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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]

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.

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]

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]

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.

Estimating animal populations is essential for conservation. Censusing large congregations is especially important since these are priorities for protection, but efficiently counting hundreds of thousands of moving animals remains a challenge. We developed a deep learning‐based system using consumer cameras that not only counts but also records behavioral information for large numbers of flying animals in a range of lighting conditions including near darkness. We built a robust training set without human labeling by leveraging data augmentation and background subtraction. We demonstrate this approach by estimating the size of a straw‐colored fruit bat (Eidolon helvum) colony in Kasanka National Park, Zambia with cameras encircling the colony to record evening emergence. Detection of bats was robust to deteriorating lighting conditions and changing backgrounds. Combined over five days, our population estimates ranged between 750,000 and 976,000 bats with a mean of 857,233. In addition to counts, we extracted wingbeat frequency, flight altitude, and local group polarity for 639,414 individuals. This open access method is an inexpensive but powerful approach that, in addition to radial emergences from central locations, can also be applied to unidirectional movements of flying groups, such as migratory streams of birds. [ABSTRACT FROM AUTHOR]

Compared with the forced oscillation test in a conventional wind tunnel test, the wind tunnel virtual flight test can achieve partially constrained dynamic motion of the model in the wind tunnel by manipulating the control surface deflection. In this paper, a new four-degree-of-freedom (4-DOF) wind tunnel virtual flight test device is established. Compared with the 3-DOF wind tunnel virtual flight test with rotation around the center of mass, the model adds vertical motion, enabling it to simulate free flight more realistically. In this paper, the nonlinear flight dynamics model of 4-DOF wind tunnel virtual flight test is established, the differences in dynamics characteristics between it and free flight are analyzed, and the identification model of aerodynamic derivatives based on 4-DOF wind tunnel virtual flight test is derived. To ensure the safety of the test, the longitudinal altitude control law as well as the sideslip and roll angle control laws in the lateral direction are designed. Finally, a set of aerodynamic parameter identification methods based on a 4-DOF wind tunnel virtual flight test is formed. The output error method is used to identify the aerodynamic derivatives of the test model, and the accuracy of the identification results is higher than that of the 3-DOF wind tunnel virtual flight test. [ABSTRACT FROM AUTHOR]

In the three-degree-of-freedom virtual flight test, the scaled model has no translational motion, which will result in the model's dynamics being different from that of free flight, and this difference will directly affect its effectiveness in simulating free flight and the design of the flight control law. In this paper, the longitudinal and lateral-directional dynamic characteristics of virtual flight are studied from the three aspects of modal characteristics, disturbance response characteristics, and control response characteristics. The dynamic characteristic difference between the virtual flight test and free flight is analyzed. In addition, the scaling ratio of the model will also lead to differences in the flight control law design methods of virtual flight and full-size aircraft. The control law design method of the virtual flight test is established by studying the similarity relation of control law parameters between full-size aircraft and the scaled model, and the correctness of similar design of closed-loop control law parameters of the scaled model is verified by a virtual flight test. Finally, the attitude control law parameters obtained from the virtual flight test are applied to the six-degree-of-freedom flight simulation of a full-size aircraft, and the simulation results show that the three-axis attitude control law designed based on the virtual flight test also has good performance in the six-degree-of-freedom flight of the full-size aircraft, and the response of attitude angle and angular rate are approximately the same for both. Generally, the steady aerodynamic data of the aircraft are obtained through the traditional static wind tunnel test to construct the aerodynamic model of the aircraft. The aerodynamic model established by this method may not be able to accurately simulate the actual dynamic flight. However, the actual test flight of full-size aircraft has many disadvantages, such as high cost and high risk. To simulate real flight on the ground, the method of placing the scaled model in a wind tunnel for dynamic flight is proposed, which is called wind tunnel–based virtual flight testing (WTBVFT). The scaled model is connected with the strut through a three-degree-of-freedom rotation mechanism and installed in the wind tunnel test section, so that the model has three degrees of freedom of angular motion and no translational motion. This paper analyzes the differences in dynamics between virtual flight and free flight under open-loop and closed-loop control. The results show that the virtual flight test can simulate the angular motion of free flight well and can be used to design, verify, and simulate the attitude control law of full-size aircraft well. [ABSTRACT FROM AUTHOR]

The development of novel aircraft concepts and propulsion technologies requires up-to-date physics-based methods and tools for conceptual aircraft design. In this context, a simulation model for the take-off manoeuvre is proposed in this article, to be employed in the conceptual design phase for aircraft whether of traditional or innovative configuration. The model is capable of evaluating the longitudinal dynamics, both translational and rotational, of the aircraft considered as a rigid body, and influenced by the aerodynamic effects introduced by the presence of the ground. The ground effect, indeed, induces variations in the aerodynamic forces depending on the distance and the attitude of the lifting surfaces from the ground, which may significantly influence the aeromechanical characteristics of the aircraft during the evolution of the take-off manoeuvre. The simulation model is based on the numerical solution of the equations of the dynamics of the rigid aircraft in the longitudinal plane and integrates a vortex lattice aerodynamic solver to evaluate the aerodynamic and aeromechanical characteristics of the aircraft considering the ground effect in each time-step. The proposed approach is configuration independent, as it can model the geometry, evaluate the aerodynamics, and simulate the dynamics of aircraft with any lifting architecture; furthermore, the simulation model is fast and flexible, making it effective for the conceptual phase of aircraft design. The paper proposes the description of the take-off manoeuvre of two aircraft with different airframes: one with a conventional tube-and-wing architecture and one with a box-wing lifting system. The results proposed highlight the potential of the simulation model to detect aeromechanic and dynamic differences during the development of the manoeuvre for different aircraft configurations, and to assess the significance of considering ground effect aerodynamics. [ABSTRACT FROM AUTHOR]

In the early stages of aircraft design, a scaled model of the aircraft is installed in a wind tunnel for dynamic semi-free flight to approximate real flight, and the test data are then used to identify the aerodynamic parameters. However, the absence of the translational motion of the test model makes its flight dynamics different from those in free flight, and the effect of this difference on parameter identification needs to be investigated. To solve this problem, a 3-DOF wind tunnel virtual flight test device is built to fix the test model on a rotating mechanism, and the model is free to rotate in three axes through the deflection of the control surfaces. The flight dynamics equations for the wind tunnel virtual flight test are established and expressed as a decoupled form of the free flight force and the influence of the test support frame force on the model's motions through linearization. The differences between wind tunnel virtual flight and free flight are analysed to develop a model for the identification of aerodynamic parameters. The selection of the lateral-directional excitation signal and the design method of its parameters are established based on the requirements for the identifiability of the aerodynamic derivatives, and a step-by-step method for the identification of aerodynamic force and moment derivatives is established. The aerodynamic parameter identification results of a blended wing body aircraft show that the identification method proposed in this paper can obtain results with high accuracy, and the response of the modified motion model is consistent with that of the free flight motion model. [ABSTRACT FROM AUTHOR]

The concept of thrust-based control without the employment of thrust-vectoring (TVC), already introduced in a previous work by the authors, is further developed in conjunction with an appropriate control suite, tasked with both artificial stabilization and beam-tracking navigation functions. In the paper, the fully non-linear mathematical model employed for testing the controllers in a virtual environment is outlined. Then a comparative approach is adopted in the analysis, where a standard tail-back airship with deflectable aerodynamic surfaces is employed as a baseline, and the performance of a four-thrusters layout with a thrust-based control and no TVC is assessed with respect to it. Featured test cases in forward flight include short climbs, abrupt turns, and multi-checkpoint navigation. The research supports the feasibility and adequate performance of the proposed thrust-based airship layout and control, and presents a critical analysis of the pros and cons with respect to the considered baseline airship configuration featuring standard aerodynamic control. [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.

Tilt-rotor unmanned aerial vehicles combine the advantages of multirotor and fixed-wing aircraft, offering features like rapid takeoff and landing, extended endurance, and wide flight conditions. This article provides a summary of the design, modeling, and control of a composite tilt-rotor. During modeling process, aerodynamic modeling was performed on the tilting and non-tilting parts based on the subcomponent modeling method, and CFD simulation analysis was conducted on the entire unmanned aerial vehicle to obtain its accurate aerodynamic characteristics. In the process of modeling the motor propeller, the reduction of motor thrust and torque due to forward flow and tilt angle velocity is thoroughly examined, which is usually ignored in most tilt UAV propeller models. In the controller design, this paper proposes a fusion ADRC control strategy suitable for vertical takeoff and landing of this type of tiltrotor. The control system framework is built using Simulink, and the control algorithm’s efficiency has been verified through simulation testing. Through the proposed control scheme, it is possible for the composite tiltrotor unmanned aerial vehicle to smoothly transition between multirotor and fixed-wing flight modes.

Abstract Estimating animal populations is essential for conservation. Censusing large congregations is especially important since these are priorities for protection, but efficiently counting hundreds of thousands of moving animals remains a challenge. We developed a deep learning‐based system using consumer cameras that not only counts but also records behavioral information for large numbers of flying animals in a range of lighting conditions including near darkness. We built a robust training set without human labeling by leveraging data augmentation and background subtraction. We demonstrate this approach by estimating the size of a straw‐colored fruit bat (Eidolon helvum) colony in Kasanka National Park, Zambia with cameras encircling the colony to record evening emergence. Detection of bats was robust to deteriorating lighting conditions and changing backgrounds. Combined over five days, our population estimates ranged between 750,000 and 976,000 bats with a mean of 857,233. In addition to counts, we extracted wingbeat frequency, flight altitude, and local group polarity for 639,414 individuals. This open access method is an inexpensive but powerful approach that, in addition to radial emergences from central locations, can also be applied to unidirectional movements of flying groups, such as migratory streams of birds.

With the progress of high-performance computing, computationally expensive high-fidelity methods can be applied early in the design process of an aircraft. This enables Computational Fluid Dynamics (CFD) for the assesment of flight performance, handling qualities, flight loads due to manoeuver or gusts, and stability such as flutter. Those aeroealstic analyses require coupling of disciplines to adequately address the interplay of flexibility, aerodynamic forces and inertia forces. Previous work at the German Aerospace Center have demonstrated the proof of concept for those multidisciplinary analyses. Now, UltraFLoads is a simulation suite in the FlowSimulator ecosystem to offer standardized, multidisciplinary scenarios. UltraFLoads can be used as a tool, as a library and as a framework. This work describes the architecture, the integration layers, which scenarios are available and which equations are used. Three different applications are presented. The first case involves a steady fluid–structure-coupled simulation with geometrically nonlinear deformation. The second case compares the deformation of a simulated elastic, free-flying aircraft with actual flight test data. The last application demonstrates the flight dynamic capabilities with a bank to bank maneuver. [ABSTRACT FROM AUTHOR]

The stick-free flight stability is an old-fashioned and non-progressive issue; nevertheless, it is still existent and of significant importance to the design of aircraft whose control system is reversible. The existence of the problem necessitates a deep assessment of stick-free flight stability throughout the aircraft design. Up to now, this problem has been addressed using either analytical approaches, which are only related to the static stability evaluation, or performing flight tests. In this study, the problem is handled in its entirety, from static and dynamic flight stability assessment to design criteria with a comprehensive perspective. Moreover, it is also exhibited that, contrary to what has been generally proposed in the literature, the limitation of the problem of stick-free flight stability through static stability assessment is far from being the main challenge. As a brief scope, the derivation of the control surface dynamics, a stick-free trim algorithm, and assessment rationale of the stick-free static and dynamic flight stability using a simulation approach are proposed. As a consequence, the aim is to set a broad understanding for designers related to this phenomenon and add adjunct design criteria in the design optimization process by approaching it from a modeling, simulation, and flight test perspective. [ABSTRACT FROM AUTHOR]

The design complexities of trending UAVs necessitates formulation of C ontrol L aws that are both robust and model-free besides being self-capable of handling the evolving dynamic environments. In this research, a unique intelligent control architecture is presented which aims at maximizing the glide range of an experimental UAV having unconventional controls. To handle control complexities, while keeping them computationally acceptable, a distinct RL technique namely Modified Model Free Dynamic Programming (MMDP) is proposed. The methodology is novel as RL based Dynamic Programming algorithm has been specifically modified to configure the problem in continuous state and control space domains without knowledge of the underline UAV model dynamics. Major challenge during the research was the development of a suitable reward function which helps in achieving the desired objective of maximising the glide performance. The efficacy of the results and performance characteristics, demonstrated the ability of the presented algorithm to dynamically adapt to the changing environment, thereby making it suitable for UAV applications. Non-linear simulations performed under different environmental and varying initial conditions demonstrated the effectiveness of the proposed methodology over the conventional classical approaches. [ABSTRACT FROM AUTHOR]

The accurate modelling and simulation of vehicle dynamics is a fundamental prerequisite for the design and experimental flight testing of aerospace vehicles. In the case of high-altitude supersonic sounding rockets, it is critically important to produce realistic trajectory predictions in a representative range of operational and environmental conditions as well as to produce reliable probability distributions of terminal locations. This article proposes a methodology to develop high-fidelity flight dynamics models that accurately capture aeroelastic, turbulence, atmospheric and other effects relevant to sounding rockets. The significance of establishing a high-fidelity model and of addressing such a problem in the context of developing a digital twin are discussed upfront, together with the key tools utilised in the analysis. In addition to state-of-the-art computational methods to determine the aerodynamic forces, moments and mass changes in various flight regimes (including parachute release), a detailed methodology for incorporating the dynamic aeroelastic response of the rocket is presented. The validity of the proposed method is demonstrated through a simulation case study, which utilises data from an existing rocket prototype. Results corroborate the correct implementation of the proposed algorithms and provide foundations for future research on virtual sensing and digital twin for autonomous navigation and guidance. [ABSTRACT FROM AUTHOR]

This work presents a systematic approach to analyzing the aerodynamic characteristics of tandem rotor forward autorotation considering rotor-to-rotor interference. The single-rotor computational model trimmed from a generic helicopter flight dynamics analysis program was used as the baseline model. The effectiveness of the baseline model is demonstrated by a comparison with data from wind tunnel tests performed in this work. The rotor disk angle of attack and driven moment distribution obtained by the modified model indicate the fact that the rotor acceleration is primarily caused by the higher angle of attack region of the disk. This is of great significance in the rotor blade design, in terms of the drag-to-lift ratio characteristics of the airfoil under different angle-of-attack ranges. The influence of wind speed, rotor shaft angle, and collective pitch on the steady-state rotor speed was then studied. The results show a nonlinear nature of the variation of steady rotor speed with collective pitch, which can cause a thrust control reverse problem during flight operations. To reveal the flow field details of rotor-to-rotor interference, the flow field Navier–Stokes equations of tandem rotor autorotation were solved. Computational results of both rotors' inflow velocities were considered when deriving the empirical model of interference. The refined interference model was compared to the wind tunnel test data of the tandem rotor autorotation and showed good performance. This synthetical methodology, which combines mechanism analysis with CFD-aided refinement and experiment verification, achieves a balance between computational costs and accuracy and thus can be readily applied to engineering practices. [ABSTRACT FROM AUTHOR]

A good mathematical model accurately represents the behavior of a system. Although aircraft has complex dynamics, it is feasible to develop a robust model that could be used to design and analyze some of its essential components such as flight control system and simulator. To solve security challenges in the hinterlands and maritime domains, a hybrid Unmanned Aerial Vehicle (UAV), nicknamed ‘AH-25’, was designed by the Air Force Institute of Technology, Nigeria. In this work, the development of flight dynamics models of the vehicle is presented. The AH-25 UAV has the capability of operating in both fixed-wing and Vertical Take-Off and Landing (VTOL) modes for effective operations on land and maritime spaces. Therefore, models that capture the dynamics of the two distinct modes were obtained. The fixed-wing nonlinear model which was developed from first principle using Newton-Euler method was decoupled, trimmed and linearized to set it in a good shape for critical aircraft systems designs. On the other hand, the multicopter model developed by Beihang flight control group was adapted for modelling the VTOL mode. Simulation results showed that the models’ responses are a replica of the actual aircraft operations. Also, responses of the model to perturbations indicated that it is open-loop stable. Furthermore, the mean performance metrics of the fixed-wing model’s open-loop time response to various inputs were evaluated. The model was found to have a rise time of 2 s, 1.33% steady state error and a settling time of 40 s.

This paper presents a dynamical analysis of a tethered aerostat using a multibody simulation tool called SimscapeMultibody™. Tethered aerostats are lighter than air vehicles that float in the air and are connected to the ground through the tether. Atmospheric winds are a source of disturbance to the aerostat. The wind produces drag force that causes the aerostat to displace from its mean position. This phenomenon is called blow-by. Solving the mathematical model of the motion dynamics with time step is the conventional method used for simulating the flight dynamics of the aerostat. This paper discusses the application of Simscape Multibody™ for the three-degrees-of-freedom flight dynamic simulation of spherical tethered aerostat. Simscape Multibody™, an extension of MATLAB®/Simulink, is a multibody modeling tool in which physical bodies and their interactions have to be defined to simulate the dynamics. The mathematical model of the defined system is automatically generated by the SimscapeMultibody™ solver, and the system dynamics is simulated. The simulation of dynamics using multibody modeling tools provides high fidelity simulation for the tethered aerostats.

In this paper, a quadrotor dynamic model's energy efficiency was investigated. A method for the design of the dynamic model which assures energy consumption estimation was presented. This model was developed to analyze the energy efficiency of the quadrotor during each maneuver. A medium-class quadrotor (4.689 kg) was used as a test platform. Thrust force correction factors obtained with FLIGHTLAB software were used to predict object behavior in forward flight. Model validation and long-duration flight tests in outdoor windy conditions are also presented. Monte-Carlo simulation was used to study the influence of uncertainties in model parameters on the simulation reliability. The developed model might be used for practical purposes (for example, energy-efficient coverage path planning). [ABSTRACT FROM AUTHOR]