35results about How to "Improve surge margin" patented technology

Systems and methods for turbocharging an internal combustion engine include operating two turbochargers in a series configuration for a first operating region and a parallel configuration for a second operating region. Systems and methods for controlling exhaust gas recirculation (EGR) in a turbocharged internal combustion engine provide low pressure EGR upstream of a compressor inlet for a first operating region and high pressure EGR downstream of a compressor outlet for a second operating range to further improve turbocharger operating margin and overall efficiency.

An object is to position a plurality of guide vanes for generating swirl flow at a housing inner circumferential side in front of an impeller wheel to improve a surge margin and to restrict the decrease in a choke flow rate, thereby increasing an operation range of a compressor. A compressor 19 includes a compressor housing 15, an impeller wheel 7 which compresses intake gas flowing in from an intake-air inlet 23, a swirl-flow generating part including a plurality of guide vanes 55 disposed circumferentially along an inner circumferential wall of the intake-air channel 21 between the intake-air inlet 23 and the impeller wheel 7 and which swirls the intake gas from the intake-air inlet 23 around the rotational axis, and a central intake-air flow path 59 formed inside the guide vanes 55 to allow intake gas to flow to the impeller wheel 7 without passing through the guide vanes 55.

A casing for a turbomachine rotor wheel includes a plurality of circumferential grooves, each of substantially constant section, with the section areas of the circumferential grooves decreasing from upstream to downstream on going from the first groove to the last groove. By treating the casing in this way, the efficiency of the rotor wheel is optimized and its surge margin is improved.

The invention relates to a two-dimensional blade profile optimization method for a compressor considering a low Reynolds number working performance; the optimization method is characterized by comprising the following steps of: step 1: analyzing the meridional flow performance of a prototype multistage axial flow compressor at different rotating speeds; step 2: determining a key two-dimensional blade profile section to be optimized; step 3: performing parametric fitting to the key two-dimensional blade profile in step 2; step 4: analyzing the aerodynamic performance of the two-dimensional blade profile, and analyzing the aerodynamic performance of the parameterized two-dimensional blade profile to obtain the variable attack angle performance of the blade profile under a low Reynolds numberworking condition and the variable attack angle performance under a high Reynolds number working condition; and step 5: optimizing the modeling parameters of the two-dimensional blade profile, adjusting the two-dimensional modeling parameters according to the variable attack angle performance parameters of the two-dimensional blade profile obtained in step 4 at different Reynolds numbers, using agenetic algorithm to circularly iterate steps 3 and 4 until a two-dimensional blade profile design reducing loss and increasing stall margin is obtained, and then obtaining a new two-dimensional blade profile. According to the invention, the performance optimization of the two-dimensional blade profile of the axial flow compressor at different Reynolds numbers and different attack angles is realized; the designed working performance of the compressor is improved while the aerodynamic performance of the compressor at the low rotation speed working condition is improved; and the flow efficiencyis improved while the normal starting and stable working of the compressor are ensured.

The invention discloses an axial flow-centrifugation integral bladed disc type combined compression system. An axial flow impeller and a centrifugal impeller are integrally connected to form an integral bladed disc; an axial flow stator is arranged between an axial flow rotor and the centrifugal impeller; an impeller outer ring of an axial flow gas compressor and an impeller outer ring of a centrifugal gas compressor are integrally connected to form an integral impeller outer ring; one assembling reserving cavity for assembling stator outer rings is formed in the integral impeller outer ring; the axial flow stator is divided into a plurality of equal parts in the circumferential direction; the stator outer rings equally divided are mounted in the assembling reserving cavity; the stator outer rings are limited in the radial direction through a cylindrical surface at the top of the assembling reserving cavity; and the stator outer rings are compressed through compression rings. The axial flow-centrifugation integral bladed disc type combined compression system simplifies the structure to improve the reliability of an engine and reduce the production cost; a cantilever stator is adopted to increase the surge margin of an axial flow gas compressor stage so as to further increase the surge margin of the combined compression system and expand the operational covered wire of the engine.

The invention aims to provide a control method of a titanium alloy compressor tip clearance. The control method is characterized in that an abradable sealing coating is sprayed on the inner wall of acasing; tips of rotor blades are compositely electroplated with a Ni/c-BN coating; and the abradable sealing coating and the Ni/c-BN coating are used as a pair of abradable sealing friction pair, thereby forming a novel gas circuit sealing coating structure. In the operation process of an engine, the tips of the rotor blades can cut into the abradable sealing coating on the inner wall of the casing, thus grooves capable of improving the sealing effect are reserved, ideal radial airflow gaps are formed, and the maximum pressure difference is obtained, so that the power of the engine is remarkably improved, the consumption of aviation kerosene is reduced, and the one-time test run percent of pass of the whole engine is improved.

The invention provides a casing treatment structure for improving the surge margin of a single-stage centrifugal compressor, and relates to the technical field of compressors. A surge expanding through hole is arranged on a casing ingeniously, when the air inlet amount needed by the compressor is increased, part of air enters the casing from the surge expanding through hole, so that the amount of air entering the casing from an air inlet way is reduced, and the surge margin is accordingly increased. Furthermore, the direction of the surge expanding hole is set to form a certain angle with a tangent line of the inner diameter of the air guide casing, so that airflow entering the casing from the surge expanding hole is basically the same in circumferential speed direction of airflow. The casing treatment structure is simple in process, low in product cost, safe, reliable, easy to maintain and simple in structure, and the compressor surge margin can be obviously improved.

The invention discloses a middle reheating type turbine cooler system applied to an airplane pod. By adopting middle reheating circulation, the system efficiency is obviously improved. The system comprises an air inlet system, a first-stage turbine, a second-stage turbine, a first-stage heat exchanger, a second-stage heat exchanger, an air compressor and an exhaust system. By means of the system,ram-air in the flying process of an airplane is used as a power source, and after two-stage turbine expansion cooling, a primary surface heat exchanger cools air in an electronic cabin. By means of the system, middle heat re-exchange circulation is adopted, and the unit flow energy heat exchange capability of a turbine cooler is obviously improved; a liquid circulation heat exchange system is notneeded, the primary surface heat exchanger is adopted to lower the weight of the heat exchangers, and the system weight is obviously lowered; the system has a very low additional power consumption amount, and energy consumption of the cooling system is obviously lowered; an air bearing system is adopted by the system, an oil lubricating system is not needed, the system self-weight is lowered, andmeanwhile, the system maintenance is improved.

A centrifugal compressor includes an impeller including a hub and a plurality of blades disposed on an outer peripheral surface of the hub at intervals in a circumferential direction, and a casing accommodating the impeller and defining an inlet passage for introducing air along an axial direction of the impeller. An inner peripheral surface of the inlet passage includes a narrowed portion, and an inclined portion connected to a downstream side of the narrowed portion and having a radial distance from a rotational axis of the impeller gradually increasing toward the vicinity of tips of leading edges of the blades. Further, 0.01R1≤σ≤0.1R1 is satisfied, where R1 is radial distance between a downstream end of the inclined portion and the rotational axis, R2 is radial distance between a downstream end of the narrowed portion and the rotational axis, and σ is constriction amount which is a difference between R1 and R2.

The invention provides a marine propulsion diesel engine supercharged air temperature sectional type control system, belongs to the technical field of diesel engines, and aims to overcome the defectsin the prior art, break through the matching concept of a traditional supercharger and a diesel engine and provide a new control thought for the defects of two existing control modes in the operationof a fixed-pitch propeller propulsion diesel engine. Through a sectional control method of the supercharger supercharged air temperature under different rotating speed sections, the surge margin of the supercharger under partial load of the diesel engine is increased under the condition that low-load combustion is not affected, good performance matching performance of the supercharger and the diesel engine is achieved, the comprehensive performance of the diesel engine is improved, precious experience and practice bases are provided for subsequent matching of the supercharger and the diesel engine, and the method is effective, convenient to implement and low in application cost.

An electric machine (101) for an aircraft is shown. The electric machine has a casing (104) containing electrical and electromechanical components, a shaft (106) which extends outside of the casing and which is connected to the propulsive fan, a seal (107) to seal the casing around the shaft, and a pressurization system (102) which pressurises the casing above an external pressure to prevent electrical breakdown within gas in the casing of the electric machine.

The invention discloses a centrifugal compressor. The centrifugal compressor comprises a centrifugal impeller, an air bleeding receiver, a radial diffuser, a diffuser inner receiver and a diffuser outer receiver, wherein the centrifugal impeller is used for converting mechanical energy into gas kinetic energy; the air bleeding receiver is arranged around the centrifugal impeller and forms a first airflow passage together with the centrifugal impeller; the radial diffuser is positioned at an outlet of the first airflow passage and is used for converting the gas kinetic energy converted by the centrifugal impeller into pressure energy; the diffuser inner receiver is connected with the radial diffuser and is used for positioning the radial diffuser; the diffuser outer receiver is connected with the air bleeding receiver and is used for covering the radial diffuser; the radial diffuser is in a wedge shape and comprises a leading edge, the leading edge is positioned on one end which is close to the centrifugal impeller, the leading edge comprises a tip and a root which integrally extends together with the tip, and the tip is provided with a sweepback surface which sweeps back relatively to the root. According to the centrifugal compressor provided by the invention, the shape of the leading edge of the diffuser is changed, so as to adapt to the outlet airflow of the centrifugal impeller, thereby effectively reducing the diffuser loss, extending the working range and increasing the surge margin.

The invention relates to a turbocharger pressure shell assembly and a gas compressor. The device comprises a pressing shell, the pressing shell comprises an air inlet channel, and the air inlet channel comprises an air inlet; the silencing cover is mounted at the air inlet of the pressing shell; the connecting part is arranged in the air inlet channel; the casing groove flow channel is formed between the connecting part and the inner wall of the pressing shell and between the connecting part and the silencing cover, the casing groove flow channel comprises a casing groove inlet and a casing groove outlet, the casing groove inlet is located between one end of the connecting part and the inner wall of the pressing shell, and the casing groove outlet is located between the other end of the connecting part and the inner wall of the pressing shell. And the outlet of the casing groove is positioned between the other end of the connecting part and the silencing cover. The noise reduction cover is arranged in the pressure shell, the noise reduction cover is matched with the casing groove flow channel to play a flow guide role, the velocity component of backflow gas in the circumferential direction when the backflow gas is converged into main gas flow is reduced, the velocity component in the axial direction is increased, and the efficiency of a supercharger compressor close to a surge area can be improved.