This paper presents the concept of a lifting-wing quadcopter unmanned aerial vehicle (UAV), a vertical take-off and landing vehicle (VTOL) with a rear wing, a canard at its front and four propellers. The aerodynamic surfaces are designed so that their mounting angle can be adjusted and fixed before flight, so its performance in transition flight can be studied for a combination of wing and canard mounting angles. A dynamic model using rigid-body equations of motion is presented, which is used to compute the transition flight trajectory from hover to cruise in horizontal flight. The trim conditions were computed for a range of fixed wing and canard mounting angles to study the effects of these variables on transition trajectory parameters such as required power, body pitch angle and propeller rotation speeds as a function of flight speed. Furthermore, a transition flight control algorithm is presented, which has a cascaded PID controller and a reference scheduler to switch between the proper reference states, controls and control allocation matrix. Finally, the transition control algorithm of the conceptual UAV is numerically simulated. Results show that this configuration can perform a fast and smooth transition from hover to cruise flight using the proposed flight control algorithm, substantially reducing required propulsive power in cruise of up to 64%. The application of the control algorithm made notable a transition manoeuver that consists of negatively inclining the aircraft at a negative pitch angle, initially at high intensity, and as the final cruising speed approaches, the inclination is attenuated until the equilibrium pitch angle is reached. Simultaneously with the negative inclination of the pitch angle, there is a slight drop in altitude, which is quickly resumed as the trajectory develops until the final cruising speed. Lastly, this aircraft configuration can be widely used in applications where performance gains in operations currently carried out by multicopters, which cover large distances and need long flight time, would bring great operational advantages.

This paper aims to present a vertical take-off and landing unmanned aerial vehicle (VTOL UAV) configuration and numerically simulate its flight transition from hover to cruise and from cruise to hover. It can tilt the canard and wing along with two attached propellers. Additionally, two fixed front propellers are pointing upwards. Multi-body equations of motion are derived for this concept of aircraft, which are used to compute the flight transition trajectory from hover to cruise configuration. Furthermore, a transition control algorithm based on gain scheduling is described, which stabilises the aircraft while it accelerates from hover to cruise, gradually tilting the wing along with its propellers, sequentially switching between equilibrium states, as the stability cost functions thresholds are reached. The transition control algorithm of the conceptual aircraft model is numerically simulated.

The physics behind the transitions of natural Samaras, or the bio-inspired counterparts, to steady autorotation has been unclear. Theoretical and experimental investigations explore the inertial and aerodynamic characteristics required to guarantee stable transitions of an artificial Samara-like decelerator from chaotic tumbling motions to azimuthal autorotation. A non-dimensional inertial criterion is proposed, which is in accord with experiments.

Although the problem of locked-in deep stall is well documented over many years, there currently exists no consistent procedure that can guarantee recovery. Past studies have suggested that it might be possible to rock the aircraft in pitch to destabilise the statically stable deep stall trim point, thereby gaining enough momentum to push the nose down. However, the methods used in these studies are either of preliminary or empirical nature and cannot guarantee recovery. In this paper, we use bifurcation analysis to derive a recovery manoeuvre, specifically by assessing the aircraft’s nonlinear frequency response under an elevator forcing. The ensuing nonlinear Bode plot detects unstable (divergent) solutions near resonance that contribute to a successful deep stall recovery. Moreover, the nonlinear resonant frequency is slightly lower than the result obtained using linear analysis, and time simulation shows that relying on the linear result does not lead to a successful recovery. It is also found that at the high angles of attack associated with deep stall, the frequency separation between the short period and phugoid mode is significantly reduced, leading to only one visible peak in the frequency response. This feature is also reflected in the time-domain step response.

Spinning/Rolling Airframe Missiles (RAM) mostly use an analog (ON-OFF type) control approach that deflects control surfaces (fins) at minimum and maximum positions continuously, and control is achieved by applying phase shift between the minimum and maximum deflections during a complete roll cycle. Therefore, the control signal shape changes continuously during the roll cycle. In this study, a novel single channel digital controller is designed and tested for a high spinning rate (10–20 Hz) RAM. The digital controller adjusts the amplitude of fin deflections instead of applying a phase shift. In this way, control signal shapes are predetermined in design to completely decouple yaw and pitch dynamics. At the beginning of each roll cycle, the algorithm decides on the control signal shape and amplitude to apply throughout the cycle. Delays on the actuator system and sensor measurements which might lead to instability at high spinning rates are handled effectively thanks to the predetermined control signal shapes that are changing with 90-degree intervals. Detailed geometry of a surface-to-air spinning missile is used to obtain aerodynamic coefficients for the entire flight regime (i.e. from launch to terminal phase) via Missile DATCOM. The 6-DOF flight dynamics model and the controller algorithms are built in MATLAB/Simulink environment. The proposed digital controller is tested systematically for various scenarios and the performance is compared with the conventional analog control approach. The digital controller gives better performance compared to the analog approach under the influence of servo delays and sensor noise.

An evaluation of the effects of different ice configurations on the flight characteristics of a fixed-wing aircraft is presented. Within a joint research project of German Aerospace Center (DLR) and Brazilian aircraft manufacturer Embraer simulation models of three supercooled large droplet (SLD) ice configurations were developed for one Phenom 300 prototype. A specific flight test campaign with artificial SLD ice shapes on the aircraft was conducted to gather the relevant flight data. The simulation models for the different ice configuration were obtained by system identification, and specific results for the SLD-ice simulation models are provided. The analysis of aircraft characteristics was based on the results of these high-quality simulation models: the icing-induced changes of the flight performance characteristics evaluated by the interpretation of several parameters like thrust-to-weight ratio and specific excess power. The typical flight performance degradation was found for all ice configurations with different magnitude. The change of aircraft eigenmodes was investigated in detail by analysing the system matrix of the linearised models at a specific trim points. In addition, the diverse effects found for different ice configurations (App. C and SLD) are discussed and the change of root locus is analysed. Furthermore, ice-induced changes of the handling qualities are evaluated using numerical criteria of flying qualities standard “MIL-STD-1797 A”: no significant deterioration was found for the investigated ice configurations.

Turbulence and gusts cause variations in the aerodynamic forces and moments applied to the structure of aircraft, resulting in passenger discomfort and dynamic loads on the structure that it must be designed to support. By designing Gust Load Alleviation (GLA) systems, two objectives can be achieved: first, realizing higher passenger comfort; and second, reducing the dynamic structural loads, which allows the design of lighter structures. In this paper, a methodology for designing combined feedback/feedforward GLA systems is proposed. The methodology relies on the availability of a wind profile ahead of the aircraft measured by a Doppler LIDAR sensor, and is based on $H_{\infty}$-optimal control techniques and a discrete-time preview-control problem formulation. Moreover, to allow design trade-offs between those two objectives (to achieve design flexibility) as well as to allow specification of robustness criteria, a variant of the problem using multi-channel $H_{\infty}$-optimal control techniques is introduced. The methodology developed in this paper is intended to be applied to large aircraft, e.g. transport aircraft or business jets. The simulation results show the effectiveness of the proposed design methodology in accounting for the measured wind profile to achieve the two mentioned objectives, while ensuring both design flexibility and controller robustness and optimality.

The coaxial compound helicopter with lift-offset rotors has been proposed as a concept for future high-performance rotorcraft. This helicopter usually utilizes a variable-speed rotor system to improve the high-speed performance and the cruise efficiency. A flight dynamics model of this helicopter associated with rotor speed governor/engine model is used in this article to investigate the effect of the rotor speed change and to study the variable rotor speed strategy. Firstly, the power-required results at various rotor rotational speeds are calculated. This comparison indicates that choice of rotor speed can reduce the power consumption, and the rotor speed has to be reduced in high-speed flight due to the compressibility effects at the blade tip region. Furthermore, the rotor speed strategy in trim is obtained by optimizing the power required. It is demonstrated that the variable rotor speed successfully improves the performance across the flight range, but especially in the mid-speed range, where the rotor speed strategy can reduce the overall power consumption by around 15%. To investigate the impact of the rotor speed strategy on the flight dynamics properties, the trim characteristics, the bandwidth and phase delay, and eigenvalues are investigated. It is shown that the reduction of the rotor speed alters the flight dynamics characteristics as it affects the stability, damping, and control power provided by the coaxial rotor. However, the handling qualities requirements are still satisfied with different rotor speed strategies. Finally, a rotor speed strategy associated with the collective pitch is designed for maneuvering flight considering the normal load factor. Inverse simulation is used to investigate this strategy on maneuverability in the Push-up & Pull-over Mission-Task-Element (MTE). It is shown that the helicopter can achieve Level 1 ratings with this rotor speed strategy. In addition, the rotor speed strategy could further reduce the power consumption and pilot workload during the maneuver.

An important prerequisite for the design, assessment and certification of aircraft and their associated control systems is a quantitative specification of the environment in which the aircraft is intended to operate, for example, atmospheric gust. Gust loads on aircraft may induce detrimental influences such as increased aerodynamic and structural loads, structural deformation and decreased flight dynamic performance. This paper presents a systematic and comprehensive overview of important concepts and applications of gust loads on aircraft. This overview includes a brief research background, concepts, research techniques, influences and load alleviation measures of gust. Finally, we summarise some potential improvements in the future work. It is also recommended to learn from previous experiences to avoid aviation accidents due to flight through atmospheric gusts and turbulence.

It 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.

The development of high-performance computing and computational fluid dynamics methods have evolved to the point where it is possible to simulate complete helicopter configurations with good accuracy. Computational fluid dynamics methods have also been applied to problems such as rotor/fuselage and main/tail rotor interactions, performance studies in hover and forward flight, rotor design, and so on. The GOAHEAD project is a good example of a coordinated effort to validate computational fluid dynamics for complex helicopter configurations. Nevertheless, current efforts are limited to steady flight and focus mainly on expanding the edges of the flight envelope. The present work tackles the problem of simulating manoeuvring flight in a computational fluid dynamics environment by integrating a moving grid method and the helicopter flight mechanics solver with computational fluid dynamics. After a discussion of previous works carried out on the subject and a description of the methods used, validation of the computational fluid dynamics for ship airwake flow and rotorcraft flight at low advance ratio are presented. Finally, the results obtained for manoeuvring flight cases are presented and discussed.

Some helicopter manufacturers are exploring the compound helicopter design as it could potentially satisfy the new emerging requirements placed on the next generation of rotorcraft. It is well understood that the main benefit of the compound helicopter is its ability to reach speeds that significantly surpass the conventional helicopter. However, it is possible that the introduction of compounding may lead to a vehicle with significantly different flight characteristics when compared to a conventional helicopter. One method to examine the flight dynamics of an aircraft is to create a linearised mathematical model of the aircraft and to investigate the stability derivatives of the vehicle. The aim of this paper is to examine the stability derivatives of a compound helicopter through a comparison with a conventional helicopter. By taking this approach, some stability, handling qualities and design issues associated with the compound helicopter can be identified. The paper features a conventional helicopter and a compound helicopter. The conventional helicopter is a standard design, featuring a main rotor and a tail-rotor. The compound helicopter configuration features both lift and thrust compounding. The wing offloads the main rotor at high speeds, whereas two propellers provide additional propulsive thrust as well as yaw control. The results highlight that the bare airframe compound helicopter would require a larger tailplane surface to ensure acceptable longitudinal handling qualities in forward flight. In addition, without increasing the size of the bare airframe compound helicopter’s vertical fin, the Dutch roll mode satisfies the ADS-33 level 1 handling qualities category for the majority of the flight envelope.

Aircraft damage modelling was conducted on a Boeing 747 to examine the effects of asymmetric horizontal stabiliser loss on the flight dynamics of a commercial Fly-by-Wire (FBW) aircraft. Robustness of the control system is investigated by analysing how characteristic eigenvalues move as a result of damage and comparison to the non-FBW aircraft is made. Furthermore, the extent of stabiliser loss that the system can successfully handle without loss of stability and acceptable performance is identified. The presented analysis of the results gives insightful knowledge to aid in the design of an improved FBW system with increased damage tolerance. A handling qualities evaluation is presented to provide an understanding of how the pilot perceives the damaged aircraft. The results of the study show that a generic FBW system improves robustness such that the aircraft is stable with 50% horizontal stabiliser loss. With 50% damage, the aircraft is controllable but unsafe to fly and may be unable to effectively complete its mission task.