MORPHING CHARACTERISTICS OF CHIRAL CORE AIRFOILS Alessandro Spadoni, Massimo Ruzzene School of Aerospace Engineering Georgia Institute of Technology Atlanta GA - USA Chrystel Remillat, Fabrizio Scarpa, Kevin Potter Department of Aerospace Engineering University of Bristol Bristol, UK 1
Objectives & Motivation Objective: Application of chiral geometry for the design of an airfoil with morphing characteristics; Motivation: - Morphing is an effective way to enhance performance of wings and rotor blades: improve flow conditions, minimize drag, eliminate the need for flap mechanisms, improve handling and control of aircraft. - A chiral structure provides compliance and allows continuous deformation of airfoil. 2
Chiral Geometry Negative Poisson s ratio: Estimated ν xy ~-.9; High in-plane shear modulus: G xy = E x 2(1 +υ ) xy As a result of ν xy being close to -1; Unique deformation mechanism: Allows large deflections, while material remains in elastic range; Design flexibility: Property of the assembly strongly depends on characteristic parameters of chiral geometry (r,r,l,θ); R t θ r L 3
Previous Work: Racecar wing with passive adaptive capabilities (*) Objective: Passive changes of mean camber line of the wing in response to changing incident airflow speeds Increase in MAXIMUM SPEED and better HANDLING Homogeneous material with homogenized chiral properties 5 mm EPPLER 42 Airfoil with 3 mm chord Reduction of the TAB ANGLE in response to increase in car speed; Corresponding elastic deformations are recovered when the speed decreases, so that the wing tip moves back upward; (*) Bornengo, D., Scarpa, F., Remillat, C. "Morphing airfoil concept with chiral core structure, IMechE J. Aer. Eng., 25. 4
Rationale z y dθ Τ = GJ dx x 2 d w M = EI xx dx 2 Given a Poisson s ratio of -1, G Wing does not require close section to carry torsional loads; Large decambering deformations can be sustained within the elastic range of the constitutive material; The core allows for continuous deformations which are important to maintain aerodynamic efficiency; The airfoil core can be tuned to achieve different functionalities by changing core geometric parameters; 5
Outline Present two designs resulting from parallel developments; Discuss results and provide recommendations for future research; Show related research on chiral structures. GATECH U. Of BRISTOL 6
Development of GATech s chiral airfoil 7
Overview The airfoil hosts a MACROSCOPIC chiral structure; Evaluation of compliance characteristics: Numerical analysis with steady aerodynamic loads; Experimental investigation using static loads. Eppler 42 series Eppler 42 s camber provides significant lift at low speed; Assembly is intended to conform as dictated by flow conditions camber decreases as velocity and lift increase Lift-induced drag decreases with velocity; Configuration can be modified for active morphing applications (active camber control). 8
Design Chiral core is MAPPED into airfoil; Influence of number of cells and L/R ratio is investigated; All other parameters are kept constant. Mapped to conform To airfoil profile Predefined layout Note: Core mapping facilitates meshing; Core periodicity is lost 9
Numerical Analysis CFD-FEA Coupled models Steady-state aerodynamic loads are defined by specified flow conditions; Airfoil performance is investigated through weakly-coupled structural and computational fluid dynamics (CFD) models; Air-loads and corresponding displacements are iteratively passed to the structural and fluid codes respectively until convergence is achieved z -φ 1 v 1 u 1 Timoshenko beam element z Fixed φ 2 y, v 2 Soft skin STRUCTURAL MODEL x,u 2 E = 71 MPa density = 27 Kg/m 3 ν =.33 wall t =.76 mm Out-of-plane t = 2.54 cm v u 2 1 φ 3 u 3 u 1 v 3 Isoparametric planar element E = 71 GPa density = 27 Kg/m 3 ν =.33 wall t =.76 mm Out-of-plane t = 2.54 cm 1
Steady-state Fluid Model Unstructured triangular mesh, using finite-volume Galerkin NSC2KE (*) http://www.inria.fr/rrrt/rt-164.html u r r Wall boundary, n = Tangency condition Minimum element size.1 chord, at leading and trailing edges Wake elements near trailing edge have a size of.1 chord Inflow points Outflow points Characteristic technique (*) Mohammadi, B., "Fluid Dynamics Computation with NSC2KE, " INRIA Report 164, 1994. Airfoil element size is linearly relaxed away from LE and TE by 5 Wake element size linearly relaxed up to outflow boundary Inflow and Outflow boundaries are at least 6 chords away from airfoil Total of approximately 11,5 fluid elements. 11
Steady-state Fluid Model Analysis of convergence of fluid model 1.9 Lift Coefficient 1 Residual norm 1.8 1-1 C l 1.7 1.6 1.5 log 1 (Normalized L 2 Norm) 1-2 1-3 1.4 1.3 1-4 1 2 3 4 5 6 7 8 9 1 Number of Iterations Sea level conditions Mach.45 angle of attack 2 No gravity Euler time stepping (steady-state solution) 1-5 1 2 3 4 5 6 7 8 9 1 Number of Iterations 8 iterations are used, as reasonable residual reduction target is 4 orders of magnitude 12
Structural & CFD iterations L/R Number of core cells Material properties Free-stream conditions Angle of attack Structural mesh of airfoil and chiral core are obtained ANSYS Beginning of convergence iterations MATLAB Initial Iteration Beginning of next Iteration Flow field region is discretized with an unstructured triangularelement mesh L = u T u Flow field is solved for pressure, density and velocity NSC2KE Equilibrium is solved imposing aerodynamic loads on chiral-core airfoil MATLAB L i L i 1 < 1 1 3 Deformed airfoil is splined and new profile is computed MATLAB New Iteration converges with previous one? No Yes Solution 13
Results: 3 cells Influence of L/R ratio.14 TE displacement 1.3 cm.13 Trailing-Edge Displacement [m].12.11.1.9.8.7.6.6.65.7.75.8.85.9.95 1 L/R L/R =.6 L/R =.9 14
Results: 2 cells Influence of L/R ratio.8.7 Trailing-Edge Displacement, [m].6.5.4.3.2.1 3 cells.6.65.7.75.8.85.9.95 1 L/R, [m] L/R =.6 L/R =.95 15
Results: Influence of L/R ratio Large node radius facilitates bending deformation of the ligaments, which is a main contributor of overall deformation L/R =.7 L/R =.9 16
2 nd CHIRAL-CORE AIRFOIL CONFIGURATION Lift vs. Mach number 2 Lift coefficient 2.5 x 14 Lift 1.9 1.8 1.7 L/R =.9 2 L/R =.9 C l 1.6 1.5 L l 1.5 1.4 1.3 L/R =.6 1 L/R =.6 1.2.5 1.1 1.15.2.25.3.35.4.45 M.15.2.25.3.35.4.45 M 17
Experimental Validation Static compliance tests L/R =.6 2-y cells L/R =.6 3-y cells Water-jet manufacturing L/R =.94 3-y cells 18
Experimental Validation Static compliance tests t 2.54 cm r r = 1.7 cm t =.65 mm.36 m.7 m Out-of-plane thickness = 2. cm t 2.54 cm r r =.67 cm t =.65 mm.36 m.7 m Material: Aluminum 661 T651 E = 71 GPa density = 27 Kg/m 3 ν =.33 r t r =.3 cm t =.65 mm.36 m 2.54 cm Out-of-plane thickness and chord dimensions were chosen given manufacturing restrictions.7 m 19
Experimental Validation Static compliance tests Experimental set-up Strain gages Strain-gage locations are chosen based on a preliminary FE analysis LVDT Clamped b.c. Strain gage conditioner and amplifier Vishay Measurement Group 21 System Iotech 2 series acquisition board 2
Static compliance tests L/R =.6, 3 unit cells across airfoil thickness Applied load, [N] 6 5 4 3 2 1 5 1 15 T.E. disp, [mm] P strain gauge 1 μ ε strain gauge 3 μ ε strain gauge 5 μ ε 6 2 1 4 2 1 3 5 5 1 15 T.E. disp, [mm] 6 4 2 2 1 5 1 15 T.E. disp, [mm] 5 1 15 T.E. disp, [mm] Four cuts have been used to core and skin compliance strain gauge 2 μ ε strain gauge 4 μ ε strain gauge 6 μ ε 3 2 1 5 1 15 T.E. disp, [mm] 1 5 5 1 15 T.E. disp, [mm] 1 5-5 5 1 15 T.E. disp, [mm] 21
Static compliance tests L/R =.6, 2 unit cells across airfoil thickness 1 2 3 5 Applied load, [N] 25 2 15 1 5 2 4 6 8 1 12 14 16 18 2 T.E. disp, [mm] P strain gauge 1 μ ε strain gauge 3 μ ε strain gauge 5 μ ε 1 5 15 1 4 5 1 15 2 T.E. disp, [mm] 5 4 2 5 1 15 2 T.E. disp, [mm] 5 1 15 2 T.E. disp, [mm] strain gauge 2 μ ε strain gauge 4 μ ε 15 1 5 1 5 1 15 2 T.E. disp, [mm] 5 5 1 15 2 T.E. disp, [mm] 22
Static compliance tests L/R =.6, 2 unit cells across airfoil thickness Non-linear FE models suggest more compliance within elastic range 9 8 7 Experimental results Numerical results Applied load, [N] 6 5 4 3 2 1 1 2 3 4 5 6 7 8 9 1 T.E. disp, [mm] 23
Static compliance tests L/R =.94, 3 unit cells across airfoil thickness 2 3 5 1 Applied load, [N] 1 9 8 7 6 5 4 3 2 1 2 4 6 8 1 12 T.E. disp, [mm] strain gauge 1 μ ε strain gauge 3 μ ε strain gauge 5 μ ε 2 1 2 1 2 1 4 5 1 15 T.E. disp, [mm] 5 1 15 T.E. disp, [mm] 5 1 15 T.E. disp, [mm] strain gauge 2 μ ε strain gauge 4 μ ε 1 5 1 5 1 15 T.E. disp, [mm] 5 5 1 15 T.E. disp, [mm] 24
Static compliance tests Summary Trailing-Edge displacement, [mm] 2 18 16 14 12 1 8 6 4 3y cells, L/R =.94 2y cells, L/R =.6 3y cells, L/R =.6 2 2 4 6 8 1 Applied Load, [N] The core can be designed to achieve different compliance through a change in a limited number of geometric (L/R); Significant decambering deformations can be sustained within the elastic range of the constitutive material; 25
Development of U. of Bristol s chiral airfoil 26
Design Constraints Selective Laser Sintering for core manufacturing Maximum core length is.2 m, chord.3 m; Minimum thickness is 1mm for any part of the structure: Ligaments thickness is 1 mm; Opted for solid nodes (nodes are stiff compared to the ligaments) Only a two cell deep core would fit into the aerofoil whilst maintaining a reasonable ligament aspect ratio. It was chosen not to have ligaments attached to the aerofoil skin, to avoid high point loads. A custom chiral core was created, joining only nodes onto the skin and following the curvature of the aerofoil. 27
Configuration Skin:.5mm glass fibre composite Flexible but stiff enough to maintain the aerofoil shape and prevent surface buckling. Non-symmetric lay-up [º,9º,+45º,-45º] facilitates conforming to airfoil shape Nose: Stiff, made of pine wood Chiral Honeycomb Core: Material Polyamide Duraform. Design attempts to keep L/r and R/r ratios uniform (uniform E) Node radii (r) are decreased along the chord as the ligament length (L) was constrained by the taper of the aerofoil. - Some of the internal cell angles, θ, differ from 3 - Initial chiral geometry has been used as a guideline. 28 Rubber Strip: Added to induce the chiral rotation(*) (*) Bornengo, D., Scarpa, F., Remillat, C. "Morphing airfoil concept with chiral core structure, IMechE J. Aer. Eng., 25.
Numerical analysis Static Analysis: Structural model is coupled with two flow solvers An iterative process was required to achieve convergence of the solution. Ansys FE is used for the structure part; Structural FE model is validated experimentally Flow solvers: Inviscid vortex lattice panel method coded in Matlab, assumes twodimensional (2-D) inviscid flow, and does not consider flow separation. Viscous (XFoil) combines a vortex panel method with a boundary layer model to provide a viscous analysis. It includes boundary layer growth, producing a more realistic pressure distribution, especially at higher angles. 29
Example of Results Airfoil tip deflection for increasing velocity, at 15º Inviscid and viscous predictions 3. Graph of Tip Vertical Deflection Against Velocity at an Angle of 15º 2.5 Tip Vertical Deflection (mm) 2. 1.5 1..5 Viscous Inviscid. 1 2 3 4 5 6 7 Velocity (m/s) Viscous analysis predicts a much lower deflection: pressure forces acting close to the trailing edge are reduced due to the predicted boundary layer at high angles. 3
Analysis of modified Airfoil Effects of: Higher number of cells; Reduction of ligament thickness (.2 mm) and length (6 mm); Changes produce a 2% increase in tip deflection at 8 m/s. Effect of chiral cell density on tip displacement Tip Displacement (mm) 1 8 6 4 2 5 Cell Deep 2 Cell Deep 2 4 6 8 1 Airspeed (m/s) 31
Aerodynamic testing: Wind tunnel: Maximum speed ~7m/s. Equipped for lift, drag and pitching moment measurements Visual Measurement Technique: The motion of the aerofoil is tracked by a digital camera system, connected to a laptop for real-time processing. The position of a number of targets can be tracked at a rate of 7.5Hz. Two targets could be used at once to monitor the strains. Wing Prototype 32 Prototype mounted in low turbulence wind tunnel
Examples of Results Lift Force against Displacement at Varying Incidence for the Experimental Analysis Deflection vs Speed at varying angles of incidence 35 2.5 3 25 2 15º EXP 1º EXP 5º EXP Lift (N) 2 15 1 15º 1º 5º º Deflection (mm) 1.5 1.5 15º FEA 5..5 1. 1.5 2. 2.5 Vertical Deflection (mm) 1 2 3 4 5 6 7 8 Speed (m/s) Numerical (FEA) Experimental 33
Comments Deflections increase as velocity increases. The relationship is non linear with respect to the lift force on the aerofoil: At low velocities, the rate of increase of deflection is lower than at higher velocities. This could be due to a certain amount of force required to activate the honeycomb deformation mechanism. These forces appear to be different for the varying angles of incidence, This could be explained by the varying pressure distribution around the aerofoil (with respect to angle of incidence), and the anisotropy of the chiral structure (i.e. the main bulk of the pressure force acts on a different area of the aerofoil at different angles, giving a different deflection response). This deflection at 15 incidence corresponds to: Camber change of.3%, A reduction in CL of.5, A reduction of CD of ~.2. Whilst this change is small, it proves that the morphing concept works. 34
Comments The measured strains between nodes 19 and 24, and 21 and 22 were found to be positive; Computed values giving a Poisson s ratio of approx. -.9. e 1 e 2 35
Future developments Aerolastic tailoring through wings with span-wise graded properties L/R Spars with different L/R ratios (Different chord-wise compliance) Wing with continuous variation of L/R ratio L/R 36
Related research Dynamic shape control LDS V23 Scanning head (Polytec PSV4 M2) Shaker & F. Transducer 1-lb PCB Piezotronics MATLAB Post- Processing DAQ & Signal processing Polytec PSV-4 M2 37
Related research 38
Comparison with numerical results ω = 1744 Hz ω = 225 Hz 39
Dynamic Shape Control: Motivations Control of boundary layers and flow-separation phenomena; Oscillatory camber concept by D. Munday, J. Jacob and G. Huang Vibrating airfoil skins have been found to produce similar results to synthetic jets [Munday]: postpone stall or airflow separation, reduce pressure drag, reduce wave drag. 4
Dynamic Shape Control: Motivations Hogawa H., babinsky H., Evaluation of wave drag reduction by flow control, Aerospace Science and Technology, 1, pp. 1-8, 26 Dynamic deformed shapes could be useful for reducing shock strength and thus wave drag. 41
Wave propagation in chiral networks Ω 25 2 15 1 5 Band-gap structure O A B O k-space position y [m/s] Cg 25 2 15 1 5-5 -1-15 -2-25 -3-2 -1 1 2 3 Cg X [m/s] 42
Related research: Chiral Honeycombs Flat-wise strength Collapse strength per unit weight ( el ) 3/ * / s, [KPa].75.8.85.9.95 1 13 12 11 1 9 8 7 6 5 L/R Hexagonal/Auxetic Honeycombs Global Buckling Chiral -3-2 -1 1 2 3, [deg] Thermal behavior 43
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