Variation Of Aerodynamic Parameters Research Articles

Research on the aerodynamic characteristics of electrically controlled rotor under Parallel Blade Vortex Interaction using Lattice Boltzmann Method

An electrically controlled rotor (ECR), also known as a swashplateless rotor, employs a trailing edge flap (TEF) system for primary rotor control instead of a swashplate, demonstrating the significant potential in rotor vibration and noise reduction. To investigate the aerodynamic characteristics of the blade flap segment of the ECR under parallel blade vortex interaction, an aerodynamic analysis model based on the lattice Boltzmann method (LBM) is established using the D3Q27 lattice model. The model is validated against experimental data of both the airfoil with trailing edge flap and conventional airfoil under vortex interaction, showing that the LBM can effectively predict variations in aerodynamic loads under both conditions. Based on this model, the effects of different flap deflection angles and miss distances on the aerodynamic characteristics of the ECR under parallel BVI are analyzed. The results indicate that under strong vortex interaction, a region of flow separation forms due to the entrainment effect of the vortex and adverse pressure gradient. The small-scale vortex structure upstream of the flap can also be observed and is believed to contribute to the unsteady flow phenomena such as vortex structure splitting, development, and separation on the upper surface of the flap. The different flap deflection angles mainly affect the scale and type of vortex structures developed on the flap upper surface. As the miss distance increases, the interaction effect is significantly weakened compared to strong vortex interaction. However, as the vortex moves downstream along the airfoil lower surface, it entrains vorticity from the lower surface, ultimately forming a negative pressure region on the lower surface of the flap. The different flap deflection angles will influence the structural characteristics during the downstream motion of the vortex, which changes the size of the negative pressure region, causing differences in the magnitude of the variations in aerodynamic parameters.

Variation in Zero Plane Displacement and Roughness Length for Momentum Revisited

Zero plane displacement height (d0\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$d_0$$\\end{document}) and momentum roughness length (z0m\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$z_{0m}$$\\end{document}), describe the aerodynamic characteristics of a vegetated surface. Usually, d0\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$d_0$$\\end{document} and z0m\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$z_{0m}$$\\end{document} are assumed to be constant functions of the physical characteristics of the surface. Prior evidence collected from the literature and our examination of flux tower data show that d0\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$d_0$$\\end{document} and z0m\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$z_{0m}$$\\end{document} vary in time at sites with tree and shrub canopies, but not grasslands. The conventional explanations of these variations are based on linear functions of wind velocity and friction velocity, with little theoretical basis. This study explains the variation in aerodynamic parameters by matching four analytical canopy velocity models to a logarithmic above-canopy velocity profile at canopy height. d0\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$d_0$$\\end{document} and z0m\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$z_{0m}$$\\end{document} come out as functions of 2 non-dimensional terms, the canopy momentum absorption capacity (parameter) and a (measurable) Péclet number. To test the theories of variation, we analysed the velocity profiles from Ozflux and Ameriflux sites. None of the theories could recreate d0\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$d_0$$\\end{document} and z0m\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$z_{0m}$$\\end{document} at half-hourly intervals. However, the canopy velocity models were able better to recreate the distribution of the variations in d0\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$d_0$$\\end{document} and z0m\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$z_{0m}$$\\end{document}. Additionally, the estimates of canopy momentum absorption capacity varied consistently with phenological changes in the canopies, whereas, the fitting parameters of the linear regression of using wind speed and friction velocity did not exhibit physically interpretable variations. The canopy velocity models may offer better predictions with an accurate estimation of the canopy height, a horizontally homogeneous and rigid canopy, and incorporation of the roughness sublayer.

Multiple‐missile fixed‐time integrated guidance and control design with multi‐stage interconnected observers under impact angle and input saturation constraints

AbstractIn this paper, a novel three‐dimensional fixed‐time integrated guidance and control (IGC) scheme with multi‐stage interconnected observers is proposed for cooperative attacks using multiple missiles against a maneuvering target under impact angle and input saturation constraints. External disturbances, modeling errors, and aerodynamic parameter variations are considered as system uncertainties and a three‐channel fully coupled IGC model for multiple missiles is established. The IGC system is designed optimally based on fixed‐time stability theory, sliding mode control, and the backstepping technique. Three inter‐cascaded fixed‐time disturbance observers based on an improved super‐twisting algorithm are designed to estimate and compensate for system uncertainties. Second‐order command filters are used to constrain virtual control signals, and additional filtering error subsystems are introduced to compensate for the tracking errors of filters. System stability and uniformly ultimately fixed‐time boundedness of all states are proven using the Lyapunov stability theory. Finally, the limits of the acceleration components of the maneuvering target perpendicular to the line of sight direction are derived. The effectiveness of the designed IGC scheme and the ability of multi‐stage interconnected observers to sense disturbances with each other are verified through simulations.

Approach and landing guidance using constrained model predictive static programming

The paper proposes a guidance scheme during the approach and landing phase of an unpowered vehicle to achieve high precision at touchdown by limiting the control inputs within allowable bounds using constrained Model Predictive Static Programming. The advantages of a single flight phase are incorporated into the design since the guess control history is obtained from a single segment based guidance approach. The requirement of ensuring continuity in states and guidance commands are hence avoided leading to minimization of control deviations using a simple cost function. Constrained input commands are used for determining control surface deflection based on dynamic inversion considering the longitudinal pitch rate dynamics. Comparison is carried out between the results obtained for constrained as well as unconstrained schemes for the nominal trajectory. Comparisons of load factor, dynamic pressure and flight path angle for three segment, two segment, single segment guidance strategies have been carried out with MPSP based guidance to evaluate the relative advantages offered by the scheme. Performance analysis of the proposed scheme is done based on numerous simulations with simultaneous variations in initial states, vehicle parameters and aerodynamic parameter variations. Simulation results indicate that the proposed scheme is suitable for landing on runways of shorter lengths due to minimal touchdown errors.

Research on Variable-Swept Hybrid Aerial Underwater Vehicle Plunge-Diving Control Based on Adaptive Dynamic Surface Control

The variation in aerodynamic parameters during the process of a variable sweepback hybrid aerial underwater vehicle (HAUV) affects flight stability. During the air–water trans-media locomotion, there are medium mutations and solid–liquid gas coupling phenomena, resulting in the complex dynamic process of HAUV. To ensure stable control during the trans-media process of a variable sweepback vehicle, this study proposes a neural-network-based adaptive dynamic surface control method for aircraft flight-path angle. This method aims to establish an effective control model for the entire process of air to media transition, in response to the characteristics of uncertainty and external disturbances in the process of variable backsweeping in the air and media transition. By utilizing the multibody dynamics method, the dynamic equations for variable-swept vehicles are established and transformed into a rigorous feedback system with model uncertainty. The adaptive dynamic surface method in this paper introduces a first-order filter, which overcomes the “differential explosion” problem in traditional backstepping control design through differential filtering; the unknown parameters present in the model are estimated online through adaptive laws, and the uncertain parts of the system are overcome through nonlinear damping items. By analyzing Lyapunov stability, the semi-global stability of the required closed-loop system can be obtained, and adjusting the controller parameters can make the tracking error infinitely small. Numerical simulations are conducted to illustrate the tracking control of flight-path angles for different plunge-diving angle rates and strategy of ingress. The results show that HAUV with variable-swept configuration with different strategy has a great effect on the stability of plunge-diving locomotion; the designed controller can effectively track the target trajectory and has a certain degree of robustness and adaptability.

Aerodynamic Parameter Identification of Projectile Based on Improved Extreme Learning Machine and Ensemble Learning Theory

The firing accuracy of the projectile has a positive relation with aerodynamic parameters. Due to the complex dynamic characteristics of projectiles, there is an overfitting risk when a single extreme learning machine (ELM) is used to identify the aerodynamic parameters of the projectile, and the identification results oscillate transonic region. To obtain the aerodynamic parameters of the projectile accurately, an aerodynamic parameter identification model based on ensemble learning theory and ELM optimized by improved particle swarm optimization is proposed. The improved particle swarm optimization algorithm (IPSO) with an adaptive update strategy is used to optimize the weight and threshold of ELM. Combined with the ensemble learning theory, the improved ELM neural network is regarded as a weak learner to generate a strong learner. The structural parameters of the strong learner were continuously optimized through training, and an aerodynamic parameter identification model of projectile based on ensemble learning theory is obtained. The simulation results show that the introduction of the IPSO and ensemble learning theory enables the model to exhibit excellent generalization ability. The proposed identification model can accurately describe the variation of aerodynamic parameters with the Mach number.

Field-to-View Constrained Integrated Guidance and Control for Hypersonic Homing Missiles Intercepting Supersonic Maneuvering Targets

An integrated guidance and control (IGC) scheme considering the field-of-view (FOV) constraint is proposed in this paper for hypersonic skid-to-turn (STT) missiles with a strapdown seeker intercepting a high-speed maneuvering target, which is based on the backstepping control (BC), barrier Lyapunov function (BLF), sliding mode control (SMC), dynamic surface control (DSC), and reduced-order extended state observer (ESO). First, a fifth-order strict feedback IGC model considering the rudder delay dynamics is derived, which also considers the drag effect on the axial velocity. Second, the missile guidance control system based on the BC consists of seeker, guidance, angle-of-attack, attitude, and rudder subsystems. The seeker subsystem was designed based on the BLF, and the other four subsystems were designed based on the SMC. The system-lumped disturbances, including unknown target maneuvers, unmodeled parts, perturbations caused by aerodynamic parameter variations, and external disturbances, were estimated and compensated for using the reduced-order ESO. The DSC prevented the “differential explosion” caused by virtual control commands introduced by the BC. Subsequently, the stability of the closed-loop system was strictly proven using the Lyapunov theory, and the boundedness of the FOV angle was strictly derived. Finally, the simulation results demonstrated the effectiveness and robustness of the proposed IGC scheme.

Wind tunnel measurements of the aerodynamic characteristics of a 3:2 rectangular cylinder including non-Gaussian and non-stationary features

This paper presents a wind tunnel investigation of the aerodynamic characteristics (force and pressure coefficients) of a static 3:2 rectangular cylinder in smooth flow for angles of attack between −4° and 90° at various values of Reynolds number. In contrast to much of the existing literature, this study shows clear dependence of the mean drag coefficients on Reynolds number. In addition, the variation of aerodynamic parameters with angle of attack is fully mapped for a symmetric section revealing that the peak values of the mean values of drag, lift, moment and Strouhal number do not occur at the same critical angle of attack.The present study also presents the first map for identifying the locations of the reattachment and stagnation points as well as the zones of angle of attack where the flow is separated and attached on a face of the section. The phenomenon of switching flow is observed at the angle of attack 25°, leading to strong non-stationarity and non-Gaussian distributions of the aerodynamic forces and pressure. Another phenomenon, so-called unsteady low-frequency vortex shedding, is also observed at higher angles of attack. These phenomena need to be accounted for when estimating wind loading and aeroelastic instability for these sections.

Three-Dimensional Guidance with Terminal Time Constraints for Wide Launch Envelops

This paper proposes a three-dimensional, nonlinear guidance strategy for terminal-time-constrained interception of a nonmaneuvering target. The restrictive assumption of decoupling the three-dimensional interceptor–target engagement into two separate mutually orthogonal planes is relaxed. The guidance commands are derived using sliding mode control, with a time-to-go estimate that accounts for interceptor’s heading angle errors. It is shown that even for arbitrarily wide launch envelops, interception is guaranteed at a prespecified desired impact time. Further, a weighted control allocation technique is presented to allocate suitable lateral accelerations along the pitch and the yaw directions. Finally, numerical simulations, including extensive Monte Carlo assessments, are carried out with constant speed as well as two different realistic interceptor models that consider time-varying speed owing to aerodynamic parameter variations. Guidance strategy is shown to be effective in the presence of noisy measurements, uncertainties, and system lag. The performance of the proposed strategy is shown to be superior compared with an existing one.

Aerodynamic effects of leading edge (LE) slats and slotted trailing edge (TE) flaps on NACA-2412 airfoil in prospect of optimization

Currently, various high lift devices are being employed in the general aviation industry. The performance of these high lift devices depends on various parameters, among which the geometrical or positional parameters are one of the most important as their influence is very prominent on the aerodynamics of the wing. Hence, attention is to be given to the optimization of these devices to increase their efficiency and performance. This article focuses on the optimization of high lift devices such as leading edge slats and trailing edge flaps. The optimization is achieved with four defined configurations for each high lift device and then compared their results with each other to find the best configuration; here the best performance refers to maximization of section coefficient of lift. A two-dimensional CFD (computational Fluid Dynamics) analysis performed on best three element airfoil using SolidWorks Flow Simulation tool with variation in aerodynamic parameters as lift and pitching moment. The obtained results have an accuracy of 1.09% for maximum coefficient of lift and 0.4% for landing condition respectively. Thus the optimized NACA2412 three-element airfoil gives good aerodynamic performance.

Cooperative Salvo Guidance Using Finite-Time Consensus Over Directed Cycles

A cooperative salvo guidance strategy using fixed finite-time consensus is proposed here. The consensus algorithm is based on sliding mode control. The cooperative salvo attack problem is transformed to a consensus problem over a cycle digraph. This aids in designing a guidance strategy through selection of heterogeneous gains. The proposed strategy guarantees convergence to a common impact time for all the interceptors, within a fixed time, from a wide range of initial heading angles. Subsequently, the guidance strategy is extended to cooperatively intercept a constant velocity target. The efficacy of the proposed algorithm is vindicated through simulations on both stationary and constant velocity targets. Numerical simulations are also performed on realistic missile models that account for aerodynamic parameter variations, yielding satisfactory performance.

The Spatial Variation of Aerodynamic Parameters and Wind Erosion Potential in Southwestern Tarim Basin from Shifting Desert to Oasis

The southwestern Tarim Basin of China is a primary wind erosion damage and environmental degradation region characterized by shifting sand dunes, semi-shifting sand dunes, fixed sand dunes and oasis farmlands. To determine their relative susceptibility to wind erosion under different land cover types, wind velocity profiles were measured and used to determine friction velocity (u*) and aerodynamic roughness (z0) of the four land cover types throughout the wind erosion season of 2010–2012. The erodible particles contents of surface soil mass were analyzed. The u* averaged 0.33 ms−1, 0.44 ms−1, 0.61 ms−1 and 0.81 ms−1, while z0 averaged 0.39 mm, 13.58 mm, 39.51 mm and 310.8 mm for the shifting sand dunes, semi-shifting sand dunes, fixed sand dunes and oasis farmlands, respectively. The results showed that u* and z0 were both correlated with surface roughness properties of the underlying vegetation condition and z0 decreases with increasing wind speed due to the near surface plant roughness properties and wind flow. The erodible particles content were 56.92%, 60.17%, 65.80% and 74.56% for the shifting sand dunes, semi-shifting sand dunes, fixed sand dunes and oasis farmlands, respectively. These results indicate that though the oasis farmlands have the largest erodible particles content but with the greatest u* and z0, the dynamic wind conditions were not usually able to achieve by causing the dust emission. In contrast, the shifting sand dunes have the greatest potential for wind erosion due to lowest u* and z0.

Analytical Performance Study of Fixed Speed Wind Turbine

Fixed speed wind turbines have the advantage of being robust and reliable. They allow a direct connection to the electric. The purpose of the article is to study the aerodynamic behaviour and determine the performance of a fixed speed wind turbine. The work presents an analysis method based on the theory of blade element moments (BEM). The variation of aerodynamic parameters is studied for a wide range of wind speeds. A case study is conducted for the design of a wind turbine adapted to the Adrar site which is located in the Algerian Sahara. The results obtained showed that the wind turbine has maximum efficiency just at the design speed. For speeds higher than the design speed, the efficiency is reduced by the stall effect with decreases in torque due to the fall of the lift force. At wind speeds lower than the design value, the thrust effect increases, which puts the rotor under high mechanical stress and blade rotation decreases with low efficiency.

Fault Tolerant Margins for Unmanned Aerial Vehicle Flight Safety

Consider an unmanned aerial vehicle (UAV) operation from the time it is launched until the time it is put into an autonomous flight control mode. The control input u(t) during this time duration is modeled σu(t) and assumed healthy with σ = 1. In practice, however, the control inputs are less effective with a σ-value in the interval [0, 1) (Fan et al. 2012; Jayakumar and Das 2006; Hu et al. J. Guid. Control. Dyn. 34(3), 927–932, 2011). Complete damaged condition is inferred from σ = 0. Given a stabilizing controller, a range of σ-values in the interval [0,1) for which the closed loop system would remain stable is referred as the fault-tolerant margin (FTM). Flight operations with control effectiveness beyond the FTM are catastrophic. Further, when aerodynamic parameter variations due to an uncertain atmospheric condition in an UAV are present, it is shown that the FTMs are extremely sensitive to these parameter perturbations. In this paper, the FTMs are presented. The effects of control effectiveness factor on the nonlinear UAV when it is in the stable range are investigated. An UAV model in Ashokkumar and York (2016) is considered to illustrate the FTMs as a threshold to guarantee the safe operation.

Nonlinear mathematical model of unsteady galloping force on a rectangular 2:1 cylinder

This paper aims at establishing a reliable mathematical model for unsteady galloping instability or the combined effects of VIV and galloping. Instead of only basing on galloping displacement, the unsteady galloping force (UGF) is measured directly from spring-suspended sectional model tests through an improved force measurement technique. The measured UGF is found to be able to numerically reconstitute the galloping displacement responses with acceptable accuracy. In establishing an UGF model from the measured force, an energy equivalent principle (EEP) is proposed to simplify the modeling procedure and identifying aerodynamic parameters. The obtained UGF force model contains a third-order and fifth-order velocity terms to consider aerodynamic damping nonlinearities, and the unsteady effect of galloping is modeled by considering the variation of aerodynamic parameters with reduced wind speed. It is found that the negative aerodynamic damping provided by the first-order and third-order velocity terms is the source of the power driving the development of the unsteady galloping, while the positive aerodynamic damping provided by the fifth-order velocity term acts as a stabilizer making the galloping amplitude stable finally. The effectiveness of proposed UGF model and EEP in parameters identification was validated by comparing the numerically calculated galloping amplitude with experimental results.

Uncertainty analysis and robust trajectory linearization control of a flexible air-breathing hypersonic vehicle

Flexible air-breathing hypersonic vehicles feature significant uncertainties which pose huge challenges to robust controller designs. In this paper, four major categories of uncertainties are analyzed, that is, uncertainties associated with flexible effects, aerodynamic parameter variations, external environmental disturbances, and control-oriented modeling errors. A uniform nonlinear uncertainty model is explored for the first three uncertainties which lumps all uncertainties together and consequently is beneficial for controller synthesis. The fourth uncertainty is additionally considered in stability analysis. Based on these analyses, the starting point of the control design is to decompose the vehicle dynamics into five functional subsystems. Then a robust trajectory linearization control (TLC) scheme consisting of five robust subsystem controllers is proposed. In each subsystem controller, TLC is combined with the extended state observer (ESO) technique for uncertainty compensation. The stability of the overall closed-loop system with the four aforementioned uncertainties and additional singular perturbations is analyzed. Particularly, the stability of nonlinear ESO is also discussed from a Liénard system perspective. At last, simulations demonstrate the great control performance and the uncertainty rejection ability of the robust scheme.

Polynomial Control for Air-to-Air Missiles Based on Coefficient Diagram Methods

Based on Coefficient Diagram Methods (CDMs), a polynomial-control scheme for air-to-air missiles is proposed. According to the analysis of CDM-based polynomial control theory and missile dynamics, two overload-loop controllers with different structures for air-to-air missiles are designed. First, the structure and order of the controller and its corresponding characteristic polynomial of the closed-loop control system are defined following the coefficient diagram. Then, parameters of the polynomial are determined based on the desired control specifications. Numerical simulation studies for different controllers under various conditions, including the variation of aerodynamic parameters, different controller structures and output disturbances, are implemented to explore the characteristics of different control systems. Simulated results verify that the controller designed via CDMs has better robustness in comparison with the traditional three-loop autopilot.

DSC‐backstepping based robust adaptive LS‐SVM control for near space vehicle's reentry attitude

PurposeThe purpose of this paper is to propose a robust control scheme for near space vehicle's (NSV's) reentry attitude tracking problem under aerodynamic parameter variations and external disturbances.Design/methodology/approachThe robust control scheme is composed of dynamic surface control (DSC) and least squares support vector machines (LS‐SVM). DSC is used to design a nonlinear controller for HSV; then, to increase the robustness and improve the control performance of the controller. LS‐SVM is presented to estimate the lumped uncertainties, including aerodynamic parameter variations and external disturbances. The stability analysis shows that all closed‐loop signals are bounded, with output tracking error and estimate error of LS‐SVM weights exponentially converging to small compacts.FindingsSimulation results demonstrate that the proposed method is effective, leading to promising performance.Originality/valueFirst, a robust control scheme composed of DSC and adaptive LS‐SVM is proposed for NSV's reentry attitude tracking problem under aerodynamic parameter variations and external disturbances; second, the proposed method can achieve more favorable tracking performances than conventional dynamic surface control because of employing LS‐SVM to estimate aerodynamic parameter variations and external disturbances.

Numerical analysis of ice accretion effects at super-cooled large droplet conditions on airfoil aerodynamics

Changes in flow field around NACA23012 airfoil from a clean condition to a super-cooled large droplet (SLD) condition were simulated, and variations in aerodynamic parameters were calculated using FLUENT. In the case of numerical simulation for a clean airfoil, flow field characteristics simulated agreed well with theory analysis, indicating that turbulence models and parameters setting are feasible. Aerodynamic parameters for iced airfoil were calculated using the same method and agreed with those measured test data under the same environment in icing wind tunnels by S. Lee. Conclusion is made that the numerical simulation is valid, and it can be an alternative to study ice accretion effects at the SLD condition on airfoil aerodynamics, leading to reduction in research cycle time and cost.

High-Performance, Robust, Bank-to-Turn Missile Autopilot Design

Subjects related to a robust multivariable autopilot design are examined in this paper. First, a canonical robust control design formulation is introduced and is illustrated by formulating an integrated autopilot design problem. This formulation addresses the considerations of missile command following, model parameter variations, actuator dynamics, flexible dynamics, and parasitic feedback effects. Then, three robust autopilot designs for the HAVE DASH II missile system are executed. The controllers are solved using the generalized Hamiltonian approach which unifies a class of robust control designs in the same framework in terms of the formulation, data structure, and solution algorithm. The simulation shows that the designs achieve good response against significant kinematic and inertia couplings and aerodynamic parameter variations.