Abstract: A scramjet engine includes first and second flow path forming members and first and second fuel injection devices. A flow path formed between the first and second flow path forming members includes a turbulence forming region where compressed air is introduced and a combustion region located downstream thereof. The second flow path forming member is formed with a protrusion in the turbulence formation region. The first fuel injection device is configured to inject fuel into the compressed air via a first fuel nozzle. The second flow path forming member is formed with a cavity located in the combustion region. The second fuel injection device is configured to inject fuel into the compressed air via a second fuel nozzle. The cavity is provided with an inclined surface connected to a bottom surface. An inclination of the inclined surface is adjusted so that a shock wave is generated in the combustion region.

Abstract: A mixing tube with multiple shapes is provided, allowing additional injection of gas in order to keep the flow detached from the second shape in an ascent phase and to bring about, in a descent phase, a controlled detachment as a result of the change of slope between the two shapes. A propulsion nozzle for an engine of a spacecraft or aircraft is provided including such a mixing tube and a method for controlling the speed transition of the propulsion gases in such a nozzle in accordance with the altitude. Also, a method is provided for vectorising the thrust in such a nozzle by radial and asymmetrical injection of gas and a control method which prevents re-attachment of the jet to the second shape of such a propulsion nozzle for an engine of a spacecraft when it is in the take-off or landing phase.

Abstract: A Brayton cycle engine including a longitudinal wall extended along a lengthwise direction. The longitudinal wall defines a gas flowpath of the engine. An inner wall assembly is extended from the longitudinal wall into the gas flowpath. The inner wall assembly defines a detonation combustion region in the gas flowpath upstream of the inner wall assembly.

Abstract: A Brayton cycle engine and method for operation. The engine includes an inner wall assembly and an upstream wall assembly each extended from a longitudinal wall into a gas flowpath. An actuator adjusts a depth of the detonation combustion region into the gas flowpath between the inner wall assembly and the upstream wall assembly. The engine flows an oxidizer through the gas flowpath and the inner wall captures a portion of the oxidizer. The engine further adjusts the captured flow of oxidizer via the upstream wall and flows a first flow of fuel to the captured flow of oxidizer to produce rotating detonation gases. The engine flows the detonation gases downstream and to mix with the flow of oxidizer, and flows and burns a second flow of fuel to the detonation gases/oxidizer mixture to produce thrust.

Abstract: A Brayton cycle engine including an inner wall assembly defining a detonation combustion region upstream thereof extended from a longitudinal wall into a gas flowpath. An actuator adjusts a depth of the detonation combustion region into the gas flowpath. A method for operating the engine includes flowing an oxidizer through the gas flowpath; capturing a portion of the flow of oxidizer via the inner wall; flowing a first flow of fuel to the captured flow of oxidizer; producing a rotating detonation gases via a mixture of the first flow of fuel and the captured flow of oxidizer; flowing at least a portion of the detonation gases downstream to mix with the flow of oxidizer; flowing a second flow of fuel to the mixture of detonation gases and oxidizer; and burning the mixture of the second flow of fuel and the detonation gases/oxidizer mixture.

Abstract: A scramjet engine has a first passage forming member and a second passage forming member. A passage is formed between a first surface and a second surface. The passage has an upstream zone, a combustion zone and a downstream zone. A cavity of a concave shape is provided on the first surface in the combustion zone. The first passage forming member has a convex section located in the upstream zone, a first fuel injection section configured to inject fuel into the passage from a first fuel nozzle provided for the convex section, and a second fuel injection section configured to inject fuel to the cavity. The second passage forming member has a third fuel injection section configured to inject fuel to a direction toward the first surface from the second surface in the passage through a second fuel nozzle provided in the downstream zone.

Abstract: An active flow control system for generating pitch control moments for an aircraft during flight. The system includes a nozzle disposed proximate the aft end of the aircraft. The nozzle is configured to discharge a gas stream in the aftward direction. A pressurized air system includes a pressurized air source and one or more injectors configured to selectively inject pressurized air into the nozzle to influence the path of the gas stream. Based upon which injectors are injecting pressurized air into the nozzle, the gas stream exits the nozzle generating no pitch control moment, generating a pitch down control moment or generating a pitch up control moment.

Abstract: A gas turbine engine has an engine core and an annular by-pass duct, within a surrounding nacelle. An exhaust nozzle of the nacelle includes a selectively deployable noise-reduction section on an inner surface thereof. The noise-reduction section includes an inflatable envelope comprising a fixed outer wall and a displaceable inner wall. At least projections on the inner wall are inwardly displaced when the envelope is pressurized and retracted radially outwardly when the envelope is de-pressurized. When the envelope is pressurized, the inner wall includes portions that project to form a rough surface on the inner surface of the nacelle exhaust nozzle. This thickens the boundary layer and reduces the speed of the gas flow at the outer radius of the nozzle, thus reducing the differential velocity with the ambient air, which reduces the gradient throughout the shear layer thus reducing the noise level of the engine.

Abstract: A noise reducing device in a jet engine that has a cylindrical casing, a cylindrical partition wall that is inserted in the casing while protruding partially from a trailing edge of the casing, and a compressor that compresses air that is taken into the cylindrical partition wall, with the inside of the cylindrical partition wall serving as a duct in which a core stream of high-speed air flows, and the space between the cylindrical partition wall and the casing serving as a duct in which a bypass stream of low-speed air flows, and being connected with a wing of an airplane by a pylon that has a projection portion that extends beyond the casing to the downstream of the core stream and the bypass stream, the noise reducing device includes a nozzle, disposed at the pylon on the downstream of the core stream, that injects a fluid toward a noise generation source that is produced from the mutual approach of an ambient air stream that is produced outside of the bypass stream and the core stream.

Abstract: The thrust vectoring apparatus comprises: a housing defining a primary outlet for emitting the primary jet; Coanda surfaces extending from opposing regions of said housing, and radially spaced from the primary outlet such that a step is defined between each Coanda surface and the primary outlet; ducts leading from a fluid source to secondary outlets; and flow control means operable to control the mass flow through the secondary outlets. When the jet engine operates to exhaust a primary jet through the primary outlet, low pressure regions are formed in the vicinity of the steps. Each secondary outlet is located adjacent one of the Coanda surfaces so as to emit a secondary flow into a low pressure region. On activation of the secondary flow by the flow control means, the primary jet is entrained by the Coanda surface opposing the Coanda surface adjacent said the secondary outlet from which the secondary flow has been emitted.

Abstract: In one embodiment, a nozzle of a gas turbine engine may be provided having a coanda injector and a fluidic injector which operate together to provide for a change in exhaust flow direction. The fluidic injector may be coincident with or downstream of the coanda injector and both may be used in high pressure ratio operations of the nozzle. The fluidic injector may be positioned opposite the coanda injector and, when activated, may provide for a region of separated flow on the same side of the nozzle as the fluidic injector. The coanda injector may provide additional momentum to an exhaust flow flowing through the nozzle and may encourage the flow to stay attached on the coanda injector side of the nozzle.

Abstract: A regulator system includes a slidable ramp intermediate a secondary flow path and a primary flow path of a gas turbine engine nozzle section.

Abstract: An improved bypass turbojet engine includes a particularly shaped hot flow nozzle arranged about a particularly shaped type of boss. The boss is bulbous in shape and is connected by a first connecting surface to a turbine casing. The hot flow nozzle is formed as a single skin that extends in a rearward direction from an engine cowl, with the hot flow nozzle being connected to the engine cowl along a second connecting surface. The hot flow nozzle and the bulbous boss are positioned relative to one another to delimit a nominal nozzle throat section and a nominal outlet section for hot flow from a hot flow generator. The hot flow nozzle single skin is also approximately bulbous in shape and widens from the second connecting surface in a direction toward the engine cowl's external wall.

Abstract: A valve system intermediate a secondary flow path and a primary flow path to selectively communicate secondary airflow into the primary gas flow path and control airflow injected from a higher pressure plenum into a lower pressure flowpath.

Abstract: A gas turbine engine is disclosed which comprises a bypass duct, a core engine, and a fluid mixing arrangement. The fluid mixing arrangement is configured to mix a bypass flow of fluid and a secondary flow of fluid, the secondary flow of fluid being drawn from the core engine. The arrangement comprises a flow-duct terminating with an outlet and being arranged to direct said secondary flow through the outlet and into the bypass flow. The arrangement is characterized by the provision of a delta-shaped wing in the region of the outlet, said wing extending at least partially across the duct and being configured to generate lift from said secondary flow effective to produce at least one trailing vortex extending into said bypass flow. The fluid mixing arrangement can be used as a ventilation arrangement or as part of a bleed valve arrangement in the gas turbine engine.

Abstract: In one aspect, the present disclosure is directed an apparatus configured to diffuse a flow of bleed air. The apparatus having an inlet collar configured to receive the flow of bleed air in a direction substantially along a longitudinal axis of the apparatus. The apparatus further having an end wall longitudinally spaced apart from the inlet collar and configured to block the flow of bleed air in a direction substantially along the longitudinal axis. The apparatus also having a first diffuser wall spaced concentrically relative to a second diffuser wall, each of the first and second diffuser walls positioned between the inlet collar and the end wall and including a plurality of perforations configured to permit the flow of bleed air to exit the apparatus at an angle relative to the longitudinal axis.

Abstract: A by-pass turbojet including a thrust reverser is disclosed. The thrust reverse includes a deflector device which deflects all or part of the primary stream gas in such a manner that the deflected primary stream gas encounters the secondary stream, thereby reducing the injection speed of the secondary stream in an aft direction, and thus generating the thrust-reversal effect. The secondary stream escapes from the aft end of the nacelle. In the thrust-reversal position, the deflector device is inscribed radially substantially within the section of the primary nozzle cowl at the aft end of the nacelle. Such a thrust reverser is particularly simple in design, inexpensive, and makes it possible to avoid having moving parts present on the outside portion of the nacelle, thus simplifying the design of the turbojet.

Abstract: An air bleed device for cooling components in a turbine engine, including an annular conduit having a substantially rectangular cross-section formed in a housing and having a radially internal wall swept by an airflow is disclosed. The device includes an air inlet orifice, and a flap valve for controlling the airflow entering through the orifice, formed by a plate borne by a maneuvering member mobile in translation parallel to the axis of the orifice between a position in which the plate closes off the orifice and a position in which the plate opens the orifice.

Abstract: Air bleed device for cooling components in a turbine engine, including an annular conduit (18) having a radially internal portion swept by an airflow (22) and which comprises at least one air inlet orifice (20) formed in an upstream radial wall (38) of the conduit (18), a flap valve (40) for controlling the airflow entering through the orifice (20) and of which the flapper (41) includes a plate (42) applied against the wall (38) and capable of being moved by sliding on said wall (38) by a maneuvering member (50) mobile in translation parallel to the wall (38) between a position in which the plate (42) closes off the orifice (20) and a position in which the plate (42) opens said orifice.

Abstract: A valve system intermediate a secondary flow path and a primary flow path to selectively communicate secondary airflow into the primary gas flow path and control airflow injected from a higher pressure plenum into a lower pressure flowpath.

Abstract: Device (10) for the tapping of air for the cooling of flaps of a turbojet nozzle, comprising an annular duct (18) having a radially internal wall (22) swept by a stream of air (24) and which comprises at least one air inlet orifice (20), a flap valve (40) for controlling the flow rate of air entering through the orifice (20), formed of an elastically deformable metal plate (42) of which a downstream end is fixed on an edge of the orifice (20), and of which an upstream end can be displaced by a manoeuvring member (46) mobile in translation parallel to the axis (52) of the orifice between a position where the plate (42) seals this orifice and a position where the plate (42) opens this orifice.

Abstract: An aircraft jet engine including a wall centered about a longitudinal axis and surrounding a stream of gas ejected at a downstream end of the wall in the direction of the axis, primary ducts distributed at the periphery of the downstream end of the wall configured on command each to eject a jet of primary fluid to interact with the ejected stream of gas, the primary ducts of each pair converging towards one another near the downstream end of the wall so that the primary jets ejected form two sides of a triangle that meet at the vertex thereof in a view projected onto a plane perpendicular to a transverse plane, and secondary ducts each associated with each pair of primary ducts and configured on command to eject a jet of secondary fluid directed into the triangle formed by the primary jets.

Abstract: A vectoring nozzle with external actuation generates thrust vectoring by applying mechanical or fluidic actuation, or both, on the nozzle deck, external sidewalls, and/or air vehicle aft body to produce changes in the aft body flowfield and/or exhaust plume. An external mechanical sidewall may be integrated into a nozzle deck or side walls without the need for engine bleed to supply fluid injectors. An external fluidic vectoring system uses injectors or plasma devices located aft of the nozzle exit to vector the exhaust plume with no external moving parts. Elements of both mechanical and fluidic systems may be combined for a given application.

Abstract: A method of using one or more microjets to create and/or control oblique shock waves. The introduction of microjet flow into a supersonic stream creates an oblique shock wave. This wave can be strengthened—by increasing microjet flow rate or the use of many microjets in an array—in order to form an oblique shock. Such an oblique shock can be used to decelerate flow in a jet aircraft engine inlet in a controlled fashion, thus increasing pressure recovery and engine efficiency while reducing flow instability. Adjusting the pressure ratio across the microjet actually alters the angle of the oblique shock. Thus, the use of microjets allows decelerating shock waves in an inlet engine to be properly positioned and controlled. Microjet arrays can also be used to ameliorate shock waves created by external aircraft surfaces, such as sensor pods and weapons.

Abstract: A nozzle assembly for a gas turbine engine includes a door axially slidable relative to a passage in communication with a secondary flow path to regulate the secondary flow though the passage.

Abstract: An InfraRed Suppression System (IRSS) includes a double-walled exhaust duct and double walled septum. The double walled structure provides a flow path for secondary (non-exhaust gas) cooling air flow to slots along the surfaces exposed to exhaust gases. Exhaust gas flow past these slots draws cooling air across the duct and septum surfaces. This flow cools the surfaces close to the slots and mixes with exhaust gasses downstream to cool the exhaust plume. The septum in the shape of a helix rotating 360° fully blocks line-of-sight to the turbine exit from the aft of the aircraft. The exhaust duct is intended to be shrouded by an aerodynamic fairing which provides an aerodynamic contour which minimizes aerodynamic impact on the aircraft.

Abstract: On a gas-turbine engine, in particular an aircraft engine, with a fan, a fan casing and a fan duct as well as with high-pressure and low-pressure turbines arranged behind each other in flow direction in the casing of the engine, the auxiliaries connected are to be operated with only a small increase in engine power. For this purpose in the flow direction of the airflow (20) exiting from the high-pressure turbine (15) a bleeding mechanism is provided for bleeding radially into the fan duct (4) at least part of the airflow (20) leaving the high-pressure turbine (15) on the upstream side of the low-pressure turbine (16).

Abstract: An outlet nozzle for a jet engine has a gas intake and a gas outlet, at least one opening being provided through the nozzle boundary wall for the injection of a fluid into the nozzle in a direction perpendicular to one plane, for the purpose of controlling the direction of a gas flow from the jet engine so that vectoring occurs substantially tangentially in this plane. A first opening configuration is provided through the nozzle boundary wall on a first side of the nozzle center line and a second opening configuration is provided through the nozzle boundary wall on a second side of the nozzle center line. At least one of said first and second opening configurations has a substantial extension in the axial direction of the nozzle, in order to permit selective injection of the fluid through said opening configuration to a varying extent in the axial direction of the nozzle.

Abstract: A nozzle assembly for a turbine engine has a convergent portion for converging fluid flow from a turbine engine core. A divergent portion for diverging fluid flow from the turbine engine core is in fluid communication with the convergent portion. A conduit is in communication with a bypass fluid flow. The conduit has a thrust vectoring port located near the divergent portion.

Abstract: An apparatus for reducing heating effects of an exhaust plume of a jet engine on an impinged surface includes fluid injectors disposed adjacent and aimed into an exhaust plume zone that's to be occupied by an exhaust plume when the engine is running. A flow generator transmits fluid flow into such an exhaust plume through the injectors. Each injector emits fluid in at least two divergent directions to increase the cross-sectional area of the exhaust plume by forming fluidic lobes in the exhaust plume.

Abstract: A valve system intermediate a secondary flow path and a primary flow path to selectively communicate secondary airflow into the primary gas flow path and control airflow injected from a higher pressure plenum into a lower pressure flowpath.

Abstract: A method and system for thrust vectoring a primary fluid flow from an exhaust nozzle of a jet engine that significantly increases the non-axial force able to be generated by a flight control surface associated with the nozzle. In one implementation the method involves placing a flight control element having a movable portion adjacent a downstream edge of the nozzle. A secondary fluid flow is created adjacent a surface of the flight control element that influences a boundary layer of the primary fluid flow over the flight control element. This causes the primary fluid flow to generate a force that is directed non-parallel (i.e., non-axial) to a longitudinal axis of the nozzle. In one specific implementation a plurality of slots are formed in the flight control surface, and the flight control surface is formed by an airfoil. In another implementation the flight control surface is formed on an interior wall of the nozzle at a downstream edge of the nozzle.

Abstract: A valve system intermediate a secondary flow path and a primary flow path to selectively communicate secondary airflow into the primary gas flow path and control airflow injected from a higher pressure plenum into a lower pressure flowpath.

Abstract: A jet engine noise suppression system uses a plurality of microjets that are located on a frame downstream of the exhaust port of a jet engine. Each microjet issues a microjet flow, either liquid or gas, directed at the main jet flow issued by the engine. The mass of the combined microjet flow is substantially smaller than the mass of the jet engine flow. The angle of at which each microjet issues its flow may be variable. The microjets may be located on a pop-up head of a blast deflector or on a upstanding rack.

Abstract: A rocket nozzle includes a dual-bell nozzle and a gas introducing section configured to introduce gas into space surrounded by the dual-bell nozzle. Combustion gas flows in the space. The dual-bell nozzle includes a first stage nozzle bell-shaped and surrounding an upstream portion of the space, and a second stage nozzle bell-shaped and surrounding a downstream portion of the space. The dual-bell nozzle has an inflection point between the first stage nozzle and the second stage nozzle. The gas introducing section includes a gas inlet provided to an inner wall surface of the first stage nozzle. The gas is introduced into the space from the gas inlet.

Abstract: Systems and methods for passively directing aircraft engine nozzle flow are disclosed. One system includes an aircraft nozzle attachable to an aircraft turbofan engine, with the nozzle including a first flow path wall bounding a first flow path and being positioned to receive engine exhaust products, and a second flow path wall bounding a second flow path and being positioned to receive engine bypass air. The first flow path wall is positioned between the first and second flow paths, and the second flow path wall is positioned between the second flow path and an ambient air flow path. Multiple flow passages can be positioned in at least one of the first and second flow path walls to passively direct gas from a corresponding flow path within the flow path wall through the flow path wall to a corresponding flow path external to the flow path wall.

Abstract: In an air mixing arrangement wherein a primary fluid is introduced through an opening in a wall to be mixed with a secondary fluid flowing along the wall surface, the opening is airfoil shaped with its leading edge being orientated at an attack angle with respect to the secondary fluid flow stream so as to thereby enhance the penetration and dispersion of the primary fluid stream into the secondary fluid stream. The airfoil shaped opening is selectively positioned such that its angle of attack provides the desired lift to optimize the mixing of the two streams for the particular application. In one embodiment, a collar is provided around the opening to prevent the secondary fluid from contacting the surface of the wall during certain conditions of operation. Multiple openings maybe used such as the combination of a larger airfoil shaped opening with a smaller airfoil shaped opened positioned downstream thereof, or a round shaped opening placed upstream of an airfoil shaped opening.

Abstract: An exhaust nozzle assembly provides fluidic thrust vectoring of the primary stream to enhance aircraft maneuverability. The exhaust nozzle assembly includes a cooling air passage that supplies cooling air. A vector opening injects cooling airflow into the exhaust passage to the primary stream. A cover plate is movable to partially block the vector openings and adjust the location of injection of the cooling airflow into the exhaust passage.

Abstract: A plurality of hatches are distributed on the rear periphery of an engine cowl. The hatches are opened only when turbojet engine power is greater than cruising power. In an open position, the hatches draw hot air jets passing through an intermediate chamber before ejecting the jets through longitudinally-distributed communication components.

Abstract: A supersonic combustion apparatus including a fixed geometric nozzle having a converging area, throat, and a diverging area, at least one movable combination of a fuel injector and an oxygen injector where the combined fuel injector and the oxygen injector is located within the divergent area of the fixed geometric nozzle, and an exit plane adjacent and downstream to the diverging area. The exit plane Mach speed is varied by heat addition in the diverging area by introduction of a combustible fuel through the fuel injector and oxygen through the oxygen injector and then axially aligning and positioning the combination of the fuel injector and the oxygen injector along the length of the diverging area to obtain a stabilized flame at the exit plane.

Abstract: A non-invasive system, method, and apparatus for control input prediction and state verification of an aircraft's fluidic vectoring exhaust is disclosed. The control system derives a desired vector state, then predicts and sets the fluidic injection input required to produce the desired vector state. A vectored state verification routine is used to determine the resulting vector state for feedback to the control system.

Abstract: Disposed at or toward the trailing edge of one or more nozzles associated with a jet engine are injection ports which can selectively be made to discharge a water stream into a nozzle flow stream for the purpose of increasing turbulence in somewhat of a similar fashion as mechanically disposed chevrons have done in the known art. Unlike mechanically disposed chevrons of the known art, the fluid flow may be secured thereby increasing the engine efficiency. Various flow patterns, water pressures, orifice designs or other factors can be made operative to provide desired performance characteristics.

Abstract: The thrust vectoring apparatus comprises: a housing defining a primary outlet for emitting the primary jet; Coanda surfaces extending from opposing regions of said housing, and radially spaced from the primary outlet such that a step is defined between each Coanda surface and the primary outlet; ducts leading from a fluid source to secondary outlets; and flow control means operable to control the mass flow through the secondary outlets. When the jet engine operates to exhaust a primary jet through the primary outlet, low pressure regions are formed in the vicinity of the steps. Each secondary outlet is located adjacent one of the Coanda surfaces so as to emit a secondary flow into a low pressure region. On activation of the secondary flow by the flow control means, the primary jet is entrained by the Coanda surface opposing the Coanda surface adjacent said the secondary outlet from which the secondary flow has been emitted.

Abstract: The invention relates to a bypass turbomachine comprising a gas turbine engine provided with a fan disposed on a longitudinal axis of the turbomachine, and an annular nacelle surrounding the engine so as to define a cold stream flow channel, the nacelle having an upstream end surrounding the fan of the engine and a stationary downstream end forming a gas exhaust nozzle, the nozzle having a throat section that corresponds to its smallest cross-section. The turbomachine further comprises means for taking in air from the cold stream flow channel upstream from the nozzle throat section, and means for injecting the taken-in air downstream from the nozzle throat section so as to open the throat section artificially.

Abstract: A fluid vectoring nozzle comprises flow vectoring means suitable for selectively producing a fluid dynamic throat in the nozzle in a plane oblique to the axis of the nozzle, such that in operation fluid flow passing through the throat is rotated about an axis parallel to and coincident with the plane of the throat. That is to say, the present invention comprises flow vectoring means suitable for selectively producing a fluid dynamic throat in the nozzle in a plane oblique to the axis of the nozzle, such that, in operation, fluid passing through the fluid dynamic throat is turned towards an angle perpendicular to the plane of the throat. Fluid injection means are provided to inject a control fluid into the nozzle through a perforate region provided in at least one nozzle wall, thereby generating a fluid dynamic restriction which initiates the formation of the fluid dynamic throat.

Abstract: The present invention relates to a thrust orienting nozzle, shaped in such a way as to divide a principal propulsion gas flow coming from at least one gas generator into a first flow and a second flow for an ejection in a first half-nozzle and in a second half-nozzle and comprising at least one of following two piloting means: a means of dividing the principal flow into each of the two half-nozzles, and a means of orienting the thrust vector produced by each of the two half-nozzles. The invention applies in particular to the yaw control of an aircraft without a vertical tail unit.

Abstract: Exhaust nozzle 14 for an engine of a flying craft, comprising a tubular body with two outlet ducts 16, 18 and boxes 26, 28 for injecting gas into the outlet ducts for the purpose of thrust vectoring, these boxes having windows 38 aligned with slots 39 formed in the outlet ducts, and controlled means for adjusting the outflow of gas injected into the outlet ducts.

Abstract: An exhaust nozzle assembly provides fluidic thrust vectoring of the primary stream to enhance aircraft maneuverability. The exhaust nozzle assembly includes a cooling air passage that supplies cooling air. A vector opening injects cooling airflow into the exhaust passage to the primary stream. A cover plate is movable to partially block the vector openings and adjust the location of injection of the cooling airflow into the exhaust passage.

Abstract: An embodiment of the invention is a technique to suppress noise in a jet engine. A noise suppressor includes an exhaust duct to exhaust a first stream from an air stream taken from an inlet. The exhaust duct has an exit end. The air stream is divided into the first stream and a second stream. The second stream has an axis. A stream director attached to vicinity of the exit end directs the first stream to a direction that is away from or sideward relative to the axis.

Abstract: A housing for a turbocharger forms a housing jacket (6, 7) which surrounds a rotor space (15) for receiving a turbine rotor (18). The housing jacket consists of sheet metal at least in part. A connection pipe (4?) provides for communication with an exhaust gas manifold (3, 4) of a combustion motor (20). To reduce heat losses, it is not only the housing jacket (6, 7, 22) which surrounds the rotor space (15) and consists of sheet metal, but at least the connection pipe (4?) for communication with the exhaust gas manifold (3, 4) is shaped of sheet metal and is in heat conductive connection with the sheet metal of the housing jacket (6, 7, 22).