A rotor assembly may be used in a gas turbine. The rotor assembly includes rotor segments arranged in succession in an axial direction and interconnected in the axial direction by a tie-rod, and a rotor segment disposed forwardly in the axial direction having a first contact surface and a rotor segment disposed rearwardly in the axial direction having a second contact surface. The first and second contact surfaces are at least partially in contact with each other, are substantially annular in shape, and extend in a radial direction and in a circumferential direction. The first contact surface and/or the second contact surface extend at least partially obliquely relative to the radial direction. An angle is formed between the first contact surface and the second contact surface when viewed in a sectional plane defined by the axial direction and the radial direction.

A rotor assembly may be used in a gas turbine. The rotor assembly includes rotor segments arranged in succession in an axial direction and interconnected in the axial direction by a tie-rod, and a rotor segment disposed forwardly in the axial direction having a first contact surface and a rotor segment disposed rearwardly in the axial direction having a second contact surface. The first and second contact surfaces are at least partially in contact with each other, are substantially annular in shape, and extend in a radial direction and in a circumferential direction. The first contact surface and/or the second contact surface extend at least partially obliquely relative to the radial direction. An angle is formed between the first contact surface and the second contact surface when viewed in a sectional plane defined by the axial direction and the radial direction.

A fan module for an aircraft turbomachine includes a fan with a disc carrying fan blades and a rotor of an electric machine. The rotor has an annular shape and is mounted coaxially downstream of the fan. The module further includes an annular support for the rotor with a downstream end fixed to the rotor and with an upstream end fixed to the fan disc. The support is fixed to the disc by shear bolts configured to break when a torque transmitted by the disc to the support exceeds a predetermined threshold.

A fan module for an aircraft turbomachine includes a fan with a disc carrying fan blades and a rotor of an electric machine. The rotor has an annular shape and is mounted coaxially downstream of the fan. The module further includes an annular support for the rotor with a downstream end fixed to the rotor and with an upstream end fixed to the fan disc. The support is fixed to the disc by shear bolts configured to break when a torque transmitted by the disc to the support exceeds a predetermined threshold.

A turbine shaft of a turbomachine extending along a longitudinal axis (A) includes a main tubular portion with a length L1 and a maximum outer main diameter. The shaft further includes a secondary tubular segment with a length L2 and a maximum outer secondary diameter greater than the maximum outer main diameter. The secondary tubular segment extends radially from the shaft such that the main tubular portion extends on either side of the secondary tubular segment along the longitudinal axis (A). The secondary tubular segment includes a weak point that is configured to break the shaft into two distinct parts when the value of a tangential stress applied to the shaft exceeds a predetermined threshold value, and thus cause retraction of the turbine.

A turbine shaft of a turbomachine extending along a longitudinal axis (A) includes a main tubular portion with a length L1 and a maximum outer main diameter. The shaft further includes a secondary tubular segment with a length L2 and a maximum outer secondary diameter greater than the maximum outer main diameter. The secondary tubular segment extends radially from the shaft such that the main tubular portion extends on either side of the secondary tubular segment along the longitudinal axis (A). The secondary tubular segment includes a weak point that is configured to break the shaft into two distinct parts when the value of a tangential stress applied to the shaft exceeds a predetermined threshold value, and thus cause retraction of the turbine.

In a turbine designing and manufacturing method attendant on a material change of a rotor disk of a turbine rotor, a temperature rise time ratio is determined which is a desired ratio of a temperature rise time of the temperature of the rotor disk from a first temperature to a second temperature after the material change to the temperature rise time before the material change. An inter-surface distance between surfaces on upstream and downstream sides of the rotor disk after the material change is determined, and a shape of the rotor disk after the material change is determined based on the inter-surface distance. The turbine is designed based on the determined shape of the rotor disk. After the material change in the shape determined in the designing process, the rotor disk and the turbine are manufactured based on the result of the designing process.

A compressor rotor for turbomachinery, such as a compressor, is provided. Disclosed embodiments can benefit from seal elements that may be arranged to inhibit passage onto respective hirth couplings of process fluid being processed by the compressor. A seal element may be affixed to adjacent rotor components (e.g., adjacent impeller bodies) by way of a slip or interference fit connection to one of the adjacent components and may be affixed to the other adjacent rotor component by way of a elastically flexible frustoconical inner surface of the seal element that permits the seal element to be placed in a spring-loaded condition, which generates a biasing force to circumferentially clamp onto a frustoconical outer surface of the other adjacent rotor component. This arrangement is conducive to user-friendly assembly/disassembly of the seal elements with respect to the adjacent rotor components.

Mini-disks of gas turbine engines are provided having an axially extending portion extending axially with respect to an axis of the engine, the axially extending portion configured to engage with a hub arm of a compressor of the engine, a radially extending portion extends radially with respect to the axis, the radially extending portion configured to engage with an attachment of a turbine disk of the gas turbine engine, an intermediate portion extending between the axially extending portion and the radially extending portion, and at least one mini-disk connector configured to engage with a portion of the turbine disk of the gas turbine engine to prevent radial movement of the mini-disk during operation.

A gas turbine engine includes a compressor section including a first compressor, a turbine section including a first turbine and a second turbine, a first shaft and a second shaft, the first shaft interconnecting the first turbine and the second compressor, and a geared architecture. The first shaft is supported on a first bearing in an overhung manner. A performance ratio is between 0.5 and 1.5.