Abstract: A propulsion plant operating on the basis of a catalytic and/or chemical decomposition of a liquid propellant, is equipped with a so-called aspirator ring (19) that covers the exit openings of the injection pipes (11), thereby screening these openings toward the decomposition chamber. The aspirator ring (19) is further enclosed by an aspirator screen (19A). The aspirator ring is constructed as a vaulted channel (20) in which the discharge bores (21) are arranged in a pattern that assures a uniform distribution of the propellant into the catalyst bed in the decomposition chamber. The aspirator screen (19A) is also formed with a channel and both the channel in the screen (19A) and the channel in the aspirator ring (19) preferably have a semicircular cross-section. The catalyst bed is divided into two separate beds so that the respective housing can also be divided so that subassemblies of the thruster may be built separately and then assembled.

Abstract: A gas turbine engine fuel injection assembly includes a combustor case surrounding a combustor having a dome. An arcuate manifold is disposed inside the combustor case adjacent to the dome, and includes a plurality of fuel injectors disposed in flow communication therewith for receiving fuel for injection into the combustor. An inlet stem is joined in flow communication with the manifold for channeling fuel thereto, and includes an inner fitting disposed in a mounting port in the combustor case. A mounting adaptor is joined to the combustor case at the mounting port, and circumferentially engages the inner fitting to restrain torsional movement thereof. An outer fitting extends in the adaptor and engages the inner fitting in flow communication therewith for supplying fuel thereto.

Abstract: The object of the invention is to create a simple and compact fuel feed for a gas turbine having an annular combustion chamber. In addition, the operational reliability of the gas turbine is to be improved.According to the invention, this is achieved when only one central main-gas feed line (15) and one central pilot-gas feed line (16) are connected to the annular combustion chamber (1). A main-gas ring line (13) and a pilot-gas ring line (14) are formed in the interior of the gas-turbine casing (5) or a burner hood (3) known per se. These ring lines (13, 14) are connected upstream to the main-gas feed line (15) and the pilot-gas feed line (16) respectively and downstream to the annular combustion chamber (1). (FIG.

Abstract: A fuel injection system for a gas turbine engine combustion chamber comprises at least one pilot fuel injector having a fuel discharge passage communicated to the combustion chamber and at least one secondary fuel injector having a fuel discharge passage communicated to the combustion chamber and a fuel purging passage communicated to a higher pressure region upstream of the combustion chamber. A fuel flow control device supplies metered fuel flow to the pilot fuel injector under all regimes of engine operation and to the secondary fuel injector under certain, intermittent regimes of engine operation requiring fuel flow above a selected level. A shuttle valve is movable to a first position to provide fuel flow communication between the fuel control device and the fuel discharge passage of the secondary fuel injector for fuel injection to the combustion chamber.

Abstract: In a gas turbine having a plurality of combustors, each having an annular array of outer fuel nozzles arranged about a center fuel nozzle. A method for operating the combustors wherein the annular array is supplied with fuel from a diffusion manifold and at least one pre-mix manifold, and further wherein the center nozzle is supplied with fuel from a center premix manifold includes the steps of:a) at start-up, supplying some of the annular array nozzles with diffusion fuel;b) at part speed supplying pre-mix fuel to one of the nozzles in the annular array;c) at full speed, part load, transferring the annular array nozzles receiving diffusion fuel to pre-mix fuel;d) as load is increased supplying pre-mix fuel to the center nozzle; and thene) supplying additional pre-mix to all nozzles as load increases.

Abstract: This invention pertains to small air breathing gas turbine engines (1). Engines of the invention employ improved combustion systems, including improved fuel injection systems and improved combustor design. These improvements enable the engine to idle at speeds as low as one-tenth the maximum operating speed of the engine, which gives the engine an increased range of operating speeds. The improved designs generally provide for pre-mix assemblies intimate with the fuel injectors, a typically flame-free primary "A" mixing zone in the combustion chamber, and a primary "B" combustion zone, in addition to the secondary combustion zone and the dilution zone. In preferred embodiments, the bearing assembly is located between the compressor and the turbine wheel. A cooling cavity is provided between the turbine wheel and the bearing assembly, thus protecting the bearings from the turbine heat.

Abstract: In a gas turbine having a plurality of combustors, each having an annular array of outer fuel nozzles arranged about a center axis, an improved construction is provided wherein each combustor has a center nozzle located on the center axis. Various methods of operating the improved combustor are also disclosed.

Abstract: A gas turbine combustor having a pre-mixing zone and a combustion zone separated by a barrier plate. A plurality of concentrically arranged annular passages, into which lean mixtures of fuel and air are introduced, are disposed in the pre-mixing zone. Each of the annular passages has a discharge end in which a plurality of swirlers are circumferentially arranged that allow fuel/air mixtures flowing through the annular passages to enter the combustion zone, the swirlers acting as flame holders. A toroidal fuel manifold upstream of each of the annular passages allows the fuel/air ratio in each annular passage to be individually controlled. The firing temperature of the combustor is increased by sequentially supplying fuel to each annular passage in turn, with the innermost passages being supplied with fuel first. A pilot fuel assembly is centrally disposed with respect to the annular passages and has swirler vanes and an ignitor.

Abstract: In a gas-operated, flame-stabilizing premixing burner for a combustion chamber, the combustion air can be introduced at least approximately tangentially into a premixing space (130). The fuel is injected via a plurality of nozzles (117) lined up in the longitudinal direction of the premixing space and is intensively mixed with the combustion air prior to ignition. The nozzles (117) are subdivided into at least two groups having a separate fuel feed (120, 121) in each case.

Abstract: A flow control valve for multiport injection of pressurized liquid through an array of injection nozzles for a gas turbine engine is connected by nozzle ports to the engine. In accordance with a first embodiment, the valve includes a slidable valve spool which progressively opens and closes a plurality of radially opening first ports. In accordance with a second embodiment the valve employs a rotatable disc which progressively opens and closes a plurality of axially opening first ports. Both embodiments utilize non-circular apertures to effect opening of the first ports. A plurality of second ports communicate with a chamber containing the valve spool or rotatable disc, the second ports communicating with individual compensators that are also in communication with the first ports and nozzle ports. The compensators connect the first ports to the nozzle ports through a variable flow valve which is biased to the fully open opposition against the pressure of liquid in the second ports.

Abstract: The present invention has an object to provide a gas turbine combustor which is able to effect stable combustion in a wide range of fuel flow rate. A burner 1 is provided with fuel nozzles 31 and 32. When a fuel flow rate is small, diffusion flame is formed with fuel supplied from the nozzle 31 with a ring-shaped flame stabilizer 11. Next, fuel is supplied from the nozzle 32 to mix with air, reach to the flame stabilizer 11 and be held by the diffusion flame already formed, whereby stable premixed flames are formed in the flame stabilizer 11 from a range of low fuel air ratio. Further, when flame is propagated from the burner 1 to the burner 2, a fuel air ratio at the outer periphery side of the burner 1 is locally raised by the fuel supplied from the nozzle 31, whereby the combustion stability can be raised in a wide range of fuel flow rate and propagation of flame to adjacent burners becomes easy.

Abstract: An annular combustor structure particularly suited to staged combustion has a faired centrebody incorporating a diffuser in the head of the combustor. The structure carries fuel injector mountings for pilot and main stage combustion which receive fuel from respective manifolds running around the hollow interior of the structure. There are two annular chambers each containing a fuel manifold, and cooling air passes through the chambers before entering the combustion volume. The manifolds are connected to an external source of fuel by conduits which pass radially through the combustion chamber outer casing walls.

Abstract: A fuel injection assembly for a combustion chamber of a gas turbine engine is disclosed in which a plurality of premixing chambers are in communication with the combustion chamber and adjacent pairs of the premixing chambers are connected to a single prevapoization chamber by prevaporization conduits. Fuel is injected into the prevaporization chamber via a fuel injector and is mixed with oxidizer passing into the prevaporization chamber such that the fuel/oxidizer mixture has a richness exceeding the stoichiometric ratio. The fuel/oxidizer mixture passes into the prevaporization conduits through a diaphragm opening and, upon entering the premixing chamber, is mixed with additional oxidizer such that the fuel/oxidizer mixture in each premixing chamber has a richness less than the stoichiometric ratio.

Abstract: An air fuel mixer is disclosed having a mixing duct, a set of inner and outer counter-rotating swirlers adjacent the upstream end of the mixing duct, and a hub separating the inner and outer swirlers to allow independent rotation thereof, wherein high pressure air from a compressor is injected into the mixing duct through the swirlers to form an intense shear region and fuel is injected into the mixing duct so that the high pressure air and the fuel is uniformly mixed therein so as to produce minimal formation of pollutants when the fuel/air mixture is exhausted out the downstream end of the mixing duct into the combustor and ignited. In addition, the mixing duct of the mixer includes slots, preferably with flares associated therewith, at its downstream end which modify the shape and direction of mixing eddies emanating therefrom.

Abstract: A gas turbine engine combustion chamber has primary and secondary combustion zones. An annular secondary fuel and air mixing duct surrounds the primary combustion zone. The annular secondary fuel and air mixing duct is defined at its radially outer extremity by an annular wall. The annular wall is at least partially formed by an annular manifold. The annular manifold has a number of radially inwardly extending fuel injectors which inject fuel into the secondary fuel and air mixing duct. The annular fuel manifold is mechanically isolated from the remaining portion of the annular wall by an annular gap. The annular fuel manifold is supported from the combustor casing by a fuel supply pipe which is secured to the combustor casing and the annular fuel manifold.

Abstract: A structure having radial holes extending from a ring passage communicated with a revolution lead fittings having spiral grooves and fuel lead fittings having truncated cone shaped holes and straight holes, and further provided with an outer fixing ring having spray holes and hemispherical recesses.

Abstract: In a fuel staging system, metered fuel from a fuel metering unit is directed into a fuel inlet line coupled to a sequence valve. A minimum pressure valve maintains the pressure of the fuel flowing through the fuel inlet line at a minimum pressure above the low pressure of the system in order to ensure sufficient force margins to stroke the sequence valve. A main fuel manifold is coupled downstream of the sequence valve, and a plurality of main fuel nozzles are each coupled to the main fuel manifold through a respective main nozzle shut-off valve. A first set of pilot nozzles is coupled to the main fuel manifold through the sequence valve, and a second set of pilot nozzles is coupled to the main fuel manifold through the sequence valve. At low engine speeds, in the first and/or second pilot open positions, fuel flows to either pilot nozzle through the main fuel manifold, and the main fuel nozzles are isolated from the main fuel manifold by the main nozzle shut-off valves.

Abstract: Lean direct injection is used in a gas turbine combustor to reduce NO.sub.x emissions. The combustor has a multi-hole perforated plate disposed at the head end of the combustion chamber. The plate has a plurality of independent fuel jets and air jets for separately injecting fuel and air, respectively, into the combustion chamber in amounts which produce a lean fuel-air equivalence ratio. The perforated plate permits the fuel and air to be injected separately without any mixing outside of the combustion chamber. Fuel is delivered to the fuel jets by a fuel manifold having a branch connected to each one of the fuel jets. Air can be delivered to the air jets by an air plenum chamber located upstream of the perforated plate and in fluid communication with the air jets.

Abstract: A fuel distribution system for a gas-turbine engine is disclosed which ensures homogenous fuel supply to all fuel injectors in order to avoid local fuel oversupplies that might cause local overheating of the turbine blades. The fuel system compensates for the deleterious load factors experienced by the aircraft and the gas turbine engine that affect the fuel distribution to the fuel injectors. The system supplies fuel to a plurality of fuel injectors arranged in an annular array so as to inject fuel into an annular combustion chamber of the gas turbine engine. The system has a closed loop, generally annular fuel connector line and a plurality of fuel distributors connected in parallel to the fuel connector line so that fuel passes into a central chamber of each of the distributors. Each fuel distributor has end chambers on opposite ends of the central chamber in which are movably located pistons whose positions within the end chambers open or close a fuel outlet in communication with each of the end chambers.

Abstract: A gas turbine engine fuel distribution assembly includes a fuel manifold having a fuel distribution channel therein, and a plurality of fuel distribution valves disposed inside the manifold. Each valve is disposed in the manifold for providing a bypass channel therearound for allowing unobstructed flow of the fuel through the distribution channel in turn to each of the valves. Each valve has an inlet disposed in flow communication with the distribution channel for receiving a portion of the fuel therefrom, and an outlet extending laterally through the manifold for providing controlled fuel flow therefrom.

Abstract: A variable friction force damper utilizes geometric variations in a sliding surface to vary the frictional force. In particular, variation in thickness of the member on which the sliding surface is formed produces a corresponding variation in the displacement of a compression spring. This variation in spring displacement produces a variation in the compression force exerted by the spring, i.e., the force with which the sliding member is pressed against the sliding surface. This in turn causes a variation in the frictional force which the sliding surface exerts on the sliding member. The result is a variable friction force damper having a variable damping capability which allows a controlled response at variable input levels.

Abstract: An improved flameholder for a thrust augmentation combustor includes a toroidal pilot gutter and inner and outer radial gutters, all having U-shaped cross-section and formed from a single piece of metal. The one-piece formation enables design improvements and variations which were not possible in the multiple piece construction of the prior art.

Abstract: A gas turbine combustor having a pre-mixing zone and a combustion zone separated by a barrier plate. A plurality of concentrically arranged annular passages, into which lean mixtures of fuel and air are introduced, are disposed in the pre-mixing zone. Each of the annular passages has a discharge end in which a plurality of swirlers are circumferentially arranged that allow fuel/air mixtures flowing through the annular passages to enter the combustion zone, the swirlers acting as flame holders. A toroidal fuel manifold upstream of each of the annular passages allows the fuel/air ratio in each annular passage to be individually controlled. The firing temperature of the combustor is increased by sequentially supplying fuel to each annular passage in turn, with the innermost passages being supplied with fuel first. A pilot fuel assembly is centrally disposed with respect to the annular passages and has swirler vanes and an ignitor.

Abstract: A combustor apparatus is provided for a gas turbine engine including a circumferential ring of injectors. A fuel manifold provides a source of fuel for the injectors. The combustor apparatus includes a plurality of first fuel flow paths between the manifold and all of the injectors. Fuel is provided to all of the injectors via these first flow paths. At least one second fuel flow path including an electronically switchable valve is coupled between the manifold and a selected one of the injectors to provide additional fuel thereto at certain times when special engine operating conditions, such as demand for initial ignition, crossfire and altitude relight are sensed.

Abstract: A fuel delivery system (10) incorporates a fuel splitter valve assembly (30) including a first valve (40) having a selectively positionable spool (42) disposed therein for controllably metering fuel flow therethrough for delivery to a first combustion chamber (22) and a second valve (50) operative to pass a second portion of the fuel flow to a second combustion chamber (24) while maintaining at least a desired minimum pressure upstream of the first valve (40) whenever the second combustion chamber is in operation, and to shut-off fuel flow to the second combustion chamber whenever the second combustion chamber is not in operation.

Abstract: A fuel delivery system for a turbine engine having a primary fuel flow circuit and a secondary fuel flow circuit. The primary fuel flow circuit has a unstaged-first manifold and a staged-second manifold. A control valve and flow meter regulate the amount of fuel flow to the unstaged and staged manifolds with a first mode of operation providing fuel to the unstaged manifold only, and a second mode of operation providing fuel to both the staged and unstaged manifolds. When high power is required of the engine, a secondary fuel circuit is provided with a secondary manifold which is connected to a distribution valve which opens when fuel flow is at a predetermined level. The invention allows fuel to be efficiently atomized at various levels of powered operation by having a control center which controls the opening and closing of the distribution valve and control valve based upon fuel flow rates fuel/air ratios and air start core speed.

Abstract: In a gas turbine engine, fuel is supplied to a fuel metering valve which is connected to a shutoff valve. The shutoff valve supplies fuel to a manifold. Fuel flow between the manifold and main fuel nozzles is controlled by fuel staging valves while the pilot nozzle is connected directly to the manifold by a fuel line which may include an orifice. The orifice is designed to match the flow in the pilot nozzle with the flow in main nozzles when the fuel staging valve is on.

Abstract: A system for supplying fuel to a plurality of fuel injectors of a gas turbine engine is disclosed having first and second fuel distributors and a fuel transfer device which enables fuel to flow from the fuel source to the first and second fuel distributors as a function of the fuel flow rate. The fuel transfer device enables fuel to flow through the first fuel distributor under low fuel flow conditions. The first fuel distributor sequentially supplies fuel to each of a plurality of groups of fuel injectors when the fuel flow rate is below a predetermined threshold value. When the fuel flow exceeds a high threshold value, the fuel transfer device enables fuel to flow through the second fuel distributor which supplies fuel to all of the fuel injectors of the engine.

Abstract: A fuel injector system for a combustor of a gas turbine engine includes first and second fuel injectors rendered operative to discharge fuel to the combustor during a high power regime of engine operation and rendered non-operative during a lower power regime of engine operation. The first and second fuel injectors include respective first and second fuel discharge passages in fuel flow communication to one another and to the combustor via associated fuel discharge lips to sustain a flame region. The first and second fuel injectors are operatively associated with respective first and second air discharge means having air discharge lips for discharging air to the combustor for sustaining the flame region therein. When the fuel injectors are rendered non-operative, different pneumatic pressures are established at the fuel discharge lips to purge fuel from the fuel injectors to the combustor.

Abstract: A support system for coupling elements of different temperatures while reducing thermal and dynamic stress. The support system couples a fuel manifold to a combustor in a gas turbine engine and includes first and second elongated brackets connected, respectively, to the combustor and to the fuel manifold. Distal ends of each of the brackets are positioned in overlapping relationship. A slot is formed in one of the overlapping distal ends and an aperture formed in the other aligned generally with the slot. A fastener extends through the aperture and slot compressing the ends against each other with a preselected force such that the ends cannot separate but can exhibit a sliding motion in the slot direction. The compressive force is controlled by a spring member positioned between one end of the fastener and an adjacent surface of one of the first and second brackets. A second disk fits around the stud similar to a large flat washer and the spring member fits around the stud between the two disks.

Abstract: A manifold system for a marine or industrial gas turbine engine, the system being independent of the engine. An annular steam manifold and an annular gaseous fuel manifold are provided transversely surrounding and substantially coaxial with the engine. The manifolds have a plurality of outlets, each connected by a flexible metal hose to its respective one of the engines' fuel nozzles. Each manifold is supported by a pair of stanchions mounted on the engine's foundation assembly and a pair of brackets attached to each stanchion. The brackets provide limited vertical, lateral, axial and tilt adjustments of the manifolds.

Abstract: A fuel leakage protection system is disclosed for a gas turbine engine having a fuel system that has a plurality of fuel nozzles that are spaced around a combustor section, a plurality of fuel tubes that are connected to the fuel nozzles individually, two generally semi-circular fuel manifolds that are located aft of the fuel nozzles and that ar connected to the fuel tubes and a branch manifold that is connected to the fuel manifolds.The fuel leakage protection system comprises a collar assembly encapsulating each fuel supply fitting connecting a fuel tube to a fuel nozzle or the branch manifold to a fuel manifold to trap leakage fuel. A fuel collection manifold collects leakage fuel from the collar assemblies via a plurality of fuel return tubes that are connected to the manifold at one end and fluidly connected to the collar assemblies at the other end.

Abstract: A burner for a gas turbine combustor has a perforated plate facing the combustor and a plurality of gaseous fuel premixing tubes conveying a lean premixed fuel through the plate. A gas pilot extending through the plate projects jets of fuel parallel to the plate between the tube locations nearest the pilot. Ignition stability at turndown is maintained.

Abstract: A fuel drain can assembly positioned between the fuel manifold and the fuel nozzles of a gas turbine engine which utilizes detachable connectors rotatable into locking engagement without special tooling.

Abstract: A fuel delivery system for a dual annular combustor having a pilot stage manifold, a first main stage manifold, and a second main stage manifold which are connected to a three-position staging valve. The staging valve is controlled by a fuel/air digital electronic control (FADEC) which directs the staging valve to a closed (staged) position, to a partially open (partially staged) position, or to an open (unstaged) position in accordance with a staging-valve-fuel-to-air-ratio control schedule.

Abstract: A fuel manifold is disclosed for providing fuel to a combustor disposed radially inside an annular casing. The manifold is disposed inside the casing and includes an arcuate manifold tube, a thermal insulation layer surrounding the tube, and a cover layer surrounding the insulation layer which is substantially rigid for protecting the insulation layer from physical damage.

Abstract: A compressor stall recovery apparatus senses a compressor stall condition and cuts off the supply of fuel to selected ones of the nozzles of the combustor until after the stall condition has abated.

Abstract: A fuel injection system for use in the combustor section of a gas turbine engine is provided with first and second manifolds of toroidal configuration. First outlet nozzles of a plurality of fuel injectors associated with the first manifold are positioned at spaced circumferential locations lying in a common plane at similar radial distances from the centerline of the engine. Second outlet nozzles of a plurality of fuel injectors associated with the second manifold are similarly positioned but evenly spaced from the first outlet nozzles. Fittings for the fuel injectors are joined by means of flexible tubing and are similarly joined to a fuel inlet fitting for connection to a source of fuel. By providing a pair of separate manifolds and flexible tubing interconnecting the fittings of adjacent pairs of fitting for the injectors, outlet nozzles for the fuel injectors can achieve a more dense configuration, their assembly and maintenance can be substantially simplified, and vibratory stresses are minimized.

Abstract: A gas turbine engine (10) incorporates a fuel manifold (12) that distributes fuel to fuel injectors (14) that are mounted to a support casing (16). When a problem is encountered within the fuel injector (14), the problem injector (14) may be disconnected from the fuel manifold (12) and removed without disturbing the fuel manifold (12).

Abstract: In order to provide a simple, low-cost fuel injection system, particularly for a dual zoned annular combustor (10), a single fuel manifold (42) is utilized having first and second manifold portions (42a and 42b) each with a respective plurality of openings (44a and 44b) through which fuel may flow into first and second fuel injection zones (34 and 36). The annular combustor (10) typically includes a combustor housing (12) comprising an axially extending sleeve (14) disposed about a longitudinally extending axis (16) and an annular liner (18) defining a combustion chamber (26) and being spaced from the housing (12) and the sleeve (14) to form a compressed air flow path (28) therebetween.

Abstract: A rigid fuel manifold system for use in a gas turbine engine which is resistant to vibrations from operation of the engine. The system comprises two fuel feed manifolds, a fuel drain manifold and support brackets which attach these manifolds to one another and provide a structure with a high natural frequency above the 1/rev of the engine operating range. Structural support brackets that permit easy assembly and low installation stress are also provided.

Abstract: The high cost of a fuel injection system in a turbine engine may be substantially reduced by utilizing a combination manifold/fuel injector (80) that is formed of a flattened tube (82) of annular shape and disposed within the case (58) of the engine in substantially surrounding relation to an annular combustor (44). The flattened tube includes fuel injecting apertures (88, 100) which need not be precision formed but which are aligned with the flared ends (74) of inlet tubes (70) extending into the combustor (44) from the plenum defined by the space between the combustor (44) and the case (58).

Abstract: A gas turbine engine 10 having a rotor 12 with turbine blades 14 and a nozzle 16 adjacent the turbine blades 14 for directing hot gases of combustion at the turbine blades 14. The engine 10 also includes an annular combustor 18 about the rotor 12 and an annular combustor housing 30 substantially surrounding the combustor 18 in generally concentric spaced relation. A fuel injection system is operatively associated with the combustor 18 for injecting fuel into a combustion space 28 to be combusted with air from a compressor 34. The fuel injection system may include a plurality of circumferentially spaced fuel injectors 68 disposed in an outer wall 20 of the combustor 18 together with a generally oval shaped manifold 70 in fluid communication with a primary fuel source and the fuel injectors 68. A turbine shrould 36 is provided to extend radially outward from the rotor to the outer wall of the combustor on the side of the nozzle 16 opposite the combustion space 28.

Abstract: Manifold head effects at low fuel flows in a fuel injected air breathing turbine are minimized by utilizing fuel injectors having fuel injecting tubes (66) with open ends (70) for fuel injection and provided with elongated capillary tubes (88) upstream thereof and connected to receive fuel from a fuel manifold (48) having an inlet (56) while uniform, relatively low velocity fuel exit flow from the ends (70) the injecting tubes (66) is achieved through the use of internal impingement surfaces (96, 102, 106, 110, 124). Pressure loss differences contributing to nonuniform fuel flow and resulting from some fuel injecting tubes (66) be more distant from the manifold inlet (56) are minimized by shortening the length of the capillary tubes (88) furthest from the manifold inlet (56).

Abstract: The expense of fabricating an annular combustor (10) for a gas turbine is minimized by providing a combustor housing (12) including an axially extending sleeve (14) and an annular liner (18) disposed within the housing (12) and about the sleeve (14). The annular liner (18) has concentric inner and outer axially elongated walls (20, 22) spaced from the sleeve (14) and the housing (12), respectively, and also has a radially extending wall (24) spaced from the housing (12) and interconnecting the inner and outer walls (20, 22) at one end to define a combustion chamber (26). The liner (18) to spaced from the housing (12) the sleeve (14) to define a compressed air flow path (28) extending from a radially outer compressed air inlet (30) to a radially inner compressed air outlet (32) in communication with the combustion chamber (26) axially remote from the radially extending wall (24).

Abstract: A dual manifold fuel supply system for the combustor of gas generator in a turbofan engine wherein alternating circumferentially aligned fuel injectors are alternately supplied with fuel by a first or second fuel manifold. Fuel flow to both manifolds are controlled by a fuel control, preferably a full authority digital electronic control, and the second manifold has an off valve interposed between it and the fuel control.

Abstract: A gas turbine engine 10 having a rotor 12 with turbine blades 14 and a nozzle 16 adjacent the turbine blades 14 for directing hot gases of combustion at the turbine blades 14. The engine 10 also includes an annular combustor 18 about the rotor 12 and an annular combustor housing 30 substantially surrounding the combustor 18 in generally concentric spaced relation. A fuel injection system is operatively associated with the combustor 18 for injecting fuel into a combustion space 28 to be combusted with air from a compressor 34. The fuel injection system may include a plurality of circumferentially spaced fuel injectors 68 disposed in an outer wall 20 of the combustor 18 together with a generally oval shaped manifold 70 in fluid communication with a primary fuel source and the fuel injectors 68. A turbine shroud 36 is provided to extend radially outward from the rotor to the outer wall of the combustor on the side of the nozzle 16 opposite the combustion space 28.

Abstract: A fuel manifold (70) for a gas turbine engine is formed from a plurality of spaced bodies (10) and interconnecting conduits (72). The bodies (10) are adapted to provide fluid communication between the conduits (72) and a corresponding plurality of nozzle assemblies (62) which are removably secured to the bodies. The nozzle assemblies (62) can thereby be detached from a structure (58) to which the manifold (70) is secured, while all components (10, 72) of the manifold (70) remain secured to the structure (58) and to each other. A fail-safe sealing member (32) adapted for use with the fuel manifold/nozzle assembly combination (70, 62) is also disclosed.

Abstract: A gas trubine combustor includes a main body with a combustion portion into which mixture of fuel and air is supplied. The gas mixture is burned through a catalytic reaction at a catalyst body arranged in the main body. The gas mixture passed through the catalyst body is introduced into a gas phase combustion portion provided in the main body. Between the catalyst body and the gas phase combustion portion is arranged a dividing unit which has a plurality of branch passages. The gas mixture passed through the catalyst body is divided by the dividing unit into a plurality of gas streams flowing through the branch passages. Each of the gas streams is mixed with fuel supplied from a fuel supply tube.

Abstract: A fuel gas injector for a gas turbine engine employs a plurality of closely spaced parallel venturi tubes disposed in a pair of spaced-apart header plates. The venturi tubes are brazed to the header plates and the perimeters of the header plates are sealed to form a plenum into which pressurized gaseous fuel is supplied. Orifices lead from the plenum to throats of the venturi tubes thereby injecting the gaseous fuel at right angles into the high-velocity air stream existing at the the throats of the venturi tubes. High shear is imposed on the injected fuel for providing complete mixing with the air. The high air velocity in the throats of the venturi tubes avoids flashback and flameholding. The combined flow from the plurality of venturi tubes mixes downstream thereof to provide a uniform velocity and fuel-air mixture across the flow field. This flow field is suitable for use in a catalyst bed which may be disposed downstream of the venturi tubes.