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This study utilizes Automatic Dependent Surveillance–Broadcast (ADS-B) data sourced by the OpenSky Network to curate a dataset aimed at enhancing the precision of aerial suppressant drop predictions in wildland firefighting. By amalgamating ADS-B data with Automated Telemetry Unit (ATU) drop information, this research constructs a reliable base for analyzing the spatial aspects of aerial firefighting operations. Using sequential machine learning models, specifically Long Short-Term Memory (LSTM) networks and 1D Convolutional Neural Networks (1DCNN), the study interprets complex flight dynamics to predict drop locations. The dataset, covering 2017 to 2023, is labeled and segmented to reflect accurate suppressant release events, facilitating the distinction between drop and non-drop activities in fixed-wing aircraft. The LSTM model demonstrated strong predictive performance with an F1 score of 0.922, effectively identifying suppressant drop events with high accuracy. This model's reliable predictions can significantly improve situational awareness in real-time aerial firefighting operations, enabling more informed decision-making and better coordination of resources during wildfire events. [ABSTRACT FROM AUTHOR]

In the present paper, a flight dynamics model is adopted to represent the trim and stability characteristics of a side-by-side helicopter in hovering conditions. This paper develops a numerical representation of the rotorcraft behavior and proposes a set of guidelines for trimming and linearizing the highly coupled rotor dynamics derived by the modeling approach. The trim algorithm presents two nested loops to compute a solution of the steady-state conditions averaged around one blade's revolution. On the other hand, a 38-state-space linear representation of the helicopter and rotor dynamics is obtained to study the effects of flap, lead–lag, and inflow on the overall stability. The results are compared with an analytical framework developed to validate the rotorcraft stability and compare different modeling approaches. The analysis showed that non-uniform inflow modeling led to a coupled longitudinal inflow–phugoid mode which made the vehicle prone to dangerous instabilities. The flap and lead–lag dynamics introduced damping in the system and can be considered beneficial for rotor dynamics. [ABSTRACT FROM AUTHOR]

Unmanned aerial vehicles (UAVs) have seen a rise in use during the last few years. Such aircrafts are now a convenient way to complete dangerous, dirty, and tedious tasks. Given that their operation involves a control problem which is non-linear and coupled, it is difficult to analyse. This paper presents the modeling and control of a fixed-wing unmanned aircraft as a contribution to this field. The system's flight dynamics is derived using Newton's second law of motion. The system is designed to have a non-linear Proportional Integral Derivative (NPID) controller and a higher-order sliding mode controller (HOSMC). When simulating the system using MATLAB Simulink software, an external disturbance was added to test the robustness of the controllers. Five performance indices which include mean square error (MSE), integral time square error (ITSE), integral absolute error (IAE), integral time absolute error (ITAE), and integral square error (ISE), were used to compare the controllers performance. These indices are used to provide a numerical assessment of the two controllers' performance. The outcomes demonstrate that the roll, pitch, and yaw states performed better than the super-twisting sliding mode controller. On the airspeed control, the non-linear PID performed better than the super-twisting sliding mode controller. [ABSTRACT FROM AUTHOR]

Advanced Aerial Mobility (AAM) platforms are poised to begin high-density operations in urban areas nationwide. This new category of aviation platforms spans a broad range of sizes, from small package delivery drones to passenger-carrying vehicles. Unlike traditional aircraft, AAM vehicles operate within the urban boundary layer, where large structures, such as buildings, interrupt the flow. This study examines the response of a package delivery drone, a general aviation aircraft, and a passenger-carrying urban air mobility aircraft through an urban wind field generated using Large Eddy Simulations (LES). Since it is burdensome to simulate flight dynamics in real-time using the full-order solution, reduced-order wind models are created. Comparing trajectories for each aircraft platform using full-order or reduced-order solutions reveals little difference; reduced-order wind representations appear sufficient to replicate trajectories as long as the spatiotemporal wind field is represented. However, examining control usage statistics and time histories creates a stark difference between the wind fields, especially for the lower wing-loading package delivery drone where control saturation was encountered. The control saturation occurrences were inconsistent across the full-order and reduced-order winds, advising caution when using reduced-order models for lightly wing-loaded aircraft. The results presented demonstrate the effectiveness of using a simulation environment to evaluate reduced-order models by directly comparing their trajectories and control activity metrics with the full-order model. This evaluation provides designers valuable insights for making informed decisions for disturbance rejection systems. Additionally, the results indicate that using Reynolds-averaged Navier–Stokes (RANS) solutions to represent urban wind fields is inappropriate. It was observed that the mean wind field trajectories fall outside the 95% confidence intervals, a finding consistent with the authors' previous research. [ABSTRACT FROM AUTHOR]

This article examines the flight dynamics and trajectory analysis of a parachute–payload system deployed from a C-17 aircraft. The aircraft is modeled with an open cargo door, extended flaps, and four turbo-fan engines operating at an altitude of 2000 feet Above Ground Level (AGL) and an airspeed of 150 knots. The payloads consist of simplified CONEX containers measuring either 192 inches or 240 inches in length, 9 feet in width, and 5.3 feet in height, with their mass and moments of inertia specified. At positive deck angles, gravitational forces cause these payloads to begin a gradual descent from the rear of the aircraft. For aircraft at zero deck angle, a ring-slot parachute with approximately 20% geometric porosity is utilized to extract the payload from the aircraft. This study specifically employs the CREATE-AV Kestrel simulation software to model the chute-payload system. The extraction and suspension lines are represented using Kestrel's Catenary capability, with the extraction line connected to the floating confluence points of the CONEX container and the chute. The chute and payload will experience coupled motion, allowing for an in-depth analysis of the flight dynamics and trajectory of both elements. The trajectory data obtained will be compared to that of a payload (without chute and cables) exiting the aircraft at positive deck angles. An adaptive mesh refinement technique is applied to accurately capture the engine exhaust flow and the wake generated by the C-17, chute, and payloads. Friction and ejector forces are estimated to align the exit velocity and timing with those recorded during flight testing. The results indicate that the simulation of extracted payloads aligns with expected trends observed in flight tests. Notably, higher deck angles result in longer distances from the ramp, leading to increased exit velocities and reduced payload rotation rates. All payloads exhibit clockwise rotation upon leaving the ramp. The parachute extraction method yields significantly higher exit velocities and shorter exit times, while the payload-chute acceleration correlates with the predicted drag of the chute as demonstrated in prior studies. [ABSTRACT FROM AUTHOR]

Introduction. A flight visibility model is presented that takes into account the nature of the underlying surface, the height of the flat terrain, and the dynamics of the aircraft flight due to changes in altitude and ground speed during aerial work. Materials and Methods. Hourly microring weather maps of the Central Russian Plain of the Moscow air hub were used. The research methodology is based on constructing a flight visibility model, taking into account the factors influencing it, including the terrain and the nature of the underlying surface. Results and their Discussion. Flight visibility significantly depends on the nature of the underlying surface, as well as on the elevation of the terrain relative to the departure point at a distance of up to 150 km. The influence of the elevation of the terrain relative to the departure point is 30% more significant than the influence of the nature of the underlying surface. Conclusions. The use of the constructed flight visibility model, which takes into account the influencing physical and geographical conditions, the nature of the underlying surface, the height of the relief of flat terrain under low clouds, will make it possible to make weather-dependent decisions through the use of an intelligent meteorological system, which leads to the improvement of meteorological support for flights during aviation operations.

Abstract Unmanned aerial vehicles (UAVs) have seen a rise in use during the last few years. Such aircrafts are now a convenient way to complete dangerous, dirty, and tedious tasks. Given that their operation involves a control problem which is non-linear and coupled, it is difficult to analyse. This paper presents the modeling and control of a fixed-wing unmanned aircraft as a contribution to this field. The system’s flight dynamics is derived using Newton’s second law of motion. The system is designed to have a non-linear Proportional Integral Derivative (NPID) controller and a higher-order sliding mode controller (HOSMC). When simulating the system using MATLAB Simulink software, an external disturbance was added to test the robustness of the controllers. Five performance indices which include mean square error (MSE), integral time square error (ITSE), integral absolute error (IAE), integral time absolute error (ITAE), and integral square error (ISE), were used to compare the controllers performance. These indices are used to provide a numerical assessment of the two controllers’ performance. The outcomes demonstrate that the roll, pitch, and yaw states performed better than the super-twisting sliding mode controller. On the airspeed control, the non-linear PID performed better than the super-twisting sliding mode controller.

During the operation of single-rotor helicopters, aviation accidents quite frequently occur in the form of an unintentional turn or even a yaw rotation, causing, as a rule, a ground collision. Numerous researchers of this problem consider the loss of helicopter tail-rotor effectiveness due to wind effects as one of its possible causes. There is even a special term – the Loss of Tail Rotor Effectiveness (LTE) in the foreign literature. Hence, this paper deals with an attempt to determine the capacity of an unintentional single-rotor helicopter yaw rotation occurrence due to wind effects (the impact of the main rotor on the tail rotor was not considered in this paper). To solve this issue, theoretical methods (analytical calculations and computational experiments) were used. To carry out analytical calculations and computational experiments, a mathematical model of the Mi-8MTV helicopter yaw rotation dynamics was developed, on the basis of which a software package integrating the LTE module (for modeling the dynamics of rotational yaw motion of the helicopter) and OGL (for helicopter motion visualization) was created. Analytical calculations revealed that the yaw angular acceleration value monitored in-flight during an unintentional rotation can manifest itself due to the tail rotor thrust loss in the vortex-ring state. But for the development of an unintentional rotation to angles and angular velocities recorded in real flights, that kind of tail rotor thrust loss should occur during the entire turn. In computational experiments using the mentioned above software package, conditions failed to be created for that kind of thrust loss during the entire turn. Consequently, those yaw angles and angular velocities, that occurred in flights, could not be reached. The tail rotor, when blown by the wind in the investigated range of wind velocities (from 1 to 20 m/s) does not lose its effectiveness to such an extent that an unintentional rotation cannot be stopped by means of the tail rotor.

Traditional flight simulation models often operate on the premise of a steady atmosphere, overlooking the complexities of actual atmospheric dynamics and the flight safety risks posed by wind disturbances, such as turbulence. To Address this oversight, the present study introduces a method for generating three-dimensional atmospheric turbulence based on spatial correlation functions. This method, rigorously validated against correlation and spectral benchmarks, guarantees isotropic properties in the synthesized turbulence fields. Through interpolation techniques, the model integrates the spatial atmospheric turbulence into the flight simulation framework effectively. The paper highlights the application of this model by examining the impact of atmospheric turbulence on the precise flight dynamics of quadcopter UAVs during aerial refueling operations. The findings demonstrate the model's pertinence not only to UAVs but also to the broader spectrum of aircraft and their operational procedures. [ABSTRACT FROM AUTHOR]

This paper explores the influence of engine modernization on the dynamic stability and performance of a general aviation aircraft. Utilizing an integral mathematical model, the study conducts a comparative analysis of the I-23 Manager piston-engine aircraft and its modified I-31T turboprop version, examining changes in aircraft dynamics depending on the power plant type. The modernization necessitated a redesign of the nose section of the fuselage, resulting in alterations to the external shape and flight properties of the aircraft. The research evaluates various dynamic stability parameters, including phugoid, short period, Dutch roll, roll, and spiral modes, under different flight conditions. Results indicate minimal changes in aerodynamic characteristics due to the engine type, yet significant improvements are observed in efficiency, noise reduction, and operational costs. The impact of the propulsion unit on the dynamic stability of the light aircraft was assessed as insignificant, suggesting that the strategy of modernizing an existing piston-driven aircraft by switching to a turboprop drive is indeed promising. With appropriate initial design assumptions, a modern turbine aircraft with strong flight qualities can be efficiently modernized in this way, without compromising the good flying properties of the existing plane. The outcomes are validated against flight tests, reinforcing the viability of integrating more sustainable and efficient propulsion systems into light aircraft. This study may therefore inform future design and regulatory decisions, providing a perspective on the implications of engine upgrades in the general aviation sector. [ABSTRACT FROM AUTHOR]

Investigating aircraft flight dynamics often requires dynamic wind tunnel testing. This paper proposes a non-contact, off-board instrumentation method using vision-based techniques. The method utilises a sequential process of Harris corner detection, Kanade–Lucas–Tomasi tracking, and quaternions to identify the Euler angles from a pair of cameras, one with a side view and the other with a top view. The method validation involves simulating a 3D CAD model for rotational motion with a single degree-of-freedom. The numerical analysis quantifies the results, while the proposed approach is analysed analytically. This approach results in a 45.41% enhancement in accuracy over an earlier direction cosine matrix method. Specifically, the quaternion-based method achieves root mean square errors of 0.0101 rad/s, 0.0361 rad/s, and 0.0036 rad/s for the dynamic measurements of roll rate, pitch rate, and yaw rate, respectively. Notably, the method exhibits a 98.08% accuracy for the pitch rate. These results highlight the performance of quaternion-based attitude estimation in dynamic wind tunnel testing. Furthermore, an extended Kalman filter is applied to integrate the generated on-board instrumentation data (inertial measurement unit, potentiometer gimbal) and the results of the proposed vision-based method. The extended Kalman filter state estimation achieves root mean square errors of 0.0090 rad/s, 0.0262 rad/s, and 0.0034 rad/s for the dynamic measurements of roll rate, pitch rate, and yaw rate, respectively. This method exhibits an improved accuracy of 98.61% for the estimation of pitch rate, indicating its higher efficiency over the standalone implementation of the direction cosine method for dynamic wind tunnel testing. [ABSTRACT FROM AUTHOR]

In this paper, the performance of Model-Free Adaptive Control (MFAC) has been investigated on a novel and specific moving-mass controlled (MMC) flying robot system. The novel one-degree-of-freedom (1 DOF) MMC flying robot test bed presented in this paper has highly nonlinear and slow dynamics with a variable center of gravity (CoG) and moment of inertia. This makes the control of this system a challenging problem. One of the solutions to this challenge is the use of data-driven control methods, in particular, MFAC. This controller uses a data-driven model to control the system using only input and output (I/O) data. This paper compares this data-driven controller with proportional-integral-derivative (PID) control, and Linear Quadratic Regulator (LQR) as two model-free and model-based controllers which are widely used controllers in industry. The results of the comparison show that in the various scenarios applied, MFAC has a clear superiority over the PID and LQR, and its adaptive structure gives more freedom of action in the implementation of different scenarios and the presented noise. The results are obtained using the Integral Time Absolute Error (ITAE) criteria and the mean maximum error has also been compared in a Monte Carlo analysis. For a more detailed study, the amount of control energy consumption was also compared, which showed a clear superiority of the MFAC. Also, the robustness of the controller was demonstrated by introducing uncertainty in the plant parameters and by running 100 Monte Carlo simulations with random initial conditions. Finally, despite the PID controller, the MFAC followed the desired scenarios well and compared to LQR consumed less energy. The results demonstrate that the MFAC outperformed the PID and LQR controllers in the presence of random initial conditions and noise in terms of mean maximum error (70.4 %) , mean ITAE (91 %) , and energy consumption (46 %) . [ABSTRACT FROM AUTHOR]

This paper investigates the roll maneuvers of a very flexible, transonic high-aspect-ratio-wing aircraft with flared folding wingtips, representing a potential future commercial transport configuration. The computational study leverages a fully coupled nonlinear aeroelastic-flight dynamics framework for simulating roll maneuvers commanded by deflecting control surfaces inboard or outboard of the flared folding wingtip hinge. Releasing the flared folding wingtips during flight makes the aircraft roll faster only when deflecting outboard-of-hinge leading-edge control surfaces. However, the achieved roll angles and rates are much lower than the ones obtained by deflecting inboard-of-hinge trailing-edge control surfaces, whose roll control effectiveness is not impacted or even degrades when releasing the flared folding wingtips. These trends are explained by analyzing the role of sweep-induced washout effects in the very flexible, transonic high-aspect-ratio wing, revealing physical mechanisms not identified in previous studies on straight or swept wings with moderate flexibility. Higher wing out-of-plane bending stiffness increases the effectiveness of flared folding wingtips in enhancing roll maneuvers at low dynamic pressure but plays a slight role at high dynamic pressure. This study highlights the need to consider the impact of washout effects for different flight conditions and levels of wing flexibility when assessing the ability of flared folding wingtips to enhance the roll maneuverability of transonic high-aspect-ratio-wing aircraft. [ABSTRACT FROM AUTHOR]

Tailless Unmanned Aerial Vehicles (UAVs) offer several advantages over their conventional counterparts, including enhanced maneuverability, higher durability, and better stealth. However, they are challenging in their stability and control due to the absence of stabilizing and control surfaces that typically exist in conventional empennage. A crucial process for analyzing the tailless UAVs’ stability/performance characteristics and determining/validating their flight control parameters is done via the modeling and simulation of flight dynamics. This paper presents a comprehensive and systematic procedure for investigating the flight dynamics of a tailless UAV, including modeling, simulation, analytical verification, and flight testing, while also explaining the interconnections among these elements. It also addresses the common challenge of limited accessibility of UAV essential data through using diverse analytical, empirical, and experimental methods. First, a rapid and effective first-principles modeling approach is introduced to simulate the nonlinear six-degree-of-freedom flight dynamics of a small tailless UAV case study. The modeling process follows a modular framework where well-defined experiments and commercial of-the-shelf software, tools, and sensors are employed to build the necessary sub-models, including geometric, mass–inertia, aerodynamic, propulsion, and actuator models. Then, all sub-models are integrated into a simulation environment to allow the prediction of the UAV dynamic response obtained from the given control inputs. The developed flight dynamic model is subjected to a thorough verification process to ensure its integrity and proper functionality by comparing the simulated trim and natural flight modes with the calculated analytical results. Finally, a set of specific flight tests are performed to validate the developed simulation model and verify relevant performance characteristics for the case-study UAV. The results show that the proposed approach provides a systematic and straightforward method for examining the flight dynamics of small tailless UAVs with reasonable accuracy. [ABSTRACT FROM AUTHOR]

A helicopter landing flow field calculation method based on Navier Stokes equations is established, which can be applied to the calculation of aerodynamic forces on the rotor, fuselage and tail rotor in the landing area. The method uses a cut-body grid for meshing, an implicit coupling solver and a dual-time scheme to simulate the flow field during the landing process. To improve the computational efficiency, the main rotor and tail rotor are modeled as momentum sources in the flow field. The method also employs a polynomial fitting technique to correct the results of the nonlinear flight dynamics model of the helicopter and achieve trim calculations in complex flow fields. The method is applied to the UH-60A helicopter and the SFS2 ship as case studies. First, a single-helicopter trim analysis coupled CFD method is performed to verify its validity. Then, the helicopter safe landing criteria are introduced and used to calculate the theoretical ship helicopter operational limitation for this helicopter/ship combination. [ABSTRACT FROM AUTHOR]

In the present paper, a flight dynamics model is adopted to represent the trim and stability characteristics of a side-by-side helicopter in hovering conditions. This paper develops a numerical representation of the rotorcraft behavior and proposes a set of guidelines for trimming and linearizing the highly coupled rotor dynamics derived by the modeling approach. The trim algorithm presents two nested loops to compute a solution of the steady-state conditions averaged around one blade’s revolution. On the other hand, a 38-state-space linear representation of the helicopter and rotor dynamics is obtained to study the effects of flap, lead–lag, and inflow on the overall stability. The results are compared with an analytical framework developed to validate the rotorcraft stability and compare different modeling approaches. The analysis showed that non-uniform inflow modeling led to a coupled longitudinal inflow–phugoid mode which made the vehicle prone to dangerous instabilities. The flap and lead–lag dynamics introduced damping in the system and can be considered beneficial for rotor dynamics.

This article examines the flight dynamics and trajectory analysis of a parachute–payload system deployed from a C-17 aircraft. The aircraft is modeled with an open cargo door, extended flaps, and four turbo-fan engines operating at an altitude of 2000 feet Above Ground Level (AGL) and an airspeed of 150 knots. The payloads consist of simplified CONEX containers measuring either 192 inches or 240 inches in length, 9 feet in width, and 5.3 feet in height, with their mass and moments of inertia specified. At positive deck angles, gravitational forces cause these payloads to begin a gradual descent from the rear of the aircraft. For aircraft at zero deck angle, a ring-slot parachute with approximately 20% geometric porosity is utilized to extract the payload from the aircraft. This study specifically employs the CREATE-AV Kestrel simulation software to model the chute-payload system. The extraction and suspension lines are represented using Kestrel’s Catenary capability, with the extraction line connected to the floating confluence points of the CONEX container and the chute. The chute and payload will experience coupled motion, allowing for an in-depth analysis of the flight dynamics and trajectory of both elements. The trajectory data obtained will be compared to that of a payload (without chute and cables) exiting the aircraft at positive deck angles. An adaptive mesh refinement technique is applied to accurately capture the engine exhaust flow and the wake generated by the C-17, chute, and payloads. Friction and ejector forces are estimated to align the exit velocity and timing with those recorded during flight testing. The results indicate that the simulation of extracted payloads aligns with expected trends observed in flight tests. Notably, higher deck angles result in longer distances from the ramp, leading to increased exit velocities and reduced payload rotation rates. All payloads exhibit clockwise rotation upon leaving the ramp. The parachute extraction method yields significantly higher exit velocities and shorter exit times, while the payload-chute acceleration correlates with the predicted drag of the chute as demonstrated in prior studies.

Advanced Aerial Mobility (AAM) platforms are poised to begin high-density operations in urban areas nationwide. This new category of aviation platforms spans a broad range of sizes, from small package delivery drones to passenger-carrying vehicles. Unlike traditional aircraft, AAM vehicles operate within the urban boundary layer, where large structures, such as buildings, interrupt the flow. This study examines the response of a package delivery drone, a general aviation aircraft, and a passenger-carrying urban air mobility aircraft through an urban wind field generated using Large Eddy Simulations (LES). Since it is burdensome to simulate flight dynamics in real-time using the full-order solution, reduced-order wind models are created. Comparing trajectories for each aircraft platform using full-order or reduced-order solutions reveals little difference; reduced-order wind representations appear sufficient to replicate trajectories as long as the spatiotemporal wind field is represented. However, examining control usage statistics and time histories creates a stark difference between the wind fields, especially for the lower wing-loading package delivery drone where control saturation was encountered. The control saturation occurrences were inconsistent across the full-order and reduced-order winds, advising caution when using reduced-order models for lightly wing-loaded aircraft. The results presented demonstrate the effectiveness of using a simulation environment to evaluate reduced-order models by directly comparing their trajectories and control activity metrics with the full-order model. This evaluation provides designers valuable insights for making informed decisions for disturbance rejection systems. Additionally, the results indicate that using Reynolds-averaged Navier–Stokes (RANS) solutions to represent urban wind fields is inappropriate. It was observed that the mean wind field trajectories fall outside the 95% confidence intervals, a finding consistent with the authors’ previous research.

Launched on 2 June 2003 and arriving at Mars on 25 December 2003 after a 7-month interplanetary cruise, Mars Express was the European Space Agency's first mission to arrive at another planet. After more than 20 years in orbit, the spacecraft and science payload remain in good health and the mission has become the second oldest operational planetary orbiter after Mars Odyssey. This contribution summarizes the Mars Express mission operations, science planning and data archiving systems, processes, and teams that are necessary to run the mission, plan the scientific observations, and execute all necessary commands. It also describes the data download, the ground processing and distribution to the scientific community for the study and analysis of Mars sub-surface, surface, atmosphere, magnetosphere, and moons. This manuscript also describes the main challenges throughout the history of the mission, including several potentially mission-ending anomalies. We summarize the evolution of the ground segment to provide new capabilities not envisaged before launch, whilst simultaneously maintaining or even increasing the quality and quantity of scientific data generated. [ABSTRACT FROM AUTHOR]

Sounding rockets constitute a class of rocket with a generally simple layout, being composed of a cylindrical center-body, a nosecone, a number of fins placed symmetrically around the longitudinal axis (usually three or four), and possibly a boat-tail. This type of flying craft is typically not actively controlled; instead, a passive stabilization effect is obtained through suitable positioning and sizing of the fins. Therefore, in the context of dynamic performance analysis, the margin of static stability is an index of primary interest. However, the classical approach to static stability analysis, which consists in splitting computations in two decoupled domains, namely, around the pitch and yaw axis, provides a very limited insight to the missile performance for this type of vehicle due to the violation of the classical assumptions of planar symmetry and symmetric flight conditions commonly adopted for winged aircraft. To tackle this issue, this paper introduces a method for analyzing static stability through a novel index, capable of more generally assessing the level of static stability for sounding rockets, exploiting the same information on aerodynamic coefficients typically required for more usual (i.e., decoupled) static stability analyses, and suggests a way to assess the validity and shortcoming of the method in each case at hand. [ABSTRACT FROM AUTHOR]

The present study aimed to investigate the capability of mid-fidelity aerodynamic solvers in performing a preliminary evaluation of the static and dynamic stability derivatives of aircraft configurations in their design phase. In this work, the mid-fidelity aerodynamic solver DUST, which is based on the novel vortex particle method (VPM), was used to perform simulations of the static and dynamic motion conditions of the Stability And Control CONfiguration (SACCON): an unmanned combat aerial vehicle geometry developed by NATO's Research and Technology Organisation (RTO), which is used as a benchmark test case in the literature for the evaluation of aircraft stability derivatives. Two different methods were exploited to extract the dynamic stability derivative values from the aerodynamic coefficient time histories that were calculated with DUST. The results for the mid-fidelity approach were in good agreement with the obtained experimental data, as well as with the results obtained using more demanding high-fidelity CFD simulations. This demonstrates its suitability when implemented in DUST for predicting the static and dynamic behavior of airloads in different conditions, as well as in reliably predicting the values of stability derivatives, with the advantage of requiring limited computational effort with respect to classical high-fidelity numerical approaches and the use of wind tunnel tests. [ABSTRACT FROM AUTHOR]

This study presents the feasibility of implementing commercial-off-the-shelf reciprocating piston compressors for a developed active flow control (AFC) actuation framework to quantify the aerospace relevant performance ramifications of control architecture and compressor operational choices. Three architectures are studied: supply valve metering, exit area metering, and a combined approach. All concepts are studied under varying compressor operation schedules. The analysis framework in this study integrates internal pneumatic actuation and discharge dynamics, an experimentally calibrated compressor pressure and thermal dynamics model, three feedback control architectures, and flight dynamics models. The framework is implemented in simulation to provide a user-friendly tool for linking AFC architecture choices to achievable flight trajectories. Actuator performance is evaluated using actuation time, output, compressor duty cycle, and specific energy consumption. Aircraft tracking performance is evaluated as usable time and slalom centerline deviation. The analysis indicates that a supply-volume-based metering approach is comparatively inefficient concerning an exit-area-based metering, resulting in high flight tracking error. Exit area metering provides the best efficiency and run time with some structural drawbacks, while the combined approach provides the best flight tracking performance at the expense of additional complexity. These results inform the choice of onboard AFC hardware choices. [ABSTRACT FROM AUTHOR]

Rapid faults diagnosis and isolation in flight control systems is vital to evade undesirable effects on the environment, humans as well as on the system itself. In this research, a new strategy based on an online dynamical system is proposed for sensors fault diagnosis and isolation of quadrotor unmanned aerial vehicle (UAV). In the presented strategy, the multi-layer neural network (MLNN) is adopted as an observer for sensor faults diagnosis of quadrotor unmanned aerial vehicle (UAV). For the improvement of faults detection accuracy and effectiveness, two new adaptive weight-updating approaches are used in this study. Lyapunov function theory-based approaches are proposed for updating the learning rate parameters of a multi-layer neural network (MLNN). The key purpose of utilizing these approaches is to attain the global minima for nonlinear functions without enhancement of computational power to enhance the accuracy of faults diagnosis and isolation. The proposed fault detection (FD) approaches are applied to quadrotor unmanned aerial vehicle (UAV), the simulated results demonstrate that the proposed approaches are having the capability of rapid faults diagnosis and isolation of sensors compared to the conventional neural network and the techniques used in the literature (Chen et al. in IEEE Trans Syst Man Cybern Syst Hum 46:260–270, 2016a, b, c, d; Payam et al. in J Intell Robot Syst 90:473–484, 2018). [ABSTRACT FROM AUTHOR]

This research introduces a flight controller for a high-performance aircraft, able to follow randomly generated sequences of waypoints, at varying altitudes, in various types of scenarios. The study assumes a publicly available six-degree-of-freedom (6-DoF) rigid aeroplane flight dynamics model of a military fighter jet. Consolidated results in artificial intelligence and deep reinforcement learning (DRL) research are used to demonstrate the capability to make certain manoeuvres AI-based fully automatic for a high-fidelity nonlinear model of a fixed-wing aircraft. This work investigates the use of a deep deterministic policy gradient (DDPG) controller agent, based on the successful applications of the same approach to other domains. In the particular application to flight control presented here, the effort has been focused on the design of a suitable reward function used to train the agent to achieve some given navigation tasks. The trained controller is successful on highly coupled manoeuvres, including rapid sequences of turns, at both low and high flight Mach numbers, in simulations reproducing a prey–chaser dogfight scenario. Robustness to sensor noise, atmospheric disturbances, different initial flight conditions and varying reference signal shapes is also demonstrated. [ABSTRACT FROM AUTHOR]

Structural modes for large aircraft occur at low frequencies and must be accounted for in the flight dynamics modeling and control system design process. An analysis is performed on various linear representations of flexible aircraft with an application to the lateral/directional flexible dynamic response of the notional Large Civil Tiltrotor (LCTR2) in hover. The goal of the work is to compare the treatment of the coupling between the rigid-body and structural states within various model structures and to develop conversions between the various model forms. The analysis shows that mean-axis and other simplified analytical representations of flexible aircraft, originally developed for fixed-wing aircraft, are also able to correctly capture the dynamic response of a tiltrotor. First, a linear model is obtained from a multibody simulation using numerical perturbation methods. This is compared with a mean-axis model, where structural dynamics are appended onto rigid-body dynamics, as would be obtained if the rigid-body aircraft response were the starting point in an analytical development of the structural coupling. Comparisons are also given with a model that would be obtained from system identification using flight-test data. Key stability and control derivatives are compared to highlight similarities and differences between the various models. The results from the analysis show the multibody model treats the structural coupling in a significantly different way than an analytical model buildup. For example, the structural coupling to roll rate response varies up to two orders of magnitude between the models, yet the overall response is identical. This work shows analysis of the structural/rigid-body coupling developed for analytical models is also valid for models developed from a multibody analysis. Lastly, structural flexibility will be shown to alter the characteristics of the lateral hovering cubic of the LCTR2 by increasing damping of an unstable oscillatory mode. [ABSTRACT FROM AUTHOR]

The paper presents the development process of the solid rocket boosters (SRBs) separation system of the ILR-33 AMBER 2K rocket. A redesign of the system was required due to the development of new, larger SRBs. The main system requirements were transmission of forces and moments between the SRBs and the main stage, execution of the separation process at a given moment in flight and mechanical integration simplification. A set of aerodynamics calculations were performed. With the use of computational fluid dynamics software, forces acting on the booster during separation for several angles of attack, as well as the critical booster deflection angle, have been determined. Next, a mathematical model was created to define the load spectrum acting on the system during the flight and separation phases, covering both static and dynamic loads. All the internal and external force sources were considered. A series of motion dynamics simulations were conducted for representative flight cases. Then, the system operational parameters were verified with the use of dedicated ground test facilities. Necessary calibrations of the mathematical model were then implemented, leading to a high level of confidence with the empirical data obtained, thereby leading to a successful system qualification for the flight campaign. [ABSTRACT FROM AUTHOR]

The article provides information on the species composition (231 species belonging to 20 genera) and trophic relationships of Apidae family bees distributed in Azerbaijan. In 2 stationary locations (Gazakh region, Demirchilar village and Nakhchivan MR, Ordubad region, Aghdara), the daily activity of foraging of striped bees and its dependence on abiotic factors were studied. During the research period, 5 types of striped bees were recorded in the 1st station and 2 types in the 2nd station. It was determined that the most active flight of the striped bees is observed between 1200 and 1700 hours. They can already work when the air temperature is 8-10 °C, but the optimal temperature for stable flight is 22-24 °C. They are also more active in high humidity conditions. [ABSTRACT FROM AUTHOR]

Flying insects exhibit outperforming stability and control via continuous wing flapping even under severe disturbances in various conditions of wind gust and turbulence. While conventional linear proportional derivative (PD)-based controllers are widely employed in insect-inspired flight systems, they usually fail to deal with large perturbation conditions in terms of the 6-DoF nonlinear control strategy. Here we propose a novel wing kinematics-based controller, which is optimized based on deep reinforcement learning (DRL) to stabilize bumblebee hovering under large perturbations. A high-fidelity Open AI Gym environment is established through coupling a CFD data-driven aerodynamic model and a 6-DoF flight dynamic model. The control policy with an action space of 4 is optimized using the off-policy Soft Actor–Critic (SAC) algorithm with automating entropy adjustment, which is verified to be of feasibility and robustness to achieve fast stabilization of the bumblebee hovering flight under full 6-DoF large disturbances. The 6-DoF wing kinematics-based DRL control strategy may provide an efficient autonomous controller design for bioinspired flapping-wing micro air vehicles. [ABSTRACT FROM AUTHOR]

The article considers the restrictions algorithms for the trajectory parameters of aircraft motion. Three groups of algorithms are under consideration. They are distinguished by the volume of airborne database about terrain, modes of aircraft operation. The first group (the full SFITA mode) uses the maximum digital cartographic information (DCI) about the terrain. In the SFITA algorithms, the terrain is approximated in the form of planes in space in some predictable, pre-emptive area in the flight path direction. Terrain following is carried out in space in the assigned direction. The equation of the adjacent plane is solved. The distance, the approach speed, and the time to reach the plane are enumerated. The net acceleration to reduce the approach speed to the plane of restriction is calculated under the specified constraint controls. The required time is calculated to reduce the approach speed towards the obstacle to zero. Having equated the expressions for the time of reaching and the required time, transition towards the distance to the plane of restriction, on which it is imperative to utilize constraint controls. The second group (the major SFITA mode) uses the DCI about the terrain in the direction of track in a pre-emptive area. The terrain is approximated by a line in the plane. The specified normal overload is used as a means of control. Terrain following is conducted in the vertical or slant planes. Subsequently, the same procedures are used as in the first group. The third group (the minimum SFITA mode) does not use DCI. The radio and barometric pressure altimeters are used as information systems about the terrain. The given SFITA mode is selected only for flights in a flat terrain. The algorithm includes the similar procedures as in the first and in the second groups. The analytical analysis, confirming the adaptive properties of algorithmic support based on the fundamental law of uniformly retarded motion, is given. The considered algorithms efficiency is confirmed by a comprehensive amount of simulation. The presented algorithms can become the foundation for developing Russian TAWS analogues.

The mathematical model was established for the flight process of the molten blast furnace slag after the centrifugal graining, and the model was discretely solved by the Runge-Kutta method. The results show that, the flight distance of the slag droplets along the x direction is proportional to the diameter and the initial velocity of droplets. Due to the air flow resistance and gravity, the velocity of droplets decreases with time, and then increases slightly. Critical impact velocity of the droplets was proposed by analyzing the excess rebound energy of the slag droplets after collision with wall. The results indicate that, the critical impact velocity is an interval, and the larger the droplets, the lower the bounds. The experiments were performed with the initial droplet velocities of 10, 12, and 14 m·s‒1. The results indicate that, the actual falling distance of the slag droplets is greater than the theoretical values, because the initial velocity of the slag droplets is less than the linear velocity of the granulation plate. Meanwhile, the impact velocity of the slag droplets is between the upper and lower bound, thus no adhesion occurs.

Korea Pathfinder Lunar Orbiter (KPLO), also known as Danuri, was successfully launched on 4 Aug. from Cape Canaveral Space Force Station using a Space-X Falcon-9 rocket. Flight dynamics (FD) operational readiness was one of the critical parts to be checked before the flight. To demonstrate FD software’s readiness and enhance the operator’s contingency response capabilities, KPLO FD specialists planned, organized, and conducted four simulations and two rehearsals before the KPLO launch. For the efficiency and integrity of FD simulation and rehearsal, different sets of blind test data were prepared, including the simulated tracking measurements that incorporated dynamical model errors, maneuver execution errors, and other errors associated with a tracking system. This paper presents the simulation and rehearsal results with lessons learned for the KPLO FD operational readiness checkout. As a result, every functionality of FD operation systems is firmly secured based on the operation procedure with an enhancement of contingency operational response capability. After conducting several simulations and rehearsals, KPLO FD specialists were much more confident in the flight teams’ ability to overcome the challenges in a realistic flight and FD software’s reliability in flying the KPLO. Moreover, the results of this work will provide numerous insights to the FD experts willing to prepare deep space flight operations.