NACA M17 AIRFOIL (m17-il) Xfoil prediction polar at RE=50,000 Ncrit=5
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Airfoil: NACA M17 AIRFOIL (m17-il) Reynolds number: 50,000 Max Cl/Cd: 35.59 at α=8.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m17-il-50000-n5.txt Download as CSV file: xf-m17-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M17 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4308 0.10887 0.10263 -0.0019 1.0000 0.0894
-8.250 -0.4333 0.10683 0.10069 -0.0064 1.0000 0.0925
-8.000 -0.4369 0.10519 0.09913 -0.0128 1.0000 0.0935
-7.750 -0.4226 0.09897 0.09296 -0.0099 1.0000 0.0965
-7.500 -0.4129 0.09515 0.08918 -0.0105 1.0000 0.1004
-7.250 -0.4060 0.09187 0.08591 -0.0137 1.0000 0.1043
-7.000 -0.4010 0.09008 0.08412 -0.0215 1.0000 0.1080
-6.750 -0.3904 0.08544 0.07951 -0.0228 1.0000 0.1096
-6.500 -0.3782 0.08106 0.07519 -0.0215 1.0000 0.1134
-6.250 -0.3668 0.07896 0.07301 -0.0267 1.0000 0.1223
-6.000 -0.3571 0.07463 0.06877 -0.0263 1.0000 0.1263
-5.750 -0.3486 0.07190 0.06608 -0.0267 1.0000 0.1335
-5.250 -0.2976 0.06598 0.05968 -0.0378 0.9623 0.1527
-5.000 -0.2738 0.06054 0.05430 -0.0388 0.9468 0.1572
-4.500 -0.1980 0.05104 0.04375 -0.0486 0.9150 0.0962
-4.000 -0.1382 0.04409 0.03580 -0.0513 0.8869 0.0778
-3.750 -0.1128 0.04155 0.03290 -0.0516 0.8737 0.0775
-3.500 -0.0870 0.03934 0.03022 -0.0516 0.8614 0.0797
-3.250 -0.0624 0.03727 0.02779 -0.0513 0.8491 0.0801
-3.000 -0.0360 0.03546 0.02544 -0.0508 0.8373 0.0786
-2.750 -0.0107 0.03369 0.02333 -0.0504 0.8261 0.0782
-2.500 0.0152 0.03211 0.02139 -0.0498 0.8159 0.0779
-2.250 0.0416 0.03070 0.01959 -0.0493 0.8060 0.0779
-2.000 0.0688 0.02946 0.01801 -0.0489 0.7956 0.0782
-1.750 0.0965 0.02837 0.01656 -0.0485 0.7864 0.0788
-1.500 0.1234 0.02737 0.01543 -0.0482 0.7772 0.0815
-1.250 0.1514 0.02674 0.01466 -0.0482 0.7676 0.0869
-1.000 0.1828 0.02605 0.01363 -0.0482 0.7598 0.0914
-0.750 0.2141 0.02545 0.01277 -0.0485 0.7503 0.0937
-0.500 0.2426 0.02483 0.01208 -0.0483 0.7424 0.0972
-0.250 0.2691 0.02442 0.01158 -0.0479 0.7338 0.1023
0.000 0.2947 0.02409 0.01113 -0.0472 0.7259 0.1094
0.250 0.3195 0.02377 0.01077 -0.0465 0.7182 0.1220
0.500 0.3436 0.02343 0.01060 -0.0460 0.7102 0.1549
1.000 0.4392 0.02141 0.01028 -0.0538 0.6951 1.0000
1.250 0.4635 0.02168 0.01035 -0.0530 0.6880 1.0000
1.500 0.4878 0.02206 0.01059 -0.0526 0.6800 1.0000
1.750 0.5120 0.02236 0.01077 -0.0519 0.6733 1.0000
2.000 0.5360 0.02278 0.01111 -0.0515 0.6655 1.0000
2.250 0.5600 0.02311 0.01136 -0.0508 0.6588 1.0000
2.500 0.5838 0.02357 0.01180 -0.0503 0.6516 1.0000
2.750 0.6077 0.02396 0.01215 -0.0497 0.6446 1.0000
3.000 0.6313 0.02442 0.01260 -0.0491 0.6377 1.0000
3.250 0.6549 0.02490 0.01310 -0.0486 0.6306 1.0000
3.500 0.6786 0.02533 0.01356 -0.0480 0.6242 1.0000
3.750 0.7016 0.02591 0.01421 -0.0475 0.6165 1.0000
4.000 0.7256 0.02629 0.01460 -0.0467 0.6106 1.0000
4.250 0.7477 0.02699 0.01545 -0.0464 0.6021 1.0000
4.500 0.7721 0.02731 0.01581 -0.0455 0.5964 1.0000
4.750 0.7933 0.02811 0.01676 -0.0451 0.5872 1.0000
5.000 0.8171 0.02849 0.01724 -0.0442 0.5805 1.0000
5.500 0.8604 0.02970 0.01879 -0.0428 0.5620 1.0000
5.750 0.8851 0.02978 0.01898 -0.0416 0.5537 1.0000
6.000 0.9055 0.03035 0.01975 -0.0406 0.5419 1.0000
6.250 0.9268 0.03078 0.02040 -0.0396 0.5307 1.0000
6.500 0.9506 0.03092 0.02072 -0.0384 0.5210 1.0000
6.750 0.9732 0.03120 0.02123 -0.0373 0.5104 1.0000
7.000 0.9930 0.03183 0.02213 -0.0362 0.4984 1.0000
7.250 1.0137 0.03235 0.02298 -0.0351 0.4868 1.0000
7.500 1.0358 0.03266 0.02357 -0.0339 0.4754 1.0000
7.750 1.0567 0.03249 0.02369 -0.0321 0.4573 1.0000
8.000 1.0744 0.03031 0.02145 -0.0280 0.4055 1.0000
8.250 1.0830 0.03043 0.02154 -0.0252 0.3440 1.0000
8.500 1.0851 0.03148 0.02203 -0.0219 0.2479 1.0000
8.750 1.0748 0.03433 0.02415 -0.0190 0.1659 1.0000
9.000 1.0612 0.03762 0.02700 -0.0166 0.1090 1.0000
9.250 1.0484 0.04110 0.03018 -0.0149 0.0846 1.0000
9.500 1.0400 0.04456 0.03351 -0.0140 0.0729 1.0000
9.750 1.0351 0.04783 0.03679 -0.0132 0.0661 1.0000
10.000 1.0308 0.05111 0.04011 -0.0127 0.0610 1.0000
10.250 1.0274 0.05435 0.04336 -0.0121 0.0570 1.0000
10.500 1.0276 0.05720 0.04636 -0.0115 0.0528 1.0000
10.750 1.0274 0.06012 0.04934 -0.0109 0.0494 1.0000
11.000 1.0273 0.06296 0.05221 -0.0100 0.0470 1.0000
11.250 1.0345 0.06502 0.05440 -0.0085 0.0451 1.0000
11.500 1.0450 0.06682 0.05643 -0.0068 0.0428 1.0000
11.750 1.0553 0.06877 0.05855 -0.0053 0.0402 1.0000
12.000 1.0662 0.07073 0.06063 -0.0038 0.0377 1.0000
12.250 1.0878 0.07218 0.06222 -0.0011 0.0359 1.0000
12.500 1.0985 0.07501 0.06540 0.0000 0.0351 1.0000
12.750 1.1023 0.07857 0.06930 0.0004 0.0346 1.0000
13.000 1.0997 0.08276 0.07380 0.0001 0.0342 1.0000
13.250 1.0924 0.08746 0.07879 -0.0009 0.0339 1.0000
13.500 1.0818 0.09265 0.08425 -0.0025 0.0337 1.0000
13.750 1.0689 0.09831 0.09016 -0.0046 0.0337 1.0000
14.000 1.0541 0.10449 0.09656 -0.0074 0.0337 1.0000
14.250 1.0380 0.11123 0.10350 -0.0107 0.0339 1.0000
14.500 1.0212 0.11862 0.11106 -0.0147 0.0341 1.0000
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Polar data table (+)
Polar graphs
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