You now have a wing design and now you need to determine the c.g of the wing. We are going to assume the wing has a uniform thickness thereby allowing for an easier calculation of the C.G. or centroid of the tapered wing. As an example: If the tapered wing is broken down into components, you have 2 triangles and one rectangle, area A, B and C. Each area has their own C.G., but the combination of the area will result in a different centroid. In this example, the tapered wing has leading edge and tailing edge sweep angles the same. Toe determine the centroid, we need to work through each area (triangles and rectangle) and ten sum up the areas and average distances. The table below summarizes the calculations. This will be reviewed in class. The X-axis is horizontal and the y-axis is vertical: axis Aerodynamics Used for Centroid Example 10-f B 5-ft 2.5 X-axis 20-ft The centroid or C.G. can be determined: Segment Area x-bar * y-bar A B x-barly-bar Area Area 100 Total 150 2825 6.7 9.2 167.5 230 10 5 1000 500 6.7 1.7 167.5 42.5 1335 772.5 C.G. Location Y-bar 5 X-bar 9 Where Y-bar = 772.5/150 = 5.1 and X-bar = 1335/150 = 8.9 1. Following this example and what was described in class, determine the centroid for your tapered wing. 2. Using this formula, calculate the MAC for your wing: M.A.C. = 2/3*(Cr+Ct - [(Cr * Ct)/Cr + Ct)] where Cr = Root chord length and Ct = Tip Chord length