For this assignment, you will use the same aircraft from the Module 1 and 3 homework assignments. You may work in either the SI or BGS system, but you must be consistent. 1. Using the same aircraft from the Module 1 and 3 homework assignments, provide the following a. The aircraft b. The cruise altitude and the corresponding standard atmosphere values at that altitude c. The wing area (if you have a multi-wing aircraft, only provide the information for one wing) d. The average chord e. The wing span f. The cruise speed g. The takeoff weight 2. Calculate the Reynolds number based on the cruise speed, cruise altitude, and average chord. 3. Calculate the aspect ratio AR. 4. Assume that the airfoil lift curve slope ao is 0.11 /deg and the wing efficiency e is 0.92. Calculate the wing lift curve slope. 5. If the aspect ratio was 25% larger, calculate the new wing lift slope. If the aspect ratio was 25% smaller, calculate the new wing lift slope. Explain the effect of aspect ratio on the wing lift slope. 6. Assume that the airfoil used for your wing is the NACA 23012. From the airfoil data in the appendix and using the Reynolds number that most closely matches the Reynolds number from question 2, provide the following a. The airfoil L=0 b. The airfoil Cmax c. The airfoil a stall 7. In Excel, replot the profile drag coefficient of the airfoil ca for the NACA 23012 from a lift coefficient of 0 to the maximum in the drag polar plot. All you are doing is reading off the Ca values from the airfoil plot, putting them in Excel, and replotting them. 8. Calculate and plot the induced drag coefficient CD for C = 0 to the maximum C, used in the previous question. Assume that e = e. Plot this curve on the same figure as the previous question. For this problem, we are going to assume that the in problem 7 (the airfoil c) is the same as the CL (wing CL) used in this problem. 9. Calculate and plot the total drag coefficient of the wing. Plot this curve on the same figure as the previous question. Just like problem 8, we are going to assume that the c in problem 7 (the airfoil c) is the same as the CL (wing C) used in problem 8. of the airfoil. Calculate 10. Assume that the CL max of your wing is 90% of the Cmax a. CLmax of the wing b. Stall speed of your aircraft at takeoff weight and at standard sea level conditions 11. Figure 5.68 in the book shows an example of how flaps change the lift curve. Use this figure to answer the following a. What is the ACL between no flaps and 8 = 15 at = 0? b. What is the AC between no flaps and 8 = 50 at = 0? C. Assume that these ACLs hold for your aircraft. What is the stall speed of your aircraft at 15 flaps and 50 flaps at takeoff weight and at standard sea level conditions? Hint: CLmax (15% flap) = CLmax (question 10) + ACL (15% flap) 12. Assume a rectangular wing has a Mer of 0.66. Plot the theoretical upper limit Mcr versus sweep angle for sweep angles from 0 deg to 60 deg by 5 deg increments. Explain the major penalties of sweeping wings. 13. For this question, you will be comparing a cylinder to the NACA 23012. For a cylinder, Ca = 1.0 for laminar boundary layer and c = 0.25 for turbulent boundary layer. The critical Reynolds number for a cylinder is 300,000. You must work with BGS units. The Reynolds number of a cylinder: Re = pVd The drag of a cylinder: D = pv2dcd The d is the cylinder diameter a. For a cylinder at sea level with a diameter of 18 in. that is traveling at 150 ft/s, what is the Reynolds number? Is the boundary layer laminar or turbulent? What is ca for this cylinder. b. Calculate the drag of the cylinder. c. From question 7, what is the minimum c of the NACA 23012? d. Using the minimum ca of the NACA 23012, what is the drag of the airfoil that is at sea level traveling at the same velocity of the cylinder and has the same maximum physical thickness of the cylinder? The NACA 23012 has a maximum thickness percentage of 12%. The airfoil and the cylinder will have the same maximum physical thickness. e. What is the chord and physical thickness of a NACA 23012 airfoil that has the same drag as the cylinder found in part b? f. From parts d and e, what can you say about the importance of streamlining? 396 -10 CHAPTER 5 Airfoils, Wings, and Other Aerodynamic Shapes CL 2.5 8 2.0 1.5 1.0 0.5 = 0 1=0 With flaps No flaps 8 = 50 S 8 = 15 8=0 With flaps No flaps 1 10 15 20 a (degrees) Figure 5.68 Illustration of the effect of flaps on the lift curve. The numbers shown are typical of a modern medium- range jet transport. 11,900 m Density 3.1687-10 kg/m 1.) a.) Airbus 0320 b.) Temp 216.66 k Typical Crusing Altitud 1.9706.10 N/m Pressure c.) Wing Area (5) (5) 122.6m d.) Carg = 3.42 Caug $ c.) Wing Span (6) = 35.8 m F.) v = 233.33 m/s 5.) Take off weight = 774725. 35 N 2) M = Mo To (+)" (+++C) = 288.16 M= 1.7894 10's by. 1.7894 x 105 kg. M.S 216.66 KIS M.S 288.16 = 0.535057 Rel= e. V.Cag = her= 472.58 264 C = 110.56 k " 216.66 K + 110.56K 288.16 K + 110.56k -1 3.168710 kg/m 233.33 m/s 3.43m 0.535057 3.) AR = 2 (35.8 m) 122.6 m 4.) 90 = 0.11/deg a = e = 0.92 a = 1+ = 10.45383 ao 57.3 20 , 0.11/deg 1+ 57.3.0.11/deg TT 0.92 10.45383 5.) AR 25% larger new AR = AR + 25% AR = 10.45383 + AR = 13.067 0.25 10.45383 ' (new) a = 0.11 1+ 53.7.0.11 . 13.067 All 25% smaller AR = 2 (new) a = - 10.45383 25% 10.45383 0.1 1+ 53.7-0.11 7.84 Ideg = 7.84