A gas turbine engine includes a fan section, a compressor section, and a turbine section. The fan section has a plurality of vane assemblies spaced circumferentially about an engine axis. The vane assemblies each include an airfoil extending between a leading edge and a trailing edge, a control rod extending through the airfoil, and a mechanism driven by the control rod to change the shape of the airfoil. A vane system for a gas turbine engine is also disclosed.

A compressor or pump stage is provided. The compressor or pump stage at least comprising a central shaft (8) and one rotor (3), where the axis of rotation of the rotor (3) is the central shaft (8) and where the rotor comprises a number, n, of rows of impellers (5) arranged at an outer perimeter of the rotor with an axial distance between neighbouring rows of impellers (5), where n={2, 3, 4...}.

An adjustable guide vane for a compressor, in particular a high-pressure compressor, of a gas turbine, in particular an aircraft gas turbine is described, said vane comprising a radially outer storage section, a radially inner storage section, and a vane section, which extends in the radial direction between the outer storage section and the inner storage section, wherein the outer and the inner storage sections are designed in such a way that the adjustable guide vane can be taken up rotatably about a vane axis in the compressor, and wherein the radially inner storage section is designed like a journal (cone-shaped) and has a casing surface that revolves relative to the vane axis, said surface being of convex shape. In this way, it is provided that the radius of curvature of the convex casing surface is at least double the maximum diameter of the storage section.

An adjustable guide vane for a compressor, in particular a high-pressure compressor, of a gas turbine, in particular an aircraft gas turbine is described, said vane comprising a radially outer storage section, a radially inner storage section, and a vane section, which extends in the radial direction between the outer storage section and the inner storage section, wherein the outer and the inner storage sections are designed in such a way that the adjustable guide vane can be taken up rotatably about a vane axis in the compressor, and wherein the radially inner storage section is designed like a journal (cone-shaped) and has a casing surface that revolves relative to the vane axis, said surface being of convex shape. In this way, it is provided that the radius of curvature of the convex casing surface is at least double the maximum diameter of the storage section.

A gas turbine engine includes a fan section that includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor is a five-stage low pressure compressor. The low pressure turbine is four-stage low pressure turbine. A high spool includes a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine-stage high pressure compressor. The high pressure turbine is a two-stage high pressure turbine.

A gas turbine engine includes a fan section that includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor is a five-stage low pressure compressor. The low pressure turbine is four-stage low pressure turbine. A high spool includes a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine-stage high pressure compressor. The high pressure turbine is a two-stage high pressure turbine.

An aero-propulsion system includes a drive shaft, a low-pressure compressor, a fan shaft driving a fan, a reduction device that couples the drive shaft and the fan shaft, and an inlet channel which extends between the fan and the low-pressure compressor, the inlet having a predetermined mean radius, a ratio between a mean radius of the inlet channel and the mean radius of the low-pressure compressor, on the one hand, and the reduction ratio of the reduction mechanism, on the other hand, being less than 0.35.

The present invention relates to a rotating blade for a turbomachine, having a blade element and a shroud, wherein the shroud is profiled at a peripheral side, i.e., when observed in a tangential section, has a contact flank that is oriented axially, at least proportionally, and has a free flank that is also oriented axially, at least proportionally, and is in fact opposite the contact flank, wherein, when observed in the tangential section, an intermediate segment between the contact flank and the free flank has a first curvature in a first transition segment to the contact flank and has a second curvature in a second transition segment to the free flank, wherein the first curvature is greater than the second curvature.

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan shaft configured to drive a fan, a support configured to support at least a portion of the fan shaft, the support defining a support transverse stiffness and a support lateral stiffness, a gear system coupled to the fan shaft, and a flexible support configured to at least partially support the gear system. The flexible support defines a flexible support transverse stiffness with respect to the support transverse stiffness and a flexible support lateral stiffness with respect to the support lateral stiffness. The input defines an input transverse stiffness with respect to the support transverse stiffness and an input lateral stiffness with respect to the support lateral stiffness.

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan shaft configured to drive a fan, a support configured to support at least a portion of the fan shaft, the support defining a support transverse stiffness and a support lateral stiffness, a gear system coupled to the fan shaft, and a flexible support configured to at least partially support the gear system. The flexible support defines a flexible support transverse stiffness with respect to the support transverse stiffness and a flexible support lateral stiffness with respect to the support lateral stiffness. The input defines an input transverse stiffness with respect to the support transverse stiffness and an input lateral stiffness with respect to the support lateral stiffness.