Abstract: The invention relates to a blade component, a compressor or turbine stage of a gas turbine, in particular a gas turbine engine characterized in that the blade component includes at least two structural elements which can be connected together by means of a connection method, in particular sintering, and that the at least one connection face of the at least two structural elements lies on a face, wherein in particular the normal vector has, for at least a part of the face, a component perpendicular to the radial orientation of the blade component. The invention also concerns a method for manufacturing a blade component, and a gas turbine with a blade component.

Abstract: A turboshaft gas turbine engine comprises, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a free power turbine. The expansion ratios of the turbines are dependent upon an outlet temperature of the combustor, and an expansion relationship between the turbines is defined as the ratio of expansion ratios of the turbines over a running range of non-dimensional power outputs of the gas turbine engine. A second derivative of the expansion relationship is from 0.6 to 1.0.

Abstract: A turboshaft gas turbine engine comprises, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a free power turbine. The gas-generator turbine is a single stage turbine having a specific work of from 200 to 350 kilojoules per kilogram, said specific work being the power output of the turbine per unit mass flow therethrough.

Abstract: A turboshaft gas turbine engine comprises, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a free power turbine. The expansion ratios of the turbines are dependent upon an outlet temperature of the combustor, and an expansion relationship between the turbines is defined as the ratio of expansion ratios of the turbines over a running range of non-dimensional power outputs of the gas turbine engine. A second derivative of the expansion relationship is from 0.6 to 1.0.

Abstract: A control method/system for a hybrid-electric aircraft propulsion system includes a generator, a propulsor, a controller and an electric storage unit. The generator including a gas turbine with blades separated from a casing by a clearance and driving an electric generator. An electric motor drives the propulsor. The controller controls the turbine and supply of electric power between the motor, the storage unit and the generator in response to a thrust demand and cooperates with a clearance controller: receives a command for a change in thrust demand; determines an operational profile that minimizes a function comprising a measure of fuel supplied to the turbine, transferring electric power from/to the storage unit, a difference between measures of current and demanded thrust and clearance over a time period; and operates the motor, turbine and storage unit according to the operational profile over the time period.

Abstract: The present invention relates to a cooled turbine blade for a gas-turbine engine having at least one cooling duct (14) extending radially, relative to a rotary axis of the gas-turbine engine, inside the airfoil and air-supply ducts (12) issuing into said cooling duct, characterized in that the cooling duct (14) extends into the blade root (6) in order to generate close to the wall a cooling airflow moved at high circumferential velocity and radially in helical form and that in the area of the blade root (6) at least one nozzle-shaped air-supply duct (12) issues into the cooling duct (14) tangentially or with a tangential velocity component.

Abstract: In impingement air cooling of gas turbine components, cooling air velocity packs of a certain amplitude and a given frequency are applied to impingement air openings, with intervallic annular swirl structures being formed which penetrate a cross-flow and hit a component to be cooled with high intensity, thus providing for efficient cooling. In order to obtain annular swirl structures with optimum cooling effect, the Strouhal number, which is determined by a ratio of amplitude, frequency of the velocity packs and size of impingement air cooling openings, ranges between 0.2 and 2.0, and preferably between 0.8 and 1.2.

Abstract: The present invention relates to a cooled turbine blade for a gas-turbine engine having at least one cooling duct (14) extending radially, relative to a rotary axis of the gas-turbine engine, inside the airfoil and air-supply ducts (12) issuing into said cooling duct, characterized in that the cooling duct (14) extends into the blade root (6) in order to generate close to the wall a cooling airflow moved at high circumferential velocity and radially in helical form and that in the area of the blade root (6) at least one nozzle-shaped air-supply duct (12) issues into the cooling duct (14) tangentially or with a tangential velocity component.

Abstract: Passive shroud cooling of rotor blades (5) is accomplished by a multitude of cooling-air flows introduced into the hot-gas flow upstream of the rotor blades on the outer circumference, these cooling-air flows, depending on the aerodynamic and geometrical conditions, being orientated at a defined radial angle and a circumferentially related tangential angle such that the cooling air primarily and essentially hits the thermally highly loaded bottom surface (11) of the outer shrouds (6) allover. In the apparatus for the performance of the method, cooling air passages (9) tangentially and radially orientated to the hot-gas flow are provided in an area of a casing section (3) which protrudes beyond the stator blades (1) in the downstream direction, this casing section (3) interrupting the transport of cooling air in the axial direction.

Abstract: In impingement air cooling of gas turbine components, cooling air velocity packs of a certain amplitude and a given frequency are applied to impingement air openings, with intervallic annular swirl structures being formed which penetrate a cross-flow and hit a component to be cooled with high intensity, thus providing for efficient cooling. In order to obtain annular swirl structures with optimum cooling effect, the Strouhal number, which is determined by a ratio of amplitude, frequency of the velocity packs and size of impingement air cooling openings, ranges between 0.2 and 2.0, and preferably between 0.8 and 1.2.

Abstract: Passive shroud cooling of rotor blades (5) is accomplished by a multitude of cooling-air flows introduced into the hot-gas flow upstream of the rotor blades on the outer circumference, these cooling-air flows, depending on the aerodynamic and geometrical conditions, being orientated at a defined radial angle and a circumferentially related tangential angle such that the cooling air primarily and essentially hits the thermally highly loaded bottom surface (11) of the outer shrouds (6) allover. In the apparatus for the performance of the method, cooling air passages (9) tangentially and radially orientated to the hot-gas flow are provided in an area of a casing section (3) which protrudes beyond the stator blades (1) in the downstream direction, this casing section (3) interrupting the transport of cooling air in the axial direction.