NASA/AMES 63A108 MOD C AIRFOIL (a63a108c-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/AMES 63A108 MOD C AIRFOIL (a63a108c-il) Reynolds number: 1,000,000 Max Cl/Cd: 69.48 at α=7.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-a63a108c-il-1000000.txt Download as CSV file: xf-a63a108c-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/AMES 63A108 MOD C AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.7369 0.11018 0.10858 0.0382 1.0000 0.0126
-9.500 -0.9548 0.04413 0.04185 -0.0023 1.0000 0.0092
-9.250 -0.9861 0.03098 0.02763 -0.0019 1.0000 0.0091
-9.000 -0.9742 0.02694 0.02308 -0.0012 1.0000 0.0092
-8.750 -0.9592 0.02321 0.01884 -0.0005 1.0000 0.0094
-8.500 -0.9380 0.02128 0.01672 -0.0001 1.0000 0.0097
-8.250 -0.9127 0.02058 0.01596 0.0000 1.0000 0.0099
-8.000 -0.8873 0.01977 0.01505 0.0001 1.0000 0.0102
-7.750 -0.8619 0.01879 0.01394 0.0003 1.0000 0.0106
-7.500 -0.8363 0.01765 0.01262 0.0005 1.0000 0.0109
-7.250 -0.8101 0.01671 0.01153 0.0006 1.0000 0.0112
-7.000 -0.7831 0.01599 0.01067 0.0007 1.0000 0.0115
-6.750 -0.7574 0.01467 0.00922 0.0009 1.0000 0.0118
-6.500 -0.7305 0.01392 0.00844 0.0009 1.0000 0.0122
-6.250 -0.7028 0.01339 0.00787 0.0009 1.0000 0.0126
-6.000 -0.6750 0.01285 0.00729 0.0008 1.0000 0.0130
-5.750 -0.6469 0.01235 0.00673 0.0008 1.0000 0.0135
-5.500 -0.6185 0.01191 0.00624 0.0007 1.0000 0.0139
-5.250 -0.5907 0.01121 0.00553 0.0005 1.0000 0.0147
-5.000 -0.5618 0.01088 0.00520 0.0003 1.0000 0.0154
-4.750 -0.5328 0.01056 0.00485 0.0001 1.0000 0.0163
-4.500 -0.5038 0.01007 0.00434 -0.0002 1.0000 0.0171
-4.250 -0.4744 0.00973 0.00401 -0.0006 1.0000 0.0181
-4.000 -0.4446 0.00944 0.00371 -0.0010 1.0000 0.0192
-3.750 -0.4147 0.00906 0.00335 -0.0014 1.0000 0.0208
-3.250 -0.3553 0.00867 0.00296 -0.0020 0.9704 0.0252
-3.000 -0.3302 0.00852 0.00280 -0.0012 0.9582 0.0285
-2.750 -0.3048 0.00836 0.00264 -0.0004 0.9475 0.0328
-2.500 -0.2791 0.00818 0.00248 0.0003 0.9381 0.0399
-2.250 -0.2526 0.00798 0.00234 0.0008 0.9283 0.0540
-2.000 -0.2264 0.00772 0.00219 0.0013 0.9169 0.0857
-1.750 -0.1999 0.00740 0.00206 0.0017 0.9046 0.1406
-1.500 -0.1730 0.00692 0.00191 0.0018 0.8926 0.2383
-1.250 -0.1455 0.00623 0.00175 0.0016 0.8818 0.3976
-1.000 -0.1182 0.00571 0.00167 0.0016 0.8715 0.5312
-0.750 -0.0905 0.00543 0.00162 0.0018 0.8579 0.6108
-0.500 -0.0624 0.00525 0.00159 0.0019 0.8442 0.6665
-0.250 -0.0346 0.00510 0.00157 0.0022 0.8292 0.7183
0.000 -0.0068 0.00501 0.00154 0.0025 0.8109 0.7576
0.250 0.0208 0.00492 0.00152 0.0028 0.7900 0.7970
0.500 0.0480 0.00485 0.00149 0.0032 0.7637 0.8322
0.750 0.0736 0.00474 0.00148 0.0041 0.7315 0.8835
1.000 0.0971 0.00460 0.00138 0.0057 0.6900 0.9640
1.250 0.1338 0.00475 0.00129 0.0038 0.6186 1.0000
1.500 0.1625 0.00511 0.00132 0.0036 0.5344 1.0000
1.750 0.1913 0.00552 0.00139 0.0032 0.4428 1.0000
2.000 0.2202 0.00595 0.00148 0.0028 0.3553 1.0000
2.250 0.2492 0.00631 0.00157 0.0024 0.2868 1.0000
2.500 0.2781 0.00662 0.00167 0.0021 0.2372 1.0000
2.750 0.3071 0.00690 0.00178 0.0018 0.1969 1.0000
3.000 0.3361 0.00715 0.00189 0.0015 0.1653 1.0000
3.250 0.3650 0.00740 0.00201 0.0013 0.1396 1.0000
3.500 0.3938 0.00764 0.00214 0.0010 0.1187 1.0000
3.750 0.4227 0.00786 0.00228 0.0008 0.1018 1.0000
4.000 0.4515 0.00811 0.00243 0.0006 0.0867 1.0000
4.250 0.4802 0.00834 0.00260 0.0004 0.0747 1.0000
4.500 0.5089 0.00858 0.00278 0.0002 0.0651 1.0000
4.750 0.5375 0.00883 0.00297 0.0000 0.0575 1.0000
5.000 0.5661 0.00907 0.00318 -0.0001 0.0512 1.0000
5.250 0.5947 0.00932 0.00340 -0.0003 0.0459 1.0000
5.500 0.6231 0.00958 0.00364 -0.0004 0.0414 1.0000
5.750 0.6515 0.00987 0.00390 -0.0006 0.0376 1.0000
6.000 0.6798 0.01015 0.00417 -0.0007 0.0343 1.0000
6.250 0.7080 0.01042 0.00443 -0.0008 0.0315 1.0000
6.500 0.7361 0.01074 0.00476 -0.0009 0.0290 1.0000
6.750 0.7639 0.01113 0.00513 -0.0010 0.0268 1.0000
7.000 0.7920 0.01142 0.00545 -0.0011 0.0252 1.0000
7.250 0.8195 0.01188 0.00591 -0.0012 0.0236 1.0000
7.500 0.8472 0.01221 0.00628 -0.0012 0.0225 1.0000
7.750 0.8747 0.01259 0.00667 -0.0013 0.0215 1.0000
8.000 0.9016 0.01315 0.00725 -0.0013 0.0205 1.0000
8.250 0.9289 0.01351 0.00766 -0.0013 0.0196 1.0000
8.500 0.9559 0.01393 0.00812 -0.0013 0.0189 1.0000
8.750 0.9823 0.01448 0.00868 -0.0013 0.0182 1.0000
9.000 1.0081 0.01519 0.00945 -0.0012 0.0177 1.0000
9.250 1.0345 0.01569 0.01003 -0.0012 0.0172 1.0000
9.500 1.0605 0.01624 0.01065 -0.0011 0.0167 1.0000
9.750 1.0862 0.01681 0.01127 -0.0009 0.0163 1.0000
10.000 1.1114 0.01744 0.01194 -0.0008 0.0159 1.0000
10.250 1.1351 0.01837 0.01292 -0.0006 0.0155 1.0000
10.500 1.1582 0.01937 0.01403 -0.0003 0.0151 1.0000
10.750 1.1827 0.02002 0.01478 0.0000 0.0148 1.0000
11.000 1.2065 0.02073 0.01559 0.0003 0.0144 1.0000
11.250 1.2295 0.02156 0.01653 0.0006 0.0141 1.0000
11.500 1.2522 0.02236 0.01741 0.0009 0.0138 1.0000
11.750 1.2749 0.02306 0.01816 0.0013 0.0135 1.0000
12.000 1.2958 0.02399 0.01914 0.0017 0.0131 1.0000
12.250 1.3106 0.02585 0.02116 0.0026 0.0128 1.0000
12.500 1.3323 0.02652 0.02195 0.0030 0.0125 1.0000
12.750 1.3515 0.02750 0.02306 0.0036 0.0123 1.0000
13.000 1.3688 0.02864 0.02433 0.0043 0.0120 1.0000
13.250 1.3840 0.02994 0.02576 0.0052 0.0118 1.0000
13.500 1.3973 0.03131 0.02726 0.0061 0.0116 1.0000
13.750 1.4087 0.03274 0.02881 0.0072 0.0115 1.0000
14.000 1.4171 0.03424 0.03041 0.0084 0.0113 1.0000
14.250 1.4166 0.03604 0.03233 0.0105 0.0112 1.0000
14.500 1.4133 0.03833 0.03475 0.0117 0.0111 1.0000
14.750 1.4075 0.04140 0.03796 0.0117 0.0110 1.0000
15.000 1.3992 0.04538 0.04208 0.0102 0.0110 1.0000
15.250 1.3854 0.05106 0.04794 0.0067 0.0109 1.0000
15.500 1.3635 0.05968 0.05678 -0.0002 0.0109 1.0000
15.750 1.3125 0.07590 0.07333 -0.0123 0.0111 1.0000
16.000 1.2096 0.10150 0.09930 -0.0277 0.0115 1.0000
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Polar data table (+)
Polar graphs
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