RAF 27 AIRFOIL (raf27-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAF 27 AIRFOIL (raf27-il) Reynolds number: 200,000 Max Cl/Cd: 43.88 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf27-il-200000-n5.txt Download as CSV file: xf-raf27-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAF 27 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.7880 0.08147 0.07778 -0.0188 1.0000 0.0119
-12.250 -0.8234 0.06914 0.06527 -0.0282 1.0000 0.0116
-12.000 -0.8509 0.06072 0.05663 -0.0343 1.0000 0.0113
-11.750 -0.8752 0.05442 0.05010 -0.0376 1.0000 0.0112
-11.500 -0.8936 0.04994 0.04541 -0.0386 1.0000 0.0113
-11.250 -0.9097 0.04633 0.04157 -0.0379 1.0000 0.0113
-11.000 -0.9233 0.04333 0.03833 -0.0356 1.0000 0.0115
-10.750 -0.9344 0.04072 0.03547 -0.0323 1.0000 0.0116
-10.500 -0.9404 0.03825 0.03273 -0.0290 1.0000 0.0118
-10.250 -0.9402 0.03581 0.03000 -0.0262 1.0000 0.0121
-10.000 -0.9351 0.03370 0.02758 -0.0238 1.0000 0.0127
-9.750 -0.9263 0.03184 0.02543 -0.0216 1.0000 0.0132
-9.500 -0.9150 0.03018 0.02346 -0.0195 1.0000 0.0138
-9.250 -0.9057 0.02806 0.02115 -0.0173 1.0000 0.0147
-9.000 -0.8916 0.02674 0.01970 -0.0155 1.0000 0.0154
-8.750 -0.8760 0.02556 0.01839 -0.0138 1.0000 0.0162
-8.500 -0.8599 0.02435 0.01704 -0.0121 1.0000 0.0171
-8.250 -0.8430 0.02327 0.01580 -0.0104 1.0000 0.0184
-8.000 -0.8248 0.02238 0.01475 -0.0088 1.0000 0.0200
-7.750 -0.8090 0.02118 0.01345 -0.0069 1.0000 0.0215
-7.500 -0.7922 0.02023 0.01245 -0.0052 1.0000 0.0233
-7.250 -0.7740 0.01946 0.01157 -0.0036 1.0000 0.0257
-7.000 -0.7544 0.01886 0.01083 -0.0021 1.0000 0.0288
-6.750 -0.7380 0.01791 0.00985 -0.0002 1.0000 0.0316
-6.500 -0.7195 0.01724 0.00913 0.0014 1.0000 0.0351
-6.250 -0.6996 0.01673 0.00852 0.0028 1.0000 0.0389
-6.000 -0.6818 0.01602 0.00782 0.0045 1.0000 0.0439
-5.750 -0.6617 0.01556 0.00727 0.0059 1.0000 0.0493
-5.500 -0.6426 0.01501 0.00674 0.0074 1.0000 0.0561
-5.250 -0.6227 0.01456 0.00627 0.0089 1.0000 0.0652
-5.000 -0.6030 0.01409 0.00581 0.0104 1.0000 0.0778
-4.750 -0.5838 0.01360 0.00537 0.0119 1.0000 0.0952
-4.500 -0.5649 0.01309 0.00497 0.0134 1.0000 0.1216
-4.250 -0.5470 0.01251 0.00464 0.0150 1.0000 0.1727
-4.000 -0.5285 0.01204 0.00438 0.0165 1.0000 0.2225
-3.750 -0.5025 0.01165 0.00414 0.0165 0.9979 0.2671
-3.500 -0.4666 0.01124 0.00390 0.0144 0.9915 0.3203
-3.250 -0.4318 0.01080 0.00372 0.0125 0.9854 0.3841
-3.000 -0.4009 0.01038 0.00356 0.0116 0.9780 0.4516
-2.750 -0.3698 0.01004 0.00342 0.0108 0.9707 0.5096
-2.500 -0.3378 0.00971 0.00333 0.0099 0.9640 0.5663
-2.250 -0.3085 0.00945 0.00326 0.0097 0.9551 0.6158
-2.000 -0.2747 0.00927 0.00318 0.0086 0.9487 0.6576
-1.750 -0.2434 0.00911 0.00310 0.0081 0.9392 0.6877
-1.500 -0.2096 0.00898 0.00301 0.0070 0.9308 0.7122
-1.250 -0.1747 0.00885 0.00294 0.0057 0.9220 0.7355
-1.000 -0.1414 0.00874 0.00288 0.0048 0.9111 0.7583
-0.750 -0.1066 0.00864 0.00283 0.0036 0.8995 0.7808
-0.250 -0.0354 0.00849 0.00276 0.0011 0.8679 0.8269
0.000 0.0000 0.00847 0.00274 0.0000 0.8486 0.8487
0.250 0.0354 0.00849 0.00276 -0.0011 0.8269 0.8679
0.500 0.0714 0.00855 0.00278 -0.0024 0.8044 0.8853
0.750 0.1066 0.00864 0.00283 -0.0036 0.7809 0.8995
1.000 0.1414 0.00874 0.00288 -0.0048 0.7584 0.9111
1.250 0.1747 0.00885 0.00294 -0.0057 0.7357 0.9221
1.500 0.2095 0.00898 0.00301 -0.0070 0.7123 0.9308
1.750 0.2433 0.00911 0.00310 -0.0080 0.6875 0.9393
2.000 0.2747 0.00927 0.00318 -0.0086 0.6574 0.9487
2.250 0.3085 0.00945 0.00326 -0.0097 0.6159 0.9552
2.500 0.3378 0.00971 0.00333 -0.0099 0.5660 0.9640
2.750 0.3699 0.01003 0.00342 -0.0108 0.5107 0.9708
3.000 0.4012 0.01036 0.00357 -0.0117 0.4544 0.9780
3.250 0.4320 0.01078 0.00372 -0.0126 0.3875 0.9854
3.500 0.4667 0.01123 0.00390 -0.0144 0.3206 0.9916
3.750 0.5026 0.01165 0.00414 -0.0165 0.2673 0.9980
4.000 0.5283 0.01204 0.00437 -0.0165 0.2224 1.0000
4.250 0.5468 0.01251 0.00464 -0.0149 0.1717 1.0000
4.500 0.5647 0.01308 0.00496 -0.0134 0.1219 1.0000
4.750 0.5836 0.01360 0.00537 -0.0118 0.0949 1.0000
5.000 0.6028 0.01409 0.00581 -0.0103 0.0776 1.0000
5.250 0.6225 0.01455 0.00627 -0.0089 0.0654 1.0000
5.500 0.6424 0.01501 0.00674 -0.0074 0.0562 1.0000
5.750 0.6615 0.01556 0.00726 -0.0059 0.0493 1.0000
6.000 0.6815 0.01602 0.00782 -0.0045 0.0438 1.0000
6.250 0.6993 0.01674 0.00852 -0.0028 0.0390 1.0000
6.500 0.7193 0.01724 0.00913 -0.0013 0.0352 1.0000
6.750 0.7379 0.01790 0.00984 0.0002 0.0316 1.0000
7.000 0.7546 0.01881 0.01080 0.0021 0.0285 1.0000
7.250 0.7739 0.01946 0.01157 0.0036 0.0256 1.0000
7.500 0.7922 0.02023 0.01244 0.0052 0.0233 1.0000
7.750 0.8090 0.02117 0.01344 0.0069 0.0215 1.0000
8.000 0.8246 0.02241 0.01478 0.0089 0.0200 1.0000
8.250 0.8429 0.02330 0.01583 0.0104 0.0185 1.0000
8.500 0.8600 0.02436 0.01704 0.0121 0.0171 1.0000
8.750 0.8761 0.02555 0.01838 0.0138 0.0161 1.0000
9.000 0.8917 0.02676 0.01972 0.0155 0.0154 1.0000
9.250 0.9059 0.02806 0.02115 0.0173 0.0147 1.0000
9.500 0.9159 0.03007 0.02334 0.0194 0.0139 1.0000
9.750 0.9261 0.03194 0.02549 0.0215 0.0133 1.0000
10.000 0.9355 0.03369 0.02757 0.0237 0.0127 1.0000
10.250 0.9407 0.03581 0.02999 0.0261 0.0122 1.0000
10.500 0.9409 0.03826 0.03274 0.0289 0.0119 1.0000
10.750 0.9346 0.04081 0.03557 0.0322 0.0117 1.0000
11.000 0.9242 0.04333 0.03833 0.0355 0.0115 1.0000
11.250 0.9106 0.04634 0.04158 0.0377 0.0114 1.0000
11.500 0.8946 0.04995 0.04542 0.0385 0.0113 1.0000
11.750 0.8728 0.05493 0.05064 0.0373 0.0115 1.0000
12.000 0.8521 0.06071 0.05662 0.0341 0.0113 1.0000
12.250 0.8243 0.06920 0.06533 0.0279 0.0116 1.0000
12.500 0.7887 0.08168 0.07799 0.0185 0.0119 1.0000
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