Induced drag is created due to downwash. Downwash is produced by the wing-tip vortices  during flight. Aspect ratio of the wing is,

$$AR=\frac{b^{2}}{S}=\frac{\left ( 10 \right )^{2}}{15}=6.67$$

At, standard sea level density of air is  $1.225\,kg/m^{3}$

Velocity of airplane is $50\,m/s$

Dynamic pressure,

$$q_{\infty} = \frac{1}{2}\rho _{\infty}V_{\infty}^{2}=\frac{1}{2}\times1.225\times\left ( 50 \right )^{2}=1531.25\,N/m^{2}$$

Lift slop of the airfoil of the wing is $0.102 \,per\,degree = 5.8442\,per\,radian$

Lift produced by the wing is,

$$L=\frac{1}{2}\rho _{\infty}V_{\infty}^{2}SC_{L}=\rho _{\infty}SC_{L}$$ $$\Rightarrow C_{L}=\frac{L}{\rho _{\infty}S}$$

L = Weight of the airplane = $1000\times9.8 = 9800\,N$

Therefore,

$$C_{L} = \frac{9800}{1531.25\times15}=0.427$$

Lift slope of the wing will be,

$$a = \frac{a_{0}}{1 + \frac{a_{0}}{\pi AR}\left ( 1+\tau \right ) }=\frac{5.8442}{1+\frac{5.8442}{\pi\left ( 6.67 \right )}\left ( 1+0.1 \right )}=4.472\,per\,radian = 0.0781\,per\,degree$$

Also, Coefficient of lift is given as

$$C_{L}=a\left ( \alpha -\alpha _{L=0} \right )\Rightarrow 0.427=0.0781\left ( \alpha -\left ( -2.5^{\circ} \right ) \right )\Rightarrow \alpha = 2.97\,degree$$

Induced drag is given as

$$D_{i}=q_{\infty}SC_{D,i}$$ $$C_{D,i}=\frac{C_{L}^{2}}{\pi e AR}=\frac{\left ( 0.427 \right )^{2}}{\pi \left ( 0.6 \right )\left ( 6.67 \right )}=0.0145$$

Therefore, $D_{i} = \left ( 1531.25 \right )\left ( 15 \right )\left ( 0.0145 \right ) = 333.05\,N$