مهندسی هوا فضا - دینامیک پرواز و کنترل
دانشگاه صنعتی شریف، تهران، ایران
مهندسی مکانیک - تبدیل انرژی
دانشگاه تهران، تهران، ایران
مهندسی مکانیک - طراحی جامدات
دانشگاه صنعتی امیرکبیر، تهران، ایران
سید حمید جلالی نائینی متولد سال 1349 در تهران تحصیلات متوسطه خود را در دبیرستان دکتر هشترودی گذرانده و در سال 1368 با کسب رتبه 520 در کنکور سراسری در رشته مهندسی مکانیک (گرایش طراحی جامدات) در دانشگاه صنعتی امیرکبیر (پلی تکنیک تهران) پذیرفته شد. وی مقطع کارشناسی ارشد را در مهندسی مکانیک (تبدیل انرژی) در سه ترم در سال 1375 اخذ نمود. پس از آن دوران خدمت وظیفه را در دانشگاه هوایی شهید ستاری با درجه ستوان یکمی انجام داد. از سال 1377 تا 1383 در سازمانها و پژوهشکده های متعددی از جمله پژوهشکده هوافضا (پژوهشگاه هوافضای فعلی) وابسته به وزارت علوم، تحقیقات و فناوری و سازمان فضایی ایران وابسته به وزارت ارتباطات و فناوری اطلاعات خدمت نمود. در سال 1383 در مقطع دکترای دانشکده هوافضای دانشگاه صنعتی شریف پذیرفته شده و پس از 7 ترم در سال 1386 از رساله دکترای خود دفاع نمود. پس از آن، کماکان در پژوهشکده ماهواره وابسته به سازمان فضایی ایران خدمت کرده تا در سال 1388 به عنوان عضو هیأت علمی دانشکده مهندسی مکانیک دانشگاه تربیت مدرس پذیرفته شد. حوزه تحقیقاتی ایشان، هدایت وسایل پروازی و کنترل وضعیت ماهواره بویژه با عمگرهای دو وضعیتی (روشن-خاموش) است. تاریخ علم و فناوری و سیاستگذاری علم و فناوری نیز جزو علاقمندیهای حوزه پژوهش ایشان قرار دارد. وی دو دوره عضو اصلی هیأت مدیره انجمن هوافضای ایران بوده است.
In this paper, the performance of the spacecraft attitude control system is enhanced using model-based disturbance feedback control (DFC) strategy in the presence of disturbance. This control strategy is applied to a single-axis spacecraft attitude control with thruster, reaction wheel, and magnetic torqrod actuators, separately. An anti-windup observer-based modified PI-D is utilized for each actuation system as a main controller. The controller gains are tuned using genetic algorithm when the time average of absolute value of pointing error is chosen as an objective function. The performance of DFC with the modified PI-D controller is investigated under disturbance and model uncertainties. The numerical simulation shows tha
An approximate solution of two-body problem is proposed using weighted combination of linear and inversely cubic gravity models. First, two approximations are introduced for the weight coefficients. To enhance the accuracy of the solution, a modified weight coefficient is suggested by a linear combination of the two primary weight coefficients. The accuracy and computational burden of the combined method are compared to those of power series solution, including the Lagrangian coefficients and the recursive formulation introduced by Turner et al. The comparison is made with the same computational burden by setting an appropriate number of terms of power series solution. The advantage among the methods depends on the computatio
In this paper, a modified PI-D controller with pulse width pulse frequency (PWPF) modulator with on-off thruster for a rigid satellite is tuned under uncertainties in a quasi- normalized form. A modified proportional-integral-derivative (PID) based on observer method is used as a controller. Uncertainty is considered on thruster model, thrust level, and external disturbance parameters. Thruster is modeled with a constant delay followed by a second-order binomial transfer function. The modulator update frequency, the input to the on- off thruster, is limited to 50 Hz. Controller gains are tuned and optimized using a multi-objective genetic algorithm. The simulation results of the deterministic optimization show that the rise time is reduced
A closed-loop optimal line-of-sight (LOS) guidance law is derived for second-order minimum and nonminimum phase control systems with two zeros. To reduce the effect of wind, the square of the velocity component normal to the LOS is added to performance index. The solution is normalized and obtained in one-dimension. The performance analysis is carried out using one-dimensional simulation and 3DOF flight simulation in the vertical plane in the presence of wind, and compared to zero-lag and first-order optimal guidance schemes. Moreover, the effects of weighting coefficient, zero location, and acceleration limit are investigated.
This paper presents a comprehensive study on the performance analysis of 8 conceptual guidance laws for exoatmospheric interception of ballistic missiles. The problem is to find the effective thrust direction of interceptor for interception of short-to-super range ballistic missiles. The zero-effort miss and the generalized required velocity concept are utilized for interception of moving targets. By comparison of the 8 conceptual guidance laws, the thrust direction is suggested to be in the direction of generalized velocity-to-begained, or constant velocity-to-be-gained direction, rather than to be in the direction along zero-effort miss, or that of linear optimal solution for long-to-super range interception. Even for short coasting range
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-, fourt
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
In this paper, a closed-loop optimal line-of-sight guidance law for rst-order control systems is derived for stationary targets. The problem is solved for the onedimensional case using normalized equations to obtain normalized guidance gains and performance curves. Three sets of normalized equations are introduced and discussed using di erent normalizing factors. The performances of the guidance laws are compared in normalized forms with zero-lag optimal guidance and a rst-order optimal scheme with steady-state gains using a second-order control system. Normalized miss distance analysis shows that the miss distance of the rst-order guidance law is smaller than the two mentioned schemes for small total ight times.? 2016 Sharif University of
This paper deals with miss distance analysis of single-lag Optimal Guidance Law (OGL) using linearized equations of motion in normalized form. The radome refraction error, acceleration limit, constant target acceleration, and arbitrary-order binomial guidance and control system are considered in the formulation. In addition, body rate feedback is utilized in the OGL formulation as a well-known classical compensation method of radome e ect for proportional navigation guidance. The numerical solution of normalized equations produces normalized miss distance curves, which are useful tools for guidance designer for analysis and design of guidance parameters for an allowable miss distance and acceleration limit. Moreover, a modi ed rst-order gui
An optimal fuel guidance strategy with acceleration limit is derived for interception of maneuvering targets. The guidance/control system is assumed as a linear time-varying arbitraryorder and is identical in each channel. An approximate model for drag acceleration is also considered in the solution. The optimal fuel strategy has some detrimental effects in most practical situations, if it is used entirely to the end. This strategy can be used at the beginning of the guidance phase for large initial heading errors in order to quickly reduce it with minimum effort. An explicit guidance law is then developed for minimum and nonminimum phase autopilots in order to modify the undesirable effects of the optimal fuel guidance law.
Purpose – The purpose of this paper is to develop a novel solution for the predicted error and introduces a systematic method to develop optimal and explicit guidance strategies for different missions.Design/methodology/approach – The predicted error is derived from its basic definition through analytical dynamics. The relations are developed for two classes of systems. First, for systems in which the acceleration commands are truncated at a specified time. Second, for systems in which the corrective maneuvers are cut off at a specified time. The predicted error differential equation is obtained in a way that allows for derivation of several optimal and explicit guidance schemes.Findings – The effect of tangential acceleration in conj
In this paper, midcourse guidance strategies are derived considering approximate models for drag and thrust in the preseI1 (: e of gravity and autopilot dynalnics. The guidance/(tontrol sys-tein is assunied as a linear tiine-varying a11d of arbitrary-order. The zero-effort 111iss differential equation is obtained in a Way that three classes of guidance laws, 11a111ely optiinal fuel, optimal energy, and closed-loop explicit strategy, can be derived by available 111etl1ods for 1ni11i111u1n a11d nonnlinimum phase autopilots. The resulting guidance laws have the structure of a guidance gain multiplied by the zero-effort miss. Several thrust/drag models are also considered and discussed to obtain the effective navigation ratio and zero-effort mi
OPTIMAL control theory has been utilized to derive modern guidance laws with improved performance. This is an attempt to replace the well-known proportional navigation (PN). The requirement for better performance has led to the development of optimal guidance laws (OGLs) by the consideration of the target future maneuver and the dynamics of an interceptor and its target from perfect to time-invariant high-order autopilots. 1− 4 The desired performance index in the exoatmosphere is usually the amount of fuel required for corrective maneuvers. If the fuel consumption were minimized, the solution would be mathematically intractable, 1 especially with trajectory constraints. Explicit guidance laws (EGLs) may be developed to deal with this obj
A generalized three-dimensional line-of-sight guidance with lead angle (GLOS) is presented for maneuvering targets. In the proposed guidance law, we define a desired line-of-sight (LOS) from the tracker by taking the target future maneuvers into account. Because of the limitation of the tracker field-of-view (FOV), the lead angles in the azimuth and elevation channels must be fractions of the azimuth and elevation angles for straight collision trajectory. Simulation results show that GLOS has better performance, ie, lower miss distance and total control effort, than the conventional LOS guidance with lead angle.
A generalized method to derive two classes of explicit midcourse guidance laws followed by a coasting phase is presented for a given interceptor maneuvering proﬁle and against maneuvering targets. This method is developed for a linear high-order time-variant autopilot Without acceleration limit. In the ﬁrst class, the acceleration command is truncated at a time, but the achieved commanded acceleration continues depending on systems dynamic. In the second class, the command and its achieved acceleration are cut off at the same time. Moreover, the optimal strategies with quadratic performance index are derived by using the Schwartz inequality and compared with corresponding proposed guidance laws. Simulation results show that the proposed
The optimal exoatmospheric strategy for N-ﬁXedinterval steering periods is developed for intercep-tion of maneuvering targets. In addition, analytical solution of a class of optimal guidance laWs is obtained With/Without acceleration limit. Further-more, a generalived nletllod to derive closed-loop steering connnand is presented for a given intercep-tor maneuvering proﬁle. Finally, a modiﬁed guidance is proposed to l1ave a better acceleration proﬁle and less fuel co11sun1ptio11 than that of the optimal guidance With the least integral of the control effort squared.
The three-dimensional equations of a modified line-of-sight guidance are derived for maneuvering targets. In this modified guidance, the pursuer is fired with a lead angle at the target, which decreases with time to reduce the lateral acceleration requirements. We assume an ideal case that the pursuer is always on the desired direction from the tracker without any error. In addition, a closedform solution is derived for a nonmaneuvering target by assuming the total pursuer acceleration to be equal to the required acceleration. Furthermore, an approximate solution for a pursuer with an arbitrarily time-varying velocity is obtained for small pursuer velocity-to-beam angle. Numerical solution indicates that the approximate solution is appropri
IN line-of-sight (LOS) trajectory, the pursuer always lies on the line between the target tracker and the target. If the pursuer is always on the tracker-target LOS, then the pursuer will surely hit the target. In other words, in LOS guidance, the pursuer maneuvers so as to be on the LOS between the target tracker and the target. 1 6 A closed-form solution of a LOS trajectory for nonmaneuvering targets and a pursuer with constant velocity is not available. In 1955, Locke presented a 10-term series solution of this, for pursuer rangevspursuerangularposition. 1 In addition, Macfadzeanclaimed in Ref. 3 that there is no closed-form solution for LOS trajectory, even with simplifying assumptions such as those made by Locke. In this Note, the clos
The closed-form solution of the three-dimensional line-of-sight (LOS) guidance is derived for the ideal case which a pursuer is always on the instantaneous line between the target tracker and the target. Then, some significant characteristics such as intercept time, cumulative velocity increment, initial condition for interception and the effect of acceleration limit are investigated and discussed. In addition, the equivalent effective navigation ratio for the LOS guidance is derived. Finally, the solutions for a special case of maneuvering target for two cases are presented.
In this paper, a trajectory shaping guidance scheme is proposed using optimal control theory based on timevarying weighting function in performance index. This guidance law satisfies the final position and dive angle in the presence of aerodynamical force and gravity. The time-varying weighting function is chosen to shape the commanded acceleration for trajectory constraints such as a zero acceleration at handover and midway trajectory constraints as well.
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