Abstract: Described is a blade for a high-speed turbine stage of an aircraft gas turbine, in particular of an aircraft engine, the blade including a radially inner blade root, and an airfoil extending radially outwardly from the blade root. It is provided that the blade be shroudless and that the airfoil have a radially outer end portion that is positionable opposite a rub surface when the blade is in an installed state, and that the airfoil have a radially inner chord length that is at least 1.1 times, preferably at least 1.2 times, in particular at least 1.3 times a radially outer chord length, the inner chord length being measured at the airfoil directly above the blade root, and the outer chord length being measured at the airfoil in the region of or below the end portion.

Abstract: The present invention relates to a rotor blade (20) for arrangement in a gas duct (2) of a turbomachine (1), having a rotor blade airfoil (23), which, viewed in a tangential section, has a blade airfoil profile (24) with a leading edge radius RVK and a rotor blade airfoil thickness d, wherein the blade airfoil profile (24) is thickened, at least in sections, specifically the blade airfoil thickness d is specified, in relation to the front edge radius RVK, such that (2d/Rvk2)?d?5.5.

Abstract: The present invention relates to a rotor blade arrangement for a turbomachine, with a rotor blade which has a sealing tip radially on the outside, and with a seal arrangement, wherein the seal arrangement forms a radially inwardly open cavity, in which the sealing tip is arranged, to which end the seal arrangement has a first sealing element, namely a first seal carrier with a first run-in coating, and a second sealing element, wherein the first run-in coating delimits the cavity radially on the outside, and the second sealing element delimits the cavity in an axial direction, and wherein the first and the second sealing element are assembled.

Abstract: A rotor blade (20) for placement in a gas channel (3) of a turbomachine (1), including a rotor blade airfoil (23) which, in relation to a flow in the gas channel (3), includes a front edge (23a) and a rear edge (23b) downstream therefrom, as well as a suction side (41) and a pressure side (42). The rotor blade airfoil (23) is provided with an inclination toward the suction side (41) over at least one section (45.1) of its radial rotor blade airfoil height (45). The inclination is set in such a way that during operation a centrifugal force bending moment (46), which effectuates the centrifugal force on the rotor blade airfoil (23) due to the inclination, is greater than a gas force bending moment (47) that acts on the rotor blade airfoil (23) due to the circulation around the rotor blade airfoil (23) in the gas channel (3).

Abstract: Rotor blade (20) to be arranged in a gas conduit (3) of a turbomachine (1), having a rotor blade airfoil (23), which radially inwardly has a chord length Si, radially outwardly has a chord length Sa, and in a radial position rx inbetween has a chord length Sx, the chord length S in the radial position rx being at least equal to the chord length Si radially inwardly (Si<Sx), and the chord length Sa radially outwardly corresponding at most 0.9 times the chord length Sx in the radial position rx inbetween (Sa<0.9 Sx).

Abstract: Described is a blade for a high-speed turbine stage of an aircraft gas turbine, in particular of an aircraft engine, the blade including a radially inner blade root, a radially outer shroud, and an airfoil extending between the blade root and the shroud. It is provided that the outer shroud have only a single sealing element, which projects radially from the shroud, in particular only a single sealing fin.

Abstract: An axially divided inner ring is provided for fastening to, in particular adjustable, guide vanes (50) of a turbomachine, in particular a compressor or turbine stage of a gas turbine, which includes a first partial ring (10) and a second partial ring (20), which is supported in the axial direction directly or indirectly on two axially facing support surfaces (11, 12) of the first partial ring (10) and is fixed on the first partial ring (10) in the radial direction with the aid of multiple alignment pins (40) distributed in the circumferential direction.

Abstract: Described is a blade for a high-speed turbine stage of an aircraft gas turbine, in particular of an aircraft engine, the blade including a radially inner blade root, and an airfoil extending radially outwardly from the blade root. It is provided that the blade be shroudless and that the airfoil have a radially outer end portion that is positionable opposite a rub surface when the blade is in an installed state, and that the airfoil have a radially inner chord length that is at least 1.1 times, preferably at least 1.2 times, in particular at least 1.3 times a radially outer chord length, the inner chord length being measured at the airfoil directly above the blade root, and the outer chord length being measured at the airfoil in the region of or below the end portion.

Abstract: Described is a blade for a high-speed turbine stage of an aircraft gas turbine, in particular of an aircraft engine, the blade including a radially inner blade root, a radially outer shroud, and an airfoil extending between the blade root and the shroud. It is provided that the outer shroud have only a single sealing element, which projects radially from the shroud, in particular only a single sealing fin.

Abstract: An axially divided inner ring is provided for fastening to, in particular adjustable, guide vanes (50) of a turbomachine, in particular a compressor or turbine stage of a gas turbine, which includes a first partial ring (10) and a second partial ring (20), which is supported in the axial direction directly or indirectly on two axially facing support surfaces (11, 12) of the first partial ring (10) and is fixed on the first partial ring (10) in the radial direction with the aid of multiple alignment pins (40) distributed in the circumferential direction.

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 housing section of a turbine engine compressor stage or a turbine engine turbine stage that, in particular, has a closed and annular-shaped, radially outer casing. The radially outer casing has radially inwardly extending webs that are angled at a slant relative to the radius.

Abstract: A damping device for being situated between a housing wall of a housing of a thermal gas turbine and a casing ring is provided. The casing ring includes an area radially internal with regard to a rotation axis of a rotor of the thermal gas turbine and facing rotating moving blades of the gas turbine. The damping device includes at least sectionally a porous damping structure. A method for manufacturing this type of damping device as well as to a thermal gas turbine, in particular an aircraft engine, in which this type of damping device is situated in a housing of the gas turbine between a housing wall and a casing ring are also provided.

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 damping device for being situated between a housing wall of a housing of a thermal gas turbine and a casing ring is provided. The casing ring includes an area radially internal with regard to a rotation axis of a rotor of the thermal gas turbine and facing rotating moving blades of the gas turbine. The damping device includes at least sectionally a porous damping structure. A method for manufacturing this type of damping device as well as to a thermal gas turbine, in particular an aircraft engine, in which this type of damping device is situated in a housing of the gas turbine between a housing wall and a casing ring are also provided.

Abstract: A housing section of a turbine engine compressor stage or a turbine engine turbine stage that, in particular, has a closed and annular-shaped, radially outer casing. The radially outer casing has radially inwardly extending webs that are angled at a slant relative to the radius.

Abstract: Rotor blades (1) are held by means of toothed roots (3) in correspondingly profiled axial grooves (4) of a rotor wheel rim (5). Between each root end and the base of an axial groove (4), there is formed a radial gap (S) accommodating a rivet (6). Each rivet (6) is centrally guided on two inserts (9, 9') resting against the base of the axial groove (4) in the radial gap (S) and having bent or angled end parts (10, 10') resting against the front and rear faces of the wheel rim (5). Seating wedges (11, 11') surround the upset head (8) and the set head (7) of the rivet (6) and are axially and radially clamped to wedge-shaped complementary surfaces (G, G') of the root end and of the respective insert (9, 9') by riveting. As a result, an improved radial and axial clamping of the rotor blades is achieved.

Abstract: A device for axially securing a moving blade and for eliminating rotor unbalances is provided. The moving blade has a blade base. A wheel disk has circumferentially distributed axial grooves. The blade base is anchored in one of the axial grooves wherein the axial groove encloses an indentation formed between a bottom of the axial groove and an end of the blade base. A filler piece operatively arranged on the bottom of the axial groove, and a securing plate having two ends, are insertable into the indentation. The securing plate is arranged to bridge a gap formed between the end of the blade base and the filler piece along an axial length of the indentation with the two ends projecting beyond the indentation. The securing plate is bent at each of its two ends in a mutually opposite manner with respect to end faces of the filler piece, wheel disk and blade base. One rotor unbalance of the blade is compensated by at least one filler piece.