We are going to do some analysis of the wing used on the Global Hawk RQ$4$B UAV. The Global Hawk usually operates at high altitude and flies at high subsonic Mach numbers. However, since we have not covered compressible flow yet, we will assume the Global Hawk is flying at low speeds at sea level. The specifications we will use are the following:

Altitude: sea level $($assume standard atmospheric conditions$)$

Wing area S $=685$ ft$2$

Wing span b $=130\mathrm{.}9$ ft

Cruising air speed Ucruise $=195$ knots $($about Mach $0\mathrm{.}3,$ so that the flow is incompressible$)$

Landing speed Ulanding $=85$ knots

Empty weight Mempty $=14,950$ lb

Fuel capacity Mfuel $=17,300$ lb

Maximum take$-$off gross weight Mmto $=32,250$ lb

AR $=25\mathrm{.}014$

$($i$)$ C$\_$L in cruise $=\mathrm{.}366$

$($ii$)$ C$\_$L in landing $=1\mathrm{.}925$

alpah$\_$zl $=-5\mathrm{.}3$

C$\_$L$(0)=\mathrm{.}597$

Assume the wing is elliptically loaded and has no twist.

$$ Determine the induced angle of attack, $\backslash$alpha i$,$ and the induced drag coefficient, CDi, for the two conditions $($i$)$ and $($ii$).$ Comment on the relative magnitude of the induced drag coefficient at landing versus cruise conditions.

now only consider case $($i$)$

$$ Determine the wing lift curve slope, and the wing angle of attack, $\backslash$alpha $,$ needed to produce the required lift coefficient.

$$ Use the drag polar found in part $($c$)$ to estimate the drag coefficient, cd$,$ for the airfoil section. If we approximate the wing drag coefficient as CD $=$ CD$0+$ CDi $,$ and make the approximation that CD$0=$ cd $($based on assuming cd$($y$)$ is constant across the span of the wing$),$

$$ determine the total wing drag coefficient, CD$,$ and the ratio of the induced drag to the total drag, CDi$/$CD$.$ Comment on the significance of the induced drag at these

conditions.