For the airfoil in Problem 4.11, calculate the value of the circulation around the airfoil.

Data from Problem 4.11:

Consider again the NACA 2412 airfoil discussed in Problem 4.10. The airfoil is flying at a velocity of 60 m/s at a standard altitude of 3 km (see Appendix D). The chord length of the airfoil is 2 m. Calculate the lift per unit span when the angle of attack is 4°. 

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Altitude hg, m h, m -5,000 -5,004 -4,900 -4,904 -4,800 -4,804 -4,700 -4,703 -4,600 -4,603 -4,500 -4,503 -4,400 -4,403 -4,300 -4,303 -4,200 -4,203 -4,100 -4,103 -4,000 -4,003 -3,900 -3,902 -3,800 -3,802 -3,700 -3,702 -3,600 -3,602 -3,500 -3,502 -3,400 -3,402 -3,300 -3,302 -3,200 -3,202 -3,100 -3,102 -3,000 -3,001 -2,900 -2,901 -2,800 -2,801 -2,700 -2,701 -2,600 -2,601 -2,500 -2,501 -2,400 -2,401 -2,300 -2,301 -2,200 -2,201 -2,100 -2,101 -2,000 -2,001 -1,900 -1,901 -1,800 -1,801 -1,700 -1,701 -1,600 -1,600 -1,500 -1,500 -1,400 -1,400 -1,300 -1,300 -1,200 -1,200 -1,100 -1,100 -1,000 -900 -800 -700 -600 -500 -400 -300 -200 -100 - 1,000 -900 -800 -700 -600 -500 -400 -300 -200 -100 Temperature T, K 320.69 320.03 319.38 318.73 318.08 317.43 316.78 316.13 315.48 314.83 314.18 313.53 312.87 212.22 311.57 310.92 310.27 309.62 308.97 308.32 307.67 307.02 306.37 305.72 305.07 304.42 303.77 303.12 302.46 301.81 301.16 300.51 299.86 299.21 298.56 297.91 297.26 296.61 295.96 295.31 294.66 294.01 293.36 292.71 292.06 291.41 290.76 290.11 289.46 288.81 Pressure p, N/m² 1.7761 +5 1.7587 1.7400 1.7215 1.7031 1.6848 1.6667 1.6488 1.6311 1.6134 1.5960 +5 1.5787 1.5615 1.5445 1.5277 1.5110 1.4945 1.4781 1.4618 1.4457 1.4297 +5 1.4139 1.3982 1.3827 1.3673 1.3521 1.3369 1.3220 1.3071 1.2924 1.2778 +5 1.2634 1.2491 1.2349 1.2209 1.2070 1.1932 1.1795 1.1660 1.1526 1.1393 +5 1.1262 1.1131 1.1002 1.0874 1.0748 1.0622 1.0498 1.0375 1.0253 Density p, kg/m³ 1.9296 +0 1.9145 1.8980 1.8816 1.8653 1.8491 1.8330 1.8171 1.8012 1.7854 1.7698 +0 1.7542 1.7388 1.7234 1.7082 1.6931 1.6780 1.6631 1.6483 1.6336 1.6189 +0 1.6044 1.5900 1.5757 1.5615 1.5473 1.5333 1.5194 1.5056 1.4918 1.4782 +0 1.4646 1.4512 1.4379 1.4246 1.4114 1.3984 1.3854 1.3725 1.3597 1.3470 +0 1.3344 1.3219 1.3095 1.2972 1.2849 1.2728 1.2607 1.2487 1.2368