GOE 506 AIRFOIL (goe506-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 506 AIRFOIL (goe506-il) Reynolds number: 1,000,000 Max Cl/Cd: 81.1 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe506-il-1000000.txt Download as CSV file: xf-goe506-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 506 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-16.250 -0.5535 0.08664 0.08444 -0.0959 0.9905 0.0120
-16.000 -0.5812 0.07682 0.07446 -0.1032 0.9884 0.0122
-15.750 -0.6083 0.06773 0.06519 -0.1105 0.9863 0.0122
-15.500 -0.6318 0.05934 0.05661 -0.1177 0.9843 0.0121
-15.250 -0.6521 0.05214 0.04924 -0.1235 0.9799 0.0121
-15.000 -0.6494 0.04784 0.04484 -0.1280 0.9759 0.0122
-14.750 -0.6482 0.04282 0.03967 -0.1338 0.9728 0.0125
-14.500 -0.6438 0.03832 0.03502 -0.1393 0.9688 0.0127
-14.250 -0.6380 0.03418 0.03071 -0.1445 0.9631 0.0128
-14.000 -0.6185 0.03146 0.02784 -0.1495 0.9597 0.0131
-13.750 -0.6005 0.02862 0.02483 -0.1545 0.9556 0.0132
-13.500 -0.5860 0.02546 0.02146 -0.1593 0.9499 0.0134
-13.250 -0.5621 0.02347 0.01935 -0.1633 0.9454 0.0138
-13.000 -0.5358 0.02243 0.01824 -0.1658 0.9406 0.0141
-12.750 -0.5152 0.02163 0.01737 -0.1666 0.9345 0.0143
-12.500 -0.4927 0.02076 0.01641 -0.1676 0.9292 0.0146
-12.250 -0.4783 0.02018 0.01574 -0.1665 0.9234 0.0149
-12.000 -0.4645 0.01967 0.01514 -0.1650 0.9176 0.0154
-11.750 -0.4477 0.01905 0.01442 -0.1641 0.9125 0.0157
-11.500 -0.4410 0.01851 0.01381 -0.1609 0.9070 0.0159
-11.250 -0.4317 0.01800 0.01321 -0.1580 0.9015 0.0161
-11.000 -0.4208 0.01735 0.01243 -0.1555 0.8961 0.0166
-10.750 -0.4183 0.01683 0.01189 -0.1512 0.8909 0.0170
-10.500 -0.4057 0.01642 0.01143 -0.1487 0.8856 0.0172
-10.250 -0.3851 0.01620 0.01116 -0.1476 0.8809 0.0177
-10.000 -0.3697 0.01587 0.01079 -0.1455 0.8766 0.0181
-9.750 -0.3534 0.01550 0.01037 -0.1435 0.8722 0.0185
-9.500 -0.3338 0.01522 0.01003 -0.1421 0.8680 0.0192
-9.250 -0.3136 0.01487 0.00959 -0.1408 0.8633 0.0196
-9.000 -0.2942 0.01468 0.00936 -0.1392 0.8587 0.0199
-8.750 -0.2830 0.01385 0.00846 -0.1364 0.8537 0.0206
-8.500 -0.2613 0.01357 0.00815 -0.1354 0.8492 0.0213
-8.250 -0.2408 0.01333 0.00790 -0.1341 0.8446 0.0219
-8.000 -0.2204 0.01305 0.00759 -0.1327 0.8393 0.0226
-7.750 -0.1983 0.01278 0.00725 -0.1316 0.8342 0.0234
-7.500 -0.1769 0.01255 0.00698 -0.1303 0.8287 0.0239
-7.250 -0.1560 0.01226 0.00664 -0.1290 0.8231 0.0244
-7.000 -0.1399 0.01162 0.00593 -0.1269 0.8180 0.0254
-6.750 -0.1185 0.01136 0.00566 -0.1257 0.8129 0.0264
-6.500 -0.0962 0.01113 0.00540 -0.1246 0.8072 0.0272
-6.000 -0.0511 0.01068 0.00487 -0.1224 0.7959 0.0288
-5.750 -0.0274 0.01055 0.00469 -0.1216 0.7896 0.0293
-5.500 -0.0087 0.01006 0.00413 -0.1198 0.7838 0.0308
-5.250 0.0137 0.00982 0.00386 -0.1187 0.7779 0.0319
-5.000 0.0370 0.00964 0.00362 -0.1177 0.7726 0.0330
-4.750 0.0608 0.00947 0.00342 -0.1169 0.7677 0.0340
-4.500 0.0846 0.00933 0.00324 -0.1160 0.7618 0.0350
-4.250 0.1071 0.00917 0.00301 -0.1148 0.7544 0.0364
-4.000 0.1299 0.00899 0.00280 -0.1137 0.7475 0.0386
-3.750 0.1523 0.00888 0.00264 -0.1125 0.7395 0.0407
-3.500 0.1755 0.00876 0.00250 -0.1115 0.7319 0.0441
-3.250 0.1941 0.00840 0.00227 -0.1096 0.7244 0.0894
-3.000 0.2108 0.00787 0.00213 -0.1076 0.7171 0.1956
-2.500 0.2560 0.00777 0.00205 -0.1053 0.7006 0.2228
-2.250 0.2779 0.00776 0.00201 -0.1040 0.6918 0.2323
-2.000 0.2992 0.00773 0.00198 -0.1026 0.6799 0.2406
-1.750 0.3190 0.00774 0.00194 -0.1008 0.6647 0.2470
-1.500 0.3345 0.00778 0.00192 -0.0981 0.6402 0.2545
-1.250 0.3249 0.00812 0.00195 -0.0901 0.5676 0.2582
-1.000 0.3212 0.00852 0.00209 -0.0835 0.5135 0.2635
-0.750 0.3309 0.00873 0.00220 -0.0797 0.4857 0.2737
-0.500 0.3451 0.00889 0.00231 -0.0770 0.4630 0.2871
-0.250 0.3584 0.00906 0.00242 -0.0740 0.4397 0.3049
0.000 0.3725 0.00923 0.00254 -0.0713 0.4143 0.3229
0.250 0.3852 0.00943 0.00268 -0.0683 0.3890 0.3482
0.500 0.4001 0.00957 0.00281 -0.0658 0.3712 0.3744
0.750 0.4155 0.00967 0.00294 -0.0634 0.3579 0.4089
1.000 0.4326 0.00970 0.00306 -0.0614 0.3491 0.4498
1.250 0.4487 0.00975 0.00319 -0.0591 0.3414 0.4926
1.500 0.4647 0.00969 0.00332 -0.0568 0.3361 0.5616
1.750 0.4779 0.00962 0.00348 -0.0540 0.3308 0.6471
2.000 0.4867 0.00947 0.00366 -0.0501 0.3263 0.7552
2.250 0.6395 0.00944 0.00407 -0.0777 0.3167 0.9536
2.500 0.6871 0.00977 0.00433 -0.0823 0.3116 0.9793
2.750 0.7230 0.00997 0.00451 -0.0843 0.3090 0.9902
3.000 0.7611 0.01019 0.00470 -0.0867 0.3066 0.9966
3.250 0.8034 0.01039 0.00486 -0.0902 0.3034 0.9995
3.500 0.8264 0.01056 0.00500 -0.0895 0.3007 1.0000
3.750 0.8408 0.01072 0.00515 -0.0870 0.2984 1.0000
4.000 0.8551 0.01092 0.00532 -0.0845 0.2954 1.0000
4.250 0.8725 0.01107 0.00546 -0.0826 0.2936 1.0000
4.500 0.8907 0.01121 0.00561 -0.0809 0.2921 1.0000
4.750 0.9092 0.01137 0.00577 -0.0792 0.2900 1.0000
5.000 0.9275 0.01154 0.00594 -0.0776 0.2871 1.0000
5.250 0.9454 0.01175 0.00613 -0.0759 0.2846 1.0000
5.500 0.9631 0.01198 0.00635 -0.0742 0.2819 1.0000
5.750 0.9810 0.01224 0.00658 -0.0726 0.2786 1.0000
6.000 1.0015 0.01241 0.00677 -0.0715 0.2769 1.0000
6.250 1.0218 0.01260 0.00697 -0.0703 0.2731 1.0000
6.500 1.0411 0.01284 0.00720 -0.0690 0.2686 1.0000
6.750 1.0594 0.01314 0.00747 -0.0676 0.2632 1.0000
7.000 1.0800 0.01336 0.00771 -0.0666 0.2598 1.0000
7.250 1.0983 0.01369 0.00798 -0.0653 0.2508 1.0000
7.500 1.1171 0.01402 0.00828 -0.0640 0.2398 1.0000
8.000 1.1317 0.01581 0.00963 -0.0583 0.1732 1.0000
8.250 1.1473 0.01634 0.01013 -0.0568 0.1664 1.0000
8.500 1.1641 0.01683 0.01061 -0.0554 0.1627 1.0000
8.750 1.1814 0.01730 0.01108 -0.0542 0.1603 1.0000
9.000 1.1979 0.01782 0.01159 -0.0528 0.1570 1.0000
9.250 1.2141 0.01837 0.01214 -0.0515 0.1535 1.0000
9.500 1.2332 0.01877 0.01257 -0.0506 0.1520 1.0000
9.750 1.2521 0.01918 0.01302 -0.0497 0.1502 1.0000
10.000 1.2693 0.01969 0.01355 -0.0486 0.1478 1.0000
10.250 1.2855 0.02028 0.01413 -0.0474 0.1447 1.0000
10.500 1.3000 0.02096 0.01481 -0.0461 0.1404 1.0000
10.750 1.3183 0.02143 0.01532 -0.0452 0.1378 1.0000
11.000 1.3350 0.02200 0.01587 -0.0442 0.1305 1.0000
11.250 1.3487 0.02277 0.01658 -0.0429 0.1181 1.0000
11.500 1.3327 0.02542 0.01888 -0.0383 0.0653 1.0000
11.750 1.3127 0.02852 0.02177 -0.0337 0.0215 1.0000
12.000 1.3200 0.02986 0.02313 -0.0320 0.0180 1.0000
12.250 1.3301 0.03103 0.02436 -0.0307 0.0168 1.0000
12.500 1.3382 0.03238 0.02574 -0.0294 0.0156 1.0000
12.750 1.3436 0.03399 0.02741 -0.0278 0.0143 1.0000
13.000 1.3527 0.03534 0.02881 -0.0267 0.0140 1.0000
13.250 1.3602 0.03685 0.03038 -0.0256 0.0134 1.0000
13.500 1.3664 0.03850 0.03209 -0.0244 0.0129 1.0000
13.750 1.3716 0.04029 0.03394 -0.0233 0.0124 1.0000
14.000 1.3740 0.04239 0.03611 -0.0222 0.0121 1.0000
14.250 1.3720 0.04497 0.03878 -0.0210 0.0116 1.0000
14.500 1.3721 0.04742 0.04131 -0.0200 0.0113 1.0000
14.750 1.3751 0.04966 0.04363 -0.0193 0.0112 1.0000
15.000 1.3760 0.05220 0.04625 -0.0187 0.0109 1.0000
15.250 1.3744 0.05510 0.04924 -0.0182 0.0107 1.0000
15.500 1.3732 0.05805 0.05227 -0.0180 0.0105 1.0000
15.750 1.3686 0.06149 0.05580 -0.0178 0.0102 1.0000
16.000 1.3652 0.06493 0.05933 -0.0179 0.0101 1.0000
16.250 1.3571 0.06908 0.06359 -0.0182 0.0099 1.0000
16.500 1.3506 0.07311 0.06770 -0.0187 0.0097 1.0000
16.750 1.3388 0.07790 0.07260 -0.0194 0.0096 1.0000
17.000 1.3248 0.08308 0.07790 -0.0203 0.0095 1.0000
17.250 1.3103 0.08842 0.08334 -0.0213 0.0094 1.0000
17.500 1.2932 0.09418 0.08923 -0.0226 0.0093 1.0000
17.750 1.2785 0.09963 0.09478 -0.0238 0.0092 1.0000
18.000 1.2630 0.10534 0.10060 -0.0253 0.0092 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 506 AIRFOIL (goe506-il)