NACA M23 AIRFOIL (m23-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M23 AIRFOIL (m23-il) Reynolds number: 1,000,000 Max Cl/Cd: 117.87 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m23-il-1000000-n5.txt Download as CSV file: xf-m23-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M23 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4139 0.10073 0.09802 0.0213 0.5754 0.0045
-7.750 -0.4100 0.09733 0.09461 0.0189 0.5718 0.0046
-7.500 -0.4052 0.09386 0.09113 0.0163 0.5683 0.0048
-7.250 -0.3958 0.08988 0.08712 0.0130 0.5650 0.0050
-7.000 -0.3852 0.08542 0.08265 0.0093 0.5619 0.0053
-6.750 -0.3749 0.07905 0.07624 0.0043 0.5591 0.0058
-6.500 -0.3583 0.07558 0.07272 0.0014 0.5551 0.0060
-6.250 -0.3397 0.07291 0.07000 -0.0010 0.5510 0.0062
-6.000 -0.3204 0.07005 0.06711 -0.0034 0.5477 0.0065
-5.750 -0.3002 0.06655 0.06354 -0.0060 0.5442 0.0070
-5.500 -0.2791 0.05946 0.05633 -0.0105 0.5414 0.0081
-5.250 -0.2561 0.05689 0.05369 -0.0123 0.5372 0.0084
-5.000 -0.2325 0.05470 0.05142 -0.0139 0.5335 0.0087
-4.750 -0.2084 0.05233 0.04898 -0.0153 0.5297 0.0092
-4.500 -0.1820 0.04592 0.04237 -0.0179 0.5268 0.0111
-4.250 -0.1568 0.04432 0.04069 -0.0189 0.5228 0.0113
-4.000 -0.1313 0.04270 0.03899 -0.0197 0.5191 0.0117
-3.750 -0.1054 0.04085 0.03706 -0.0205 0.5153 0.0123
-3.500 -0.0760 0.03468 0.03059 -0.0214 0.5123 0.0148
-3.250 -0.0499 0.03351 0.02933 -0.0218 0.5085 0.0150
-3.000 -0.0236 0.03256 0.02831 -0.0223 0.5049 0.0154
-2.750 0.0031 0.03132 0.02699 -0.0226 0.5012 0.0159
-2.500 0.0306 0.02939 0.02493 -0.0229 0.4973 0.0167
-2.250 0.0592 0.02636 0.02166 -0.0227 0.4938 0.0179
-2.000 0.0883 0.02148 0.01638 -0.0219 0.4911 0.0196
-1.750 0.1154 0.01983 0.01457 -0.0218 0.4881 0.0199
-1.500 0.1429 0.01823 0.01280 -0.0217 0.4843 0.0202
-1.250 0.1706 0.01702 0.01144 -0.0217 0.4803 0.0205
-0.750 0.2264 0.01427 0.00828 -0.0215 0.4738 0.0215
-0.500 0.2547 0.01331 0.00714 -0.0215 0.4701 0.0221
-0.250 0.2831 0.01262 0.00631 -0.0215 0.4661 0.0227
0.000 0.3115 0.01207 0.00562 -0.0217 0.4622 0.0231
0.250 0.3400 0.01158 0.00503 -0.0218 0.4591 0.0235
0.500 0.3686 0.01115 0.00452 -0.0219 0.4556 0.0236
0.750 0.3970 0.01077 0.00406 -0.0221 0.4515 0.0238
1.000 0.4254 0.01049 0.00371 -0.0223 0.4475 0.0240
1.250 0.4538 0.01027 0.00346 -0.0224 0.4444 0.0243
1.500 0.4820 0.00998 0.00313 -0.0226 0.4409 0.0244
1.750 0.5100 0.00973 0.00285 -0.0227 0.4370 0.0244
2.000 0.5380 0.00954 0.00262 -0.0228 0.4330 0.0245
2.250 0.5660 0.00937 0.00244 -0.0229 0.4297 0.0246
2.500 0.5941 0.00924 0.00232 -0.0231 0.4262 0.0249
2.750 0.6222 0.00917 0.00224 -0.0233 0.4224 0.0252
3.000 0.6505 0.00916 0.00221 -0.0235 0.4184 0.0255
3.250 0.6784 0.00902 0.00207 -0.0237 0.4150 0.0268
3.500 0.7066 0.00897 0.00202 -0.0239 0.4113 0.0279
3.750 0.7350 0.00897 0.00203 -0.0242 0.4076 0.0288
4.000 0.7633 0.00901 0.00207 -0.0245 0.4036 0.0295
4.250 0.7917 0.00902 0.00210 -0.0248 0.4002 0.0301
4.500 0.8201 0.00905 0.00215 -0.0251 0.3963 0.0313
4.750 0.8484 0.00911 0.00222 -0.0255 0.3925 0.0323
5.000 0.8767 0.00919 0.00229 -0.0258 0.3848 0.0328
5.250 0.9048 0.00932 0.00239 -0.0262 0.3727 0.0337
5.750 0.9606 0.00951 0.00271 -0.0269 0.3505 0.1453
6.250 1.0184 0.00864 0.00338 -0.0286 0.3090 0.9900
6.500 1.0526 0.00976 0.00408 -0.0314 0.2355 0.9939
6.750 1.0905 0.01342 0.00658 -0.0377 0.0163 0.9975
7.000 1.1302 0.01382 0.00702 -0.0410 0.0128 0.9993
7.250 1.1579 0.01419 0.00741 -0.0416 0.0112 0.9997
7.500 1.1836 0.01464 0.00789 -0.0419 0.0097 1.0000
7.750 1.2053 0.01510 0.00839 -0.0414 0.0087 1.0000
8.000 1.2268 0.01550 0.00882 -0.0408 0.0081 1.0000
8.250 1.2477 0.01595 0.00929 -0.0402 0.0074 1.0000
8.500 1.2676 0.01645 0.00981 -0.0394 0.0068 1.0000
8.750 1.2855 0.01710 0.01050 -0.0385 0.0062 1.0000
9.000 1.3026 0.01776 0.01120 -0.0374 0.0058 1.0000
9.250 1.3190 0.01840 0.01191 -0.0363 0.0056 1.0000
9.500 1.3335 0.01914 0.01270 -0.0350 0.0053 1.0000
9.750 1.3459 0.01998 0.01359 -0.0336 0.0050 1.0000
10.000 1.3554 0.02107 0.01474 -0.0325 0.0048 1.0000
10.250 1.3563 0.02251 0.01624 -0.0304 0.0046 1.0000
10.500 1.3632 0.02403 0.01779 -0.0296 0.0044 1.0000
10.750 1.3689 0.02590 0.01972 -0.0291 0.0041 1.0000
11.000 1.3692 0.02843 0.02236 -0.0286 0.0039 1.0000
11.250 1.3745 0.03049 0.02449 -0.0283 0.0039 1.0000
11.500 1.3784 0.03272 0.02681 -0.0280 0.0038 1.0000
11.750 1.3808 0.03512 0.02928 -0.0277 0.0037 1.0000
12.000 1.3820 0.03767 0.03191 -0.0274 0.0036 1.0000
12.250 1.3821 0.04033 0.03466 -0.0271 0.0035 1.0000
12.500 1.3813 0.04313 0.03754 -0.0269 0.0034 1.0000
12.750 1.3799 0.04599 0.04048 -0.0266 0.0033 1.0000
13.000 1.3776 0.04901 0.04358 -0.0265 0.0033 1.0000
13.250 1.3765 0.05198 0.04662 -0.0264 0.0032 1.0000
13.500 1.3748 0.05510 0.04982 -0.0265 0.0031 1.0000
13.750 1.3740 0.05818 0.05299 -0.0266 0.0030 1.0000
14.000 1.3730 0.06132 0.05620 -0.0268 0.0030 1.0000
14.250 1.3716 0.06455 0.05950 -0.0270 0.0029 1.0000
14.500 1.3711 0.06770 0.06273 -0.0272 0.0028 1.0000
14.750 1.3706 0.07090 0.06599 -0.0276 0.0028 1.0000
15.000 1.3691 0.07423 0.06939 -0.0279 0.0027 1.0000
15.250 1.3681 0.07754 0.07276 -0.0283 0.0027 1.0000
15.500 1.3660 0.08104 0.07634 -0.0288 0.0026 1.0000
15.750 1.3635 0.08456 0.07993 -0.0292 0.0026 1.0000
16.000 1.3594 0.08836 0.08381 -0.0298 0.0025 1.0000
16.250 1.3536 0.09235 0.08789 -0.0303 0.0024 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M23 AIRFOIL (m23-il)