Compute the following geometric properties of the wing box: a. Cross-section Area b. Centroid c. Moments...

 

 

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Compute the following geometric properties of the wing box: a. Cross-section Area b. Centroid c. Moments of inertia of cross-section Iyy, I_zz and Lyz about the centroid all the given info is listed below Given: At an operating RPM of 225 rpm each blade has to generate 8,700 lbs of lift. The blade has a length of 30 ft from center of rotation to tip. The non-tapered lifting airfoil on the blade extends from 4 ft to 30 ft. The chord length is 32 inches. The aerodynamic center of pressure is at 27.8% of chord length(x/c) from leading edge along the length and 2% of chord length along thickness (in level flight). For shear force and bending moment calculations assume the blade attachment to the hub at is clamped Assume drag loads are proportional to the lift (0.7%) The blade cross-section uses a NACA 8-H-12 airfoil. All bending loads are carried by the center wing box (figure 1), that is be approximated by a hexagonal section for preliminary analysis and design. The center wing box is a hollow thin wall multicell beam structure. Figure 1 shows the cross-section of the structural wing box that carries the axial, bending and torsion loads. For preliminary analysis we will neglect curvatures of the skin. The leading edge, trailing edge and center spars are located at 10%, 70% and 30% of the chord length from the leading edge in the chord-wise direction. For the preliminary sizing of the wing assume that only the wing box structure (wing skins and spar and stringers) carry the wing loads. For the preliminary analysis assume the wing skin thickness is 0.1 inch. Spar web thickness is 0.25 inch, stringers that act as corner braces to attach wing skin and spar webs are all 2.5 in^2 in area. The spars and ribs are made from Al6061-T6 alloy, and the wing skin is Al7075-T76 alloy. Figure 1: A D Astr B E Z C F Compute the following geometric properties of the wing box: a. Cross-section Area b. Centroid c. Moments of inertia of cross-section Iyy, I_zz and Lyz about the centroid all the given info is listed below Given: At an operating RPM of 225 rpm each blade has to generate 8,700 lbs of lift. The blade has a length of 30 ft from center of rotation to tip. The non-tapered lifting airfoil on the blade extends from 4 ft to 30 ft. The chord length is 32 inches. The aerodynamic center of pressure is at 27.8% of chord length(x/c) from leading edge along the length and 2% of chord length along thickness (in level flight). For shear force and bending moment calculations assume the blade attachment to the hub at is clamped Assume drag loads are proportional to the lift (0.7%) The blade cross-section uses a NACA 8-H-12 airfoil. All bending loads are carried by the center wing box (figure 1), that is be approximated by a hexagonal section for preliminary analysis and design. The center wing box is a hollow thin wall multicell beam structure. Figure 1 shows the cross-section of the structural wing box that carries the axial, bending and torsion loads. For preliminary analysis we will neglect curvatures of the skin. The leading edge, trailing edge and center spars are located at 10%, 70% and 30% of the chord length from the leading edge in the chord-wise direction. For the preliminary sizing of the wing assume that only the wing box structure (wing skins and spar and stringers) carry the wing loads. For the preliminary analysis assume the wing skin thickness is 0.1 inch. Spar web thickness is 0.25 inch, stringers that act as corner braces to attach wing skin and spar webs are all 2.5 in^2 in area. The spars and ribs are made from Al6061-T6 alloy, and the wing skin is Al7075-T76 alloy. Figure 1: A D Astr B E Z C F