Abstract: A disclosed airfoil health monitoring system and method obtains a signal comprising a waveform indicative of an airfoil path with a sensor. Features of the waveform are determined and compared waveform characteristics indicative of a vibrational mode. A vibrational mode of the airfoil may then be determined based on the comparison between the predetermined waveform characteristics and the obtained waveform indicative of the airfoil path.

Abstract: A turbine blade for a gas turbine engine includes an airfoil that includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface extending from a platform in a radial direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord. A circumferential coordinate is scaled by a local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading and trailing edges at the span location.

Abstract: A turbine blade for a gas turbine engine includes an airfoil that includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface that extends from a platform in a radial direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord. A circumferential coordinate is scaled by a local axial chord, and a span location, wherein the local axial chord corresponds to a width of the airfoil between the leading and trailing edges at the span location.

Abstract: A gas turbine engine having an engine casing extending circumferentially about an engine centerline axis; and a compressor section, a combustor section, and a turbine section within said engine casing. At least one of said compressor section and said turbine section includes at least one airfoil and at least one seal member adjacent to the at least one airfoil, wherein a tip of the at least one airfoil is metal having a thin film ceramic coating and the at least one seal member is coated with an abrasive.

Abstract: A gas turbine engine is provided, which includes a first compressor rotor coupled to a first shaft, a second compressor rotor downstream of the first compressor and coupled to a second shaft. A combustor is disposed downstream of the compressor rotors. A turbine is disposed downstream of the combustor and coupled to the second shaft. A gear is configured to be driven by the second shaft. The second shaft is configured for driving the first shaft through the gear, whereby the turbine drives the first compressor rotor at a different rotational speed than the turbine.

Abstract: A turbomachine combustor nozzle includes a first plate member having a first plurality of openings, and a second plate member having a second plurality of openings that are configured and disposed to be in alignment with the first plurality of openings. A plurality of nozzle members extends through corresponding ones of the first plurality of openings and the second plurality of openings. Each of the plurality of nozzle members includes a solid core.

Abstract: A flame detector in photonic communication with a no-flame region of a combustor of a gas turbine may emit a signal when a photon is detected. A controller may be arranged to receive a signal from the flame detector and may determine whether a flame presence in the no-flame region is indicated responsive to the signal.

Abstract: A fan drive gear system for a gas turbine engine includes a carrier supporting circumferentially arranged gears within gear mount sections spaced circumferentially about a periphery. A shelf is disposed between each of the gear mounting sections. A torque frame is attached to the carrier and includes circumferentially arranged finger sections. Each of the finger sections includes first and second ends that are spaced apart to receive the shelf. A pin extends through openings in the shelf and first and second ends to secure the torque frame to the carrier.

Abstract: A gas turbine engine rotor section includes a rotor body with a ledge extending axially from a location on the rotor body. The ledge defines a radially inner surface radially inwardly of the ledge, and a hub extends axially from the rotor, and beyond the ledge. The hub has a radially outer surface spaced from the ledge radially inner surface. A first distance is defined between the radially inner surface of the ledge and the radially outer surface of the hub. A knife edge seal has at least one pointed knife seal portion at a radially outer end. A radially inwardly extending arm portion extends from the seal, and an axially inwardly extending portion extends axially inwardly from the radially inwardly extending portion. The axially inwardly extending portion has a radially outer face and a radially inner face, which are spaced by a second distance. The second distance is less than the first distance.

Abstract: A combustion system including a combustor; a combustor liner disposed within the combustor is provided. At least one primary fuel nozzle is provided to provide fuel to a primary combustion zone disposed proximate to the upstream end of the combustor liner. A transition duct is coupled to the downstream end of the combustor liner. A secondary nozzle assembly is disposed proximate to the downstream end of the combustor to provide fuel to a secondary combustion zone at locations predetermined to reduce peak thermal loads on the surface area of the transition duct.

Abstract: An example airfoil for a gas turbine engine includes a body having a first surface extending from a first edge to a second edge and a cavity disposed in the body. A first cover is at least partially disposed within the cavity. The first cover includes a first portion cooperates with a corresponding second portion. A second cover covers the first cover and forms at least a portion of the first surface with the body. The first cover is disposed between the body and the second cover. The first cover and the second cover have a different coefficient of thermal expansion than the body.

Abstract: A disclosed gas turbine engine includes a first fan section including a plurality of fan blades rotatable about an axis, a compressor in fluid communication with the first fan section, a combustor in fluid communication with the compressor and a first turbine section in fluid communication with the combustor. The first turbine section includes a low pressure turbine that drives the first fan section. A second fan section is supported between the first fan section and the compressor and is driven by a second turbine section disposed between the second fan section and the compressor for driving the second fan section.

Abstract: A rotor body rotates about an axis of rotation. A ledge provides a holding structure for holding blades. A plurality of blades are positioned beneath the ledge. load slot is sized to allow a mount portion of the blades to be moved radially inwardly of the ledge. The blades are moved circumferentially to have the mounted structure radially inwardly of the ledge. A lock slot is positioned on one circumferential side of the load slot. The lock slot is formed to receive a lock, and the lock is partially received within a portion of at least one of the blades, to lock the blades within the rotor, and a shield slot on a second circumferential side of the load slot. The shield slot is sized to be different from the lock slot such that a lock cannot be inadvertently positioned within the shield slot.

Abstract: A turbine blade is disclosed. The turbine blade includes a platform, an airfoil extending from one side of the platform, a root extending radially from another side of the platform, and a pocket located beneath the platform. The pocket is defined by a plurality of walls, and a pad is disposed in a corner of the pocket. The pad includes three pad corners and three sides connecting the three pad corners, wherein each side extends along a different one of the plurality of walls.

Abstract: A liner for a gas turbine engine includes a liner defining an inner surface exposed to exhaust gases and a duct spaced radially outward of the liner. A plurality of hanger assemblies is disposed within the radial space between the liner and the duct for supporting the liner relative to the duct. Each of the hanger assemblies includes a cable having a first end attached to the duct and a second end attached to the liner.

Abstract: A method of protecting operating margin of the gas turbine engine includes calculating an aerodynamic distortion of air entering an inlet of a gas turbine engine that has a compressor section with variable vanes that are movable subject to a control parameter. The control parameter is selectively modified in response to the aerodynamic distortion.

Abstract: The present disclosure describes a turbine blade for a turbine section of a gas turbine engine. The turbine blade includes an airfoil, a platform extending from one side of the airfoil, a root extending from the platform, and at least one purging fin. The at least one purging fin is connected to the root and an underside of the platform, and the at least one purging fin extends along a wall of the root.

Abstract: A lubrication system for a fan drive gear system includes a main lubrication system and an auxiliary lubrication system. The auxiliary lubrication system including a collection channel disposed about the fan drive gear system for collecting expelled lubricant and an auxiliary pump including an inlet receiving lubricant from the collection channel and an outlet in communication with an auxiliary passage. The auxiliary pump supplies lubricant through the auxiliary passages to a bearing passage for communicating lubricant to a bearing. The auxiliary passages include a reservoir after the outlet of the auxiliary pump and before the bearing passage for storing lubricant.

Abstract: A turbomachine combustor includes a combustor cap having a cap surface and a wall that define, at least in part, a resonator volume. A plurality of injection nozzle members extend from the cap surface. Each of the plurality of injection nozzle members include an inner nozzle member and a plurality of outer nozzle members. An adjustable conduit extends through the wall into the resonator volume. The adjustable conduit includes an internal passage having a dimensional parameter. A combustor dynamics mitigation system is operably connected to the combustor cap. The combustor dynamics mitigation system includes a controller configured and disposed to control one a size of the resonator volume and the dimensional parameter of the adjustable conduit to modify combustor dynamics in the combustor.

Abstract: A combustion turbine engine that includes: a compressor; a combustor that receives fuel from a fuel line; a turbine; a heat exchange portion comprising a portion of the fuel line in heat transfer relationship with a heat source for heating the fuel; a rapid heating value meter disposed to test the heating value of the fuel that is configured to provide heating value test results within approximately 1 minute; a cold leg bypass comprising a fuel line that bypasses the heat exchange portion, the cold leg bypass being connected to the fuel line at an upstream fork and at a fuel mixing junction; and valves for controlling the fuel being directed through the heat exchange portion and the fuel being direct through the cold leg bypass; wherein the length of fuel line between the fuel mixing junction and the combustor is less than 20 meters.

Abstract: A nacelle structure for a gas turbine engine assembly includes a fan case, an inlet, and a noise attenuation device. The fan case is configured to be disposed about a fan section of the gas turbine engine, which fan section has a diameter D. The inlet is attached to the fan case and extends axially forward of the fan case. A hilite of the inlet is spaced a distance L from a region in which the fan section is configured to be disposed. A ratio L/D is less than about 0.6. The noise attenuation structure covers a portion of an inner surface of the fan case and the inlet.

Abstract: A fan blade having a body with a dovetail and an airfoil extending radially outwardly. The airfoil includes a pair of skins spaced to form an internal core, which define a pressure side and a suction side, and extending from a radially inner end to a radially outer tip. The core receives a plurality of braided tubes, with the tubes extending with at least a component in a radially outward direction. A fan and an engine are also described.

Abstract: An airfoil includes an airfoil body that has a leading edge and a trailing edge and a first sidewall and a second sidewall that is spaced apart from the first sidewall. The first sidewall and the second sidewall join the leading edge and the trailing edge and at least partially define a cavity in the airfoil body. A damper member is enclosed in the cavity and is loose within the cavity.

Abstract: A gas turbomachine inlet system includes a duct member having an inlet portion fluidically coupled to an outlet portion through an intermediate portion. The inlet portion, outlet portion, and intermediate portion define a fluid flow zone. A throttling system is arranged in the duct member at one of the inlet portion, outlet portion and intermediate portion. The throttling system is configured and disposed to selectively establish a pressure drop through the fluid flow zone. A method of controlling inlet pressure drop through an inlet system for a gas turbomachine is also described herein.

Abstract: An airfoil includes an airfoil body that has a leading edge and a trailing edge and a first sidewall and a second sidewall that is spaced apart from the first sidewall. The first sidewall and the second sidewall join the leading edge and the trailing edge and at least partially define a cavity in the airfoil body. A damper member is enclosed in the cavity and is free-floating within the cavity.

Abstract: A turbomachine system includes a rotor that defines a longitudinal axis of the turbomachine system. A first blade is coupled to the rotor, and the first blade has first and second laminated plies. A first band is coupled to the first blade and is configured to resist separation of the first and second laminated plies.

Abstract: An airfoil includes a body that has a platform, an airfoil extending outwardly from a side of the platform and a root extending outwardly from another side of the platform. A member is connected in an articulated joint to the body.

Abstract: A turbomachine with at least one combustion chamber and at least one turbine which lies downstream of it. The turbine has at least one support element and a lining element connected to the support element and formed as a heat shield. A gap sealed by means of a seal, is formed between the at least one combustion chamber and the turbine, an exchangeable insert is arranged in the gap on the end face on the support element and protects at least the support element against hot gases which penetrate into the gap, and against oxidation which is associated with it.

Abstract: A gas turbine engine has a fan at an axially outer location. The fan rotates about an axis of rotation. The fan delivers air into an outer bypass duct, and across a booster fan positioned radially inwardly of the outer bypass duct. The booster fan delivers air into a radially middle duct, and across a cold turbine into a radially inner core duct being directed into a compressor. From the compressor, air flows axially in a direction back toward the fan through a combustor section, and across an exhaust of the turbine section as directed into the middle duct. A gear reduction drives the fan from a fan drive turbine section. A method is also disclosed.

Abstract: A nose cone for a turbofan gas turbine engine includes a central tip, an outer perimeter and a substantially conical outer wall extending therebetween which encloses a cavity therewithin. The outer wall includes an inner substrate layer facing the cavity and an outer layer which overlies and at least partially encloses the inner substrate layer. The outer layer is composed entirely of a nanocrystalline metal forming an outer surface of the nose cone.

Abstract: A rotor includes a disk that has slots circumferentially arranged around its periphery. Blades include respective roots that are mounted in respective ones of the slots. The roots are smaller than the slots such that there are circumferential gaps between the roots and circumferential sides of the slots. Wedges are respectively located within the circumferential gaps. The wedges are free floating with regard to the blades and the disk.

Abstract: A Brayton cycle heat engine includes a compressor, a combustion chamber, and a turbine. The Brayton cycle heat engine also includes a heat exchanger positioned at least partially within the combustion chamber. The heat exchanger is configured to deliver thermal energy to the combustion chamber from an external source, heating air entering the combustion chamber from the compressor, where the air exits the combustion chamber and drives the turbine.

Abstract: An airfoil for a gas turbine engine includes a substrate and a sheath providing an edge. A cured adhesive secures the sheath to the substrate. The cured adhesive has a fillet that extends beyond the edge that includes a mechanically worked finished surface. A method of manufacturing the airfoil includes the steps of securing a sheath to a substrate with adhesive, curing the adhesive, and mechanically removing a portion of the adhesive extending beyond the sheath.

Abstract: A gas turbine engine has a fan section, a gear arrangement configured to drive the fan section, a compressor section, including both a low pressure compressor section and a high pressure compressor section. A turbine section is configured to drive the compressor section and the gear arrangement. An overall pressure ratio provided by the combination of a pressure ratio across the low pressure compressor section and a pressure ratio across the high pressure compressor section is greater than about 35. The pressure ratio across a first of the low and high pressure compressor sections is between about 3 and about 8. The pressure ratio across a second of the low and high pressure compressor sections is between about 7 and about 15. The fan is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct.

Abstract: An arrangement (100) for delivering combustions gas from a plurality of combustors onto a first row of turbine blades along respective straight gas flow paths, including: a hoop structure (104) at a downstream end of the arrangement and defining at least part of an annular chamber (24); and a plurality of discrete ducts (102), each disposed between a respective combustor and the hoop structure (104). Each duct (102) is secured to the hoop structure (104) at a respective duct joint (116). The hoop structure (104) includes a quantity of hoop segments (105, 130, 132) that is less than a quantity of ducts (102).

Abstract: A gas turbine engine system includes a fan section, a low pressure compressor section downstream of the fan section, a first engine core downstream from the low pressure compressor section, a second engine core downstream from the low pressure compressor section, and a flowpath control mechanism configured to selectively restrict fluid flow through the second engine core. The first engine core includes a first engine core compressor section, a first engine core combustor downstream of the first engine core compressor section, and a first engine core turbine section downstream of the first engine core combustor. The second engine core includes a second engine core compressor section, a second engine core combustor downstream of the second engine core compressor section, and a second engine core turbine section downstream of the second engine core combustor.

Abstract: The utility model relates to hybrid drive for vehicles that significantly decreases the consumption of fuel because it uses a turbine motor (7), which is turned on solely when the charge of the accumulator batteries (3) drops and always works at best rpm. The device increases the distance covered by the vehicle without need for frequent recharge of the batteries from the power supply network. Furthermore, the turbine motor is lighter than the petrol piston engine and in the same time more efficient. The device includes installed accumulator batteries (3) that supply power to at least one electric motor coupled via transmission to the drive wheels of the vehicle, and fuel tank (4) supplying fuel to turbine motor (7) coupled via reducer (11) to generator (6) that is connected to the accumulator batteries. Control unit (9) monitors the charge of the accumulator batteries (3) and turns on and off the turbine motor (7).

Abstract: A stacked laminate component for a turbine engine that may be used as a replacement for one or more metal components is provided. The stacked laminate component can have a body formed by a process of stacking and laminating layers to define a radially inner surface along the hot gas path. The layers can be substantially orthogonal to the radially inner surface. The layers can be at least a first layer of a first material and a second layer of a second material. At least the first material is a ceramic matrix composite. The second material can have a higher thermal conductivity or a higher creep strength than the first material.

Abstract: The combustor of the present invention is provided with a combustor basket to which air A is supplied from outside, first nozzles which are installed in a plural number annularly along an inner circumference of the combustor basket to individually supply premixed gas M of the air (A) with a fuel (f) into the combustor basket, and a transition piece, a base end of which is connected to the combustor basket, and which burns the premixed gas (M) supplied from the first nozzles to form a flame front (F) which spreads to an outer circumference toward the leading end in an axial direction. Each of the first nozzles supplies the premixed gas (M) by changing the fuel concentration around the center axis of the first nozzle so that the flame front (F) is made uniform in temperature in the axial direction.

Abstract: A fan section for a gas turbine engine has a fan rotor with a plurality of fan blades. A plurality of exit guide vanes are positioned to be downstream of the fan rotor. The fan rotor is driven through a gear reduction relative to a turbine section. The exit guide vanes are desired to address resultant sound from interaction of wakes from the fan blades across exit guide vanes. A gas turbine engine incorporating a fan section is also disclosed.

Abstract: Gas turbine engines and related systems involving blade outer air seals are provided. In this regard, a representative blade outer air seal segment for a set of rotatable blades includes: a blade arrival end; and a blade departure end; each of the blade arrival end and the blade departure end being angularly offset with respect to a longitudinal axis about which the blades rotate.

Abstract: A turbomachine includes a housing that defines a flow path, and a stage arranged within the housing. The stage includes a plurality of rotating airfoil members and a first plurality of stationary airfoil members. A second plurality of stationary airfoil members is arranged directly adjacent to the first plurality of stationary airfoil members. A flow improvement system is associated with each of the first and second pluralities of stationary airfoil members. The flow improvement system establishes a predetermined clocking of each of the first plurality of stationary airfoil members relative to each of the second plurality of stationary airfoil members to improve flow characteristics along the flow path.

Abstract: Vane assemblies for gas turbine engines and methods for assembling vane assemblies are disclosed. The vane assemblies may include at least one shroud having at least one vane-receiving portion, at least one vane having at least one end portion received in the vane-receiving portion, and at least one sealing member having an uncompressed cross-section that is substantially circular. The sealing member(s) are disposed between and in contact with the end portion of the vane and the vane-receiving portion of the shroud.

Abstract: A compressor compresses air to produce compressed air. A mixture of fuel and the compressed air is combusted in a combustor to produce combustion gas. The combustion gas is supplied to a turbine to obtain rotational power. High-temperature gas accumulated in a space partitioned by an exhaust-side bearing portion that rotatably supports a turbine shaft and an exhaust diffuser is discharged through an exhaust gas passage. The high-temperature gas is sucked into the exhaust gas passage by exhaust gas flowing in the exhaust diffuser.

Abstract: A gas turbine system including a compressor section, a combustor section and a turbine section, wherein the compressor section includes a first wheel and a second wheel having a center bore extending axially therethrough, and wherein the first wheel and the second wheel are relatively adjacent each other. Also included is a gap disposed between the first wheel and the second wheel, wherein airflow is directed radially inward within the gap toward the center bore of the second wheel with or without the help of impellers extending radially. The compressor section further includes an airflow manipulation device disposed within the gap, comprising of at least one or multiple vanes which may start in radial direction but extend at least partially into the center bore.

Abstract: A gas turbine engine includes a seal assembly that is supported by a member at a joint. The seal assembly includes a seal support having a radial flange secured to the joint. A first bend adjoins the radial flange to a first leg, which is oriented generally in an axial direction. A second bend adjoins the first leg to a second leg, which is conical in shape. A seal is supported by the second leg.

Abstract: A gas turbine engine includes a very high speed low pressure turbine such that a quantity defined by the exit area of the low pressure turbine multiplied by the square of the low pressure turbine rotational speed compared to the same parameters for the high pressure turbine is at a ratio between about 0.5 and about 1.5. The high pressure turbine is mounted on the low pressure turbine with an intermediate bearing.

Abstract: An example method of controlling performance of gearbox of a gas turbine engine includes establishing a gear characteristic of a plurality of double helical gears each disposed about a respective axis in a gearbox. Performance of the plurality of double helical gears is controlled by selecting a circumferential offset distance between a first plurality of gear teeth spaced apart from a second plurality of gear teeth on each of the plurality of double helical gears in response to the established gear characteristic.

Abstract: A gas turbine engine includes high and low pressure turbines. A mid turbine frame is arranged axially between the high and low pressure turbines. A bearing is operatively supported by a support structure. An inner case is secured to the support structure and includes a first conical member and a bearing support to which the bearing is mounted. The bearing support includes a second conical member that is secured to the first conical member at a joint. The first and second conical members are arranged radially inward of the joint.

Abstract: A gas turbine engine includes a fan rotatable about an axis, a compressor section, a combustor in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor. The turbine section includes a fan drive turbine and a second turbine. The second turbine is disposed forward of the fan drive turbine. The fan drive turbine includes at least three rotors and at least one rotor having a bore radius (R) and a live rim radius (r), and a ratio of r/R is between about 2.00 and about 2.30. A speed change system is driven by the fan drive turbine for rotating the fan about the axis.