BOEING-VERTOL V23010-1.58 AIRFOIL (v23010-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL V23010-1.58 AIRFOIL (v23010-il) Reynolds number: 1,000,000 Max Cl/Cd: 94.11 at α=8.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-v23010-il-1000000-n5.txt Download as CSV file: xf-v23010-il-1000000-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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5687 0.08494 0.08315 -0.0126 1.0000 0.0139
-9.750 -0.5770 0.07870 0.07693 -0.0170 1.0000 0.0140
-9.500 -0.6168 0.06614 0.06432 -0.0300 1.0000 0.0139
-9.250 -0.6406 0.06186 0.05996 -0.0298 1.0000 0.0140
-9.000 -0.6550 0.05712 0.05509 -0.0296 1.0000 0.0140
-8.750 -0.6616 0.05293 0.05077 -0.0288 1.0000 0.0141
-8.500 -0.6616 0.04942 0.04712 -0.0277 1.0000 0.0142
-8.250 -0.6579 0.04621 0.04378 -0.0263 1.0000 0.0144
-8.000 -0.6523 0.04308 0.04049 -0.0247 1.0000 0.0147
-7.750 -0.6491 0.03909 0.03629 -0.0225 1.0000 0.0150
-7.500 -0.6807 0.02255 0.01841 -0.0159 0.9982 0.0161
-7.250 -0.6571 0.01990 0.01544 -0.0161 0.9932 0.0164
-7.000 -0.6283 0.01871 0.01412 -0.0167 0.9875 0.0165
-6.750 -0.5944 0.01761 0.01291 -0.0184 0.9802 0.0167
-6.500 -0.5575 0.01664 0.01183 -0.0206 0.9677 0.0168
-6.250 -0.5213 0.01579 0.01087 -0.0226 0.9495 0.0170
-6.000 -0.4900 0.01508 0.00999 -0.0233 0.9111 0.0171
-5.750 -0.4664 0.01459 0.00928 -0.0223 0.8674 0.0173
-5.500 -0.4431 0.01405 0.00856 -0.0212 0.8337 0.0174
-5.250 -0.4206 0.01362 0.00788 -0.0200 0.7794 0.0176
-5.000 -0.3967 0.01319 0.00726 -0.0191 0.7502 0.0178
-4.750 -0.3716 0.01274 0.00670 -0.0185 0.7325 0.0179
-4.500 -0.3464 0.01226 0.00612 -0.0179 0.7187 0.0181
-4.250 -0.3209 0.01186 0.00564 -0.0173 0.7031 0.0182
-4.000 -0.2955 0.01148 0.00517 -0.0167 0.6838 0.0184
-3.750 -0.2702 0.01115 0.00474 -0.0160 0.6575 0.0185
-3.500 -0.2453 0.01088 0.00433 -0.0153 0.6239 0.0186
-3.250 -0.2200 0.01067 0.00400 -0.0147 0.5939 0.0188
-3.000 -0.1943 0.01052 0.00374 -0.0141 0.5669 0.0190
-2.750 -0.1685 0.01037 0.00348 -0.0136 0.5405 0.0192
-2.500 -0.1427 0.01022 0.00323 -0.0131 0.5110 0.0193
-2.250 -0.1171 0.01013 0.00300 -0.0125 0.4740 0.0194
-2.000 -0.0925 0.00992 0.00263 -0.0119 0.4349 0.0198
-1.750 -0.0670 0.00983 0.00241 -0.0113 0.3986 0.0201
-1.500 -0.0410 0.00977 0.00225 -0.0109 0.3700 0.0204
-1.250 -0.0145 0.00972 0.00212 -0.0105 0.3469 0.0208
-1.000 0.0120 0.00969 0.00201 -0.0101 0.3223 0.0211
-0.750 0.0385 0.00968 0.00190 -0.0097 0.2979 0.0215
-0.500 0.0653 0.00966 0.00181 -0.0094 0.2811 0.0218
-0.250 0.0922 0.00964 0.00174 -0.0091 0.2681 0.0222
0.000 0.1193 0.00962 0.00167 -0.0088 0.2590 0.0226
0.250 0.1465 0.00960 0.00162 -0.0085 0.2510 0.0230
0.500 0.1738 0.00956 0.00157 -0.0082 0.2452 0.0238
0.750 0.2010 0.00954 0.00154 -0.0080 0.2403 0.0252
1.000 0.2284 0.00954 0.00152 -0.0077 0.2360 0.0264
1.250 0.2560 0.00952 0.00151 -0.0075 0.2334 0.0282
1.500 0.2834 0.00950 0.00153 -0.0073 0.2299 0.0365
1.750 0.3105 0.00948 0.00154 -0.0070 0.2254 0.0590
2.000 0.3300 0.00838 0.00148 -0.0059 0.2210 0.4626
2.250 0.3529 0.00792 0.00152 -0.0049 0.2186 0.6223
2.500 0.3772 0.00771 0.00158 -0.0041 0.2158 0.7041
2.750 0.4018 0.00759 0.00165 -0.0032 0.2123 0.7675
3.000 0.4264 0.00751 0.00173 -0.0023 0.2086 0.8190
3.250 0.4504 0.00744 0.00182 -0.0013 0.2053 0.8754
3.500 0.4800 0.00740 0.00195 -0.0013 0.2028 0.9363
3.750 0.5199 0.00749 0.00207 -0.0038 0.1998 0.9614
4.000 0.5543 0.00760 0.00216 -0.0051 0.1960 0.9701
4.250 0.5899 0.00773 0.00225 -0.0067 0.1916 0.9762
4.500 0.6228 0.00786 0.00236 -0.0078 0.1877 0.9821
4.750 0.6558 0.00797 0.00247 -0.0089 0.1840 0.9859
5.000 0.6884 0.00810 0.00258 -0.0099 0.1783 0.9894
5.250 0.7197 0.00825 0.00270 -0.0106 0.1725 0.9929
5.500 0.7521 0.00844 0.00282 -0.0116 0.1608 0.9960
5.750 0.7850 0.00868 0.00297 -0.0129 0.1458 0.9986
6.000 0.8153 0.00895 0.00316 -0.0135 0.1329 1.0000
6.250 0.8388 0.00915 0.00334 -0.0126 0.1268 1.0000
6.500 0.8622 0.00938 0.00354 -0.0117 0.1206 1.0000
6.750 0.8860 0.00958 0.00374 -0.0109 0.1162 1.0000
7.000 0.9097 0.00981 0.00395 -0.0101 0.1117 1.0000
7.250 0.9332 0.01006 0.00418 -0.0093 0.1062 1.0000
7.500 0.9575 0.01025 0.00438 -0.0086 0.1034 1.0000
7.750 0.9815 0.01047 0.00459 -0.0078 0.1003 1.0000
8.000 1.0051 0.01073 0.00484 -0.0071 0.0961 1.0000
8.250 1.0292 0.01095 0.00508 -0.0063 0.0927 1.0000
8.500 1.0531 0.01119 0.00532 -0.0056 0.0878 1.0000
8.750 1.0762 0.01151 0.00560 -0.0048 0.0804 1.0000
9.000 1.0979 0.01198 0.00598 -0.0039 0.0655 1.0000
9.250 1.1182 0.01259 0.00649 -0.0028 0.0500 1.0000
9.500 1.1387 0.01320 0.00702 -0.0017 0.0395 1.0000
9.750 1.1605 0.01365 0.00748 -0.0008 0.0348 1.0000
10.000 1.1825 0.01408 0.00791 0.0001 0.0317 1.0000
10.250 1.2045 0.01451 0.00835 0.0010 0.0295 1.0000
10.500 1.2259 0.01498 0.00883 0.0019 0.0272 1.0000
10.750 1.2480 0.01537 0.00925 0.0027 0.0262 1.0000
11.000 1.2695 0.01580 0.00972 0.0036 0.0250 1.0000
11.250 1.2902 0.01629 0.01023 0.0045 0.0238 1.0000
11.500 1.3103 0.01681 0.01078 0.0055 0.0226 1.0000
11.750 1.3312 0.01724 0.01126 0.0064 0.0219 1.0000
12.000 1.3512 0.01771 0.01178 0.0074 0.0211 1.0000
12.250 1.3704 0.01824 0.01234 0.0085 0.0202 1.0000
12.500 1.3885 0.01882 0.01295 0.0098 0.0194 1.0000
12.750 1.4052 0.01945 0.01363 0.0111 0.0186 1.0000
13.000 1.4228 0.01998 0.01421 0.0124 0.0181 1.0000
13.250 1.4390 0.02055 0.01485 0.0138 0.0174 1.0000
13.500 1.4531 0.02118 0.01552 0.0155 0.0168 1.0000
13.750 1.4630 0.02185 0.01623 0.0179 0.0161 1.0000
14.000 1.4706 0.02264 0.01707 0.0204 0.0155 1.0000
14.250 1.4784 0.02348 0.01797 0.0227 0.0151 1.0000
14.500 1.4866 0.02435 0.01893 0.0246 0.0148 1.0000
14.750 1.4935 0.02539 0.02003 0.0264 0.0144 1.0000
15.000 1.4990 0.02662 0.02133 0.0279 0.0139 1.0000
15.250 1.5032 0.02806 0.02285 0.0291 0.0136 1.0000
15.500 1.5056 0.02981 0.02468 0.0300 0.0132 1.0000
15.750 1.5063 0.03194 0.02689 0.0304 0.0128 1.0000
16.000 1.5046 0.03454 0.02959 0.0302 0.0125 1.0000
16.250 1.5014 0.03755 0.03271 0.0296 0.0122 1.0000
16.500 1.4957 0.04116 0.03644 0.0283 0.0121 1.0000
16.750 1.4876 0.04539 0.04080 0.0264 0.0119 1.0000
17.000 1.4729 0.05086 0.04641 0.0236 0.0119 1.0000
17.250 1.4512 0.05768 0.05339 0.0199 0.0118 1.0000
17.500 1.4196 0.06624 0.06214 0.0154 0.0119 1.0000
17.750 1.3810 0.07581 0.07189 0.0107 0.0120 1.0000
18.000 1.3388 0.08596 0.08219 0.0057 0.0120 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL V23010-1.58 AIRFOIL (v23010-il)