Maneuver Loads Research Articles

Active maneuver load alleviation for a pitching wing via spanwise-distributed camber morphing

This paper presents the design and verification of a nonlinear model inversion (NMI) controller for the maneuver load alleviation of a pitching oscillating wing based on spanwise-distributed active camber morphing. Recurrent neural networks (RNNs) are used to predict nonlinear and unsteady aerodynamic forces due to wing's large amplitude pitching maneuver, and a fully connected neural network is introduced to build the dynamic inversion of the aeroelastic system for control law design. The inversed system is concatenated with a PI controller to assemble a nonlinear active controller. The controller is first utilized in an offline environment for a 1DoF pitching finite-span wing with spanwise-distributed active camber morphing and then verified in CFD-based fluid-structure-control coupling simulation. The results show that the offline controller could eliminate the maneuver load. In the online CFD-based fluid-structure-control simulation, the bending moment can be alleviated by 38%.

Dynamic modeling and response analysis of misaligned rotor system with squeeze film dampers under maneuver loads

Dynamic modeling and response analysis of misaligned rotor system with squeeze film dampers under maneuver loads

Vibration characteristics of dual-rotor system with staggered-type double elastic ring squeeze film dampers under maneuvering flight

The vibration of the rotor system is significantly intensified during maneuvering flight, rendering traditional elastic ring squeeze film dampers (ERSFDs) insufficient or unstable. To improve the applicability of dampers, a novel structure of staggered-type double elastic ring squeeze film dampers (SDERSFD) is proposed in this paper. Through the synergistic effect of double elastic rings, the damping effect is significantly enhanced, and the circumferential stiffness anisotropy of the damper is improved, which is more suitable for vibration reduction requirements under complex working conditions. Utilizing the Reynolds equation, the internal and external pressure field control equations of the double elastic rings are established. The finite difference method is then employed to solve the dynamic characteristics of the damper. Additionally, a finite element model of the dual-rotor system with SDERSFDs considering complex dynamic factors such as unbalanced force, oil film force, gravity, and additional inertial force caused by maneuvering flight is established for the study of nonlinear rotor dynamic characteristics and the analysis of SDERSFDs vibration reduction effect. Based on the model, the influence of various factors such as maneuvering speed, maneuvering radius, rotor speed and unbalance on the transient response of dive-pull up condition is analyzed. The results reveal that during dive-pull up, the vibration of the rotor system intensifies significantly. When compared with the ERSFDs, the SDERSFDs effectively reduce the vibration amplitudes of the rotor system under maneuvering flight condition, achieving the maximum vibration reduction ratio of 48.9%. Furthermore, the SDERSFD exhibits robustness against large maneuvering loads, enhancing the stability of the rotor system under extreme conditions.

Impact of Dynamic Wheel Loading on Flexible Pavement Responses for Non-Free Rolling Conditions

Since 2000, pavement design methodologies have transitioned from empirical to mechanistic-empirical procedures. However, the complexities of loading patterns, contact stress distribution, material characterization, vehicle maneuvering, and dynamic load amplification are still not fully considered, despite their significant effect on pavement performance. In this study, state-of-the-art numerical models were used to investigate the changes in critical pavement responses derived from the combined effect of roughness-induced dynamic loading and vehicle maneuvering. A decoupled vehicle–tire–pavement interaction approach composed of a random process to generate artificial road roughness profiles, a mechanical full truck model, and three-dimensional finite element tire and flexible pavement models allow the prediction of the impact of road roughness on vehicle dynamics. In addition, the study quantifies the effect of dynamic loading on contact stresses, and consequently, on pavement critical responses. Because of the expected increase in axle weight from truck electrification, overweight scenarios were also considered. Changes in load history and distribution of strain fields were assessed for two typical pavement structures (thin and thick). The combined effect of roughness-induced dynamic loading and vehicle maneuvering greatly altered the critical pavement responses associated with bottom-up fatigue cracking, near-surface cracking, and rutting. Under the most adverse conditions, the critical tensile strains at the bottom of the asphalt concrete increased up to 125%, the shear strain increased up to 100%, and the compressive strain escalated to 120% when compared with the reference cases. Higher temperatures exacerbated the impact of dynamic wheel loading and vehicle maneuvering.

Fluid–Structure Coupled Analysis of Maneuver Load Alleviation on a Large Transport Aircraft

Fluid–structure coupled simulations are performed for quasi-steady pitching maneuvers of a large transport aircraft with a load alleviation system consisting of trailing edge flaps and droop noses at the leading edge. For the fluid part of the simulation, a RANS approach is used, whereas the structural simulation is based on a linear modal model. The selected maneuvers are located on the lift boundary of the maneuver envelope at maximum load factor, as not only are the structural loads high here, but also redistributing lift for load alleviation is especially challenging as the wing is close to stall. For a target value of 33% reduction in wing bending loads, which is derived from the CS-25, the effects of applying load alleviation on wing loads, deformations, and aircraft maneuverability are investigated. It is found that the desired reduction of wing bending loads can be achieved for all maneuvers considered, while controlling the torsional moment turns out to be more difficult. A loss of maneuverability due to load alleviation is observed for only one of the maneuvers, while for the other maneuvers, maneuverability could even be increased. To assess the influence of the elasticity of the wing, additional simulations of the rigid wing are carried out and compared with the elastic results. Finally, other potential applications of the considered flap system are also explored. Here, for example, a cruise drag reduction of up to 1.8% is possible by adjusting the spanwise lift distribution.

Air and Structural Loads Analysis of a 5-Ton Class Rotorcraft in a Pull-Up Maneuver Using CFD/CSD Coupled Approach

The air and structural loads of a 5-ton class light helicopter (LH) rotor in a 2.24 g pull-up maneuver are investigated using a coupling between the computational structural dynamics (CSD) and computational fluid dynamics (CFD) methods. The LH rotor is characterized by a five-bladed system with elastomeric bearings and inter-bladed dampers. The periodic trim solution along with the converged CFD/CSD delta airloads obtained in steady-level flight (advance ratio of 0.287) are used to perform the transient CSD maneuver analysis. The resulting vehicle attitude angles and velocity profiles of the aircraft are then prescribed in the quasi-static (QS) CFD maneuver analysis. It is demonstrated that the present QS approach provides an effective means for the maneuver loads’ analysis. The important flow behaviors such as BVI (blade–vortex interaction)-induced oscillations and the negative pitching moment peaks met in maneuver flight are captured nicely with the proposed method. Either the vortex trajectories or the surface pressure distributions are examined to identify the sources of the oscillations. A loose CFD/CSD coupling (LC) is used to predict the blade elastic motions, structural moments, and pitch link loads at the specified maneuver revolution of the rotor and also to correlate these with the transient CSD-based predictions. A reasonable correlation is obtained. The LC results show more pronounced 5P (five per revolution) oscillations on the structural response than those of the CSD-based methods.

Integrated design of lateral-directional control law and vertical tail load for civil aircraft

According to new requirements of airworthiness regulation CS25.353, the advanced lateral-directional P-Beta control law design is studied in this paper. The LQR method based on the tracking command model is used to design the lateral-directional control law parameters. At the same time, the vertical tail load model of aircraft for yaw maneuver is introduced into the lateral-directional control law design model. The aim is to build the integrated design of control law and flight load. Finally, the simulation results show that the P-Beta control law can not only improve Dutch roll damping and coordination turn characteristics, but it can also reduce the yaw maneuver load by 18.5%. Thus, the structure weight of the aircraft’s vertical tail can be reduced.

Transient state analysis of a rub-impact rotor system during maneuvering flight

Maneuvering flight substantially affects the dynamic behavior of rotors; particularly, such flight may cause rubbing between a rotor and stator, which is one of the most serious damages in aircraft engines. In this paper, a nonlinear dynamic model for describing the dynamic characteristics of a rub-impact rotor system during maneuvering flight is established based on the Lagrange equations. Subsequently, numerical simulations employing the Newmark method are performed, delving into the detailed discussion of the influence of parameters such as rotational speed and maneuvering flight on the transient and steady-state responses of the rotor system. The effect mechanism of maneuver load and its coupling with rub impact is revealed. The results show that the impact response induced by maneuvering flight is more obvious in the subcritical state than in the supercritical state. The additional stiffness and damping are also induced; in particular, the additional damping has a coupling effect. Moreover, the rub impact imposes an additional constraint on the rotor system, thereby weakening the influence of the maneuver load and becoming the major factor that determines the dynamic behavior of the rotor system at high speeds.

Aeroelastic Analysis of Actuated Adaptive Wingtips Based on Pressure Actuation

Folding wingtips address the challenges posed by high-aspect-ratio wings, such as airport conformity and increased wing root bending moment. Actuated adaptive wingtips extend the functionalities of folding wingtips by using a stiffness-adaptive aeroelastic hinge that is actively adjustable in flight. The objective of this paper is the aeroelastic analysis of a wing equipped with an adaptive-stiffness hinge. While the structural design of the wingtip actuator based on pressure-actuated cellular structures (PACS) was developed in a previous study, in this study the authors verify the concept of actuated adaptive wingtips through aeroelastic analysis. This study shows that the investigated PACS actuator, structurally designed from glass-fiber-reinforced plastic, is capable of bearing the loads acting on the wingtips of a Cessna Citation X. The adaptive-stiffness hinge, positioned between 86.7 and 91.2% of the semispan, reduces the wing root bending moment by up to 7.8% in a 2.5g maneuver load case, while keeping the wing straight in cruise. A further increase in load alleviation potential can be achieved in the future by extending the actuator’s operating envelope and thus increasing its load-bearing capacity. The functional verification of the actuated adaptive wingtip concept by means of aeroelastic analysis forms the basis for the manufacturing and testing of a functional prototype.

Research on Aero-engine Maneuvering Load Distribution Based on POT Model

Research on Aero-engine Maneuvering Load Distribution Based on POT Model

Parametric aeroelastic modeling, maneuver loads analysis using CFD methods and structural design of a fighter aircraft

The DLR Future Fighter Demonstrator (FFD) is a highly agile, two-seated aircraft with twin-engines equipped, a reheat and a design flight speed extending into the supersonic regime (up to Ma=2.0). Based on a given conceptual design, the presented work focuses on the aeroelastic modeling, including structures, masses and aerodynamics. Using the models, a comprehensive loads analysis with 688 maneuver load cases, covering the whole flight envelope, is performed. Comparing the results obtained from aerodynamic panel methods (VLM and ZONA51) with higher fidelity results obtained from CFD, the necessity of CFD based maneuver loads analysis in preliminary design of such fighter configuration is shown, as it leads to physically different as well as higher loads. The rigorous application of CFD is a heavy burden during the preliminary design, but this work demonstrates that it is doable as of today. Finally, the model is subject to structural optimization, demonstrating that the differences in loads result in a heavier primary structural net mass with ≈3.3t for the approach based on aerodynamic panel methods and ≈4.1t for the CFD based approach. Because all remaining models are unchanged, this difference in mass can be clearly attributed to the physical differences in the flow solutions obtained from the panel methods and CFD.

Vibration Behavior of Dual-Rotor Caused by Maneuver Load and Intershaft Bearing Defect

This paper concentrates on the vibration amplitude modulation and frequency modulation behavior of a dual-rotor system induced by the maneuver load and defective intershaft bearing. The motion equations of a dual-rotor system are presented by considering the intershaft bearing force with a local defect of the outer raceway and the maneuver load of a barrel-roll flight. The fourth-order Runge–Kutta method is employed to study the vibration responses of the dual-rotor system, which are reflected by the time history curves, the bearing contact force curves, and the Hilbert envelope spectra. It is shown that the rotor responses display impulse vibrations due to the defective intershaft bearing; and the corresponding Hilbert envelope spectra contain defect characteristic frequencies, multiple defect frequencies, side frequencies, and modulation frequencies. The results reveal the frequency modulation behavior of the impulse vibrations induced by the rotating speed ratio and the amplitude modulation behavior influenced by the maneuver load, the bearing defect span, and the rotating speed. Finally, defective intershaft bearing experiments under maneuvering flight are conducted, confirming the availability of the numerical vibration amplitude modulation and frequency modulation behavior. This analysis provides a theoretical foundation for bearing state detection during flight maneuvers.

Dynamic modeling and response analysis of rub-impact rotor system with squeeze film damper under maneuvering load

With the development of the aero-engine rotor system towards accurate modeling, it is necessary to establish a general rotor system model considering many important strong nonlinear factors. In this paper, based on the finite element method, a 6-node rub-impact rotor system model including the newly-built blade-casing rub-impact force, climbing maneuvering load, nonlinear Hertz contact force of rolling bearing, rotor eccentric unbalance force, gravity field and squeeze film damper oil film force is established. The dynamic responses of the system under different rub-impact stiffness, different oil film clearance and different bearing clearance are analyzed in detail by means of bifurcation diagram, three-dimensional spectrum diagram and axis trajectory diagram. Through comparative analysis, it is found that the change of the oil film clearance has the greatest influence on the rub-impact response of the system, followed by the bearing clearance, and the change of the rub-impact stiffness has the least influence on the rub-impact response of the system. The study of this paper can provide a theoretical basis for the rub-impact response of the rub-impact rotor system with squeeze film damper under climbing maneuvering flight.

A novel compilation method of comprehensive mission spectrum of aero-engine maneuvering load based on use-related mission segment

Comprehensive Mission Spectrum (CMS) of an aero-engine can reflect the usage characteristics of the engine. It can provide load input for engine life prediction and accelerated mission test. In this paper, a novel compilation method of CMS of aero-engine maneuvering load based on mission segment is proposed. Firstly, the use-related Typical Mission Segment (TMS) of maneuvering load is divided and identified according to spectral characteristics. Secondly, the mathematical model of different kinds of TMS are established based on stochastic process theory. Finally, the CMS of maneuvering load is compiled based on TMS. The proposed method can accurately quantify the compilation of CMS. The compiled CMS shows good agreement with the original load spectrum. According to damage consistency inspection, the compiled CMS is consistent with the damage caused by the original load spectrum in terms of low cycle fatigue.

Potential Estimation of Load Alleviation and Future Technologies in Reducing Aircraft Structural Mass

In recent years, load alleviation technologies have been more widely used in transport aircraft. For aircraft already in service, load alleviation can contribute in extending the fatigue life, or enable small configurational changes. If load alleviation is considered in the aircraft design process, the structural mass of the aircraft can be reduced. This paper investigates various maneuver and gust load alleviation algorithms as well as potential future technologies regarding flight operation, turbulence forecast and material science, and it evaluates the mass reduction that can be achieved. In doing so, a long-range transport aircraft was taken as the reference, and the considered load case conditions were 1-cos gusts, maneuvers and quasi-steady landing. Based upon the loads, the composite structure of the lifting surfaces was optimized, while the secondary masses as well as the wing planform were kept unchanged. With all technologies implemented, a reduction of the wing box mass by 26.5% or 4.4% of the operating empty mass could be achieved.

Load Alleviation of Flexible Aircraft by Dynamic Control Allocation

As wing designs aim for higher aerodynamic efficiency, the underlying aircraft structure becomes more flexible, requiring additional features to alleviate the loads encountered from gusts and maneuvers. While alleviating loads, it is desirable to minimize the deviations from the original flight trajectory. In this work, a dynamic control allocation method that exploits redundant control effectors for maneuver and gust load alleviation is proposed for flexible aircraft. The control architecture decouples the two objectives of load alleviation and rigid-body trajectory tracking by exploiting the null space between the input and the rigid-body output. A reduced-dimensional null space input is established, which affects the flexible output (but not the rigid-body output) when passed through a null space filter to generate incremental control signals. This null space input is determined by model predictive control to maintain the flexible output of the aircraft within specified values, thereby achieving load alleviation. Numerical simulations are used to illustrate the operation of this load alleviation system on nonlinear models. It is shown that the proposed load alleviation system can successfully avoid the violation of load bounds in the presence of both gust disturbances and maneuvers with minimal effect on the trajectory tracking performance.

Global aero-structural design optimization of composite wings with active manoeuvre load alleviation

In the scope of the DLR project VicToria (Virtual Aircraft Technology Integration Platform), an integrated process for aero-structural wing optimization based on high fidelity simulation methods is continuously developed and applied. Based upon a parametric geometry, flight performance under transonic flight conditions and manoeuvre loads are computed by solving the Reynolds-averaged Navier–Stokes equations. Structural mass and elastic characteristics of the wing are determined from structural sizing of the composite wing box for essential manoeuvre load cases using computational structural mechanics. Static aeroelastic effects are considered in all flight conditions and active manoeuvre load alleviation is integrated in the process. Global aero-structural wing optimizations are successfully performed for wings with and without active manoeuvre load alleviation. The active manoeuvre load alleviation is introduced with a simplified modelling of control surface deflections using a mesh deformation technique. The minimization of the fuel consumption for three typical flight missions represents the objective function. Wing optimizations are performed for variable and constant wing planform parameters as well as for wings with conventional composite wing box structure and for more flexible wings. The latter is accomplished by introducing modifications of the structural concept and the strain allowable. A significant mass reduction of the optimized wing box is obtained for wings with active manoeuvre load alleviation, resulting in a drop in fuel consumption of about 3%. For wing optimizations with the more flexible wing concept, the active manoeuvre load alleviation shows an additional reduction of the fuel consumption in the order of 2%. The wings with active manoeuvre load alleviation results in optimized wing geometries with increased aspect ratio and reduced taper ratio.

Nonlinear dynamics of flexible diaphragm coupling’s rotor system during maneuvering flight

In this paper, the flexible multi-diaphragm coupling which is used as flexible power transmission shaft of the aeroengine accessory is taken as the research object, and the coupling stiffness matrix and axial nonlinear stiffness of the diaphragms are considered in the coupling rotor system. On this basis, in order to consider the influence of aircraft maneuvering load, in non-inertial system the bending-pendular-axial coupled differential equations of flexible diaphragm coupling were established by Lagrange method. The modal characteristics of the flexible diaphragm coupling were analyzed and compared with the finite element solutions, and the correctness of model is verified. Runge-Kutta method is used to solve and analyze the influence of different maneuvering flight conditions on the vibration characteristics of the flexible diaphragm coupling. The research indicates that the coupling between diaphragm’s axial and radial stiffness leads to the right shift of resonant region, the increase of resonance peak value, and the nonlinear characteristics of amplitude-frequency curve such as jump and multi-value. In the non-inertial system, only the installation distance a of the flexible diaphragm coupling along the wingspan leads to the increase of the axial deformation offset of the flexible diaphragm coupling in the rolling flight state. The increase of climbing or diving angular velocity makes the flexible diaphragm coupling’s vibration changes from single period to multi-period, bifurcation or chaos state; With the increase of diving angular velocity and rolling angular velocity, the axial critical speed gradually increases; Each flight condition not only affects the vibration characteristics, but also causes the axial, radial and angular deformation of the flexible diaphragm coupling to a certain extent. This study provided a theoretical basis and method for the design and analysis of diaphragm coupling.

Coandă-Type Flow Actuation for Load Alleviation

Circulation control has the potential to improve current gust and maneuver load alleviation schemes, with the ultimate goal of reducing wing structure weight and saving fuel. A Coandă-type flow actuator is integrated on the pressure side of a two-dimensional airfoil model, and its load alleviation performance and dynamic behavior are determined experimentally. Quasi-steady force and wake measurements are conducted in a wind tunnel to quantify the effect of the actuator on lift and drag. Lift decrements of are reached for steady blowing, exceeding typical gust-induced wing loads in magnitude. For pulsed blowing, only moderate mass flow reductions of up to 25% are achieved, caused by discontinuous flow turning around the Coandă surface. Unsteady force and pressure measurements characterize the dynamic lift response to the actuation. The total onset time of the lift response, normalized by the ratio of chord length to freestream velocity, is for actuator activation and for actuator deactivation. These values are below typical onset times for gust loads (). The present results show a high-lift reduction potential and fast dynamic behavior of the Coandă-type flow actuator, making it a viable candidate for fast and effective load alleviation.

Unsteady Aeroelastic Analysis of the Semi Aeroelastic Hinge Including Local Geometric Nonlinearities

A recent consideration in aircraft design is the use of folding wing tips, with the aim of enabling higher aspect ratio wings with less induced drag but also meeting airport gate limitations. Of particular interest is the concept of using in flight free-floating wing tips in order to reduce aircraft gust and maneuver loads. This study investigates the effects of local geometric nonlinearities on the dynamic stability, postflutter behavior and gust response of floating wing tips. A multibody formulation is introduced to account for finite rotations of rigid folding wing tips attached through hinges on a flexible airframe structure including also aerodynamic follower forces for the folding wing tip components. It is found that the postflutter behavior is characterized by super-critical limit cycle oscillations and no subcritical instability was observed for the analyzed cases. Moreover, the wing tip gust response can vary significantly when geometric nonlinearities are accounted for whereas a small impact was observed on the main airframe structure.