Abstract: A monolithic combustor apparatus comprises an outer casing comprising a forward flange, a fuel manifold disposed on the outer casing and defining an annular chamber extending perimetrically around the outer casing, a combustor liner disposed within the outer casing, the combustor liner defining an annular combustion chamber, a first annular plenum disposed between the outer casing and the combustor liner, an inner liner disposed radially from the combustor liner, a first inner flange extending forward from the combustor liner, and a second inner flange extending radially inward from the first inner flange.

Abstract: A system includes an internal manifold ring defining at least one circumferentially extending flow channel. A plurality of feed arms extends outward from the manifold ring. A circumferentially segmented outer ring can be supported from the feed arms, outboard of the feed arms. Each feed arm can include a plurality of branches extending therefrom, wherein each branch is in fluid communication with the manifold ring through a respective one of the feed arms. A plurality of nozzles can be included, each nozzle connected to a respective one of the branches, wherein the system is devoid of nozzles radially inward from the manifold ring.

Abstract: A monolithic combustor apparatus comprises an outer casing comprising a forward flange, a fuel manifold disposed on the outer casing and defining an annular chamber extending perimetrically around the outer casing, a combustor liner disposed within the outer casing, the combustor liner defining an annular combustion chamber, a first annular plenum disposed between the outer casing and the combustor liner, an inner liner disposed radially from the combustor liner, a first inner flange extending forward from the combustor liner, and a second inner flange extending radially inward from the first inner flange.

Abstract: The invention relates to an aviation turbine engine fan module including a de-icing system (10) for de-icing an inlet cone (1) and comprising a sheath (30) placed inside an inside space defined upstream by the inlet cone, said sheath comprising a first duct (38) having at least one hot air admission orifice (42), said first duct being configured to convey hot air from a bearing enclosure (22) of the engine towards a wall of the inlet cone in order to heat it from the inside, the sheath further comprising a second duct (40) having at least one outlet situated downstream from the admission orifice of the first duct, said second duct being configured to discharge air from the first duct towards the downstream end of the engine. The invention also provides a method of de-icing a turbine engine inlet cone.

Abstract: An afterburner for use with a gas turbine engine includes a plurality of vanes distributed downstream of a turbine of the gas turbine engine. The vanes can include one or more exit apertures through which hot combustion flow from a pilot can be injected. The exit apertures can be protrusions or slots in some forms. In some embodiments, cooling passages are arranged around the exit apertures. An upstream vane portion can be positioned to inject fuel to be combusted via interaction with hot flow that is discharged through the exit apertures.

Abstract: The present embodiment improves the durability of an afterburner and yet suppresses a reduction in the engine efficiency of an aircraft engine. A ring-shaped cooling channel through which cooling air flows is formed between the outer peripheral surface of a liner and the inner peripheral surface of a rear duct. A plurality of cooling holes for blowing the cooling air along the inner peripheral surface of the liner are formed penetrating the liner. A ring-shaped annulus flame-holding member on the inner peripheral surface of the liner is provided concentrically with respect to a plurality of radial flame-holding members. The inner diameter of the annulus flame-holding member decreases in the downstream direction. The annulus flame-holding member functions as a throttle ring which throttles the flowing of a main flow of a mixed gas inside the liner.

Abstract: The present embodiment sufficiently ensures the ignition stability and the flame-holding property of an afterburner while suppressing a reduction in the efficiency of an aircraft engine. A flame holder is disposed directly downstream of an injection hole of a fuel injector in a liner. The flame holder comprises: a ring-shaped annulus flame-holding member which is provided on the inner circumferential surface of the liner and is capable of propagating a flame in the circumferential direction; and a plurality of radial flame-holding members which are radially disposed inwards of the annulus flame-holding member and are capable of propagating the flame in the radial direction. A guide ring is provided inwards of the radial flame-holding members, and a ring-shaped guide channel that guides a fuel-containing mixed gas in the downstream direction is formed between the outer peripheral surface of the guide ring and the inner peripheral surface of the annulus flame-holding member.

Abstract: A flame-holder device for a reheat channel of a turbojet, the device including an arm in the form of a trough defining a cavity and a heat shield fastened in the cavity of the arm. The flame-holder device further includes a fastener plate including a first leg integrally formed with the fastener plate and a second leg removably mounted on the plate, the arm being fastened to the first and second legs via fastener members.

Abstract: Disclosed is a liquid-gas mixer including a central passageway provided about an injector axis, and first and second gas passageways. The first gas passageway is radially outward of the central passageway and radially inward of the second gas passageway. A turbulator is provided between the first and second gas passageways, and includes a plurality of first disturbance generators and a plurality of second disturbance generators. The first and second disturbance generators are provided about the turbulator in an alternating arrangement.

Abstract: A system includes a gas turbine enclosure. The system also includes a gas turbine engine disposed in the gas turbine enclosure, wherein the gas turbine engine is configured to output an exhaust flow. The system further includes an exhaust driven eductor configured to draw an air flow through and out of the gas turbine engine enclosure using the exhaust flow. The system yet further includes an exhaust stack coupled to the gas turbine enclosure, wherein the exhaust stack is configured to output a mixed flow of the exhaust flow and the air flow. The system still further includes a diffuser plate disposed within the exhaust stack, wherein the diffuser plate is configured to provide a homogenous flow distribution for the mixed flow downstream of the diffuser plate.

Abstract: An afterburner is disclosed for use with a gas turbine engine and, in one form, is structured to receive a bypass air. The afterburner can be situated in a bypass duct and can be a toroidal combustor or a can combustor. In one embodiment, the afterburner includes a combustor arranged to receive bypass air and a plurality of vanes distributed downstream of a turbine of the gas turbine engine. The vanes can include one or more exit apertures through which hot combustion flow from the afterburner combustor can be injected. The exit apertures can be protrusions or slots in some forms. In one embodiment, a cooling passage is arranged around the exit apertures. An upstream vane portion can be positioned to inject fuel to be combusted via interaction with hot flow that is discharged through the exit apertures.

Abstract: A gas turbine engine augmentor includes at least one fluid based augmentor initiator defining a chamber in flow communication with a source of air and a source of fuel. The chamber includes a plurality of ejection openings in flow communication with an exhaust flowpath. The at least one fluid based augmentor initiator is devoid of any exhaust flowpath protrusions thereby minimizing any pressure drops and loss of thrust during dry work phase of operation. The source of fuel is operable for injecting fuel into the chamber such that at least a portion of the fuel flow is ignited at the plurality of ejection openings to produce a plurality of fuel-rich hot jets radially into the exhaust flowpath.

Abstract: One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique reheat system for a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and reheat systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Abstract: An aircraft engine includes an afterburner which has a flame stabilizer. The flame stabilizer maintains a flame generated from a mixed gas of a combustion gas and air. The flame stabilizer includes multiple gutters each configured to generate a flame stabilization area for the flame on its downstream side. Each gutter is formed from: a curved apex section having a stagnation point; and flat plate-shaped side surface sections integrally formed on the respective two sides of the apex section. Each gutter has a V-shaped cross-sectional shape which is opened to the downstream side. At least one through-hole is formed only in each side surface section of each gutter.

Abstract: An augmentor for a gas turbine engine includes an augmentor spray bar with a spray bar outlet, and an augmentor spray bar tip. The spray bar tip includes a tip body, a tip bushing and a tip support strut. The tip body includes a first flow passage extending therethrough between a tip inlet and a tip outlet, wherein the tip inlet is connected to the spray bar outlet. The tip bushing includes a bushing bore. The tip body extends through the bushing bore defining a second flow passage between the tip body and the tip bushing. The tip support strut connects the tip body to the tip bushing across the second flow passage.

Abstract: A flame-holder device for an afterburner of a turbofan, the turbofan including a first inner annular casing and a second inner annular casing, these defining a passage for a primary flow, and an outer annular casing defining with the first inner annular casing a passage for a secondary flow, including at least one arm-support made of metallic material configured to be fixed to the outer casing by an upper plate and at least one flame-holder arm of a one-piece structure of composite material including two mutually attached walls arranged to define a groove whose profile is approximately a V. The walls support, on their upper part situated in the secondary flow, a fixing mechanism for fixing to the arm-support. The upper parts are approximately planar and are positioned facing each other.

Abstract: The tube (60) includes first and second portions (66, 58) that each have mating spherical sections (82, 84) at adjacent ends (72, 74) of each portion (66, 68). The second portion (68) includes a spring sleeve (92) configured to surround an outlet end (72) of the first portion (66). A spring (102) is captured inside the spring sleeve (92) and secured between a retaining clip (100) and a shoulder (88) protruding from the exterior surface of an outlet end (72) of the first portion (66) to maintain fluid seals between the spherical sections (82, 84). The spherical sections (82, 84) permit slidable motion relative to each other thereby facilitating rotational and angular motion between the tube portions (66, 68). One of the two portions (66, 68) may include a telescope mount (110) to permit limited axial motion of the tube (60).

Abstract: A gas turbine or rocket engine hot section includes a first duct case, a second duct case, a plurality of vanes arranged about an axial centerline, and an igniter located with a first of the plurality of vanes. The first of the plurality of vanes extends axially between a leading edge and a flame holder surface at a trailing edge. The flame holder surface extends radially between a first vane end connected to the first duct case and a second vane end connected to the second duct case. The flame holder surface includes a first section that tapers towards the first vane end, and a second section that tapers away from the first section and towards the second vane end.

Abstract: An afterburner chamber includes a burner ring having housed therein an annular fuel strip that is pierced by orifices and that is mounted inside an annular thermal protection tube, and an ignition spark plug penetrating into the inside of the burner ring. The protection tube and the burner ring have orifices for passing air. The chamber also includes a device for ventilating the orifices of the fuel strip that are situated in the immediate vicinity of the spark plug.

Abstract: A flame holder for an afterburner duct of a jet engine, including an arm in the form of a gutter forming a cavity, a shield for protecting the cavity of the arm from heat, and an air supply baffle housed in the cavity, is disclosed. The arm, the protective shield, and the air supply baffle are held together by a one-piece spacer shoe.

Abstract: Integrally formed, multi zone fuel manifolds for gas turbine engines and methods of construction are provided. The fuel manifold includes a plurality of annular channel members positioned adjacent each other, aligned and coupled together at the inner and outer peripheral sides thereof, with each pair of channel members defining a separate fluid conduit or passage within the manifold assembly. The fuel manifold can be constructed of any number of stacked and secured channel members to provide a compact multi zone fuel manifold having the desired number of fluid conduits or passages with a minimum number of sealed joints.

Abstract: A gas turbine reheat system is provided with channel-shaped gutters, with a wall portion spaced from each gutter to form a space therebetween. The space is supplied with cooling air conveniently from the engine by-pass duct.

Abstract: In a by-pass gas turbine engine having a lobed type by-pass air and turbine exhaust mixer unit, the reheat gutters are arranged in the mixer in the turbine exhaust duct. The gutters extend radially and are substantially Y-shaped, each of the two legs of the Y extending radially along the sides of a mixer lobe.

Abstract: A flame holder for an augmentor includes geometry to produce large radial variations in heat release to prevent coupling with modes within the exhaust duct that produce combustor screech. A surface of the flame holder varies axially in a radial direction to produce a radial phase variation in heat released from the flame holder. The radially changing surface produces radially non-uniform and unsteady heat release that does not couple with acoustic modes within the exhaust duct.

Abstract: A flame-holder device for the afterburning of a bypass turbojet, said turbojet comprising first and second internal annular casings, defining a passageway for a main flow, and an external annular casing defining with said first internal annular casing a passageway for a bypass flow, comprising: at least one metal arm support designed to be attached to said external casing by means of a top plate, at least one flame-holder arm, having a trough shape making a cavity, designed to be attached to the arm support, and at least one screen for the heat protection of the arm cavity, the arm support and the heat-protection screen being in one piece.

Abstract: A device for mounting a dividing wall for separating the main air stream and the bypass air in a bypass turbojet engine afterburner is disclosed. The device includes an attachment unit which attaches the upstream end of the wall to a guide vane casing and a supporting unit which supports the downstream end of the wall provided on flame holder arms of the afterburner radially on the inside of the dividing wall.

Abstract: A device attaching an afterburner to a turbojet casing. The device comprises a transverse plate for attachment to the casing and at least two side brackets for supporting two legs of the afterburner. The brackets and the two legs extend longitudinally. The afterburner includes a flame-arrestor arm with the two legs.

Abstract: A turbine engine trailing edge box for an afterburner has corrugations formed to guide and support a spraybar. The trailing edge box is formed of sheet metal to create corrugations in the walls. The spraybar is supported and guided by the corrugations. The corrugations are shaped and sized to hold the spraybar, but to prevent shifting of the spraybar during engine operation. In addition, the corrugations reinforce the walls of the trailing edge box to provide structural support.

Abstract: A flameholder arm for an afterburner of a turbomachine, particularly a jet engine, is disclosed. The arm includes a ventilation duct which is centered and immobilized inside the arm independent of the heatshield. The duct also includes a transverse lip which includes an orifice for centering and guiding a fuel injection harness. The transverse lip prevents the harness from being positioned in any way incorrectly in the arm.

Abstract: An afterburner chamber (10) comprising a burner ring (64) having housed therein an annular fuel strip (28) that is pierced by orifices (32, 34) and that is mounted inside an annular thermal protection tube (56), and an ignition spark plug penetrating into the inside of the burner ring (64), the protection tube (56) and the burner ring (64) having orifices (46, 47) for passing air, the chamber also including means for ventilating the orifices (32, 34) of the fuel strip that are situated in the immediate vicinity of the spark plug.

Abstract: A flameholder for an augmentor rotates about an axis radially positioned with a turbine engine. The flameholder is in a streamlined position when the augmentor is not operating to minimize pressure loss and radar cross-section and a turbulent position when the augmentor is operating. A linkage connects the flameholder to an actuator to control the rotational position of the flameholder on the axis. Bypass airflow from the fan cools the flameholders and linkages. The length of the linkages can be varied from one flameholder to another to create asymmetric heat release within the augmentor to avoid screech.

Abstract: A gas turbine engine fueling system includes a number of spray bars having conduits extending through associated vanes. A number of fuel injector nozzles are along each conduit. Each of the nozzles are positioned to discharge an associated fuel stream from one of the sides of the associated vane. A number of wear members are each mounted relative to the associated one of the nozzles for a range of motion relative thereto. The wear members movably cooperate with the associated vane to accommodate operating deflection and/or tolerances of the spray bars and vanes.

Abstract: The invention relates to a flame holder (10) for an afterburner duct of a jet engine, comprising an arm (100) in the form of a gutter forming a cavity, a shield (110) for protecting the cavity of the arm (100) from heat, and an air supply baffle (120) housed in the cavity. The arm (100), the protective shield (110) and the air supply baffle (120) are held together by a one-piece spacer shoe (150).

Abstract: The turbojet of the invention includes a combustion chamber, a channel for heating the gas stream, where the heating channel includes at least one device for injection of fuel into the gas stream, which includes an open chamber, with a U-shaped section, having at least one wall within which extend fuel-injection means, which inject the fuel in at least one direction. It is characterised by the fact that a cooling jacket is provided in the chamber, alongside the wall forming the base of its U-section, and the fuel injection device includes protection means interposed between the fuel-injection means and the wall, in a fuel-injection direction. Thus, as a result of the invention, the fuel injection device, which is bathed in a very hot environment, is protected against thermal shocks due to the projection of colder fuel onto its walls, and its life expectancy is therefore increased.

Abstract: A gas turbine engine augmentor radial fuel spray bar has a counterswirling spray bar heat shield. Two embodiments of the heat shield include one with a cambered airfoil cross-section and another with a twisted airfoil cross-section and may have varying or constant degree of camber or twist, respectively, along a radial length of the spray bar heat shield. The spray bar may have one or more spray bar fuel tubes within the heat shield, openings in the heat shield, and fuel holes in the tubes operable for injecting fuel through the openings. A gas turbine engine augmentor having a plurality of circumferentially spaced apart radial flameholders may incorporate a plurality of the augmentor radial fuel spray bars with one or more of the augmentor radial fuel spray bars circumferentially disposed between one or more circumferentially adjacent pairs of the radial flameholders.

Abstract: The after-burner device receives a “hot” central primary flow from the turbine of the turbojet and a “cold” peripheral secondary flow, and it has an after-burner ignition zone situated in the secondary flow that reaches the after-burner device. A fraction of the primary flow is brought into the after-burner ignition zone in order to raise the temperature in said zone to a value that is higher than that of the secondary flow so as to encourage after-burner ignition.

Abstract: A fuel shield is configured for use in the afterburner of a turbofan aircraft engine. The shield includes wings obliquely joined together at a nose to conform with the leading edge region of a flameholder vane. A hood is joined to the wings and extends obliquely therefrom to conform with a supporting outer shell of the flameholder.

Abstract: A gas turbine engine augmentor has a centerbody within a gas flowpath from upstream to downstream. The augmentor has upstream and downstream shell sections, a downstream rim of the upstream shell section meeting an upstream rim of the downstream shell section shell section. A plurality of vanes are positioned in the gas flowpath outboard of the centerbody. An augmentor spray bar fuel conduit extends through the centerbody and a first of the vanes to deliver fuel to the centerbody. A seal is mounted to the spray bar and positioned in a recess extending from at least one of the downstream rim of the upstream shell section and upstream rim of the downstream shell section shell section. The seal has a first portion and a second portion engaging the first portion in a backlocked interfitting.

Abstract: A fuel shield is configured for use in the afterburner of a turbofan aircraft engine. The shield includes wings obliquely joined together at a nose, with each of the wings including an offset mounting tab at a proximal end thereof. The wings and tabs are configured to complement a flameholder vane around its leading edge, with the tabs contacting the vane sidewalls to offset the wings outwardly therefrom and form a thermally insulating gap therebetween.

Abstract: A flame holder for an augmentor includes geometry to produce large radial variations in heat release to prevent coupling with modes within the exhaust duct that produce combustor screech. A surface of the flame holder varies axially in a radial direction to produce a radial phase variation in heat released from the flame holder. The radially changing surface produces radially non-uniform and unsteady heat release that does not couple with acoustic modes within the exhaust duct.

Abstract: A fuel shield is configured for use in the afterburner of a turbofan aircraft engine. The shield includes wings obliquely joined together at a nose to conform with the leading edge region of a flameholder vane. A hood is joined to the wings and extends obliquely therefrom to conform with a supporting outer shell of the flameholder.

Abstract: A device for feeding air and fuel to a burner ring in an after-burner chamber, including a plurality of flame-holder arms extending radially about the axis of the chamber from an outer casing, and a burner ring made up of ring sectors mounted substantially end to end on the flame-holder arms, the ring sectors containing fuel feed rail and air manifold boxes.

Abstract: A gas turbine engine augmentor includes an annular trapped vortex cavity pilot having a cavity forward wall, a cavity radially outer wall, and a cavity aft wall, an annular cavity therebetween, and cavity fuel injector tubes operably disposed through the outer wall into the cavity. Circumferentially spaced apart radial flameholders with integral spraybars and/or radial spraybars interdigitated with the radial flameholders radially inwardly into an exhaust flowpath of the augmentor just forward and upstream of the trapped vortex cavity pilot at a radially outer portion of a combustion zone of the exhaust flowpath. An annular trapped dual vortex cavity pilot version is operable for producing trapped dual counter-rotating inner and outer vortices of fuel and air mixtures.

Abstract: A turbofan gas turbine engine augmenter includes a fuel/air swirler disposed between an axially extending bypass flowpath and an axially extending exhaust flowpath. A swirler inlet is axially open to and positioned substantially normal to the bypass flowpath and a swirler outlet is open to and positioned substantially parallel to the exhaust flowpath. A swirl chamber within the fuel/air swirler is between the swirler inlet and the swirler outlet. A swirl axis of the fuel/air swirler extends through the swirler outlet substantially normal to the exhaust flowpath. An air swirler may be centered about the swirl axis within the fuel/air swirler. The air swirler may be a louvered or have a plurality of swirling vanes. The swirler inlet may be radially offset with respect to the swirl axis. An air scoop may lead from the swirler inlet to a rounded swirler housing within which the air swirler is disposed.

Abstract: A method facilitates assembling a gas turbine engine with a flameholder. The method comprises coupling at least one turning vane between a radially outer casing and a radially inner casing to form a flameholder, forming at least two slots that extend substantially radially through the outer and inner casings, coupling at least one fuel injector to the flameholder, and coupling the flameholder within the augmenter.

Abstract: The invention relates to a burner ring for turbofan jet engines. The ring includes an upstream annular envelope which forms a channel which is open axially in the downstream direction, and a manifold for injecting fuel into the channel. The ring includes multiple sectors connected together. Each sector has a fuel inlet which is connected to the fuel injection manifold. Part of the upstream annular envelope is in the core flow. Each sector has a connector which receives the fuel inlet and a ventilation duct which extends through the channel and upstream of the injection manifold. Each sector is fitted with an inlet for a bypass air, which air is then emitted by the ventilation duct to cool the injection manifold. A sector of downstream annular envelope is arranged downstream of the injection manifold to protect this manifold.

Abstract: A support arm (13) is intended to be installed in a radial position in an afterburner device of a bypass turbojet. The device comprises first and second inner annular casings defining a passage for a primary flow and an outer annular casing defining together with the first inner annular casing a passage for a secondary flow. The arm (13) comprises a monobloc structure made of composite material including two integral walls (14, 15), on the one hand, designed to define a groove (16) having a substantially V-shaped profile and, on the other hand, including first end parts (17) joined together and adapted to define a foot (18) and second end parts (19) adapted each to define at least one flange (20, 21) intended to be positively connected to the outer annular casing.

Abstract: A turbine engine augmentor has a centerbody within a gas flowpath. A flameholder is positioned in the flowpath outboard of the centerbody. A burner within the centerbody has an outlet for expelling combustion products of a first pilot fuel. A pilot fuel conduit has an outlet positioned to introduce a supplemental pilot fuel to the expelled combustion products.

Abstract: A gas turbine engine augmentor nozzle has an inlet for connection to an augmentor fuel conduit and an outlet for expelling a spray of fuel. A passageway between the inlet and outlet is at least partially bounded by outlet end surface portions diverging from each other. The nozzle may be used as a replacement for a non-divergent nozzle and may reorient a fuel jet centerline toward radial.

Abstract: A turbine engine has a centerbody within a gas flowpath and a downstream tailcone and a pilot at an upstream end of the tailcone. A flameholder is positioned in the flowpath outboard of the centerbody. The pilot has a first surface diverging in a downstream direction.