NACA M11 AIRFOIL (m11-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M11 AIRFOIL (m11-il) Reynolds number: 500,000 Max Cl/Cd: 84.84 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m11-il-500000.txt Download as CSV file: xf-m11-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M11 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5136 0.08580 0.08349 -0.0176 1.0000 0.0239
-8.750 -0.5155 0.08189 0.07960 -0.0199 1.0000 0.0245
-8.500 -0.5218 0.07756 0.07531 -0.0230 1.0000 0.0249
-8.250 -0.5331 0.07336 0.07112 -0.0253 1.0000 0.0250
-8.000 -0.5382 0.06836 0.06609 -0.0282 1.0000 0.0262
-7.500 -0.5360 0.05722 0.05462 -0.0319 1.0000 0.0279
-7.250 -0.5373 0.05245 0.04964 -0.0306 1.0000 0.0280
-7.000 -0.5580 0.04366 0.04056 -0.0279 1.0000 0.0286
-6.750 -0.5516 0.04125 0.03811 -0.0260 1.0000 0.0293
-6.500 -0.5417 0.03977 0.03660 -0.0241 1.0000 0.0299
-6.250 -0.5324 0.03823 0.03500 -0.0218 1.0000 0.0307
-6.000 -0.5243 0.03652 0.03318 -0.0192 1.0000 0.0319
-5.750 -0.5178 0.03435 0.03083 -0.0159 1.0000 0.0339
-5.500 -0.4508 0.01470 0.01080 -0.0229 0.9904 0.0392
-5.250 -0.4203 0.01312 0.00916 -0.0246 0.9862 0.0409
-5.000 -0.3882 0.01133 0.00714 -0.0262 0.9830 0.0444
-4.750 -0.3631 0.00921 0.00461 -0.0264 0.9761 0.0506
-4.500 -0.3301 0.00824 0.00366 -0.0283 0.9721 0.0539
-4.250 -0.3328 0.01705 0.01124 -0.0218 0.9732 0.0376
-4.000 -0.2966 0.01533 0.00934 -0.0235 0.9704 0.0378
-3.750 -0.2650 0.01386 0.00773 -0.0243 0.9635 0.0384
-3.500 -0.2300 0.01281 0.00660 -0.0257 0.9578 0.0385
-3.250 -0.1979 0.01196 0.00570 -0.0265 0.9488 0.0388
-3.000 -0.1649 0.01120 0.00489 -0.0275 0.9401 0.0392
-2.750 -0.1361 0.01063 0.00429 -0.0275 0.9271 0.0403
-2.500 -0.1091 0.01019 0.00381 -0.0271 0.9122 0.0414
-2.250 -0.0834 0.00980 0.00336 -0.0264 0.8963 0.0422
-2.000 -0.0586 0.00949 0.00298 -0.0255 0.8793 0.0432
-1.750 -0.0338 0.00924 0.00266 -0.0247 0.8622 0.0442
-1.500 -0.0089 0.00904 0.00239 -0.0238 0.8451 0.0454
-1.250 0.0161 0.00888 0.00215 -0.0230 0.8285 0.0466
-1.000 0.0413 0.00877 0.00195 -0.0222 0.8125 0.0478
-0.750 0.0664 0.00862 0.00175 -0.0214 0.7972 0.0526
-0.500 0.0913 0.00846 0.00162 -0.0206 0.7824 0.0715
-0.250 0.1095 0.00757 0.00151 -0.0189 0.7680 0.3186
0.000 0.1096 0.00581 0.00147 -0.0130 0.7547 0.8217
0.250 0.2231 0.00610 0.00197 -0.0311 0.7410 0.9677
0.500 0.2890 0.00649 0.00226 -0.0390 0.7257 0.9872
0.750 0.3607 0.00652 0.00219 -0.0484 0.7091 0.9993
1.000 0.3867 0.00654 0.00213 -0.0481 0.6935 1.0000
1.250 0.4091 0.00657 0.00209 -0.0469 0.6778 1.0000
1.500 0.4316 0.00661 0.00207 -0.0459 0.6623 1.0000
2.000 0.4769 0.00672 0.00205 -0.0437 0.6261 1.0000
2.250 0.4996 0.00679 0.00206 -0.0426 0.6056 1.0000
2.500 0.5224 0.00689 0.00207 -0.0415 0.5850 1.0000
2.750 0.5450 0.00701 0.00210 -0.0404 0.5586 1.0000
3.000 0.5678 0.00714 0.00214 -0.0394 0.5311 1.0000
3.250 0.5908 0.00728 0.00221 -0.0384 0.5050 1.0000
3.500 0.6141 0.00743 0.00229 -0.0374 0.4822 1.0000
3.750 0.6373 0.00760 0.00239 -0.0365 0.4563 1.0000
4.000 0.6603 0.00781 0.00250 -0.0355 0.4253 1.0000
4.250 0.6830 0.00805 0.00263 -0.0346 0.3894 1.0000
4.500 0.7051 0.00838 0.00280 -0.0335 0.3482 1.0000
4.750 0.7260 0.00884 0.00304 -0.0323 0.2945 1.0000
5.000 0.7456 0.00946 0.00334 -0.0310 0.2226 1.0000
5.250 0.7638 0.01028 0.00377 -0.0295 0.1445 1.0000
5.500 0.7828 0.01100 0.00422 -0.0281 0.0931 1.0000
5.750 0.8033 0.01157 0.00467 -0.0268 0.0684 1.0000
6.000 0.8245 0.01204 0.00510 -0.0256 0.0559 1.0000
6.250 0.8452 0.01257 0.00561 -0.0243 0.0468 1.0000
6.500 0.8661 0.01306 0.00609 -0.0230 0.0399 1.0000
6.750 0.8865 0.01359 0.00666 -0.0217 0.0346 1.0000
7.000 0.9058 0.01422 0.00730 -0.0202 0.0300 1.0000
7.250 0.9251 0.01486 0.00801 -0.0187 0.0274 1.0000
7.500 0.9452 0.01538 0.00858 -0.0173 0.0252 1.0000
7.750 0.9634 0.01607 0.00929 -0.0157 0.0232 1.0000
8.000 0.9777 0.01718 0.01048 -0.0135 0.0216 1.0000
8.250 0.9974 0.01771 0.01109 -0.0121 0.0204 1.0000
8.500 1.0156 0.01837 0.01182 -0.0105 0.0192 1.0000
8.750 1.0333 0.01907 0.01256 -0.0089 0.0182 1.0000
9.000 1.0482 0.02008 0.01360 -0.0069 0.0173 1.0000
9.250 1.0575 0.02206 0.01571 -0.0042 0.0166 1.0000
9.500 1.0736 0.02307 0.01686 -0.0024 0.0162 1.0000
9.750 1.0898 0.02396 0.01788 -0.0007 0.0156 1.0000
10.000 1.1048 0.02500 0.01904 0.0012 0.0149 1.0000
10.250 1.1192 0.02589 0.02003 0.0030 0.0142 1.0000
10.500 1.1321 0.02692 0.02115 0.0050 0.0137 1.0000
10.750 1.1430 0.02823 0.02257 0.0071 0.0133 1.0000
11.000 1.1524 0.02960 0.02405 0.0093 0.0130 1.0000
11.250 1.1558 0.03169 0.02630 0.0122 0.0127 1.0000
11.500 1.1475 0.03525 0.03016 0.0161 0.0124 1.0000
12.000 1.1325 0.04048 0.03584 0.0226 0.0123 1.0000
12.250 1.1243 0.04297 0.03855 0.0248 0.0122 1.0000
12.500 1.1174 0.04524 0.04103 0.0261 0.0121 1.0000
12.750 1.1054 0.04872 0.04472 0.0267 0.0120 1.0000
13.000 1.0907 0.05282 0.04904 0.0264 0.0119 1.0000
13.250 1.0728 0.05789 0.05431 0.0251 0.0121 1.0000
13.500 1.0541 0.06348 0.06010 0.0225 0.0120 1.0000
13.750 1.0358 0.06958 0.06639 0.0192 0.0119 1.0000
14.000 1.0063 0.07811 0.07506 0.0149 0.0122 1.0000
14.250 0.9818 0.08661 0.08371 0.0096 0.0123 1.0000
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Polar data table (+)
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