NACA M7 AIRFOIL (m7-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M7 AIRFOIL (m7-il) Reynolds number: 200,000 Max Cl/Cd: 63.51 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m7-il-200000.txt Download as CSV file: xf-m7-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M7 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-14.000 -0.5046 0.16913 0.16625 0.0532 1.0000 0.0129
-13.750 -0.5007 0.16684 0.16394 0.0522 1.0000 0.0131
-7.000 -0.5408 0.09682 0.09408 0.0356 0.8807 0.0176
-6.750 -0.5343 0.09364 0.09081 0.0338 0.8608 0.0183
-6.500 -0.5252 0.09034 0.08742 0.0317 0.8452 0.0188
-6.250 -0.5128 0.08704 0.08402 0.0292 0.8315 0.0194
-6.000 -0.4954 0.08406 0.08092 0.0258 0.8189 0.0199
-5.750 -0.4730 0.08155 0.07824 0.0218 0.8073 0.0204
-5.500 -0.4512 0.07877 0.07530 0.0192 0.7968 0.0206
-5.250 -0.4310 0.07578 0.07214 0.0178 0.7874 0.0207
-5.000 -0.4196 0.07020 0.06651 0.0170 0.7779 0.0211
-4.750 -0.4107 0.06573 0.06201 0.0176 0.7691 0.0218
-4.500 -0.3952 0.06234 0.05851 0.0174 0.7604 0.0226
-4.250 -0.3752 0.05919 0.05525 0.0166 0.7512 0.0236
-4.000 -0.3536 0.05622 0.05211 0.0159 0.7435 0.0248
-3.750 -0.3293 0.05329 0.04903 0.0152 0.7347 0.0264
-3.500 -0.2987 0.05101 0.04650 0.0144 0.7269 0.0287
-3.250 -0.2609 0.05025 0.04533 0.0141 0.7190 0.0297
-3.000 -0.2500 0.04463 0.03973 0.0140 0.7118 0.0312
-2.750 -0.2299 0.04209 0.03708 0.0142 0.7043 0.0335
-2.500 -0.2026 0.03998 0.03475 0.0143 0.6965 0.0377
-2.250 -0.1691 0.03795 0.03228 0.0149 0.6899 0.0415
-2.000 -0.1503 0.03500 0.02934 0.0148 0.6822 0.0443
-1.750 -0.1145 0.03562 0.02942 0.0160 0.6753 0.0532
-1.500 -0.0961 0.03091 0.02477 0.0157 0.6682 0.0567
-1.250 -0.0685 0.02931 0.02286 0.0163 0.6615 0.0688
-1.000 -0.0403 0.02809 0.02133 0.0168 0.6544 0.0812
-0.750 -0.0163 0.02589 0.01911 0.0170 0.6476 0.0886
-0.500 0.0112 0.02443 0.01746 0.0173 0.6407 0.1014
-0.250 0.0389 0.02336 0.01615 0.0178 0.6338 0.1229
0.000 -0.6931 0.05561 0.05181 0.0505 0.6886 0.0250
0.250 0.1068 0.01840 0.01020 0.0205 0.6215 0.0476
0.500 0.1368 0.01717 0.00855 0.0215 0.6163 0.0393
0.750 0.1663 0.01599 0.00728 0.0217 0.6086 0.0385
1.000 0.1937 0.01534 0.00649 0.0223 0.6029 0.0411
1.250 0.2205 0.01452 0.00570 0.0228 0.5956 0.0416
1.500 0.2464 0.01399 0.00514 0.0235 0.5897 0.0443
1.750 0.2736 0.01374 0.00485 0.0238 0.5830 0.0518
2.000 0.3839 0.01193 0.00513 0.0068 0.5736 1.0000
2.250 0.4112 0.01199 0.00513 0.0068 0.5657 1.0000
2.500 0.4369 0.01199 0.00497 0.0074 0.5565 1.0000
2.750 0.4622 0.01192 0.00475 0.0080 0.5452 1.0000
3.000 0.4884 0.01188 0.00466 0.0084 0.5335 1.0000
3.250 0.5148 0.01194 0.00469 0.0086 0.5252 1.0000
3.500 0.5407 0.01199 0.00465 0.0090 0.5175 1.0000
3.750 0.5672 0.01205 0.00473 0.0092 0.5088 1.0000
4.000 0.5930 0.01212 0.00474 0.0096 0.5013 1.0000
4.250 0.6193 0.01215 0.00479 0.0098 0.4910 1.0000
4.500 0.6452 0.01214 0.00482 0.0102 0.4786 1.0000
4.750 0.6710 0.01214 0.00482 0.0105 0.4654 1.0000
5.000 0.6968 0.01211 0.00478 0.0109 0.4487 1.0000
5.250 0.7226 0.01212 0.00478 0.0112 0.4299 1.0000
5.500 0.7487 0.01217 0.00486 0.0115 0.4074 1.0000
5.750 0.7747 0.01230 0.00500 0.0116 0.3779 1.0000
6.000 0.8008 0.01261 0.00518 0.0116 0.3257 1.0000
6.250 0.8263 0.01512 0.00643 0.0092 0.1005 1.0000
6.500 0.8514 0.01624 0.00740 0.0088 0.0737 1.0000
6.750 0.8762 0.01694 0.00822 0.0088 0.0685 1.0000
7.000 0.9003 0.01782 0.00919 0.0087 0.0651 1.0000
7.250 0.9233 0.01889 0.01036 0.0086 0.0632 1.0000
7.500 0.9457 0.01980 0.01137 0.0088 0.0625 1.0000
7.750 0.9671 0.02079 0.01246 0.0092 0.0620 1.0000
8.000 0.9877 0.02181 0.01358 0.0098 0.0614 1.0000
8.250 1.0079 0.02280 0.01465 0.0105 0.0603 1.0000
8.500 1.0276 0.02386 0.01579 0.0114 0.0595 1.0000
8.750 1.0471 0.02501 0.01702 0.0124 0.0591 1.0000
9.000 1.0669 0.02623 0.01832 0.0135 0.0586 1.0000
9.250 1.0868 0.02750 0.01974 0.0145 0.0574 1.0000
9.500 1.1058 0.02880 0.02110 0.0153 0.0547 1.0000
9.750 1.1239 0.03206 0.02435 0.0163 0.0514 1.0000
10.000 1.1410 0.03263 0.02526 0.0172 0.0488 1.0000
10.250 1.1569 0.03399 0.02681 0.0181 0.0446 1.0000
10.500 1.1710 0.03754 0.03044 0.0188 0.0402 1.0000
10.750 1.1780 0.03944 0.03281 0.0202 0.0365 1.0000
11.000 1.1863 0.04154 0.03516 0.0214 0.0327 1.0000
11.250 1.1930 0.04423 0.03796 0.0224 0.0300 1.0000
11.750 1.1753 0.05375 0.04816 0.0249 0.0275 1.0000
12.000 1.1634 0.05564 0.05029 0.0262 0.0268 1.0000
12.250 1.1514 0.05861 0.05350 0.0262 0.0260 1.0000
12.500 1.1386 0.06236 0.05748 0.0254 0.0253 1.0000
12.750 1.1244 0.06673 0.06206 0.0239 0.0248 1.0000
13.000 1.1085 0.07168 0.06721 0.0219 0.0245 1.0000
13.250 1.0900 0.07739 0.07312 0.0192 0.0246 1.0000
13.500 1.0695 0.08378 0.07969 0.0159 0.0248 1.0000
13.750 1.0472 0.09093 0.08698 0.0121 0.0255 1.0000
14.000 1.0240 0.09873 0.09491 0.0078 0.0263 1.0000
14.250 0.9999 0.10731 0.10356 0.0035 0.0273 1.0000
14.500 0.8828 0.15682 0.15318 -0.0239 0.0739 1.0000
14.750 0.8772 0.16308 0.15941 -0.0280 0.0687 1.0000
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