NACA 66 (p51htip-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 66 (p51htip-il) Reynolds number: 500,000 Max Cl/Cd: 67.04 at α=3° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-p51htip-il-500000-n5.txt Download as CSV file: xf-p51htip-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 66
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
------ -------- --------- --------- -------- -------- --------
-11.500 -0.5422 0.08222 0.07978 -0.0543 0.9100 0.0095
-11.250 -0.5993 0.06502 0.06245 -0.0663 0.8961 0.0093
-11.000 -0.6416 0.05720 0.05442 -0.0681 0.8848 0.0092
-10.750 -0.6715 0.05222 0.04923 -0.0666 0.8751 0.0093
-10.500 -0.7028 0.04734 0.04409 -0.0630 0.8662 0.0094
-10.000 -0.7577 0.03583 0.03173 -0.0530 0.8522 0.0098
-9.750 -0.7715 0.02984 0.02507 -0.0486 0.8467 0.0104
-9.500 -0.7658 0.02650 0.02123 -0.0462 0.8427 0.0108
-9.250 -0.7520 0.02421 0.01857 -0.0446 0.8386 0.0111
-9.000 -0.7342 0.02261 0.01668 -0.0435 0.8350 0.0113
-8.750 -0.7156 0.02122 0.01510 -0.0426 0.8318 0.0116
-8.250 -0.6719 0.01977 0.01343 -0.0415 0.8261 0.0122
-8.000 -0.6490 0.01901 0.01257 -0.0410 0.8231 0.0125
-7.750 -0.6260 0.01827 0.01172 -0.0405 0.8202 0.0129
-7.500 -0.6027 0.01754 0.01087 -0.0401 0.8176 0.0134
-7.250 -0.5790 0.01691 0.01012 -0.0396 0.8153 0.0140
-7.000 -0.5548 0.01636 0.00945 -0.0392 0.8130 0.0147
-6.750 -0.5313 0.01565 0.00865 -0.0388 0.8105 0.0151
-6.500 -0.5084 0.01497 0.00794 -0.0383 0.8079 0.0159
-6.250 -0.4841 0.01457 0.00751 -0.0379 0.8054 0.0165
-6.000 -0.4598 0.01416 0.00704 -0.0375 0.8031 0.0173
-5.750 -0.4356 0.01375 0.00656 -0.0371 0.8009 0.0182
-5.500 -0.4112 0.01336 0.00611 -0.0367 0.7990 0.0190
-5.250 -0.3860 0.01305 0.00575 -0.0365 0.7969 0.0197
-5.000 -0.3627 0.01253 0.00522 -0.0359 0.7948 0.0211
-4.750 -0.3374 0.01224 0.00492 -0.0357 0.7927 0.0224
-4.500 -0.3120 0.01196 0.00462 -0.0355 0.7908 0.0239
-4.250 -0.2863 0.01171 0.00433 -0.0353 0.7890 0.0253
-4.000 -0.2604 0.01149 0.00407 -0.0352 0.7874 0.0263
-3.750 -0.2359 0.01111 0.00364 -0.0347 0.7858 0.0284
-3.500 -0.2101 0.01087 0.00338 -0.0346 0.7843 0.0306
-3.250 -0.1837 0.01067 0.00317 -0.0345 0.7825 0.0330
-3.000 -0.1570 0.01050 0.00299 -0.0345 0.7806 0.0357
-2.750 -0.1308 0.01028 0.00279 -0.0344 0.7787 0.0434
-2.500 -0.1049 0.01003 0.00262 -0.0343 0.7771 0.0645
-2.250 -0.0823 0.00944 0.00241 -0.0337 0.7754 0.1680
-2.000 -0.0625 0.00859 0.00217 -0.0329 0.7738 0.3366
-1.750 -0.0500 0.00723 0.00189 -0.0307 0.7722 0.6119
-1.500 -0.0272 0.00700 0.00202 -0.0296 0.7709 0.7261
-1.250 -0.0004 0.00702 0.00208 -0.0293 0.7696 0.7547
-1.000 0.0267 0.00705 0.00217 -0.0291 0.7679 0.7756
-0.750 0.0542 0.00709 0.00224 -0.0291 0.7663 0.7893
-0.500 0.0818 0.00714 0.00231 -0.0291 0.7648 0.8004
-0.250 0.1102 0.00717 0.00233 -0.0293 0.7634 0.8052
0.000 0.1389 0.00718 0.00232 -0.0296 0.7619 0.8080
0.250 0.1677 0.00719 0.00231 -0.0300 0.7603 0.8107
0.500 0.1963 0.00720 0.00233 -0.0303 0.7588 0.8129
0.750 0.2249 0.00721 0.00234 -0.0306 0.7571 0.8151
1.000 0.2535 0.00723 0.00236 -0.0309 0.7549 0.8173
1.250 0.2817 0.00723 0.00240 -0.0312 0.7516 0.8198
1.500 0.3101 0.00723 0.00242 -0.0314 0.7480 0.8225
1.750 0.3385 0.00723 0.00243 -0.0317 0.7443 0.8255
2.000 0.3666 0.00721 0.00242 -0.0318 0.7389 0.8279
2.250 0.3939 0.00714 0.00238 -0.0317 0.7283 0.8303
2.500 0.4210 0.00709 0.00235 -0.0315 0.7148 0.8329
2.750 0.4481 0.00707 0.00233 -0.0314 0.6994 0.8359
3.000 0.4733 0.00706 0.00224 -0.0308 0.6568 0.8391
3.250 0.4805 0.00812 0.00245 -0.0271 0.4716 0.8430
3.500 0.4898 0.00923 0.00296 -0.0242 0.3299 0.8471
3.750 0.5024 0.01014 0.00339 -0.0219 0.2156 0.8518
4.250 0.5332 0.01160 0.00417 -0.0183 0.0568 0.8604
4.500 0.5540 0.01197 0.00447 -0.0172 0.0387 0.8646
4.750 0.5764 0.01225 0.00476 -0.0165 0.0332 0.8690
5.000 0.5977 0.01257 0.00509 -0.0155 0.0287 0.8732
5.250 0.6194 0.01283 0.00540 -0.0145 0.0271 0.8776
5.500 0.6405 0.01312 0.00574 -0.0135 0.0253 0.8827
5.750 0.6609 0.01345 0.00610 -0.0124 0.0236 0.8879
6.000 0.6772 0.01391 0.00660 -0.0105 0.0213 0.8939
6.250 0.6957 0.01420 0.00694 -0.0090 0.0206 0.9006
6.500 0.7125 0.01453 0.00735 -0.0072 0.0198 0.9076
6.750 0.7292 0.01495 0.00782 -0.0054 0.0189 0.9158
7.000 0.7454 0.01535 0.00829 -0.0036 0.0181 0.9244
7.250 0.7623 0.01578 0.00877 -0.0019 0.0174 0.9342
7.500 0.7790 0.01629 0.00933 -0.0004 0.0167 0.9460
7.750 0.7960 0.01716 0.01025 0.0008 0.0159 0.9599
8.000 0.8228 0.01768 0.01086 0.0002 0.0154 0.9721
8.250 0.8512 0.01833 0.01158 -0.0009 0.0148 0.9875
8.500 0.8716 0.01900 0.01230 -0.0004 0.0143 1.0000
8.750 0.8891 0.01967 0.01303 0.0007 0.0138 1.0000
9.000 0.9068 0.02038 0.01380 0.0018 0.0135 1.0000
9.250 0.9244 0.02111 0.01458 0.0029 0.0131 1.0000
9.500 0.9421 0.02179 0.01530 0.0039 0.0128 1.0000
9.750 0.9592 0.02257 0.01612 0.0049 0.0124 1.0000
10.000 0.9756 0.02357 0.01717 0.0060 0.0121 1.0000
10.250 0.9923 0.02482 0.01853 0.0070 0.0118 1.0000
10.500 1.0095 0.02584 0.01968 0.0080 0.0116 1.0000
10.750 1.0261 0.02698 0.02096 0.0090 0.0115 1.0000
11.000 1.0417 0.02806 0.02219 0.0101 0.0111 1.0000
11.250 1.0560 0.02919 0.02346 0.0112 0.0107 1.0000
11.500 1.0688 0.03058 0.02501 0.0125 0.0105 1.0000
11.750 1.0796 0.03210 0.02670 0.0139 0.0102 1.0000
12.000 1.0889 0.03360 0.02838 0.0153 0.0100 1.0000
12.250 1.0960 0.03534 0.03029 0.0169 0.0098 1.0000
12.500 1.1019 0.03694 0.03203 0.0184 0.0097 1.0000
12.750 1.1057 0.03878 0.03404 0.0199 0.0095 1.0000
13.000 1.1085 0.04064 0.03603 0.0213 0.0094 1.0000
13.250 1.1087 0.04280 0.03835 0.0227 0.0093 1.0000
13.500 1.1067 0.04521 0.04092 0.0240 0.0092 1.0000
13.750 1.1058 0.04745 0.04329 0.0250 0.0091 1.0000
14.000 1.1008 0.05023 0.04622 0.0258 0.0090 1.0000
14.250 1.0916 0.05369 0.04986 0.0264 0.0089 1.0000
14.500 1.0781 0.05780 0.05415 0.0266 0.0089 1.0000
14.750 1.0626 0.06243 0.05896 0.0260 0.0088 1.0000
15.000 1.0432 0.06811 0.06486 0.0246 0.0088 1.0000
15.250 1.0176 0.07554 0.07253 0.0216 0.0088 1.0000
15.500 0.9960 0.08318 0.08034 0.0175 0.0088 1.0000
15.750 0.9539 0.09685 0.09431 0.0090 0.0089 1.0000
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