BOEING VERTOL V13006-.7 AIRFOIL (v13006-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: BOEING VERTOL V13006-.7 AIRFOIL (v13006-il) Reynolds number: 100,000 Max Cl/Cd: 35.32 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-v13006-il-100000-n5.txt Download as CSV file: xf-v13006-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING VERTOL V13006-.7 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.6288 0.10037 0.09539 0.0163 1.0000 0.0499
-8.250 -0.6296 0.09610 0.09117 0.0116 1.0000 0.0515
-8.000 -0.6316 0.09099 0.08608 0.0020 1.0000 0.0526
-7.750 -0.6285 0.08608 0.08104 -0.0052 1.0000 0.0531
-7.500 -0.6234 0.08168 0.07645 -0.0091 1.0000 0.0533
-7.250 -0.6161 0.07579 0.07058 -0.0103 1.0000 0.0539
-7.000 -0.6063 0.07117 0.06604 -0.0097 1.0000 0.0549
-6.750 -0.5952 0.06729 0.06213 -0.0101 1.0000 0.0560
-6.500 -0.5827 0.06335 0.05810 -0.0113 1.0000 0.0572
-6.250 -0.5623 0.05646 0.05076 -0.0140 1.0000 0.0382
-6.000 -0.5489 0.05203 0.04621 -0.0145 1.0000 0.0350
-5.500 -0.5056 0.04258 0.03565 -0.0149 1.0000 0.0278
-5.250 -0.4873 0.03865 0.03146 -0.0148 1.0000 0.0274
-5.000 -0.4665 0.03536 0.02778 -0.0144 1.0000 0.0272
-4.750 -0.4433 0.03288 0.02479 -0.0138 1.0000 0.0282
-4.500 -0.4226 0.02957 0.02121 -0.0135 1.0000 0.0296
-4.250 -0.3988 0.02729 0.01861 -0.0129 1.0000 0.0302
-4.000 -0.3737 0.02513 0.01608 -0.0122 1.0000 0.0302
-3.750 -0.3484 0.02325 0.01390 -0.0115 1.0000 0.0304
-3.500 -0.3229 0.02160 0.01199 -0.0107 1.0000 0.0310
-3.250 -0.2973 0.02016 0.01039 -0.0100 1.0000 0.0319
-3.000 -0.2719 0.01898 0.00910 -0.0093 1.0000 0.0339
-2.750 -0.2467 0.01805 0.00808 -0.0086 1.0000 0.0374
-2.500 -0.2217 0.01707 0.00700 -0.0078 1.0000 0.0392
-2.250 -0.1980 0.01597 0.00592 -0.0069 1.0000 0.0413
-2.000 -0.1743 0.01513 0.00508 -0.0060 1.0000 0.0451
-1.750 -0.1497 0.01460 0.00448 -0.0053 1.0000 0.0533
-1.500 -0.1252 0.01396 0.00386 -0.0046 1.0000 0.0663
-1.250 -0.1012 0.01285 0.00335 -0.0041 1.0000 0.1757
-1.000 -0.0395 0.01023 0.00331 -0.0093 1.0000 1.0000
-0.750 -0.0183 0.01020 0.00313 -0.0081 1.0000 1.0000
-0.500 0.0029 0.01019 0.00300 -0.0069 1.0000 1.0000
-0.250 0.0242 0.01020 0.00291 -0.0057 1.0000 1.0000
0.000 0.0457 0.01023 0.00288 -0.0046 1.0000 1.0000
0.250 0.0674 0.01028 0.00288 -0.0036 1.0000 1.0000
0.500 0.0915 0.01035 0.00292 -0.0032 0.9983 1.0000
0.750 0.1376 0.01042 0.00298 -0.0071 0.9780 1.0000
1.000 0.1826 0.01045 0.00302 -0.0106 0.9459 1.0000
1.250 0.2245 0.01047 0.00303 -0.0131 0.9026 1.0000
1.500 0.2601 0.01052 0.00304 -0.0141 0.8534 1.0000
1.750 0.2891 0.01062 0.00305 -0.0137 0.7996 1.0000
2.000 0.3147 0.01080 0.00309 -0.0125 0.7422 1.0000
2.250 0.3390 0.01105 0.00318 -0.0112 0.6831 1.0000
2.500 0.3631 0.01137 0.00329 -0.0099 0.6222 1.0000
2.750 0.3875 0.01174 0.00346 -0.0089 0.5619 1.0000
3.000 0.4119 0.01213 0.00368 -0.0080 0.5056 1.0000
3.250 0.4364 0.01256 0.00398 -0.0072 0.4523 1.0000
3.500 0.4600 0.01313 0.00425 -0.0064 0.3849 1.0000
3.750 0.4833 0.01378 0.00452 -0.0057 0.3089 1.0000
4.000 0.5075 0.01437 0.00488 -0.0052 0.2537 1.0000
4.250 0.5310 0.01516 0.00528 -0.0048 0.1643 1.0000
4.500 0.5531 0.01657 0.00603 -0.0043 0.0578 1.0000
4.750 0.5770 0.01766 0.00703 -0.0037 0.0423 1.0000
5.000 0.6013 0.01862 0.00811 -0.0030 0.0378 1.0000
5.250 0.6255 0.01954 0.00921 -0.0024 0.0353 1.0000
5.500 0.6492 0.02057 0.01042 -0.0017 0.0334 1.0000
5.750 0.6724 0.02173 0.01169 -0.0009 0.0320 1.0000
6.000 0.6954 0.02305 0.01311 -0.0001 0.0309 1.0000
6.250 0.7178 0.02463 0.01476 0.0005 0.0292 1.0000
6.500 0.7399 0.02670 0.01700 0.0012 0.0276 1.0000
6.750 0.7634 0.02823 0.01880 0.0019 0.0267 1.0000
7.000 0.7860 0.03024 0.02112 0.0027 0.0263 1.0000
7.250 0.8073 0.03254 0.02380 0.0034 0.0259 1.0000
7.500 0.8269 0.03518 0.02686 0.0042 0.0257 1.0000
7.750 0.8441 0.03819 0.03035 0.0050 0.0256 1.0000
8.000 0.8585 0.04163 0.03431 0.0057 0.0255 1.0000
8.250 0.8693 0.04551 0.03868 0.0063 0.0256 1.0000
8.500 0.8758 0.04984 0.04347 0.0066 0.0257 1.0000
8.750 0.8776 0.05453 0.04858 0.0066 0.0258 1.0000
9.000 0.8739 0.05956 0.05397 0.0059 0.0258 1.0000
9.250 0.8642 0.06497 0.05967 0.0042 0.0259 1.0000
9.500 0.8480 0.07069 0.06556 0.0013 0.0262 1.0000
9.750 0.8307 0.07851 0.07348 -0.0061 0.0266 1.0000
10.000 0.8175 0.08788 0.08285 -0.0140 0.0270 1.0000
10.250 0.8085 0.09573 0.09065 -0.0189 0.0274 1.0000
10.500 0.8020 0.10244 0.09733 -0.0224 0.0276 1.0000
10.750 0.7973 0.10859 0.10345 -0.0252 0.0278 1.0000
11.000 0.7936 0.11457 0.10940 -0.0279 0.0279 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING VERTOL V13006-.7 AIRFOIL (v13006-il)