Synthetic-Waypoint Guidance Algorithm for Following a Desired Flight Trajectory

Hardware in the loop simulation for homing missiles

This paper describes rapid prototyping of a guidance and control system for missiles to improve its overall design process eectiveness. A rapid prototyping design process is developed based on MATLAB/Simulink/RTW and Skyfly. An air-to-surface missile is used as a design example. A high fidelity 6-DOF missile simulation model is built for rapid prototyping test and a trimmed simulation model is presented as validation and linear equations of motion are provided for guidance and control analysis. Simulation study results indicate that the proposed rapid prototyping process is practically feasible for eectively developing a guidance and control system of an air-to-surface missile.

When I was a child, my father worked for the Naval Ordinance Plant in Forest Park, Illinois as an electrical engineer. He traveled frequently, flying out of Midway Airport or taking the train. For short trips during the winter, the train was almost as fast and more reliable. He traveled to Seattle, and these trips involved two or more stops to refuel. After O’Hare Airport opened, DC-6 and Lockheed Tri-Star planes began to support nonstop flights to Seattle with significant time savings. When 707s and DC-8s arrived, the long flights became much shorter. My family would take our father to the airport to say goodbye and head to the observation deck to watch his plane take off. Plane delays because of “mechanical” problems were more often than not issues with mechanical gyros. These delays were substantial, and he would join us on the observation deck to pass the time. Because of weather conditions on these flights, safe, efficient travel depended heavily on the aircraft guidance system. My first flight on an airplane was to Seattle in 1970 during a college Christmas break. It was a red eye on a 747 repositioning flight with about as many crew as passengers on board. Being a math and physics major, I thought at the time it would be wonderful to play a part in the design of this type of plane.

Steel boned carbides are reliable guides for China's space flight programme

In this paper, the dynamical model in a matrix second-order nonlinear form with respect to the three Euler angles is firstly established for the guidance system of a rigid missile, which is complete in the sense that no approximation is taken, and automatically turns out to be a fully-actuated one. Then, with the help of a recently proposed general parametric design approach for general fully-actuated second-order nonlinear systems, a direct parametric approach for missile guidance design via proportional plus derivative feedback is proposed, which gives a complete parametrization of the pair of feedback gains, and allows usage of the established complete model but not a simplified one. The approach possesses two important features. Firstly, with the proposed controller parametrization, the missile guidance system, though highly nonlinear, can be turned into a constant linear system with desire eigenstructure. Secondly, in such a design there are still degrees of freedom which may be further utilized to improve the system performance. An example is considered to demonstrate the use of the proposed approach.

The main problem in bringing autonomy to any vehicle lies in the design of a suitable guidance law. For truly autonomous operation, the vehicle needs to have a reliable Navigation, Guidance and Control (NGC) system of which the guidance system is the key element that generates suitable trajectories to be followed. In this review paper, various guidance laws found in the literature and their relevance to autonomous underwater vehicles (AUVs) are discussed. Since existing guidance laws for underwater vehicles have emulated from tactical airborne missile systems, a number of approaches for the missile guidance systems are considered. Finally, potential guidance strategies for AUVs are proposed.

Impact point prediction guidance of ballistic missile in high maneuver penetration condition

With the development and validation of flight control devices, electronic equipment is widely used in current flight vehicles, which results in a series of control signals in the discrete-time domain. However, many mature integrated guidance and control schemes are designed in the continuous-time domain, and some results of them may not be preserved in the discrete-time domain especially for the sliding mode control. To address this issue, a chattering-free discrete-time sliding mode control scheme is proposed for the integrated guidance and control systems of missiles. By applying the Euler’s discretization method, an approximated discrete-time model with matched and mismatched disturbances is first derived for integrated guidance and control system. Then, the estimates of the disturbances are obtained by using discrete-time generalized proportion integral observers. Integrating the estimated disturbances into the sliding mode variable, a discrete-time sliding mode controller is established by utilizing a modified chattering-free reaching law. The reachability of quasi-sliding mode band, states boundedness, and output convergence performance are analyzed rigorously. Finally, comparative numerical simulations are conducted to verify the effectiveness of the proposed control scheme in the presence of unknown target maneuvers and parameter uncertainties.

The results of development of formalized model of process of formation of decisive function of combined correlation-extreme navigation system of unmanned aerial vehicles are presented. The development of the model is carried out taking into account the state of the operational and tactical situation, as well as the effect of changing the geometry of the sighting. The peculiarity of the developed model consists in the formation of a decisive function using a set of reference images, which is necessary for the formation of a decisive function, the set of reference images is built taking into account changes in navigation parameters of an unmanned aerial vehicle and angular orientation parameters. The originality of the formalized model consists in the determined order of formation of the decisive function of the combined correlation-extreme navigation system, which, unlike the known ones, consists in step-by-step determination of coordinates of the reference object by clarifying the navigation parameters of unmanned aerial vehicles with subsequent refinement of the angular orientation parameters according to the following sequence: determining the height of formation of the current image, determining the angles of view and determining the angular orientation parameters. Each of the individual reference images corresponds to its own multidimensional matrix of luminance values of the corresponding pixels or radio-thermal temperature. This takes into account the area of objects located on the sighting surface, which leads to refinement of the coordinates of the reference object in the current image. Changes in the spatial position and orientation of unmanned aerial vehicles, geometric properties of the sighting surface, as well as taking into account variations in informative features of the sighting surface objects depending on external factors and the influence of the medium on the propagation of signals in the optical and radiometric ranges of waves and obstacles, which will reduce or completely eliminate the discrepancy of coordinates of the reference object to the current image determined in advance in the reference image.

This paper describes the development of a proposed framework of metrics for the evaluation of the performance of aircraft guidance systems. The methodologies and metrics developed remain generally agnostic to whether or not the aircraft is manned. Although more complicated missions such as autonomous exploration/search, ferry, surveillance, multi-agent collaboration, and manned flight may be addressed at a later time, A-B flight scenarios are chosen to study the proposed metrics. The proposed metrics will form building blocks for the more complicated missions. Metrics development has thus far generally focused on NOE flight, and in particular on the observability of the vehicle throughout its mission. That is, a formulation of probability of detection by potential and generally unknown threats in the mission area will be the main metric. Secondary metrics provide insight into the vehicle's trajectory quality in terms of safety and comfort, experienced by both humans and machines are described as well. Scalability of the benchmarking system is also important and benchmarking should be general enough to allow guidance algorithms to be graded independently of the vehicle platform, for instance. Non-dimensionalization metrics will address this concern.

In this paper, the performance of two distinct classes of feedback guidance algorithms is evaluated for a spacecraft rendezvous problem utilizing a continuous low-thrust propulsion system. They are the DG (Differential Geometric) and ZEM/ZEV (Zero-Effort-Miss/Zero- Effort-Velocity) feedback guidance algorithms. Even though these two guidance algorithms do not attempt to minimize the onboard fuel consumption or ΔV directly, the ΔV requirement is used as a measure of their orbital rendezvous performance for various initial conditions and a wide range of the rendezvous time (within less than one orbital period of the target vehicle). For the DG guidance, the effects of its guidance parameter and terminal time on the closed-loop performance are evaluated by numerical simulations. For the ZEM/ZEV guidance, its nearfuel- optimality is further demonstrated for a rapid, short-range orbital rendezvous, in comparison with the corresponding open-loop optimal solutions. Furthermore, the poor ΔV performance of the ZEM/ZEV guidance for a slow, long-range orbital rendezvous is remedied by simply adding an initial drift phase. The ZEM/ZEV feedback guidance algorithm and its appropriate variants are then shown to be a simple practical solution to a non-impulsive rendezvous problem, in comparison with the DG guidance as well as the open-loop optimal guidance.

Optimal explicit guidance for three-dimensional launch trajectory

The Operational Procedure Information Model, presented in this paper, provides a formalism for the specification of the behavior of operationally embedded reactive systems found in aircraft guidance and navigation systems. The information model assigns semantic interpretations of the operational procedure construct to the elements of a finite state machine. The operational procedure construct captures the embedded operational behavior of the system over all the missions in the life‐cycle. The finite state machine captures the reactive behavior of the system.The information model, captured in a database and interrogated through a graphical user‐interface, can be used for simulation, analysis, and the generation of code and documentation.

A dynamic model of a two-axis rate gyro-stabilized platform-based seeker with cross-coupling, mass imbalance, and disturbance torque is developed on the basis of the working principle of seeker two-loop steady tracking theory; coordinate transformations are used to analyze the effects of seeker servo control mode on missile guidance and control systems. Frequency domain is used to identify the servo motor transfer function. Furthermore, a block diagram of the two-gimbal-coupled system is developed, and the coupling characteristics of gimbal angle are analyzed with different missile body inputs. Simulation results show that the analysis conforms with the actual movement rule of seeker gimbal and optical axis, and cross-coupling exists between the two gimbals. The lag compensation network can increase the open loop gain and increase the capacity for disturbance rate rejection. Simulations validate the theory and technology support for developing the seeker servo control model in engineering.

Real-world applications of computational intelligence can enhance the fault detection and identification capabilities of a missile guidance and control system. A simulation of a bank-to- turn missile demonstrates that actuator failure may cause the missile to roll and miss the target. Failure of one fin actuator can be detected using a filter and depicting the filter output as fuzzy numbers. The properties and limitations of artificial neural networks fed by these fuzzy numbers are explored. A suite of networks is constructed to (1) detect a fault and (2) determine which fin (if any) failed. Both the zero order moment term and the fin rate term show changes during actuator failure. Simulations address the following questions: (1) How bad does the actuator failure have to be for detection to occur, (2) How bad does the actuator failure have to be for fault detection and isolation to occur, (3) are both zero order moment and fine rate terms needed. A suite of target trajectories are simulated, and properties and limitations of the approach reported. In some cases, detection of the failed actuator occurs within 0.1 second, and isolation of the failure occurs 0.1 after that. Suggestions for further research are offered.