BOEING-VERTOL VR-9 AIRFOIL (vr9-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: BOEING-VERTOL VR-9 AIRFOIL (vr9-il) Reynolds number: 100,000 Max Cl/Cd: 27.43 at α=2.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-vr9-il-100000-n5.txt Download as CSV file: xf-vr9-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL VR-9 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.6361 0.10248 0.09766 0.0127 1.0000 0.0409
-8.750 -0.6352 0.09840 0.09363 0.0104 1.0000 0.0423
-8.500 -0.6360 0.09413 0.08941 0.0076 1.0000 0.0438
-6.750 -0.5805 0.04728 0.04253 -0.0147 1.0000 0.0327
-6.500 -0.5768 0.04096 0.03598 -0.0156 1.0000 0.0290
-6.250 -0.5989 0.04815 0.04234 -0.0152 1.0000 0.0252
-6.000 -0.5809 0.04307 0.03662 -0.0145 1.0000 0.0208
-5.750 -0.5646 0.03914 0.03243 -0.0141 1.0000 0.0199
-5.500 -0.5456 0.03540 0.02815 -0.0134 1.0000 0.0194
-5.250 -0.5247 0.03198 0.02424 -0.0124 1.0000 0.0191
-5.000 -0.5021 0.02893 0.02065 -0.0114 1.0000 0.0198
-4.750 -0.4797 0.02625 0.01749 -0.0106 1.0000 0.0229
-4.500 -0.4552 0.02402 0.01491 -0.0097 1.0000 0.0249
-4.250 -0.4298 0.02172 0.01227 -0.0086 1.0000 0.0257
-4.000 -0.4049 0.01975 0.01010 -0.0075 1.0000 0.0270
-3.750 -0.3808 0.01880 0.00911 -0.0069 1.0000 0.0332
-3.500 -0.3568 0.01760 0.00778 -0.0059 1.0000 0.0356
-3.250 -0.3348 0.01628 0.00640 -0.0048 1.0000 0.0379
-3.000 -0.3114 0.01545 0.00544 -0.0039 1.0000 0.0412
-2.750 -0.2873 0.01484 0.00458 -0.0031 1.0000 0.0466
-2.500 -0.2633 0.01423 0.00397 -0.0024 1.0000 0.0581
-2.250 -0.2392 0.01366 0.00341 -0.0016 1.0000 0.0789
-2.000 -0.2213 0.01163 0.00292 -0.0008 1.0000 0.4264
-1.750 -0.2034 0.01070 0.00283 0.0015 1.0000 0.6517
-1.500 -0.1816 0.01035 0.00262 0.0033 1.0000 0.7338
-1.250 -0.1573 0.01014 0.00249 0.0045 1.0000 0.7945
-1.000 -0.1184 0.00996 0.00243 0.0030 1.0000 0.8963
-0.750 -0.0734 0.00990 0.00229 -0.0005 1.0000 0.9519
-0.500 -0.0363 0.00986 0.00215 -0.0028 1.0000 0.9823
-0.250 0.0026 0.00985 0.00207 -0.0055 0.9912 1.0000
0.000 0.0420 0.00988 0.00205 -0.0082 0.9603 1.0000
0.250 0.0807 0.00991 0.00205 -0.0106 0.9281 1.0000
0.500 0.1167 0.00995 0.00209 -0.0124 0.8976 1.0000
0.750 0.1500 0.01000 0.00213 -0.0135 0.8680 1.0000
1.000 0.1800 0.01008 0.00220 -0.0139 0.8393 1.0000
1.250 0.2075 0.01018 0.00232 -0.0137 0.8128 1.0000
1.500 0.2338 0.01030 0.00250 -0.0131 0.7877 1.0000
1.750 0.2590 0.01044 0.00268 -0.0123 0.7586 1.0000
2.000 0.2819 0.01069 0.00281 -0.0108 0.7023 1.0000
2.250 0.3034 0.01106 0.00295 -0.0089 0.6206 1.0000
2.500 0.3223 0.01187 0.00303 -0.0068 0.4313 1.0000
2.750 0.3377 0.01420 0.00358 -0.0057 0.0650 1.0000
3.000 0.3617 0.01488 0.00426 -0.0050 0.0478 1.0000
3.250 0.3865 0.01541 0.00496 -0.0043 0.0431 1.0000
3.500 0.4108 0.01604 0.00576 -0.0036 0.0404 1.0000
3.750 0.4346 0.01681 0.00670 -0.0028 0.0386 1.0000
4.000 0.4575 0.01780 0.00783 -0.0019 0.0373 1.0000
4.250 0.4793 0.01935 0.00943 -0.0008 0.0354 1.0000
4.500 0.5043 0.02019 0.01046 0.0000 0.0298 1.0000
4.750 0.5283 0.02186 0.01233 0.0009 0.0277 1.0000
5.000 0.5525 0.02393 0.01459 0.0018 0.0249 1.0000
5.250 0.5752 0.02639 0.01723 0.0024 0.0211 1.0000
5.500 0.5991 0.02908 0.02046 0.0036 0.0205 1.0000
5.750 0.6212 0.03222 0.02416 0.0048 0.0202 1.0000
6.000 0.6413 0.03570 0.02817 0.0059 0.0204 1.0000
6.250 0.6589 0.03955 0.03245 0.0068 0.0208 1.0000
6.500 0.6765 0.04414 0.03776 0.0082 0.0238 1.0000
6.750 0.6891 0.04875 0.04276 0.0087 0.0251 1.0000
7.000 0.6994 0.05325 0.04760 0.0088 0.0254 1.0000
7.250 0.7069 0.05791 0.05256 0.0086 0.0254 1.0000
7.500 0.7116 0.06263 0.05751 0.0079 0.0255 1.0000
7.750 0.7144 0.06714 0.06216 0.0070 0.0260 1.0000
8.000 0.7159 0.07155 0.06665 0.0061 0.0263 1.0000
8.250 0.7163 0.07599 0.07114 0.0051 0.0265 1.0000
8.500 0.7132 0.08056 0.07577 0.0036 0.0266 1.0000
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Polar data table (+)
Polar graphs
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