NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M10 AIRFOIL (m10-il) Reynolds number: 100,000 Max Cl/Cd: 45.26 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m10-il-100000-n5.txt Download as CSV file: xf-m10-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M10 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4576 0.11385 0.10916 0.0031 1.0000 0.0406
-9.750 -0.4611 0.11048 0.10582 0.0009 1.0000 0.0412
-9.500 -0.5800 0.11726 0.11234 0.0160 1.0000 0.0332
-9.250 -0.5750 0.11373 0.10883 0.0150 1.0000 0.0348
-9.000 -0.5715 0.10999 0.10511 0.0133 1.0000 0.0357
-8.750 -0.5686 0.10605 0.10120 0.0114 1.0000 0.0357
-8.500 -0.5666 0.10184 0.09703 0.0090 1.0000 0.0347
-8.250 -0.5672 0.09670 0.09193 0.0040 1.0000 0.0293
-8.000 -0.5642 0.09259 0.08785 0.0028 1.0000 0.0283
-7.750 -0.5620 0.08829 0.08358 -0.0001 1.0000 0.0276
-7.500 -0.5566 0.08356 0.07886 -0.0040 1.0000 0.0270
-7.250 -0.5494 0.07862 0.07392 -0.0081 1.0000 0.0265
-7.000 -0.5401 0.07352 0.06878 -0.0122 1.0000 0.0261
-6.750 -0.5288 0.06829 0.06346 -0.0161 1.0000 0.0259
-6.500 -0.5149 0.06302 0.05807 -0.0196 1.0000 0.0265
-6.250 -0.4985 0.05757 0.05243 -0.0229 1.0000 0.0277
-6.000 -0.4801 0.05188 0.04646 -0.0254 1.0000 0.0287
-5.750 -0.4610 0.04622 0.04045 -0.0270 1.0000 0.0290
-5.500 -0.4405 0.04067 0.03443 -0.0279 1.0000 0.0297
-5.250 -0.4200 0.03576 0.02900 -0.0282 1.0000 0.0315
-5.000 -0.3988 0.03458 0.02773 -0.0283 1.0000 0.0352
-4.750 -0.3751 0.03086 0.02348 -0.0281 1.0000 0.0374
-4.500 -0.3493 0.02770 0.01954 -0.0275 1.0000 0.0415
-4.250 -0.3257 0.02534 0.01688 -0.0273 1.0000 0.0448
-4.000 -0.3007 0.02373 0.01500 -0.0269 1.0000 0.0478
-3.750 -0.2750 0.02244 0.01330 -0.0264 1.0000 0.0528
-3.500 -0.2491 0.02087 0.01136 -0.0258 1.0000 0.0541
-3.250 -0.2237 0.01948 0.00973 -0.0251 1.0000 0.0556
-3.000 -0.1992 0.01827 0.00847 -0.0246 1.0000 0.0579
-2.750 -0.1749 0.01743 0.00758 -0.0240 1.0000 0.0603
-2.500 -0.1508 0.01685 0.00691 -0.0233 1.0000 0.0649
-2.250 -0.1270 0.01638 0.00633 -0.0226 1.0000 0.0690
-2.000 -0.1039 0.01565 0.00564 -0.0218 1.0000 0.0708
-1.750 -0.0811 0.01512 0.00511 -0.0211 1.0000 0.0732
-1.500 -0.0583 0.01473 0.00472 -0.0203 1.0000 0.0768
-1.250 -0.0218 0.01429 0.00424 -0.0224 0.9908 0.0848
-1.000 0.0172 0.01375 0.00382 -0.0249 0.9795 0.1128
-0.750 0.0524 0.01234 0.00355 -0.0273 0.9679 0.4100
-0.500 0.1029 0.01089 0.00357 -0.0307 0.9622 1.0000
-0.250 0.1406 0.01091 0.00342 -0.0329 0.9445 1.0000
0.000 0.1772 0.01093 0.00331 -0.0347 0.9264 1.0000
0.250 0.2106 0.01096 0.00324 -0.0357 0.9064 1.0000
0.500 0.2415 0.01100 0.00319 -0.0361 0.8859 1.0000
0.750 0.2698 0.01107 0.00318 -0.0359 0.8651 1.0000
1.000 0.2966 0.01115 0.00319 -0.0354 0.8444 1.0000
1.250 0.3225 0.01125 0.00324 -0.0347 0.8241 1.0000
1.500 0.3480 0.01137 0.00331 -0.0339 0.8040 1.0000
1.750 0.3735 0.01149 0.00339 -0.0331 0.7851 1.0000
2.000 0.3991 0.01163 0.00354 -0.0324 0.7652 1.0000
2.250 0.4248 0.01178 0.00368 -0.0317 0.7461 1.0000
2.500 0.4504 0.01194 0.00384 -0.0310 0.7273 1.0000
2.750 0.4761 0.01212 0.00404 -0.0303 0.7072 1.0000
3.000 0.5017 0.01230 0.00427 -0.0296 0.6873 1.0000
3.250 0.5268 0.01248 0.00448 -0.0287 0.6612 1.0000
3.500 0.5509 0.01268 0.00462 -0.0276 0.6236 1.0000
3.750 0.5743 0.01293 0.00477 -0.0262 0.5732 1.0000
4.000 0.5977 0.01326 0.00493 -0.0251 0.5108 1.0000
4.250 0.6210 0.01372 0.00518 -0.0240 0.4369 1.0000
4.500 0.6435 0.01444 0.00556 -0.0231 0.3403 1.0000
4.750 0.6611 0.01638 0.00637 -0.0222 0.1325 1.0000
5.000 0.6825 0.01786 0.00751 -0.0216 0.0761 1.0000
5.250 0.7056 0.01893 0.00862 -0.0209 0.0598 1.0000
5.500 0.7277 0.02013 0.00992 -0.0201 0.0494 1.0000
5.750 0.7489 0.02144 0.01127 -0.0193 0.0406 1.0000
6.000 0.7706 0.02282 0.01281 -0.0182 0.0360 1.0000
6.250 0.7919 0.02429 0.01433 -0.0174 0.0310 1.0000
6.500 0.8142 0.02613 0.01631 -0.0164 0.0282 1.0000
6.750 0.8378 0.02819 0.01864 -0.0154 0.0262 1.0000
7.000 0.8607 0.03016 0.02090 -0.0145 0.0234 1.0000
7.250 0.8821 0.03211 0.02311 -0.0138 0.0211 1.0000
7.500 0.9019 0.03478 0.02609 -0.0128 0.0201 1.0000
7.750 0.9188 0.03814 0.02982 -0.0118 0.0194 1.0000
8.000 0.9318 0.04230 0.03444 -0.0106 0.0190 1.0000
8.250 0.9406 0.04671 0.03941 -0.0093 0.0186 1.0000
8.500 0.9495 0.04994 0.04317 -0.0079 0.0180 1.0000
8.750 0.9535 0.05377 0.04750 -0.0066 0.0174 1.0000
9.000 0.9524 0.05803 0.05217 -0.0054 0.0171 1.0000
9.250 0.9460 0.06245 0.05692 -0.0047 0.0169 1.0000
9.500 0.9343 0.06680 0.06153 -0.0042 0.0169 1.0000
9.750 0.9179 0.07115 0.06605 -0.0044 0.0170 1.0000
10.000 0.9008 0.07644 0.07148 -0.0071 0.0171 1.0000
10.250 0.8845 0.08291 0.07805 -0.0118 0.0173 1.0000
10.500 0.8698 0.09046 0.08567 -0.0176 0.0176 1.0000
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Polar data table (+)
Polar graphs
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