Light Aircraft Aircraft Design

8/10/2019 Light Aircraft Aircraft Design

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T.O. over 50' = 326 ft, fixed pitch prop Landing dist. =

T.O. distance = 39 ft, constant speed prop Land over 50'=

T.O. over 50' = 268 ft, constant speed prop

T.O. Speed= 18 mph = 15 kts

Weight and Balance Estimate:

Ultimate Load Facto 6 g's start of flap/semispan 0.08

Horizontal Tail Area: 1.50 sq ft end of flap/semispan: 0.49

Vertical Tail Area: 0.75 sq ft airfoil alphaCL0: -2.0

Max fuel capacity: 1 gallons Sug. wing incidence 3.7

Wing root t/c: 18 % chord (for ca Wing incidence: 1.0

Fuselage Length: 2 ft (from firewall Wing taper ratio: 1

# front passengers: 0

# rear passengers: 0

Fuselage constructi 1 1 - tube & fabric, 2 - metal, 3 - composite

Wing construction: 1 1 - fabric external brace, 2 - metal external brace, 3 -

Tail construction: 1 1 - fabric external brace, 2 - metal cantilever, 3 - com

Propeller type: 1 1 - wooden fixed pitch, 2 - metal fixed pitch

Landing gear type: 1 1 - taildragger, 2 - tricycle gear

Fuselage structure Weight Estimate (lbs) Wing structur

Welded steel tube & fabric covered 0.4 Fabric covered

Metal fuselage 0.4 Metal covered

Composite fuselage 0.4 Metal cantilever

Cockpit controls 0.0 Composite cant

Control system 0.0

Seat (per passenger) 0.0 Propulsion gr

Interior 0.0 Cowling

Paint (complete airplane) 0.0 Engine mountFuel system co

Tail structure Fuel tanks

Fabric covered external braced hor. 1.2 Wooden fixed p

Metal horizontal tail 1.3 Metal fixed pitc

Composite horizontal tail 1.5

Fabric covered external braced vert. 0.6 Landing gear

Metal vertical tail 0.7 Main wheels

Composite vertical tail 0.8 Nose wheel

Tail wheel

Electrical and instrumentation group Main gear legs

Instruments/Avionics 0.5 Nose gear leg

Battery 0.0 Tail wheel legElectrical system 0.1

Mean aerodynamic chord (MAC) = 12.00 inches (for a monoplane wing)

Fuselage station of MAC 1/4 chord: 12.00 inches

Empty Aircraft Weight and Balance



Item Estimated Weight Weight (lbs) F.S.

1 Wing 5.9 5.9 15.00

2 Fuselage 0.4 0.4 15.00

3 Hor. tail 1.2 1.2 28.00

4 Vert. tail 0.6 0.6 28.00

5 Main gear whe 0.0 0.0 13.00

6 Main gear wheel pants 8.0 13.00

7 Main gear legs 0.1 0.1 13.00

8 Nose gear whe 0.0 0.0 0.00

9 Nose gear wheel pants 0.0 0.00

10 Nose gear leg 0.0 0.0 0.00

11 Tail wheel 0.0 0.0 28.00

12 Tail wheel leg 0.0 0.0 28.00

13 Cockpit control 0.0 0.0 15.00

14 Control system 0.0 0.0 18.00

15 Engine (including accessories) 0.5 8.00

16 Spinner 3.0 0.00

17 Propeller 1.5 1.5 0.00

18 Engine mount 0.0 0.0 9.00

19 Cowling 0.0 0.0 8.00

20 Battery 0.0 0.0 95.00

21 Fuel tanks 0.7 0.7 10.00

22 Fuel system 0.0 0.0 10.00

23 Inst/Avionics 0.5 0.5 10.00

24 Electrical syste 0.1 0.1 10.00

25 Interior 0.0 0.0 0.00

26 Front seats 0.0 0.0 0.00

27 Back seats 0.0 0.0 0.00

28 Paint 0.0 0.0 0.00

Totals 22.4 12.0

Payload

Empty weight 22 lbs

Front passenger 0 lbs

Rear passenger 0 lbs

Baggage 0 lbs

Fuel (6 lbs/gallon for 1 lbs

Oil (7.5 lbs/gallon) 0.1 lbs

Payload 1 lbs

Gross Weight = 24 lbs

Gross Weight from cell C15 = 5 lbs

Forward CG Condition

Item Weight (lbs) F.S.

1 Empty weight 22 12.0

2 Front passenger 0 0

3 Rear passenger 0 0

4 Baggage 0 0

5 Fuel (6 lbs/gallon for 1 12



6 Oil (7.5 lbs/gallon) 0.1 12

Total 24 12.0

Total payload = 1 lbs

Gross weight = 24 lbs (for the given payload)

Hor C.G. position = 12.0 in = 25.0% MAC

Ver C.G. position = 5.9 in = -67.8% MAC (negative

Horizontal angle between wheel and aft fuselage (or 14.1 degrees (min. a

Vertical angle between CG and main gear axle (taild 25.0 degrees ( 13 de

Aft CG Condition

Item Weight (lbs) F.S.

1 Empty weight 22 12.0

2 Front passenger 0 0

3 Rear passenger 0 0

4 Baggage 0 0

5 Fuel (6 lbs/gallon for 1 12

6 Oil (7.5 lbs/gallon) 0.1 12

Total 24 12.0

Total payload = 1 lbs

Gross weight = 24 lbs (for the given payload)

Hor C.G. position = 12.0 in = 25.0% MAC

Ver C.G. position = 5.9 in = -67.8% MAC (negative

Horizontal angle between wheel and aft fuselage (or 14.1 degrees (min. a

Vertical angle between CG and main gear axle (taild 25.0 degrees ( 13 de

Vertical angle between CG and main gear axle (tric N/A degrees (must

Min. vertical angle between CG and main gear axle N/A degrees (minim

Aircraft Preliminary Performance Estimate and Sizing Spreadsheet

Modified for use on R/C aircraft.

Based on methods and information presented in:

Sport Aviation Magazine, May, 2000, by Neal Willford

"Technical Aerodynamics" by K.D. Wood

"Aircraft Design" by K.D. Wood

"Engineering Aerodynamics" by W.S. Diehl

"Airplane Performance, Stability and Control" by Perkins and Hage

"Preliminary Design Processes" by Herb Rawdon



Background calculations

approximately 1.3 to 1.4 Cdo = 0.083

if no flaps enter same value as flaps up Lp = 10.00

approximately 1.8 for plain flaps, 2.0 for slotted Lt = 64.10

sq ft, flaps down Ls = 0.17

bhp 74.6 Watts lambda = 20.02

Wing AR = 6.00

noplane) Lt cnsspd = 56.19

ane) lamda cnsspd= 16.79

a biplane. Enter 0 for a monoplane) Cs 3bl = 0.52

get from engine manufacturer's information L/Dmax = 6.83

gallons Prop/body int= 0.96

minutes

chord/Diameter @ 75% prop radius

.03 concrete, .05 short grass, 0.1 long grass

ratio of takeoff speed to stall speed (1.15 to 1.2)

en subtract .03 if using a wooden propeller)

Fixed Pitch Propeller Performance Propeller advance ratio, J = 0.15

max ROC = 222 fpm T (fixed pitch)= 2

Abs. Ceiling = 12906 feet Tc (fixed pitch)= 1

Service Ceiling 7092 feet T (constant speed)= 2

Constant Speed Propeller Performance Tc (constant speed)= 2

max ROC = 284 fpm

Abs. Ceiling = 14828 feet R = 1

Service Ceiling 9607 feet Dc = 0Cl at Vmax = 0.44 Xt fixed pitch= 9

Estimated Propeller Ht fixed pitch= 1

Pitch = 0.4 inches Xt constant speed= 10

Ht constant speed= 1



37 feet, flaps down (1.15xVstall)

455 feet, flaps down (1.15xVstall)

0 if no flaps

0 if no flaps Wing root chord= 12.00

airfoil zero lift angle (typically -2 to -4 deg) Wing tip chord= 12.00

degrees (suggested wing incidence at the MAC) aw = 0.077

degrees (wing incidence at the MAC) delta alfa0 = -4.00

(tip chord/root chord) approx delta Clflap= 0.78

flaps up stall angle = 13.4

flaps down stall angle = 14.1

metal cantilever, 4 - composite cantilever

osite cantilever

Weight Estimate (lbs)

external braced 5.9

xternal braced 6.5

wing 1.1

ilever wing 1.3

up

0.0

0.0ponents 0.0

0.7

itch propeller 1.5

propeller 4.6

0.0

0.0

0.0

0.1

0.0

0.0



W.L. HZ mom. VT mom.

14.00 88 83

12.00 5 4

12.00 33 14

14.00 16 8

8.00 0 0

8.00 104 64

9.00 1 1

0.00 0 0

0.00 0 0

0.00 0 0

10.00 0 0

10.00 0 0

10.00 0 0

13.00 0 0

11.00 4 6

11.00 0 33

11.00 0 16

11.00 0 0

11.00 0 0

10.00 4 0

11.00 7 7

11.00 0 0

11.00 5 6

11.00 1 1

12.00 0 0

11.00 0 0

11.00 0 0

13.00 0 0

10.8 270 244


6.0 269 135

0 0 0

0 0 0

0 0 0

3 12 3



3 1 0

5.9 283 138

below wing)

ngle req'd to keep tail from hitting ground on landing)

g minimum at forward CG location, about 25 deg at aft)


6.0 269 135

0 0 0

0 0 0

0 0 0

3 12 3

3 1 0

5.9 283 138

below wing)

ngle req'd to keep tail from hitting ground on landing)

g minimum at forward CG location, about 25 deg at aft)

e greater than the angle below for a tricycle gear)

um req'd angle to keep from tipping on tail at landing)



lbs at .7Vto

lbs at Vto

lbs at .7Vto

lbs at Vto

lbs at .7Vto

lbs at Vtoft

ft

ft

ft



inches

inches

(CL/deg) wing lift slope

degrees, delta for zero alpha due to flaps

approx increase in Cl of 2-d wing section due to flap deflection

degrees

degrees



SD7062 GOE 386 AIRFOIL EPPLER E1212 AIRFOIL

1 0 1 -0.0016 1 0

0.99646 0.00027 0.95034 -0.00497 0.99621 0.00008

0.98593 0.00092 0.90057 -0.00834 0.98512 0.00006

0.96864 0.00163 0.80102 -0.01509 0.96735 -0.00038

0.94491 0.00216 0.70148 -0.02184 0.9434 -0.00118

0.91518 0.0023 0.60194 -0.02859 0.91355 -0.00231

0.87998 0.00189 0.5024 -0.03534 0.8782 -0.00392

0.8399 0.00087 0.40286 -0.04209 0.83789 -0.00611

0.79557 -0.00079 0.30331 -0.04884 0.79324 -0.00892

0.74768 -0.00307 0.20369 -0.05429 0.74489 -0.01234

0.69692 -0.00593 0.15379 -0.05577 0.69355 -0.01635

0.64403 -0.00927 0.10366 -0.05387 0.63994 -0.02089

0.58975 -0.01295 0.07841 -0.05018 0.5848 -0.02587

0.53477 -0.0168 0.05316 -0.0466 0.52889 -0.0312

0.47979 -0.02064 0.0276 -0.03823 0.47299 -0.03672

0.42547 -0.02426 0.01447 -0.02892 0.41785 -0.04229

0.37238 -0.02747 0.00001 0 0.36422 -0.0477

0.32105 -0.0301 0.01007 0.03588 0.31285 -0.05273

0.27199 -0.03206 0.0214 0.05315 0.26443 -0.05712

0.2257 -0.03327 0.04479 0.07694 0.21963 -0.06055

0.18265 -0.03367 0.06855 0.09525 0.1791 -0.06247

0.14328 -0.03319 0.09262 0.10888 0.14296 -0.06215

0.10798 -0.03179 0.14122 0.12957 0.11083 -0.05933

0.0771 -0.02948 0.19032 0.1428 0.0825 -0.05429

0.05099 -0.02623 0.28963 0.15304 0.058 -0.04733

0.02991 -0.02205 0.38971 0.15182 0.03738 -0.03884

0.01414 -0.01701 0.49062 0.13836 0.02081 -0.02927

0.00425 -0.01115 0.592 0.11804 0.00858 -0.019110.00027 -0.00327 0.69356 0.09502 0.00152 -0.00892

0.00103 0.00711 0.79543 0.06743 0 0

0.00562 0.01871 0.8976 0.03546 0.0001 0.00239

0.0141 0.03081 0.94872 0.01892 0.00342 0.01552

0.0265 0.043 1 0.0016 0.01072 0.02949

0.04272 0.05491 0.02185 0.04371

0.06269 0.06616 0.03668 0.05773

0.08637 0.07647 0.05506 0.07116

0.11363 0.08562 0.07682 0.0836

0.14432 0.0934 0.10175 0.09455

0.17825 0.09965 0.12989 0.10351

0.21521 0.10428 0.16137 0.110310.25496 0.10722 0.19609 0.11495

0.2972 0.10846 0.23389 0.11735

0.34159 0.10801 0.27463 0.11756

0.38779 0.10592 0.31804 0.1157

0.4354 0.10229 0.36383 0.11193

0.48402 0.09723 0.41164 0.10643

0.53322 0.09089 0.46109 0.09947



0.58254 0.08347 0.51172 0.09133

0.63154 0.07517 0.563 0.08233

0.67973 0.06626 0.61437 0.07282

0.72662 0.05702 0.6652 0.06312

0.77166 0.04778 0.7148 0.05357

0.81427 0.03885 0.76246 0.04443

0.85381 0.03049 0.80747 0.03595

0.88967 0.02289 0.8491 0.02829

0.92127 0.01615 0.88664 0.02155

0.94818 0.01036 0.91947 0.01575

0.97004 0.00571 0.94698 0.01076

0.98634 0.00242 0.96898 0.00634

0.99652 0.00057 0.98556 0.00275

1 0 0.99625 0.00062

1 0

-0.05

0

0.05

0.1

0.15

0 0.2 0.4 0.6 0.8 1

Y - A x i s

X-Axis

SD7062 Airfoil

-0.1

0

0.1

0.2

0 0.2 0.4 0.6 0.8 1 Y - A x i s

X-Axis

GOE 386 Airfoil

-0.05

0

0.05

0.1

0.15

0 0.2 0.4 0.6 0.8 1 Y - A x i s

Eppler E1212 Airfoil



-0.1

.

X-Axis

-0.04-0.02

00.020.040.060.080.1

0 0.2 0.4 0.6 0.8 1

Y

X

SD 7032



SD7032

1 0

0.9967 0.0003

0.98684 0.00113

0.97054 0.00226

0.94797 0.0035

0.91942 0.00458

0.88534 0.00526

0.84635 0.00535

0.80313 0.00485

0.75634 0.00379

0.70664 0.00224

0.65469 0.0003

0.60112 -0.0019

0.54659 -0.0043

0.49176 -0.00678

0.43724 -0.00922

0.38364 -0.01152

0.33154 -0.01363

0.28153 -0.01547

0.2342 -0.01699

0.1901 -0.0181

0.14974 -0.01867

0.11351 -0.01862

0.0818 -0.01787

0.05491 -0.01635

0.03308 -0.01403

0.01649 -0.01088

0.00532 -0.007010.00038 -0.00223

0.00115 0.00448

0.00606 0.01293

0.01502 0.02206

0.02812 0.03145

0.04524 0.04078

0.06627 0.04976

0.09105 0.05809

0.11948 0.06548

0.15146 0.07182

0.18671 0.07703

0.22499 0.080960.26604 0.08359

0.30953 0.08493

0.35506 0.085

0.40222 0.08385

0.45058 0.08154

0.49967 0.07816

0.54902 0.07381



0.59812 0.06861

0.64646 0.0627

0.69356 0.0562

0.73892 0.04925

0.78208 0.04199

0.82264 0.0346

0.86021 0.02731

0.89436 0.02041

0.92464 0.0142

0.95054 0.00894

0.97155 0.00485

0.98712 0.00204

0.99674 0.00048

1 0

1.2

1.2

1.2



1.2



GOE 386 Eppler

Alpha CL CM CD Alpha CL CM CD

-4 0.1261 -0.106 0.01679 -4 -0.1071 -0.063 0.01567

-2 0.3416 -0.104 0.0158 -2 0.1133 -0.06 0.01548

0 0.5625 -0.103 0.016 0 0.3331 -0.056 0.0143

2 0.7709 -0.098 0.01471 2 0.5504 -0.055 0.01416

4 1.0046 -0.101 0.01528 4 0.7794 -0.054 0.01406

6 1.1362 -0.064 0.01688 6 0.9797 -0.05 0.01651

8 1.2801 -0.073 0.02051 8 1.1709 -0.045 0.01996

10 1.4267 -0.066 0.02524 10 1.3431 -0.038 0.02316

12 1.5611 -0.059 0.03114 12 0.6936 -0.025 0.03033

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

-5 0 5 10 15

C

L

Alpfa

CL vs. Alpfa GOE386

-0.2

0

0.2

0.40.6

0.8

1

1.2

1.4

1.6

-5 0 5 10 15

C

L

Alpha

Cl vs. Alpfa Eppler1212

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

0 0.01 0.02 0.03 0.04

C

L

CD

Cl vs CD

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

0 0.01 0.02 0.03 0.04

C L

CD

CL vs. CD



SD7062 SD7032


-4 0.0192 -0.091 0.01735 -4 0.0128 -0.102 0.0193

-2 0.2331 -0.087 0.01115 -2 0.2376 -0.1 0.01026

0 0.4405 -0.082 0.01029 0 0.4478 -0.093 0.00815

2 0.6703 -0.081 0.01069 2 0.6652 -0.091 0.00877

4 0.884 -0.079 0.01198 4 0.8749 -0.088 0.01035

6 1.089 -0.076 0.01402 6 1.076 -0.085 0.01268

8 1.2754 -0.071 0.01703 8 1.2475 -0.079 0.01669

10 1.4319 -0.063 0.02059 10 1.388 -0.069 0.02232

12 1.5274 -0.049 0.02681 12 1.418 -0.051 0.03535

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

-5 0 5 10 15

C

L

Alpha

Cl vs. Alpha SD7062

0

0.2

0.4

0.60.8

1

1.2

1.4

1.6

1.8

0 0.01 0.02 0.03

C

L

CD

CL vs. CD

-0.2

0

0.2

0.40.6

0.8

1

1.2

1.4

1.6

-5 0 5 10 15

C L

Alpha

Cl vs. Alpha SD7062

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

0 0.01 0.02 0.03 0.04

C L

CD

CL vs. CD



GOE 386 Eppler


-4 0.122 -0.104 0.01414 -4 -0.1118 -0.062 0.01164

-2 0.3369 -0.102 0.01263 -2 0.1117 -0.06 0.01103

0 0.5609 -0.101 0.01212 0 0.3381 -0.056 0.01106

2 0.7738 -0.099 0.01147 2 0.5602 -0.056 0.01101

4 0.8595 -0.093 0.01075 4 0.7718 -0.053 0.01099

6 1.1291 -0.063 0.01396 6 0.9969 -0.052 0.01216

8 1.2816 -0.074 0.01745 8 1.1994 -0.048 0.01384

10 1.4226 -0.066 0.02222 10 1.3838 -0.043 0.01647

12 1.5441 -0.059 0.02879 12 1.5151 -0.031 0.02039

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

-5 0 5 10 15

C L

Alpfa

CL vs. Alpfa GOE386

-0.2

0

0.2

0.4

0.60.8

1

1.2

1.4

1.6

-5 0 5 10 15

C L

Alpha

Cl vs. Alpfa Eppler1212

0

0.2

0.4

0.6

0.8

11.2

1.4

1.6

1.8

0 0.01 0.02 0.03 0.04

C L

CD

Cl vs CD

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

0 0.01 0.02 0.03

C L

CD

CL vs. CD



SD7062 SD7032


-4 0.1747 -0.062 0.14676 -4 0.0201 -0.099 0.0125

-2 0.1217 -0.003 0.08059 -2 0.2379 -0.096 0.00641

0 0.458 -0.085 0.00622 0 0.4433 -0.091 0.00604

2 0.663 -0.079 0.00605 2 0.6676 -0.091 0.00674

4 0.8938 -0.081 0.00924 4 0.8812 -0.09 0.00827

6 1.1049 -0.079 0.01089 6 1.0858 -0.088 0.01045

8 1.2986 -0.075 0.01328 8 1.2549 -0.082 0.01565

10 1.4668 -0.068 0.01621 10 1.398 -0.072 0.02109

12 1.5805 -0.055 0.0212 12 1.479 -0.056 0.02936

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

-5 0 5 10 15

C L

Alpha

Cl vs. Alpha SD7062

0

0.2

0.4

0.6

0.8

11.2

1.4

1.6

1.8

0 0.05 0.1 0.15 0.2

C L

CD

CL vs. CD

-0.2

0

0.2

0.40.6

0.8

1

1.2

1.4

1.6

-5 0 5 10 15

C L

Alpha

Cl vs. Alpha SD7062

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

0 0.01 0.02 0.03 0.04

C L

CD

CL vs. CD

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Light Aircraft Aircraft Design