Aerodynamic Characteristics Aerodynamic Characteristics of a NACA 4412 Airfoilof a NACA 4412 Airfoil

Presented By: David HeffleyPresented By: David HeffleyMentor: Dr. Van TreurenMentor: Dr. Van Treuren

ScholarScholar’’s Days DayJanuary 26, 2007January 26, 2007

OverviewOverview

ObjectiveObjectiveTheoryTheoryApparatusApparatusExperimental ComparisonExperimental ComparisonResultsResultsSummarySummaryRecommendationsRecommendations

ObjectiveObjective

Study the lift and drag forces on a NACA 4412 Study the lift and drag forces on a NACA 4412 airfoilairfoilResolve discrepancy in wind tunnel dataResolve discrepancy in wind tunnel dataDevelop experimental techniques for an airfoilDevelop experimental techniques for an airfoilCompare wind tunnel dataCompare wind tunnel data

Force Balance to Pressure DistributionForce Balance to Pressure DistributionBaylor data to published NACA dataBaylor data to published NACA data

NACA 4412 AirfoilNACA 4412 Airfoil4 digit code used to describe airfoil shapes4 digit code used to describe airfoil shapes1st digit 1st digit -- maximum camber in percent chordmaximum camber in percent chord2nd digit 2nd digit -- location of maximum camber along chord line (from leading edge)location of maximum camber along chord line (from leading edge) in tenths of in tenths of

chordchord3rd and 4th digits 3rd and 4th digits -- maximum thickness in percent chordmaximum thickness in percent chordNACA 4412 with a chord of 6NACA 4412 with a chord of 6””

Max camber: 0.24Max camber: 0.24”” (4% x 6(4% x 6””))Location of max camber: 2.4Location of max camber: 2.4”” aft of leading edge (0.4 x 6aft of leading edge (0.4 x 6””))Max thickness: 0.72Max thickness: 0.72”” (12% x 6(12% x 6””))

Mean camber line

Chord line

Chord

x=0 x=c

Max thickness

Max camber

Leading edge Trailing edge

x

z

TheoryTheoryLift, Drag and Angle of AttackLift, Drag and Angle of Attack Stall AngleStall Angle

ViscousMomentumVc

===µρReNumber Reynolds

α

Lift

Drag∞V

RelativeWind

TheoryTheory

SV

LCl2

21 ρ

=SV

DCd2

21 ρ

=

Dyn

StatLocalP P

PPC

−=

∫ −=1

0

)()(cxdCCC PUPLY∫

−

−=cy

cy

PAPFX cydCCC )()(

Relates lift and drag forces to the velocity

Direct Method (Force Balance)

Pressure Distribution (Pressure Ported Airfoil)

Relates local pressure on an airfoil to the velocity

αα sincos XYl CCC −= αα cossin XYd CCC +=

Experimental ApparatusExperimental ApparatusBaylor University Wind TunnelBaylor University Wind Tunnel

24” by 24” Test Section Test Range: 0 – 150 ft/s Open loop tunnel

Experimental ApparatusExperimental Apparatus

Force BalanceForce Balance Pressure Tapped AirfoilPressure Tapped Airfoil

18 pressure ports-18 to 20 Degrees

Both NACA 4412 airfoils are 24” wide with a 6”

chord length

-8 to 20 Degrees

Experimental ComparisonExperimental Comparison

Re = 3,000,000Re = 3,000,00054 pressure ports54 pressure portsVariable density wind Variable density wind tunneltunnel2424”” chord lengthchord length

Re = 150,000Re = 150,00018 pressure ports18 pressure portsConstant density Constant density wind tunnelwind tunnel66”” chord lengthchord length

NACA Baylor University

ResultsResults

Stall angleStall angle11 degrees for 150,000 Re (Baylor)11 degrees for 150,000 Re (Baylor)15 degrees for 3,000,000 Re (NACA)15 degrees for 3,000,000 Re (NACA)

Lift coefficient agrees within 2% of NACA Lift coefficient agrees within 2% of NACA published datapublished dataNoticeable inaccuracies in drag coefficient data Noticeable inaccuracies in drag coefficient data from the pressure ported airfoilfrom the pressure ported airfoilDrag coefficient is Re dependentDrag coefficient is Re dependent

Aerodynamic CurvesAerodynamic Curves

Lift CurveLift Curve Drag CurveDrag Curve

α

Cl Cd

Cl

Higher Re Curve

Lift CurveLift CurveCl v α

-0.90

-0.70

-0.50

-0.30

-0.10

0.10

0.30

0.50

0.70

0.90

1.10

1.30

1.50

1.70

-20 -16 -12 -8 -4 0 4 8 12 16 20 24

Angle of Attack (Degrees)

Coe

ffici

ent o

f Lift

NACA Report563NACA Report824Force Balance

Pressure

Lift Pressure DistributionLift Pressure Distribution10 degrees CP vs. x/c

-4

-3

-2

-1

0

1

-0.1 0.1 0.3 0.5 0.7 0.9 1.1x/c

CP

Exp Lower SurfaceExp Upper SurfaceNACA 563 Lower SurfaceNACA 563 Upper Surface

Drag CurveDrag CurveCD v CL

0

0.005

0.01

0.015

0.02

0.025

0.03

0.035

0.04

0.045

-0.75 -0.25 0.25 0.75 1.25Coefficient of Lift

Coe

ffici

ent o

f Dra

g

NACA 563NACA 824Force BalancePressure

Drag Pressure DistributionDrag Pressure Distribution

10 degrees CP vs. y/c

-4

-3

-2

-1

0

1

-0.04 -0.02 0 0.02 0.04 0.06 0.08 0.1 0.12y/c

CP

Exp Lower SurfaceExp Upper SurfaceNACA 563 Lower SurfaceNACA 563 Upper Surface

CCDD vs. Reynolds Numbervs. Reynolds Number

Munson, B. R., Young, D. F., and Okiishi, T. H., 2006, Fundamentals of Fluid Mechanics

SummarySummary

ObjectivesObjectivesStudy airflow over an airfoil Study airflow over an airfoil Resolve discrepancy in previous wind tunnel dataResolve discrepancy in previous wind tunnel dataCompare wind tunnel dataCompare wind tunnel data

ResultsResultsStall angle is a function of the Reynolds numberStall angle is a function of the Reynolds numberLift coefficient relates closely to published dataLift coefficient relates closely to published dataInsufficient pressure ports to accurately map the pressure Insufficient pressure ports to accurately map the pressure distribution for drag coefficientdistribution for drag coefficientDrag coefficient highly dependent on Reynolds numberDrag coefficient highly dependent on Reynolds number

RecommendationsRecommendations

Further experimentsFurther experimentsNACA 0012 (Double the pressure ports)NACA 0012 (Double the pressure ports)Utilize BaylorUtilize Baylor’’s 3D printers 3D printer

Develop lift and drag curves for future Develop lift and drag curves for future experiments to referenceexperiments to reference

QuestionsQuestions