RAF 30 AIRFOIL (raf30-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RAF 30 AIRFOIL (raf30-il) Reynolds number: 50,000 Max Cl/Cd: 28.69 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf30-il-50000-n5.txt Download as CSV file: xf-raf30-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAF 30 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.750 -0.7104 0.09903 0.09116 -0.0247 1.0000 0.0704
-12.500 -0.7516 0.08711 0.07918 -0.0323 1.0000 0.0694
-12.250 -0.7924 0.07759 0.06953 -0.0381 1.0000 0.0684
-12.000 -0.8255 0.07074 0.06248 -0.0412 1.0000 0.0679
-11.750 -0.8503 0.06569 0.05722 -0.0422 1.0000 0.0679
-11.500 -0.8696 0.06162 0.05293 -0.0417 1.0000 0.0682
-11.250 -0.8847 0.05819 0.04926 -0.0403 1.0000 0.0689
-11.000 -0.8955 0.05516 0.04597 -0.0380 1.0000 0.0696
-10.750 -0.9021 0.05233 0.04285 -0.0355 1.0000 0.0705
-10.500 -0.9036 0.04958 0.03974 -0.0332 1.0000 0.0719
-10.250 -0.8996 0.04713 0.03699 -0.0312 1.0000 0.0739
-10.000 -0.8862 0.04536 0.03524 -0.0300 1.0000 0.0765
-9.750 -0.8742 0.04351 0.03324 -0.0286 1.0000 0.0794
-9.500 -0.8594 0.04151 0.03097 -0.0272 1.0000 0.0821
-9.250 -0.8415 0.03957 0.02870 -0.0260 1.0000 0.0851
-9.000 -0.8237 0.03798 0.02713 -0.0249 1.0000 0.0889
-8.750 -0.8078 0.03660 0.02567 -0.0235 1.0000 0.0941
-8.500 -0.7890 0.03515 0.02403 -0.0223 1.0000 0.1000
-8.250 -0.7729 0.03377 0.02270 -0.0208 1.0000 0.1059
-8.000 -0.7562 0.03243 0.02124 -0.0191 1.0000 0.1141
-7.750 -0.7440 0.03113 0.01999 -0.0171 1.0000 0.1249
-7.500 -0.7327 0.02985 0.01878 -0.0149 1.0000 0.1391
-7.250 -0.7224 0.02857 0.01762 -0.0125 1.0000 0.1581
-7.000 -0.7129 0.02737 0.01657 -0.0100 1.0000 0.1799
-6.750 -0.7028 0.02631 0.01567 -0.0075 1.0000 0.2076
-6.500 -0.6926 0.02538 0.01491 -0.0050 1.0000 0.2414
-6.250 -0.6834 0.02453 0.01430 -0.0022 1.0000 0.2845
-6.000 -0.6744 0.02378 0.01381 0.0007 1.0000 0.3308
-5.500 -0.6508 0.02269 0.01306 0.0064 1.0000 0.4111
-5.250 -0.6363 0.02230 0.01275 0.0090 1.0000 0.4443
-5.000 -0.6210 0.02199 0.01249 0.0114 1.0000 0.4751
-4.750 -0.6056 0.02174 0.01227 0.0139 1.0000 0.5048
-4.500 -0.5902 0.02154 0.01206 0.0163 1.0000 0.5335
-4.250 -0.5753 0.02136 0.01188 0.0189 1.0000 0.5619
-4.000 -0.5597 0.02127 0.01182 0.0214 1.0000 0.5891
-3.750 -0.5444 0.02123 0.01180 0.0240 1.0000 0.6163
-3.500 -0.5298 0.02119 0.01175 0.0267 1.0000 0.6432
-3.250 -0.5138 0.02120 0.01177 0.0291 1.0000 0.6680
-3.000 -0.4983 0.02119 0.01174 0.0315 1.0000 0.6911
-2.750 -0.4815 0.02121 0.01173 0.0336 1.0000 0.7115
-2.500 -0.4578 0.02125 0.01173 0.0342 0.9967 0.7325
-2.250 -0.4188 0.02138 0.01177 0.0318 0.9863 0.7542
-2.000 -0.3785 0.02154 0.01186 0.0295 0.9762 0.7734
-1.750 -0.3382 0.02168 0.01194 0.0271 0.9661 0.7921
-1.500 -0.2960 0.02183 0.01202 0.0243 0.9563 0.8100
-1.250 -0.2519 0.02198 0.01212 0.0213 0.9467 0.8273
-1.000 -0.2088 0.02211 0.01221 0.0184 0.9362 0.8438
-0.750 -0.1583 0.02228 0.01233 0.0142 0.9277 0.8592
-0.500 -0.1083 0.02242 0.01244 0.0100 0.9178 0.8736
-0.250 -0.0556 0.02252 0.01253 0.0053 0.9082 0.8869
0.000 0.0000 0.02253 0.01253 0.0000 0.8990 0.8990
0.250 0.0557 0.02252 0.01253 -0.0053 0.8869 0.9083
0.500 0.1083 0.02242 0.01244 -0.0100 0.8736 0.9178
0.750 0.1583 0.02228 0.01233 -0.0141 0.8592 0.9276
1.000 0.2087 0.02211 0.01220 -0.0184 0.8437 0.9362
1.250 0.2519 0.02198 0.01211 -0.0213 0.8272 0.9467
1.500 0.2960 0.02183 0.01201 -0.0243 0.8099 0.9563
1.750 0.3382 0.02168 0.01194 -0.0271 0.7920 0.9660
2.000 0.3785 0.02154 0.01186 -0.0295 0.7734 0.9762
2.250 0.4188 0.02138 0.01177 -0.0318 0.7542 0.9863
2.500 0.4578 0.02125 0.01172 -0.0342 0.7326 0.9967
2.750 0.4815 0.02121 0.01173 -0.0336 0.7115 1.0000
3.000 0.4983 0.02119 0.01173 -0.0315 0.6910 1.0000
3.250 0.5139 0.02120 0.01177 -0.0291 0.6680 1.0000
3.500 0.5298 0.02119 0.01175 -0.0267 0.6432 1.0000
3.750 0.5444 0.02123 0.01180 -0.0240 0.6163 1.0000
4.000 0.5598 0.02127 0.01182 -0.0214 0.5891 1.0000
4.250 0.5753 0.02136 0.01188 -0.0189 0.5618 1.0000
4.500 0.5902 0.02153 0.01206 -0.0163 0.5334 1.0000
4.750 0.6056 0.02174 0.01227 -0.0139 0.5047 1.0000
5.000 0.6211 0.02199 0.01249 -0.0114 0.4751 1.0000
5.250 0.6363 0.02230 0.01275 -0.0090 0.4442 1.0000
5.500 0.6509 0.02269 0.01306 -0.0064 0.4111 1.0000
5.750 0.6639 0.02317 0.01339 -0.0037 0.3743 1.0000
6.000 0.6743 0.02378 0.01380 -0.0007 0.3304 1.0000
6.250 0.6834 0.02453 0.01430 0.0022 0.2843 1.0000
6.500 0.6926 0.02537 0.01491 0.0050 0.2413 1.0000
6.750 0.7028 0.02631 0.01567 0.0075 0.2073 1.0000
7.000 0.7129 0.02737 0.01659 0.0100 0.1798 1.0000
7.250 0.7223 0.02858 0.01762 0.0125 0.1580 1.0000
7.500 0.7328 0.02984 0.01878 0.0149 0.1393 1.0000
7.750 0.7440 0.03113 0.02000 0.0171 0.1246 1.0000
8.000 0.7563 0.03243 0.02124 0.0191 0.1141 1.0000
8.250 0.7730 0.03377 0.02269 0.0207 0.1059 1.0000
8.500 0.7892 0.03514 0.02402 0.0222 0.1001 1.0000
8.750 0.8080 0.03660 0.02568 0.0235 0.0942 1.0000
9.000 0.8240 0.03799 0.02714 0.0249 0.0890 1.0000
9.250 0.8416 0.03956 0.02869 0.0260 0.0852 1.0000
9.500 0.8595 0.04151 0.03096 0.0272 0.0822 1.0000
9.750 0.8744 0.04352 0.03325 0.0285 0.0794 1.0000
10.000 0.8864 0.04538 0.03526 0.0300 0.0765 1.0000
10.250 0.8998 0.04715 0.03702 0.0312 0.0739 1.0000
10.500 0.9035 0.04960 0.03976 0.0332 0.0719 1.0000
10.750 0.9024 0.05232 0.04284 0.0354 0.0706 1.0000
11.000 0.8961 0.05515 0.04596 0.0380 0.0697 1.0000
11.250 0.8849 0.05819 0.04926 0.0402 0.0689 1.0000
11.500 0.8700 0.06162 0.05293 0.0417 0.0682 1.0000
11.750 0.8506 0.06569 0.05722 0.0421 0.0678 1.0000
12.000 0.8259 0.07074 0.06248 0.0411 0.0678 1.0000
12.250 0.7927 0.07765 0.06958 0.0380 0.0684 1.0000
12.500 0.7523 0.08713 0.07920 0.0322 0.0694 1.0000
12.750 0.7112 0.09905 0.09117 0.0246 0.0704 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAF 30 AIRFOIL (raf30-il)