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Aerodynamics PDR 2Aerodynamics PDR 2

Ashley Brawner

Neelam Datta

Xing Huang

Jesse Jones

Team 2: Balsa to the Wall

Matt NegilskiMike Palumbo

Chris SelbyTara Trafton

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OverviewOverview

Design Point Airfoil Selection Component Drag Buildup Drag Polar AR trade study (CL)max Approximation

(Cl)max method

(CL)max Raymer method Flap analysis

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The Design PointThe Design Point

Weight 5.5 [lbs]

Dihedral Angle 0°

Speed 110 [ft/sec]

Horizontal Tail Span

1.5 [ft]

Planform area based on approximated (CL)max and weight estimate

Dihedral angle of 0° taken from Roskam

Design speed decreased from 150 ft/sec

Designed to high speed mission

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Airfoil Selection: Airfoil Selection: Main WingMain Wing

Wing Section NACA 1408

Gives approximate 2D Cl needed for dash Relatively thin for minimizing drag Thick enough for structural strength

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XFOIL: Drag Polar

-0.6

-0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

1.1

1.2

0 0.005 0.01 0.015 0.02 0.025 0.03

Cd

Cl

Jones_main1

NACA 1408

MH 30

MH 64

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Airfoil Selection: Airfoil Selection: TailTail

Tail Sections Horizontal Stabilizer

Symmetric with low Cd over a wider range of a.o.a. compared to other similar airfoils

Symmetric Jones airfoil (≈8% t/c) Vertical Stabilizer

NACA 0006

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Various airfoils Cd vs. alpha

0.004

0.005

0.006

0.007

0.008

0.009

0.01

0.011

0.012

0.013

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5

alpha [deg]

Cd

Symmetric Jones (7% t/c)

NACA 0006

NACA 0007

NACA 0008

Symmetric Jones (6% t/c)

Symmetric Jones (8% t/c)

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Drag Build-up Method (Raymer)Drag Build-up Method (Raymer)

Cfc = Component skin friction coefficient

FFc = Component form factor

Qc = Component interference effects

Swet,c = Component wetted area

Sref = Wing planform

ref

c cwetccfcD S

SQFFCC

,0

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Component Coefficient of frictionComponent Coefficient of friction

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Drag Build-up Method resultsDrag Build-up Method results

5AR 5AR

5AR

5AR sec]/[110 ftV Inputs:

11-1 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9Drag Polar

CD

CL

Drag PolarDrag Polar

20707.0253. LD CC

iDDD CCC ,0

eARk

1

20 LDD kCCC

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AR Trade studyAR Trade study

elseallARfC

RaymerARfeeAR

k

kCC

CCC

D

LiD

iDDD

,

)(

1

0

2,

,0

LD

D

CARfC

ARfC

constelseall

,

.

0

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AR Trade studyAR Trade studyLines of constant C

D (An AR trade study)

AR

CL

0.02550.026 0.0265

0.027 0.0275 0.0280.029

0.03

0.035

0.04

0.05

0.06

0.08

0.10.12

3 4 5 6 7 8 9 10 11 120

0.2

0.4

0.6

0.8

1

1.2

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(C(Cll))maxmax Approximation Approximation

Compare XFOIL with Abbott & Doenhoff wind tunnel data

Conclusion αClmax ≈ 0.8αClmax(XFOIL)

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NACA 1408 Cl vs. alphaRe = 9,000,000

-1

-0.5

0

0.5

1

1.5

2

-10 -5 0 5 10 15 20 25

alpha [deg]

Cl Experimental

Numerical

XFOILClCl maxmax 8.0

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NACA 1408 Cl vs. alphaRe = 6,000,000

-1

-0.5

0

0.5

1

1.5

2

-10 -5 0 5 10 15 20 25

alpha [deg]

Cl Experimental

Numerical XFOILClCl maxmax 8.0

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NACA 1408 Cl vs. alphaRe = 3,000,000

-1

-0.5

0

0.5

1

1.5

2

-10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 16 18 20

alpha [deg]

Cl Experimental

Numerical

XFOILClCl maxmax 8.0

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NACA 1408 Cl vs. alphaRe = 500,000

0

0.2

0.4

0.6

0.8

1

1.2

0 2 4 6 8 10 12

alpha [deg]

Cl

XFOILClCl maxmax 8.0

Clmax = .9412 (clean)

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Flap analysisFlap analysis

Use (CL)max approximation from Raymer Ads

Use XFOIL to find (Cl)max with flaps Observation -

Flapped (Cl)max follows linear trend Determine maximum achievable (CL)max Find flap configuration that acheives optimal

(CL)max

max,max_max, Ll

LflappedlL C

C

CCC

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XFOIL: NACA 1408 Cl vs. alphaRe= 500,000

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

0 2 4 6 8 10 12

alpha [deg]

Cl

Clean

Flap: x/c=0.8, 20 degrees

Flap: x/c=0.8, 30 degrees

Flap: x/c=0.75, 20 degrees

Flap: x/c=0.75, 30 degrees

Flap: x/c=0.85, 20 degrees

Flap: x/c=0.85, 30 degrees

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NACA 1408 Clmax vs. alphaRe = 500,000

y = -0.1113x + 1.8911

R2 = 0.997

y = 0.0975x + 0.1765

R2 = 0.9998

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

0 1 2 3 4 5 6 7 8 9

alpha [deg]

Clm

ax

Clmax with flaps

Naca 1408 (clean)

Linear (Clmax with flaps)

Linear (Naca 1408 (clean))

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Flap analysis: (continued)Flap analysis: (continued)

Use linear fit lines to find a Δ(Cl)max and then find Δ(CL)max with the following equation from Raymer: ads

The ratio blank is based on the intial sizing of the wing area and tail span and is assumed to remain constant

HLflapped

ll

LL S

SC

C

CC

cosmax,max,

S

S flapped

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Flap Geometry:Flap Geometry:

flap hinge location (x/c) = 0.8 maximum flap deflection = 35° constant (cf/c) flap

(CL)max (w/ flaps) = 1.06

(Cl)max (w/o flaps) = 0.85

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Summary TableSummary Table

(CL)max (w/ flaps) 1.06

(CL)max (w/o flaps) 0.84

CD0 0.0253

AR 6

b 5 [ft]

croot 16.24 [in]

ctip 7.35 [in]

troot 1.3 [in]

ttip 0.6 [in]

Flap location (x/c) 0.8

Maximum flap deflection 35°

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Questions?Questions?