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In this paper, the possible impact effects of orbital bombardment systems and their feasibility are studied. These effects are the projectile penetration into concrete and steel targets and seismic effects. The equations of motion for the re-entry of a projectile and the penetration were solved numerically. The projectile penetration is modelled using the Alekseevskii–Tate model. By varying the altitude (h), projectile length (L), manoeuvre velocity (ΔV) and the target properties, the flight time (t), earthquake magnitude (M) and penetration depth (P) are calculated. The calculations show that the impact of a tungsten alloy rod with a length of 8 m and a 0.4 m diameter results in an earthquake with a seismic magnitude of only 2.5 on the Richter scale. For concrete, the optimal result is obtained for a projectile with a length of 0.56 m. It penetrates 1.79 m with a minimal ΔV trajectory. These results show that a kinetic orbital bombardment system is not feasible without major technological developments, the impact angle being a bottleneck of the concept. Moreover, one has to accept very high costs. Without any means to change the attitude of the projectile, using ICBMs or bombers shows a better penetration performance than re-entry. Weapons in orbit may provide a strategic advantage. However, they are restricted by international space laws. Impact angle of the projectile is a bottleneck for kinetic orbital bombardment. Larger impact angles can be achieved, but at the expense of a larger mass-to-orbit. A hypersonic drag device may be used to optimise the impact angle and thus improve the system. Alternative projectile delivery methods (Bomber, Intercontinental Ballistic Missile (ICBM)) show better performance for both steel and concrete targets. Essentially, only penetration phenomena matter because the seismic effects are not significant. Therefore, orbital bombardment systems don't even resemble weapons of mass destruction (WMD). Given their limited effect, destroying a particular target requires a guidance and flight control system, which, given the high velocities, may not be feasible. [ABSTRACT FROM AUTHOR]

This paper investigates a problem of autonomous aerial refueling point parallel rendezvous guidance and flight control subject to exogenous wind field disturbances. The tanker and uncrewed receiver aircraft approach the common refueling rendezvous point according to their respective flight plans, ultimately maintaining a fixed altitude and the same speed. According to the relative position of refueling target point and uncrewed receiver, a new continuous terminal sliding-mode guidance law combined with the prescribed performance control theory is designed under the end angle and speed constraints. Meanwhile, a finite-time observer is proposed to estimate the parameter uncertainties and wake vortex disturbances during refueling rendezvous. Based on the estimated information, a backstepping double-integral sliding-mode controller is developed for the uncrewed receiver flight control system. In the process of the guidance and flight control system, the integral variable is introduced to improve the steady-state performance. The designed guidance and control system can enable the uncrewed receiver to achieve the refueling rendezvous task within finite time and improve the efficiency of uncrewed aerial refueling. Comparative numerical simulations are provided to demonstrate the effectiveness of the proposed guidance and control scheme. [ABSTRACT FROM AUTHOR]

The use of non-piloted eVTOL aircraft in non-segregated airspace requires reliable and deterministic automatic flight guidance systems for the aircraft to remain predictable to all the users of the airspace and maintain a high level of safety. In this paper we present a 3D trajectory generation module based on clothoid transition segments in the horizontal plane and high order polynomial transition segments in the vertical plane. The expressions of the coefficients of the polynomial are derived offline are used to generate the trajectory online, making the system capable of running in real time without requiring enormous computational power. For the horizontal plane, we focus on the flyby transition, and therefore present a thorough analysis of the flyby geometry and the limitations linked to this geometry and the construct of three-segment trajectory generation around a fixed turn rate. We present feasible solutions for these limitations, and show simulation results for the combined horizontal and vertical plane concepts, allowing the system to generate complex 3D trajectories. [ABSTRACT FROM AUTHOR]

This paper presents a mission data-handling algorithm that provides a data link interface between a flight control computer (FCC) and a ground control station (GCS) using the protocol standard STANAG 4586, AEP-84 Volume II. A subset of defined messages is used for the transmission of mission and area definition data. Functions for decoding uplink messages, processing received data, and encoding downlink messages are developed. Due to a modular software architecture, these functions can be used both for the onboard FCC, as well as for the GCS using additional logic, to invert the transmission direction. Furthermore, high level consistency and validation checks are performed on the transmitted data before forwarding the mission to the flight guidance functions. Successful transmission of mission and area data between the FCC and the GCS is performed using a high-fidelity model-in-the-loop simulation of a demo mission for a medical evacuation eVTOL aircraft. [ABSTRACT FROM AUTHOR]

Automatic flight control system （AFCS） is one of the most important safety‑critical systems in modern aircraft, and flight guidance control system （FGCS） is an important part of it. There are dozens of flight modes in FGCS, and its mode conversion logic is very complex, which is prone to pattern confusion and other problems in the conversion of various modes, making it difficult to verify their safety and correctness. However, formal modeling and verification of safety‑critical systems can improve the correctness and safety of the system by using formal methods in computer science. This paper takes the automatic flight mode conversion logic of typical FGCS as the research object, uses the software tool ART （Avionics requirement tool） independently developed by the author team to carry out formal modeling and verification, and compares the verification ability and efficiency with the Design Verifier tool in Matlab/Simulink. The case study results show that it is feasible to model and verify the automatic flight mode conversion logic of FGCS using formal methods. Meanwhile, our software platform has more complete verification capability and better verification efficiency. [ABSTRACT FROM AUTHOR]

Copyright of Electronics & Control Systems is the property of National Aviation University and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Copyright of Przegląd Elektrotechniczny is the property of Przeglad Elektrotechniczny and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

This paper develops a mapping approach to explore the relationship between different layers of a multilayer air transport network composed of airway, route, and flight network. A two-step methodology is adopted to investigate the hierarchical structure and mapping relationship of the integrated network. First, the relationship between airway and route network is characterized by a multisource multisink shortest path method based on a generalized incidence matrix. Second, the relationship between route and flight network is formulated by a two-dimension array. A case study of an en route airspace in Lanzhou air traffic control area in China verifies the feasibility of the proposed two-step methodology. [ABSTRACT FROM AUTHOR]

The article describes the condition of the missile program in the U.S., Soviet Union, and other countries. According to Eileen Galloway of the U.S. Committee on Armed Services, the development of guided missiles in the country may be categorized by mission, systems of propulsion, and its system of guidance. Meanwhile, the Soviets have tested five of the six basic types which includes its intercontinental ballistic missile. Canada is also training its armed forces to use missiles and conducting some independent development of guided missiles.

The article provides information on the development of the different kinds of missiles in the U.S. Armed Forces. The missiles have three principal components which include propulsion, guidance, and payload system. The relative efficiency of propellants is measured by the specific impulse which is defined as the thrust produced by consuming the propellants at the rate of one pound per second. Missiles are also classified depending on the way they interact. Some of the types include the cruise missiles, aerodynamically controlled missiles, and ballistic missiles.

Abstract The paper proposes a method to minimize stage-by-stage initial unbalance in the aero engine assembly of multistage rotors based on the connective assembly model. The process of the mass eccentric deviation propagation in the assembly is analyzed. The initial unbalance of the final assembly is improved by properly selecting the assembly orientations of multistage rotors. A constrained nonlinear programming model is extracted from the optimal assembly strategy by choosing the initial unbalance as the objective function, and choosing the assembly orientations as the nonlinear constraint. The globally optimized solution of the constrained nonlinear programming model is solved using a genetic algorithm. The proposed method is used on an experimental set-up with the horizontal balancing machine and the effectiveness of the proposed method is verified by the assembly of the multistage rotors using the optimal assembly strategy. Compared to the direct assembly strategy, which the assembly orientations without consideration, the initial unbalance of final assembly using the optimal assembly strategy are reduced by 45.8%, 63.1% and 72.7% for two, three and four rotors assembly, respectively. The proposed method can improve the assembly quality of multistage rotors in the aero engine assembly and be used for assembly guidance, tolerance allocation, and so on, especially for the assembly with a large number of rotors. [ABSTRACT FROM AUTHOR]

Abstract A novel three dimensional aiming point guidance law is presented in this paper, which eliminates the need for gravity compensation, resulting in engagements occurring with zero acceleration commands, in turn leading to smaller overall control effort and an excellent zero effort miss behavior. The conditions necessary for collision in the presence of gravity are derived using model-based prediction and computationally efficient shooting method. This is followed by employing differential-geometric guidance philosophy for achieving collision with the target. The efficacy of the proposed guidance law is then verified using simulation studies for terminal phase ballistic missile interception, and is compared with proportional navigation, aiming point guidance and compensated-weave guidance. Results indicate that the proposed guidance law ensured collision with desired level of accuracy and outperformed the aforementioned guidance laws used for comparison, in terms of miss distance, zero effort miss behavior and control effort. [ABSTRACT FROM AUTHOR]

This paper presents an impact time control guidance law that does not perform explicit time-to-go estimation. In order to satisfy the interception and the desired impact time constraint simultaneously, a sliding surface variable formulated as a sum of the relative range and the desired time-to-go is defined, weighted by two nonzero weighting functions. Then, the achievement of the sliding mode satisfies the following equivalence: The relative range is zero if and only if the elapsed time equals to the desired impact time. It means that both the interception and impact time control can be satisfied at the same time in the sliding mode. The impact time control guidance law is derived to enforce the defined surface variable to the sliding mode. Because the law is designed based on the capture condition without separate time-to-go estimation process, the achievement of the sliding mode always guarantees the interception of the target at the desired impact time. In addition, the proposed law can be applied to an engagement considering nonstationary targets in a straightforward manner because the corresponding time-to-go estimation is not needed. Simulation results demonstrate that the proposed guidance law has better performance in comparison with the existing guidance law, in respect of satisfying the desired impact time constraint against a nonstationary target. [ABSTRACT FROM AUTHOR]

This paper proposes a guidance law that achieves the desired terminal impact time without violating a seeker's field-of-view (FOV) limits. In order to derive the guidance law, kinematic conditions for impact time control are defined, and the backstepping control technique is applied for the satisfaction of the conditions. As a virtual control input for the backstepping structure, the missile lead angle, which represents the angle between the line-of-sight (LOS) and missile velocity vector, is used and its magnitude is limited by a prescribed limit. Then, the seeker's look angle can also be confined within a specific range because the seeker look angle is mainly determined by the difference between the LOS and the velocity vector. This capability to confine the seeker's look angle with achieving the desired impact time is the main contribution of the paper. To evaluate the performance of the proposed law, numerical simulation is conducted. The result demonstrates that the proposed guidance law enables the missile to simultaneously achieve the interception and desired impact time without violating the prespecified FOV limits. [ABSTRACT FROM AUTHOR]

This paper proposes a guidance strategy, which caters to both impact angle and impact time terminal constraints. This guidance scheme is based on switching between impact time and impact angle guidance laws. Unlike the existing impact time guidance laws, the proposed guidance strategy takes into account the curvature of the trajectory due to requirement of impact angle. The guidance law is derived using sliding mode control theory with the switching surface based on impact time error. The interceptor first corrects its course to nullify the impact time error and then aims to achieve interception with desired impact angle. In order to reduce transitions between the two guidance laws, a novel hysteresis loop is introduced in the switching conditions. The guidance law is initially designed for stationary targets, and later it is extended to constant velocity targets using the notion of predicted interception point. In order to validate the efficacy of the proposed guidance strategy, simulation results are presented with constant as well as realistic time-varying speed interceptor models for different engagement scenarios against stationary and constant velocity targets. The performance of the guidance law is evaluated under noisy measurements and the presence of system lag and its performance is compared with other existing guidance laws. [ABSTRACT FROM AUTHOR]

Existing literature has shown that the control force at the nose could cause dynamic instability for controlled projectiles. To lower the adverse impact on the dual-spin projectile with fixed canards under the premise of meeting guidance system requirements, the influence of control moment provided by a motor on the flight stability is analyzed in this paper. Firstly, the effect of the rolling movement on stability is analyzed based on the stability criterion derived using the Hurwitz stability theory. Secondly, the evaluation parameters combining the features of different control periods that could assess the variation of stability features after the motor torque are obtained. These effective formulas are used to indicate that, to reduce the flight instability risks, the stabilized rolling speed of roll speed keeping period should be as small as possible; the variation trend of motor torque during the rolling speed controlling period and the roll angle of the forward body during roll angle switching period are recommended corresponding to the projectile and trajectory characteristics. Moreover, detailed numerical simulations of 155mm dual-spin projectile are satisfactory agreement with the theoretical results. [ABSTRACT FROM AUTHOR]

Abstract This paper proposes a novel cooperative guidance strategy, which aims to intercept one highly maneuvering target with multiple inferior missiles. In the scenario of interest, both the missiles and target are assumed to have bounded maneuverability, and the guidance goal is to make the joint reachable set of interceptors cooperatively cover the target maneuvering range. Under this guidance scheme, a preprogrammed covering strategy and an adaptive covering strategy are designed for the missile teams without and with communication capability respectively. The former attempts to specify different subsets of the target maneuvering range to different missiles as the desirable reachable sets, while the latter aims to coordinate the desirable reachable sets of different missiles dynamically according to the changing engagement situation. Considering the disadvantage of inferior maneuverability, the inherent limitations of the proposed guidance scheme are discussed. Numerical simulations with different target maneuvering modes demonstrate the prominent performance improvements of cooperative strategy over the traditional guidance laws. [ABSTRACT FROM AUTHOR]

In this paper, a cubic spline guidance law is proposed for intercepting a stationary target at a desired impact angle in surface-to-surface engagement scenarios. An inverse method is used, which represents the altitude as a cubic polynomial of the downrange. The paper addresses many shortcomings of earlier work in this area. In particular, an explicit guidance expression is derived, which makes the guidance law effective and accurate in the presence of disturbances and uncertainties. It is also shown that the guidance command can be obtained using a single cubic spline polynomial even for impact angles greater than and equal to $\pi / 2$ , while resulting in substantial improvement in terms of lateral acceleration and length of the trajectory. The paper also obtains an analytically derived capture region. The proposed guidance law is conceptually simple, independent of time-to-go, computationally inexpensive, and does not need linearization. [ABSTRACT FROM AUTHOR]

A novel cooperative defensive guidance law is presented for a two-on-two engagement. Instead of classical strategies in which evasive maneuvers are performed or additional agents (e.g., defender missiles) are deployed, the target pair lures the pursuing missiles into collision. The optimal cooperative strategy is solved using state-dependent Riccati equation method. Linearized kinematics and arbitrary-order linear adversaries' dynamics are assumed in the guidance law derivation. Imperfect information is assumed on the relative states and the guidance laws employed by the missile pair. Guidance strategies that the pursuing missiles may employ are assumed to belong to a finite set of linear guidance laws. In addition to the proposed cooperative defensive strategy, a decentralized estimation scheme based on the multiple-model adaptive estimation approach is also presented. Guidance law and estimation performance are demonstrated using nonlinear simulations. Simulation results show the viability of the proposed guidance law and highlight the sensitivity of the guidance law to range measurement accuracy. [ABSTRACT FROM AUTHOR]

In this paper, a new guidance strategy is proposed for intercepting stationary and constant-velocity moving targets with desired impact time and angle via a virtual target approach, subject to nonlinear engagement kinematics and lateral-acceleration limits. The guidance procedure is divided into two stages by introducing a virtual target. Specifically, for the first stage, the nonsingular terminal sliding-mode guidance law is employed to intercept the virtual target with a specified impact angle in finite time. Furthermore, to achieve the desired impact angle on the real target, the proportional navigation guidance law is used in the second stage to keep the missile traveling with invariant flight-path angle. The impact time and angle on the real target are simultaneously controlled by selecting a proper guidance parameter whose optimal value is obtained via an optimization routine with a small number of iterations. This technique is user/designer-friendly in that it does not involve the time-to-go estimation procedure. In addition, less control energy is required in the overall guidance process. Numerical simulations with comparisons are conducted to validate the effectiveness and feasibility of the proposed guidance strategy in different engagement scenarios. [ABSTRACT FROM AUTHOR]

Full-state single-loop and full-state two-loop autopilot-guidance architectures are derived under a linear quadratic differential game formulation. In the full-state single-loop case, the guidance command is injected directly to the actuator, whereas in the full-state two-loop case, it is the input to the autopilot loop. To prevent impractical end-game scenarios, where the states diverge to unacceptable values, a cost function that includes appropriate running cost terms on some of the states is proposed. The conditions for obtaining an equivalence relation between the full-state single-loop and full-state two-loop architectures are derived under a linear quadratic differential game formulation and the proposed cost function. Under such a formulation, the two full-state architectures are identical if and only if the number of guidance commands matches the number of available controllers. The guidance laws performance is illustrated using an interceptor missile having forward and aft controls in linear and nonlinear settings, while considering two types of evasion strategies. The first strategy is a linear controller based on the linear quadratic differential game solution. The second strategy is a “bang–bang” controller based on the optimal evasion solution. It is shown that the linear evasion strategy may not be suitable to represent a realistic evading strategy. In addition, the conditions for the existence of a saddle point solution are analyzed for the two full-state guidance laws. [ABSTRACT FROM AUTHOR]

Two guidance laws are discussed in this paper. The first one is the Vector guidance law. This guidance law when equipped with the appropriate gains has the capability to hit a target at the desired impact angle. A parametric method to find the optimal gains of this guidance law which will maximize the impact velocity and keep the miss distance and impact angle errors within bounds is developed in this research. Further, it is seen that the separation in the upper and lower bounds increases with increase in one of the gain values. Also, it is found that, only one of the gain values is independent and that the other dependent gain value is related through a simple straight-line expression. Next guidance law to be discussed is the Diveline guidance law. This law uses multiple divelines to hit a target at the desired impact angle. In the present research, the capability of this Diveline guidance law using a single diveline is analyzed. A method to derive the Diveline guidance law from the Vector guidance law is given in this study. The miss distance and impact angle errors evolving because of reducing the maximum acceleration limit is studied using simulations. Finally, three methods to increase the capture region (i.e. bounds on the set of initial states to achieve zero errors) of a guidance law is discussed. [ABSTRACT FROM AUTHOR]

This paper investigates the problem of salvo attack guidance law design for multiple missiles against a stationary target considering communication delay. The proposed guidance law requires no information on time-to-go and consists of two guidance stages. More specifically, a simple decentralized midcourse guidance law is proposed to provide desired initial conditions for the latter phase. As for the second guidance phase, all missiles are switched to classical pure proportional navigation guidance law. Considering the fact that the communication delay is inevitable in real applications, the corresponding allowed maximum upper bound of the communication delay among all interceptors is also derived analytically. Nonlinear numerical simulations clearly confirm the effectiveness of the proposed formulation. [ABSTRACT FROM AUTHOR]

This paper addresses energy-optimal guidance and control of satellite relative motion for formation flying and rendezvous using impulsive maneuvers. To meet the requirements of future multisatellite missions, closed-form solutions of the inverse relative dynamics are sought in arbitrary orbits. Time constraints dictated by mission operations and relevant perturbations acting on the formation are accounted for by splitting the optimal reconfiguration into guidance (long-term) and control (short-term) layers. Both problems are cast in relative orbit element space, which allows the simple inclusion of secular and long-periodic perturbations through a state-transition matrix and the translation of the energy-optimal optimization into a minimum-length path-planning problem. From the choice of state variables, both guidance and control problems can be solved in closed form, leading to optimal, predictable maneuvering schemes for simple onboard implementation. Besides generalizing previous work, this paper finds five new in-plane and out-of-plane closed-form solutions to the relative motion control problem in unperturbed eccentric and perturbed near-circular orbits. A delta-v lower bound is formulated, which provides insight into the optimality of the control solutions. Finally, the functionality and performance of the new maneuvering schemes are rigorously assessed through numerical integration of the equations of motion and systematic comparison with primer vector optimal control. [ABSTRACT FROM AUTHOR]