44results about How to "Reduce blending loss" patented technology

The invention discloses a blade hole type prewhirl nozzle for a prewhirl cooling system, and belongs to the field of application of aero-engine prewhirl cooling systems. The blade hole type prewhirl nozzle is characterized in that a blade pressure face and a blade suction face are adopted for forming inclined hole runners deflecting in the circumferential direction of a nozzle ring at equal intervals, deflecting molded faces of the nozzle runners are in a gradually-contracted shape, a minimum laryngeal area is achieved at an outlet of the prewhirl nozzle, the accelerating performance and the deflecting performance of the prewhirl nozzle are effectively guaranteed, and the flowing loss of a main flow is small. Blade type solid blocks are distributed between adjacent runners of the prewhirl nozzle, under the condition that the laryngeal area of the nozzle is definite, the ratio of the blade height to the pitch of the prewhirl nozzle can be flexibly adjusted by additionally arranging the blade type solid blocks, and the long and narrow performance of the sections of the runners is effectively improved, so that the secondary flow loss of the end walls in the runners is reduced, the flow coefficient and the prewhirl efficiency of the prewhirl nozzle are improved, and the system temperature drop effect is further improved.

The invention belongs to the technical field of aviation and space, and particularly relates to a high pressure annular jet pump suitable for a pumping pressure type supply system. The high pressure annular jet pump comprises an inlet section pipe, a mixing section pipe, a throat pipe, a diffusion section pipe and an outlet section pipe which are sequentially arranged in the flowing direction of main fluid. The high pressure annular jet pump is characterized by further comprising step section pipes which are sequentially arranged between the inlet section pipe and the mixing section pipe; each step section pipe comprises a convergence section pipe and a straight section pipe; each straight section pipe is provided with a spraying nozzle for providing high pressure fluid. The high pressure annular jet pump is compact in structure, light in weight, good in anti-cavitation performance and capable of meeting the requirement for normal work under the working conditions of high rotating speed and low inlet pressure.

The invention relates to a treatment casing applied to an axial-flow air compressor in the field of impeller machinery, in particular to an unsteady standing vortex type treatment casing, which can be used for improving the stability margin of the axial-flow air compressor without reducing efficiency. The unsteady standing vortex type treatment casing comprises an air jet groove, a bridge-type air flow channel and an air suction groove, wherein the air jet groove is arranged at the position of a suction surface of a blade, and the air suction groove is arranged at the position of a pressure surface of the blade. The treatment casing is designed to suppress leakage flow generated on the suction surface of a front edge of the blade, the air suction groove is exactly close to the pressure surface in the rear middle of the blade when the blade sweeps over, and the air jet groove is close to the suction surface of the front edge of the blade. Therefore, differential pressure is formed between the air suction groove and the air jet groove to drive fluids to spray from the downstream of the blade to the upstream of the blade, and the air jet groove is positioned close to the suction surface of the front edge of the blade and exactly aligned to a leakage vortex generating area of the blade, so that leakage flow can be suppressed, and further the stall margin of the air compressor is increased. Moreover, as an air jet direction is consistent with a main flow direction, mixing loss of an air jet area and a main flow area is decreased, the momentum of a boundary layer of the wall surface of the casing is increased, and positive effects are achieved for the efficiency of the air compressor.

A turbine cooling cascade with a vortex structure relates to a turbine cooling cascade. The invention settles the problems of flowing loss, blending loss and easy burning of blade leading edge and end wall surface of blade leading edge caused by the existence of horse shoe vortex structure in the prior blade cascade. According to the blade (2) of the invention, a pressure surface (2-2) next to the blade leading edge (2-1), a suction surface (2-3) of blade (2) next to the blade leading edge (2-1) and the blade leading edge (1-1) of wall (1) are respectively provided with a plurality of elongated slots (3) which are distributed in row. One side wall of each elongated slot (3) is provided with a plurality of cold air injection openings (4). Furthermore the positions of cold air injection opening (4) are next to the base of elongated slot (3). The turbine cooling cascade with vortex structure according to the invention reduces the flowing loss at the inner part of flow passage, reduces the bigger blending loss of air currents with different flowing velocities, provides better thermal protection function for the end wall of blade and the blade surface, effectively protects the end wall adjacent with the blade leading edge and the high-temperature area on the surface of blade, and increases the usage factor of cold air.

The invention relates to the technical field of gas turbines, in particular to a gas turbine rim sealing structure. The structure comprises a static vane connected to a static vane platform and a movable vane connected to a movable vane platform, wherein an airtight cool air channel is formed between the static vane platform and the movable vane platform, blowing holes are formed in the peripheral direction of the static vane platform and communicated with a static vane cooling channel, an outlet of each blow hole is located on the downstream side of the static vane platform and communicated with the airtight cool air channel, and the blow holes are used for guiding cooling air in the cooling channel out and making cooling air form an air curtain at the position of an outlet of the airtight cool air channel. The invention further relates to a rim sealing method. By the adoption of the method, cooling air in the static vane cooling channel can be guided out from the downstream side of the static vane platform, and the air curtain can be formed by cooling air at the position of the outlet of the airtight cool air channel. The air curtain can improve the airtight performance of the gas turbine rim sealing structure, reduce airtight cool air amount, reduce the influence of airtight cool air on aerodynamic efficiency of main current, and improve overall efficiency.

The invention discloses a centrifugal compressor diffusor structure with blades being integrated with a case and a hub. The centrifugal compressor diffusor structure with the blades being integrated with the case and the hub belongs to the technical field of middle-small aero-engine compressors, and is characterized in that a radial diffuser and an axial diffuser are connected into a whole, and aleading edge is designed into a sharp edge, so that the loss is reduced, and the outer diameter of each diffuser is reduced; and the blades are further integrated with the case of each diffuser and the hub, so that the distortion degrees of the blades of each diffuser are reduced, and the blades are convenient to design and process. The centrifugal compressor diffusor structure with the blades being integrated with the case and the hub provided by the invention can be directly applied to small aero-engines, the outer diameter of the engine can be reduced without increasing the processing costand reducing the compressor performance, and the thrust-weight ratio of the engine is improved.

The invention belongs to the field of gas turbine engines, and relates to a turbine disk cavity sealing structure with bypass gas entraining. The turbine disk cavity sealing structure with bypass gasentraining comprises an inner support ring assembly provided with an inner support ring gas inlet hole, a front baffle plate assembly, a pre-rotation nozzle, a first gas passage for transferring a first path of cold gas output by the pre-rotation nozzle, a second gas passage for transferring a second path of cold gas output by the pre-rotation nozzle, and a third gas passage arranged between the inner support ring assembly and the front baffle plate assembly, wherein the third gas passage comprises a first sealing device arranged between the inner support ring assembly and the front baffle plate assembly, a sub gas passage and a bypass gas entraining passage; a third path of cold gas output by the pre-rotation nozzle flows into the sub gas passage and the bypass gas entraining passage through the first sealing device; and gas output by the third gas passage flows toward a high-pressure turbine movable blade.

The invention discloses a turbulent baffle heat transfer enhancing device with jet impingement function, and relates to the heat transfer technology. The heat transfer enhancing structure with impingement and turbulence functions utilizes a zigzag channel arranged along a blade expanding direction, controls the section surface of a circulation part by regulating a baffle structure in the channel to achieve the aim of enhancing heat transfer, and is suitable to be arranged on a trailing edge of a turbine blade of a gas turbine with narrow space and high heat load to form a closed cooling structure. The jet baffle turbulent structure of the invention has simple structure and high heat exchange capacity, can meet the cooling requirement of the trailing edge of the turbine blade of the gas turbine through high-efficiency impingement cooling function, and can effectively cool the high-temperature trailing edge of the turbine blade on the premise of closed cycle of a cooling medium, reduce mixing loss caused by cooling, and make the temperature on the trailing edge of the blade distributed more evenly compared with other open cooling structures applied to the trailing edge of the turbine blade.

The invention discloses a pressure expander blade, a compressor structure and a compressor. The pressure expander blade comprises a blade body, and a cavity is formed in the blade body. Air supplementing holes are formed in the blade body. According to the pressure expander blade, the compressor structure and the compressor, jet flow can be formed on the suction surface of the pressure expander blade by means of air supplementing through the pressure expander blade of the hollow structure and the air supplementing holes in the back portion of the pressure expander blade so as to blow down a low-speed and low-energy area formed on the suction surface, airflow mixing losses caused by air supplementing are reduced, and therefore aerodynamic efficiency of the centrifugal compressor is improved, and the operation range of the compressor can be widened while the aerodynamic performance of compressor design points is improved.

The invention relates to a closed centripetal turbine cover-machine case cavity structure for restraining leakage losses, and the closed centripetal turbine cover-machine case cavity structure is suitable for a compressed air energy storage system. An expansion runner, a seal structure and a leakage flow mixing restraining structure are sequentially arranged in a narrow and long cavity formed between a front wheel cover and a machine case. The circulation area of the expansion runner is gradually increased in the flowing direction. The leakage flow mixing restraining structure is a shielding sheet from a cavity outlet to a seal outlet. On the basis of leakage flow control through a conventional seal structure, the inlet leakage flow can be reduced through the cavity expansion runner, leakage flow and main flow mixing losses are further reduced through leakage flow mixing restraining, finally, the pneumatic efficiency and the acting capability of a turbine are improved, and the rate ofutilization of energy through the turbine is increased. The closed centripetal turbine cover-machine case cavity structure can be widely applied to aerospace, transportation, petrochemical engineering, energy power and other multiple fields besides compressed air energy storage and is particularly suitable for high-pressure radial-flow turbines.

The invention provides a solar heating system based on dissipation-reduction optimization and a control system of the solar heating system. The solar heating system comprises a solar heat collector, a heat storing water tank, an indoor heat dissipater, a heat collecting circulating pump, a heat supply circulating pump, connecting pipelines, connecting valves, a digital temperature sensor, a flowmeter, a frequency converter and a controller. According to the solar heating system based on dissipation-reduction optimization and the control system of the solar heating system, optimized design is conducted on the integrated form, an operation mode and a control method of the links of solar heat collecting, heat storing, heat supplementing and heat supplying from the perspective of energy dissipation, and the common problems that the solar heating system conducts reverse heat transferring through the heat collector, control targets of supply and storage of the heat collecting volume cannot be coordinated, the heat of the heating side is transmitted reversely to the water tank, the control target of heat supply of the water tank cannot be coordinated, simultaneous operation of heat supplementing and heat storing cannot be achieved, and heat storing is ineffective in application of solar opto-thermal heating are solved. Meanwhile, the solar heating system based on dissipation-reduction optimization and the control system of the solar heating system can maximally utilize the passive heat storage characteristic of a building, the requirement for heat storing volume is reduced effectively, and the mixing loss in the heat storing water tank is effectively reduced.

The invention provides a novel axial flow turbine suitable for pulse pressure charging. The novel axial flow turbine comprises a volute, stators, a rotor, a rotating shaft, a seal ring and a bearing. A gas inlet is formed in the front end of the volute, and a pressure expander is arranged at the rear end of the volute. An exhaust port is fixedly formed in the upper end of the pressure expander. The stators are each in an annular shape and located on the inner wall of the gas inlet of the volute. The front end of the rotating shaft is in an arc shape and located on the portion, close to the rear side, between the stators. The rear end of the rotating shaft is fixed to the rear side wall of the volute through the bearing. The rotor is fixed to the portion, close to the front end, on the rotating shaft. The front end of the pressure expander is located on the rear side of the rear side wall of the rotor. According to the novel axial flow turbine suitable for pulse pressure charging, work of the stators and work of the rotor are closer to the quasi-steady state, and therefore the efficiency of the rotor is improved; and in addition, flowing at the outlet of the rotor is uniform in the circumferential direction, and the mixing loss generated on the downstream part of the rotor due to circumferential-direction ununiformity is also reduced to some extent.

The invention discloses a compressor air supplementing structure and a compressor. The compressor air supplementing structure comprises a first diffuser, a first volute and a backflow device; the output end of the first diffuser is connected with the input end of the backflow device through the first volute, and the first volute is provided with an air supplementing opening. According to the compressor air supplementing structure, volute air supplementing is adopted as the air supplementing manner, and air supplementing refrigerants enter the compressor through an inlet of the first volute, flow in the circumferential direction of the first volute, and meanwhile enter a main flow zone through an inlet of the backflow device continuously. As the air supplementing refrigerants flow in the circumferential direction, and the air supplementing refrigerants have a high rotating speed which is the same as the main flow speed, mixing losses of supplemented air and main flow can be effectively reduced, and the aerodynamic efficiency of the centrifugal compressor is further improved.

The invention discloses a steam supplementing chamber structure of a steam turbine cylinder, and belongs to the technical field of steam turbines. The steam supplementing chamber structure of the steam turbine cylinder comprises an annular main steam flow channel formed between the cylinder and a rotor; a mixing ring cavity is formed in the cylinder in the circumferential direction of the cylinder; the outer side of the mixing ring cavity communicates with a steam supplementing channel; the inner side of the mixing ring cavity communicates with the main steam flow channel through a flow guide ring cavity; the flow guide ring cavity inclines towards the main steam flow flowing direction of the main steam flow channel; a plurality of flow guide partition plates are arranged in the flow guide ring cavity in the circumferential direction of the flow guide ring cavity; and the flow guide ring cavity is divided into a plurality of sub-cavities by the plurality of flow guide partition plates. According to the invention, the mixing degree and the mixing loss of the supplemented steam and the main steam flow can be reduced, and the influence of the supplemented steam on the stability of the rotor can be reduced.

The invention provides a static scroll plate assembly, a scroll compressor and refrigeration equipment. The static scroll plate assembly comprises a plate body, an enthalpy increasing channel and an enthalpy spraying channel, wherein the plate body is provided with a compression channel; the enthalpy increasing channel is arranged on the plate body, and the enthalpy increasing channel is used forcommunicating with an enthalpy increasing pipeline of the scroll compressor; and the enthalpy spraying channel is arranged on the bottom wall of the plate body, the enthalpy spraying channel communicates with the compression channel and the enthalpy increasing channel, the enthalpy spraying channel comprises a flow blocking part, and the flow blocking part is used for preventing gas in the compression channel from flowing back to the enthalpy increasing channel. The gas in the compression channel is prevented from flowing back to the enthalpy increasing channel through the flow blocking part,and spreading of the gas flow flowing back to the enthalpy increasing pipeline from the compression channel can be effectively attenuated, so that the pressure pulsation intensity in the enthalpy increasing pipeline is weakened, the technical effect of reducing the pressure pulsation level in the enthalpy increasing pipeline is achieved, and the problems that the joint between the enthalpy increasing pipeline and a shell of the scroll compressor is broken and a valve plate is broken due to the fact that the refrigerant gas in an enthalpy increasing pipeline generates strong pulsation in the prior art are solved.

The invention discloses a design for a transonic axial fan blade angular seam casing treatment structure, and belongs to the technical field of turbo machinery. According to the design for the transonic axial fan blade angular seam casing treatment structure, a plurality of narrow seams are formed in a fan casing in the axial direction and are uniformly distributed in the circumferential direction to form an axial seam casing treatment structure, and the width of the axial seam casing treatment structure covers the whole axial chord length of blade tips; and opening surfaces of the axial seam casing treatment structure rotate by a certain angle in the direction of the blade installation angle so as to form the blade angular seam casing treatment structure. The design for the transonic axial fan blade angular seam casing treatment structure has the advantages that the air-entraining direction and the air-jetting direction of the blade angular seam casing treatment structure are matched with the air flow direction in channels of the blade tips, and therefore the air-entraining amount and the air-jetting amount can be increased, and the mixing loss of the air flow can be reduced; and the stability extension capability of the blade angular seam casing treatment can be enhanced, and the blade angular seam casing treatment structure further improves the comprehensive stall margin and meanwhile reduces the efficiency loss.

The invention discloses a double-wall cooling device and application, and belongs to the field of aero-engines. Comprising a cold side flat plate, a hot side corrugated plate and hollow turbulent flow columns, the cold side flat plate is located on one side of a cold air channel, the hot side corrugated plate is located on one side of a high-temperature fuel gas channel, and an inverted hill-shaped middle channel with the sectional area changing periodically is formed between two layers of wall plates; a plurality of hollow turbulent flow columns are arranged between the cold side flat plate and the hot side corrugated plate, the cold air channel is communicated with the high-temperature fuel gas channel, and the connecting positions of the hollow turbulent flow columns and the hot side corrugated plate are located on the wave crests of the hot side corrugated plate. Hollow column turbulent flow cooling and external air film cooling are combined, the purpose of high cooling efficiency is achieved by using less cold air, meanwhile, the internal thermal stress of the structure is reduced, and the flow resistance of the structure is reduced. And meanwhile, the defect that a traditional corrugated plate cold air channel is large in flow resistance is overcome, and the problem that outflow cannot be achieved due to insufficient cold air total pressure of a downstream heat screen and a spray pipe heat screen is effectively solved.

The invention discloses a turbine cascade blade top structure with an array type DBD plasma excitation layout and belongs to the technical field of turbine active control. The problem that due to theeffects of the two side pressure differential of the pressure face and the suction face of a blade, energy loss is caused by leakage flowing formed at the blade top can be solved. The turbine cascadeblade top structure comprises a blade top and a plurality of plasma exciters. The multiple plasma exciters are arranged on a plurality of contour lines on the upper surface of the blade top in an array mode. The contour lines are lines formed from the front edge point of the blade top to the tail edge point. The length direction of a bare electrode of the plasma exciter on each contour line is parallel to the corresponding contour line, and the length of the bare electrode on each contour line is smaller than or equal to the length of the corresponding contour line. The turbine cascade blade top structure is an active flowing control blade top structure.