FX 60-177 AIRFOIL (fx60177-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: FX 60-177 AIRFOIL (fx60177-il) Reynolds number: 100,000 Max Cl/Cd: 37.32 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx60177-il-100000-n5.txt Download as CSV file: xf-fx60177-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: FX 60-177 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
------ -------- --------- --------- -------- -------- --------
-14.250 -0.5769 0.10180 0.09624 -0.0666 1.0001 0.0157
-14.000 -0.6026 0.09440 0.08874 -0.0695 1.0001 0.0155
-13.750 -0.6279 0.08795 0.08217 -0.0717 1.0001 0.0154
-13.500 -0.6524 0.08234 0.07643 -0.0729 1.0001 0.0153
-13.250 -0.6760 0.07748 0.07141 -0.0733 1.0001 0.0153
-13.000 -0.6984 0.07322 0.06699 -0.0730 1.0001 0.0152
-12.750 -0.7195 0.06944 0.06304 -0.0720 1.0001 0.0152
-12.500 -0.7391 0.06605 0.05947 -0.0704 1.0001 0.0152
-12.250 -0.7568 0.06301 0.05625 -0.0684 1.0001 0.0153
-12.000 -0.7728 0.06024 0.05329 -0.0661 1.0001 0.0153
-11.750 -0.7804 0.05716 0.04995 -0.0651 0.9988 0.0155
-11.500 -0.7782 0.05396 0.04640 -0.0654 0.9960 0.0158
-11.250 -0.7704 0.05156 0.04384 -0.0658 0.9934 0.0161
-11.000 -0.7617 0.04955 0.04170 -0.0658 0.9899 0.0165
-10.750 -0.7510 0.04772 0.03973 -0.0659 0.9867 0.0170
-10.500 -0.7378 0.04601 0.03785 -0.0661 0.9842 0.0179
-10.250 -0.7259 0.04436 0.03596 -0.0653 0.9806 0.0190
-10.000 -0.7131 0.04286 0.03435 -0.0647 0.9768 0.0202
-9.750 -0.6988 0.04165 0.03308 -0.0647 0.9735 0.0216
-9.500 -0.6814 0.04038 0.03164 -0.0645 0.9707 0.0236
-9.250 -0.6736 0.03922 0.03043 -0.0626 0.9655 0.0251
-9.000 -0.6597 0.03815 0.02931 -0.0620 0.9616 0.0273
-8.750 -0.6411 0.03712 0.02817 -0.0619 0.9587 0.0307
-8.500 -0.6368 0.03638 0.02741 -0.0594 0.9524 0.0338
-8.250 -0.6212 0.03547 0.02643 -0.0588 0.9484 0.0384
-8.000 -0.6004 0.03456 0.02546 -0.0591 0.9454 0.0437
-7.750 -0.5913 0.03359 0.02449 -0.0572 0.9401 0.0478
-7.500 -0.5774 0.03267 0.02351 -0.0561 0.9352 0.0527
-7.250 -0.5584 0.03157 0.02238 -0.0561 0.9319 0.0582
-7.000 -0.5342 0.03061 0.02137 -0.0569 0.9293 0.0678
-6.750 -0.5279 0.02980 0.02060 -0.0543 0.9224 0.0768
-6.500 -0.5063 0.02884 0.01974 -0.0545 0.9186 0.0957
-6.250 -0.4807 0.02774 0.01890 -0.0557 0.9159 0.1374
-6.000 -0.4669 0.02660 0.01815 -0.0549 0.9103 0.2006
-5.750 -0.4462 0.02498 0.01732 -0.0562 0.9061 0.3274
-5.500 -0.4189 0.02420 0.01720 -0.0571 0.9032 0.4474
-5.250 -0.3905 0.02444 0.01771 -0.0564 0.9009 0.5129
-5.000 -0.3746 0.02480 0.01813 -0.0538 0.8948 0.5503
-4.750 -0.3508 0.02539 0.01870 -0.0523 0.8905 0.5854
-4.500 -0.3200 0.02593 0.01914 -0.0524 0.8876 0.6153
-4.250 -0.2875 0.02648 0.01957 -0.0526 0.8854 0.6324
-4.000 -0.2722 0.02681 0.01979 -0.0505 0.8779 0.6461
-3.750 -0.2472 0.02753 0.02044 -0.0486 0.8737 0.6551
-3.500 -0.2143 0.02793 0.02071 -0.0489 0.8709 0.6657
-3.250 -0.1953 0.02803 0.02069 -0.0479 0.8635 0.6773
-3.000 -0.1705 0.02841 0.02100 -0.0465 0.8592 0.6813
-2.750 -0.1374 0.02843 0.02089 -0.0475 0.8564 0.6879
-2.500 -0.1165 0.02845 0.02081 -0.0468 0.8494 0.6949
-2.250 -0.0909 0.02861 0.02090 -0.0459 0.8442 0.6993
-2.000 -0.0516 0.02824 0.02040 -0.0490 0.8413 0.7085
-1.750 -0.0356 0.02851 0.02063 -0.0465 0.8339 0.7115
-1.500 -0.0097 0.02856 0.02062 -0.0460 0.8289 0.7155
-1.250 0.0266 0.02830 0.02028 -0.0481 0.8260 0.7220
-1.000 0.0467 0.02841 0.02035 -0.0470 0.8186 0.7268
-0.750 0.0731 0.02837 0.02027 -0.0463 0.8130 0.7308
-0.500 0.1105 0.02800 0.01983 -0.0481 0.8099 0.7362
-0.250 0.1319 0.02801 0.01980 -0.0478 0.8006 0.7414
0.000 0.1616 0.02779 0.01956 -0.0478 0.7963 0.7442
0.250 0.1971 0.02737 0.01911 -0.0488 0.7935 0.7476
0.500 0.2178 0.02741 0.01913 -0.0483 0.7826 0.7533
0.750 0.2551 0.02686 0.01855 -0.0499 0.7791 0.7574
1.000 0.2709 0.02700 0.01871 -0.0477 0.7689 0.7607
1.250 0.3035 0.02659 0.01829 -0.0483 0.7646 0.7641
1.500 0.3303 0.02644 0.01813 -0.0486 0.7556 0.7683
1.750 0.3692 0.02589 0.01757 -0.0510 0.7495 0.7723
2.000 0.3907 0.02575 0.01746 -0.0496 0.7399 0.7747
2.250 0.4220 0.02528 0.01700 -0.0498 0.7329 0.7773
2.500 0.4457 0.02515 0.01689 -0.0492 0.7213 0.7805
2.750 0.4833 0.02457 0.01631 -0.0508 0.7137 0.7839
3.250 0.5374 0.02426 0.01604 -0.0510 0.6891 0.7904
3.500 0.5651 0.02399 0.01581 -0.0508 0.6774 0.7930
3.750 0.5968 0.02363 0.01545 -0.0513 0.6643 0.7956
4.000 0.6282 0.02338 0.01520 -0.0520 0.6485 0.7987
4.250 0.6614 0.02319 0.01499 -0.0534 0.6307 0.8021
4.500 0.6933 0.02298 0.01476 -0.0542 0.6125 0.8046
4.750 0.7214 0.02283 0.01458 -0.0540 0.5934 0.8066
5.000 0.7491 0.02277 0.01447 -0.0540 0.5732 0.8087
5.250 0.7753 0.02282 0.01449 -0.0539 0.5515 0.8113
5.500 0.8025 0.02294 0.01453 -0.0541 0.5290 0.8143
5.750 0.8301 0.02316 0.01466 -0.0546 0.5053 0.8174
6.000 0.8535 0.02347 0.01490 -0.0543 0.4817 0.8199
6.250 0.8726 0.02378 0.01515 -0.0530 0.4607 0.8220
6.500 0.8912 0.02416 0.01549 -0.0518 0.4405 0.8245
6.750 0.9104 0.02457 0.01589 -0.0509 0.4211 0.8273
7.000 0.9315 0.02503 0.01631 -0.0503 0.4034 0.8303
7.250 0.9538 0.02556 0.01682 -0.0503 0.3854 0.8333
7.500 0.9723 0.02610 0.01733 -0.0494 0.3678 0.8357
7.750 0.9871 0.02663 0.01785 -0.0477 0.3517 0.8382
8.000 1.0030 0.02721 0.01842 -0.0464 0.3361 0.8409
8.250 1.0206 0.02784 0.01906 -0.0455 0.3215 0.8437
8.500 1.0395 0.02851 0.01974 -0.0449 0.3077 0.8463
8.750 1.0592 0.02924 0.02049 -0.0447 0.2946 0.8490
9.000 1.0742 0.02996 0.02120 -0.0434 0.2830 0.8514
9.250 1.0881 0.03065 0.02196 -0.0420 0.2708 0.8540
9.500 1.1031 0.03141 0.02278 -0.0408 0.2596 0.8566
9.750 1.1182 0.03226 0.02365 -0.0399 0.2490 0.8591
10.000 1.1337 0.03318 0.02461 -0.0392 0.2382 0.8616
10.250 1.1506 0.03413 0.02564 -0.0388 0.2280 0.8640
10.500 1.1630 0.03512 0.02662 -0.0376 0.2197 0.8662
10.750 1.1759 0.03599 0.02766 -0.0363 0.2106 0.8687
11.000 1.1870 0.03704 0.02877 -0.0350 0.2023 0.8715
11.250 1.1987 0.03816 0.02999 -0.0340 0.1932 0.8744
11.500 1.2106 0.03939 0.03129 -0.0332 0.1850 0.8771
11.750 1.2219 0.04077 0.03272 -0.0326 0.1771 0.8796
12.000 1.2306 0.04204 0.03410 -0.0313 0.1698 0.8819
12.250 1.2382 0.04343 0.03559 -0.0301 0.1624 0.8842
12.500 1.2462 0.04491 0.03719 -0.0290 0.1555 0.8868
12.750 1.2538 0.04650 0.03892 -0.0282 0.1485 0.8896
13.000 1.2610 0.04827 0.04075 -0.0275 0.1427 0.8924
13.250 1.2683 0.05007 0.04272 -0.0270 0.1364 0.8952
13.500 1.2702 0.05200 0.04468 -0.0258 0.1316 0.8980
13.750 1.2748 0.05395 0.04681 -0.0250 0.1257 0.9008
14.000 1.2774 0.05622 0.04920 -0.0244 0.1201 0.9036
14.250 1.2790 0.05872 0.05175 -0.0241 0.1152 0.9063
14.750 1.2804 0.06395 0.05732 -0.0236 0.1051 0.9121
15.000 1.2790 0.06678 0.06025 -0.0234 0.1011 0.9153
15.250 1.2783 0.06981 0.06348 -0.0236 0.0965 0.9185
15.500 1.2750 0.07330 0.06706 -0.0242 0.0927 0.9216
15.750 1.2706 0.07685 0.07066 -0.0247 0.0895 0.9245
16.000 1.2673 0.08033 0.07437 -0.0252 0.0859 0.9276
16.250 1.2622 0.08421 0.07842 -0.0261 0.0827 0.9309
16.500 1.2561 0.08838 0.08271 -0.0273 0.0800 0.9344
16.750 1.2494 0.09260 0.08698 -0.0287 0.0775 0.9379
17.000 1.2419 0.09695 0.09156 -0.0299 0.0749 0.9420
17.250 1.2338 0.10163 0.09644 -0.0317 0.0725 0.9463
17.750 1.2161 0.11138 0.10648 -0.0358 0.0684 0.9559
18.000 1.2084 0.11615 0.11133 -0.0382 0.0664 0.9621
18.250 1.1985 0.12138 0.11672 -0.0409 0.0646 0.9703
|
Polar data table (+)
Polar graphs
<< Back to FX 60-177 AIRFOIL (fx60177-il)