1681results about "Turbine/propulsion air intakes" patented technology

A method of generating energy in a power plant (30) having a gas turbine (29), includes a first step a gas containing air (1) is compressed in a first compressor (2) of the gas turbine (29), a second step the compressed gas (3, 3a, 3b; 5; 7a, 7b) is fed to a combustion process with the addition of fuel (8) in a combustor (23), a third step the hot flue gas (9) from the combustor (23) is expanded in an expander or a turbine (10), driving a generator (18), of the gas turbine (29) while performing work, and a fourth step a partial flow of the expanded flue gas (11) is recirculated to the inlet of the first compressor (2) and admixed with the gas containing air (1). Carbon dioxide (CO₂) is separated from the compressed gas (3, 3a, 3b; 5; 7a, 7b) in a CO₂ separator (6) before the third step. In such a method, the overall size and energy costs are reduced by virtue of the fact that, to permit increased CO₂ concentrations in the CO₂ separator (6), not more than about 70% of the carbon dioxide contained in the compressed gas (3, 3a, 3b; 5, 5a, 5b; 7a, 7b) is removed from the compressed gas (3, 3a, 3b; 5, 5a, 5b; 7a, 7b).

Embodiments relate to hydraulic fracturing equipment powered by one or more natural gas turbine generators. Natural gas from a supply line is released via a valve into a turbine gas line. The turbine gas line includes one or more regulators to reduce the pressure of the natural gas stream in the turbine gas line to a pressure or pressure range optimum for one or more gas compressors. The gas compressors increase the pressure of the natural gas stream, which is then directed to one or more natural gas turbine generators. The natural gas turbine generators combust the natural gas to produce electricity, which powers electric hydraulic fracturing equipment.

An inlet air treatment system. The inlet air treatment system includes a housing with more than two symmetric sides. Each of the sides includes a number of air filters and a number of spray arrays. A central air deflector is positioned within the housing.

An airfoil and method of fabricating an airfoil including a first and a second side coupled together at a leading and a trailing edge and extending there between. The airfoil includes a plurality of first chord sections having a first chord length and extending outward from one of the first side or second side of the airfoil at the leading edge and a plurality of second chord sections having a second chord length and extending outward from the one of the first side or the second side of the airfoil at the leading edge. The leading edge including spaced-apart wave-shaped projections defining a waveform. The configuration defining a three-dimensional crenulated airfoil configured to facilitate desensitization of an airfoil unsteady pressure response to at least one impinging upstream generated wake or vortex by decorrelating spatially and temporally and reducing in amplitude an unsteady pressure caused by interaction of the airfoil with the upstream generated wake or vortex.

A system including an air intake subsystem having a housing is provided. A coalescer having a frame is disposed inside the housing. The system includes a mounting bracket, and a release mechanism coupled to the frame and the mounting bracket. The release mechanism is adapted to selectively release the frame.

The invention discloses a sound insulation cabin of a turbine engine. The sound insulation cabin is arranged on the turbine engine in a sleeving mode, and comprises a cabin body, an intake noise reduction unit and a ventilation noise reduction unit; the intake noise reduction unit and the ventilation noise reduction unit are arranged on the cabin body, and the periphery of the cabin body is filledwith sound insulation materials; and the intake noise reduction unit is used for intake noise reduction of the turbine engine, and the ventilation noise reduction unit is used for noise reduction ofa ventilation system of the turbine engine. The sound insulation cabin has the beneficial effects that the intake noise reduction unit is additionally arranged at an air inlet of the turbine engine, and intake noise reduction of the turbine engine is achieved; the ventilation noise reduction unit is additionally arranged in the transmission direction of the turbine engine, and the ventilation andcooling noise reduction for the turbine engine is achieved; the periphery of the cabin body is filled with the sound insulation materials, and the overall noise reduction for the periphery of the turbine engine is achieved.

In an air intake structure (14) of an aircraft engine pod, the link or connection between the air intake lip (15), the front reinforcing frame (16) and the acoustic panel (17) is implemented in such a way that the lip (15) can be dismantled without breaking the link between the frame (16) and the panel (17). Moreover, the internal, rear part (15a) of the lip (15) normally covers the front part (17a) of the panel (17), as well as the members (22, 24) ensuring the link between the latter and the frame (16).

A low drag ducted ram air turbine generator and cooling system is provided. The ducted ram air turbine generator and cooling system has reduced drag while extracting dynamic energy from the air stream during the complete range of intended flight operating regimes. A centerbody/valve tube having an aerodynamically shaped nose is slidably received in a fairing and primary structure to provide a variable inlet area. An internal nozzle control mechanism attached to the valve tube positions nozzle control doors to provide variable area nozzles directing air flow to the turbine stator and rotor blades to maintain optimum generator efficiency. An alternate embodiment includes an annular internal nozzle having interleaved panels to modulate the air flow to the turbine.

The invention discloses continuous large-power turbine fracturing equipment. The large-power turbine fracturing equipment comprises a turbine engine, a reduction box, a transmission shaft and a plunger pump, wherein the turbine engine and the reduction box are arranged on the same straight line; the reduction box is connected to the plunger pump through a transmission shaft; and an angle of the transmission shaft ranges from 2 degrees to 4 degrees. The continuous large-power turbine fracturing equipment has the beneficial effects that: a chassis T1 material is selected, so that a stable working platform is provided for the equipment; the turbine engine and the reduction box are on the same straight line, the transmission shaft is arranged between the reduction box and the plunger pump, andthe angle of the transmission shaft ranges from 2 degrees to 4 degrees, so that stable and efficient transmission of the turbine engine is guaranteed, and a fault occurrence rate is reduced; a lubricating system driven by an auxiliary power system guarantees that the turbine engine, the reduction box and the plunger pump work under a proper environment; dual lubricating systems guarantee that theplunger pump realizes continuous operation under power of 5000 HP or higher; and the technical means finally can meet continuous large-power operation requirements of the fracturing equipment.

In a gas turbine power generation system having a gas turbine engine and a generator connected to the engine, maintenance ease is enhanced by forming the housing to have openable maintenance faces at two of its faces. Installation space is reduced and a compact configuration is achieved by providing a partition that divides the interior space of the housing into two regions into an upper bay and a lower bay, mounting the engine in the upper bay and mounting air intake duct in the lower bay at a location directly under the engine. Noise is reduced by using a partition to define two separate spaces (bays) and mounting the engine and the air intake duct in the upper and lower bays. Noise is also reduced by preventing engine rotation noise from escaping to the outside through the air intake duct. Moreover, it achieves a reduction in the amount of dedicated space required by constituting the air intake duct of a duct section having an air inlet at a plane coincident with that of the maintenance opening and a filter-housing section for removably housing an air filter, thereby enabling a common space to be utilized for that required in front of the maintenance opening and the air inlet.

An aircraft propulsion system includes a gas turbine engine having a fan section, at least one row of FLADE fan blades disposed radially outwardly of and drivingly connected to the fan section, the row of FLADE fan blades radially extending across a FLADE duct circumscribing the fan section, an engine inlet including a fan inlet to the fan section and an annular FLADE inlet to the FLADE duct. A fixed geometry inlet duct is in direct flow communication with the engine inlet. The fan section may include only a single direction of rotation fan or alternatively axially spaced apart first and second counter-rotatable fans in which the FLADE fan blades are drivingly connected to one of the first and second counter-rotatable fans. The row of FLADE fan blades may be disposed between rows of variable first and second FLADE vanes.

Within an intake for a gas turbine engine provision is provided through stator vanes 25, 45 in the stator whereby back pressure from a necessary obstruction 34, 46 can be utilised to balance forward pressure variations caused by intake droop or crosswinds in order to reduce those forward pressure detriments for more efficient engine operation according to a desired objective regime. Typically the flow through the intake 20 is analysed and then an appropriate positioning of the stator vanes 25, 45 determined in order to provide approximate balance between the forward pressures and back pressures. Normally, a combination of camber variation of stator vanes 45a and stagger variation of stator vanes 45b are utilised in order to achieve a desired momentum balance around the circumference of the intake 20. It will be understood that both the forward pressures and the back pressures are differentially variable about the circumference such that one opposes the other.

A wall configuration of an axial-flow machine which can reduce the secondary flow loss is provided. A trough is formed between a blade and another blade in the blade row and extends in at least an axial direction of the blade row. The region where the trough is formed is axially between a leading edge and a trailing edge of the blade. A center line of the trough has a curvature in the same direction as a camber line of the blade. A maximum amplitude of the trough is located adjacent to an axial center of the blade or located axially between the axial center and the leading edge of the blade.

The invention discloses a hypersonic aerocraft and air inlet internal and external waverider integrated design method, and relates to a near space aerocraft. An aerodynamics characteristic is firstly appointed, and then a design scheme meeting the characteristic is inferred backwards; a three-dimensional shock wave curved surface in a complex shape is appointed, the change rule of the transverse curvature center is obtained, and a series of basic flow fields meeting the needs of the waverider design are inferred backwards according to the change rule; flow lines of different curvature centers and different radial positions are traced in every basic flow field in the circumferential direction; a waverider device capable of producing the appointed complex three-dimensional shock wave curved surface is obtained finally, namely the integrated design scheme is obtained. The advantages of a waveriders and an internal waverider air inlet are kept, the integrated design of the two high-performance devices is achieved, the waverider model with high lift-drag ratio and the scheme of the air inlet with full-flow capture can be obtained at the same time, and accordingly the overall performance of the aerocraft is improved.

A gas turbine comprises a compressor for compressing a gas supplied therein and discharging the compressed gas, a combustor in which the discharged gas from the compressor and a fuel are combusted, a turbine to be driven by a combustion gas from the combustor, and a water injection unit which injects water into the gas to be supplied to the compressor, thereby lowering the temperature of the gas to be introduced into the compressor to a temperature lower than the atmospheric temperature, and causing water droplets having been injected in the gas and within the compressor to be vaporized while flowing down therein, wherein the quantity of water spray injection is controlled while monitoring operational conditions of the gas turbine.

A bleed air duct that preferably includes an inlet section configured to include a flush scoop and a louver. The louver is located and configured such that in the desired operating flow range of the duct, the fluid entering the flush scoop is disturbed and as a result creates a low pressure region downstream of the louver. The low pressure region substantially eliminates the generation of any pressure pulses and acoustic resonance also known as Helmholtz resonance.

The present invention relates to a nacelle (1) for a turbojet engine comprising an air intake structure (4) capable of ducting a flow of air towards a fan and a central structure (5) intended to surround the said fan and to which the air intake structure is attached, the central structure comprising a casing (9) intended to surround the fan and an external structure (13), characterized in that the air intake structure comprises, on the one hand, at least one internal panel (41) attached to the central structure via the casing and therewith forming a fixed structure and, on the other hand, at least one longitudinal external panel (40) attached removably to the fixed structure and incorporating an air intake lip (4a), the said external panel being removable between an operating position in which the external panel is aerodynamically continuous with the external structure of the central section and the air intake lip provides aerodynamic continuity with the internal panel of the air intake structure, and a maintenance position in which the external panel is separated from the external structure of the central section and the air intake lip is separated from the internal panel of the air intake structure.

Embodiments of the invention relate to a supersonic inlet having a cowl lip configured to capture the conic shock and exhibit a zero or substantially zero cowl angle. The inlet may be configured to employ a relaxed isentropic compression surface and an internal bypass. The nacelle bypass may prevent flow distortions, introduced by the capture of the conic shock, from reaching the turbomachinery, thereby allowing the cowl angle to be reduced to zero or substantially zero. Such a cowl angle may reduce the inlet's contribution to the overall sonic boom signature for a supersonic aircraft while allowing for an increase in engine pressure recovery and a subsequent improvement in generated thrust by the engine.