Abstract: This invention relates to a blade of a turbomachine, more particularly an axial-flow high-pressure turbine, having at least one integrated cooling air duct issuing at the downstream back of the blade root when viewed in the direction of the working gas conducted by the blade. This invention also relates to a rotor disk of a turbomachine having several blades designed in accordance with the present invention. The cooling air duct in accordance with this invention branches off from a cooling air chamber provided in the blade root, from which chamber a cooling air duct issuing at the surface of the blade conducting the working gas is supplied with cooling air, where the cooling air duct issuing at the back of the blade root deflects the cooling air flow through it and provides an essentially continuous passage. The cooling air-duct is preferably made to converge in the direction from the cooling air chamber toward its outlet port and is shaped in accordance with design rules for axial-flow reaction turbines.

Abstract: A casing construction is provided on a gas turbine having a fan downstream of which a precompressor is situated, the precompressor casing being detachably connected with the casing of the compressor which follows the precompressor—viewed in the flow direction. An upstream end section of the precompressor casing is additionally linked to the compressor casing by at least two tension-rod-type tension devices which are situated essentially diametrically opposite one another with respect to the longitudinal axis of the gas turbine. Steel cables are preferably used as the tension devices and are fastened at least on one of the two casings by means of a tension-nut-type end piece. As a result, the precompressor casing can be constructed of a non-high-strength material and, in addition, can be linked by means of several pairs of tension-rod type tension devices to the casing of the compressor.

Abstract: An axially stepped annular combustion chamber, especially of an aircraft gas turbine, has an essentially independent main combustion chamber 5' as well as an independent pilot burner chamber 5. An appropriate design of internal limiting walls 6a, 6b of pilot burner chamber 5 ensures that the combustion gases enter the main burner zone 5' essentially in the radial direction. This ensures optimum mixing of the fuel with air in this main combustion zone and/or main combustion chamber 5', thus minimizing exhaust emissions and ensuring optimum temperature distribution at combustion chamber outlet 8. Internal limiting wall 6a can have a deflecting section 12 or outer wall section 6b can run at an angle to pilot burner lengthwise axis 3a, so that the cross section of pilot burner zone 5 is reduced in the flow direction.

Abstract: In connection with a turbine rotor disk in whose disk grooves air-cooled turbine blades have been inserted, at least two cooling air channels, respectively extending from the same disk front face, terminate in each disk groove. The outlet openings of two cooling air channels in each disk groove preferably lie essentially next to each other in a common sectional plane, which is normal to the disk axis. Because of this it is possible to supply a larger cooling air flow without drastically increasing the weakening, or respectively the stress of the disk in the groove bottom.

Abstract: A fuel injection device for a combustion chamber of a gas turbine with a liquid-cooled injection nozzle having a coolant tube which surrounds a fuel-conducting tube at a distance and which terminates in an annular chamber in the vicinity of the nozzle exit opening, or which constitutes this annular chamber which directly surrounds the fuel-conducting tube, wherein a separating wall element which surrounds the fuel-conducting tube is provided inside the coolant tube upstream of the annular chamber, viewed in the flow direction of the fuel, which divides the interior of the coolant tube into two chamber segments, wherein the first chamber segment is connected with a feed conduit and the second chamber segment with a removal conduit for the coolant.

Abstract: A turboprop engine having an oil cooler which is disposed in the engine pod and upon which cooling air acts. The oil cooler is disposed in a flow duct which at the rear opens into the environment and at the front can be connected alternately to an air-inlet opening or to the feed duct of the compressor of the aircraft gas turbine.

Abstract: The holding of a burner collar as well as of a heat shield in the end wall of an annular combustion chamber is described. An annularly constructed holding part is provided which, in segments, has flanks between which the burner collar is held by lugs arranged thereon. The holding part is placed on a ring attachment of the heat shield and, together with the heat shield, is screwed to the end wall.

Abstract: A tandem turbine-blade cascade for a turbine, turbo-engine or power engine includes at least two rows of blades disposed substantially directly in line with one another in the rotor or stator. The second row of blades has a larger number of blades than the first row of blades and the second row of blades are distributed nonuniformly. The spacing (s.sub.1) between two blades of the first row of blades is substantially equal to the sum (s.sub.2) of two successive blade spacings (s.sub.12, s.sub.22) in the second row of blades and the distribution quotient (s.sub.12 /s.sub.22) of the blade spacings in the second row is in the range between 0.4 and 1.0.

Abstract: A heat shield for the head area of a combustion chamber has as usual a through-hole for the burner. A continuous collar with air passage holes projects from the back side of the heat shield at the edge of the through-hole. Cooling air can flow through the holes into a ring-shaped channel arranged between the heat shield and the burner, then into the combustion chamber. This cool air flow lies as a cool air film on the surface of the heat shield. For that purpose, the cool air flow or cool air film swirls in the same direction as the combustion air supplied through the burner. To generate this swirling motion, the air passage holes in the collar are inclined in the radial direction. The heat shield is further provided with appropriate inclined effusion holes.

Abstract: In connection with a turbine rotating disk with disk grooves, constituted by disk fingers, for receiving turbine blades, as well as with measures for guiding a cooling air flow from a chamber located in front of the disk to one behind the disk, a cooling air transfer channel is provided in at least some of the disk fingers, which respectively starts at the front disk face, and makes a transition into a cooling air exhaust channel, also essentially extending in the radial direction in the disk finger, whose outlet opening on the rear disk face side lies closer to the disk axis than the disk ring section, which has the disk grooves and is widened in the disk axial directions. By means of this it is possible to convey a larger cooling air flow which, in an advantageous manner, indirectly aids the cooling of the disk in the area of the disk grooves.

Abstract: The invention concerns a shut-off valve member which can be moved by spring force into a position in which it interrupts the fuel line to the gas turbine. The shut-off valve member is held by a likewise pre-tensioned scissor-type mechanism in the position in which it opens the fuel line. If a faulty gas turbine component acts on the scissor-type mechanism, the shut-off valve member is released and moves abruptly into the position in which it interrupts the fuel supply. The gas turbine is thus switched off safely and reliably in the event of failure of a component, for example the turbine shaft.

Abstract: A combustion chamber of a gas turbine with a ring-shaped head section, in which a front plate for receiving burners is provided, which are supplied with a primary airflow via annularly arranged air inlet openings in the head section, wherein the stagnation points being formed on the exterior of the head section (1) in respect to the inflow to the air inlet openings form a ring, within which the air inlet openings are located, and wherein furthermore the front plate has heat shields provided in the combustion chamber burner section and fixed in place by means of bolt connections. In addition, mounting openings for the bolt connections (11) are provided in the head section, wherein the backup points being formed on the exterior of the head section in respect to the inflow to the mounting openings form a ring, which is not located outside the ring formed by the backup points of the air inlet openings.

Abstract: A flow-guiding body is designed as a pointed, substantially conical molded shell. The projection of its base surface is formed by a straight line and by a curve that interconnects the ends of the straight line. The curve forms no significant angles. The molded shell faces with its point the fluid flow that hits its outer side and may be used as a mixing element for gaseous fuel and air, as an air sprayer with flame-holder, as a mixing element for admixed air in combustion chambers, as a swirling element or as a shell-shaped air sprayer combined with a fuel film generator or a fuel pressure spraying nozzle.

Abstract: A heat shield for a gas turbine annular combustion chamber having a plurality of effusion holes (5), the central axes of which are inclined towards the heat shield surface and over which cooling air can penetrate from the rear to apply a film of cooling air to the hot surface. The surface is subdivided into sectors (7) and transition areas (10) between the sectors, the central axes of the effusion holes essentially being arranged in parallel to each other in a given sector or transition area. In addition, the central axes (6) of the effusion holes (5) located in each surfaces sector (7) are parallel to one another and extend substantially toward the associated corner area (8) and, in sections, extending approximately in a direction the same as the fuel combustion air swirl (4) in this sector (7).

Abstract: In order to cool the hot surface of the heat shield that surrounds the burner of a gas turbine annular combustion chamber as efficiently as possible, especially in the vicinity of the burner throughflow opening, a cooling air stream that escapes through a ridge that runs around the edge of the throughflow opening is guided by a guide rib in the direction of the throughflow opening. This guide rib is aligned essentially parallel to the ridge and on the combustion chamber side has an end that is bent in such fashion that the cooling air stream that flows into the gap between the ridge and the guide rib is deflected thereby in the direction of the hot surface of the heat shield.

Abstract: The boundary layer stimulation by blowing flow fluid into the proximity of the separation region on turbine blades is disclosed. In order to prevent a laminar boundary layer separation, the blowing-in takes place in the form of intermittent fluid pulses. A particularly advantageous arrangement for an axial turbomachinery rotor disk is provided for this purpose, on the end face of which a stationary ring is provided which has at least one passage opening with which inlet openings for the blade fluid ducts which are provided in the disk which passes by and by way of which the blowing into the proximity of the separation region takes place, become periodically congruent.

Abstract: A fixed bearing for a radial or diagonal flow compressor is guided in a damping bush whose axial position can be fixed by means of a reconciling plate to establish a desired running gap with respect to the compressor housing. The damping bush has a receiving device for a supporting element which rests against the reconciling plate, such that the supporting element and the reconciling plate are removable when the damping bush is disposed in the compressor housing. When the rotor is mounted, the required thickness of the reconciling disk can be determined and this reconciling disk can then be inserted without removing the rotor or the damping bush for the adjustment of the desired running gap.

Abstract: An aircraft gas turbine has a barrier air flow produced by the fan or a low-pressure compressor which passes continuously through the compressor bearing chamber, while the turbine bearing chamber is supplied with barrier air by the high-pressure compressor. The barrier air flow drawn from the turbine bearing chamber passes into an ejector which is also connected to the compressor bearing chamber so that, when the pressure is insufficient, the barrier air flow is drawn-off by the ejector.

Abstract: A cooled turbine blade is described that has shroud band segment in which three cooling air branch ducts extend essentially perpendicularly to the longitudinal axis of the blade as well as in the circumferential direction of the blade reinforcing band. Cooling air bores branch off the cooling air branch ducts and lead out on the surface of the shroud band segment and form film cooling holes there. A first branch duct is supplied with cooling air by a blade cooling duct which is situated in the leading area, while a second branch duct is connected with another blade cooling duct 13 and the third branch duct is connected with the second branch duct 14.

Abstract: The invention concerns an aerodynamic plain bearing for shafts (1) rotating at high speeds, with a bearing housing (4) having a cylindrical bore (3) and with a tangentially attached, elastic and damping structure (2) having segments (2" or 2a), which is supported at a plurality of circumferential points (2') in bearing housing (4) and forms a multiple converging running gap (6) with shaft (1) and damping gap (5) with bearing housing (4). A minimum number of components are used and radial installation space is also minimized. Segments (2" or 2a) of the structure are thin-walled, elastic sleeve sections with a uniform or profiled wall thickness and a low coefficient of friction, as from titanium nitrite, and simultaneously act as damping and spring elements.