RAE 101 AIRFOIL (rae101-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAE 101 AIRFOIL (rae101-il) Reynolds number: 200,000 Max Cl/Cd: 44.2 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae101-il-200000-n5.txt Download as CSV file: xf-rae101-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAE 101 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
------ -------- --------- --------- -------- -------- --------
-12.250 -0.8145 0.07775 0.07397 -0.0142 1.0000 0.0136
-12.000 -0.8518 0.06635 0.06238 -0.0228 1.0000 0.0133
-11.750 -0.8767 0.05916 0.05497 -0.0277 1.0000 0.0132
-11.500 -0.8993 0.05349 0.04907 -0.0304 1.0000 0.0132
-11.250 -0.9185 0.04910 0.04444 -0.0310 1.0000 0.0132
-11.000 -0.9326 0.04586 0.04098 -0.0298 1.0000 0.0134
-10.750 -0.9448 0.04300 0.03787 -0.0273 1.0000 0.0135
-10.500 -0.9492 0.04045 0.03504 -0.0249 1.0000 0.0139
-10.250 -0.9484 0.03787 0.03214 -0.0228 1.0000 0.0144
-10.000 -0.9436 0.03535 0.02924 -0.0209 1.0000 0.0150
-9.750 -0.9342 0.03320 0.02668 -0.0191 1.0000 0.0155
-9.500 -0.9232 0.03084 0.02407 -0.0175 1.0000 0.0159
-9.250 -0.9088 0.02908 0.02216 -0.0162 1.0000 0.0164
-9.000 -0.8921 0.02767 0.02061 -0.0151 1.0000 0.0170
-8.750 -0.8744 0.02632 0.01912 -0.0139 1.0000 0.0176
-8.500 -0.8559 0.02503 0.01767 -0.0127 1.0000 0.0185
-8.250 -0.8368 0.02380 0.01627 -0.0116 1.0000 0.0194
-8.000 -0.8165 0.02282 0.01510 -0.0105 1.0000 0.0207
-7.750 -0.7982 0.02155 0.01376 -0.0093 1.0000 0.0221
-7.500 -0.7786 0.02061 0.01278 -0.0081 1.0000 0.0234
-7.250 -0.7586 0.01973 0.01182 -0.0070 1.0000 0.0248
-7.000 -0.7384 0.01888 0.01088 -0.0058 1.0000 0.0265
-6.750 -0.7179 0.01813 0.01002 -0.0046 1.0000 0.0287
-6.500 -0.6984 0.01732 0.00921 -0.0034 1.0000 0.0315
-6.250 -0.6776 0.01668 0.00850 -0.0022 1.0000 0.0345
-6.000 -0.6567 0.01606 0.00780 -0.0010 1.0000 0.0377
-5.750 -0.6364 0.01541 0.00717 0.0002 1.0000 0.0432
-5.500 -0.6151 0.01487 0.00656 0.0014 1.0000 0.0493
-5.250 -0.5940 0.01434 0.00603 0.0025 1.0000 0.0580
-5.000 -0.5729 0.01381 0.00555 0.0037 1.0000 0.0714
-4.750 -0.5521 0.01327 0.00511 0.0048 1.0000 0.0914
-4.500 -0.5317 0.01271 0.00468 0.0060 1.0000 0.1242
-4.250 -0.5122 0.01207 0.00429 0.0072 1.0000 0.1746
-4.000 -0.4937 0.01136 0.00393 0.0085 1.0000 0.2490
-3.750 -0.4754 0.01075 0.00372 0.0100 1.0000 0.3375
-3.500 -0.4402 0.01035 0.00354 0.0081 0.9916 0.4054
-3.250 -0.4043 0.01008 0.00341 0.0062 0.9831 0.4535
-3.000 -0.3701 0.00985 0.00328 0.0048 0.9729 0.4913
-2.750 -0.3358 0.00966 0.00316 0.0035 0.9619 0.5248
-2.500 -0.3014 0.00948 0.00306 0.0021 0.9503 0.5562
-2.250 -0.2670 0.00932 0.00297 0.0009 0.9374 0.5857
-2.000 -0.2331 0.00917 0.00288 -0.0002 0.9228 0.6122
-1.750 -0.1998 0.00904 0.00279 -0.0012 0.9060 0.6353
-1.500 -0.1675 0.00893 0.00270 -0.0018 0.8876 0.6559
-1.250 -0.1363 0.00885 0.00262 -0.0022 0.8680 0.6747
-1.000 -0.1074 0.00878 0.00254 -0.0021 0.8468 0.6921
-0.750 -0.0796 0.00873 0.00249 -0.0018 0.8257 0.7093
-0.500 -0.0527 0.00870 0.00245 -0.0013 0.8049 0.7275
-0.250 -0.0263 0.00868 0.00242 -0.0007 0.7844 0.7460
0.000 0.0000 0.00868 0.00241 0.0000 0.7649 0.7649
0.250 0.0263 0.00868 0.00242 0.0007 0.7460 0.7844
0.500 0.0527 0.00870 0.00245 0.0013 0.7275 0.8049
0.750 0.0796 0.00873 0.00249 0.0018 0.7094 0.8256
1.000 0.1074 0.00878 0.00254 0.0021 0.6921 0.8468
1.250 0.1363 0.00885 0.00262 0.0022 0.6748 0.8680
1.500 0.1674 0.00893 0.00270 0.0018 0.6558 0.8876
1.750 0.1998 0.00904 0.00279 0.0012 0.6354 0.9061
2.000 0.2330 0.00917 0.00288 0.0002 0.6122 0.9228
2.250 0.2669 0.00932 0.00297 -0.0009 0.5858 0.9374
2.500 0.3014 0.00948 0.00306 -0.0021 0.5564 0.9504
2.750 0.3358 0.00966 0.00316 -0.0035 0.5251 0.9620
3.000 0.3702 0.00985 0.00328 -0.0048 0.4909 0.9730
3.250 0.4043 0.01008 0.00340 -0.0062 0.4532 0.9832
3.500 0.4403 0.01035 0.00354 -0.0081 0.4055 0.9917
3.750 0.4752 0.01075 0.00372 -0.0100 0.3363 1.0000
4.000 0.4936 0.01135 0.00393 -0.0085 0.2503 1.0000
4.250 0.5120 0.01207 0.00429 -0.0072 0.1745 1.0000
4.500 0.5316 0.01271 0.00468 -0.0059 0.1243 1.0000
4.750 0.5520 0.01327 0.00511 -0.0048 0.0912 1.0000
5.000 0.5728 0.01381 0.00555 -0.0036 0.0715 1.0000
5.250 0.5939 0.01434 0.00603 -0.0025 0.0582 1.0000
5.500 0.6150 0.01488 0.00656 -0.0014 0.0492 1.0000
5.750 0.6363 0.01541 0.00716 -0.0002 0.0432 1.0000
6.000 0.6566 0.01606 0.00780 0.0010 0.0378 1.0000
6.250 0.6775 0.01668 0.00850 0.0022 0.0345 1.0000
6.500 0.6984 0.01732 0.00921 0.0034 0.0316 1.0000
6.750 0.7180 0.01813 0.01002 0.0046 0.0286 1.0000
7.000 0.7384 0.01889 0.01089 0.0058 0.0266 1.0000
7.250 0.7586 0.01973 0.01182 0.0070 0.0249 1.0000
7.500 0.7787 0.02060 0.01277 0.0081 0.0234 1.0000
7.750 0.7983 0.02155 0.01376 0.0092 0.0221 1.0000
8.000 0.8164 0.02288 0.01515 0.0105 0.0208 1.0000
8.250 0.8369 0.02382 0.01629 0.0116 0.0195 1.0000
8.500 0.8561 0.02502 0.01766 0.0127 0.0185 1.0000
8.750 0.8746 0.02634 0.01914 0.0139 0.0177 1.0000
9.000 0.8924 0.02769 0.02064 0.0150 0.0170 1.0000
9.250 0.9090 0.02909 0.02217 0.0162 0.0165 1.0000
9.500 0.9237 0.03082 0.02406 0.0175 0.0160 1.0000
9.750 0.9348 0.03318 0.02666 0.0190 0.0155 1.0000
10.000 0.9441 0.03533 0.02921 0.0208 0.0150 1.0000
10.250 0.9489 0.03785 0.03211 0.0227 0.0144 1.0000
10.500 0.9497 0.04045 0.03504 0.0248 0.0139 1.0000
10.750 0.9441 0.04318 0.03806 0.0273 0.0136 1.0000
11.000 0.9337 0.04582 0.04094 0.0297 0.0134 1.0000
11.250 0.9185 0.04921 0.04457 0.0309 0.0133 1.0000
11.500 0.8981 0.05377 0.04937 0.0302 0.0133 1.0000
11.750 0.8761 0.05941 0.05523 0.0274 0.0134 1.0000
12.000 0.8516 0.06657 0.06260 0.0225 0.0133 1.0000
12.250 0.8172 0.07737 0.07358 0.0144 0.0138 1.0000
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Polar data table (+)
Polar graphs
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