Abstract: A turbofan aircraft engine having a primary duct (C), including a combustion chamber (BK), a first turbine (HT) disposed downstream of the combustion chamber, a compressor (HC) disposed upstream of the combustion chamber and coupled (W1) to the first turbine, and a second turbine (L) disposed downstream of the first turbine and coupled (W2) via a speed reduction mechanism (G) to a fan (F) for feeding a secondary duct (B) of the turbofan aircraft engine. A square of a ratio of a maximum blade diameter (DF) of the fan to a maximum blade diameter (DL) of the second turbine is at least 3.5, in particular at least 4.

Abstract: The invention relates to a turbofan aircraft engine that comprises a primary duct including a combustion chamber; a first turbine disposed downstream of the combustion chamber; a compressor disposed upstream of the combustion chamber and coupled to the first turbine; and a second turbine disposed downstream of the first turbine and coupled to a fan for feeding a secondary duct of the turbofan aircraft engine. The bypass ratio of the inlet area of the secondary duct to the inlet area of the primary duct is at least 7 and the second turbine comprises at least two stages. The mean outer radius of the last stage of the second turbine divided by the length of the second turbine is at least 1.4.

Abstract: The invention relates to a turbofan aircraft engine that comprises a primary duct including a combustion chamber; a first turbine disposed downstream of the combustion chamber; a compressor disposed upstream of the combustion chamber and coupled to the first turbine; and a second turbine disposed downstream of the first turbine and coupled to a fan for feeding a secondary duct of the turbofan aircraft engine. The bypass ratio of the inlet area of the secondary duct to the inlet area of the primary duct is at least 7 and the second turbine comprises at least two stages. For the first stage the mean radius r of a stator vane expressed in inch divided by the number of stator vanes is at least 0.18.

Abstract: A turbofan aircraft engine has at least one stage pressure ratio is at least 1.5, and a quotient of the total blade count divided by 110 is less than a difference ([(p1/p2)?1]) of the total pressure ratio minus one, and the total pressure ratio is greater than 4.5, and the turbine has at least two and no more than five turbine stages; and/or a product (An2) of an exit area (AL) of the second turbine and a square of a rotational speed of the second turbine at the design point is at least 4.5·1010 [in2·rpm2], and a blade tip velocity (uTIP) of at least one turbine stage of the second turbine at the design point is at least 400 meters per second. A jet and method are also provided.

Abstract: A turbofan aircraft engine having a primary duct (C), including a combustion chamber (BK), a first turbine (HT) disposed downstream of the combustion chamber, a compressor (HC) disposed upstream of the combustion chamber and coupled (W1) to the first turbine, and a second turbine (L) disposed downstream of the first turbine and coupled (W2) via a speed reduction mechanism (G) to a fan (F) for feeding a secondary duct (B) of the turbofan aircraft engine. A square of a ratio of a maximum blade diameter (DF) of the fan to a maximum blade diameter (DL) of the second turbine is at least 3.5, in particular at least 4.

Abstract: A rotor of a turbine-type machine, in particular a gas turbine rotor, is disclosed. The rotor has a rotor base body, where the rotor base body has a groove extending in the circumferential direction of the rotor base body, and has multiple rotor blades, or rotor blade segments, or an integrally bladed, circumferentially closed rotor blade ring. The rotor blades, or rotor blade segments, or rotor blade ring are/is anchored in the groove extending in the circumferential direction of the rotor base body by a blade base or ring base. The groove has a profiled groove wall leg on only one side, with the blade base of the rotor blades or the rotor blade segments, or ring base of the rotor blade ring, being in contact with the profiled groove wall leg with a correspondingly profiled supporting flank.

Abstract: A rotor of a turbine-type machine, in particular a gas turbine rotor, is disclosed. The rotor has a rotor base body, where the rotor base body has a groove extending in the circumferential direction of the rotor base body, and has multiple rotor blades or rotor blade segments. The rotor blades or rotor blade segments are each anchored in the groove extending in the circumferential direction of the rotor base body by a blade base. The groove has a profiled groove wall leg on only one side, with the blade base of the rotor blades or the rotor blade segments being in contact with the profiled groove wall leg with a correspondingly profiled supporting flank.

Abstract: A rotor of a turbine-type machine, in particular a gas turbine rotor, is disclosed. The rotor has a rotor base body, where the rotor base body has a groove extending in the circumferential direction of the rotor base body, and has multiple rotor blades, or rotor blade segments, or an integrally bladed, circumferentially closed rotor blade ring. The rotor blades, or rotor blade segments, or rotor blade ring are/is anchored in the groove extending in the circumferential direction of the rotor base body by a blade base or ring base. The groove has a profiled groove wall leg on only one side, with the blade base of the rotor blades or the rotor blade segments, or ring base of the rotor blade ring, being in contact with the profiled groove wall leg with a correspondingly profiled supporting flank.

Abstract: A rotor of a turbine-type machine, in particular a gas turbine rotor, is disclosed. The rotor has a rotor base body, where the rotor base body has a groove extending in the circumferential direction of the rotor base body, and has multiple rotor blades or rotor blade segments. The rotor blades or rotor blade segments are each anchored in the groove extending in the circumferential direction of the rotor base body by a blade base. The groove has a profiled groove wall leg on only one side, with the blade base of the rotor blades or the rotor blade segments being in contact with the profiled groove wall leg with a correspondingly profiled supporting flank.