Abstract: An air inlet manifold and a gas turbine including the same, capable of preventing transmission of vibration from the gas turbine to the air inlet manifold, while at the same time preventing air leakage, and the air inlet manifold is installed to be spaced apart from the compressor and relatively movable, as the anti-vibration pad in a compressed state is interposed between the air inlet manifold and the compressor.

Abstract: A centrifugal heat dissipation fan including a housing and an impeller is provided. The housing has at least one inlet disposed along an axis and at least one first outlet and a second outlet located in different radial directions, wherein the first outlet and the second outlet are opposite to and separated from each other. The impeller is disposed in the housing along the axis. A heat dissipation system of an electronic device is also provided.

Abstract: The invention relates to a stabilizer channel, in particular of a radial compressor or diagonal compressor, having an annular stabilizer chamber which encloses a main flow channel in the intake region of a compressor wheel and is delimited from the main flow channel by an annular bridge. The annular stabilizer chamber is bladeless and is connected to the main flow channel via a downstream inlet channel and an upstream outlet channel. A plurality of flow directing elements are arranged in the downstream inlet channel. The downstream inlet channel is arranged between an upstream part of the annular bridge and a downstream part of the annular bridge. The invention furthermore relates to a compressor, in particular a radial compressor or diagonal compressor, comprising the stabilizer channel according to the invention and to a turbocharger comprising the compressor.

Abstract: A fan guard for a fan container includes a housing and a plurality of wings. The housing has a hollow interior defined by a cylindrical inner surface. The housing extends longitudinally between a first housing end and a second housing end. The plurality of wings is positioned within the hollow interior of the housing. Each wing of the plurality of wings extends radially, from a center of symmetry of the cylindrical inner surface to the cylindrical inner surface. Each wing of the plurality of wings is radially curved between the first housing end and the second housing end.

Abstract: A gas turbine engine includes a primary flow path fluidly connecting a compressor section, a combustor section and a turbine section. A cooling air flowpath is positioned radially outward of the primary flowpath. A first seal spans from an inner diameter of the cooling air flowpath to an outer diameter of the cooling air flowpath. The first seal includes at least one axial convolution and a plurality of pass through features. A bifurcation feature is positioned downstream of the first seal. The bifurcation feature is configured to direct a portion of an airflow through the first seal into a cooling air passage. A second seal contacts the bifurcation and the first seal, such that fluid radially inward of the bifurcation feature.

Abstract: A fan is provided, including: a body including a fan motor unit, which includes an impeller, a motor and an air-output three-way seat and a nozzle connected to an air outlet of the body, wherein the nozzle receives the airflow from the body and emit the airflow, the nozzle includes a nozzle body in a shape of half a frame; the air-output three-way seat includes at least one air inlet, a first air outlet, a second air outlet, and an air-splitting wall body for splitting the airflow passing through the air inlet and then guiding the airflow to the first air outlet and the second air outlet, respectively. The fan changes the moving direction of the airflow therein, and decreases the entire height and volume.

Abstract: A stator vane provided rearward of a rotor blade, includes: an airfoil body having an airfoil profile. The maximum airfoil thickness position of the airfoil body on the airfoil profile satisfies following conditions at least on a tip side of the airfoil body: (a) on a plane expanded in a circumferential direction in which the stator vane is arranged, the position is located in a first region which is close to a trailing edge of the airfoil body from an intersection of the airfoil body and a line which is parallel to an extension line of a camber line of the rotor blade at a trailing edge of the rotor blade and passes through a leading edge of another stator vane adjacent in the circumferential direction, and (b) the position is located in a second region having a chord ratio from 0.2 to 0.8.

Abstract: In accordance with at least one aspect of this disclosure, embodiments of fluid pumps, pump cases, valve bodies, and volutes are disclosed herein. Embodiments of methods for manufacturing fluid pumps, pump cases, valve bodies, and volutes are also disclosed herein. Embodiments can include additive manufacturing, for example. Certain embodiments can include additively manufacturing a pump case in a tilted orientation, utilizing only coincidental support structure having a unique shape, and with teardrop shaped volute and/or valve body.

Abstract: The invention relates to a rotor for a fluid pump, in particular for use in the medical sphere, the rotor being compressible for bringing to the place of use and thereafter being expandable. The compressibility is assisted by the provision of cavities, in particular also production of the rotor at least partially from a foam.

Abstract: Discharge grate intended to be mounted inside or at the outlet of a conduit of a discharge valve of a turbine engine of an aircraft, the discharge grate comprising an upstream face intended to receive a gas flow, a downstream face parallel to the upstream face and intended to deliver the gas flow received on the upstream face, and orifices passing through the perforated plate from the upstream face to the downstream face and intended to convey the gas flow through the perforated plate. The discharge grate comprises for each orifice of the perforated plate a tubular channel, coaxial with the orifice with which it is associated, and projecting from the downstream face of the perforated plate.

Abstract: A nozzle guide vane for a gas turbine engine having a combined side wall thickness value which varies within a cavity region so as to provide a point with a maximum value of combined side wall thickness, which is advantageous for capturing debris travelling through the engine core.

Abstract: A turbine vane for a gas turbine engine has an airfoil including leading and trailing edges joined by spaced-apart pressure and suction sides to provide an external airfoil surface. The surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil defined by a set of Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location.

Abstract: Assembly including a rotor disc, an adjacent stator and a plurality of sealing elements secured to the rotor disc, the stator including an inner platform and a root bearing at least one abradable element configured to cooperate with the sealing elements, the sealing elements being placed in an enclosure formed by the abradable element, the enclosure being open to the inside and delimited axially by an upstream abradable edge and a downstream abradable edge, the enclosure being delimited radially by an outer abradable edge, at least one of the sealing elements including a first lip configured to cooperate with the upstream abradable edge or the downstream abradable edge, and a second, separate lip configured to cooperate with the outer abradable edge.

Abstract: A vane for a turbine engine turbine includes a blade and at least one platform that radially extends the blade, and a system for reducing vortices including at least one circuit including a duct extending from at least one air intake orifice formed in the platform upstream of the leading edge as far as at least one air exhaust slot formed on the trailing edge.

Abstract: A method of operating an aircraft engine having a bladed rotor coupled thereto during an icing condition is provided. The method comprises controlling the aircraft engine to alternatingly achieve an ice accretion rotational speed of the bladed rotor at which ice accretion on blades of the bladed rotor occurs, and a reduced rotational speed of the bladed rotor to promote ice shedding from the blades of the bladed rotor. The reduced rotational speed is lower than the ice accretion rotational speed.

Abstract: A diagonal fan having an electric motor, a housing and a diagonal impeller generating a diagonal flow which is deflected into an axial flow direction. The diagonal impeller has impeller blades, an air inlet and an air outlet, wherein the housing forms a flow channel for an airflow generated by the diagonal impeller, which has a non-rotationally symmetric axial section and a cylindrical axial section axially directly adjacent, as seen in the flow direction, wherein an air outlet-side radial outer end of the diagonal impeller is arranged in the cylindrical axial section of the flow channel of the housing and an air gap is provided between the radial outer end and the housing, and wherein the non-rotationally symmetric axial section of the flow channel is arranged in a region of the flow channel which is adjacent to the air-inlet side of the air gap in an axial plane with the diagonal impeller.

Abstract: Assembly for axial turbomachine, in particular for an aircraft turbojet, the assembly comprising: an annular casing with an internal surface (40); an annular row of stator baffles (26) with at least one stator baffle (26) comprising an airfoil (50) which extends radially from a fixing platform (34), the fixing platform (34) being fixed to the casing and having a polygonal outline; characterised in that it further comprises a gasket (80) comprising a frame, the outline of which conforms to the polygonal outline of the fixing platform (34), the frame being in radial contact with the fixing platform (34) and the casing in order to ensure a seal.

Abstract: A turbine housing for a radial/mixed flow turbine, the turbine having a rotor comprising a plurality of blades disposed about a hub, the turbine housing comprising: a shroud configured to at least partially enshroud the rotor wherein a clearance is provided between the shroud and blade tips of the plurality of blades, the shroud and the hub defining a main passage of fluid flow in the turbine; and at least one injection opening provided in the shroud, the at least one injection opening configured to allow injection of a secondary flow of fluid into at least the clearance between the shroud and the blade tips.

Abstract: A rotary device includes: a housing; an impeller disposed in the housing; a rotor disposed in the housing and configured to drive the impeller; and a labyrinth seal disposed between the impeller and the rotor in the housing and configured to control an amount of air injected through the impeller to cool the rotor. A flow path opening that penetrates through the rotor is formed inside the rotor along a rotational axis of the rotor and the labyrinth seal comprises a teeth portion having a predetermined number of steps provided in the labyrinth seal.

Abstract: The invention is related to a high gamma prim nickel based superalloy, its use and a method of manufacturing of turbine engine components by welding, 3D additive manufacturing, casting and hot forming, and the superalloy comprises 9.0-10.5 wt. % Cr, 20-22 wt. % Co, 1.0-1.4 wt. % Mo, 5.0-5.8 wt. % W, 2.0-6.0 wt. % Ta, 3.0-6.5 wt. % Al, 0.2-0.5 wt. % Hf, 0.01-0.16 wt. % C, 1.5-3.5 wt. % Re, 0-1.0 Ge wt. %, 0-0.2 wt. % Y, 0-1 wt. % Si, 0-0.015 wt. % B and nickel with impurities to balance.

Abstract: Disclosed is a compressor including: a rotatably-mounted rotor disk including a slot in an outer circumference thereof; and a blade including a root member connected to fix to the slot of the rotor disk, an airfoil with a leading edge facing introduced air and a trailing edge, and a platform formed in between the root member and the airfoil to stably support the airfoil over the rotor disk, the blade further including first and second bumps being respectively formed on each end side of the platform to face the introduced air.

Abstract: Systems and methods of conditioning inlet air flow in a turbine engine. Where distortions in uniformity of inlet air flow are caused at least in part by the interaction of the air flow with the air inlet duct, a method of adaptively removing the distortions prior to the compressor stage comprises determining the distortion in the airflow; exposing the airflow to a plurality of correction vanes; and positioning the plurality of correction vanes based at least upon the determined distortion. An inlet conditioner system comprises an adaptable conditioning grid located within an air passage; a sensor suite configured to sense a characteristic of the airflow within the air passage; and a control system operably connected to the sensor suite and the adaptable conditioning grid. The control system may be adapted to configure the adaptable conditioning grid based on a sensed characteristic.

Abstract: A compressor of a gas turbine engine includes a rotor and a stator located downstream of the rotor. The stator has a plurality of vanes each with an airfoil extending span-wise between a root proximate an inner hub of the stator and a tip. A fillet is disposed at the leading edge of the root of the airfoil, and extends between a pressure side surface of the airfoil and the inner hub.

Abstract: Wind funnel and gas combustion turbine systems are disclosed. Air travels through a wind funnel where it is compressed, and then flows into a compression section of a gas turbine that is fueled by a hydrocarbon fuel source such as natural gas. Compressed air from the wind funnel enters the compression section of the gas turbine through one or more air inlets toward multiple rotating turbine blades in an air feed direction having a component normal to air impact surfaces of the turbine blades. In addition to the compressed air from the wind funnel, ambient air may be introduced into the compression section of the gas combustion turbine.

Abstract: An assembly adapted for use in a gas turbine engine or other engine has a carrier component and a blade track segment. The assembly includes a mounting system for coupling the blade track segment to the carrier component. In an illustrative embodiment, the assembly is a turbine shroud segment for blocking gasses from passing over turbine blades included in the gas turbine engine.

Abstract: A centrifugal compressor includes impellers arranged in a plurality of stages in a direction of an axis line so that a working fluid flowing from one side inlet in the direction of the axis line is pumped outward in a radial direction, and a casing that surrounds the impellers, and that has a return channel through which the working fluid discharged from the impeller on a front stage side between the impellers adjacent to each other is guided inward in the radial direction so as to be introduced to the impeller on a rear stage side, and a plurality of return vanes disposed at an interval in a circumferential direction inside the return channel. The return vane is configured so that the thickness of a hub side in a leading edge side region is thicker than the thickness of a shroud side.

Abstract: Rotor blades and turbomachines are provided. A rotor blade includes a platform and an airfoil extending radially from a root proximate the platform to a tip. The airfoil includes a leading edge and a trailing edge. The airfoil further includes a pressure side surface extending between the leading edge and the trailing edge. The airfoil also includes a suction side surface disposed opposite the pressure side surface and extending between the leading edge and the trailing edge. The leading edge, the trailing edge, the pressure side surface, and the suction side surface collectively define a first profile at the tip and a second profile radially inward from the tip. The first profile includes a suction side overhang, a pressure side overhang, and a pressure side underhang relative to the second profile.

Abstract: An inlet for a flight vehicle engine, such as for a supersonic or hypersonic engine, includes an internal flow diverter to divert boundary layer flow. The flow diverter is configured to minimize disruption to flow outside the diverted boundary by being configured through use of a flow field that is also used to configure the walls of the inlet. The flow field that is used to configure an inlet-creating shape and a diverter-creating shape has the same flow generator, contraction ratio, compression ratio, mass capture ratio, pressure ratio between entrance and exit, and/or Mach number, for example. The internal diverter may be configured so as to allow arbitrary selection of a leading edge shape for the internal diverter, for example to use a shape that helps avoid radar detection.

Abstract: A mixing connector includes an intake passage, an EGR passage that fetches a portion of exhaust gas exhausted from an engine body to use as EGR gas, and that returns the EGR gas to the intake passage, and a merging section that connects the EGR passage to the intake passage so that longitudinal directions of the intake passage and the EGR passage intersect each other. An upstream side region located on an inlet port side of the intake passage from the merging section on an opposite wall surface configuring an inner surface of the intake passage and located on a side opposite to the merging section includes a first wall surface and a second wall surface which are sequentially arranged at an interval from the merging section side toward the inlet port, and a third wall surface projecting inward of the first wall surface between the first and second wall surfaces.

Abstract: A recirculation stage of a radial turbomachine, the recirculation stage has an inner delimiting contour and an outer delimiting contour. The stator vane stage has stator vanes, wherein a difference between a vane construction angle at the leading edge and a vane construction angle at a downstream position defines a redirection angle for each point on the camber line of a respective profile cross section, wherein the stator vanes extend at least along part of the third section, wherein the trailing edges are arranged in the third section, wherein at the trailing edges in the center of the span width the redirection angle is in each case greater than the average overall redirection angle, wherein at both ends of the span width at in each case at least 10% of the span width in each case the redirection angle is smaller than the average overall redirection angle.

Abstract: A combustion chamber arrangement includes an annular combustion chamber, a plurality of lean burn fuel injectors and a stage of turbine nozzle guide vanes. The stage of turbine nozzle guide vanes is arranged at the downstream end of the annular combustion chamber and includes a plurality of circumferentially spaced vanes extending between and being secured to an annular inner wall and an annular outer wall. The annular inner wall has an outer surface which has a plurality of outer surface portions and each portion is positioned between a pair of circumferentially adjacent vanes and the annular outer wall has an inner surface which has a plurality of inner surface portions and each portion is positioned between a pair of circumferentially adjacent vanes.

Abstract: This disclosure relates to the field of display fabrication technologies, and discloses a fine mask support frame, a fine mask, and a method for fabricating the same. The fine mask support frame includes: a plurality of bezels surrounding a mask area, wherein adjustment openings are arranged on at least one pair of bezels arranged opposite to each other, and at least one adjustment piece is arranged on each of the adjustment openings, wherein the adjustment piece is configured to adjust a shape of a corresponding adjustment opening to thereby adjust deformation of the mask area.

Abstract: The present invention includes: a turbine wheel 24 that is attached to a rotating shaft 14 and that has a plurality of turbine blades in the circumferential direction; scroll passages 15Aa, 15Ab that are formed in a spiral shape on the outside of the turbine wheel 24, and that are formed by being divided into a plurality of sections in the circumferential direction, the passages communicating with each other at the position of the turbine wheel 24 disposed between respective tongue sections 15Ac, 15Ad; and threads 51 that are provided on an back facing surface 41 disposed so as to oppose the back surface of the turbine wheel 24, the back surface being on the axially opposite side from the turbine blade side, and the linear parts extending from starting points 51a near the tongue sections 15Ac, 15Ad toward the rotating shaft 14 side so as to control the passage of fluid between the back surface and the back facing surface 41.

Abstract: A volute casing includes a chamber, an outlet passage, and a cutwater to direct fluid to the outlet passage. The cutwater includes an inner surface, an outer surface and a leading edge joining the inner and outer surfaces and has a cross-sectional contour in a midplane perpendicular to the axial direction. The cross-sectional contour includes a starting point at the leading edge, and a minimum point on the inner surface, the starting point defined by a tangent to the leading edge, the tangent intersecting the central axis, and the minimum point defined by a location, at which the inner surface has a minimum distance from the central axis. A straight profile chord located in the cross-sectional contour, and extending from the cutwater starting point to the cutwater minimum point, has a maximum orthogonal distance from the inner surface being at most 15% of the length of the profile chord.

Abstract: A turbocharger includes: a turbine wheel; a hub-side wall surface and a shroud-side wall surface which face each other and which together form a flow path for exhaust gas to flow into the turbine wheel; and a nozzle vane which has a shroud-side end surface facing the shroud-side wall surface and a hub-side end surface facing the hub-side wall surface and which is rotatably disposed in the flow path. The nozzle vane has a pressure surface which is, at least on a shroud side, oblique toward a downstream side of the exhaust gas with a distance in a vane height direction from the shroud-side end surface.

Abstract: A turbocharger includes a turbine wheel, a hub-side wall surface and a shroud-side wall surface which face each other and which together form a flow path for exhaust gas to flow into the turbine wheel, and a plurality of nozzle vanes each of which is rotatably disposed on one of the hub-side wall surface or the shroud-side wall surface in the flow path. A throat formed between adjacent two of the plurality of nozzle vanes has a narrower throat width at a position of a first end surface of each of the adjacent two nozzle vanes facing the other of the hub-side wall surface or the shroud-side wall surface than a throat width at a middle position between the first end surface and a second end surface on an opposite side to the first end surface of each of the adjacent two nozzle vanes.

Abstract: An article for service at high temperatures comprises a substrate comprising a first region and a second region; and a coating disposed over the substrate. The coating comprises a first portion disposed over the first region of the substrate and a second portion disposed over the second region of the substrate. The coating includes a layer comprising a ceramic material and further including a quantity of surface-connected voids, and a protective agent is disposed within at least some of the surface-connected voids of the layer. Within the first portion, the agent is present within the layer at a concentration of less than or equal to 4 percent by volume of the layer exclusive of the voids, while, within the second portion, the agent is present within the layer at a concentration up to the carrying capacity of the layer. Methods for forming the article are also presented.

Abstract: An airfoil includes a core structure that has an airfoil portion that defines at least a portion of an airfoil profile. The airfoil portion has first and second opposed ends. A support wall extends from the first end. An endwall panel is attached to the support wall adjacent the airfoil portion.

Abstract: A centrifugal compressor, in which a plurality of impellers that rotate along with a rotation shaft to pump fluid by using centrifugal force are disposed at a plurality of stages along an axial direction to compress, in a stepwise manner, the fluid sucked in from a suction port, the centrifugal compressor including: a closed impeller including a plurality of vanes disposed in a radial manner about the rotation shaft, and a shroud covering the plurality of vanes from radially outside; and an open impeller including a second plurality of vanes without a shroud. The closed impeller is disposed at a rearmost stage in a fluid flow direction. The open impeller is disposed at a stage located immediately after the suction port in the fluid flow direction.

Abstract: A cooling arrangement for a turbine engine includes a cooling source and first and second structures. A ceramic-based composite cooling tube fluidly provides a fluid connection between the first and second structures. The cooling tube is configured to transfer a cooling fluid from the cooling source to the second structure.

Abstract: A blower for providing a supply of air at positive pressure in the range of approximately 2 cm H2O to 30 cm H2O includes a motor, at least one impeller, and a stationary component. The stationary component includes an inlet and an outlet. The motor, the impeller, the inlet and outlet are co-axial.

Abstract: An impeller is provided, which includes an upper plate, blades, and a lower plate. The blades are arranged between the upper plate and lower plate. The upper plate includes an upper surface and a lower surface, the blades are integrally formed with the upper plate by injection molding, and the blades are located at the lower surface of the upper plate. The blades include first blades. Each of the first blades includes one arc, a combination of multiple arcs, or a combination of multiple arcs and straight lines. The first blades are uniformly distributed along a circumference of the upper plate, and the first blade includes a first tail, and the first tail of the first blade make a first side of the first blade have no intersection with an outer edge of the upper plate.

Abstract: A gas turbine exhaust diffuser is configured to be at least partially disposed within an exhaust collector. The diffuser also includes an inner wall configured to extend to a back wall of the exhaust collector, an outer wall circumferentially disposed about the inner wall along a portion of the collector, and an inlet between the upstream end of the inner and outer walls. The inlet is configured to receive an exhaust flow from a gas turbine engine. An outlet for the exhaust is located between the inner wall and a downstream end of the outer wall. The outer wall is asymmetrically shaped so that an opening between the outlet and the back wall of the exhaust collector is larger on a first lateral side of the diffuser adjacent an exhaust exit of the collector than an opposite lateral side of the diffuser.

Abstract: Aspects of the disclosure are directed to a system for an engine having an axial centerline, comprising: a diffuser case, a turbine case, and a turbine vane support, where the diffuser case and the turbine case are coupled to one another via a substantially radially oriented flange, where the turbine vane support includes a heat shield for the flange, and where the turbine vane support includes a radially outward projecting tab that couples to the turbine case via a radial interference fit.

Abstract: When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P0 situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P1. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine.

Abstract: A turbine blade for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil including leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface extending from a platform in a radial direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil described by a set of Cartesian coordinates.

Abstract: A doubler attachment system is provided for fixedly coupling a strut to an engine casing. The strut includes at least one flanged end configured to mate with the engine casing. The doubler attachment system includes a doubler plate having a first radial surface and a second opposing radial surface. The first radial surface is configured to mate with and conform to a contour of the at least one flanged end of the strut. The doubler attachment system further includes at least two bolts fixedly attached to the doubler plate at the first radial surface. The at least two bolts extend radially away from the first radial surface and toward the engine casing.

Abstract: Eductor systems and methods improve eduction by increasing swirl. An eductor housing defines a primary plenum channeling a flow stream from an inlet opening to an exit opening, and defines a secondary plenum separated from the primary plenum and channeling another flow stream from a duct inlet to exit slots. A gas flow stream is delivered into the eductor housing and is directed through the exit opening so that the gas flow stream educes the flow stream through the inlet opening and the exit opening. The second flow stream is delivered to the secondary plenum through a duct connected at an angle to induce a swirl in the secondary plenum to effect a static pressure reduction at the exit opening that draws the flow stream from the inlet opening through the exit opening.

Abstract: A method for modeling an offset portion of a blade of a non-ducted propeller is provided. The method includes: parametrizing a C1-class curve representing a deformation of the blade characterizing the offset, according to a position along a section at a given height in the blade, the curve intersecting consecutively through a first bend control point, a central control point, and a second bend control point, the first and second bend control points defining the extent of the blade section, the parametrization being implemented according to a first deformation parameter defining the abscissa of the central control point, a second parameter of deformation defining the ordinate of the second bend point, and a third deformation parameter defining the angle of the tangent to the curve at the second bend control point; optimizing one of the deformation parameters; and plotting the values of the optimized parameters on an interface.

Abstract: A rotating circular aerofoil system that includes a disc body that has an aerofoil shape in cross section. The aerofoil shape has an angle of attack that is symmetrical. The disc body's includes a circular inner edge which forms a circular center opening and the disc body's leading edge. The disc body also includes an outer edge that forms the trailing edge. Intersecting ribs extend across the center opening and include a shaft opening that attaches to a rotating shaft. Formed on the top and bottom surfaces of the disc body are evenly spaced apart, extending T-shaped or L-shaped fins. In one embodiment, an air blower or impeller is attached to a hollow shaft linked to an air pump. When the air pump is activated, air is delivered to the air blower and impeller that distributes the air over the top surface of the disc body to improve lift.