Abstract: An apparatus for supplying inlet gas 29 and fuel 17 to a gas turbine engine 10, the apparatus comprising: a heat exchanger 28 for transferring thermal energy from the inlet gas 29 to a fuel 17 for the gas turbine engine 10, the heat exchanger 28 comprising a first input for receiving fuel 17, a second input for receiving inlet gas 29, a first output for providing inlet gas 31 after heat exchange to the gas turbine engine 10 and a second output for providing fuel 33 after heat exchange to the gas turbine engine 10.

Abstract: A gas turbine engine simulator comprising a simulator rotor disc which has substantially the same maximum external dimensions as a rotor disc and is manufactured from a material which has a density of less than 220 kg/m3. The simulator rotor disc is manufactured from a foamed plastic material with a closed cell structure. The simulator rotor disc is provided with a cavity in flow communication with a source of simulator coolant fluid, at least one flow outlet, at least one heater unit and at least one thermocouple mounted within said cavity for the measurement of simulator coolant fluid temperature within said cavity.

Abstract: A magnetic coil 32 for an electrical machine is constructed from two sub-conductors 34A, 34B, each in the form of laminae of conducting material such as copper sheet, coated with an insulating material. At one end of the coil, the sub-conductors extend by different distances, to provide space at 39 for a transposition of the sub-conductors.

Abstract: A combustion chamber inner casing (24) has a flange (28) on its downstream end. A retaining seal ring (40) is bolted to the flange (28) and traps a flanged seal ring (46) against the flange (28) in a manner that allows the seal ring (46) to move by expansion, relative to seal ring (40).

Abstract: A variable geometry pre-mixing fuel injector (50) for injecting a fuel/air mix in a downstream direction, comprising: an air inlet (60); a fuel inlet (58) positioned downstream of the air inlet (60); a duct (56) extending at least downstream of the fuel inlet (58) to define a fuel and air pre-mixing zone (62), that narrows to form an opening (64); and means for varying the flow of fuel/air mix from the pre-mixing zone (62) through the opening (64).

Abstract: A securing assembly comprises a first rotatable member defining a first recess. The assembly also includes a second rotatable member defining a second recess. The first recess has an engagement formation. Securing means are locatable in the first and second recesses for securing the first and second rotatable members to each other. An urging formation is provided on one or both of the securing means and the second rotatable member for urging the securing means into engagement with the engagement formation when the first and second rotatable members are rotated.

Abstract: A closure panel arrangement is provided in which a closure panel 6 can only be retained when appropriately aligned with engagement between clasp hooks 8 and pins 9 is achieved along with association between latch hooks 10 and pins 11. If there is not appropriate association and alignment then a bias 12, 22, 32, 42 creates an outward bias force upon the closure panel 6 whereby the panel 6 will continue to fall out. In such circumstances, the panel 6 cannot remain in position due to simply key fit interference/friction association of sufficient strength for non-operational conditions but which will fail during operation due to vibration or stresses.

Abstract: A turbofan gas turbine engine (10) comprises a fan rotor (32) carrying a plurality of radially extending fan blades (34). A fan blade containment assembly (38) surrounds the fan blades (34) and the fan blade containment assembly (38) comprises a generally cylindrical, or frustoconical, metal casing (40). The casing (40) comprises an annular member (54) positioned axially upstream of the tip (37) of the fan blade (34). The annular member (54) extends in a radially inwardly and axially downstream direction from the metal casing (40) towards the tip (37) of the fan blade (34). A set of acoustic panels (72C) is also provided. A set of acoustic panels (72B) is secured to the metal casing (40) by resilient fingers (80, 84) on the upstream and downstream ends (82,86) of the acoustic panels (72B) which locate on the annular member (54) and the acoustic panels (72C).

Abstract: A fluid flow control system (20) comprises a pumping assembly (22) for pumping a fluid (23) through an orifice (30), and a pressure measuring assembly (32) for measuring, and providing information on, the change in pressure across the orifice (30). The fluid flow control system (20) also includes a controller (36, 38) to adjust the pump assembly (22) in accordance with the aforesaid information and vary the rate of flow of the pumped fluid.

Abstract: Correcting unbalance with regard to rotors (11, 101) and shafts is important in order to improve performance of those rotors (11, 101) or shafts within a practical machine or engine. Previously, result orientated iterative balancing of the rotor or shaft has been achieved in which mass has been added or possibly intrusive and distorted adjustments in drag and stiffness presented through bearings or alteration of the rotor (11, 101) itself have been utilised to improve rotor balance performance. In the present invention, a dynamic bearing 13 is utilised in order to provide force components with some components in phase with rotor deflection in order to simulate negative stiffness and/or some force components in phase with rotor velocity in order to simulate negative drag upon the rotor (11, 101). In such circumstances, a more accurate determination of rotor unbalance is achieved to enable known techniques of perturbed boundary condition analysis for balance correction to be implemented more accurately.

Abstract: An aerofoil assembly, for example a bladed rotor assembly (40B) comprises a rotor (42) carrying a plurality of rotor blades (44), at least one of the rotor blades (44) having a coating (46) on the surface of the rotor blade (44). At least one of the rotor blades (44) has a coating (46) having a different thickness, a different area of contact with the surface of the rotor blade (44), a different position of contact on the surface of the rotor blade (44), a different shape of contact on the surface of the rotor blade (44) and/or a different composition compared to at least one of the other rotor blades (44). The coating (46) is applied in a non-uniform manner to reduce the vibration level of the rotor blade (44), or rotor blades (44), with the highest vibration response for a given excitation by changing the bladed rotor assembly (40B) mode shapes and the relative vibration of the rotor blades (44).

Abstract: A seal (40) provides sealing between two pressure zones (60, 62) and a rotating and non-rotating member (42, 52) by providing a first sealing member (46) between two seal lands (48, 50). A second seal (64) is mounted between the non-rotating member (52, 58) and one of the seal lands (50, 66) so as to control pressure around the seal lands (48, 50).

Abstract: Methods for estimating performance of and/or detecting faults in components of a multi-component system, where the performance of each component is defined by one or more performance parameters x related to measurement parameters z that can be expressed as a function of the performance and operating parameters defining an operating condition of the system. The methods include: defining a series of fault classes corresponding to possible outcomes of faulty components; creating an initial population of strings for each fault class, each including a plurality of elements corresponding to the performance and operating parameters, values being assigned to the string elements which represent estimated values of said parameters and are constrained only to indicate fault affected values for performance parameters of the fault affected component of the respective class; and optimising an objective function which gives a measure of the consistency between measured and calculated values of the measurement parameters.

Abstract: A method of securing a releasable component in an assembled engine by inserting a conduit through an aperture in a casing of the engine, the conduit being carried by a borescope and directing the borescope to the releasable component, and supplying an adhesive onto the releasable component through the conduit.

Abstract: A method to recover wax from wax patterns used in investment casting using a lost wax technique without damaging incorporated ceramic cores is disclosed that includes subjecting the wax pattern to specific conditions. These conditions include chilling a casting tool including the wax pattern and ceramic cores such that a brittle transition temperature for the wax material is approached. Once chilled the wax pattern is subjected to brittle fracture release of the wax from the underlying ceramic cores without damage to those cores.

Abstract: A spacing arrangement for a gas turbine engine 10. The spacing arrangement being arranged such that as a rotor blade 26 of a compressor 13 rotates, the rotor disc 28 will move outwards and forwards to maintain a substantially constant gap 36 between the rotor blade 26 and an inclined casing 20.

Abstract: Continuous engine monitoring with regard to gas turbine engines is difficult particularly when the engine is utilised for propulsion of an aircraft. Nevertheless, by use of transfer passages 4 to an arrangement located within the tail cone 9 of an engine, it is possible through a transparent test cell 6 and appropriate heating to generate sufficient infra red radiation as well as incandescent visible light for detectors 11, 16 to determine various constituent parts of the gas flow. Thus a photo detector 11 can provide through electrical signals, the proportion of soot/smoke particles in the gas flow through their incandescence, whilst use of filters 17 for particular spectral lines in an infra red radiation spectrum will allow detectors 16 to provide an indication through an electrical signal the proportion and presence of other constituents such as NO, CO and unburnt hydrocarbons in an exhaust gas flow.

Abstract: A sealing arrangement, for example between adjacent nozzle guide vane platform segments (2A, 2B) in a gas turbine engine, comprises a seal (12) which extends across a gap (13) between the platform segments (2A, 2B) and engages, at opposite ends, in grooves (8A, 8B). The seal (12) comprises a web element (14) having recesses (20A, 20B) at its opposite edges, which receive sealing elements (16A, 16B). The sealing elements (16A, 16B) may comprise a braided amorphous silica yarn or other filamentary temperature-resistant material.

Abstract: A fan rotor arrangement comprises a fan rotor (24) and a plurality of fan blades (26). Each fan blade (26) comprises a root portion (36) and an aerofoil portion (38). Each aerofoil portion (38) has a leading edge (44), a trailing edge (46) and a tip (48). Concave pressure surface (50) and convex suction surface (52) extend from the leading edge (44) to the trailing edge (46) of each aerofoil portion (38). An annular wall (54) surrounds the fan rotor (24) and fan blades (26) and an inner surface (56) of the annular wall (54) has a circumferentially extending groove (58). The circumferentially extending groove (58) is arranged axially, or chordally, between the leading edges (44) and the trailing edges (46) at the tips (48) of the aerofoil portions (38) of the fan blades (26).

Abstract: A seal system comprises a rotor and a stator. First and second cavities are defined between the rotor and the stator. A plurality of seals for inhibiting a flow of gas through the cavities are provided. Relative motion between the rotor and the stator can cause a flow of gas through the cavities via the seals. The. seals are arranged such that an increase in pressure in one of the first and second cavities is offset by a decrease in pressure in the other of the first and second cavities.