Flight Dynamics Model Research Articles

CFD Simulation of a Finned Smart Bullet with Microactuator

The geometric configuration of the micro-actuators is known to have a large impact on control authority. This study describes a finned smart bullet (FSB) to support the requirements of various missions. The flow control system consists of micro-actuator located at the rear body of FSB, alter the flow field in the rear region of FSB to create asymmetric pressure distributions and thus produce aerodynamic control forces and moments. This work seeks to derive an optimal mass centre (MC) for FSB and explore the FSB configuration with rear micro-actuator. Computational fluid dynamics simulations are performed to study the MC position for FSB stable flight and review the aerodynamic characteristics FSB configuration with actuator. For control trajectory of the FSB, the further work will explore the FSB best configuration to finish the projectile dynamic model for standard 6DOF projectile flight dynamic model in order achieve flight control.

Research on the Flight Control Strategy of a New Concept Fuel-electric Hybrid Multi-rotor UAV

In order to solve the shortcomings of conventional electric multi-rotor UAVs, such as low load, short range and endurance, a new concept of fuel-electric hybrid multi-rotor UAV, in which the lift and attitude control are separated, is presented in this paper. Firstly, to comprehend the control characteristics of this UAV and also to design its flight control law, the flight dynamics model of this kind of UAV is established. Subsequently, considering the control redundancy in the height channel of the main rotors and the auxiliary rotors of UAV, a power switch control strategy based on vertical acceleration is proposed. Further, in view of the control characteristics of the aircraft, according to the backstepping -PID control method, the attitude and speed controller are designed. The numerical simulation and the flight test of the prototype are carried out respectively in order to verify the control strategy. Both theoretical simulation and flight test results show that the designed UAV controller can satisfy the desired control effect.

Achieving bioinspired flapping wing hovering flight solutions on Mars via wing scaling

Achieving atmospheric flight on Mars is challenging due to the low density of the Martian atmosphere. Aerodynamic forces are proportional to the atmospheric density, which limits the use of conventional aircraft designs on Mars. Here, we show using numerical simulations that a flapping wing robot can fly on Mars via bioinspired dynamic scaling. Trimmed, hovering flight is possible in a simulated Martian environment when dynamic similarity with insects on earth is achieved by preserving the relevant dimensionless parameters while scaling up the wings three to four times its normal size. The analysis is performed using a well-validated 2D Navier–Stokes equation solver, coupled to a 3D flight dynamics model to simulate free flight. The majority of power required is due to the inertia of the wing because of the ultra-low density. The inertial flap power can be substantially reduced through the use of a torsional spring. The minimum total power consumption is 188 W kg−1 when the torsional spring is driven at its natural frequency.

Augmented flight dynamics model for pilot workload evaluation in tilt-rotor aircraft optimal landing procedure after one engine failure

An augmented flight dynamics model is developed to extend the existing flight dynamics model of tilt-rotor aircraft for optimal landing procedure analysis in the event of one engine failure. Compared with the existing flight dynamics model, the augmented model involves with more pilot control information in cockpit and is validated against the flight test data. Based on the augmented flight dynamics model, the optimal landing procedure of XV-15 tilt-rotor aircraft after one engine failure is formulated into a Nonlinear Optimal Control Problem (NOCP), solved by collocation and numerical optimization method. The time histories of pilot controls in cockpit during the optimal landing procedure are obtained for the evaluation of pilot workload. An evaluation method which can synthetically quantify the pilot workload in time and frequency domains is proposed with metrics of aggressiveness and cutoff frequencies of pilot controls. The scale of the pilot workload is compared with those of the shipboard landing procedures, bob-up/bob-down and dash/quick-stop maneuvers of UH-60 helicopter. The results show that the aggressiveness of pilot collective and longitudinal controls for the tilt-rotor aircraft optimal landing procedure after one engine failure are higher than those for UH-60 helicopter shipboard landing procedures up to the condition of sea state 4, while the pilot cutoff frequency of collective control is lower than that of the bob-up/bob-down maneuver but the pilot cutoff frequency of longitudinal control is higher than that of the dash/quick-stop maneuver. The evaluated pilot workload level is between Cooper–Harper HQR Level 2 and Level 3.

A Three-axis PD Control Model for Bumblebee Hovering Stabilization

Flight stabilization in insects is normally achieved through a closed-loop system integrating the internal dynamics and feedback control. Recent studies have reported that flight instability may exist in most flying insects but how insects achieve the flight stabilization still remains poorly understood. Here we propose a control model specified for bumblebee hovering stabilization by applying a three-axis PD (proportional-derivative)-controller to a free-flying bumblebee computational model with six Degrees of Freedom (DoFs). Morphological and kinematic models of a realistic bumblebee in hovering are built up based on measurements whereas a versatile bio-inspired dynamic flight simulator is employed in simulations. A simplified flight dynamic model is further developed as a fast model for control parameter tuning. Our results demonstrate that the stabilizing control model is capable of achieving the hovering stabilization with small perturbations in terms of 6-DoF, implying that the simplified linear algorithms can still work reasonably for bumblebee hovering. A further sensitivity analysis of the control parameters reveals that yaw control via manipulating pitch angle of the wing is mostly sensitive, implicating that bumblebee may utilize alternative yaw control strategies.

Flight Load Assessment for Light Aircraft Landing Trajectories in Windy Atmosphere and Near Wind Farms

This work focuses on the wake encounter problem occurring when a light, or very light, aircraft flies through or nearby a wind turbine wake. The dependency of the aircraft normal load factor on the distance from the turbine rotor in various flight and environmental conditions is quantified. For this research, a framework of software applications has been developed for generating and controlling a population of flight simulation scenarios in presence of assigned wind and turbulence fields. The JSBSim flight dynamics model makes use of several autopilot systems for simulating a realistic pilot behavior during navigation. The wind distribution, calculated with OpenFOAM, is a separate input for the dynamic model and is considered frozen during each flight simulation. The aircraft normal load factor during wake encounters is monitored at different distances from the rotor, aircraft speeds, rates of descent and crossing angles. Based on these figures, some preliminary guidelines and recommendations on safe encounter distances are provided for general aviation aircraft, with considerations on pilot comfort and flight safety. These are needed, for instance, when an accident risk assessment study is required for flight in proximity of aeolic parks. A link to the GitHub code repository is provided.

Aerodynamics and Flight Dynamics Simulation of Basic Finner Supersonic Flight in Aeroballistic Experiment

This paper presents some results of aerodynamics and flight dynamics modeling of Basic Finner supersonic flight in an aeroballistic experiment [1]. Three dimensional surface based on the experiment description was formed. Unstructured tetrahedral volume mesh is formed after it. Aerodynamics computations were done by software package for CFD numerical computation of free form high-speed vehicle ug3D developed in the Institute for Problems in Mechanics. Obtained from CFD calculations aerodynamic coefficients became initial data for unguided Basic Finner flight dynamics calculations. Flight dynamics computations were done by developed solver MODIN for typical high speed vehicles flight dynamics calculations [2]. Six degrees of freedom motion without considering Earth rotation and form are adopted. Results of experimental and obtained numerical aerodynamic coefficients and flight parameters data comparison are also presented.

Flight dynamic coupling analysis of a bio-inspired elastic-wing aircraft

ABSTRACTIt is a challenge for small, fixed-wing aerial vehicles to maintain flight stability under gusts. Inspired by the geometric features and the structural dynamic characteristics of the gliding bird wing, an elastic wing with similar characteristics was designed and optimised for use as part of unmanned aerial vehicle. A flight dynamic model, which includes the coupling of the longitudinal flight modes and the aeroelastic modes of the flexible wing, was built to analyse the mechanisms of specific coupling for the structural characteristics of the wing design, and how these specific couplings affect flight dynamics. The results showed that the bio-inspired elastic wing effectively allows alleviation of the gust response of the prototype through coupling effects of the short period and the first aeroelastic mode, even with a considerable frequency gap. These effects become more significant when the airspeed becomes larger. The conclusions of this research can facilitate further development of bird-sized unmanned aerial vehicles to extend their applications and make these vehicles more adaptive for flight in complex atmospheric environments.

A Flight Control System for Small Unmanned Aerial Vehicle

The program adaptation of the controller for the flight control system (FCS) of an unmanned aerial vehicle (UAV) is considered. Linearized flight dynamic models depend mainly on the true airspeed of the UAV, which is measured by the onboard air data system. This enables its use for program adaptation of the FCS over the full range of altitudes and velocities, which define the flight operating range. FCS with program adaptation, based on static feedback (SF), is selected. The SF parameters for every sub-range of the true airspeed are determined using the linear matrix inequality approach in the case of discrete systems for synthesis of a suboptimal robust H∞-controller. The use of the Lagrange interpolation between true airspeed sub-ranges provides continuous adaptation. The efficiency of the proposed approach is shown against an example of the heading stabilization system.

Roll performance assessment of a light aircraft: Flight simulations and flight tests

A methodology is presented for the roll performance assessment of a light aircraft. The study is based both on flight simulations and flight tests, focusing on the accurate determination of the lateral control power. The chosen test airplane was a Tecnam P92, a two-seat ultralight aircraft which was specifically equipped with a lightweight and accurate instrumentation for the planned set of flight data measurements. A matrix of flight experiments for the test campaign was established with the support of 6-degree-of-freedom simulations, implementing a carefully constructed baseline dynamic model of the aircraft. The article discusses first the general problem of a reliable evaluation of aircraft roll performance indicators, i.e. the estimation of the aerodynamic derivatives that mainly influence the airplane ability to roll. Next, the results of extensive flight test activities are presented. The analysis of several roll maneuvers performed at different flight speeds and with different aileron maximum deflections showed interesting rolling characteristics for this non-aerobatic aircraft. One notable finding was a clear nonlinear dependency of the aileron efficiency index on aileron deflection amplitude. A control power derivative extracted from flight data in the form of lookup table was used to correct the baseline flight dynamics model. Flight simulations outputs based on the updated model showed a satisfactory agreement with experimental time histories. According to this, the present effort proposes a new method to estimate the aileron control derivative in whole the flight envelope for light aircraft.

Biplane-Quadrotor Tail-Sitter UAV: Flight Dynamics and Control

This paper presents the development of 6-degrees-of-freedom (6-DOF) flight dynamics model and nonlinear control design for a novel hybrid biplane-quadrotor unmanned aerial vehicle capable of efficient operation in both vertical takeoff and landing (VTOL) mode and forward flight mode. The development of vehicle dynamics model considers a complete description of flight dynamics, wing aerodynamics, and prop wash modeling. Rotor and wing aerodynamics during hover, transition, and forward flight are the key components of the vehicle dynamics equations and are systematically validated with experimental data available in the literature. The rotor aerodynamic loads calculated using blade element theory and uniform inflow–based model is found to offer adequate fidelity during both axial and edgewise flights of the rotor. The rigid body rotation of 90° about pitch axis during transition from hover to forward flight and back necessitates the use of quaternions for representation of the attitude of the vehicle to avoid singularity associated with Euler angles for such maneuvers. A comprehensive nonlinear control design using dynamic inversion approach is developed for the whole flight regime that includes climb/descent, hover, transition, and forward flight. The performance of the proposed control design is demonstrated through numerical simulations on the 6-DOF model of the vehicle.

Flight Dynamics Modeling and Control of a Novel Catapult Launched Tandem-Wing Micro Aerial Vehicle With Variable Sweep

In this paper, the variable sweep is used as a replacement for the conventional control surfaces for flight control on a tandem-wing micro aerial vehicle (MAV). This configuration allows the MAV to be folded into and launched from a tubular catapult while also retaining an adequately high effectiveness and a simple structure. As new control inputs, the four sweep angles of the MAV are planned as symmetric and asymmetric morphing to eliminate the reactive forces between the airfoils and fuselage during morphing. The aerodynamic characteristics of the MAV are presented through numerical simulations to study the effects of variable sweep morphing. An accurate nonlinear multibody dynamic model of the variable sweep MAV is established using the Kane method. Next, open-loop dynamic responses of sweep morphing based on the nonlinear dynamic model are analyzed and compared with the responses caused by the elevons control mode to examine the control effectiveness. Moreover, a flight control law based on sweep control is designed and verified using a height and yaw tracking simulation with different actuator response rates. The results show that the sweep control mode produces weaker coupling between the longitudinal and lateral dynamics than the elevons control mode and that the innovative approach for flight control is feasible and effective.

Active disturbance rejected predictive functional control for space vehicles with RCS

Reaction control system (RCS) is a powerful and efficient actuator for space vehicles attitude control, which is typically characterized as a pulsed unilateral effector only with two states (off/on). Along with inevitable internal uncertainties and external disturbances in practice, this inherent nonlinear character always hinders space vehicles autopilot from pursuing precise tracking performance. Compared to most of pre-existing methodologies that passively suppress the uncertainties and disturbances, a design based on predictive functional control (PFC) and generalized extended state observer (GESO) is firstly proposed for three-axis RCS control system to actively reject that with no requirement for additional fuel consumption. To obtain a high fidelity predictive model on which the performance of PFC greatly depends, the nonlinear coupling multiple-input multiple-output (MIMO) flight dynamics model is parameterized as a state-dependent coefficient form. And based on that, a MIMO PFC algorithm in state space domain for a plant of arbitrary orders is deduced in this paper. The internal uncertainties and external disturbances are lumped as a total disturbance, which is estimated and cancelled timely to further enhance the robustness. The continuous control command synthesised by above controller-rejector tandem is finally modulated by pulse width pulse frequency modulator (PWPF) to on-off signals to meet RCS requirement. The robustness and feasibility of the proposed design are validated by a series of performance comparison simulations with some prominent methods in the presence of significant perturbations and disturbances, as well as measurement noise.

Flight Dynamics Models and Control of Paragliders and Hang Gliders

Flight Dynamics Models and Control of Paragliders and Hang Gliders

Research on Trajectory Planning and Tracking of Hexa-copter

In this paper, the flight control problem of hexa-copter is studied in detail from threedimensional trajectory planning to tracking. Then the cubic spline interpolation method is used to generate the trajectory by using these time marked waypoints. The flight trajectory curve produced by this method is smooth, twice differentiable, and it is easy to control implementation. The flight dynamics model of the UAV has the characteristics of multi-input multi-output, strong coupling, under-actuation, severe nonlinearity and external environmental disturbance. In order to improve the accuracy of flight trajectory and the stability of attitude control, a multi-loop sliding mode variable structure control method is proposed to achieve the hexa-copter flight trajectory tracking. The simulation results show that this method can track the predetermined flight trajectory and keep the attitude stability of the UAV normally.

Helicopter flight simulation trim in the coordinated turn with the hybrid genetic algorithm

Helicopter trim models are multivariate nonlinear equations and it is difficult to determine these initial trim points comparable to flight conditions. To solve this question, a hybrid genetic algorithm is presented in this paper, that combines the quick convergence ability of the quasi-Newton method and the advantages of genetic algorithm, such as global convergence. The trim control vector and the constraint conditions were established in the coordinated-turn based on the helicopter flight dynamic model. The coordinated turn flight of a UH-60 A helicopter was taken as an example to simulate on the experimental platform. Comparisons were made between the trim results and flight test data and there is a good agreement among them, and the efficiency of the algorithm presented is verified. It is a general method that can be applied to trim the helicopter of different flight conditions.

Towards real-time pilot-in-the-loop CFD simulations of helicopter/ship dynamic interface

This study presents the development of computationally efficient coupling of Navier–Stokes Computational Fluid Dynamics (CFD) with a helicopter flight dynamics model with the ultimate goal of real-time simulation of airwake effects in the helicopter/ship Dynamic Interface (DI). The flight dynamics model is free to move within a computational domain, where the main rotor forces are converted to source terms in the momentum equations of the CFD solution using an actuator disk model. Simultaneously, the CFD solver calculates induced velocities that are fed back to the simulation and affect the aerodynamic loads in the flight dynamics. The CFD solver models the inflow, ground effect and interactional aerodynamics in the flight dynamics simulation, and these calculations can be coupled with the solution of the external flow (e.g., ship airwake effects). The simulation framework for fully-coupled pilot-in-the-loop (PIL) flight dynamics/CFD is demonstrated for a simplified shedding wake. Initial tests were performed with 0.38 million structured grid cells running on 352 processors and showed near-real-time performance. Improvements to the coupling interface are described that allow the simulation run at near-real-time execution speeds on currently available computing platforms. Improvements in computing hardware are expected to allow real-time simulations.

Continuous-Time State-Space Unsteady Aerodynamic Modeling for Efficient Loads Analysis

The present paper proposes a continuous-time state-space formulation of the unsteady vortex lattice method, which is derived through a discretization of the governing advection equation for transport of vorticity in the wake. A continuous-time system is obtained by only discretizing the advection equation in space, while retaining the derivative with respect to time. The discretization in space is based on the discontinuous Galerkin method. The present method can be applied to any arbitrary nonuniform wake discretization and can be extended to higher-order panel methods or a nonflat wake shape. The method is extended to compressible flows by applying the Prandtl–Glauert transformation. The time-dependent terms in the small disturbance potential equation are neglected. Thus, incompressible flow solution procedures are applied with minimum modifications to unsteady compressible problems. The benefits are demonstrated by applying the model to the gust analysis of a general aircraft wing, varying the time step, and introducing a nonuniform wake discretization, resulting in a reduced model size for a given accuracy. The resulting continuous-time state-space model can be used for efficient loads analysis of general aircraft wings including the effects of compressibility and allows for easy integration with structural or flight dynamic models for efficient aero(servo)elastic analyses.

Flight dynamic modeling and control for a telescopic wing morphing aircraft via asymmetric wing morphing

In this paper, a sliding mode flight controller is formulated in order to enhance the lateral maneuverability for a tailless telescopic wing morphing aircraft by using additional asymmetric wing telescoping. Based on the nonlinear time-varying equations of motion and the approximate aerodynamic model obtained by wind tunnel tests, a nonlinear time-varying model with coupling between aerodynamic parameters and control inputs is established and the open-loop dynamic response characteristics are analyzed. According to the difficulty of precise aerodynamic modeling during both sides of wing telescoping, the sliding mode control approach is used to design the control law to track maneuver reference command, which has a low requirement for modeling precision and is suitable for the nonlinear time-varying system. In addition, the control inputs are allocated to asymmetric wing telescoping and aerodynamic control surfaces based on the minimum energy method. The closed-loop simulation results of a roll agility maneuvering called “T90 maneuvering” are presented. The results show that a larger roll angular velocity is produced and the control surfaces in the maneuvering flight can maintain a large control margin, which means the maneuverability is significantly improved and the control burden of the aerodynamic control surfaces has been reduced compared with the aircraft without use of telescopic wings; moreover, the robustness of the sliding mode controller is verified by simulation of aerodynamic perturbations.

Flight Dynamics Modeling of Dual-Spin Guided Projectiles

This paper presents a complete nonlinear parameter-dependent mathematical model, as well as a procedure for computing the quasi -linear parameter-varying ( q -LPV) model of a class of spin-stabilized canard-controlled guided projectiles. The proposed projectile concept possesses a so-called dual-spin configuration: the forward section contains the necessary control and guidance software and hardware, whereas the aft roll-decoupled and rapidly spinning section contains the payload. Wind-axes instead of body-axes variables, as is the case in the existing literature, are preferred for the modeling of the highly coupled pitch/yaw airframe nonlinear dynamics, since they are better suited to control synthesis. A q -LPV model, approximating these nonlinear dynamics around the system equilibrium manifold and capturing their dependence on diverse varying flight parameters, is analytically obtained. In addition, a detailed stability analysis specific to this kind of weapons is performed throughout their large flight envelope using the aforementioned q -LPV model. Furthermore, a study is conducted in order to quantify the influence of reducing the dimension of the flight parameter vector on the exactness of the q -LPV model. Finally, the critical influence on the pitch/yaw-dynamics of the nose-embedded sensor position, and of uncertainty on the various static and dynamic aerodynamic coefficients as well as the aerodynamic angles, is shown.