Abstract: In a radiation pyrometer in a gas turbine engine, non-uniformly distributed changes in transparency of an optical element in the system, e.g. contamination of pyrometer objective lens L.sub.1 by combustion products from the turbine can be automatically detected. Lens L.sub.1 focusses radiation from target T onto an additional lens L.sub.2, which in turn focusses it onto the end E of the fibre-optic bundle F. Considered reciprocally, lens L.sub.2 images the end E of fibre-optic F into L.sub.1, so that although each optical fibre in end E receives radiation from all parts of target T, each fibre is imaged into a corresponding discrete area of L.sub.1. Fibre-optic bundle F comprises two (or more) sub-bundles F.sub.1 and F.sub.2 forming complementary sub-areas of end E and each sub-bundle is provided with its own photodetector P.sub.1, P.sub.2 and associated pre-amplifier A.sub.1,A.sub.2 adjusted so that their outputs A.sub.1 S.sub.1 and A.sub.2,S.sub.2 are matched at the non-contamination condition.

Abstract: A gas turbine engine infra-red radiation suppressor comprising an exhaust nozzle which is surrounded by a duct having a plurality of diffuser vanes at its downstream end. The exhaust nozzle and duct constitute an ejector which is powered by the exhaust efflux from the exhaust nozzle and draws ambient air into the duct through a gap between the exhaust nozzle and the duct. The air serves to cool the exhaust efflux and thereby reduce its infra-red radiation emission, and the diffuser vanes prevent direct line of sight of the hot exhaust nozzle.

Abstract: A combustion system comprising a tubular structure having an upstream end and defining within it a longitudinal passage. A tubular wall surrounding the structure and defining with the structure a further longitudinal passage, of annular cross section. The tubular structure having longitudinally extending, circumferentially spaced staves and longitudinally spaced coaxial hoops for the passage of fluid, the hoops being shaped to deflect substantially axially flowing fluid either inwards or outwards through the openings for discharge either inside or outside the assembly respectively.

Abstract: A gas turbine engine combustion apparatus for vaporizing fuel comprises a flame tube having an upstream end wall in which are located a number of flow guides having associated fuel ducts. In operation, fuel and air flow through the U-shaped duct in each guide and the fuel flows onto the inside of the end wall which acts as a vaporizing surface. The exterior of the end wall is cooled by a flow of air from the engine compressor. The invention avoids the need of vaporizing tubes within the flame tube.

Abstract: A method of vacuum moulding a composite material in which laid-up prepregs uncured resin and reinforcing filaments are placed on a heated former and covered with a sheet of a microporous film material a breather material and an impermeable membrane. The volume enclosed by the membrane is evacuated and the temperature of the mould increased to firstly outgas the resin and mould the prepregs to the shape of the former and finally cure the resin. The arrangement ensures excellent access of the prepregs to the vacuum during the moulding operation, thereby in turn ensuring that the resultant moulded article has very low porosity as a result of gas or vapor inclusions or is porosity-free.

Abstract: A method of manufacturing a mould for use in the investment casting of a hollow gas turbine engine blade in which a ceramic core is coated with a surface modifying material prior to the application of wax thereto to define a pattern of the object to be cast. The surface modifying material adheres to the core and provides a textured surface to which the wax keys, thereby preventing the wax lifting from the core.

Abstract: A method of treating filaments in order to increase their wettability by molten magnesium or an alloy thereof. The filaments are treated by coating them with particulate molybdenum trioxide, chromic oxide, ferric oxide or nickel oxide and subsequently infiltrating them with molten magnesium or an alloy thereof.

Abstract: An exhaust nozzle for a gas turbine engine comprising a duct 17(c) extending along an axis and having at a dowstream end thereof a mechanism for varying the geometry and area of the nozzle. The mechanism 28 comprises an axially translatable member 28 which has a face 29 extending in a direction transverse to the axis of the duct 17(c) against which pressurized gases flowing through the duct acts to balance out the loads on the flaps. A plurality of flaps 38,43 are spaced circumferentially around the axis of the duct. The flaps 38 are pivotally attached to the axially translatable member 28 and are provided with a cam follower 40 that co-operates with a cam 37. The cams 37 are fixed relative to the duct 17(c) and thereby defines the attitude of each flap 38 relative to the member 28. Seal plates 41 and 45 cover the gaps between adjacent flaps 38 and 43 to the seal plates 45. A plurality of struts interconnect the member 28 and the second flaps of the pivotal attachment of the flaps 43 to the flap 38.

Abstract: To enable shroud segments in a gas turbine rotor blade stage to operate at high temperatures with an adequate margin of safety, the shroud segments are mounted and cooled such that they are thrust outwards against seatings on surrounding high strength support structure by the gas pressure in the turbine passage, the shroud segments being provided with strengthening ribs or the like so that the outward thrust is transferred to the support structure through a plurality of load paths distributed over the outer sides of the shroud segments as necessary to avoid overstressing.

Abstract: A device is provided for passing a fluid flow across a static or rotating barrier wherein there is relative rotation between the flow and the barrier and wherein it is necessary to pass the maximum of flow through the device while keeping the diameter of the device as small as possible.One embodiment of the invention shown in FIG. 2 is a device for use in a gas turbine engine for supplying cooling air across a bolted joint to a space behind a cover plate 6 for cooling an engine disc. The device comprises a cylindrical receiver 11 having helical slots 13 running through it from one axial end face to the other, the helix angle being the same as the angle of the relative velocity vector between the flow and the slots.The helical shape of the slots ensures that the radius of the axis of each slot is constant along the length of the slots and allows greater slot height and hence greater flow area.

Abstract: A bypass gas turbine aeroengine has an exhaust mixer of the lobed type. The trailing edges of the confronting sides of each lobe are cut back to form notches or scallops in the sides of the lobe. To reduce aerodynamic losses and engine length and weight, the final turbine stage within the engine core is without outlet guide vanes so that the turbine exhaust gas stream issuing from the turbine has a substantial swirl component of velocity. Correction back to an axial direction of mean flow at the downstream end of the exhaust mixer nozzle is achieved by making the area of the scallop in one of the confronting sides of each lobe greater than the area of the scallop in the other side, the arrangement being such that sufficient of the turbine exhaust stream effuses from the larger scallops in directions contrary to the swirl component of velocity to produce the desired correction.

Abstract: An anti-icing system for a rotating blunt nose bullet of a gas turbine engine uses a rotating thermosyphon effect to create a circulation within a chamber connected for rotation with the nose bullet which draws hot air from a downstream region of the engine forwards to impinge on the nose bullet over a small region close to the engine axis. This avoids complications with separate hot air supplies or electrical heaters. As shown in FIGS. 3 and 4 of the drawings, in one embodiment of the invention the chamber consists simply of a tube closed off by the nose bullet and in combination with an engine main shaft which extends to the turbine region of the engine to draw hot gas from this region. Alternatively, the chamber may extend only to the downstream end of the first compressor.

Abstract: An exhaust mixer of the multi-lobed type for a mixed stream turbofan. In order to promote cortical mixing between the turbine exhaust stream and the fan air stream, the flow surfaces defining the lobes--and the troughs between the lobes--are twisted between their upstream and downstream ends so that opposing sides of each lobe and each trough comprise flow surfaces twisted in opposing senses. Noise suppressing jet propulsion nozzles can be produced with a similar construction.

Abstract: Apparatus for casting in a vacuum has a single evacuable furnace chamber 10 in which the charge melting and mould heating takes place. The charge is melted in a crucible 24 by an induction coil 20, while the mould is heated by resistance heaters 18 which are made from a FECRALLOY (Registered Trade Mark) material and can be operated in air or vacuum. In operation an empty mould 16 and a charged crucible 24 are positioned in the furnace 14 and the furnace chamber is sealed. The heaters 20 and 18 are operated while the furnace is pumped down. A control device 40 ensures that the charge is not melted until the appropriate vacuum level has been reached. After pouring the filled mould is withdrawn into a withdrawal zone 25 and as soon as the casting has solidified to the extent that air can be admitted to the chamber without detriment, the vacuum is released and the mould and charge replaced.

Abstract: Measurements of net thrust of an aeroengine on a thrust balance in a test cell must be corrected to gross thrust and it is also advantageous to detect changes in the airflows into and through the test cell of a magnitude sufficient to alter significantly the thrust correction factor. The invention can accomplish both these objectives by obtaining measures of the static pressures p.sub.i and p.sub.e in planes i and e respectively of the test cell using pressure probes 29 and 31, and utilizing the measures of p.sub.i and p.sub.e(a) in a function of the formD=f(p.sub.i, p.sub.e)where D is the thrust correction necessary to convert the net thrust reading on the thrust balance to gross thrust, and(b) to monitor variations of p.sub.i and p.sub.e with respect to the net thrust and detect such of these variations as are outside predetermined limits of variation, such detection indicating the occurence of changes in the characteristics of the airflows sufficient to alter significantly the value of D.

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.

Abstract: In order to combat flutter in rotor blades rotating within a duct bounded by a duct wall, the duct wall incorporates cavities tuned for resonance to a known flutter frequency. The cavities, which may comprise tubes immediately underlying a facing sheet of the duct wall, are arrayed around the circumference of the duct wall and communicate with the duct through openings in the duct wall near the tips of the blades. Flutter energy from the blades is transferred directly to the cavities by pressure waves and produces resonance in the cavities so that the flutter is damped aerodynamically. The invention is applicable to fans in turbofan engines.

Abstract: A fully aerobatic oil system for an aircraft gas turbine engine includes two re-circulatory oil systems the pumps for which have inlets 19 and 31a disposed at locations in the engine such that when one inlet is starved of oil, due to the aircraft attitude changing, oil will be supplied to the other one. In a specific embodiment, the breather of the oil system 30 (FIG. 2) is positioned in the engine gearbox and provides the lowest pressure point in the system. The breather is of the type described in UK Pat. No. 1,508,212 which includes a rotating compartment filled with RETIMET (Regd. Trade Mark). The bearing chambers are pressurised with air from the engine and the air/oil mixture in each chamber is vented via a first series of vents 24 to the breather where the air and oil are separated. The breather is used as a pump to pump the separated oil back to the oil tank via a scavenge line 35 which includes a relief valve 38.

Abstract: To minimize leakage of high pressure compressor air from the high pressure side of an aixal flow compressor rotor back to the low pressure side by the outside of a ring of shroud segments, the shroud segments, or at least their front and rear edge portions, are made relatively thin and springy and are retained to a carrier ring so that those edge portions are held sprung against circumferentially extending sealing lands on the supporting structure. It is arranged that the bending in the shroud segments is effective to accommodate predetermined manufacturing tolerances in the shroud segments and the supporting structure, thus avoiding unwanted clearances at the sealing lands. Alternatively, or in addition, the bending can accommodate reductions in bending due to differential expansion between the components during operation of the compressor rotor at elevated temperatures.

Abstract: A fuel control system for a gas turbine engine comprises a main fuel controller and an overspeed limiter arranged to prevent overspeed of the engine in the case of a fault in the main controller. To ensure that the limiter is tested regularly, the logic capacity of the main controller (which is preferably a digital electronic unit) is used to vary the parameters of the limiter each time the engine is started up. This variation is such as to cause the limiter to operate and produce an observable reduction in the engine speed which is detected by the main controller and uses as an indication of correct limiter operation.