Cross-Flow Instability: Flow diagnostics and control of swept wing boundary layers

The research presented in this booklet focusses on the cross-flow instability. Applying traditional and advanced flow diagnostics, the boundary layer evolution is studied in detail. The topology and evolution of both primary and secondary instability mechanisms is revealed with unprecedented detail for experimental research paving the way for new advanced-diagnostics investigations. Important confirmations of the outcomes of past experimental, numerical and theoretical studies are achieved together with the description of a newly-reported flow phenomenon. The latter consists of a low frequency motion of the "stationary" primary vortices. While this phenomenon is considered not relevant for the transition evolution, it is deemed important for experimental investigations as it encompasses very high levels of turbulent kinetic energy.Advanced flow control experiments based on alternating current dielectric barrier discharge plasma actuators are also performed following different instability control approaches. The primary instability is conditioned by the external forcing either in the wavenumber spectrum (by inducing selected spanwise modes) or in intensity (by weakening or enhancing the cross-flow velocity). The secondary instability modes are conditioned in the frequency spectrum and phase.These efforts achieved the intended scopes. Although, when selected stationary modes were forced, the boundary layer fluctuations were enhanced. These fluctuations can directly cause the turbulent breakdown vanishing the beneficial effect of the performed instability control. The cross-flow forcing, making use of newer actuators reaching higher frequencies, resulted successful yielding transition promotion or delay depending on the forcing direction.Aerodynamic

Three-dimensional organisation of primary and secondary crossflow instability

An experimental investigation of primary and secondary crossflow instability developing in the boundary layer of a 45° swept wing at a chord Reynolds number of 2.17 × 106 is presented. Linear stability theory is applied for preliminary estimation of the flow stability while surface flow visualisation using fluorescent oil is employed to inspect the topological features of the transition region. Hot-wire anemometry is extensively used for the investigation of the developing boundary layer and identification of the statistical and spectral characteristics of the instability modes. Primary stationary, as well as unsteady type-I (z-mode), type-II (y-mode) and type-III modes are detected and quantified. Finally, three-component, three-dimensional measurements of the transitional boundary layer are performed using tomographic particle image velocimetry. This research presents the first application of an optical experimental technique for this type of flow. Among the optical techniques, tomographic velocimetry represents, to date, the most advanced approach allowing the investigation of spatially correlated flow structures in three-dimensional fields. Proper orthogonal decomposition (POD) analysis of the captured flow fields is applied to this goal. The first POD mode features a newly reported structure related to low-frequency oscillatory motion of the stationary vortices along the spanwise direction. The cause of this phenomenon is only conjectured. Its effect on transition is considered negligible but, given the related high energy level, it needs to be accounted for in experimental investigations. Secondary instability mechanisms are captured as well. The type-III mode corresponds to low-frequency primary travelling crossflow waves interacting with the stationary ones. It appears in the inner upwelling region of the stationary crossflow vortices and is characterised by elongated structures approximately aligned with the axis of the stationary waves. The type-I secondary instability consists instead of significantly inclined structures located at the outer upwelling region of the stationary vortices. The much narrower wavelength and higher advection velocity of these structures correlate with the higher-frequency content of this mode. The results of the investigation of both primary and secondary instability from the exploited techniques agree with and complement each other and are in line with existing literature. Finally, they present the first experimental observation of the secondary instability structures under natural flow conditions.Aerodynamic

Conditioning of unsteady cross-flow instability modes using dielectric barrier discharge plasma actuators

In this study, experiments are performed towards the identification and measurement of unsteady modes occurring in a transitional swept wing boundary layer. These modes are generated by the interaction between the primary stationary and travelling cross-flow instabilities or by secondary instability mechanisms of the stationary cross-flow vortices and have a crucial role in the laminar-to-turbulent breakdown process. Detailed hot-wire measurements were performed at the location of stationary instability amplitude-saturation. In order to deterministically capture the spatio-temporal evolution of the unsteady modes, measurements were phase- and frequency-conditioned using concurrent forcing by means of a dielectric barrier discharge plasma actuator mounted upstream of the measurement domain. The actuator effect, when positioned sufficiently upstream the secondary modes onset, was tuned such to successfully condition the high-frequency type-I and the low-frequency type-III modes without modifying the transition evolution. Two primary stationary cross-flow vortices of different amplitude were measured, revealing the effect of base-flow variations on the growth of travelling instabilities. The response of these two stationary waves to the naturally occurring and forced fluctuations was captured at different chordwise positions. Additionally, the deterministic conditioning of the instability phase to the phase of the actuation allowed phase-averaged reconstruction of the spatio-temporal evolution of the unsteady structures providing valuable insight on their topology. Finally, the effect of locating the actuator at a more downstream position, closer to the type-I mode branch-I, resulted in laminar-to turbulent breakdown for the high-frequency actuation while the low-frequency forcing showed milder effects on the transition evolution.Aerodynamic

Flow visualization of swept wing boundary layer transition

In this work the flow visualization of the transition pattern occurring on a swept wing in a subsonic flow is presented. This is done by means of fluorescent oil flow technique and boundary layer hot-wire scans. The experiment was performed at Reynolds number of 2:15 . 106 and at angle of attack of -3º. At these conditions, three different flows are investigated: a natural transition case and two other ones where instead the transition mechanism, the stationary cross-flow waves, was forced with discrete roughness elements. Previously published results on similar flows were confirmed for two of the three tested configurations (natural transition and boundary layer forcing at the wavelength of the dominant stationary mode) while, for the third one (where a sub-critical wavelength was forced), some discrepancies are observed. A parametric study on the effect of the Reynolds number and the angle of attack is also presented.Aerospace Structures & Computational MechanicsAerospace Engineerin

Time-resolved PIV investigation of the secondary instability of cross-flow vortices

Time-resolved PIV measurements of the secondary instability modes of cross-flow vortices are presented. Measurements are performed on a large scale 45o swept wing at chord Reynolds number of 2.17 million in a low turbulence wind-tunnel facility. Using acquisition frequencies of 20 kHz, the present study is the first experimental demonstration of spatio-temporally resolved measurements of these structures. Statistical and spectral analysis reveals a fluctuating velocity field, strongly conditioned in space by the primary stationary cross-flow vortex. The flow structures related to the type-I high-frequency instability and type-III are captured by Proper Orthogonal Decomposition of the instantaneous flow-fields. Their temporal evolution is analysed showing good agreement with previous studies thus confirming that POD is correctly representing the flow structures of the relevant instability modes. The low frequency meandering oscillation of the stationary vortices, first reported by Serpieri &Kotsonis (2016b), is observed and characterised.Aerodynamic

High-speed PIV analysis of trailing edge aeroacoustics

Tonal noise generated by airfoils observed at low to moderate Reynolds numbers is related to laminar boundary layer instabilities, which has lead to the term laminar boundary layer instability noise. The particular features of the acoustic spectrum have been discussed and a number of theories have been proposed in literature over the past 50 years. Previous research suggests that the appearance of tonal noise is related to a feedback between the acoustic waves scattered at the trailing edge and the receptive part of the boundary layer [4]. Reported studies have been performed on the basis of hot-wire anemometry, laser doppler velocimetry, phase-locked PIV, acoustic measurements, numerical simulations or theoretical models. In recent years, PIV has become an alternative for the investigation of aeroacoustic sources. In particular the aeroacoustic sources for trailing edge noise have been investigated by Schröder et al. [10] using time-resolved PIV and Shannon and Morris [11] based on phase-locked PIV. Nakano et al. [8] focussed on the subject of laminar boundary layer instability noise on an airfoil and related their occurrence on the pressure side to noise emissions based on a correlation based technique. In the present study, planar high-speed PIV is performed simultaneously with acoustic far-field measurements. This combination allows to associate features of the acoustic emissions to events in the source field near the trailing edge for a better understanding of the tonal noise generation on an airfoil. In the past a wide range of Reynolds numbers has been investigated for different airfoil models, most notably the NACA0012 which is also selected for the present study. In particular, it is found that a periodic amplitude modulation of rapidly growing instabilities on the pressure side of the airfoil is responsible for the occurrence of multiple tones for the present configuration.Aerodynamics, Wind Energy and PropulsionAerospace Engineerin

Conditioning of cross-flow instability modes using dielectric barrier discharge plasma actuators

In the current study, selective forcing of cross-flow instability modes evolving on a swept wing at is achieved by means of spanwise-modulated plasma actuators, positioned near the leading edge. In the perspective of laminar flow control, the followed methodology holds on the discrete roughness elements/upstream flow deformation (DRE/UFD) approach, thoroughly investigated by e.g. Saric et al. (AIAA Paper 1998-781, 1998), Malik et al. (J. Fluid Mech., vol. 399, 1999, pp. 85-115) and Wassermann & Kloker (J. Fluid Mech., vol. 456, 2002, pp. 49-84). The possibility of using active devices for UFD provides several advantages over passive means, allowing for a wider range of operating numbers and pressure distributions. In the present work, customised alternating current dielectric barrier discharge plasma actuators have been designed, manufactured and characterised. The authority of the actuators in forcing monochromatic stationary cross-flow modes at different spanwise wavelengths is assessed by means of infrared thermography. Moreover, quantitative spatio-temporal measurements of the boundary layer velocity field are performed using time-resolved particle image velocimetry. The results reveal distinct steady and unsteady forcing contributions of the plasma actuator on the boundary layer. It is shown that the actuators introduce unsteady fluctuations in the boundary layer, amplifying at frequencies significantly lower than the actuation frequency. In line with the DRE/UFD strategy, forcing a sub-critical stationary mode, with a shorter wavelength compared to the naturally selected mode, results in less amplified primary vortices and related fluctuations, compared to the critical forcing case. The effect of the forcing on the flow stability is further inspected by combining the measured actuators body force with the numerical solution of the laminar boundary layer and linear stability theory. The simplified methodology yields fast and computationally cheap estimates on the effect of steady forcing (magnitude and direction) on the boundary layer stability.Aerodynamic

High-speed PIV analysis of trailing edge aeroacoustics

Tonal noise generated by airfoils observed at low to moderate Reynolds numbers is related to laminar boundary layer instabilities, which has lead to the term laminar boundary layer instability noise. The particular features of the acoustic spectrum have been discussed and a number of theories have been proposed in literature over the past 50 years. Previous research suggests that the appearance of tonal noise is related to a feedback between the acoustic waves scattered at the trailing edge and the receptive part of the boundary layer [4]. Reported studies have been performed on the basis of hot-wire anemometry, laser doppler velocimetry, phase-locked PIV, acoustic measurements, numerical simulations or theoretical models. In recent years, PIV has become an alternative for the investigation of aeroacoustic sources. In particular the aeroacoustic sources for trailing edge noise have been investigated by Schröder et al. [10] using time-resolved PIV and Shannon and Morris [11] based on phase-locked PIV. Nakano et al. [8] focussed on the subject of laminar boundary layer instability noise on an airfoil and related their occurrence on the pressure side to noise emissions based on a correlation based technique. In the present study, planar high-speed PIV is performed simultaneously with acoustic far-field measurements. This combination allows to associate features of the acoustic emissions to events in the source field near the trailing edge for a better understanding of the tonal noise generation on an airfoil. In the past a wide range of Reynolds numbers has been investigated for different airfoil models, most notably the NACA0012 which is also selected for the present study. In particular, it is found that a periodic amplitude modulation of rapidly growing instabilities on the pressure side of the airfoil is responsible for the occurrence of multiple tones for the present configuration

Experimental investigation of aerofoil tonal noise generation

The present study investigates the mechanisms associated with tonal noise emission from a NACA 0012 aerofoil at moderate incidence (0; 1; 2 and 4 angle of attack) and with Reynolds numbers ranging from 100 000 to 270 000. Simultaneous time-resolved particle image velocimetry (PIV) of the aeroacoustic source region near the trailing edge and acoustic measurements in the far field are performed in order to establish the correspondence between the flow structure and acoustic emissions. Results of these experiments are presented and analysed in view of past research for a number of selected cases. Characteristics of the acoustic emission and principal features of the average flow field agree with data presented in previous studies on the topic. Time-resolved analysis shows that downstream convecting vortical structures, resulting from growing shear layer instabilities, coherently pass the trailing edge at a frequency equal to that of the dominant tone. Therefore, the scattering of the vortical structures and their associated wall pressure fluctuations are identified as tone generating mechanisms for the cases investigated here. Moreover, wavelet analysis of the acoustic pressure and velocity signals near the trailing edge show a similar periodic amplitude modulation which is associated with multiple tonal peaks in the acoustic spectrum. Periodic amplitude modulation of the acoustic pressure and velocity fluctuations on the pressure side are also observed when transition is forced on the suction side, showing that pressure-side events alone can be the cause.Aerodynamics, Wind Energy & PropulsionAerospace Engineerin

Secondary crossflow instability through global analysis of measured base flows

A combined experimental and numerical approach to the analysis of the secondary stability of realistic swept-wing boundary layers is presented. Global linear stability theory is applied to experimentally measured base flows. These base flows are three-dimensional laminar boundary layers subject to spanwise distortion due to the presence of primary stationary crossflow vortices. A full three-dimensional description of these flows is accessed through the use of tomographic particle image velocimetry (PIV). The stability analysis solves for the secondary high-frequency modes of type I and type II, ultimately responsible for turbulent breakdown. Several pertinent parameters arising from the application of the proposed methodology are investigated, including the mean flow ensemble size and the measurement domain extent. Extensive use is made of the decomposition of the eigensolutions into the terms of the Reynolds-Orr equation, allowing insight into the production and/or destruction of perturbations from various base flow features. Stability results demonstrate satisfactory convergence with respect to the mean flow ensemble size and are independent of the handling of the exterior of the measurement domain. The Reynolds-Orr analysis reveals a close relationship between the type I and type II instability modes with spanwise and wall-normal gradients of the base flow, respectively. The structural role of the in-plane velocity components in the perturbation growth, topology and sensitivity is identified. Using the developed framework, further insight is gained into the linear growth mechanisms and later stages of transition via the primary and secondary crossflow instabilities. Furthermore, the proposed methodology enables the extension and enhancement of the experimental measurement data to the pertinent instability eigenmodes. The present work is the first demonstration of the use of a measured base flow for stability analysis applied to the swept-wing boundary layer, directly avoiding the modelling of the primary vortices receptivity processes.AerodynamicsEducation A