RG 8 AIRFOIL (rg8-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RG 8 AIRFOIL (rg8-il) Reynolds number: 1,000,000 Max Cl/Cd: 92.8 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rg8-il-1000000-n5.txt Download as CSV file: xf-rg8-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RG 8 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
------ -------- --------- --------- -------- -------- --------
-14.000 -0.8758 0.06195 0.05974 -0.0555 1.0000 0.0057
-13.750 -0.9014 0.05303 0.05066 -0.0617 1.0000 0.0057
-13.500 -0.9262 0.04573 0.04319 -0.0666 1.0000 0.0057
-13.250 -0.9475 0.04017 0.03747 -0.0695 1.0000 0.0057
-13.000 -0.9601 0.03656 0.03373 -0.0705 1.0000 0.0058
-12.750 -0.9703 0.03401 0.03108 -0.0701 1.0000 0.0058
-12.500 -0.9890 0.03217 0.02916 -0.0670 1.0000 0.0059
-12.250 -0.9920 0.03014 0.02698 -0.0656 0.9997 0.0059
-12.000 -0.9708 0.02836 0.02505 -0.0676 0.9981 0.0060
-11.750 -0.9495 0.02670 0.02323 -0.0691 0.9965 0.0062
-11.500 -0.9259 0.02528 0.02166 -0.0706 0.9952 0.0064
-11.250 -0.9005 0.02400 0.02022 -0.0720 0.9942 0.0066
-11.000 -0.8787 0.02273 0.01882 -0.0725 0.9925 0.0067
-10.750 -0.8549 0.02166 0.01761 -0.0731 0.9907 0.0068
-10.500 -0.8292 0.02071 0.01654 -0.0738 0.9890 0.0070
-10.250 -0.8013 0.02005 0.01580 -0.0747 0.9876 0.0071
-10.000 -0.7725 0.01943 0.01509 -0.0757 0.9864 0.0072
-9.750 -0.7444 0.01836 0.01390 -0.0769 0.9852 0.0073
-9.500 -0.7171 0.01678 0.01214 -0.0782 0.9841 0.0077
-9.250 -0.6868 0.01591 0.01117 -0.0796 0.9833 0.0079
-9.000 -0.6658 0.01529 0.01049 -0.0787 0.9796 0.0081
-8.750 -0.6385 0.01475 0.00990 -0.0791 0.9773 0.0084
-8.500 -0.6095 0.01417 0.00926 -0.0798 0.9754 0.0086
-8.250 -0.5792 0.01364 0.00868 -0.0807 0.9739 0.0089
-8.000 -0.5482 0.01317 0.00815 -0.0818 0.9726 0.0093
-7.750 -0.5165 0.01268 0.00761 -0.0829 0.9716 0.0097
-7.500 -0.4914 0.01222 0.00709 -0.0826 0.9680 0.0099
-7.250 -0.4651 0.01176 0.00657 -0.0825 0.9641 0.0102
-7.000 -0.4350 0.01132 0.00608 -0.0832 0.9616 0.0104
-6.750 -0.4037 0.01092 0.00563 -0.0841 0.9597 0.0106
-6.500 -0.3717 0.01053 0.00519 -0.0852 0.9581 0.0107
-6.250 -0.3490 0.01021 0.00484 -0.0842 0.9515 0.0108
-6.000 -0.3182 0.00971 0.00427 -0.0850 0.9480 0.0113
-5.750 -0.2839 0.00928 0.00380 -0.0866 0.9456 0.0120
-5.500 -0.2523 0.00896 0.00345 -0.0875 0.9409 0.0127
-5.250 -0.2185 0.00868 0.00314 -0.0889 0.9360 0.0137
-5.000 -0.1816 0.00841 0.00284 -0.0909 0.9315 0.0144
-4.750 -0.1484 0.00819 0.00259 -0.0921 0.9235 0.0152
-4.500 -0.1136 0.00798 0.00233 -0.0936 0.9151 0.0171
-4.250 -0.0817 0.00774 0.00209 -0.0946 0.9036 0.0239
-4.000 -0.0519 0.00747 0.00188 -0.0951 0.8901 0.0451
-3.500 0.0044 0.00712 0.00155 -0.0954 0.8592 0.0866
-3.250 0.0314 0.00695 0.00143 -0.0952 0.8425 0.1134
-3.000 0.0580 0.00680 0.00132 -0.0950 0.8253 0.1439
-2.750 0.0844 0.00670 0.00122 -0.0947 0.8077 0.1685
-2.500 0.1107 0.00661 0.00114 -0.0944 0.7896 0.1967
-2.250 0.1368 0.00649 0.00106 -0.0941 0.7715 0.2335
-2.000 0.1626 0.00623 0.00098 -0.0939 0.7526 0.3085
-1.750 0.1881 0.00594 0.00090 -0.0936 0.7332 0.3987
-1.500 0.2142 0.00584 0.00088 -0.0933 0.7142 0.4508
-1.250 0.2406 0.00582 0.00088 -0.0930 0.6946 0.4870
-1.000 0.2670 0.00586 0.00089 -0.0927 0.6748 0.5098
-0.750 0.2936 0.00591 0.00090 -0.0924 0.6552 0.5292
-0.500 0.3202 0.00598 0.00092 -0.0921 0.6344 0.5450
0.000 0.3735 0.00615 0.00098 -0.0915 0.5935 0.5694
0.250 0.4001 0.00625 0.00102 -0.0912 0.5729 0.5798
0.500 0.4268 0.00635 0.00107 -0.0909 0.5525 0.5896
0.750 0.4534 0.00645 0.00112 -0.0907 0.5314 0.5999
1.000 0.4800 0.00658 0.00118 -0.0904 0.5105 0.6101
1.250 0.5066 0.00669 0.00125 -0.0901 0.4892 0.6196
1.500 0.5330 0.00683 0.00133 -0.0898 0.4674 0.6293
1.750 0.5595 0.00697 0.00141 -0.0895 0.4454 0.6396
2.000 0.5858 0.00711 0.00150 -0.0892 0.4246 0.6488
2.250 0.6123 0.00726 0.00160 -0.0890 0.4035 0.6559
2.500 0.6385 0.00743 0.00170 -0.0886 0.3822 0.6633
2.750 0.6649 0.00759 0.00181 -0.0884 0.3621 0.6704
3.000 0.6909 0.00778 0.00194 -0.0880 0.3404 0.6779
3.250 0.7168 0.00798 0.00208 -0.0877 0.3176 0.6855
3.500 0.7425 0.00819 0.00223 -0.0873 0.2955 0.6942
3.750 0.7684 0.00839 0.00238 -0.0870 0.2756 0.7027
4.250 0.8195 0.00884 0.00273 -0.0861 0.2343 0.7217
4.500 0.8445 0.00910 0.00293 -0.0857 0.2118 0.7326
5.000 0.8931 0.00975 0.00340 -0.0845 0.1588 0.7563
5.250 0.9173 0.01007 0.00367 -0.0839 0.1363 0.7696
5.500 0.9412 0.01040 0.00395 -0.0832 0.1156 0.7847
5.750 0.9652 0.01070 0.00422 -0.0826 0.0988 0.8018
6.000 0.9880 0.01108 0.00455 -0.0817 0.0778 0.8213
6.250 1.0108 0.01142 0.00488 -0.0808 0.0617 0.8433
6.500 1.0329 0.01174 0.00521 -0.0798 0.0490 0.8707
6.750 1.0520 0.01198 0.00553 -0.0780 0.0390 0.9149
7.000 1.0750 0.01227 0.00584 -0.0771 0.0301 1.0000
7.250 1.0982 0.01269 0.00622 -0.0764 0.0227 1.0000
7.500 1.1210 0.01315 0.00664 -0.0756 0.0153 1.0000
7.750 1.1426 0.01370 0.00714 -0.0746 0.0083 1.0000
8.000 1.1653 0.01413 0.00758 -0.0738 0.0067 1.0000
8.250 1.1877 0.01458 0.00805 -0.0729 0.0056 1.0000
8.500 1.2103 0.01498 0.00849 -0.0721 0.0052 1.0000
8.750 1.2325 0.01540 0.00895 -0.0712 0.0050 1.0000
9.000 1.2541 0.01585 0.00944 -0.0703 0.0047 1.0000
9.250 1.2751 0.01634 0.00998 -0.0692 0.0044 1.0000
9.500 1.2954 0.01686 0.01055 -0.0680 0.0042 1.0000
9.750 1.3149 0.01742 0.01115 -0.0668 0.0040 1.0000
10.000 1.3331 0.01805 0.01184 -0.0653 0.0037 1.0000
10.250 1.3496 0.01876 0.01263 -0.0635 0.0035 1.0000
10.500 1.3631 0.01953 0.01348 -0.0612 0.0034 1.0000
10.750 1.3777 0.02014 0.01416 -0.0591 0.0033 1.0000
11.000 1.3917 0.02079 0.01488 -0.0570 0.0033 1.0000
11.250 1.4047 0.02149 0.01565 -0.0548 0.0032 1.0000
11.500 1.4174 0.02220 0.01644 -0.0527 0.0031 1.0000
11.750 1.4291 0.02298 0.01729 -0.0504 0.0030 1.0000
12.000 1.4395 0.02385 0.01824 -0.0481 0.0030 1.0000
12.250 1.4505 0.02469 0.01914 -0.0460 0.0028 1.0000
12.500 1.4588 0.02572 0.02026 -0.0437 0.0028 1.0000
12.750 1.4661 0.02685 0.02148 -0.0414 0.0027 1.0000
13.000 1.4727 0.02806 0.02278 -0.0392 0.0026 1.0000
13.250 1.4785 0.02938 0.02418 -0.0371 0.0026 1.0000
13.500 1.4830 0.03085 0.02576 -0.0351 0.0025 1.0000
13.750 1.4870 0.03244 0.02744 -0.0333 0.0025 1.0000
14.000 1.4881 0.03436 0.02947 -0.0316 0.0024 1.0000
14.250 1.4888 0.03643 0.03165 -0.0302 0.0024 1.0000
14.500 1.4888 0.03871 0.03403 -0.0291 0.0024 1.0000
14.750 1.4876 0.04125 0.03668 -0.0283 0.0023 1.0000
15.000 1.4842 0.04420 0.03975 -0.0278 0.0023 1.0000
15.250 1.4789 0.04758 0.04325 -0.0277 0.0023 1.0000
15.500 1.4749 0.05103 0.04681 -0.0281 0.0023 1.0000
15.750 1.4679 0.05510 0.05101 -0.0290 0.0022 1.0000
16.000 1.4581 0.05984 0.05589 -0.0304 0.0022 1.0000
16.250 1.4481 0.06491 0.06109 -0.0322 0.0022 1.0000
16.500 1.4357 0.07068 0.06700 -0.0347 0.0022 1.0000
16.750 1.4192 0.07747 0.07394 -0.0379 0.0022 1.0000
17.000 1.4014 0.08482 0.08145 -0.0416 0.0022 1.0000
17.250 1.3851 0.09222 0.08899 -0.0455 0.0022 1.0000
17.500 1.3638 0.10086 0.09778 -0.0502 0.0022 1.0000
17.750 1.3429 0.10969 0.10676 -0.0551 0.0022 1.0000
18.000 1.3220 0.11877 0.11597 -0.0603 0.0022 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RG 8 AIRFOIL (rg8-il)