Hinge Moment Research Articles

Investigation on bending collapse behavior of similar and dissimilar Aluminum-Steel hat-section sandwich beams jointed by structural adhesive: An experimental, numerical and theoretical study

A significant feature about the application of adhesive joints in various structures is their ability to join dissimilar materials, thus effectively maintaining the structural integrity regardless varying loads. In this paper, the bending collapse behavior of adhesively bonded hat section sandwich beam (HSSB) structures with using of aluminum and steel were investigated. Four configurations of HSSB with different materials for the upper hat beam and lower plate were subjected to quasi-static three-point bending tests as part of the experimental investigation. In comparing the specific absorbed energy (SEA) of the beams, a sample of a hat section sandwich beam with a steel upper hat and an aluminum lower plate had a higher SEA value despite the load-displacement and bending collapse behavior similar to the hat section sandwich beam with a steel upper hat and a steel lower plate. subsequently, a finite element model was developed for bending collapse analysis and a theoretical prediction based on the Kecman model was presented to investigate the hinge moment on HSSB. Based on this, the findings of the numerical study and the predicted theoretical model were in good agreement with the experimental results. The effect of adhesive thickness, the identical mass, and the thickness of the upper hat and lower plate were examined for the purpose of performing an in-depth analysis of the influencing parameters in HSSB structures.

A Hinge Moment Alleviation Control Strategy for Morphing Tail Aircraft Based on a Data-Driven Method

Morphing airplane technology is currently a focal point of research. For morphing airplanes, besides effective morphing strategies and control schemes, the hinge moment at the root of the vertical tail during morphing is a critical factor influencing flight safety. To prevent failure in tail morphing due to excessive hinge moments, this paper analyzes the hinge moment characteristics of the variable vertical tail structure in high-speed flight, based on a flying wing model from the China Aerodynamics Research and Development Center. The proposed adaptive morphing tail hinge moment reduction (AMTHR) method is model-free, utilizing real-time data to dynamically adjust the rudder and reduce hinge moments without requiring prior knowledge of system dynamics. This method utilizes the concept of extremum-seeking control by introducing periodic perturbations to the system and adjusting the control input based on their impact on the output. This approach drives the output toward an extremum point, enabling real-time reduction of the vertical tail hinge moment. Finally, the simulation analysis is carried out under the conditions of no wind and gust disturbance, and the effect of this method on the load reduction of the tail hinge moment is verified.

Prediction model of aircraft hinge moment: Compressed sensing based on proper orthogonal decomposition

By hinge moment, we mean the aerodynamic torque exerted on the rudder shaft by the airflow passing through the aircraft control surface, with obtaining high-precision results often relying on wind tunnel tests. Due to the complex aerodynamic balance insulation and installation errors that must be considered in cryogenic wind tunnels, the main method for calculating hinge moments is to directly integrate surface pressure distribution information. However, it is usually difficult to arrange enough pressure taps, resulting in the accuracy failing to meet expectations. Combining the sparse wind tunnel test data and low-precision computational fluid dynamics results, this paper introduces the compressed sensing based on proper orthogonal decomposition (CS-POD) method and presents the sub-Ma model and the full-Ma model for predicting hinge moments. The number of sensors and sensor positions are determined based on the sparsity of the numerical simulations and basis functions. Then, the CS algorithm solves the basis coefficients. Finally, the hinge moments are obtained by integrating the reconstruction pressure distribution which is calculated by linearly combining the basis functions and basis coefficients. The result shows that the full-Ma model exhibits higher prediction accuracy with approximately five sensors under subsonic and transonic cases, reducing the relative error of the sub-Ma model by 2–10 times, even at high angles of attack. The mean reconstruction accuracy for the hinge moments is 97.6%, and for the normal forces, it is 94.3%. Therefore, adding relevant terms when the number of samples is small can effectively improve modeling accuracy.

Aerodynamic Hinge Moment Characteristics of Pitch-Regulated Mechanism for Mars Rotorcraft: Investigation and Experiments

Aerodynamic Hinge Moment Characteristics of Pitch-Regulated Mechanism for Mars Rotorcraft: Investigation and Experiments

Enhancing UAV elevator actuator model using multibody dynamics simulation

This paper proposes an improved actuator system model for UAV elevators using multibody dynamics simulation. The multibody dynamics simulation employs the Simscape Multibody, module in MATLAB coupled with Simulink to model the servo and hinge moment calculation. The actuator system comprises an electrical servo and mechanical components, including arms, push rods, horns, and the elevator. The electrical servo is modeled using a PID controller and a simplified motor model. The multibody dynamics simulation is employed to capture the dynamics of the mechanical components, coupled with the electrical servo through torque delivery to the mechanical components. The simulation is applied to the elevator of a medium altitude long endurance (MALE) UAV with a Maximum Take Off Weight of 1300 Kg. Generating these quantities provide a benefit in capturing the operational envelope of the servo to be compared to its limitations. Given the features of this simulation, it is proposed to extend the research by integrating this method with flight dynamics simulation.

Effect of facet gap on heliostat wind loading

Multi-facet heliostats are more versatile and advantageous in design compared to single facet heliostats, while generating new design challenges. Multi-facet designs can reduce manufacturing costs, allow for decentralisation of power supply, and bring the option of face down stow. Addition of a gap between two facets, changes the flow structure around the heliostat. The effect of gap between two heliostat facets is therefore investigated on wind loads through experimental investigation in a large wind tunnel. A heliostat model was placed in a simulated atmospheric boundary layer at The University of Adelaide wind tunnel. The investigation covered gap ratios between two panels (gap spacing/facet width) from 0 to 2, and panel aspect ratios (facet width/chord length) of 0.35 to 0.65, forming a complete heliostat aspect ratio of 0.7 to 1.3 with no gap. Wind load coefficients were determined using load cell measurements at the base, and application of differential pressure sensors on the surface, of the heliostat model. Findings indicate the presence of a gap, in general, increases the wind load on the heliostat, with facet aspect ratio exacerbating the results. The maximum load cases shift further into the operating range of heliostats where a change in centre of pressure, due to a change in gap ratio, shifts maximum coefficients from an elevation angle of 30° to 45° for lift force and hinge moment, and 90° to 75° for azimuthal moment. Increasing loads on elevation and azimuth drives potentially increases tracking deviations and beam misalignment, reducing concentrated solar power plant performance.

Extreme Value Analysis for Peak Heliostat Wind Load Predictions

This study investigated high-frequency load fluctuations on a rectangular heliostat model in a boundary layer wind tunnel experiment. Statistical methods for peak wind load predictions were compared with load distributions through load cell and surface pressure measurements on the heliostat model placed in two simulated atmospheric boundary layers (ABLs) representing a flat desert terrain and open country terrain. It was found that statistical peak predictions based on horizontal wind gust velocity amplitude underestimates the load coefficients in stow and at oblique azimuth angles for maximum azimuth moments. In addition, it was shown that the distributions of the centre of pressure position and the hinge moment in stow position are non-Gaussian. It is therefore recommended that the frequency and amplitude of the vertical wind velocity component must be considered for peak load predictions of hinge moments and azimuth moments.

Research of Unsteady Aerodynamic Characteristics of Electrically Controlled Rotor Airfoils with Trailing-Edge Flaps

An electrically controlled rotor (ECR), also known as a swashplateless rotor, eliminates the swashplate system to implement the primary control via the trailing-edge flaps (TEFs), which can result in enhancements in rotor performance, as well as substantial reductions in weight, drag, and cost. In this paper, the unsteady aerodynamic characteristics of the airfoil with TEF of a sample ECR under unsteady freestream condition are investigated. The CFD results are obtained with sliding and overset grid techniques that simulate the airfoil pitching and flap deflection. Comparative analysis of the aerodynamic characteristics under steady and unsteady freestream conditions at different advance ratios is conducted. At various advance ratios, the lift and drag coefficients are higher at a small angle of attack under unsteady freestream condition; however, it is the opposite at a large angle of attack. The peak values of the lift and drag coefficients show an increased trend with the increase in the advance ratio. On the contrary, the pitch moment and flap hinge moment coefficients demonstrate minor variation under unsteady freestream condition. Furthermore, the aerodynamic characteristics of airfoils become more unsteady with variation in the freestream. Therefore, the lift and drag coefficients of the ECR airfoil with TEF show significant differences between steady and unsteady freestream conditions; however, the pitch moment and the flap hinge moment coefficients show little difference.

Heliostat Wind Load Field Measurements at the University of Adelaide Atmospheric Boundary Layer Research Facility (ABLRF)

The University of Adelaide has recently commissioned a facility dedicated to investigating the atmospheric boundary layer (ABL) for the analysis of wind loads on full-scale heliostats. Wind tunnel testing is an affordable way to analyse loads on a scaled structure before committing to a full-scale design. Scale testing however has its challenges as most cases in literature fail to correctly scale the ABL when scaling a model due to the differences between the ratio of the heliostat chord to the boundary layer depth in a wind tunnel and ABL. There is a lack of direct comparison between wind tunnel and full-scale heliostat wind loads. The Atmospheric Boundary Layer Research Facility (ABLRF) consists of arrays of ultrasonic anemometers and a 1.5 aspect ratio heliostat, mounted on a 6-axis load cell, for the comparison of loads measured in the wind tunnel with a full-scale model. Preliminary results categorise the site to have a roughness of 0.01 m to 0.03 m indicating open country farmland, when compared to standards. Comparison between coefficients of lift force, drag force, and hinge moment on the heliostat model at a single elevation angle at the ABLRF and wind tunnel models in literature verify the commissioning of the site, allowing for further in-depth analysis of wind load coefficients at varying elevation and azimuth angles.

Numerical study on aerodynamic characteristics of the grid fins with different grid patterns

Grid fins are unconventional control surfaces configured by a grid of cells within an outer frame; this grid acts as multiple lifting surfaces, and the pattern of the cells strongly affects the grid fin's performance. In this study, three-dimensional simulations were performed using a computational fluid dynamic approach with a k–ω shear stress transport turbulence model to investigate the effect of grid patterns on fin aerodynamic characteristics. Three fin models were designed, including square, tri, and hexa grid patterns, to be more independently comparative; other parameters consisting of the outer frame's dimensions, internal web's thickness, chord length, dragging area, and grid cell area were kept similar between the models. The Mach numbers 0.7, 1.2, and 2.5 corresponding to subsonic, transonic, and supersonic regimes were investigated for varying angles of attack from −5° to 15° for each fin model. The aerodynamic efficiency of grid fins is appreciably improved by increasing the normal force coefficient (CN) while reducing the area force coefficient (CA) and hinge moment coefficient (CHM). The results underscore the hexa grid pattern's superiority, with CN increasing by 2.2% and CA decreasing by 1.92% compared to the square model (the original model) at Mach number 0.7. Especially, the hexa model exhibited a substantial decrease in CHM, with the largest difference being 56.8% compared to the square model at Mach 2.5.

Estimation of Load on Control Lines of Ram Air Parachute Designed for Precise Delivery Using 9-Dof Model

Controlled Aerial Delivery System (CADS) consists of Ram Air Parachute (RAP) Airborne Guidance Unit (AGU) and useful payload. AGU pulls the control lines attached to the control surfaces of RAP and provides required maneuverability to the system. Estimation of load on control lines is essential for design and development of Airborne Guidance Unit (AGU). This paper presents the mathematical methods of the estimation of load on control lines the results of which are validated by experiments. The mathematical estimate of the load on control lines was carried out by developing a model similar to the aircraft control surface hinge moment estimation. The mathematical model was incorporated in 9DOF model to arrive at real time load estimation on control lines during inflation and subsequent steady flight condition. Experimental results are obtained during flight trials by using load cells on control lines and current sensor on power line of actuators. The mathematical results are in good agreement with the experimental results. The load on control lines are within 6 to 8% of the suspended load.

Flight-Test Determination of Longitudinal Stability Using System Identification

In the present work, system identification methods were applied to the flight data from dynamic maneuvers in order to obtain an aerodynamic model of the longitudinal dynamics of a propeller-driven aircraft, including a model for the elevator hinge moment. As a result, the model allowed numerical computation of elevator deflections and stick forces required to trim the aircraft at different flight conditions, including variations of the center-of-gravity location. From these results, the stick-fixed and stick-free neutral points were computed. When compared to a conventional flight-test technique, the new method provided similar mean values for the neutral point location and reduced the associated uncertainty by about 50%. The data collected for system identification required only a fraction of the flight time used in conventional techniques. Moreover, the aerodynamic modeling revealed nonstandard dependencies, such as the linear influence of the airspeed and thrust coefficients on the aerodynamic coefficients. Therefore, propeller slipstream effects were automatically embedded into the neutral point results obtained.

Influence of the hinge moment and the telescoping of the lower limbs on jumping height of single-leg running jump in basketball

Influence of the hinge moment and the telescoping of the lower limbs on jumping height of single-leg running jump in basketball

Observer-Based Backstepping Adaptive Force Control of Electro-Mechanical Actuator with Improved LuGre Friction Model

A dynamic load simulator, which can reproduce on-ground aerodynamic hinge moment of control surface, is an essential rig for the performance and stability test of an aircraft actuation system. In this paper, an observer-based backstepping adaptive control (OBAC) strategy with an improved friction model is proposed to deal with the force tracking problem of an electro-mechanical actuator under the influence of nonlinear friction and lumped disturbances. First, The LuGre friction model is improved by introducing the load effect of electrical linear load simulator (ELLS), and both dynamic and static parameters are identified with experimental data. Then, the ELLS system is divided into a loading subsystem and actuation subsystem for backstepping controller deign. The estimation of the position disturbance is obtained using an extended state observer and used for feedforward compensation for the loading subsystem. To reject the disturbance of friction parameter uncertainties for actuation subsystem, a friction scale factor with a reasonable adaptive updating law is introduced during the friction compensation process. Finally, the stability of the whole closed-loop system is demonstrated using a Lyapunov-based method, and experiments are performed to validate the effectiveness of the developed algorithm.

Computational fluid dynamics study of the rear geometry influence on aerodynamic load coefficients of heliostats in stow position

The study of heliostat load coefficients in stow position has become more extensive because of the importance of this scenario in the design process and cost estimation of an entire Concentrating Solar Thermal Power Tower plant. When in stow position, in the presence of high-velocity wind flow, heliostats behavior can be compared to an airfoil; as the air flows at different velocities over the upper and lower sections, significant pressure imbalances need to be considered. This paper presents a Large Eddy Simulation study of how four different geometries in the rear area of a squared heliostat at a horizontal inclination affect the magnitudes of the hinge moment (CMHy), drag (CFx), lift (CFz) and overturning (CMy) coefficients. Results show maximum differences of more than 128% for peak CFx when comparing type A and B geometries at the same 45° wind orientations. Also, for types A and B at β = 180°, peak CMy and CMHy presented differences of 47% and 36%, respectively, between heliostat models. Also, considering force's change in direction (i.e. lift vs downforce), a 219% maximum difference was observed for peak CFz between type A and type D models at β = 45°. The results show how the differences in shapes and size proportions between the front and rear heliostat sections affect the stow position's aerodynamic load coefficients (ALC). It was also observed that the ratio of the chord length to the maximum vertical distance from the rear mirror surface to the farthest point of structure geometry of the heliostat is directly correlated with ALC variations for analyzed heliostats.

Impact of the aileron gap sealing on the aileron effectiveness

PurposeThe purpose of the presented aileron modification analysis is the improvement of the flight handling by eliminating adverse phenomena in the aileron area, such as aileron shaking movements, without the risk of deterioration of flow characteristics during manoeuvres. It was also crucial to reduce aileron forces acting on the control stick.Design/methodology/approachNumerical CFD analysis of the aileron system with modifications of sealing in the aileron gap area were performed. The effect of the caulking strip at the upper surface of the aileron gap was determined, as well as caulking at the entrance to the aileron gap on the bottom surface. A solution has also been proposed, consisting of completely closing the aileron gap by using a diaphragm. The three-dimensional flow analysis was carried out, allowing localization of the flow disturbances in the aileron gap at cruising speed. The result of the analysis are the aerodynamic and the hinge moment coefficients determining forces on the control stick, depending on the type of seals.FindingsIt has been shown that the use of subsequent sealing means has a direct impact on the hinge moment value. The results of the CFD analysis showed that the more closed aileron gap is, the higher aileron forces are generated on the control stick. Completely closing the flow in the aileron gap changes the character of the force generated on the control stick.Practical implicationsThrough CFD analyses of the aileron gap sealing in the PZL-130 Orlik aircraft, the impact of successive aileron gap sealing on the aileron efficiency was determined. It has been shown that simple change of the aileron gap size by the slat sealing can significantly affect the value of the forces generated on the control stick.Originality/valueThe research using CFD methods allowed to verify the impact of the particular type of aileron gap sealing on the hinge moment value and thus to determine proper sealing configuration for the PZL-130 Orlik aircraft at low computational cost.

Оценка параметров управляемого движения спускаемого аппарата с надувным тормозным устройством путем отклонения элементов конструкции

Currently, there is a constant development and improvement of space technology. One of the promising areas of space research is the delivery of a container with a payload both in an interplanetary mission and from the Earth’s orbit. The cargo could include soil samples mined on the planets of the solar system, or the results of experiments in orbit. To deliver the cargo, it is necessary to ensure controlled motion of the descent vehicle in the planet’s atmosphere. The paper considers the possibility of using inflatable devices for braking and controlling a spacecraft in the planet’s atmosphere, as well as for perturbational control. A mathematical model of the descent vehicle motion was developed, which was used to analyze the balancing angles of attack that occur when the payload is deflected. The results obtained made it possible to determine the overloads acting on the descent vehicle during controlled motion, determine the hinge moments, and calculate the parameters of the descent vehicle trajectory for various atmospheric entry conditions. Findings of the research show the possibility of using such descent vehicles and allow us to formulate design requirements for the mechanical part of the structure.

Comparative Study of Lifting Surface and CFD Methods in the Simulation of Morphing Process of Folding Wing

A folding wing aircraft can autonomously change its configuration in flight to respond to different flying environments. Hinge moment during morphing process is an important basis for driving mechanism design and structural strength check, and its calculation depends on the simulation of the morphing process. Existing studies mostly use the simplified lifting surface method for aerodynamic modeling. In this paper, the CFD-based simulation method is studied, and the results are compared with those by the lifting surface method. First, the unsteady aerodynamic modeling method of the folding wing based on the CFD method is studied, an effective description method is given to define the wall mesh motion caused by wing folding and aileron deflection, and an unfolding–refolding strategy is proposed to improve the quality of the internal mesh in the case of large folding angles. Then, the CFD aerodynamic model is coupled through a developed coupling calculation program with the flexible multibody structure model for the simulation of a flight-morphing process. The comparison with the results of the lifting surface method shows that hinge moments obtained according to the two aerodynamic models are markedly different. Analysis of the difference shows that airfoil thickness considerably affects aerodynamic loading distributions and hinge moments of folding wing aircraft. The lifting surface method ignores airfoil thickness, which will cause large simulation errors in hinge moments.

Elevator Tab Model and Aerodynamic Hinge Moment Variation Effect on Flutter Analysis of a Small Commuter Aircraft

Investigation of flutter instability analysis is an important process during design phase and certification of an aircraft. Without proper analysis, the effect of modification at later stage of design process could come at high cost and risk. The present work examines detailed elevator tab model which could result in a more critical tab flutter speed. The present investigation of aerodynamic hinge moment variation effect on flutter analysis also shows a significant decrease in the flutter speeds. Three configurations involving elevator trim tab and balance tab fixed/free mechanism are investigated. Investigation of flutter sensitivity to tab rotational frequency and tab bending/torsion frequency was performed to determine the minimum tab rotational frequency requirement, so that it can be used as reference in design process. Variation of Mach number (for Configuration 1) and elevator aerodynamic hinge moment (all configurations) are performed. It is found that the aerodynamic variation takes effect when the mode involves both trim and balance tabs and tab elastic modes (e.g., tab torsion). It is found also that the most critical flutter case is for Configuration 3 where the trim and balance tabs are not fixed (independent). The result suggests that control surface tabs need to be modelled from the early design process, so that if any critical flutter appears and needs to be responded by a change in design, it can be implemented earlier and therefore avoid any costly risk.

Influence of Interference of the Airscrew and the High-Aspect-Ratio Wing on the Hinge Moment Deflect Control Surfaces of the Wing

Influence of Interference of the Airscrew and the High-Aspect-Ratio Wing on the Hinge Moment Deflect Control Surfaces of the Wing