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.

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 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: 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).

Abstract: A gas turbine combustion chamber with a burner 7, includes means for the supply of fuel and an atomizer 6, wherein the means for the supply of fuel are provided such that the fuel is injected in areas with maximum airflow velocity.

Abstract: A protective ring (4) for the fan blades (2, 3) of a gas turbine engine includes a material combination of metal bands and polymer-bonded fiber-weave layers interconnected in alternating arrangement. The protective ring thus formed, which is characterized by tough-elastic properties and low weight, can be used as a full containment (4?) with sufficiently large wall thickness or as a penetration containment (4?) with smaller wall thickness and with a trapping layer arranged on the outer circumference. As a full containment, the protective ring has impact-absorption capacity, as a penetration containment it prevents a further destruction of the broken-through protective ring.

Abstract: A combustion chamber suitable for a gas turbine engine is provided with at least one Helmholtz resonator having a resonator cavity and a damping tube in flow communication with the chamber interior. The damping tube is provided with at least one cooling hole extending through its wall for improved damping and cooling.

Abstract: A fluid-flow machine includes at least one rotor and a free number of stators flown by a fluid, with at least one blade thereof positioned on throat-confining surfaces provided with both a device for fluid removal from the flow path and a device for fluid supply into the flow path (bi-functionality). The machine includes at least one line associated with the device for fluid removal for returning the removed fluid to an upstream position in the flow path and at least one further line associated with the device for fluid supply for supplying the fluid from a further downstream position in the flow path.

Abstract: A hollow turbine blade cooled with compressor air is divided into a cooling air chamber (6) and into impingement air cooling chambers (8, 9) by inner, supporting partitions (3, 4). The cooling air is conveyed from the cooling air chamber into the impingement air cooling chamber via impingement air channels (7) provided in the partitions. The impingement air channels are concave with regard to the adjacent outer wall (2) of the blade airfoil (1) and arranged completely in the hot area near the outer wall and, in addition, have an oblong or elliptical cross-section whose longitudinal axis agrees with the radial orientation of the turbine blade. By reduced stress concentration in the area of the impingement air channels, the fatigue and creep characteristics are improved and life is increased.

Abstract: A gas turbine combustion chamber includes combustion chamber tiles 3 attached to a supporting structure 6 of the gas turbine combustion chamber, with each tile possessing at least one dilution air hole 4 which is flush with a dilution air hole of the supporting structure 6, wherein a diameter of the dilution air hole of the supporting structure 6 is considerably larger than a diameter 14 of the dilution air hole 4 of the combustion chamber tile 3.