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FLOW AROUND AIRFOILS IN DIFFERENT FLOW
REGIMESTeam Members:
Chandrashekar VishwanadhaJibi Ninan Varughese
Krishna PNitya Kalva
ContentsIntroductionProcedureResults: Comparison of different casesConclusionReferences
Objective
To analyze the flow around different airfoils with varying angle of attacks and in different flow regimes
Coordinates of
Airfoil
MATLAB for Gmsh
script generatio
n
Gmsh code
Mesh generation using Gmsh
CFD Analysis
in ANSYS Fluent
Procedure
Meshed Airfoil GeometryAsymmetric Air-foil Symmetric Air-foil
• NACA 0012 airfoil• Max thickness 12% at 30% chord• Max camber 0% at 0% chord
• NACA 6409 airfoil• Max thickness 9% at 29.3% chord• Max camber 6% at 39.6% chord
CFD AnalysisSetup
Solver Boundary conditions
• Type: Density-based• Steady State• 2-D planar space• Absolute velocity
formulation
• Implicit• 2nd Order upwind• Solution steering to
dynamically change CFL• Under-relaxation factor:
0.5
Flow conditions• Inviscid flow• Ideal gas• Energy Equation: On• Po = 1 (atm.), To = 300
( K)
Pressure far field
Pressure far field
Pre
ssu
re f
ar
field P
ressu
re fa
r field
wall
CasesCase 1: Symmetrical Wedge
Mach no = 1.5Mach no = 3
Case 2: Varying Mach numbers @ α = 0o
Symmetrical airfoil (M = 0.2, 1, 3, 7) Asymmetrical airfoil (M = 0.2, 1, 3, 7)
Case 3: Varying Angle of attack @ M =3Symmetrical airfoil (α = 5o, 10o, 15o)Asymmetrical airfoil (α = -15o, 5o, 15o)
ResultsCase 1: Symmetrical Wedge
Mach contour – Mach 1.5 • Detached shock as
expected from the theta- beta-M diagram
Mach contour- Mach 3 • Attached shock• M2= 2.23 which is
approximately equal to theoretical value of (2.26)
• M3=3.64 which is approximately equal to theoretical value 3.65
Wedge: Varying Mach Number
Wedge
Velocity vector plot for Mach 1.4• Sudden change in direction of
flow at bow shock• Gradual change in direction of
flow in Prantdl Meyer expansion
Pressure plot for Mach 1.4
ResultsCase 2: Varying Mach Number
Symmetric Airfoil: Varying Mach NumberLow Subsonic: M =0.2 Transonic: M = 1
Supersonic: M = 3 Hypersonic: M = 7
Mach number contours
Symmetric Airfoil: Varying Mach NumberPressure contours
Low Subsonic: M =0.2 Transonic: M = 1
Supersonic: M = 3 Hypersonic: M = 7
Asymmetric Airfoil: Varying Mach NumberLow Subsonic: M =0.2 Transonic: M = 1
Supersonic: M = 3 Hypersonic: M = 7
Mach number contours
Asymmetric Airfoil: Varying Mach NumberPressure contours
Low Subsonic: M =0.2 Transonic: M = 1
Supersonic: M = 3 Hypersonic: M = 7
ResultsCase 3: Varying Angle of Attack
Symmetric Airfoil: Varying Angle of Attackα = 5o α = 10o
α = 15o
Mach number contours
• Symmetric airfoils produce no lift at zero angle of attack
• velair, top > vel air, bottom
• So, according to Bernoulli's principle (ideal conditions):
Pressuretop < Pressurebottom
• This produces lift!
Slip line
Symmetric Airfoil: Varying Angle of Attackα = 5o α = 10o
α = 15o
Pressure contours
• As angle of attack increases, the differential pressure between top and bottom surfaces increases, thus increasing the lift
Asymmetric Airfoil: Varying Angle of Attackα = 5o α = 15o
α = -15o
Mach number contours
• Asymmetric airfoils produce lift even at zero angle of attack because of difference in surface areas of top and bottom surfaces
• Positive α: velair, top > vel air,
bottom
• Negative α: velair, top < vel air,
bottom
α = 5o α = 15o
α = -15o
Pressure contours
• As angle of attack increases, the differential pressure between top and bottom surfaces increases, thus increasing the lift
• Airfoil is being pushed down during negative angle of attack i.e. lift is negative
Asymmetric Airfoil: Varying Angle of Attack
Lift and drag coefficients
0 5 10 15 20 25 300
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.088
0.09
0.092
0.094
0.096
0.098
0.1Cd,Cl vs AOA
C_lift_symmetric
Angle of attack [degrees]
C_l
ift
C_d
rag
0 1 2 3 4 5 6 7 8
-0.2
-0.1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0
0.02
0.04
0.06
0.08
0.1
0.12
0.14
0.16
0.18
Cl Cd
Cd, Cl vs Mach Number
Mach number
C_l
ift
C_d
rag
ConclusionThe wedge was perfectly symmetrical and thus
produced no lift (C_lift ~ E-06) with zero angle of attack. It also helped in investigating attached shocks and helped in validating the fluent solver
Flow over symmetrical and asymmetrical airfoils was studied under different sonic conditions and angles of attack
The optimal operating conditions observed in fluent agreed with theoretically specified values
ReferencesModern Compressible Flow: With Historical
Perspective. John D. Anderson, JR, 3rd Edition http://airfoiltools.com/search/indexfluent_13.0_workshop02-airfoil.pdfhttp://en.wikipedia.org/wiki/Airfoil
Thank You!Questions?
AppendixHypersonic Flow