649 results about "Compressor blade" patented technology

An airfoil for use as rotor blades in compressors for turbomachines, such as gas turbine engines. The airfoil includes increased forward sweep and forward dihedral effective to reduce losses generated by interaction of tip clearance flow, secondary flows and passage shocks.

The present invention discloses a method and system for augmenting shaft output of stationary gas turbines that can be used in multiple modes of operation. The system comprises a washing unit (25, 27, 28) adapted to inject a spray (26) of atomized liquid so as to impinge on the compressor blades (12) in order to wet said blades (12), thereby obtaining a release of fouling material from said blades (12); and at least one liquid injection unit (21, 23, 24, 29, 210, 212, 214, 215, 216) adapted to inject a spray (22, 211, 213) of atomized liquid into an air stream of said turbine duct (101) or at the gas turbine (10) in order to increase a mass flow of said air flow, wherein the power output from said gas turbine engine can be augmented. With the invention follows also benefits such as fuel savings and improved environmental performance by reduction of emissions.

Systems and methods are provided for varying the exhaust gas recycle circuit of low emission gas turbines. In one or more embodiments, the systems and methods incorporate alternatives to the use of a direct contact cooler. In the same or other embodiments, the systems and methods incorporate alternatives intended to reduce or eliminate the erosion or corrosion of compressor blades due to the presence of acidic water droplets in the recycled gas stream.

The present invention thus provides an improved method for coating turbine engine components. The method utilizes a cold high velocity gas spray technique to coat turbine blades, compressor blades, impellers, blisks, and other turbine engine components. These methods can be used to coat a variety of surfaces thereon, thus improving the overall durability, reliability and performance of the turbine engine itself. The method includes the deposition of powders of alloys of nickel and aluminum wherein the powders are formed so as to have an amorphous microstructure. Layers of the alloys may be deposited and built up by cold high velocity gas spraying. The coated items displayed improved characteristics such as hardness, strength, and corrosion resistance.

In one exemplary embodiment, a gas turbine engine includes a turbine and a high pressure compressor. The high pressure compressor includes a last stage having a last stage compressor blade and a last stage vane. The gas turbine engine includes a first flow path through which bleed air flows to the turbine and a second flow path through which air from the last stage of the high pressure compressor flows. The bleed air in the first flow path exchanges heat with a portion of the air in the second flow path in a heat exchanger to cool the air in the second flow path. The cooled air in the second flow path is returned to the high pressure compressor to cool the high pressure compressor.

A motor-assisted turbocharger is composed of a turbine driven by energy of exhaust gas, a compressor rotated by the turbine and a rotary electric machine for assisting rotation of the compressor. The turbine, the compressor blade and the rotary electric machine are connected to each other by a common rotating shaft. The compressor blade made of a material including a magnetic material faces an inner surface of a housing in which a magnetic sensor is embedded. A magnetic field in an air gap between the compressor blade and the magnetic sensor changes according to rotation of the compressor blade. The magnetic sensor detects changes in the magnetic field to thereby detect the rotational speed of the compressor blade. In place of the magnetic field, other physical amounts in the air gap, such as pressure, sound frequencies, capacitance or the like may be used for detecting the rotational speed.

A turbofan engine includes a combined stage rotor mounted for common rotation with a shaft, a compressor stage, and a turbine stage. Each of the plurality of fan blades of the combined stage rotor are contiguous with a respective compressor blade. The combined stage rotor is located at a junction between a fan bypass duct and a core duct. The fan bypass stage portion includes a plurality of the fan blades which communicate incoming airflow into fan bypass airflow through the fan bypass duct. The compressor stage portion includes a plurality of the compressor blades which communicated airflow into core airflow through the core duct. The plurality of fan blades extend substantially perpendicular to the plurality of compressor blades such that the plurality of compressor blade portions extend at least partially into the core duct and preferably will include splitters for enhanced efficiency.

A gas turbine engine comprises a fan and a turbine section having a first turbine rotor. The first turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from a fan drive turbine rotor. The compressor rotor has a number of compressor blades in at least one of a plurality of rows of the compressor rotor. The blades operate at least some of the time at a rotational speed. The number of compressor blades in at least one row and the rotational speed are such that the following formula holds true for at least one row of the compressor rotor:

(the number of blades×the rotational speed)/(60 seconds/minute)≧5500 Hz; and

• • the rotational speed being in revolutions per minute.

A compressor module and a method of designing a gas turbine engine are also disclosed.

A coating system and process capable of providing erosion and corrosion-resistance to a component, particularly a steel compressor blade of an industrial gas turbine. The coating system includes a metallic sacrificial undercoat on a surface of the component substrate, and a ceramic topcoat deposited by thermal spray on the undercoat. The undercoat contains a metal or metal alloy that is more active in the galvanic series than iron, and electrically contacts the surface of the substrate. The ceramic topcoat consists essentially of a ceramic material chosen from the group consisting of mixtures of alumina and titania, mixtures of chromia and silica, mixtures of chromia and titania, mixtures of chromia, silica, and titania, and mixtures of zirconia, titania, and yttria.

A turbine or compressor blade assembly includes a blade fixed to a dovetail section attachable to a wheel. The dovetail section has a dovetail shaped to fit in a correspondingly shaped slot in the wheel. A dovetail platform serves as an interface between the blade and the dovetail. An undercut fillet radius is formed at an intersection of the dovetail platform and a dovetail pressure surface, where the undercut radius has a multi-part profile shape configured to attenuate edge of contact stresses. An additional feature is the area where the undercut radius transitions into the P-cut area at the forward end (leading edge) of the dovetail.

A compressor blade (26) comprises an aerofoil (36) and a root (38). The aerofoil (36) comprises a concave wall (40) extending from a leading edge (44) to a trailing edge (46) and a convex wall (42) extending from the leading edge (44) to the trailing edge (46). The aerofoil (36) defines at least one internal chamber (48). The root (38) is connected to the aerofoil (36) and the root (38) has a base (50) remote from the aerofoil (38) and at least one aperture (52) extends from the base (50) to the at least one internal chamber (48) in the aerofoil (36). The dimensions, shape and position of the least one aperture (52) are selected such that the root (38) is deformable. This reduces the weight of the containment region of a fan casing (28).

The invention relates to a CAD (Computer-Aided Design) aided design method of a controllable diffused blade profile with curvature continuity, which belongs to the field of design of a compressor blade. The method is characterized by comprising the steps of respectively solving a leading edge point by using the point of tangency of a pressure surface, a suction surface and a small circuit at the leading edge; fitting with a cubic spline to solve the curvature of each data point on the suction face and the pressure surface; solving a side profile of the suction face and a side profile of the pressure surface of the leading edge with curvature continuity by a cubic Bezier curve; giving the curvature distribution of the suction face by a B spline curve to solve a control point of the B spline curve, so that the deviation between the profile of the suction face with curvature continuity and the original profile of the suction face corresponding to curve distribution is minimum; and solving the position of the small circle at a tail edge of the optimized blade profile. Compared with the original controllable diffused blade profile, the range of the angle of attack is expanded by 2.5 degrees, and the pitot loss in a low loss area is reduced.

A compressor blade for an aircraft engine includes a blade core made of a fiber compound material and a metallic enclosing structure. The enclosing structure is of a multi-part design and includes blanks (4, 5) attached to the blade core by a metallic weave (6) on the suction side and on the pressure side, with the two blanks being firmly connected at the aerodynamically shaped leading edge of the compressor blade by to a leading-edge former (3). Depending on the specific loads applied on the pressure side and on the suction side, welds attaching the blanks to the leading-edge former (8, 9) are offset relative to each other and the material thicknesses of the components of the enveloping structure are matched to each other.

A turbocharger monitor including a sensor adapted and positioned with respect to a turbocharger housing to sense passing compressor blades, generating both rotor rotation information and rotor displacement information. A controller is configured to receive the generated information from the blade sensor, and to establish rotor rotation speed and an indicator of bearing condition.

A method and system augments shaft output of gas turbine engines that can be used in multiple modes of operation. The system comprises a washing unit capable of injecting atomized water into the gas turbine engine, thereby obtaining a release of fouling material from the at least one compressor blade; and at least one water injection unit capable of injecting atomized water into the air stream of the gas turbine engine's inlet duct or at the gas turbine, under the control of a computational fluid dynamic model, in order to increase a mass flow of said air flow, wherein the power output from said gas turbine engine can be augmented.

The invention discloses an efficient and compact pre-cooling heat exchanger for a pre-cooling gas sucking type engine. The efficient and compact pre-cooling heat exchanger for the pre-cooling gas sucking type engine comprises a pre-cooling heat exchanger body in the shape of a similar cylinder; the cross section of the similar cylinder is in the shape of a ring different in diameter; the longitudinal section of the similar cylinder is formed by two symmetrically arranged arc segments; the center angle of each arc segment is 45 degrees; the pre-cooling heat exchanger body comprises a cold flow outlet pipe, a cold flow inlet pipe and a plurality of heat exchanging pipes; each heat exchanging pipe is in the shape of an arc; the arc-shaped center angle of each heat exchanging pipe is 45 degrees; and the heat exchanging pipes are distributed into the similar cylinder in a crossed manner. The efficient and compact pre-cooling heat exchanger for the pre-cooling gas sucking type engine can have the engineer realizability on the premise that the heat exchanging effect is guaranteed, so that air incoming flow in front of a gas compressor can be pre-cooled, the high temperature over-running problem of gas compressor blades is solved, the engine thrust is increased, the flight envelop is extended, the welding difficulty is effectively reduced, the frosting blockage problem is effectively solved, the engineer realizability is achieved under the current technology, and after the efficient and compact pre-cooling heat exchanger for the pre-cooling gas sucking type engine is applied to the pre-cooling gas sucking type engine, the engine performance can be substantially improved.

A counter-rotating blade stage in lieu of a stator stage may compensate for relatively low rotational speed of a gas turbine engine spool. A first spool may have at least one compressor blade stage and at least one turbine blade stage. A combustor is located between the at least one compressor blade stage and the at least one turbine blade stage along a core flowpath. The at least one counter-rotating compressor blade stage is interspersed with the first spool at least one compressor blade stage. A transmission couples the at least one additional compressor blade stage to the first spool for counter-rotation about the engine axis.