NACA M6 AIRFOIL (m6-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA M6 AIRFOIL (m6-il) Reynolds number: 1,000,000 Max Cl/Cd: 97.01 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m6-il-1000000.txt Download as CSV file: xf-m6-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M6 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.8607 0.05811 0.05606 -0.0336 1.0000 0.0214
-12.000 -0.8919 0.05197 0.04970 -0.0356 1.0000 0.0214
-11.750 -0.9338 0.04636 0.04381 -0.0327 1.0000 0.0214
-11.500 -0.9768 0.04051 0.03759 -0.0270 1.0000 0.0216
-11.250 -0.9798 0.03754 0.03445 -0.0244 1.0000 0.0218
-11.000 -0.9682 0.03613 0.03296 -0.0230 1.0000 0.0221
-10.750 -0.9548 0.03486 0.03162 -0.0216 1.0000 0.0223
-10.500 -0.9364 0.03411 0.03085 -0.0207 1.0000 0.0225
-10.250 -0.9191 0.03315 0.02983 -0.0197 1.0000 0.0227
-10.000 -0.9041 0.03171 0.02827 -0.0184 1.0000 0.0230
-9.750 -0.8902 0.02994 0.02632 -0.0169 1.0000 0.0232
-9.500 -0.8757 0.02807 0.02425 -0.0154 1.0000 0.0235
-9.250 -0.8587 0.02650 0.02249 -0.0141 1.0000 0.0239
-9.000 -0.8400 0.02512 0.02092 -0.0130 1.0000 0.0244
-8.750 -0.8186 0.02362 0.01921 -0.0124 0.9932 0.0247
-8.500 -0.7914 0.02231 0.01770 -0.0128 0.9677 0.0250
-8.000 -0.7498 0.02100 0.01601 -0.0100 0.9194 0.0254
-7.750 -0.7278 0.02066 0.01548 -0.0087 0.9011 0.0256
-7.500 -0.7097 0.01799 0.01253 -0.0073 0.8854 0.0261
-7.250 -0.6863 0.01705 0.01150 -0.0066 0.8702 0.0266
-7.000 -0.6619 0.01638 0.01073 -0.0060 0.8553 0.0268
-6.750 -0.6369 0.01575 0.01000 -0.0055 0.8404 0.0271
-6.500 -0.6116 0.01519 0.00934 -0.0050 0.8253 0.0274
-6.000 -0.5599 0.01414 0.00809 -0.0041 0.7954 0.0280
-5.750 -0.5338 0.01366 0.00753 -0.0038 0.7812 0.0283
-5.500 -0.5075 0.01321 0.00701 -0.0034 0.7680 0.0286
-5.250 -0.4811 0.01280 0.00651 -0.0031 0.7555 0.0290
-5.000 -0.4545 0.01245 0.00608 -0.0028 0.7436 0.0294
-4.750 -0.4274 0.01215 0.00573 -0.0026 0.7324 0.0299
-4.500 -0.4005 0.01185 0.00537 -0.0024 0.7221 0.0302
-4.250 -0.3734 0.01159 0.00504 -0.0022 0.7120 0.0304
-4.000 -0.3463 0.01128 0.00469 -0.0020 0.7027 0.0307
-3.750 -0.3209 0.01074 0.00408 -0.0015 0.6938 0.0312
-3.500 -0.2943 0.01036 0.00367 -0.0012 0.6851 0.0318
-3.250 -0.2673 0.01011 0.00337 -0.0010 0.6768 0.0324
-3.000 -0.2398 0.00988 0.00312 -0.0009 0.6686 0.0330
-2.750 -0.2122 0.00971 0.00290 -0.0008 0.6608 0.0338
-2.500 -0.1844 0.00952 0.00270 -0.0007 0.6532 0.0347
-2.250 -0.1565 0.00940 0.00253 -0.0007 0.6454 0.0356
-2.000 -0.1284 0.00926 0.00238 -0.0007 0.6382 0.0364
-1.750 -0.1007 0.00909 0.00217 -0.0006 0.6307 0.0384
-1.500 -0.0726 0.00895 0.00203 -0.0006 0.6237 0.0408
-1.250 -0.0445 0.00883 0.00191 -0.0006 0.6164 0.0458
-1.000 -0.0177 0.00854 0.00179 -0.0004 0.6096 0.0889
-0.750 0.0092 0.00825 0.00171 -0.0003 0.6024 0.1469
-0.500 0.0362 0.00803 0.00165 -0.0002 0.5954 0.2013
-0.250 0.0619 0.00760 0.00158 0.0001 0.5885 0.3098
0.000 0.0755 0.00606 0.00137 0.0025 0.5819 0.7036
0.250 0.0956 0.00557 0.00143 0.0045 0.5755 0.8589
0.500 0.1219 0.00555 0.00151 0.0052 0.5681 0.9021
0.750 0.1490 0.00562 0.00157 0.0056 0.5612 0.9235
1.000 0.1768 0.00568 0.00164 0.0059 0.5536 0.9368
1.250 0.2045 0.00580 0.00172 0.0061 0.5461 0.9479
1.500 0.2317 0.00592 0.00183 0.0066 0.5383 0.9589
1.750 0.2624 0.00610 0.00196 0.0063 0.5301 0.9664
2.000 0.2993 0.00632 0.00215 0.0046 0.5203 0.9730
2.250 0.3457 0.00657 0.00235 0.0006 0.5091 0.9754
2.500 0.3828 0.00672 0.00243 -0.0014 0.4993 0.9773
2.750 0.4161 0.00680 0.00250 -0.0026 0.4910 0.9794
3.000 0.4461 0.00693 0.00259 -0.0031 0.4828 0.9830
3.250 0.4814 0.00699 0.00262 -0.0048 0.4730 0.9837
3.500 0.5149 0.00707 0.00266 -0.0061 0.4616 0.9843
3.750 0.5481 0.00716 0.00271 -0.0074 0.4515 0.9850
4.000 0.5813 0.00723 0.00276 -0.0087 0.4418 0.9858
4.250 0.6141 0.00733 0.00282 -0.0099 0.4298 0.9867
4.500 0.6462 0.00745 0.00289 -0.0110 0.4150 0.9877
4.750 0.6778 0.00758 0.00297 -0.0120 0.3994 0.9888
5.000 0.7092 0.00772 0.00306 -0.0129 0.3832 0.9901
5.250 0.7399 0.00789 0.00317 -0.0138 0.3641 0.9914
5.500 0.7697 0.00809 0.00330 -0.0144 0.3453 0.9929
5.750 0.7995 0.00832 0.00344 -0.0151 0.3222 0.9939
6.000 0.8310 0.00860 0.00360 -0.0162 0.2962 0.9945
6.250 0.8624 0.00889 0.00379 -0.0174 0.2689 0.9952
6.500 0.8930 0.00928 0.00403 -0.0184 0.2378 0.9960
6.750 0.9231 0.00969 0.00430 -0.0194 0.2079 0.9968
7.000 0.9524 0.01016 0.00461 -0.0203 0.1772 0.9977
7.250 0.9813 0.01067 0.00496 -0.0210 0.1458 0.9986
7.500 1.0081 0.01150 0.00550 -0.0217 0.0958 0.9995
7.750 1.0332 0.01207 0.00595 -0.0217 0.0749 1.0000
8.000 1.0528 0.01254 0.00634 -0.0205 0.0606 1.0000
8.250 1.0721 0.01299 0.00673 -0.0192 0.0490 1.0000
8.500 1.0911 0.01343 0.00712 -0.0179 0.0422 1.0000
8.750 1.1102 0.01380 0.00749 -0.0165 0.0388 1.0000
9.000 1.1281 0.01425 0.00794 -0.0149 0.0356 1.0000
9.250 1.1468 0.01458 0.00830 -0.0135 0.0339 1.0000
9.500 1.1640 0.01501 0.00872 -0.0118 0.0319 1.0000
9.750 1.1792 0.01554 0.00928 -0.0097 0.0299 1.0000
10.000 1.1966 0.01588 0.00965 -0.0081 0.0288 1.0000
10.250 1.2127 0.01628 0.01008 -0.0062 0.0276 1.0000
10.500 1.2270 0.01678 0.01059 -0.0041 0.0264 1.0000
10.750 1.2373 0.01748 0.01134 -0.0014 0.0249 1.0000
11.000 1.2530 0.01786 0.01176 0.0004 0.0242 1.0000
11.250 1.2673 0.01831 0.01224 0.0024 0.0233 1.0000
11.500 1.2801 0.01882 0.01278 0.0046 0.0224 1.0000
11.750 1.2872 0.01949 0.01347 0.0076 0.0216 1.0000
12.000 1.2876 0.02059 0.01464 0.0111 0.0207 1.0000
12.250 1.3008 0.02127 0.01537 0.0126 0.0203 1.0000
12.500 1.3129 0.02213 0.01628 0.0139 0.0197 1.0000
12.750 1.3244 0.02311 0.01732 0.0151 0.0191 1.0000
13.000 1.3357 0.02420 0.01846 0.0160 0.0185 1.0000
13.250 1.3451 0.02551 0.01981 0.0169 0.0179 1.0000
13.500 1.3488 0.02741 0.02177 0.0178 0.0173 1.0000
13.750 1.3544 0.02926 0.02369 0.0184 0.0168 1.0000
14.000 1.3648 0.03073 0.02524 0.0187 0.0165 1.0000
14.250 1.3733 0.03243 0.02700 0.0189 0.0161 1.0000
14.500 1.3802 0.03430 0.02895 0.0191 0.0157 1.0000
14.750 1.3861 0.03633 0.03105 0.0192 0.0154 1.0000
15.000 1.3909 0.03851 0.03329 0.0192 0.0150 1.0000
15.250 1.3934 0.04095 0.03579 0.0192 0.0147 1.0000
15.500 1.3927 0.04379 0.03870 0.0190 0.0144 1.0000
15.750 1.3871 0.04721 0.04221 0.0187 0.0142 1.0000
16.000 1.3751 0.05142 0.04652 0.0183 0.0138 1.0000
16.250 1.3756 0.05435 0.04954 0.0177 0.0137 1.0000
16.500 1.3757 0.05738 0.05267 0.0171 0.0136 1.0000
16.750 1.3742 0.06070 0.05607 0.0164 0.0134 1.0000
17.000 1.3719 0.06419 0.05966 0.0155 0.0132 1.0000
17.250 1.3683 0.06790 0.06346 0.0145 0.0130 1.0000
17.500 1.3642 0.07175 0.06740 0.0134 0.0129 1.0000
17.750 1.3596 0.07574 0.07147 0.0121 0.0127 1.0000
18.000 1.3547 0.07985 0.07567 0.0108 0.0125 1.0000
18.250 1.3490 0.08410 0.08000 0.0094 0.0123 1.0000
18.500 1.3423 0.08855 0.08454 0.0078 0.0122 1.0000
18.750 1.3357 0.09310 0.08917 0.0061 0.0120 1.0000
19.000 1.3278 0.09788 0.09403 0.0042 0.0119 1.0000
19.250 1.3185 0.10288 0.09911 0.0023 0.0118 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M6 AIRFOIL (m6-il)