Alstom Technology Limited

Ansaldo Energia Switzerland Ag

Ansaldo Energia IP UK Limited

General Electric Technology Gmbh

Marco Micheli

Wolfgang Kappis

Luis Puerta

Gabriel Dunkel

Sasha Savic

Edward Len Miller

Wilfried Rick

Erik Peter Boldt

Daniel Glesti

Thomas Palkovich

Wifried Rick

Tihomir Stojkovic

Erik Boldt

Luis Federico Puerta

A method is disclosed for injecting water into a multistage axial compressor of a gas turbine. With low equipment cost, a significant power enhancement can be achieved, even under changing boundary conditions, by water being injected at a plurality of points along the axial compressor, and by the injected water mass flow being controlled at the individual injection points in accordance with ambient conditions and operating parameters of the gas turbine in such a way that an evened-out loading in the individual stages of the axial compressor can be created.

Method for injecting water into a multistage axial compressor of a gas turbine

An axial compressor of a fluid-flow machine includes a flow path disposed between a rotor shaft and a relatively stationary housing wall. The flow path extends in an axial direction of the rotor shaft concentrically with the rotor shaft and the housing wall. A plurality of compressor stages are axially arranged in sequence along the flow path in the axial direction. Each of the compressor stages includes a rotor blades row and a guide vane row disposed after the rotor blades row in the axial direction. The flow path and the compressor stages operating therein are penetrable by a mass flow of a fluid to be compressed in a flow direction during operation of the compressor. A return is configured to return a part of the mass flow from a compressor discharge.

Axial compressor for fluid-flow machines

The axial compressor has a two-stage guide vane cascade at the discharge-side end of the rotor. The guide vanes of the second stage of the cascade are staggered in the circumferential direction in relation to the guide vanes of the first stage in such a way that vortex streamers created by the guide vanes of the first stage cannot impinge upon the guide vanes of the second stage.

Axial compressor

A compressor component having an improved airfoil profile so as to eliminate previously known vibratory issues in the blade tip is disclosed. By altering the airfoil profile throughout its span, the natural frequency of the airfoil is altered so as to not coincide with a critical engine order of the compressor. Further, the present invention provides a novel airfoil profile in accordance with the coordinates of Table 1.

Compressor blade

The present disclosure provides an improved first stage airfoil for a compressor blade having a unique chord length (CD), stagger angle (&#x3b3;) and camber angle (&#x394;&#x3b2;). The stagger angle (&#x3b3;) and camber angle (&#x394;&#x3b2;) provide improved aerodynamics while the chord length (CD) provides for reduced airfoil weight.

Airfoil for a compressor blade

An axial compressor is disclosed for compressing air, such as for a gas turbine. The axial compressor can include a rotor, having a multiplicity of rotor blades, and rotor casing. The casing and rotor can form an annular gas passage. Outside the gas passage an annular and concentric bleed air chamber, can include at least one bleed air slot in functional communication with the gas passage. At least one injection device having at least one nozzle which can inject the fluid into the gas passage via the bleed air slot.

Axial compressor with an injection device for injecting a fluid

A rotor for a turbomachine includes at least one moving blade row including a plurality of moving blades disposed adjacent to one another in a circumferential direction with respect to an axis of rotation. Each one of the plurality of moving blades includes a blade leaf and a blade root having an outer face, the outer face of the blade root including a blade root curved end wall contour adjacent the blade leaf in the circumferential direction. The blade root curved end wall contour has a concave curvature in an axial sectional plane of the rotor. At least one rotor shaft includes a reception slot for a respective one of the at least one moving blade row extending in the circumferential direction. Each of the plurality of moving blades is inserted into the reception slot via a respective blade root.

Turbomachine rotor

The present disclosure provides an improved first stage airfoil for a compressor blade having a unique chord length (CD) stagger angle (&#x3b3;) and camber angle (&#x394;&#x3b2;). The stagger angle (&#x3b3;) and camber angle (&#x394;&#x3b2;) provide improved aerodynamics while the chord length (CD) provides for reduced airfoil weight.

The present disclosure provides an improved second stage airfoil for a compressor blade having a unique chord length (CD) stagger angle (&#x3b3;) and camber angle (&#x394;&#x3b2;). The stagger angle (&#x3b3;) and camber angle (&#x394;&#x3b2;) provide improved aerodynamics while the chord length (CD) provides for reduced airfoil weight.

An improved second stage airfoil for a compressor blade having a unique chord length (CD), a stagger angle (&#x3b3;), and camber angle (&#x394;&#x3b2;). The stagger angle (&#x3b3;) and camber angle (&#x394;&#x3b2;) provide improved aerodynamics while the chord length (CD) provides for reduced airfoil weight.

The disclosure provides blades, and the modification thereof, for stages 18-22 of an axial compressor wherein the blades have reduced susceptibility to tip cracking. The blades and blades manufactured by the provided method have a thickened profile that results in reduced stress in response to multi frequency impulses and can have increased frequency response of the chord wise bending mode.

Blade for an axial compressor and manufacturing method thereof

When a turbocompressor is being started and/or accelerated, in particular a compressor in a gas-turbine set, the position (VIGV) of a row of variable inlet guide vanes is varied dynamically; in particular, the row of inlet guide vanes is closed during acceleration.

Method for operating a turbocompressor

A method for modifying a multistage compressor () involves exchanging the rotor blades of the first compressor rotor blade row (LA) for modified rotor blades which have an identical blade leaf profile () to the original rotor blades and the blade angle (B') of which is different from the blade angle (B&#x2032;) of the original rotor blades. Furthermore, the blades of at least one further blade row (LAN, LEN) arranged downstream of the second compressor stage are exchanged for modified blades which have an identical blade leaf profile () to the original blades and the blade angle (B&#x2032;, B&#x2033;) of which is different from the blade angle (B&#x2032;, B&#x2033;) of the original blades. The method makes it possible to increase the mass flow of a compressor and essentially maintain the stability reserve against stall.

Method for modifying a multistage compressor

In a method of modifying a turbocompressor, the compressor blades are replaced by new compressor blades in such a way that the delivered mass flow of the compressor increases. At the same time, the capacities for blowing off during the start-up are increased.

Growing community of inventors

Wettingen, Switzerland

Marco Micheli