RAE 101 AIRFOIL (rae101-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: RAE 101 AIRFOIL (rae101-il) Reynolds number: 100,000 Max Cl/Cd: 36.4 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae101-il-100000-n5.txt Download as CSV file: xf-rae101-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.7471 0.08116 0.07601 -0.0150 1.0000 0.0254
-11.000 -0.7755 0.07206 0.06680 -0.0221 1.0000 0.0250
-10.750 -0.8008 0.06542 0.06001 -0.0265 1.0000 0.0248
-10.500 -0.8235 0.06040 0.05480 -0.0283 1.0000 0.0245
-10.250 -0.8434 0.05648 0.05067 -0.0277 1.0000 0.0245
-10.000 -0.8588 0.05298 0.04691 -0.0258 1.0000 0.0246
-9.750 -0.8672 0.04932 0.04292 -0.0241 1.0000 0.0247
-9.500 -0.8696 0.04586 0.03909 -0.0223 1.0000 0.0251
-9.250 -0.8668 0.04275 0.03557 -0.0205 1.0000 0.0261
-9.000 -0.8602 0.03988 0.03220 -0.0187 1.0000 0.0273
-8.750 -0.8497 0.03722 0.02906 -0.0170 1.0000 0.0283
-8.500 -0.8353 0.03459 0.02627 -0.0159 1.0000 0.0293
-8.250 -0.8183 0.03259 0.02405 -0.0148 1.0000 0.0303
-8.000 -0.7999 0.03075 0.02201 -0.0137 1.0000 0.0316
-7.750 -0.7804 0.02912 0.02016 -0.0126 1.0000 0.0337
-7.500 -0.7597 0.02772 0.01847 -0.0115 1.0000 0.0364
-7.250 -0.7400 0.02600 0.01668 -0.0104 1.0000 0.0387
-7.000 -0.7203 0.02468 0.01531 -0.0093 1.0000 0.0411
-6.750 -0.7000 0.02357 0.01410 -0.0082 1.0000 0.0447
-6.500 -0.6803 0.02245 0.01289 -0.0069 1.0000 0.0491
-6.250 -0.6616 0.02139 0.01184 -0.0055 1.0000 0.0537
-6.000 -0.6416 0.02053 0.01085 -0.0042 1.0000 0.0598
-5.750 -0.6231 0.01957 0.00990 -0.0028 1.0000 0.0686
-5.500 -0.6041 0.01866 0.00901 -0.0014 1.0000 0.0796
-5.250 -0.5847 0.01782 0.00821 0.0000 1.0000 0.0981
-5.000 -0.5660 0.01690 0.00746 0.0013 1.0000 0.1280
-4.750 -0.5484 0.01587 0.00677 0.0027 1.0000 0.1861
-4.500 -0.5324 0.01478 0.00623 0.0044 1.0000 0.2851
-4.250 -0.5145 0.01413 0.00596 0.0060 1.0000 0.3803
-4.000 -0.4947 0.01374 0.00575 0.0075 1.0000 0.4413
-3.750 -0.4743 0.01345 0.00558 0.0091 1.0000 0.4870
-3.500 -0.4538 0.01322 0.00541 0.0106 1.0000 0.5253
-3.250 -0.4331 0.01303 0.00528 0.0121 1.0000 0.5593
-3.000 -0.4126 0.01286 0.00518 0.0136 1.0000 0.5901
-2.750 -0.3923 0.01271 0.00510 0.0152 1.0000 0.6194
-2.500 -0.3722 0.01259 0.00502 0.0167 1.0000 0.6471
-2.250 -0.3445 0.01250 0.00498 0.0167 0.9953 0.6741
-2.000 -0.3053 0.01243 0.00495 0.0144 0.9837 0.6994
-1.500 -0.2279 0.01227 0.00480 0.0102 0.9589 0.7431
-1.250 -0.1893 0.01219 0.00473 0.0082 0.9458 0.7643
-1.000 -0.1511 0.01212 0.00468 0.0064 0.9316 0.7842
-0.750 -0.1132 0.01206 0.00464 0.0047 0.9164 0.8051
-0.500 -0.0754 0.01202 0.00461 0.0030 0.9003 0.8255
-0.250 -0.0371 0.01199 0.00460 0.0014 0.8835 0.8453
0.000 0.0000 0.01198 0.00459 0.0000 0.8649 0.8648
0.250 0.0371 0.01199 0.00460 -0.0014 0.8453 0.8835
0.500 0.0754 0.01202 0.00461 -0.0030 0.8255 0.9003
0.750 0.1132 0.01206 0.00464 -0.0047 0.8051 0.9164
1.000 0.1510 0.01212 0.00468 -0.0064 0.7842 0.9316
1.250 0.1892 0.01219 0.00472 -0.0082 0.7643 0.9458
1.500 0.2279 0.01227 0.00479 -0.0102 0.7431 0.9590
1.750 0.2665 0.01235 0.00486 -0.0123 0.7222 0.9715
2.000 0.3053 0.01243 0.00494 -0.0144 0.6994 0.9838
2.250 0.3445 0.01249 0.00498 -0.0167 0.6741 0.9954
2.500 0.3720 0.01259 0.00502 -0.0167 0.6472 1.0000
2.750 0.3922 0.01271 0.00510 -0.0151 0.6194 1.0000
3.000 0.4125 0.01285 0.00518 -0.0136 0.5902 1.0000
3.250 0.4330 0.01302 0.00528 -0.0120 0.5595 1.0000
3.500 0.4536 0.01322 0.00541 -0.0105 0.5253 1.0000
3.750 0.4742 0.01345 0.00558 -0.0090 0.4871 1.0000
4.000 0.4945 0.01374 0.00575 -0.0075 0.4412 1.0000
4.250 0.5143 0.01413 0.00596 -0.0060 0.3801 1.0000
4.500 0.5322 0.01479 0.00623 -0.0043 0.2848 1.0000
4.750 0.5483 0.01587 0.00677 -0.0027 0.1861 1.0000
5.000 0.5660 0.01690 0.00746 -0.0013 0.1280 1.0000
5.250 0.5847 0.01782 0.00821 0.0001 0.0981 1.0000
5.500 0.6040 0.01866 0.00901 0.0014 0.0794 1.0000
5.750 0.6230 0.01956 0.00990 0.0028 0.0686 1.0000
6.000 0.6416 0.02052 0.01085 0.0042 0.0599 1.0000
6.250 0.6616 0.02139 0.01184 0.0055 0.0536 1.0000
6.500 0.6804 0.02244 0.01289 0.0069 0.0491 1.0000
6.750 0.7001 0.02356 0.01410 0.0081 0.0447 1.0000
7.000 0.7203 0.02467 0.01531 0.0093 0.0411 1.0000
7.250 0.7401 0.02600 0.01669 0.0104 0.0387 1.0000
7.500 0.7598 0.02771 0.01847 0.0115 0.0364 1.0000
7.750 0.7805 0.02910 0.02014 0.0126 0.0336 1.0000
8.000 0.8001 0.03074 0.02200 0.0137 0.0317 1.0000
8.250 0.8185 0.03259 0.02405 0.0148 0.0304 1.0000
8.500 0.8355 0.03460 0.02624 0.0159 0.0293 1.0000
8.750 0.8501 0.03719 0.02904 0.0170 0.0284 1.0000
9.000 0.8606 0.03991 0.03221 0.0186 0.0274 1.0000
9.250 0.8671 0.04277 0.03559 0.0204 0.0262 1.0000
9.500 0.8699 0.04591 0.03914 0.0222 0.0253 1.0000
9.750 0.8676 0.04934 0.04295 0.0240 0.0248 1.0000
10.000 0.8595 0.05296 0.04690 0.0257 0.0246 1.0000
10.250 0.8436 0.05654 0.05073 0.0276 0.0245 1.0000
10.500 0.8236 0.06049 0.05489 0.0282 0.0246 1.0000
10.750 0.8016 0.06542 0.06000 0.0264 0.0247 1.0000
11.000 0.7773 0.07193 0.06666 0.0221 0.0251 1.0000
11.250 0.7491 0.08093 0.07578 0.0151 0.0255 1.0000
11.500 0.7217 0.09229 0.08716 0.0069 0.0260 1.0000
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Polar data table (+)
Polar graphs
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