Abstract: A pulverized solid fuel nozzle assembly 34 includes a fuel feed pipe 38 and a nozzle tip 36 pivotally secured relative to the fuel feed pipe 38. The fuel feed pipe 38 includes a generally cylindrical shell 99 having a round outlet end 102 and a bulbous protrusion 106 disposed around a perimeter of the round outlet end 102. The nozzle tip 36 includes an inner shell 40 having a round inlet end 108 arranged in concentric relationship with the round outlet end 102 of the generally cylindrical shell 99. The round inlet end 108 is disposed around the bulbous protrusion 106 for forming a seal between the inner shell 40 and the fuel feed pipe 38. The nozzle tip 36 also includes an outer shell 39 arranged in coaxial relationship with the inner shell 40, and an annular air channel 42 disposed between the inner and outer shells 40, 39. The nozzle tip 36 is pivotable about at least one axis 52, 250 for directing a stream of pulverized solid fuel from the inner shell 40.

Abstract: The blades of a turbomachine comprise airfoils, which are bent such that the lean angle (?), defined as the angle which the stacking line of the airfoil includes with the radial direction, and measured in the direction of rotation (?), is variable along the width of the flow channel and decreases from the hub towards the housing.

Abstract: The invention relates to a gas-cooled electrical device (13), especially a gas-cooled generator, which comprises at least a rotor (1), a stator (2) and also a stator casing (3) which encloses this stator, and an end-winding space (5). In this case, passages for a cooling gas are provided both in the rotor (1) and in the stator (2), wherein the end-winding space (5), especially connectors (7) and/or output conductors (7) which are arranged therein, are impinged by at least some of the cooling gas flow (8) which issues from the rotor (1). In order to be able to achieve an especially concentrated and controlled cooling, at least the connectors (7) and the output conductors (7) are arranged in the working area of at least one guiding device (14) which guides the cooling gas flow (8).

Abstract: A device for improving the protection of a component against surface discharges includes a non-woven substrate made of fibers and configured to be electrically insulating and a coating containing silicon carbide (SiC) disposed on the substrate.

Abstract: A guide vane for a gas turbine includes a vane body having a leading edge and a trailing edge and a shroud extending at least between the leading edge and the trailing edge. The shroud has a first side wall and a second side wall extending essentially radially and in a longitudinal direction of the gas turbine. The first side wall has a groove disposed in a region of the trailing edge extending in a longitudinal direction of the shroud and is configured to receive a sealing plate. The first side wall also has a clearance extending from the groove in a region of the trailing edge, wherein a depth of the clearance in a circumferential direction of the gas turbine is equal to a depth of the groove, and wherein a width of the clearance in the longitudinal direction of the shroud is between one to three times the depth of the clearance.

Abstract: An atomizer system 160 for use in a spray dry absorber 130, the atomizer system 160 includes a support member 161, an atomizing disc 162 rotatably coupled to the support member 161, a first feed line 163 for delivery of a reagent slurry 163a to a delivery area 165 and a second feed line 164 for delivery of a dilution liquid 164a to the delivery area 165. The delivery area 165 is positioned within the atomizer system 160 to reduce scale buildup generated by interaction of the reagent slurry 163a and the dilution liquid 164a.

Abstract: A gas turbine includes a turbine housing; a rotatable turbine rotor with turbine blades; and at least one heat shield segment for a stator. The heat shield segment is disposed radially between the rotor and the housing and attached to the housing and has a profile with a curved section in at least one region in an axial direction of the gas turbine and a radial outer surface. The heat shield segment includes a rib, a first boss, and a second boss disposed on the radial outer surface.

Abstract: A turbomachine includes a turbine stage having a plurality of blades including a first and a second blade disposed adjacent to each other, each blade having a radial longitudinal axis and a blade tip and configured to twist around the radial longitudinal axis in an operating state in a twist direction, wherein the first and the second blades are pretwisted in a direction of twist in a non-operating state, and wherein a connecting element interconnects the first and the second blades to each other in a circumferential direction in a region of the respective blade tip so as to maintain the pretwisting.

Abstract: A method of dismantling a turbomachine includes opening a turbomachine casing in the region of a combustion chamber, releasing a top half shell from a bottom half shell, axially displacing the top half shell and thereafter removing the top half shell in essentially the radial direction, rotationally fixedly connecting an auxiliary tool to the rotor of the turbomachine, connecting the auxiliary tool to the bottom half shell via axially displaceable adapter slides arranged on said auxiliary tool, axially displacing the bottom half shell to release the bottom half shell from axial connecting contours, rotating the bottom half shell with the rotor by approximately 180° until the bottom half shell assumes essentially the position of the removed top half shell, and releasing the bottom half shell and removal of said bottom half shell.

Abstract: The present disclosure relates to a method for improving aerdynamic stability of a working fluid flow through a compressor of a turbomachine, in particular through a compressor of gas turbine used for power production, particularly against rapidly changing aero speed of the compressor. The method comprises to introduce a first water mass flow to the working fluid flow of the compressor. Furthermore, the disclosure relates to a turbomachine, in particular a gas turbine, which can be driven according to the above method.

Abstract: An axial flow turbine blade and diaphragm construction comprises an annulus of static turbine blades 20 and inner and outer spacer rings 40, 42 having apertures 41, 43 shaped to accommodate inner and outer platform portions 22, 23 of the blades, whereby the platform portions and the spacer rings together form inner and outer port walls of the turbine diaphragm. A particular feature of the blades is that the inner and outer platform portions 22 and 23 have straight side edges 22B, 22C and 23B, 23C, which are joined to each other by curved leading edges 22D and 23D that in plan view have a shape that follows the edge of the corner fillet 24 in the region of the leading edge 26 of the aerofoil 21.

Abstract: A modification method for reducing emissions from an annular shaped combustor of a gas turbine plant, having uniformly spaced circumferentially mounted premix burners (20), includes the steps of: removing at least one burner (20), thereby disrupting the spatial uniformity of the remaining the burners (20); and modifying the combustor air distribution system so as to compensate for the increased burner pressure drop of the remaining burners, thus enabling the modified combustor to operate at a load equivalent to the unmodified combustor. Emission reduction is enabled by the increase in the gas velocity of the burner for a given load further enabled by the flame stabilizing effect of disrupting the spatial uniformity.

Abstract: A gas turbine vane to improve vane performance by addressing known failure mechanisms. A cooling circuit to the trailing edge of a vane airfoil is fed from the outer diameter platform, which prevents failure due to an oxidized and eroded airfoil trailing edge. The gas turbine includes an outer diameter platform, a hollow airfoil and an inner diameter platform with a plurality of cooling tubes extending radially through the airfoil. The cooling tubes are open at the outer diameter end and closed with covers at the inner diameter end. The inner diameter platform is also cooled and includes a meterplate for a portion of the cooling passageway and includes an undercut to improve thermal deflections of the inner diameter platform.

Abstract: A shaft sealing arrangement includes at least two seal rings radially enclosing a rotating shaft so as to define a clearance between the shaft and the at least two seal rings. A sealing oil supply channel is disposed between the at least two seal rings and a pump is configured for pumping a sealing oil through the sealing oil supply channel in a radial direction towards the shaft so that the sealing oil subsequently flows through the clearance in a first axial direction toward a first side of the sealing arrangement and in a second opposite axial direction toward a second side of the sealing arrangement. The sealing oil is collected on the first and second sides and at least partially reintroduced into the sealing oil supply channel.

Abstract: A gas turbine engine component having an axial pin stop assembly that reduces stress around a seal pin damper slot between a platform and a buttress is disclosed. The seal pin damper slot is extended to the leading edge of the buttress and an axial pin stop seal extends into the pin damper slot to limit movement of the seal pin damper while relieving stress concentration at the interface between the buttress and platform.

Abstract: A gas turbine engine component having shaped cooling holes that further enhances the cooling of a desired region while reducing stress levels in and around the cooling holes is disclosed. The cooling holes are generally elliptically-shaped and diffuse from a cooling fluid supply side to a discharge side and are oriented on the turbine component to reduce stress concentrations while directing the cooling fluid to a desired surface or location. The elliptical cooling holes have openings in the surface that have high points that are concentric and planar.

Abstract: A compressor diaphragm for a gas turbine engine having improved wear capability, manufacturability, and assembly techniques is disclosed. The diaphragm includes a shiplap-type joint at an outer vane platform for connecting to adjacent vane assemblies and a clamshell-like assembly of a seal box secures and seals regions around the inner vane platform of the compressor diaphragm so as to reduce wear between the seal box and the vane assemblies. The inner platform of the diaphragm segments are fastened to each other through circumferentially-oriented fasteners at the inner diameter platform.

Abstract: A gas turbine blade having reduced stresses around a regions where the airfoil training edge joins the platform is disclosed. The lower stresses are achieved due to an undercut region proximate the blade trailing edge being placed in the platform. The undercut region has a compound fillet radius formed from two single and tangent radii where the first radius closer to the airfoil is larger than the second radius such that a ratio of R1/R2 is approximately 2.0 to 2.5.

Abstract: A combustion chamber arrangement is described for operating a gas turbine, with a combustion chamber wall which encloses the combustion chamber space and in the region of the combustion chamber outlet encloses a flow passage for hot gases which develop inside the combustion chamber, has a combustion chamber wall edge which freely terminates in the axial flow direction of the hot gases. The combustion chamber wall edge is formed with a profile which blocks or at least inhibits a diffuser effect when a cooling air flow, which is guided axially through the flow passages into the annular spatial area, flows over the combustion chamber wall edge.

Abstract: A method of removing unwanted coating material from cooling passages of a turbine component includes determining coordinates of the position and orientation of cooling passages at the surface of the turbine component and determining reference points in the region of the cooling passages. After coating of the turbine component, the reference points are measured once again and the thickness of the coating is calculated. In a basic processing program, the data for the position, passage orientation and coating thicknesses and also CAD data for the cooling passages are interlinked and a laser processing program is automatically adapted for each individual cooling passage. Using the laser processing program, a pulsed laser is guided over disk-shaped volume segments of the unwanted coating material and the material is removed in the process.