NACA M5 AIRFOIL (m5-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: NACA M5 AIRFOIL (m5-il) Reynolds number: 100,000 Max Cl/Cd: 45.98 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m5-il-100000-n5.txt Download as CSV file: xf-m5-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M5 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.6071 0.09320 0.08862 0.0098 1.0000 0.0373
-8.250 -0.6116 0.08803 0.08347 0.0058 1.0000 0.0368
-8.000 -0.6132 0.08269 0.07812 0.0021 1.0000 0.0365
-7.750 -0.6120 0.07719 0.07257 -0.0014 1.0000 0.0364
-7.500 -0.6073 0.07209 0.06737 -0.0041 1.0000 0.0373
-7.250 -0.6004 0.06678 0.06191 -0.0067 1.0000 0.0389
-7.000 -0.5921 0.06081 0.05570 -0.0088 1.0000 0.0402
-6.750 -0.5822 0.05462 0.04918 -0.0101 1.0000 0.0408
-6.500 -0.5702 0.04789 0.04182 -0.0105 1.0000 0.0425
-6.250 -0.5546 0.04552 0.03945 -0.0105 1.0000 0.0448
-6.000 -0.5354 0.04331 0.03708 -0.0104 1.0000 0.0468
-5.750 -0.5165 0.03954 0.03292 -0.0101 1.0000 0.0480
-5.500 -0.4956 0.03648 0.02946 -0.0098 1.0000 0.0507
-5.250 -0.4734 0.03331 0.02575 -0.0092 1.0000 0.0532
-5.000 -0.4500 0.03047 0.02237 -0.0087 1.0000 0.0540
-4.750 -0.4251 0.02823 0.01965 -0.0083 1.0000 0.0548
-4.500 -0.3999 0.02627 0.01733 -0.0081 1.0000 0.0563
-4.250 -0.3750 0.02484 0.01584 -0.0082 1.0000 0.0588
-4.000 -0.3490 0.02357 0.01439 -0.0081 1.0000 0.0604
-3.750 -0.3230 0.02225 0.01285 -0.0079 1.0000 0.0612
-3.500 -0.2884 0.02099 0.01138 -0.0095 0.9773 0.0625
-3.250 -0.2520 0.01988 0.01008 -0.0112 0.9559 0.0642
-3.000 -0.2178 0.01896 0.00899 -0.0124 0.9360 0.0663
-2.750 -0.1877 0.01824 0.00814 -0.0126 0.9151 0.0684
-2.500 -0.1604 0.01757 0.00741 -0.0122 0.8953 0.0707
-2.250 -0.1354 0.01706 0.00689 -0.0115 0.8772 0.0754
-2.000 -0.1105 0.01670 0.00647 -0.0106 0.8595 0.0817
-1.750 -0.0862 0.01625 0.00599 -0.0097 0.8433 0.0873
-1.500 -0.0616 0.01589 0.00558 -0.0087 0.8280 0.0945
-1.250 -0.0367 0.01554 0.00521 -0.0078 0.8135 0.1058
-1.000 -0.0114 0.01519 0.00492 -0.0071 0.7997 0.1304
-0.750 0.0135 0.01456 0.00469 -0.0065 0.7867 0.2246
-0.500 0.1039 0.01283 0.00514 -0.0177 0.7759 0.9963
-0.250 0.1344 0.01286 0.00498 -0.0182 0.7620 1.0000
0.000 0.1593 0.01289 0.00485 -0.0175 0.7485 1.0000
0.250 0.1842 0.01293 0.00475 -0.0169 0.7355 1.0000
0.500 0.2092 0.01298 0.00467 -0.0162 0.7228 1.0000
0.750 0.2343 0.01304 0.00462 -0.0156 0.7107 1.0000
1.000 0.2591 0.01311 0.00459 -0.0149 0.6990 1.0000
1.250 0.2841 0.01319 0.00458 -0.0143 0.6875 1.0000
1.500 0.3096 0.01327 0.00461 -0.0138 0.6754 1.0000
1.750 0.3349 0.01337 0.00466 -0.0133 0.6638 1.0000
2.000 0.3602 0.01347 0.00471 -0.0127 0.6526 1.0000
2.250 0.3854 0.01359 0.00479 -0.0121 0.6419 1.0000
2.500 0.4108 0.01371 0.00490 -0.0116 0.6301 1.0000
2.750 0.4363 0.01384 0.00504 -0.0111 0.6184 1.0000
3.000 0.4617 0.01398 0.00520 -0.0106 0.6069 1.0000
3.250 0.4870 0.01412 0.00535 -0.0101 0.5956 1.0000
3.500 0.5122 0.01428 0.00551 -0.0095 0.5842 1.0000
3.750 0.5377 0.01444 0.00575 -0.0090 0.5715 1.0000
4.000 0.5631 0.01461 0.00598 -0.0085 0.5585 1.0000
4.250 0.5884 0.01479 0.00622 -0.0080 0.5446 1.0000
4.500 0.6135 0.01495 0.00646 -0.0074 0.5288 1.0000
4.750 0.6384 0.01512 0.00666 -0.0067 0.5116 1.0000
5.000 0.6633 0.01528 0.00690 -0.0061 0.4886 1.0000
5.250 0.6876 0.01544 0.00705 -0.0053 0.4578 1.0000
5.500 0.7116 0.01567 0.00727 -0.0046 0.4216 1.0000
5.750 0.7354 0.01601 0.00755 -0.0039 0.3800 1.0000
6.000 0.7587 0.01650 0.00793 -0.0033 0.3276 1.0000
6.250 0.7805 0.01732 0.00847 -0.0027 0.2515 1.0000
6.750 0.8156 0.02084 0.01084 -0.0015 0.0728 1.0000
7.000 0.8344 0.02228 0.01222 -0.0007 0.0569 1.0000
7.250 0.8540 0.02348 0.01353 0.0002 0.0469 1.0000
7.500 0.8713 0.02495 0.01511 0.0014 0.0415 1.0000
7.750 0.8896 0.02618 0.01644 0.0024 0.0366 1.0000
8.000 0.9052 0.02776 0.01808 0.0036 0.0332 1.0000
8.250 0.9231 0.02924 0.01972 0.0049 0.0310 1.0000
8.500 0.9411 0.03079 0.02140 0.0061 0.0288 1.0000
8.750 0.9584 0.03227 0.02294 0.0071 0.0265 1.0000
9.000 0.9744 0.03438 0.02507 0.0082 0.0247 1.0000
9.250 0.9927 0.03645 0.02743 0.0093 0.0238 1.0000
9.500 1.0093 0.03890 0.03019 0.0105 0.0230 1.0000
9.750 1.0231 0.04159 0.03324 0.0118 0.0222 1.0000
10.000 1.0332 0.04429 0.03630 0.0132 0.0212 1.0000
10.250 1.0400 0.04693 0.03932 0.0146 0.0202 1.0000
10.500 1.0435 0.04958 0.04225 0.0160 0.0194 1.0000
10.750 1.0423 0.05248 0.04543 0.0174 0.0190 1.0000
11.000 1.0342 0.05556 0.04877 0.0193 0.0188 1.0000
11.250 1.0226 0.05905 0.05252 0.0202 0.0187 1.0000
11.500 1.0088 0.06313 0.05685 0.0200 0.0186 1.0000
11.750 0.9929 0.06790 0.06186 0.0186 0.0186 1.0000
12.000 0.9747 0.07348 0.06768 0.0161 0.0187 1.0000
12.250 0.9532 0.08020 0.07462 0.0122 0.0189 1.0000
12.500 0.9271 0.08859 0.08322 0.0067 0.0192 1.0000
12.750 0.8969 0.09916 0.09391 -0.0006 0.0196 1.0000
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Polar data table (+)
Polar graphs
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