121results about How to "Reduce aerodynamic losses" patented technology

The tangential standing vortex burning chamber in annular flame tube structure consists of inner and outer boxes and inner and outer flame tubes separately inside the inner and the outer boxes. The outer flame tube has one circumferential head including oil nozzle and flame stabilizing tube, and one main burning hole in the downstream of the head, and the central line of the head and the main burning hole is in some angle to the circumference tangent. The inner and outer flame tubes have inner and outer mixing holes separately in front of their combustors to ensure the excellent outlet temperature distribution, and inclined cooling holes in the walls and inclined to the main burning holes. The present invention has raised liquid atomizing combustion efficiency, reduced combustor outlet temperature distribution coefficient and expanded stable operation range of combustor.

A bucket for use on a steam turbine rotor wheel, the bucket comprising a shank portion and an airfoil portion, the airfoil portion having a radially outer tip with a tip cover adapted to be engaged, in use, by a similar tip cover on an adjacent bucket, wherein a radial step is formed in the tip cover and the airfoil portion along a leading edge of the airfoil portion.

The invention provides a structure suitable for a thrust augmentation type turbofan engine and capable of realizing backward radar and infrared comprehensive stealth. The structure comprises a shielding support plate, an inner cone and a trailing edge blowing flame stabilizer, and is simple in structure, high in reliability, small in weight gain and small in aerodynamic loss, and the radar and infrared stealth performance can be remarkably improved only by virtue of the self structure.

The invention discloses a discrete air film cooling hole shape and relates to the technology of gas turbines. In the discrete air film cooling hole shape, a cylindrical hole outlet sinks for a depth from a cooled surface and is provided with an expanding structure starting at the center of the cylindrical hole outlet, the expanding structure is in bilateral symmetry with a cylindrical hole central line as the symmetrical axis, the sink depth H is 0.2-0.5 times of the diameter of a cylindrical hole, the expanding width W is 3-4 times of the diameter of the cylindrical hole, and the expanding length B is 1.0-2.0 times of the diameter of the cylindrical hole. The hole shape comprises: 1) a fundamental form; and 2) an expanded boss arranged in the space between the lower stream of an air outlet edge of the cylindrical hole and an air outlet edge of the expanding structure on the basis of the fundamental form. The discrete air film cooling hole shape has the advantages of being high in crosswise average air film cooling efficiency, good in cooling effect in high blowing ratio, less in pneumatic loss and easy to achieve. The discrete air film cooling hole shape is used in air film cooling of the gas turbines and is suitable for pressure surface, suction surface and end wall discrete hole air film cooling of turbine blades.

A compressor cascade experimental device includes an even number of blades and an odd number of suction blades. An even number of blades and an odd number of suction blades are installed on the mounting surfaces of the first grid plate and the second grid plate, and the suction blades are located at the center of the first grid plate and the second grid plate. There is an end wall suction slot between each mounting slot. The installation groove of the suction vane and the suction groove of the end wall are provided with through gas suction slots. The first suction hood and the second suction hood are respectively installed at the centers of the outer surfaces of the first grid plate and the second grid plate, and the central air hole of the suction cover corresponds to the gas suction slit and the end wall suction groove respectively. In the present invention, through the combined suction of the blade surface of the compressor cascade and the end wall of the cascade, and connecting the second suction cover and the first suction cover to the suction equipment, the surface of the blade and the surface of the blade under different gas suction quantities can be simultaneously performed. The combined suction experiment of the end wall of the cascade obtains the optimal suction volume of the compressor cascade and improves the pressure ratio and efficiency of the compressor cascade.

The invention discloses a coaxial opposite-rotating dual-rotating-wing duct type vertical take-off and landing aircraft. The coaxial opposite-rotating dual-rotating-wing duct type vertical take-off and landing aircraft is composed of a loading bin, a duct receiver, a flow guiding rudder, an opposite-rotating driving motor and two opposite-rotating rotors. The loading bin is located above the duct receiver which is of an annular cavity structure, opposite-rotating rotor blades are wrapped so that extra front edge suction force can be generated, and the power requirement of the aircraft is lowered while the lifting force of the aircraft is increased. An installing frame is fixed to the lower end of the inner wall of the duct receiver. The flow guiding rudder is in butt joint with the installing frame. The opposite-rotating driving motor is installed on the upper portion of the installing frame. The two opposite-rotating rotors are fixedly connected with an output shaft of the opposite-rotating driving motor, and an undercarriage is embedded in the lower portion of the outer side wall of the duct receiver. Coaxial opposite-rotating dual rotating wings counteract self-rotating moment, and stable control over the posture of the aircraft is facilitated. Pneumatic loads of each stage of rotor blades are reduced through the opposite-rotating rotors, pneumatic loss can be reduced easily, meanwhile the opposite-rotating rotor blades reduce rotational flow loss, pneumatic efficiency is greatly improved, and time of endurance is prolonged.

The invention discloses a turbocharger composite nozzle device, comprising a turbine wheel and a nozzle ring supporting disc mounted coaxially therewith. The nozzle ring supporting disc is provided with a plurality of groups of composite nozzle blades each comprising a movable blade and a fixed blade, the fixed blade is fixedly connected with the nozzle ring supporting disc and the movable blade is in transmission connection with a power transmission device; in case of other working conditions of an engine, a transmission mechanism adjusts the position of the movable blade continuously according to the change of the working conditions of the engine, and the control for waste gas energy is realized by changing the area of volute nozzle outlet, in order to improve aerodynamic performance ofthe turbocharger, especially aerodynamic performance of the engine under the working condition of small flow, thus better matching with the engine is reached.

The invention discloses a turbofan engine and belongs to the technical field of design of an aero turbofan engine. The engine comprises a second afterburner (2) and a first afterburner (1), wherein the second afterburner (2) comprises an afterburner inner spraying pipe and a main force exerting cavity; the afterburner inner spraying pipe is fixedly connected with the back end of an inner cylinder;the main force exerting cavity is positioned in the afterburner inner spraying pipe; a through hole is formed in the afterburner inner spraying pipe; the first afterburner (1) comprises an afterburner outer spraying pipe and an auxiliary force exerting cavity; the afterburner outer spraying pipe is fixedly connected with the back end of an outer cylinder; the auxiliary force exerting cavity is positioned between the afterburner outer spraying pipe and the afterburner inner spraying pipe; the first afterburner (1) also comprises a baffle; the baffle is controlled by an adjusting mechanism to close or open the through hole in the afterburner inner spraying pipe; and the opening section of the tail end of the afterburner outer spraying pipe is adjustable. The length and the weight of the afterburner are reduced by adjusting the afterburning condition of the main force exerting cavity and the auxiliary force exerting cavity.

An article, such as a gas turbine engine turbine blade, includes a film cooling system having a trench recessed within a wall of the component, a passage that opens into the trench for discharging a coolant flow into the trench, and a guide member within the trench for influencing the coolant flow discharge from the passage.

The invention discloses a turbine blade body structure. A two-dimension blade profile of a blade body is a crescent closed curve defined by an inner arc curve and a back arc curve; the blade body has a characteristic parameter chord length CH, and an axial width CX; the blade body is formed by twisting and superimposing multiple two-dimension blade profiles according to a certain rule, and has a characteristic parameter blade body height H; the chord length CH of the two-dimension blade profile accords with a parabolic distribution rule from the root area to the top area in the blade height direction; a trend is that the chord lengths of the root area and the top area of the blade body are high, and the chord length of the middle area of the blade body is low; and the axial width CX of the two-dimension blade profile substantially remains unchanged. The turbine blade body structure can allow the turbine blade to simultaneously give consideration to different main loss sources of the root area, the top area and the middle area, thereby effectively and reliably improving the aerodynamic performance of the turbine blade and reducing aerodynamic loss.

The invention provides a turbine blade which has a pressure face, a suction face, a blade top surface and at least one wing rib. The pressure face and the suction face are opposite and extend between a front edge and a tail edge of the blade, the blade top surface connecting the pressure face with the suction face is arranged at the pointed end of the blade, and the wing ribs extend from the pointed end of the blade and include at least one radial wing rib radially extending outwards from the blade top surface and / or at least one peripheral wing rib peripherally extending out from the pressure face and / or the suction face. The turbine blade with the wing ribs at the pointed end can beneficially combine respective advantages of the radial wing ribs and the peripheral wing ribs to realize effect of reducing leakage loss by different degrees.

The invention provides a novel binary plug type spray pipe. The novel binary plug type spray pipe comprises a supporting component, a cone body and a mixer (5). The supporting component is a cylindrical component which is formed by sequentially connecting a cylinder (1), a circular-to-square transition section (2) and a spray pipe contraction section (3) end to end, an inlet of the supporting component is round, and an outlet of the supporting component is rectangular. The mixer (5) is an annular component which is connected to the axis of an inner cavity of the supporting component in a sleeving mode. The cone body is a double-layer thin-wall component which is connected to the axis of an inner ring of the mixer (5) in a sleeving mode and is composed of a wedge-shaped plug cone and a central cone which is in butt joint with the wedge-shaped plug cone and has a circular section, and a supporting structure is arranged between the two thin walls of the central cone. According to the novel binary plug type spray pipe provided by the invention, the central cone and the plug cone are integrally designed, transition of the whole pneumatic flow path is smooth, and pneumatic loss is small; high-temperature components in an engine can be completely shielded, and the novel binary plug type spray pipe has a good stealth capability; and meanwhile, the novel binary plug type spray pipe has the advantages of being simple in structure, short, and low in weight.

A high pressure turbine (HPT) blade includes a substantially arcuate trailing edge including one or more fluid injection elements disposed therein. Each fluid injection element injects a fluid such as air into a desired trailing edge region of the HPT blade or vane to enhance mixing out of the wakes generated via the HPT blade or vane. The enhanced mixing out reduces HPT / LPT interaction losses and / or the axial gap between the HPT and LPT components in a gas turbine engine. The interaction losses include unsteady thermal wake segregation effects that lead to unexpected heat-up of endwalls (planforms and blade / vane-tips) of downstream blades / vanes, and further include aerodynamic losses, both transonic and subsonic.

A high pressure turbine (HPT) blade includes a substantially arcuate trailing edge including one or more fluid injection elements disposed therein. Each fluid injection element injects a fluid such as air into a desired trailing edge region of the HPT blade or vane to enhance mixing out of the wakes generated via the HPT blade or vane. The enhanced mixing out reduces HPT / LPT interaction losses and / or the axial gap between the HPT and LPT components in a gas turbine engine. The interaction losses include unsteady thermal wake segregation effects that lead to unexpected heat-up of endwalls (planforms and blade / vane-tips) of downstream blades / vanes, and further include aerodynamic losses, both transonic and subsonic.

The invention provides a very-high-pressure intercooling cycle turbofan engine. The engine comprises fans, a high-pressure air compressor, an intercooler, a very-high-pressure air compressor, a combustion chamber, a circular rotating fan transition section, a very-high-pressure turbine, a high-pressure turbine, a low-pressure turbine, a low-pressure shaft, a high-pressure shaft and a very-high-pressure shaft. The fans, the high-pressure air compressor, the intercooler, the very-high-pressure air compressor, the combustion chamber, the circular rotating fan transition section, the very-high-pressure turbine, the high-pressure turbine and the low-pressure turbine are connected in sequence. The very-high-pressure turbine drives the very-high-pressure air compressor to operate through the very-high-pressure shaft. The high-pressure turbine drives the high-pressure air compressor to operate through the high-pressure shaft. The low-pressure turbine drives the fans to operate through the low-pressure shaft. The rotation central axis of the very-high-pressure air compressor, and the axis of the combustion chamber are both parallel to the axis of the engine but not coincide with the axis ofthe engine. The high-pressure air compressor is composed of a multi-stage high-load axial flow air compressor and a single-stage high-through-flow centrifugal / oblique flow air compressor.

The invention discloses a novel slotted air film cooling structure for guide blades of turbines of gas turbine engines. The novel slotted air film cooling structure comprises a blade base, air film slots, connecting ribs and stepped planes, and is characterized in that the air film slots are oblique slots which are forwardly oblique along a mainstream direction, are communicated with a tail cooling air cavity and are equidistantly arrayed along the blade height direction, outlets of the air film slots are separated from a blade tail edge, and the connecting ribs and the air film slots are alternately distributed. The novel slotted air film cooling structure has the advantages that uniform and consistent cooling air films can be transversely formed by the novel blade structure, and outflow air films are high in stability; an eddy flow effect of outflow cooling air can be weakened, the cooling air films are large in coverage area, and accordingly the novel slotted air film cooling structure is high in cooling efficiency and simple in machining.

The invention discloses a novel engine oil supply support plate receiver structure and an engine comprising the structure. An oil supply support plate receiver is fixedly welded with a low-pressure turbine guider; multiple support plate blades in the oil supply support plate receiver are nonuniformly arranged in the peripheral direction to prevent resonance of an engine turbine part under high-speed operation condition; lubricating oil pipelines are arranged in at least one support plate blade, and are positioned on the lower side in the oil supply support plate receiver; each support plate blade is a three-dimensional torsional structure on space; at least two of multiple section blade shapes in different radial positions of the same support plate blade are different, so that each support plate blade can adapt to upstream incoming airflow angle and provide a proper flowing field to a downstream turbine, and pneumatic loss is obviously reduced; and the maximum thickness positions of different section blade shapes of each support plate blade are arranged in the same axial section, and are positioned on the same radial angle line, so that the support plate blades with the three-dimensional torsional structures can be smoothly arranged on the lubricating oil pipelines.

The invention discloses an oblique flow and centrifugal combined compressor. The oblique flow and centrifugal combined compressor comprises an oblique flow compressor, a centrifugal compressor and a backflow device. The backflow device is in a gooseneck shape and is connected between the oblique flow compressor and the centrifugal compressor. The backflow device comprises a backflow device curve section channel, a backflow device blade and a backflow device slewing section channel which are connected in sequence. An inlet of the backflow device curve section channel is connected with an outlet of an oblique flow diffuser. The backflow device curve section channel is in an arc shape, so that the backflow device blade is close to the front side of the oblique flow diffuser, and an outlet of the backflow device slewing section channel is connected with an inlet of a centrifugal impeller. On the basis that the axial length of the oblique flow and centrifugal combined compressor is shortened, the whole flow channel structure is simplified, large-angle turns of airflow in the backflow device channels are decreased, pneumatic loss is small, and then the overall efficiency of the combined compressor is improved. Moreover, the axial length of the combined compressor is further shortened, and the oblique flow and centrifugal combined compressor is more compact in structure and has more excellent rotor dynamic characteristics.

The invention relates to the technical field of turbine engines, in particular to a middle receiver of a turbine engine. The middle receiver comprises a receiver outer ring and a receiver inner ring,wherein the receiver outer ring sleeves the receiver inner ring; the axial central line of the receiver outer ring, the axial center line of the receiver inner ring and the axis of an engine shaft aresuperposed; a gap is formed between the receiver outer ring and the receiver inner ring; a plurality of guide blades are arranged between the receiver outer ring and the receiver inner ring; and theguide blades are uniformly distributed in the peripheral direction around the axis of the engine shaft for guiding the direction of gas flow. According to the middle receiver of the turbine engine provided by the invention, the middle receiver integrates the action of a turbine guider, and the turbine guider does not need to additionally arrange for guiding gas flow, so that engine components canbe reduced, and therefore, the engine part quantity and engine weight are reduced. The engine is more compact in layout, and integral length of the engine is shortened.

The invention discloses an adjusting mechanism for inlet guide blades of a centrifugal blower. The adjusting mechanism comprises an inner volute, an outer volute, a shield cap, a front-section guide blade assembly and a rear-section guide blade assembly. During working, an actuator drives one of rear-section guide blade rotating shafts to rotate through a coupling, the rear-section guide blade rotating shaft rotates to drive a driving ring to rotate, accordingly, all the guide blade rotating shafts are driven to rotate, and the purpose of adjusting the rotating angle of the rear-section guideblades is achieved. According to the adjusting mechanism for the inlet guide blades of the centrifugal blower, a traditional single-blade flow guiding mode is broken through, by adopting the mode thatthe front-section guide blades are fixed and the rear-section guide blades rotate, adjusting of the tandem-cascade-type guide blades is achieved, and flowing separation of an air flow at the guide blades is weakened; and under the situation of ensuring adjustability of the flow direction and the flow amount of the air flow, the aerodynamic loss brought by the guide blades is reduced, the efficiency of the whole mechanism is improved, and energy consumption is reduced.

The invention discloses a rectifying and noise lowering exhaust hood. The rectifying and noise lowering exhaust hood mainly comprises a non-return valve, an air duct, a shell assembly, an oil net anda smoke collecting hood; the oil net is located on the center portion of the smoke collecting hood; the shell assembly is connected above the smoke collecting hood and the oil net; the air duct is located in the shell assembly; the non-return valve is arranged above the air duct; an impeller and a motor are arranged in the air duct; air inflow rings are arranged on the two sides of the impeller; and the positions of air inlets in the air duct are provided with a rectifying cover. According to the rectifying and noise lowering exhaust hood, due to the fact that the rectifying cover is arrangedat the positions of the air inlets of the air duct, local vortexes are reduced, air flow is made smooth, and the overall performance of a smoke machine is improved.

The turbine blade (10) has a leaf blade (16) through which a hot gas is flowable. A throttle element is equipped with two projections at respective openings (28) with respect to a flow direction of a channel. Each projection is attached to one of two surfaces (34,36) arranged in an inner-facing manner.

An article, such as a gas turbine engine turbine blade, includes a film cooling system having a trench recessed within a wall of the component, a passage that opens into the trench for discharging a coolant flow into the trench, and a guide member within the trench for influencing the coolant flow discharge from the passage.

The invention relates to a variable nozzle turbocharger regulating mechanism integrated on a turbine shell, belonging to the technical field of variable nozzle turbochargers used for vehicles. The variable nozzle turbocharger regulating mechanism comprises nozzle ring fixed blades, nozzle ring adjustable blades, a mounting disc, driven gears, a driven gear ring, a driving gear, a double-rocker driving mechanism, a split-type turbine shell and a turbine shell cover board. The driven gears and the nozzle ring adjustable blades are respectively installed on the circumferences of the front and the back side faces of the mounting disc through mounting shafts of the driven gears and the nozzle ring adjustable blades. The driven gears are connected at the inner edge of the driven gear ring which is used for controlling the nozzle ring adjustable blades to revolve synchronously, and through the driven gear ring, the nozzle ring adjustable blades on the circumference can revolve synchronously. For the purpose of adapting to air motion law, the initial installation angles of the nozzle ring adjustable blades are arranged to be unequal along the circumferential direction. The width of the nozzle ring is controlled by the nozzle ring fixed blades, and the nozzle ring adjustable blades are driven by the double-rocker driving mechanism, so the accuracy and the uniformity of the adjustment of opening sizes of the nozzle ring adjustable blades are ensured.

The invention provides a variable geometry turbine with vortex elimination hole structures in the front of movable blades. The variable geometry turbine comprises a case and hubs, wherein adjustable static blades and the movable blades are uniformly installed between the case and the hubs in the circumferential direction, upper rotating shafts and lower rotating shafts are installed on the upper end faces and the lower end faces of the adjustable static blades respectively, the upper rotating shafts are embedded into the case, the lower rotating shafts are embedded into the correspondinghub, the movable blades are installed on the correspondinghub, the curved vortex elimination hole structures communicating with blade pressure surfaces and suction surfaces are formed in the front of the movable blades in the radial direction, and load distribution of blade profiles of the movable blades is of a rear-loading type. According to the variable geometry turbine, separation vortexes in the front of the blades can be reduced or even eliminated underthe off-design condition so that aerodynamic loss in channels of the movable blades can be obviously reduced, transverse differential pressure in front of the blades can be reduced underthe design condition, formation of secondary flow can be delayed, and then loss of the secondary flow can be reduced. The variable geometry turbine hasgood characteristic underall operating conditions on the whole.

The invention relates to the field of structural design of aero-engine jet pipes, in particular to a jet pipe capable of reducing features of infrared signals and radar signals, and aims at solving the problem that the pneumatic loss of an existing binary hidden jet pipe is large. The jet pipe comprises three barrel segments, namely, a circular-semicircular section segment, a semicircular-crescent-semicircular section segment and a semicircular section outlet segment which are connected in sequence. A conical protrusion is arranged on the inner wall of the bottom of the semicircular-crescent-semicircular section segment, and used for covering a high-temperature component in an engine, and the outer surface, exposed to the interior of the semicircular-crescent-semicircular section segment, of the conical protrusion is arranged to be streamline. According to the jet pipe capable of reducing the features of the infrared signals and the radar signals, the pneumatic loss of the jet pipe can be reduced, meanwhile the conical protrusion can completely cover the high-temperature component in the engine, the features of the infrared signals and the radar signals of an engine exhaust system can be greatly reduced, and an aircraft can be helped to meet the control requirement for the low features of the infrared signals and the radar signals.

An object of the invention to provide a rotor blade and an axial-flow compressor. The rotor blade is suitable for being mounted on a rotor drum of a rotor of the axial-flow compressor, and includes ablade body, a tenon and a margin plate; for two rotor blades that can be circumferentially adjacent to the rotor drum, a first groove structure of a suction surface portion of the margin plate of oneof the rotor blades and a second groove structure of a pressure surface portion of the margin plate of the other rotor blade are combined and spliced to form a flow channel wall surface between the two rotor blades disposed adjacently in the circumferential direction, and the flow channel wall surface is provided with a plurality of grooves uniformly distributed in a wave shape in the circumferential direction; the grooved extend in a direction from a leading edge to a trailing edge, and the depth of the grooves gradually increases along an extending direction; and end points at which the grooves extend are located on the side wall surface of the margin plate. Presence of the grooves makes the thickness of a boundary layer reduce, thereby reducing a blending and mixing area of interstage leakage and the boundary layer, and so, pneumatic loss caused by the interstage leakage is reduced, and the compressor efficiency and surge margin are improved.

The invention discloses a turbine movable blade pressure surface and a compound angle film hole layout structure at the top, is suitable for an aircraft engine or a ground gas turbine, and relates to the field of aircraft engine turbine blade cooling. Aiming at deflection of film outlet currents at the top of the blade pressure surface in a rotating state from blade root to blade tip to different extents, the detailed subarea compound angle design is performed; the turbine movable blade pressure surface is divided into five different blade areas of a blade root area, a blade root-blade middle transition area, a blade middle area, a blade middle-blade tip transition area and a blade tip area; and in consideration of the blade top, each blade area is provided with different compound angle film holes, so that the technical problem of no facilitation of wall attachment of films due to deflection of the film outlet currents of the pressure surface to the blade tip under the influence of the Coriolis Force pointed to the blade tip, the centrifugal floating force pointed to the blade tip and the secondary channel current and caused by leakage of the currents from top gaps is solved, the film covering is more uniform, and the cooling efficiency is improved.

The invention discloses a gas turbine movable vane guard ring structure with grate teeth for tight sealing. The gas turbine movable vane guard ring structure comprises a guard ring base body, hook structures arranged at the two ends of the guard ring base body and a guard ring molded surface arranged on the bottom surface of the guard ring base body, and the guard ring molded surface is provided with a grate tooth structure; and the grate tooth structure comprises at least one grate tooth arranged on the guard ring molded surface, the guard ring molded surface is provided with grooves matchedwith the grate teeth, one end of each grate tooth is connected into the corresponding groove in a matched manner, and the other end of each grate tooth and the surface of the guard ring molded surfaceare welded. Cooling air flows out from a guard ring and the upstream adjacent fixed vane end wall, and the cooling air covering and protecting effect is formed on the guard ring molded surface. By means of the guard ring structure, under the condition that an existing movable vane top structure is not changed, the guard ring molded surface is provided with the grate tooth structure, and thereforewhen main flow fuel gas flows through a movable vane top gap, the leakage flow of the main flow fuel gas at the movable vane top gap is reduced, the pneumatic loss of the vane top gap is reduced, andthe performance of a gas turbine is improved.