RG 14 10% AIRFOIL (rg1410-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: RG 14 10% AIRFOIL (rg1410-il) Reynolds number: 200,000 Max Cl/Cd: 63.53 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rg1410-il-200000.txt Download as CSV file: xf-rg1410-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: RG 14 10% 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5151 0.09986 0.09617 -0.0290 1.0000 0.0630
-10.000 -0.5319 0.09345 0.08984 -0.0361 1.0000 0.0635
-9.750 -0.5507 0.08641 0.08282 -0.0432 1.0000 0.0636
-9.500 -0.5257 0.08546 0.08189 -0.0333 1.0000 0.0656
-9.250 -0.5210 0.08277 0.07922 -0.0327 1.0000 0.0669
-9.000 -0.5237 0.07904 0.07553 -0.0340 1.0000 0.0681
-8.750 -0.5365 0.07405 0.07060 -0.0374 1.0000 0.0689
-8.500 -0.5584 0.06940 0.06596 -0.0391 1.0000 0.0692
-8.250 -0.5724 0.06505 0.06158 -0.0397 1.0000 0.0707
-8.000 -0.5868 0.06057 0.05698 -0.0397 1.0000 0.0727
-7.750 -0.6582 0.04050 0.03542 -0.0354 1.0000 0.0406
-7.500 -0.6563 0.03633 0.03096 -0.0328 1.0000 0.0398
-7.250 -0.6508 0.03286 0.02711 -0.0301 1.0000 0.0395
-7.000 -0.6399 0.03074 0.02455 -0.0276 1.0000 0.0406
-6.750 -0.6266 0.02917 0.02260 -0.0252 1.0000 0.0413
-6.500 -0.6138 0.02583 0.01886 -0.0231 1.0000 0.0417
-6.250 -0.5973 0.02333 0.01610 -0.0215 1.0000 0.0426
-6.000 -0.5788 0.02192 0.01459 -0.0200 1.0000 0.0441
-5.750 -0.5597 0.02107 0.01365 -0.0186 1.0000 0.0465
-5.500 -0.5398 0.02008 0.01251 -0.0172 1.0000 0.0487
-5.250 -0.5192 0.01909 0.01135 -0.0157 1.0000 0.0504
-5.000 -0.4985 0.01815 0.01026 -0.0144 1.0000 0.0521
-4.750 -0.4787 0.01703 0.00920 -0.0132 1.0000 0.0559
-4.500 -0.4584 0.01651 0.00865 -0.0119 1.0000 0.0607
-4.250 -0.4385 0.01574 0.00784 -0.0106 1.0000 0.0655
-4.000 -0.4188 0.01516 0.00731 -0.0094 1.0000 0.0734
-3.750 -0.3804 0.01442 0.00669 -0.0119 0.9954 0.0957
-3.500 -0.3424 0.01376 0.00623 -0.0144 0.9900 0.1370
-3.250 -0.3033 0.01318 0.00598 -0.0173 0.9847 0.2019
-3.000 -0.2679 0.01249 0.00577 -0.0194 0.9782 0.3043
-2.750 -0.2318 0.01168 0.00566 -0.0217 0.9723 0.4592
-2.500 -0.2022 0.01093 0.00563 -0.0220 0.9646 0.6190
-2.250 -0.1694 0.01064 0.00587 -0.0220 0.9585 0.7669
-2.000 -0.1391 0.01063 0.00593 -0.0216 0.9500 0.8318
-1.750 -0.1020 0.01066 0.00596 -0.0225 0.9443 0.8749
-1.500 -0.0693 0.01070 0.00598 -0.0224 0.9363 0.9059
-1.250 -0.0244 0.01076 0.00598 -0.0248 0.9318 0.9310
-1.000 0.0330 0.01083 0.00596 -0.0299 0.9299 0.9472
-0.750 0.0955 0.01084 0.00589 -0.0363 0.9283 0.9587
-0.500 0.1553 0.01080 0.00579 -0.0423 0.9226 0.9682
-0.250 0.2165 0.01061 0.00555 -0.0487 0.9172 0.9755
0.000 0.2714 0.01039 0.00529 -0.0540 0.9085 0.9838
0.250 0.3296 0.01006 0.00493 -0.0599 0.8985 0.9900
0.500 0.3787 0.00977 0.00461 -0.0641 0.8821 0.9967
0.750 0.4145 0.00957 0.00434 -0.0656 0.8592 1.0000
1.000 0.4379 0.00948 0.00417 -0.0645 0.8356 1.0000
1.250 0.4595 0.00944 0.00405 -0.0630 0.8111 1.0000
1.500 0.4802 0.00944 0.00397 -0.0615 0.7865 1.0000
1.750 0.5010 0.00945 0.00388 -0.0599 0.7622 1.0000
2.000 0.5210 0.00949 0.00383 -0.0582 0.7360 1.0000
2.250 0.5409 0.00955 0.00381 -0.0565 0.7102 1.0000
2.500 0.5610 0.00964 0.00381 -0.0549 0.6863 1.0000
2.750 0.5808 0.00974 0.00384 -0.0532 0.6618 1.0000
3.000 0.6000 0.00987 0.00389 -0.0514 0.6360 1.0000
3.250 0.6186 0.01003 0.00396 -0.0495 0.6091 1.0000
3.500 0.6368 0.01021 0.00404 -0.0475 0.5812 1.0000
3.750 0.6545 0.01042 0.00414 -0.0455 0.5526 1.0000
4.000 0.6723 0.01064 0.00427 -0.0435 0.5250 1.0000
4.250 0.6900 0.01088 0.00444 -0.0415 0.4962 1.0000
4.500 0.7077 0.01114 0.00461 -0.0395 0.4663 1.0000
4.750 0.7254 0.01144 0.00480 -0.0375 0.4356 1.0000
5.000 0.7434 0.01177 0.00503 -0.0356 0.4019 1.0000
5.250 0.7612 0.01216 0.00530 -0.0337 0.3645 1.0000
5.500 0.7790 0.01262 0.00560 -0.0319 0.3235 1.0000
5.750 0.7966 0.01318 0.00597 -0.0301 0.2794 1.0000
6.000 0.8144 0.01382 0.00640 -0.0284 0.2389 1.0000
6.250 0.8329 0.01447 0.00690 -0.0269 0.2066 1.0000
6.500 0.8517 0.01516 0.00745 -0.0255 0.1811 1.0000
6.750 0.8704 0.01589 0.00806 -0.0241 0.1605 1.0000
7.000 0.8886 0.01672 0.00878 -0.0226 0.1437 1.0000
7.250 0.9077 0.01749 0.00950 -0.0213 0.1291 1.0000
7.500 0.9282 0.01811 0.01016 -0.0202 0.1164 1.0000
7.750 0.9486 0.01874 0.01083 -0.0190 0.1053 1.0000
8.000 0.9679 0.01955 0.01164 -0.0177 0.0959 1.0000
8.250 0.9858 0.02051 0.01255 -0.0164 0.0875 1.0000
8.500 1.0057 0.02118 0.01333 -0.0152 0.0796 1.0000
8.750 1.0234 0.02234 0.01450 -0.0138 0.0731 1.0000
9.000 1.0418 0.02310 0.01532 -0.0125 0.0667 1.0000
9.250 1.0591 0.02442 0.01668 -0.0111 0.0613 1.0000
9.500 1.0765 0.02516 0.01752 -0.0097 0.0559 1.0000
9.750 1.0925 0.02676 0.01913 -0.0084 0.0513 1.0000
10.000 1.1084 0.02757 0.02013 -0.0068 0.0471 1.0000
10.250 1.1229 0.02872 0.02128 -0.0053 0.0435 1.0000
10.500 1.1367 0.03059 0.02333 -0.0037 0.0402 1.0000
10.750 1.1485 0.03172 0.02464 -0.0016 0.0374 1.0000
11.000 1.1580 0.03275 0.02568 0.0005 0.0350 1.0000
11.250 1.1678 0.03565 0.02873 0.0020 0.0328 1.0000
11.500 1.1725 0.03738 0.03077 0.0047 0.0316 1.0000
11.750 1.1753 0.03936 0.03300 0.0072 0.0303 1.0000
12.000 1.1770 0.04130 0.03514 0.0094 0.0292 1.0000
12.250 1.1785 0.04289 0.03683 0.0114 0.0280 1.0000
12.500 1.1808 0.04484 0.03882 0.0128 0.0270 1.0000
12.750 1.1741 0.04875 0.04291 0.0142 0.0261 1.0000
13.000 1.1616 0.05239 0.04682 0.0155 0.0259 1.0000
13.250 1.1489 0.05571 0.05040 0.0161 0.0256 1.0000
13.500 1.1332 0.05982 0.05477 0.0158 0.0254 1.0000
13.750 1.1150 0.06471 0.05991 0.0145 0.0253 1.0000
14.000 1.0956 0.07033 0.06576 0.0122 0.0253 1.0000
14.250 1.0739 0.07676 0.07241 0.0088 0.0253 1.0000
14.500 1.0512 0.08408 0.07991 0.0047 0.0255 1.0000
14.750 1.0276 0.09217 0.08816 -0.0003 0.0257 1.0000
15.000 1.0001 0.10178 0.09793 -0.0065 0.0259 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RG 14 10% AIRFOIL (rg1410-il)