In this paper, the problem of Guidance in the Mid-Course of a projectile’ s flight has been addressed to meet the requirements and limitations of the homing phase. These requirements and constraints are usually for low cost systems that are required to increase their range, including the angle of priority (angle of view) and the flight path angle (the angle of impact) of the projectile along with the amount of control effort obtained in the Mid-Course. Obviously, if these requirements are met, the success of the projectile mission will increase. Therefore, considering the importance of the problem and despite the various and scattered works that have been done so far, in this research, first the Mid-Course Guidance optimization problem of the projectile based on a performance function consisting of the control effort criterion, and also the final values of view angle and flight angle in the Mid-Course is defined as the penalty functions, respectively. Then, using the particle swarm optimization algorithm, optimal acceleration commands are generated. The simulation results compared with other methods show the proper functioning of the algorithm and its ability to solve more complex models.

In this paper, analytical solution of the generalized collision course (GGG) is presented considering approximate models for drag and thrust in the presence of gravity. The GGG is near optimal for a considerable flight time toward the end. Therefore, Guidance laws based on GGG only need modification for the first stage of flight. The proposed Guidance law is a modification of midcourse strategies based on the GGG and its implementation issues. A recursive relation for estimation of time-to-go for the GGG is presented in order to reduce the onboard computational burden. Two other recursive relations for time-to-go are obtained for optimal Guidance laws. The relations can be used for both midcourse and terminal applications. For an aerodynamically controlled interceptor, the Guidance law produces the commanded acceleration in the direction normal to its velocity vector or approximately normal to its body axis. Simulation results show that the presented strategy is superior to the midcourse Guidance laws based on GGG for a greater final velocity.

In [1], an optimal midcourse Guidance law for close distances, where the difference of gravity for interceptor and ballistic missile is negligible, was introduced. There, a closed form solution, based on an optimization problem, was found, with very good performance for close distances but degraded performance in real problems with unequal gravity for missile and interceptor. In this paper, the difference of gravity is taken into account by considering a spherical gravity model. A new equation to express the relative motion between missile and interceptor is used to derive a "Near Optimal Guidance Law". The results found using the new derivation are similar to those in [1] but it provides a Guidance law with matrix coefficients. Next, the performance of the Guidance law is improved using Kepler's algorithm. This modified approach results in an almost perfect intercept, even for large distances between missile and interceptor.

In this paper, generalized model predictive spread control methodis developed to consider the intermediate constraints on system states and system inputs. Because of using the orthogonal basis functions and thus reducing the computational burden, this new method which is named constrained generalized model predictive spread control can be used in online implementation of a finite-time constrained optimal control problem. For demonstrating the performance of the proposed technique, in this paper an interceptor midcourse Guidance problem is formulated to reach a desired point in space. Several constraints are considered such as: hard and soft intermediate and terminal constraints on system states and different constraints on input acceleration command to the interceptor in different time intervals of the trajectory. It is shown that in all the above situations, the proposed method could produce the optimal Guidance commands such that all the interceptor trajectory constraints are satisfied.

In this paper, a fuzzy logic Guidance algorithm is presented for the ascending phase of satellite launch vehicles in the presence of wind effects. In this algorithm, the midcourse constraints including maximum allowable angle of attack at the maximum dynamic pressure and the product of the dynamic pressure and angle of attack, as well as constraints on the final altitude and flight-path angle are considered. The algorithm uses a Mamdani- type fuzzy controller with centroid defuzzification. Maximizing and minimizing set methods to reduce wind effect, while satisfying the midcourse and final constraints. Simulation results show that the presented algorithm improves the performance of the satellite launch vehicle, satisfying the constraints within the maximum allowable estimation error on wind speed.

In this paper a new method for Guidance of solid propellant and aerodynamic control missile is introduced. In proposed method after achieving maximum acceleration، flight path angle of missile changes in order to impact to target. In this method flight information such as time history of magnitude of velocity and position vectors are saved in the table before firing. For calculating difference of nominal and disturbed parameters new concepts are used. Corrected values of nominal parameters are used in Guidance after firing to guaranty impact to target. The flight simulation of proposed method is compared with functional Guidance method and the effect of the proposed method is shown. Design verification is done in base of simulation in presence of flight disturbances.

An Integrated Fuzzy Guidance (IFG) law for a surface to air homing missile is introduced. The introduced approach is a modification of the well-known Proportional Navigation Guidance (PNG) law. The IFG law enables the missile to approach a high maneuvering target while trying to minimize control effort as well as miss distance in a two-stage flight. In the first stage, while the missile is far from the intended target, the IFG tends to have low sensitivity to the target maneuvering seeking to minimize the overall control effort. When the missile gets closer to the target, a second stage is started and IFG law changes tactic by increasing that sensitivity attempting to minimize the miss distance.A Fuzzy-Switching Point (FSP) controller manages the transition between the two stages.The FSP is optimized based on variety of scenarios; some of which are discussed in the paper. The introduced scheme depends on line-of-sight angle rate, closing velocity, and target-missile relative range. The performance of the new IFG law is compared with other Guidance laws. The results show a relative superiority in wide variety of flight conditions.

An optimal explicit Guidance law that maximizes terminal velocity is developed for the reentry of a vehicle to a fixed target. The equations of motion are reduced with differential flatness approach and acceleration commands are related to the parameters of trajectory. An optimal trajectory is determined by solving a real-coded genetic algorithm. For online trajectory generation, optimal trajectory is approximated. The approximated trajectory is compared with the pure proportional navigation and genetic algorithm solutions. The near optimal terminal velocity solution compares very well with these solutions. The approach robustness is examined by Monte Carlo simulation. Other advantages such as trajectory representation with minimum parameters, applicability to any reentry vehicle configuration and any control scheme, and Time-to-Go independency make this Guidance approach more favorable.