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Instability of the three-dimensional boundary layer on a yawed circular cylinder placed in a uniform flow is investigated experimentally by introducing acoustic disturbances from a point near the attachment line. To exemplify the flow dominated by streamline-curvature instability rather than crossflow instability, which has been often observed in many swept-wing flows, is the aim here. In upstream regions of the disturbance wedge originating from the point source, both streamline-curvature and crossflow disturbances are superposed on each other and yield complicated amplitude distributions. A newly proposed method enables the decomposition of the distorted amplitude distribution into contributions from the two instability modes. Detailed observations, however, show that the crossflow mode decays with the distance from the source much faster than the streamline-curvature mode and allows the latter to be dominant in a region further downstream. A fundamental characteristic of the streamline-curvature instability wave is confirmed by examining its phase distribution in the spanwise and normal directions. Wave numbers and spatial growth rates are in good agreement with theoretical predictions. [ABSTRACT FROM AUTHOR]

Flight testing of an unmanned and scaled supersonic experimental airplane NEXST-1 is performed by the Japan Aerospace Exploration Agency to validate computational-fluid-dynamics-based aerodynamic design technology. This is the first attempt to apply the natural laminar flow wing concept to a supersonic vehicle. The concept of the natural laminar flow wing is validated by measuring the surface pressure and the transition locations of the boundary layer. In this paper, the results of the transition measurement are introduced and compared with the numerically predicted results. The transition locations detected experimentally are in good agreement with the predicted locations, and the natural laminar flow effect is confirmed in the aerodynamic design conditions of the supersonic experimental airplane NEXST-1. [ABSTRACT FROM AUTHOR]

Experimental investigation of the boundary-layer transition on an axisymmetric nose model at 1- and 2-deg incidence was conducted at Mach 1.2. The configuration of the model is the forward part of a Sears-Haack body defined to have minimum wave drag caused by volume at 0-deg incidence in supersonic flow. Transition locations were obtained with small surface roughness on the order of 0.1 μ using an infrared camera. An unsteady pressure transducer was applied to investigate instability mechanisms that lead to transition. These results show that the most aft transition occurred on the leeward ray rather than on the windward rays as observed on sharp cones. Tollmien-Schlichting instability dominated the transition process on the windward ray, and crossflow instability was assumed to dominate on the side similar to the transition on the sharp cones. However, transition occurred more aft on the leeward ray than for the case of the sharp cones, and the transition front was determined by turbulent wedges formed as a consequence of the more forward transition on the side. The disturbance believed to be a traveling crossflow wave was observed at 1-deg incidence, and its measured frequency was in good agreement with the maximum amplified frequency, which was calculated using a compressible linear eN code. [ABSTRACT FROM AUTHOR]

Focuses on a study which detailed the measurement of transition characteristics of a supersonic natural laminar flow wing using hot-film sensors and an infrared camera at Mach 2. Vicinity in which the pressure falls very rapidly; Reason for conducting tests in two wind tunnels; Definition of transition location.