19 results about "Inlet distortion" patented technology

The invention belongs to the field of test fluid mechanics, and discloses a flow correction method considering air inlet distortion and a calibration device. The flow calibration device comprises a standard flow spray pipe, an air inlet section, a measurement section, a rectification section and an exhaust spray pipe which are sequentially connected from front to back. The flow correction method considering the air inlet distortion comprises the steps that distortion air inlet flow correction is conducted; flow and distortion intensity evaluation is carried out; training a neural network; carrying out an air inlet channel test; calculating an air inlet flow distortion parameter; and correcting a flow calculation result. According to the flow correction method and the calibration device considering the air inlet distortion, the air inlet distortion, the total pressure loss and the Reynolds number influence are comprehensively considered, technical support can be provided for high-precision flow monitoring of a large-curvature aircraft air inlet performance evaluation test, and the method and the calibration device have high engineering practical value.

The invention provides a ground vortex intensity prediction method under a crosswind condition, and the method comprises the following steps: S1, judging whether a ground vortex is generated or not according to a ground vortex intensity prediction model under the crosswind condition, and calculating a dimensionless circulation under a corresponding working condition; s2, according to an air inlet distortion index prediction model, calculating the size of an air inlet total pressure distortion index caused by suction of a ground vortex; and S3, according to a CFD calculation result, determining the model output parameter precision and aerodynamic characteristics reflected by the model output parameter precision so as to verify the reliability of the model. The invention provides a reliable ground vortex intensity prediction model for 90-degree crosswind conditions. The ground vortex intensity prediction model not only can qualitatively judge whether a ground vortex is generated, but also can quantitatively calculate ground vortex intensity and air inlet total pressure distortion.

The invention discloses a rapid and high-precision adjustable conical inserting plate type air inlet distortion device which comprises a control module, a distortion generator body, an air inlet flange, an air inlet pipeline, an air outlet pipeline and an air outlet flange. A control module is mounted on the top surface of the distortion generator main body; one end of the air inlet pipeline is connected with the front surface of the distortion generator main body; one end of the air outlet pipeline is connected with the back surface of the distortion generator main body; an upper blanking cap is installed on the top face of the valve body, a lower blanking cap is installed on the bottom face of the valve body, and a movable inserting plate is installed in the valve body. The lower end of the bracket is connected with the top surface of the upper blanking cap; the upper end of the lead screw is connected with the output end of the servo motor, and the middle part of the lead screw is provided with a lead screw nut; and the connecting sleeve sleeves the outer side of the lead screw. According to the invention, a traditional hydraulic mechanism can be replaced, rapid movement and accurate positioning of the plugboard can be realized, and the reliability of the measured distortion index can be improved.

The invention provides an engine air inlet distortion suppression device and method, the engine air inlet distortion suppression device is arranged on an engine air inlet pipeline, the air inlet pipeline extends in the axial direction of the engine air inlet pipeline, the air inlet pipeline is provided with a transverse section perpendicular to the axial direction of the engine air inlet pipeline, and the engine air inlet distortion suppression device is characterized in that the suppression device is arranged on the transverse section and comprises a flow expander, circumferential blades and radial blades; the center of the flow expander coincides with the center of the transverse section, the projection area of the flow expander on the transverse section is adjustable in the radial direction of the transverse section, and the radial direction vertically extends from the center of the transverse section to the peripheral side of the transverse section; the circumferential blades are distributed around the flow expander and can rotate around the circumferential direction, so that the projection area of the circumferential blades on the transverse section is adjustable in the radial direction, and the circumferential direction is the direction around the axial direction on the transverse section; the radial blades are distributed in the radial direction of the transverse section and can rotate in the radial direction, so that the projection area of the radial blades on the transverse section is adjustable.

The application provides a wing-body fusion back support engine reverse thrust device capable of inhibiting air intake distortion, which comprises a ring-shaped reverse thrust vane located on the inner side of a back support engine nacelle cover and reverse thrust flow guiding plates located on the two side surfaces of a back support engine nacelle support. When reverse thrust is needed, with the increase of the unfolding angle of the flow guiding plates, the hindering effect of the flow guiding plates on the forward movement of the reverse thrust air flow is more and more significant, so that the air intake quality of the turbofan engine during the reverse thrust work is ensured; in addition, the flow guiding plates also have the interference effect on the air flow similar to that of a spoiler, so that the form resistance is increased, which is beneficial to the faster deceleration of the airplane. The application effectively solves the adaptability problem of the reverse thrust device and the wing-body fusion back support engine, inhibits the occurrence of air intake distortion, strengthens the braking effect of the reverse thrust device, and guarantees the flight safety of the wing-body fusion back support engine layout.

The invention belongs to the technical field of engine nacelle production, and particularly relates to an automatic welding system for an inner supporting framework for engine nacelle production, which comprises a frame, a fixed clamping assembly, a movable clamping assembly, an outer seam welding assembly, a cover ring pushing assembly and an inner seam welding assembly. The frame is composed of a cylinder part and an end plate. The fixed clamping assembly is used for clamping the inner heat dissipation pipe section and driving the inner heat dissipation pipe section to rotate in the circumferential direction. The movable clamping assembly clamps the outer radiating pipe section and achieves axial displacement of the outer radiating pipe section. The outer seam welding assembly is responsible for welding reinforcement and outer seam welding of the spoiler and the middle supporting pipe section; the cover ring pushing assembly adsorbs the cover ring piece and axially pushes the cover ring piece; and the inner seam welding assembly clamps the middle supporting pipe section and performs axial displacement, and inner seam welding is completed. Automatic welding of the inner supporting framework is achieved, the welding efficiency is remarkably improved, meanwhile, the inner supporting framework has the air inlet fairing supporting function, the heating function and the low-speed air inlet distortion control function, and the problems that in the prior art, efficiency is low, and the function is single are solved.

The application belongs to the technical field of engine nacelle production, and particularly relates to an automatic welding system for an inner support framework for engine nacelle production, which comprises a frame, a fixed clamping assembly, a movable clamping assembly, an outer seam welding assembly, a cover ring pushing assembly and an inner seam welding assembly. The frame is composed of a cylinder part and an end plate. The fixed clamping assembly is used for clamping an inner heat dissipation pipe section and driving the same to rotate circumferentially. The movable clamping assembly clamps an outer heat dissipation pipe section and realizes axial displacement of the same. The outer seam welding assembly is responsible for welding and reinforcing of a spoiler and an intermediate support pipe section and outer seam welding. The cover ring pushing assembly adsorbs a cover ring and pushes the same axially. The inner seam welding assembly clamps the intermediate support pipe section and realizes axial displacement of the same, thereby completing inner seam welding. The application realizes automatic welding of the inner support framework, significantly improves welding efficiency, and enables the inner support framework to have the functions of air inlet rectifier cover support, heating and low-speed air inlet distortion control, thereby solving the problems of low efficiency and single function of the prior art.

A method for a gas turbine engine includes calculating an inlet airflow distortion value at a reference area within an airflow duct and in front of a fan of the gas turbine engine. The inlet airflow distortion value is indicative of an air pressure of an air vortex at the reference area caused by a set of crosswind conditions. The calculating is based on a ratio between an inlet airflow static air pressure in the reference area and a total ambient air pressure. The inlet airflow distortion value is compared to a threshold corresponding to a permissible amount of inlet airflow distortion at the reference area for the set of crosswind conditions, and, based on the comparison indicating that the inlet airflow distortion exceeds the threshold, an inlet airflow distortion notification is provided. A system for a gas turbine engine and a method for a gas turbine engine are also disclosed.

This invention belongs to the field of aero-engine performance simulation technology, specifically disclosing a data-driven method for predicting compressor flow field and performance. The method includes: acquiring operational data of the target equipment under different operating conditions using a volumetric force model and forming a snapshot matrix; decomposing the snapshot matrix and extracting the spatial mode matrix and modal coefficient matrix of the inlet and outlet flow fields, and constructing a training set using these matrices; inputting the training set into a BP neural network to continuously update the mapping relationship between the inlet and outlet flow field modal coefficient vectors; inputting the sampled data into the BP neural network to perform flow field reconstruction and predict the performance parameters of the target equipment. The method has the following advantages: the volumetric force model reduces data acquisition costs, and the POD method performs data dimensionality reduction and feature extraction, enabling accurate prediction of complex distorted flow fields and performance parameters even under small sample conditions; the data-driven flow field prediction method accurately captures high-dimensional flow field features, overcomes the problem of predicting nonlinear characteristics of the flow field under inlet distortion, improves prediction accuracy, reduces computational costs, and optimizes performance prediction efficiency.

The invention discloses an aero-engine air inlet distortion simulation device, and relates to the technical field of engine air inlet distortion simulation devices.The aero-engine air inlet distortion simulation device comprises an air duct and multiple sets of fan blades arranged in the air duct, a base ring is arranged in the air duct, the multiple sets of fan blades are located at the horizontal left side position of the base ring, and the base ring is provided with a speed pressing mechanism for adjusting the horizontal right side air speed of the base ring; the rapid pressing mechanism comprises a rotating assembly used for driving the base ring to freely rotate in the vertical direction, four groups of wind passing areas which are annularly symmetrical from the axis and are distributed in a divergent mode are arranged on the base ring, and wind adjusting assemblies used for adjusting the wind speeds of the right sides of the four groups of wind passing areas correspondingly are arranged on the base ring; according to the invention, by arranging the speed pressing mechanism, the purpose of adjusting and simulating the airflow speed at different positions of an engine air inlet channel can be achieved by independently controlling the rotating speed of a single set of fan blades and the independent adjusting process of the air outlet speed of four groups of air passing areas, and meanwhile, the response speed of the gust simulation process is further improved; the simulation experiment efficiency is improved.

The invention provides a high-degree-of-freedom distortion index simulation device, and belongs to the technical field of aero-engines, the high-degree-of-freedom distortion index simulation device comprises a total pressure distortion plate positioning switching assembly, a turbulent flow plugboard adjusting assembly and a fixed pipeline, the total pressure distortion plate positioning switching assembly comprises a movable motor base, a distortion plate socket, a distortion plate and a base, a gear on the movable motor base is connected with a rack below a distortion plate socket in a meshed mode, the distortion plate socket is connected to the base through a guide rail and a lead screw, a groove is formed in the distortion plate socket, and a distortion plate is connected into the groove in a sliding mode through a pin shaft. The turbulent flow inserting plate adjusting assembly comprises a fan-shaped inserting plate and sealing sliding blocks, the sealing sliding blocks are arranged at the two ends of the fan-shaped inserting plate, a T-shaped groove is formed in the fan-shaped inserting plate, a guide block is slidably connected into the T-shaped groove, and the guide block is fixed to an annular plate of a fixed pipeline. The high-degree-of-freedom distortion index simulation device is adopted, and the problems that a traditional air inlet distortion test is low in efficiency, high in cost, limited in evaluation accuracy and poor in engineering applicability are solved.

The invention provides an air inlet distortion device. The air inlet distortion device comprises a first pipeline and a second pipeline which are sequentially arranged in the axial direction of the distortion device. The first pipeline is sleeved with the first part of the flexible sealing sleeve, the second pipeline is sleeved with the second part of the flexible sealing sleeve, and the joint position between the first pipeline and the second pipeline is wrapped with the third part of the flexible sealing sleeve; at least one tightening piece is connected with the first part of the flexible sealing sleeve in a sleeving manner, so that the flexible sealing sleeve is tightened and fixed to the first pipeline; at least one tightening piece is connected with the second part of the flexible sealing sleeve in a sleeving mode so that the flexible sealing sleeve can be tightened and fixed to the second pipeline. And the flexible buffer gasket is arranged between the first pipeline and the second pipeline. In conclusion, through a flexible connection structure formed by the flexible sealing sleeve and the flexible buffer gasket, mechanical damage can be avoided while the air tightness of the distortion device is ensured.

The invention discloses an automatic adjusting device for a radial air inlet angle of an axial flow compressor, and belongs to the field of axial flow compressors. Comprising a large rotation module, a small rotation module, a telescopic adjustment module and a flow deflector rotation module. The large rotary module comprises a device base and an annular gear ring fixed on the device base; the small rotary module drives a circumferential angle positioning gear on the small rotary module to be meshed with the annular gear ring through a steering engine, so that the small rotary module can do circular motion along the large rotary module; the telescopic adjusting module is installed on the small rotation module and drives a telescopic plate to do radial reciprocating motion through gear-rack transmission so as to adjust the radial action range of the flow deflector. The flow deflector rotating module is installed at the tail end of the telescopic plate and drives the flow deflector to rotate through at least one air cylinder so as to change the flow area of the air inlet channel. According to the method, the core problem that local and dynamic air inlet distortion cannot be dealt with in the traditional technology is solved.

The invention provides an ice throwing device which comprises a support assembly and an ice throwing assembly, and the ice throwing assembly comprises an ice storage box arranged on the support assembly and provided with an ice storage cavity and an ice outlet communicated with the ice storage cavity; the ejection assembly is arranged in the ice storage cavity and is opposite to the ice outlet; the stop assembly is movably arranged in the ice storage cavity and located between the ejection assembly and the ice outlet; an ice storage area defined between the ejection assembly and the stop assembly is used for placing ice blocks; the stop assembly can be located at the stop position so as to stop ice blocks in the ice storage area. The ejection assembly is used for driving the ice blocks in the ice storage area to be ejected out through the ice outlet when the stop assembly leaves the stop position. Potential energy generated by compression of the ejection assembly is used for adjusting the initial speed when the ice blocks are ejected out of the ice outlet, then the flight path of the ice blocks is adjusted, the precision of the ice blocks projected into the engine is improved, the test uncertainty is reduced, and the risk of air inlet distortion of the engine is reduced.

The application relates to a modeling method of an engine with inlet distortion based on parallel compressor theory, comprising an air path model of the engine, a matching method of the air path model and application. On the basis of an original double-shaft split-fan engine model, three calling models of a distortion splitter model, a parallel fan model and a mixing chamber model are added, and then a new air path model matching of the engine is carried out. The application realizes accurate control of an airplane under special working conditions. The application is based on a parallel compressor model, takes a double-shaft split-fan engine with a large bypass ratio as a research object, and establishes an aero-engine simulation model considering inlet distortion.

The invention discloses an aero-engine inlet distortion total pressure flow field reconstruction method, and belongs to the technical field of engine inlet flow field total pressure distortion flow field reconstruction. The method comprises the following steps: collecting total pressure data, wall static pressure data and flight parameters of an engine inlet cross section AIP, and constructing a training data set; performing normalization processing on the training data set; constructing a flow field reconstruction network model, wherein the flow field reconstruction network model comprises an S-Pyraformer model and a physical information neural network (PINN) model; defining a PINN loss function which comprises data fitting loss and physical equation residual loss, and guiding the network to meet the physical law of the flow field while fitting data through the loss function; training the flow field reconstruction network model by using a training data set, and optimizing network parameters; and performing total pressure flow field reconstruction on the new input parameters by using the trained model, and outputting a reconstruction result and a precision evaluation index thereof. According to the method, the problems of precision, real-time performance and physical consistency in aero-engine inlet distortion flow field reconstruction are solved.

The invention provides an engine slow state air inlet distortion evaluation method, and belongs to the technical field of aero-engine air inlet performance calculation, and the method comprises the steps: constructing an engine slow state numerical calculation model, the model comprises an air inlet channel outlet rear section and a throttling cone, the air inlet channel outlet rear section comprises a straight section and an expansion section, the throttling cone is arranged in the expansion section, and the rotating speed change of the engine is simulated by changing the axial position of the throttling cone, so that the throttling area is changed, the outlet flow of the air inlet channel is controlled, and the slow state simulation of the engine is realized; grid division and solver selection are carried out on the engine slow state numerical calculation model, boundary conditions are set, engine slow state steady flow field solving is carried out, and flow field parameters of an air inlet channel outlet simulation section are obtained; and discretizing an air inlet outlet simulation section which is the same as the real air inlet outlet section, and calculating an air inlet steady-state circumferential distortion index when the engine is in a slow state based on discrete point flow field parameters of the air inlet outlet simulation section.

The application relates to a non-axisymmetric smooth transition circumferential groove casing treatment design method, and belongs to the technical field of internal flow control of a turbomachine; the inlet distortion condition is real inlet distortion with radial non-uniformity and circumferential non-uniformity, and has a greater influence on the stability margin of a duct fan; the non-axisymmetric circumferential groove casing treatment design method is non-axisymmetric for the low-energy fluid distribution of the duct fan at a stall condition under the inlet distortion, deep grooves are formed in blade channels with serious blockage, shallow grooves are formed in blade channels with weak blockage, and a linear smooth transition is adopted between the deep grooves and the shallow grooves. The non-axisymmetric circumferential groove casing treatment design method is applied to the stability expansion of the duct fan under the inlet distortion condition, and can greatly improve the stability margin of the duct fan.