Abstract: The present invention relates to a cooling device for a jet engine having an axial compressor with several compressor stages including a rotor with rotor blades, a stator with stator vanes and an annulus. In order to reduce the temperature of the components at the outlet of the high-pressure compressor by simple measures and hence to increase the efficiency of a jet engine, a slot-like branch opening surrounding the rotor for a cooling airflow diverted from the main airflow into a first cavity upstream of the rotor is provided upstream of the last compressor stage of the axial compressor, with passage openings being arranged in the rotor for passing on the diverted cooling airflow from the first cavity into a second cavity downstream of the rotor.

Abstract: The present invention relates to a gas-turbine balancing device with at least one annular groove provided on the outer circumference of at least one intermediate or high-pressure rotor of an intermediate or high-pressure compressor, with at least one balancing element being arranged in the annular groove, and with the balancing element being moveable in the circumferential direction along the annular groove and provided with a fixing device.

Abstract: A gas-turbine engine with at least one compressor and at least one bleed-air tapping device, which includes an annular duct in a radially outer wall of a flow duct, and with an annular closing element, which is arranged in the region of the annular duct and can be moved in a substantially axial direction from a closed position to an open position, with the closing element having an annular flow divider projection which in the open position projects in the flow duct.

Abstract: The present invention relates to a cooling device for a jet engine having an axial compressor with several compressor stages including a rotor with rotor blades, a stator with stator vanes and an annulus. In order to reduce the temperature of the components at the outlet of the high-pressure compressor by simple measures and hence to increase the efficiency of a jet engine, a slot-like branch opening surrounding the rotor for a cooling airflow diverted from the main airflow into a first cavity upstream of the rotor is provided upstream of the last compressor stage of the axial compressor, with passage openings being arranged in the rotor for passing on the diverted cooling airflow from the first cavity into a second cavity downstream of the rotor.

Abstract: The present invention relates to a gas-turbine engine with at least one compressor and at least one bleed-air tapping device, which includes an annular duct in a radially outer wall of a flow duct, and with an annular closing element, which is arranged in the region of the annular duct and can be moved in a substantially axial direction from a closed position to an open position, with the closing element having an annular flow divider projection which in the open position projects in the flow duct.

Abstract: The present invention relates to a gas-turbine balancing device with at least one annular groove provided on the outer circumference of at least one intermediate or high-pressure rotor of an intermediate or high-pressure compressor, with at least one balancing element being arranged in the annular groove, and with the balancing element being moveable in the circumferential direction along the annular groove and provided with a fixing device.

Abstract: An apparatus (3) retains bladed rotor disks (4, 6, 9) of a jet engine which engine includes an axial-flow compressor with several bladed rotor disks (4, 6, 9), at least one turbine section, at least one shaft and at least one tie-rod (12, 47), with at least two bladed rotor disks at the compressor inlet side (2) being designed as blisks (6), each being a single component integrally combining a rotor disk and circumferentially distributed rotor blades (5). So that apparatus for the retention of blisks can easily be mounted and demounted, the blisks (6) are releasably retained via the tie-rod (12, 47) and the other rotor disks (4, 9) are firmly attached to each other separately and independently of the tie-rod (12, 47).

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: An apparatus (3) retains bladed rotor disks (4, 6, 9) of a jet engine which engine includes an axial-flow compressor with several bladed rotor disks (4, 6, 9), at least one turbine section, at least one shaft and at least one tie-rod (12, 47), with at least two bladed rotor disks at the compressor inlet side (2) being designed as blisks (6), each being a single component integrally combining a rotor disk and circumferentially distributed rotor blades (5). So that apparatus for the retention of blisks can easily be mounted and demounted, the blisks (6) are releasably retained via the tie-rod (12, 47) and the other rotor disks (4, 9) are firmly attached to each other separately and independently of the tie-rod (12, 47).

Abstract: An assembly fixture for the separable connection of a first component 1 and a second component 2, the first component 1 having an essentially cylindrical end-face recess 3, with the second component 2 having an essentially cylindrical protrusion 4 which fits into the recess 3, with the protrusion 4 of the second component 2 being conical at a side opposite the cylindrical side, with the first component 1 being provided with a conical face at a side facing the cylindrical side of the recess 3, and with a conical clamping body 5 which, together with the protrusion 4 of the second component 2, can be fitted into the recess 3 of the first component 1 and clamped against the first component 1.

Abstract: For precision balancing of the rotor (3) that, in fully assembled condition of a gas turbine engine, is encompassed by an inner and an outer case, inserts (7) with a through hole (11) and a female threaded section are positioned at an even spacing in at least one circumferential groove (6) of the rotor disks (5) that receives the roots of the rotor blades (4). A temporarily installed guide tube (15) that runs through the openings (13, 14) connects to the through hole that in a certain rotor position is in line with the openings (13, 14) in the inner and outer cases (1, 2) for inserting and screwing in a balancing screw (16).

Abstract: An assembly fixture for the separable connection of a first component 1 and a second component 2, the first component 1 having an essentially cylindrical end-face recess 3, with the second component 2 having an essentially cylindrical protrusion 4 which fits into the recess 3, with the protrusion 4 of the second component 2 being conical at a side opposite the cylindrical side, with the first component 1 being provided with a conical face at a side facing the cylindrical side of the recess 3, and with a conical clamping body 5 which, together with the protrusion 4 of the second component 2, can be fitted into the recess 3 of the first component 1 and clamped against the first component 1.