19 results about "Stagnation point" patented technology

The invention relates to a real-time measurement method for an enthalpy value in an ablation process of a thermal protection material in an electric arc wind tunnel test, is used for a ground test of an electric arc wind tunnel thermal protection system, and belongs to the research field of ground aerodynamic thermal tests of aircrafts. The non-equilibrium temperature (rotation, rotation and electron) and the concentration of atomic nitrogen and atomic oxygen of a stagnation point area of a stagnation point model of the heat-proof material are obtained, online measurement of the local center stagnation point enthalpy value of the heat-proof material is achieved, the whole-process real-time monitoring of the material ablation process is achieved, the enthalpy value simulation range of various types of arc wind tunnels such as tubular, segmented and laminated arc wind tunnels can be covered, and the performance of the heat-proof material is improved. The method has very wide applicability; and rapid and long-time measurement of the whole ablation process of the heat-proof material can be realized, the response frequency can be up to 100Hz, and the measurement period is greater than or equal to 1h.

A raised structure for reducing frictional drag due to viscosity of a flow toward an object in a direction defining an acute angle with a leading edge of the object. The raised structure includes raised bodies configured to be provided on a surface of the object at the leading edge on a downstream side of a stagnation point of the flow. A height of each raised body changes along a convex curve, and the raised bodies are arranged in an array to define a first uneven shape changing periodically in a first cross section configured to have a constant distance from the stagnation point and be orthogonal to the surface, and a second uneven shape changing in a second cross section configured to be orthogonal to a line composed of the stagnation point and the first cross section, the second uneven shape having concave and convex portions that change periodically.

The present invention relates to a sailing boat comprising an aerodynamic profile and a system for determining the characteristics of the incident air flow on its leading edge. In order to be able to precisely determine the position of the stagnation point at different regions along the leading edge of the aerodynamic profile, a system comprises: rows (S1 to S5) of pressure sensors (16) distributed on either side of the leading edge (14) and virtually forming a pattern (M1 to M5) spaced apart from each other in the form of a simple broken line; and a computer (20) connected to the pressure sensors (16). The computer (20) determines the corresponding stagnation point position along each of the patterns (M1 to M5), the stagnation point position being defined by the abscissa of the curve (a1 to a5) and by the height (z1 to z5) evaluated based on the corresponding height data from the pressure sensors (16) in the corresponding rows (S1 to S5), for which the pressure P*, interpolated based on the pressure measurements provided by the pressure sensors (16) in the corresponding rows (S1 to S5), is at a maximum.

The present application discloses a method, apparatus, device, and medium for determining boundary conditions at an aircraft engine inlet, relating to the field of fluid mechanics. The method comprises: during computational fluid dynamics numerical simulation of an aircraft, collecting air flow rates from each grid surface at the engine inlet boundary to determine the current actual flow rate at the engine inlet; determining a critical specific flow density based on stagnation point information in the current flow field, and determining a target velocity for the grid surface based on the critical specific flow density, target flow rate, and current actual flow rate; determining the local static pressure of the grid surface based on the critical velocity and target velocity, determining whether the gas flow corresponding to each grid surface is supersonic, and determining a target pressure value for the grid surface based on the judgment result and the local static pressure; and determining target density and target velocity component information for the grid surface based on the body center value of the flow field unit to determine the target boundary conditions. The present application can efficiently and accurately determine the boundary conditions of an aircraft engine inlet and improve robustness.

The present invention discloses a sweat-cooled leading edge thermal protection structure coupled with reverse steam injection, which belongs to the field of aircraft thermal protection technology, and includes a nose cone leading edge body, wherein the nose cone leading edge body includes a porous end cap, a cooling cavity, a front connecting pipe, a rear connecting pipe, an impermeable solid wall, a water injection pipeline, an internal cooling flow channel, a vapor-liquid separator, a front cooling pipe, a one-way throttle valve, a steam ejector and a reverse steam nozzle. The present invention uses liquid water to sweat-cool the leading edge non-stagnation area, and performs thermal protection on the stagnation area in the form of reverse steam injection. Under the protection of a double air film, uniform and efficient cooling of the stagnation and non-stagnation areas is achieved, effectively avoiding steam blockage and heat transfer deterioration in the stagnation area when the two areas sweat uniformly, as well as cooling transition in the non-stagnation area. The reverse steam injection can further push the bow shock wave of the leading edge away from the surface of the structure, reduce the aerodynamic pressure on the outer surface of the structure, and reduce the flow resistance of the liquid sweat-cooling medium.

The application discloses a flow parameter measuring device and method suitable for a wide range of angles of attack, and the measuring device comprises a ball head, a connecting rod, a pressure sensor and a data processing system; the pressure distribution on the surface of the ball head can be measured through the pressure sensor arranged on the ball head; the windward area is determined by the pressure data processing system, and the position of the stagnation point and the flow dynamic pressure are determined according to the position and the pressure measurement value of the multiple measuring points with the maximum pressure measurement value in the windward area, so that the flow speed direction and size are obtained. The measuring device of the application has no moving parts, is simple in structure and high in reliability, and solves the problem of quickly and accurately obtaining flow parameters when the moving device moves in a wide range of angles of attack.

The application discloses a high-efficiency lift-increasing flow control device suitable for a large-thickness airfoil, wherein a gas flow injection member is arranged on a pressure surface and located upstream of a free flow stagnation point, and adopts an inwardly inclined discharge angle of 0-5 degrees; a gas flow recovery member is arranged at the rear of a suction surface. A circulating driving member sucks low-momentum boundary layer fluid at a trailing edge, discharges the pressurized fluid from the injection member, and forms a zero-mass closed loop. In a positive angle of attack working condition, the injected gas exchanges momentum with the flow upstream of the stagnation point, further pushes the stagnation point position to the pressure surface, and expands the range of the negative pressure area at the front edge of the suction surface; meanwhile, the inwardly inclined discharge angle utilizes the wall adhesion effect of the large-curvature front edge to make the injected gas stably flow along the front edge, and forms a concentrated and strong negative pressure concentration area. The device can greatly increase the maximum lift coefficient, delay the stall angle of attack, and has a simple structure and no exposed moving parts, thereby providing a high-efficiency lift-increasing solution for a short-takeoff-and-landing aircraft.

The present application belongs to the technical field of hypersonic wind tunnel equipment construction, and discloses a design method for a hypersonic silent nozzle suction flow channel and a suction device. The design method is based on boundary layer thickness equivalent extrapolation for suction flow channel design. Numerical simulation is used with the nozzle outlet and the contraction section boundary layer thickness as the design input parameters. One-dimensional pipe flow theory is adopted, the lip vertex is the stagnation point, and there is no flow separation before the suction throat. The virtual interface between the inner and outer flows is reconstructed to ensure that the stagnation point position of the suction lip front edge is relatively fixed in a wide range of operation under the single-point design condition. The wall boundary layer is effectively sucked, and the upstream turbulent boundary layer disturbance of the nozzle is eliminated. The suction device is arranged upstream of the nozzle throat, and an annular suction pipe is arranged. The suction ring cavity is connected to the collection branch pipe, and the end of each collection branch pipe is inserted into the collection ring pipe. The incoming flow disturbance is suppressed, the flow is stabilized, the wide range operation requirement is met, and the engineering practical value is achieved.

The application relates to an arc wind tunnel test coaxial calorimeter heat measurement correction method and system, and belongs to the field of aerodynamic heat test research. n The application can realize accurate measurement of the ground test heat flow of the arc wind tunnel; the arc wind tunnel test coaxial calorimeter heat measurement correction method can be suitable for heat measurement of arc wind tunnel flat plate model test and stagnation point model test, and the heat flow test is suitable for very wide, covering 10kW / m 2 -30000kW / m 2 , and has a wide application range.

The invention provides a circular water film-based low-pressure nozzle impact performance optimization method, which comprises the following steps of: establishing a coordinate system by taking a jet flow impact point as an original point, and selecting at least four parameters from geometric parameter vectors of a nozzle to construct a joint objective function, the combined target value adopts a weighted product form of stagnation point pressure, water film area and jet flow hitting power; constructing a test matrix, changing at least four geometric parameters at the same time, and obtaining multi-dimensional response data of the joint target value through numerical simulation; building an explicit second-order model of the joint target value by using multivariate nonlinear regression; performing global optimization by adopting a genetic algorithm or a response surface method, and outputting an optimal geometric parameter combination which maximizes the joint target value; the improved nozzle is obtained through machining according to the optimal geometric parameter combination; and if the hitting power coefficient of the improved nozzle is greater than that of the original nozzle, the optimization is completed. According to the method, parallel search can be carried out in a multi-dimensional parameter space, and the optimization period is remarkably shortened.

This application discloses a surface pressure-based method for sensing gust flow fields, relating to the field of gust flow field sensing. The method fits the airfoil leading-edge parabola to determine the flow model around the airfoil leading-edge parabola; establishes the relationship between the abscissa of the stagnation point and the effective angle of attack of the incoming flow; constructs a residual function between the measured pressure and the fitted pressure; uses dynamic pressure sensors distributed on the aircraft wing surface to sense the current gust flow field, while simultaneously fitting the current gust flow field using the airfoil leading-edge parabola flow model; based on the measured and fitted pressures, and using the residual function, employs a nonlinear least squares method to obtain the current combined velocity of the incoming flow and the abscissa of the current stagnation point; thereby calculating the current effective angle of attack of the incoming flow, the horizontal and vertical velocities of the current gust flow, and the frequency of the current gust. This application can measure and sense different cross-sections and unsteady flow parameters of an aircraft wing.

The invention discloses a boundary condition determination method, device and equipment for an aircraft engine inlet and a medium, and relates to the field of hydrodynamics, and the method comprises the steps: collecting the air flow of each grid surface of the engine inlet boundary in the process of carrying out computational fluid dynamics numerical simulation on an aircraft, and determining the current actual flow of the engine inlet; the critical specific flow density is determined based on stagnation point information in the current flow field, and the target speed of the grid surface is determined based on the critical specific flow density, the target flow and the current actual flow; determining the geostatic pressure of the grid surface based on the critical speed and the target speed, and determining the target pressure value of the grid surface based on the judgment result and the geostatic pressure by judging whether the gas flow corresponding to each grid surface belongs to the supersonic speed or not; and determining target density and target velocity component information of the grid surface based on the body center value of the flow field unit so as to determine a target boundary condition. According to the method, the engine inlet boundary condition of the aircraft can be efficiently and accurately determined, and the robustness is improved.

The invention relates to an ultrahigh-enthalpy unbalanced airflow stagnation point enthalpy measuring method which is used for a ground test of an electric arc wind tunnel thermal protection system, and particularly relates to a development-based dual-band laser absorption spectrum measuring method which belongs to the field of aircraft ground aerodynamic thermal test research. Measuring the non-equilibrium temperature of the surface airflow of the water-cooling stagnation point model and the component concentration of atomic nitrogen and oxygen to obtain the internal energy enthalpy and chemical enthalpy of the stagnation point model, thereby realizing the measurement of the superhigh-enthalpy non-equilibrium airflow stagnation point enthalpy; the method provided by the invention can complement the existing method, and fills the gap of enthalpy value measurement in the interval of 50-100MJ / kg in the aerodynamic thermal test; in addition, the method is not limited to measurement of airflow parameters in a test debugging link, and can also realize whole-process flow field parameter monitoring of an aerodynamic heat assessment test of an aircraft thermal protection material.

The invention relates to a coaxial calorimeter heat measurement correction method and system for an arc wind tunnel test, belongs to the field of aerodynamic heat test research, and provides a coaxial heat measurement heat diffusion coefficient temperature correction function alpha (T), and heat flow correction calculation is carried out according to a temperature change curve T (t) along with time and the heat diffusion coefficient temperature correction function alpha (T). Therefore, accurate measurement of the ground test heat flow of the arc wind tunnel can be realized. The coaxial calorimeter heat measurement correction method for the arc wind tunnel test can be suitable for heat measurement of an arc wind tunnel flat plate model test and a stagnation point model test, the heat flow test is very wide in application range and covers 10 kW / m < 2 > to 30000 kW / m < 2 >, and the application range is wide.

The invention belongs to the technical field of aerospace engineering, and particularly relates to a computational fluid mechanics-based aerodynamic thermal environment rapid prediction method, which comprises the following steps of: generating a surface grid according to an aerodynamic configuration; generating a pneumatic grid based on the surface grid, and then carrying out non-viscous flow field calculation based on a CFD method to obtain a flow field variable; searching all surface units according to the characteristic that the stationary point speed is 0, and determining the position of a stationary point; coordinates (x, y, z) are used as independent variables, and an epsilon-curve is obtained through calculation around stationary points; using coordinates (x, y, z) as independent variables, adopting a Gaussian kernel function to realize surface streamline tracking until the surface streamline is intersected with the epsilon-curve, and then calculating along the streamline to obtain a shape factor; and based on an axial symmetry comparison idea, predicting the aerodynamic thermal environment along the surface streamline by using an aerodynamic thermal environment calculation formula. In the method, due to the fact that the hybrid grid and the polyhedral grid are higher in calculation efficiency, the calculation cost can be remarkably reduced when aerodynamic thermal environment prediction is carried out based on the method.

The invention relates to the field of thermal protection material ablation assessment tests, in particular to a method for determining a cold wall catalytic coefficient of a thermal protection material, and the method comprises the steps: S100, calibrating a flow field; s200, preparing a heat-proof material: arranging a coaxial calorimeter at a stationary point of a coating probe, and mounting the coating probe subjected to pre-coating by adopting the heat-proof material on a feeding mechanism; s300, collecting a stationary point response temperature; s400, the cold wall stagnation point heat flux density is calculated; s500, obtaining the mole fractions of the components: obtaining the mole fractions of the components corresponding to the three flow fields by utilizing the calibrated incoming flow total enthalpy and incoming flow total pressure according to an NASA high-temperature air component library; s600, obtaining a stagnation point heat flow; s700, obtaining a stationary point heat flow curved surface diagram; s800, obtaining a matching curve; and S900, obtaining the cold wall catalytic coefficient.

A neural network-based method for predicting heat flux on the wall of a cylinder under hypersonic rarefied flow. In the offline phase, a parameter setting standard is obtained through a Direct Simulation Monte Carlo algorithm verified for heat flux independence. The neural network is then trained using normalized DSMC simulation samples based on this standard. In the online phase, for the problem of flow around a cylinder under hypersonic argon flow, the multi-point heat flux coefficient of the cylinder wall is obtained based on the existing connections within the neural network and the corresponding training parameters for a wide range of known values ​​of the incoming flow Knudsen number, temperature ratio, and Mach number. This method enables rapid and accurate prediction of the heat flux coefficient of the cylinder wall under hypersonic argon flow under a wide range of rarefied flow conditions. Twenty-one test results demonstrate that the stagnation point heat flux error is within 1.5%, and the maximum wall heat flux error does not exceed 4.8%. The calculation speed is close to that of the theoretical formula.

The present application belongs to the technical field of wind tunnel test, and discloses a high-enthalpy shock tunnel flow field diagnosis device and method. The main body of the flow field diagnosis device is a double-layer square tube structure embedded inside and outside. The inner layer center cavity is an inner flow channel, and a 16° wedge is fixed in the inner flow channel. The front section of the outer layer square tube body is provided with an outer baffle, and the rear section is a straight section. The double-layer square tube structure corresponding to the 16° wedge is provided with left-right symmetrical optical windows, and the flow field density change of the 16° wedge is observed through a schlieren system. The front section of the outer layer square tube body is provided with a TDLAS measurement unit, the rear section is provided with a ball head stagnation point heat flow assembly and a pitot pressure assembly. The flow field diagnosis method comprises installing the flow field diagnosis device, performing flow field measurement test, and calculating flow field parameters. The flow field diagnosis device and method are verified by multiple test means, and the free stream parameters of the high-enthalpy shock tunnel are given, thereby providing technical support for developing hypersonic aircraft aerodynamic force and aerodynamic heat characteristic test and air-breathing power characteristic evaluation test.

The invention discloses a gust flow field sensing method based on surface pressure, and relates to the field of gust flow field sensing. According to the method, an airfoil front edge parabola is fitted, and an airfoil front edge parabola streaming model is determined; establishing a relationship between a stationary point position abscissa and an incoming flow effective attack angle; constructing a residual function between the measured pressure intensity and the fitting pressure intensity; dynamic pressure sensors distributed on the surface of an aircraft wing are used for conducting flow sensing on the current gust flow field, and meanwhile, a wing section front edge parabola streaming model is adopted for fitting the flow of the current gust flow field; according to the measurement pressure intensity and the fitting pressure intensity, based on a residual function, a nonlinear least square method is adopted, and the current incoming flow resultant velocity and the current stationary point position abscissa are obtained; and calculating the effective attack angle of the current incoming flow, the horizontal speed and the vertical speed of the current gust incoming flow and the frequency of the current gust. According to the invention, different sections and unsteady flow parameters of aircraft wings can be measured and sensed.