Cambered plate C=8% T=5% R=1.6 (cp-080-050-gn) Xfoil prediction polar at RE=1,000,000 Ncrit=1
| Details | Polar file |
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Airfoil: Cambered plate C=8% T=5% R=1.6 (cp-080-050-gn) Reynolds number: 1,000,000 Max Cl/Cd: 64.23 at α=8.5° Description: Mach=0 Ncrit=1 Source: Xfoil prediction Download polar: xf-cp-080-050-gn-1000000-n1.txt Download as CSV file: xf-cp-080-050-gn-1000000-n1.csv |
XFOIL Version 6.96
Calculated polar for: Cambered plate C=8% T=5% R=1.6
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 1.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.0642 0.09221 0.08944 -0.1277 0.8938 0.0233
-8.750 -0.0530 0.08932 0.08651 -0.1301 0.8873 0.0237
-8.500 -0.0441 0.08632 0.08348 -0.1323 0.8806 0.0245
-8.000 -0.0211 0.08176 0.07886 -0.1356 0.8663 0.0250
-7.750 -0.0102 0.07968 0.07674 -0.1368 0.8585 0.0252
-7.500 0.0009 0.07775 0.07479 -0.1379 0.8522 0.0254
-7.250 0.0088 0.07539 0.07240 -0.1390 0.8437 0.0260
-7.000 0.0142 0.07344 0.07030 -0.1392 0.8145 0.0264
-6.750 0.0240 0.07099 0.06794 -0.1407 0.8288 0.0267
-6.500 0.0278 0.06885 0.06572 -0.1409 0.8091 0.0273
-6.250 0.0337 0.06639 0.06330 -0.1420 0.8138 0.0278
-6.000 0.0342 0.06529 0.06200 -0.1401 0.7678 0.0281
-5.750 0.0391 0.06341 0.06014 -0.1402 0.7671 0.0283
-1.750 0.1583 0.01527 0.00757 -0.1429 0.0363 0.0421
-1.500 0.1805 0.01437 0.00647 -0.1419 0.0353 0.0427
-1.250 0.2040 0.01389 0.00586 -0.1409 0.0346 0.0431
-1.000 0.2279 0.01355 0.00542 -0.1399 0.0340 0.0434
-0.750 0.2520 0.01331 0.00510 -0.1390 0.0334 0.0437
-0.500 0.2760 0.01309 0.00481 -0.1379 0.0331 0.0440
-0.250 0.3002 0.01295 0.00461 -0.1369 0.0328 0.0443
0.000 0.3241 0.01278 0.00438 -0.1359 0.0325 0.0446
0.250 0.3483 0.01275 0.00432 -0.1348 0.0323 0.0448
0.500 0.3723 0.01271 0.00425 -0.1337 0.0321 0.0450
0.750 0.3961 0.01266 0.00417 -0.1325 0.0318 0.0452
1.000 0.4195 0.01265 0.00413 -0.1313 0.0315 0.0454
1.250 0.4421 0.01260 0.00405 -0.1298 0.0313 0.0456
1.500 0.4644 0.01259 0.00401 -0.1283 0.0311 0.0458
1.750 0.4865 0.01256 0.00395 -0.1268 0.0307 0.0460
2.000 0.5087 0.01255 0.00392 -0.1253 0.0304 0.0463
2.250 0.5309 0.01256 0.00390 -0.1238 0.0301 0.0465
2.500 0.5530 0.01258 0.00390 -0.1222 0.0300 0.0468
2.750 0.5750 0.01260 0.00391 -0.1207 0.0299 0.0470
3.000 0.5969 0.01265 0.00393 -0.1191 0.0299 0.0472
3.250 0.6190 0.01270 0.00397 -0.1176 0.0298 0.0475
3.500 0.6411 0.01277 0.00403 -0.1160 0.0297 0.0479
3.750 0.6631 0.01285 0.00410 -0.1145 0.0297 0.0481
4.000 0.6849 0.01295 0.00419 -0.1129 0.0296 0.0482
4.250 0.7066 0.01306 0.00429 -0.1113 0.0295 0.0483
4.500 0.7282 0.01318 0.00441 -0.1097 0.0294 0.0484
4.750 0.7497 0.01330 0.00453 -0.1081 0.0293 0.0485
5.000 0.7711 0.01344 0.00467 -0.1064 0.0293 0.0485
5.250 0.7923 0.01359 0.00482 -0.1048 0.0292 0.0487
5.500 0.8133 0.01374 0.00498 -0.1031 0.0290 0.0488
5.750 0.8340 0.01391 0.00515 -0.1013 0.0290 0.0488
6.000 0.8548 0.01409 0.00534 -0.0996 0.0289 0.0491
6.250 0.8755 0.01426 0.00553 -0.0979 0.0288 0.0493
6.500 0.8959 0.01446 0.00574 -0.0961 0.0286 0.0498
6.750 0.9160 0.01466 0.00595 -0.0943 0.0285 0.0500
7.000 0.9359 0.01488 0.00618 -0.0925 0.0283 0.0502
7.250 0.9560 0.01509 0.00640 -0.0907 0.0281 0.0504
7.500 0.9758 0.01533 0.00665 -0.0888 0.0279 0.0504
7.750 0.9952 0.01557 0.00691 -0.0870 0.0277 0.0508
8.000 1.0144 0.01584 0.00718 -0.0851 0.0276 0.0508
8.250 1.0335 0.01611 0.00747 -0.0832 0.0275 0.0509
8.500 1.0527 0.01639 0.00777 -0.0813 0.0273 0.0511
8.750 1.0713 0.01669 0.00809 -0.0794 0.0272 0.0515
9.000 1.0899 0.01700 0.00841 -0.0774 0.0270 0.0517
9.250 1.1081 0.01733 0.00877 -0.0755 0.0270 0.0519
9.500 1.1266 0.01766 0.00911 -0.0736 0.0268 0.0519
9.750 1.1447 0.01801 0.00948 -0.0716 0.0267 0.0522
10.000 1.1623 0.01838 0.00988 -0.0697 0.0266 0.0524
10.250 1.1801 0.01876 0.01027 -0.0677 0.0265 0.0526
10.500 1.1974 0.01917 0.01071 -0.0658 0.0265 0.0526
10.750 1.2151 0.01957 0.01113 -0.0638 0.0263 0.0530
11.000 1.2319 0.02000 0.01159 -0.0618 0.0262 0.0534
11.250 1.2489 0.02045 0.01207 -0.0599 0.0261 0.0544
11.500 1.2656 0.02093 0.01258 -0.0579 0.0260 0.0553
11.750 1.2816 0.02143 0.01312 -0.0559 0.0259 0.0559
12.000 1.2975 0.02196 0.01368 -0.0539 0.0258 0.0564
12.250 1.3136 0.02249 0.01424 -0.0520 0.0257 0.0568
12.500 1.3291 0.02307 0.01486 -0.0500 0.0256 0.0578
12.750 1.3442 0.02366 0.01549 -0.0480 0.0255 0.0596
13.000 1.3596 0.02426 0.01612 -0.0461 0.0254 0.0610
13.250 1.3747 0.02488 0.01680 -0.0442 0.0253 0.0670
13.500 1.3824 0.02476 0.01774 -0.0410 0.0252 0.8023
14.250 1.4905 0.02740 0.02083 -0.0490 0.0249 1.0000
14.500 1.5046 0.02820 0.02166 -0.0471 0.0247 1.0000
14.750 1.5183 0.02904 0.02252 -0.0453 0.0245 1.0000
15.000 1.5318 0.02989 0.02341 -0.0435 0.0245 1.0000
15.250 1.5447 0.03080 0.02435 -0.0417 0.0244 1.0000
15.500 1.5575 0.03173 0.02532 -0.0399 0.0242 1.0000
15.750 1.5703 0.03267 0.02630 -0.0382 0.0241 1.0000
16.000 1.5821 0.03370 0.02737 -0.0364 0.0241 1.0000
16.250 1.5938 0.03474 0.02846 -0.0346 0.0240 1.0000
16.500 1.6057 0.03578 0.02953 -0.0330 0.0238 1.0000
16.750 1.6174 0.03685 0.03064 -0.0313 0.0237 1.0000
17.000 1.6289 0.03796 0.03178 -0.0298 0.0235 1.0000
17.250 1.6392 0.03918 0.03306 -0.0281 0.0235 1.0000
17.500 1.6511 0.04028 0.03418 -0.0267 0.0232 1.0000
17.750 1.6607 0.04158 0.03554 -0.0251 0.0232 1.0000
18.000 1.6711 0.04283 0.03682 -0.0237 0.0230 1.0000
18.250 1.6807 0.04417 0.03820 -0.0223 0.0229 1.0000
18.500 1.6898 0.04558 0.03966 -0.0209 0.0229 1.0000
18.750 1.6985 0.04704 0.04117 -0.0195 0.0228 1.0000
19.000 1.7064 0.04859 0.04278 -0.0182 0.0227 1.0000
19.250 1.7127 0.05033 0.04459 -0.0169 0.0226 1.0000
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Polar data table (+)
Polar graphs
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