COMMAND GUIDANCE SYSTEMS SIMULATION : AIRFRAME ANALYSIS

Abstract

A complete six-degrees of freedom flight simulation model for anti-aircraft command guided missile system is developed. A computer code that solves the model is constructed with BORLANDC. Modular concept is considered in the code development. The non-linear differential equations that describe the model are solved by Runge-kutta 4 method. The integration step is chosen small enough that the numerical errors are negligible. The aerodynamic non-linear coefficients that describe the missile airframe are calculated by the aid of standard NASA curves. The missile is roll angle stabilized throughout the flight. It is controlled in the lateral planes via two pairs of rear control fins. The pitch and yaw control channels are identical except for constant gravity bias added to the pitch channel. The missile airframe is deeply investigated. The step response of the airframe to unit step fin deflection is obtained. The results indicate that the airframe can be accurately represented by a second order lag system. The weathercock natural frequency and the damping coefficient are obtained by transforming the time-domain data to the frequency domain by the Fast Fourier Transform. The obtained results show that the missile airframe is heavily underdamped. As well, the airframe bandwidth increases by increasing the missile speed. Doubling the missile speed fairly doubles the weathercock frequency at the expense of aerodynamic damping coefficient reduction.