Control Surface Hinge Moments Research Articles

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.

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.

Comparison of flutter calculation methods based on ground vibration test result

PurposeA low-cost but credible method of low-subsonic flutter analysis based on ground vibration test (GVT) results is presented. The purpose of this paper is a comparison of two methods of immediate flutter problem solution: JG2 – low cost software based on the strip theory in aerodynamics (STA) and V-g method of the flutter problem solution and ZAERO I commercial software with doublet lattice method (DLM) aerodynamic model and G method of the flutter problem solution. In both cases, the same sets of measured normal modes are used.Design/methodology/approachBefore flutter computation, resonant modes are supplied by some non-measurable but existing modes and processed using the author’s own procedure. For flutter computation, the modes are normalized using the aircraft mass model. The measured mode orthogonalization is possible. The flutter calculation made by means of both methods are performed for the MP-02 Czajka UL aircraft and the Virus SW 121 aircraft of LSA category.FindingsIn most cases, both compared flutter computation results are similar, especially in the case of high aspect wing flutter. The Czajka T-tail flutter analysis using JG2 software is more conservative than the one made by ZAERO, especially in the case of rudder flutter. The differences can be reduced if the proposed rudder effectiveness coefficients are introduced.Practical implicationsThe low-cost methods are attractive for flutter analysis of UL and light aircraft. The paper presents the scope of the low-cost JG2 method and its limitations.Originality/valueIn comparison with other works, the measured generalized masses are not used. Additionally, the rudder effectiveness reduction was implemented into the STA. However, Niedbal (1997) introduced corrections of control surface hinge moments, but the present work contains results in comparison with the outcome obtained by means of the more credible software.

Nature of Freeplay-Induced Aeroelastic Oscillations

Time- and frequency-domain analyses are used to revisit a well studied case of an aeroelastic wing with control-surface freeplay nonlinearity. The dynamic response in general is presented as dependent on the stability of two linear systems, with zero and nominal control-surface rotational stiffness, and values of two angles, freeplay and preload. The principal new findings include the following. A physical explanation is given to the bifurcation airspeeds and frequencies of the oscillations for the system with both zero and nonzero aerodynamic preload of the control surface. For the latter and more complex case, it is shown and explained why a certain amount of the preload is not only always destabilizing but may also be the reason for the otherwise unexpected higher-frequency oscillations. A conclusion thus is drawn that the aircraft control-surface freeplay analyses based only on zero and high quenching aerodynamic preloads are nonconservative. All results are compared with the experimental and other published theoretical analyses for the considered model. Finally, a suggestion is given for future experiments with a goal to observe the response changes with the preload and to determine the stabilizing control-surface hinge moment.

Force control system design for aerodynamic load simulator

Abstract A dynamic load simulator which can reproduce on-ground the aerodynamic hinge moment of control surface is an essential rig for the performance and stability test of an aircraft actuation system. The hinge moment varies widely over the flight envelope depending on the specific flight condition and maneuvering status. To replicate the wide spectrum of this hinge moment variation within some accuracy bounds, a force controller is designed based on the quantitative feedback theory (QFT). A dynamic model of load actuation system is derived, and verified through the experiment. The efficacy of QFT force controller is verified by a numerical simulation, in which combined aircraft dynamics, flight control law and hydraulic actuation system dynamics of aircraft control surface are considered.

Dynamic Characteristic Analysis of Aerodynamic Load Simulator English

A dynamic load simulator(DLS) which can reproduce on-ground the aerodynamic hinge moment of control surface is an essential rig for the performance and stability test of aircraft actuation system. By setting up load actuator as counter acting with the control surface driving actuator and designing an appropriate force control system for load actuator, DLS can be mechanized. Obtaining an accurate mathematical model for the DLS is the first step to successfully design an aerodynamic load replicati on system. Two theoretical models are presented and tested for their validities with the experimental results, which turns out to be not successful. An alternative way of using system identification approaches in investigated to develop a good nominal model for DLS dynamics, and suitable uncertainty bounds for this nominal model are proposed with the consideration of experimental results.

Dynamic characteristic analysis and force loop design for the aerodynamic load simulator

A dynamic load simulator which can reproduce, on-ground, the aerodynamic hinge moment of control surface, is an essential rig for performance and stability test of aircraft actuation system. The hinge moment varies widely in the flight envelope depending on specific flight condition and maneuvering status. To replicate the wide spectrum of this hinge moment variation within some accuracy bounds, a force controller is designed based on the Quantitative Feedback theory (QFT). A dynamic model of load actuation system is derived, and compared with the experimental results. Through this comparison, a nominal plant model of load actuation system with some uncertainty bounds is developed. The efficacy of QFT force controller is verified by numerical simulation, in which hydraulic actuation system dynamics of aircraft control surface is considered.

Benchmark Active Controls Technology Model Aerodynamic Data

The Benchmark Active Controls Technology (BACT) model is a part of the Benchmark Models Program (BMP). The BMP is a NASA Langley Research Center program that includes a series of models which were used to study different aeroelastic phenomena and to validate computational fluid dynamics codes. The primary objective of BACT testing was to obtain steady and unsteady loads, accelerations, and aerodynamic pressures due to control surface activity in order to calibrate unsteady CFD codes and active control design tools. Three wind-tunnel tests in the Transonic Dynamics Tunnel (TDT) have been completed. The first and parts of the second and third tests focused on collecting open-loop data to define the model''s aeroservoelastic characteristics, including the flutter boundary across the Mach range. It is this data that is being presented in this paper. An extensive database of over 3000 data sets was obtained. This database includes steady and unsteady control surface effectiveness data, including pressure distributions, control surface hinge moments, and overall model loads due to deflections of a trailing edge control surface and upper and lower surface spoilers.

Variable Stiffness Spar Approach for Aircraft Maneuver Enhancement Using ASTROS

An innovative variable stiffness spar (VSS) approach is studied for improving aircraft roll performance. In this concept some of the existing wing spars are replaced by the adaptive-structure VSS to control the stiffness as a function of Mach number and altitude. The VSS stiffness scheduling is designed to maximize the roll rate while satisfying flutter, control surface hinge moment, and maximum deflection constraints. The VSS mechanism consists of segmented spar having articulated joints at the connections with wing ribs and an electrical actuator capable of rotating the spar. The wing stiffness provided by the spar varies sinusoidally as a function of the rotation angle. The objective of the present study is to explore when and how to best apply this concept and assess its payoffs in terms of performance gains. The F/A-18 pre-roll-modification aircraft was selected as the baseline aircraft for its low torsional wing stiffness and available flight data. The multidisciplinary design optimization software ASTROS * was used tier performing the analyses in the Mach number range of M = 0.8-1.2 at altitudes up to 35,000 ft (40,668 m). Results show that VSS can amplify the aeroelastic forces and significantly enhance roll performance of aireraft.

Static aeroelastic tailoring for oblique wing lateral trim

Due to its asymmetrical configuration, an oblique wing aircraft may exhibit a roll-trim imbalance. To correct this imbalance, control surfaces, such as ailerons, might be deflected or built-in twists might be added. This paper explores an alternative method of achieving oblique wing roll trim. This is the concept of applying composite tailoring to the oblique wing for lateral trim. The practical use of this concept to trim an oblique wing is demonstrated through analysis of a realistic wing by a static aeroelastic computational procedure. The computational method includes the full-potential transonic aerodynamic code FLO22 and a Ritz structural plate program that is used to model the stiffness created by symmetrical, but unbalanced, advanced composite wing skins. Analysis results indicate that asymmetric composite tailoring reduces the aileron deflection needed for roll equilibrium of the oblique wing. Furthermore, the use of aeroelastic tailoring for lateral trim can reduce control surface hinge moments and drag, resulting in performance benefits when compared to an untailored wing. At the same time, however, wing-skin stresses are greater when tailoring is used, leading to potential design tradeoffs. As such, an integrated design approach would be required to evaluate the true impact of aeroelastic tailoring on the overall performance of an oblique wing.

Integrated design of structures

Abstract : It is shown that for highly sophisticated, naturally unstable airplanes flying supersonically a joint strategy to lay out the flight control system while minimizing design loads must be adopted. The selection of control surface geometry must be performed utilizing all possibilities from overall structural optimization including aeroelastic tailoring for primary carbon fibre structures. In the proposed design philosophy the behavior of the elastic airplane structure must be introduced and optimized in the very early design stage. It is shown in the paper that the required control surface hinge moments can be reduced by optimizing mass penalties and efficiencies. Minimizing installed hydraulic power supply has also a beneficial effect on engine performance at low speed, high altitudes.