A light aircraft with a mass of 6,360 kg is flown at constant altitude at a...

  

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A light aircraft with a mass of 6,360 kg is flown at constant altitude at a constant speed of 250 km/h. The aircraft's wings are based on the NACA 23015 aerofoil section with a total plan area of 22 m2 and an aspect ratio of 10 (based on overall wing span and total plan area). The variations of lift and drag coefficients with angle of attack for the 2-Dimensional aerofoil section are shown in Figure 1. Assume that the aircraft's wings provide all the required lift. Use: p= 1.2 kg/m3 and v= 1.50 x 10-5 m²/s. For the above flight condition, determine: (a) the lift generated by the aircraft's wings and hence the lift coefficient of the wings; (b) the drag on the wings, including the induced drag; (c) the power supplied by the aircraft's propulsion system if the drag on the remainder of the aircraft is 420 N; (d) the geometric angle of attack of the wings; (e) any assumptions that were required to solve this problem. [3 marks] [6 marks] [2 marks] [3 marks] [1 mark ] 18 CLra = 1.72 1.6 1.4 CL 1.0 0.020 0.8 0.016 Cp 0.6- 0.012 NACA 23015 0.4- 0.008 000 0.2 0.004 NACA 23015 8 12 16 20 4 8 12 16 20 Angie of attack, a (deg) Angle of attack, a (deg) (a) Lift Coefficient (b) Drag Coefficient Figure 1. Lift and drag coefficient curves for the 2-Dimensional (infinite span) NACA 23015 section wing. 2. A light aircraft with a mass of 6,360 kg is flown at constant altitude at a constant speed of 250 km/h. The aircraft's wings are based on the NACA 23015 aerofoil section with a total plan area of 22 m2 and an aspect ratio of 10 (based on overall wing span and total plan area). The variations of lift and drag coefficients with angle of attack for the 2-Dimensional aerofoil section are shown in Figure 1. Assume that the aircraft's wings provide all the required lift. Use: p= 1.2 kg/m3 and v= 1.50 x 10-5 m²/s. For the above flight condition, determine: (a) the lift generated by the aircraft's wings and hence the lift coefficient of the wings; (b) the drag on the wings, including the induced drag; (c) the power supplied by the aircraft's propulsion system if the drag on the remainder of the aircraft is 420 N; (d) the geometric angle of attack of the wings; (e) any assumptions that were required to solve this problem. [3 marks] [6 marks] [2 marks] [3 marks] [1 mark ] 18 CLra = 1.72 1.6 1.4 CL 1.0 0.020 0.8 0.016 Cp 0.6- 0.012 NACA 23015 0.4- 0.008 000 0.2 0.004 NACA 23015 8 12 16 20 4 8 12 16 20 Angie of attack, a (deg) Angle of attack, a (deg) (a) Lift Coefficient (b) Drag Coefficient Figure 1. Lift and drag coefficient curves for the 2-Dimensional (infinite span) NACA 23015 section wing. 2.