NACA 63-015A AIRFOIL (n63015a-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 63-015A AIRFOIL (n63015a-il) Reynolds number: 50,000 Max Cl/Cd: 28.88 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63015a-il-50000.txt Download as CSV file: xf-n63015a-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63-015A 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.7044 0.10730 0.09997 -0.0218 1.0000 0.1414
-12.000 -0.7059 0.10116 0.09384 -0.0235 1.0000 0.1402
-11.750 -0.8157 0.08347 0.07602 -0.0358 1.0000 0.1263
-11.500 -0.8340 0.07797 0.07044 -0.0366 1.0000 0.1251
-11.250 -0.8578 0.07302 0.06536 -0.0365 1.0000 0.1240
-11.000 -0.8821 0.06877 0.06094 -0.0348 1.0000 0.1230
-10.750 -0.9025 0.06487 0.05680 -0.0325 1.0000 0.1224
-10.500 -0.9162 0.06108 0.05272 -0.0301 1.0000 0.1224
-10.250 -0.9251 0.05754 0.04883 -0.0276 1.0000 0.1231
-10.000 -0.9302 0.05422 0.04510 -0.0249 1.0000 0.1245
-9.750 -0.9322 0.05116 0.04155 -0.0221 1.0000 0.1262
-9.500 -0.9200 0.04775 0.03795 -0.0207 1.0000 0.1292
-9.250 -0.9007 0.04497 0.03507 -0.0198 1.0000 0.1340
-9.000 -0.8902 0.04255 0.03223 -0.0177 1.0000 0.1398
-8.750 -0.8675 0.03992 0.02961 -0.0171 1.0000 0.1482
-8.500 -0.8471 0.03756 0.02709 -0.0160 1.0000 0.1589
-8.250 -0.8225 0.03536 0.02492 -0.0153 1.0000 0.1743
-8.000 -0.7991 0.03332 0.02306 -0.0144 1.0000 0.1959
-7.750 -0.7792 0.03128 0.02131 -0.0130 1.0000 0.2261
-7.500 -0.7678 0.02930 0.01979 -0.0103 1.0000 0.2677
-7.250 -0.7669 0.02752 0.01870 -0.0059 1.0000 0.3262
-7.000 -0.7706 0.02669 0.01878 0.0003 1.0000 0.4104
-6.750 -0.7673 0.02760 0.02014 0.0071 1.0000 0.4945
-6.500 -0.7414 0.03037 0.02298 0.0124 1.0000 0.5533
-6.250 -0.7183 0.03245 0.02495 0.0171 1.0000 0.5931
-6.000 -0.6787 0.03527 0.02757 0.0205 1.0000 0.6259
-5.750 -0.6060 0.03936 0.03131 0.0211 1.0000 0.6567
-5.500 -0.5144 0.04279 0.03425 0.0184 1.0000 0.6865
-5.250 -0.4703 0.04318 0.03438 0.0178 1.0000 0.7106
-5.000 -0.4181 0.04330 0.03425 0.0155 1.0000 0.7327
-4.750 -0.3931 0.04274 0.03356 0.0160 1.0000 0.7517
-4.500 -0.3711 0.04209 0.03282 0.0165 1.0000 0.7689
-4.250 -0.3498 0.04142 0.03208 0.0170 1.0000 0.7848
-4.000 -0.3320 0.04074 0.03135 0.0178 1.0000 0.7995
-3.750 -0.3239 0.04010 0.03066 0.0199 1.0000 0.8131
-3.500 -0.3313 0.03951 0.03008 0.0241 1.0000 0.8261
-3.250 -0.3325 0.03898 0.02952 0.0273 1.0000 0.8392
-3.000 -0.3050 0.03846 0.02892 0.0260 1.0000 0.8513
-2.750 -0.2961 0.03792 0.02835 0.0274 1.0000 0.8631
-2.500 -0.2987 0.03740 0.02780 0.0304 1.0000 0.8755
-2.250 -0.2801 0.03696 0.02730 0.0300 1.0000 0.8872
-2.000 -0.2569 0.03654 0.02683 0.0287 1.0000 0.8985
-1.750 -0.2450 0.03614 0.02638 0.0292 1.0000 0.9104
-1.500 -0.2324 0.03580 0.02600 0.0294 1.0000 0.9229
-1.250 -0.1959 0.03555 0.02570 0.0254 1.0000 0.9337
-1.000 -0.1690 0.03533 0.02545 0.0229 1.0000 0.9454
-0.750 -0.1418 0.03518 0.02526 0.0201 1.0000 0.9577
-0.500 -0.1101 0.03511 0.02516 0.0165 1.0000 0.9701
-0.250 -0.0699 0.03509 0.02512 0.0110 1.0000 0.9821
0.000 0.0000 0.03510 0.02513 0.0000 0.9919 0.9919
0.250 0.0699 0.03509 0.02512 -0.0110 0.9821 1.0000
0.500 0.1100 0.03510 0.02516 -0.0164 0.9701 1.0000
0.750 0.1417 0.03517 0.02525 -0.0201 0.9577 1.0000
1.000 0.1690 0.03532 0.02544 -0.0229 0.9454 1.0000
1.250 0.1959 0.03554 0.02569 -0.0254 0.9337 1.0000
1.500 0.2324 0.03579 0.02599 -0.0294 0.9230 1.0000
1.750 0.2449 0.03612 0.02637 -0.0291 0.9104 1.0000
2.000 0.2569 0.03653 0.02681 -0.0287 0.8985 1.0000
2.250 0.2802 0.03695 0.02728 -0.0301 0.8873 1.0000
2.500 0.2985 0.03739 0.02778 -0.0304 0.8755 1.0000
2.750 0.2960 0.03791 0.02833 -0.0274 0.8632 1.0000
3.000 0.3050 0.03844 0.02890 -0.0260 0.8513 1.0000
3.250 0.3327 0.03896 0.02951 -0.0274 0.8392 1.0000
3.500 0.3311 0.03950 0.03006 -0.0241 0.8262 1.0000
3.750 0.3238 0.04008 0.03064 -0.0198 0.8131 1.0000
4.000 0.3321 0.04073 0.03133 -0.0178 0.7995 1.0000
4.250 0.3499 0.04140 0.03206 -0.0170 0.7848 1.0000
4.500 0.3713 0.04207 0.03281 -0.0165 0.7690 1.0000
4.750 0.3934 0.04272 0.03354 -0.0160 0.7517 1.0000
5.000 0.4184 0.04328 0.03423 -0.0156 0.7327 1.0000
5.250 0.4707 0.04316 0.03436 -0.0178 0.7106 1.0000
5.500 0.5148 0.04277 0.03423 -0.0184 0.6865 1.0000
5.750 0.6062 0.03934 0.03129 -0.0212 0.6567 1.0000
6.000 0.6779 0.03531 0.02761 -0.0205 0.6260 1.0000
6.250 0.7183 0.03244 0.02494 -0.0171 0.5931 1.0000
6.500 0.7413 0.03037 0.02298 -0.0124 0.5533 1.0000
6.750 0.7672 0.02760 0.02013 -0.0071 0.4946 1.0000
7.000 0.7706 0.02668 0.01877 -0.0003 0.4105 1.0000
7.250 0.7669 0.02752 0.01870 0.0059 0.3263 1.0000
7.500 0.7678 0.02930 0.01978 0.0103 0.2677 1.0000
7.750 0.7792 0.03128 0.02131 0.0130 0.2260 1.0000
8.000 0.7991 0.03332 0.02306 0.0144 0.1959 1.0000
8.250 0.8226 0.03536 0.02492 0.0153 0.1742 1.0000
8.500 0.8472 0.03756 0.02709 0.0159 0.1589 1.0000
8.750 0.8676 0.03992 0.02961 0.0171 0.1482 1.0000
9.000 0.8902 0.04255 0.03223 0.0177 0.1398 1.0000
9.250 0.9008 0.04498 0.03507 0.0198 0.1340 1.0000
9.500 0.9202 0.04776 0.03795 0.0207 0.1292 1.0000
9.750 0.9323 0.05116 0.04155 0.0221 0.1261 1.0000
10.000 0.9303 0.05423 0.04511 0.0249 0.1245 1.0000
10.250 0.9252 0.05755 0.04884 0.0276 0.1231 1.0000
10.500 0.9164 0.06109 0.05273 0.0301 0.1224 1.0000
10.750 0.9028 0.06489 0.05682 0.0324 0.1224 1.0000
11.000 0.8824 0.06881 0.06097 0.0347 0.1230 1.0000
11.250 0.8581 0.07307 0.06541 0.0364 0.1240 1.0000
11.500 0.8344 0.07803 0.07049 0.0365 0.1251 1.0000
11.750 0.8163 0.08354 0.07609 0.0357 0.1263 1.0000
12.000 0.7066 0.10127 0.09395 0.0233 0.1402 1.0000
12.250 0.7052 0.10740 0.10008 0.0216 0.1414 1.0000
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Polar data table (+)
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