Instability Waves In Boundary Layers Research Articles

Interaction of second-mode disturbances with an incipiently separated compression-corner flow

Abstract

Characterization of Freestream Disturbances in Conventional Hypersonic Wind Tunnels

Although low-disturbance (“quiet”) hypersonic wind tunnels are believed to provide more reliable extrapolation of boundary-layer transition behavior from ground to flight, the presently available quiet facilities are limited to Mach 6, moderate Reynolds numbers, low freestream enthalpy, and subscale models. As a result, only conventional (“noisy”) wind tunnels can reproduce both Reynolds numbers and enthalpies of hypersonic flight configurations and must therefore be used for flight vehicle test and evaluation involving high Mach number, high enthalpy, and larger models. This paper outlines the recent progress and achievements in the characterization of tunnel noise that have resulted from the coordinated effort within the AVT-240 specialists group on hypersonic boundary-layer transition prediction. The new experimental measurements cover a range of conventional wind tunnels with different sizes and Mach numbers from 6 to 14 and extend the database of freestream fluctuations within the spectral range of boundary-layer instability waves over commonly tested models. New direct numerical simulation datasets elucidate the physics of noise generation inside the turbulent nozzle wall boundary layer, characterize the spatiotemporal structure of the freestream noise, and account for the propagation and transfer of the freestream disturbances to a Pitot-mounted sensor.

Experimental investigation of an aeroacoustic feedback mechanism on a two-dimensional side mirror model

The tonal self-noise emission of a two-dimensional blunt-body side mirror model is investigated experimentally in order to characterize the underlying source mechanism. The two-dimensional side mirror design is closely related to an automotive side mirror model11Early development state of an AUDI A7 side mirror. which generates tonal self-noise. The examined flow field and tonal noise emission of the two-dimensional side mirror indicates the existence of an aeroacoustic feedback mechanism comprising convective boundary layer instability waves, their scattering to acoustic waves at the trailing edge and the acoustic receptivity at a dominant feedback point. Fundamental evidence of this concept is found in the typical variation of the discrete-frequency noise emission with the free-stream velocity as well as a laminar flow separation in the vicinity of the trailing edge featuring massively amplified discrete-frequency modes. The existence of an aeroacoustic feedback loop is finally confirmed by the systematic variation of the feedback loop resonance length, which is a crucial loop parameter.

The effect of a cavity on airfoil tones

Abstract The presence of a cavity in the pressure surface of an airfoil has been found via experiment to play a role in the production of airfoil tones, which was attributed to the presence of an acoustic feedback loop. The cavity length was sufficiently small that cavity oscillation modes did not occur for most of the investigated chord-based Reynolds number range of 70,000–320,000. The airfoil tonal noise frequencies varied as the position of the cavity was moved along a parallel section at the airfoil's maximum thickness: specifically, for a given velocity, the frequency spacing of the tones was inversely proportional to the geometric distance between the cavity and the trailing edge. The boundary layer instability waves considered responsible for the airfoil tones were only detected downstream of the cavity. This may be the first experimental verification of these aspects of the feedback loop model for airfoil tonal noise.

Direct simulations of trailing-edge noise generation from two-dimensional airfoils at low Reynolds numbers

The aeroacoustic sound generated from the flow around two NACA four-digit airfoils is investigated numerically, at relatively low Reynolds numbers that do not prompt boundary-layer transition. By using high-order finite-difference schemes to discretize compressible Navier–Stokes equations, the sound scattered on airfoil surface is directly resolved as an unsteady pressure fluctuation. As the wavelength of an emitted noise is shortened compared to the airfoil chord, the diffraction effect on non-compact chord length appears more noticeable, developing multiple lobes in directivity. The instability mechanism that produces sound sources, or unsteady vortical motions, is quantitatively examined, also by using a linear stability theory. While the evidence of boundary-layer instability waves is captured in the present result, the most amplified frequency in the boundary shear layer does not necessarily agree with the primary frequency of a trailing-edge noise, when wake instability is dominant in laminar flow. This contradicts the observation of other trailing-edge noise studies at higher Reynolds numbers. However, via acoustic disturbances, the boundary-layer instability may become more significant, through the resonance with the wake instability, excited by increasing a base-flow Mach number. Evidence suggests that this would correspond to the onset of an acoustic feedback loop. The wake-flow frequencies derived by an absolute-instability analysis are compared with the frequencies realized in flow simulations, to clarify the effect of an acoustic feedback mechanism, at a low Reynolds number.

Vortex Shedding and Noise Radiation from a Slat Trailing Edge

Vortex shedding and the associated noise radiation from a trailing edge were experimentally investigated for a leading-edge slat of a multi-element airfoil at stowed chord Reynolds number Re 2.1 x 10 5 , however, acoustic feedback becomes pronounced between the trailing-edge noise and the boundary-layer instability waves on the suction surface, so that multiple spectral peaks appear both in the velocity fluctuations and sound pressure. At and around the Reynolds number for the first appearance of tonal noise, Re = 1.9 x 10 5 , both of the instability modes coexist. Beyond Re = 2.1 x 10 5 , the boundary-layer instability waves excited by the acoustic feedback evolve into high-intensity vortices before reaching the trailing edge and suppress the absolute instability of the wake through diminishing the reversed-flow region in the wake.

On the Generation of Discrete Frequency Tones by the Flow around an Aerofoil

Tonal noise, the self-induced discrete frequency noise generated by aerofoils, is investigated. It is heard from an aerofoil placed in streams at low Mach number flows when inclined at a small angle to the stream. It is a piercing whistle, commonly up to 30 dB above the background noise level. Previous authors have attributed tonal noise to a feedback loop consisting of a coupling between laminar boundary-layer instability waves and sound waves propagating in the free stream. Boundary-layer measurements have shown the presence of tonal noise is closely related to the existence of a region of separated flow close to the trailing edge of the aerofoil. An analysis of the linear stability of the boundary-layer flow over the aerofoil is presented. The amplification of the instability waves is shown to be controlled almost entirely by the region of separated flow close to the trailing edge. In light of these new experimental and theoretical results the suitability of the aero-acoustic feedback model is discussed.

2:1 resonance for modulated boundary-layer instability waves: New experimental results showing quadratic breakdown.

2:1 resonance for modulated boundary-layer instability waves: New experimental results showing quadratic breakdown.

Application of the Spectral Multidomain Method to the Navier-Stokes Equations

An application of the spectral multidomain method to the two-dimensional, time-dependent, incompressible Navier-Stokes equations is presented. The governing equations are discretized on a non-staggered, stretched mesh with a mixed finite difference/Chebyshev method and are integrated by a time-splitting procedure. The accuracy of the multidomain method is demonstrated for two different channel flow configurations, one with a backward-facing step and the other with a rectangular step. The method is then applied to the investigation of the effects of an isolated, two-dimensional roughness element on the spatial development of instability waves in boundary layers.