45 results about "Inlet distortion" patented technology

The invention discloses an aerial turbofan engine control method based on a surge margin estimation model. The modeling of the provided surge margin estimation model is divided into two parts, namely an undistorted model during conventional flight and a loss model during super-maneuver flight, wherein the undistorted model is realized by screening the optimal model input and modeling through a nonlinear fitting method based on a surge margin feature selection algorithm; the loss model is realized by estimating the aircraft engine inlet distortion degree based on an online attack angle prediction model and calculating to obtain the surge margin loss during distortion. The stability of an engine is predicted in real time by utilizing the estimation model, a turbine fall ratio control instruction is subjected to margin loss compensation on the basis that a conventional control loop of the engine is not changed, and high-stability control is realized. The aircraft engine can be guaranteed to work stably and efficiently, and the surge margin of the aircraft engine is controlled to be 11-13 percent.

The invention belongs to the technical field of aero-turbine engines, particularly relates to an air inlet flow field distortion simulation device used for aero-turbine engine tests, and particularly relates to an adjustable fan-shaped plate flow field distortion simulator. The distortion simulator is applied to an engine air inlet distortion resistance test and an engine surge test. Compared with the exiting distortion simulator, the device has the following advantages that the adjustable distortion simulator adopts the adjustable fan-shaped plate mode according to the use requirements of a turbofan engine, and the angle of the fan-shaped plate is adjustable. The structural form is domestically used for the first time; the state of distortion simulation is changed through mechanical transmission, and the structural form and the sealing performance of the device are better than those of other forms of adjustable simulators; and the test efficiency is high, and multiple engine air inlet distribution state tests can be completed in one test so that the test cost can be effectively reduced.

The invention discloses a flow distortion simulation device which comprises a tank body, a tube body, a motor, a worm, a bevel gear assembly, a worm gear, a lead screw, a guiding rod, a linear bearing, a gear nut, an inserting plate, a displacement detection assembly and a controller. The tube body is arranged on the tank body, the worm is arranged in the tank body, the bevel gear assembly is respectively connected with the motor and the worm, the worm gear is arranged on the worm, the lead screw is meshed with the worm gear, the linear bearing is arranged on the guiding rod, the gear nut is arranged on the guiding rod and is in screw-thread fit with the lead screw, the inserting plate is respectively connected with the linear bearing and the gear nut, and the controller is respectively connected with the displacement detection assembly and the motor. The flow distortion simulation device has the advantages of being capable of meeting requirements of a turbofan engine inlet distortion resistant test and an engine surge test, small in occupied space, continuous and adjustable in distortion index, convenient to install and operate, high in simulation accuracy and degree of automation, capable of improving the test efficiency, reducing the test cost, reducing the workloads of people and the like.

The invention belongs to the field of aerodynamic stability of small turbofan engines, and in particular relates to an adjustable fan-shaped inlet pressure distortion simulation board of an engine. The adjustable fan-shaped inlet pressure distortion simulation board comprises a positioning fan-shaped plate, a plurality of deformation fan-shaped plates and a transmission fan-shaped plate. The circumferential movement of the simulation board is utilized to shield a local area within a flow channel, so that a range of a blockage ratio of the channel reaches 0 to 0.17, and the range can meet the use requirements for air inlet distortion resistance tests and surge tests of most of engines.

A tip turbine engine according to the present invention includes a plurality of independently variable inlet guide vanes for the fan and/or for the compressor. An actuator is operatively coupled to each of the flaps, such that each actuator can selectively vary the flap of its associated inlet guide vane. In one embodiment, the inlet guide vanes each include a pivotably mounted flap that is variable independently of the flaps of at least some of the other inlet guide vanes. In another embodiment, the inlet guide vanes each include at least one fluid outlet or nozzle directing pressurized air, as controlled by the associated actuator, to control inlet distortion.

The invention provides a total pressure distortion generator. The total pressure distortion generator comprises a frame body, vertical streamline hollow clamping plates, blocking partition plates, distortion elements, push rods and an air cylinder assembly, wherein the vertical streamline hollow clamping plates are installed between a front plate and a rear plate of the frame body in parallel; thevertical streamline hollow clamping plates are installed at intervals; each blocking partition plate is installed between every two adjacent vertical streamline hollow clamping plates so as to enablethe shape of the overall airflow inlet and outlet surface formed by each airflow channel between the assembled vertical streamline hollow clamping plates to correspond to the shape of the needed airflow cross section; each distortion element is installed between every two adjacent vertical streamline hollow clamping plates and used for adjusting the airflow to form air inlet distortion; each pushrod is arranged in an inner space of the vertical streamline hollow clamping plate, and each push rod is connected with one corresponding distortion element; and each cylinder in the cylinder assembly is connected with a corresponding push rod. According to the invention, a steady-state and dynamic distortion ratio adjustable total pressure distortion simulation function can be realized.

The invention discloses a depth-adjustable inlet distortion device of an aircraft engine. The distortion device is installed between an inlet channel and a horn mouth of the engine and mainly comprises two quick release rings, a plug plate base and a plug plate. One quick release ring is connected with the horn mouth, the other quick release ring is fixedly connected with the inlet channel through a sealing ring, the circular end of the lower portion of the plug plate base and the quick release rings are fixedly connected into an integral structure, the upper portion of the plug plate base is of a square structure, the width of the plug plate base is the same as the diameter of the inlet channel, a square plug plate groove is formed in the plug plate base, the plug plate groove is provided with a plug plate sealing block and a positioning clamp, the plug plate can be fixed after being inserted in the plug plate groove, the plug plate is fixed at any position in the plug plate groove, then, shielding of the plug plate in the inlet channel at different depths is achieved, and multiple sets of distortion indexes are produced. The plug plate depth can be adjusted, the inlet indexes of the engine are changed, and engine stability examination is carried out under different distortion indexes; and a clamping ring structure is adopted, so that work is reliable, disassembly and assembly are simple, and the test run efficiency can be improved.

The invention discloses a compressor stability boundary judgment method considering the influence of intake total pressure distortion, which comprises the following steps of obtaining the maximum stable state aerodynamic load Hmax of each stage of a compressor and the dimensionless diffusion length of each stage of rotor blade, and obtaining the maximum stable state aerodynamic load Hef (theta) ofeach stage of the compressor; solving the effective stable state pneumatic load of the critical distortion angle of the gas compressor under the gas inlet distortion condition; if the load is smallerthan Hmax, determining that the gas compressor is still in a stable working state, if the load is larger than Hmax, determining that the gas compressor enters an unstable state, wherein the working point of the last stable working state is considered as a stable boundary point under the current rotating speed; and repeating the steps to finally obtain the stable boundary points at different rotating speeds, and connecting the stable boundary points to obtain the stable boundary of the gas compressor. The method can quickly and accurately predict the stable boundary of the gas compressor underthe condition of inlet gas total pressure distortion, can be used in the initial stage of gas compressor design, can reduce the development cycle of the gas compressor, and reduces the development cost and risk.

The invention relates to an airworthiness certification method for influences of different flying angles on stall/surge characteristics of an aero-engine. The method includes the steps that firstly, a change rule curve map of the distortion range AP and the distortion degree PP of total pressure air inlet distortion characteristic parameters at different flying angles along with change of Ma numbers is established according to a numerical calculation method; then, according to the map and a one-dimensional model, a general characteristic curve and a surge margin of the aero-engine can be calculated; otherwise, starting from the stall/surge airworthiness standard of the aero-engine, a method for determining the limit flying angle of the aero-engine is proposed; finally, in combination with relevant files of CAAC rules, airworthiness certification processes for stall/surge of the aero-engine are formulated.

The invention discloses a method for identifying the balanced polar curve of a two-engine airplane through a ground taxiing test. The variable quantity of inlet distortion caused by change of the rotation speeds of engines is ignored; a runway for the test is horizontal, and the angle of attack and the angle of pitch are the same in size; taxiing state points of two different angles of pitch are selected to be subjected to synchronous accelerating/decelerating and asynchronous accelerating/decelerating taxiing tests when the two engines are at the different rotation speeds, and the accelerating/decelerating speed is worked out through accelerating/decelerating motion graph lines; and accordingly, profile drag and lift-to-drag factors are worked out, and the balanced polar curve CD-CL of the airplane is obtained through fitting. According to the method, inaccuracy of data of an engine bed test is overcome, the true balanced polar curve of the airplane is obtained, meanwhile, the aerialflight test cost is reduced, and the test flight risk of maiden flight is reduced; and the balanced polar curve of the airplane is identified on the basis that the data of the engine bed test are notutilized, a foundation is built for reasonably making a test flight plane of maiden flight in the later period, and the good practicality is achieved.

The invention discloses a ramjet circumferential air inlet distortion restraint grating and a design method thereof, relating to ramjets. The ramjet peripheral air inlet distortion restraint grating is provided with three distortion restraint grating pieces which are sequentially sleeved and laminated together in series, wherein the three distortion restraint grating pieces consist of concentric circular rings and radial banners with consistent distribution rules; the position of the first grating piece is immovable relative to the ramjet; the second grating piece and the third grating piece can rotate around a shaft which is concentric to the first grating piece; the second grating piece and the third grating piece are connected with the first grating piece by virtue of a sleeve shaft to serve as a whole so as to be arranged at the inlet of a circular section of the ramjet combustion chamber; and the second grating piece and the third grating piece are respectively controlled to rotate by virtue of an independent servo rotating motor. The design method comprises the following steps: determining the maximum circumferential distortion value DC (60); determining the minimum grating sector angle thetamin needed for restraining the maximum circumferential distortion value DC (60); determining the grating sector angles; determining the number of radial banners of the gratings; and determining the number of concentric circular rings of the gratings.

The invention provides an inlet distortion constraint device and a centrifugal compressor provided with the same, and relates to the technical field of an inlet airflow rectifying device of the centrifugal compressor. The inlet distortion constraint device comprises a flow guide pipe, wherein flow stabilizing plates used for rectifying turbulent airflow are arranged in an inner cavity of the flowguide pipe in the flow guiding direction. According to the inlet distortion constraint device and the centrifugal compressor provided with the same, the problems of efficiency attenuation of a gas compressor and increase of shafting fluctuation due to introduction of inlet airflow into the centrifugal compressor in a compact system in a traditional technology are solved; and the problems of complicated construction, high cost and even increase of operating failure rate of the compressor due to arrangement of complex parts are solved.

The invention discloses a method for predicting dynamic stall process of an aviation axial flow compressor under complex air inlet distortion. The method comprises the steps that a blade wall fixing effect is described through a distributed force source, and a dynamic relation between a local blade element runner force source and an airflow parameter is effectively constructed based on a loss coefficient and a backward angle concept. The method has the capability of quantitatively describing the dynamic stall evolution process of the multi-stage axial flow compressor under complex air inlet distortion. Compared with a traditional two-dimensional calculation method and a high-order three-dimensional CFD calculation method, the method has the advantages that the calculation efficiency can be considered, and meanwhile, the aerodynamic instability characteristics of the aviation axial flow compressor under complex air inlet distortion disturbance including rotational flow and radial/circumferential space combination distortion characteristics are effectively reflected, and a technical support is provided for high-fidelity aerodynamic stability design of an advanced aero-engine.

The invention discloses a method, system for calculating the aerodynamic stability of a turboshaft engine, equipment and a medium. The method comprises the steps: dividing a calculation model into a plurality of control bodies, and then calculating a control equation set for each control body to obtain the aerodynamic parameters of each control body; and combining an instability criterion to predict whether the pneumatic instability of the engine occurs under the given air inlet distortion condition, so that the influence of the air inlet pressure distortion on the pneumatic stability of the turboshaft engine can be accurately evaluated. A calculation model adopted by the method only focuses on parameters of inlet and outlet sections of the gas compressor, the whole combined gas compressor can be modularized, and static pressure, static temperature, mass flow and circumferential speed are adopted as variables to establish the equation set to describe flow in each control body. On one hand, all variables can be obtained through conversion of the equation set formula based on the four basic variables, on the other hand, the four basic variables can be conveniently obtained through a test means, and the data obtaining process is simplified.

The invention provides a stationary blade distortion-resistant axial flow compressor and an axial flow compressor stationary blade distortion-resistant method.The cascade consistency of a stationary blade distortion area of the compressor is greater than that of a non-distortion area, and the cascade consistency of the non-distortion area is the original consistency. In the design, stationary blades of a stationary blade axial flow compressor are non-uniformly distributed, and the cascade consistency is increased in a stationary blade area influenced by air inlet distortion fluid without changing the blade consistency of original stationary blades in other areas, so that the aim of regulating and controlling a non-uniform flow field by using the non-uniformly arranged stationary blades is achieved, the performance (flow, pressure ratio, efficiency and the like) of the air compressor is ensured not to change under the condition of a uniform incoming flow, and the aerodynamic performance and the stability margin under the condition of air inlet distortion are improved at the same time.

The invention provides a radial air inlet distortion restraining grating of a ramjet engine and a designing method of the radial air inlet distortion restraining grating, and relates to a ramjet engine. The radial air inlet distortion restraining grating of the ramjet engine is provided with a first distortion restraining grating plate and a second distortion restraining grating plate which are stacked together coaxially, wherein the first distortion restraining grating plate consists of flexible spiral plates which surround a center drum; the second distortion restraining grating plate consists of radial scrolls of which the number is fixed; the geometric position of the second distortion restraining grating plate is fixed relative to the ramjet engine; and the first distortion restraining grating plate and the second distortion restraining grating plate are arranged on an inlet of a circular combustion chamber of the ramjet engine. The designing method of the radial air inlet distortion restraining grating comprises the following steps of determining a maximum radial distortion value DIS_R according to incoming flow conditions of different aircrafts; determining the minimum area ARmin of a ring for restraining the maximum radial distortion value DIS_R; determining the number of turns of the spiral plates of the first distortion restraining grating plate; and determining the number of the radial scrolls of the second distortion restraining grating plate. The number of the turns of the spiral plates is equal to the quotient of the area of the circle of the inlet of the circular combustion chamber and the ARmin, and the number of the radial scrolls is three times greater than that of the spiral plates of the first distortion restraining grating plate.

The invention belongs to the technical field of engine control, and particularly relates to an active stability augmentation control method and device for an aero-engine. The method comprises the following steps: S1, acquiring aircraft attitude parameters; s2, according to the aircraft attitude parameters, determining that the aircraft flies in a large-overload and high-maneuverability state; s3, according to the height and speed of the aircraft and the attitude parameters of the aircraft, determining an identification margin insufficient working condition, wherein the identification margin insufficient working condition represents the residual stability margin condition of the engine; s4, executing nozzle diameter amplification control, gas compressor adjustable stator blade angle decrease control, gas release valve opening control and stress application oil supply amount increase control under the corresponding working conditions according to the working conditions with the insufficient mark margin. According to the method, by identifying the high air inlet distortion state, the stability augmentation control measures are implemented in a targeted mode, the thrust of the engine cannot be greatly reduced, and meanwhile it is guaranteed that the thrust in other states in the working envelope cannot be reduced.

The invention belongs to the pressure, temperature and speed testing technical field, and discloses a combined dynamic probe measuring the turbine machine inlet distortion flow field; the probe comprises a probe head portion and a pole (1); the probe head portion comprises a temperature sensor mount hole (2), a pressure hole (3), a hot line support (4) and a hot line (5) welded on the hot line support; a pressure sensor is arranged in a position close to the pressure hole (3) in the probe head portion; the pressure hole (3) axis is vertical to the hot line (5); a lead wire (7) is led out through a pole inner channel; the temperature sensor mount hole (2) and the pressure hole (3) are in a same circumferential position and equal in sizes; compared with a conventional pressure probe and a temperature probe, the combined dynamic probe can simultaneously measure the inflow total pressure and temperature, speed, direction and speed pulsation dynamic parameters through calibrations, and is suitable for measuring the turbine machine inlet distortion flow field total temperature, total pressure and turbulence.

The invention provides a method, equipment and medium for predicting the stability boundary of an air compressor under the condition of air inlet distortion, which comprises the following steps: calculating a blade domain by adopting a distributed force source volume force model, calculating a non-blade domain by adopting a Reynolds average method, giving a distortion condition required by a target condition by adopting an inlet boundary condition, and calculating the stability boundary of the air compressor under the condition of air inlet distortion. Therefore, the spatial distribution of the compressor load is obtained. According to the method, the flow stability is analyzed and predicted based on the understanding of the gas compressor load space distribution and the critical load, it is considered that when the load of the gas compressor existing area reaches the local critical load, instability can be triggered, and high prediction precision is achieved; in addition, due to the fact that the calculation method based on the distributed force source volume force model is adopted, rapid obtaining of the air compressor load space distribution under the air inlet distortion condition can be achieved, and therefore the efficiency of stable boundary prediction work can be greatly improved.

The invention provides a test piece mechanism of a high-pressure compressor. The mechanism is used for simulating the working environment of a whole turbofan engine and comprises a high-pressure compressor body, an inlet support plate and a transition section flow channel are arranged at the inlet of the high-pressure compressor body, a distortion device is arranged at the upstream of the inlet support plate and the transition section flow channel, and the distortion device is positioned in an outer flow channel of the inlet support plate and the transition section flow channel. The inventionprovides a high-pressure compressor test structure form capable of accurately simulating a complete machine environment. The high-pressure compressor test structure form comprises an inlet distortiondevice form, a high-pressure compressor body and an outlet simulation combustion chamber form. Meanwhile, a set of high-pressure compressor inlet distortion device structure scheme is provided, an inlet pressure profile can be generated, and the profile is consistent with the inlet pressure profile of the high-pressure compressor in the whole machine environment.