RG 15A-1.8/11.0 AIRFOIL (rg15a111-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RG 15A-1.8/11.0 AIRFOIL (rg15a111-il) Reynolds number: 1,000,000 Max Cl/Cd: 90.05 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rg15a111-il-1000000-n5.txt Download as CSV file: xf-rg15a111-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RG 15A-1.8/11.0 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
------ -------- --------- --------- -------- -------- --------
-15.750 -0.8632 0.09280 0.09052 -0.0344 1.0000 0.0041
-15.500 -0.9634 0.07032 0.06782 -0.0473 1.0000 0.0040
-15.250 -1.0194 0.05497 0.05222 -0.0578 1.0000 0.0039
-15.000 -1.0476 0.04595 0.04299 -0.0648 1.0000 0.0039
-14.750 -1.0689 0.03978 0.03663 -0.0687 1.0000 0.0039
-14.500 -1.0732 0.03662 0.03336 -0.0699 1.0000 0.0039
-14.250 -1.0855 0.03311 0.02969 -0.0705 1.0000 0.0039
-14.000 -1.0829 0.03142 0.02791 -0.0699 1.0000 0.0040
-13.750 -1.0920 0.02922 0.02557 -0.0682 1.0000 0.0040
-13.500 -1.0983 0.02784 0.02407 -0.0650 1.0000 0.0040
-13.250 -1.1024 0.02671 0.02284 -0.0612 1.0000 0.0041
-13.000 -1.0967 0.02565 0.02168 -0.0588 1.0000 0.0041
-12.750 -1.0881 0.02470 0.02063 -0.0566 1.0000 0.0042
-12.500 -1.0796 0.02365 0.01948 -0.0544 1.0000 0.0042
-12.250 -1.0713 0.02249 0.01820 -0.0520 1.0000 0.0043
-12.000 -1.0610 0.02140 0.01699 -0.0499 1.0000 0.0044
-11.750 -1.0482 0.02047 0.01597 -0.0480 1.0000 0.0046
-11.500 -1.0244 0.01957 0.01498 -0.0482 0.9992 0.0048
-11.250 -0.9966 0.01880 0.01413 -0.0490 0.9979 0.0050
-11.000 -0.9681 0.01810 0.01335 -0.0499 0.9965 0.0052
-10.750 -0.9391 0.01741 0.01259 -0.0509 0.9952 0.0054
-10.500 -0.9092 0.01678 0.01188 -0.0520 0.9941 0.0057
-10.250 -0.8822 0.01615 0.01118 -0.0524 0.9923 0.0059
-10.000 -0.8542 0.01559 0.01056 -0.0529 0.9902 0.0061
-9.750 -0.8252 0.01505 0.00995 -0.0536 0.9883 0.0063
-9.500 -0.7956 0.01443 0.00925 -0.0545 0.9864 0.0067
-9.250 -0.7650 0.01392 0.00871 -0.0555 0.9849 0.0072
-9.000 -0.7334 0.01350 0.00827 -0.0566 0.9837 0.0077
-8.750 -0.7042 0.01315 0.00789 -0.0572 0.9816 0.0083
-8.500 -0.6778 0.01277 0.00747 -0.0571 0.9777 0.0088
-8.250 -0.6480 0.01243 0.00708 -0.0577 0.9749 0.0092
-8.000 -0.6173 0.01194 0.00655 -0.0586 0.9725 0.0099
-7.500 -0.5598 0.01122 0.00579 -0.0594 0.9650 0.0112
-7.250 -0.5294 0.01091 0.00546 -0.0601 0.9608 0.0119
-7.000 -0.4965 0.01060 0.00510 -0.0612 0.9575 0.0126
-6.750 -0.4669 0.01034 0.00482 -0.0617 0.9513 0.0130
-6.500 -0.4349 0.00991 0.00435 -0.0628 0.9453 0.0143
-6.250 -0.4033 0.00960 0.00401 -0.0637 0.9376 0.0152
-5.750 -0.3417 0.00911 0.00341 -0.0650 0.9161 0.0171
-5.500 -0.3134 0.00890 0.00314 -0.0651 0.9028 0.0182
-5.250 -0.2862 0.00869 0.00289 -0.0649 0.8889 0.0203
-5.000 -0.2593 0.00852 0.00266 -0.0647 0.8751 0.0227
-4.750 -0.2330 0.00830 0.00244 -0.0644 0.8608 0.0314
-4.500 -0.2069 0.00810 0.00225 -0.0640 0.8464 0.0447
-4.250 -0.1806 0.00796 0.00209 -0.0637 0.8319 0.0545
-4.000 -0.1542 0.00784 0.00194 -0.0634 0.8171 0.0650
-3.750 -0.1280 0.00768 0.00180 -0.0631 0.8021 0.0810
-3.500 -0.1018 0.00754 0.00167 -0.0627 0.7872 0.1008
-3.250 -0.0753 0.00742 0.00155 -0.0624 0.7724 0.1179
-3.000 -0.0489 0.00729 0.00144 -0.0622 0.7579 0.1398
-2.750 -0.0225 0.00715 0.00134 -0.0619 0.7433 0.1668
-2.500 0.0041 0.00706 0.00126 -0.0616 0.7285 0.1896
-2.250 0.0305 0.00692 0.00118 -0.0614 0.7133 0.2228
-2.000 0.0566 0.00674 0.00110 -0.0611 0.6973 0.2713
-1.750 0.0829 0.00659 0.00104 -0.0609 0.6816 0.3163
-1.500 0.1091 0.00642 0.00098 -0.0606 0.6655 0.3706
-1.000 0.1617 0.00614 0.00091 -0.0601 0.6334 0.4746
-0.750 0.1882 0.00602 0.00089 -0.0599 0.6166 0.5266
-0.500 0.2147 0.00594 0.00090 -0.0596 0.5999 0.5763
-0.250 0.2414 0.00594 0.00092 -0.0594 0.5817 0.6114
0.000 0.2683 0.00598 0.00095 -0.0591 0.5637 0.6352
0.250 0.2953 0.00605 0.00098 -0.0589 0.5449 0.6527
0.500 0.3223 0.00612 0.00102 -0.0587 0.5252 0.6671
0.750 0.3492 0.00622 0.00107 -0.0585 0.5052 0.6795
1.250 0.4030 0.00643 0.00118 -0.0580 0.4651 0.7018
1.500 0.4299 0.00655 0.00124 -0.0578 0.4455 0.7117
1.750 0.4567 0.00667 0.00132 -0.0576 0.4262 0.7221
2.000 0.4834 0.00679 0.00140 -0.0573 0.4062 0.7324
2.250 0.5101 0.00692 0.00148 -0.0571 0.3873 0.7419
2.500 0.5367 0.00706 0.00158 -0.0568 0.3685 0.7520
2.750 0.5632 0.00720 0.00168 -0.0566 0.3501 0.7621
3.000 0.5893 0.00737 0.00180 -0.0563 0.3284 0.7722
3.250 0.6152 0.00756 0.00193 -0.0559 0.3042 0.7829
3.500 0.6410 0.00776 0.00206 -0.0555 0.2826 0.7931
3.750 0.6670 0.00794 0.00220 -0.0552 0.2631 0.8015
4.250 0.7190 0.00830 0.00250 -0.0545 0.2309 0.8191
4.500 0.7447 0.00849 0.00266 -0.0541 0.2153 0.8287
5.000 0.7948 0.00894 0.00304 -0.0532 0.1795 0.8510
5.250 0.8186 0.00924 0.00328 -0.0525 0.1559 0.8638
5.500 0.8424 0.00951 0.00350 -0.0517 0.1372 0.8791
5.750 0.8661 0.00972 0.00372 -0.0509 0.1242 0.8971
6.000 0.8885 0.00993 0.00394 -0.0498 0.1121 0.9200
6.250 0.9119 0.01017 0.00419 -0.0490 0.0984 0.9544
6.500 0.9446 0.01049 0.00447 -0.0503 0.0859 1.0000
6.750 0.9691 0.01080 0.00474 -0.0498 0.0763 1.0000
7.000 0.9935 0.01111 0.00502 -0.0493 0.0676 1.0000
7.250 1.0178 0.01143 0.00532 -0.0488 0.0603 1.0000
7.500 1.0414 0.01180 0.00564 -0.0481 0.0522 1.0000
7.750 1.0653 0.01214 0.00596 -0.0476 0.0451 1.0000
8.000 1.0883 0.01254 0.00633 -0.0469 0.0378 1.0000
8.250 1.1109 0.01297 0.00671 -0.0461 0.0311 1.0000
8.500 1.1338 0.01336 0.00710 -0.0454 0.0274 1.0000
8.750 1.1563 0.01377 0.00750 -0.0446 0.0239 1.0000
9.000 1.1791 0.01413 0.00788 -0.0439 0.0219 1.0000
9.250 1.2009 0.01456 0.00831 -0.0430 0.0196 1.0000
9.500 1.2226 0.01498 0.00875 -0.0422 0.0177 1.0000
9.750 1.2445 0.01537 0.00917 -0.0413 0.0165 1.0000
10.000 1.2653 0.01581 0.00963 -0.0403 0.0148 1.0000
10.250 1.2852 0.01630 0.01013 -0.0392 0.0131 1.0000
10.500 1.3052 0.01676 0.01062 -0.0381 0.0116 1.0000
10.750 1.3233 0.01732 0.01119 -0.0367 0.0099 1.0000
11.000 1.3412 0.01785 0.01175 -0.0353 0.0084 1.0000
11.250 1.3558 0.01845 0.01237 -0.0333 0.0068 1.0000
11.500 1.3694 0.01905 0.01301 -0.0312 0.0058 1.0000
11.750 1.3814 0.01973 0.01373 -0.0289 0.0049 1.0000
12.000 1.3943 0.02037 0.01443 -0.0268 0.0045 1.0000
12.250 1.4061 0.02109 0.01520 -0.0246 0.0041 1.0000
12.500 1.4167 0.02188 0.01604 -0.0225 0.0037 1.0000
12.750 1.4265 0.02274 0.01697 -0.0203 0.0035 1.0000
13.000 1.4351 0.02370 0.01800 -0.0181 0.0033 1.0000
13.250 1.4428 0.02474 0.01912 -0.0160 0.0031 1.0000
13.500 1.4513 0.02577 0.02024 -0.0141 0.0030 1.0000
13.750 1.4584 0.02694 0.02149 -0.0123 0.0030 1.0000
14.000 1.4642 0.02827 0.02290 -0.0106 0.0029 1.0000
14.250 1.4700 0.02965 0.02438 -0.0091 0.0028 1.0000
14.500 1.4735 0.03131 0.02612 -0.0076 0.0027 1.0000
14.750 1.4764 0.03311 0.02802 -0.0064 0.0026 1.0000
15.000 1.4785 0.03508 0.03009 -0.0054 0.0026 1.0000
15.250 1.4791 0.03732 0.03243 -0.0047 0.0025 1.0000
15.500 1.4787 0.03980 0.03501 -0.0042 0.0025 1.0000
15.750 1.4772 0.04253 0.03785 -0.0041 0.0024 1.0000
16.000 1.4734 0.04568 0.04111 -0.0042 0.0023 1.0000
16.250 1.4693 0.04905 0.04459 -0.0047 0.0024 1.0000
16.500 1.4623 0.05299 0.04865 -0.0057 0.0023 1.0000
16.750 1.4550 0.05720 0.05298 -0.0070 0.0023 1.0000
17.000 1.4443 0.06212 0.05804 -0.0089 0.0022 1.0000
17.250 1.4309 0.06774 0.06380 -0.0113 0.0022 1.0000
17.500 1.4157 0.07392 0.07012 -0.0142 0.0022 1.0000
17.750 1.3995 0.08056 0.07690 -0.0175 0.0022 1.0000
18.000 1.3827 0.08755 0.08402 -0.0211 0.0022 1.0000
18.250 1.3605 0.09581 0.09243 -0.0255 0.0022 1.0000
18.500 1.3374 0.10450 0.10128 -0.0303 0.0022 1.0000
18.750 1.3135 0.11364 0.11057 -0.0355 0.0022 1.0000
19.000 1.2890 0.12313 0.12019 -0.0410 0.0022 1.0000
19.250 1.2629 0.13320 0.13041 -0.0469 0.0022 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RG 15A-1.8/11.0 AIRFOIL (rg15a111-il)