Trimmed Flight Conditions Research Articles

Fixed-Time Aircraft Spin Recovery Control Based on a Nonsingular Integral Terminal Sliding Mode Approach

This paper introduces a control strategy to recover from various spin motions to level trim flight conditions, utilizing a fixed-time nonsingular terminal sliding mode approach to minimize altitude loss. Firstly, an integral terminal sliding mode surface is constructed to guarantee the tracking errors of the spin recovery system converge to the origin within a fixed time and avoid the singularity problem. Next, a fixed-time type reaching law is proposed to weaken the chattering phenomenon and guarantee terminal sliding hyperplanes converge to zero in a fixed time. Furthermore, the stability analysis indicates the fixed-time convergence stability of the spin recovery control system and provides an accurate calculation of the settling time. Finally, comparative simulations are performed to validate the effectiveness and superiority of the developed spin recovery control scheme.

Validation of Variable Rotor Speed Trim Using Computational Fluid Dynamics

A multirotor trim algorithm is developed and validated using the HPCMP CREATETM -AV Helios rotorcraft analysis and simulation framework. Validation data were sourced from experiments in which a quadrotor with xed-pitch rotors was trimmed for steady hover and forward flight in the NASA Langley Low-Speed Aeroacoustic Wind Tunnel. For each ight condition, the aircraft attitudes were fixed so that the only control variables were the rotation rates of the four rotors, which were varied until the residual loads were minimal. The resultant trim states were rst replicated in the computational framework to verify the aerodynamic solver. Then, the trim algorithm was validated in Helios by constraining the aircraft attitudes and searching for the rotor speeds that minimized the residual loads. The analysis demonstrated excellent agreement between the predicted and measured trim states. Finally, free-flight trim cases were simulated to quantify the effect of the trim constraints and verify the experimentally trimmed ight conditions. The predicted free-flight trim state showed reasonable agreement to the constrained case, with negligible change in the residual loads, indicating trim was achieved in both the experiments and the simulations.

Optimal Performance Trim Procedure for Multirotor Vertical Takeoff and Landing Vehicles

The present paper proposes a general approach capable of determining the set of trim commands, aerodynamic control surface actuation, and/or engine thrust regulation that allow specific trimmed steady-state flight conditions of nonconventional vertical takeoff and landing vehicles under given performance constraints. Exploiting the peculiarity of these vehicles to have redundant controls, namely more controls than those strictly required to define steady-state flight conditions, their trim problem is recast into a constrained minimization of a cost function. This approach is of particular interest in that it allows the identification of control settings for multirotor vehicles that, at the same time, guarantee trimmed flight conditions and optimization of selected target functions such as, for instance, performance or noise emission. The proposed approach is numerically applied to quadcopter and hexacopter configurations for three flight conditions (level forward flight, climb, and coordinate turn), and three strategies of trim by different trim variables (rotor angular velocity, blade collective pitch, and combination of angular velocity and blade collective pitch). The effect of the introduced cost function is investigated, comparing the outcomes of minimum control effort and minimum torque strategies. The results achieved demonstrate the capability of the proposed algorithm to exploit redundant controls of multirotor vehicles for identifying selected optimal trim operative conditions.

Conceptual Sizing of Vertical Lift Package Delivery Platforms

Conceptual sizing and performance estimation of four vertical takeoff and landing (VTOL) configurations for a package delivery mission is presented in this Paper. The multifidelity VTOL design framework HYDRA is used to size a notional quadrotor, hexarotor, quadrotor biplane tailsitter (QBiT), and a lift-augmented tricopter for weight classes of 10, 15, 20, and 25 kg. Sizing is performed using a combination of physics-based empty weight models for the airframe, rotor blades, and wings, along with statistical models for predicting motor and battery weights. A gradient-based optimization methodology was implemented to identify the best designs for each configuration under each weight class. An optimization-based longitudinal trim methodology was used to identify the appropriate shaft power settings for each vehicle at various trimmed flight conditions along the mission profile. Representative payload delivery and return scenarios were constructed from different cruise speeds and ranges to identity a vehicle design in each configuration that can complete the highest number of payload drop missions successfully. The hexarotor performed better than the quadrotor in terms of requiring lower installed power and deliver heavier payload packages for a given radius of action. Wing-borne designs such as the QBiT and tricopter are capable of delivering packages in a short time owing to full/partial conversion to airplane mode during cruise flight but may require careful fine-tuning of the trimline and flight control system to successfully transition between hover and cruise flight modes.

Trim Flight Conditions for a Low-Boom Aircraft Design Under Uncertainty

The purpose of this paper is to investigate the noise generation of a low-boom aircraft design in flight trim conditions under uncertainty. The deflection of control surfaces to maintain a trimmed flight state has the potential to change the perceived loudness at the ground. Furthermore, the uncertainties associated with the control surface deflections can complicate the overall uncertainty quantification. Incorporation of the uncertainties in the prediction of perceived sound levels during the design phase can lead to improved robustness. In this paper, a brief review of low-boom flight trim research is presented. Realistic flight trim conditions requiring control surface deflection are integrated into the current research efforts for uncertainty quantification and vehicle design. In addition, a generalized set of procedures for the characterization of uncertainties in flight trim conditions are introduced. In a case study of the application of these procedures, a 5 dB average difference in the perceived level of loudness was found between clean (no deflections) and trimmed configurations. Also, uncertainties attributable to control surface deflection were found to account for, on average, over 50% of the total near-field uncertainty. Uncertainty discretization methods implemented were able to give more insight into the overall global variances.

Aerodynamic Modeling, System Identification, and Analysis of Iced Aircraft Configurations

To characterize changed aircraft dynamics due to icing an incremental model extension ( model) is analytically derived, and its parameters are estimated. This model extension is formulated as a separate module in the aerodynamic part of the aircraft simulation and can be used in existing simulation models. Within a cooperation of DLR, German Aerospace Center and Embraer flight data of a light business jet with different ice accumulation cases are used to identify the parameters of a clean and a model describing icing effects. The model parameters are interpreted with respect to the aerodynamic degradation caused by icing. An analysis of different asymmetric icing cases is performed using the identified model. Trimmed flight conditions deliver information about the maximum expectable additional icing-induced moments and the necessary control surface deflections compensating the disturbance. The ability to cover various icing effects is shown by dynamic simulations reducing local icing effects similar to ice shedding situations in flight.

An optimization study on load alleviation techniques in gliders using morphing camber

This study explores wing morphing for load alleviation as a means to reduce the required wing structural weight without compromising aircraft performance. A comparative study between the lift-to-drag ratio (L/D) performance of a fixed wing glider (FWG) and a cambered morphing wing glider (CMWG) is presented. Both aircraft are aero-structurally optimized for the best L/D for a given speed and payload mass. A combination of lifting-line theory and 2D viscous calculations is used for the aerodynamics and an equivalent beam model is employed for the structural analysis. Pull-up and -down maneuvers at 25 m/s and near stall angle of attack are assumed as critical load cases. Results of the FWG optimization are shown for several trimmed flight conditions with varying mass and velocity. Results are compared to the ones from the CMWG optimization and conclusions are drawn on the improvement in the L/D ratio throughout the flight envelope and on potential reductions in the wing structural mass due to the load alleviation strategy. The wing camber adaptation provides significant performance gains in a large range of flight speeds with negligible penalties in the low speeds range. However, maneuverability is penalized.

Fault tolerant heading control system design for Turac unmanned aerial vehicle

In this study, a fault tolerant heading control system is designed for a one-third scale fixed wing vertical takeoff-and-landing unmanned aerial vehicle, Turac. A nonlinear six degrees-of-freedom (DoF) mathematical model is obtained and linearized at the calculated trim flight condition. A proportional heading control system is designed as a nominal horizontal flight controller. Detection and isolation of the faults that can occur during flight are performed by Kalman filters which are designed individually for each sensor output. After the fault isolation process the obtained fault data is fed to the reconfigurable Kalman filter. Then the feedback signal from the faulty sensor is blocked and the estimated output from the reconfigurable Kalman filter is fed to the control system. So, the closed-loop system could follow the reference signal without updating the controller’s parameters. Simulation studies are performed on the closed-loop system for faulty sensor situations.

Effect of body aerodynamics on the dynamic flight stability of the hawkmoth Manduca sexta

This study explores the effects of the body aerodynamics on the dynamic flight stability of an insect at various different forward flight speeds. The insect model, whose morphological parameters are based on measurement data from the hawkmoth Manduca sexta, is treated as an open-loop six-degree-of-freedom dynamic system. The aerodynamic forces and moments acting on the insect are computed by an aerodynamic model that combines the unsteady panel method and the extended unsteady vortex-lattice method. The aerodynamic model is then coupled to a multi-body dynamic code to solve the system of motion equations. First, the trimmed flight conditions of insect models with and without consideration of the body aerodynamics are obtained using a trim search algorithm. Subsequently, the effects of the body aerodynamics on the dynamic flight stability are analysed through modal structures, i.e., eigenvalues and eigenvectors in this case, which are based on linearized equations of motion. The solutions from the nonlinear and linearized equations of motion due to gust disturbances are obtained, and the effects of the body aerodynamics are also investigated through these solutions. The results showed the important effect of the body aerodynamics at high-speed forward flight (in this paper at 4.0 and 5.0 m s−1) and the movement trends of eigenvalues when the body aerodynamics is included.

Multiple time scale analysis of aircraft longitudinal dynamics with aerodynamic vectoring

The paper analyzes the longitudinal dynamics of sustained high angle-of-attack flights performed by a small agile Unmanned Aerial Vehicle (UAV) equipped with an aerodynamic vectoring feature. This unique feature is achieved through variable-incidence wing that allows the angle of incidence of the wing with respect to fuselage to be decoupled. The aerodynamics of the UAV is estimated through wind-tunnel experiments. Subsequently, the dynamics of the UAV motion are studied for trim conditions across the complete spectrum of velocity between hover and forward cruise flight. The short-period eigenvalue migration, as the trimmed flight conditions are changed, studied, and the peculiarities observed from this linear analysis are highlighted. Further, the Multiple Time Scales method in conjunction with bifurcation theory is used to obtain the approximate solutions of the multiple degree-of-freedom nonlinear equations of motion. The explicit analytical results obtained are useful to identify the key parameters affecting the dynamics and stability of the aircraft’s longitudinal motion. The conditions leading to limit cycle responses in the vicinity of the stall regime are also highlighted.

Ornithopter Flight Simulation Based on Flexible Multi-Body Dynamics

This paper introduces a flight simulation of an ornithopter (flapping-wing air vehicle) based on the flexible multi-body dynamics, a refined flapping-wing aerodynamic model and the fluid-structure interaction approach. A simulated ornithopter was modeled using the multi-body dynamics software, MSC.ADAMS, where the flexible parts can be included by importing a finite element model built in the finite element analysis software, ANSYS. To model the complex aerodynamics of flapping-wing, an improved version of modified strip theory was chosen. The proposed integrative simulation framework of ornithopter was validated by the wind tunnel test data reported in the literature. A magpie-sized model ornithopter was numerically designed and simulated to have the longitudinal trim flight condition. We observed a limit-cycle-oscillation of flight state variables, such as pitch attitude, altitude, flight speed, during the trimmed flight of the model ornithopter. Under the trimmed condition of free flight of the model ornithopter, we fixed all the degrees of freedom at the center of gravity to measure the constraint forces and moment. The concept of the “zero moment point” is introduced to explain the physics of ornithopter trimmed longitudinal flight.

Aircraft Spin Recovery, with and without Thrust Vectoring, Using Nonlinear Dynamic Inversion

The present paper addresses the problem of spin recovery of an aircraft as a nonlinear inverse dynamics problem of determining the control inputs that need to be applied to transfer the aircraft from a spin state to a level trim flight condition. A stable, oscillatory, flat, left spin state is first identified from a standard bifurcation analysis of the aircraft model considered, and this is chosen as the starting point for all recovery attempts. Three different symmetric, level-flight trim states, representative of high, moderate, and low-angle-of-attack trims for the chosen aircraft model, are computed by using an extended-bifurcation-analysis procedure. A standard form of the nonlinear dynamic inversion algorithm is implemented to recover the aircraft from the oscillatory spin state to each of the selected level trims. The required control inputs in each case, obtained by solving the inverse problem, are compared against each other and with the standard recovery procedure for a modern, low-aspect-ratio, fuselage heavy configuration. The spin recovery procedure is seen to be restricted because of limitations in control surface deflections and rates and because of loss of control effectiveness at high angles of attack. In particular, these restrictions adversely affect attempts at recovery directly from high-angle-of-attack oscillatory spins to low-angleof-attack trims using only aerodynamic controls. Further, two different control strategies are examined in an effort to overcome difficulties in spin recovery because of these restrictions. The first strategy uses an indirect, two-step recovery procedure in which the airplane is first recovered to a high- or moderate-angle-of-attack level-flight trim condition, followed by a second step where the airplane is then transitioned to the desired low-angle-of-attack trim. The second strategy involves the use of thrust-vectoring controls in addition to the standard aerodynamic control surfaces to directly recover the aircraft from high-angle-of-attack oscillatory spin to a low-angle-of-attack level-flight trim state. Our studies reveal that both strategies are successful, highlighting the importance of effective thrust management in conjunction with suitable use of all of the aerodynamic control surfaces for spin recovery strategies.

Non-linear anti-windup for manual flight control

In this paper we address the windup problem arising from actuator magnitude and rate saturation for the linearized longitudinal short-period dynamics of a tailless aircraft model around any trim flight condition. The proposed anti-windup scheme allows for more aggressive manoeuvres than the standard “command limiting” approach. Moreover, the compensation law guarantees stability of the controlled aircraft for any pilot command and enforces flight quality specifications whenever they are achievable within the given control constraints. “Pilot-in-the-loop” simulations confirm the effectiveness of the control strategy.

Hingeless-Rotor Aeromechanical Stability in Axial and Forward Flight With Wake Dynamics

A finite-state wake model is used to investigate aeromechanical stability of hingeless-rotor helicopters in the ground-contact, hover and trimmed-night conditions. The investigation covers three items: (1) the convergence of the damping with increasing number of wake harmonics for the lag regressing, and body pitch and roll modes; (2) a parametric study of the damping over a range of thrust level, advance ratio and number of blades; and (3) correlations, primarily with the damping and frequency measurements of these lag and body modes. The convergence and parametric studies are conducted in the hover and trimmed-flight conditions; they include predictions from the widely used dynamic inflow model. The correlations are conducted in the ground-contact conditions and include predictions from the dynamic inflow and vortex models; recently, this vortex model is proposed for the axial-flight conditions and is used to investigate the coupled free vibrations of rotor flapping and body modes. The convergence and parametric studies show that a finite-state wake model that goes well beyond the dynamic inflow model is required for fairly converged damping, Moreover, the correlations from the finite-state wake, dynamic inflow and vortex models are generally satisfactory.

Anticipated Spin Susceptibility Characteristics of the A-10 Aircraft

This paper presents the results of a real-time six-degree-of-freedom computer analysis of the spin entry, spin, and recovery characteristics of the U.S. Air Force A-10 close air-support aircraft. The primary emphasis of this study was directed toward determination of the spin susceptibility of the A-10 aircraft during l-g trim and accelerated flight conditions. The computer analysis was based upon 640 wind-tunnel test hours on 1/10and 1/20scale models of the prototype A-10 aircraft at high angles of attack. The analysis approach followed guidelines of the U.S. Air Force specification, MIL-S-83691A, Stall/Post-Stall/Spin Flight Test Demonstration Requirements. The results indicated that three spin modes exist for the A-10 aircraft but that only one of these spin modes (a low angle-of-attack spin mode) could be entered with misapplied control inputs from l-g trim or accelerated flight conditons. The exception to this statement was the fully extended speedbrake configuration, where two spin modes were obtained from a l-g trim flight condition. Recoveries from all spin modes were shown to be excellent except for the high angle-of-attack spin mode, where these results were inconclusive.