Abstract: Gas turbine unit having an axis parallel to the main gas flow, the gas turbine unit comprising: a blade having a tip; a stator heat shield having an inner surface facing the blade tip; wherein the inner surface of the stator heat shield and the blade tip define a variable clearance depending on the applied thermal condition; wherein the blade tip is configured to have a cylindrical shape along the axial direction in a hot running condition starting from a conical shape along the axial direction in a cold starting condition.

Abstract: The application describes a method of operating a gas turbine during a cool-down phase. The gas turbine provides a compressor, a combustor downstream of the compressor, and a turbine downstream of the combustor, with the turbine providing a turbine vane carrier. The method includes feeding a flow of cooling air from the compressor to the turbine vane carrier, measuring a temperature of the flow of cooling air and measuring a temperature of the turbine vane carrier. In the method, the flow of cooling air is fed at a first flow rate when the temperature of the turbine vane carrier is lower than the temperature of the cooling air, and the flow of cooling air is fed at a second flow rate when the temperature of the turbine vane carrier is higher than the temperature of the cooling air, wherein the first flow rate is higher than the second flow rate.

Abstract: The invention pertains to a seal system for a passage between a turbine stator and a turbine rotor, including: a first arm extending radially outwards from the turbine rotor and toward the first seal arranged on the stator, and terminating short of the first seal thereby creating a first gap between the first seal and the first arm. The seal system further includes a second seal arranged on the turbine stator, and a second arm extending axially from the turbine rotor towards the second seal base, and terminating short of the second seal thereby creating a second gap between the second seal and the second arm. The invention further refers to a gas turbine including such a seal system.

Abstract: Gas turbine unit having an axis parallel to the main gas flow, the gas turbine unit comprising: a blade having a tip; a stator heat shield having an inner surface facing the blade tip; wherein the inner surface of the stator heat shield and the blade tip define a variable clearance depending on the applied thermal condition; wherein the blade tip is configured to have a cylindrical shape along the axial direction in a hot running condition starting from a conical shape along the axial direction in a cold starting condition.

Abstract: A gas turbine includes a compressor, an annular combustion chamber, and a turbine, a combustion chamber shell of the combustion chamber adjoining the turbine inlet in a transition region in order to introduce the hot gases generated in the combustion chamber into the downstream turbine such that a thermal expansion-induced relative movement between the combustion chamber and the turbine inlet is possible. Combustion chamber shell support elements distributed on the periphery come into contact with a conical contour on the shaft cover due to the thermal expansion that occurs during operation and are supported on said contour. An improvement with respect to loading and service life is achieved in that the conical contour and the machine axis form an angle that allows the combustion chamber shell support elements to slide onto the conical contour.

Abstract: The application describes a method of operating a gas turbine during a cool-down phase. The gas turbine provides a compressor, a combustor downstream of the compressor, and a turbine downstream of the combustor, with the turbine providing a turbine vane carrier. The method includes feeding a flow of cooling air from the compressor to the turbine vane carrier, measuring a temperature of the flow of cooling air and measuring a temperature of the turbine vane carrier. In the method, the flow of cooling air is fed at a first flow rate when the temperature of the turbine vane carrier is lower than the temperature of the cooling air, and the flow of cooling air is fed at a second flow rate when the temperature of the turbine vane carrier is higher than the temperature of the cooling air, wherein the first flow rate is higher than the second flow rate.

Abstract: The invention pertains to a seal system for a passage between a turbine stator and a turbine rotor, including: a first arm extending radially outwards from the turbine rotor and toward the first seal arranged on the stator, and terminating short of the first seal thereby creating a first gap between the first seal and the first arm. The seal system further includes a second seal arranged on the turbine stator, and a second arm extending axially from the turbine rotor towards the second seal base, and terminating short of the second seal thereby creating a second gap between the second seal and the second arm. The invention further refers to a gas turbine including such a seal system.

Abstract: A gas turbine includes a compressor, an annular combustion chamber, and a turbine, a combustion chamber shell of the combustion chamber adjoining the turbine inlet in a transition region in order to introduce the hot gases generated in the combustion chamber into the downstream turbine such that a thermal expansion-induced relative movement between the combustion chamber and the turbine inlet is possible. Combustion chamber shell support elements distributed on the periphery come into contact with a conical contour on the shaft cover due to the thermal expansion that occurs during operation and are supported on said contour. An improvement with respect to loading and service life is achieved in that the conical contour and the machine axis form an angle that allows the combustion chamber shell support elements to slide onto the conical contour.