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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
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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
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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
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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
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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
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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
W.L. HZ mom. VT mom.
6.0 269 135
0 0 0
0 0 0
0 0 0
3 12 3
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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)
W.L. HZ mom. VT mom.
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)
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lbs at .7Vto
lbs at Vto
lbs at .7Vto
lbs at Vto
lbs at .7Vto
lbs at Vtoft
ft
ft
ft
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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
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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
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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
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-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
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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
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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
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1.2
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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
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SD7062 SD7032
Alpha CL CM CD Alpha CL CM CD
-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
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GOE 386 Eppler
Alpha CL CM CD Alpha CL CM CD
-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
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SD7062 SD7032
Alpha CL CM CD Alpha CL CM CD
-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