Need help to solve this matlab question. Need to find kn .

Fyi in the program below ( bold fid = fopen('F15_01_00','r') ) . This file contain this information.

Formula of Kn shown below

% Section 1 : read in aircraft data %% m=20411.66; %mass in kg; s=56.485; %wing area in m2; b=13.045; %wing span m c=4.862; %mac m Ix=38911.97517; %Ixx kg-m3 Iy=223845.5435; %Iyy kg-m2 Iz=254758.1928; %Izz kg-m2 Ixz=-705.025334; %Ixz kg-m2 fid = fopen('F15_01_00','r') % %

ma=fscanf(fid,'%*s %f',1) %Mach number h=fscanf(fid,'%*s %f',1) %altitude ft v0=fscanf(fid,'%*s %f',1) %velocity m/s q=fscanf(fid,'%*s %f',1) %dynamic pressure theta0=fscanf(fid,'%*s %f',1) % trimed theta degree % cd_0=fscanf(fid,'%*s %f',1) %CD0 cl_0=fscanf(fid,'%*s %f',1) %CL0 cd_u=fscanf(fid,'%*s %f',1) %Cdu cl_u=fscanf(fid,'%*s %f',1) %CLu cm_u=fscanf(fid,'%*s %f',1) %cmu

cd_alpha=fscanf(fid,'%*s %f',1) %clapha per deg cl_alpha=fscanf(fid,'%*s %f',1) %cdalpha per deg cm_alpha=fscanf(fid,'%*s %f',1) %cmalpha per deg

cz_alpha_dot=fscanf(fid,'%*s %f',1) cm_alpha_dot=fscanf(fid,'%*s %f',1)

cz_q=fscanf(fid,'%*s %f',1) %clq cm_q=fscanf(fid,'%*s %f',1) %cmq

cx_de=fscanf(fid,'%*s %f',1) %cdde cz_de=fscanf(fid,'%*s %f',1) %clde cm_de=fscanf(fid,'%*s %f',1) %cmde

CY_beta =fscanf(fid,'%*s %f',1); Cl_beta =fscanf(fid,'%*s %f',1); CN_beta =fscanf(fid,'%*s %f',1); CY_p =fscanf(fid,'%*s %f',1); Cl_p =fscanf(fid,'%*s %f',1); CN_p =fscanf(fid,'%*s %f',1); CY_r =fscanf(fid,'%*s %f',1); Cl_r =fscanf(fid,'%*s %f',1); CN_r =fscanf(fid,'%*s %f',1); CY_delta_a =fscanf(fid,'%*s %f',1); CY_delta_r =fscanf(fid,'%*s %f',1); Cl_delta_a =fscanf(fid,'%*s %f',1); Cl_delta_r =fscanf(fid,'%*s %f',1); CN_delta_a =fscanf(fid,'%*s %f',1); CN_delta_r =fscanf(fid,'%*s %f',1);

% fclose(fid); % % g=9.81; pi=3.1415926; theta0=theta0*pi/180; % Section 1a : Compute the longitudinal stability margin kn=

M Alt(ft) U(m/s) q(N/m2) alpha_trim CDO CLO CDu CLu CMu CDalpha CLalpha CMalpha CZ_alfa_dot CM_alfa_dot 0.5 1000.00000 169.559 17113.18747 0.18105 0.01468 0.20708 0.00E+00 3.73E-06 -3.53E-06 0.37257 4.8706 -0.268819 -17.2322 -11.887 17.2322 3.8953 -0.00438308 -0.57295 -0.69281 -0.97403 -0.13345 0.12996 CZY q CX_de CZ_de CM_de CY_beta CL_beta CN_beta CY_P CL_P CN_P CY_r CL_r CN_r CY_da CY_dr CL_da CL_dr CN_da CN_dr -0.2 -0.033721 0 0.1099 -0.40471 -1.15E-03 -0.15041 0.026356 -2.39E-03 0.0021917 -0.0697631 Calculate the longitudinal static stability margin of the aircraft. Is the aircraft statically stable? (5 marks) Criteria/Condition for Longitudinal Static Stability From Equation (2.12): de dn dn dCm a1 da = (h ho) v, (3 (1 - de a3 + az dClw dClw aC Lw (2.13) For static stability :