbine engine. The ceramic powder includes powder particles each having an inner core formed of a first material and an outer region formed of a second material.
|Coating, coating layer system, coated superalloy component|
Coatings as may be used in a gas turbine are provided. A nickel based coating may include 15 to 40 wt % cobalt, 10 to 25 wt % chromium, 5 to 15 wt % aluminum, 0.05 to 1 wt % yttrium and/or at least one of elements from lanthanum series, 0.05 to 8 wt % ruthenium or iron, 0 to 1 wt % iridium, 0.05 to 5 wt % molybdenum, 0 to 3 wt % silicon, 0 to 5 wt % tantalum, 0 to 2 wt % hafnium, unavoidable impurities, and a balance of nickel.
|Method of protecting a surface|
A method of masking part of a surface of a wall of a gas turbine component including at least one area having cooling holes defined therein, the method including applying a viscous curable masking compound to the part of the surface over an entirety of each of the at least one area, including blocking access to the cooling holes from the surface by applying the masking compound over the cooling holes without completely filling the cooling holes with the masking compound, and forming a respective solid masking element completely covering each of the at least one area and the cooling holes defined therein by curing the masking compound.. .
|Turbine engine comprising a metal protection for a composite part|
A gas turbine engine including at least a first composite part configured for mounting in a second metal part of the engine, the first composite part including an interface surface configured to be in surface contact with the second metal part, the engine including a metal protection removably mounted on the first composite part and configured to cover the interface surface.. .
A blade at least part of which is arranged in use to rotate outside of a casing of a gas turbine engine. The blade has a main body, a tip fence and a blend region defining a curve between the main body and the fence.
|Modular blade or vane for a gas turbine and gas turbine with such a blade or vane|
The invention relates to a modular blade or vane for a gas turbine, which includes the modular components of a platform element with a planar or contoured surface defining a platform level and a through-opening therein, and an airfoil, extending through the platform element. The airfoil includes a load carrying structure extending along a longitudinal axis of the airfoil, having a root portion for fastening on a blade or vane carrier of the gas turbine, having a tip portion, and having at least one interior passage, extending from the root portion to the tip portion of the airfoil.
|Bumper for synchronizing ring of gas turbine engine|
A synchronizing assembly for a gas turbine engine has a synchronizing ring. A bumper assembly has a cradle and a bumper held within the cradle.
|Geared architecture for high speed and small volume fan drive turbine|
A gas turbine engine includes a flex mount for a fan drive gear system. A very high speed fan drive turbine drives the fan drive gear system..
|Method for setting a gear ratio of a fan drive gear system of a gas turbine engine|
A gas turbine engine includes a fan section including a fan rotatable about an axis. A speed reduction device is connected to the fan.
|Cooling system of ring segment and gas turbine|
A cooling system of ring segment is provided with: a collision plate that has a plurality of small holes; a cooling space that is enclosed by the collision plate and a main body of the segment body; a first cavity that arranged is the upstream end portion of the segment body in the flow direction of the combustion gas so as to be perpendicular to the axial direction of a rotating shaft; a first cooling passage that communicates from the cooling space to the first cavity; and a second cooling passage that communicates from the first cavity to a fire combustion gas d gas space in the downstream end portion of the segment body in the flow direction of the combustion gas.. .