NACA 64-110 AIRFOIL (n64110-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 64-110 AIRFOIL (n64110-il) Reynolds number: 1,000,000 Max Cl/Cd: 69.3 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64110-il-1000000.txt Download as CSV file: xf-n64110-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-110 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.8064 0.05192 0.05001 -0.0349 1.0000 0.0083
-10.500 -0.8544 0.04488 0.04271 -0.0339 1.0000 0.0082
-10.250 -0.9124 0.03462 0.03179 -0.0293 1.0000 0.0082
-10.000 -0.9151 0.03036 0.02711 -0.0273 1.0000 0.0085
-9.750 -0.8997 0.02888 0.02543 -0.0264 1.0000 0.0087
-9.500 -0.9026 0.02338 0.01932 -0.0242 1.0000 0.0092
-9.250 -0.8820 0.02257 0.01844 -0.0237 1.0000 0.0096
-9.000 -0.8593 0.02213 0.01796 -0.0234 1.0000 0.0099
-8.750 -0.8381 0.02112 0.01682 -0.0229 1.0000 0.0102
-8.500 -0.8164 0.02005 0.01561 -0.0223 1.0000 0.0106
-8.250 -0.7949 0.01878 0.01419 -0.0217 1.0000 0.0109
-8.000 -0.7725 0.01768 0.01295 -0.0211 1.0000 0.0114
-7.750 -0.7494 0.01679 0.01195 -0.0205 1.0000 0.0118
-7.500 -0.7256 0.01618 0.01125 -0.0201 1.0000 0.0121
-7.250 -0.7063 0.01451 0.00941 -0.0189 1.0000 0.0128
-7.000 -0.6796 0.01366 0.00853 -0.0193 0.9960 0.0136
-6.750 -0.6458 0.01310 0.00793 -0.0210 0.9873 0.0142
-6.500 -0.6120 0.01246 0.00725 -0.0226 0.9782 0.0149
-6.250 -0.5794 0.01189 0.00660 -0.0239 0.9672 0.0156
-6.000 -0.5515 0.01141 0.00605 -0.0241 0.9528 0.0162
-5.750 -0.5261 0.01113 0.00570 -0.0237 0.9382 0.0168
-5.500 -0.5041 0.01032 0.00477 -0.0227 0.9239 0.0180
-5.250 -0.4796 0.00993 0.00432 -0.0221 0.9116 0.0193
-5.000 -0.4542 0.00963 0.00396 -0.0217 0.8999 0.0204
-4.750 -0.4281 0.00937 0.00364 -0.0214 0.8888 0.0218
-4.500 -0.4016 0.00918 0.00339 -0.0212 0.8783 0.0230
-4.250 -0.3758 0.00876 0.00287 -0.0208 0.8681 0.0262
-4.000 -0.3490 0.00855 0.00263 -0.0207 0.8580 0.0298
-3.750 -0.3219 0.00834 0.00238 -0.0206 0.8484 0.0336
-3.500 -0.2950 0.00811 0.00213 -0.0204 0.8391 0.0414
-3.250 -0.2681 0.00781 0.00190 -0.0204 0.8292 0.0663
-3.000 -0.2424 0.00718 0.00162 -0.0203 0.8196 0.1652
-2.750 -0.2184 0.00618 0.00128 -0.0203 0.8104 0.3500
-2.500 -0.1928 0.00556 0.00112 -0.0203 0.8009 0.4863
-2.250 -0.1653 0.00538 0.00105 -0.0203 0.7921 0.5358
-2.000 -0.1375 0.00530 0.00100 -0.0203 0.7831 0.5658
-1.750 -0.1094 0.00522 0.00095 -0.0204 0.7738 0.5872
-1.500 -0.0814 0.00518 0.00092 -0.0204 0.7648 0.6086
-1.250 -0.0536 0.00511 0.00090 -0.0204 0.7554 0.6367
-1.000 -0.0257 0.00506 0.00091 -0.0204 0.7465 0.6669
-0.750 0.0024 0.00505 0.00090 -0.0204 0.7379 0.6810
-0.500 0.0308 0.00503 0.00090 -0.0205 0.7290 0.6917
-0.250 0.0591 0.00504 0.00089 -0.0206 0.7205 0.7015
0.000 0.0875 0.00505 0.00089 -0.0207 0.7116 0.7107
0.250 0.1158 0.00504 0.00090 -0.0208 0.7029 0.7195
0.500 0.1441 0.00507 0.00091 -0.0209 0.6941 0.7286
0.750 0.1724 0.00508 0.00093 -0.0209 0.6837 0.7378
1.000 0.2007 0.00509 0.00096 -0.0210 0.6734 0.7466
1.500 0.2563 0.00519 0.00100 -0.0210 0.6335 0.7652
1.750 0.2839 0.00528 0.00103 -0.0209 0.6067 0.7746
2.000 0.3119 0.00535 0.00107 -0.0209 0.5887 0.7844
2.250 0.3394 0.00544 0.00114 -0.0208 0.5654 0.7937
2.500 0.3671 0.00553 0.00121 -0.0208 0.5416 0.8029
2.750 0.3943 0.00569 0.00129 -0.0208 0.5074 0.8126
3.000 0.4195 0.00611 0.00144 -0.0205 0.4189 0.8222
3.250 0.4414 0.00716 0.00182 -0.0199 0.2452 0.8324
3.500 0.4647 0.00800 0.00218 -0.0196 0.1232 0.8430
3.750 0.4891 0.00857 0.00250 -0.0192 0.0552 0.8534
4.000 0.5151 0.00885 0.00275 -0.0189 0.0396 0.8637
4.250 0.5413 0.00911 0.00303 -0.0186 0.0325 0.8745
4.750 0.5921 0.00969 0.00371 -0.0178 0.0240 0.8984
5.000 0.6173 0.00990 0.00399 -0.0172 0.0221 0.9111
5.250 0.6416 0.01015 0.00428 -0.0165 0.0202 0.9250
5.500 0.6627 0.01077 0.00501 -0.0153 0.0179 0.9417
5.750 0.6868 0.01099 0.00529 -0.0145 0.0173 0.9614
6.000 0.7186 0.01137 0.00573 -0.0156 0.0165 0.9846
6.250 0.7472 0.01184 0.00624 -0.0161 0.0157 1.0000
6.500 0.7727 0.01233 0.00676 -0.0159 0.0151 1.0000
6.750 0.7979 0.01284 0.00729 -0.0157 0.0145 1.0000
7.000 0.8222 0.01345 0.00793 -0.0153 0.0139 1.0000
7.250 0.8406 0.01521 0.00983 -0.0141 0.0129 1.0000
7.500 0.8662 0.01560 0.01026 -0.0140 0.0126 1.0000
7.750 0.8904 0.01628 0.01102 -0.0136 0.0123 1.0000
8.000 0.9141 0.01707 0.01189 -0.0131 0.0119 1.0000
8.250 0.9371 0.01804 0.01295 -0.0126 0.0115 1.0000
8.500 0.9599 0.01898 0.01400 -0.0121 0.0111 1.0000
8.750 0.9825 0.01991 0.01503 -0.0115 0.0108 1.0000
9.000 1.0050 0.02066 0.01585 -0.0111 0.0104 1.0000
9.250 1.0277 0.02120 0.01643 -0.0107 0.0101 1.0000
9.500 1.0478 0.02228 0.01759 -0.0100 0.0097 1.0000
9.750 1.0534 0.02672 0.02250 -0.0079 0.0091 1.0000
10.000 1.0731 0.02744 0.02334 -0.0071 0.0089 1.0000
10.250 1.0878 0.02906 0.02518 -0.0059 0.0087 1.0000
10.500 1.0971 0.03140 0.02779 -0.0042 0.0085 1.0000
10.750 1.1002 0.03427 0.03095 -0.0021 0.0083 1.0000
11.000 1.0940 0.03784 0.03485 0.0007 0.0081 1.0000
11.250 1.0633 0.04257 0.03992 0.0057 0.0081 1.0000
11.500 1.0310 0.04764 0.04528 0.0084 0.0080 1.0000
11.750 0.9878 0.05492 0.05285 0.0084 0.0081 1.0000
12.000 0.9460 0.06347 0.06162 0.0052 0.0082 1.0000
12.250 0.9114 0.07266 0.07098 -0.0007 0.0085 1.0000
12.500 0.7249 0.10565 0.10430 -0.0224 0.0099 1.0000
12.750 0.7364 0.10666 0.10533 -0.0235 0.0096 1.0000
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Polar data table (+)
Polar graphs
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