Patent application number | Description | Published |
20090053068 | STRUCTURES FOR DAMPING OF TURBINE COMPONENTS - Disclosed is a coating for gas turbine components including at least one material having vibration-damping properties. Further disclosed is an airfoil of a gas turbine having damped vibrational characteristics including an airfoil substrate and a coating applied to the airfoil substrate including at least one material having vibration-damping properties. A method of damping vibration of a gas turbine component includes applying a coating including at least one material having damping properties to the turbine component. | 02-26-2009 |
20100028133 | TURBOMACHINE COMPONENT DAMPING STRUCTURE AND METHOD OF DAMPING VIBRATION OF A TURBOMACHINE COMPONENT - A turbomachine component includes a main body having a surface, and a damping structure mounted to the surface of the main body. The damping structure is formed from a material having a temperature dependent damping characteristic. | 02-04-2010 |
20100104446 | FABRICATED HYBRID TURBINE BLADE - A blade for a turbine includes: a support structure of high strength material as a central portion of the blade, the support structure including a root and a body extending to a tip, the body providing a leading edge, a trailing edge and mounting section; and a honeycomb skin attached to the mounting section for providing a lightweight airfoil portion of the blade. A method and a turbine are included. | 04-29-2010 |
20100183446 | TURBINE BLADE OR VANE WITH IMPROVED COOLING - Disclosed is a turbine blade or vane including a blade or vane body including a leading edge and a trailing edge, a plurality of cooling openings disposed along the trailing edge, a first width of the trailing edge, the first width being disposed across the cooling openings, and a second width of the trailing edge the second width being disposed between the cooling openings, wherein the second width is smaller than the first width. | 07-22-2010 |
20110243711 | INTERIOR COOLING CHANNELS - Cooling channels through the interior of a machine component that include: a first set of cooling channels, the first set of cooling channels including a plurality of parallel channels that reside in a first plane; a second set of cooling channels, the second set of cooling channels including a plurality of parallel channels that reside in a second plane. Along a longitudinal axis, the cooling channels of the first and second set of cooling channels may include an alternating diverging-converging configuration, the alternating diverging-converging configuration creating a series of broader chamber sections connected by a series of narrower throat sections. The first set of cooling channels and the second set of cooling channels may be configured such that, when viewed from the side, a crisscrossing pattern with a plurality of intersections is formed. The first plane resides in spaced relation to the second plane, with the first plane being offset from the second plane such that a plurality of the chamber sections of the first set of cooling channels connect to a plurality of the chamber sections of the second set of cooling channels. | 10-06-2011 |
20120308399 | TURBINE NOZZLE SLASHFACE COOLING HOLES - A turbine vane or blade segment includes at least one airfoil extending radially outwardly from a radially inner band. The radially inner band is formed with substantially axially-extending side edges, and the at least one airfoil is formed with a leading edge, a trailing edge, a pressure side, a suction side, and an internal cooling plenum. The pressure side faces one of the substantially axially-extending side edges and the leading edge is located proximate to that one edge. A plurality of cooling passages are formed in the radially inner band in fluid communication with the internal plenum and exiting the inner band via a plurality of exit holes in the one of the substantially axially-extending side edges. The plurality of exit holes are confined to a region along the one of the substantially axially-extending side edges where static pressure P | 12-06-2012 |
20140030073 | CLOSED LOOP COOLING SYSTEM FOR A GAS TURBINE - A system for removing heat from a gas turbine generally includes a plurality of stationary nozzles arranged in an annular array within the gas turbine. Each of the plurality of stationary nozzles may include a radially outer platform and a radially extending cooling passage. The radially extending cooling passage may have an inlet that extends generally axially and circumferentially across a portion of the radially outer platform. A closed loop cooling coil may extend continuously circumferentially around the radially outer platform of two or more of the plurality of stationary nozzles. The closed loop cooling coil may be disposed circumferentially across and outside of the inlet of the radially extending cooling passage of each of the two or more stationary nozzles, and a cooling medium may flow through the cooling coil and out of the gas turbine so as to remove heat from the gas turbine. | 01-30-2014 |
20140286762 | TURBINE AIRFOIL ASSEMBLY - A turbine airfoil assembly has an airfoil with an inner wall, an outer wall, a leading edge and a trailing edge. The airfoil has one or more chambers extending in a substantially chordwise direction of the airfoil. An insert has a plurality of impingement holes, and the insert is configured to be inserted within one of the chambers. The insert is configured to cool the airfoil via the plurality of impingement holes. A chambering element is attached only to the insert, the chambering element is configured to provide an increased cooling gas pressure inside a boundary area defined by the chambering element relative to an area outside the boundary area. A gap exists between the inner wall of the airfoil and the chambering element, and the gap allows cooling gas to exit the boundary area and enter the area outside the boundary area. | 09-25-2014 |