VLM Based Stability & Control Evaluation Code icase = 3 MACH NUMBER = 0.20000 Sref = 400.00000 c ref = 11.70000 b ref = 36.00000 HEIGHT ABOVE GROUND = 8.00000 X-cg = 32.40000 Z-cg = 1.00000 CL0 = -0.07784 Cm0 = 0.01807 Cdi(at alpha = 0) = 0.00047 CL-alpha = 4.93908 Cm-alpha = -0.40949 Cm/CL = -0.08291 Cdi/CL^2 = 0.06005 CL (at alpha = 5) = 0.35318 Cdi(at alpha = 5) = 0.00749 Lift-curve slope of tail due to downwash of wing = 0.48045 CL-q = 8.88228 CM-q = -5.86071 CL-delta [2 1] = 0.07212 Cm-delta [2 1] = 0.14026 CL-delta [2 2] = 1.15303 Cm-delta [2 2] = 0.12812 CL-delta [3 1] = 0.03362 Cm-delta [3 1] = 0.01560 CL-delta [3 2] = 0.67143 Cm-delta [3 2] = -0.33240 CL-delta [4 1] = 0.94745 Cm-delta [4 1] = -1.13186 Cl-delta [2 1] = -0.00830 Cn-delta [2 1] = -0.00146 Cl-delta [2 2] = -0.17614 Cn-delta [2 2] = -0.00436 Cl-delta [3 1] = -0.00661 Cn-delta [3 1] = -0.00030 Cl-delta [3 2] = -0.17322 Cn-delta [3 2] = -0.00054 Cl-delta [4 1] = -0.10780 Cn-delta [4 1] = 0.02525 Cy-beta = -0.52457 Cn-beta = 0.07900 Cl-beta = -0.10010 Cy-r = 0.46587 Cn-r = -0.18577 Cl-r = 0.08320 Cl-p = -0.43447 Cn-p = -0.11396 Cy-delta [2 2] = 0.10781 Cl-delta [2 2] = 0.02099 Cn-delta [2 2] = -0.04148 Cy-delta [4 2] = 0.10821 Cl-delta [4 2] = 0.02097 Cn-delta [4 2] = -0.04219