EPPLER 638 AIRFOIL (e638-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 638 AIRFOIL (e638-il) Reynolds number: 200,000 Max Cl/Cd: 68.2 at α=9.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e638-il-200000-n5.txt Download as CSV file: xf-e638-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 638 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3064 0.10447 0.10115 -0.0205 1.0000 0.0207
-9.500 -0.3036 0.10107 0.09780 -0.0220 1.0000 0.0212
-9.250 -0.3014 0.09765 0.09443 -0.0235 1.0000 0.0215
-9.000 -0.2808 0.09099 0.08769 -0.0352 0.8863 0.0222
-8.750 -0.2622 0.08591 0.08235 -0.0423 0.8183 0.0223
-8.500 -0.2605 0.08188 0.07818 -0.0458 0.7793 0.0223
-8.250 -0.2650 0.07791 0.07412 -0.0490 0.7530 0.0224
-8.000 -0.2694 0.07456 0.07066 -0.0501 0.7312 0.0224
-7.750 -0.2717 0.07123 0.06720 -0.0507 0.7136 0.0224
-7.500 -0.2702 0.06782 0.06367 -0.0513 0.6980 0.0224
-7.250 -0.2705 0.06337 0.05914 -0.0504 0.6841 0.0172
-7.000 -0.2670 0.06000 0.05564 -0.0506 0.6709 0.0170
-6.750 -0.2606 0.05723 0.05275 -0.0503 0.6577 0.0160
-6.500 -0.2554 0.05219 0.04743 -0.0505 0.6481 0.0172
-6.250 -0.2466 0.04827 0.04326 -0.0497 0.6377 0.0173
-6.000 -0.2359 0.04456 0.03927 -0.0486 0.6277 0.0174
-5.750 -0.2238 0.04170 0.03619 -0.0474 0.6177 0.0171
-5.500 -0.2103 0.03867 0.03289 -0.0460 0.6079 0.0169
-5.250 -0.1955 0.03554 0.02944 -0.0444 0.5986 0.0168
-4.750 -0.1631 0.02867 0.02166 -0.0402 0.5813 0.0170
-4.250 -0.1224 0.02449 0.01672 -0.0372 0.5639 0.0185
-4.000 -0.0992 0.02359 0.01561 -0.0364 0.5555 0.0198
-3.750 -0.0754 0.02189 0.01350 -0.0352 0.5471 0.0209
-3.500 -0.0503 0.02035 0.01156 -0.0341 0.5392 0.0216
-3.250 -0.0243 0.01923 0.01006 -0.0333 0.5310 0.0226
-3.000 0.0009 0.01841 0.00920 -0.0328 0.5232 0.0244
-2.750 0.0268 0.01778 0.00843 -0.0323 0.5157 0.0262
-2.500 0.0531 0.01703 0.00752 -0.0317 0.5089 0.0275
-2.250 0.0791 0.01651 0.00685 -0.0311 0.5018 0.0287
-2.000 0.1034 0.01587 0.00621 -0.0304 0.4954 0.0311
-1.750 0.1283 0.01544 0.00573 -0.0297 0.4884 0.0327
-1.500 0.1528 0.01508 0.00528 -0.0289 0.4826 0.0343
-1.250 0.1779 0.01481 0.00494 -0.0282 0.4767 0.0363
-1.000 0.2022 0.01450 0.00461 -0.0275 0.4709 0.0397
-0.750 0.2272 0.01433 0.00434 -0.0268 0.4658 0.0428
-0.500 0.2526 0.01416 0.00412 -0.0262 0.4598 0.0472
-0.250 0.2776 0.01399 0.00395 -0.0255 0.4545 0.0616
0.000 0.2983 0.01338 0.00382 -0.0243 0.4502 0.2165
0.500 0.4996 0.01223 0.00445 -0.0527 0.4341 0.9976
0.750 0.5314 0.01224 0.00438 -0.0538 0.4289 1.0000
1.000 0.5545 0.01231 0.00435 -0.0529 0.4249 1.0000
1.500 0.6010 0.01248 0.00436 -0.0514 0.4169 1.0000
1.750 0.6242 0.01256 0.00440 -0.0506 0.4127 1.0000
2.000 0.6473 0.01267 0.00443 -0.0498 0.4089 1.0000
2.250 0.6705 0.01280 0.00447 -0.0490 0.4058 1.0000
2.500 0.6938 0.01290 0.00457 -0.0482 0.4019 1.0000
2.750 0.7169 0.01302 0.00467 -0.0474 0.3980 1.0000
3.000 0.7399 0.01316 0.00477 -0.0466 0.3944 1.0000
3.250 0.7627 0.01333 0.00486 -0.0457 0.3913 1.0000
3.500 0.7859 0.01348 0.00503 -0.0449 0.3878 1.0000
3.750 0.8088 0.01363 0.00520 -0.0441 0.3843 1.0000
4.000 0.8316 0.01379 0.00535 -0.0433 0.3808 1.0000
4.250 0.8543 0.01398 0.00549 -0.0424 0.3777 1.0000
4.500 0.8771 0.01418 0.00569 -0.0416 0.3746 1.0000
4.750 0.8998 0.01436 0.00593 -0.0407 0.3712 1.0000
5.000 0.9224 0.01455 0.00615 -0.0399 0.3678 1.0000
5.250 0.9448 0.01476 0.00635 -0.0390 0.3646 1.0000
5.500 0.9673 0.01499 0.00656 -0.0381 0.3618 1.0000
5.750 0.9896 0.01521 0.00685 -0.0372 0.3585 1.0000
6.000 1.0118 0.01543 0.00714 -0.0363 0.3551 1.0000
6.250 1.0338 0.01566 0.00740 -0.0354 0.3517 1.0000
6.500 1.0559 0.01590 0.00766 -0.0345 0.3486 1.0000
6.750 1.0779 0.01616 0.00795 -0.0336 0.3457 1.0000
7.000 1.0994 0.01641 0.00831 -0.0327 0.3419 1.0000
7.250 1.1208 0.01666 0.00863 -0.0317 0.3381 1.0000
7.500 1.1422 0.01690 0.00891 -0.0307 0.3345 1.0000
7.750 1.1634 0.01718 0.00921 -0.0297 0.3310 1.0000
8.000 1.1838 0.01744 0.00961 -0.0286 0.3264 1.0000
8.250 1.2043 0.01770 0.00997 -0.0275 0.3222 1.0000
8.500 1.2249 0.01797 0.01024 -0.0264 0.3183 1.0000
8.750 1.2444 0.01826 0.01066 -0.0252 0.3134 1.0000
9.000 1.2636 0.01853 0.01104 -0.0240 0.3082 1.0000
9.250 1.2828 0.01881 0.01135 -0.0227 0.3036 1.0000
9.500 1.3011 0.01914 0.01183 -0.0213 0.2981 1.0000
9.750 1.3189 0.01944 0.01223 -0.0199 0.2926 1.0000
10.000 1.3361 0.01978 0.01262 -0.0184 0.2874 1.0000
10.250 1.3524 0.02013 0.01313 -0.0168 0.2805 1.0000
10.500 1.3672 0.02050 0.01353 -0.0150 0.2742 1.0000
10.750 1.3816 0.02091 0.01409 -0.0132 0.2664 1.0000
11.000 1.3938 0.02136 0.01460 -0.0111 0.2596 1.0000
11.250 1.4049 0.02187 0.01522 -0.0089 0.2515 1.0000
11.500 1.4113 0.02244 0.01585 -0.0059 0.2439 1.0000
11.750 1.4163 0.02311 0.01661 -0.0030 0.2355 1.0000
12.000 1.4207 0.02393 0.01751 -0.0003 0.2262 1.0000
12.250 1.4224 0.02499 0.01862 0.0022 0.2171 1.0000
12.500 1.4218 0.02636 0.02004 0.0044 0.2067 1.0000
12.750 1.4199 0.02806 0.02179 0.0061 0.1954 1.0000
13.000 1.4156 0.03019 0.02396 0.0073 0.1844 1.0000
13.250 1.4089 0.03276 0.02657 0.0082 0.1739 1.0000
13.500 1.3992 0.03586 0.02969 0.0086 0.1640 1.0000
13.750 1.3887 0.03925 0.03313 0.0087 0.1548 1.0000
14.000 1.3781 0.04280 0.03674 0.0086 0.1464 1.0000
14.250 1.3638 0.04687 0.04085 0.0083 0.1391 1.0000
14.500 1.3518 0.05082 0.04486 0.0078 0.1315 1.0000
14.750 1.3381 0.05503 0.04912 0.0072 0.1258 1.0000
15.000 1.3255 0.05927 0.05342 0.0065 0.1188 1.0000
15.250 1.3124 0.06372 0.05792 0.0055 0.1133 1.0000
15.500 1.3019 0.06801 0.06227 0.0045 0.1071 1.0000
15.750 1.2889 0.07270 0.06699 0.0033 0.1020 1.0000
16.000 1.2812 0.07682 0.07118 0.0022 0.0964 1.0000
16.250 1.2702 0.08144 0.07584 0.0008 0.0916 1.0000
16.500 1.2623 0.08571 0.08018 -0.0004 0.0865 1.0000
16.750 1.2530 0.09028 0.08479 -0.0019 0.0814 1.0000
17.000 1.2445 0.09480 0.08936 -0.0034 0.0766 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 638 AIRFOIL (e638-il)