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- 9/6/2010
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Document Sample
Aircraft Name: USAGE
Version No: v 1.0 Yellow is for user input
Date Modified: Blue is for output (formulae)
Designed by:
Contact:
Aircraft Parameters:
A/C Gross Weight: 5 lbs Flaps up Clmax: 1.4
Desired stall speed: 20 knots, flaps up Flaps down Clmax: 1.8
Desired stall speed: 15 knots, flaps down
Min Wing Area = 2.6 sq ft, flaps up Min Wing Area = 3.6
Flat plate area: 0.50 sq ft Max horsepower: 0.1
Total wing area: 6.0 sq ft Max prop RPM: 20700
Wingspan: 6.0 ft (upper wingspan for a biplane or wingspan for a monoplane)
Lower wingspan: 0 ft (lower wingspan for a biplane. Enter 0 for a monoplane)
Wing gap: 0 ft (distance between upper and lower wing if the a/c is a biplane. Enter 0
estimated k1 = 1.00 engine bsfc:
biplane span factor 0.5
Rectangular wing: 1 1, yes, 0 no Max fuel (gal): 0.01
max fus width: 0.4 feet Reserve fuel: 2
fus max height: 0.4 feet Prop W.R.: 0.066
est airplane 'e'= 0.83 mu = 0.1
Propeller Diameter: 9 inches Vto/Vstall 1.15
Est Prop efficiency= 0.45
Prop efficiency: 0.78 ** iterate until equals estimated prop efficiency (then subtract .03 if
Estimated Sea Level Standard Day Performance:
Vmax = 27 mph = 24 knots
V best ROC = 17 mph = 15 knots
Vmax L/D = 17 mph = 15 knots
V min pwr = 13 mph = 11 knots
Vstall, clean = 15 mph = 13 knots
Vstall, flaps = 13 mph = 12 knots
Max range = 47 statute miles= 41 nautical miles
Wing loading= 0.83 lbs/sq ft
Power loading = 50.00 lbs/horsepower
Suggested Peak Efficiency 2 Blade Propeller Diameter = 8 inches
Estimated Takeoff and Landing Performance:
T.O. distance = 55 ft, fixed pitch prop
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 Factor: 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 Wing incidence:
% chord (for cantilever wing) 1.0
Fuselage Length: 2 Wing
ft (from firewall aft) taper ratio: 1
# front passengers: 0
# rear passengers: 0
Fuselage construction: 1 1 - tube & fabric, 2 - metal, 3 - composite
Wing construction: 1 1 - fabric external brace, 2 - metal external brace, 3 - metal cantilever, 4
Tail construction: 1 1 - fabric external brace, 2 - metal cantilever, 3 - composite cantilever
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 structure
Welded steel tube & fabric covered fuselage 0.4 Fabric covered external braced wi
Metal fuselage 0.4 Metal covered external braced win
Composite fuselage 0.4 Metal cantilever wing
Cockpit controls 0.0 Composite cantilever wing
Control system 0.0
Seat (per passenger) 0.0 Propulsion group
Interior 0.0 Cowling
Paint (complete airplane) 0.0 Engine mount
Fuel system components
Tail structure Fuel tanks
Fabric covered external braced hor. tail 1.2 Wooden fixed pitch propeller
Metal horizontal tail 1.3 Metal fixed pitch propeller
Composite horizontal tail 1.5
Fabric covered external braced vert. tail 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 leg
Electrical 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 (lbs)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 wheels 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 wheel 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 controls 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 system 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 gasoline) 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 gasoline) 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 below wing)
Horizontal angle between wheel and aft fuselage (or tailwheel) = 14.1 degrees (min. angle req'd to keep
Vertical angle between CG and main gear axle (taildragger) = 25.0 degrees ( 13 deg minimum at forw
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 gasoline) 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 below wing)
Horizontal angle between wheel and aft fuselage (or tailwheel) = 14.1 degrees (min. angle req'd to keep
Vertical angle between CG and main gear axle (taildragger) = 25.0 degrees ( 13 deg minimum at forw
Vertical angle between CG and main gear axle (tricycle gear) = N/A degrees (must be greater than the
Min. vertical angle between CG and main gear axle (tricycle gear) = N/A degrees (minimum req'd angle to k
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
pan for a monoplane) Lt cnsspd = 56.19
for a monoplane) lamda cnsspd= 16.79
ng if the a/c is 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)
efficiency (then 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 = 0
Cl 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
al brace, 3 - metal cantilever, 4 - composite cantilever
ever, 3 - composite cantilever
ng structure Weight Estimate (lbs)
bric covered external braced wing 5.9
etal covered external braced wing 6.5
etal cantilever wing 1.1
mposite cantilever wing 1.3
opulsion group
0.0
0.0
el system components 0.0
0.7
ooden fixed pitch propeller 1.5
etal fixed pitch 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
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
AC (negative below wing)
grees (min. angle req'd to keep tail from hitting ground on landing)
grees ( 13 deg 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
AC (negative below wing)
grees (min. angle req'd to keep tail from hitting ground on landing)
grees ( 13 deg minimum at forward CG location, about 25 deg at aft)
grees (must be greater than the angle below for a tricycle gear)
grees (minimum 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 Vto
ft
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 -0.01911
0.00027 -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 0.11031
0.25496 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
SD7062 Airfoil
0.15
0.1
Y-Axis
0.05
0
0 0.2 0.4 0.6 0.8 1
-0.05
X-Axis
GOE 386 Airfoil
0.2
0.1
Y-Axis
0
0 0.2 0.4 0.6 0.8 1
-0.1 X-Axis
Eppler E1212 Airfoil
0.15
0.1
Y-Axis
0.05
0
-0.05 0 0.2 0.4 0.6 0.8 1
-0.05 0 0.2 0.4 0.6 0.8 1
-0.1 X-Axis
SD 7032
0.1
0.08
0.06
0.04
Y
0.02
0
-0.02 0 0.2 0.4 0.6 0.8 1
-0.04
X
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 -0.00701
0.00038 -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 0.08096
0.26604 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
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
CL vs. Alpfa GOE386 Cl vs. Alpfa Eppler1212
1.8 1.6
1.6 1.4
1.4 1.2
1.2 1
1 0.8
CL
CL
0.8
0.6
0.6
0.4
0.4
0.2
0.2
0 0
-0.2 0 -5 -0.2 0 5 10 15
-5 5
Alpfa 10 15 Alpha
Cl vs CD CL vs. CD
1.8 1.6
1.6 1.4
1.4 1.2
1.2 1
1 0.8
CL
CL
0.8 0.6
0.6 0.4
0.4
0.2
0.2
0
0
-0.2 0 0.01 0.02 0.03 0.04
0 0.01 0.02 0.03 0.04 CD
CD
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
Cl vs. Alpha SD7062 Cl vs. Alpha SD7062
1.8 1.6
1.6 1.4
1.4 1.2
1.2 1
1 0.8
CL
CL
0.8
0.6
0.6
0.4
0.4
0.2 0.2
0 0
-5 -0.2 0 5 10 15 -5 -0.2 0 5 10 15
Alpha Alpha
CL vs. CD CL vs. CD
1.8 1.6
1.6 1.4
1.4 1.2
1.2 1
1
CL
0.8
CL
0.8
0.6
0.6
0.4
0.4
0.2 0.2
0 0
0 0.01 0 0.01 0.02 0.03 0.04
CD 0.02 0.03 CD
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
CL vs. Alpfa GOE386 Cl vs. Alpfa Eppler1212
1.8 1.6
1.6 1.4
1.4 1.2
1.2 1
1 0.8
CL
CL
0.8
0.6
0.6
0.4
0.4
0.2
0.2
0
0
-0.2 0 -5 -0.2 0 5 10 15
-5 5
Alpfa 10 15 Alpha
Cl vs CD CL vs. CD
1.8 1.6
1.6 1.4
1.4 1.2
1.2 1
1 0.8
CL
CL
0.8 0.6
0.6
0.4
0.4
0.2
0.2
0
0
-0.2 0 0.01 0.02 0.03
0 0.01 0.02 0.03 0.04 CD
CD
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
Cl vs. Alpha SD7062
Cl vs. Alpha SD7062 1.6
1.8 1.4
1.6 1.2
1.4 1
1.2
0.8
CL
1
0.6
CL
0.8
0.6 0.4
0.4 0.2
0.2 0
0 -5 -0.2 0 5 10 15
Alpha
-5 -0.2 0 5 10 15
Alpha
CL vs. CD
1.6
CL vs. CD 1.4
1.8 1.2
1.6 1
1.4
CL
0.8
1.2
0.6
1
CL
0.8 0.4
0.6 0.2
0.4 0
0.2 0 0.01 0.02 0.03 0.04
CD
0
0 0.05 0.1
CD 0.15 0.2
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