CAST 10-2/DOA 2 AIRFOIL (cast102-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: CAST 10-2/DOA 2 AIRFOIL (cast102-il) Reynolds number: 100,000 Max Cl/Cd: 45.83 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-cast102-il-100000.txt Download as CSV file: xf-cast102-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: CAST 10-2/DOA 2 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4062 0.10279 0.09819 -0.0365 1.0000 0.1139
-9.750 -0.4373 0.09900 0.09449 -0.0386 1.0000 0.1154
-9.500 -0.4694 0.09462 0.09019 -0.0407 1.0000 0.1158
-9.250 -0.4261 0.09293 0.08848 -0.0336 1.0000 0.1229
-9.000 -0.4397 0.08953 0.08513 -0.0332 1.0000 0.1268
-8.750 -0.4685 0.08511 0.08079 -0.0344 1.0000 0.1290
-8.500 -0.5011 0.08048 0.07621 -0.0353 1.0000 0.1295
-8.250 -0.5320 0.07686 0.07262 -0.0340 1.0000 0.1294
-8.000 -0.5627 0.07398 0.06975 -0.0311 1.0000 0.1292
-7.750 -0.5948 0.07091 0.06664 -0.0286 1.0000 0.1297
-7.500 -0.6323 0.06847 0.06400 -0.0260 1.0000 0.1309
-7.250 -0.6159 0.06316 0.05888 -0.0246 1.0000 0.1362
-7.000 -0.6233 0.06016 0.05583 -0.0225 1.0000 0.1421
-6.750 -0.6914 0.06716 0.06161 -0.0225 1.0000 0.1473
-6.500 -0.6743 0.06149 0.05638 -0.0213 1.0000 0.1529
-6.250 -0.6718 0.05852 0.05327 -0.0200 1.0000 0.1665
-6.000 -0.6652 0.05569 0.05038 -0.0186 1.0000 0.1823
-5.750 -0.6566 0.05305 0.04773 -0.0171 1.0000 0.2013
-5.000 -0.6324 0.04577 0.04036 -0.0120 1.0000 0.3019
-4.750 -0.6208 0.04304 0.03800 -0.0086 1.0000 0.3375
-4.500 -0.6057 0.04059 0.03559 -0.0067 1.0000 0.3681
-4.250 -0.5078 0.03571 0.02694 -0.0138 1.0000 0.1009
-4.000 -0.4787 0.03252 0.02341 -0.0131 1.0000 0.0847
-3.750 -0.4511 0.03026 0.02079 -0.0123 1.0000 0.0768
-3.500 -0.4255 0.02992 0.02003 -0.0110 1.0000 0.0733
-3.250 -0.3997 0.02822 0.01820 -0.0105 1.0000 0.0738
-3.000 -0.3743 0.02709 0.01697 -0.0099 1.0000 0.0737
-2.750 -0.3491 0.02620 0.01599 -0.0092 1.0000 0.0732
-2.500 -0.3236 0.02508 0.01490 -0.0087 1.0000 0.0736
-2.250 -0.2983 0.02402 0.01395 -0.0083 1.0000 0.0751
-2.000 -0.2736 0.02330 0.01336 -0.0082 1.0000 0.0799
-1.750 -0.2484 0.02288 0.01288 -0.0080 1.0000 0.0854
-1.500 -0.2219 0.02239 0.01234 -0.0082 1.0000 0.0914
-1.250 -0.1947 0.02204 0.01198 -0.0085 1.0000 0.1058
-1.000 -0.1686 0.02020 0.01290 -0.0074 1.0000 0.7450
-0.750 -0.1661 0.02075 0.01364 0.0001 1.0000 0.8169
-0.500 -0.1646 0.02106 0.01402 0.0074 1.0000 0.8618
-0.250 -0.1634 0.02119 0.01419 0.0144 1.0000 0.9019
0.000 -0.1548 0.02125 0.01423 0.0193 1.0000 0.9365
0.250 -0.1100 0.02190 0.01479 0.0167 1.0000 0.9807
0.500 -0.0618 0.02243 0.01521 0.0113 1.0000 0.9948
0.750 -0.0396 0.02242 0.01515 0.0104 1.0000 1.0000
1.000 -0.0334 0.02212 0.01482 0.0123 1.0000 1.0000
1.250 -0.0161 0.02215 0.01480 0.0122 1.0000 1.0000
1.500 0.0310 0.02284 0.01542 0.0066 0.9920 1.0000
1.750 0.0938 0.02369 0.01622 -0.0016 0.9759 1.0000
2.000 0.1474 0.02428 0.01677 -0.0078 0.9615 1.0000
2.250 0.1988 0.02469 0.01719 -0.0133 0.9469 1.0000
2.500 0.2498 0.02496 0.01747 -0.0184 0.9319 1.0000
2.750 0.3006 0.02505 0.01761 -0.0231 0.9166 1.0000
3.000 0.3516 0.02493 0.01755 -0.0275 0.9010 1.0000
3.250 0.4013 0.02462 0.01733 -0.0314 0.8857 1.0000
3.500 0.4502 0.02412 0.01693 -0.0348 0.8707 1.0000
3.750 0.4982 0.02341 0.01635 -0.0378 0.8556 1.0000
4.000 0.5483 0.02239 0.01550 -0.0406 0.8401 1.0000
4.250 0.6020 0.02091 0.01420 -0.0435 0.8237 1.0000
4.500 0.6343 0.01970 0.01316 -0.0427 0.7997 1.0000
4.750 0.6682 0.01818 0.01178 -0.0416 0.7683 1.0000
5.000 0.6914 0.01701 0.01069 -0.0387 0.7098 1.0000
5.250 0.7497 0.01636 0.00856 -0.0406 0.4555 1.0000
5.500 0.7653 0.01768 0.00918 -0.0386 0.3758 1.0000
5.750 0.7857 0.01877 0.00989 -0.0376 0.3305 1.0000
6.000 0.8088 0.01978 0.01065 -0.0370 0.2994 1.0000
6.250 0.8335 0.02079 0.01144 -0.0368 0.2758 1.0000
6.500 0.8575 0.02164 0.01228 -0.0365 0.2544 1.0000
6.750 0.8808 0.02248 0.01305 -0.0360 0.2350 1.0000
7.000 0.9018 0.02326 0.01372 -0.0353 0.2156 1.0000
7.250 0.9210 0.02391 0.01440 -0.0342 0.1966 1.0000
7.500 0.9415 0.02467 0.01521 -0.0333 0.1797 1.0000
7.750 0.9603 0.02544 0.01600 -0.0321 0.1618 1.0000
8.000 0.9779 0.02640 0.01689 -0.0308 0.1422 1.0000
8.250 0.9927 0.02757 0.01811 -0.0288 0.1179 1.0000
8.500 1.0101 0.02895 0.01933 -0.0275 0.1003 1.0000
8.750 1.0322 0.03051 0.02096 -0.0267 0.0889 1.0000
9.000 1.0557 0.03233 0.02296 -0.0261 0.0810 1.0000
9.250 1.0805 0.03427 0.02474 -0.0263 0.0748 1.0000
9.500 1.1022 0.03644 0.02735 -0.0252 0.0713 1.0000
9.750 1.1226 0.03880 0.03002 -0.0242 0.0683 1.0000
10.000 1.1403 0.04087 0.03228 -0.0231 0.0656 1.0000
10.250 1.1588 0.04417 0.03564 -0.0227 0.0630 1.0000
10.500 1.1629 0.04675 0.03877 -0.0195 0.0618 1.0000
10.750 1.1641 0.05007 0.04257 -0.0163 0.0611 1.0000
11.000 1.1601 0.05366 0.04660 -0.0128 0.0609 1.0000
11.250 1.1505 0.05738 0.05072 -0.0090 0.0609 1.0000
11.500 1.1346 0.06095 0.05461 -0.0047 0.0610 1.0000
11.750 1.1155 0.06476 0.05871 -0.0008 0.0612 1.0000
12.000 1.0941 0.06888 0.06309 0.0022 0.0615 1.0000
12.250 1.0714 0.07339 0.06782 0.0044 0.0619 1.0000
12.500 1.0481 0.07837 0.07299 0.0055 0.0623 1.0000
12.750 1.0254 0.08391 0.07867 0.0055 0.0627 1.0000
|
Polar data table (+)
Polar graphs
<< Back to CAST 10-2/DOA 2 AIRFOIL (cast102-il)