فهرست مطالب seyed hamid jalali naini

This paper presents a modification to a type of Pulse-width Pulse-Frequency (PWPF) Modulator utilized an integrator block. In this modulator that called here as "Integral Pulse-Width Pulse-Frequency (IPWPF)," an integrator is used instead of the first-order low-pass filter. To improve the performance of the control system, the modulator is modified by using a logical circuit in order to reset the output of the integrator. In this logical circuit, if the error signal becomes less than a specified small value, the integrator will be reset, that is, "Small Error-Reset Integrator (SE-RI)." The modification is applied to the stabilization and pointing modes. In stabilization mode, the control gain is obtained analytically such that the angular rate of the satellite becomes zero or less than a specific percentage of its initial value by a single pulse. Simulation results show that the performance of the modified IPWPF is comparable with that of PWPF in pointing mode.

In this paper, the effect of system time delay on the miss distance of proportional navigation guidance law is studied for high-order binomial guidance and control systems. For this purpose, a linearized model is utilized in one dimension. The normalized miss distance due to heading error (HE), step, ramp, and parabolic target maneuvers, and seeker noise is computed using normalized adjoint equations. The glint, range-independent, and (semi-) active range-dependent noises are considered for the seeker, and the results are verified by straight-forward method. Moreover, the effects of system time lag, system time constant, effective navigation ratio, system order up to 30 in proportional navigation miss distance are also investigated. An extremum effective navigation ratio is also obtained to minimize the worst case miss distance with respect to the final time. Furthermore, approximate miss distance formulas are presented based on worst case analysis on final time using curve fitting for mentioned miss distance sources. Finally, the approximate formulas of steady state nondimensional coefficients due to seeker noise in terms of effective navigation coefficient, for example for system orders of 5 and 10, are obtained.

his paper suggests arobust optimization algorithm for the design of the satellite attitude control system in order to increase the robustness of the performance under uncertainties. A single-axis on-off attitude control with rigid dynamics is considered using Schmitt-Trigger and PID controller. The model uncertainties include the moment of inertia, thrust level, thruster delay and theexternal disturbance amplitude.A weighted combination of expected value and standard deviation of pointing error is considered as an objective function for the robust optimization. The numerical solutions show that the robust optimization reduces the variations of the objective function, i.e. it increases the robustness of the system performance compared to the deterministic optimization.

In this paper, the performance of a single-axis attitude control with pulse-width pulse-frequency (PWPF) modulation is enhanced using a modified proportional-integral-derivative (PID) controller for a rigid satellite with on-off thruster actuators. For this purpose, the well-known observer-based PID approach is utilized. The on-off thruster actuator is modeled with a constant delay followed by a second-order binomial transfer function. The modulator update frequency is limited to 40 Hz as an input to the on-off thruster actuators. In this study, the design criteria of pointing accuracy, overshoot of the attitude response, fuel consumption, and the number of thruster firings are considered for a step external disturbance (with different values). The parameters of the observer-based PID controller are tuned using parametric search method. Simulation results show that the fuel consumption and settling time of the observer-based approach are considerably decreased with respect to those of PID controller with PWPF modulator. Moreover, the overshoot of the observer-based approach is omitted. Finally, the robustness of the observer-based modified PID controller is investigated in presence of uncertainties in satellite moment of inertia and thrust level of on-off actuators.

In this paper, an explicit formulation of optimal line-of-sight strategy is derived in closed-loop for integrated guidance and control (IGC) system without consideration of fin deflection limit. The airframe dynamics is modeled by a second-order nonminimum phase transfer function, describing short period approximation. In the derivation of our optimal control problem, the actuator is assumed to be perfect and without limitation on fin deflection, whereas fin deflection limit is applied for the performance analysis of the presented optimal IGC solution. The problem geometry is assumed in one dimension and the final position and final time are fixed. The formulation is obtained in four different normalized forms to give more insight into the design and performance analysis of the optimal IGC strategy. In addition, guidance gains are obtained analytically in explicit form for steady-state solution. In overall, the performance of IGC is better than that of IGC with steady-state gains, but have more computational burden; however, it is reasonable for now-a-day microprocessors. Curve fitting or look-up table may be used instead for an implementation of optimal IGC strategy. Moreover, parametric study of nondimensional IGC parameters is carried out, such as weighing factor, dc gain, and short period frequency. Finally, the performance of the both IGC strategies is evaluated with airframe model uncertainties.

In this paper, an explicit optimal line-of-sight guidance law for second-order binomial control systems is derived in closed-loop without acceleration limit. The problem geometry is assumed in one dimension and the final time and final position are fixed. The formulation is normalized in three forms to give more insight into the design and performance analysis of the guidance law. The computational burdun of the guidance law is reasonable for now-a-day microprocessors; however curve fitting or look-up table may be used for the implementation of the second-order optimal guidance law. The performance of the second-order optimal guidance law is compared in normalized forms with zero-lag and first-order optimal guidance laws using third-, fourth-, and sixth-order binomial control systems with/without acceleration limit. Moreover, the effect of the final time, the equivalent time constant of the vehicle control system, the vehicle-to-target line-of-sight weighting factor in cost function, and acceleration limit are investigated. Normalized miss distance analysis shows that the miss distance of the second-order guidance law is smaller than the two mentioned schemes for small total flight times, especially with large maneuvering capability.

In this paper, a closed
loop strategy in the vertical plane is derived in order to determine the thrust direction of a launch vehicle in terms of velocities
to
be
gained The two velocities-to-be-gained are utilized, here, for a given altitude and zero vertical speed in a specied nal time The formulation is obtained for constant gravity assumption, but it works when the velocities
to-be-gained are obtained for a spherical-Earth model via explicit or implicit relations