A stator vane assembly for a compressor of an aircraft turbine engine includes an inner shroud, an outer shroud, and stator vanes. The stator vanes are attached only to the inner shroud and are in non-immobilizing mechanical contact with the outer shroud.

Structures for reducing forward loads in compressors are described. A compressor includes inner and outer circumferential support structures. The inner circumferential support structure includes an aft-most and forward spacer arms. The compressor also includes two stages, each including a vane having a root positioned at the outer circumferential support structure and a tip positioned radially inward from the root, and a rotor extending radially from the spacer arm adjacent to the vane. An intersection of the rotors and spacer arms defines centrally located points. An arrangement of a first line extending through the points forms an angle with a second line parallel to a longitudinal centerline and extending through the tips of the vanes. The angle is greater than 0 and less than or equal to 45.

An apparatus comprises a variable pitch blade configured to connect with an endwall of an open rotor engine. The variable pitch blade defines a chamber therein on a bottom edge thereof. A retractable edge member pivotally connects within the chamber to move between a first position wherein at least a portion of the retractable edge member is located within the chamber and a second position wherein a portion of the retractable edge member extends downward from the bottom edge of the variable pitch blade to block a gap between the bottom edge of the variable pitch blade and the endwall.

A stator vane for a gas turbine stator vane stage is provided that includes an airfoil having leading and trailing edges, a vane tip, suction and pressure side surfaces, and at least one aero passage. The leading and trailing edges are chordwise spaced apart. The vane tip is spanwise spaced apart from a radial base end. The suction side surface extends chordwise between the leading and trailing edges, and extends spanwise between the radial base end and the vane tip. The pressure side surface extends chordwise between the leading and trailing edges, and extends spanwise between the radial base end and the vane tip. The at least one aero passage extends through the airfoil between the suction and pressure side surfaces, and is disposed proximate and spanwise separated from the vane tip. The stator vane is configured to be cantilevered with the vane tip being unsupported.

A gas turbine engine includes a ceramic matrix composite (CMC) vane arc segments that are arranged in a circumferential row. Each of the CMC vane arc segments includes an airfoil section that defines first and second side walls, leading and trailing ends, and first and second radial ends. At the first radial end, the airfoil section has a single-sided platform that and the second side wall has a bearing surface. The single-sided platform of each of the CMC vane arc segments in the circumferential row is situated to bear against the bearing surface of the next of the CMC vane arc segments in the circumferential row.

The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.