Abstract: Fan downstream guide vane profiles have an optimized form of skeleton line angle distribution in an area situated between an upper and a lower limitation as well as a specific thickness distribution superimposed on the respective skeleton line angle distribution. Such guide vanes are characterized by lower pressure losses and a larger working range than the known downstream guide vanes, thereby reducing fuel consumption of the engine and increasing the operating stability thereof.

Abstract: The present invention relates to a premix burner of a combustion chamber of a gas turbine with at least one annular duct for supplying air and fuel, including a radially outer and a radially inner combustion chamber wall relative to a burner central axis and with at least one swirler arranged in the duct, said swirler including several flow-guiding elements distributed around the circumference of the duct cross-section, characterized in that at least one radially inner duct wall is provided in the area of the flow-guiding elements with a concave recess of the annular groove type.

Abstract: A combustion chamber tile of a gas turbine includes a bolt for mounting the combustion chamber tile on a combustion chamber wall, with the combustion chamber tile being designed substantially plate-like and having on one side at least one mounting element, on which the bolt, which is provided as a separate component, is positively anchored.

Abstract: A method and an apparatus is disclosed for the operation of a turboprop aircraft engine provided with pusher propellers. In order to reduce thermal loading of the pusher propellers impaired by the hot exhaust-gas flow of the engine and increase the service life of the pusher propellers, cold air from the environment outside of the aircraft engine is fed into, and mixed with, the hot exhaust-gas flow passing the pusher propellers and their connecting structure before the hot exhaust-gas flow reaches the pusher propellers.

Abstract: A guide blade ring (1) arranged downstream of the fan in the bypass duct of a turbofan engine for untwisting the airflow generated by the fan includes guide stator vane support struts (3) integrated into the guide blade ring at regular intervals respectively in place of a guide stator vane, the guide stator vane support struts (3) having a larger thickness and chord length than the guide stator vanes (2). A flow channel (4), respectively defined between a suction side (7) of the guide stator vane support strut and a guide stator vane (2.1) adjacent thereto on the suction side, is expanded by a geometric modification of the guide stator vane (2.1) with respect to the remaining guide stator vanes (2) in accordance with the respective configuration of the guide stator vane support strut.

Abstract: In a jet engine with compressor air circulation provided for stabilizing the flow conditions, the compressed hot compressor air tapped from the flow path of the compressor (1) is first cooled in the bypass duct (3) and then resupplied to the upstream compressor via line (12). This reduces the thermal load of the blades and improves efficiency.

Abstract: On a turbofan engine, at least one of the downstream guide vanes (1) of the downstream guide vane assembly (2) arranged behind the fan in the bypass duct, and a fairing element (3) arranged behind a downstream guide vane, are provided as a combined—one-piece and aerodynamically shaped—vane and fairing element (4) functioning as both a downstream guide vane and a fairing element for installations arranged in the bypass duct or an aerodynamically shaped supporting strut. The one-piece configuration of a fairing element with upstream vane, i.e. the integration of fairing elements provided with a specific outer contour into the downstream guide vane assembly, results in lower pressure losses and reduced fuel consumption as well as reduced pressure effect on the fan and, consequently, increased operating stability of the fan, higher fan efficiency and reduced sound emission.

Abstract: The present invention relates to an assembly of an axial turbomachine, comprising at least one outlet guide vane of a compressor and a diffuser arranged downstream of the outlet guide vane in the flow direction. It is provided that the outlet guide vane is connected to the compressor and that the diffuser is connected to the combustion chamber, without there being a direct mechanical connection between the diffuser and the outlet guide vane. The invention furthermore relates to a method for manufacturing an assembly of this type.

Abstract: Axial-flow compressor including, within a compressor casing (4), at least one rotor (2) of rotor blades (3) connected to a drive shaft (1) and a stator (5) held on the casing inner wall and, associated to the rotor gap (6) between the blade tips and the casing inner wall, a flow pulse generator (7) for stabilizing the rotor gap flow, characterized in that the flow pulse generator (7) includes pulse channels (7a) arranged on the inner wall of the casing and extending upstream of the rotor (2) and tapering in flow direction to accelerate the wall-near flow (9), with the shape and size of the pulse channels (7a) being determined by circumferentially spacedly disposed, successive separators (7b) attached without gap on the compressor casing inner wall. The flow pulse generator (7) so designed, which is easily manufacturable, improves the stabilization of the rotor gap (6) flow, extends the operating range of the compressor and increases the surge limit.

Abstract: The aerodynamically shaped, symmetrical cross-sectional profile for the struts or the fairing (7) of struts and service lines in the bypass duct and the core-flow duct of a turbofan engine is defined by a course of the local thickness (T) and a position of maximum thickness (PMT) on both sides over the central chord (C) extending from the leading edge (LE) to the trailing edge (TE) of the strut or fairing (7). In order to avoid flow separations the course of the local thickness (T) between the position of maximum thickness (PMT) and a trailing-edge position (PTE), which is at 95 percent of the chord length, has a first turning point (WP1) with a trailing-edge thickness (TET) and a subsequent second turning point (WP2).

Abstract: In the area of the support struts and/or the aerodynamic fairings downstream of the stator vanes, the cross-section of the bypass duct of a turbofan engine is enlarged such that the pressure variations caused by the stagnation effect of the installations and reacting on the fan are reduced, enabling the fan to be operated with more efficiency and stability and finally the losses of the overall system and the fuel consumption to be reduced. The cross-sectional enlargement is accomplished by modifying the course of the wall in a limited area, actually by gradually enlarging the flow cross-section in the bypass duct in the axial and in the circumferential direction, with this enlargement being confined to the area around the leading edge of the support struts and/or the aerodynamic fairings.

Abstract: A device for mixing fuel and air of a jet engine includes at least one air-carrying duct and at least one further fuel-carrying duct. Guide elements extending over the duct height and distributed over the circumference of the duct are provided at least in the air-carrying duct, in the area of which guide elements a twist can be imparted to the air flowing in the air-carrying duct in order to improve mixing between the air and the fuel downstream of the ducts. The guide elements are provided at a trailing edge with at least one downstream extending projection and/or upstream extending recess designed at least approximately or at least in some areas with a sharp edge.

Abstract: A casing (2) includes at least one casing structure (casing treatment) for stabilizing a flow in an area of blade tips of rotor blades (4) in a fluid-flow machine, with the casing structure (casing treatment) being provided in at least one stage on an inner circumference of the casing (2). To provide a casing which improves compressor stability, is simply designed, features low weight and operates reliably without heating-up fluid in the fluid-flow machine, the casing structure is designed as a duct (20) which includes a first end (21) and a second end (22), with the first end (21) issuing into the interior of the casing (2) in the area of the blade tips of a rotor blade row and with the second end (22) being closed.

Abstract: A jet engine has a bypass duct limited by an inner wall and an outer wall and inside which a fluid flows. Between the inner and outer walls of the bypass duct a support unit is provided that includes strut-like support elements connected at opposite ends to the inner and outer walls, respectively. Central longitudinal planes of the support elements describe in the areas of the support elements facing the inner wall a positive acute angle with an engine axis, and in the areas of the support elements facing the outer wall a negative acute angle with the engine axis. Flow cross-sections are each enlarged in the area between the side surfaces of the support elements each describing an acute angle with the walls, starting from the areas facing the fluid flow in the direction of the areas of the support elements facing away from the fluid flow.

Abstract: A hub cone (20, 30, 40) is provided for an aircraft engine having a propeller (2, 32, 42) or a blower (fan) (3) enveloped by a casing (5). To reduce flow losses of the hub cone and increase efficiency of the engine, a contour of the hub cone is described by the following equation: S(x)=Rmax*{1?[(x?Lmax)/Lmax]2}1/M; where: S(x) is a shape of the cone defined along the machine axis x; Rmax is a maximum extension (26) of the cone in the radial direction; Lmax is a maximum extension (25) of the cone in the direction of the machine axis x and M is a quantity describing the shape S(x).

Abstract: A gas-turbine axial compressor has a casing 3 and a rotatable shaft 6, which form an annular duct, in which at least one stator 2 and one rotor 4 are arranged. A shroud 21 is arranged at a free end of the stator 2 vanes, which extends over part of the axial length of the stator 2. A rotor platform 23 of at least one rotor 4 extends axially beneath a further part of the axial length of the stator 2, at which no shroud 21 is arranged.

Abstract: At a propeller (1) or a fan (2) of an aircraft engine, part of the inflowing air—at the air stagnation point forming at the tip (8a) of the hub cone (8)—is conducted into an interior of a hub cone via air inlet openings (9) and, via air outlet openings (10) in an area with minimum static pressure at the downstream end of the hub cone, on a circumference of the latter and at a velocity essentially corresponding to the velocity of the air inflow, is injected into a thick boundary layer on the hub cone, essentially in the direction of flow, thereby accelerating the boundary layer to the velocity of the air inflow. This enables the inflow of air also to the root areas (4a, 6a) of the fan blades/propeller blades (4, 6), to be effected at an aerodynamically favorable, less steep inflow angle.

Abstract: A gas-turbine compressor has a casing (1) in which a rotor hub (2) is rotatably borne, and a compressor duct (9) being disposed between the casing (1) and the rotor hub (2), in which at least one rotor (4), which is rotatable about a machine axis (5), and one stator (5) are arranged. Recesses (12) of a bleed-air tapping device (6) are provided in the casing (1), which—in at least one circumferential area (11)—are arranged circumferentially to each other recess (12) and include a circumferential leading edge (16) in the circumferential direction and a circumferential trailing edge (17), each of which includes an identical angle ?E with the surface (18) of the casing (1).

Abstract: For the rotor and stator blades of turbomachines, more particularly of gas-turbine engines, an airfoil design is provided with a defined area of a skeleton line angle distribution for skeleton lines of airfoil sections near the gap. With the distribution of the dimensionless skeleton line angles (?) over the chord length (l) in a certain area between two limiting curves (7, 8) according to the present invention, and the corresponding course of the skeleton lines in a blade portion extending up to 30 percent of the blade height, a uniformed pressure distribution is ensured, minimizing disturbances and losses due to the influence of the gap.

Abstract: A bleed air outlet provided in the bypass duct of a turbofan engine includes a bleed air tube protruding into the bypass duct and a cover having a plurality of air outlet openings provided in the top of the cover, with the cover being conceived as an elongate, essentially oval, shell-like aerodynamic fairing element (3) which, while extending in the longitudinal direction of the bypass duct, is provided with a longitudinally convexly curved and transversely essentially straight top (10) and sidewalls (11) arcuately extending from the latter to an inner bypass duct wall. Such a design of the bleed air outlet reduces pressure losses, acoustic emissions and vibrations in the bypass duct.