WORTMANN FX 66-17AII-182 AIRFOIL (AS TESTED AT NASA) (fx6617ai-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: WORTMANN FX 66-17AII-182 AIRFOIL (AS TESTED AT NASA) (fx6617ai-il) Reynolds number: 50,000 Max Cl/Cd: 4.72 at α=9.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx6617ai-il-50000.txt Download as CSV file: xf-fx6617ai-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: WORTMANN FX 66-17AII-182 AIRFOIL (AS TESTED AT N
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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.1878 0.11617 0.11074 -0.0613 0.9508 0.3202
-10.000 -0.1656 0.11105 0.10553 -0.0661 0.9413 0.3223
-9.750 -0.2607 0.09002 0.08461 -0.0885 0.9351 0.1573
-9.500 -0.3809 0.07448 0.06894 -0.1041 0.9204 0.1322
-9.250 -0.3791 0.06958 0.06391 -0.1071 0.9110 0.1297
-9.000 -0.3961 0.06606 0.06020 -0.1076 0.8998 0.1273
-8.750 -0.4129 0.06233 0.05613 -0.1081 0.8909 0.1247
-8.500 -0.4264 0.05930 0.05269 -0.1073 0.8818 0.1226
-8.250 -0.4324 0.05674 0.04968 -0.1061 0.8739 0.1214
-8.000 -0.4285 0.05435 0.04691 -0.1052 0.8665 0.1215
-7.750 -0.4133 0.05167 0.04379 -0.1054 0.8610 0.1219
-7.500 -0.4141 0.05050 0.04235 -0.1027 0.8541 0.1219
-7.250 -0.3928 0.04828 0.03976 -0.1028 0.8492 0.1229
-7.000 -0.3756 0.04678 0.03799 -0.1020 0.8442 0.1248
-6.750 -0.3700 0.04615 0.03717 -0.0996 0.8384 0.1276
-6.500 -0.3462 0.04493 0.03545 -0.0995 0.8338 0.1332
-6.250 -0.3110 0.04327 0.03393 -0.1003 0.8301 0.1421
-6.000 -0.3149 0.04356 0.03413 -0.0963 0.8261 0.1457
-5.750 -0.3058 0.04327 0.03394 -0.0937 0.8225 0.1530
-5.500 -0.2883 0.04272 0.03347 -0.0922 0.8190 0.1682
-5.250 -0.2630 0.04159 0.03255 -0.0920 0.8157 0.1975
-5.000 -0.2657 0.04113 0.03245 -0.0888 0.8133 0.2296
-4.750 -0.2759 0.03965 0.03270 -0.0848 0.8124 0.3697
-4.500 -0.2879 0.04333 0.03694 -0.0743 0.8119 0.5509
-4.250 -0.2911 0.04628 0.03981 -0.0664 0.8114 0.5920
-4.000 -0.2920 0.04855 0.04199 -0.0596 0.8111 0.6223
-3.750 -0.2913 0.05041 0.04374 -0.0534 0.8117 0.6501
-3.500 -0.2921 0.05251 0.04581 -0.0455 0.8136 0.6830
-3.250 -0.4713 0.04285 0.03408 -0.0420 0.9998 0.2130
-3.000 -0.4619 0.04324 0.03712 -0.0359 0.9998 0.5840
-2.750 -0.4615 0.04559 0.03940 -0.0278 0.9998 0.6364
-2.500 -0.4580 0.04716 0.04085 -0.0213 0.9998 0.6721
-2.250 -0.4543 0.04829 0.04188 -0.0153 0.9998 0.7031
-2.000 -0.4557 0.04928 0.04282 -0.0074 0.9998 0.7420
-1.750 -0.4618 0.04976 0.04330 0.0019 0.9998 0.7806
-1.500 -0.4639 0.04985 0.04330 0.0094 0.9998 0.8185
-1.250 -0.4606 0.04972 0.04303 0.0147 0.9998 0.8488
-1.000 -0.4499 0.04956 0.04269 0.0173 0.9998 0.8692
-0.750 -0.4336 0.04956 0.04250 0.0179 0.9998 0.8861
-0.500 -0.4157 0.04967 0.04241 0.0180 0.9998 0.9029
-0.250 -0.3906 0.05019 0.04271 0.0165 0.9977 0.9200
0.000 -0.3509 0.05158 0.04386 0.0117 0.9903 0.9358
0.250 -0.3028 0.05363 0.04566 0.0048 0.9822 0.9492
0.500 -0.2530 0.05564 0.04743 -0.0027 0.9708 0.9606
0.750 -0.2069 0.05735 0.04894 -0.0094 0.9580 0.9710
1.000 -0.1613 0.05911 0.05052 -0.0161 0.9455 0.9803
1.250 -0.1173 0.06099 0.05223 -0.0227 0.9343 0.9911
1.500 -0.0778 0.06311 0.05418 -0.0284 0.9245 1.0002
2.000 -0.0484 0.06369 0.05454 -0.0297 0.8968 1.0002
2.250 -0.0304 0.06442 0.05514 -0.0311 0.8853 1.0002
2.500 0.0104 0.06724 0.05780 -0.0367 0.8784 1.0002
2.750 0.0284 0.06771 0.05818 -0.0379 0.8653 1.0002
3.000 0.0502 0.06891 0.05928 -0.0399 0.8544 1.0002
3.250 0.0922 0.07189 0.06212 -0.0455 0.8474 1.0002
3.500 0.1077 0.07253 0.06270 -0.0463 0.8354 1.0002
3.750 0.1364 0.07471 0.06477 -0.0494 0.8278 1.0002
4.000 0.1669 0.07668 0.06667 -0.0527 0.8172 1.0002
4.250 0.1835 0.07792 0.06785 -0.0537 0.8071 1.0002
4.500 0.2223 0.08098 0.07081 -0.0582 0.8002 1.0002
4.750 0.2338 0.08180 0.07160 -0.0583 0.7889 1.0002
5.000 0.2764 0.08560 0.07531 -0.0633 0.7834 1.0002
5.250 0.2819 0.08587 0.07556 -0.0624 0.7715 1.0002
5.500 0.3140 0.08908 0.07871 -0.0657 0.7663 1.0002
5.750 0.3268 0.09002 0.07962 -0.0658 0.7547 1.0002
6.000 0.3426 0.09191 0.08148 -0.0666 0.7470 1.0002
6.250 0.3679 0.09418 0.08372 -0.0684 0.7384 1.0002
6.500 0.3801 0.09592 0.08544 -0.0687 0.7308 1.0002
6.750 0.4057 0.09834 0.08784 -0.0705 0.7223 1.0002
7.000 0.4152 0.09996 0.08947 -0.0704 0.7144 1.0002
7.250 0.4395 0.10244 0.09194 -0.0720 0.7067 1.0002
7.500 0.4501 0.10434 0.09384 -0.0720 0.6999 1.0002
7.750 0.4709 0.10657 0.09608 -0.0731 0.6912 1.0002
8.000 0.4851 0.10895 0.09848 -0.0736 0.6854 1.0002
8.250 0.4988 0.11066 0.10022 -0.0739 0.6763 1.0002
8.500 0.5350 0.11522 0.10481 -0.0769 0.6724 1.0002
8.750 0.5256 0.11488 0.10450 -0.0746 0.6615 1.0002
9.000 0.5583 0.11900 0.10868 -0.0771 0.6570 1.0002
9.250 0.5508 0.11924 0.10896 -0.0753 0.6476 1.0002
9.500 0.5790 0.12278 0.11256 -0.0771 0.6419 1.0002
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Polar data table (+)
Polar graphs
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