Abstract: A gas turbine combustor, in particular for an aircraft engine, has at least one chamber, through which combustion gas flows in use, and which is defined by a lateral wall with channeling for feeding cooling air into the chamber and cooling the lateral wall; part of the channeling is defined by at least one double wall for guiding a relative stream of cooling air into the chamber in a tangential direction with respect to the lateral wall.

Abstract: A hinge device for a rotary member of an aircraft engine having a conduit housing the rotary member and for conducting a stream of gas, and a cavity outside the conduit and for receiving a mass of air having, in use, a higher pressure than the conduit; the device having a seat formed in a platform interposed between the cavity and the conduit, and a pin integral with the rotary member and engaging the seat to rotate about an axis; the pin is cooled by a stream of air, which flows inside a passage having an inlet which comes out inside the cavity, an outlet which comes out inside the conduit, and at least one number of channels formed outside the pin and spaced angularly apart; and part of the passage is formed on a collar portion fitted to the pin to calibrate the flow of air inside the passage.

Abstract: A method of producing and assembling a cooling device inside an axial-flow gas turbine blade; the airfoil profile of the blade has an inner surface defining a chamber, and is connected to a supporting structure by two opposite end pins having respective openings for the passage of cooling air and which come out inside the chamber; the method provides for forming an insert having a number of holes and defined by a first and at least a second body separate from each other, and each of a size approximating but no larger than that of at least one of the openings; the bodies are inserted successively through the openings in the pins, and are positioned inside the chamber to direct a relative stream of air through each hole on to the inner surface of the airfoil profile.

Abstract: A blade for a stator of a variable-geometry turbine, particularly for aircraft engines; the blade has an airfoil profile hinged to a structure of the stator to rotate about an axis, and having a pressure front wall and a suction rear wall; the blade also has two end walls, which are located at opposite ends of the airfoil profile, cooperate in sliding manner with the structure of the stator, and are cooled, in use, by air flowing through a number of holes; the outlets of the holes are located close to the outer edges between the end walls and the front wall to generate a tangential stream of cooling air by virtue of the pressure difference acting on the front and rear walls.

Abstract: A double-wall blade for a turbine, particularly for aeronautical applications, has a streamlined lateral wall, which extends along an axis, surrounds the axis itself, and is defined by an inner wall and an outer wall facing and integral with each other; the blade is provided with inner channeling, which has a cavity into which, in use, cooling air is admitted, and a number of ducts, which are provided tangentially between the inner wall and outer wall, are separated from one another by baffles which are transverse to the axis and have respective intakes which are separate from one another and communicate with the cavity in order, in use, each to have a corresponding flow of cooling air pass through them.

Abstract: A gas turbine combustor, in particular for an aircraft engine, has at least one chamber, through which combustion gas flows in use, and which is defined by a lateral wall with channeling for feeding cooling air into the chamber and cooling the lateral wall; part of the channeling is defined by at least one double wall for guiding a relative stream of cooling air into the chamber in a tangential direction with respect to the lateral wall.