37 results about "Transonic" patented technology

The invention discloses an aerodynamic configuration structure of a precision guidance assembly and a design method, and belongs to the technical field of precision guidance, an assembly precursor adopts continuous three-section curve design: a supersonic speed section adopts a parabola profile, is provided with ribs on the surface, and is used for optimizing shock wave control under a supersonic speed working condition; the transonic speed section adopts a hyper-ellipsoid curved surface, ribs are arranged on the surface of the transonic speed section, and metamaterials are attached to the ribs and used for restraining transonic speed separation; the subsonic speed section adopts arc transition, and ribs are arranged on the surface and used for improving the lift force efficiency of the control surface; wherein the curvatures at the joints of the three sections of curves are continuous. According to the method, supersonic velocity-transonic velocity-subsonic velocity three-section collaborative design is realized, and meanwhile, a boundary layer is improved from a microscopic level by combining a collaborative drag reduction design of a microscopic rib and metamaterial technology, and the performance of a guidance assembly is optimized and improved.

The invention provides an atmospheric data analysis method based on observability analysis, belongs to the field of aircraft navigation, guidance and control, and aims at solving the problem that the precision of traditional atmospheric data measurement is reduced under the condition of a complex flow field or a sensor fault. In the prior art, due to lack of strict theoretical support, estimation errors and convergence problems may occur in a transonic stage. According to the method, an atmospheric data state space model is established through flight dynamics, and atmospheric data estimation is carried out by adopting an extended Kalman filter based on the model; in the estimation process, observability analysis is executed to verify the reliability of an estimation result; and finally, optimizing the system model according to an analysis result. According to the method, the dependence of an existing method on an empirical model is reduced, the convergence and the stability of an estimation result are ensured, and the adaptability and the robustness of a system in a complex environment are remarkably improved.

The application discloses a high-pressure-ratio transonic axial flow compressor, which comprises a hub, front-row rotor blades, rear-row rotor blades and stator blades; the rear-row rotor blades are arranged between the front-row rotor blades and the stator blades; a plurality of front-row rotor blades and a plurality of rear-row rotor blades are uniformly arranged on the hub along a circumferential direction; and the front-row rotor blades and the rear-row rotor blades one-to-one correspond to form tandem rotors. The application realizes shock wave supercharging through the front-row rotor blades, and the rear-row rotor blades complete large airflow deflection; meanwhile, the application adopts an adsorption type stator to reduce or prevent airflow separation in a blade channel, thereby significantly improving a compressor stage pressure ratio and improving aerodynamic performance of the compressor. The application can realize a higher stage pressure ratio, and in the case of high inlet Mach number, high load and large airflow deflection angle, the application overcomes the problem that a high stage pressure ratio and efficiency are difficult to realize due to the fact that flow separation in the compressor cannot be controlled in the prior art.

This paper discloses an empirical formula blade generation system for designing subsonic and transonic turbine stages. The values of the design variables used in the parameterized system are calculated by the relevant equations. Iteration is required to modify the design variables. The parameterized system then generates the geometries and blocks of the stator and rotor. Using these geometries and blocks, a mesh is automatically formed using an ICEM CFD script. The turbine stage performance is then evaluated using a CFD solver. The calculation results are analyzed to identify future improvements. The final blade design is obtained through multiple rounds of repeated training and residual iteration. Finally, a highly efficient subsonic and transonic turbine stage is produced.

The present application discloses a tail wing control mechanism, which relates to the field of aircraft. A power piece is installed on the fuselage, and its output end is fixedly connected to the first rocker arm assembly and the second rocker arm assembly; a rotating mechanism, one end of which is rotatably connected to the first rocker arm assembly, and the other end of which is fixedly connected to the stabilizer surface of the tail wing; a locking mechanism is arranged between the first rocker arm assembly and the rotating mechanism; wherein, when the rotating mechanism is in a locked state, the rotating mechanism transmits the driving force of the first rocker arm assembly to the stabilizer surface, so that the deflection direction of the stabilizer surface is always opposite to the deflection direction of the rudder surface; the end of the second rocker arm assembly away from the output end of the power piece can be rotatably passed through the stabilizer surface on the front side of the rudder surface, and is fixedly connected to the rudder surface. The present application replaces the simple increase of the rudder surface area with mechanical linkage, avoids the problem of rudder effect overload during low-speed flight, takes into account the full-speed range control requirements, and effectively solves the contradiction between high-speed or low-speed rudder effect of transonic aircraft.

The application discloses a kind of launch vehicle fairing inner noise environment fast prediction method, comprising: obtaining the total noise of the core first stage engine nozzle center jet of launch vehicle;The distance from the characteristic interface of fairing to the interface where the core first stage engine nozzle is calculated;The jet outer noise at the characteristic interface of fairing is calculated;Using the aerodynamic noise empirical formula, the maximum aerodynamic outer noise of fairing in transonic flight phase is calculated;Based on the jet outer noise at the characteristic interface of fairing and the maximum aerodynamic outer noise of fairing in transonic flight phase, the maximum sound pressure level spectrum envelope of fairing is calculated;Based on the maximum sound pressure level spectrum envelope of fairing, combined with the sound insulation amount of fairing cabin section, the inner noise of fairing is calculated.The method can quickly predict the inner noise environment of new research launch vehicle fairing, solve the problem of new research launch vehicle fairing inner noise environment condition design.

An aircraft capable of transonic flight is equipped with a slotted airfoil to significantly improve the cruise performance during flight. The aircraft includes a plurality of slotted airfoils uniquely configured to maximize natural laminar flow across the aircraft wings. In addition to reduced wing drag, the unique configuration minimizes shocks at various conditions over the course of a flight. Each slotted airfoil includes a fore element, an after element, and a slot disposed between the fore element and aft element. The aft element is maneuverable relative to the fore element. The slot is configured to promote a substantially constant air flow field onto the aft element during all flight angles of attack.

The application discloses a kind of aerodynamic configuration structures and design methods of precision guidance assembly, belong to the field of precision guidance technology, assembly precursor uses continuous three curve design: supersonic section, uses parabolic profile, surface is provided with rib, for optimizing shock control under supersonic working condition;Transonic section, uses hyperelliptic surface, surface is provided with rib and is pasted on rib with super material, for inhibiting transonic separation;Subsonic section: using circular arc transition, surface is provided with rib, for improving rudder surface lift efficiency;Wherein, curvature is continuous at three curve junctions.The application realizes supersonic-transonic-subsonic three-section coordinated design, simultaneously combines the coordinated drag reduction design of micro rib and super material technology, improves boundary layer from micro level, optimizes and improves the performance of guidance assembly.

The invention provides a method and device for evaluating the comprehensive cooling efficiency of a turbine guide vane by using reverse heat transfer, and the method comprises the following steps: measuring a local convective heat transfer coefficient of the outer surface of a pneumatic test piece blade, and defining the local convective heat transfer coefficient as an external convective heat transfer coefficient without the influence of an internal cooling flow; acquiring a convective heat transfer coefficient between an internal cooling flow of the heat transfer test piece and a wall surface, and defining the convective heat transfer coefficient as an internal cooling coefficient; in the transonic speed blade grid wind tunnel, introducing mainstream gas into the outer surface of the heat transfer test piece and introducing heated internal cooling gas into an internal cooling cavity at the same time, so as to form a temperature difference through reverse heat transfer; measuring the convective heat transfer coefficient of the outer surface of the blade under the condition; and combining the Bit number determined by the external convective heat transfer coefficient and the wall thickness-heat conduction ratio, the ratio of the external convective heat transfer coefficient to the internal cooling coefficient, and the ratio of the correction coefficient to the internal cooling coefficient according to a preset function relationship to obtain the dimensionless comprehensive cooling efficiency. The problem that a traditional high-temperature test is difficult to implement is solved.

In order to solve the problems that in the prior art, calculation precision and efficiency are difficult to consider at the same time during supersonic jet machining, and detached shock waves cannot be accurately captured, the invention provides a finite element simulation method for the stagnation flexible constraint flow field characteristics of a jet impact area. The finite element simulation method comprises the steps that firstly, a geometric model of a jet area and an outer watershed is established, and preliminary solving is conducted; extracting jet flow axis pressure gradient characteristics to determine detached shock wave interval coordinates; then reconstructing a block geometric model comprising an upper jet flow region, a detached shock wave region, a stagnation region and an outer watershed according to the coordinates, and implementing differential local grid encryption on the shock wave region and the stagnation region; and finally, carrying out second iterative solution based on the Sutherland viscosity law to obtain final flow field distribution. According to the method, the thickness and the position of the shock wave layer are accurately captured while the number of the grids is controlled through the block encryption strategy, the contradiction between the calculation economy and the numerical precision under the transonic flow condition is solved, and the method is suitable for numerical simulation of the complex supersonic gas jet flow field containing the detached shock wave and the stagnation region.

The application discloses a method for suppressing transonic buffeting load and response based on agent cooperation, and relates to the technical field of aeroelasticity, and aims at solving the problem of transonic buffeting of a wing of an aircraft. The method comprises the following steps: firstly, calculating the elastic wing tip acceleration response based on a one-way fluid-solid coupling method; secondly, configuring a multi-target reinforcement learning agent, taking the variable-camber trailing edge and the telescopic micro-bulge as control mechanisms, taking the maintenance of the aerodynamic lift-drag ratio and the suppression of the vibration response of the structure as targets, and letting the agent adopt a proximal policy optimization algorithm to be trained to obtain an optimal control law capable of cooperatively driving the telescopic micro-bulge and the variable-camber trailing edge; and thirdly, deploying the trained cooperative control law on the target wing, and observing the suppression effect of the wing buffeting through quantitative evaluation of the slowing rate. The application trains the agent through a deep reinforcement learning method, so that the agent can guide the telescopic micro-bulge to suppress the diffusion in the separation zone, reduce the shock motion amplitude, drive the variable-camber trailing edge to regulate the flow at the trailing edge and compensate the overall lift fluctuation, and cooperatively suppress the buffeting load and response.

The invention relates to the technical field of aircrafts, in particular to an aircraft propulsion system and an aircraft, the aircraft propulsion system comprises a cartridge receiver, a propulsion device, a control device and an energy storage device; the propelling device comprises at least one stage of magnetic levitation rim motor set, a combustion chamber and an exhaust nozzle. The combustion chamber is located at the exhaust end of the magnetic levitation rim motor set and communicates with the exhaust end. The exhaust nozzle is located at the air outlet end of the combustion chamber and communicates with the air outlet end. The magnetic levitation rim motor set comprises at least one magnetic levitation rim motor, all the magnetic levitation rim motors are electrically connected with the energy storage device, and the control device is electrically connected with the energy storage device and all the magnetic levitation rim motors. And the control device is used for controlling each magnetic levitation rim motor to switch between a power generation state for supplying power to the energy storage device and an electric propulsion state for pressurizing air entering from the air inlet end according to the flight speed of the aircraft, so that the thrust requirement of the aircraft during high-speed flight is met, and the transonic flight capability of the aircraft is realized.

The application belongs to the technical field of wind tunnel test, and particularly relates to a transonic fan blade engine wind tunnel test system. By introducing a high-precision data synchronous recording device, it is ensured that the displacement signals captured by the vibration sensing component and the pressure data obtained by the dynamic pressure acquisition unit can be completely aligned in time. At the same time, the system also integrates an environmental parameter monitor, which can track and compensate the changes of physical quantities such as temperature, humidity and air pressure in real time, thereby significantly improving the accuracy of experimental data. Finally, the addition of an analysis platform and a result visualization interface realizes in-depth analysis of the blade vibration mode, provides intuitive and detailed data support for researchers, and greatly improves the understanding of the dynamic behavior of the blade under transonic conditions.

The application discloses a transonic buffet control structure based on a wing trailing edge air vent, and belongs to the technical field of aircraft flow control. The buffet control structure is an air vent arranged in the wing trailing edge. One end of the air vent is communicated with the upper surface of the airfoil, and the other end is communicated with the blunt trailing edge of the airfoil. The aperture of the air vent and the thickness of the trailing edge of the airfoil are of the same order of magnitude. The application can control the buffet while reducing the influence on the lift-drag characteristics of the original wing.

The invention provides a vehicle nose cone transonic speed reverse design method based on a bending shock wave theory and a vehicle nose cone. The method comprises the steps that incoming flow parameters are determined and designed according to a flight envelope; determining a curved shock wave surface shape function and a discrete shock wave molded line according to head cone geometric constraints, and solving discrete point aerodynamic parameters and first-order derivatives based on a curved shock wave theory; solving a subsonic velocity flow field control equation by adopting a time domain characteristic line method; and obtaining external flow field information by simultaneous equations, and reversely generating a nose cone wall surface. According to the method, through the innovative logic of'presetting the shock wave form-reversely deriving the wall surface ', the bending shock wave theory and the time domain characteristic line method are fused, the transonic shock wave structure is accurately controlled, the wall surface pressure distribution is optimized, the flow separation and shock wave oscillation are inhibited, the design efficiency is improved while the design precision is ensured, and the method is suitable for large-scale popularization and application. Reliable technical support is provided for rapid optimization design of the nose cone of the vehicle, and the method is suitable for aerodynamic design of nose cones of various transonic vehicles.

The invention provides a treatment casing for an ultrahigh-load transonic-speed serial rotor and a gas compressor, and relates to the technical field of pneumatic design of gas compressors. The serial rotor comprises a front-row rotor and a rear-row rotor which are sequentially arranged in the axial direction. The treatment casing comprises a casing body and a plurality of casing channels. The casing body is arranged outside the serial rotor in a surrounding mode. The multiple casing channels are arranged on the outer wall face of the casing body at intervals in the circumferential direction of the casing body, and each casing channel comprises a suction section, a connecting section and a spraying section. The suction section extends outwards from the outer wall face of the casing body and is suitable for sucking high-pressure air flow at the tail edge of a rear row rotor, the connecting section extends in the axial direction, and one end of the connecting section is connected with the suction section. The injection section extends from the other end of the connecting section to the axis of the casing body and is suitable for spraying the high-pressure airflow from the suction section to the front edge of the rear row rotor, and the included angle between the airflow sprayed out of the injection section and the inner wall face of the casing body is larger than 55 degrees and smaller than 70 degrees.

The invention discloses an airfoil profile stability optimization method and system for inhibiting transonic buffeting, and relates to the technical field of aircraft safety. According to the method, the geometric morphology of the airfoil profile is accurately described through CST parameterization, the grid quality and the calculation precision after the airfoil profile is deformed are guaranteed by combining the radial basis function dynamic grid technology, the flow stability is quantified by relying on the pre-solution gain, efficient iteration is conducted through the nonlinear conjugate gradient method, transonic buffeting can be remarkably restrained, and the flow stability can be remarkably improved; and the aerodynamic performance can be optimized on the premise that engineering constraints such as no reduction of the airfoil thickness and no reduction of the lift force are met, experience trial and error and violent search of traditional optimization are avoided, the efficiency and reliability of transonic airfoil design are greatly improved, and the method is suitable for high-performance design requirements of core components such as aerospace vehicle airfoils.

The invention belongs to the technical field of aviation aircraft design, and particularly relates to an aerodynamic configuration structure of a supercritical wing, the wing is generated by supercritical airfoil profile control of five specific stations in the spanwise direction, and 45% of the stations are reference airfoils. The front edge radius, the thickness and the mounting angle of the inner side (0% and 20% station) airfoil profile are sequentially larger than those of the reference airfoil profile; the leading edge radius of the outboard (70%, 100% station) airfoil profile is significantly larger than that of the reference airfoil profile, the mounting angle is negative, and the absolute values are sequentially increased. And front and rear edge points of each section are connected by using secondary splines to form front and rear edges of the wing, so that a three-dimensional molded surface is generated. Through differential design of spanwise airfoil parameters, a large-area separation area under a large angle of attack can be actively controlled to appear in the middle of the airfoil firstly, control failure caused by stalling of an outer wing and sudden lift drop caused by stalling of an inner wing are effectively avoided, and the stability of the airfoil is improved while the transonic cruise efficiency is maintained. And the stall safety of the upper single-wing transport plane is obviously improved.

The purpose of this invention is to provide a one-dimensional prediction method for the characteristics of a multi-stage axial-flow transonic turbine based on dimensionless numbers, comprising the following steps: turbine characteristic calculation; angle-of-attack loss calculation; total pressure loss calculation using a loss model; lag angle calculation using a lag angle model; transonic position determination; transonic cross-sectional characteristic parameter calculation; iterating the above process until the required accuracy is met, thus completing the characteristic prediction at the operating point. This invention's one-dimensional prediction method for the aerodynamic characteristics of a multi-stage axial-flow transonic turbine based on dimensionless numbers can plot turbine characteristic curves using turbine inlet velocity, inlet mass flow rate, and total expansion ratio as the abscissas; it can predict turbine design and non-design operating conditions; the method simultaneously considers the influence of angle of attack and lag angle, and incorporates a loss model; this method can simultaneously predict the characteristics of subsonic and transonic turbines; furthermore, this invention improves the accuracy of turbine characteristic prediction and shortens the design cycle of axial-flow turbines.

Apparatus and articles of manufacture are disclosed of an aircraft comprising: a brace to support a wing of the aircraft, the brace including a first end coupled to the wing and a second end coupled to a fuselage of the aircraft, the brace including: a strut extending between the first end and the second end; and an airfoil having a cavity containing the strut.

A design and evaluation method for a hypersonic wind tunnel visualization nozzle involves the field of hypersonic wind tunnel technology. It addresses the problem that traditional flow field performance evaluation methods have difficulty capturing detailed defects in the three-dimensional flow field. First, the nozzle profile is designed based on the improved Sivells method, and the axial Mach number distribution is constructed in combination with transonic theory. Second, the glass material is thermoplasticized using high-temperature flame spray technology to achieve precise control of the nozzle curvature. Subsequently, tracer particle images are collected synchronously using a dual-pulse laser and a high-speed camera, and the three-dimensional velocity field is reconstructed and key flow field parameters are extracted using a cross-correlation algorithm. Finally, numerical simulation and genetic algorithms are combined to optimize the nozzle profile parameters, and iterative corrections are made until the flow field uniformity meets the standard. This method breaks through the spatial limitations of traditional probe measurements and realizes the dynamic capture of the full-domain flow field. At the same time, by quantitatively correlating geometric errors and flow field distortion, the reliability of wind tunnel test data and the efficiency of design optimization are significantly improved.

The invention relates to a method for identifying shock wave motion frequency and growth rate, which comprises the following steps of: performing modal decomposition on airfoil profile surrounding flow field pressure data of transonic airfoil profile forced oscillation by using a single-frequency modal decomposition method to obtain frequency and modal characteristics of a surrounding flow field during transonic airfoil profile forced oscillation; the single-frequency modes are sequenced according to the energy ratio from large to small, the first-order mode is shock waves formed on the two sides of the airfoil profile under forced oscillation, and the frequency of the first-order mode is consistent with the motion frequency of the shock waves; the second-order mode reflects frequency doubling structures formed on the two sides of the airfoil profile in the swing process of shock waves on the two sides of the airfoil profile; the third-order mode reflects a slow evolution structure of the flow field from the initial state to the periodic pulsation state, and the frequency of the third-order mode is the frequency corresponding to the total physical duration of all the flow fields participating in mode decomposition; and identifying the shock wave growth rate according to the time coefficient amplitude of each order modal. The frequency and growth rate of the shock wave can be accurately captured, and the form of the shock wave is given.

The application belongs to the technical field of engine fan and compressor processing machine casing subsonic rotor blade design, and particularly relates to an engine fan and compressor processing machine casing subsonic rotor blade design method. The design is designed without a processing machine casing at a design rotating speed, and a shock wave form at a transonic rotor blade tip is designed as a near stall point form. On this basis, the design adopts a processing machine casing, and at the design rotating speed, the shock wave form at the transonic rotor blade tip is degraded from the near stall point form to a near highest efficiency point form, so as to ensure the efficiency of the fan and the compressor in the engine at the design rotating speed and avoid the efficiency decline caused by the adoption of the processing machine casing.

A fluid-dynamically effective component (2) comprises a mechanically supported component root (3), a free-ending component tip (6), a span (5) between the component root (3) and component tip (6), a negative pressure side (7), a positive pressure side (8), a component profile (14), and a contour bulge (12) on the positive pressure side (8). The contour bulge (12) deforms the component profile (14) relative to a basic profile shape (15), is limited to a region of the local profile depth (18), and extends over at least an outer region of 80% to 90% of the span (5), but is restricted to a region between 10% and 30% of the span (5).The component profile (14) is designed for an inflow (9) in the transonic velocity range, in which a lower pressure than in the overpressure side (8) and a supersonic flow region ending with a compression shock are formed on the low-pressure side (7), but no supersonic flow region ending with a compression shock is formed on the low-pressure side, and in which the contour bump (12) counteracts flutter of the component (2).

The present invention proposes a supersonic air film nozzle and a design method, device, and near-space waverider aircraft, including obtaining the supersonic air film nozzle outlet pressure of the high-speed waverider aircraft under pressure matching conditions, determining the supersonic air film mass flow rate, the supersonic air film Mach number, the supersonic air film nozzle throat height, and the nozzle outlet height of the supersonic air film nozzle, obtaining the transonic solution of the supersonic air film nozzle, and setting the Mach number distribution of the supersonic air film nozzle axis. The supersonic air film nozzle transonic solution, the nozzle outlet, and the Mach number distribution of the supersonic air film nozzle axis are used as boundary conditions to determine the nozzle inviscid profile, solve the boundary layer displacement thickness, obtain the final nozzle profile, and then obtain a complete supersonic air film nozzle. The wave is completely eliminated in the nozzle, shock waves are fundamentally eliminated, and the distribution of flow field parameters such as pressure, temperature, and velocity is controllable.

The application relates to the field of aerospace technology, in particular to a wide working domain grid fin integrated with a shock control structure, which comprises an external frame and a plurality of wing elements located in the external frame, the inside of the side wall of each wing element is provided with a pressure stabilization cavity and a ventilation cavity in communication with the pressure stabilization cavity, the pressure stabilization cavity is close to the opening of the wing element, and the ventilation cavity is close to the bottom of the wing element; the left and right wall surfaces of the side wall of each wing element are provided with a bleed air area and an exhaust area, the bleed air area is located at the leading edge of the wing element, the exhaust area is located at the bottom of the wing element, the left and right wall surfaces of the side wall of each wing element are provided with bleed air holes in communication with the bleed air area and the pressure stabilization cavity, and the left and right wall surfaces of the side wall of each wing element are provided with exhaust holes in communication with the exhaust area and the ventilation cavity. The application can widen the working speed domain, improve the transonic performance, expand the working attack angle range, enhance the maneuverability, reduce the flight resistance, improve the economy, and enhance the reliability of the structure and control.

The invention belongs to the technical field of aviation aircraft design, and discloses an aerodynamic design method suitable for middle stall control in the spanwise direction of a supercritical wing. According to the method, five specific supercritical airfoil profiles are configured in the spanwise direction of the airfoil, and the spanwise positions of the five specific supercritical airfoil profiles are 0%, 20%, 45%, 70% and 100% half-span respectively. Wherein the 45% half-span part is a reference airfoil profile. The front edge radius, thickness and mounting angle of the inner side (0%, 20%) airfoil profile are increased relative to the reference airfoil profile, the front edge radius of the outer side (70%, 100%) airfoil profile is greatly increased relative to the reference airfoil profile, and the mounting angle of the outer side (70%, 100%) airfoil profile is reduced, so that the control of the airfoil airflow separation position under the high incidence angle is realized. According to the method, a large-area separation area at a large angle of attack can be firstly controlled to occur in the middle of the wing in the spanwise direction, so that the risk of plane rolling and control failure caused by stalling of an outer wing firstly is avoided, and meanwhile, sudden drop of lift force caused by stalling of an inner wing is also prevented. The design can be completed by only using five airfoil profiles, and the stall safety of the upper single-wing transport plane is remarkably improved while the transonic speed cruise efficiency is ensured.

This invention discloses a transonic unmanned aerial vehicle (UAV) structure, including a fuselage, a nose, and a wing mechanism. The nose is located at the front end of the fuselage, and the wing mechanism is located on both the fuselage and the nose. The wing mechanism includes a canard, terminal wings, air guides, a dorsal engine nacelle, a dorsal air intake, winglets, connecting wings, a first air guide slot, a rear wing, a second air guide slot, a bottom wing, and an outwardly convex air guide surface. This solution, through the optimized design and integrated molding of multiple wing surfaces and air guide structures, improves airflow conditions, solves existing defects such as abnormal air intake and structural flutter, enhances aerodynamic efficiency, flight stability, and transonic flight performance, expands payload installation space and load-bearing capacity, and comprehensively optimizes the overall performance of the UAV.

This invention discloses a method for determining the initial value line of a unilateral expansion nozzle based on an artificial neural network, belonging to the field of aero-engines. The method involves the following steps: acquiring CFD calculation data of the exit transonic region of an asymmetric convergent nozzle under different aerodynamic / geometric parameters; extracting the contour line positions where the Mach number is greater than 1 from the CFD calculation data and discretizing these positions point-by-point; determining the neural network model architecture, setting optimization objectives and constraints; establishing a mapping between aerodynamic / geometric parameters and the distribution of contour line positions at the exit of the asymmetric convergent nozzle; verifying the prediction accuracy of the trained model; and using the trained model to predict the position of the contour line in the transonic region of the exit of an asymmetric convergent nozzle within a certain operating condition range for an asymmetric convergent nozzle, with the obtained contour lines serving as the initial value lines of the initial expansion domain of the unilateral expansion nozzle. This invention improves the efficiency and accuracy of determining the initial value line position of a unilateral expansion nozzle.

The application discloses an elevator and all-moving tail wing integrated control method and device, relates to the transonic aircraft field, and comprises the following steps: acquiring the critical Mach number, the first Mach number interval and the second Mach number interval of a target aircraft; acquiring the target control ratio of the elevator and the all-moving tail wing of the target aircraft in a preset Mach number range; when the speed of the target aircraft is in the first Mach number interval and the elevator is controlled, determining the target deflection angle of the all-moving tail wing based on the target control ratio; when the speed of the target aircraft is in the second Mach number interval and the all-moving tail wing is controlled, determining the target deflection angle of the elevator based on the target control ratio; and controlling the target aircraft to switch between the elevator control and the all-moving tail wing control according to the target deflection angle of the all-moving tail wing or the target deflection angle of the elevator. The application can solve the problems of the low control precision of the all-moving tail wing at subsonic speed and the insufficient controllability of the elevator at transonic speed in the prior art, and improve the flight control performance of the aircraft.