M6 (65%) (m665-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: M6 (65%) (m665-il) Reynolds number: 100,000 Max Cl/Cd: 44.76 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m665-il-100000-n5.txt Download as CSV file: xf-m665-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: M6 (65%)
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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5483 0.08928 0.08423 -0.0127 1.0000 0.0341
-8.500 -0.5497 0.08524 0.08024 -0.0144 1.0000 0.0328
-8.250 -0.5561 0.08075 0.07579 -0.0169 1.0000 0.0319
-8.000 -0.5625 0.07648 0.07153 -0.0183 1.0000 0.0309
-7.750 -0.5666 0.07165 0.06667 -0.0201 1.0000 0.0299
-7.500 -0.5703 0.06624 0.06117 -0.0215 1.0000 0.0286
-7.250 -0.5776 0.05743 0.05197 -0.0227 1.0000 0.0261
-6.750 -0.5659 0.04941 0.04343 -0.0204 1.0000 0.0255
-6.500 -0.5596 0.04552 0.03926 -0.0187 1.0000 0.0254
-6.250 -0.5518 0.04183 0.03522 -0.0167 1.0000 0.0254
-6.000 -0.5422 0.03841 0.03141 -0.0144 1.0000 0.0255
-5.750 -0.5320 0.03501 0.02757 -0.0120 1.0000 0.0258
-5.500 -0.5198 0.03205 0.02428 -0.0098 1.0000 0.0264
-5.250 -0.5044 0.03029 0.02236 -0.0082 1.0000 0.0279
-5.000 -0.4872 0.02876 0.02063 -0.0065 1.0000 0.0301
-4.750 -0.4693 0.02687 0.01837 -0.0045 1.0000 0.0323
-4.500 -0.4498 0.02532 0.01636 -0.0026 1.0000 0.0353
-4.250 -0.4313 0.02351 0.01435 -0.0011 1.0000 0.0385
-4.000 -0.4111 0.02227 0.01296 0.0004 1.0000 0.0409
-3.750 -0.3899 0.02092 0.01143 0.0018 1.0000 0.0424
-3.500 -0.3690 0.01979 0.01015 0.0033 1.0000 0.0440
-3.250 -0.3492 0.01883 0.00909 0.0050 1.0000 0.0460
-3.000 -0.3289 0.01804 0.00820 0.0064 0.9992 0.0481
-2.750 -0.2926 0.01707 0.00725 0.0043 0.9893 0.0530
-2.500 -0.2560 0.01643 0.00659 0.0022 0.9786 0.0631
-2.250 -0.2202 0.01579 0.00594 0.0005 0.9674 0.0758
-2.000 -0.1850 0.01505 0.00534 -0.0012 0.9558 0.1057
-1.750 -0.1504 0.01411 0.00490 -0.0029 0.9443 0.2158
-1.500 -0.0445 0.01238 0.00563 -0.0165 0.9751 0.9732
-1.250 0.0261 0.01231 0.00529 -0.0253 0.9791 1.0000
-1.000 0.0690 0.01223 0.00505 -0.0287 0.9572 1.0000
-0.750 0.1121 0.01215 0.00481 -0.0321 0.9354 1.0000
-0.500 0.1530 0.01206 0.00460 -0.0349 0.9120 1.0000
-0.250 0.1905 0.01200 0.00441 -0.0369 0.8880 1.0000
0.000 0.2233 0.01196 0.00426 -0.0378 0.8641 1.0000
0.250 0.2509 0.01197 0.00416 -0.0376 0.8404 1.0000
0.500 0.2768 0.01200 0.00410 -0.0371 0.8191 1.0000
0.750 0.3013 0.01205 0.00408 -0.0362 0.7984 1.0000
1.000 0.3254 0.01212 0.00407 -0.0353 0.7796 1.0000
1.250 0.3499 0.01219 0.00409 -0.0345 0.7620 1.0000
1.500 0.3738 0.01228 0.00414 -0.0336 0.7442 1.0000
1.750 0.3975 0.01238 0.00423 -0.0327 0.7271 1.0000
2.000 0.4210 0.01249 0.00432 -0.0317 0.7103 1.0000
2.250 0.4443 0.01261 0.00444 -0.0306 0.6934 1.0000
2.500 0.4674 0.01275 0.00455 -0.0294 0.6756 1.0000
2.750 0.4901 0.01288 0.00471 -0.0282 0.6564 1.0000
3.000 0.5129 0.01303 0.00487 -0.0271 0.6368 1.0000
3.250 0.5357 0.01319 0.00505 -0.0259 0.6181 1.0000
3.500 0.5585 0.01336 0.00525 -0.0247 0.5989 1.0000
3.750 0.5808 0.01353 0.00548 -0.0234 0.5754 1.0000
4.000 0.6021 0.01370 0.00565 -0.0218 0.5446 1.0000
4.250 0.6220 0.01393 0.00576 -0.0200 0.4984 1.0000
4.500 0.6400 0.01430 0.00587 -0.0179 0.4264 1.0000
4.750 0.6574 0.01492 0.00614 -0.0159 0.3402 1.0000
5.000 0.6717 0.01605 0.00663 -0.0139 0.2158 1.0000
5.250 0.6839 0.01773 0.00753 -0.0119 0.0945 1.0000
5.500 0.7002 0.01894 0.00855 -0.0101 0.0629 1.0000
5.750 0.7180 0.01988 0.00951 -0.0084 0.0507 1.0000
6.000 0.7352 0.02087 0.01052 -0.0067 0.0443 1.0000
6.250 0.7527 0.02180 0.01159 -0.0049 0.0409 1.0000
6.500 0.7696 0.02279 0.01268 -0.0031 0.0384 1.0000
6.750 0.7857 0.02393 0.01389 -0.0012 0.0365 1.0000
7.000 0.8013 0.02536 0.01533 0.0007 0.0351 1.0000
7.250 0.8192 0.02684 0.01688 0.0023 0.0342 1.0000
7.500 0.8387 0.02816 0.01837 0.0037 0.0330 1.0000
7.750 0.8579 0.02952 0.01993 0.0051 0.0313 1.0000
8.000 0.8763 0.03102 0.02165 0.0065 0.0297 1.0000
8.250 0.8943 0.03286 0.02371 0.0080 0.0290 1.0000
8.500 0.9109 0.03489 0.02600 0.0095 0.0284 1.0000
8.750 0.9256 0.03708 0.02848 0.0112 0.0279 1.0000
9.000 0.9378 0.03947 0.03119 0.0130 0.0275 1.0000
9.250 0.9472 0.04205 0.03412 0.0150 0.0272 1.0000
9.500 0.9533 0.04484 0.03726 0.0172 0.0270 1.0000
9.750 0.9556 0.04782 0.04059 0.0195 0.0269 1.0000
10.000 0.9537 0.05099 0.04409 0.0220 0.0269 1.0000
10.250 0.9471 0.05424 0.04766 0.0245 0.0269 1.0000
10.500 0.9344 0.05755 0.05125 0.0272 0.0270 1.0000
10.750 0.9150 0.06082 0.05474 0.0298 0.0271 1.0000
11.000 0.8927 0.06486 0.05900 0.0305 0.0273 1.0000
11.250 0.8678 0.07003 0.06437 0.0289 0.0275 1.0000
11.500 0.8405 0.07676 0.07126 0.0249 0.0279 1.0000
11.750 0.8095 0.08600 0.08061 0.0181 0.0284 1.0000
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Polar data table (+)
Polar graphs
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