Journal of Aerospace Engineering

A new procedure for synthesis of nonlinear proportional and nonlinear rate-feedback controllers for use with unstable nonlinear systems with application to a direct drive inverted pendulum is presented. The approach is to stabilize the nonlinear system followed by generation of the corresponding describing function models at various operating regimes of interest. With the known stabilizing controller and the stabilized frequency domain models, the frequency domain models of the unstable plant are algebraically extracted. A computer-aided design technique is used, and a set of proportional plus rate feedback controllers for the set of obtained open-loop frequency domain models is designed. The table of rate feedback gains as a function of rate feedback block input amplitudes is treated as the describing function model of the unknown nonlinear rate feedback gain; describing function inversion is used to obtain the nonlinear rate feedback gain. One linear proportional controller at an arbitrary operating regime of interest is designed to force the closed-loop system comprising the linear proportional gain and the nonlinear rate feedback to mimic the desired behavior. In continuation, the describing function model of the closed-loop system is generated, and the desired open-loop frequency domain behavior is extracted. A set of proportional gains at various operating regimes is designed, followed by describing inversion to determine the nonlinear proportional gain. Finally, the closed-loop system comprising the nonlinear proportional gain and the nonlinear plant with the nonlinear rate feedback gain is tested for stability and robustness performance. The procedure is experimentally verified using a direct drive inverted pendulum, and it is demonstrated that the synthesized controller performs favorably to a linear proportional-integral-derivative (PID) controller, and its performance is competitive with two other nonlinear controllers. The primary differences between this research and prior work are that the procedure is developed for unstable systems and the developed procedure is experimentally verified.

Teleoperation rendezvous and docking (RVD) can be used as a supporting technology for unmanned spacecrafts, or as a technology to dock the chaser with a noncooperative target. Teleoperation RVD has attracted little attention in the scientific community in the past according to a literature survey, while it is of application significance to China’s space station project. This paper outlines the concept design of a novel teleoperation RVD system as well as the controller. Experiments to verify the feasibility of the novel concept design focus on the communication bandwidth, the impact of time delay within the control loop, and a predictive display approach to reduce the impact of time delay. The results show that the chaser is controllable even though the time delay is as large as 6 s during ground-based teleoperating closed-loop control, i.e., the newly designed system is feasible. In addition, the control approach based on predictive display demonstrates an increase in effectiveness within teleoperation RVD. Experimental results indicate that the success rate of this approach increases by 14% compared to the traditional approach that uses only visual feedbacks.

This paper discusses flying objects’ adaptive control with direct application to the flight of helicopters. Two new automatic adaptive control systems are suggested: the former is used for pitch angle control, while the latter is used for control of helicopter pitch angle and velocity; this second system is an extension of the first one. The adaptive control is based on the dynamic inversion principle and the use of neural networks. The two adaptive control systems have reference models, linear dynamic compensators, linear observers, and neural networks. The adaptive components of the automatic control laws compensate for the approximation errors of the dynamic model’s nonlinear functions. The used actuators are linear or nonlinear. To eliminate the neural networks’ adapting difficulties, a pseudo-control hedging (PCH) block is inserted in the adaptive system; it limits the adaptive pseudo-control by means of a component that represents the estimation error of the actuator dynamics. Thus, the PCH block moves back the reference model—i.e., it introduces a reference model response correction with respect to the actuator position estimation; the signal provided by the PCH block represents a reference model’s additional input. For the two new automatic adaptive control systems, a technical computing environment is used to obtain time characteristics of the adaptive systems with linear and nonlinear actuators. Phase trajectories of the two adaptive control systems with nonlinear actuators express the convergence of the nonlinear systems to stable limit cycles.

In this paper, the steady two-dimensional boundary layer flow of chemical reaction magnetohydrodynamics (MHD) viscous fluid over a shrinking sheet with suction/injection is studied. The governing equations for the problem are changed to dimensionless ordinary differential equations by similarity transformation. The resulting nonlinear differential equations are solved for velocity and concentration profiles using the homotopy perturbation and the finite difference methods. Graphical results have been presented for velocity and concentration profiles for various physical parameters of interest. To the best of our knowledge, this type of analytical solution for chemical reaction MHD viscous fluid over a shrinking sheet with suction/injection is presented for the first time in the literature.

Electrodynamic tether systems orbiting the Earth are prone to libration instability because of periodic changes in the geomagnetic field, plasma density, and lunisolar gravitational attractions in addition to nonperiodic changes resulting from the irregularity of the geomagnetic field, inhomogeneity of the Earth, and solar pressures. The long-term orbital and libration dynamics of a bare electrodynamic tether in deorbiting obsolete satellites is investigated by considering space environmental perturbations of current-induced electrodynamic force, atmospheric drag, Earth’s oblateness, irregularity of the geomagnetic field, variable space plasma density, solar radiation pressure, and lunisolar gravitational attractions. The electrodynamic tether is assumed to be rigid and the tethered spacecraft is modeled as a lumped mass. The study shows by numerical simulation that the out-of-plane libration is the primary source of libration instability in inclined orbits, which destabilizes the in-plane libration through nonlinear modal coupling. Accordingly, a simple stability criterion for current on/off switching control is derived from the libration energy of the tether to stabilize the out-of-plane libration by limiting the roll angle amplitude to a preset range. This in turn stabilizes the in-plane libration. The control requires only the feedback of the maximum roll angle with a minimum interval for current on/off switching imposed to avoid excessive current switching. The effectiveness of the control strategy has been demonstrated by analyzing the libration dynamics of electrodynamic tether with and without the current regulation in deorbiting satellites. Numerical results show that this approach is very effective in stabilizing both in-plane and out-of-plane libration of a tethered system subjected to periodic and nonperiodic perturbations.

This study considers the aeroelastic instability of a subsonic wing with an external fuel tank store containing sloshing fuel. The wing is modeled using Euler-Bernoulli beam and the fuel sloshing is simulated by a mass-spring-damper system as an equivalent mechanical model. An extension of the Peters finite-state aerodynamic model along with the strip theory is utilized to obtain the external work due to the unsteady lift and pitching moment. Also, the aerodynamic loads on the store are calculated using the slender body theory. Using Lagrange equations, the governing equations of motion are derived and solved by the Rayleigh-Ritz method to determine the flutter speed of the wing/store. The effect of fuel sloshing on the aeroelastic instability of the wing is investigated by several numerical studies and some conclusions are outlined. The obtained results show that when the fuel-sloshing frequency in the external tank is between the frequencies of the fundamental bending and torsional modes of the wing, the sloshing/aeroelasticity coupling occurs. Numerical studies illustrate that the fundamental frequency and damping of sloshing in the tank are the main parameters determining the increase or decrease of the flutter speed and a proper tuning of these parameters can considerably extend the flutter-safe speed range.

This paper studied the attitude tracking control of a spacecraft using variable speed control moment gyros (VSCMGs) as the actuator. The dynamics of the spacecraft with uncertainties was established. These uncertainties included the moment of spacecraft inertia, the configuration of the VSCMGs, and the moment of the rotator inertia. To solve the problem, the uncertainties were classified into two types. Type I included installation misalignment angles and some part of the rotator inertia, which could impact the character of the output of the VSCMGs. The rest of the uncertainties were categorized as Type II, which could impact the response of the spacecraft attitude control. Different parameter estimation methods were designed for the uncertainties of the two types. For Type-I uncertainties, their relationship with the output-torque error was established based upon which the estimation of Type-I uncertainties was designed. Based on the Lyapunov theory, the estimation of Type-II uncertainties and the adaptive attitude tracking controller were proposed. In this way, the estimation value of Type-I uncertainties could achieve the actual value, and the impact of Type-II uncertainties imposed on the system could be counteracted by the proposed attitude controller. The simulation results showed that the proposed methods could reduce output torque errors. Therefore, attitude tracking performance would be improved.

This research examines the thermal radiation effects on the steady flow of Maxwell fluid over a nonisothermal stretched surface. An incompressible fluid fills the semi-infinite porous medium. The thermal conductivity is taken in a time-dependent fashion. Boundary layer approximations are used for the momentum and energy equations. The governing nonlinear partial differential equations are reduced to ordinary differential equations by employing a similarity transformation. The solutions of velocity and temperature are presented by using the homotopy analysis method (HAM). The variations of various interesting embedded parameters of velocity and temperature are displayed and discussed. The values of the local Nusselt number have been compared to the existing numerical solution in a limited sense.

A new systematic procedure for synthesis of nonlinear lead-lag controllers for highly nonlinear unstable plants is developed. The procedure is to stabilize the unstable plant followed by generation of sinusoidal-input describing function (SIDF) models at various operating regimes of interest. The unstable plant frequency domain behavior is extracted, and a set of linear lead-lag controllers are designed that would force the open-loop behavior of the system (comprised of the lead-lag controller and the nonlinear plant) and mimic that of desired. Finally, a describing function inversion is used, and the nonlinear functions describing the values of the lead-lag coefficients as a function of the error signal are obtained. Experimental results are presented for verification, and bounded-input bounded-output stability is demonstrated by successful generation of the describing function models of the closed-loop system comprised of the synthesized nonlinear lead-lag controller and the nonlinear plant.

A study of the dynamic control of a rotating blade subjected to external excitations is presented. A tapered thin-walled beam of closed cross-section contour with fiber-reinforced composite material is used for present study. Since a rotating blade may be exposed to a severe environment such as high angular velocity, large external load, and so on, robust control methodology is implemented to extend the blade’s life and improve its efficiency. To verify the performance of the state estimator, a sliding mode observer is introduced, and an associated robustness test is conducted in case of model uncertainty. A robust control methodology using sliding mode control in conjunction with a sliding mode observer is implemented, and its performance toward reducing the flapping dynamic responses of a rotating blade under initial conditions and various external loadings such blast load and distributed load is demonstrated. Moreover, its control performance is compared with that of the conventional linear quadratic Gaussian implementation.

The gust response of a typical section is investigated in terms of both high-fidelity computational fluid dynamics (CFD) and low-fidelity semianalytical solutions of the aerodynamic flow, enabling the suitability of the two approaches in the preliminary design of a small, flexible-winged unmanned air vehicle (UAV) to be assessed. A sinusoidal vertical gust acts as the aerodynamic perturbation and the aeroelastic response is provided for different aerofoil shapes, spring stiffnesses, wind gust intensities, and modeling approaches. For attached flow, the predicted low-fidelity gust response is found to agree well with the corresponding high-fidelity gust response; for separated flow, the low-fidelity model is unable to predict the strong oscillation of the typical section, and suitable tuning of its response is needed. Three different methods are explored for correcting the low-fidelity results based on just a few high-fidelity ones. The good agreement found between the high-fidelity and the tuned low-fidelity results obtained shows that a multifidelity-metamodel-based strategy is suitable for efficiently correcting reduced-order models.

Orbit design is a very important and fundamental task in the mission analysis and design of a satellite. In the present paper, a multiobjective optimization tool is developed for the repeating sun-synchronous orbit design of remote-sensing satellites with attitude maneuver capability. The gravitational second zonal harmonic is taken into consideration and the objective functions can be chosen among revisit time and off-nadir resolution, whose formulations are presented in this paper, as well as cross-track maneuver angle and orbital lifetime. Some applicable constraints for maximum attitude maneuver and off-nadir resolution deterioration with respect to at-nadir resolution are also considered in the optimization. Finally, through various case studies, the performance of the developed multiobjective optimization tool is successfully represented.

In this paper, the optimal time-domain finite element method is applied to the flutter suppression of a nonlinear two-dimensional typical wing section. The aeroelastic governing equations are based on the inclusion of stiffness nonlinearity in pitching motion and on the quasi-steady aerodynamics. The flutter suppression problem is formulated as a general optimization problem with equality constraints that are functions of state variables. Using the variational approach, the optimality conditions are derived and the resulting equations are discretized in time-domain. Then, by setting out the discrete equations, a set of nonlinear algebraic equations is generated, and through the Newton–Raphson method, the optimum answer is attained. The numerical results are presented in which the performance of the nonlinear optimal control system designed by the time-domain finite element technique for nonlinear aeroelastic wing sections is illustrated.

Impinging jets on the curves surfaces and flat plate are frequently used in industrial practice requiring studying the subatmospheric zone. It is important to understand the structure of behavior impinging jet on the flow fields. Same of prior researchers have concluded that a recirculation zone occurs on the impingement flat plate. This recirculation zone creates subatmospheric pressure at the impingement surface. This study has been made to investigate the effects of nozzle to plate spacing, the Reynolds number and changing of the surface shape on subatmospheric region which occurs on these impingement surfaces. The subatmospheric region could be very important in engineering application. It is hoped that this study will provide data for the further research on this subject. In the present study the air jets from convergent circular nozzle of 10-mm exit diameter have been allowed to impinge normally on a flat plate, hemispherical, and conical surfaces. Experiments were done with varying the stagnation pressure from 1.2 to 2 bars. Reynolds number based on the nozzle exit diameter for this range of stagnation pressure was changed from $7.5\times {10}^5$ to $3\times {10}^6$. The distance between the nozzle exit plane and the test surface was varied from 10 to 80 cm. Measurements of static and total pressure were made by the help of pressure taps made on the test surfaces and a pitot tube. The numerical simulations were performed by the solution of Reynolds—Averaged Navier-Stokes equations with a $v^{\prime 2}\text{-}f$ model. Behaviors of the impinging jets were compared with those of a free jet. The flow field was found to be dependent on the stagnation pressure, spacing between the nozzle and impinging surface, and the impinging surface geometry. A comparison of the predicted behavior of the jet with those obtained from experimental results gives a good agreement.

A numerical study is carried out to investigate a 7.62-mm bullet turning at the trajectory apex, particularly when fired vertically. The bullet model includes an aerodynamic model, which basically covers angles of attack up to 180°. Computational fluid dynamics are used to estimate the aerodynamic properties at the high angles of attack. The role of aerodynamic moment, and particularly Magnus phenomena, in turning at the apex is studied. A positive Magnus moment tilts the total aerodynamic moment vector clockwise and generally drives the bullet turning nose down. The bullet turning was detected through a flat spin–type maneuver. The Magnus-moment oscillatory behavior and/or the Magnus-moment-caused bullet instability may occur during the turning at the trajectory apex. These were found to have a negligible effect on the bullet-turning launch angle, but they may contribute considerably to the bullet falling-phase behavior. Possibilities to affect the bullet descending part of the flight are also shortly considered.

The ignition characteristics of a flush wall, liquid-kerosene-fueled scramjet combustor are tested in a directly connected ground test rig. The total temperature and Mach number of the vitiated airflow at the exit of the nozzle are 1,505 K and 3.0, respectively. The ignition process is accomplished by air throttling and a micro hot gas generator (MHGG). Without any cavity or backward steps, the flame holding is also accomplished by the MHGG. By controlling the air-throttling pressure and overlap time of the air throttling and the MHGG, different ignition results can be achieved. This paper divides these different results into four ignition modes and provides a detailed description of the characteristics of each mode. In addition, this paper summarizes the experiment data and proves that the ignition laws have some relationship to the fuel–air equivalence ratio.

A brand-new concept of solar-powered airplane is brought forward in this paper. The special characteristics of such airplanes are sun seeking and eternal flight, which can consistently fly under the local daytime. The foreseen applications of the airplane are first discussed. The possibility of sun-seeking eternal flight is then analyzed and a $\mathrm{\Lambda }$ wing configuration airplane is proposed considering the high flying latitude and low solar elevation angle. The principles of realizing sun-seeking eternal flight are subsequently discussed. A case study of sun-seeking eternal flight is finally presented, which reveals the flight envelop of the special solar-powered airplane. The result shows the solar-powered airplane can realize sun-seeking eternal flight between the latitude 65° and 76°. In sum, the sun-seeking eternal flight of a solar-powered airplane is believed to be wide utilization waiting for further research.