Journal of Aerospace Engineering

A cruising autopilot is developed for a nonlinear flexible aircraft based on the model-following method. The autopilot controller, called the outer loop, is designed for a rigid body approximation of the flexible aircraft. The model-following control, named the inner loop, approximates the input-output behavior of the flexible aircraft to a rigid body. The autopilot is applied to the flexible aircraft with a model-following inner loop. This inner loop controls the flexible modes and the outer loop, which are the rigid body dynamics. The model-following control is a suboptimal linear quadratic formulation with output feedback. This technique can stabilize the flexible modes and keep their amplitudes acceptable. The autopilot is a speed and altitude tracker with a wing leveler. It is designed via a robust $H_{\infty }$ with static output feedback. An aircraft model with nine flexible modes is used to evaluate the method. It is a conceptual medium size jet, like an Embraer $190\text{/}190$ or Boeing $737–200/300$. Simulations are performed in the presence of a stochastic Dryden turbulence model combined with a 1-cosine (1-cos) shape wind gust. The model-following inner loop is validated via a comparison with an internal loop of flexible mode stabilization, designed via a standard linear quadratic regulator with output feedback.

In this research, the previous design procedure for single-variable dual-range linear controllers for highly nonlinear systems is extended to the multivariable case. The procedure is based on generating the sinusoidal input describing function models of the plant followed by selection of two of these models for optimization of the dual-range linear controller. A single multivariable controller is designed by optimization to achieve a stable closed-loop system that satisfies a set of user-defined performance measures. Finally, the design is verified by digital simulation of the nonlinear plant and linear controller. Stability is demonstrated by successful generation of the sinusoidal input describing function models of the final closed-loop system. The procedure developed and the associated software are applied to bank-angle control of an unmanned aerial vehicle (UAV) with a discontinuous nonlinear term. It is shown that performance of a dual-range linear controller is superior to that of both a single-range and a multirange linear controller, and it competes with two other nonlinear controllers.

After being initially launched into the parking orbit, a spacecraft usually implements several maneuvers to insert its nominal orbit as the geostationary orbit for global communication, a Halo orbit to survey the solar wind, and an adjacent orbit to rendezvous with other spacecraft. The transfer problem between two different trajectories is very important in academic research and practical engineering. Trajectory correction maneuvers are necessary to keep actual flight trajectories near the nominal ones because of the errors produced by control maneuvers, measurements, launching, and modeling. However, an inappropriate strategy costs more time and fuel and may even result in mission failure. This study investigates optimal maneuver strategies from the stochastic control viewpoint to track interesting and typical trajectories, including the Halo-transfer and Lambert rendezvous orbits. This study proposes an improved correction algorithm to eliminate modeling errors for reference transfer trajectories. Moreover, this study develops an innovative approach to obtain the optimal correction maneuver strategy that uses the stochastic control theory in determining corrections and adopts the genetic-algorithm and Monte Carlo joint simulation to yield the schedules of maneuvers. Finally, numerical simulations indicate that the proposed approach has potential applications in the tracking of reference trajectories in closed-loop correction maneuvers.

The increasingly cluttered environment in space is placing a premium on future spacecraft and satellites that are capable of tracking and estimating the trajectory of unknown space debris autonomously without consistent communication with ground stations. In this paper, a vision-based debris trajectory–tracking method in close proximity is presented using two cameras onboard of two satellites in a formation. To differentiate the target debris from other clutters, a data-association technique is investigated. A two-stage nonlinear robust controller is developed to adjust the attitude of the satellites such that the target debris can be maintained within the field of view of the onboard cameras. Capabilities of the proposed estimation and control methods are validated in the simulations.

In this research, the applicability and feasibility of using a spherical-shaped platform as an experimental facility that emulates the dynamics of on-orbit conditions in the laboratory environment is investigated. Indeed, this will permit evaluation of path planning and feedback control algorithms for performing precise satellite maneuvers on an actual system in local situ. The correct operation of this system requires that rolling motion be maintained; nonetheless, some practical issues may occur because of passage to surfaces with different roughnesses, causing an irregular transient performance from a sliding to pure rolling state or vice versa. Noting that an impulse in momentum is verified on reaching the separating limits of those regions, nevertheless the designed attitude controller is observed to be able to reestablish its function despite such disturbances and change of operational conditions. This study is a preliminary step toward final concretization by the build of a prototype in the near future.

In this article, the use of a magnetic attitude control system (MACS) in preparing a spacecraft for orbit circularization without the use of high-level torque-restoring actuators is investigated. The main task of the attitude control subsystem is to maneuver the spacecraft so that its spin axis direction aligns with the direction of its velocity at the apogee place and to execute the spacecraft spin up for orbit circularization. The attitude control system is required to provide a desired attitude for maneuvering, such that the relevant maneuvers can be accomplished sequentially and successively at the apogee point. The nutation angle control system has been designed to maintain the nutation angle to less than 10° during orbital maneuvers. The magnitude of thrust force, duration of thrust application in each maneuver, and spin rate are chosen in such a way that the desired attitude of the satellite is not affected by maneuver-induced disturbances. This selection has been finalized through repeated simulations and by considering all the relevant factors and disturbances. The final selection is a thrust of 3 N using a cold gas propulsion system with an application duration of 100 s at the apogee point and a satellite spin rate of 20 rpm. The whole scenario required for the execution of orbit circularization from the initial stage of spacecraft separation from the carrier rocket has been simulated. Moreover, a scenario has been designed and implemented for control of the spin axis direction so that if the objective becomes the control of the spin axis in a short time, in addition to the magnetorquer of the longitudinal axis, one of the magnetorquers of the transverse axis can also be used. In this situation, the other magnetorquer of the transverse axis is simultaneously used for control of the spin rate. Also, for accelerating and optimizing control of the spin axis direction in the power acquisition mode, a scheme has been presented, in which spin axis maneuvering is minimized between two given attitudes. In addition, the wobble angle that is produced due to the existence of a dynamic imbalance in the spacecraft is never eliminated by this active damping method. Computer simulation results of attitude, pointing, and orbital maneuver controls are provided to demonstrate the validity of the design capability.

A novel approach for discrete multibody system modeling is presented. The approach segments the multibody system into a series of subsystems with kinematic constraints. The equations of motion of the unconstrained subsystems are either easier to obtain or already exist. The Udwadia-Kalaba equation is then introduced to calculate the constraint forces due to the constraints. The approach does not require any auxiliary variables, such as Lagrange multipliers, and does not need to use any projection. It can systematically reach the model via clustering of the segments.

The performance of a star tracker is closely linked to the characteristics of the guide star catalog. It is an essential part of design to estimate the number of guide stars observed by the star tracker in any given boresight direction. This paper reports on the star covering region evaluation method, which is used to find ranges of orientation where the star tracker can observe a star. With this method, a conservative number of guide stars observed by the star tracker in any given boresight direction can be quantitatively evaluated and credible ranges of orientation can be found, where the star tracker can observe enough guide stars to perform autonomous star identification; thus, verifying the design of the star tracker. The celestial sphere is divided into 131,072 approximately congruent spherical triangles, such that the star’s covering region can be approximately depicted by triangles in which the circumcenters are close enough to the star. Then, the stars observed by the tracker, in which the boresight is located anywhere inside the circumcircle of the triangle, can be found by identifying where the triangle is located in a given star’s covering region. A guide star catalog containing 2,873 stars was generated and the star covering region evaluation was performed. It was found that the star tracker with a circular field of view (FOV) radius of 4° could measure at least three guide stars when its boresight direction was located in 63,554 spherical triangles. The evaluation results for 10,000 random orientations were used to compare the counting numbers of practically observed stars. The mean error is 0.6415 and the maximum error is 5, which results from the stars not considered because they are on the edge of the FOV. The results of a simulation have validated the method.

A methodology for modeling and analysis of satellite antenna systems is presented considering the effects of the joint clearance and reflector flexibility in the system. The joint clearance model is established based on a thorough geometric description of the eccentricity vector. The contact-impact forces are evaluated based on a Hertz contact theory and a modified Coulomb’s friction law. The fixed-interface component-mode synthesis method and Lagrange’s equations are used to achieve a lower-order dynamic model by modal truncation. Furthermore, in the process of dynamic modeling, the changes of topology were expressed as constraint variations; namely, adding or deleting the variable structure constraint in the dynamic model. Finally, the multibody dynamics of satellite antenna are analyzed by considering a flexible reflector with ideal and nonideal joints. Thus, the simulation results can predict the dynamic characteristics of satellite attitude and antenna pointing with the effects of clearance and reflector flexibility. The conclusions have important theoretical and significant practical engineering value in the analysis and control of satellite antenna pointing accuracy.

In this paper, experimental and numerical studies are conducted to study the damage and failure behavior of a laminated T700/BA9912 composite under low-velocity impact. The three-dimensional (3D) Hashin damage criterion and the cohesive zone model are used in the finite-element analysis, and the numerical simulation reveals the delamination, matrix damage, and fiber breakage process in the material. To show the capacity and fidelity of the numerical model, a drop-weight test and a nondamage inspection (NDI) are performed, and the model predictions are compared with the test results. The numerical model is able to predict the load versus impact time responses of the specimens. The predicted size and shape of the delamination area show excellent agreement with the NDI result, and the predicted through-thickness matrix failure directly below the impactor and the fiber breakage in the bottom layers of the laminates are in good agreement with the experimental observations.

Buckling and postbuckling behavior of laminated composite slender beams is presented. The nonlinear model is introduced by using the exact expression of the curvature. The material of each layer of the beam is assumed to be linearly elastic. The formulations are based on Timoshenko’s stability theory including nonlinearity. A numerical solution of elliptical integral form is employed to determine buckling loads and postbuckling equilibrium paths of composite beams. Results from the proposed method are theoretically exact from small to very large curvatures and transverse and longitudinal displacements for slender beams under axial compressive loads. The numerical illustrations concern the postbuckling response of laminated beams with different types of boundary conditions, geometric parameters, and stacking sequences. The results reveal that the geometric parameters, boundary conditions, and stack sequences have a significant effect on the postbuckling behavior of laminated composite beams.

Impact responses and damage induced by drop-weight impact on two types of glass-laminated aluminum-reinforced epoxy (GLARE) fiber-metal laminates (FMLs) were studied experimentally and numerically. Indenters with various shapes and sizes were used under different impact energies. For line-nose Charpy indenters, the effect of the angle between the indenter and the fiber direction was also investigated. Both the nondestructive ultrasonic and mechanical sectioning techniques were adopted to evaluate impact damage in the laminates. The results showed that GLARE 3 (cross-ply) offers higher impact resistance than GLARE 2 (unidirectional). The first failure at low-velocity impact occurred as delamination between the nonimpacted-side aluminum and the adjacent fiber-epoxy layer, then was followed by a visible crack in the outer aluminum layer at the nonimpacted side. More severe local damages appeared with smaller indenters, indicating that the energy dissipated mainly through delamination and cracks for smaller-size indenters. On the other hand, larger global deflection occurred if larger-size indenters were used, implying more energy might have been absorbed owing to extensive global deformation of the FMLs. Finite-element simulations were carried out using an explicit finite-element program. The damage patterns, histories of impact force, energy, and deflection, as well as the dynamic contact stiffness, were obtained numerically. Good agreement was obtained between experimental results and finite-element simulations.

Granular beds made of activated alumina and magnesium oxide used for recycling/processing water in space life-support systems are usually compressed using a spring mechanism for effective functioning. The spring force is a critical factor in providing appropriate compaction to the entire length of the granular bed, and therefore, it is important to investigate and predict the axial stress and displacement along the bed. This paper presents results from experimental, numerical (finite-element), and analytical models of a full-scale dry granular material bed of activated alumina packed in a cylinder. The finite-element (FE) model, when used with proper contact model and properties, predicted both the axial stress and the displacement along the length of the granular bed well. Janssen’s one-dimensional model did not provide as accurate predictions for stress because of simplifications associated with the model. Because the system is long and slender, it is very important to properly model the wall friction in predicting the behavior of the granular bed. All the assessments used in this study indicated that the axial stress decays very steeply along the bed.

The flexibility of spacesuits made of soft material is limited by high-pressure differences across the garment, which can induce strong tensile stress and stiffen soft materials. This paper describes a new single-axis pneumatic joint design for the enhancement of spacesuit flexibility. The joint is based on the implementation of an isotensoid shape that can theoretically carry interior pressure without circumferential stress. By introducing excess skin material into the isotensoid along a circumferential direction, local slack regions are generated, which enable expansion and contraction, resulting in a flexion motion with low resistance torque. Joint prototypes were fabricated from silicone rubber–coated Vectran woven fabric (Kuraray, Tokyo, Japan) and were subjected to flexion tests using a custom-made robotic arm. The torque measurement results showed that the increase in torque was low until the flexion angle reached 80° and rose rapidly at higher angles. It was concluded that the presence of excess skin can effectively provide flexibility at low angles and that increased torque at higher angles should be addressed in future design improvements.

A unique shield system for lunar dust has been developed using electrostatic force to prevent dust from entering into bearings and mechanical seals of equipment used for lunar exploration. A single-phase rectangular high voltage is applied to insulated parallel-plate electrodes printed on the edges of the gap in the mechanical sealing part. It was demonstrated that more than 70% of the dust was repelled from the gap with very little power, and it was predicted by numerical calculations performed using the distinct element method that the shielding performance of the system would improve further in the low-gravity and vacuum environment of the Moon. This technology is expected to increase the reliability of equipment used in long-term manned and unmanned activities on the lunar surface.

Measuring regolith density in outer space has always been a difficult problem. In this paper, a bulk density measuring method for planetary regolith based on vibration is proposed. The bulk density measurement of planetary regolith being sampled is realized mainly by vibration signal processing and general regression neural network identification. The method makes full use of the special structure of a mini coiling-type sampler, which utilizes the flexible coiling spring as its sampling arm. To begin, vibration signals are acquired and their trends are removed. Then, multiresolution wavelet analysis is used to filter out the signals that are not related to bulk density. After that, two of the power spectrum estimation methods, Periodogram and Welch, are introduced to extract the density-related features of the vibration signals and reduce the dimensions of them. The multiclass classification support vector machine is then adopted to verify the relationship between the bulk densities of regoliths and feature vectors. Finally, the general regression neural network is used to identify the bulk densities. Experiments show that by utilizing the coiling-type sampler the regolith bulk density measurement can achieve an accuracy of 89.22%. After comparison of the two density-related feature extracting methods, it is found that the Welch method is the better of the two for performing the density measurement.

This paper provides insights into the stability of flexible orthotropic rectangular plates in supersonic flow by revealing the amplitude-speed dependency behavior for such panels in precritical and postcritical flight conditions. Flutter and postflutter behaviors are discussed to highlight the important effect of the structural and aerodynamic nonlinearities inherently present for panels in high-speed flows. The influence of transversal shear on stability is also illustrated. Depending on the plate’s geometrical parameters, soft- and hard-type flutter behaviors are possible, and a graphical interpretation is provided along with pertinent conclusions.

This paper presents the nonlinear aeroelastic analysis of slender wings using a nonlinear structural model coupled with a nonlinear unsteady aerodynamic model. A high-aspect ratio and flexibility are the specific characteristics of this type of wing. Wing flexibility, coupled with a long wingspan can lead to large deflections during normal flight operation of an aircraft; therefore, a wing in vertical/forward-afterward/torsional motion using a third-order form of nonlinear flexible Euler-Bernoulli beam equations is used for the structural modeling. An unsteady nonlinear aerodynamic considering static stall based on the Wagner function is used for determination of aerodynamic loading of the wing. Combining these two types of formulations yields fully nonlinear integro/differential aeroelastic equations. Using Galerkin’s method, a mode summation technique, and a numerical method, the governing equations are solved to predict the nonlinear aeroelastic response of a wing in the stall region. The results obtained are compared with the results given in the literature. In addition, the effects of the static stall characteristics of some typical airfoils on the aeroelastic behavior of a flexible wing are investigated. The results indicate that in a range of static stall angles of attacks there is a specific angle of attack in which the limit cycle oscillation boundary speed has a maximum value.

A comprehensive numerical analysis was conducted to study the internal flow development in an integrated rocket-ramjet (IRR) propulsion system. The study consists of two parts: transition from the rocket booster to the ramjet sustainer and combustion dynamics during ramjet operation. The physical model of concern includes the entire IRR flow path, extending from the leading edge of the inlet center body through the exhaust nozzle. The theoretical formulation is based on the Farve-averaged conservation equations of mass, momentum, energy, and species concentration in axisymmetric coordinates and accommodates finite-rate chemical kinetics and variable thermophysical properties. Turbulence closure is achieved using a low-Reynolds-number $\kappa \text{-}\epsilon$ two-equation model. The governing equations are solved numerically by means of a finite-volume preconditioned flux-differencing scheme capable of treating a chemically reacting flow over a wide range of Mach numbers. Various important physiochemical processes involved in the transition from the booster to the sustainer phase are investigated systemically. Emphasis is placed on the flow interactions between the inlet diffuser and combustor. The effects of operation timing on the flow evolution, fuel spread, ignition, and flame development are studied.

The flow and combustion dynamics in the ramjet sustainer of an integrated rocket-ramjet (IRR) engine are investigated systematically. The physical model includes the entire engine flowpath, from the freestream in front of the inlet to the exit of the exhaust nozzle. The flowfield obtained previously is used as the initial condition for the present analysis, so that the complete operation history of the engine can be obtained. Various physiochemical processes dictating the engine internal flow development and flame behavior are examined. Emphasis is placed on the interactions between the inlet and combustion dynamics, as well as the overall engine performance. The key mechanisms for driving and sustaining the flow oscillations are explored. In addition, the acoustic wave structures in the engine flowpath are identified.

This work investigates the effects of nanoparticles in the peristaltic flow of Johnson-Segalman fluid. The fluid is considered in an asymmetric channel. The relevant nonlinear equations are modeled by using mass, momentum, energy, and concentration laws. The solutions for velocity, temperature, nanoparticle phenomenon, and pressure gradient are derived. The results obtained are analyzed for various parameters of interest involved in the problem.

This research focuses on the numerical studies used to combined direct perturbed-force (DPF) and updated geometry (UG) technologies to characterize the vibration behaviors of a wrinkled membrane-inflated arch. The predicted wrinkles are in good agreement with experimental measurements. These wrinkling results are then introduced into modal analysis to predict the mode shapes and associated frequencies of the inflated arch. The effects of the wrinkles on the vibration characteristics of the inflated arch are then analyzed. The initial wrinkling location and load are found to be very much dependent on the inflated pressure. The vibration characteristics of the inflated arch are very sensitive to the wrinkles. The wrinkles should be effectively eliminated and controlled to increase the structural frequencies and obtain the required vibration characteristics of the membrane-inflated arch.