Abstract: A gas turbine reheat system is provided with channel-shaped gutters, with a wall portion spaced from each gutter to form a space therebetween. The space is supplied with cooling air conveniently from the engine by-pass duct.
Abstract: In a by-pass gas turbine engine having a lobed type by-pass air and turbine exhaust mixer unit, the reheat gutters are arranged in the mixer in the turbine exhaust duct. The gutters extend radially and are substantially Y-shaped, each of the two legs of the Y extending radially along the sides of a mixer lobe.
Abstract: Elbow nozzles in VTOL aircraft engines, are highly stressed when operating, in that they act as high pressure vessels. The invention provides a means for varying the outlet area of such a nozzle without weakening its structure. A single flap is inserted inside the nozzle and connected to the nozzle extremity for pivoting about an axis which lies across the outlet between a position wherein its edges seal against the nozzle side and so restrict the nozzle outlet area and, a position wherein it lies parallel with a wall of the nozzle so that exhaust gas can pass both sides and out of a bigger outlet.
Abstract: A nozzle guide vane for a gas turbine engine comprises a hollow ceramic vane member which includes an aerofoil portion and ceramic platforms. The ceramic vane member is supported by engagement of its interior surface with external supporting surfaces on a metallic load-bearing spine adapted to be supported from fixed structure of the engine.
Abstract: An aircraft has a wing 10 supporting at its underside a gas turbine engine 13. A jet pipe 14 of the engine has an exhaust nozzle 15 of rectangular flow area defined by upper and lower panels 16,17 and end walls 18. The upper panel, which is a part common to the nozzle and to a trailing edge portion of the wing, comprises a pair of flaps 20,21 which are pivotal between a position of alignment with the wing 10 and a position in which they are inclined to the wing. A part of the gas flowing through the jet pipe is diverted through a passage 25 into a passage 26 defined between the flaps 20,21 so as to entrain and accelerate air 26A flowing over the wing 10. This increases the lift of the wing at the rear thereof and also increases the thrust developed at the nozzle 15. This in turn helps to compensate for a nose-up moment on the aircraft arising from the thrust 13D2 of the engine passing forward of the centre of gravity 9A of the aircraft.
November 23, 1981
Date of Patent:
November 2, 2010
Ralph Murch Deńning, Clifford Stanley Woodward
Abstract: An aircraft having a gas turbine engine with vectorable cold and hot nozzles. The aircraft is provided with a ducting in the wing which extends from an inlet facing downwards to an outlet in the upper surface of the wing directed rearwardly. The cold vectorable nozzles of the engine are designed so that they can be swung from a vertical downward position to provide vertical lift to a horizontal position directed rearwardly to produce forward thrust, to a position where they are directed upwardly and discharged into the ducting 24. In this way, the air discharged through the outlet of the ducting 24 enhances the aerodynamic lift produced by the wing at slower speeds by inducing the air flow over the wing. In addition, the forward component of the thrust produced in the ducting provides a forward thrust on the aircraft, whilst additional forward thrust is produced by swiveling the hot nozzles.
Abstract: A variable configuration final nozzle assembly for a combined rocket/ramjet engine consists of a fixed annular plug 20 together with upstream and downstream arrays of flaps 10,14 pivoted to the wall of a jet pipe 1 at their upstream and downstream ends 12,16 respectively. In the rocket mode of operation the upstream and downstream flaps are pivoted inwardly and locked in contact with the annular plug to form convergent and divergent portions 2,4 of a nozzle for the rocket exhaust which passes through the plug where a nozzle throat is defined by the internal surface of the plug. In the ramjet mode of operation the flaps lie alongside the jet pipe wall and ramjet exhaust gas flows through and around the plug. A second throat 50 is defined for the flow through the annular nozzle surrounding the plug and flow conditions in both parts of the flow are matched.
Abstract: A nozzle is provided which is capable of varying the direction of discharge of a fluid. As shown in FIG. 2 the nozzle 11 has sidewalls defining a circular opening 12 and a plug 14 mounted within the nozzle for transverse movement in the plane of the nozzle outlet. Thus a component of velocity can be imparted to the jet in any direction around 360.degree. in the plane of the nozzle outlet.Operation of the nozzle is by means of a plurality of piston-in-cylinder devices 18,20 acting between the plug 14 and the nozzle sidewalls. The plug is mounted on an arm 15 in a universal joint 16.
Abstract: A variable area nozzle for a turbomachine comprising, a duct 25, an axially movable shroud 28 located at the downstream end of the duct 25, at least one circumferential array of plates 30 mounted at their upstream ends on the shroud, and an outer wall 34 surrounding the shroud 28 and the plates 30. The outer wall 34 extends downstream beyond the downstream end of the duct 25 and a plurality of links 36 which are pivotally attached to the outer wall 34 downstream of the downstream end of the duct 25 are provided. Each link 36 is pivotally attached to one of the plates 30. An actuator means 50 which is operable on the shroud 28 to move the shroud 28 along the duct 25 is provided. Axial movement of the shroud cause the plates 30 and the links 36 to pivot about the point of attachment of the links 36 to the outer wall 34 to vary the area of the nozzle. Further circumferential array of plates 37,39 may be provided downstream of the plates 30.
Abstract: A gas turbine engine of the type comprising a first compressor 4, a second compressor 6, a combustor 8 and turbines 10 for driving the compressors 4 and 6 is provided with an air intake 16, leading to the second compressor 6. The flow from the first compressor 4 is discharged through vectorable nozzles 12,14. The nozzles 12,14 are movable between a first position where they discharge to ambient and a second position where they register with the intake 16 and the flow from the first compressor 4 discharges to the second compressor 6.
Abstract: A method of manufacturing a bladed disc from a fiber reinforced composite material which comprises injecting a mixture of short fibers and a matrix material into a dividable die so configured as to define the form of a disc having integral radially extending aerofoil blades. The mixture is injected into the die at such a location that the short fibers within the aerofoil-defining portions of the die are radially aligned. The die contains two support rings which are coaxially disposed within the bladed disc after injection molding so as to provide radial support for the aerofoil blades of the thus formed bladed disc. The method may be modified to provide the bladed disc with an integral shroud.
Abstract: A nozzle guide vane assembly for a gas turbine engine comprises inner and outer porous sheet metal platform rings and a plurality of aerofoils extending between the rings, each of the aerofoil portions having mounting means at its inner and outer extremities by which the aerofoils, and thus the rings, are supported from fixed engine structure.
Abstract: A digital noise generator for use in aeroengine noise testing can accurately and repeatably reproduce a standard noise spectrum so that differing noise analysis systems can be calibrated to a common standard. The noise generator comprises a pseudo-random binary sequence generator which is fed with a low clock frequency derived from a high frequency crystal oscillator. The generator produces a pseudo-random noise single bit signal which is fed to the most significant bit input of a digital to analogue converter, whose other inputs are fed with simulated tone data residing within an EPROM, the pseudo-random and tone noise thereby being digitally summed before conversion to an output analogue signal of desired spectral characteristics.
December 23, 1985
Date of Patent:
June 16, 1987
John D. Brown, Peter Gladdish, Michael A. McCormick
Abstract: An apparatus for use in the manufacture of brush seals which comprise a plurality of filaments of brush material sandwiched between two sideplates. The apparatus comprises a workplate onto which is mounted one of the sideplates of the seal. A plurality of first clamping members clamp against the sideplate. Each member can be lifted to enable a tuft of brush material to be inserted and clamped against the sideplate. To enable more than one layer of tufts to be built up on the sideplate a second clamping means is provided to enable the first clamping members to be released. An indexing mechanism allows the first clamping members to be moved relative to the clamped tufts and when the assembly of tufts is complete a second sideplate is clamped onto the tufts and first sideplate. The apparatus also includes means for cutting a prewound ribbon of adhesively bonded wires into short lengths to form the tufts.
Abstract: A gas turbine jet propulsion engine has a rectangular exhaust gas outlet 14. A pair of U-shaped tracks 21,22 one on each side of the outlet and which extend downstream of it, support rollers carrying the plates 24a etc. of a pair of articulated deflectors 24,26, one above and one below the outlet. The deflectors are translatable along the tracks between the final exhaust discharge opening of the engine. The configurations include a convergent divergent nozzle, a thrust vectoring nozzle and thrust reversing outlets.
Abstract: A reinforced tubular core for casting gas turbine engine blades with cooling air passages therein is disclosed. A method of casting is also disclosed in which the blades are directionally solidified to produce columnar grained or single crystal blades and in which non-linear passages can be produced. The problem in producing such articles is that the moulds and cores used in the casting process are held at temperatures in excess of 1500.degree. for long periods and presently used Silica cores deform during the process. Stronger cores of Alumina or Silicon Nitride cannot be easily bent were believed to be non-leachable from the casting. The present invention provides a core having a tubular silica sheath with a solid alumina rod inside it for support. The sheath can be bent and the straight alumina rods can be inserted from opposite ends of the sheath.
May 22, 1985
Date of Patent:
January 20, 1987
David Mills, Anthony T. Lindahl, Alan D. Kington
Abstract: A method is described for the manufacture of articles in defect immunized materials in which the defects are eliminated, or broken up and oriented in such manner as to minimize their harmful effects on the article. Referring to FIG. 1e, a rotor disc for a gas turbine engine is formed to an approximate shape by stacking together "sticks" 5 of material in an evacuated container and bonding them together by a hot isostatic pressing process. The "sticks" 5 are produced by extruding a starting body, for example, of powder material, to produce an elongation of up to twenty times and then cutting them to the appropriate length. By this means any non-metallic inclusions in the powder are broken up, inspection of the sticks and rejection of defective ones becomes easier, and the sticks can be oriented so that the effects of any remaining defects can be minimized.
Abstract: It is known to apply coatings which have desired characteristics, e.g., wear resistance, thermal barrier capabilities anti-oxidation, to articles the material of which does not have the desired characteristic, wherein the coating is applied directly to the article. Some article materials, e.g. titanium are highly reactive which immediately obviates several known methods of applying coatings. The invention uses a member (12) which would normally be preformed to mate with the article surface to be coated. The member (12) is manufactured from cheap, non-reactive material, as a base to which the coating (14) is applied by the most suitable of any of the known means. The member (12) is then placed against the article with the coating (14) therebetween and the whole hot isostatically pressed. The member (12) is afterwards removed, e.g. by machining or etching.
Abstract: The present invention provides a method for manufacturing a thin-walled ceramic casting mould which is particularly suitable for casting directionally solidified articles in which rapid cooling of the cast material is required. The mould is transfer moulded around a disposable pattern material and includes an integral core. The mould has an outer wall thickness of the order of 0.5 mm to 2.0 mm. The mould and core may be made of the same or different ceramic materials chosen for their strength or thermal conductivity. The method further provides for a series of high temperature disposable supports embedded in the disposable pattern material, providing support for the outer wall of the mould during the firing process. The disposable supports are disposed of at a temperature at which the mould has acquired self-supporting strength.
Abstract: A compound helicopter shown in FIG. 1 of the drawings has wings 12 in addition to a helicopter rotor 14 and has twin powerplants 16 each including a low pressure compressor 18, a gas generator 20, a power turbine 22 driven by the gas generator and connected through a gearbox 32 to drive the helicopter rotor, and a variable area final propulsion nozzle 24 which receives the exhaust from the power turbine. Augmentor wing flaps 28 are provided on the wings and fed with air from the low pressure compressor for providing additional lift and thrust from the wings. The flaps 28 are pivotally mounted on the trailing edge of the wing and are movable to a position where the trailing edges of the flaps 28 obturate the flow through the gap between the flaps 28. In this position the flaps provide a means of decelerating the forward speed of the aircraft.