NACA 63-015A AIRFOIL (n63015a-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA 63-015A AIRFOIL (n63015a-il) Reynolds number: 200,000 Max Cl/Cd: 53.23 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63015a-il-200000.txt Download as CSV file: xf-n63015a-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-15.250 -0.5034 0.14403 0.14051 -0.0137 1.0000 0.0744
-15.000 -0.6472 0.16174 0.15789 0.0121 1.0000 0.0683
-14.750 -0.8803 0.09236 0.08807 -0.0300 1.0000 0.0377
-14.500 -0.9014 0.08555 0.08114 -0.0331 1.0000 0.0374
-14.250 -0.9247 0.07913 0.07456 -0.0359 1.0000 0.0372
-14.000 -0.9474 0.07328 0.06854 -0.0381 1.0000 0.0370
-13.750 -0.9692 0.06802 0.06308 -0.0395 1.0000 0.0368
-13.500 -0.9891 0.06334 0.05818 -0.0401 1.0000 0.0367
-13.250 -1.0069 0.05917 0.05378 -0.0400 1.0000 0.0367
-13.000 -1.0221 0.05547 0.04983 -0.0392 1.0000 0.0367
-12.750 -1.0345 0.05216 0.04627 -0.0378 1.0000 0.0368
-12.500 -1.0433 0.04916 0.04301 -0.0360 1.0000 0.0368
-12.250 -1.0485 0.04641 0.03998 -0.0338 1.0000 0.0370
-12.000 -1.0498 0.04383 0.03714 -0.0316 1.0000 0.0371
-11.750 -1.0474 0.04149 0.03452 -0.0293 1.0000 0.0373
-11.500 -1.0416 0.03941 0.03217 -0.0272 1.0000 0.0375
-11.250 -1.0328 0.03760 0.03008 -0.0252 1.0000 0.0379
-11.000 -1.0094 0.03469 0.02701 -0.0254 1.0000 0.0386
-10.750 -0.9853 0.03281 0.02513 -0.0256 1.0000 0.0397
-10.500 -0.9677 0.03156 0.02384 -0.0247 1.0000 0.0410
-10.250 -0.9499 0.03036 0.02255 -0.0236 1.0000 0.0425
-10.000 -0.9308 0.02914 0.02120 -0.0227 1.0000 0.0439
-9.750 -0.9122 0.02807 0.01997 -0.0215 1.0000 0.0451
-9.500 -0.8879 0.02624 0.01823 -0.0215 1.0000 0.0472
-9.250 -0.8721 0.02522 0.01724 -0.0200 1.0000 0.0493
-9.000 -0.8573 0.02432 0.01628 -0.0183 1.0000 0.0519
-8.750 -0.8445 0.02324 0.01518 -0.0162 1.0000 0.0550
-8.500 -0.8332 0.02230 0.01432 -0.0140 1.0000 0.0586
-8.250 -0.8201 0.02160 0.01354 -0.0118 1.0000 0.0630
-8.000 -0.8149 0.02054 0.01258 -0.0085 1.0000 0.0678
-7.750 -0.8053 0.01994 0.01193 -0.0055 1.0000 0.0743
-7.500 -0.8035 0.01909 0.01121 -0.0015 1.0000 0.0821
-7.250 -0.8022 0.01842 0.01061 0.0028 1.0000 0.0918
-7.000 -0.8023 0.01783 0.01009 0.0071 1.0000 0.1044
-6.500 -0.7633 0.01560 0.00847 0.0073 0.9917 0.1976
-6.250 -0.7338 0.01398 0.00760 0.0051 0.9837 0.3238
-6.000 -0.7025 0.01320 0.00728 0.0035 0.9747 0.4237
-5.750 -0.6626 0.01295 0.00713 0.0007 0.9693 0.4794
-5.500 -0.6272 0.01283 0.00701 -0.0010 0.9606 0.5134
-5.250 -0.5868 0.01275 0.00695 -0.0035 0.9555 0.5419
-5.000 -0.5551 0.01273 0.00691 -0.0042 0.9454 0.5635
-4.750 -0.5173 0.01270 0.00686 -0.0060 0.9397 0.5819
-4.500 -0.4886 0.01268 0.00678 -0.0060 0.9286 0.5965
-4.250 -0.4573 0.01265 0.00669 -0.0065 0.9202 0.6106
-4.000 -0.4286 0.01265 0.00668 -0.0064 0.9103 0.6241
-3.750 -0.4010 0.01270 0.00672 -0.0060 0.9005 0.6371
-3.500 -0.3727 0.01271 0.00670 -0.0056 0.8917 0.6491
-3.250 -0.3475 0.01271 0.00665 -0.0049 0.8809 0.6606
-3.000 -0.3199 0.01271 0.00662 -0.0045 0.8726 0.6706
-2.750 -0.2939 0.01270 0.00658 -0.0039 0.8620 0.6796
-2.500 -0.2678 0.01268 0.00649 -0.0034 0.8531 0.6892
-2.250 -0.2406 0.01263 0.00642 -0.0030 0.8438 0.6963
-2.000 -0.2146 0.01260 0.00634 -0.0025 0.8346 0.7049
-1.750 -0.1873 0.01255 0.00626 -0.0022 0.8261 0.7120
-1.500 -0.1610 0.01254 0.00621 -0.0018 0.8168 0.7201
-1.250 -0.1339 0.01249 0.00612 -0.0015 0.8088 0.7276
-1.000 -0.1074 0.01249 0.00612 -0.0012 0.7997 0.7354
-0.750 -0.0804 0.01245 0.00603 -0.0009 0.7920 0.7433
-0.500 -0.0537 0.01246 0.00606 -0.0006 0.7830 0.7509
-0.250 -0.0268 0.01243 0.00599 -0.0003 0.7755 0.7592
0.000 0.0000 0.01245 0.00604 0.0000 0.7667 0.7667
0.250 0.0268 0.01243 0.00598 0.0003 0.7592 0.7755
0.500 0.0537 0.01246 0.00606 0.0006 0.7509 0.7830
0.750 0.0804 0.01245 0.00603 0.0009 0.7433 0.7920
1.000 0.1074 0.01249 0.00612 0.0012 0.7354 0.7997
1.250 0.1339 0.01249 0.00612 0.0015 0.7276 0.8088
1.500 0.1610 0.01254 0.00622 0.0018 0.7201 0.8168
1.750 0.1873 0.01255 0.00626 0.0022 0.7120 0.8261
2.000 0.2146 0.01260 0.00634 0.0025 0.7049 0.8346
2.250 0.2406 0.01263 0.00642 0.0030 0.6964 0.8438
2.500 0.2678 0.01268 0.00649 0.0034 0.6892 0.8531
2.750 0.2939 0.01270 0.00658 0.0039 0.6796 0.8620
3.000 0.3199 0.01271 0.00662 0.0045 0.6706 0.8726
3.250 0.3475 0.01271 0.00665 0.0049 0.6606 0.8809
3.500 0.3727 0.01271 0.00670 0.0056 0.6491 0.8917
3.750 0.4010 0.01270 0.00672 0.0060 0.6371 0.9005
4.000 0.4285 0.01265 0.00668 0.0064 0.6241 0.9103
4.250 0.4573 0.01265 0.00669 0.0065 0.6106 0.9202
4.500 0.4886 0.01268 0.00678 0.0060 0.5965 0.9286
4.750 0.5173 0.01270 0.00685 0.0060 0.5819 0.9397
5.000 0.5551 0.01273 0.00691 0.0042 0.5635 0.9455
5.250 0.5868 0.01275 0.00695 0.0035 0.5418 0.9556
5.500 0.6272 0.01282 0.00701 0.0010 0.5134 0.9606
5.750 0.6625 0.01294 0.00713 -0.0007 0.4794 0.9694
6.000 0.7026 0.01320 0.00728 -0.0035 0.4237 0.9748
6.250 0.7339 0.01398 0.00759 -0.0051 0.3236 0.9837
6.500 0.7634 0.01559 0.00847 -0.0073 0.1976 0.9917
6.750 0.7955 0.01712 0.00950 -0.0100 0.1270 0.9989
7.000 0.8021 0.01782 0.01008 -0.0071 0.1044 1.0000
7.250 0.8021 0.01842 0.01060 -0.0027 0.0918 1.0000
7.500 0.8034 0.01909 0.01120 0.0015 0.0821 1.0000
7.750 0.8053 0.01994 0.01193 0.0055 0.0743 1.0000
8.000 0.8149 0.02054 0.01258 0.0085 0.0678 1.0000
8.250 0.8202 0.02159 0.01354 0.0118 0.0630 1.0000
8.500 0.8333 0.02230 0.01432 0.0140 0.0586 1.0000
8.750 0.8447 0.02324 0.01518 0.0162 0.0549 1.0000
9.000 0.8576 0.02432 0.01628 0.0182 0.0519 1.0000
9.250 0.8725 0.02522 0.01724 0.0200 0.0493 1.0000
9.500 0.8883 0.02624 0.01823 0.0214 0.0472 1.0000
9.750 0.9126 0.02807 0.01997 0.0214 0.0451 1.0000
10.000 0.9313 0.02915 0.02120 0.0226 0.0439 1.0000
10.250 0.9503 0.03037 0.02256 0.0236 0.0424 1.0000
10.500 0.9682 0.03157 0.02385 0.0246 0.0410 1.0000
10.750 0.9858 0.03282 0.02514 0.0255 0.0397 1.0000
11.000 1.0100 0.03470 0.02703 0.0253 0.0386 1.0000
11.250 1.0333 0.03761 0.03010 0.0251 0.0379 1.0000
11.500 1.0421 0.03943 0.03219 0.0271 0.0375 1.0000
11.750 1.0479 0.04151 0.03455 0.0292 0.0373 1.0000
12.000 1.0503 0.04386 0.03717 0.0315 0.0371 1.0000
12.250 1.0491 0.04644 0.04002 0.0337 0.0370 1.0000
12.500 1.0439 0.04919 0.04304 0.0359 0.0368 1.0000
12.750 1.0350 0.05220 0.04631 0.0377 0.0368 1.0000
13.000 1.0228 0.05551 0.04987 0.0391 0.0367 1.0000
13.250 1.0076 0.05922 0.05383 0.0399 0.0367 1.0000
13.500 0.9897 0.06340 0.05824 0.0399 0.0367 1.0000
13.750 0.9698 0.06810 0.06316 0.0393 0.0368 1.0000
14.000 0.9482 0.07337 0.06863 0.0379 0.0370 1.0000
14.250 0.9252 0.07924 0.07469 0.0356 0.0372 1.0000
14.500 0.9022 0.08567 0.08127 0.0328 0.0374 1.0000
14.750 0.8809 0.09252 0.08824 0.0297 0.0377 1.0000
15.000 0.6488 0.16188 0.15803 -0.0124 0.0683 1.0000
15.250 0.5024 0.14366 0.14014 0.0139 0.0744 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 63-015A AIRFOIL (n63015a-il)