FX 61-163 AIRFOIL (fx61163-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: FX 61-163 AIRFOIL (fx61163-il) Reynolds number: 100,000 Max Cl/Cd: 47.75 at α=10.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx61163-il-100000.txt Download as CSV file: xf-fx61163-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: FX 61-163 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.3843 0.10360 0.09908 -0.0453 1.0001 0.1414
-10.000 -0.3698 0.10117 0.09667 -0.0438 1.0001 0.1477
-9.750 -0.3933 0.09613 0.09179 -0.0474 1.0001 0.1558
-9.500 -0.3755 0.09405 0.08973 -0.0451 1.0001 0.1627
-7.500 -0.4693 0.04247 0.03594 -0.0805 0.9369 0.0641
-7.250 -0.4260 0.03837 0.03086 -0.0835 0.9310 0.0545
-7.000 -0.3925 0.03461 0.02683 -0.0852 0.9216 0.0528
-6.750 -0.3512 0.03159 0.02346 -0.0876 0.9155 0.0518
-6.500 -0.3143 0.02948 0.02105 -0.0888 0.9065 0.0526
-6.250 -0.2756 0.02798 0.01927 -0.0903 0.8986 0.0554
-6.000 -0.2421 0.02585 0.01729 -0.0905 0.8910 0.0586
-5.750 -0.2110 0.02464 0.01613 -0.0904 0.8826 0.0625
-5.500 -0.1810 0.02362 0.01504 -0.0903 0.8743 0.0685
-5.250 -0.1600 0.02255 0.01404 -0.0899 0.8642 0.0823
-5.000 -0.1441 0.01925 0.01202 -0.0916 0.8559 0.2552
-4.750 -0.1329 0.02202 0.01625 -0.0835 0.8470 0.6130
-4.500 -0.1098 0.02643 0.02058 -0.0738 0.8414 0.6457
-4.250 -0.0864 0.02992 0.02397 -0.0651 0.8340 0.6744
-4.000 -0.0492 0.03167 0.02550 -0.0610 0.8290 0.6967
-3.750 -0.0273 0.03162 0.02529 -0.0598 0.8217 0.7072
-3.500 -0.0060 0.03134 0.02484 -0.0594 0.8150 0.7187
-3.250 0.0165 0.03122 0.02455 -0.0586 0.8086 0.7285
-3.000 0.0398 0.03104 0.02424 -0.0578 0.8020 0.7366
-2.750 0.0611 0.03068 0.02370 -0.0578 0.7968 0.7468
-2.500 0.0813 0.03062 0.02358 -0.0564 0.7896 0.7532
-2.250 0.1006 0.03033 0.02315 -0.0561 0.7841 0.7622
-2.000 0.1234 0.03022 0.02296 -0.0552 0.7784 0.7689
-1.750 0.1383 0.03008 0.02275 -0.0542 0.7720 0.7773
-1.500 0.1669 0.02983 0.02236 -0.0542 0.7676 0.7835
-1.250 0.1751 0.02988 0.02240 -0.0526 0.7606 0.7911
-1.000 0.2005 0.02968 0.02212 -0.0522 0.7553 0.7964
-0.750 0.2217 0.02955 0.02189 -0.0522 0.7505 0.8033
-0.500 0.2339 0.02962 0.02197 -0.0503 0.7432 0.8086
-0.250 0.2605 0.02941 0.02167 -0.0504 0.7385 0.8142
0.000 0.2721 0.02960 0.02185 -0.0493 0.7318 0.8203
0.250 0.2929 0.02953 0.02175 -0.0483 0.7259 0.8248
0.500 0.3231 0.02929 0.02141 -0.0491 0.7220 0.8299
0.750 0.3287 0.02980 0.02198 -0.0474 0.7138 0.8351
1.000 0.3547 0.02963 0.02176 -0.0471 0.7090 0.8391
1.250 0.3741 0.02981 0.02193 -0.0466 0.7032 0.8440
1.500 0.3923 0.03016 0.02228 -0.0468 0.6960 0.8488
1.750 0.4230 0.02981 0.02189 -0.0469 0.6922 0.8525
2.000 0.4286 0.03055 0.02271 -0.0448 0.6832 0.8573
2.250 0.4616 0.03037 0.02249 -0.0461 0.6784 0.8614
2.500 0.4751 0.03091 0.02308 -0.0451 0.6707 0.8656
2.750 0.4991 0.03092 0.02309 -0.0448 0.6645 0.8697
3.000 0.5381 0.03050 0.02263 -0.0464 0.6613 0.8734
3.250 0.5446 0.03170 0.02395 -0.0454 0.6506 0.8780
3.500 0.5776 0.03125 0.02348 -0.0457 0.6470 0.8815
3.750 0.5826 0.03244 0.02477 -0.0441 0.6367 0.8857
4.000 0.6202 0.03207 0.02441 -0.0455 0.6326 0.8895
4.250 0.6292 0.03330 0.02575 -0.0447 0.6226 0.8936
4.500 0.6613 0.03291 0.02538 -0.0450 0.6181 0.8973
4.750 0.6722 0.03400 0.02657 -0.0441 0.6082 0.9013
5.000 0.7094 0.03359 0.02621 -0.0454 0.6034 0.9051
5.250 0.7487 0.03290 0.02552 -0.0464 0.5999 0.9087
5.500 0.7520 0.03421 0.02698 -0.0445 0.5885 0.9131
5.750 0.7973 0.03325 0.02605 -0.0462 0.5858 0.9168
6.000 0.8003 0.03499 0.02796 -0.0450 0.5737 0.9214
6.250 0.8451 0.03371 0.02670 -0.0461 0.5710 0.9252
6.500 0.8488 0.03516 0.02832 -0.0445 0.5588 0.9302
6.750 0.8900 0.03425 0.02747 -0.0458 0.5536 0.9343
7.000 0.9058 0.03465 0.02801 -0.0447 0.5439 0.9392
7.250 0.9575 0.03296 0.02637 -0.0466 0.5409 0.9433
7.500 0.9651 0.03409 0.02768 -0.0453 0.5289 0.9489
7.750 0.9808 0.03446 0.02820 -0.0442 0.5188 0.9551
8.000 1.0318 0.03256 0.02636 -0.0459 0.5133 0.9602
8.250 1.0486 0.03296 0.02697 -0.0452 0.5027 0.9669
8.500 1.1086 0.03055 0.02457 -0.0479 0.4969 0.9724
8.750 1.1318 0.03045 0.02467 -0.0478 0.4846 0.9821
9.000 1.1631 0.02985 0.02422 -0.0485 0.4721 0.9999
9.250 1.1989 0.02898 0.02347 -0.0495 0.4587 0.9999
9.500 1.2337 0.02814 0.02273 -0.0505 0.4436 0.9999
9.750 1.2651 0.02773 0.02241 -0.0512 0.4281 0.9999
10.000 1.2960 0.02744 0.02217 -0.0519 0.4111 0.9999
10.250 1.3150 0.02760 0.02242 -0.0514 0.3900 0.9999
10.500 1.3297 0.02785 0.02268 -0.0502 0.3665 0.9999
10.750 1.3360 0.02883 0.02373 -0.0487 0.3422 0.9999
11.000 1.3419 0.03009 0.02498 -0.0475 0.3179 0.9999
11.250 1.3443 0.03182 0.02671 -0.0466 0.2939 0.9999
11.500 1.3431 0.03413 0.02903 -0.0460 0.2695 0.9999
11.750 1.3399 0.03684 0.03167 -0.0456 0.2454 0.9999
12.000 1.3344 0.03998 0.03473 -0.0455 0.2223 0.9999
12.250 1.3277 0.04345 0.03814 -0.0456 0.2000 0.9999
12.500 1.3199 0.04708 0.04163 -0.0457 0.1799 0.9999
12.750 1.3119 0.05081 0.04524 -0.0457 0.1607 0.9999
13.000 1.3036 0.05469 0.04905 -0.0458 0.1428 0.9999
13.250 1.2957 0.05852 0.05278 -0.0459 0.1272 0.9999
13.500 1.2895 0.06226 0.05644 -0.0460 0.1134 0.9999
13.750 1.2849 0.06590 0.06006 -0.0461 0.1013 0.9999
14.000 1.2827 0.06931 0.06343 -0.0462 0.0911 0.9999
14.250 1.2797 0.07280 0.06684 -0.0465 0.0823 0.9999
14.500 1.2732 0.07707 0.07122 -0.0478 0.0748 0.9999
14.750 1.2687 0.08101 0.07520 -0.0486 0.0678 0.9999
15.000 1.2632 0.08512 0.07927 -0.0497 0.0615 0.9999
15.250 1.2599 0.08908 0.08338 -0.0504 0.0554 0.9999
15.500 1.2612 0.09189 0.08600 -0.0499 0.0486 0.9999
15.750 1.2548 0.09665 0.09107 -0.0511 0.0442 0.9999
16.000 1.2560 0.09981 0.09415 -0.0513 0.0399 0.9999
16.250 1.2578 0.10317 0.09768 -0.0509 0.0371 0.9999
16.500 1.2552 0.10746 0.10221 -0.0518 0.0353 0.9999
16.750 1.2512 0.11197 0.10693 -0.0531 0.0340 0.9999
17.000 1.2464 0.11668 0.11182 -0.0548 0.0330 0.9999
17.250 1.2418 0.12137 0.11665 -0.0567 0.0320 0.9999
17.500 1.2389 0.12577 0.12114 -0.0584 0.0311 0.9999
17.750 1.2326 0.13092 0.12642 -0.0606 0.0304 0.9999
18.000 1.2190 0.13773 0.13345 -0.0644 0.0302 0.9999
18.250 1.2023 0.14553 0.14150 -0.0696 0.0304 0.9999
18.500 1.1226 0.17121 0.16792 -0.0887 0.0349 0.9999
18.750 1.0848 0.18838 0.18518 -0.1004 0.0373 0.9999
|
Polar data table (+)
Polar graphs
<< Back to FX 61-163 AIRFOIL (fx61163-il)