NACA 63012A AIRFOIL (n63012a-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 63012A AIRFOIL (n63012a-il) Reynolds number: 500,000 Max Cl/Cd: 58.08 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63012a-il-500000.txt Download as CSV file: xf-n63012a-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63012A 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
------ -------- --------- --------- -------- -------- --------
-13.250 -0.9092 0.07303 0.07015 -0.0230 1.0000 0.0165
-13.000 -0.9268 0.06723 0.06418 -0.0266 1.0000 0.0164
-12.750 -0.9420 0.06241 0.05919 -0.0290 1.0000 0.0164
-12.500 -0.9564 0.05815 0.05476 -0.0304 1.0000 0.0163
-12.250 -0.9708 0.05427 0.05070 -0.0309 1.0000 0.0163
-11.750 -1.0007 0.04656 0.04258 -0.0288 1.0000 0.0162
-11.500 -1.0121 0.04425 0.04004 -0.0262 1.0000 0.0163
-11.250 -1.0268 0.04021 0.03568 -0.0230 1.0000 0.0164
-11.000 -1.0327 0.03593 0.03106 -0.0205 1.0000 0.0166
-10.750 -1.0257 0.03330 0.02825 -0.0189 1.0000 0.0170
-10.500 -1.0135 0.03154 0.02636 -0.0176 1.0000 0.0173
-10.250 -0.9998 0.02976 0.02440 -0.0162 1.0000 0.0175
-10.000 -0.9841 0.02813 0.02262 -0.0150 1.0000 0.0178
-9.750 -0.9668 0.02662 0.02097 -0.0139 1.0000 0.0182
-9.500 -0.9484 0.02520 0.01938 -0.0129 1.0000 0.0185
-9.250 -0.9292 0.02384 0.01790 -0.0119 1.0000 0.0189
-9.000 -0.9093 0.02261 0.01654 -0.0109 1.0000 0.0194
-8.750 -0.8887 0.02165 0.01546 -0.0099 1.0000 0.0200
-8.500 -0.8675 0.02088 0.01459 -0.0090 1.0000 0.0206
-8.250 -0.8462 0.02018 0.01378 -0.0081 1.0000 0.0210
-8.000 -0.8304 0.01826 0.01178 -0.0064 1.0000 0.0217
-7.750 -0.8128 0.01727 0.01077 -0.0049 1.0000 0.0224
-7.500 -0.7944 0.01657 0.01005 -0.0034 1.0000 0.0231
-7.250 -0.7763 0.01595 0.00941 -0.0018 1.0000 0.0240
-7.000 -0.7585 0.01537 0.00878 -0.0001 1.0000 0.0249
-6.750 -0.7407 0.01487 0.00824 0.0016 1.0000 0.0260
-6.500 -0.7234 0.01445 0.00778 0.0034 1.0000 0.0270
-6.250 -0.7119 0.01361 0.00693 0.0062 1.0000 0.0290
-6.000 -0.6781 0.01306 0.00635 0.0045 0.9968 0.0318
-5.750 -0.6420 0.01241 0.00567 0.0023 0.9928 0.0363
-5.500 -0.6062 0.01189 0.00515 0.0003 0.9877 0.0438
-5.250 -0.5695 0.01121 0.00458 -0.0021 0.9837 0.0629
-5.000 -0.5372 0.01050 0.00410 -0.0035 0.9761 0.1061
-4.750 -0.5053 0.00959 0.00361 -0.0052 0.9690 0.1928
-4.500 -0.4792 0.00856 0.00313 -0.0056 0.9575 0.3164
-4.250 -0.4542 0.00782 0.00283 -0.0055 0.9445 0.4242
-4.000 -0.4284 0.00745 0.00267 -0.0051 0.9311 0.4884
-3.750 -0.4026 0.00725 0.00254 -0.0046 0.9175 0.5273
-3.500 -0.3767 0.00712 0.00243 -0.0041 0.9039 0.5542
-3.250 -0.3505 0.00704 0.00233 -0.0036 0.8908 0.5763
-3.000 -0.3246 0.00696 0.00226 -0.0031 0.8782 0.5976
-2.750 -0.2984 0.00691 0.00219 -0.0026 0.8657 0.6173
-2.500 -0.2719 0.00688 0.00214 -0.0022 0.8534 0.6366
-2.250 -0.2454 0.00685 0.00210 -0.0018 0.8420 0.6530
-2.000 -0.2187 0.00684 0.00207 -0.0014 0.8314 0.6681
-1.750 -0.1917 0.00684 0.00203 -0.0012 0.8205 0.6817
-1.500 -0.1647 0.00682 0.00201 -0.0009 0.8100 0.6941
-1.250 -0.1375 0.00681 0.00198 -0.0007 0.8002 0.7047
-1.000 -0.1100 0.00681 0.00194 -0.0005 0.7901 0.7143
-0.750 -0.0825 0.00680 0.00193 -0.0004 0.7802 0.7240
-0.500 -0.0551 0.00680 0.00191 -0.0002 0.7710 0.7328
-0.250 -0.0275 0.00681 0.00189 -0.0001 0.7610 0.7426
0.000 0.0000 0.00679 0.00190 0.0000 0.7516 0.7516
0.250 0.0275 0.00681 0.00189 0.0001 0.7426 0.7610
0.500 0.0551 0.00680 0.00191 0.0002 0.7328 0.7710
0.750 0.0826 0.00680 0.00193 0.0004 0.7239 0.7801
1.000 0.1101 0.00681 0.00194 0.0005 0.7143 0.7901
1.250 0.1375 0.00681 0.00198 0.0007 0.7047 0.8002
1.500 0.1647 0.00682 0.00201 0.0009 0.6941 0.8099
1.750 0.1918 0.00684 0.00203 0.0011 0.6817 0.8205
2.000 0.2188 0.00684 0.00207 0.0014 0.6681 0.8314
2.250 0.2455 0.00685 0.00210 0.0018 0.6531 0.8420
2.500 0.2721 0.00688 0.00214 0.0022 0.6366 0.8534
2.750 0.2985 0.00691 0.00219 0.0026 0.6173 0.8656
3.000 0.3247 0.00696 0.00226 0.0030 0.5975 0.8782
3.250 0.3506 0.00704 0.00233 0.0036 0.5763 0.8908
3.500 0.3768 0.00712 0.00243 0.0041 0.5541 0.9039
3.750 0.4027 0.00725 0.00254 0.0046 0.5273 0.9175
4.000 0.4285 0.00745 0.00267 0.0051 0.4883 0.9311
4.250 0.4542 0.00782 0.00283 0.0055 0.4242 0.9445
4.500 0.4792 0.00856 0.00313 0.0056 0.3154 0.9575
4.750 0.5052 0.00959 0.00361 0.0052 0.1930 0.9691
5.000 0.5373 0.01050 0.00410 0.0035 0.1062 0.9762
5.250 0.5696 0.01121 0.00458 0.0020 0.0629 0.9838
5.500 0.6063 0.01189 0.00515 -0.0003 0.0438 0.9878
5.750 0.6420 0.01241 0.00567 -0.0023 0.0362 0.9929
6.000 0.6782 0.01306 0.00635 -0.0045 0.0318 0.9969
6.250 0.7113 0.01360 0.00692 -0.0061 0.0291 1.0000
6.500 0.7229 0.01443 0.00776 -0.0033 0.0270 1.0000
6.750 0.7403 0.01487 0.00824 -0.0015 0.0260 1.0000
7.000 0.7582 0.01536 0.00877 0.0002 0.0249 1.0000
7.250 0.7760 0.01594 0.00940 0.0019 0.0240 1.0000
7.500 0.7943 0.01656 0.01005 0.0034 0.0231 1.0000
7.750 0.8129 0.01727 0.01077 0.0049 0.0224 1.0000
8.000 0.8306 0.01826 0.01179 0.0063 0.0217 1.0000
8.250 0.8464 0.02022 0.01382 0.0080 0.0210 1.0000
8.500 0.8679 0.02088 0.01459 0.0089 0.0206 1.0000
8.750 0.8892 0.02165 0.01546 0.0098 0.0200 1.0000
9.000 0.9098 0.02262 0.01655 0.0108 0.0194 1.0000
9.250 0.9298 0.02386 0.01791 0.0117 0.0189 1.0000
9.500 0.9491 0.02521 0.01940 0.0127 0.0185 1.0000
9.750 0.9675 0.02665 0.02099 0.0138 0.0182 1.0000
10.000 0.9850 0.02814 0.02262 0.0148 0.0178 1.0000
10.250 1.0006 0.02980 0.02444 0.0160 0.0175 1.0000
10.500 1.0145 0.03157 0.02638 0.0173 0.0173 1.0000
10.750 1.0267 0.03334 0.02829 0.0187 0.0169 1.0000
11.000 1.0335 0.03602 0.03115 0.0203 0.0166 1.0000
11.250 1.0273 0.04034 0.03583 0.0228 0.0164 1.0000
11.500 1.0125 0.04440 0.04021 0.0260 0.0163 1.0000
11.750 1.0017 0.04664 0.04267 0.0285 0.0162 1.0000
12.000 0.9856 0.05044 0.04669 0.0300 0.0162 1.0000
12.500 0.9575 0.05828 0.05490 0.0300 0.0163 1.0000
12.750 0.9433 0.06254 0.05933 0.0286 0.0164 1.0000
13.000 0.9281 0.06739 0.06434 0.0261 0.0164 1.0000
13.250 0.9102 0.07327 0.07040 0.0225 0.0165 1.0000
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