RAF 30 AIRFOIL (raf30-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 30 AIRFOIL (raf30-il) Reynolds number: 100,000 Max Cl/Cd: 41.72 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf30-il-100000.txt Download as CSV file: xf-raf30-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.7953 0.08460 0.07896 -0.0399 1.0000 0.0787
-12.250 -0.7997 0.07897 0.07332 -0.0410 1.0000 0.0778
-12.000 -0.8258 0.07294 0.06719 -0.0432 1.0000 0.0772
-11.750 -0.8497 0.06793 0.06206 -0.0439 1.0000 0.0765
-11.500 -0.8739 0.06364 0.05761 -0.0433 1.0000 0.0758
-11.250 -0.8993 0.06009 0.05387 -0.0408 1.0000 0.0757
-11.000 -0.9199 0.05691 0.05049 -0.0372 1.0000 0.0754
-10.750 -0.9348 0.05365 0.04696 -0.0339 1.0000 0.0755
-10.500 -0.9443 0.05048 0.04347 -0.0307 1.0000 0.0758
-10.250 -0.9487 0.04744 0.04007 -0.0276 1.0000 0.0761
-10.000 -0.9487 0.04450 0.03676 -0.0246 1.0000 0.0766
-9.750 -0.9445 0.04175 0.03362 -0.0219 1.0000 0.0771
-9.500 -0.9379 0.03949 0.03094 -0.0191 1.0000 0.0781
-9.250 -0.9251 0.03681 0.02802 -0.0173 1.0000 0.0800
-9.000 -0.9072 0.03488 0.02605 -0.0162 1.0000 0.0825
-8.750 -0.8893 0.03310 0.02411 -0.0147 1.0000 0.0848
-8.500 -0.8705 0.03141 0.02222 -0.0132 1.0000 0.0873
-8.250 -0.8529 0.03009 0.02059 -0.0114 1.0000 0.0906
-8.000 -0.8296 0.02818 0.01877 -0.0108 1.0000 0.0945
-7.750 -0.8084 0.02686 0.01745 -0.0097 1.0000 0.0987
-7.500 -0.7881 0.02572 0.01613 -0.0082 1.0000 0.1039
-7.250 -0.7689 0.02432 0.01493 -0.0069 1.0000 0.1111
-7.000 -0.7515 0.02314 0.01376 -0.0050 1.0000 0.1204
-6.750 -0.7371 0.02192 0.01267 -0.0027 1.0000 0.1340
-6.500 -0.7279 0.02054 0.01157 0.0004 1.0000 0.1594
-6.250 -0.7224 0.01912 0.01059 0.0040 1.0000 0.2068
-6.000 -0.7153 0.01810 0.00996 0.0074 1.0000 0.2621
-5.750 -0.7054 0.01739 0.00948 0.0105 1.0000 0.3140
-5.500 -0.6945 0.01682 0.00913 0.0135 1.0000 0.3598
-5.250 -0.6829 0.01638 0.00889 0.0164 1.0000 0.4025
-5.000 -0.6709 0.01608 0.00875 0.0193 1.0000 0.4448
-4.750 -0.6590 0.01590 0.00871 0.0223 1.0000 0.4868
-4.500 -0.6464 0.01583 0.00878 0.0253 1.0000 0.5252
-4.250 -0.6332 0.01583 0.00888 0.0282 1.0000 0.5604
-4.000 -0.6191 0.01585 0.00896 0.0309 1.0000 0.5913
-3.750 -0.6046 0.01587 0.00900 0.0334 1.0000 0.6194
-3.500 -0.5892 0.01590 0.00907 0.0358 1.0000 0.6443
-3.250 -0.5742 0.01594 0.00914 0.0382 1.0000 0.6685
-3.000 -0.5589 0.01601 0.00924 0.0405 1.0000 0.6915
-2.750 -0.5446 0.01608 0.00932 0.0428 1.0000 0.7152
-2.500 -0.5229 0.01629 0.00959 0.0440 0.9977 0.7399
-2.250 -0.4821 0.01676 0.01009 0.0419 0.9881 0.7695
-2.000 -0.4422 0.01712 0.01044 0.0398 0.9785 0.7965
-1.750 -0.3962 0.01745 0.01076 0.0365 0.9702 0.8168
-1.500 -0.3559 0.01763 0.01092 0.0343 0.9599 0.8358
-1.250 -0.3108 0.01787 0.01112 0.0312 0.9508 0.8536
-1.000 -0.2578 0.01814 0.01135 0.0266 0.9429 0.8694
-0.750 -0.2067 0.01839 0.01157 0.0224 0.9335 0.8836
-0.500 -0.1423 0.01860 0.01174 0.0157 0.9268 0.8955
-0.250 -0.0773 0.01879 0.01192 0.0090 0.9178 0.9046
0.000 0.0002 0.01881 0.01192 0.0000 0.9119 0.9120
0.250 0.0772 0.01879 0.01191 -0.0089 0.9045 0.9178
0.500 0.1424 0.01859 0.01173 -0.0157 0.8955 0.9268
0.750 0.2069 0.01839 0.01157 -0.0224 0.8836 0.9336
1.000 0.2579 0.01814 0.01135 -0.0266 0.8694 0.9429
1.250 0.3108 0.01787 0.01113 -0.0312 0.8537 0.9509
1.500 0.3558 0.01763 0.01091 -0.0343 0.8357 0.9599
1.750 0.3960 0.01745 0.01076 -0.0365 0.8168 0.9701
2.000 0.4423 0.01711 0.01043 -0.0398 0.7965 0.9786
2.250 0.4821 0.01675 0.01008 -0.0419 0.7694 0.9881
2.500 0.5230 0.01629 0.00959 -0.0440 0.7398 0.9978
2.750 0.5445 0.01608 0.00932 -0.0428 0.7153 1.0000
3.000 0.5588 0.01601 0.00925 -0.0405 0.6916 1.0000
3.250 0.5741 0.01594 0.00914 -0.0381 0.6685 1.0000
3.500 0.5892 0.01590 0.00907 -0.0358 0.6444 1.0000
3.750 0.6045 0.01587 0.00900 -0.0334 0.6194 1.0000
4.000 0.6189 0.01585 0.00896 -0.0308 0.5911 1.0000
4.250 0.6330 0.01583 0.00887 -0.0281 0.5600 1.0000
4.500 0.6463 0.01583 0.00877 -0.0253 0.5251 1.0000
4.750 0.6588 0.01590 0.00870 -0.0223 0.4865 1.0000
5.000 0.6708 0.01608 0.00875 -0.0193 0.4449 1.0000
5.250 0.6828 0.01638 0.00889 -0.0164 0.4024 1.0000
5.500 0.6945 0.01681 0.00912 -0.0134 0.3597 1.0000
5.750 0.7053 0.01739 0.00948 -0.0105 0.3138 1.0000
6.000 0.7153 0.01810 0.00996 -0.0074 0.2618 1.0000
6.250 0.7223 0.01913 0.01059 -0.0040 0.2066 1.0000
6.500 0.7279 0.02054 0.01157 -0.0004 0.1590 1.0000
6.750 0.7372 0.02191 0.01267 0.0027 0.1340 1.0000
7.000 0.7514 0.02314 0.01377 0.0050 0.1203 1.0000
7.250 0.7689 0.02432 0.01493 0.0069 0.1110 1.0000
7.500 0.7881 0.02572 0.01614 0.0082 0.1038 1.0000
7.750 0.8085 0.02686 0.01745 0.0097 0.0988 1.0000
8.000 0.8297 0.02818 0.01878 0.0108 0.0945 1.0000
8.250 0.8529 0.03008 0.02059 0.0114 0.0906 1.0000
8.500 0.8705 0.03141 0.02223 0.0132 0.0873 1.0000
8.750 0.8893 0.03310 0.02412 0.0147 0.0847 1.0000
9.000 0.9073 0.03489 0.02607 0.0162 0.0825 1.0000
9.250 0.9252 0.03681 0.02802 0.0173 0.0800 1.0000
9.500 0.9380 0.03952 0.03095 0.0190 0.0782 1.0000
9.750 0.9446 0.04176 0.03363 0.0219 0.0771 1.0000
10.000 0.9490 0.04451 0.03676 0.0246 0.0766 1.0000
10.250 0.9491 0.04743 0.04006 0.0276 0.0762 1.0000
10.500 0.9440 0.05041 0.04340 0.0308 0.0755 1.0000
10.750 0.9345 0.05358 0.04689 0.0340 0.0753 1.0000
11.000 0.9197 0.05684 0.05042 0.0372 0.0752 1.0000
11.250 0.8982 0.06000 0.05379 0.0409 0.0753 1.0000
11.500 0.8744 0.06362 0.05759 0.0432 0.0758 1.0000
11.750 0.8502 0.06794 0.06207 0.0438 0.0765 1.0000
12.000 0.8269 0.07294 0.06719 0.0431 0.0772 1.0000
12.250 0.8051 0.07868 0.07301 0.0412 0.0779 1.0000
12.500 0.7957 0.08464 0.07900 0.0397 0.0787 1.0000
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Polar data table (+)
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