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1 – 10 of 489Xiaowei Shao, Mingxuan Song, Jihe Wang, Dexin Zhang and Junli Chen
The purpose of this paper is to present a method to achieve small satellite formation keeping operations by using the differential lift and drag to control the drift caused by J2…
Abstract
Purpose
The purpose of this paper is to present a method to achieve small satellite formation keeping operations by using the differential lift and drag to control the drift caused by J2 perturbation in circular or near-circular low earth orbits (LEOs).
Design/methodology/approach
Each spacecraft is equipped with five large flat plates, which can be controlled to generate differential accelerations. The aerodynamic lift and drag acting on a flat plate is calculated by the kinetic theory. To maintain the formation within tracking error bounds in the presence of J2 perturbation, a nonlinear Lyapunov-based feedback control law is designed.
Findings
Simulation results demonstrate that the proposed method is efficient for the satellite formation keeping and better accuracy advantage in comparison with classical approaches via the fixed maximum differential aerodynamic acceleration.
Research limitations/implications
Because the aerodynamic force will reduce drastically as the orbital altitude increases, the formation keeping control strategy for small satellites presented in this paper should be limited to the scenarios when satellites are in LEO.
Practical implications
The formation keeping control method in this paper can be applied to solve satellite formation keeping problem for small satellites in LEO.
Originality/value
This paper proposes a Lyapunov control strategy for satellite formation keeping considering both lift and drag forces, and simulation results show better performance with high accuracy under J2 perturbation.
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The purpose of this paper is to develop a physics-based model for UAV propellers that is capable of predicting all aerodynamic forces and moments in any general forward flight…
Abstract
Purpose
The purpose of this paper is to develop a physics-based model for UAV propellers that is capable of predicting all aerodynamic forces and moments in any general forward flight condition such as no flow, pure axial flow and pure side flow etc.
Design/methodology/approach
The methodology adopted in this paper is the widely used Blade Element Momentum Theory (BEMT) for propeller model development. The difficulty arising from the variation of induced flow with blade’s angular position is overcome by supplementing the BEMT with the inflow model developed by Pitt and Peters. More so, high angle of attack aerodynamics is embedded in the simulation as it is likely for the blades to stall in general forward flight, for example during extreme aerobatics/maneuvers.
Findings
The validity of the model is demonstrated via comparison with experiments as well as with other existing models. It is found that one of the secondary forces is negligible while the other is one order of magnitude less than the primary static thrust, and as such may be neglected depending on the level of accuracy required. On the other hand, both secondary moments must be considered as they are of similar order of magnitude as the primary static torque.
Research limitations/implications
The paper does not consider the swirl component of the induced flow under the assumption that it is negligible compared to the axial component.
Originality/value
This paper fulfills the identified need of a propeller model for general forward flight conditions, and aims to fill this void in the existing literature pertaining to UAVs.
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Tandralee Chetia, Dhayalan Rajaram and Kumaran G. Sreejalekshmi
Flapping-wing vehicles show various advantages as compared to fixed wing vehicles, making flapping-wing vehicles' study necessary in the current scenario. The present study aims…
Abstract
Purpose
Flapping-wing vehicles show various advantages as compared to fixed wing vehicles, making flapping-wing vehicles' study necessary in the current scenario. The present study aims to provide guidelines for fixing geometric parameters for an initial engineering design by a simple aerodynamic and flight dynamic parametric study.
Design/methodology/approach
A mathematical analysis was performed to understand the aerodynamics and flight dynamics of the micro-air vehicle (MAV). Only the forces due to the flapping wing were considered. The flapping motion was considered to be a combination of the pitching and plunging motion. The geometric parameters of the flapping wing were varied and the aerodynamic forces and power were observed. Attempts were then made to understand the flight stability envelope of the MAV in a forward horizontal motion in the vertical plane with similar parametric studies as those conducted in the case of aerodynamics.
Findings
From the aerodynamic study, insights were obtained regarding the interaction of design parameters with the aerodynamics and feasible ranges of values for the parameters were identified. The flapping wing was found to have neutral static stability. The flight dynamic analysis revealed the presence of an unstable oscillatory mode, a stable fast subsidence mode and a neutral mode, in the forward flight of the MAV. The presence of unstable modes highlighted the need for active control to restore the MAV to equilibrium from its unstable state.
Research limitations/implications
The study does not take into account the effects of control surfaces and tail on the aerodynamics and flight dynamics of the MAV. There is also a need to validate the results obtained in the study through experimental means which shall be taken up in the future.
Practical implications
The parametric study helps us to understand the extent of the impact of the design parameters on the aerodynamics and stability of the MAV. The analysis of both aerodynamics and dynamic stability provides a holistic picture for the initial design. The study incorporates complex mathematical equations and simplifies such to understand the aerodynamics and flight stability of the MAV from an engineering perspective.
Originality/value
The study adds to already existing knowledge on the design procedures of a flapping wing.
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Farid Shahmiri and Fariborz Saghafi
This paper aims to focus on mathematical model development issues, necessary for a better prediction of dynamic responses of articulated rotor helicopters.
Abstract
Purpose
This paper aims to focus on mathematical model development issues, necessary for a better prediction of dynamic responses of articulated rotor helicopters.
Design/methodology/approach
The methodology is laid out based on model development for an articulated main rotor, using the theories of aeroelastisity, finite element and state‐space represented indicial‐based unsteady aerodynamics. The model is represented by a set of nonlinear partial differential equations for the main rotor within a state‐space representation for all other parts of helicopter dynamics. The coupled rotor and fuselage formulation enforces the use of numerical solution techniques for trim and linearization calculations. The mathematical model validation is carried out by comparing model responses against flight test data for a known configuration.
Findings
Improvements in dynamic prediction of both on‐axis and cross‐coupled responses of helicopter to pilot inputs are observed.
Research limitations/implications
Further work is required for investigation of the unsteady aerodynamics, a state‐space representation, within various compatible dynamic inflow models to describe the helicopter response characteristics.
Practical implications
The results of this work support ongoing research on the development of highly accurate helicopter flight dynamic mathematical models. These models are used as engineering tools both for designing new aerial products such as modernized agile helicopters and optimization of the old version products at minimum time and expense.
Originality/value
Provides further information on the mathematical model development problems associated with advanced helicopter flight dynamics research.
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Rongqi Shi and Weiyu Wan
This paper aims to clarify the flight dynamics characteristics and improve the flight performance for large-scale morphing aircrafts. With specific focus on the effects of…
Abstract
Purpose
This paper aims to clarify the flight dynamics characteristics and improve the flight performance for large-scale morphing aircrafts. With specific focus on the effects of morphing on mass distribution, aerodynamics and flight stability, the study aims to develop the dynamic model, outline the morphing strategies design and evaluate the flight stability in transient stage of morphing.
Design/methodology/approach
The mode of relaxing the rigidity condition was opted, which introduced the functions of position of center of mass and moments of inertia with respect to the morphing parameters, and yielded a parameter-dependent flight dynamics model. The morphing strategies were designed by optimizing the morphing parameters with the corresponding performance metric of each mission segment, where the aerodynamics was estimated concurrently by DATCOM. Based on the decoupled and linearized longitudinal parameter-dependent model, the flight stability in transient stage of morphing was evaluated based on Hurwitz rules, with the stability condition proposed.
Findings
The research suggests that the longitudinal flight stability in transient stage of morphing can be evaluated by the relationship of aerodynamic pitching moment derivatives and the effects of morphing on the mass distribution, which results in a constraint on the morphing rate.
Research limitations/implications
The aerodynamics is computed under quasi-steady aerodynamic assumption in low morphing rate and only the longitudinal flight stability is analyzed. Therefore, researchers are encouraged to evaluate the lateral stability and aerodynamics in high morphing rate.
Practical implications
The paper includes implications for the improvement of the flight performance of a multi-mission morphing aircraft and the design of the flight control system.
Originality/value
Methods of dynamic modeling and morphing strategies design are proposed for large-scale morphing aircrafts, and the condition of flight stability in transient stage of morphing is obtained. The results provide reference to research works in the field of dynamics and control of large-scale morphing aircrafts.
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Farid Shahmiri and Fariborz Saghafi
The purpose of this paper is to examine the cross‐coupled responses of a coupled rotor‐fuselage flight dynamic simulation model, including a finite‐state inflow aerodynamics and a…
Abstract
Purpose
The purpose of this paper is to examine the cross‐coupled responses of a coupled rotor‐fuselage flight dynamic simulation model, including a finite‐state inflow aerodynamics and a coupled flap‐lag and torsion flexible blade structure.
Design/methodology/approach
The methodology is laid out based on model development for an articulated main rotor, using the theories of aeroelastisity, finite element and finite‐state inflow formulation. The finite‐state inflow formulation is based on a 3D unsteady Euler‐based concepts presented in the time domain. The most advantages of the model are the capability of modeling dynamic wake effects, tip losses and skewed wake aerodynamics. This is, in fact, a special type of the inflow model relating inflow states, to circulatory blade loadings through a set of first‐order differential equations. A non‐iterative solution of the differential equations has practically altered the model into a simple and direct formulation appending properly to the rest of the helicopter mathematical model. A non‐linear distribution of the induced velocity over the rotor disc is finally obtained by the use of both Legendre polynomials and higher‐harmonic functions. Ultimately, validations of the theoretical results show that the on‐axis response, direct reaction to the pilot input, has a good accuracy both quantitatively and qualitatively against flight test data, and the off‐axis response, cross‐coupled or indirect reaction to the pilot input are improved by this approach of modeling.
Findings
Improvements in dynamic prediction of both trim control settings and dynamic cross‐coupled responses of helicopter to pilot inputs are observed.
Research limitations/implications
Further work is required for investigation of the augmented finite state inflow model, including the wake rotation correction factors to describe helicopter maneuvering flight characteristics.
Practical implications
The results of this work support the future researches on design and development of advanced flight control system, incorporating a high bandwidth with low‐phase delay to control inputs and also high levels of dynamic stability within minimal controls cross coupling.
Originality/value
This paper provides detailed characteristics on the mathematical integration problems associated with the advanced helicopter flight dynamics research.
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A. Arun Kumar, S.R. Viswamurthy and R. Ganguli
This paper aims to validate a comprehensive aeroelastic analysis for a helicopter rotor with the higher harmonic control aeroacoustic rotor test (HART‐II) wind tunnel test data.
Abstract
Purpose
This paper aims to validate a comprehensive aeroelastic analysis for a helicopter rotor with the higher harmonic control aeroacoustic rotor test (HART‐II) wind tunnel test data.
Design/methodology/approach
Aeroelastic analysis of helicopter rotor with elastic blades based on finite element method in space and time and capable of considering higher harmonic control inputs is carried out. Moderate deflection and coriolis nonlinearities are included in the analysis. The rotor aerodynamics are represented using free wake and unsteady aerodynamic models.
Findings
Good correlation between analysis and HART‐II wind tunnel test data is obtained for blade natural frequencies across a range of rotating speeds. The basic physics of the blade mode shapes are also well captured. In particular, the fundamental flap, lag and torsion modes compare very well. The blade response compares well with HART‐II result and other high‐fidelity aeroelastic code predictions for flap and torsion mode. For the lead‐lag response, the present analysis prediction is somewhat better than other aeroelastic analyses.
Research limitations/implications
Predicted blade response trend with higher harmonic pitch control agreed well with the wind tunnel test data, but usually contained a constant offset in the mean values of lead‐lag and elastic torsion response. Improvements in the modeling of the aerodynamic environment around the rotor can help reduce this gap between the experimental and numerical results.
Practical implications
Correlation of predicted aeroelastic response with wind tunnel test data is a vital step towards validating any helicopter aeroelastic analysis. Such efforts lend confidence in using the numerical analysis to understand the actual physical behavior of the helicopter system. Also, validated numerical analyses can take the place of time‐consuming and expensive wind tunnel tests during the initial stage of the design process.
Originality/value
While the basic physics appears to be well captured by the aeroelastic analysis, there is need for improvement in the aerodynamic modeling which appears to be the source of the gap between numerical predictions and HART‐II wind tunnel experiments.
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Fathi Jegarkandi Mohsen, Salezadeh Nobari Ali, Sabzehparvar Mahdi, Haddadpour Hassan and Tavakkoli Farhad
The purpose of this paper is to investigate the aeroelastic behavior of a supersonic flight vehicle flying at moderate angles of attack using global analytic nonlinear aerodynamic…
Abstract
Purpose
The purpose of this paper is to investigate the aeroelastic behavior of a supersonic flight vehicle flying at moderate angles of attack using global analytic nonlinear aerodynamic model.
Design/methodology/approach
Aeroelastic behavior of a supersonic flight vehicle flying at moderate angles of attack is considered, using nonlinear aerodynamics and linear elastodynamics and structural models. Normal force distribution coefficient over the length of the vehicle and pitching moment coefficient are the main aerodynamic parameters used in the aeroelastic modeling. It is very important to have closed form analytical relations for these coefficients in the model. They are generated using global nonlinear multivariate orthogonal modeling functions in this work. Angle of attack and length of the vehicle are selected as independent variables in the first step. Local variation of angle of attack is applied to the analytical model and due to its variation along the body of the vehicle, equations of motion are finalized. Mach number is added to the independent variables to investigate its role on instability of the vehicle and the modified model is compared with the previous one in the next step. Thrust effect on the aeroelastic stability of the vehicle is analyzed at final stage.
Findings
It is shown that for the vehicles having simple configurations and low length to diameter ratios flying at low angles of attack, assuming normal force distribution coefficient linear relative to α is reasonable. It is concluded that vehicle's velocity and thrust has not great effect on its divergence dynamic pressure.
Originality/value
Based on the constructed model, a simulation code is generated to investigate the aeroelastic behavior of the vehicle. The resultant code is verified by investigating the static aeroelastic stability margin of the vehicle presented in the references. Mach number effect on the aeroelastic behavior of the vehicle is considered using modified aerodynamic model and is compared with the results. Data base for identifying aerodynamic coefficients is conducted using CFD code.
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Shawn S. Keshmiri, Edward Lan and Richard Hale
The purpose of this paper is to evaluate the accuracy of linear and quasi-steady aerodynamic models of aircraft aerodynamic models when a small unmanned aerial system flies in the…
Abstract
Purpose
The purpose of this paper is to evaluate the accuracy of linear and quasi-steady aerodynamic models of aircraft aerodynamic models when a small unmanned aerial system flies in the presence of strong wind and gust at a high angle of attack and a high sideslip angle.
Design/methodology/approach
Compatibility analysis were done to improve the quality of recorded flight test data. A robust method called fuzzy logic modeling is used to set up the aerodynamic models. The reduced frequency is used to represent the unsteadiness of the flow field according to Theodorsen’s theory. The work done by the aerodynamic moments on the motions is used as the criteria of stability.
Findings
In portions of flight, aircraft’s stability and control derivatives were unstable and nonlinear functions of airflow angles and angular rates. The roll angle had an important effect on unsteadiness of directional oscillatory damping derivatives. The pilot-induced oscillation and wing rock possibilities were investigated and dismissed so that the lateral directional oscillatory motion was classified as a nonlinear Dutch roll oscillation. Major modeling enhancements or real-time parameter identification are required for the control of a small unmanned aerial system in off-nominal conditions. The robustness tests of all-weather autopilot systems must be done with consideration of sign change.
Originality/value
Oscillatory damping derivatives were reconstructed using flight test data and the inadequacy of engineering level software in predicting this type of instability observed and demonstrated for a flight in the presence of wind shear and external disturbances.
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In this informal symposium, presided over by R. D. Kelly, United Air Lines, after talks, rather than the reading of papers, the pilots concerned assembled on the rostrum and…
Abstract
In this informal symposium, presided over by R. D. Kelly, United Air Lines, after talks, rather than the reading of papers, the pilots concerned assembled on the rostrum and answered questions. They were: