“…It should be noted that excitation of several modes of the secondary instability in a three-dimensional supersonic boundary layer with frequencies of about 1 MHz was theoretically predicted in [20,22]. Available hot-wire anemometers cannot measure oscillations with such frequencies, but the presence of the secondary instability mode with the maximum growth at 100 kHz at M = 3.5, which was predicted in [20], is consistent with the experimental data obtained at M = 3.5 [36].…”
Section: Experimental Results and Their Analysissupporting
Results of experiments aimed at studying the linear and nonlinear stages of the development of natural disturbances in the boundary layer on a swept wing at supersonic velocities are presented. The experiments are performed on a swept wing model with a lens-shaped airfoil, leadingedge sweep angle of 45 • , and relative thickness of 3%. The disturbances in the flow are recorded by a constant-temperature hot-wire anemometer. For determining the nonlinear interaction of disturbances, the kurtosis and skewness are estimated for experimentally obtained distributions of the oscillating signal over the streamwise coordinate or along the normal to the surface. The disturbances are found to increase in the frequency range from 8 to 35 kHz in the region of their linear development, whereas enhancement of high-frequency disturbances is observed in the region of their nonlinear evolution. It is demonstrated that the growth of disturbances in the high-frequency spectral range (f > 35 kHz) is caused by the secondary instability.
“…It should be noted that excitation of several modes of the secondary instability in a three-dimensional supersonic boundary layer with frequencies of about 1 MHz was theoretically predicted in [20,22]. Available hot-wire anemometers cannot measure oscillations with such frequencies, but the presence of the secondary instability mode with the maximum growth at 100 kHz at M = 3.5, which was predicted in [20], is consistent with the experimental data obtained at M = 3.5 [36].…”
Section: Experimental Results and Their Analysissupporting
Results of experiments aimed at studying the linear and nonlinear stages of the development of natural disturbances in the boundary layer on a swept wing at supersonic velocities are presented. The experiments are performed on a swept wing model with a lens-shaped airfoil, leadingedge sweep angle of 45 • , and relative thickness of 3%. The disturbances in the flow are recorded by a constant-temperature hot-wire anemometer. For determining the nonlinear interaction of disturbances, the kurtosis and skewness are estimated for experimentally obtained distributions of the oscillating signal over the streamwise coordinate or along the normal to the surface. The disturbances are found to increase in the frequency range from 8 to 35 kHz in the region of their linear development, whereas enhancement of high-frequency disturbances is observed in the region of their nonlinear evolution. It is demonstrated that the growth of disturbances in the high-frequency spectral range (f > 35 kHz) is caused by the secondary instability.
“…18 This research is a continuation of a series of experimental studies of the origin of turbulence on the swept wing (Figure 1) in a supersonic flow. 14,16,17,19,20 The results of these studies will be useful in the creation of modern engineering methods for predicting the position of the laminar-turbulent transition. Figure 1.Swept wing model: (a) top view; (b) aerofoil.…”
Section: Introductionmentioning
confidence: 99%
“…These data will make it possible to carry out a correct comparison with the results of calculations on the linear theory of stability and on direct numerical simulation where experimental conditions are modeled, and to solve the problem of the applicability of calculation methods for such a complex flow.The results of the first hot-wire measurements in a supersonic boundary layer on a swept wing were presented in Ermolaev et al 18 This research is a continuation of a series of experimental studies of the origin of turbulence on the swept wing ( Figure 1) in a supersonic flow. 14,16,17,19,20 The results of these studies will be useful in the creation of modern engineering methods for predicting the position of the laminar-turbulent transition. leading edge bluntness is about 0.4 mm.…”
The paper is devoted to an experimental and theoretical study of effect of moderate angle-of-attack variation on disturbances evolution and laminar-turbulent transition in a supersonic boundary layer on swept wing at Mach 2. Monotonous growth of the transition Reynolds numbers with angle of attack increasing from −2° to 2.7° is confirmed. For the same conditions, calculations based on linear stability theory are performed. The experimental and computational results show a favourable comparison.
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