11 results about "Stagnation point" patented technology

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 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 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.