Abstract: A turbomachine blade is repaired by making a leading edge cutback between two points (A) and (E) where point (A) lies within a range of relative blade height (h/H) of 0?(h/H)point A<0.85, and point (E) lies within a range of 0.95<(h/H)point E?1.0, with (h) being a blade height coordinate which originates at an end of the blade closer to starting point (A), and (H) being the blade height. A cutback trace is defined by (s/SAE) and (c/Cm), with (c) being a local cutback length in meridional flow direction, (Cm) being a meridional chord length, (s) being a local cutback length in radial direction, and (SAE) being a radial cutback length, where (c) gradually increases from point (A) to its maximum at 0.60<s/SAE<0.95, with (c/Cm) there being ?0.25, and then decreasing to ?0.05.
Abstract: An assembly fixture for the separable connection of a first component 1 and a second component 2, the first component 1 having an essentially cylindrical end-face recess 3, with the second component 2 having an essentially cylindrical protrusion 4 which fits into the recess 3, with the protrusion 4 of the second component 2 being conical at a side opposite the cylindrical side, with the first component 1 being provided with a conical face at a side facing the cylindrical side of the recess 3, and with a conical clamping body 5 which, together with the protrusion 4 of the second component 2, can be fitted into the recess 3 of the first component 1 and clamped against the first component 1.
Abstract: An inner shroud (6) for the stator blades (3) of the compressor of a gas turbine is a plastic component and includes certain additives, which may be concentrated in partial areas for specific functionality, and/or pre-manufactured inserts for sealing, ensuring sliding properties, wear protection and/or setting heat expansion. Such an inner shroud made of plastic material fulfills all requirements usually imposed on inner shrouds made of metal, while being simply producible and having low weight.
Abstract: A hollow fan blade for the fan of an aircraft engine includes a blade base section (2) and a blade tip section (5) which are assembled at opposing faying surfaces (6, 6?) by a joining process. Starting from the respective faying surface, cavities (7, 9) are produced in the two blade portions (2, 5) which are dimensioned in accordance with the loads occurring in the respective blade areas. The joining weld (4) is in a low-loaded area. The blade base section and the blade tip section can be constructed of solid material with formed-in cavities. The blade tip section can also be a sheet-metal structure.
Abstract: A blade retention arrangement axially fixes blades (1) to a disk (2). A profiled blade root (3) extending from the airfoil is inserted into a conformal axial slot (5) in the disk (2) and is axially secured by means of a split retaining ring (4), with the retaining ring (4) at least partly engaging a groove (6) on the blade root (3). The disk (2) is provided with a circumferential annular groove (7) which has a radially inward seating face for the retaining ring (4) and the retaining ring (4) is elastically deformable at least in the axial direction.
Abstract: This invention relates to a method for the manufacture of a gas-turbine combustion chamber which consists of individual wall sections produced by casting. To make up the combustion chamber, the wall sections are joined by laser welding. Preferably, the individual wall sections are segments of the annular or circular combustion chamber, with the casting material of the wall sections being a high-temperature nickel-base casting alloy.
Abstract: On a fixture for electro-chemical machining for the production of long, curved cavities (11) in a component (12), the working electrode (3) of the electro-chemical machining tool (1) is connected to a guide body (1a) which moveably rests against the inner walls of the respective cavity by means of power-actuated guide elements (2), with the guide body being connected to a flexurally soft and torsionally stiff guide linkage (5) coupled to a feed and rotary drive. Control of the movement of the guide body is accomplished in dependence of at least one wall thickness measured with a measuring device (13, 14) during feed movement.
Abstract: On a compressor with compressor blades, a flow transition fixation mechanism (4) is provided on the suction side (2), approximately parallel to the leading edge (3) and upstream of the compression shocks acting upon the blade, which prevents the transition point from the laminar to the turbulent boundary layer flow from oscillating, thus suppressing oscillation of the compression shocks and their coupling effect with the natural frequencies of the compressor blades.
Abstract: An inexpensively manufacturable arrangement for the cooling of thermally highly loaded components, for example the combustion chamber of a gas turbine, comprises cavities (4) provided on the thermally loaded inner surface/air exit side (3) of the wall (1) of the component which are supplied with a cooling medium from the outer surface/air inlet side (5) via cooling air openings (6) and which are covered by a heat resistant wire mesh (7) of a certain permeability. The cavity, which acts as pressure accumulator, forms, in connection with the wire mesh, a uniform pressurized-air film (8) for the cooling and heat shielding of the wall surface.
Abstract: In a fault detection logic for an engine controller for the detection of too low or too high a thrust relative to an intended positive or negative acceleration, the positive or negative error of the engine pressure ratio and the positive or negative acceleration of an engine shaft are simultaneously compared with a specified threshold, with the simultaneously determined transgression of both thresholds safely identifying a loss of thrust control. Disturbances caused by engine surge or signal noise are eliminated.
Abstract: A valve housing 1 with valve seat 10 incorporated in a fluid line features a closing element 2 with a valve disk 11 and a valve stem 12 extending through and sealed in the housing wall which, on the outer side, is locked by means of a locking pin 17 actuated by a solenoid. Owing to the geometry of the valve disk and the annular gap 20 as well as the action of a spring element 23, the inlet-side pressure acting upon the closing element 2 exceeds the outlet-side pressure, as a result of which the fluid flow via the annular gap is interrupted in only a few milliseconds merely by the permanently present, higher mechanical pressure if the valve stem is released merely by the movement of the locking pin effected by a small and lightweight solenoid 18, actuated on the basis of a signal indicating an event.
Abstract: A tube-type vortex reducer for the conduction of cooling air in a compressor 1 of a gas turbine with at least one radial secondary air tube 2 arranged in a disk interspace 3, includes a deflector arranged in a discharge area of the secondary air tube for the deflection of the secondary air flow into an axial direction or away from an axial direction.
December 4, 2002
Date of Patent:
January 9, 2007
Rolls-Royce Deutschland Ltd & Co KG
Stefan Hein, Manuela Stein, Dieter Peitsch
Abstract: A compressor blade for an aircraft engine includes a blade core made of a fiber compound material and a metallic enclosing structure. The enclosing structure is of a multi-part design and includes blanks (4, 5) attached to the blade core by a metallic weave (6) on the suction side and on the pressure side, with the two blanks being firmly connected at the aerodynamically shaped leading edge of the compressor blade by to a leading-edge former (3). Depending on the specific loads applied on the pressure side and on the suction side, welds attaching the blanks to the leading-edge former (8, 9) are offset relative to each other and the material thicknesses of the components of the enveloping structure are matched to each other.
Abstract: In a gas turbine, the compressor load is reduced and the turbine load is increased for controlled bearing load reduction by forming a low-pressure chamber (8) that is shielded from the compressor air at the compressor (1) and a high-pressure chamber (9) that is supplied with compressor air at the turbine along the high-pressure shaft (2). The pressures can be controlled within specific limits. Controlled bearing load reduction allows high speeds and a long service life of the bearings.
Abstract: A combustion chamber of a gas turbine includes starter film cooling of a combustion chamber wall 4 and several circularly arranged burners 7 with local maxima and minima in the intensity of the starter film 3 being provided around the circumference of the combustion chamber wall 4.
Abstract: On a high-pressure turbine rotor wheel, where a blade neck interspace (14) exists between two adjacent turbine blades (4) which is defined by the disk lobe (2) and the opposite blade necks (5) and blade platforms (6), part of the air supplied by the high-pressure cooling channel (9) is passed via a diverter cooling channel (10) into the blade neck interspace (14) to effectively cool the blades and the disk rim (1) in this area and to simultaneously supply the low-pressure channels (12) with cooler air and displace the hot sealing air. Thus, the life of the turbine rotor wheel is increased with simple means.
Abstract: A high-pressure turbine of a gas-turbine engine provided with a turbine disk (1) carrying rotor blades (2) and with stator blades (3), and with a lateral wheel cavity (5) being formed between the turbine disk (1) and the stator ring (4), wherein a seal (8) is provided in axial direction between stator blade platforms (6) and rotor blade platforms (7) and which is arranged radially outwardly adjacent a main gas duct (9).
Abstract: A mechanical-electronic shut-off system detects a shaft failure and initiates the shut-off of the fuel supply. It features, on the free end of a reference shaft (2) connected the to energy-consuming end of the respective shaft (1), an axially moveable signal trip element (13, 14) held under pre-load (12) whose locking arrangement (17,18) is released via a radial driver arrangement (17, 19) by rotary movement in the event of a shaft failure. The resultant relative rotation of the shaft (1) enables the signal trip element to move towards a sensor (21) or a switching element. An electric signal so produced instantly interrupts the further supply of fuel by means of an electronic control and avoids or controls a dangerous overspeed condition of the failed shaft.
Abstract: The exchange, removal or addition, as applicable, of media and/or power between the individual shafts of an engine, between individual engines and between the engines and the aircraft provides for additional degrees of freedom, enabling engine parameters to be addressed in terms of a reduction or avoidance of negative resonances or beats. It also provides an ability to alter thrust from engines of a multi-engined aircraft to reduce rudder trim.
Abstract: A tube-type vortex reducer conducts cooling air in a compressor (1) of a gas turbine. Radial secondary air tubes (2) are arranged in a disk interspace (5) and attached to a compressor disk (3) with their radial outward end sections. The end sections of the secondary air tubes (2) are attached to a separate carrier ring (6), which is connected to a compressor disk (3).