Abstract: A gas turbine engine includes a housing with a duct wall that defines a generally annular and axially elongated hot gas flow path for passage of combustion gas. The engine further includes a contoured shroud mounted within the internal engine cavity and defining a hot gas recirculation pocket for receiving hot gas ingested from the hot gas flow path through the annular space and for recirculating the ingested hot gas back through the annular space to the hot gas flow path. The contoured shroud includes a base wall extending radially inwardly from the duct wall, an inboard wall extending from the base wall in an axial direction toward the rotor, and an end wall extending from the inboard wall in a radially outward direction. The end wall terminates in a circumferentially extending free edge disposed in proximity to the annular space. The end wall defines an opening.

Abstract: The present invention discloses a turbo molecular pump with improved blade structures. The turbo molecular pump comprises a rotor and a stator, wherein the rotor includes five rotor blade assemblies and the stator includes five stator blade assemblies, and wherein the blade number and the blade angle of each rotor blade assembly and stator blade assembly are adjusted to optimization, so as to enhance the pumping speed and the stability of the turbo molecular pump as well as to reduce the difficulty for manufacturing the turbo molecular pump.

Abstract: A fan system with enhanced air flow-static pressure characteristics and reduced fan noise compared to the related art is provided. The number of duct blades of a duct is the same as the number of stationary blades of an axial flow fan located in front of the duct, and the duct blades correspond to the stationary blades respectively. An end surface of a rear end portion of each stationary blade and an end surface of a front portion of a duct blade corresponding to the stationary blade have the same shape, and they align together and contact each other to form one composite stationary blade, with a discharge port of each axial flow fan communicating with an inlet port of a duct housing located behind the axial flow fan.

Abstract: The present invention provides a gas turbine rotor having a first axial compression stage for compressing air in a first direction and a second axial compression stage for compressing the air in a second direction, the second direction being generally opposite the first direction.

Abstract: A two-stage high pressure turbine includes a second stage blade having an airfoil with a profile substantially in accordance with at least an intermediate portion of the Cartesian coordinate values of X, Y and Z set forth in Table 2. The X and Y values are distances, which when smoothly connected by an appropriate continuing curve, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape.

Abstract: A steam turbine is provided having a rotor which rotates around a machine axis, and a stator which concentrically encompasses the rotor with clearance, between which an annular passage is formed which is exposed to throughflow by steam in the axial direction and in which a multiplicity of rotor blades, which are fastened on the rotor, and fixed stator blades are arranged in a plurality of stages, wherein the stator blades of the last stage have a sweep with a sweep angle which changes in sign over the relative blade height.

Abstract: A power plant includes a first turbine having a plurality of turbine stages. The first turbine discharges a first fluid flow. A second turbine is operatively coupled to the first turbine. The second turbine receives the first fluid flow from the first turbine and a second fluid flow. An interface member is mounted between the first and second turbines. The interface member includes a main body portion having an inner surface, an inlet portion and an outlet portion that combine to form a fluid flow path that extends between the first and second turbines. The interface member further includes an input member having an input passage that leads the second fluid into the fluid flow path. The interface member also includes a baffle plate that extends toward the fluid flow path. The baffle plate establishes a flow gap between the input passage and the inlet and outlet portions.

Abstract: A vacuum pump comprises a first pumping section 106, and, downstream therefrom, a second pumping section 108. The pump comprises a first pump inlet 120 through which fluid can enter the pump and pass through both the first and second pumping sections towards a pump outlet, and a second pump inlet 122 through which fluid can enter the pump and pass through only the second pumping section towards the outlet. The second pumping section 108 comprises at least one turbo-molecular pumping stage 109a, 109b and, downstream therefrom, an externally threaded rotor 109c.

Abstract: The invention relates to a compressor for a turbine engine including a casing (4), at least one compressor stage consisting of a stationary blade (2) impeller and a mobile blade (1) impeller positioned upstream from said stationary blade (2) impeller, and cavities (5) made in said casing opposite the through-path of the mobile blades (1), said cavities having a length L2 measured axially and being shifted upstream relative to the blades (1) so as to generate an overlap with a length L1, characterised in that the lengths L1 and L2 are respectively between 35% and 50% and between 80% and 90% of the axial chord Cax measured at the outer end of the blades (1), and in that the cavities (5) do not in communication with one another.

Abstract: A method of designing a multistage turbine for a turbomachine in which each turbine stage comprises a stator row and a rotor row each made up of an annular row of airfoils, wherein, for all of the stator or rotor rows, it consists in simultaneously modifying the shapes of the airfoils of said rows to straighten out the wakes from the trailing edges of said airfoils, then in angularly positioning the rows in such a manner that the wakes from the airfoils of the stator (or rotor) airfoils impact against the leading edges of the stator (or rotor, respectively) airfoils of the rows situated downstream, in order to achieve multistage aerodynamic coupling simultaneously over the turbine as a whole.

Abstract: A serial axial fan unit includes first and second motors with their base portions, i.e., first and second base portions axially facing each other. A motor gap is arranged axially between the first and second base portions. An axial length of the motor gap is preferably in a range from approximately 0.3 mm to approximately 2.0 mm. This configuration reduces transmissions of vibration of each of the first and second motors to the other, thereby reducing vibration interference between the first and second motors.

Abstract: A heat dissipation fan includes a housing, a first rotor, a second rotor, a base and a plurality of static blades. The first rotor has a shaft and a plurality of rotor blades. The second rotor is coupled to the first rotor and has a plurality of rotor blades. The base is disposed in the housing for supporting the first and second rotors. The static blades are disposed between the housing and the base, wherein a rear portion of each static blade extends along an axial line of the heat dissipation fan for improving the working efficiency of the second rotor.

Abstract: A rotor wheel is provided and includes a body having first and second opposing faces and portions recessed from a plane of the first face to define therein an annular groove and a plurality of tributary grooves, the annular groove being receptive of fluid from an external source and formed to direct the fluid to flow along an annular flow path, and the plurality of tributary grooves being receptive of the fluid from the annular groove and respectively formed to direct the fluid to flow sequentially along radial and axial tributary flow paths while preventing inter-tributary groove fluid communication.

Abstract: A pressure activated flow path seal for a steam turbine is disclosed. In one embodiment a gap closure component is located about a rotary component and the stationary component of the steam turbine. A pressure differential activates the gap closure component to seal or reduce the radial clearance of a steam leakage path between the rotary component and the stationary component.

Abstract: A nickel base alloy is provided which includes the following components by weight: Co: 2.75 to 3.25% Cr: 11.5 to 12.5% Mo: 2.75 to 3.25% Al: 3.75 to 4.25% Ti: 4.1 to 4.9% Ta: 1.75 to 2.25% C: 0.006 to 0.04% B: ?0.01% Zr: ?0.01% Hf: ?1.25% Nb: ?1.25% balance Ni.

Abstract: Gas turbine engine systems and related methods involving vane-blade count ratios greater than unity are provided. In this regard, a representative turbine stage for a gas turbine engine includes a first set of rotatable blades operative to be positioned downstream of and adjacent to a first set of vanes, a number of blades of the first set of blades being less than a number of vanes of the first set of vanes.

Abstract: A steam turbine is provided that is configured for post-modification for operation in a carbon capture mode. The turbine includes a turbine rotor, a turbine casing and a plurality of turbine stages. In an initial configuration of the turbine, the turbine rotor and turbine casing are each longer, by respective lengths, than is necessary to accommodate the plurality of turbine stages. The lengths are sufficient to accommodate at least one further turbine stage at an exit of the turbine during the post-modification, such that after modification, the turbine operates with an increased expansion ratio and an increased volumetric flow rate at the exit.

Abstract: An axial fan unit includes two or more impellers accommodated in a housing. Each impeller has a plurality of blades disposed about a center axis and is rotated by an associated motor. The rotation of the impellers creates a flow of air generally flowing along the center axis. A plurality of ribs are disposed between the impellers and connected to the housing. The ribs support an associated motor. At least one of an air-inlet side edge and an air-outlet side edge of each rib is inclined with respect to a direction that is perpendicular or substantially perpendicular to the center axis, such that it gets closer to one of an air-inlet side end and an air-outlet side end of the axial fan unit as it moves away from the center axis.

Abstract: A heating apparatus and a system that utilize friction to generate thermal energy. The heating apparatus comprises a housing unit, a plurality of heating chambers for generating heat, an actuating unit, a shaft; and at least one blade unit. Each heating chamber comprises a stationary disc member, a rotating disc member, and a medium disposed between the stationary disc member and the rotating disc member. The actuating unit drives the rotating disc member in the heating chamber to generate thermal energy by friction among the stationary disc member, the rotating disc member, and the medium. Thermal energy generation is controlled by rotating speed, diameter of the disc member, and contact pressure between the rotating disc member and the stationary disc member. Tube members can also be used in the heating chambers.

Abstract: The present invention is a high efficiency turbine system which can increase a pressure differential between an upstream location and a downstream location. The turbine system includes a propeller attached to a shaft, which can both be located in a shroud. The shroud includes a projection, such as a brim, which protrudes inward and/or outward from the shroud. The projection includes brim units arranged asymmetrically and/or in repeating patterns to generate various vortex and swirl patterns. The brim units can have a different size, shape, width, and/or height than an adjacent brim unit. The brim units can be arranged in a non-parallel manner and can be rotatable. Furthermore, the brim units can form brim groups which can be arranged asymmetrically and/or in repeating patterns to generate various swirl patterns. The turbine system can also be used in a renewable energy system, which can be used to power electronic devices.

Abstract: A gas turbine engine includes a compressor, combustor, and high pressure turbine operatively joined together. A first interstage bleed circuit is joined in flow communication between a first preultimate stage of the compressor and hollow blades in the turbine to provide thereto pressurized primary air. A second interstage bleed circuit is joined in flow communication between a second preultimate stage of the compressor and the turbine blades to provide thereto pressurized secondary air at a lower pressure than the primary air.

Abstract: An axial-flow compressor for a gas turbine engine has a rotor drum (2) in thermally lower loaded first compressor stages (3 to 6) which includes a one-piece ring, or rotor rings (7 to 10) attached to one another. Fiber belts (18, 21) are wound onto these rings close to the rotor blades and include carbon fibers embedded in a high-temperature resistant polymer matrix. As the rotor disks can be dispensed with, since their function will be assumed by the fiber belts, the compressor features low weight, requires limited space only, and, in addition, can be produced cost-effectively.

Abstract: A blower unit for a portable blower, having a drive shaft that is rotatably driven by a drive motor. Disposed in the blower unit is an axial fan that is provided with at least one fan wheel driven by the drive shaft and at least guide wheel fixedly disposed in the housing. To achieve a simple construction and a simple manufacture, the housing has at least two housing sections and is divided approximately parallel to the axis of rotation of the drive shaft.

Abstract: A heat dissipation module includes a housing, and a fan including an outer frame, a first rotor, a base, a second rotor and a driving device. The first rotor includes a shaft, a first hub and a plurality of first rotor blades disposed around the first hub. The base is disposed in the outer frame. The second rotor includes a plurality of second rotor blades, and is coupled to the shaft of the first rotor and disposed next to the first rotor. The driving device is supported by the base for driving the first rotor and the second rotor simultaneously.

Abstract: A counter-rotating axial-flow fan is provided that is capable of increasing effect of cooling a stator. One or more through-holes 83 penetrating support frame bodies (21, 75) in the axial direction are formed in support frame bodies 21 and 75. One or more vent holes 57a for introducing air, sucked from a suction port 19, into a cup-like member 51 are formed in a bottom wall portion 57 of a cup-like member 51 of a first impeller 9. One or more vent holes 111a for discharging air, introduced into an internal space of a second motor 61, to the outside are formed in a bottom wall portion 111 of a cup-like member 105 of a second impeller 63.

Abstract: A method of fabricating a turbine rotor is provided. The method includes fabricating a plurality of substantially cylindrical disks. Fabricating each disk includes fabricating a substantially cylindrical body and extending a bore substantially concentrically through the body. The method also includes coupling at least two of the plurality of disks together to form a rotor having a bore extending axially therethrough.

Abstract: A wind turbine comprises a tubular casing having an inlet opening, an outlet opening, an outer surface generating a pressure decrease, an inner surface presenting a convergent section joined to the inlet opening, a divergent section joined to the outlet opening and to the convergent section by a throat, and a propeller mounted rotating with respect to the tubular casing in proximity to the throat. It is coupled to a first generating machine. It comprises another propeller mounted rotating with respect to the tubular casing, placed upstream from the first propeller in the convergent section.

Abstract: An arrangement for supporting a shaft of a vacuum pump includes a first bearing for supporting the shaft at one of its end, a second bearing for supporting the shaft at another of its ends and formed as a rolling bearing, a holder for receiving the second bearing and including an oscillating base, a fixation member secured in the vacuum pump housing, metallic connection members for connecting the oscillating base with the fixation member and axial springs arranged between the oscillating base and the fixation member and providing for the axial stiffness of the holder which is greater than its radial stiffness thereof and greater than the axial stiffness of the first bearing.

Abstract: A stator vane for a turbo molecular pump has stator vane halves each having inner and outer rim portions and radially arranged stator blades connected integrally between the inner and outer rim portions. The inner rim portion of each stator vane half has a pair of inner rim ends and the outer rim portion of each stator vane half has a pair of outer rim ends. The stator vane halves are disposed in abutment with one another along an abutment line to form an annular body with the inner rim ends of one of the stator vane halves being disposed in confronting relation with the respective inner rim ends of the other of the stator vane halves and with the outer rim ends of the one of the stator vane halves being disposed in confronting relation with the respective outer rim ends of the other of the stator vane halves.

Abstract: A low-pressure turbine of a gas turbine is disclosed. The turbine comprises a number of stages arranged one behind the other in an axial manner in the flow-through direction of the turbine. Each stage is formed from a fixed vane ring having a number of vanes and from a rotating blade ring having a number of blades. Each stage is characterized by a characteristic value vane-to-blade ratio that indicates the ratio of the number of vanes to the number of blades within a stage. One of the stages of the turbine is designed in such a manner that, in the event of noise-critical conditions of the turbine, the characteristic value vane-to-blade ratio of this stage is between a lower cut-off limit for mode k=?1 of the blade-passing frequency (BPF) of said stage and an upper cut-off limit for the mode k=?2 of the blade-passing frequency (BPF) of this stage.

Abstract: A hybrid gas compressor has at least one rotor and shroud which define a compressor gas path extending from an inlet to an outlet. The compressor includes at least two compression stages and one diffusion stage between the inlet and outlet, the compression stages including respective circumferential arrays of blades extending from the rotor and the diffusion stage including a circumferential array of vanes between the compression stages. Blade aerodynamic loadings may be controlled, particularly in the last stage, to provide desired compression characteristics across the compressor. A bleed outlet is optionally located between the inlet and outlet for bleeding from the gas path.

Abstract: A method of assembling a gas turbine engine is provided. The method includes coupling at least one turbine nozzle segment within the gas turbine engine. The at least one turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band that includes an aft flange and a radial inner surface. The method also includes coupling at least one turbine shroud segment downstream from the at least one turbine nozzle segment, wherein the at least one turbine shroud segment includes a leading edge and a radial inner surface, and coupling a cooling fluid source in flow communication with the at least one turbine nozzle segment such that cooling fluid channeled to each turbine nozzle outer band aft flange is directed at an oblique discharge angle towards the leading edge of the at least one turbine shroud segment.

Abstract: A rotor assembly for a gas turbine engine, the rotor assembly comprises at least two rotors defining a cavity therebetween. A first rotor defines a cooling air inlet in its radially inward portion. A second rotor defines a cooling air outlet in its radially outward portion, such that the cooling air passes radially outwardly through the cavity.

Abstract: A turbomolecular pump (12) has a circular intake opening (16) distal of an inlet rotor stage (18). The intake opening (16) comprises at least two opening sections (41, 42, 43) that are separated from each other.

Abstract: A method for flow optimization in multi-stage turbine-type machines in which the inflow of a third of three consecutive blade rings is optimized, the first and the third blade ring having the same number of blades being situated on the same unit, rotor or stator, the second blade ring being situated on the other of the two units, rotor or stator, and an operating state, occurring during a high proportion of the operating time, being selected by ascertaining or predefining the appropriate operating parameters. In this operating state, the maxima of the obstruction, periodically occurring in the area of the outlet edges of the blade profiles of the second blade ring, are deflected onto the inlet edges of the blade profiles of the third blade ring within a predefined tolerance angle; the positions or the geometries of blade profiles of at least one of the three blade rings are modified as needed.

Abstract: A turbine assembly for a gas turbine engine. The turbine assembly includes at least one stator assembly including a radially inner band and at least one stator vane that extends radially outward from the inner band. The stator vane includes an airfoil having a root portion adjacent to the inner band and a tip portion. The airfoil also includes at least one lean directional change that is defined between the root portion and the tip portion. The turbine assembly also includes at least one turbine blade assembly that includes at least one rotor blade. The blade assembly is coupled in flow communication with the stator assembly such that an axial spacing is defined therebetween. The axial spacing defined adjacent to the at least one lean directional change is wider than the axial spacing defined adjacent to the root portion.

Abstract: A method for modifying a multistage compressor (101) involves exchanging the rotor blades of the first compressor rotor blade row (LA1) for modified rotor blades which have an identical blade leaf profile (121) to the original rotor blades and the blade angle (B?11) of which is different from the blade angle (B?10) of the original rotor blades. Furthermore, the blades of at least one further blade row (LAN, LEN) arranged downstream of the second compressor stage are exchanged for modified blades which have an identical blade leaf profile (125, 126) to the original blades and the blade angle (B?N1, B?N1) of which is different from the blade angle (B?NO, B?NO) of the original blades. The method makes it possible to increase the mass flow of a compressor and essentially maintain the stability reserve against stall.

Abstract: A method and apparatus to reduce the axial thrust in rotary machines such as compressors, centrifugal pumps, turbines, etc. includes providing additional peripheral restrictive means (7) attached at the peripheral portion of the disk forming the subdividing means (4) on the side facing the rotating rotor (2). An additional ring element at the periphery of the subdividing means forms additional radial (11) and axial restrictive means (15). Such peripheral restrictive means (7, 11 and 15) function as sealing dams, which combined with the outward flow induced by the rotating impeller, form self-pressurizing hydrodynamic bearings in the axial and radial planes, improving rotordynamic stability. Additionally, a stationary ring element in the center of the cavity forms a seal with the rotor, reducing leakage to suction.

Abstract: A gas turbine engine rotor stack includes one or more longitudinally outwardly concave spacers. Outboard surfaces of the spacers may be in close facing proximity to inboard tips of vane airfoils. The airfoils have dihedral and sweep.

Abstract: A transition duct for routing a gas flow from a combustor to the first stage of a turbine section in a combustion turbine engine has an internal passage from an inlet to an outlet that is offset from the inlet in the longitudinal, radial and tangential directions. The offset outlet and the curved internal passage discharge the gas flow toward the first stage blade array at an angle in the tangential direction relative to the longitudinal direction. This angled discharge can be presented directly to the blades, thus avoiding the need for first stage vanes and the associated costs and complexity.

Abstract: A stator-rotor assembly which includes at least one interface region between the stator and rotor is described. At least one stator or rotor surface in the interface region includes a pattern of concavities. The concavities restrict gas flow through a gap between the stator and the rotor. Various turbomachines which can contain such a stator-rotor assembly are also described. The disclosure also discusses methods to restrict gas flow through gaps in a stator-rotor assembly, utilizing the concavities.

Abstract: A method for cooling a shroud segment of a gas turbine engine includes providing a turbine shroud assembly including a shroud segment having a leading edge defining a forward face. A turbine nozzle is coupled to the turbine shroud assembly such that a gap is defined between an aft face of an outer band of the turbine nozzle and the forward face, wherein a lip formed on the aft face is positioned radially inwardly with respect to the gap and extends substantially axially downstream from the gap. Cooling air is directed into the gap. Cooling air exiting the gap impinges against the lip. Post impingement cooling air is directed at the forward face to facilitate forming a film cooling layer on the shroud segment. The film cooling layer is shielded from combustion gases flowing through the gas turbine engine.

Abstract: A serial axial fan includes first and second axial fans. The first axial fan preferably includes a first motor portion, a first impeller, and a first housing. The second axial fan preferably includes a second motor portion, a second impeller, and a second housing. In the first impeller, an angle defined by a rotational plane of the first impeller with a chord of each blade of the first impeller on an imaginary cylindrical surface centered about a central axis increases as the radius of the imaginary cylindrical surface increases.

Abstract: A turbine (100) of a turbine installation, especially a steam turbine of a steam turbine installation, includes at least one radial or diagonal turbine stage (120) with radial or diagonal inflow and axial outflow, and also at least one axial turbine stage (121-125) with axial inflow and axial outflow. The at least one radial or diagonal turbine stage (120) forms the first stage of the turbine (100) and the at least one axial turbine stage (121-125) is arranged downstream of the radial or diagonal turbine stage (121) as an additional stage of the turbine. The at least one radial or diagonal turbine stage (120) has a higher temperature resistance than the at least one axial turbine stage (121-125). The turbine (100) makes it possible to significantly increase the process temperature of the steam turbine installation, wherein measures for increasing the temperature resistance need only to be adopted for components of the radial or diagonal turbine stage (120).

Abstract: To reduce secondary flow loss and to improved turbine efficiency, a section located radially outward of a border section 28 of a stationary blade 21 is bent in the rotational direction of a rotor. Thus, even if combustion gas leaks from a tip clearance between an end wall of a casing and a tip portion of a rotor blade, and a stagnation line 35 near a tip portion 22 is situated in the side of a back surface 24, because a section located radially outward of the border section 28 is bent in the rotational direction of the rotor, the stagnation line 35 is also situated toward the rotational direction of the rotor. Therefore, the stagnation lines 35 formed at various heights in the heightwise direction of the stationary blade 21 are generally aligned in the rotational direction of the rotor. Thus, fluctuation of pressure distribution in the heightwise direction of the stationary blade 21, of the combustion gas flowing into the stationary blade 21 can be reduced.

Abstract: A method for cooling a shroud segment of a gas turbine engine is provided. The method includes providing a turbine shroud assembly including a shroud segment having an inner surface and a leading edge that is substantially perpendicular to the inner surface, and coupling a turbine nozzle to the turbine shroud segment such that a gap is defined between an aft edge of an outer band of the turbine nozzle and the leading edge. The method also includes directing cooling air into the gap, and directing the cooling air in the gap through at least one cooling hole extending between the leading edge and the inner surface.

Abstract: An axial flow turbine provided with a stage composed of a turbine nozzle and a turbine rotor blade arranged in an axial flow direction. Both end portions of a nozzle blade of the turbine nozzle are supported by a diaphragm inner ring and a diaphragm outer ring, and a flow passage is formed to have its diameter expanded from an upstream stage to a downstream stage. In such axial flow turbine, trailing edges at ends of the nozzle blade supported by the diaphragm inner ring and the diaphragm outer ring are curved as a curvature to an outlet side, and an intermediate portion between the trailing edges is formed to be straight.

Abstract: A method of assembling a gas turbine engine is provided. The method includes coupling at least one turbine nozzle segment within the gas turbine engine. The at least one turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band that includes an aft flange and a radial inner surface.

Abstract: A fan apparatus includes a ducted housing with an axis of rotation including a first section and a second section. The fan apparatus further includes a first plurality of vanes being radially coupled within the first section and a second plurality of vanes being radially coupled within the second section. Additionally, the fan apparatus includes an impeller with a plurality of blades rotatably disposed within the ducted housing between the first plurality of vanes and the second plurality of vanes for rotary motion about the axis of rotation. The first plurality of vanes are stationary in a first arrangement to guide an inflow of air towards the impeller. The plurality of blades are rotary in a second arrangement to have a rotational direction substantially opposite to the inflow of air guided by the first arrangement. The second plurality of vanes are stationary in a third arrangement to diffuse an outflow of air substantially along the axis of rotation.

Abstract: Aspects of the invention are directed to a system for improving engine performance by reducing vane tip clearances in the compressor section of a turbine engine. According to aspects of the invention, an abrasive material can be attached to a rotor disc. Thus, if a vane tip contacts the rotor disc during engine operation, the vane tip is worn away by the abrasive material. The system can reduce the risk of substantial component damage resulting from a vane tip rubbing event. The abrasive material can be provided in the form of an insert that can be removably attached to the rotor disc. Thus, the insert can be readily removed and replaced when necessary.