Abstract: A fluid-flow machine has at least one row of blades 5 having blade ends moving relative to one of a hub 3 and a casing 1, with a gap 11 positioned therebetween. At least one groove 7 extends essentially in a circumferential direction of the machine is in an area of the gap 11 along at least part of the circumference, with the extension of the groove 7 in the circumferential direction being large as compared to the extension of the groove 7 in the meridional flow direction. A cross-sectional area of the groove 7, in meridional view of the fluid-flow machine, essentially departs from a parallelogrammic shape and, due to its contour, is inclined in an upstream direction. A centroid of the groove cross-sectional area is provided upstream of the center of the groove aperture 12 on the main flow path.

Abstract: On a fuel supply system for a gas-turbine engine with a high-pressure supply line and a fuel metering unit, the fuel distribution system connected to staged or non-staged burners includes only one single fuel line (3) and the valve arrangements (7) connected to this fuel line and directly associated to individual burners (4), with the valve arrangements (7) each having several cutoff valves (8 to 10), which are—independently of each other—individually or simultaneously electrically settable to a closed or open limit position, with the cutoff valves each featuring a specific, yet different cross-sectional area. The simply designed, low-wear and low-maintenance fuel supply and distribution system, avoids coking in the one fuel line in the hot engine section, and enables a binary-coded fuel supply to the respective burner. With a sufficient number of valves in each arrangement, the fuel metering unit can be omitted.

Abstract: A gas turbine includes at least one multi-stage compressor unit 2, with the compressor unit 2 including several, independent compressor modules 3-5 which, independently of each other, are rotatably borne on a drive shaft 1 and are each engageable with the drive shaft 1 by variable transmission 9-11.

Abstract: A turbofan engine (1) includes at least one apparatus (30) for driving at least one generator (22, 35), with the engine including a pre-compressor (4), at least one compressor (10) and at least two engine shafts (13, 15) rotatably arranged in an engine casing. At least one first generator (22) is coupled via an auxiliary gearbox (21) with one of the engine shafts (13) and is electrically connected to at least one accessory (23). In order to generate electrical power to such an amount in a turbofan engine (1) that stable operation of the accessories (23), in particular the compressors, is maintained also in the low-energy range, the apparatus (30) includes at least one auxiliary turbine (31) arranged between the pre-compressor (4) and the compressor (10).

Abstract: On a platform cooling arrangement for the nozzle guide vane stator of a gas turbine arranged downstream of the combustion chamber, one or several parallel row(s) of cooling-air ejection ducts (10) are arranged continuously or in groups on the circumference. The cooling-air ejection ducts are angled relative to the axial direction at an angle (?) to produce a vortex structure on the surface of the platform (2) which, on the one hand, reduces mixing of the cooling air jets (11) with the hot gas flow (8) and, on the other hand, ensures complete cooling of the area of boundary layer separation (12) downstream of a boundary layer separation line (13) up to the suction side (14) of the adjacent nozzle guide vane (1).

Abstract: A fluid flow machine includes a main flow path in which at least one row of blades (3, 4) is arranged, and a blade shroud (2) which forms a confinement of the main flow path in the area of a blade row (3, 4) and is arranged in a recessed cavity (10) of a surrounding physical structure (11). The cavity (10) so formed between the shroud (2) and the surrounding physical structure (11) is provided with an annular connection (14) to the main flow path at least on the outflow side of the blade shroud (2), and with the blade shroud (2) being penetrated by at least one secondary flow path (16) which, commencing at a location of high pressure on the shroud (2), issues at a surface wetted by the main flow.

Abstract: A turboprop engine (1) includes an engine nacelle (3) and at least one bleed air line (25) on the low-pressure compressor (4) and at least one ejector (21) formed by a cooling air duct (24) and a nozzle (22) to create a cooling air flow within the engine nacelle during critical ground idle operation (controlled or uncontrolled), and without undesirably increasing fuel consumption or disturbing the work cycle of the engine (1). The ejector (21) is arranged within the engine nacelle (3) in the forward part of the turboprop engine (1), with the cooling air duct (24) appertaining to the ejector (21) connecting at least one air intake (23) disposed on the periphery of the engine nacelle (3) with the interior of the engine nacelle (3), and with the at least one nozzle (22) being arranged in the cooling air duct (24).

Abstract: On a burner for a gas-turbine combustion chamber which comprises a lean premix burner with centrally integrated stabilizing burner, a core air annulus (11) accommodating the atomizer nozzle (10) of the stabilizing burner is concentrically surrounded by a main air annulus (4) supplying the weak air-fuel mixture. In the adjacent issuing areas of the main air annulus and the core air annulus, a flame stabilization ring (13), which is heated by the combustion gases and whose cross-sectional surface increases in area toward the combustion chamber (5), is provided to produce an approximately hollow-cylindrical hot-gas recirculation zone (17) originating at the flame stabilization ring which ensures a stable flame formation throughout the range of operating conditions of the gas turbine.

Abstract: In the case of an ignition system for a gas turbine engine, the power (P1, P2) intended for the two ignition units (1, 2) is supplied via a power supply bus (10, 14) each comprising a standard bus (11, 11?; 15, 15?) and an emergency bus (12, 12?; 16, 16?), and a relay (13, 17) connected to the respective power supply line to the corresponding ignition unit. In the event of failure of power supply in a supply line for power P1 or P2, the corresponding relay will automatically (mechanically) switch over to the respective other power supply line, thus ensuring that the function of the ignition unit is maintained, even if—for whatever reasons—power supply to one or the other ignition unit is interrupted.

Abstract: A turbomachine with fluid removal, having at least one stator and at least one downstream rotor, with the stator being provided with stationary blades (1) and the rotor comprising several rotor blades attached to a rotating shaft, with a casing (2) confining the passage of fluid through the rotor and the stator in the outward direction, wherein, in an area of at least one blade (1) of the stator, a provision for fluid removal is provided adjacent to a suction side (3) of the blade (1).

Abstract: A passive internal control system for rotor blade tip clearance of a high-pressure turbine is based on inner rings that expand in a radial direction under the influence of heat and radially adjust the clearance delimited by liner segments (9) on the casing side. It comprises a radially expandable U-shaped downstream inner ring (10) mounted to inner platforms (7b) of guide vanes (7) on a side where the rotor does not have a static bearing, said ring providing expansion compensation in the axial and peripheral directions and forming a torsion box (8) with struts (13, 14) that absorbs rolling and tilting moments.

Abstract: A method for the manufacture of highly loadable components by precision forging, e.g., of gas turbine compressor blades with different stress zones, initially provides for the production of a preform in a one-step forming operation, e.g., casting or sintering, with further deformable volume zones with a certain volume concentration in the highly loaded areas. The preform is subsequently finish-forged at a deformation rate which is limited by the respective volumes and is merely directed to the load and strength-specific recrystallization and structural requirements in the critical stress zones of the component.

Abstract: A generator housing (1) of an engine generator is connected to a gearbox casing (17) via a cover plate (23) that bears the rotor shaft (5). The cover plate has a two-piece design and includes a bearing retainer ring (13) for a rotor shaft bearing (11) made of a rigid material and fastened only to the gearbox casing, and a heat-conducting ring (21) connected with both the gearbox casing and the generator housing. A centering cylinder (18) molded to the bearing retainer ring centers the generator housing and thus the stator (3) and rotor (4) while a spacer ring (20) and/or an adjustable spacer ring (10) ensure the exact axial positioning of the rotor or bearing retainer ring. The generator of this design is safe to operate and has a long service life.

Abstract: A tube-type vortex reducer for the conduction of cooling air in a compressor (1) of a gas turbine, having radial secondary air tubes (2) arranged in a disk interspace (5) and attached to a compressor disk (3) at their radially outward end sections (33). Radially inward sections (30) of the secondary air tubes (2) are located fittingly and radially outward in recesses (31) of locating pads (32) of a compressor disk (3), and the radially outward sections (33) of the secondary air tubes (2) are carried radially shiftable in recesses (34) of locating arms (35) of the compressor disk (3) and are secured against radially inward movement by locking elements (36).

Abstract: A fuel injection device for a gas turbine includes an airflow passage 1 whose walls 2 are provided with at least one fuel opening 3 for the injection of fuel into the airflow, with the center axes 4 of the fuel openings 3 being inclined at least in a circumferential direction.

Abstract: Single-crystal components of gas turbine engines, like turbine blades, are inspected in the installed state within the engine for cracking in certain critical areas using longitudinal ultrasonic waves. A first, rough orientation of the ultrasonic waves onto the critical area is accomplished under camera-visual control using an ultrasonic probe whose shape conforms to the respective component area and which, therefore, can be form-fitted to the component. For fine-positioning of the ultrasonic waves in the critical area, a reference signal is generated at a component-specific geometrical contour adjacent to the critical area by second ultrasonic waves emitted at a local distance to the first ultrasonic waves. The presence of this signal ensures the safe, disturbance-free detection of cracks in the critical blade area by means of longitudinal sonic waves. The invention includes an apparatus for the performance of the method.

Abstract: On a two or multi-stage turbine, expansion rings are provided in all stages at the sides of the rotors for passive, continuous running gap control whose thermal expansion and contraction behavior corresponds to that of the rotors and which are connected to radially moveable upstream and downstream stator vanes (5, 7). The downstream stator vanes are assembled to the upstream stator vanes via a bridge (16; 17, 18) which is axially and circumferentially fixed and located flexibly in the radial direction on the outer casing (10) of the turbine. The stator vanes (6) arranged between the rotor disks are integrally connected to the bridge or, as separate components, held axially and circumferentially on the bridge to take up rolling and tilting moments.

Abstract: Passive shroud cooling of rotor blades (5) is accomplished by a multitude of cooling-air flows introduced into the hot-gas flow upstream of the rotor blades on the outer circumference, these cooling-air flows, depending on the aerodynamic and geometrical conditions, being orientated at a defined radial angle and a circumferentially related tangential angle such that the cooling air primarily and essentially hits the thermally highly loaded bottom surface (11) of the outer shrouds (6) allover. In the apparatus for the performance of the method, cooling air passages (9) tangentially and radially orientated to the hot-gas flow are provided in an area of a casing section (3) which protrudes beyond the stator blades (1) in the downstream direction, this casing section (3) interrupting the transport of cooling air in the axial direction.

Abstract: A device for simulating sound produced by certain equipment, for example a rotor-stator arrangement of a turbomachine, or for generation of opposing sound fields for active sound control, including active sound reduction and active sound amplification, comprises flow obstacles (2) provided in a flow duct (1) flown by a fluid at which vortices (5, 6) are shed at a certain frequency depending on the shape and size of the flow obstacles and the velocity of flow. The quantity and spatial arrangement of the flow obstacles is selected such that a periodically spatially and temporally changing pressure field for the excitation of a sound field (8) of a certain modal content is produced by the entirety of the vortices shed. This sound field reacts synchronizingly on the vortex shedding. The resonant circuit so formed, whose vortex shedding frequency is in the range of the resonant frequency of the sound field to be excited, is the sound source.

Abstract: The compressor blades of an aircraft engine are, in at least one natural-vibration critical area, designed such that at the blade leading edge (6), the leading edge shock wave (14) attaches to the leading edge (6), as a result of which a laminar boundary layer flow (7) on the suction side (13) quickly transitions into a turbulent boundary layer flow (9) which is kept constant and prevented from re-lamination by the further, continuous curvature of the suction side. Therefore, the transition, whose periodic movement is also suppressed, cannot communicate with a suction-side compression shock (10), preventing the compression shock from augmenting the natural vibrations of the blade occurring under certain flight conditions. The blade leading edge can, for example, be designed as an ellipse with a semi-axis ratio equal to or smaller than 1:4.