FX 66-182 AIRFOIL (fx66182-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: FX 66-182 AIRFOIL (fx66182-il) Reynolds number: 500,000 Max Cl/Cd: 104.36 at α=8.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx66182-il-500000-n5.txt Download as CSV file: xf-fx66182-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: FX 66-182 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-16.000 -0.4996 0.10973 0.10720 -0.0600 0.9264 0.0080
-15.750 -0.6069 0.09232 0.08969 -0.0600 0.9545 0.0076
-15.500 -0.6569 0.07693 0.07400 -0.0703 0.9471 0.0076
-15.250 -0.6657 0.06294 0.05958 -0.0880 0.9226 0.0076
-15.000 -0.6373 0.05295 0.04890 -0.1073 0.8463 0.0077
-14.750 -0.6457 0.04965 0.04520 -0.1081 0.7869 0.0077
-14.500 -0.6553 0.04639 0.04165 -0.1087 0.7549 0.0078
-14.250 -0.6627 0.04334 0.03835 -0.1091 0.7342 0.0079
-14.000 -0.6650 0.04090 0.03570 -0.1093 0.7169 0.0080
-13.750 -0.6661 0.03865 0.03327 -0.1093 0.7023 0.0080
-13.500 -0.6674 0.03641 0.03083 -0.1091 0.6902 0.0082
-13.250 -0.6653 0.03454 0.02881 -0.1088 0.6788 0.0081
-13.000 -0.6615 0.03285 0.02702 -0.1085 0.6689 0.0083
-12.750 -0.6552 0.03142 0.02548 -0.1081 0.6594 0.0085
-12.500 -0.6475 0.03010 0.02405 -0.1076 0.6511 0.0086
-12.250 -0.6389 0.02887 0.02271 -0.1071 0.6433 0.0088
-12.000 -0.6303 0.02765 0.02138 -0.1064 0.6363 0.0089
-11.750 -0.6193 0.02665 0.02029 -0.1058 0.6298 0.0091
-11.500 -0.6085 0.02561 0.01914 -0.1051 0.6242 0.0093
-11.250 -0.5975 0.02462 0.01805 -0.1043 0.6189 0.0097
-11.000 -0.5872 0.02361 0.01693 -0.1034 0.6143 0.0099
-10.750 -0.5763 0.02266 0.01589 -0.1023 0.6100 0.0102
-10.500 -0.5675 0.02170 0.01482 -0.1009 0.6057 0.0105
-10.250 -0.5582 0.02099 0.01401 -0.0990 0.6016 0.0108
-10.000 -0.5464 0.02051 0.01348 -0.0972 0.5979 0.0111
-9.750 -0.5296 0.02003 0.01294 -0.0961 0.5941 0.0115
-9.500 -0.5119 0.01955 0.01240 -0.0951 0.5903 0.0119
-9.250 -0.4934 0.01908 0.01184 -0.0942 0.5869 0.0127
-8.750 -0.4548 0.01812 0.01071 -0.0924 0.5813 0.0141
-8.500 -0.4338 0.01771 0.01027 -0.0918 0.5786 0.0147
-8.250 -0.4117 0.01736 0.00987 -0.0913 0.5758 0.0157
-8.000 -0.3899 0.01696 0.00940 -0.0907 0.5731 0.0166
-7.750 -0.3667 0.01668 0.00902 -0.0902 0.5705 0.0174
-7.250 -0.3247 0.01565 0.00790 -0.0888 0.5658 0.0189
-7.000 -0.3014 0.01526 0.00747 -0.0884 0.5632 0.0197
-6.750 -0.2776 0.01492 0.00708 -0.0881 0.5606 0.0205
-6.500 -0.2533 0.01462 0.00672 -0.0877 0.5583 0.0213
-6.250 -0.2288 0.01432 0.00636 -0.0874 0.5562 0.0224
-6.000 -0.2051 0.01395 0.00596 -0.0871 0.5541 0.0240
-5.500 -0.1544 0.01343 0.00534 -0.0867 0.5504 0.0259
-5.250 -0.1284 0.01319 0.00506 -0.0866 0.5485 0.0267
-5.000 -0.1021 0.01297 0.00480 -0.0865 0.5463 0.0275
-4.750 -0.0760 0.01274 0.00455 -0.0864 0.5441 0.0294
-4.500 -0.0497 0.01254 0.00433 -0.0863 0.5419 0.0324
-4.250 -0.0234 0.01233 0.00410 -0.0862 0.5400 0.0396
-4.000 0.0023 0.01205 0.00388 -0.0862 0.5382 0.0599
-3.750 0.0276 0.01171 0.00366 -0.0861 0.5365 0.0981
-3.500 0.0532 0.01132 0.00344 -0.0861 0.5349 0.1480
-3.250 0.0778 0.01077 0.00318 -0.0860 0.5331 0.2291
-3.000 0.1023 0.01018 0.00292 -0.0860 0.5311 0.3241
-2.750 0.1265 0.00961 0.00273 -0.0859 0.5290 0.4370
-2.500 0.1533 0.00948 0.00278 -0.0858 0.5272 0.5012
-2.250 0.1814 0.00952 0.00284 -0.0859 0.5255 0.5313
-2.000 0.2096 0.00960 0.00290 -0.0860 0.5237 0.5510
-1.750 0.2379 0.00970 0.00294 -0.0862 0.5218 0.5622
-1.500 0.2664 0.00975 0.00298 -0.0864 0.5201 0.5693
-1.250 0.2951 0.00979 0.00300 -0.0866 0.5184 0.5753
-1.000 0.3238 0.00985 0.00301 -0.0869 0.5167 0.5809
-0.750 0.3523 0.00988 0.00305 -0.0871 0.5150 0.5848
-0.500 0.3808 0.00991 0.00306 -0.0873 0.5132 0.5874
-0.250 0.4093 0.00995 0.00306 -0.0876 0.5114 0.5901
0.000 0.4377 0.01000 0.00306 -0.0878 0.5096 0.5929
0.250 0.4661 0.01006 0.00307 -0.0880 0.5079 0.5954
0.500 0.4943 0.01012 0.00309 -0.0883 0.5062 0.5975
0.750 0.5228 0.01014 0.00314 -0.0885 0.5048 0.5995
1.000 0.5512 0.01018 0.00319 -0.0888 0.5032 0.6014
1.250 0.5795 0.01022 0.00324 -0.0890 0.5015 0.6036
1.500 0.6078 0.01027 0.00329 -0.0893 0.4997 0.6058
1.750 0.6360 0.01032 0.00334 -0.0895 0.4979 0.6079
2.000 0.6642 0.01038 0.00338 -0.0897 0.4960 0.6099
2.250 0.6922 0.01045 0.00343 -0.0900 0.4942 0.6121
2.500 0.7200 0.01051 0.00350 -0.0901 0.4926 0.6140
2.750 0.7479 0.01059 0.00358 -0.0903 0.4910 0.6158
3.000 0.7758 0.01064 0.00368 -0.0905 0.4893 0.6177
3.250 0.8036 0.01069 0.00378 -0.0907 0.4873 0.6199
3.500 0.8313 0.01076 0.00387 -0.0909 0.4851 0.6222
3.750 0.8589 0.01083 0.00397 -0.0911 0.4831 0.6245
4.000 0.8864 0.01090 0.00406 -0.0912 0.4811 0.6265
4.250 0.9138 0.01099 0.00416 -0.0914 0.4794 0.6284
4.500 0.9411 0.01107 0.00427 -0.0915 0.4777 0.6303
4.750 0.9682 0.01118 0.00440 -0.0916 0.4760 0.6324
5.000 0.9954 0.01125 0.00456 -0.0917 0.4742 0.6349
5.250 1.0224 0.01134 0.00471 -0.0918 0.4721 0.6374
5.500 1.0493 0.01143 0.00486 -0.0919 0.4699 0.6399
5.750 1.0759 0.01152 0.00499 -0.0919 0.4672 0.6422
6.000 1.1019 0.01160 0.00510 -0.0918 0.4640 0.6444
6.250 1.1276 0.01170 0.00524 -0.0916 0.4607 0.6467
6.500 1.1533 0.01176 0.00541 -0.0915 0.4565 0.6491
6.750 1.1781 0.01184 0.00554 -0.0912 0.4511 0.6519
7.000 1.2024 0.01196 0.00567 -0.0907 0.4465 0.6550
7.250 1.2276 0.01206 0.00586 -0.0905 0.4418 0.6582
7.500 1.2518 0.01217 0.00605 -0.0901 0.4368 0.6611
7.750 1.2745 0.01231 0.00623 -0.0895 0.4302 0.6641
8.000 1.2971 0.01244 0.00644 -0.0888 0.4219 0.6674
8.250 1.3180 0.01263 0.00665 -0.0878 0.4116 0.6709
8.500 1.3325 0.01296 0.00693 -0.0857 0.3911 0.6744
8.750 1.3354 0.01353 0.00738 -0.0816 0.3603 0.6780
9.000 1.3241 0.01466 0.00829 -0.0753 0.3168 0.6821
9.250 1.3081 0.01616 0.00959 -0.0690 0.2726 0.6867
9.500 1.2844 0.01824 0.01147 -0.0625 0.2330 0.6915
9.750 1.2679 0.02041 0.01357 -0.0580 0.2054 0.6971
10.000 1.2514 0.02303 0.01614 -0.0544 0.1802 0.7033
10.250 1.2374 0.02587 0.01894 -0.0517 0.1585 0.7097
10.500 1.2251 0.02882 0.02187 -0.0495 0.1385 0.7171
10.750 1.2137 0.03183 0.02487 -0.0477 0.1206 0.7255
11.000 1.2014 0.03502 0.02804 -0.0460 0.1023 0.7358
11.250 1.1944 0.03779 0.03085 -0.0447 0.0887 0.7495
11.750 1.1733 0.04371 0.03701 -0.0415 0.0583 0.8488
12.000 1.1792 0.04684 0.04025 -0.0434 0.0444 1.0000
12.250 1.1776 0.04958 0.04298 -0.0429 0.0361 1.0000
12.500 1.1784 0.05217 0.04557 -0.0425 0.0308 1.0000
12.750 1.1816 0.05455 0.04797 -0.0422 0.0256 1.0000
13.000 1.1814 0.05734 0.05074 -0.0420 0.0181 1.0000
13.250 1.1840 0.05987 0.05328 -0.0418 0.0155 1.0000
13.500 1.1867 0.06245 0.05589 -0.0417 0.0142 1.0000
13.750 1.1905 0.06494 0.05842 -0.0417 0.0133 1.0000
14.000 1.1941 0.06749 0.06102 -0.0417 0.0125 1.0000
14.250 1.1991 0.06992 0.06351 -0.0418 0.0120 1.0000
14.500 1.2048 0.07229 0.06595 -0.0419 0.0116 1.0000
14.750 1.2088 0.07491 0.06864 -0.0421 0.0112 1.0000
15.000 1.2130 0.07751 0.07130 -0.0423 0.0108 1.0000
15.250 1.2161 0.08030 0.07414 -0.0426 0.0104 1.0000
15.500 1.2188 0.08316 0.07707 -0.0430 0.0101 1.0000
15.750 1.2184 0.08647 0.08045 -0.0434 0.0095 1.0000
16.000 1.2227 0.08922 0.08327 -0.0439 0.0094 1.0000
16.250 1.2289 0.09172 0.08584 -0.0443 0.0090 1.0000
16.500 1.2325 0.09458 0.08877 -0.0449 0.0087 1.0000
16.750 1.2356 0.09754 0.09180 -0.0455 0.0084 1.0000
17.000 1.2395 0.10040 0.09472 -0.0462 0.0079 1.0000
17.250 1.2412 0.10364 0.09803 -0.0470 0.0078 1.0000
17.500 1.2435 0.10679 0.10123 -0.0478 0.0075 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 66-182 AIRFOIL (fx66182-il)