Abstract: A combustor for a gas turbine engine comprising a combustion chamber wall having formed therein at least one hole for admitting air into the combustion chamber and at least one air intake chute aligned with said hole. The air intake chutes are attached to the combustor wall in regions which, during operation, will be regions of low stress concentration.

Abstract: An outer combustor wall (36) includes an outer support shell (40) surrounding an outer combustion liner (42). The shell (40) includes a first radial aft flange (46), and the liner (42) includes a second radial aft flange (48) axially adjoining the first flange. An annular retainer (50) and seal (62) are joined to the first flange for slidingly trapping the second flange therebetween for thermally floating the liner on the shell.

Abstract: An auxiliary power unit (APU) includes a rotary fuel slinger, and a single-spool, two-stage turbine. The rotary fuel slinger receives a rotational drive force from the turbine, and a supply of fuel from a fuel source. The rotary fuel slinger injects the fuel into a combustor as a continuous “fuel sheet.” As a result, fuel mixing and atomization inside the combustor is improved, which reduces emissions and pattern factor, which in turn increases turbine life.

Abstract: A jet engine using exhaust gas to get the propulsive force by rotating a fan within the exhaust gas burned in and emitted from the engine includes a body, a burner installed to the body to inject and burn the fuel in compressed air, a high-pressure turbine for giving rotational force by high-pressure exhaust gas from the burner, a low-pressure turbine for giving rotational force by low-pressure exhaust from the high-pressure turbine, a rotary shaft combined to a rotational center of the high-pressure and low-pressure turbines, and a propulsive force providing unit which is rotating together with the rotary shaft to change component of lateral velocity of the exhaust gas emitted through the turbines from the burner to be directed rearward.

Abstract: A conical helical design for a turbine combustor scroll utilizes as much cavity of the combustor housing as possible by adding an axial shift and an irregular cross sectional shape in the scroll without adversely effecting aerodynamic performance. The axial shift region of the combustor scroll extends the cross-sectional area centroid of the scroll beyond the scroll's discharge area B-width. The resulting scroll design allows for the use of a high performance engine with a larger combustor while reducing the weight of the system by making the combustor housing as small as possible. Furthermore, the scroll design increases the air velocity for convection cooling by reducing the gap between the scroll and the housing. The turbine scroll of the present invention is useful in engines for which high performance is required, such as certain high performance aircraft.

Abstract: An integrated gas turbine compressor and rotary fuel injector is provided which includes a rotary cup combustor having an atomizing element for dispersal and distribution of fuel throughout the combustor. The combustor is positioned in close proximity to the compressor impeller of a gas turbine engine having a centrifugal compressor. The rear face of the impeller, or the rear face of an attachment to the impeller, is contoured to receive the fuel supply. The contour terminates at a lip disposed at a larger radius from the centerline shaft than conventional shaft-mounted slingers. The larger radius of fuel release provides better fuel atomization due to larger rim speeds, thereby improving combustor performance.

Abstract: An annular combustor that has angled effusion holes through at least one surface of the combustor liner with the angle of the effusion holes oriented to cause the flow of air through the holes to facilitate swirling of the fuel and air within the combustor. The effusion holes thereby facilitate efficient cooling of the combustor liner combined with superior fuel/air mixing within the combustor.

Abstract: A combustor baffle includes an annular splashplate having a center mounting tube therein for receiving a carbureted stream. A semi-ferrule is offset both laterally and transversely from the tube, and is open laterally outwardly therefrom. The semi-ferrule cooperates with an adjoining semi-ferrule for defining a port in which a pilot injector or igniter may be mounted.

Abstract: A combustor system for a miniature gas turbine engine includes fuel injection via orifices which direct fuel into fuel-air injection tubes which feed a fuel-rich mixture of fuel and air into a leading end of the combustor liner to form a primary burning region. Fuel system pressures are kept low and controlled by control of the fuel injection port size and number. Fuel breakup is via airblast and tube wall impingement. The fuel-air injection tubes are directed circumferentially, radially outward and toward the front end of the combustor. Air is fed into the combustor such that two distinct burning regions are created. Each region approximates a “well-stirred reactor” and the combination of the two regions results in an efficient use of combustion volume.

Abstract: A method for determining a target exhaust temperature for a gas turbine including: determining a target exhaust temperature based on a compressor pressure condition; determining a temperature adjustment to the target exhaust temperature based on at least one parameter of a group of parameters consisting of specific humidity, compressor inlet pressure loss and turbine exhaust back pressure; and adjusting the target exhaust temperature by applying the temperature adjustment.

Abstract: An annular combustor has an annular inner shell, an annular outer shell and a dome end wall connecting the inner shell with the outer shell. A dam is positioned between the dome end wall and the exit end and extends radially from at least one of the inner shell and the outer shell to provide a reduced flow area passageway within the combustion chamber.

Abstract: An outer combustor wall (36) includes an outer support shell (40) surrounding an outer combustion liner (42). The shell (40) includes a first radial aft flange (46), and the liner (42) includes a second radial aft flange (48) axially adjoining the first flange. An annular retainer (50) and seal (62) are joined to the first flange for slidingly trapping the second flange therebetween for thermally floating the liner on the shell.

Abstract: An annular combustion chamber for a gas turbine is disclosed, which permits a sealed air cooling, such that high powers and efficiencies can be achieved with low NOx emissions. The two-piece annular combustion chamber housing surrounds a thin-walled sealed liner, which guides the hot gas escaping from the burners to the turbine inlet in a gas-tight manner. A variable annular chamber is located between the liner and the annular combustion chamber housing, in which the cooling air is collected as a counter-current to the hot gas and fed to the burners (closed air cooling). The surfaces of the liner are variably cooled with efficient baffle cooling devices and drilled patterns in. The thermally-elastic liner constructed from submerged arc-welded rings is axially fixed at the flange only and for the remainder provided with reinforcing circumferential beading. The freely expanding liner is otherwise concentrically fixed to the cold annular combustion chamber housing by means of elastic tensioning elements.

Abstract: A combustor system for a miniature gas turbine engine includes fuel injection via orifices which direct fuel into fuel-air injection tubes which feed a fuel-rich mixture of fuel and air into a leading end of the combustor liner to form a primary burning region. Fuel system pressures are kept low and controlled by control of the fuel injection port size and number. Fuel breakup is via airblast and tube wall impingement. The fuel-air injection tubes are directed circumferentially, radially outward and toward the front end of the combustor. Air is fed into the combustor such that two distinct burning regions are created. Each region approximates a “well-stirred reactor” and the combination of the two regions results in an efficient use of combustion volume.

Abstract: A combustion chamber (4) of a gas turbine (1), the combustion space (24) of which is bounded by an annular combustion chamber inner wall (28) and a combustion chamber outer wall,(26), in which in order to generate a working medium (M) a supplied fuel is brought into reaction with supplied combustion air, and the combustion chamber wall (25) of which is provided on the inside with a lining formed from a plurality of heat shield elements (38), with the or each heat shield element (38) together with the combustion chamber wall (25) forming an inner space (40) to which a cooling medium (K) can be applied, is to be designed so as to provide a high level of system efficiency at the same time as having a comparatively simple structure and it should also be possible to disassemble the combustion chamber inner wall (28) in a time-saving manner.

Abstract: A jet engine using exhaust gas to get the propulsive force by rotating a fan within the exhaust gas burned in and emitted from the engine includes a body, a burner installed to the body to inject and burn the fuel in compressed air, a high-pressure turbine for giving rotational force by high-pressure exhaust gas from the burner, a low-pressure turbine for giving rotational force by low-pressure exhaust from the high-pressure turbine, a rotary shaft combined to a rotational center of the high-pressure and low-pressure turbines, and a propulsive force providing unit which is rotating together with the rotary shaft to change component of lateral velocity of the exhaust gas emitted through the turbines from the burner to be directed rearward.

Abstract: An attachment and sealing system for securing a combustor cap assembly to a combustion chamber, while providing effective cooling to the combustor cap assembly aft end, is disclosed. The combustor cap assembly is secured within a combustion chamber by a plurality of pins such that the aft end region of the combustor cap assembly is cooled by a fluid medium that is injected through a plurality of first holes in the combustion chamber wall. A generally annular seal, which encompasses the combustor cap assembly, has a plurality of raised ridges, which are in sealing contact with the combustion chamber wall. Multiple embodiments are disclosed regarding the length of the generally annular seal and position and orientation of the plurality of first cooling holes in the combustion chamber wall.

Abstract: The invention relates to an annular combustor (13) for a gas turbine (10), into which combustor (13) burners (14, 15) open on an inlet side, and which combustor (13) extends in the axial direction from the inlet side to an outlet side (33) and is lined on the insides with cooled liner segments (16, 17) for protection from the hot gases.

Abstract: A method enables a gas turbine engine multi-domed combustor including an outer liner and an inner liner that define a combustion chamber therebetween to be assembled. The method comprises coupling a first dome including a heat shield that includes an annular endbody that extends a first distance axially from the heat shield to the combustor outer liner, and coupling a second dome including a heat shield that includes an annular endbody that extends a second distance axially from the heat shield to the first dome, such that the second dome is radially aligned with respect to the first dome, and wherein the second dome second distance is less than the first dome first distance.

Abstract: A flame-passage device (10) for non-annular gas turbine combustion chambers comprises a tubular body (11), which is provided with a plurality of cooling holes (12), for cooling on the swirl-cooling type. The tubular body (11) is inserted in a flanged pipe (27), which connects the cases (31, 32) of two successive combustion chambers (17, 18), and has a first end (13) with a cylindrical shape, and a second end (14) with an oval shape, wherein the second, oval end (14) is provided with three rings (23, 24, 25), for anchorage of the corresponding combustion chamber (17) to the case (31).

Abstract: A gas turbine combustion system and method used for generating electrical power includes a compressor that receives and compresses air. A first stage turbine nozzle is flowise connected to the compressor and receives a portion of the compressed air from the compressor within a first air flow. A torus configured combustion chamber is positioned around the first stage turbine nozzle and receives a portion of the compressed air from the compressor within a second air flow that is passed through the combustion chamber where air and fuel are mixed and combusted. The air is discharged at the first stage turbine nozzle to mix with the first air while achieving a dry low NOx combustion.

Abstract: Air tubes are circumferentially spaced around the outer liner of an annular combustor for a microturbine engine in proximity to and downstream of a dam extending into the combustion chamber for reducing the emission of NOX. The air tubes are dimensioned so that the length to passage diameter is such that a swirling motion of the air injected into the combustion zone is normal to the center line of the annular combustor.

Abstract: A combustor for a gas turbine engine having a longitudinal axis extending therethrough, including an inner liner, an outer liner spaced radially from the inner liner, an annular dome connected to the inner and outer liners, a plurality of air/fuel mixers connected to the dome and circumferentially spaced within the dome, an outer cowl connected to the outer liner, and an inner cowl connected to the inner liner, wherein a combustion chamber is formed by the inner liner, the outer liner and the dome. The dome further includes a first end connected to a liner of the combustor, a second end spaced radially from the first end, a first portion extending rearwardly from the first end adjacent the liner, a second portion extending rearwardly from the second end and flaring radially outward from the longitudinal axis, and a nonlinear third portion connecting the first and second dome portions.

Abstract: The invention relates to an annular combustion chamber (4) for a gas turbine (1) wherein the annular combustion chamber (4) extends in an axial direction (A), encloses a combustor (7), and has on its inside facing the combustor (7) a bearing structure (26) on which a lining element (10) secured to this lines the annular combustion chamber (4). The object is to disclose an annular combustion chamber (4) with a lining element (10) that meets the mechanical requirements while at the same time taking account of the system's maintenance-friendliness.

Abstract: A fabricated cowl for a double annular combustor of a gas turbine engine having a longitudinal axis extending therethrough, including an outer annular portion, an inner annular portion, and a middle annular portion. A plurality of circumferentially spaced radial slots are formed in at least one of the annular portions. The outer annular portion includes a plurality of circumferentially spaced windows of desired size formed therein, wherein a plurality of outer radial members connecting the outer annular portion and the middle annular portion are provided between adjacent windows therein. Similarly, the inner annular portion includes a plurality of circumferentially spaced windows of desired size formed therein, wherein a plurality of inner radial members connecting the inner annular portion and the middle annular portion are provided between adjacent windows therein.

Abstract: An integrated gas turbine compressor and rotary fuel injector is provided which includes a rotary cup combustor having an atomizing element for dispersal and distribution of fuel throughout the combustor. The combustor is positioned in close proximity to the compressor impeller of a gas turbine engine having a centrifugal compressor. The rear face of the impeller, or the rear face of an attachment to the impeller, is contoured to receive the fuel supply. The contour terminates at a lip disposed at a larger radius from the centerline shaft than conventional shaft-mounted slingers. The larger radius of fuel release provides better fuel atomization due to larger rim speeds, thereby improving combustor performance.

Abstract: A process is described for the operation of an annular combustion chamber and also an annular combustion chamber is shown with numerous circularly arranged premix burners, in which a fuel-air mixture is produced before it is ignited and the fuel-air mixture is used as a hot gas stream to drive at least one turbine stage of a gas turbine plant. At least one premix burner is operated such that it has a spatial mixing profile deviating within the fuel-air mixture from all other premix burners.

Abstract: A gas turbine includes an annular combustion chamber and an outer wall of an annular combustion chamber. A straining ring is arranged on the outer wall of the annular combustion chamber and enables oscillations of the outer wall to be damped via friction. The effects of combustion oscillations produced by damaging vibrations of the annular combustion chamber are thus reduced. A method is further for damping an oscillation of an outer wall of an annular combustion chamber.

Abstract: A combustion chamber arrangement for gas turbines includes a multiplicity of individual combustion chambers which open out in a common annular combustion chamber and are arranged in the shape of a circle. The arrangement is such that it can be cooled highly efficiently using convective cooling, while also as far as possible dispensing with the use of impingement cooling. In such a combustion chamber arrangement, the annular combustion chamber, in the transition region to the individual combustion chambers, is of a height which fluctuates periodically and which is minimal in the region between adjacent individual combustion chambers.

Abstract: An offset annular combustion chamber for an airplane engine gas turbine, the chamber comprising a pilot head having a plurality of nozzle systems distributed on a pilot head chamber end wall interconnecting an inner longitudinally-extending side wall of the chamber to a pilot head outer longitudinally-extending side wall, and a take-off head that is radially and axially offset from the pilot head and comprising a plurality of nozzle systems distributed on a take-off head chamber end wall interconnecting the pilot head outer longitudinally-extending side wall and a take-off head outer longitudinally-extending side wall, the pilot head having at least N substantially identical nozzle systems with overall permeability PA adapted to lighting and to speeds close to idling, and the take-off head having at least 2N substantially identical nozzle systems of overall permeability PB, where PB is greater than or equal to PA, the permeability PA lying in the range 10% to 40% of the total air flow rate penetrating into th

Abstract: The invention relates to a combustion-chamber arrangement for gas turbines, having an annular combustion chamber connected to at least one burner and leading into a turbine space, an annular-combustion-chamber wall which defines the annular combustion chamber being of double-shell construction, and an inner shell of the annular-combustion-chamber wall being composed of lining elements releasably arranged on an outer shell of the annular-combustion-chamber wall, and a gap space through which a cooling medium can flow being formed between the inner shell and the outer shell of the annular-combustion-chamber wall.

Abstract: An annular combustor suitable for use in a gas turbine engine having a plurality of fuel injectors and a plurality of apertures circumferentially arranged within the inner and outer liners of the combustion chamber is disclosed. Further, each of the apertures of the inner liner are disposed on one side of the centerline of each of the fuel injectors and each of the apertures of the outer liner are disposed on the opposing side of the centerline of each of the fuel injectors. In this way, air exhausted from the apertures advantageously opposes the direction of swirl of the fuel and air mixture that is injected from the fuel injectors thereby providing enhanced mixing of the fuel and air within the combustion chamber, increased residence time and the reduction of hot streaks in the inner and outer liners.

Abstract: Air tubes are circumferentially spaced around the outer liner of an annular combustor for a microturbine engine in proximity to and downstream of a dam extending into the combustion chamber for reducing the emission of NOX. The air tubes are dimensioned so that the length to passage diameter is such that a swirling motion of the air injected into the combustion zone is normal to the center line of the annular combustor.

Abstract: A recuperated gas turbine engine is disclosed, comprising: a heat exchanger, and gas turbine (including compressor, can-type combustor and turbine); wherein, said heat exchanger comprises a compressed air passageway and a turbine exhaust gas passageway adjacent to each other within the casing which extend spirally throughout said heat exchanger and towards an inner cylindrical chamber in which said combustor is positioned approximately to the center of said casing. Improved engine fuel efficiency is achieved by preheating the compressed air before it reaches the combustor with the higher-temperature exhaust gas. A can-type combustor is used for alleviating heat-dissipation issues to improve efficiency of the combustion. A concentric back-to-back rotor arrangement significantly shortens the length of a conventional engine turbine rotor which improves on the operational stability of a gas turbine engine.

Abstract: A seal structure for sealing a gap between a combustor liner and a neighboring structure adjacent to the combustor liner, includes an annular sealing member mounted on the neighboring structure so as to be in contact with an annular outer surface of the combustor liner to seal the gap between the combustor liner and the neighboring structure. The annular sealing member includes a plurality of sealing segments which are arranged in an annular form as a whole.

Abstract: The present invention relates to a turbine engine, wherein compressor rotor blades and turbine rotor blades set around the interior circular surface of a hollow revolving shaft while compressor stator blades and turbine stator blades are set around the exterior circular surface of an inner engine case inside the hollow revolving shaft. By using the present invention, the configuration of the turbine engine can be compact, and the volume can be substantially reduced. Besides, the stiffness of the rotor blades can be increased, and efficiency of the compressor, the burner, the turbine, and the whole engine can be increased.

Abstract: An annular combustor (10) has an annular inner shell (15), an annular outer shell (20) and a dome end wall (30) connecting the inner shell (15) with the outer shell (20). A dam (45) is positioned between the dome end wall (30) and the exit end and extends radially from at least one of the inner shell (15) and the outer shell (20) to provide a reduced flow area passageway within the combustion chamber (35).

Abstract: An engine core structure for a gas turbine engine includes an outer annular shaft body, a turbine rotor body, turbine rotor blades radially connected between the outer annular shaft body and the turbine rotor body to hold the turbine rotor body in the rear section of the outer annular shaft body, the outer annular shaft body including slots through which compressed air passes in a radially inward direction, and a can type combustor mounted in a receiving clamber inside the outer annular shaft body to enlarge the diameter of the core shaft, to avoid vibration due to resonance, to save space, to eliminate dissipation of heat, and to improve the thermal efficiency of the gas turbine engine.

Abstract: In a turbine engine with an annular recuperator surrounding the turbine, exhaust gas is directed from the turbine to the recuperator by a generally curved exhaust dome. A vortex disrupter structure extends from the exhaust dome to a point distal of the turbine to evenly distribute the exhaust gas entering the recuperator and sustain diffusion of the exhaust gas to increase the expansion ratio across the turbine.

Abstract: An annular one-piece liner for a combustor of a gas turbine engine, including a first end adjacent to an upstream end of the combustor, a second end adjacent to a downstream end of the combustor, and a plurality of corrugations between the first and second ends, each corrugation having an amplitude and a wavelength between an adjacent corrugation, wherein the amplitude of the corrugations and/or the wavelength between adjacent corrugations is variable from the first end to the second end.

Abstract: The invention relates to a heat-shield brick (1), in particular for lining a combustion chamber wall, comprising a hot side (3) that can be exposed to a hot medium, a wall side (5) that lies opposite said hot side (3) and a peripheral side (7) that lies adjacent to the hot side (3) and the wall side (5) and that has a peripheral lateral face (9). A tensioning element (11), pre-stressed in the peripheral direction is provided on the peripheral side (7), whereby a compressive stress is generated perpendicularly to the peripheral lateral face. The invention also relates to a combustion chamber comprising a combustion chamber lining, which has heat-shield bricks of this type and to a gas turbine comprising a combustion chamber.

Abstract: An annular combustor suitable for use in a gas turbine engine having a plurality of fuel injectors and a plurality of apertures circumferentially arranged within the inner and outer liners of the combustion chamber is disclosed. Further, each of the apertures of the inner liner are disposed on one side of the centerline of each of the fuel injectors and each of the apertures of the outer liner are disposed on the opposing side of the centerline of each of the fuel injectors. In this way, air exhausted from the apertures advantageously opposes the direction of swirl of the fuel and air mixture that is injected from the fuel injectors thereby providing enhanced mixing of the fuel and air within the combustion chamber, increased residence time and the reduction of hot streaks in the inner and outer liners.

Abstract: A combustor for a gas turbine engine includes inner and outer liners with a row 112, 114 of dilution air holes penetrating through each liner. The row 114 of holes in the outer liner comprise at least a set of large size, major outer holes 116 and may also include a set of smaller size minor outer holes 118 circumferentially intermediate neighboring pairs of the major outer holes. The row 112 of holes in the inner liner include dilution air holes circumferentially offset from the major outer holes and may also include a set of minor holes 122 circumferentially intermediate major inner holes 120. The major and minor holes admit respective major and minor jets of dilution air into the combustor. The unique distribution of major and minor holes and the corresponding major and minor dilution air jets helps to minimize NOx emissions and regulates the spatial temperature profile of the exhaust gases discharged from the combustor.

Abstract: In a turbine engine with an annular recuperator surrounding the turbine, exhaust gas is directed from the turbine to the recuperator by a generally curved exhaust dome. A vortex disrupter structure extends from the exhaust dome to a point distal of the turbine to evenly distribute the exhaust gas entering the recuperator and sustain diffusion of the exhaust gas to increase the expansion ratio across the turbine.

Abstract: An engine core rotor shaft structure for gas turbine engine is disclosed, which includes an outer annular shaft body, an inner annular shaft body, high pressure turbine (HPT) rotor blades radially connected between the outer annular shaft body and the inner annular shaft body to hold the inner annular shaft body in the rear section of the outer annular shaft body, and a can type combustor mounted in a receiving chamber inside the outer annular shaft body to enlarge the diameter of the core shaft, to avoid vibration due to resonance, to save space occupation, to eliminate dissipation of heat, and to improve the thermal efficiency of the gas turbine engine.

Abstract: A combustor liner is provided having first and second annular bands which define an overlapping circumferential joint area, wherein a weld is disposed in the joint area encompassing substantially all of the axial length of the joint area. A method for producing such a combustor liner is also provided.

Abstract: In a combustion chamber arrangement, especially an annular combustion chamber arrangement for a gas turbine, one or more burners has, at its mouth, a deflecting device by which a combustion chamber arrangement is deflected. This achieves the effect of acoustic detuning, whereby the formation of a combustion oscillation is suppressed.

Abstract: The present invention deals with a turboreactor comprising at least a precombustion chamber, a combustion chamber, means for directing combustion gases from the precombustion chamber into the combustion chamber, an air compressor, and means for dividing the air flow from the air compressor while guiding it in the combustion chamber. The turboreactor comprises advantageously a contrarotating stator. The invention relates also to a process in which the blades are heated and pushed in a thermoplastic layer of the body of the rotor and/or stator.

Abstract: The present invention relates to a turbine engine, wherein compressor rotor blades and turbine rotor blades set around the interior circular surface of a hollow revolving shaft while compressor stator blades and turbine stator blades are set around the exterior circular surface of an inner engine case inside the hollow revolving shaft. By using the present invention, the configuration of the turbine engine can be compact, and the volume can be substantially reduced. Besides, the intnesity of the rotor blades can be increased, and efficiency of the compressor, the burner, the turbine, and the whole engine can be increased.

Abstract: A combustion system for a power generating gas turbine engine which includes at least a combustion chamber with a annular fuel manifold at one end of the combustion chamber and a passageway having a narrow throat downstream of the fuel manifold whereby air passes around the fuel manifold and mixes with fuel and is diffused through the passageway into the burn zone defined in the combustion chamber in an ultimate location.