BOEING-VERTOL V23010-1.58 AIRFOIL (v23010-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: BOEING-VERTOL V23010-1.58 AIRFOIL (v23010-il) Reynolds number: 500,000 Max Cl/Cd: 78 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-v23010-il-500000-n5.txt Download as CSV file: xf-v23010-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL V23010-1.58 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5570 0.07774 0.07526 -0.0204 1.0000 0.0193
-9.250 -0.5659 0.07199 0.06951 -0.0258 1.0000 0.0191
-9.000 -0.5784 0.06788 0.06537 -0.0284 1.0000 0.0189
-8.750 -0.5935 0.06437 0.06181 -0.0280 1.0000 0.0188
-8.500 -0.6036 0.06008 0.05741 -0.0281 1.0000 0.0187
-8.250 -0.6112 0.05543 0.05260 -0.0277 1.0000 0.0188
-8.000 -0.6209 0.04913 0.04599 -0.0263 1.0000 0.0192
-7.750 -0.6179 0.04549 0.04215 -0.0247 1.0000 0.0192
-7.500 -0.6138 0.04169 0.03812 -0.0228 1.0000 0.0192
-7.250 -0.6070 0.03825 0.03444 -0.0207 1.0000 0.0192
-7.000 -0.5995 0.03451 0.03040 -0.0182 1.0000 0.0194
-6.750 -0.5902 0.03124 0.02684 -0.0157 1.0000 0.0194
-6.500 -0.5799 0.02842 0.02377 -0.0133 1.0000 0.0195
-6.250 -0.5528 0.02641 0.02158 -0.0143 0.9954 0.0197
-6.000 -0.5206 0.02542 0.02052 -0.0160 0.9878 0.0200
-5.750 -0.4873 0.02415 0.01913 -0.0177 0.9789 0.0203
-5.500 -0.4548 0.02198 0.01671 -0.0190 0.9682 0.0205
-5.250 -0.4221 0.01988 0.01434 -0.0202 0.9549 0.0206
-5.000 -0.3876 0.01823 0.01247 -0.0216 0.9337 0.0208
-4.500 -0.3287 0.01595 0.00967 -0.0218 0.8574 0.0213
-4.250 -0.3045 0.01518 0.00869 -0.0208 0.8219 0.0215
-4.000 -0.2807 0.01452 0.00781 -0.0197 0.7841 0.0217
-3.750 -0.2565 0.01395 0.00704 -0.0188 0.7592 0.0219
-3.500 -0.2315 0.01346 0.00640 -0.0180 0.7406 0.0222
-3.250 -0.2061 0.01304 0.00588 -0.0173 0.7248 0.0226
-3.000 -0.1807 0.01264 0.00538 -0.0167 0.7074 0.0229
-2.750 -0.1556 0.01226 0.00491 -0.0159 0.6874 0.0230
-2.500 -0.1305 0.01194 0.00448 -0.0152 0.6617 0.0231
-2.250 -0.1066 0.01153 0.00397 -0.0143 0.6305 0.0234
-2.000 -0.0824 0.01125 0.00357 -0.0135 0.5988 0.0237
-1.750 -0.0577 0.01107 0.00327 -0.0128 0.5678 0.0241
-1.500 -0.0326 0.01094 0.00302 -0.0121 0.5364 0.0245
-1.250 -0.0073 0.01085 0.00280 -0.0115 0.5003 0.0250
-1.000 0.0179 0.01081 0.00261 -0.0109 0.4580 0.0254
-0.750 0.0431 0.01082 0.00245 -0.0103 0.4166 0.0260
-0.500 0.0688 0.01082 0.00231 -0.0098 0.3833 0.0265
-0.250 0.0948 0.01084 0.00221 -0.0093 0.3557 0.0272
0.000 0.1209 0.01082 0.00211 -0.0089 0.3318 0.0286
0.250 0.1473 0.01083 0.00204 -0.0085 0.3102 0.0299
0.500 0.1739 0.01085 0.00198 -0.0081 0.2947 0.0313
0.750 0.2005 0.01087 0.00195 -0.0078 0.2820 0.0337
1.250 0.2538 0.01088 0.00194 -0.0071 0.2623 0.0593
1.500 0.2701 0.00943 0.00184 -0.0055 0.2569 0.5296
1.750 0.2902 0.00896 0.00192 -0.0039 0.2524 0.6924
2.000 0.3121 0.00871 0.00202 -0.0024 0.2486 0.7882
2.250 0.3360 0.00861 0.00212 -0.0011 0.2434 0.8514
2.500 0.3642 0.00864 0.00224 -0.0009 0.2381 0.9010
2.750 0.4121 0.00876 0.00241 -0.0048 0.2335 0.9542
3.000 0.4517 0.00890 0.00253 -0.0072 0.2293 0.9699
3.250 0.4873 0.00903 0.00262 -0.0088 0.2251 0.9762
3.500 0.5201 0.00921 0.00274 -0.0098 0.2205 0.9836
3.750 0.5556 0.00931 0.00284 -0.0114 0.2169 0.9878
4.000 0.5889 0.00943 0.00295 -0.0126 0.2130 0.9923
4.250 0.6225 0.00957 0.00306 -0.0138 0.2088 0.9962
4.500 0.6567 0.00974 0.00319 -0.0152 0.2044 0.9996
4.750 0.6824 0.00985 0.00332 -0.0147 0.2013 1.0000
5.000 0.7063 0.00998 0.00345 -0.0139 0.1973 1.0000
5.250 0.7301 0.01012 0.00358 -0.0130 0.1928 1.0000
5.500 0.7538 0.01031 0.00374 -0.0122 0.1883 1.0000
5.750 0.7781 0.01044 0.00390 -0.0114 0.1837 1.0000
6.000 0.8022 0.01061 0.00405 -0.0106 0.1769 1.0000
6.250 0.8263 0.01079 0.00422 -0.0099 0.1689 1.0000
6.500 0.8501 0.01101 0.00440 -0.0091 0.1591 1.0000
6.750 0.8737 0.01125 0.00460 -0.0083 0.1483 1.0000
7.000 0.8973 0.01151 0.00483 -0.0075 0.1396 1.0000
7.250 0.9204 0.01182 0.00509 -0.0067 0.1318 1.0000
7.500 0.9438 0.01210 0.00537 -0.0059 0.1256 1.0000
7.750 0.9667 0.01245 0.00568 -0.0050 0.1191 1.0000
8.000 0.9903 0.01273 0.00598 -0.0043 0.1141 1.0000
8.250 1.0134 0.01306 0.00632 -0.0035 0.1098 1.0000
8.500 1.0366 0.01339 0.00666 -0.0027 0.1060 1.0000
8.750 1.0604 0.01366 0.00698 -0.0020 0.1022 1.0000
9.000 1.0833 0.01402 0.00734 -0.0013 0.0969 1.0000
9.250 1.1064 0.01436 0.00771 -0.0006 0.0918 1.0000
9.500 1.1290 0.01475 0.00809 0.0002 0.0835 1.0000
9.750 1.1506 0.01523 0.00852 0.0010 0.0711 1.0000
10.000 1.1698 0.01596 0.00915 0.0021 0.0549 1.0000
10.250 1.1888 0.01670 0.00984 0.0033 0.0450 1.0000
10.750 1.2277 0.01798 0.01116 0.0054 0.0358 1.0000
11.000 1.2461 0.01868 0.01189 0.0066 0.0330 1.0000
11.250 1.2653 0.01927 0.01254 0.0077 0.0313 1.0000
11.500 1.2833 0.01991 0.01324 0.0089 0.0296 1.0000
11.750 1.2997 0.02067 0.01404 0.0103 0.0281 1.0000
12.000 1.3153 0.02142 0.01486 0.0118 0.0269 1.0000
12.250 1.3311 0.02210 0.01562 0.0132 0.0260 1.0000
12.500 1.3451 0.02284 0.01645 0.0148 0.0250 1.0000
12.750 1.3553 0.02366 0.01733 0.0170 0.0241 1.0000
13.000 1.3613 0.02461 0.01834 0.0196 0.0233 1.0000
13.250 1.3656 0.02571 0.01951 0.0221 0.0225 1.0000
13.500 1.3725 0.02672 0.02063 0.0241 0.0220 1.0000
13.750 1.3778 0.02793 0.02194 0.0258 0.0215 1.0000
14.000 1.3816 0.02937 0.02348 0.0272 0.0209 1.0000
14.250 1.3834 0.03112 0.02533 0.0283 0.0203 1.0000
14.500 1.3829 0.03329 0.02760 0.0288 0.0199 1.0000
14.750 1.3800 0.03597 0.03039 0.0288 0.0195 1.0000
15.000 1.3738 0.03931 0.03384 0.0280 0.0191 1.0000
15.250 1.3639 0.04350 0.03817 0.0265 0.0189 1.0000
15.500 1.3507 0.04860 0.04342 0.0240 0.0186 1.0000
15.750 1.3370 0.05414 0.04912 0.0211 0.0185 1.0000
16.000 1.3173 0.06088 0.05602 0.0174 0.0185 1.0000
16.250 1.2934 0.06837 0.06366 0.0135 0.0184 1.0000
16.500 1.2703 0.07564 0.07108 0.0099 0.0183 1.0000
16.750 1.2454 0.08319 0.07874 0.0061 0.0184 1.0000
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Polar data table (+)
Polar graphs
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