naca 4412
DESCRIPTION
Aerodynamic CharacteristicsTRANSCRIPT
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