7 results about "Stagnation point" patented technology

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.

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.

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.