BOEING-VERTOL VR-12 AIRFOIL (vr12-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL VR-12 AIRFOIL (vr12-il) Reynolds number: 500,000 Max Cl/Cd: 73.05 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-vr12-il-500000.txt Download as CSV file: xf-vr12-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL VR-12 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5045 0.09470 0.09259 -0.0106 1.0000 0.0201
-9.250 -0.5066 0.08827 0.08618 -0.0162 1.0000 0.0201
-9.000 -0.5092 0.08132 0.07926 -0.0235 1.0000 0.0202
-8.750 -0.5157 0.07430 0.07219 -0.0307 1.0000 0.0202
-8.500 -0.5221 0.06837 0.06617 -0.0337 1.0000 0.0202
-8.250 -0.5254 0.06260 0.06026 -0.0357 1.0000 0.0202
-8.000 -0.5550 0.05001 0.04738 -0.0367 0.9946 0.0165
-7.750 -0.5369 0.04740 0.04467 -0.0383 0.9873 0.0167
-7.500 -0.5280 0.04158 0.03854 -0.0387 0.9779 0.0153
-7.250 -0.5627 0.02485 0.02015 -0.0307 0.9607 0.0168
-7.000 -0.5401 0.02532 0.02070 -0.0298 0.9529 0.0172
-6.750 -0.5198 0.02546 0.02083 -0.0284 0.9456 0.0179
-6.500 -0.5049 0.02258 0.01737 -0.0256 0.9372 0.0199
-6.250 -0.4888 0.02083 0.01534 -0.0234 0.9305 0.0210
-6.000 -0.4644 0.02078 0.01534 -0.0229 0.9236 0.0219
-5.750 -0.4417 0.02057 0.01506 -0.0218 0.9175 0.0235
-5.500 -0.4198 0.01933 0.01344 -0.0202 0.9115 0.0255
-5.250 -0.3961 0.01880 0.01298 -0.0197 0.9050 0.0268
-5.000 -0.3728 0.01836 0.01247 -0.0186 0.8999 0.0283
-4.750 -0.3469 0.01733 0.01125 -0.0179 0.8937 0.0294
-4.500 -0.3213 0.01703 0.01076 -0.0170 0.8882 0.0307
-4.250 -0.2974 0.01533 0.00893 -0.0161 0.8838 0.0321
-4.000 -0.2706 0.01460 0.00820 -0.0158 0.8780 0.0332
-3.750 -0.2451 0.01407 0.00762 -0.0151 0.8717 0.0346
-3.500 -0.2194 0.01360 0.00709 -0.0144 0.8641 0.0362
-3.250 -0.1939 0.01308 0.00650 -0.0135 0.8557 0.0373
-3.000 -0.1675 0.01269 0.00605 -0.0129 0.8479 0.0381
-2.750 -0.1434 0.01182 0.00512 -0.0119 0.8401 0.0395
-2.500 -0.1183 0.01130 0.00460 -0.0112 0.8317 0.0415
-2.250 -0.0926 0.01095 0.00421 -0.0105 0.8233 0.0432
-2.000 -0.0662 0.01062 0.00386 -0.0100 0.8146 0.0448
-1.750 -0.0400 0.01038 0.00355 -0.0093 0.8066 0.0466
-1.500 -0.0131 0.01009 0.00323 -0.0089 0.7971 0.0485
-1.250 0.0130 0.00978 0.00288 -0.0083 0.7893 0.0520
-1.000 0.0403 0.00957 0.00266 -0.0080 0.7803 0.0567
-0.750 0.0668 0.00929 0.00244 -0.0075 0.7718 0.0801
-0.500 0.0785 0.00734 0.00218 -0.0050 0.7635 0.5620
-0.250 0.1110 0.00626 0.00236 -0.0051 0.7544 0.9216
0.000 0.1661 0.00662 0.00265 -0.0103 0.7421 0.9614
0.250 0.2316 0.00715 0.00308 -0.0177 0.7269 0.9861
0.500 0.2992 0.00733 0.00314 -0.0259 0.7047 0.9980
0.750 0.3345 0.00734 0.00303 -0.0275 0.6806 1.0000
1.000 0.3614 0.00737 0.00291 -0.0274 0.6458 1.0000
1.250 0.3888 0.00754 0.00280 -0.0273 0.5850 1.0000
1.500 0.4169 0.00800 0.00280 -0.0277 0.4888 1.0000
1.750 0.4445 0.00851 0.00289 -0.0281 0.4041 1.0000
2.000 0.4713 0.00893 0.00298 -0.0283 0.3429 1.0000
2.250 0.4976 0.00922 0.00307 -0.0283 0.3095 1.0000
2.500 0.5237 0.00943 0.00316 -0.0281 0.2876 1.0000
2.750 0.5497 0.00962 0.00325 -0.0279 0.2713 1.0000
3.000 0.5755 0.00981 0.00335 -0.0277 0.2583 1.0000
3.250 0.6011 0.01001 0.00346 -0.0274 0.2469 1.0000
3.500 0.6266 0.01017 0.00357 -0.0270 0.2365 1.0000
3.750 0.6520 0.01034 0.00370 -0.0267 0.2276 1.0000
4.000 0.6769 0.01055 0.00384 -0.0262 0.2190 1.0000
4.250 0.7018 0.01071 0.00398 -0.0257 0.2120 1.0000
4.500 0.7263 0.01092 0.00415 -0.0252 0.2048 1.0000
4.750 0.7509 0.01111 0.00433 -0.0247 0.1986 1.0000
5.000 0.7752 0.01131 0.00449 -0.0241 0.1922 1.0000
5.250 0.7991 0.01157 0.00472 -0.0235 0.1859 1.0000
5.500 0.8235 0.01173 0.00489 -0.0229 0.1804 1.0000
5.750 0.8471 0.01199 0.00510 -0.0222 0.1749 1.0000
6.000 0.8705 0.01226 0.00537 -0.0215 0.1704 1.0000
6.250 0.8945 0.01246 0.00560 -0.0208 0.1664 1.0000
6.500 0.9177 0.01272 0.00584 -0.0201 0.1622 1.0000
6.750 0.9399 0.01309 0.00618 -0.0192 0.1573 1.0000
7.000 0.9638 0.01326 0.00640 -0.0186 0.1540 1.0000
7.250 0.9869 0.01351 0.00667 -0.0178 0.1507 1.0000
7.500 1.0093 0.01382 0.00697 -0.0169 0.1475 1.0000
7.750 1.0300 0.01431 0.00743 -0.0158 0.1436 1.0000
8.000 1.0533 0.01450 0.00770 -0.0151 0.1416 1.0000
8.250 1.0759 0.01476 0.00801 -0.0142 0.1390 1.0000
8.500 1.0980 0.01504 0.00833 -0.0134 0.1364 1.0000
8.750 1.1194 0.01539 0.00868 -0.0124 0.1337 1.0000
9.000 1.1390 0.01592 0.00919 -0.0111 0.1304 1.0000
9.250 1.1610 0.01617 0.00952 -0.0103 0.1281 1.0000
9.500 1.1832 0.01637 0.00980 -0.0094 0.1250 1.0000
9.750 1.2045 0.01663 0.01008 -0.0085 0.1215 1.0000
10.000 1.2234 0.01710 0.01053 -0.0072 0.1177 1.0000
10.250 1.2439 0.01742 0.01093 -0.0061 0.1149 1.0000
10.500 1.2651 0.01766 0.01126 -0.0052 0.1118 1.0000
10.750 1.2850 0.01797 0.01160 -0.0040 0.1081 1.0000
11.000 1.3009 0.01856 0.01216 -0.0024 0.1037 1.0000
11.250 1.3219 0.01877 0.01249 -0.0014 0.1007 1.0000
11.500 1.3408 0.01909 0.01287 -0.0002 0.0962 1.0000
11.750 1.3548 0.01970 0.01345 0.0017 0.0911 1.0000
12.000 1.3741 0.01999 0.01385 0.0028 0.0869 1.0000
12.250 1.3873 0.02050 0.01437 0.0049 0.0821 1.0000
12.500 1.3990 0.02109 0.01501 0.0071 0.0777 1.0000
12.750 1.4125 0.02167 0.01561 0.0088 0.0729 1.0000
13.000 1.4231 0.02248 0.01644 0.0108 0.0682 1.0000
13.250 1.4358 0.02322 0.01723 0.0124 0.0640 1.0000
13.500 1.4439 0.02432 0.01832 0.0143 0.0595 1.0000
13.750 1.4550 0.02526 0.01935 0.0157 0.0563 1.0000
14.000 1.4631 0.02645 0.02058 0.0173 0.0529 1.0000
14.250 1.4646 0.02818 0.02233 0.0192 0.0494 1.0000
14.500 1.4761 0.02923 0.02352 0.0202 0.0473 1.0000
14.750 1.4825 0.03071 0.02508 0.0213 0.0447 1.0000
15.000 1.4838 0.03269 0.02712 0.0225 0.0424 1.0000
15.250 1.4827 0.03498 0.02950 0.0235 0.0403 1.0000
15.500 1.4901 0.03659 0.03124 0.0241 0.0388 1.0000
15.750 1.4935 0.03862 0.03338 0.0246 0.0372 1.0000
16.000 1.4931 0.04114 0.03599 0.0249 0.0357 1.0000
16.250 1.4882 0.04425 0.03919 0.0249 0.0344 1.0000
16.500 1.4766 0.04834 0.04339 0.0244 0.0333 1.0000
16.750 1.4719 0.05184 0.04703 0.0236 0.0325 1.0000
17.000 1.4654 0.05580 0.05113 0.0225 0.0317 1.0000
17.250 1.4552 0.06054 0.05602 0.0207 0.0310 1.0000
17.500 1.4402 0.06630 0.06192 0.0182 0.0304 1.0000
17.750 1.4202 0.07310 0.06888 0.0151 0.0301 1.0000
18.000 1.3962 0.08080 0.07673 0.0114 0.0298 1.0000
18.250 1.3690 0.08909 0.08518 0.0075 0.0297 1.0000
18.500 1.3411 0.09753 0.09376 0.0036 0.0296 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL VR-12 AIRFOIL (vr12-il)