NACA M19 AIRFOIL (m19-il) Xfoil prediction polar at RE=500,000 Ncrit=9
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Airfoil: NACA M19 AIRFOIL (m19-il) Reynolds number: 500,000 Max Cl/Cd: 102.54 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m19-il-500000.txt Download as CSV file: xf-m19-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M19 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
------ -------- --------- --------- -------- -------- --------
-7.500 -0.2806 0.08601 0.08336 -0.0003 0.7173 0.0144
-7.250 -0.2767 0.08312 0.08046 -0.0025 0.7111 0.0146
-7.000 -0.2740 0.08014 0.07748 -0.0051 0.7045 0.0147
-6.750 -0.2672 0.07675 0.07404 -0.0088 0.6988 0.0148
-6.500 -0.2560 0.07293 0.07019 -0.0131 0.6925 0.0149
-6.250 -0.2426 0.06890 0.06610 -0.0167 0.6869 0.0149
-6.000 -0.2352 0.06320 0.06039 -0.0175 0.6815 0.0151
-5.750 -0.2279 0.05912 0.05629 -0.0164 0.6753 0.0155
-5.500 -0.2145 0.05577 0.05288 -0.0179 0.6697 0.0161
-5.250 -0.1979 0.05213 0.04921 -0.0204 0.6636 0.0168
-5.000 -0.1795 0.04844 0.04542 -0.0231 0.6583 0.0174
-4.750 -0.1584 0.04474 0.04166 -0.0260 0.6526 0.0186
-4.500 -0.1234 0.04185 0.03862 -0.0312 0.6470 0.0199
-4.250 -0.0956 0.03835 0.03497 -0.0343 0.6423 0.0200
-4.000 -0.0702 0.03472 0.03124 -0.0364 0.6365 0.0201
-3.750 -0.0617 0.04925 0.04538 -0.0383 0.6433 0.0202
-3.500 -0.0426 0.04377 0.03982 -0.0400 0.6378 0.0209
-3.250 -0.0211 0.04137 0.03734 -0.0407 0.6319 0.0215
-3.000 0.0045 0.03921 0.03506 -0.0417 0.6263 0.0227
-2.750 0.0337 0.03702 0.03275 -0.0428 0.6202 0.0251
-2.500 0.0733 0.03600 0.03145 -0.0438 0.6148 0.0275
-2.250 0.1038 0.03376 0.02905 -0.0444 0.6090 0.0277
-2.000 0.1331 0.03128 0.02636 -0.0449 0.6036 0.0278
-1.750 0.1558 0.02685 0.02173 -0.0457 0.5989 0.0294
-1.500 0.1820 0.02544 0.02023 -0.0461 0.5928 0.0305
-1.250 0.2094 0.02417 0.01881 -0.0464 0.5873 0.0327
-1.000 0.2400 0.02283 0.01728 -0.0464 0.5816 0.0364
-0.750 0.2723 0.02241 0.01661 -0.0460 0.5760 0.0382
-0.500 0.2978 0.01916 0.01311 -0.0465 0.5712 0.0408
-0.250 0.3252 0.01831 0.01220 -0.0467 0.5650 0.0430
0.000 0.3536 0.01748 0.01119 -0.0467 0.5593 0.0469
1.000 0.4664 0.01440 0.00760 -0.0470 0.5365 0.0686
1.500 0.5225 0.01345 0.00649 -0.0471 0.5248 0.0854
1.750 0.5531 0.01132 0.00392 -0.0460 0.5197 0.0456
2.000 0.5807 0.01095 0.00347 -0.0459 0.5145 0.0449
2.250 0.6084 0.01069 0.00323 -0.0459 0.5084 0.0460
2.500 0.6358 0.01044 0.00292 -0.0458 0.5029 0.0446
2.750 0.6634 0.01025 0.00273 -0.0458 0.4968 0.0436
3.000 0.6910 0.01014 0.00260 -0.0458 0.4903 0.0432
3.250 0.7187 0.01009 0.00255 -0.0459 0.4839 0.0434
3.500 0.7466 0.01006 0.00253 -0.0459 0.4777 0.0446
3.750 0.7743 0.01009 0.00254 -0.0460 0.4716 0.0477
4.000 0.8019 0.01004 0.00256 -0.0461 0.4606 0.0871
4.250 0.8443 0.00851 0.00277 -0.0499 0.4456 1.0000
4.500 0.8708 0.00861 0.00284 -0.0497 0.4332 1.0000
4.750 0.8970 0.00876 0.00291 -0.0496 0.4137 1.0000
5.000 0.9229 0.00900 0.00303 -0.0495 0.3848 1.0000
5.250 0.9477 0.00954 0.00330 -0.0494 0.3294 1.0000
5.500 0.9610 0.01281 0.00514 -0.0492 0.0545 1.0000
5.750 0.9848 0.01350 0.00574 -0.0489 0.0221 1.0000
6.000 1.0095 0.01395 0.00628 -0.0486 0.0200 1.0000
6.250 1.0336 0.01453 0.00695 -0.0482 0.0183 1.0000
6.500 1.0565 0.01529 0.00784 -0.0477 0.0171 1.0000
6.750 1.0765 0.01646 0.00917 -0.0470 0.0157 1.0000
7.000 1.0952 0.01766 0.01048 -0.0462 0.0149 1.0000
7.250 1.1146 0.01865 0.01156 -0.0454 0.0145 1.0000
7.500 1.1316 0.01983 0.01283 -0.0443 0.0141 1.0000
7.750 1.1459 0.02122 0.01431 -0.0429 0.0138 1.0000
8.000 1.1573 0.02281 0.01600 -0.0412 0.0139 1.0000
8.250 1.1665 0.02465 0.01789 -0.0391 0.0142 1.0000
8.500 1.1836 0.02576 0.01902 -0.0376 0.0165 1.0000
9.250 1.1801 0.01911 0.01297 -0.0262 0.0213 1.0000
9.500 1.1968 0.02084 0.01480 -0.0249 0.0196 1.0000
9.750 1.2172 0.02320 0.01720 -0.0243 0.0180 1.0000
10.000 1.2471 0.02721 0.02119 -0.0250 0.0163 1.0000
10.250 1.3248 0.04547 0.03956 -0.0306 0.0149 1.0000
10.500 1.2650 0.03816 0.03272 -0.0221 0.0149 1.0000
10.750 1.2623 0.03782 0.03256 -0.0190 0.0142 1.0000
11.000 1.2559 0.03893 0.03382 -0.0157 0.0136 1.0000
11.250 1.2487 0.04097 0.03604 -0.0132 0.0131 1.0000
11.500 1.2404 0.04367 0.03891 -0.0111 0.0128 1.0000
11.750 1.2300 0.04675 0.04217 -0.0094 0.0125 1.0000
12.000 1.2176 0.05022 0.04581 -0.0080 0.0123 1.0000
12.250 1.2035 0.05405 0.04980 -0.0068 0.0122 1.0000
12.500 1.1878 0.05822 0.05414 -0.0060 0.0120 1.0000
12.750 1.1708 0.06269 0.05877 -0.0054 0.0119 1.0000
13.000 1.1523 0.06744 0.06367 -0.0051 0.0119 1.0000
13.250 1.1327 0.07241 0.06879 -0.0052 0.0118 1.0000
13.500 1.1123 0.07752 0.07405 -0.0056 0.0118 1.0000
13.750 1.0904 0.08274 0.07942 -0.0063 0.0119 1.0000
14.000 1.0675 0.08812 0.08494 -0.0074 0.0119 1.0000
14.250 1.0437 0.09362 0.09058 -0.0089 0.0119 1.0000
14.500 1.0181 0.09908 0.09618 -0.0106 0.0121 1.0000
14.750 0.9869 0.10395 0.10119 -0.0123 0.0123 1.0000
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Polar data table (+)
Polar graphs
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