NACA 642-015A AIRFOIL (n64015a-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA 642-015A AIRFOIL (n64015a-il) Reynolds number: 100,000 Max Cl/Cd: 43.17 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64015a-il-100000.txt Download as CSV file: xf-n64015a-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 642-015A 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.6264 0.12887 0.12365 -0.0075 1.0000 0.1618
-12.250 -0.7671 0.07011 0.06457 -0.0482 1.0000 0.0776
-12.000 -0.9045 0.07581 0.06979 -0.0397 1.0000 0.0771
-11.750 -0.8913 0.07015 0.06409 -0.0402 1.0000 0.0755
-11.500 -0.9016 0.06595 0.05975 -0.0389 1.0000 0.0746
-11.250 -0.9160 0.06220 0.05583 -0.0366 1.0000 0.0735
-11.000 -0.9286 0.05853 0.05192 -0.0339 1.0000 0.0724
-10.750 -0.9384 0.05470 0.04777 -0.0310 1.0000 0.0710
-10.500 -0.9454 0.05099 0.04365 -0.0280 1.0000 0.0698
-10.250 -0.9468 0.04757 0.03980 -0.0251 1.0000 0.0690
-10.000 -0.9414 0.04457 0.03643 -0.0227 1.0000 0.0688
-9.750 -0.9301 0.04204 0.03362 -0.0208 1.0000 0.0695
-9.500 -0.9164 0.03995 0.03132 -0.0192 1.0000 0.0713
-9.250 -0.9022 0.03794 0.02902 -0.0174 1.0000 0.0733
-9.000 -0.8860 0.03598 0.02675 -0.0156 1.0000 0.0750
-8.750 -0.8680 0.03424 0.02469 -0.0140 1.0000 0.0765
-8.500 -0.8383 0.03162 0.02212 -0.0146 1.0000 0.0797
-8.250 -0.8174 0.03026 0.02075 -0.0137 1.0000 0.0843
-8.000 -0.7980 0.02910 0.01940 -0.0122 1.0000 0.0890
-7.750 -0.7715 0.02726 0.01778 -0.0121 1.0000 0.0951
-7.500 -0.7561 0.02634 0.01677 -0.0100 1.0000 0.1026
-7.250 -0.7429 0.02503 0.01571 -0.0079 1.0000 0.1115
-7.000 -0.7362 0.02402 0.01482 -0.0045 1.0000 0.1213
-6.750 -0.7346 0.02317 0.01408 -0.0003 1.0000 0.1336
-6.500 -0.7367 0.02230 0.01337 0.0043 1.0000 0.1490
-6.250 -0.7403 0.02130 0.01259 0.0091 1.0000 0.1742
-6.000 -0.7484 0.01972 0.01161 0.0141 1.0000 0.2323
-5.750 -0.7628 0.01767 0.01103 0.0202 1.0000 0.4279
-5.500 -0.7567 0.01777 0.01145 0.0244 1.0000 0.5407
-5.250 -0.7436 0.01818 0.01187 0.0273 1.0000 0.5851
-5.000 -0.7193 0.01882 0.01248 0.0283 0.9971 0.6192
-4.750 -0.6788 0.01973 0.01331 0.0266 0.9890 0.6498
-4.500 -0.6379 0.02055 0.01404 0.0247 0.9813 0.6746
-4.250 -0.5991 0.02129 0.01471 0.0234 0.9729 0.6944
-4.000 -0.5603 0.02235 0.01575 0.0229 0.9655 0.7136
-3.750 -0.5190 0.02353 0.01689 0.0224 0.9586 0.7329
-3.500 -0.4799 0.02412 0.01739 0.0211 0.9514 0.7471
-3.250 -0.4432 0.02416 0.01731 0.0192 0.9434 0.7595
-3.000 -0.3914 0.02462 0.01767 0.0154 0.9394 0.7664
-2.750 -0.3660 0.02432 0.01727 0.0150 0.9289 0.7766
-2.500 -0.3119 0.02453 0.01738 0.0104 0.9250 0.7820
-2.250 -0.2917 0.02411 0.01687 0.0106 0.9142 0.7920
-2.000 -0.2423 0.02420 0.01690 0.0068 0.9097 0.7962
-1.750 -0.2135 0.02410 0.01674 0.0061 0.9010 0.8033
-1.500 -0.1798 0.02381 0.01639 0.0044 0.8939 0.8104
-1.250 -0.1428 0.02383 0.01638 0.0027 0.8872 0.8154
-1.000 -0.1238 0.02352 0.01601 0.0033 0.8782 0.8250
-0.750 -0.0766 0.02350 0.01597 0.0000 0.8744 0.8287
-0.500 -0.0579 0.02355 0.01601 0.0011 0.8639 0.8361
-0.250 -0.0247 0.02333 0.01577 -0.0001 0.8586 0.8434
0.000 0.0000 0.02354 0.01600 0.0000 0.8497 0.8497
0.250 0.0247 0.02333 0.01577 0.0001 0.8434 0.8586
0.500 0.0579 0.02355 0.01601 -0.0011 0.8361 0.8639
0.750 0.0766 0.02350 0.01597 0.0000 0.8287 0.8744
1.000 0.1237 0.02352 0.01601 -0.0033 0.8250 0.8782
1.250 0.1427 0.02383 0.01638 -0.0027 0.8154 0.8872
1.500 0.1798 0.02381 0.01639 -0.0044 0.8104 0.8939
1.750 0.2135 0.02409 0.01674 -0.0061 0.8033 0.9011
2.000 0.2424 0.02420 0.01690 -0.0068 0.7962 0.9097
2.250 0.2917 0.02410 0.01687 -0.0106 0.7920 0.9142
2.500 0.3119 0.02452 0.01738 -0.0104 0.7820 0.9250
2.750 0.3659 0.02432 0.01726 -0.0150 0.7766 0.9289
3.000 0.3911 0.02461 0.01767 -0.0154 0.7664 0.9394
3.250 0.4432 0.02416 0.01730 -0.0192 0.7595 0.9435
3.500 0.4799 0.02412 0.01739 -0.0211 0.7471 0.9515
3.750 0.5190 0.02352 0.01689 -0.0224 0.7329 0.9587
4.000 0.5603 0.02235 0.01574 -0.0229 0.7136 0.9655
4.250 0.5991 0.02128 0.01471 -0.0234 0.6944 0.9730
4.500 0.6378 0.02055 0.01404 -0.0247 0.6746 0.9813
4.750 0.6788 0.01972 0.01331 -0.0266 0.6498 0.9891
5.000 0.7193 0.01882 0.01248 -0.0283 0.6193 0.9972
5.250 0.7434 0.01818 0.01187 -0.0273 0.5852 1.0000
5.500 0.7565 0.01776 0.01145 -0.0244 0.5409 1.0000
5.750 0.7628 0.01767 0.01103 -0.0202 0.4291 1.0000
6.000 0.7484 0.01971 0.01160 -0.0141 0.2329 1.0000
6.250 0.7402 0.02129 0.01258 -0.0090 0.1744 1.0000
6.500 0.7366 0.02229 0.01336 -0.0043 0.1490 1.0000
6.750 0.7345 0.02316 0.01407 0.0003 0.1337 1.0000
7.000 0.7362 0.02401 0.01481 0.0045 0.1214 1.0000
7.250 0.7428 0.02502 0.01570 0.0079 0.1115 1.0000
7.500 0.7561 0.02634 0.01677 0.0100 0.1026 1.0000
7.750 0.7715 0.02725 0.01778 0.0121 0.0952 1.0000
8.000 0.7980 0.02910 0.01940 0.0121 0.0890 1.0000
8.250 0.8174 0.03026 0.02074 0.0137 0.0843 1.0000
8.500 0.8383 0.03162 0.02212 0.0146 0.0797 1.0000
8.750 0.8680 0.03424 0.02469 0.0140 0.0765 1.0000
9.000 0.8860 0.03598 0.02674 0.0156 0.0750 1.0000
9.250 0.9023 0.03795 0.02902 0.0174 0.0733 1.0000
9.500 0.9165 0.03996 0.03132 0.0191 0.0713 1.0000
9.750 0.9302 0.04204 0.03362 0.0208 0.0695 1.0000
10.000 0.9415 0.04458 0.03644 0.0226 0.0688 1.0000
10.250 0.9470 0.04757 0.03981 0.0250 0.0690 1.0000
10.500 0.9456 0.05100 0.04366 0.0279 0.0698 1.0000
10.750 0.9386 0.05472 0.04779 0.0310 0.0710 1.0000
11.000 0.9289 0.05856 0.05195 0.0338 0.0724 1.0000
11.250 0.9164 0.06224 0.05586 0.0365 0.0736 1.0000
11.500 0.9022 0.06600 0.05980 0.0388 0.0746 1.0000
11.750 0.8920 0.07021 0.06415 0.0400 0.0755 1.0000
12.000 0.9050 0.07583 0.06982 0.0396 0.0771 1.0000
12.250 0.6637 0.10804 0.10285 0.0190 0.1074 1.0000
12.500 0.6283 0.12891 0.12369 0.0073 0.1618 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 642-015A AIRFOIL (n64015a-il)