BOEING-VERTOL VR-15 AIRFOIL (vr15-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL VR-15 AIRFOIL (vr15-il) Reynolds number: 500,000 Max Cl/Cd: 70.85 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-vr15-il-500000.txt Download as CSV file: xf-vr15-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL VR-15 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.6057 0.09138 0.08928 0.0069 1.0000 0.0181
-8.500 -0.6076 0.08538 0.08330 0.0004 1.0000 0.0183
-8.250 -0.6114 0.07944 0.07734 -0.0056 1.0000 0.0184
-8.000 -0.6092 0.07369 0.07152 -0.0098 1.0000 0.0185
-7.750 -0.6051 0.06817 0.06590 -0.0126 1.0000 0.0185
-7.000 -0.6112 0.04508 0.04214 -0.0150 1.0000 0.0145
-6.750 -0.5946 0.04350 0.04050 -0.0148 1.0000 0.0149
-6.500 -0.6050 0.02892 0.02489 -0.0103 1.0000 0.0117
-6.250 -0.6018 0.02019 0.01499 -0.0055 1.0000 0.0135
-6.000 -0.5844 0.01861 0.01317 -0.0037 1.0000 0.0147
-5.750 -0.5650 0.01882 0.01342 -0.0023 1.0000 0.0156
-5.500 -0.5479 0.01841 0.01291 -0.0003 1.0000 0.0172
-5.250 -0.5210 0.01716 0.01140 -0.0003 0.9981 0.0196
-5.000 -0.4854 0.01757 0.01192 -0.0025 0.9955 0.0217
-4.750 -0.4493 0.01768 0.01193 -0.0042 0.9929 0.0247
-4.500 -0.4175 0.01559 0.00961 -0.0053 0.9897 0.0271
-4.250 -0.3823 0.01508 0.00908 -0.0071 0.9869 0.0294
-4.000 -0.3469 0.01425 0.00816 -0.0086 0.9841 0.0314
-3.750 -0.3142 0.01350 0.00731 -0.0095 0.9792 0.0328
-3.500 -0.2814 0.01319 0.00692 -0.0104 0.9737 0.0341
-3.250 -0.2548 0.01162 0.00531 -0.0101 0.9673 0.0368
-3.000 -0.2292 0.01113 0.00481 -0.0094 0.9594 0.0386
-2.750 -0.2057 0.01077 0.00443 -0.0083 0.9504 0.0403
-2.500 -0.1842 0.01049 0.00410 -0.0065 0.9399 0.0423
-2.250 -0.1637 0.01022 0.00378 -0.0045 0.9280 0.0439
-2.000 -0.1419 0.00993 0.00344 -0.0028 0.9156 0.0449
-1.750 -0.1190 0.00957 0.00302 -0.0014 0.9036 0.0466
-1.500 -0.0956 0.00921 0.00262 -0.0001 0.8915 0.0514
-1.250 -0.0710 0.00898 0.00235 0.0009 0.8798 0.0570
-1.000 -0.0482 0.00835 0.00207 0.0021 0.8685 0.1571
-0.750 -0.0372 0.00641 0.00180 0.0050 0.8571 0.6307
-0.500 0.0025 0.00559 0.00202 0.0033 0.8475 0.9372
-0.250 0.0486 0.00586 0.00222 -0.0001 0.8371 0.9713
0.000 0.1311 0.00629 0.00255 -0.0115 0.8249 1.0000
0.250 0.1573 0.00626 0.00242 -0.0110 0.8083 1.0000
0.500 0.1836 0.00624 0.00232 -0.0106 0.7903 1.0000
0.750 0.2100 0.00626 0.00222 -0.0102 0.7704 1.0000
1.000 0.2368 0.00627 0.00214 -0.0100 0.7484 1.0000
1.250 0.2637 0.00631 0.00208 -0.0097 0.7234 1.0000
1.500 0.2907 0.00637 0.00203 -0.0095 0.6912 1.0000
1.750 0.3179 0.00650 0.00197 -0.0093 0.6430 1.0000
2.000 0.3454 0.00682 0.00194 -0.0093 0.5585 1.0000
2.250 0.3733 0.00737 0.00201 -0.0097 0.4484 1.0000
2.500 0.4007 0.00788 0.00213 -0.0100 0.3582 1.0000
2.750 0.4274 0.00825 0.00225 -0.0101 0.3065 1.0000
3.250 0.4797 0.00875 0.00248 -0.0097 0.2523 1.0000
3.500 0.5056 0.00895 0.00262 -0.0094 0.2352 1.0000
3.750 0.5314 0.00915 0.00276 -0.0091 0.2208 1.0000
4.000 0.5571 0.00935 0.00291 -0.0088 0.2082 1.0000
4.250 0.5826 0.00955 0.00309 -0.0084 0.1966 1.0000
4.500 0.6080 0.00977 0.00326 -0.0081 0.1856 1.0000
4.750 0.6334 0.00999 0.00345 -0.0076 0.1760 1.0000
5.000 0.6587 0.01019 0.00365 -0.0072 0.1679 1.0000
5.250 0.6837 0.01046 0.00388 -0.0068 0.1587 1.0000
5.500 0.7089 0.01062 0.00405 -0.0063 0.1494 1.0000
5.750 0.7339 0.01085 0.00429 -0.0059 0.1417 1.0000
6.000 0.7587 0.01110 0.00451 -0.0054 0.1333 1.0000
6.250 0.7837 0.01130 0.00475 -0.0049 0.1256 1.0000
6.500 0.8082 0.01158 0.00502 -0.0044 0.1169 1.0000
6.750 0.8328 0.01182 0.00525 -0.0039 0.1069 1.0000
7.000 0.8573 0.01210 0.00553 -0.0034 0.0958 1.0000
7.250 0.8813 0.01244 0.00586 -0.0028 0.0827 1.0000
7.500 0.9046 0.01290 0.00624 -0.0022 0.0682 1.0000
7.750 0.9270 0.01351 0.00678 -0.0014 0.0552 1.0000
8.000 0.9495 0.01409 0.00737 -0.0007 0.0472 1.0000
8.250 0.9716 0.01471 0.00804 0.0002 0.0411 1.0000
8.500 0.9935 0.01534 0.00871 0.0011 0.0362 1.0000
8.750 1.0143 0.01611 0.00954 0.0021 0.0327 1.0000
9.000 1.0364 0.01663 0.01013 0.0029 0.0301 1.0000
9.250 1.0570 0.01736 0.01089 0.0039 0.0278 1.0000
9.500 1.0723 0.01879 0.01244 0.0056 0.0255 1.0000
9.750 1.0930 0.01942 0.01320 0.0067 0.0247 1.0000
10.000 1.1125 0.02016 0.01404 0.0078 0.0234 1.0000
10.250 1.1315 0.02093 0.01490 0.0089 0.0222 1.0000
10.500 1.1493 0.02178 0.01582 0.0102 0.0210 1.0000
10.750 1.1637 0.02301 0.01714 0.0118 0.0200 1.0000
11.000 1.1694 0.02541 0.01973 0.0144 0.0190 1.0000
11.250 1.1824 0.02676 0.02125 0.0161 0.0186 1.0000
11.500 1.1963 0.02788 0.02254 0.0177 0.0182 1.0000
11.750 1.2076 0.02921 0.02404 0.0194 0.0177 1.0000
12.000 1.2167 0.03065 0.02563 0.0213 0.0172 1.0000
12.250 1.2227 0.03208 0.02721 0.0235 0.0167 1.0000
12.500 1.2251 0.03351 0.02878 0.0260 0.0162 1.0000
12.750 1.2268 0.03514 0.03054 0.0280 0.0158 1.0000
13.000 1.2276 0.03699 0.03251 0.0295 0.0155 1.0000
13.250 1.2262 0.03925 0.03491 0.0306 0.0152 1.0000
13.500 1.2226 0.04195 0.03775 0.0310 0.0149 1.0000
13.750 1.2169 0.04516 0.04110 0.0308 0.0147 1.0000
14.000 1.2081 0.04910 0.04518 0.0299 0.0145 1.0000
14.250 1.1959 0.05391 0.05016 0.0279 0.0144 1.0000
14.500 1.1814 0.05968 0.05611 0.0247 0.0143 1.0000
14.750 1.1646 0.06645 0.06306 0.0204 0.0143 1.0000
15.000 1.1464 0.07410 0.07089 0.0153 0.0143 1.0000
15.250 1.1258 0.08261 0.07957 0.0098 0.0144 1.0000
15.500 1.0990 0.09258 0.08973 0.0035 0.0145 1.0000
15.750 1.0456 0.10860 0.10603 -0.0059 0.0151 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL VR-15 AIRFOIL (vr15-il)