NASA SC(2)-0614 AIRFOIL (sc20614-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NASA SC(2)-0614 AIRFOIL (sc20614-il) Reynolds number: 100,000 Max Cl/Cd: 28.81 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-sc20614-il-100000.txt Download as CSV file: xf-sc20614-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0614 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
------ -------- --------- --------- -------- -------- --------
-12.000 -0.7085 0.07960 0.07334 -0.0618 1.0000 0.1014
-11.750 -0.7239 0.07444 0.06815 -0.0623 1.0000 0.1005
-11.500 -0.7526 0.06944 0.06310 -0.0621 1.0000 0.0995
-11.250 -0.7876 0.06545 0.05903 -0.0599 1.0000 0.0984
-11.000 -0.8214 0.06157 0.05502 -0.0577 1.0000 0.0974
-10.750 -0.8444 0.05765 0.05088 -0.0561 1.0000 0.0970
-10.500 -0.8577 0.05404 0.04701 -0.0545 1.0000 0.0972
-10.250 -0.8630 0.05063 0.04330 -0.0532 1.0000 0.0978
-10.000 -0.8621 0.04735 0.03967 -0.0522 1.0000 0.0986
-9.750 -0.8557 0.04423 0.03615 -0.0515 1.0000 0.0996
-9.500 -0.8443 0.04138 0.03286 -0.0511 1.0000 0.1008
-9.250 -0.8292 0.03881 0.02986 -0.0509 1.0000 0.1026
-9.000 -0.8115 0.03684 0.02796 -0.0502 1.0000 0.1054
-8.750 -0.7923 0.03551 0.02659 -0.0495 1.0000 0.1087
-8.500 -0.7719 0.03388 0.02470 -0.0492 1.0000 0.1122
-8.250 -0.7497 0.03238 0.02280 -0.0491 1.0000 0.1154
-8.000 -0.7298 0.03083 0.02141 -0.0483 1.0000 0.1196
-7.750 -0.7077 0.02982 0.02033 -0.0478 1.0000 0.1246
-7.500 -0.6838 0.02885 0.01908 -0.0477 1.0000 0.1296
-7.250 -0.6631 0.02761 0.01805 -0.0467 1.0000 0.1351
-7.000 -0.6395 0.02681 0.01717 -0.0464 1.0000 0.1416
-6.750 -0.6165 0.02584 0.01624 -0.0458 1.0000 0.1483
-6.500 -0.5930 0.02513 0.01559 -0.0454 1.0000 0.1563
-6.250 -0.5696 0.02433 0.01486 -0.0449 1.0000 0.1650
-6.000 -0.5450 0.02373 0.01428 -0.0447 1.0000 0.1762
-5.750 -0.5211 0.02299 0.01376 -0.0444 1.0000 0.1896
-5.500 -0.4955 0.02231 0.01327 -0.0444 1.0000 0.2092
-5.250 -0.4674 0.02142 0.01277 -0.0453 1.0000 0.2437
-5.000 -0.4259 0.01943 0.01258 -0.0499 1.0000 0.4443
-4.750 -0.4175 0.02082 0.01443 -0.0440 1.0000 0.5634
-4.500 -0.4012 0.02204 0.01559 -0.0406 1.0000 0.6016
-4.250 -0.3888 0.02327 0.01678 -0.0362 1.0000 0.6221
-4.000 -0.3771 0.02444 0.01793 -0.0317 1.0000 0.6384
-3.750 -0.3655 0.02553 0.01898 -0.0273 1.0000 0.6527
-3.500 -0.3533 0.02650 0.01992 -0.0231 1.0000 0.6667
-3.250 -0.3392 0.02729 0.02068 -0.0196 1.0000 0.6804
-3.000 -0.3233 0.02796 0.02131 -0.0168 1.0000 0.6942
-2.750 -0.3163 0.02868 0.02203 -0.0113 1.0000 0.7036
-2.500 -0.3043 0.02922 0.02256 -0.0075 1.0000 0.7159
-2.250 -0.2913 0.02969 0.02302 -0.0040 1.0000 0.7297
-2.000 -0.2787 0.03012 0.02345 -0.0005 1.0000 0.7445
-1.750 -0.2642 0.03049 0.02379 0.0023 1.0000 0.7595
-1.500 -0.2548 0.03069 0.02401 0.0064 1.0000 0.7724
-1.250 -0.2458 0.03078 0.02411 0.0106 1.0000 0.7860
-1.000 -0.2342 0.03081 0.02414 0.0139 1.0000 0.8002
-0.750 -0.2195 0.03079 0.02412 0.0161 1.0000 0.8137
-0.500 -0.1862 0.03104 0.02433 0.0139 0.9967 0.8240
-0.250 -0.1503 0.03103 0.02431 0.0120 0.9887 0.8305
0.000 -0.1030 0.03138 0.02463 0.0074 0.9805 0.8371
0.250 -0.0602 0.03149 0.02475 0.0032 0.9705 0.8422
0.500 -0.0261 0.03138 0.02465 0.0015 0.9603 0.8467
0.750 0.0188 0.03151 0.02480 -0.0026 0.9509 0.8511
1.000 0.0665 0.03162 0.02493 -0.0073 0.9395 0.8554
1.250 0.1180 0.03131 0.02466 -0.0117 0.9236 0.8588
1.500 0.1728 0.03069 0.02409 -0.0160 0.9066 0.8610
1.750 0.2207 0.03004 0.02351 -0.0193 0.8907 0.8631
2.000 0.2677 0.02927 0.02282 -0.0221 0.8747 0.8660
2.250 0.3224 0.02807 0.02171 -0.0256 0.8573 0.8686
2.500 0.3809 0.02658 0.02033 -0.0293 0.8399 0.8707
2.750 0.4354 0.02500 0.01887 -0.0321 0.8227 0.8731
3.000 0.4841 0.02313 0.01712 -0.0332 0.8015 0.8754
3.250 0.5189 0.02182 0.01588 -0.0326 0.7682 0.8772
3.500 0.5494 0.02081 0.01489 -0.0314 0.7126 0.8790
3.750 0.5871 0.02038 0.01342 -0.0306 0.4959 0.8807
4.000 0.5933 0.02205 0.01385 -0.0272 0.3204 0.8830
4.250 0.6098 0.02344 0.01453 -0.0260 0.2465 0.8855
4.500 0.6343 0.02453 0.01526 -0.0261 0.2104 0.8876
4.750 0.6585 0.02523 0.01583 -0.0258 0.1892 0.8895
5.000 0.6856 0.02604 0.01649 -0.0260 0.1744 0.8914
5.250 0.7149 0.02695 0.01720 -0.0267 0.1623 0.8934
5.500 0.7454 0.02774 0.01808 -0.0275 0.1526 0.8956
5.750 0.7785 0.02898 0.01915 -0.0290 0.1444 0.8978
6.000 0.8091 0.02985 0.02017 -0.0298 0.1370 0.9003
6.250 0.8423 0.03132 0.02144 -0.0314 0.1304 0.9030
6.500 0.8665 0.03205 0.02246 -0.0308 0.1253 0.9060
6.750 0.8938 0.03305 0.02353 -0.0310 0.1200 0.9088
7.000 0.9244 0.03484 0.02524 -0.0322 0.1153 0.9114
7.250 0.9498 0.03625 0.02706 -0.0321 0.1116 0.9144
7.500 0.9756 0.03769 0.02869 -0.0322 0.1077 0.9174
7.750 1.0004 0.03905 0.03002 -0.0323 0.1042 0.9207
8.000 1.0212 0.04144 0.03271 -0.0319 0.1016 0.9236
8.250 1.0384 0.04385 0.03562 -0.0307 0.0999 0.9264
8.500 1.0537 0.04661 0.03883 -0.0297 0.0982 0.9291
8.750 1.0656 0.04924 0.04184 -0.0282 0.0965 0.9320
9.000 1.0768 0.05159 0.04445 -0.0267 0.0947 0.9357
9.250 1.0923 0.05389 0.04689 -0.0261 0.0931 0.9398
9.500 1.0897 0.05803 0.05152 -0.0236 0.0933 0.9433
9.750 1.0597 0.06413 0.05834 -0.0191 0.0956 0.9464
10.000 1.0280 0.06959 0.06424 -0.0154 0.0977 0.9496
10.250 1.0020 0.07488 0.06982 -0.0134 0.0995 0.9531
10.500 0.9869 0.08017 0.07527 -0.0131 0.1010 0.9565
10.750 0.9986 0.08490 0.08002 -0.0135 0.1023 0.9608
11.000 0.7440 0.13761 0.13341 -0.0529 0.1657 0.9500
11.250 0.7696 0.14020 0.13604 -0.0510 0.1631 0.9545
|
Polar data table (+)
Polar graphs
<< Back to NASA SC(2)-0614 AIRFOIL (sc20614-il)