GRUMMAN K-2 AIRFOIL (k2-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GRUMMAN K-2 AIRFOIL (k2-il) Reynolds number: 100,000 Max Cl/Cd: 25.76 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-k2-il-100000-n5.txt Download as CSV file: xf-k2-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GRUMMAN K-2 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.7526 0.07192 0.06549 -0.0473 1.0000 0.0308
-9.750 -0.7710 0.06799 0.06146 -0.0471 1.0000 0.0307
-9.500 -0.7905 0.06458 0.05794 -0.0456 1.0000 0.0306
-9.250 -0.8104 0.06175 0.05500 -0.0429 1.0000 0.0305
-9.000 -0.8269 0.05878 0.05188 -0.0403 1.0000 0.0304
-8.750 -0.8382 0.05562 0.04849 -0.0380 1.0000 0.0304
-8.500 -0.8454 0.05241 0.04501 -0.0357 1.0000 0.0304
-8.250 -0.8476 0.04929 0.04158 -0.0336 1.0000 0.0304
-8.000 -0.8458 0.04624 0.03819 -0.0316 1.0000 0.0304
-7.750 -0.8400 0.04331 0.03488 -0.0298 1.0000 0.0305
-7.500 -0.8305 0.04056 0.03173 -0.0282 1.0000 0.0307
-7.250 -0.8172 0.03818 0.02906 -0.0269 1.0000 0.0310
-7.000 -0.8010 0.03626 0.02698 -0.0258 1.0000 0.0314
-6.750 -0.7835 0.03451 0.02504 -0.0248 1.0000 0.0319
-6.500 -0.7647 0.03302 0.02342 -0.0238 1.0000 0.0327
-6.250 -0.7452 0.03161 0.02182 -0.0229 1.0000 0.0340
-6.000 -0.7245 0.03012 0.02004 -0.0218 1.0000 0.0355
-5.750 -0.7034 0.02871 0.01847 -0.0209 1.0000 0.0366
-5.500 -0.6822 0.02751 0.01726 -0.0200 1.0000 0.0376
-5.250 -0.6606 0.02639 0.01609 -0.0190 1.0000 0.0388
-5.000 -0.6387 0.02533 0.01494 -0.0180 1.0000 0.0403
-4.750 -0.6164 0.02440 0.01390 -0.0170 1.0000 0.0425
-4.500 -0.5947 0.02348 0.01309 -0.0164 1.0000 0.0450
-4.250 -0.5720 0.02267 0.01224 -0.0157 1.0000 0.0481
-4.000 -0.5490 0.02177 0.01134 -0.0152 1.0000 0.0508
-3.750 -0.5251 0.02103 0.01063 -0.0149 1.0000 0.0554
-3.500 -0.5002 0.02028 0.00989 -0.0149 1.0000 0.0610
-3.250 -0.4745 0.01961 0.00920 -0.0149 1.0000 0.0687
-3.000 -0.4481 0.01889 0.00857 -0.0151 1.0000 0.0815
-2.750 -0.4208 0.01805 0.00799 -0.0155 1.0000 0.1120
-2.500 -0.3906 0.01649 0.00744 -0.0175 1.0000 0.2779
-2.250 -0.3648 0.01513 0.00757 -0.0177 1.0000 0.5419
-2.000 -0.3591 0.01573 0.00863 -0.0110 1.0000 0.6420
-1.750 -0.3404 0.01625 0.00918 -0.0082 1.0000 0.6952
-1.500 -0.3226 0.01703 0.00995 -0.0050 1.0000 0.7380
-1.000 -0.2937 0.01852 0.01138 0.0031 1.0000 0.7854
-0.750 -0.2779 0.01877 0.01160 0.0063 1.0000 0.7935
-0.500 -0.2542 0.01890 0.01165 0.0071 1.0000 0.8041
-0.250 -0.2369 0.01908 0.01180 0.0098 1.0000 0.8137
0.000 -0.2169 0.01919 0.01188 0.0116 1.0000 0.8235
0.250 -0.1952 0.01925 0.01192 0.0128 1.0000 0.8321
0.500 -0.1709 0.01923 0.01188 0.0132 1.0000 0.8381
0.750 -0.1490 0.01920 0.01184 0.0143 1.0000 0.8429
1.000 -0.1225 0.01919 0.01182 0.0140 1.0000 0.8479
1.250 -0.0878 0.01926 0.01189 0.0120 0.9974 0.8519
1.500 -0.0466 0.01937 0.01202 0.0092 0.9890 0.8543
1.750 -0.0027 0.01944 0.01213 0.0058 0.9787 0.8567
2.000 0.0441 0.01938 0.01212 0.0020 0.9649 0.8588
2.250 0.0909 0.01915 0.01198 -0.0016 0.9485 0.8609
2.500 0.1353 0.01871 0.01162 -0.0043 0.9262 0.8630
2.750 0.1788 0.01815 0.01115 -0.0067 0.8999 0.8650
3.000 0.2071 0.01771 0.01082 -0.0062 0.8703 0.8677
3.250 0.2372 0.01725 0.01046 -0.0057 0.8342 0.8700
3.500 0.3306 0.01670 0.00939 -0.0160 0.6390 0.8682
3.750 0.3591 0.01760 0.00940 -0.0157 0.4506 0.8700
4.000 0.3806 0.01833 0.00959 -0.0146 0.3362 0.8727
4.250 0.4041 0.01892 0.00984 -0.0140 0.2609 0.8758
4.500 0.4289 0.01951 0.01015 -0.0139 0.2055 0.8788
4.750 0.4546 0.02007 0.01050 -0.0139 0.1649 0.8813
5.000 0.4770 0.02043 0.01078 -0.0130 0.1382 0.8837
5.250 0.5000 0.02088 0.01109 -0.0122 0.1125 0.8865
5.500 0.5241 0.02139 0.01150 -0.0118 0.0866 0.8891
5.750 0.5492 0.02202 0.01206 -0.0116 0.0644 0.8915
6.000 0.5747 0.02272 0.01270 -0.0115 0.0505 0.8942
6.250 0.5971 0.02342 0.01338 -0.0106 0.0436 0.8967
6.500 0.6189 0.02415 0.01417 -0.0095 0.0390 0.8992
6.750 0.6418 0.02496 0.01504 -0.0089 0.0352 0.9015
7.000 0.6655 0.02583 0.01595 -0.0085 0.0324 0.9038
7.250 0.6896 0.02698 0.01718 -0.0081 0.0308 0.9063
7.500 0.7144 0.02823 0.01861 -0.0078 0.0294 0.9087
7.750 0.7361 0.02941 0.01992 -0.0068 0.0283 0.9111
8.000 0.7585 0.03073 0.02134 -0.0062 0.0275 0.9133
8.250 0.7811 0.03240 0.02324 -0.0057 0.0266 0.9154
8.500 0.8026 0.03429 0.02553 -0.0049 0.0255 0.9177
8.750 0.8230 0.03601 0.02754 -0.0043 0.0243 0.9202
9.000 0.8410 0.03748 0.02917 -0.0034 0.0235 0.9228
9.250 0.8561 0.03940 0.03132 -0.0021 0.0228 0.9253
9.500 0.8669 0.04232 0.03478 -0.0004 0.0226 0.9278
9.750 0.8735 0.04571 0.03869 0.0014 0.0224 0.9303
10.000 0.8752 0.04952 0.04300 0.0033 0.0223 0.9329
10.250 0.8699 0.05335 0.04727 0.0056 0.0223 0.9360
10.500 0.8588 0.05740 0.05170 0.0078 0.0224 0.9396
10.750 0.8424 0.06168 0.05630 0.0094 0.0225 0.9432
11.000 0.8224 0.06674 0.06164 0.0093 0.0226 0.9463
11.250 0.8003 0.07273 0.06786 0.0067 0.0227 0.9493
11.500 0.7807 0.08035 0.07565 0.0006 0.0229 0.9521
|
Polar data table (+)
Polar graphs
<< Back to GRUMMAN K-2 AIRFOIL (k2-il)