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Abstract

A method is provided for detecting flow bifurcation during sonic flow conditions to identify shock location. The method includes setting the initial test conditions, either by setting up an altitude and airspeed in a test aircraft or by setting up an initial flow condition in a wind tunnel. The method for detecting flow bifurcation during sonic flow conditions to identify shock location requires an array of hot films be attached to an aerodynamic surface where shock is expected to impinge. These hot-films are connected to a bank of Constant Voltage Anemometers (CVA), one CVA for each hot film. After reaching the altitude where the testing is to be done, (or attaining the reference condition in a wind tunnel), a pilot initiates an auto-initializing sequence before the onset of a shock, typically at a low speed condition. This sets the current in each sensor to identical values and sets the sensor temperature above ambient temperature of the reference conditions. An auto zero sequence is then initiated to provide zero output current and create a reference point before the onset of the shock. A shock is then generated by linear acceleration of by angle of attack changes. After generating the shock, voltage output of the hot film sensors and the temperatures of the hot films are evaluated. The location of the shock is shown by the hot-film having the highest temperature.

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