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 course of the leading edges of turbomachine components, such as rotor blades and stator vanes is defined mathematically exactly and repeatedly as well as aerodynamically advantageously by the respective axial coordinate in the direction of the machine axis in relation to the blade height in percent, extending from the blade tip as per equation (1): axial ? ? coordinate ? [ % ? ? blade ? ? height ] = 1 5 ? ? extension ? [ % ? ? blade ? ? height ] ? ? tan ? ? sweep tip ( 1 - ? - 5 ? ( 100 ? % - blade ? ? height ? [ % ] ) extention ? [ % ? ? blade ? ? height ] ) and extending from the blade hub as per equation (2): axial ? ? coordinate ? [ % ? ? blade ? ? height ] = 1 5 ? ? extension ? [ % ? ? blade ? ? height ] ? ? tan ? ? sweep hub ( 1 - ? - 5 ? ? blade ? ? height ? [ % ] extention ? [ % ? ? blade ? ? height ] ) where sweep

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: An aerodynamically shaped supporting and fairing element (2) is arranged in the bypass duct of a gas-turbine engine and includes a supporting structure (3, 4) of aerodynamically shaped leading-edge and trailing-edge parts (7 and 8) connected by a connecting web (9) to form a recess (10) which extends on both sides of and longitudinally to the supporting structure. A fairing inlay (11) having an aerodynamically shaped outer contour and flushly adjoining the leading-edge part and the trailing-edge part (7, 8) is detachably fitted, with the fairing inlay being made in the same material or in a lighter material having reduced sound reflection or even a sound-absorbing effect. The connecting web (9) can include cavities (14). Such a supporting and fairing element features reduced weight, generates less noise and can easily be maintained and repaired.

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 comprises guide stator vane support struts (3) integrated into the guide blade ring at regular intervals respectively in place of a guide stator vane, said 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. Due to the integration of the support struts into the guide blade ring in place of the guide stator vanes and the inventive embodiment thereof, the weight of the engine and the flow losses can be reduced.

Abstract: On an axial-flow compressor, more particularly one for an aircraft gas-turbine engine, including at least one rotor disposed in a casing and having compressor blades extending from a rotor hub as well as one stator, a slot-type recess (14) is formed in the compressor blades (7), with the recess originating at the trailing edge (13) of the compressor blades (7) and immediately adjacing the rotor hub (6). This enables secondary flows occurring in the transition area between rotor hub and compressor blades and, consequently, rotor losses to be reduced, thereby providing for increased compressor efficiency and reduced fuel consumption.

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: On a propeller, more particularly one for aircraft applications, an efflux slot (6) originating at the propeller blade trailing edge (5) is provided in the transition area between propeller blade (3) and hub (2) for improving the secondary airflow and dimensioned such that precisely the air boundary layer, but not the main flow, is allowed to flow off. Propeller losses are reduced and propeller thrust and efficiency are increased.

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: With a turbofan engine the inner sidewall (2) of the bypass duct (3) features, downstream of the flow divider (11), a straight, conically widening form which, downstream of the supports, struts and guide vanes (7), smoothly transits into a circular-arc transition area (22). The form of the engine fairing (1) is adaptable to the shape of the inner sidewall (2), so that the bypass duct (3) has a conically flaring shape. In the interior of the engine, the gearbox (14) and/or other installations are arranged in a space existing between the inner sidewall (2) and a separating wall (4) confining the core-flow duct (5). The conical design of the bypass duct with the circular-arc transition area enables the gearbox and other installations to be arranged in the engine interior and, furthermore, is aeroacoustically and aerodynamically favorable.

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 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 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 compressor for a gas turbine, in particular an aircraft gas turbine, has a rotor hub carrying rotor blades, a stator equipped with stator vanes, a shroud associated to the stator vanes, and an arrangement providing sealing between the shroud and rotor hub to prevent leakage. To achieve almost complete suppression of leakage air and, concurrently, simplification of design and manufacture, the sealing arrangement (20) between the shroud (8) and rotor hub (3) is formed by a discharge arrangement (23) for leakage air.

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 includes 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: A gas-turbine axial compressor has a casing 3 and a rotatable shaft 5, 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 further part of the axial length of the stator 2, at which no shroud 21 is arranged, radially adjoins a rotor platform 28 formed by a rotor hub 8.

Abstract: A centrifugal compressor includes a rotor 2, which is rotatably borne in a casing 1. A fluid gap flow between the casing 1 and the rotor 2 is extracted via cutouts in 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 (I) 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 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.