Abstract: A turbine bucket airfoil has a conical fillet about the intersection of the airfoil tip and tip shroud having a nominal profile in accordance with coordinate values of X and Y, offset 1, offset 2 and Rho set forth in Table I. The shape parameters of offset 1, offset 2 and Rho define the configuration of the fillet at the specified X and Y locations about the fillet to provide a fillet configuration accommodating high localized stresses. The fillet shape may be parabolic, elliptical or hyperbolic as a function of the value of the shape parameter ratio of D1 D1 + D2 at each X, Y location where Dl is a distance between an intermediate point along a chord between edge points determined by offsets O2 and O2 and a shoulder point on the fillet surface and D2 is a distance between the shoulder point and an apex location at the intersection of the airfoil tip and tip shroud.

Abstract: An air moving apparatus for generating cooling airflow is provided that includes a noise reduction system for reducing noise generated by a fan. The air moving apparatus includes a fan having a rotatable hub and a plurality of blades mounted to the hub for rotating about an axis of rotation to provide pressurized airflow. A sensor is situated on a surface of at least one fan blade for sensing airflow characteristics of the air flowing over the fan blade. An actuator, also situated on the surface of the fan blade, changes the characteristic of the airflow over the fan blade in response to the sensed airflow characteristic.

Abstract: In various embodiments, the present invention provides a means for improving gas turbine engine performance by applying fluid flow control to the inter-turbine duct joining a high-pressure turbine spool and an associated low-pressure turbine spool, allowing the low-pressure turbine spool to have a relatively larger diameter than the high-pressure turbine spool. One or more unobstructed fluid flow paths between one or more boundary layer suction ports disposed within the upstream end of the outer-body surface of the inter-turbine duct and the suction side of the associated low-pressure turbine nozzle are provided. Advantageously, the natural static pressure difference between these points results in a self-aspirating assembly. The fluid flow control provided by the respective suction and blowing forces generated allows for a relatively larger diameter low-pressure turbine spool and/or relatively fewer low-pressure turbine nozzles to be used than is possible with conventional systems, assemblies, and methods.

Abstract: A diffuser includes an opening along the diffuser wall for preventing or delaying boundary layer separation of the main flow. The opening may include a curved passageway having a curvature convex relative to the main flow for utilizing the Coanda effect. A gas turbine may include the diffuser with an opening and a centerbody with an opening, each for providing a secondary flow of fluid into the diffuser. The gas turbine may direct fluid from an upstream turbine stage to the opening. A steam turbine may include the diffuser and may situate the opening downstream the diffuser inlet but upstream the location of boundary separation.

Abstract: A flow directing device of a gas turbine engine, comprising: an airfoil having a leading edge, trailing edge, suction side and pressure side; a wall abutting the airfoil; and a fillet between the airfoil and wall. The fillet has an enlarged section at the leading edge, along the suction and pressure sides, and towards the trailing edge. The device could be part of a vane segment. In addition to eliminating a horseshoe vortex, the device also reduces heat load on the airfoil by directing the cooler gas from the proximal end of the airfoil to the hotter gas at the medial section of the airfoil.

Abstract: A vane assembly for a gas turbine engine includes at least one vane that includes a first body, a second body, and a passageway. The first body includes a first sidewall and a second sidewall that are connected at a leading edge and a trailing edge. The passageway extends between the second body and the first body leading edge.

Abstract: In order to provide a stall prediction method for an axial flow compressor, a high response pressure sensor is provided in a duct wall adjacent to the leading edge of the rotor blade and a rotating stall is predicted by computing autocorrelation of pressure data obtained as a time history data. The initiation of rotating stall is determined when the autocorrelation of the time history data starts dropping rapidly. When it is determined that the rotating stall is imminent, generation of the rotating stall is prevented by taking necessary countermeasures such as to reduce the fuel supply.

Abstract: The present invention is a device (100) for at least partially stabilizing an unstable fluid flow within a flow channel (103) by capturing at least a portion of the unstable fluid within a vaneless diffuser having a diffuser slot (104). The present invention also includes maintaining and harnessing a substantial portion of the energy contained in the fluid as it flows through the diffuser in order to utilize the fluid to improve the condition of the flow field. An example of a beneficial use includes discharging the diffuser effluent into the flow at other points critical to instability, hence reducing the overall instability of the flow channel.

Abstract: A turbine includes a row of blades each having an integral airfoil, platform, and dovetail. Each platform has opposite first and second side edges corresponding with the opposite pressure and suction sides of the airfoil. The platform first side edge is radially higher than the platform second side edge continuously between opposite forward and aft edges of the platform. In this way, adjacent platforms define a down step therebetween for preventing obstruction of combustion gases flowable downstream thereover.

Abstract: The present invention relates to an endwall shape for reducing shock strength on transonic turbomachinery airfoils which define at least one flow passage. The endwall shape includes a non-axisymmetric trough which extends from a leading portion of the at least one flow passage to a point near a trailing edge portion of the at least one flow passage.

Abstract: A ventilator housing of a ventilator has a flow passage having an intake opening and an outlet opening. Outflow contouring devices are arranged about the outlet opening and are configured to at least reduce a vortex formation within a shearing layer that is generated during operation of the ventilator so as to surround an actual airflow of the ventilator and that is located between the airflow and surrounding air. The outflow contouring devices enlarge the proportion of uninterrupted air flow relative to the entire flow cross-section of the flow passage.

Abstract: An axial or centrifugal flow compressor having arrays of blades (16) extending across a working medium flowpath (18) includes a casing treatment for enhancing the compressor's fluid dynamic stability. The casing treatment features a circumferentially extending compartment (62), typically comprising a voluminous pressure compensation chamber (64) and a single passage (66) circumferentially coextensive with the chamber, for establishing fluid communication between the chamber and the flowpath. The voluminous character of the compartment attenuates the circumferential pressure difference across the tips of heavily loaded compressor blades (16), making the compressor less susceptible to tip vortex induced instabilities. In one embodiment of the invention, a passage (66) is oriented so that any fluid discharged from the passage enters the flowpath with a streamwise directional component.

Abstract: The present invention relates to an endwall shape for reducing shock strength on transonic turbomachinery airfoils which define at least one flow passage. The endwall shape includes a non-axisymmetric trough which extends from a leading portion of the at least one flow passage to a point near a trailing edge portion of the at least one flow passage.

Abstract: A blade for an axial-flow turbine includes an intrados producing a positive pressure between a leading edge and a trailing edge, and an extrados producing a negative pressure. The intrados is formed at its rear portion with a flat surface portion connected to the trailing edge, and the extrados has a curved surface portion formed at least at a portion corresponding to the flat surface portion. The trailing edge of the turbine blade is pointed at its end. The angle of intersection between the intrados and the extrados at the trailing edge is a right angle or an acute angle. Thus, it is possible to inhibit the flowing of a gas from the intrados at the trailing edge toward the extrados and to decrease the degree of curvature of the extrados at the trailing edge portion to reduce the flow speed, thereby minimizing a shock wave generated at the trailing edge portion to reduce the pressure loss and enhance the performance of the turbine.

Abstract: In a cooling structure for a gas turbine, cooling air diffusion holes are formed from inner surface to outer surface of a platform so as to open from high pressure side blade surface of a moving blade offset in a direction toward low pressure side blade surface of adjacent moving blade confronting the high pressure side blade surface, in a direction of primary flow.

Abstract: A diffuser for slowing down a fluid and a method for operating a diffuser is described. A channel of the diffusor has an inlet with a smaller flow cross section than a flow cross section of an outlet and at least one opening for receiving an energizing fluid to be transported selectively into the channel. Pulsations of the impact interface are suppressed effectively at all of the operating points by injecting the energization fluid. Pressures of the fluid moving in the diffusor are measured, and amplitudes and frequencies of the measured pressures are evaluated. Energizing fluid is fed into the diffusor if the amplitudes within a predetermined frequency band exceed a threshold value. The utilization ratio of the inventive diffuser is considerably improved as a result of such a measure.

Abstract: A cast assembly for a gas turbine engine compressor, the compressor including a plurality of axial flow stators and a rotor having a plurality of axial flow blades, each of the blades having a tip section and a leading edge. The assembly comprises a stator outer platform, a boss, a passage, a blade outer shroud and a circumferential slot. A boss is formed in the platform and includes a passage. The passage includes an inlet and an outlet. The inlet is disposed in the circumferential slot. The outlet being disposed upstream of the leading edge of the blades. The passage circumferentially converges inward from the inlet to the outlet. The outer shroud joins with said outer platform during the assembly of the compressor such that the circumferential slot is formed downstream and proximate to the leading edge of the blades. The circumferential slot is in flow communication with the inlet and converges axially.

Abstract: For improving characteristic uprising at the right-hand side of head-flow rate characteristic of a turbo machine, as well as for suppressing increase of vibration and/or noises thereof, a plurality of first grooves 24 in direction of gradient in pressure of fluid are formed on an inner flow surface of a casing, for connecting an inlet side of blades of an impeller and an area on the inner flow surface of the casing where the impeller blades reside in, over an inner circumference of the casing. Also, second grooves 25 in a circumferential direction are formed in an area on the inner flow surface of the casing where the impeller blades reside in, for communicating the first grooves in the circumferential direction of the casing.

Abstract: The present invention relates to a bleed deflector for a gas turbine engine. The bleed deflector comprises an inlet portion for receiving bleed air from the engine, a perforated top plate for discharging the bleed air into a duct, and an aerodynamically shaped strut for positioning the perforated top plate above an inner wall of the duct to allow dilution air to flow beneath the top plate.

Abstract: In an exhaust turbocharger for an internal combustion engine including a turbine and a compressor, which is driven by the turbine, wherein the compressor comprises a compressor impeller in a flow duct, a transfer component with flow passages extends radially around the compressor impeller in the axial direction of the impeller and provides for communication with the flow duct. To provide for variable adjustment of the effective flow cross-section of the transfer duct, a transfer ring is provided which comprises two component rings, which are rotatable relative to one another.

Abstract: A gas turbine engine outlet guide vane assembly has annular inner and outer end walls, a flowpath between the inner and outer end walls, outlet guide vanes radially disposed between the inner and outer end walls, and boundary layer energizing means for energizing boundary layers using secondary flow to mix free stream flow into the boundary layers along the inner and outer end walls and suction and pressure sides of the vanes. The boundary layer energizing means includes having the vanes circumferentially leaned in a direction that the suction sides face. The boundary layer energizing means also includes swept leading and/or trailing edges of the vanes that extend radially between the inner and outer end walls. The swept leading and/or trailing edges may be curved inwardly into the vanes from the outer end walls to leading and/or trailing edge points respectively between the end walls.

Abstract: A turbo-type machine, comprising a casing 2 storing an impeller within an inside thereof, and a plural number of grooves formed on an inner surface of the casing in a direction of pressure gradient of fluid, wherein the grooves 124 are provided from 80 to 150 pieces around a periphery on the inner surface of the casing, and a total width of the grooves all around the inner surface of the casing is from 30% to 50% of a peripheral length on the inner surface of the casing. Also, the area on cross-section of the groove 124 within an area where the blades exist is made larger than that of the groove outside where the blades exist, or openings 127 are drilled on a bottom surface of the grooves, each penetrating through thickness of the casing, or the groove is made in a two-layer structure in the direction of pressure gradient of fluid.

Abstract: A turbo machine comprising: an impeller having a plurality of blades therewith; a casing having a flow surface defined therein and being positioned with the impeller therein; and a plurality of grooves being formed in the flow surface of the casing, for connecting between an inlet side of said impeller and an area on the flow surface of the casing in which the blades of said impeller reside. Each groove has a length of at least part of which is oriented in an axial direction of the casing, a width measured in a circumferential direction, and a depth. The width of each groove is equal to or greater than the depth thereof.

Abstract: Tip treatment bars 16 of a rotor casing 2 for a gas turbine engine are subjected to a process which induces compressive stress within the bars 16. The process may be a laser shock peening process applied to opposite sides 20 of each bar 16. The processing of the tip treatment bars 16 renders them less susceptible to fatigue failure.

Abstract: A gas turbine engine exhaust nozzle includes an outlet for discharging exhaust, and a plurality of circumferentially spaced apart stub airfoils disposed adjacent the outlet. The airfoils shed vortices for mixing exhaust flow and attenuating noise and reducing infrared signature with minimum performance loss.

Abstract: The cylindrical compressor casing of a gas turbine engine has a plurality of radially inwardly and axially rearwardly opening anti-surge grooves disposed on a radially inner surface thereof whereby foreign objects ingested into the engine and entering the grooves are free to move axially rearwardly of the engine.

Abstract: A pump, comprising: an impeller 1 having blades; and a casing 121 for storing the impeller therein, on an inner surface of which, confronting to the impeller, are formed plural numbers of shallow grooves 124 in a direction of pressure gradient of fluid, around a periphery thereof, wherein the fluid being increased up in pressure by the blades 122 flows within the grooves in a reverse direction, directing to an upstream side, so as to spout out at a place where re-circulations occur when the flow rate is low in amount.

Abstract: A stall precursor detector system for an axial flow compressor having at least one optical sensor to detect the deflection of a rotating airfoil of the compressor, means for comparing the deflection with a predetermined value, the predetermined value being indicative of onset of a stall in the compressor; and means for identifying onset of a stall and producing a stall onset signal if the deflection is greater than the predetermined value. A method for detecting a stall onset in an axial flow compressor system comprising (a) providing at least one optical sensor to measure the deflection of an airfoil, the deflection caused by a rotating stall cell; (b) comparing the deflection as measured in step (a) with a predetermined value indicative of a stall; and (c) identifying onset of a stall if the measured deflection is greater than the predetermined value.

Abstract: A rotor assembly for a gas turbine engine operates with reduced circumferential rim stress. The rotor assembly includes a rotor including a plurality of rotor blades extending radially outward from an annular rim. A root fillet extends circumferentially around each blade between the blades and rim. The rim includes an outer surface including a plurality of concave indentations extending between adjacent rotor blades and forming a compound radius. Each indentation extends from a leading edge of the rotor blades towards a trailing edge of the rotor blades.

Abstract: A swirl generator for creating bulk swirl in a compressor plenum in the opposite direction to the direction of rotation of the compressor rotor. The generator includes a continuous flow passageway formed by an inlet duct and a compressor forming an angle in the range of about 0°-180° to one another. A non-uniform velocity gradient is created by forming an area of non-uniform cross-sectional area within the flow passageway. The non-uniform area may comprise a ramp with a tapered end line, a rounded line of intersection, a fillet or by lengthening one side wall of the inlet duct relative to the opposite side wall. Preferably, the non-uniform cross sectional area uniformly tapers from ones sidewall of the flow passageway to the opposite sidewall.

Abstract: In a centrifugal compressor, a housing is provided which has an annular treatment cavity in a shroud wall. The shroud wall has a first opening providing communication between an installed zone of an impeller and the treatment cavity, and a second opening providing communication between the treatment cavity and a zone upstream of the first opening so as to circulate air a during a low-flow-rate operation. Louvers are arranged in the second opening in circumferentially equidistantly spaced apart relationship with each other and inclined reversely to a rotative direction of the impeller so that the air a sucked to the treatment cavity is, when discharged through the second opening, subjected to directional flow guide action with a whirling direction reverse to the rotative direction of the impeller.

Abstract: A power control system for driving a vessel (which may be manned or unmanned) is used to provide a series of pulse trains to the system wherein at least one pulse within each series of pulse trains provides a differing amount of power. Preferably, a portion of the series of pulse trains which has a duration of 10% of the duration of the pulse train delivers more than 20% of the total power to the drive member for the vessel which the drive member receives each cycle. In this way, the noise signature of a vessel is reduced or is varied to match the noise signature of another vessel or wildlife of a natural phenomenon.

Abstract: An aerofoil (22,24), preferably of a high lift, highly loaded design, for an axial flow turbo machine (10). The aerofoil having a span, a leading edge (LE), a trailing edge (TE) and a cambered sectional profile comprising a pressure surface (30,72) and a suction surface (28,70) extending between the leading edge (LE) and trailing edge (TE). The aerofoil (22,24) having at least one aerofoil cross bleed passage (36,37,78,80) defined in the aerofoil (22,24) which extends from the pressure surface (30,72) through the aerofoil (22,24) to the suction surface (28,70). The at least one passage (36,37,78,80) preferably disposed generally at a location on the suction surface (28,70) at which boundary layer separation from the suction surface (28,70) would normally occur.

Abstract: For improving characteristic uprising at the right-hand side of head-flow rate characteristic of a turbo machine, as well as for suppressing increase of vibration and/or noises thereof, a plurality of first grooves 24 in direction of gradient in pressure of fluid are formed on an inner flow surface of a casing, for connecting an inlet side of blades of an impeller and an area on the inner flow surface of the casing where the impeller blades reside in, over an inner circumference of the casing. Also, second grooves 25 in a circumferential direction are formed in an area on the inner flow surface of the casing where the impeller blades reside in, for communicating the first grooves in the circumferential direction of the casing.

Abstract: A compression system endwall bleed system having a plurality of bleed slots formed in a mechanical housing downstream of a rotating shrouded rotor. In one embodiment, the plurality of bleed slots bleed off a separated tip boundary layer to relieve back pressure associated with this blockage. A sealing structure downstream of the shrouded rotor is utilized to minimize working fluid leakage ahead of the plurality of bleed slots.

Abstract: A flow by-pass system is provided in a downward-discharging exhaust hood of a steam turbine to by-pass a small percentage of the total steam flow from the top portion of the exhaust hood to the vicinity of the condenser and in that way relieve excess pressure in the top portion incident to the more convoluted path of the main portion of steam passing from the top to the bottom and to decrease energy loss caused by friction and thus to improve turbine efficiency. The flow by-pass system includes by-pass conduits within the front portion of the exhaust hood extending from the top portion to the vicinity of the condenser.

Abstract: A method for inhibiting radial transfer of core gas flow away from a center radial region and toward the inner and outer radial boundaries of a core gas flow path within a gas turbine engine is provided that includes the steps of: (1) providing a flow directing structure that includes an airfoil that abuts a wall surface, said airfoil having a leading edge, a pressure side, and a suction side; and (2) increasing the velocity of the core gas flow in the area where the leading edge of the airfoil abuts the wall. Increasing the velocity of the core gas flow in the area where the leading edge of the airfoil abuts the wall impedes the formation of a pressure gradient along the surface of the airfoil that forces core gas from the center region of the core gas toward the wall. The apparatus includes apparatus for diverting core gas flow away from the area where the airfoil abuts the wall.

Abstract: In a cooling structure for a gas turbine, cooling air diffusion holes are formed from inner surface to outer surface of a platform so as to open from high pressure side blade surface of a moving blade offset in a direction toward low pressure side blade surface of adjacent moving blade confronting the high pressure side blade surface, in a direction of primary flow.

Abstract: To reduce profile losses in the aerofoils of low pressure turbines of axial flow reaction turbines, the wake-passing effects on a row of aerofoils downstream of a first row of aerofoils are enhanced by providing a region of roughness on part of the suction surface of each aerofoil of the downstream row. The region of roughness has its leading edge between the location of the geometric throat on the suction surface and a location 75% of the surface perimeter from the leading edge. The region extends over at least 3% of the suction surface perimeter. The use of such a region of roughness is particularly effective in improving the performance of very high-lift aerofoils (coefficient of lift of 1.1 or greater).

Abstract: A gas turbine engine casing includes a tip treatment ring 20 which comprises end supports 18 which support tip treatment bars 16. Each tip treatment bar 16 is mounted at its ends within the respective end supports 18 by damping elements 22, which isolate the bar 16 from the end supports 18. The elements 22 damp vibrations induced in the tip treatment bars 16 so inhibiting the initiation of high cycle fatigue cracking.

Abstract: A method for reducing blade vibration in turbo-machinery is described. The method involves inserting flow obstructions and/or gas injections upstream of vibrating blades in such a manner that power flow into the blades is reduced by means of cancellation within a modal power integral.

Abstract: A gas turbine engine rotor assembly such as a fan or compressor includes fan blades spaced axially from stator vanes inside an annular duct. A controlled air blowing system is secured to the annular duct and adapted to blow a hot air flow in a direction opposite to the rotation of the fan blades into the annular duct between the fan blades and the stator vanes, to impart an opposite rotational momentum with respect to the air compressed by the rotor blades, in a asymmetrical pattern to interfere with or destroy a symmetrical pattern of the blade rotation-wake of the compressed air, thereby reducing the strength or preventing the generation of the spinning mode so that the fan-stator interaction tones are significantly reduced.

Abstract: The invention relates to a radial compressor for producing a compressed gas flow, having a casing, an impeller which is arranged therein has a hub which carries blades. A ring is arranged in the flow duct concentrically with the hub, and is accommodated in a recess in the inner surface of the casing. The ring has wall slits on its inner surface. An annular gap is provided between that end face of the ring which is directed downstream and an end face of the recess which corresponds thereto and is directed upstream.

Abstract: Existing pressure oscillations created by axial or centrifugal fans in a diverging shroud are utilized to power a passive, acoustic jet, the nozzle of which directs high momentum flux gas particles essentially tangentially into the boundary layer of the flow in a diffuser, or a duct, the fluid particles in the resonant chamber of the passive acoustic jet being replenished with low momentum flux particles drawn from the fluid flow in a direction normal to the surface, thereby to provide a net time averaged flow of increased momentum flux particles to defer, even eliminate, the onset of boundary layer separation in the diffuser or duct. The passive acoustic jet is used in the vicinity of fan blade tips to alleviate undesirable flow effects in the tip region, such as leakage.

Abstract: A method of positioning arrays of stages of a turbine, particularly for aircraft engines; the turbine having an axis and a first, a second and a third array of blades arranged consecutively, and of which the first and third array have a relative velocity about the axis with respect to the second array; the method including the steps of determining the paths of wakes moving inside gaps defined by the blades in the second array and towards the third array; distinguishing, for each wake, a first and at least one second zone differing from each other; selecting one of the first and second zone; and regulating the position of the third array with respect to the first array by aligning the leading edges of the blades in the third array with the paths of the selected wake zones.

Abstract: Apparatus for reducing acoustic radiation which occurs during the operation of turbomachinery includes a shroud having a generally cylindrical inner surface and a coaxial rotor having a plurality of blades extending generally radially outwardly at equally spaced circumferential locations. The inner surface of the shroud is circumferentially contoured such that the tip clearance between each of said blades and the inner surface is caused to vary in a periodic manner upon rotation of said rotor. This creates a new periodic unsteady pressure field which is substantially equal to, and out of phase with, an existing periodic pressure field resulting from nonuniform inflow into the blades and results in the radiation from the turbomachinery of reduced blade rate frequency tones. The inner surface of the shroud may have a circumferential sinusoidal contour whose periodicity is an integral multiple of the number of the blades.

Abstract: A turbo machine comprising: an impeller having a plurality of blades therewith; a casing having a flow surface defined therein and being positioned with the impeller therein; and a plurality of grooves being formed in the flow surface of the casing, for connecting between an inlet side of said impeller and an area on the flow surface of the casing in which the blades of said impeller reside. Each groove has a length of at least part of which is oriented in an axial direction of the casing, a width measured in a circumferential direction, and a depth. The width of each groove is equal to or greater than the depth thereof.

Abstract: Turbine nozzles and turbine moving blades of an axial-flow turbine are provided which are capable of reducing a secondary flow loss with a simple structure. A nozzle blade passage formed by nozzle blades, an outer diaphragm ring and an inner diaphragm ring is structured in such a manner that the shapes of inner and outer walls of the nozzle blades are made to be irregular so that stepped portions (h1 at the root and h2 at the tip) each having curvature R are formed. The nozzle blades are formed in such a manner that ends (Zr, Zp and Zt) of an trailing edge of the nozzle blades are positioned at the most downstream position at the central portion of the nozzle blades. Moreover, relationships Zt<Zr<Zp are satisfied. Similarly to the nozzle blade passage, the moving blades are formed in such a manner that stepped portions h3 and h4 each having curvature R are formed in the moving blade passage.

Abstract: A controlled passive porosity system includes a lifting device for operation in a multi-phase liquid, the lifting device having a low pressure surface with a plurality of holes and a high pressure surface with a plurality of holes. A common plenum is disposed within the lifting device and allows fluid communication between liquid adjacent the high pressure surface and liquid adjacent the low pressure surface. When cavitation occurs adjacent to the low pressure surface, a series of valves associated with the holes allows selective and localized fluid communication between the surfaces. The fluid communication has the effect of raising the pressure of the liquid in the area of cavitation, thereby eliminating the cavitation. In an alternate embodiment, the lifting device includes a first plenum, a second plenum, and an intermediate fluid passage, the fluid passage fluidly connecting the first plenum and the second plenum.

Abstract: A gas turbine blade (10) which has a rounded trailing edge (18), is provided with a row of side by side arranged ceramic fibres (20) along the trailing edge (18). During operation of the turbine blade (10), the rounded shape of trailing edge (18) causes gasflows to break from the rounded edge before reaching the edge extremity. The presence of the fibres (20) prevent the formation of vortices in the gasflow, and thereby improve turbine efficiency.