airfoil analysis & comparison - group builder...2011/01/25 · higher aoa and a sharper stall,...

Airfoil Analysis & Comparison

Overview

� Airfoil selection

� Airfoil Designations

� Xfoil� Xfoil

� Reynolds Number (aviation’s dirty lie)

� Aerobatic airfoils & geometry

� Data

� Some Thoughts & Summary

Airfoil Selection

� For the purposes of selecting an aerobatic airfoil (and plain old curiosity), several existing airfoils were analyzed using Xfoil:

Munk M6 (flat-wing pitts; S-1C & S-1D)�Munk M6 (flat-wing pitts; S-1C & S-1D)

�NACA 0012 (Pitts lower wing) & 0015

�NACA 63A015 (Pitts upper wing)

�Eppler 472 (Extra, Edge*, MX2*)

�Clark Y (older Pipers, many others)

Airfoil Selection

� In addition, ‘Akro1’ series of airfoils was developed.

� Airfoils in this series consist of elliptical nose section with straight section with straight segments to a blunt trailing edge.

� Geometry is based on appearance of high performance acro airfoils

Airfoil Designations

� NACA 4-digit series:

� 1st #: max camber in

% of chord

� Decathlon uses

NACA 1412

� 1 = 1% camber

� 2nd #: position of max

camber in tenths of

chord

� 3rd/4th #: percentage

of max thickness

� 4 = max chamber @

40% chord

� 12 = 12% thickness


� NACA 6-series- first ‘successful’ laminar flow airfoils

� 1st #: series designation (always 6)

� 2nd #: position of min. pressure in tenths of chord (how far back laminar flow can go)

� Pitts upper wing: 63A015

� 6 = NACA 6-series

� 3 = laminar flow as far back as 30% chord

� A = part of rear airfoil replaced with straight line segment (fabric laminar flow can go)

� Special letters & numbers: don’t ask!

� 3rd #: lift coefficient in tenths at which airfoil was designed for low drag

� 4th/5th #: max thickness in # of chord

with straight line segment (fabric covered aircraft)

� 0 = designed for Cl = 0 (symmetric airfoil)

� 15 = 15% thickness


� Munk M6: Munk’s 6th airfoil?�Max Munk was German immigrant, one of

America’s greatest aerodynamicists of 20’s, 30’s, 40’s. Published a lot of material.30’s, 40’s. Published a lot of material.

� Clark Y: Y name airfoils?�Clark Y is in fact a Göttingen Gö 398

�Clark Y (german airfoil) is the basis for the NACA 4 and 5 digit series. Thank a German.


� Akro1-series

� 1st #: position of max

thickness in %

� 2nd #: max thickness in %

� Best design: Akro1-20-

15-10

� Akro1 = 1st series

� 20 = max thickness is at � 2nd #: max thickness in %

chord

� 3rd #: thickness of blunt

trailing edge in tenths of

% (sorry!)

� 20 = max thickness is at

20% chord

� 15 = 15% max thickness

� 10 = 1% thick trailing

edge

Profiles

Profiles

Akro1-20-15-10

0.3

0.4

0.5

-0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

0 0.2 0.4 0.6 0.8 1 1.2

Series1

Xfoil

� Program used for analysis is Xfoil:� “Panel” code with boundary

layer analysis.

� Generally accurate until just past stall past stall

� (unfortunately, we are interested in post-stall region)

� Analysis is not 100% correct, but comparisons between airfoils are very accurate.

� Xfoil has some design capacity.

� Never use airfoil that hasn’t been tested in a wind tunnel

Reynolds Number (Re)

� Dirty Lie: Airfoils stall at same AoA, properties don’t vary with airspeed.

� Airfoil properties vary with Reynolds number:� CL max, AoAstall increase, CD decreases with increasing Re. L max stall D

� Higher Re is better.

� Re increases linearly with airspeed, decreases with altitude.

� Re = ρVc/µ� ρ: density

� V: velocity

� c: characteristic length (chord, in our case)

� µ: fluid viscosity

Airfoil Data

� Performance calculated at two Reynolds numbers:

�Re = 2,250,000 (3 ft chord near stall speed)�Re = 2,250,000 (3 ft chord near stall speed)

�Re = 10,500,000 (4.5 ft chord near 250 mph)

� Airfoils sometimes grouped into 12% group and 15% group.

Aerobatic Airfoils

� Aerobatic airfoils are (ideally) tailored for low drag at high AoA, instead of low drag at cruise (low AoA)

For unlimited aerobatics, sharp stall is � For unlimited aerobatics, sharp stall is desired. For training and recreation, gradual stall is better.

� Important characteristics for aerobatics are CL vs. AoA, L/D, and CL vs. CD.

Important Airfoil Geometry

� Symmetric

� Forward max thickness increases CL max.

� NACA/Clark Y/M6 airfoils have highest CL max at 12% thickness. thickness.

� More decreases max lift, but improves structural weight.

� A blunt trailing edge keeps stagnation (separation) point attached to trailing edge to much higher AoA than sharp T.E.

� Steeper pressure recovery both increases max lift and sharpens stall.

Cl vs. AoA (classic airfoils)

Cl vs. AoA (Re = 2.25 million)

1.2

1.4

1.6

1.8


2

2.5

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

-5 0 5 10 15 20 25

AoA

Cl

Clark Y

Munk M6

NACA 0012

NACA 0015

NACA 63A015

-0.5

0

0.5

1

1.5

-10 -5 0 5 10 15 20 25

AoA

Cl

Clark Y

Munk M6

NACA 0012

NACA 0015

NACA 63A015

Cl vs. AoA (newer airfoils)


1.4

1.6

1.8

2


2

2.5

0

0.2

0.4

0.6

0.8

1

1.2

0 5 10 15 20 25 30

AoA

Cl

NACA 0012

Eppler 472

Eppler 472 (mod T.E.)

Akro1-20-15-10

0

0.5

1

1.5

0 5 10 15 20 25 30

AoA

Cl

NACA 0012

Eppler 472


Akro1-20-15-10

L/D vs. AoA (classic airfoils)

L/D vs. AoA (Re = 2.25 million)

120

140

160


120

140

160

180

-20

0

20

40

60

80

100

-5 0 5 10 15 20 25

AoA

L/D

Clark Y

Munk M6

NACA 0012

NACA 0015

NACA 63A015

-20

0

20

40

60

80

100

120

-10 -5 0 5 10 15 20 25

AoA

L/D

Clark Y

Munk M6

NACA 0012

NACA 0015

NACA 63A015

L/D vs. AoA (newer airfoils)


100

120


200

250

0

20

40

60

80

0 5 10 15 20 25 30

AoA

L/D

NACA 0012

Eppler 472


Akro1-20-15-10

0

50

100

150

0 5 10 15 20 25 30

AoA

L/D

NACA 0012

Eppler 472


Akro1-20-15-10

Cl vs. Cd (classic airfoils)

Cl vs Cd (Re = 2.25 million)

1.2

1.4

1.6

1.8


2

2.5

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

0 0.02 0.04 0.06 0.08 0.1 0.12

Cd

Cl

Clark Y

Munk M6

NACA 0012

NACA 0015

NACA 63A015

-0.5

0

0.5

1

1.5

0 0.01 0.02 0.03 0.04 0.05 0.06

Cd

Cl

Clark Y

Munk M6

NACA 0012

NACA 0015

NACA 63A015

Cl vs. Cd (newer airfoils)


1.4

1.6

1.8

2


2

2.5

0

0.2

0.4

0.6

0.8

1

1.2

1.4

0 0.05 0.1 0.15 0.2 0.25

Cd

Cl

NACA 0012

Eppler 472


Akro1-20-15-10

0

0.5

1

1.5

0 0.05 0.1 0.15 0.2 0.25

Cd

Cl

NACA 0012

Eppler 472


Akro1-20-15-10

Some thoughts

� We lose a lot of performance with no camber.

� The NACA 0012 and 0015 seem almost identical, and the 0015 seems to have slightly greater lift, going against the general statement that 12% thickness has the greatest lift.greatest lift.

� The Eppler airfoil (without modification) has lower drag at higher AoA and a sharper stall, but not much higher CLmax.

� Adding a blunt trailing edge to the Eppler and Akro1 airfoils resulted in drastic improvements.� It is likely that blunting the trailing edges of the NACA/Munk M6

airfoils would have the same effect.

Some thoughts

� Some of the data associated with the Akro1 looks suspicious. It’s superiority to the other airfoils is possibly bad data.

� The Munk has superior performance to the other classic airfoils at low AoA, agreeing with the general belief that airfoils at low AoA, agreeing with the general belief that the flatwing Pitts climbs and cruises better. At high AoA, there is no appreciable difference with the other airfoils.

� The 63A015 looks lousy. From personal design experience, the 6-series has much lower drag than the 4 and 5 digit series. Explanations could be the low (30%) extent of laminar flow and the modification (63A015)

In Conclusion

� The data largely agreed with existing data and experience.

� There is a good probability of bad data in some places.places.

� Arguments could be made either way about how large/small differences between airfoils are.� That said, I will be using the Eppler 472 with the blunt

trailing edge. To hell with designing an airfoil from scratch.

airfoil analysis & comparison - group builder...2011/01/25 · higher aoa and a sharper stall,...

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