Abstract: Bi-directional nacelle ventilation and cooling systems for use with aircraft and related methods are disclosed. An example apparatus includes a passageway to fluidly couple an opening formed in a nacelle of an aircraft engine and an engine compartment of the nacelle. The opening provides an inlet into the engine compartment when passive airflow is available to vent or cool the engine compartment and the opening to provide an outlet from the engine compartment when forced air is needed to vent or cool the engine compartment. A fan is positioned in the passageway to provide the forced air when the passive air is unavailable.

Abstract: Disclosed is a dental handpiece structure including a head having a turbine chamber, a rotor installed in the turbine chamber and having a transmission shaft, and a transmission element composed of a plurality of turbine vanes. A high-pressure air entering into the turbine chamber can push each turbine vane to drive and rotate the rotor, and a breach is upwardly formed between two adjacent turbine vanes. A head cap is covered onto the turbine chamber, and a circular groove is formed at the head cap and between an end surface facing the turbine chamber and the transmission element, and the circular groove and each breach are arranged opposite to each other for collecting the high-pressure air guided out from each breach to enhance the cutting force of the dental hand-piece.

Abstract: A gas turbine engine including a stator vane directing hot combustion gases onto rotor blades is provided. The stator vane includes a platform disposed at the side of the vane radially inward/outward with respect to the axis of rotation of the engine, the platform having a trailing edge portion downstream with respect to the flow of gases past the stator vane. A support and cooling arrangement is included for directing a cooling fluid to an upstream end of a radially inwardly/outwardly facing side of the trailing edge portion of the platform, the arrangement also directing the cooling fluid to flow over the side in a generally axial direction to a downstream end of the side, the cooling fluid cooling the trailing edge portion as it flows over the side, wherein turbulators are included to increase heat transfer from the trailing edge portion as the cooling fluid flows over the side.

Abstract: A rotary engine powered by external heat with a gas that is reactive to temperature sealed inside a casing where a portion, of the sealed in gas is heated and expands within a chamber on one side of the engine and is separated by a partition from another portion of the gas which is cooled and contracts within a chamber on the opposite side of the engine as the heated gas pushes the partition which rotates a rotor while simultaneously the cooled gas pulls the partition and every 180° of rotation the heated gas moves over the rotor to the cooling side of the engine while the cooled gas moves under the rotor through a valve to the heating side of the engine.

Abstract: An evaporative cooling system for combustion gas turbine system has an array of cooling media including first and second cooling media types. The first cooling media type has a first maximum air velocity rating, and the second cooling media type has a second maximum air velocity rating that greater than the first maximum air velocity rating.

Abstract: A turbine shroud segment is metal injection molded (MIM) about an insert having a cooling air cavity covered by an impingement plate. The insert is held in position in an injection mold and then the MIM material is injected in the mold to form the body of the shroud segment about the insert.

Abstract: A heat exchanger for a gas turbine engine includes first and second opposing sides enclosing a cavity. A first set of fins is supported on the first side and arranged outside the cavity. The fins have leading and trailing edges respectively including first and second heights. The first height is less than the second height. In one application, the heat exchanger is arranged in a gas turbine engine. A core is supported relative to a fan case. The core includes a core nacelle and a fan case. A fan duct is provided between the core nacelle and the fan case. A heat exchanger includes fins arranged in a fan duct. The fins are oriented such that the shorter leading edge faces into the airflow.

Abstract: A cooling system of ring segment is provided with: a collision plate that has a plurality of small holes; a cooling space that is enclosed by the collision plate and a main body of the segment body; a first cavity that arranged is the upstream end portion of the segment body in the flow direction of the combustion gas so as to be perpendicular to the axial direction of a rotating shaft; a first cooling passage that communicates from the cooling space to the first cavity; and a second cooling passage that communicates from the first cavity to a fire combustion gas d gas space in the downstream end portion of the segment body in the flow direction of the combustion gas.

Abstract: An internally coolable component for a gas turbine with at least one cooling duct, on the inner surface of which swirl elements are arranged in the form of turbulators which extend transversely with respect to the main flow direction of a coolant is provided. In order to reduce pressure losses for the coolant in the cooling duct, pins with different heights are set up between the turbulators.

Abstract: An exhaust-gas turbocharger (1) having a turbine housing (2) which has a turbine inlet (3) and a turbine outlet (4) and which connects a wastegate duct (5) between the turbine housing inlet (3) and the turbine housing outlet (4), which wastegate duct can be opened and closed via a shut-off element (6). The shut-off element is in the form of a piston (6) which is guided in a longitudinally displaceable manner in an interior space (14) of a guide (7), and the wastegate duct (5) opens into the interior space (14) transversely with respect to the longitudinal central line (L) of the guide (7).

Abstract: An aircraft jet engine with a longitudinal axis, including a wall enclosing a gas flow ejected from a downstream end of the wall in the longitudinal axis, plural ducts being distributed around the periphery of the downstream end of the wall each including a terminal section, including an outlet opening. Each duct can eject a fluid jet through the outlet opening thereof. The ducts are configured to eject the fluid jets essentially parallel to each other, each fluid jet ejected through the corresponding outlet opening forming a lateral angle with the longitudinal axis in a projection view in a plane with the longitudinal axis.

Abstract: A compressor (22; 260; 300; 400) has a housing assembly (140) with a suction port (24), a discharge port (26), and an economizer port (58). An impeller (154) is mounted to be driven in at least a first condition so as to draw a main flow of fluid through the suction port and discharge the fluid from the discharge port. A diffuser (220; 302; 402) has a plurality of diffuser passages (222; 310, 308; 403). Each diffuser passage has an inlet (224) positioned to receive fluid from the impeller and an outlet (220) downstream of the inlet in the first condition. One or more passages (230, 231, 232, 234, 236, 268; 320, 330, 332; 412) are positioned to draw an economizer flow of fluid from the economizer port and deliver the economizer flow downstream of the impeller inlet (212) but upstream of the diffuser passage outlets (226).

Abstract: A cooling system for an electronics chassis includes a plurality of centrifugal blowers arranged to motivate cooling air through the electronics chassis. The centrifugal blowers are arranged in one or more sets, each having blowers oriented with respective inlets in mutual facing relationship. The orientation, positioning, and alignment of the centrifugal blowers facilitates a compact arrangement of the plurality of blowers that achieves increased aerodynamic efficiencies to reduce noise output and energy consumption.

Abstract: A propulsion system and method includes an annular exhaust nozzle about an axis radially outboard of the annular cooling flow nozzle and ejecting an exhaust flow through an annular exhaust nozzle about an axis radially outboard of the annular cooling flow nozzle.

Abstract: An axial compressor comprising liquid drop feed means for feeding liquid drops to an operating fluid of the compressor, a casing for forming a flow path through which the operating fluid flows down and a plurality of stages, each of which is composed of one continuous rotor blade row and one continuous stator vane row, the axial compressor being structured so that the liquid drops evaporate inside the compressor, characterized in that: the casing is provided with a cavity therein, and the cavity is formed by an outer casing and an inner casing which is enclosing a periphery of the rotor blade rows at the plurality of stages and forming internally a flow path of the operating fluid, and a flow path is provided for feeding the operating fluid to the cavity on a downstream side of a region forming the cavity of the inner casing.

Abstract: A differentially pumped mass spectrometer system comprises a mass spectrometer having first and second pressure chambers through which, during use, ions are conveyed along a path. A pump assembly for differentially evacuating the chambers is attached to the mass spectrometer. The pump assembly comprises a housing attached to the mass spectrometer and a cartridge inserted into the housing. The cartridge has a plurality of inlets each for receiving fluid from a respective pressure chamber and a pumping mechanism for differentially pumping fluid from the chambers. The cartridge is inserted into the housing such that the pumping mechanism is inclined relative to the ion path, but with the cartridge protruding into the mass spectrometer to such an extent that at least one of the inlets at least partially protrudes into its respective chamber without crossing the ion path.

Abstract: A turbine exhaust cylinder for an industrial gas turbine engine with an external blower that delivers cooling air to a heat shield jacket surrounding an outer diameter cylinder to provide impingement cooling for the cylinder. Spent impingement cooling air is collected as passed through a space formed between a fairing and a strut to provide cooling to these parts. The cooling air then provides cooling for the inner diameter cylinder before being discharged into the turbine exhaust gas or out from the turbine exhaust cylinder.

Abstract: A manufacturing method includes providing a substrate with an outer surface and at least one interior space and machining the substrate to selectively remove a portion of the substrate and define one or more cooling supply holes therein. Each of the one or more cooling supply holes is in fluid communication with the at least one interior space. The method further includes disposing an open cell porous metallic layer on at least a portion of the substrate. The open cell porous metallic layer is in fluid communication with the one or more cooling supply holes. A coating layer is disposed on the open cell porous metallic layer. The coating layer having formed therein one or more cooling exit holes in fluid communication with the open cell porous metallic layer. The substrate, the one or more cooling supply holes, the open cell porous metallic layer and the cooling exit holes providing a cooling network for a component.

Abstract: A covering device for an integrally bladed rotor base body of a turbomachine for preventing or reducing a cooling air stream that flows from a high-pressure side to a low-pressure side through channels that are formed between adjacent rotor blades, a plurality of cover elements being provided that are individually insertable into the channels and have at least one peripheral sealing surface. Also an integrally bladed rotor base body having a covering device of this kind, a method for producing such an integrally bladed rotor base body, as well as a turbomachine having such an integrally bladed rotor base body.

Abstract: There is provided an exhaust section of a gas turbine and a gas turbine that can cool a strut and the periphery of a bearing without decreasing the efficiency of the gas turbine. A structure of an exhaust section of a gas turbine has: a casing in which a gas path part is formed; a bearing part that rotatably supports moving blades of a turbine section; a strut that extends inwardly from the casing and supports the bearing part; an opening formed in the casing; and a cooling flow channel that extends from the opening toward the bearing part along the strut to guide air at a lower temperature than an exhaust gas flowing in the gas path part to the gas path part and opens into the gas path part at a position downstream of a last-stage moving blade of the turbine section.

Abstract: A blade outer air seal (BOAS) segment has a plurality of cooling passages axially extending through a platform of the BOAS segment. The passages each have an inlet defined in a radially outer surface and an exit defined on a leading edge of the platform. At least one of the passages positioned close to respective circumferential sides of the platform extends linearly from one of the inlets and is skewed away from the axial direction to cool a corner area of the platform.

Abstract: A gas turbine engine, including: a plurality of blades (60) assembled into an annular row of blades and partly defining a hot gas path (26) and a cooling fluid path (24), wherein the cooling fluid path extends from a rotor cavity (22) to the hot gas path; an angel wing assembly (99) disposed on a side (74) of a base (76) of the row of blades; and pumping features (130) distributed about the angel wing assembly configured to impart, at a narrowest gap (42) of the cooling fluid path, motion to a flow of cooling fluid flowing there through. The plurality of pumping features, the angel wing assembly, and the base of the row of blades are effective to produce a helical motion to the flow of cooling fluid as it enters the hot gas path.

Abstract: A gas turbine engine component includes a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe diverging longitudinally from the metering section and a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section.

Abstract: A variable stator vane in a gas turbine engine has a journal provided with a circumferential recess. An aperture extends from the recess to a face of the journal directed towards the aerofoil portion of the vane. In operation of the engine, heated air, for example from a downstream compressor stage of the engine is admitted to the recess and flows through the aperture to, for example, the pressure surface of the aerofoil portion. The aerofoil portion is thereby heated, so preventing ice accretion on the vane.

Abstract: A gas turbine engine includes a diffuser section supplying cooling fluid, a rotatable shaft, shaft cover structure disposed about the rotatable shaft, support structure, and a cooling fluid channel between the shaft cover structure and the support structure. The support structure provides structural support for the shaft cover structure and receives cooling fluid from the diffuser section. One of the diffuser section and the support structure comprises a plurality of apertures through which at least a portion of the cooling fluid passes. The cooling fluid channel is in fluid communication with the apertures in the one of the diffuser section and the support structure for supplying cooling fluid to a blade disc structure located in a turbine section of the engine.

Abstract: A ring segment for a gas turbine engine includes a panel and a cooling system. The cooling system receives cooling fluid from an outer side of the panel for cooling the panel and includes at least one cooling fluid supply passage, at least one serpentine cooling passage, and at least one cooling fluid discharge passage. The cooling fluid supply passage(s) receive the cooling fluid from the outer side of the panel and deliver the cooling fluid to a first cooling fluid chamber within the panel. The serpentine cooling passage(s) receive the cooling fluid from the first cooling fluid chamber, wherein the cooling fluid provides convective cooling to the panel as it passes through the serpentine cooling passage(s). The cooling fluid discharge passage(s) discharge the cooling fluid from the cooling system.

Abstract: A gas turbine having a turbine stage is presented. The gas turbine has a guide vane that can be operated cooled by a cooling medium, and a cooling medium feed device for feeding the cooling medium into the interior of the guide van, which, in the region of its rear edge, has on the pressure side at least one cooling medium outlet opening, through which the cooling medium can flow out of the interior of the guide vane into the main flow. The cooling medium feed device has a mass flow control device for controlling the mass flow through the at least one cooling medium outlet opening, with which the mass flow through the at least one cooling medium outlet opening can be increased during operation of the gas turbine under part-load compared to operation of the gas turbine under full load.

Abstract: A microchannel cooled turbine component includes a first portion of the microchannel cooled turbine component having a substrate surface. Also included is a second portion of the microchannel cooled turbine component comprising a substance that is laser fused on the substrate surface. Further included is at least one microchannel extending along at least one of the first portion and the second portion, the at least one microchannel formed and enclosed upon formation of the second portion.

Abstract: An end-wall component of the mainstream gas annulus of a gas turbine engine has a cooling arrangement including one or more circumferentially extending rows of ballistic cooling holes through which, in use, dilution cooling air is jetted into the mainstream gas to reduce the mainstream gas temperature adjacent the end-wall. A portion of the cooling holes are first cooling holes angled such that the direction of the dilution cooling air jetted therethrough has, on entry into the mainstream gas annulus, a component in one tangential direction. A portion of the cooling holes are second cooling holes angled such that the direction of the dilution cooling air jetted therethrough has, on entry into the mainstream gas annulus, a component in the opposite tangential direction. The first and second cooling holes are arranged such that the cooling air from jets having entry components in opposing tangential directions collide and coalesce.

Abstract: The present application thus provides an active cooling system for a compressor and a turbine of a gas turbine engine. The active cooling system may include an air extraction pipe extending from the compressor to the turbine, a heat exchanger positioned about the air extraction pipe, and an external air cooling system in communication with the air extraction pipe.

Abstract: The invention relates to a compressor unit, in particular, for submarine application, comprising an electric motor. The transported medium for compression, in particular, natural gas for transport, not only frequently contains various aggressive chemical compounds but is also carrier of various condensates which hinder compression and in particular lead to increased wear of the compressor. On assembly aggressive sea water can also enter the compressor unit. The invention provides a solution to the above, wherein the rotation axis is arranged vertically during operation and the housing comprises a drain at the lower axial end. The invention further relates to an assembly method for a compressor unit, in which the compressor unit is filed above water with an incompressible fluid, transported to a submarine operating position, connectors are connected to the inlet and the outlet and the fluid removed from the compressor unit through the drain.

Abstract: An impingement plate adapted to reduce thermally-induced strains and stresses that may damage the plate or its attachment to a second component. The plate includes an interior region having cooling holes, a peripheral wall surrounding the interior region and projecting out of the plane of the interior region, a peripheral flange surrounding the peripheral wall and lying in a plane spaced apart from the plane of the interior region, and one or more through-thickness rib. One such rib may be disposed in the interior region, project away from and out of the plane of the interior region, and linearly extend across the interior region. Alternatively or in addition, one such rib may be disposed between the peripheral wall and flange and project out of the plane of the flange.

Abstract: A turbine engine includes a turbine, a compressor for compressing air and a combustor for receiving the compressed air through an inlet passage and operable to burn fuel therewith to deliver hot exhaust gas to the turbine. Also included is a wheel space defined proximate to the combustor. Further included is a cooling air passage extending between the compressor and the wheel space. Yet further included is a valve assembly having a valve member disposed in the cooling air passage and operable to admit a cooling air to the wheel space in response to a condition therein.

Abstract: A turbine engine including an intermediate space defined between outer and inner portions of the turbine engine. A flow energizer is provided including a flow body located within the intermediate space and including an inlet port, an outlet port and a flow passage extending within the flow body between the inlet and outlet ports. The inlet port receives a flow of a first medium located within the intermediate space and the flow body injects an energizing flow of a second medium to a portion of the first medium within the flow body to create an energized flow of a mixed medium from the outlet portion, the energized flow of mixed medium creates a flow of the first medium adjacent to the flow body within the intermediate space.

Abstract: A rotor disc 2 is provided with blade receiving recesses 40 at its outer periphery. The disc 2 has an internal cavity 34 for conveying cooling air to blades 4 retained in the recesses 40. Each blade-receiving recess 40 intersects the cavity 34 to provide communication between the cavity 34 and the recesses 40. The disc 2 may be made from separate disc components 12, 14 which define the cavity 34 between them.

Abstract: An exhaust fan chamber is defined with a belt cover and a fan cover provided over the belt cover, and an exhaust fan is accommodated in the exhaust fan chamber. The fan cover has an air intake port for sucking air within an engine cover into the exhaust fan chamber and an exhaust outlet port for discharging air within the exhaust fan chamber to outside of the exhaust fan chamber. The exhaust outlet port is in communication with an exhaust opening formed in an upper portion of the engine cover so that the air discharged through the exhaust outlet port is discharged to outside of the engine cover through the exhaust opening.

Abstract: A device for directing gas impingement to an inner casing of a gas turbine may include a plate configured for attachment to the outer surface of the inner casing. The plate has a first surface opposing the inner casing when the plate is attached to an area of the inner casing and a second surface opposite the first surface. The plate defines a plurality of holes through the plate from the first surface to the second surface. The holes are arranged with a predetermined non-uniform distribution in the plate corresponding to a desired preferential impingement pattern for providing non-uniform heat transfer from the area during operation of the gas turbine so as to control temperature of the inner casing across the area. Various options and modifications are possible. Related gas turbine assemblies are also disclosed.

Abstract: Thermal control is provided for a gas turbine casing by supplying thermal control gas from a compressor to a space between an outer casing and an inner casing, and transferring the thermal control gas from the space through the opening in the inner casing via a plurality of holes defined through a plate attached to an outer surface of the inner casing. The holes are arranged with a predetermined non-uniform distribution corresponding to a desired preferential impingement pattern for providing non-uniform heat transfer. A gas turbine thermal control assembly includes structure providing preferential heat transfer from the inner casing during operation of the gas turbine via a thermal control gas flow path from radially outside of the inner casing into the interior of the gas turbine.

Abstract: A cooling arrangement for a gas turbine engine. The cooling arrangement comprises a discharge channel for air flow from a compressor, a first cooling channel and at least one aperture providing communication between the flow of air through the discharge channel and the first cooling channel. A restrictor device in the aperture regulates the flow of air between the discharge channel and the first cooling channel. The restrictor device deforms to vary air flowing through the aperture in response to a physical condition of the engine. This physical condition of the engine may be that of the temperature of air flowing through the discharge channel, the restrictor device responding to regulate the flow of air based on that temperature. The restrictor device may be a two-way shape memory alloy.

Abstract: A stator vane for an industrial turbine, the vane includes a serpentine flow cooling circuit for cooling of the airfoil, and two separate purge air channels to supply air to the rim cavities. The purge channels are formed as separate channels from the serpentine flow channels so that the purge air is not heated by the hot metal and has smooth surfaces so that pressure losses is minimal.

Abstract: A turbine housing (430) includes a support housing (440) with an annular shape. The support housing (432) includes an outer surface (433) and an inner surface (434). The inner surface (434) is located radially inward from the outer surface (433). A plurality of cooling holes (431) extols through the support housing(440) from the outer surface (433) to the inner surface (434). The cooling holes (431) include a stadium shape. The stadium shape includes a rectangle with a top and bottom of a first length and ends capped with semicircles with a radius of a second length. The cooling holes (431) are disposed at an angle (98) from thirty to sixty degrees. The angle (98) is between a line (96) projected to the outer surface (433) from an axis of the turbine housing (430) and a line (97) extending through a center of each cooling hole (431) parallel to the first length.

Abstract: An information handling system cooling fan has a vibration filter that tunes out predetermined fan vibration frequencies to reduce fan vibration transfer from the cooling fan to an information handling system chassis in a targeted frequency range. For example, a leaf spring and damper are biased between the cooling fan and a fan bay support so that stiffness of the spring attenuates selected vibration frequency ranges of the cooling fan bay subsystem, such as frequency ranges where the vibrations have the greatest amplitude or frequency ranges where the vibrations tend to damage other components of the information handling system. In one embodiment, a vibration source generates out-of-phase vibrations that cancel cooling fan vibrations to further reduce overall system vibration output.

Abstract: A gas turbine engine includes a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe diverging longitudinally and laterally from the metering section, a second lobe diverging longitudinally and laterally from the metering section, an upstream end located at the outlet, a trailing edge located at the outlet opposite the upstream end and generally opposite first and second sidewalls. Each sidewall has an edge extending along the outlet between the upstream end and the trailing edge. Each edge diverges laterally from the upstream end and converges laterally before reaching the trailing edge.

Abstract: A low-pressure steam turbine includes an inner casing, an outer casing arranged outside the inner casing so as to cover the inner casing, a heat carrier heating channel between the inner casing and the outer casing so that a heat carrier flows therethrough, a heat carrier inlet passage for introducing the heat carrier into the heat carrier heating channel, and a heat carrier chamber in the inside of at least one of stationary blades to receive the heat carrier that has passed through the heat carrier heating channel. The at least one of stationary blades is heated by the heat carrier which has been heated by passing through the heat carrier heating channel.

Abstract: A turbine ring sector comprising a wall presenting an inside face and an outside face, in which the inside face defines an axially oriented airflow passage through which gas flows; and a multiply-perforated metal sheet situated on the side of the wall that is opposite from the airflow passage is disclosed. The multiply-perforated metal sheet has a bottom, and the space between the outside face of the wall and the bottom of the multiply-perforated metal sheet defines a gap. The variation in the value of the gap over the entire axial extent thereof decreases from a first maximum value towards a minimum value and then increases from the minimum value to a second maximum value.

Abstract: A ring segment for a gas turbine engine includes an outer panel defining a structural body for the ring segment. An outer side of an inner panel is attached to an inner side of the outer panel at an interface, and an inner side of the inner panel defines a portion of a hot gas path through the gas turbine engine. An outer side of the outer panel, opposite from the interface, is in communication with a source of cooling air. A plurality of impingement holes extend through the outer panel from the outer side to the inner side of the outer panel for directing impingement air to the outer side of the inner panel. The outer and inner panels define a plurality of flow channels at the interface for effecting convective cooling of the outer panel along the flow channels between the outer and inner panels.

Abstract: A gas turbine engine turbine diaphragm (460) includes an inner cylindrical portion (461), a mounting portion (463), and a disk portion (462). The mounting portion (463) is located radially outward from the inner cylindrical portion (461). The disk portion (462) extends radially between the inner cylindrical portion (461) and the mounting portion (463). The disk portion (462) includes a plurality of angled holes (464). Each angled hole (464) follows a vector which is angled in at least one plane. A component of the vector is located on a plane perpendicular to a radial extending from an axis of the diaphragm (460). The component of the vector is angled relative to an axial direction of the diaphragm (460).

Abstract: A system for cooling a generator mounted in the tail-cone of an engine. The system comprises a fairing, which receives through an inlet thereof air from a bypass duct and directs the bypass air towards a cavity of the tail-cone for cooling the generator. The bypass air is then expelled through an outlet of a support strut positioned in fluid communication with the tail-cone cavity. The fairing inlet and the strut outlet are both positioned in a plane substantially perpendicular to a longitudinal plane of the engine. In this manner, circulation of the bypass air through the fairing, the tail-cone cavity, and the strut may be achieved. The bypass air directed through the fairing further enables cooling of service lines accommodated in the fairing. A lobe mixer is further used to direct the fairing and shield the latter from core exhaust.

Abstract: Exhaust apparatus for exhausting “dirty” exhaust gases accept a core flow of such exhaust gases and combine that with an annularly-surrounding “rooftop” flow of ambient air for diluting the exhaust gases as well as expelling the diluted flow in a forcibly expelled plume in order to ensure that the “effective” expulsion distance of the expelled diluted flow is at least the physical length of the exhaust apparatus plus the gains gotten from efflux velocity and flowrate.

Abstract: A turbine wheel for a continuous-flow machine includes a centrally arranged hub with circumferentially arranged blades and a pressure equalizing channel. The hub is configured to enclose a cavity, and a resulting principle axis of inertia of the turbine wheel coincides with an axis of rotation of the turbine wheel. The pressure equalizing channel is configured to fluidically connect the cavity to at least one axial end face of the hub and to an environment of the turbine wheel. A diameter of the pressure equalizing channel is smaller than a diameter of the cavity. In one embodiment, the turbine wheel is for an exhaust gas turbocharger.