Hossam Hendy

The Rate-Gyro autopilot is acknowledged as a straightforward, highly dependable, and cost-efficient method for controlling tactical missiles. Its simplicity lies in its single-loop control topology, obviating the need for acceleration sensing. Notably, the airframes of most missiles exhibit pronounced oscillatory responses due to a low damping factor and relatively high natural frequency. Consequently, the flight control system serves the dual purpose of enhancing the missile's damping characteristics and aligning with guidance commands. Rapid autopilot responses are imperative for timely target interception, provided they do not compromise the overall robustness and stability of the flight system. Traditionally, the design of the Rate-Gyro autopilot hinges on two principal factors, namely the damping factor and open-loop crossover frequency. These factors are crucial for achieving resilient and stable behavior aligned with the desired guidance commands. During the autopilot gains design process, certain constraints may imperil the stability of the flight control system. Factors such as the interplay between open crossover frequency, phase margin, actuator dynamics, navigation sensor, and other delays in the control loop are pertinent considerations. Neglecting these aspects in the selection of the desired flight damping factor or autopilot gains may lead to instability in the flight system. This paper introduces an effective technique for the selection of gains and provides an explicit expression for autopilot gains design. The proposed technique incorporates considerations for the relationships among the aforementioned parameters.

In the current era, which is characterized by increasing demand for high-precision location and navigation capabilities, various industries, including those involved in intelligent vehicle systems, logistics, augmented reality, and more, heavily rely on accurate location information to optimize processes and deliver personalized experiences. In this context, the integration of Global Navigation Satellite System (GNSS) and inertial sensor technologies in smartphones has emerged as a critical solution to meet these demands. This research paper presents an algorithm that combines a GNSS with a modified downdate algorithm (MDDA) for satellite selection and integrates inertial navigation systems (INS) in both loosely and tightly coupled configurations. The primary objective was to harness the inherent strengths of these onboard sensors for navigation in challenging environments. These algorithms were meticulously designed to enhance performance and address the limitations encountered in harsh terrain. To evaluate the effectiveness of these proposed systems, vehicular experiments were conducted under diverse GNSS observation conditions. The experimental results clearly illustrate the considerable improvements achieved by the recommended tightly coupled (TC) algorithm when integrated with MDDA, in contrast to the loosely coupled (LC) algorithm. Specifically, the TC algorithm demonstrated a remarkable reduction of over 90% in 2D position root mean square error (RMSE) and a 75% reduction in 3D position RMSE when compared to solutions utilizing the weighting matrix provided by Google with all visible satellites. These findings underscore the substantial advancements in precision resulting from the integration of GNSS and INS technologies, thereby unlocking the full potential of transformative applications in the realm of intelligent vehicle navigation.

Conventionally, the Three-Loop Autopilot is dominated by three main parameters namely the damping factor, time constant and the open-loop crossover frequency, to achieve the desired performance. In developing the autopilot gains, the same gains as the conventional approach can also be obtained using the general optimality theory. The basic idea of setting the design parameters and the feedback gains using methods of optimal control revolves around the proper selection of the performance index for which the control is optimized. In this paper, an explicit formula for autopilot gains is driven by taking into account the relations between the desired performance and the open-loop frequency, phase margin and time constant without the need to adjust the LQR weights. Since each set of design parameters selection doesn’t guarantee the existence of the optimal control law, the optimization criteria for the three-loop autopilot are derived to know the set of design parameters for which the control is optimized. Finally, an optimal criterion to select the design parameters and therefore the autopilot gains is developed, where the time constant is set to the designer objective while the open-loop crossover frequency and the phase margin as design constraints.

Recently, there has been a renewed interest in the development of battle tanks armoring, which made old generations of anti-tank missiles are no longer able to counteract modern tanks by hitting armor in the front or the side sections of the tank. Recent developments directed toward designing the antitank missiles to hit the tank from the upper ceiling which is considered one of the weakest points in the tank armor. The new generations of anti-tank missiles are developed with the ability to hit tanks from its upper section (top-attack). So, the work in this paper, focuses on developing the antitank missile guidance law to generate a top-attack trajectory considering (impact angle, missile maneuverability, miss distance) constraints, the accuracy of the generated trajectory is evaluated through two main parameters impact angle and miss distance, a polynomial function based trajectory is proposed. Further extended analysis is carried out for evaluating missile motion parameters via a 6-DOF model. The proposed algorithm is simulated utilizing Matlab/Simulink with different scenarios.

The initial alignment of the Inertial Navigation Systems (INS) has significant importance in recent years. It is needed to ensure the accuracy and reliability of the system. For the initial alignment of Strap-down INS (SINS), Navigation Mathematical Software (NMSW) implements an efficient technique. These techniques provide a solution for navigation problems such that the inertial velocity, and coordinate estimation in addition to the inertial sensor's calibration. In this paper, the navigation algorithm for the closed-loop system is modified to feedback continuously in alignment. The NMSW is testing, and a typical trajectory is calculated based on the block diagram of the hardware-in-the-loop (HIL) flight simulation. Results show that the initial calculated variance setting of Azimuth angle error will significantly affect the speed of the initial alignment. The process of simulation and computation is highlighted related to the trajectory simulator data from the equation of motions of the high-speed flying vehicle.

Low Earth Orbit satellite(LEO) missions have potentially attractive features such as low launch cost, and wide remote sensing applications. There are three vital missions during the satellites life time including orbit transfer from the parking to the operating orbit, maintaining the operating orbit and de-orbiting mission to a disposal orbit at the end of the satellite lifetime according to the standard disposal orbit law. The mathematical model has been implemented to study various LEO mission orbits including the three vital missions. The mathematical model evaluates the required performance parameters, velocity budget, number of manoeuvers, duration and consumed propellant mass for each mission. The mathematical algorithm is carried out using Matlab/Simulink, the complete mathematical flight dynamic model has been verified through comparisons with the generated mission scenarios from Satellite Tool kit (STK), the graphical user interface is designed to display the model. The carried out algorithm can be applied in satellite mission design to help predicting the parameters for each mission to minimize the risks and intensive tests. it can be considered as powerful tool to design and analysis LEO satellite missions.

Vibration effects for space crafts has a great influence during the launching phase and also after launching. Active vibration isolation techniques have good results to protect such systems from vibration problems in wide frequency bands. But, these active systems increase the total weight of space craft. That will increase the need of power supplies, which may affect the mission life time due to batteries and solar cell charging consuming. Passive control for such problem is designed using Transfer Matrix Method for Multibody Systems (MSTMM) that can provide an easy method to analyze a passive vibration isolation technique with simple design. The main objective of this paper is to study the different models effect in isolation design using MSTMM and passive control techniques to conserve the power consumption for enlarging the power batteries life. That means longer mission intervals for such systems. Simulation results are given to demonstrate the effectiveness of the offered technique.