53results about How to "Reduce vibration stress" patented technology

The present invention is a steam turbine rotor blade assembly including: an airfoil; a shroud provided at a tip of the airfoil; a blade root (dovetail) projecting toward a radially internal circumferential side of a turbine rotor and fitted to a root attachment provided on an outer circumferential portion of the turbine rotor; a platform provided between the airfoil and the blade root; a pin provided between the blade root and the root attachment; a bore formed between respective surfaces of the shrouds facing each other and included in the respective adjacent rotor blades; and a bar-like member provided in the bore.

A loudspeaker arrangement comprises a loudspeaker driver (201) mounted in an enclosure (203). The loudspeaker driver (201) comprises a speaker frame (209) which forms an outer support (221) and an inner support (213). An active radiator (205) is mounted on the inner support (213) and a passive radiator (207) is mounted between and on the inner support (213) and the outer support (221). The loudspeaker driver (201) may be mounted by the outer support (221) being fixed to the enclosure (203). The passive radiator (205) allows a low frequency extension for the loudspeaker driver (201). The integrated design of the active radiator 205 and passive radiator 207 allows the radiators (205, 207) to closely interact and behave as a single unit. The approach may reduce cost, facilitate manufacturing, improve audio quality and / or reduce vibrational stress on the enclosure.

A method is provided for determining a blade topology that reduces the effects of vibratory stress on a turbomachine blade having a plurality of discrete locations, wherein each discrete location has a thickness. The method includes the steps of creating a computational model of the blade, using the computational model to modify the thickness of at least one of the discrete locations a predetermined amount, determining a combination of a discrete location and predetermined thickness amount that reduces vibratory stress on the blade, and applying the determined combination to the computational model to create a revised blade. A turbomachine including a blade having the determined blade topology is also provided.

A compressor component having an airfoil with a profile in accordance with Table 1 is disclosed. The compressor component, such as a compressor blade, has an increased thickness over a portion of the airfoil span in order to increase stiffness. Furthermore, the airfoil has been restacked so as to induce a compressive stress in the blade root / attachment area. The increased stiffness and restacked airfoil combine to improve high-cycle fatigue capability of the compressor component.

A rotor assembly having a seal member in the root section of a rotor blade is disclosed. Various construction details are developed for blocking the flow of gases between adjacent rotor blades. In one detailed embodiment, a deformable seal member formed of a high temperature material is disposed between the root sections of adjacent rotor blades and engages the blades under operative conditions.

A stator assembly having a windage cover for structure adjacent to a cavity bounded by rotating elements in a rotary machine is disclosed. Various construction details are developed for damping vibrations in the windage cover as the windage cover bounds a cavity having swirling high velocity gases that are capable of transmitting acoustic energy and kinetic energy to adjacent structure. In one detailed embodiment, the windage cover in the uninstalled condition has diverging arms that are resiliently compressed during installation to exert a frictional force on the adjacent structure.

A vibration generator has two groups of shafts, on which at least two groups of imbalances are disposed, and which are connected with at least one drive, by means of which they are put into rotation. The shaft groups are connected with the at least one drive in such a manner that the speed of rotation of one shaft group amounts to half the speed of rotation of the other shaft group, and the ratio of the static moments of the shaft groups provided with the imbalance groups amounts to between six to one and ten to one.

A diffuser pipe for a gas turbine engine comprises a hollow pipe body including a first end, a second end fluidly connected to the first end, and at least one flattened area proximate to the second end. A ring is connected to the second end. The ring is an outlet of the diffuser pipe. At least one stiffener is disposed on the at least one flattened area. The ring and the at least one stiffener reduce vibratory stresses at the second end of the pipe body. A method of manufacturing a diffuser pipe of a gas turbine engine is also presented.

A blade for a continuous-flow machine is disclosed, especially an aircraft engine, whereby, starting from the middle section, the cross section of the blade tip is reduced with respect to the middle section, at least over a front partial section in the direction of the leading edge and over at least a rear section in the direction of the trailing edge, and a continuous-flow machine having at least one row of blades including such blades is also disclosed.

The present invention is a steam turbine rotor blade assembly including: an airfoil; a shroud provided at a tip of the airfoil; a blade root (dovetail) projecting toward a radially internal circumferential side of a turbine rotor and fitted to a root attachment provided on an outer circumferential portion of the turbine rotor; a platform provided between the airfoil and the blade root; a pin provided between the blade root and the root attachment; a bore formed between respective surfaces of the shrouds facing each other and included in the respective adjacent rotor blades; and a bar-like member provided in the bore.

A method is provided for determining a blade topology that reduces the effects of vibratory stress on a turbomachine blade having a plurality of discrete locations, wherein each discrete location has a thickness. The method includes the steps of creating a computational model of the blade, using the computational model to modify the thickness of at least one of the discrete locations a predetermined amount, determining a combination of a discrete location and predetermined thickness amount that reduces vibratory stress on the blade, and applying the determined combination to the computational model to create a revised blade. A turbomachine including a blade having the determined blade topology is also provided.

The invention belongs to the technical field of gas turbine equipment and particularly relates to a high-pressure turbine moving blade set with damping vibration attenuation blocks. According to the high-pressure turbine moving blade set with the damping vibration attenuation blocks, the damping vibration attenuation blocks are arranged in cavities between adjacent high-pressure turbine moving blade extending roots, when an engine works, the damping vibration attenuation blocks rotate along with high-pressure turbine moving blades, the centrifugal force generated by the damping vibration attenuation blocks provides normal stress for friction contact between the damping vibration attenuation blocks and the high-pressure turbine moving blades, and when the blades vibrate, the vibration quantity is consumed through contact friction between the blades and the damping vibration attenuation blocks, so that the vibration stress is reduced. The high-pressure turbine moving blade set is simple and reliable in structure, good in vibration attenuation effect, easy and convenient to regulate and control, high in environmental adaptability and free of influences of an engine temperature field and a flow field, structural parameters of the high-pressure turbine blades do not need to be greatly changed, blade damping can be effectively increased, the safety margin of working of the high-pressure turbine moving blades is increased, the service life of the turbine blades is prolonged, and the gas turbine safe operation reliability is improved.

A gas turbine engine includes a combustor having a plurality of fuel nozzles arranged circumferentially equidistant from one another about a first center. The plurality of fuel nozzles produce a plurality of pressure pulses in a fluid contained in the combustor. A turbine section is disposed downstream of the combustor and receiving the fluid from the combustor. The turbine section includes a plurality of vanes arranged circumferentially equidistant from one another about a second center. When projecting the second center onto the first center, the plurality of vanes is angularly offset in a circumferential direction relative to the plurality of fuel nozzles of a predetermined amount. The offsetting positions the vanes in flowpaths of the pressure pulses generated by the plurality of fuel nozzles.

In a silencer, holes are disposed in one row with intervals to each other along one direction, holes forming the one row are disposed in a plurality of rows in a direction perpendicular to the one direction, and positions of the holes in adjacent rows are shifted in the one direction and disposed in a zigzag pattern. A disposition interval of the holes of an outer edge hole row forming one row in at least two sides opposite to each other in the silencing porous plate is set to be larger than a disposition interval of the holes of a row inside the outer edge hole row.

In a silencer, holes are disposed in one row with intervals to each other along one direction, holes forming the one row are disposed in a plurality of rows in a direction perpendicular to the one direction, and positions of the holes in adjacent rows are shifted in the one direction and disposed in a zigzag pattern. A disposition interval of the holes of an outer edge hole row forming one row in at least two sides opposite to each other in the silencing porous plate is set to be larger than a disposition interval of the holes of a row inside the outer edge hole row.

The invention relates to the field of industrial steam turbine blades, and particularly discloses a 645 mm last stage dynamic blade for an industrial steam turbine. The 645 mm last stage dynamic blade for the industrial steam turbine comprises a blade body, a blade root and a boss lashing wire for making the blade form whole circle self locking, wherein the blade body and the blade root are integrally formed; a molded line of the blade body is a variable section twisted blade; the height H of the blade body is equal to 645 mm, the bottom radius Dr is equal to 560 mm, and the exhausting area of the blade body is 3.6 m<2>; and the center of the self-locking boss lashing wire is arranged at the position, with the relative blade height being 0.65 from bottom to top, of the blade body. The large load last stage dynamic blade, with the exhausting area being 3.6 m<2>, for the industrial steam turbine has the beneficial effects that the pneumatic efficiency is high, the strength performance is good, the exhausting area is large, and manufacturing and assembling are convenient, and can widely used for waste heat utilizing and industrial steam turbines for the PTA manufacturing process.