CLARK W AIRFOIL (clarkw-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: CLARK W AIRFOIL (clarkw-il) Reynolds number: 50,000 Max Cl/Cd: 32.22 at α=9.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-clarkw-il-50000.txt Download as CSV file: xf-clarkw-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: CLARK W AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3420 0.10855 0.10124 -0.0271 1.0000 0.2455
-8.500 -0.3453 0.10676 0.09954 -0.0262 1.0000 0.2544
-8.250 -0.3687 0.10618 0.09913 -0.0254 1.0000 0.2582
-8.000 -0.3437 0.10153 0.09444 -0.0237 1.0000 0.2685
-7.750 -0.3824 0.10263 0.09578 -0.0215 1.0000 0.2734
-7.500 -0.3524 0.09709 0.09020 -0.0203 1.0000 0.2813
-7.250 -0.3781 0.09688 0.09014 -0.0173 1.0000 0.2887
-7.000 -0.3967 0.09527 0.08869 -0.0143 1.0000 0.2925
-6.750 -0.3832 0.09191 0.08534 -0.0123 1.0000 0.3010
-6.500 -0.4349 0.09319 0.08684 -0.0097 1.0000 0.3072
-6.250 -0.4098 0.08835 0.08200 -0.0067 1.0000 0.3135
-6.000 -0.4443 0.08831 0.08211 -0.0054 1.0000 0.3237
-5.500 -0.4541 0.08326 0.07719 -0.0018 1.0000 0.3414
-5.250 -0.4409 0.07967 0.07363 0.0018 1.0000 0.3482
-4.500 -0.4163 0.05481 0.04731 -0.0317 1.0000 0.1768
-4.250 -0.3999 0.05166 0.04405 -0.0315 1.0000 0.1737
-4.000 -0.3801 0.04782 0.03993 -0.0326 1.0000 0.1690
-3.750 -0.3495 0.04165 0.03253 -0.0366 1.0000 0.1604
-3.500 -0.3255 0.03888 0.02936 -0.0371 1.0000 0.1610
-3.250 -0.3048 0.03747 0.02800 -0.0366 1.0000 0.1656
-3.000 -0.2806 0.03576 0.02592 -0.0368 1.0000 0.1723
-2.750 -0.2553 0.03399 0.02378 -0.0370 1.0000 0.1793
-2.500 -0.2309 0.03273 0.02229 -0.0369 1.0000 0.1894
-2.250 -0.2069 0.03178 0.02117 -0.0367 1.0000 0.2068
-2.000 -0.1843 0.03111 0.02061 -0.0363 1.0000 0.2314
-1.750 -0.1622 0.03073 0.02028 -0.0357 1.0000 0.2672
-1.500 -0.1424 0.03072 0.02035 -0.0350 1.0000 0.3103
-1.250 -0.1227 0.03066 0.02040 -0.0343 1.0000 0.3540
-1.000 -0.1003 0.03026 0.02002 -0.0339 1.0000 0.3936
-0.750 -0.0753 0.02988 0.01959 -0.0341 1.0000 0.4393
-0.500 -0.0496 0.02938 0.01930 -0.0344 1.0000 0.4942
-0.250 -0.0228 0.02872 0.01917 -0.0347 1.0000 0.5757
0.000 -0.0080 0.02750 0.01927 -0.0326 0.9989 1.0000
0.250 0.0583 0.02916 0.02014 -0.0413 0.9833 1.0000
0.500 0.1083 0.03037 0.02096 -0.0467 0.9663 1.0000
0.750 0.1546 0.03146 0.02176 -0.0513 0.9490 1.0000
1.000 0.1997 0.03246 0.02254 -0.0554 0.9316 1.0000
1.250 0.2437 0.03338 0.02328 -0.0591 0.9142 1.0000
1.500 0.2870 0.03420 0.02396 -0.0625 0.8969 1.0000
1.750 0.3297 0.03493 0.02458 -0.0656 0.8796 1.0000
2.000 0.3715 0.03555 0.02512 -0.0684 0.8626 1.0000
2.250 0.4064 0.03614 0.02565 -0.0699 0.8446 1.0000
2.500 0.4373 0.03673 0.02621 -0.0707 0.8262 1.0000
2.750 0.4698 0.03726 0.02672 -0.0716 0.8084 1.0000
3.000 0.5028 0.03775 0.02721 -0.0724 0.7912 1.0000
3.250 0.5358 0.03818 0.02764 -0.0731 0.7744 1.0000
3.500 0.5682 0.03860 0.02807 -0.0736 0.7582 1.0000
3.750 0.6002 0.03895 0.02846 -0.0740 0.7423 1.0000
4.000 0.6323 0.03927 0.02880 -0.0742 0.7271 1.0000
4.250 0.6695 0.03928 0.02887 -0.0748 0.7134 1.0000
4.500 0.7150 0.03874 0.02842 -0.0761 0.7014 1.0000
4.750 0.7387 0.03931 0.02903 -0.0752 0.6863 1.0000
5.000 0.7615 0.03993 0.02972 -0.0741 0.6714 1.0000
5.250 0.7824 0.04073 0.03060 -0.0730 0.6569 1.0000
5.500 0.8036 0.04154 0.03148 -0.0720 0.6429 1.0000
5.750 0.8286 0.04213 0.03214 -0.0712 0.6299 1.0000
6.000 0.8836 0.04058 0.03075 -0.0723 0.6207 1.0000
6.250 0.8958 0.04200 0.03225 -0.0705 0.6061 1.0000
6.500 0.9059 0.04368 0.03402 -0.0686 0.5919 1.0000
6.750 0.9201 0.04504 0.03549 -0.0670 0.5780 1.0000
7.000 0.9449 0.04549 0.03606 -0.0658 0.5646 1.0000
7.250 0.9847 0.04444 0.03514 -0.0650 0.5497 1.0000
7.500 1.0235 0.04334 0.03419 -0.0641 0.5333 1.0000
7.750 1.0582 0.04266 0.03364 -0.0630 0.5165 1.0000
8.000 1.0910 0.04220 0.03329 -0.0619 0.4990 1.0000
8.250 1.1326 0.04104 0.03224 -0.0613 0.4792 1.0000
8.500 1.1694 0.04008 0.03129 -0.0602 0.4562 1.0000
8.750 1.2103 0.03849 0.02958 -0.0592 0.4266 1.0000
9.000 1.2199 0.03894 0.03012 -0.0557 0.3997 1.0000
9.250 1.2380 0.03877 0.02988 -0.0528 0.3688 1.0000
9.500 1.2519 0.03886 0.02985 -0.0495 0.3363 1.0000
9.750 1.2614 0.03925 0.03008 -0.0459 0.3030 1.0000
10.000 1.2663 0.04007 0.03074 -0.0419 0.2713 1.0000
10.250 1.2683 0.04137 0.03191 -0.0380 0.2432 1.0000
10.500 1.2745 0.04282 0.03313 -0.0347 0.2188 1.0000
10.750 1.2786 0.04474 0.03504 -0.0316 0.1996 1.0000
11.000 1.2867 0.04673 0.03699 -0.0291 0.1836 1.0000
11.250 1.2954 0.04900 0.03930 -0.0268 0.1710 1.0000
11.500 1.3089 0.05147 0.04177 -0.0253 0.1601 1.0000
11.750 1.3312 0.05401 0.04417 -0.0249 0.1493 1.0000
12.000 1.3182 0.05709 0.04772 -0.0208 0.1459 1.0000
12.250 1.3142 0.06020 0.05109 -0.0181 0.1413 1.0000
12.500 1.3298 0.06329 0.05415 -0.0173 0.1348 1.0000
12.750 1.3075 0.06704 0.05827 -0.0138 0.1337 1.0000
13.000 1.2825 0.07134 0.06291 -0.0112 0.1330 1.0000
13.250 1.2537 0.07637 0.06824 -0.0096 0.1329 1.0000
13.500 1.2205 0.08236 0.07449 -0.0092 0.1334 1.0000
13.750 1.1847 0.08948 0.08183 -0.0103 0.1344 1.0000
14.000 1.1490 0.09775 0.09024 -0.0128 0.1354 1.0000
|
Polar data table (+)
Polar graphs
<< Back to CLARK W AIRFOIL (clarkw-il)