NACA 63-210 AIRFOIL (n63210-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: NACA 63-210 AIRFOIL (n63210-il) Reynolds number: 100,000 Max Cl/Cd: 42.8 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n63210-il-100000-n5.txt Download as CSV file: xf-n63210-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63-210 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.6103 0.07608 0.07105 -0.0332 1.0000 0.0264
-9.750 -0.6237 0.07035 0.06528 -0.0375 1.0000 0.0262
-9.500 -0.6380 0.06559 0.06046 -0.0403 1.0000 0.0261
-9.250 -0.6543 0.06153 0.05630 -0.0414 1.0000 0.0260
-9.000 -0.6660 0.05721 0.05181 -0.0420 1.0000 0.0260
-8.750 -0.6732 0.05282 0.04711 -0.0420 1.0000 0.0261
-8.500 -0.6763 0.04848 0.04229 -0.0416 1.0000 0.0264
-8.250 -0.6742 0.04416 0.03764 -0.0410 1.0000 0.0269
-8.000 -0.6644 0.04127 0.03457 -0.0405 1.0000 0.0280
-7.750 -0.6494 0.03966 0.03286 -0.0400 1.0000 0.0301
-7.500 -0.6362 0.03709 0.02990 -0.0391 1.0000 0.0325
-7.250 -0.6216 0.03428 0.02657 -0.0379 1.0000 0.0343
-7.000 -0.6056 0.03193 0.02376 -0.0368 1.0000 0.0365
-6.750 -0.5877 0.02999 0.02167 -0.0359 1.0000 0.0380
-6.500 -0.5687 0.02862 0.02016 -0.0350 1.0000 0.0401
-6.250 -0.5490 0.02704 0.01829 -0.0339 1.0000 0.0412
-6.000 -0.5288 0.02545 0.01649 -0.0326 1.0000 0.0413
-5.750 -0.5088 0.02409 0.01497 -0.0314 1.0000 0.0415
-5.500 -0.4893 0.02292 0.01367 -0.0300 1.0000 0.0418
-5.250 -0.4705 0.02188 0.01256 -0.0286 1.0000 0.0422
-5.000 -0.4524 0.02098 0.01159 -0.0271 1.0000 0.0427
-4.750 -0.4247 0.02005 0.01056 -0.0276 0.9958 0.0434
-4.500 -0.3898 0.01916 0.00959 -0.0294 0.9883 0.0444
-4.250 -0.3546 0.01840 0.00871 -0.0314 0.9809 0.0458
-4.000 -0.3180 0.01765 0.00786 -0.0336 0.9744 0.0479
-3.750 -0.2835 0.01700 0.00711 -0.0354 0.9662 0.0522
-3.500 -0.2480 0.01644 0.00648 -0.0373 0.9586 0.0609
-3.250 -0.2143 0.01543 0.00583 -0.0392 0.9508 0.1259
-3.000 -0.1898 0.01353 0.00549 -0.0401 0.9407 0.4460
-2.750 -0.1639 0.01321 0.00582 -0.0392 0.9311 0.6221
-2.500 -0.1355 0.01324 0.00598 -0.0385 0.9218 0.6741
-2.250 -0.1077 0.01326 0.00597 -0.0381 0.9111 0.7034
-2.000 -0.0812 0.01335 0.00605 -0.0370 0.9014 0.7353
-1.750 -0.0514 0.01334 0.00598 -0.0368 0.8932 0.7537
-1.500 -0.0226 0.01329 0.00585 -0.0368 0.8833 0.7653
-1.250 0.0072 0.01324 0.00571 -0.0370 0.8745 0.7770
-1.000 0.0365 0.01320 0.00560 -0.0369 0.8663 0.7876
-0.750 0.0636 0.01318 0.00554 -0.0365 0.8565 0.7984
-0.500 0.0912 0.01317 0.00549 -0.0362 0.8474 0.8097
-0.250 0.1187 0.01317 0.00544 -0.0359 0.8384 0.8217
0.000 0.1443 0.01318 0.00545 -0.0353 0.8280 0.8339
0.250 0.1701 0.01319 0.00546 -0.0346 0.8184 0.8458
0.500 0.1963 0.01319 0.00546 -0.0339 0.8095 0.8582
0.750 0.2214 0.01320 0.00551 -0.0332 0.7990 0.8713
1.000 0.2474 0.01321 0.00554 -0.0325 0.7896 0.8848
1.250 0.2744 0.01319 0.00554 -0.0321 0.7809 0.8988
1.500 0.3017 0.01320 0.00562 -0.0318 0.7706 0.9135
1.750 0.3312 0.01320 0.00566 -0.0320 0.7613 0.9286
2.000 0.3632 0.01319 0.00570 -0.0327 0.7521 0.9434
2.250 0.3976 0.01321 0.00582 -0.0340 0.7414 0.9584
2.500 0.4342 0.01322 0.00591 -0.0359 0.7302 0.9738
2.750 0.4708 0.01306 0.00576 -0.0372 0.7079 0.9928
3.000 0.4944 0.01302 0.00572 -0.0362 0.6809 1.0000
3.250 0.5171 0.01307 0.00576 -0.0351 0.6568 1.0000
3.500 0.5404 0.01312 0.00578 -0.0340 0.6260 1.0000
3.750 0.5621 0.01321 0.00569 -0.0325 0.5700 1.0000
4.000 0.5817 0.01359 0.00564 -0.0307 0.4697 1.0000
4.250 0.5950 0.01496 0.00591 -0.0287 0.2747 1.0000
4.500 0.6105 0.01665 0.00674 -0.0276 0.1202 1.0000
4.750 0.6304 0.01785 0.00757 -0.0268 0.0647 1.0000
5.000 0.6532 0.01867 0.00839 -0.0262 0.0538 1.0000
5.250 0.6754 0.01955 0.00932 -0.0256 0.0492 1.0000
5.500 0.6981 0.02033 0.01023 -0.0249 0.0467 1.0000
5.750 0.7203 0.02115 0.01116 -0.0242 0.0442 1.0000
6.000 0.7421 0.02202 0.01208 -0.0236 0.0414 1.0000
6.250 0.7628 0.02308 0.01316 -0.0229 0.0393 1.0000
6.500 0.7833 0.02435 0.01448 -0.0220 0.0382 1.0000
6.750 0.8060 0.02551 0.01574 -0.0214 0.0375 1.0000
7.000 0.8295 0.02682 0.01715 -0.0208 0.0369 1.0000
7.250 0.8537 0.02827 0.01873 -0.0204 0.0363 1.0000
7.500 0.8783 0.02990 0.02052 -0.0200 0.0358 1.0000
7.750 0.9026 0.03172 0.02255 -0.0195 0.0354 1.0000
8.000 0.9260 0.03378 0.02491 -0.0190 0.0351 1.0000
8.250 0.9475 0.03608 0.02753 -0.0183 0.0348 1.0000
8.500 0.9666 0.03867 0.03051 -0.0175 0.0347 1.0000
8.750 0.9826 0.04139 0.03364 -0.0164 0.0342 1.0000
9.000 0.9966 0.04371 0.03628 -0.0154 0.0326 1.0000
9.250 1.0105 0.04576 0.03851 -0.0146 0.0310 1.0000
9.500 1.0216 0.04820 0.04107 -0.0139 0.0296 1.0000
9.750 1.0225 0.05176 0.04502 -0.0123 0.0285 1.0000
10.000 1.0183 0.05532 0.04910 -0.0103 0.0279 1.0000
10.250 1.0085 0.05917 0.05334 -0.0084 0.0276 1.0000
10.500 0.9926 0.06278 0.05723 -0.0063 0.0275 1.0000
10.750 0.9745 0.06675 0.06144 -0.0055 0.0275 1.0000
11.000 0.9543 0.07138 0.06623 -0.0061 0.0274 1.0000
11.250 0.9338 0.07669 0.07172 -0.0082 0.0275 1.0000
11.500 0.9132 0.08284 0.07802 -0.0116 0.0277 1.0000
11.750 0.8917 0.09016 0.08546 -0.0164 0.0279 1.0000
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Polar data table (+)
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