BOEING-VERTOL VR-13 AIRFOIL (vr13-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL VR-13 AIRFOIL (vr13-il) Reynolds number: 1,000,000 Max Cl/Cd: 94.64 at α=8.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-vr13-il-1000000.txt Download as CSV file: xf-vr13-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL VR-13 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.5610 0.12922 0.12763 0.0179 1.0000 0.0091
-11.000 -0.5576 0.12494 0.12336 0.0161 1.0000 0.0091
-9.750 -0.8707 0.02571 0.02236 -0.0221 1.0000 0.0049
-9.500 -0.8576 0.02254 0.01879 -0.0208 1.0000 0.0049
-9.250 -0.8389 0.02047 0.01643 -0.0199 0.9984 0.0050
-9.000 -0.8128 0.01889 0.01462 -0.0203 0.9909 0.0051
-8.750 -0.7870 0.01774 0.01330 -0.0203 0.9832 0.0053
-8.500 -0.7642 0.01684 0.01225 -0.0194 0.9754 0.0055
-8.250 -0.7434 0.01611 0.01139 -0.0180 0.9677 0.0057
-8.000 -0.7231 0.01541 0.01056 -0.0164 0.9607 0.0059
-7.750 -0.7013 0.01475 0.00977 -0.0152 0.9536 0.0063
-7.500 -0.6789 0.01421 0.00910 -0.0139 0.9474 0.0066
-7.250 -0.6570 0.01329 0.00805 -0.0127 0.9408 0.0072
-7.000 -0.6339 0.01289 0.00760 -0.0116 0.9352 0.0078
-6.750 -0.6082 0.01250 0.00716 -0.0111 0.9292 0.0086
-6.500 -0.5843 0.01200 0.00658 -0.0101 0.9232 0.0097
-6.250 -0.5580 0.01183 0.00640 -0.0097 0.9173 0.0109
-6.000 -0.5315 0.01165 0.00618 -0.0092 0.9110 0.0120
-5.750 -0.5060 0.01141 0.00592 -0.0086 0.9052 0.0134
-5.500 -0.4781 0.01139 0.00592 -0.0086 0.8985 0.0144
-5.250 -0.4516 0.01130 0.00576 -0.0081 0.8921 0.0154
-5.000 -0.4252 0.01093 0.00536 -0.0077 0.8850 0.0165
-4.750 -0.3981 0.01086 0.00529 -0.0075 0.8786 0.0174
-4.500 -0.3702 0.01083 0.00525 -0.0074 0.8725 0.0187
-4.250 -0.3421 0.01084 0.00523 -0.0073 0.8657 0.0199
-4.000 -0.3162 0.01043 0.00476 -0.0068 0.8583 0.0215
-3.750 -0.2887 0.01038 0.00472 -0.0067 0.8489 0.0231
-3.500 -0.2614 0.01018 0.00450 -0.0064 0.8393 0.0244
-3.250 -0.2345 0.00994 0.00418 -0.0060 0.8304 0.0254
-3.000 -0.2068 0.00977 0.00398 -0.0058 0.8207 0.0261
-2.750 -0.1820 0.00908 0.00322 -0.0051 0.8115 0.0281
-2.250 -0.1273 0.00870 0.00278 -0.0046 0.7914 0.0313
-2.000 -0.0997 0.00855 0.00258 -0.0044 0.7823 0.0328
-1.750 -0.0719 0.00842 0.00240 -0.0042 0.7732 0.0337
-1.500 -0.0446 0.00817 0.00210 -0.0040 0.7646 0.0351
-1.250 -0.0171 0.00798 0.00185 -0.0038 0.7565 0.0373
-1.000 0.0110 0.00784 0.00171 -0.0037 0.7477 0.0398
-0.750 0.0390 0.00774 0.00158 -0.0036 0.7376 0.0425
-0.500 0.0668 0.00760 0.00143 -0.0034 0.7267 0.0530
-0.250 0.0898 0.00672 0.00128 -0.0027 0.7149 0.2901
0.000 0.1088 0.00561 0.00116 -0.0013 0.7005 0.6056
0.250 0.1243 0.00473 0.00111 0.0016 0.6842 0.8516
0.500 0.1733 0.00478 0.00130 -0.0026 0.6516 0.9552
0.750 0.2205 0.00525 0.00154 -0.0066 0.5917 0.9787
1.000 0.2673 0.00584 0.00175 -0.0108 0.5042 0.9873
1.250 0.3113 0.00639 0.00193 -0.0146 0.4203 0.9923
1.500 0.3537 0.00687 0.00206 -0.0181 0.3445 0.9955
1.750 0.4002 0.00718 0.00214 -0.0224 0.3005 0.9987
2.000 0.4360 0.00737 0.00220 -0.0243 0.2753 1.0000
2.250 0.4628 0.00749 0.00223 -0.0242 0.2585 1.0000
2.500 0.4896 0.00761 0.00227 -0.0241 0.2441 1.0000
3.000 0.5428 0.00783 0.00237 -0.0237 0.2219 1.0000
3.250 0.5692 0.00794 0.00243 -0.0235 0.2128 1.0000
3.500 0.5956 0.00807 0.00250 -0.0233 0.2037 1.0000
3.750 0.6218 0.00817 0.00258 -0.0231 0.1958 1.0000
4.000 0.6477 0.00833 0.00267 -0.0228 0.1863 1.0000
4.250 0.6735 0.00844 0.00276 -0.0225 0.1796 1.0000
4.500 0.6990 0.00861 0.00288 -0.0221 0.1720 1.0000
4.750 0.7245 0.00872 0.00299 -0.0217 0.1670 1.0000
5.000 0.7498 0.00889 0.00312 -0.0213 0.1604 1.0000
5.250 0.7751 0.00903 0.00325 -0.0209 0.1552 1.0000
5.500 0.8004 0.00919 0.00340 -0.0204 0.1497 1.0000
5.750 0.8252 0.00941 0.00358 -0.0200 0.1435 1.0000
6.000 0.8505 0.00954 0.00372 -0.0195 0.1403 1.0000
6.250 0.8754 0.00971 0.00389 -0.0190 0.1362 1.0000
6.500 0.8999 0.00995 0.00410 -0.0185 0.1311 1.0000
6.750 0.9248 0.01012 0.00429 -0.0180 0.1284 1.0000
7.000 0.9496 0.01029 0.00447 -0.0175 0.1249 1.0000
7.250 0.9740 0.01050 0.00467 -0.0170 0.1206 1.0000
7.500 0.9980 0.01077 0.00493 -0.0164 0.1162 1.0000
7.750 1.0227 0.01092 0.00512 -0.0159 0.1139 1.0000
8.000 1.0470 0.01112 0.00534 -0.0154 0.1104 1.0000
8.250 1.0707 0.01140 0.00559 -0.0148 0.1054 1.0000
8.500 1.0946 0.01162 0.00584 -0.0142 0.1022 1.0000
8.750 1.1187 0.01182 0.00607 -0.0136 0.0987 1.0000
9.000 1.1419 0.01212 0.00635 -0.0129 0.0931 1.0000
9.250 1.1653 0.01237 0.00663 -0.0123 0.0884 1.0000
9.500 1.1882 0.01268 0.00692 -0.0116 0.0817 1.0000
9.750 1.2106 0.01302 0.00725 -0.0108 0.0743 1.0000
10.000 1.2318 0.01349 0.00766 -0.0099 0.0649 1.0000
10.250 1.2522 0.01402 0.00814 -0.0089 0.0567 1.0000
10.500 1.2733 0.01446 0.00859 -0.0080 0.0527 1.0000
10.750 1.2930 0.01502 0.00913 -0.0069 0.0476 1.0000
11.000 1.3143 0.01540 0.00957 -0.0060 0.0451 1.0000
11.250 1.3344 0.01587 0.01006 -0.0050 0.0414 1.0000
11.500 1.3530 0.01647 0.01067 -0.0038 0.0366 1.0000
11.750 1.3726 0.01694 0.01116 -0.0027 0.0333 1.0000
12.000 1.3895 0.01763 0.01183 -0.0013 0.0291 1.0000
12.250 1.4079 0.01815 0.01239 -0.0001 0.0274 1.0000
12.500 1.4245 0.01879 0.01306 0.0014 0.0254 1.0000
12.750 1.4393 0.01956 0.01385 0.0030 0.0232 1.0000
13.000 1.4554 0.02018 0.01455 0.0044 0.0221 1.0000
13.250 1.4715 0.02075 0.01519 0.0058 0.0213 1.0000
13.500 1.4842 0.02143 0.01592 0.0077 0.0202 1.0000
13.750 1.4938 0.02222 0.01676 0.0100 0.0193 1.0000
14.000 1.5020 0.02318 0.01778 0.0122 0.0181 1.0000
14.250 1.5098 0.02425 0.01893 0.0141 0.0170 1.0000
14.500 1.5219 0.02507 0.01983 0.0155 0.0166 1.0000
14.750 1.5328 0.02602 0.02086 0.0168 0.0160 1.0000
15.000 1.5420 0.02713 0.02205 0.0181 0.0153 1.0000
15.250 1.5494 0.02845 0.02343 0.0194 0.0146 1.0000
15.500 1.5544 0.03001 0.02506 0.0206 0.0138 1.0000
15.750 1.5565 0.03193 0.02706 0.0217 0.0131 1.0000
16.000 1.5605 0.03376 0.02898 0.0224 0.0127 1.0000
16.250 1.5661 0.03552 0.03084 0.0229 0.0123 1.0000
16.500 1.5688 0.03764 0.03306 0.0232 0.0118 1.0000
16.750 1.5689 0.04013 0.03565 0.0233 0.0113 1.0000
17.000 1.5659 0.04311 0.03872 0.0231 0.0108 1.0000
17.250 1.5589 0.04675 0.04246 0.0224 0.0104 1.0000
17.500 1.5478 0.05119 0.04702 0.0211 0.0100 1.0000
17.750 1.5316 0.05671 0.05269 0.0189 0.0097 1.0000
18.000 1.5103 0.06360 0.05975 0.0156 0.0096 1.0000
18.250 1.4810 0.07232 0.06866 0.0110 0.0097 1.0000
18.500 1.4420 0.08309 0.07965 0.0055 0.0101 1.0000
18.750 1.3902 0.09594 0.09275 -0.0007 0.0108 1.0000
19.000 1.3296 0.11023 0.10726 -0.0076 0.0116 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL VR-13 AIRFOIL (vr13-il)