Transition in swept wing subsonic PhD Candidate: Jacopo ... · Serpieri, Jacopo Author: sbhoewar...
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Transition in swept wing subsonic flows and active control for drag reduction Motivation Modern jet aircrafts commonly use swept wings. This wing geometry features peculiar changes to the flow configuration when compared to straight wings. One of these differences relates to the mechanism that causes the flow to undergo transition from laminar to turbulent regimes. As known, turbulent flows generate higher surface friction with respect to laminar flows. Friction means drag. Drag means fuel consumption which in the end turns into operating costs and pollution. It is very important for the aeronautical community to achieve a better understanding of the phenomena causing transition on swept wings. The aim of this research project is to investigate these mechanisms and to operate advanced active techniques for flow control never deployed to date. Figure 3. Oil flow visualization of stationary CF waves and transition wedges on a model of swept wing at Re= 1.3e6 and α=-10°. Introduction The main cause of transition from laminar to turbulent regimes in swept wing subsonic flows is the crossflow instability (CF). This is the component of the flow parallel to leading edge line. This component is constant along the span and the chord for wings that behave as infinite swept wings and causes a strong distortion of the flow, which becomes three dimensional. The line traced by a particle of flow close to the wing, but still in the inviscid region, projected on the surface, will indeed be a curve. When looking at the boundary layer configuration two velocity profiles are present. To analyze them it is worth to stay in the reference system of a streamline where x t is the tangent direction, z t the normal one and y t the normal-to- wall coordinate. In this system, while the x t profile will reach the outer velocity, the z t velocity component will become null at the boundary layer edge (in fact the streamline is the actual path of the flow) thus leading to an inflection point as depicted in figure 2. The resulting boundary layer is highly unstable and might undergo transition. The CF modes, in low free stream turbulence flows, appear as stationary waves along the span. Figure 1 is a contour plot of these vortices as seen from the oncoming flow direction in a plane which as the wing- span as abscissa and the normal-to-wall direction as vertical axis for a given chordwise position. In a preliminary experimental campaign a 45° swept wing model has been installed in the LTT wind tunnel at the Low Speed Lab of the TU Delft. Some oil flow visualization was deployed to visualize the CF modes and their effect on transition (see figure 3). Furthermore free-stream measurements were performed to assess the quality and the homogeneity of the flow in the LTT facility. PhD Candidate: Jacopo Serpieri Department: AWEP Section: Aerodynamics Supervisor: Dr. M. Kotsonis Promoter: Prof. Dr. F. Scarano Start date: 1-09-2013 Funding: STW (VENI Grant) Cooperations: Fokker - DNW Type: □ Scientific Aerospace Engineering Figure 4. CAD of the wing model installed in the LTT testing chamber with contoured liners (red). Figure 2. Velocity profiles in the streamline coordinate system indicating the distortion effect due to the CF, from Saric et al. (2003). WIND Cross-Flow stationary waves Transition Wedges Laminar Region Turbulent Region Figure 1. Velocity contours describing CF stationary waves, from Saric et al. (2003). Research methodology 1. Experimental Approach The goal of this research is to investigate the CF instability. The CF waves will be investigated in all the phases: their generation, their growth, their saturation. Furthermore also the effects of boundary conditions, such as surface roughness and free stream turbulence, on their development will be studied by means of experimental campaigns deploying Hotwire Anemometry, particle image velocimetry (PIV) and oil flow visualization. The experimental campaigns will be carried in the LTT at TU Delft: a subsonic, closed circuit, low turbulence wind tunnel. The testing chamber of this tunnel has a cross section of 125 x 180 cm 2 and a length of 260cm. The model will have to feature several requirements: • Extended favourable pressure gradient along the airfoil chord (to avoid Tollmien-Schlichting instability). • Small curvature radius at the leading edge (to avoid attachment line instability). • No concave shape (to avoid Görtler instability). Once that all the other sources of instability are avoided the only source of instability will be the CF modes. A specific airfoil has been designed to fit all these requirements, modifying the NACA 66-018 geometry. To reproduce infinite swept wing flows two wall liners, shaped as the potential flow streamlines, have been designed and will be installed (see figure 4). Research methodology 2. Theoretical Approach Linear stability theory (LST) and parabolized stability equations (PSE) can be used to predict the behavior of the flow around the experimental model. 3. Flow Control Attempts to control this source of instability will pursued by means of both passive (distributed and discrete controlled roughness) and active (plasma actuators) techniques (see figure 5). Figure 5. Schematic of a Plasma Actuator, from www.temasek-lab.nus.edu. Research Partners Funding Institution
Transcript of Transition in swept wing subsonic PhD Candidate: Jacopo ... · Serpieri, Jacopo Author: sbhoewar...
Transition in swept wing subsonic
flows and active control for drag
reduction
Motivation
Modern jet aircrafts commonly use swept wings. This winggeometry features peculiar changes to the flowconfiguration when compared to straight wings. One ofthese differences relates to the mechanism that causes theflow to undergo transition from laminar to turbulentregimes.
As known, turbulent flows generate higher surface frictionwith respect to laminar flows. Friction means drag. Dragmeans fuel consumption which in the end turns intooperating costs and pollution.
It is very important for the aeronautical community toachieve a better understanding of the phenomena causingtransition on swept wings.
The aim of this research project is to investigate thesemechanisms and to operate advanced active techniques forflow control never deployed to date.
Figure 3. Oil flow visualization of stationary CF waves and transition wedges on a model of swept wing at Re= 1.3e6 andα=-10°.
Introduction
The main cause of transition from laminar to turbulent regimes in swept wing subsonic flows is the crossflow instability (CF). This is the component of the flow parallel to leading edge line. This component is constant along the span and the chord for wings that behave as infinite swept wings and causes a strong distortion of the flow, which becomes three dimensional. The line traced by a particle of flow close to the wing, but still in the inviscid region, projected on the surface, will indeed be a curve.
When looking at the boundary layer configuration two velocity profiles are present. To analyze them it is worth to stay in the reference system of a streamline where xt is the tangent direction, zt the normal one and yt the normal-to-wall coordinate. In this system, while the xt profile will reach the outer velocity, the zt velocity component will become null at the boundary layer edge (in fact the streamline is the actual path of the flow) thus leading to an inflection point as depicted in figure 2. The resulting boundary layer is highly unstable and might undergo transition.
The CF modes, in low free stream turbulence flows, appearas stationary waves along the span.
Figure 1 is a contour plot of these vortices as seen fromthe oncoming flow direction in a plane which as the wing-span as abscissa and the normal-to-wall direction asvertical axis for a given chordwise position.
In a preliminary experimental campaign a 45° swept wingmodel has been installed in the LTT wind tunnel at the LowSpeed Lab of the TU Delft. Some oil flow visualization wasdeployed to visualize the CF modes and their effect ontransition (see figure 3).
Furthermore free-stream measurements were performed toassess the quality and the homogeneity of the flow in theLTT facility.
PhD Candidate: Jacopo SerpieriDepartment: AWEPSection: AerodynamicsSupervisor: Dr. M. KotsonisPromoter: Prof. Dr. F. ScaranoStart date: 1-09-2013Funding: STW (VENI Grant)Cooperations: Fokker - DNWType: □ Scientific
Aero
space
Engin
eering
Figure 4. CAD of the wing model installed in the LTTtesting chamber with contoured liners (red).
Figure 2. Velocity profiles inthe streamline coordinate
system indicating thedistortion effect due to the
CF, from Saric et al. (2003).
WIND
Cross-Flow
stationary waves
TransitionWedges
Laminar Region
Turbulent Region
Figure 1. Velocity contours describing CF stationarywaves, from Saric et al. (2003).
Research methodology
1. Experimental Approach
The goal of this research is to investigate the CF instability.
The CF waves will be investigated in all the phases: their generation, their growth, their saturation. Furthermore also the effects of boundary conditions, such as surface roughness and free stream turbulence, on their development will be studied by means of experimental campaigns deploying Hotwire Anemometry, particle image velocimetry (PIV) and oil flow visualization.
The experimental campaigns will be carried in the LTT at TU Delft: a subsonic, closed circuit, low turbulence wind tunnel. The testing chamber of this tunnel has a cross section of 125 x 180 cm2 and a length of 260cm.
The model will have to feature several requirements:
• Extended favourable pressure gradient along the airfoil chord (to avoid Tollmien-Schlichting instability).
• Small curvature radius at the leading edge (to avoid attachment line instability).
• No concave shape (to avoid Görtler instability).
Once that all the other sources of instability are avoided the only source of instability will be the CF modes.
A specific airfoil has been designed to fit all these requirements, modifying the NACA 66-018 geometry.
To reproduce infinite swept wing flows two wall liners, shaped as the potential flow streamlines, have been designed and will be installed (see figure 4).
Research methodology
2. Theoretical Approach
Linear stability theory (LST) and parabolized stability equations (PSE) can be used to predict the behavior of the flow around the experimental model.
3. Flow Control
Attempts to control this source of instability will pursued by means of both passive (distributed and discrete controlled roughness) and active (plasma actuators) techniques (see figure 5).
Figure 5. Schematic of a Plasma Actuator, fromwww.temasek-lab.nus.edu.
Research Partners Funding Institution
