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.

Abstract: Arrays of microslats operating within the boundary layer of an aircraft wing or other surfaces are provided. The arrays consist of rows of microslats, staggered between rows, and shaped so that the trailing edge of the microslat is normal or perpendicular to the direction of expected reverse flow on the surface. For flow regions having directionally stable flow, the microslats are rectangular. Where the direction of the expected reverse flow is variable, the trailing edge of the microslats is triangular, thereby accepting an angular range of reverse flow. Each microslat operates independently through a hinge on its leading edge and an individual spring which holds the microslat flush to the aerodynamic surface until actuated by reverse flow. Where the boundary layer is expected to thicken, i.e., downstream along a surface, the arrays are formed with two layers of microslats, smaller microslats in a top layer and larger in a bottom layer.

Abstract: A broad-range, multi-directional aircraft airspeed measuring system is provided. The airspeed measuring system has multiple vortex generating probes located within a venturi section. At least one rearward facing probe and one forward facing probe are included. Additional probes can be added to extend the high speed range of the airspeed indicator. A splitter plate or plates separate flow channels from each other to provide a separate flow channel for each vortex probe. Each vortex probe has a hot film sensor and anemometer.