Abstract: A tail pipe (50) comprises: a pipe (51); an acoustic attenuator (61) that forms an acoustic space (Ss) on the outer peripheral side of the pipe (51); and a cooling air jacket (65) that forms a cooling air space (Sa) isolated from the outer space (So), which is the space on the outer peripheral side of the pipe (51). The pipe (51) has: a first air flow path (56) that is formed between the outer peripheral surface (55o) and the inner peripheral surface (55i); and an acoustic hole (59) that penetrates from the acoustic space (Ss) to a combustion space (Sc), which is a space on the inner peripheral side of the pipe (51c). The first air flow path (56) has: an inlet (56i) that faces into the cooling air space (Sa) and guides the air in the cooling air space (Sa) into the first air flow path (56); and an outlet (56o) that faces into the acoustic space (Ss) and guides the air passing through the first air flow path (56) into the acoustic space (Ss).

Abstract: A liner assembly for a combustor. The liner assembly includes a liner defining a combustion chamber of the combustor. The liner includes a looped section at a forward end of the liner. The looped section includes one or more bends such that the looped section forms a compliant joint and vibrations are dampened through the liner downstream of the looped section.

Abstract: A core for an electrical apparatus includes a plurality of electrical steel sheets having a ferromagnetic or ferrimagnetic coating applied to both sides of the electrical steel sheets. The electrical steel sheets are arranged in a stack to form a laminated stack. The ferromagnetic or ferrimagnetic coating is applied to both sides of the electrical steel sheets. The coating may comprise MnZn ferrites, NiZn ferrites, MgMnZn ferrites, CoNiZn ferrites, Co ferrites, Ni ferrites, Yttrium iron garnets (Y3Fe5O12) or other ferromagnetic or ferrimagnetic coating materials.

Abstract: A combustor liner for a gas turbine includes a liner at least partially defining a combustion chamber. The liner includes a plurality of dilution openings therethrough, at least one dilution opening of the plurality of dilution openings defined by (a) a main dilution opening portion defining a main dilution opening perimeter, and (b) a plurality of secondary dilution opening portions arranged about the main dilution opening perimeter and extending outward from the main dilution opening portion.

Abstract: A combustion chamber includes a bottom wall having at least one opening, at least one sleeve mounted upstream of the bottom wall and fixed to the bottom wall, a closing ring defining an annular groove with the sleeve and fixed to the sleeve, and at least one air and fuel injection system having an axis, mounted in the opening of the bottom wall. The injection system has an annular flange extending radially with respect to the axis and is mounted in the groove with radial clearance. A baffle is situated downstream of the bottom wall and is fixed to the sleeve and/or to the bottom wall.

Abstract: A coating fabrication method includes providing engineered granules and thermally consolidating the engineered granules on a substrate to form a silicate-resistant barrier coating. Each of the engineered granules is an aggregate of at least one refractory matrix region and at least one calcium aluminosilicate additive region (CAS additive region) attached with the at least one refractory matrix region. In the thermal consolidation, the refractory matrix region from the engineered granules form grains of a refractory matrix of the silicate-resistant barrier coating and the CAS additive region from the engineered granules form CAS additives that are dispersed in grain boundaries between the grains.

Abstract: A gas turbine combustor is equipped with a nozzle in which an air ejection passage extending along an axis and having an open distal end, and a fuel supply passage extending along the axis and having an open distal end are formed; swirling vanes provided around the nozzle so as to be twisted around the axis of the nozzle; an inner cylinder surrounding an outer periphery of the nozzle and the swirling vanes, and in which compressed air flows through an inside of the inner cylinder toward a downstream side; an outer cylinder which defines an inversion flow path, which inverts the compressed air on an outer periphery of the inner cylinder and introduces the compressed air to the inside of the inner cylinder, between the inner and outer cylinders; and an air introduction pipe having one end connected to a space on an upstream side of the compressed air from the inversion flow path, and the other end connected to the air ejection passage.

Abstract: A turbofan engine assembly including a compressor, an intermittent internal combustion engine having an inlet in fluid communication with an outlet of the compressor through at least one first passage of an intercooler, a turbine having an inlet in fluid communication with an outlet of the intermittent internal combustion engine, the turbine compounded with the intermittent internal combustion engine, a bypass duct surrounding the intermittent internal combustion engine, compressor and turbine, and a fan configured to propel air through the bypass duct and through an inlet of the compressor, wherein the intercooler is located in the bypass duct, the intercooler having at least one second passage in heat exchange relationship with the at least one first passage, the at least one second passage in fluid communication with the bypass duct.

Abstract: A system for directing cooling air into a gas turbine combustor is provided. The system comprises a transition duct coupled to a flow sleeve, where air to be used for combustor cooling and in the combustion process enters a bellmouth of the transition duct, passes through a plurality of struts within the bellmouth, and is distributed to a passage located between the combustion liner and flow sleeve.

Abstract: A fuel nozzle assembly includes a fuel plenum for delivering fuel to a combustion zone of a combustor. At least one narrow-band acoustic damper is provided in the fuel nozzle assembly. The narrow-band acoustic damper is tuned to dampen pressure oscillations within the combustor. The narrow-band acoustic damper may be a quarter-wave tube mounted to a cap plate within or adjacent to the fuel nozzle assembly.

Abstract: An endcover assembly for a combustor includes an endcover and a fuel nozzle rigidly connected to the endcover. The fuel nozzle includes a plurality of tubes that extends axially through a first plate, a fuel plenum and a second plate where each tube includes an inlet defined upstream of the first plate and an outlet disposed downstream from the second plate. The endcover assembly further includes an inlet flow conditioner having a forward end connected to the endcover and an aft end axially spaced from the forward end. The inlet flow conditioner circumferentially surrounds at least a portion of the fuel nozzle. A support plate is joined to the aft end of the inlet flow conditioner. Each tube of the fuel nozzle extends through a corresponding tube hole defined by the support plate. The fuel nozzle is connected to the support plate.

Abstract: A turbomachine includes a plurality of nozzles, and each nozzle has an airfoil. The turbomachine has opposing walls defining a pathway into which a fluid flow is receivable to flow through the pathway. A throat distribution is measured at a narrowest region in the pathway between adjacent nozzles, at which adjacent nozzles extend across the pathway between the opposing walls to aerodynamically interact with the fluid flow. The airfoil defines the throat distribution, and the throat distribution is defined by values set forth in Table 1, where the throat distribution values are within a +/?10% tolerance of the values set forth in Table 1. The throat distribution reduces aerodynamic loss and improves aerodynamic loading on each airfoil.

Abstract: An exemplary gas turbine engine component assembly includes, among other things, a case having a generally cylindrical outer surface. At least one synchronizing ring is received at least partially about the outer surface of the case. A plurality shims situated between an inner surface on the at least one synchronizing ring and the outer surface of the case at a plurality of shim locations are circumferentially spaced about the case outer surface. Each shim location includes at least one of the shims. A radial thickness of the shims at a first one of the shim locations is different than the radial thickness of the shims at a second one of the shim locations.

Abstract: A thermally dissipative article and a method of forming a thermally dissipative article are disclosed. The thermally dissipative article includes a component, a porous material formed in a layer on the component. The method of forming a thermally dissipative article includes providing a metal powder mixture and a soluble particulate mixture which forms a porous coating upon sintering and immersion in a solvent to remove the soluble particulate.

Abstract: Low dielectric constant (low-k) polyhemiaminal (PHA) and polyhexahydrotriazine (PHT) materials with cyclic aliphatic ring structures are described. The materials are formed by a method that includes heating a mixture comprising amines and paraformaldehyde. The reaction mixtures may be used to form low-k PHT prepregs, composites and dielectrics used in integrated circuits.

Abstract: An electro-optical device includes: (1) a first electrode layer; (2) a second electrode layer; and (3) a middle layer disposed between the first electrode layer and the second electrode layer. The middle layer includes a material having the formula: [AaBbXxX?x?X?x?], where A is selected from potassium, rubidium, and cesium; B is selected from germanium, tin, and lead; X, X?, and X? are independently selected from fluorine, chlorine, bromine, and iodine; a is in the range of 1 to 9; b is in the range of 1 to 5; a sum of x, x?, and x? is in the range of 1 to 9; and the material is at least one of n-doped and p-doped.

Abstract: An exoatmospheric vehicle uses a control system that includes a thrust system to provide thrust to control flight of the vehicle. A regenerative heat system is used to preheat portions of the thrust system, prior to their use in control of the vehicle. The heat for preheating may be generated by consumption of a fuel of the vehicle, such as a monopropellant fuel. The fuel may be used to power a pump (among other possibilities), to pressurize the fuel for use by thrusters of the thrust system. The preheated portions of the thrust system may include one or more catalytic beds of the thrust system, which may be preheated using exhaust gasses from the pump. The preheating may reduce the response time of the thrusters that have their catalytic beds preheated. Other thrusters of the thrust system may not be preheated at all before operation.

Abstract: A combustor generally includes a plate that extends radially and circumferentially within at least a portion of the combustor. The combustor may also include a shroud that at least partially surrounds the plate and a plurality of tubes that extend through the plate. One or more flexible couplings may at least partially surround at least some of the plurality of tubes and the one or more flexible couplings may be connected to the plate.

Abstract: A ducting arrangement (10), including: a plurality of discrete ducts (18), each defining a flow path and configured to receive a flow of combustion gases from a respective combustor can, where the plurality of discrete ducts merge to form a common duct structure; and a throat insert (50) configured to define at least part of a junction of one of the discrete ducts and the common duct structure.

Abstract: There is provided a gas turbine combustor capable of improving cooling performance of a combustion chamber thereof and reducing the amount of NOx emissions.

Abstract: An interface assembly for a combustor includes an interface housing having a channel defined by a forward wall and at least one aft wall segment, the aft wall segment operatively coupled to an aft flange of a flow sleeve. Also included is a piston ring fittingly disposed in the channel.

Abstract: A gas turbine engine having a fire wall that is configured to provide a fire resistant barrier between a first zone and a second zone in the gas turbine engine, the second zone being hotter than the first zone when the gas turbine engine is in use. The gas turbine also has an actuator that is located in the first zone and is configured to generate a mechanical force when operated, an actuatable device that is located in the second zone and is configured to be actuated by a mechanical force and a mechanical force transmitting device that extends from the actuator to the actuatable device via a hole in the fire wall. The mechanical force transmitting device is configured to, when the actuator is operated, actuate the actuatable device by transmitting a mechanical force generated by the actuator to the actuatable device.

Abstract: A fan blade for a gas turbine engine includes an airfoil that includes leading and trailing edges joined by pressure and suction sides to provide an exterior airfoil surface that extends 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 the 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 method for determining the temperature of an aggressive and/or abrasive gas is described using an acoustic transmitter and an acoustic receiver. The transmitter emits an acoustic signal with varying frequencies and the receiver extracts from the acoustic input signal a frequency of the maximum. Based on this the frequency of this maximum the temperature of the gas between transmitter and receiver is calculated.

Abstract: Acoustic damping resonators (60, 62, 76, 78) formed in pockets (32) between reinforcing ribs (30) in a grid of ribs on the outer surface of a wall (74) surrounding a combustion gas flow path (18). Each resonator has a perforated cover plate (64A-C) spanning between sides formed by the ribs (30). Film cooling exit holes (44) are provided in the wall (74) under each resonator. Resonating chambers (48A-C) of different volumes may be provided on the wall to damp different unwanted frequencies. Different sets of resonators with different volumes may be separately controlled by respective airflow control manifolds (66) via throttle valves (68, T1-T3). Control logic (70) may control the valves based on frequency/airflow response functions (80, 82) for each size of resonator to optimize damping and cooling and to lower emissions over varying engine operating conditions.

Abstract: A power generation plant including a plurality of combustor cans, one or more sensors configured to measure acoustic vibrations from the plurality of combustor cans, and a controller that includes a processor. The processor is programmed to identify a dominant tone of the measured acoustic vibrations, the dominant tone being a frequency where a highest amount of energy lies, determine, for each of the plurality of combustor cans, a frequency of the dominant tone at predefined time intervals, and determine a degradation status of at least one of the plurality of combustor cans based on the frequency of the dominant tone at each of the predefine time intervals for the at least one of the plurality of combustor cans.

Abstract: A can-annular gas turbine engine combustion arrangement (10), including: a combustor can (12) comprising a combustor inlet (38) and a combustor outlet circumferentially and axially offset from the combustor inlet; an outer casing (24) defining a plenum (22) in which the combustor can is disposed; and baffles (70) configured to divide the plenum into radial sectors (72) and configured to inhibit circumferential motion of compressed air (16) within the plenum.

Abstract: A turbomachine combustion chamber shell ring in which dilution holes in the turbomachine combustion chamber shell ring are covered with inserts defining chambers around same on an inner face of the shell ring. Ventilation holes, through the insert, induce ventilation of portions of the shell ring surrounding the dilution holes, cool the portions, and prevent crack formation.

Abstract: An inner barrel member for a gas turbomachine includes a body having an outer surface, an inner surface and one or more radial flow splitters provided on the outer surface. The one or more radial flow splitters are configured and disposed to be arranged along a combustor centerline at a combustion flow outlet radially inwardly of a transition piece.

Abstract: A gas turbine includes a combustion chamber and a microwave source to produce microwave radiation. The gas turbine is arranged to guide the microwave radiation into a cavity of the combustion chamber. Due to the microwave radiation, in the cavity of the combustion chamber, combustion in the cavity may be supported and thus lean operation of the gas turbine is made possible.

Abstract: The invention relates to a burner arrangement for using in a single combustion chamber or in a can-combustor comprising a center body burner located upstream of a combustion zone, an annular duct with a cross section area, intermediate lobes which are arranged in circumferential direction and in longitudinal direction of the center body. The lobes being actively connected to the cross section area of the annular duct, wherein a cooling air is guided through a number of pipes within the lobes to the center body and cools beforehand at least the front section of the center body based on impingement cooling. Subsequently, the impingement cooling air cools the middle and back face of the center body based on convective and/or effusion cooling. At least the back face of the center body includes on the inside at least one damper.

Abstract: A turbine engine, a turbojet, or a turboprop engine, including an annular combustion chamber defined by an inner shell and an outer shell, a turbine distributor arranged downstream of the combustion chamber, with a downstream end of the outer shell and/or of the inner shell of the chamber including a radial rim arranged opposite a radial rim of an upstream end of the distributor, a sealing mechanism including at least one strip extending between the rims to provide a seal between the combustion chamber and the distributor. The sealing strip extends axially and circumferentially between the rims and bears sealingly against free ends of the rims.

Abstract: A helical seal system includes a first component, and a second component rotatable relative to the first component. The second component extends from a high pressure portion to a low pressure portion through an intermediate portion. A helical seal is provided on the intermediate portion of the second component. The helical seal includes at least one thread component having a pitch that is configured and disposed to draw fluids from the low pressure portion toward the high pressure portion when the second component is rotated.

Abstract: A gas turbine combustor floating collar for mounting an igniter or fuel nozzle to a combustor is provided with an outer periphery in a non-circular shape, for example having at least one section thereof formed with a flat surface, such as a square or triangular shape. The outer periphery of the floating collar is complementary to and completely surrounded by an inner periphery surface of a recess of a boss affixed on the combustor, thereby preventing substantial rotation of the floating collar with respect to the boss.

Abstract: A turbine assembly is disclosed herein. The turbine assembly may include a rotor with at least one bucket extending radially therefrom. The turbine assembly also may include a stator assembly positioned adjacent to the at least one bucket. Moreover, the turbine assembly may include a seal assembly. The seal assembly may include a first flange extending in a substantially axial direction from the at least one bucket. The seal assembly also may include a second flange extending from the stator assembly in the substantially axial direction opposite the first flange. The second flange may include at least one protuberance projecting towards the first flange.

Abstract: A heat shield for a combustion chamber of a gas turbine engine is disclosed. The heat shield includes a plate portion, an inner ring, and a plurality of standoffs. The plate portion includes an annular sector shape. The inner ring extends from an inner part of the plate portion. The inner ring includes a hollow cylinder shape. The plurality of standoffs extends from the plate portion, proximate outer edges of the plate portion and in the same direction as the inner ring. The plurality of standoffs forms a plurality of scallops. Each scallop is located between adjacent standoffs.

Abstract: The invention concerns a method for manufacturing of a gas turbine engine component (37) comprising an outer ring structure (42, 47), an inner ring structure (41), and a plurality of circumferentially spaced elements (46, 46a, 46b) extending between the inner ring structure (41) and the outer ring structure (42), wherein a primary gas channel for axial gas flow is defined between the elements (46, 46a, 46b), and wherein the component (37) has an inlet side for gas entrance and an outlet side for gas outflow. The invention is characterized in that the method comprises the step of machining a one-piece metal blank as to form a one-piece part (47) comprising: a portion (46b) of each of said elements (46), wherein said portion (46b) relates to a portion of an extension length of the elements (46) between said ring structures (41, 42); and a ring-shaped member (42) that connects said element portions (46b) and that is intended to form part of one of the ring structures.

Abstract: A variable area fan nozzle comprising an array of rigid petals and a petal actuation system comprising left and right assemblies, each assembly comprising: a multiplicity of tracks attached to or integrally formed with respective petals; a curved pivoting ring segment; an actuator coupled to the ring segment; and a multiplicity of sets of cam followers spaced along the ring segment and aligned with respective tracks. Each ring segment is pivotable between first and second angular positions depending on how the state of the actuator changes. As one ring segment pivots in one direction, one set of cam followers exert inward forces on the tracks to deflect petals inward; as that ring segment pivots in the other direction, another set of cam followers exert outward forces on the tracks to deflect petals outward.

Abstract: The invention relates to a method for determining at least one firing temperature for controlling a gas turbine that comprises at least one compressor, at least one combustion chamber and at least one turbine, compressed air being drawn off at the compressor in order to cool the turbine and being fed to the turbine via at least one external cooling duct and via a control valve arranged in the cooling duct, in which method a plurality of temperatures and pressures of the working medium being measured in various positions of the gas turbine and the at least one firing temperature being derived from the measured temperatures and pressures. A more flexible and more accurate control is achieved additionally by determining the cooling air mass flow via the external cooling duct and by taking said flow into account when deriving the at least one firing temperature.

Abstract: A gas turbine engine includes a compressor, a combustor adjacent the compressor, a turbine adjacent the combustor, a shaft, a motor, a variable frequency drive, a stored energy source and a ride thru unit. The motor is coupled to the shaft. The variable frequency drive is electrically connected to the motor and to an AC power source. The ride thru unit electrically connects to the variable frequency drive, the AC power source and the stored energy source. The ride thru unit includes at least one DC to DC voltage converter.

Abstract: A support frame for assembling a combustion module for a gas turbine includes a base plate disposed at a bottom end of the support frame and a support plate that is vertically separated from the base plate by one or more vertical support members. The support plate defines an opening that is sized to allow a portion of the combustion module to pass therethrough. A support block extends vertically from the base plate towards the support plate where the support block defines one or more fastener holes for connecting an aft end of a combustion liner of the combustion module to the support block. A central support column extends vertically from the base plate towards the support plate. A horizontal support extends radially outward from the central support column to align the combustion liner with the opening in the support plate.

Abstract: A process of producing heat energy for use in heat exchange with other fluids and substances so as to impart said heat energy to said fluid or substances which includes several steps. In a first step, a mixture of fuel and oxidant is ignited in a combustion zone or zones (16, 18, 98, 122) to create a combusted fuel mix which in the form of shockwaves are carried away from the combustion zone or zones (16, 18, 98, 122) to provide a low pressure area within the combustion zone or zones (16, 18, 98, 122).

Abstract: A toroidal combustion chamber and a plurality of plates are disposed with respect to one another such that their labyrinthine walls are juxtaposed and form labyrinthine conduits leading from an exhaust outlet of the toroidal combustion chamber to an exit perpendicular to the exhaust outlet.

Abstract: A method of assembling a combustion system for a gas turbine engine includes providing a combustion chamber frame, an inner casing structure, and an outer casing structure. The method also includes mounting the combustion chamber frame between the inner casing structure and the outer casing structure such that the combustion chamber frame is coupled to the inner casing structure and the outer casing structure.

Abstract: A running-in coating for a compressor housing is disclosed. The running-in coating has at least two layers where a first layer is dimensionally stable and at least one additional layer has run-in capability.

Abstract: A combustor section for a gas turbine engine includes a combustor vane which extends at least partially into a combustion chamber.

Abstract: The present application provides a combustion nozzle for use with a gas turbine engine. The combustion nozzle may include a number of mixing tubes, an outer shell surrounding the mixing tubes, and a floating aft plate assembly. The floating plate assembly may enclose the outer shell. The mixing tubes may extend through the aft plate assembly.

Abstract: A system for pre-heating a working fluid within a combustor includes a compressor for providing the working fluid to the combustor. An outer casing is disposed downstream from the compressor. The outer casing at least partially defines a high pressure plenum that at least partially surrounds the combustor. A combustion chamber is defined within the combustor downstream from the high pressure plenum. A catalytic combustor is disposed within the high pressure plenum upstream from the combustion chamber so as to provide thermal energy to the working fluid upstream from the combustion chamber.

Abstract: A combustion chamber includes an inner wall, an outer wall surrounding the inner wall, and a flow passage between the inner wall and outer wall that permits forward flow and restricts reverse flow. The inner wall has a plurality of effusion holes in fluid communication with the inside of the combustion chamber. The outer wall has a plurality of cooling side holes in fluid communication with a cooling source. The inner wall and the outer wall define a flow passage therebetween that fluidly connects one or more of the cooling side holes with one or more of the effusion holes. The flow passage has a geometric configuration to permit forward flow of gases through the flow passage from the cooling source to the inside of the combustion chamber and to restrict reverse flow of gases through the flow passage from the inside of the combustion chamber to the cooling source. The permitted flow rate is greater than the restricted flow rate.

Abstract: A combustion liner for a gas turbine combustor includes an annular main body having a forward end axially separated from an aft end, and a transitional intersection defined between the forward end and the aft end. The main body extends continuously from the forward end to the aft end. A plurality of fuel injector passages extend radially through the main body upstream from the transitional intersection. The main body comprises a conical section having a circular cross section that diverges between the forward end and the transitional intersection, and a transition section having a non-circular cross section that extends from the transitional intersection to the aft end of the main body.