Abstract: A diffuser for decelerating a compressed gas, is provided, which includes a ring divided at circumferential positions around the ring into several arcuate sections, the complete ring is assembled from these sections. The ring defines one or more passages, the cross-section of which increases in the direction of flow of a fluid through the diffuser. The interface between adjacent sections is configured so that relative movement between the sections is prevented. The interface is configured to have a series of interlocking serrations along the axial length of the diffuser. In an embodiment, there are two arcuate sections involving two interfaces, and the peaks and troughs of the serrations of each of the two sections are arranged at an angle to a longitudinal axis of the diffuser. A small relative movement between the sections is achieved when the two angles at the two interfaces have opposite signs.

Abstract: A turbomachine comprising: an annular combustion chamber (40); a centrifugal compressor (10); an annular diffuser (20) enabling the flow of gas leaving the compressor to be diffused and directing this flow of gas towards the combustion chamber, the diffuser comprising: an upstream portion (21) oriented radially and presenting diffusion passages (22) connected to the outlet from the compressor; an intermediate portion (24) that is elbow-shaped; and a downstream portion (25) having a series of circularly spaced-apart flow-straightening vanes (26); and an outer casing (32) surrounding the outside of the combustion chamber and the downstream portion; in which the flow stream in the downstream portion (25) is delimited externally by the outer casing (32).

Abstract: The turbine outlet heated diffuser may include a cavity positioned immediately downstream of an air cycle machine turbine. The turbine outlet heated diffuser may be designed to prevent or reduce ice formation at the turbine outlet by heating the diffuser wall. The cavity may receive a heated flow from an air cycle machine compressor discharge and provide a cavity outlet flow to a turbine inlet plenum. The design of the heated diffuser is such to minimize the cycle performance impact that results from the addition of an air cycle machine icing protection feature.

Abstract: A diffuser is provided that includes a first wall, a second wall, and a first plasma actuator. The first wall has a radially-extending section and a curved section. The second wall has a radially-extending section and a curved section that are spaced apart from the radially-extending section and the curved section of the first wall such that the curved section of the second wall is disposed radially inwardly from the curved section of the first wall and a flow path is defined therebetween. The first plasma actuator is disposed on the curved section of the second wall, and the first plasma actuator is adapted to generate a first electric field to ionize a first portion of air flowing along the flow path to form a plurality of ionized species capable of inducing a second portion of air to flow in a desired direction.

Abstract: The invention relates to a gas turbine, for energy generation, with a compressor, arranged coaxially to a rotor, mounted such as to rotate, for the compression of an inlet gaseous fluid, at least partly serving for combustion of a fuel in a subsequent annular combustion chamber, with generation of a hot working medium, with an annular diffuser arranged coaxially to the rotor, between the compressor and the annular combustion chamber, for distribution and deflection of the fluid, whereby a part of the fluid is diverted as cooling fluid for the turbine stages after the combustion chamber, by means of a dividing element, arranged in the fluid flow.

Abstract: A combustor of a turbine, a method of retro-fitting a combustor of a turbine and a method of building a combustor of a turbine are each provided. The combustor includes a combustion chamber along which a dilution breach is defined, a casing perimetrically surrounding the combustion chamber so as to define an airflow between the casing and the combustion chamber, the airflow being configured to supply dilution air to the combustion chamber via the dilution breach, and an easy to adjust dilution airflow tuning part disposed on the casing and in communication with the airflow. The dilution airflow tuning part is configured to increase and/or decrease an available amount of the dilution air to be supplied to the combustion chamber.

Abstract: A plenum (210) in a gas turbine engine mid-frame section (200) comprises one or more annular flow splitters (240) spaced from an inboard annular wall (232) that partition air flow flowing from a compressor into two or more portions of flow having different vectors. This provides for an improved balancing between supplying air to compression chamber intakes more directly and to transitions to aid in convective cooling. When an annular diffuser (202) is spaced between the compressor and the plenum (210), the flow splitters (240) may provide an additional diffusion action. When no annular diffuser is so provided, the flow splitters (452, 454, 456) are effective to diffuse the air flow. Embodiments include those in which an annular diffuser (304) is relatively shorter and there is a longer axial expanse in the plenum (320) for flow splitters (350, 352, 354, 356).

Abstract: Efficiency and/or power are increased in a turbine engine by using a self-contained, passive heat transfer device, such as a heat pipe, to transfer heat from working fluid in one section of the engine to working fluid in another section of the engine.

Abstract: An auxiliary power unit intake duct for an aircraft, comprising an internal straight portion 1a, an internal first bend portion 1b and an internal second bend portion 1c, extending between an air inlet 2 and a plenum entry 3 of the auxiliary power unit, a plurality of acoustic guides vanes 4,5 located in the straight portion 1a of the intake duct 1, extending horizontally between opposite side portions of the straight portion 1a; and a plurality of curved aerodynamic guide vanes 6,7 located in the second bend portion 1c proximate to said plenum inlet 3, extending vertically between a top portion and a bottom portion of the second bend portion 1c, each of the curved aerodynamic guide vanes 6,7 being concentric with said bend curvature of the second bend portion 1c.

Abstract: A diffuser particle separator may be integrated into a gas turbine engine to remove corrosive dust and salt particles from the engine's core air flow. The air flow may pass over a series of particle accumulator entrance orifices, trapping particles in a particle accumulator while allowing the air flow to continue unimpeded. Since dust deposits may become molten at high temperatures, removal of dust from the core and secondary airflow may be critical for long-life superalloy and ceramic components, particularly those with small diameter air-cooling holes and thermal barrier coatings.

Abstract: The invention relates to a gas turbine, for energy generation, with a compressor, arranged coaxially to a rotor, mounted such as to rotate, for the compression of an inlet gaseous fluid, at least partly serving for combustion of a fuel in a subsequent annular combustion chamber, with generation of a hot working medium, with an annular diffuser arranged coaxially to the rotor, between the compressor and the annular combustion chamber, for distribution and deflection of the fluid, whereby a part of the fluid is diverted as cooling fluid for the turbine stages after the combustion chamber, by means of a dividing element, arranged in the fluid flow.

Abstract: In order to desirably prevent leakage of a working fluid from the tip clearance to reduce loss of a heat drop of the working fluid, a first corresponding blade of a second stationary blade row, which corresponds to a first reference blade of a first stationary blade row, is disposed in a position distant by 2L from the position of a lower end of a rotor blade in the direction of movement of the rotor blades when the rotor blade makes the closest approach to a lower end of the first reference blade. It should be noted that L is a value obtained by multiplying the average time T required for the high-pressure working fluid to pass through the rotor blade by the traveling speed U of the rotor blade. On the other hand, a second corresponding blade of the second stationary blade row, which corresponds to a second reference blade of the first stationary blade row, is disposed in a position of a lower end of a rotor blade when the rotor blade makes the closest approach to a lower end of the first reference blade.

Abstract: A system is provided for aerodynamically coupling air flow from a centrifugal compressor to an axial combustor. The system includes a diffuser, a deswirl assembly, combustor inner and outer annular liners, a combustor dome, and a curved annular plate. The diffuser has an inlet that communicates with the centrifugal compressor, an outlet, and a flow path that extends radially outward. The deswirl assembly has an inlet that communicates with the diffuser outlet to receive air flowing in a radially outward direction, an outlet, and a flow path configured to redirect the air in a radially inward and axial direction through the deswirl assembly outlet at an angle toward a longitudinal axis. The curved annular plate is coupled to combustor inner and outer annular liner upstream ends to form a combustor subplenum therebetween and has a first opening and a second opening formed therein, the first opening aligned with the deswirl assembly outlet to receive air discharged therefrom.

Abstract: A diffuser for a centrifugal compressor in a gas turbine engine, the diffuser including a diffuser ring having a series of bores defined therethrough to receive and direct air exiting the compressor, each bore being defined by a respective bore surface including a boride layer protecting the bore surface from erosion damage.

Abstract: A jet engine that provides for the sealing of bleeding air to the cabin, the air passing through a cavity delimited, on the one hand, by the external shell of the compressor and an annular structure connected to the shell, and, on the other hand, by the external casing of the diffuser grating and a strut connected to the external casing and to an external engine casing shell, this external casing shell being fastened to the annular structure by bolting. A tube is provided between the orifice in the strut and the outlet vent, a first end of which tube is mounted in the outlet vent by a connection which is free to rotate and prevented from translational movement, and a second end of which tube is mounted in the orifice by a connection which is free to rotate and free to move translationally.

Abstract: An ECS system includes an ACM mounted adjacent an air-liquid heat exchanger through a diffuser that contains a diffuser plate. The diffuser plate receives airflow from the ACM which strikes the diffuser plate and flows radially outward and around the diffuser plate and into the air-liquid heat exchanger to provide minimal pressure loss and proper flow distribution into the air-liquid heat exchanger with significantly less packaging space.

Abstract: A apparatus and method for improving combustion by improving at least one of temperature distribution in the combustor, pressure distribution around the combustor and combustion noise level in the combustor, by redistributing air around the combustor to modify the structure of the air flow prior to entry into the combustor.

Abstract: A plenum (210) in a gas turbine engine mid-frame section (200) comprises one or more annular flow splitters (240) spaced from an inboard annular wall (232) that partition air flow flowing from a compressor into two or more portions of flow having different vectors. This provides for an improved balancing between supplying air to compression chamber intakes more directly and to transitions to aid in convective cooling. When an annular diffuser (202) is spaced between the compressor and the plenum (210), the flow splitters (240) may provide an additional diffusion action. When no annular diffuser is so provided, the flow splitters (452, 454, 456) are effective to diffuse the air flow. Embodiments include those in which an annular diffuser (304) is relatively shorter and there is a longer axial expanse in the plenum (320) for flow splitters (350, 352, 354, 356).

Abstract: Device for guiding a stream of air entering a combustion chamber of a turbomachine, comprising a flow straightener (14) followed by a diffuser (16), one of the shrouds (40) of the flow straightener being formed as one with one wall of revolution (34) of the diffuser, the other of the shrouds (38) of the flow straightener being added and attached to the other wall of revolution (32) of the diffuser, and the vanes (42) of the flow straightener being secured by one end to one shroud (40) of the flow straightener and separated by a small clearance (46) from the other shroud (38) at their other end.

Abstract: An improved combustion air charger, such as a turbocharger or a supercharger, includes a housing (10) having a rotary shaft (18) journalled therein. At least one compressor wheel (20,22) is located on the shaft (18). The housing (10) includes an ambient inlet (30) as well as a compressed air outlet (32) and a heat exchanger (36) is located between at least one of the compressor wheels (20) and the outlet (32) and is arranged so that air flow through the heat exchanger (36) is generally in the radially inward direction.

Abstract: A turbojet diffuser disposed between a compressor and a combustion chamber, and secured to an outer upstream annular flange of an outer casing of the combustion chamber by suspension means which comprise a first frustoconical wall extending from the outer longitudinal wall of the diffuser towards the combustion chamber, and a second frustoconical wall extending towards the compressor between the first frustoconical wall and the outer casing of the combustion chamber.

Abstract: The invention relates to a device for passive control of the thermal expansion of the extension casing of a turbo-jet engine, said extension casing surrounding the casing of the high pressure compressor of the turbo-jet engine, and including a flange for attachment to an upstream flange of the casing of the combustion chamber. It is characterised in that at least one circumferential cavity is provided between both said flanges wherein circulates a gas flux tapped at the inlet of the combustion chamber. One uses thus a natural circulation generated by the differential pressure. Thanks to the device of the invention, the flange is passively controlled, and the stresses resulting from a differential expansion between the skin of the casing and its attachment flange are reduced.

Abstract: A gas turbine engine pre-diffuser 40 is generally annular, including radially inner and radially outer walls 40 and 42 and a generally cylindrical midline 48 defined between the walls. The pre-diffuser 40 includes a central member 46 which forces air flowing through the pre-diffuser 40 to separate, initially to be directed away from the midline 48 before subsequently being allowed to diffuse back towards the midline 48. The majority of the diffusion takes place on the walls of the central member and is thus in an inner region of the annulus of air ejected from the pre-diffuser to pass to the combustor. Any boundary losses therefore do not significantly effect air at the extremities of this annulus, this air being destined for the annuli of the combustor and requiring relatively high energy levels.

Abstract: A gas diffusion arrangement (22) for a gas turbine engine (10) is disclosed. The gas diffusion arrangement (22) has an inlet (26) and an outlet (28) for the gas, and comprises diffusion means (32) having an upstream region (38) and a downstream region (40). Distribution means (41) is arranged between the downstream region (40) and the outlet (28) to distribute the gas into a desired flow pattern. The area ratio of the cross-sectional area of the downstream region (40) to the cross-sectional area of the upstream region (38) is greater than 1.5 to 1.

Abstract: The invention relates to the sealing of the cavity from which air is bled off to the cabin, which cavity is delimited, on the one hand, by the external shell (6) of the compressor and an annular structure (7) connected to the shell, and, on the other hand, by the external casing (12) of the diffuser grating and a strut (13) connected to said external casing and to an external engine casing shell (14), this external casing shell (14) being fastened to the annular structure (7) by bolting together flanges, using sealing means provided between the annular structure (7) and the external casing (12) of the diffuser grating.

Abstract: A diffuser arrangement is provided in which a wall surface of the diffuser arrangement incorporates an upstream part and a downstream part with an aperture between them. There is a step displacement between the upstream part and the downstream part along with specific shaping of the leading edge of the aperture whereby fluid air flow is drawn into the aperture from a fluid flow whilst avoiding undue disturbance to that flow. The present diffuser arrangement allows incorporation within an engine without complex fabrication or structural requirements.

Abstract: A gas turbine engine combustor inlet diffuser assembly has at least one diverging annular wall including a flowpath surface bounding a diffuser flowpath. An annular blowing slot is axially located along the annular wall and an annular blowing air flowpath leads to and is in fluid communication with the blowing slot. An annular array of scoops disposed in the diffuser flowpath downstream of the blowing slot have upstream facing openings and are in fluid communication with the blowing air flowpath. The annular scoops are supported on hollow struts. Each of the hollow struts has at least one radially extending flow passage for directing blowing air from the scoop to the blowing air flowpath. The blowing slot opens in a downstream direction with respect to the diffuser flowpath. A row of blowing air compressor blades disposed across the blowing air flowpath may be used to pump up the pressure of the blowing air.

Abstract: A diffuser arrangement is provided in which a wall surface of the diffuser arrangement incorporates an upstream part and a downstream part with an aperture between them. There is a step displacement between the upstream part and the downstream part along with specific shaping of the leading edge of the aperture whereby fluid air flow is drawn into the aperture from a fluid flow whilst avoiding undue disturbance to that flow. The present diffuser arrangement allows incorporation within an engine without complex fabrication or structural requirements.

Abstract: An apparatus for providing propulsive power that utilizes a novel swirl generator for rapidly and efficiently atomizing, vaporizing, as necessary, and mixing a fuel into an oxidant. The swirl generator converts an oxidant flow into a turbulent, three-dimensional flowfield into which the fuel is introduced. The swirl generator effects a toroidal outer recirculation zone and a central recirculation zone, which is positioned within the outer recirculation zone. These recirculation zones are configured in a backward-flowing manner that carries heat and combustion byproducts upstream where they are employed to continuously ignite a combustible fuel/oxidizer mixture in adjacent shear layers. The swirl generator is compatible with the throttle range of conventional gas turbine engines, provides smooth combustion with no instabilities and minimum total pressure losses, enables significant reductions the in L/D ratio of the combustor and is readily packaged into various applications.

Abstract: A gas turbine engine combustor inlet diffuser assembly has at least one diverging annular wall including a flowpath surface bounding a diffuser flowpath. An annular blowing slot is axially located along the annular wall and an annular blowing air flowpath leads to and is in fluid communication with the blowing slot. An annular array of scoops disposed in the diffuser flowpath downstream of the blowing slot have upstream facing openings and are in fluid communication with the blowing air flowpath. The annular scoops are supported on hollow struts. Each of the hollow struts has at least one radially extending flow passage for directing blowing air from the scoop to the blowing air flowpath. The blowing slot opens in a downstream direction with respect to the diffuser flowpath. A row of blowing air compressor blades disposed across the blowing air flowpath may be used to pump up the pressure of the blowing air.

Abstract: A gas turbine engine pre-diffuser 40 is generally annular, including radially inner and radially outer walls 40 and 42 and a generally cylindrical midline 48 defined between the walls. The pre-diffuser 40 includes a central member 46 which forces air flowing through the pre-diffuser 40 to separate, initially to be directed away from the midline 48 before subsequently being allowed to diffuse back towards the midline 48. The majority of the diffusion takes place on the walls of the central member and is thus in an inner region of the annulus of air ejected from the pre-diffuser to pass to the combustor. Any boundary losses therefore do not significantly effect air at the extremities of this annulus, this air being destined for the annuli of the combustor and requiring relatively high energy levels.

Abstract: In gas turbines, compressed air is supplied via an air duct to combustion chambers and is heated there. Pressure losses in the air duct should be minimized in order to ensure good overall efficiency. This is achieved by the compressed air flowing with approximately constant velocity in the air duct from the compressor to the inlet into the combustion chamber. This is supported by the effective cross section of the air duct being almost constant over this distance.

Abstract: A combustor for a gas turbine engine that facilitates reducing combustor chamber dump pressure losses to improve combustor and engine performance is described. The combustor includes a diffuser that diffuses airflow directed into the combustor. The diffuser includes an outer wall, an inner wall, and a plurality of splitter vanes between the outer wall and the inner wall. The splitter vanes are spaced radially apart, and each is spaced radially from a respective diffuser wall such that an outer passage, an inner passage, and a central passage are defined by the splitter vanes.

Abstract: An annular diffuser for a gas turbine engine includes inner and outer walls spaced apart to define a diffuser flow path, and a support member coupling the inner wall to the outer wall while allowing independent radial displacement therebetween. The support member includes a first end portion rigidly connected to the inner wall, and a second end portion extending outside of the diffuser flowpath and coupled to the outer wall by a radially extending pin. An aerodynamically-shaped shroud member surrounds the support member to thermally isolate the support member from the diffuser flowpath, thereby shielding the support member from transient thermal loads. The shroud member is pinned to the support member at a single axial location. A combustor dome panel is attached to the downstream end portion of the shroud member and is adapted to independently support inner and outer combustor liners in spaced relation to define a combustion chamber therebetween.

Abstract: A thermal insulator between the turbine and the compressor of a gas turbine engine is provided and at the same time it enables the air that has passed through the compressor to be heated by the heat of the hot turbine. A gas turbine engine comprises a centrifugal compressor for compressing air that is drawn or sucked into the compressor, a compressed air passage supplying the air compressed by the compressor to a burner, and a turbine driven by means of combustion gas generated in the burner. The compressor and the turbine are provided on a rotating shaft adjacent to each other in the axial direction. Evaporation sections on the radially inner side of a plurality of heat pipes provided in a radial form face towards the outer periphery of the rotating shaft between the compressor and the turbine, and condensation sections on the radially outer side of the heat pipes face towards the compressed air passage.

Abstract: The present invention contemplates an apparatus for forming a combustion mixture in a gas turbine engine. In one form, the apparatus includes a diffuser having at least two flowpath structures spaced apart to define a flowpath for directing fluid flow. At least one of the flowpath structures includes a flowpath surface and a plurality of edges disposed along a trailing end portion of the flowpath surface. The edges extend perpendicularly from the flowpath surface and are arranged generally parallel to the fluid flow to define a plurality of corners. A fluid vortex is generated as the fluid flow rolls over each of the corners. A spray ring is disposed along a trailing edge of one of the flowpath structures and is integrally attached to the diffuser. The spray ring includes a plurality of fuel delivery apertures adapted to spray fuel into respective ones of the fluid vortices to form the combustion mixture.

Abstract: A device for conditioning the flow of hot gas in a catalytic combustor in preparation for entry into a catalytic reactor. The device is composed of at least one and most preferably two or more disks that are secured to a shroud so as to be disposed in a plane generally perpendicular to the hot gas flow direction. Each disk is composed of a plurality of small cells oriented so that flow channels therethrough are axially disposed. The cells linearize the gas flow and exert drag on the gas flow therethrough. This generates a static pressure gradient in the flow fields upstream and downstream of the honeycomb disk, which in turn causes flow adjustments so as to produce a more uniform axial flow field. This results in a more uniform fuel/air concentration distribution and velocity distribution at the catalytic reactor inlet.

Abstract: A combustor for a gas turbine engine that facilitates reducing combustor chamber dump pressure losses to improve combustor and engine performance is described. The combustor includes a diffuser that diffuses airflow directed into the combustor. The diffuser includes an outer wall, an inner wall, and a plurality of splitter vanes between the outer wall and the inner wall. The splitter vanes are spaced radially apart, and each is spaced radially from a respective diffuser wall such that an outer passage, an inner passage, and a central passage are defined by the splitter vanes.

Abstract: A combustor apparatus for a gas turbine engine includes a combustor liner support having an annular dome panel and a plurality of load transfer members extending axially therefrom. The dome panel maintains inner and outer combustor liners in spaced relation to define a combustion chamber. The load transfer members extend into a diffuser flowpath defined by inner and outer flowpath structures which are interconnected by a plurality of struts. Each of the load transfer members surrounds at least a portion of a corresponding strut to shield the strut from fluid flowing through the diffuser flowpath.

Abstract: A combustor arrangement (8) for a gas turbine engine (3) comprising a combustion chamber, fuel nozzles(52a,52b), and a bled diffuser (7) located upstream of said combustion chamber to, in use, direct an airflow from an upstream compressor (6) into the combustor (8). The fuel nozzles (52a,52b) arranged in use to supply fuel into the combustion chamber where it is mixed and combusted with the airflow from the compressor (6). The bled diffuser (7) adapted to bleed off a portion of said airflow from a main airflow into the combustion chamber. At least one bleed duct (46) is connected to the bled diffuser (7) to return and direct the air bled from the diffuser (7) to a main gas flow through the engine (3) at a location downstream of the fuel supply means (52a,52b). The bled air from the diffuser (7) preferably providing cooling of a part of the gas turbine engine (3), for example part of the turbine (10) or combustor outlet vane (58), downstream of the combustor (8).

Abstract: A fuel injection system for a gas turbine engine combustor, wherein the combustor includes a dome inlet module having a plurality of flow passages formed therein and at least one cavity formed in a liner downstream of said dome inlet module. The fuel injection system includes a fuel supply and a plurality of fuel injector bars positioned circumferentially around and interfacing with the inlet dome module. The fuel injector bars are in flow communication with the fuel supply, with each of the fuel injector bars further including a body portion having an upstream end a downstream end, and a pair of sides. A plurality of injectors formed in the body portion and in flow communication with the fuel supply, whereby fuel is provided to the dome inlet module flow passages and/or the cavity through the fuel injector bars.

Abstract: A fuel injection system for a gas turbine engine combustor, wherein the combustor includes a dome inlet module having a plurality of flow passages formed therein and at least one cavity formed in a liner downstream of said dome inlet module. The fuel injection system includes a fuel supply and a plurality of fuel injector bars positioned circumferentially around and interfacing with the inlet dome module. The fuel injector bars are in flow communication with the fuel supply, with each of the fuel injector bars further including a body portion having an upstream end, a downstream end, and a pair of sides. At least one injector is formed in the downstream end of the body portion and in flow communication with the fuel supply, whereby fuel is provided to the cavity through the fuel injector bars in accordance with a Rich-Quench-Lean (RQL) process.

Abstract: A deswirler system for a centrifugal compressor of a gas turbine engine that improves overall engine performance as a result of exhibiting significantly reduced friction losses. The deswirler system generally entails an annular-shaped manifold having an inlet configured to receive radially-outward flowing gas from a diffuser of the compressor, an outlet configured to discharge the gas in an axial downstream direction, and an arcuate passage therebetween. The deswirler system further includes a plurality of deswirler vanes directly within the arcuate passage and closely coupled to the diffuser.

Abstract: A three stage lean burn combustion chamber (28) comprises a primary combustion zone (36), a secondary combustion zone (40) and a tertiary combustion zone (44). Each of the combustion zones (36,40,44) is supplied with premixed fuel and air by respective fuel and air mixing ducts (76,78,80,92). The primary fuel and air mixing ducts (78,78) comprise first and second radial flow swirlers (60,62). The primary fuel injectors (64,66) inject fuel into the first and second swirlers (60,62). The primary fuel injectors (44,66) and the first and second swirlers (60,62) are arranged such the fuel to air ratio of the fuel and air flowing from the passages (61) of the first swirler (60) into the primary combustion zone (36) is different to the fuel to air ratio of the fuel and air flowing from the passages (65) of the second swirler (62) into the primary combustion zone (36).

Abstract: A coating for a portion of the flowpath of a diffuser of a centrifugal compressor. The coating covers a portion of the diffuser against which flows high velocity air from the compressor blades. The coating is a smooth, tenacious, heat-resistant powder paint that reduces the aerodynamic drag of the diffuser walls on the high velocity air thus increasing the efficiency of the compressor.

Abstract: An annular type gas turbine combustor which can regulate air flow distribution without being influenced by high temperature flames and with a desired fuel-air ratio for reducing the generation of NO.sub.x etc. A slide duct (10), for regulating air flow distribution, is disposed slidably in an axial direction at a terminal end of a diffuser (1). And as the slide duct (10) is moved, air flow supplied downstream to a swirler (4) is regulated so as to obtain a desired fuel-air ratio.

Abstract: In a gas-turbine annular combustion chamber (4) which is arranged downstream of a compressor (1) and is equipped on its front plate with at least one row of premix burners (5) arranged in an annular form, in each case a combustion-air duct (15) designed as a diffuser leads directly downstream of the compressor outlet from the guide vanes (9) of the last compressor row to each burner (5), at the downstream end of which combustion-air duct (15) at least one longitudinal-vortex generator (16) is located, at least one fuel injection means (17) being provided in or downstream of the longitudinal-vortex generator (16). A mixing duct (19) which ends in the combustion chamber (4) and has a constant height (H) and a length (L) which corresponds approximately to twice the value of the hydraulic duct diameter (D) is arranged downstream of the fuel injection means (17). The overall size of the gas turbine in the region of the combustion chamber (4) can thereby be substantially reduced.

Abstract: A gas turbine engine with first and second axial flow compressors and an intercooler therebetween is provided with a diffuser to diffuse the air flow leaving the downstream end of the first axial flow compressor at a first radial distance from the axis to the upstream end of the intercooler at a second radial distance from the axis. The diffuser comprises a first radially extending wall, a second radially extending wall and a plurality of diffuser vanes extending therebetween which define a plurality of generally radially extending diffusing passages. The vanes are wedge shaped. This arrangement provides diffusion in a short axial length.

Abstract: In a power plant which essentially comprises a compressor (2), a combustion chamber (4) and a turbine (7), an injector system (30) is provided in the region of an outlet diffuser (3) of the compressor (2). This injector system (30) comprises a ring of individual injectors (31) which are arranged on the inner circumference of the outlet diffuser (3). The propulsion nozzles of these injectors (31) are loaded with steam (10) from a waste-heat steam generator (9). The throughflow cross section of the outlet diffuser (3) at the same time forms the combining nozzle of the injectors (31).

Abstract: A gas turbine engine diffuser case construction is described that increases the amount of available discharge air for service applications compared to the prior art while reducing downstream component distortion and degradation. The diffuser case has passageways for diffused air that connect a diffused air region to a manifold. At least one orifice is located at the end of each of the passageways in the diffused air region. The size of the orifices controls the amount of air that is discharged from the diffused air region to the manifold. Diffused compressor air collects in the manifold and is then extracted from the manifold as needed for engine or service applications. The size and location of the orifices in the diffuser radially and circumferentially control the rate of discharge of the diffused air from the diffuser before the air enters the combustor region. The ability to influence the flow of air at particular locations around the combustor helps the designer to improve cooling uniformity.