NACA 65(1)-212 a=0.6 (naca651212a06-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 65(1)-212 a=0.6 (naca651212a06-il) Reynolds number: 500,000 Max Cl/Cd: 81.46 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca651212a06-il-500000.txt Download as CSV file: xf-naca651212a06-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 65(1)-212 a=0.6
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.5238 0.08557 0.08339 -0.0379 1.0000 0.0236
-10.250 -0.5323 0.07867 0.07653 -0.0431 1.0000 0.0240
-10.000 -0.5597 0.06833 0.06615 -0.0524 1.0000 0.0238
-9.750 -0.5842 0.06240 0.06014 -0.0562 1.0000 0.0236
-9.500 -0.6006 0.05748 0.05509 -0.0587 0.9760 0.0235
-9.250 -0.5991 0.05343 0.05082 -0.0610 0.9496 0.0243
-9.000 -0.6066 0.05032 0.04748 -0.0591 0.9293 0.0249
-8.750 -0.6113 0.04721 0.04412 -0.0566 0.9148 0.0256
-8.500 -0.6064 0.04681 0.04322 -0.0531 0.9021 0.0278
-8.250 -0.6037 0.04418 0.04027 -0.0509 0.8923 0.0279
-8.000 -0.6304 0.03042 0.02549 -0.0458 0.8830 0.0218
-7.750 -0.6160 0.02749 0.02224 -0.0444 0.8755 0.0220
-7.500 -0.5985 0.02537 0.01992 -0.0433 0.8697 0.0226
-7.250 -0.5786 0.02290 0.01710 -0.0422 0.8632 0.0224
-7.000 -0.5567 0.02121 0.01521 -0.0414 0.8574 0.0227
-6.750 -0.5335 0.01990 0.01373 -0.0408 0.8523 0.0231
-6.500 -0.5090 0.01884 0.01258 -0.0404 0.8466 0.0238
-6.250 -0.4844 0.01797 0.01159 -0.0399 0.8415 0.0247
-6.000 -0.4594 0.01719 0.01069 -0.0395 0.8368 0.0258
-5.750 -0.4340 0.01627 0.00967 -0.0391 0.8313 0.0265
-5.500 -0.4087 0.01555 0.00885 -0.0387 0.8266 0.0272
-5.250 -0.3832 0.01501 0.00821 -0.0383 0.8224 0.0278
-5.000 -0.3607 0.01372 0.00686 -0.0375 0.8174 0.0289
-4.750 -0.3375 0.01298 0.00610 -0.0369 0.8128 0.0308
-4.500 -0.3129 0.01254 0.00560 -0.0364 0.8087 0.0325
-4.250 -0.2877 0.01209 0.00512 -0.0361 0.8040 0.0341
-4.000 -0.2621 0.01171 0.00468 -0.0358 0.7995 0.0358
-3.750 -0.2373 0.01124 0.00413 -0.0353 0.7956 0.0380
-3.500 -0.2118 0.01085 0.00371 -0.0350 0.7919 0.0429
-3.250 -0.1851 0.01059 0.00343 -0.0349 0.7874 0.0475
-3.000 -0.1595 0.01015 0.00305 -0.0346 0.7832 0.0682
-2.750 -0.1436 0.00840 0.00249 -0.0335 0.7794 0.3731
-2.500 -0.1234 0.00759 0.00243 -0.0324 0.7756 0.5764
-2.250 -0.0968 0.00747 0.00241 -0.0322 0.7715 0.6185
-2.000 -0.0696 0.00742 0.00237 -0.0320 0.7676 0.6449
-1.750 -0.0427 0.00741 0.00237 -0.0318 0.7642 0.6734
-1.500 -0.0160 0.00740 0.00243 -0.0314 0.7606 0.6999
-1.250 0.0106 0.00742 0.00253 -0.0311 0.7566 0.7239
-1.000 0.0381 0.00748 0.00258 -0.0309 0.7529 0.7380
-0.750 0.0662 0.00754 0.00258 -0.0309 0.7495 0.7481
-0.500 0.0944 0.00756 0.00259 -0.0311 0.7462 0.7537
-0.250 0.1228 0.00757 0.00260 -0.0313 0.7422 0.7604
0.000 0.1511 0.00757 0.00258 -0.0315 0.7384 0.7664
0.250 0.1794 0.00758 0.00257 -0.0316 0.7351 0.7730
0.500 0.2077 0.00760 0.00256 -0.0318 0.7319 0.7797
0.750 0.2358 0.00756 0.00257 -0.0319 0.7279 0.7865
1.000 0.2639 0.00752 0.00255 -0.0321 0.7241 0.7938
1.250 0.2919 0.00746 0.00251 -0.0322 0.7207 0.8014
1.500 0.3198 0.00742 0.00250 -0.0323 0.7175 0.8097
1.750 0.3474 0.00733 0.00252 -0.0323 0.7136 0.8193
2.000 0.3740 0.00727 0.00262 -0.0321 0.7095 0.8302
2.250 0.3988 0.00734 0.00282 -0.0311 0.7051 0.8421
2.500 0.4220 0.00751 0.00309 -0.0297 0.6999 0.8547
2.750 0.4509 0.00759 0.00313 -0.0301 0.6904 0.8965
3.000 0.4763 0.00755 0.00308 -0.0294 0.6776 0.9073
3.250 0.5008 0.00746 0.00300 -0.0286 0.6622 0.9167
3.500 0.5246 0.00736 0.00289 -0.0275 0.6441 0.9253
3.750 0.5484 0.00730 0.00285 -0.0265 0.6266 0.9339
4.000 0.5725 0.00728 0.00285 -0.0256 0.6060 0.9421
4.250 0.5971 0.00733 0.00286 -0.0248 0.5752 0.9497
4.500 0.6189 0.00764 0.00294 -0.0237 0.5048 0.9577
4.750 0.6299 0.00934 0.00364 -0.0217 0.2918 0.9677
5.000 0.6462 0.01146 0.00468 -0.0214 0.0751 0.9758
5.250 0.6750 0.01223 0.00530 -0.0223 0.0436 0.9820
5.500 0.7073 0.01291 0.00596 -0.0238 0.0357 0.9864
5.750 0.7397 0.01343 0.00653 -0.0253 0.0320 0.9909
6.000 0.7673 0.01429 0.00741 -0.0261 0.0288 0.9970
6.250 0.7875 0.01510 0.00827 -0.0252 0.0274 1.0000
6.500 0.8005 0.01557 0.00879 -0.0227 0.0265 1.0000
6.750 0.8126 0.01613 0.00940 -0.0201 0.0257 1.0000
7.000 0.8262 0.01671 0.01002 -0.0178 0.0246 1.0000
7.250 0.8412 0.01732 0.01064 -0.0158 0.0236 1.0000
7.500 0.8563 0.01801 0.01134 -0.0140 0.0227 1.0000
7.750 0.8696 0.01913 0.01248 -0.0119 0.0218 1.0000
8.000 0.8905 0.02093 0.01434 -0.0112 0.0211 1.0000
8.250 0.9118 0.02173 0.01522 -0.0103 0.0208 1.0000
8.500 0.9349 0.02284 0.01642 -0.0098 0.0205 1.0000
8.750 0.9576 0.02395 0.01764 -0.0092 0.0200 1.0000
9.000 0.9815 0.02546 0.01929 -0.0089 0.0197 1.0000
9.250 1.0025 0.02679 0.02077 -0.0081 0.0191 1.0000
9.500 1.0216 0.02805 0.02216 -0.0072 0.0184 1.0000
9.750 1.0408 0.03012 0.02445 -0.0063 0.0182 1.0000
10.000 1.0565 0.03261 0.02720 -0.0049 0.0182 1.0000
10.250 1.0652 0.03647 0.03147 -0.0028 0.0186 1.0000
10.500 1.0651 0.04058 0.03597 0.0002 0.0193 1.0000
10.750 1.0567 0.04533 0.04103 0.0035 0.0203 1.0000
11.000 1.0116 0.05393 0.05041 0.0116 0.0252 1.0000
11.250 0.9992 0.05628 0.05291 0.0145 0.0247 1.0000
11.500 0.9832 0.05929 0.05609 0.0167 0.0243 1.0000
11.750 0.9701 0.06214 0.05906 0.0181 0.0239 1.0000
12.000 0.9476 0.06641 0.06351 0.0190 0.0239 1.0000
12.250 0.9254 0.07099 0.06824 0.0188 0.0239 1.0000
12.500 0.9080 0.07533 0.07269 0.0178 0.0235 1.0000
12.750 0.8700 0.08351 0.08106 0.0145 0.0240 1.0000
13.000 0.8328 0.09330 0.09100 0.0088 0.0247 1.0000
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Polar data table (+)
Polar graphs
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