DORNIER A-5 AIRFOIL (doa5-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: DORNIER A-5 AIRFOIL (doa5-il) Reynolds number: 100,000 Max Cl/Cd: 35.43 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-doa5-il-100000-n5.txt Download as CSV file: xf-doa5-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: DORNIER A-5 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.5409 0.09755 0.09197 -0.0604 1.0000 0.0478
-12.000 -0.5846 0.08759 0.08195 -0.0646 1.0000 0.0472
-11.750 -0.6295 0.08016 0.07444 -0.0663 1.0000 0.0467
-11.500 -0.6689 0.07496 0.06912 -0.0658 1.0000 0.0465
-11.250 -0.7057 0.07092 0.06498 -0.0637 1.0000 0.0463
-11.000 -0.7373 0.06746 0.06138 -0.0615 0.9993 0.0462
-10.750 -0.7540 0.06307 0.05670 -0.0631 0.9937 0.0463
-10.500 -0.7687 0.05962 0.05298 -0.0629 0.9868 0.0464
-10.250 -0.7729 0.05608 0.04910 -0.0630 0.9812 0.0466
-10.000 -0.7731 0.05284 0.04552 -0.0624 0.9756 0.0469
-9.750 -0.7660 0.04970 0.04200 -0.0624 0.9715 0.0473
-9.500 -0.7609 0.04714 0.03910 -0.0610 0.9661 0.0478
-9.250 -0.7483 0.04447 0.03594 -0.0608 0.9623 0.0489
-9.000 -0.7325 0.04245 0.03361 -0.0604 0.9585 0.0497
-8.750 -0.7151 0.04111 0.03221 -0.0598 0.9542 0.0505
-8.500 -0.6932 0.03970 0.03066 -0.0599 0.9512 0.0513
-8.250 -0.6669 0.03822 0.02898 -0.0606 0.9488 0.0521
-8.000 -0.6498 0.03698 0.02760 -0.0593 0.9438 0.0529
-7.750 -0.6263 0.03574 0.02618 -0.0591 0.9402 0.0538
-7.500 -0.5992 0.03455 0.02482 -0.0595 0.9376 0.0551
-7.250 -0.5697 0.03351 0.02351 -0.0602 0.9356 0.0570
-7.000 -0.5542 0.03275 0.02281 -0.0585 0.9305 0.0581
-6.750 -0.5306 0.03201 0.02207 -0.0582 0.9270 0.0597
-6.500 -0.5036 0.03124 0.02125 -0.0583 0.9246 0.0613
-6.250 -0.4748 0.03050 0.02044 -0.0587 0.9227 0.0633
-6.000 -0.4567 0.02996 0.01982 -0.0572 0.9185 0.0650
-5.750 -0.4379 0.02933 0.01926 -0.0559 0.9140 0.0670
-5.500 -0.4129 0.02872 0.01870 -0.0559 0.9108 0.0702
-5.250 -0.3847 0.02816 0.01810 -0.0564 0.9084 0.0743
-5.000 -0.3690 0.02765 0.01763 -0.0546 0.9039 0.0773
-4.750 -0.3534 0.02718 0.01720 -0.0530 0.8988 0.0809
-4.500 -0.3280 0.02666 0.01664 -0.0531 0.8953 0.0870
-4.250 -0.2984 0.02599 0.01604 -0.0541 0.8926 0.0963
-4.000 -0.2854 0.02554 0.01563 -0.0520 0.8852 0.1051
-3.750 -0.2606 0.02494 0.01510 -0.0521 0.8808 0.1229
-3.500 -0.2315 0.02418 0.01452 -0.0532 0.8779 0.1628
-3.250 -0.2220 0.02222 0.01370 -0.0522 0.8722 0.3573
-3.000 -0.2087 0.02228 0.01466 -0.0488 0.8671 0.5591
-2.750 -0.1807 0.02267 0.01507 -0.0482 0.8640 0.6063
-2.500 -0.1507 0.02325 0.01564 -0.0476 0.8616 0.6330
-2.250 -0.1382 0.02386 0.01622 -0.0444 0.8547 0.6473
-2.000 -0.1125 0.02430 0.01660 -0.0436 0.8507 0.6611
-1.750 -0.0806 0.02450 0.01670 -0.0441 0.8477 0.6743
-1.500 -0.0559 0.02504 0.01723 -0.0425 0.8436 0.6819
-1.250 -0.0371 0.02530 0.01743 -0.0410 0.8365 0.6919
-1.000 -0.0085 0.02559 0.01770 -0.0402 0.8328 0.6983
-0.750 0.0279 0.02545 0.01746 -0.0419 0.8302 0.7075
-0.500 0.0418 0.02575 0.01777 -0.0391 0.8216 0.7115
-0.250 0.0743 0.02553 0.01750 -0.0396 0.8164 0.7158
0.000 0.1159 0.02497 0.01685 -0.0421 0.8132 0.7197
0.500 0.1653 0.02470 0.01651 -0.0413 0.7986 0.7268
0.750 0.2001 0.02439 0.01619 -0.0424 0.7957 0.7297
1.000 0.2394 0.02396 0.01572 -0.0443 0.7934 0.7330
1.250 0.2497 0.02419 0.01594 -0.0417 0.7809 0.7382
1.500 0.2885 0.02367 0.01540 -0.0435 0.7770 0.7417
2.000 0.3369 0.02336 0.01512 -0.0418 0.7601 0.7488
2.250 0.3582 0.02329 0.01505 -0.0408 0.7497 0.7538
2.500 0.3913 0.02285 0.01461 -0.0415 0.7422 0.7576
2.750 0.4084 0.02286 0.01467 -0.0395 0.7312 0.7609
3.250 0.4604 0.02254 0.01442 -0.0387 0.7139 0.7701
3.500 0.4935 0.02212 0.01403 -0.0393 0.7061 0.7739
3.750 0.5087 0.02218 0.01415 -0.0370 0.6917 0.7775
4.000 0.5329 0.02195 0.01395 -0.0361 0.6764 0.7816
4.250 0.5646 0.02153 0.01351 -0.0364 0.6591 0.7860
4.500 0.5960 0.02109 0.01305 -0.0365 0.6386 0.7894
4.750 0.6237 0.02083 0.01277 -0.0361 0.6155 0.7927
5.000 0.6499 0.02068 0.01258 -0.0355 0.5869 0.7963
5.250 0.6827 0.02048 0.01223 -0.0361 0.5539 0.7998
5.500 0.7076 0.02059 0.01221 -0.0356 0.5157 0.8037
5.750 0.7279 0.02075 0.01220 -0.0341 0.4770 0.8065
6.000 0.7457 0.02107 0.01232 -0.0324 0.4427 0.8098
6.250 0.7618 0.02150 0.01259 -0.0306 0.4125 0.8139
6.500 0.7789 0.02202 0.01291 -0.0291 0.3891 0.8185
6.750 0.7943 0.02245 0.01330 -0.0272 0.3672 0.8222
7.000 0.8097 0.02290 0.01371 -0.0254 0.3470 0.8261
7.250 0.8252 0.02342 0.01418 -0.0237 0.3269 0.8306
7.500 0.8408 0.02399 0.01470 -0.0222 0.3050 0.8350
7.750 0.8531 0.02455 0.01521 -0.0201 0.2813 0.8384
8.000 0.8666 0.02512 0.01575 -0.0183 0.2551 0.8421
8.250 0.8790 0.02584 0.01635 -0.0165 0.2287 0.8458
8.500 0.8915 0.02664 0.01700 -0.0149 0.2118 0.8493
8.750 0.9048 0.02745 0.01771 -0.0133 0.1986 0.8527
9.000 0.9195 0.02817 0.01843 -0.0119 0.1882 0.8563
9.500 0.9501 0.02977 0.02005 -0.0096 0.1715 0.8646
9.750 0.9642 0.03068 0.02092 -0.0084 0.1647 0.8686
10.000 0.9805 0.03146 0.02177 -0.0074 0.1581 0.8726
10.250 0.9957 0.03237 0.02270 -0.0064 0.1517 0.8765
10.500 1.0121 0.03334 0.02365 -0.0056 0.1463 0.8802
10.750 1.0295 0.03420 0.02464 -0.0050 0.1406 0.8840
11.000 1.0454 0.03513 0.02560 -0.0042 0.1359 0.8881
11.250 1.0619 0.03609 0.02663 -0.0035 0.1311 0.8926
11.500 1.0775 0.03705 0.02774 -0.0028 0.1259 0.8972
11.750 1.0918 0.03805 0.02879 -0.0020 0.1217 0.9015
12.000 1.1069 0.03909 0.02997 -0.0014 0.1175 0.9063
12.250 1.1215 0.04018 0.03122 -0.0008 0.1132 0.9118
12.500 1.1353 0.04132 0.03245 -0.0004 0.1098 0.9187
13.000 1.1587 0.04358 0.03499 0.0010 0.1034 1.0000
13.250 1.1703 0.04500 0.03647 0.0015 0.1004 1.0000
13.500 1.1823 0.04643 0.03789 0.0018 0.0979 1.0000
13.750 1.1937 0.04808 0.03968 0.0023 0.0950 1.0000
14.000 1.2033 0.04983 0.04158 0.0027 0.0921 1.0000
14.250 1.2127 0.05156 0.04339 0.0031 0.0896 1.0000
14.500 1.2237 0.05322 0.04505 0.0034 0.0874 1.0000
14.750 1.2298 0.05537 0.04737 0.0038 0.0850 1.0000
15.000 1.2315 0.05785 0.05011 0.0042 0.0825 1.0000
15.250 1.2342 0.06027 0.05268 0.0044 0.0803 1.0000
15.500 1.2383 0.06256 0.05505 0.0045 0.0783 1.0000
15.750 1.2456 0.06463 0.05709 0.0045 0.0765 1.0000
16.000 1.2371 0.06836 0.06117 0.0044 0.0746 1.0000
16.250 1.2303 0.07209 0.06515 0.0040 0.0729 1.0000
16.500 1.2238 0.07584 0.06910 0.0033 0.0712 1.0000
16.750 1.2197 0.07940 0.07279 0.0024 0.0697 1.0000
17.000 1.2197 0.08248 0.07592 0.0014 0.0682 1.0000
17.250 1.2144 0.08644 0.07998 0.0001 0.0669 1.0000
17.500 1.1921 0.09313 0.08699 -0.0026 0.0659 1.0000
17.750 1.1664 0.10077 0.09493 -0.0062 0.0649 1.0000
18.000 1.1327 0.11035 0.10480 -0.0115 0.0642 1.0000
|
Polar data table (+)
Polar graphs
<< Back to DORNIER A-5 AIRFOIL (doa5-il)