The paper presents the relations that form the basis for the calculus models needed to study rolling guided missiles. The equations of the general motion are written within the Resal trihedral in a way that allows, after linearization, one to draw up a structural scheme of the controlled object, a scheme that will be used in the guided missile system analysis. Opposite to the flying machines with stabilized roll where the analysis is done after decoupling the channels, for the missile with rotation, the analysis of the scheme is done simultaneously on both channels, as terms used are complex terms. Furthermore, there will be presented some characteristics of this type of missile by analyzing a model close to the one of an antichariot missile with gasodynamic control.The basic notations are according to standards presented by Ni~a et al.

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Abstract-The paper presents some aspects regarding the mathematical model and performance evaluation of a three stages microlauncher with a payload up to 50 kg. This work uses two separate models dedicated for each flight phase. For the ascending phase, we will use a three degrees of freedom model in quasi-velocity frame. For the orbital phase we will use a Kepler model, and for the orbital injection a Gauss perturbing model. The results analyzed will be in quasi-velocity frame but also some orbital parameters will be presented. Using these models, the microlauncher performances will be evaluated. The novelty of the paper consists in orbital injection approach, with optimal maneuvers description. IndexTerms-Mathematical model, microlauncher performances, orbital injection, optimal maneuvers.

The paper purpose is to present some aspects regarding the calculus model and technical solutions for small satellites attitude control. Mathematical model is put in nonlinear and linear form. The linear form is used for attitude control system synthesis. The attitude control system obtained is used in nonlinear form in order to maintain desired attitude. A few numerical simulations are made for standard input and the satellite behavior is obtained. The satellite model presented will be with six DOF and use Cartesian coordinates. At this item, as novelty of the work we will use the rotation angles to describe the kinematical equations. Also this paper proposes a Fourier linearising of Trigger Schmidt element used for applying the command moment. The results analyzed will be the rotation angles of the satellite as well the rotation velocity. The conclusions will focus the comparison between results obtained using optimal attitude control system, and the possibility to use such system for small satellite.