M6 (65%) (m665-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: M6 (65%) (m665-il) Reynolds number: 500,000 Max Cl/Cd: 61.91 at α=3° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m665-il-500000-n5.txt Download as CSV file: xf-m665-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5932 0.08395 0.08163 -0.0111 1.0000 0.0080
-9.250 -0.6029 0.07839 0.07611 -0.0149 1.0000 0.0082
-9.000 -0.6265 0.07106 0.06881 -0.0204 1.0000 0.0080
-8.750 -0.6469 0.06444 0.06214 -0.0223 1.0000 0.0080
-8.500 -0.6563 0.05860 0.05619 -0.0234 1.0000 0.0080
-8.250 -0.7488 0.03079 0.02716 -0.0180 1.0000 0.0080
-8.000 -0.7540 0.02490 0.02051 -0.0136 1.0000 0.0080
-7.750 -0.7438 0.02215 0.01733 -0.0109 1.0000 0.0081
-7.500 -0.7270 0.02088 0.01586 -0.0092 1.0000 0.0082
-7.250 -0.7083 0.01992 0.01477 -0.0077 1.0000 0.0083
-7.000 -0.6880 0.01931 0.01408 -0.0064 1.0000 0.0085
-6.750 -0.6680 0.01862 0.01328 -0.0050 1.0000 0.0088
-6.500 -0.6492 0.01755 0.01203 -0.0034 1.0000 0.0091
-6.250 -0.6303 0.01649 0.01075 -0.0016 1.0000 0.0093
-6.000 -0.6081 0.01542 0.00950 -0.0006 0.9992 0.0095
-5.500 -0.5463 0.01354 0.00730 -0.0023 0.9892 0.0102
-5.250 -0.5153 0.01288 0.00651 -0.0030 0.9829 0.0107
-5.000 -0.4838 0.01215 0.00569 -0.0039 0.9762 0.0114
-4.750 -0.4516 0.01167 0.00517 -0.0050 0.9677 0.0123
-4.500 -0.4177 0.01124 0.00469 -0.0064 0.9580 0.0133
-4.250 -0.3820 0.01086 0.00423 -0.0082 0.9461 0.0141
-4.000 -0.3456 0.01048 0.00381 -0.0101 0.9306 0.0150
-3.750 -0.3109 0.01022 0.00351 -0.0116 0.9108 0.0162
-3.500 -0.2808 0.01002 0.00323 -0.0120 0.8875 0.0174
-3.250 -0.2538 0.00987 0.00296 -0.0117 0.8636 0.0184
-3.000 -0.2284 0.00971 0.00270 -0.0111 0.8407 0.0190
-2.750 -0.2035 0.00956 0.00246 -0.0103 0.8199 0.0203
-2.500 -0.1784 0.00945 0.00227 -0.0097 0.8001 0.0220
-2.250 -0.1532 0.00936 0.00209 -0.0090 0.7818 0.0249
-2.000 -0.1281 0.00924 0.00192 -0.0084 0.7647 0.0297
-1.750 -0.1029 0.00915 0.00176 -0.0077 0.7484 0.0337
-1.250 -0.0525 0.00891 0.00150 -0.0065 0.7184 0.0612
-1.000 -0.0277 0.00874 0.00140 -0.0058 0.7043 0.0955
-0.750 -0.0027 0.00859 0.00132 -0.0053 0.6909 0.1322
-0.500 0.0223 0.00845 0.00125 -0.0047 0.6779 0.1701
-0.250 0.0459 0.00815 0.00119 -0.0039 0.6657 0.2560
0.000 0.0675 0.00769 0.00112 -0.0028 0.6540 0.3895
0.250 0.0689 0.00613 0.00105 0.0029 0.6439 0.8118
0.500 0.1540 0.00620 0.00140 -0.0094 0.6291 0.9370
0.750 0.1967 0.00655 0.00170 -0.0122 0.6163 0.9617
1.000 0.2406 0.00683 0.00191 -0.0154 0.6024 0.9730
1.250 0.2762 0.00699 0.00199 -0.0171 0.5840 0.9782
1.500 0.3076 0.00707 0.00199 -0.0179 0.5634 0.9801
1.750 0.3378 0.00715 0.00201 -0.0185 0.5462 0.9823
2.000 0.3665 0.00724 0.00206 -0.0187 0.5310 0.9849
2.250 0.3962 0.00734 0.00209 -0.0191 0.5111 0.9868
2.500 0.4275 0.00743 0.00212 -0.0200 0.4893 0.9882
2.750 0.4577 0.00755 0.00217 -0.0206 0.4629 0.9899
3.000 0.4860 0.00785 0.00223 -0.0210 0.4044 0.9918
3.250 0.5129 0.00832 0.00235 -0.0212 0.3244 0.9937
3.500 0.5396 0.00890 0.00254 -0.0214 0.2337 0.9953
3.750 0.5672 0.00951 0.00279 -0.0219 0.1532 0.9968
4.000 0.5948 0.01011 0.00307 -0.0224 0.0836 0.9983
4.250 0.6235 0.01051 0.00333 -0.0229 0.0519 0.9996
4.500 0.6477 0.01090 0.00362 -0.0224 0.0299 1.0000
4.750 0.6699 0.01119 0.00390 -0.0214 0.0230 1.0000
5.000 0.6919 0.01151 0.00421 -0.0203 0.0188 1.0000
5.250 0.7140 0.01184 0.00457 -0.0192 0.0170 1.0000
5.500 0.7360 0.01215 0.00494 -0.0181 0.0161 1.0000
5.750 0.7579 0.01250 0.00533 -0.0170 0.0155 1.0000
6.000 0.7794 0.01287 0.00576 -0.0159 0.0149 1.0000
6.250 0.8005 0.01328 0.00621 -0.0147 0.0143 1.0000
6.500 0.8213 0.01372 0.00671 -0.0134 0.0137 1.0000
6.750 0.8414 0.01424 0.00728 -0.0121 0.0132 1.0000
7.000 0.8608 0.01483 0.00793 -0.0106 0.0130 1.0000
7.250 0.8787 0.01557 0.00874 -0.0089 0.0126 1.0000
7.500 0.8954 0.01647 0.00974 -0.0071 0.0123 1.0000
7.750 0.9119 0.01738 0.01074 -0.0052 0.0121 1.0000
8.000 0.9298 0.01814 0.01159 -0.0035 0.0120 1.0000
8.250 0.9470 0.01900 0.01255 -0.0018 0.0119 1.0000
8.500 0.9641 0.01988 0.01354 -0.0001 0.0118 1.0000
8.750 0.9832 0.02040 0.01414 0.0012 0.0114 1.0000
9.000 1.0001 0.02126 0.01512 0.0029 0.0112 1.0000
9.250 1.0170 0.02205 0.01601 0.0045 0.0108 1.0000
9.500 1.0325 0.02308 0.01716 0.0063 0.0106 1.0000
9.750 1.0468 0.02426 0.01849 0.0082 0.0105 1.0000
10.000 1.0617 0.02519 0.01956 0.0099 0.0102 1.0000
10.250 1.0752 0.02624 0.02074 0.0118 0.0099 1.0000
10.500 1.0887 0.02712 0.02173 0.0136 0.0096 1.0000
10.750 1.1022 0.02778 0.02245 0.0154 0.0093 1.0000
11.000 1.1134 0.02867 0.02343 0.0174 0.0091 1.0000
11.250 1.1221 0.02971 0.02456 0.0196 0.0088 1.0000
11.500 1.1242 0.03120 0.02621 0.0228 0.0087 1.0000
11.750 1.1195 0.03299 0.02815 0.0264 0.0085 1.0000
12.000 1.1169 0.03473 0.03007 0.0293 0.0083 1.0000
12.250 1.1186 0.03621 0.03174 0.0311 0.0081 1.0000
12.500 1.1145 0.03853 0.03425 0.0327 0.0081 1.0000
12.750 1.1115 0.04093 0.03683 0.0335 0.0079 1.0000
13.000 1.1035 0.04420 0.04031 0.0337 0.0078 1.0000
13.250 1.0977 0.04746 0.04375 0.0331 0.0076 1.0000
13.500 1.0840 0.05213 0.04864 0.0316 0.0075 1.0000
13.750 1.0642 0.05815 0.05488 0.0288 0.0075 1.0000
14.000 1.0518 0.06350 0.06038 0.0259 0.0074 1.0000
14.250 1.0261 0.07168 0.06876 0.0210 0.0075 1.0000
14.500 1.0032 0.08010 0.07734 0.0157 0.0075 1.0000
14.750 0.9739 0.09067 0.08804 0.0091 0.0076 1.0000
15.000 0.9356 0.10434 0.10185 0.0013 0.0079 1.0000
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Polar data table (+)
Polar graphs
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