BOEING-VERTOL VR-15 AIRFOIL (vr15-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL VR-15 AIRFOIL (vr15-il) Reynolds number: 200,000 Max Cl/Cd: 49.3 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-vr15-il-200000.txt Download as CSV file: xf-vr15-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5978 0.09844 0.09510 -0.0041 1.0000 0.0339
-8.750 -0.6005 0.09331 0.08993 -0.0088 1.0000 0.0340
-8.500 -0.6072 0.08651 0.08322 -0.0080 1.0000 0.0348
-8.250 -0.6041 0.08311 0.07983 -0.0073 1.0000 0.0355
-8.000 -0.6006 0.07942 0.07612 -0.0081 1.0000 0.0362
-7.750 -0.5966 0.07539 0.07206 -0.0099 1.0000 0.0370
-7.500 -0.5917 0.07119 0.06781 -0.0117 1.0000 0.0381
-7.250 -0.5856 0.06680 0.06334 -0.0136 1.0000 0.0394
-7.000 -0.5777 0.06225 0.05865 -0.0153 1.0000 0.0413
-6.750 -0.5644 0.05892 0.05473 -0.0167 1.0000 0.0451
-6.500 -0.5629 0.05213 0.04765 -0.0165 1.0000 0.0462
-6.250 -0.5486 0.04876 0.04453 -0.0165 1.0000 0.0487
-6.000 -0.5334 0.04625 0.04194 -0.0159 1.0000 0.0529
-5.750 -0.5227 0.04258 0.03755 -0.0140 1.0000 0.0595
-5.500 -0.5073 0.03926 0.03438 -0.0134 1.0000 0.0620
-5.250 -0.4898 0.04088 0.03533 -0.0102 1.0000 0.0719
-5.000 -0.4829 0.03462 0.02933 -0.0092 1.0000 0.0758
-4.750 -0.4727 0.03308 0.02757 -0.0066 1.0000 0.0885
-4.500 -0.4590 0.03156 0.02601 -0.0044 1.0000 0.0964
-4.250 -0.4387 0.02552 0.01877 0.0009 1.0000 0.0569
-4.000 -0.4217 0.02285 0.01580 0.0031 1.0000 0.0551
-3.750 -0.4018 0.02078 0.01342 0.0051 1.0000 0.0537
-3.500 -0.3812 0.01955 0.01195 0.0067 1.0000 0.0552
-3.250 -0.3597 0.01859 0.01076 0.0081 1.0000 0.0568
-3.000 -0.3371 0.01760 0.00962 0.0092 1.0000 0.0575
-2.750 -0.3145 0.01683 0.00873 0.0103 1.0000 0.0586
-2.500 -0.2801 0.01538 0.00734 0.0087 0.9974 0.0631
-2.250 -0.2422 0.01465 0.00659 0.0066 0.9942 0.0661
-2.000 -0.2060 0.01410 0.00603 0.0048 0.9895 0.0701
-1.750 -0.1686 0.01343 0.00542 0.0026 0.9853 0.0760
-1.500 -0.1290 0.01301 0.00498 0.0001 0.9795 0.0855
-1.250 -0.0944 0.01073 0.00440 -0.0021 0.9728 0.4772
-1.000 0.0305 0.00975 0.00484 -0.0193 0.9891 1.0000
-0.750 0.0858 0.00961 0.00458 -0.0249 0.9803 1.0000
-0.500 0.1325 0.00949 0.00437 -0.0288 0.9687 1.0000
-0.250 0.1680 0.00943 0.00424 -0.0302 0.9545 1.0000
0.000 0.1945 0.00940 0.00416 -0.0296 0.9382 1.0000
0.250 0.2165 0.00937 0.00408 -0.0280 0.9213 1.0000
0.500 0.2369 0.00932 0.00399 -0.0260 0.9047 1.0000
0.750 0.2567 0.00926 0.00388 -0.0238 0.8879 1.0000
1.000 0.2768 0.00920 0.00376 -0.0216 0.8710 1.0000
1.250 0.2983 0.00914 0.00366 -0.0198 0.8508 1.0000
1.500 0.3198 0.00908 0.00355 -0.0179 0.8298 1.0000
1.750 0.3420 0.00903 0.00345 -0.0162 0.8057 1.0000
2.000 0.3646 0.00901 0.00335 -0.0146 0.7784 1.0000
2.250 0.3877 0.00903 0.00327 -0.0131 0.7440 1.0000
2.500 0.4111 0.00910 0.00320 -0.0117 0.6962 1.0000
2.750 0.4341 0.00931 0.00310 -0.0103 0.6114 1.0000
3.000 0.4560 0.01005 0.00311 -0.0091 0.4674 1.0000
3.250 0.4792 0.01084 0.00337 -0.0086 0.3716 1.0000
3.500 0.5033 0.01143 0.00368 -0.0081 0.3255 1.0000
3.750 0.5276 0.01192 0.00400 -0.0076 0.2968 1.0000
4.000 0.5519 0.01240 0.00432 -0.0071 0.2762 1.0000
4.250 0.5762 0.01284 0.00466 -0.0066 0.2588 1.0000
4.500 0.6005 0.01326 0.00503 -0.0060 0.2434 1.0000
4.750 0.6247 0.01366 0.00538 -0.0054 0.2296 1.0000
5.000 0.6490 0.01405 0.00573 -0.0049 0.2174 1.0000
5.250 0.6734 0.01441 0.00613 -0.0043 0.2067 1.0000
5.500 0.6975 0.01487 0.00660 -0.0036 0.1975 1.0000
5.750 0.7213 0.01534 0.00702 -0.0030 0.1885 1.0000
6.000 0.7456 0.01565 0.00743 -0.0024 0.1784 1.0000
6.250 0.7693 0.01604 0.00784 -0.0018 0.1679 1.0000
6.500 0.7927 0.01641 0.00823 -0.0011 0.1570 1.0000
6.750 0.8159 0.01680 0.00862 -0.0005 0.1456 1.0000
7.000 0.8394 0.01709 0.00900 0.0002 0.1333 1.0000
7.250 0.8622 0.01749 0.00944 0.0009 0.1196 1.0000
7.500 0.8844 0.01801 0.00999 0.0017 0.1056 1.0000
7.750 0.9059 0.01867 0.01065 0.0026 0.0917 1.0000
8.000 0.9270 0.01938 0.01137 0.0036 0.0787 1.0000
8.250 0.9484 0.02008 0.01214 0.0046 0.0687 1.0000
8.500 0.9681 0.02119 0.01333 0.0058 0.0614 1.0000
8.750 0.9853 0.02272 0.01481 0.0072 0.0551 1.0000
9.000 1.0056 0.02366 0.01598 0.0083 0.0511 1.0000
9.250 1.0244 0.02490 0.01731 0.0096 0.0476 1.0000
9.500 1.0406 0.02710 0.01955 0.0109 0.0445 1.0000
9.750 1.0576 0.02882 0.02155 0.0123 0.0425 1.0000
10.000 1.0744 0.03020 0.02318 0.0137 0.0402 1.0000
10.250 1.0894 0.03189 0.02512 0.0151 0.0383 1.0000
10.500 1.1031 0.03374 0.02716 0.0166 0.0369 1.0000
10.750 1.1149 0.03587 0.02945 0.0181 0.0357 1.0000
11.000 1.1214 0.03899 0.03281 0.0199 0.0347 1.0000
11.250 1.1167 0.04357 0.03778 0.0222 0.0341 1.0000
11.500 1.1109 0.04672 0.04127 0.0246 0.0337 1.0000
11.750 1.1026 0.04919 0.04405 0.0273 0.0333 1.0000
12.000 1.0876 0.05212 0.04723 0.0298 0.0331 1.0000
12.250 1.0696 0.05584 0.05120 0.0307 0.0329 1.0000
12.500 1.0488 0.06054 0.05613 0.0299 0.0329 1.0000
12.750 1.0249 0.06656 0.06236 0.0272 0.0331 1.0000
13.000 0.9972 0.07444 0.07042 0.0222 0.0335 1.0000
13.250 0.6769 0.11609 0.11249 0.0030 0.0460 1.0000
13.500 0.6748 0.11928 0.11568 0.0023 0.0451 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL VR-15 AIRFOIL (vr15-il)