RG 14 AIRFOIL (rg14-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RG 14 AIRFOIL (rg14-il) Reynolds number: 1,000,000 Max Cl/Cd: 84.23 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rg14-il-1000000-n5.txt Download as CSV file: xf-rg14-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RG 14 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5505 0.08807 0.08645 -0.0199 1.0000 0.0035
-9.750 -0.5783 0.07787 0.07630 -0.0253 1.0000 0.0032
-9.250 -0.7674 0.03258 0.03023 -0.0406 1.0000 0.0032
-9.000 -0.7628 0.02979 0.02721 -0.0386 1.0000 0.0032
-8.750 -0.7576 0.02691 0.02403 -0.0363 1.0000 0.0033
-8.500 -0.7391 0.02454 0.02139 -0.0362 0.9991 0.0034
-8.250 -0.7132 0.02250 0.01909 -0.0373 0.9974 0.0035
-8.000 -0.6866 0.02051 0.01683 -0.0383 0.9956 0.0037
-7.750 -0.6586 0.01865 0.01469 -0.0393 0.9941 0.0038
-7.500 -0.6326 0.01731 0.01315 -0.0396 0.9918 0.0039
-7.250 -0.6051 0.01598 0.01161 -0.0400 0.9893 0.0040
-7.000 -0.5762 0.01483 0.01025 -0.0407 0.9870 0.0041
-6.750 -0.5461 0.01379 0.00905 -0.0415 0.9850 0.0042
-6.500 -0.5143 0.01310 0.00826 -0.0427 0.9833 0.0045
-6.250 -0.4856 0.01239 0.00743 -0.0430 0.9805 0.0046
-6.000 -0.4586 0.01160 0.00653 -0.0430 0.9760 0.0047
-5.750 -0.4285 0.01088 0.00570 -0.0436 0.9726 0.0048
-5.250 -0.3699 0.00982 0.00446 -0.0444 0.9631 0.0050
-5.000 -0.3386 0.00940 0.00398 -0.0453 0.9579 0.0051
-4.750 -0.3085 0.00884 0.00332 -0.0458 0.9508 0.0059
-4.500 -0.2763 0.00849 0.00291 -0.0467 0.9427 0.0066
-4.250 -0.2457 0.00822 0.00259 -0.0473 0.9309 0.0074
-4.000 -0.2163 0.00800 0.00230 -0.0476 0.9167 0.0082
-3.750 -0.1884 0.00784 0.00206 -0.0475 0.9008 0.0093
-3.500 -0.1617 0.00762 0.00185 -0.0472 0.8840 0.0177
-3.250 -0.1355 0.00743 0.00166 -0.0468 0.8662 0.0324
-3.000 -0.1093 0.00730 0.00151 -0.0464 0.8483 0.0465
-2.750 -0.0832 0.00718 0.00138 -0.0460 0.8299 0.0617
-2.500 -0.0572 0.00704 0.00124 -0.0456 0.8097 0.0855
-2.250 -0.0318 0.00684 0.00112 -0.0452 0.7895 0.1288
-2.000 -0.0061 0.00667 0.00103 -0.0448 0.7684 0.1734
-1.750 0.0196 0.00652 0.00094 -0.0444 0.7480 0.2199
-1.500 0.0456 0.00641 0.00087 -0.0441 0.7272 0.2605
-1.250 0.0710 0.00621 0.00081 -0.0437 0.7061 0.3303
-1.000 0.0968 0.00605 0.00076 -0.0434 0.6847 0.3917
-0.750 0.1224 0.00590 0.00073 -0.0430 0.6632 0.4579
-0.500 0.1480 0.00574 0.00070 -0.0426 0.6412 0.5277
-0.250 0.1729 0.00551 0.00068 -0.0422 0.6195 0.6211
0.000 0.1974 0.00531 0.00070 -0.0415 0.5973 0.7131
0.250 0.2227 0.00528 0.00075 -0.0409 0.5746 0.7703
0.500 0.2488 0.00533 0.00079 -0.0405 0.5523 0.7998
0.750 0.2750 0.00541 0.00084 -0.0400 0.5290 0.8210
1.000 0.3013 0.00551 0.00088 -0.0397 0.5050 0.8362
1.250 0.3277 0.00562 0.00093 -0.0393 0.4810 0.8485
1.500 0.3539 0.00574 0.00100 -0.0390 0.4568 0.8609
1.750 0.3799 0.00587 0.00106 -0.0386 0.4311 0.8734
2.000 0.4058 0.00600 0.00114 -0.0381 0.4067 0.8852
2.250 0.4314 0.00615 0.00122 -0.0376 0.3812 0.8964
2.500 0.4566 0.00629 0.00132 -0.0370 0.3569 0.9089
2.750 0.4813 0.00644 0.00142 -0.0363 0.3328 0.9232
3.000 0.5055 0.00661 0.00153 -0.0355 0.3056 0.9380
3.250 0.5307 0.00680 0.00164 -0.0349 0.2805 0.9507
3.500 0.5576 0.00700 0.00177 -0.0348 0.2555 0.9603
4.000 0.6156 0.00751 0.00209 -0.0356 0.2001 0.9764
4.250 0.6459 0.00775 0.00225 -0.0363 0.1785 0.9836
4.500 0.6772 0.00804 0.00244 -0.0373 0.1537 0.9899
4.750 0.7073 0.00843 0.00270 -0.0381 0.1217 0.9973
5.000 0.7336 0.00875 0.00292 -0.0380 0.1000 1.0000
5.250 0.7578 0.00904 0.00315 -0.0374 0.0843 1.0000
5.500 0.7821 0.00935 0.00339 -0.0368 0.0698 1.0000
5.750 0.8062 0.00971 0.00367 -0.0363 0.0531 1.0000
6.000 0.8299 0.01013 0.00399 -0.0356 0.0355 1.0000
6.250 0.8540 0.01052 0.00434 -0.0351 0.0244 1.0000
6.500 0.8786 0.01085 0.00465 -0.0346 0.0179 1.0000
6.750 0.9030 0.01122 0.00500 -0.0341 0.0132 1.0000
7.000 0.9271 0.01161 0.00538 -0.0335 0.0093 1.0000
7.250 0.9513 0.01199 0.00577 -0.0330 0.0066 1.0000
7.500 0.9754 0.01239 0.00618 -0.0325 0.0050 1.0000
7.750 0.9991 0.01282 0.00664 -0.0318 0.0040 1.0000
8.000 1.0226 0.01328 0.00717 -0.0312 0.0034 1.0000
8.250 1.0450 0.01387 0.00784 -0.0303 0.0029 1.0000
8.500 1.0664 0.01457 0.00866 -0.0294 0.0026 1.0000
8.750 1.0893 0.01506 0.00920 -0.0287 0.0025 1.0000
9.000 1.1114 0.01562 0.00983 -0.0279 0.0024 1.0000
9.250 1.1332 0.01619 0.01048 -0.0271 0.0023 1.0000
9.500 1.1545 0.01680 0.01116 -0.0262 0.0023 1.0000
9.750 1.1740 0.01759 0.01206 -0.0250 0.0022 1.0000
10.000 1.1937 0.01831 0.01289 -0.0239 0.0021 1.0000
10.250 1.2119 0.01915 0.01385 -0.0227 0.0021 1.0000
10.500 1.2299 0.01997 0.01477 -0.0214 0.0020 1.0000
10.750 1.2463 0.02091 0.01582 -0.0199 0.0020 1.0000
11.000 1.2617 0.02187 0.01690 -0.0183 0.0019 1.0000
11.250 1.2756 0.02290 0.01804 -0.0166 0.0019 1.0000
11.500 1.2866 0.02408 0.01937 -0.0145 0.0018 1.0000
11.750 1.2976 0.02504 0.02043 -0.0123 0.0018 1.0000
12.000 1.3045 0.02606 0.02156 -0.0096 0.0017 1.0000
12.250 1.3099 0.02718 0.02279 -0.0068 0.0017 1.0000
12.500 1.3127 0.02851 0.02425 -0.0040 0.0016 1.0000
12.750 1.3145 0.02997 0.02586 -0.0015 0.0016 1.0000
13.000 1.3138 0.03172 0.02774 0.0008 0.0016 1.0000
13.250 1.3169 0.03327 0.02938 0.0024 0.0015 1.0000
13.500 1.3185 0.03507 0.03129 0.0037 0.0015 1.0000
13.750 1.3048 0.03855 0.03497 0.0049 0.0015 1.0000
14.000 1.3016 0.04134 0.03788 0.0051 0.0015 1.0000
14.250 1.2957 0.04471 0.04138 0.0046 0.0014 1.0000
14.500 1.2765 0.05028 0.04717 0.0028 0.0015 1.0000
14.750 1.2618 0.05588 0.05293 0.0002 0.0015 1.0000
15.000 1.2542 0.06086 0.05802 -0.0025 0.0014 1.0000
15.250 1.2374 0.06780 0.06510 -0.0065 0.0014 1.0000
15.500 1.2128 0.07667 0.07415 -0.0117 0.0015 1.0000
15.750 1.1727 0.08924 0.08695 -0.0191 0.0015 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RG 14 AIRFOIL (rg14-il)