effect of airfoil-preserved undulations on wing performance3shorbagy, mohamad a., bamsan el -hadidi,...

Effect of Airfoil-Preserved Undulations on Wing Performance

Faith LoughnaneDr. Sidaard Gunasekaran

Department of Mechanical and Aerospace EngineeringUniversity of Dayton

28th Annual Student Research Symposium

Apr 3

20

20

Ohio Aerospace InstituteCleveland, OH

Outline

1

2

3

4

5

6

Motivation

Context and Objective

Background Research

Experimental Setup

Results

Conclusions and Future Work

2 of 20

Context

LIFT

WEIGHT

THRUSTDRAGParasiticInduced

Wave (N/A)

3 of 20

WINGTIP

VORTEX

Motivation

Bioinspiration: Undulated Wings/Flippers Found in Nature

• Undulations found many places in

nature: birds, bats, humpback whale

flipper

• Humpback whale flipper has been

source of inspiration for similar

studies

• Whale able to perform tight

maneuvers, may be attributed to

presence of tubercles

• Tubercles along leading edge

and partially along trailing edge

4 of 20

Objective

To investigate the effect of airfoil-preserved surface undulations on the aerodynamic

performance and balance of induced and parasite drag.

6 Undulations 9 Undulations 12 Undulations

2Leading Edge

Undulations (LEU)

NACA 0012

Profile

Trailing Edge

Undulations (TEU)

NACA 0012

Profile

Leading + Trailing Edge

Undulations (LETEU)

NACA 0012

Profile

5 of 20

Background Research

2

1Hansen, Kristy L., Kelso, Richard M., and Dally, Bassam B. "Performance variations of leading-edge tubercles for distinct airfoil profiles." AIAA Journal 49, no. 1

(2011): 185-194.2Van Nierop, E., Alben, S & Brenner, M.P., How bumps on whale flippers delay stall: an aerodynamic model, Physical review letters, PRL 100, 054502, February

2008 3Shorbagy, Mohamad A., Bamsan El-hadidi, Gamal El-Bayoumi, Osama Said, and Moatasem Fouda. "Experimental Study on BioInspired Wings With Tubercles."

In AIAA Scitech 2019 Forum, p. 0848. 2019.

Hansen et. al analyzed the effect of leading edge protrusions on a NACA 0021 wing at a

Reynolds number of 120,000 [1]

• Reducing tubercle wavelength was found to be beneficial until a certain point

Van Nierop, Alben, and Brenner (2008) studied the effect of bumps along the leading

edge of the wing [2]

• Bumps along the leading edge cause a more gradual stall which is insensitive of the

wavelength of bumps along the span

Shorbagy et. al performed an experimental study on the effect of sinusoidal bumps

along the leading and trailing edges on aerodynamic performance [3]

• Highest performing cases at low AoA had bumps along the leading edge and

partially along the trailing edge, mimics Humpback whale

6 of 20

Test Models

2

Undulation Placement

Trailing EdgeNumber of

Undulations

6

9

12

Leading Edge Leading + Trailing Edge

7 of 20

Experimental Setup (Force)

Griffin Motion

Rotary Stage

Gamma Sensor

Splitter Plate

Test Model

Inlet

X

Z

Y

Freestream speed:

35 m/s

Reynolds number:

282,000

Angle of Attack

range: -17 to 17

degrees

1 degree angle of

attack increment

8 of 20

Experimental Setup (Cross-Stream PIV)

Freestream speed:

25 m/s

Reynolds number:

201,000

Angle of Attack

range: 2 to 10

degrees

2 degree angle of

attack increment

X

Inlet

Z

Splitter

Plate

Griffin Motion

Rotary Stage

Wingtip

Vortex

1c 2c 3c 4c 5c 6c 7c 8c 9c 10c

Laser

Interrogation

Region

Collector

Imperx

B2021

Camera

Y

9 of 20

Force-Based Results

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Lif

t, C

L

Angle of Attack, Degrees

NACA 00126 LE9 LE12 LEHELMBOLDHELMBOLD AR 4.5

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Lif

t, C

L


NACA 0012

HELMBOLD

HELMBOLD AR 4.5

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Lif

t, C

L


NACA 00126 LE9 LE12 LE6 TE9 TE12 TEHELMBOLDHELMBOLD AR 4.5

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Lif

t, C

L


NACA 00126 LE9 LE12 LE6 TE9 TE12 TE6 LE + TE9 LE + TE12 LE + TEHELMBOLDHELMBOLD AR 4.5

-20

-15

-10

-5

0

5

10

15

20

25

1 3 5 7 9 11 13 15 17

% D

iffe

ren

ce,

CL


-20

-15

-10

-5

0

5

10

15

20

25

1 3 5 7 9 11 13 15 17

% D

iffe

ren

ce,

CL


-20

-15

-10

-5

0

5

10

15

20

25

1 3 5 7 9 11 13 15 17

% D

iffe

ren

ce,

CL


Coefficient of Lift

TEU wings consistently

produce more lift than

baseline wing until stall

LEU and LETEU wings

stall gradually and

have a higher angle of

maximum CL

AR 4

AR 4

AR 4

AR 4

11 of 20

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Dra

g,

CD


NACA 0012

6 LE

9 LE

12 LE

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Dra

g,

CD


NACA 0012

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Dra

g,

CD


NACA 00126 LE9 LE12 LE6 TE9 TE12 TE

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Dra

g,

CD


NACA 00126 LE9 LE12 LE6 TE9 TE12 TE6 LE + TE9 LE + TE12 LE + TE

Coefficient of Drag

-20

-10

0

10

20

30

40

50

60

70

1 3 5 7 9 11 13 15 17

% D

iffe

ren

ce,

CD


-20

-10

0

10

20

30

40

50

60

70

1 3 5 7 9 11 13 15 17

% D

iffe

ren

ce,

CD


-20

-10

0

10

20

30

40

50

60

70

1 3 5 7 9 11 13 15 17

% D

iffe

ren

ce,

CD


TEU wings show most

favorable drag

performance while LEU

wings produce higher

drag

TEU and LETEU wings

show reduction in drag

from 7° until stall

12 of 20

Induced Drag and Parasite Drag

Undulations affect the balance of induced and parasitic drag

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16

0.18

0.2

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Pa

ras

itic

Dra

g, C

D,P


0

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0.09

0.1

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Ind

uc

ed

Dra

g,

CD

,i


NACA 0012

0

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0.09

0.1

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Ind

uc

ed

Dra

g,

CD

,i


NACA 0012

6 LE

9 LE

12 LE

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16

0.18

0.2

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Pa

ras

itic

Dra

g, C

D,P


0

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0.09

0.1

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Ind

uc

ed

Dra

g,

CD

,i


NACA 0012

6 LE

9 LE

12 LE

6 TE

9 TE

12 TE

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16

0.18

0.2

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Pa

ras

itic

Dra

g, C

D,P


0

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0.09

0.1

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Ind

uc

ed

Dra

g,

CD

,i



0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16

0.18

0.2

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

of

Pa

ras

itic

Dra

g, C

D,P


13 of 20

0

2

4

6

8

10

12

0 2 4 6 8 10 12 14 16 18

Ae

rod

yn

am

ic E

ffic

ien

cy,

CL/C

D


NACA 0012

0

2

4

6

8

10

12

0 2 4 6 8 10 12 14 16 18

Ae

rod

yn

am

ic E

ffic

ien

cy,

CL/C

D


NACA 00126 LE9 LE12 LE

Aerodynamic Efficiency

-50

-40

-30

-20

-10

0

10

20

1 3 5 7 9 11 13 15 17

% D

iffe

ren

ce,

CL/C

D


0

2

4

6

8

10

12

0 2 4 6 8 10 12 14 16 18

Ae

rod

yn

am

ic E

ffic

ien

cy,

CL/C

D


NACA 00126 LE9 LE12 LE6 TE9 TE12 TE

-50

-40

-30

-20

-10

0

10

20

1 3 5 7 9 11 13 15 17

% D

iffe

ren

ce,

CL/C

D


0

2

4

6

8

10

12

0 2 4 6 8 10 12 14 16 18

Ae

rod

yn

am

ic E

ffic

ien

cy,

CL/C

D



-50

-40

-30

-20

-10

0

10

20

1 3 5 7 9 11 13 15 17

% D

iffe

ren

ce,

CL/C

D


6 TEU wing shows

higher CL/CD until

stall

Disrupting the

leading edge

results in poor

performance12 LEU 6 TE 12 LETEU

14 of 20

Cross-Stream PIV Results

Freestream

3c

X

Y

Z

0

0.2

0.4

0.6

0.8

1

1.2

0 0.2 0.4 0.6 0.8 1 1.2 1.4

Ux/U

∞

r/rc

BATCHELOR'S MODEL

0

0.2

0.4

0.6

0.8

1

1.2

0 0.2 0.4 0.6 0.8 1 1.2 1.4

Ux/U

∞

r/rc

NACA 0012 2 DEG NACA 0012 4 DEG

NACA 0012 6 DEG NACA 0012 8 DEG

NACA 0012 10 DEG BATCHELOR'S MODEL

0

0.2

0.4

0.6

0.8

1

1.2

0 0.2 0.4 0.6 0.8 1 1.2 1.4

Ux/U

∞

r/rc

NACA 0012 2 DEG NACA 0012 4 DEGNACA 0012 6 DEG NACA 0012 8 DEGNACA 0012 10 DEG 12 LE 2 DEG12 LE 4 DEG 12 LE 6 DEG12 LE 8 DEG 12 LE 10 DEGBATCHELOR'S MODEL

0

0.2

0.4

0.6

0.8

1

1.2

0 0.2 0.4 0.6 0.8 1 1.2 1.4

Ux/U

∞

r/rc

NACA 0012 2 DEG NACA 0012 4 DEGNACA 0012 6 DEG NACA 0012 8 DEGNACA 0012 10 DEG 12 LE 2 DEG12 LE 4 DEG 12 LE 6 DEG12 LE 8 DEG 12 LE 10 DEG6 TE 2 DEG 6 TE 4 DEG6 TE 6 DEG 6 TE 8 DEG6 TE 10 DEG BATCHELOR'S MODEL

0

0.2

0.4

0.6

0.8

1

1.2

0 0.2 0.4 0.6 0.8 1 1.2 1.4

Ux/U

∞

r/rc

NACA 0012 2 DEG NACA 0012 4 DEGNACA 0012 6 DEG NACA 0012 8 DEGNACA 0012 10 DEG 12 LE 2 DEG12 LE 4 DEG 12 LE 6 DEG12 LE 8 DEG 12 LE 10 DEG6 TE 2 DEG 6 TE 4 DEG6 TE 6 DEG 6 TE 8 DEG6 TE 10 DEG BATCHELOR'S MODEL

0

0.2

0.4

0.6

0.8

1

1.2

0 0.2 0.4 0.6 0.8 1 1.2 1.4

Ux/U

∞

r/rc

NACA 0012 2 DEG NACA 0012 4 DEGNACA 0012 6 DEG NACA 0012 8 DEGNACA 0012 10 DEG 12 LE 2 DEG12 LE 4 DEG 12 LE 6 DEG12 LE 8 DEG 12 LE 10 DEG6 TE 2 DEG 6 TE 4 DEG6 TE 6 DEG 6 TE 8 DEG6 TE 10 DEG 12 LE+TE 2 DEG12 LE+TE 4 DEG 12 LE+TE 6 DEG12 LE+TE 8 DEG 12 LE+TE 10 DEGBATCHELOR'S MODEL

Comparison to Batchelor’s Model

• Good agreement in

boundary

• Trends similar between

undulated wings and

baseline

η = r/rc

α = Lamb’s constant: 1.256

Deviations in

vortex core

Agrees along

boundary

16 of 20

Wingtip Vortex Circulation

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.00 0.20 0.40 0.60 0.80

Cir

cu

lati

on

(Γ

/(c*U

∞))

Coefficient of Lift, CL

NACA 0012

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.00 0.20 0.40 0.60 0.80

Cir

cu

lati

on

(Γ

/(c*U

∞))


NACA 0012

6TE

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.00 0.20 0.40 0.60 0.80

Cir

cu

lati

on

(Γ

/(c*U

∞))


NACA 0012

6TE

12LE

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.00 0.20 0.40 0.60 0.80

Cir

cu

lati

on

(Γ

/(c*U

∞))


NACA 0012

6TE

12LE

12LE+TE

Undulated wings

produce lower

circulation for a

given CL

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.00 0.02 0.04 0.06 0.08 0.10

Cir

cu

lati

on

(Γ

/(c*U

∞))

Coefficient of Drag, CD

NACA 0012

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.00 0.02 0.04 0.06 0.08 0.10

Cir

cu

lati

on

(Γ

/(c*U

∞))


NACA 0012

6TE

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.00 0.02 0.04 0.06 0.08 0.10

Cir

cu

lati

on

(Γ

/(c*U

∞))


NACA 0012

6TE

12LE

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.00 0.02 0.04 0.06 0.08 0.10

Cir

cu

lati

on

(Γ

/(c*U

∞))


NACA 0012

6TE

12LE

12LE+TE

TEU wing

produces a higher

CD for a given

circulation

Tells you “how much” rotation is occurring in the wingtip vortex

17 of 20

0.010%

0.015%

0.020%

0.025%

0.030%

0.035%

0.040%

0.045%

0.050%

0.00 0.02 0.04 0.06 0.08 0.10

Peak

(U

RM

S/W

∞)


NACA 0012

0.010%

0.015%

0.020%

0.025%

0.030%

0.035%

0.040%

0.045%

0.050%

0.00 0.02 0.04 0.06 0.08 0.10

Peak

(U

RM

S/W

∞)


NACA 0012

6TE

0.010%

0.015%

0.020%

0.025%

0.030%

0.035%

0.040%

0.045%

0.050%

0.00 0.02 0.04 0.06 0.08 0.10

Peak

(U

RM

S/W

∞)


NACA 0012

6TE

12LE

0.010%

0.015%

0.020%

0.025%

0.030%

0.035%

0.040%

0.045%

0.050%

0.00 0.02 0.04 0.06 0.08 0.10

Peak

(U

RM

S/W

∞)


NACA 0012

6TE

12LE

12LE+TE

Wingtip Vortex URMS

Wingtip vortex RMS trends agree with aerodynamic performance trends

Gives an idea of

“how much”

turbulence is

occurring in the

wingtip vortex

RMS- Root mean

square velocity

Peak RMS- in center

of wingtip vortex

18 of 20

Conclusions

21. Number of undulations does not show correlation to aerodynamic performance,

however placement of undulations is important

2. TEU and LETEU undulations show improved aerodynamic performance over LEU,

baseline wings especially at lower angles of attack

3. LEU and LETEU wings hint post-stall benefits observed in previous literature

(higher stall angle)

4. RMS has an inverse relationship with aerodynamic performance and wingtip vortex

RMS plays a role in affecting the balance of induced and parasite drag

Objective

To investigate the effect of airfoil-preserved surface undulations on the aerodynamic

performance of a wing using NACA 0012 cross-sections

19 of 20

Future Work

2

• Analyze free shear layer PIV images and obtain coefficient of drag using

momentum deficit profiles

• Conduct numerical to compare experimental results

• Perform wingtip vortex PIV at higher angles of attack where there is more

variation between lift/drag coefficients

• Model undulated wing with geometric twist distribution to result in an elliptical lift

distribution

20 of 20

Acknowledgements

2

Mentor: Dr. Sidaard Gunasekaran

Partners: Michael Mongin and Rachael Supina

UD-LSWT Technician: Jielong Cai

UD-LSWT Research Lab Group

Henry Luce Foundation

Ohio Space Grant Consortium (OSGC)

Thank You

Back-upSlides

Dimensions

2

12.7 cm.

25.4 cm.

Amplitude, A

Front View of Wing:

Wavelength, λ

Case A, cm λ, cmSurface Area,

cm2

Planform Area,

cm2

6 LE 2.54 3.91 313.06

285.35

6 TE 2.54 3.91 308.23

6 LE + TE0.635 (LE)

1.905 (TE)3.91 308.02

9 LE 2.54 2.61 319.95

9 TE 2.54 2.61 311.16

9 LE + TE0.635 (LE)

1.905 (TE)2.61 310.71

12 LE 2.54 1.96 328.00

12 TE 2.54 1.96 315.17

12 LE + TE0.635 (LE)

1.905 (TE)1.96 314.32

Wind Tunnel

Test Section

Inlet

Direction of

Freestream

Collector

Freestream Range:

5 to 40 m/s

T.I – 0.1% at 15 m/s

Span Efficiencies

Freestream Range:

5 to 40 m/s

T.I – 0.1% at 15 m/s

𝐶𝐷𝑖 =𝐶𝐿

2

𝜋𝑒𝑉𝐴𝑅

𝑒𝑣 = (1 + 𝛿 + 𝑘𝜋𝐴𝑅)−1

𝑑𝐶𝐿𝑑𝛼

=𝑎0

1 +𝑎0𝜋𝐴𝑅 (1 + 𝜏)

Span Efficiencies

Freestream Range:

5 to 40 m/s

T.I – 0.1% at 15 m/s

Vortex Wandering

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.40

0.45

0.000.050.100.150.200.250.300.35

y/c

x/c

NACA 0012 Baseline

0.049

0.045

Wingtip Vortex Center

locations at 6°


Stevens, P. R. R. J., and H. Babinsky. "Experiments to investigate lift production mechanisms on pitching flat plates." Experiments in Fluids

58, no. 1 (2016). doi:10.1007/s00348-016-2290-x.

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.40

0.45

0.00 0.05 0.10 0.15 0.20 0.25 0.30 0.35

y/c

x/c

𝐫

𝚪

r/rc

0.7

0.6

0.5

0.4

0.3

0.2

0.1

0

Cir

cu

lati

on

(m

2/s

)0 0.5 1 1.5 2 2.5 3

Experimental Data

LOV Model

𝚪 𝒓 = 𝚪𝟎 𝟏 − 𝐞𝐱𝐩 −𝒓𝟐

𝒓𝒄𝟐

𝚪𝟎 = 𝟎. 𝟔

𝐑𝟐 = 𝟎. 𝟗𝟗𝟗

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0 2 4 6 8 10 12

Cir

cu

lati

on

(Γ

/(c*U

∞))


NACA 0012

6TE

12LE

12LE+TE


Undulated wings

trend below

baseline

Undulated wings tend to produce lower circulation for a given angle of attack

29%

0.010%

0.015%

0.020%

0.025%

0.030%

0.035%

0.040%

0.045%

0.050%

0 100 200 300

Peak

(U

RM

S/W

∞)

Vortex Reynolds Number

0.010%

0.015%

0.020%

0.025%

0.030%

0.035%

0.040%

0.045%

0.050%

0 100 200 300

Peak

(U

RM

S/W

∞)


0.010%

0.015%

0.020%

0.025%

0.030%

0.035%

0.040%

0.045%

0.050%

0 100 200 300

Peak

(U

RM

S/W

∞)


0.010%

0.015%

0.020%

0.025%

0.030%

0.035%

0.040%

0.045%

0.050%

0.00 0.20 0.40 0.60 0.80

Pe

ak

(U

RM

S/W

∞)


NACA0012

0.010%

0.015%

0.020%

0.025%

0.030%

0.035%

0.040%

0.045%

0.050%

0.00 0.20 0.40 0.60 0.80

Pe

ak

(U

RM

S/W

∞)


NACA0012

6TE

0.010%

0.015%

0.020%

0.025%

0.030%

0.035%

0.040%

0.045%

0.050%

0.00 0.20 0.40 0.60 0.80

Pe

ak

(U

RM

S/W

∞)


NACA0012

6TE

12LE

Wingtip Vortex URMS

Baseline URMS

0.010%

0.015%

0.020%

0.025%

0.030%

0.035%

0.040%

0.045%

0.050%

0.00 0.20 0.40 0.60 0.80

Pe

ak

(U

RM

S/W

∞)


NACA0012

6TE

12LE

12LETE

0.010%

0.015%

0.020%

0.025%

0.030%

0.035%

0.040%

0.045%

0.050%

0 100 200 300

Peak

(U

RM

S/W

∞)


6 TE URMS

Baseline URMS

Baseline URMS

12LE URMS

× 𝟏𝟎−𝟐𝐔𝐑𝐌𝐒/𝐖∞(%)

effect of airfoil-preserved undulations on wing performance3shorbagy, mohamad a., bamsan el -hadidi,...

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