Journal of Aerospace Engineering

This paper addresses the problem of multiple unmanned aerial vehicle (UAV) rendezvous when the UAVs have to perform maneuvers to avoid collisions with other UAVs. The proposed solution consists of using velocity control and a wandering maneuver, if needed, of the UAVs based on a consensus among them on the estimated time of arrival at the point of the rendezvous. This algorithm, with a slight modification is shown to be useful in tracking stationary or slowly moving targets with a standoff distance. The proposed algorithm is simple and computationally efficient. The simulation results demonstrate the efficacy of the proposed approach.

This paper presents an automatic landing control design using adaptive, integrated guidance and control (IGC) logic. The proposed IGC design uses a combination of an adaptive output feedback inversion and backstepping techniques. The problem is formulated as an adaptive output feedback control problem for a line-of-sight-based chasing flight configuration. The design objective is to regulate the relative distance between the aircraft and the moving reference point on a landing pattern and two bearing angles maintaining turn coordination. Adaptive neural networks are trained online with available measurements to compensate for inversion error as a result of unmodeled dynamics and modeling errors of the aircraft in the design process. In addition, a reference command trajectory for the automatic landing control is designed in a way that the aircraft follows the landing pattern regardless of its initial position. The automatic landing system using IGC logic is evaluated using a sophisticated six-degrees-of-freedom nonlinear simulation program with the approach and landing scenario.

The paper focuses on the design of a robust H-infinity attitude controller for an unmanned small-scale helicopter. To take into account the salient nonlinearities, a model with six-degrees-of-freedom nonlinear dynamics and some linear approximation of the aerodynamic parts are used when extracting a linear model and performing simulations to check the performance of the designed controller. To design a robust H-infinity controller, an augmented plant is constructed by adjusting several weighting functions. Then, a robust controller is synthesized utilizing the augmented system with the weighting functions and H-infinity control methodology. Using computer simulation it is shown that the H-infinity controller works well when applied to the nonlinear model even though it is designed using a linear model approximation. Through frequency response analysis, it is shown that the proposed controller can overcome more than half of the uncertainty variations around a nominal point at the input side. The time-domain simulation with the nonlinear model demonstrates that the proposed controller is very robust in relation to the uncertainties, as was expected, overcoming large gain uncertainties and time delay in each input channel. The analysis and simulation results also show that the control system satisfies the Level 1 handling requirements, as defined in Aeronautical Design Standard ADS-33E-PRF.

Vision-based unmanned aerial vehicle (UAV) control is a core technology in global positioning system (GPS)-denied environments. However, small UAVs have difficulty processing image data in real time because of their limited payload capacity and relatively weak processor. Development of an image-processing system that is suitable for small UAVs is therefore needed. In this paper, an approach for a vision-based simultaneous localization and mapping (SLAM) system for small UAVs is suggested. For real-time localization, a method wherein Kanade-Lucas-Tomasi-based localization and tracking are performed within small UAVs in real time is adopted. At the same time, scale-invariant feature transform–based mapping and more accurate localization are performed at the ground-control station. A vision-based three-dimensional map-building method for small UAVs using a monocular camera is also proposed. The proposed method first extracts straight-line information from images and calculates their equations in three-dimensional space. It then constructs plane information from line information.

Formal verification has played an important role in demonstrating correctness of safety-critical systems. Small-scale unmanned helicopters have been increasingly developed for various purposes such as scientific exploration and commercial or defense applications. The HELISCOPE project in Korea aims to develop a small-scale unmanned helicopter and its onboard embedded computing system for flight control and real-time transmission of multimedia data. This paper shares the authors’ experience on the formal verification of the operational flight program (OFP) developed in the project. Because the OFP provides real-time controls on various sensors and actuators, demonstration of its correctness through formal verifications has been strongly recommended. This paper focuses on real-time process communications among sensing processes, a monitoring process, and a controller process. They all share a common data area. Two different formal models were developed for the OFP and verified with the SPIN and UPPAAL model checkers. While the SPIN model checker is widely used for modeling and verifying communication protocols, the real-time behavior of the OFP controller needs more advanced techniques that can handle real-time properties explicitly, i.e., timed automata and the UPPAAL verification system. The verification of the OFP found several safety-critical faults; they were all reported to development teams and fixed.

An investigation of the flow field around a double-delta wing was carried out at the Aerodynamics Analysis and Design Laboratory's (AADL) water tunnel facility. The model has a sharp leading edge and 76/40° configuration, which is representative of current fighter aircraft. The experiments were conducted at Reynolds numbers of 10,000 and 15,000. The dye injection technique was utilized to analyze the flow. A comparison of the vortical flow structure and the bursting point locations was done for a Reynolds number of 15,000 and found to be in reasonable agreement. The major objective of this experiment was to study the effect of the angle of attack (AOA) and Reynolds number on the vortical flow structure. In this flow visualization result, the dominant feature of the AOA and Reynolds number effect on the vortical flow structure over the double-delta wing has been observed. With increasing Reynolds number, the distance between the vortex trajectory and the model surface becomes smaller. As the AOA increases, the intertwining or coiling up features cannot be seen and the vortex bursting point locations move upstream. The distance between the vortical streakline and the model surface becomes larger with increasing AOA. The crossover point of the main and strake wing vortices also depends on the AOA.

The unsteady wake characteristics of a butterfly-shaped wing undergoing flapping and feathering motions are studied. A 4-bar quick return mechanism is used to generate the flapping motion. To achieve considerable feathering motion, the leading edge of the wing is fixed to the flapping arm of the model, while the trailing edge of the wing is kept free to bend and deform out of weight and aerodynamic loading. The flapping frequency and wing size considered for this study cover rotational Reynolds numbers of 6,040–10,080 for a zero advance ratio. The three-dimensional (3D) (stereoscopic) time-resolved (TR) velocity field is measured to understand the dynamics of the lift and thrust generation. The conservation of momentum principle has been used to estimate the thrust force from the particle image velocimetry (PIV) results. The mean velocity profiles in the wake show the frequency dependence ejection of the trailing edge vortex, which is believed to be effectively used for thrust production and maneuvering with various sized butterflies. The unsteady lift generation mechanism is discussed with the help of the measured variation of circulation with time. The thrust estimated from the PIV indicates positive generation throughout the entire cycle of flapping. An analytical approach for lift and thrust estimation has been considered for better understanding of the expected nature and variation of these aerodynamic forces.

The purpose of this work is to design a new layout for a quadrotor unmanned aerial vehicle (UAV) for experimental customized control-system validation. Geometrical and structural properties have been computed, and commercial off-the-shelf equipment has been selected. The flight control system, both hardware and software, has been custom developed; the same applies for the mathematical model describing the flight dynamics. Two proportional-derivative (PD) controllers have been designed and experimentally tested. A different control strategy with a state feedback controller has been implemented, assigning a robust eigenstructure and the desired dynamic specifications. An adaptive algorithm, which guarantees a uniformly bounded response for both the inputs and the outputs, has been presented. An L1 adaptive controller has been studied and implemented, including the state predictor and the adaptive and control laws. Tests with the PD controller have shown slowly stable response even in the event of numerical and vibrational noise. In conclusion, the L1 adaptive controller permits one to obtain optimal response characteristics, and the low-pass filter integrated in control law, combined with a Kalman filter, can reduce noise and correlate data between different sensors.

This paper presents aerodynamic design and characterization of a new quad tilt-wing unmanned aerial vehicle [Sabancı University unmanned aerial vehicle (SUAVI)] through wind tunnel tests and provides experimental data for the design of similar aerial platforms. SUAVI is capable of vertical takeoff and landing (VTOL) and horizontal flight, and it can perform both indoor and outdoor surveillance. Aerodynamic design of the vehicle directly affects its operational performance, including flight stability and flight duration in vertical, transition, and horizontal flight modes. Selection of the propulsion system and determination of the shape of the fuselage and the wings are done in an optimal manner by taking several aerodynamic criteria into account. Flow simulations reveal that the rear wings are affected by the downwash of the front wings. To solve this problem, the rear wings are placed at a higher incidence angle than the front wings. Wind tunnel tests are performed to measure the lift and drag forces and pitching moments for level flight in the entire speed range. Furthermore, the rear and front motor throttle settings and the wing incidence angle combinations for the nominal flight are measured and tabulated. To eliminate undesired spanwise air flows at the wing tips, several winglets with different shapes and sizes are introduced, and the optimum winglet is developed by several simulations and experiments. Some of the results presented in this paper are novel contributions to the literature and can be used in the design of new hybrid unmanned aerial vehicles (UAVs).

A method is presented to construct computationally efficient reduced-order models (ROMs) of three-dimensional aerodynamic flows around commercial aircraft components. The method is based on the proper orthogonal decomposition (POD) of a set of steady snapshots, which are calculated using an industrial solver based on some Reynolds averaged Navier-Stokes (RANS) equations. The POD-mode amplitudes are calculated by minimizing a residual defined from the Euler equations, even though the snapshots themselves are calculated from viscous equations. This makes the ROM independent of the peculiarities of the solver used to calculate the snapshots. Also, both the POD modes and the residual are calculated using points in the computational mesh that are concentrated in a close vicinity of the aircraft, which constitute a much smaller number than the total number of mesh points. Despite these simplifications, the method provides quite good approximations of the flow variables distributions in the whole computational domain, including the boundary layer attached to the aircraft surface and the wake. Thus, the method is both robust and computationally efficient, which is checked considering the aerodynamic flow around a horizontal tail plane, in the transonic range $0.4\le \text{Mach number}\le 0.8$, $-3°\le \text{angle of attack}\le 3°$.

The potential positive influence of close ground proximity on the aerodynamic performance of a wing has long been established; however, flight at a Mach number that causes shock waves to form between the wing and the ground plane would have significant aerodynamic consequences. A numerical study to investigate the effects of extremely low-level flight at subsonic Mach numbers up to 0.9 was conducted involving the symmetric, swept ONERA M6 wing. The aerodynamic coefficients as well as the pressure distributions and tip vortex behavior were analyzed for cases involving the angle of incidence, ground clearance, and Mach number as variables. It was found that lift was augmented with decreasing ground clearance in the majority of cases with the wing at positive incidence, although the critical Mach number for the wing, and in particular for the lower surface, was reduced at very low clearances. The earlier emergence of a lower shock was accompanied by a considerable increase in drag, although this was offset in part by a marked decrease in the strength of the tip vortex when the wing was within one chord length of the ground plane. As a result of the more pronounced shock structures and separation regions that occur above a freestream Mach number of 0.8, there is no net aerodynamic efficiency to be gained from flight with this wing in close ground proximity compared with unbounded flight at altitude.

The development and formation of the leading-edge vortices due to the change in the angle of attack, $\alpha$, and yaw angle, $\theta$, for a unique cross-flow plane at a dimensionless distance of $x/C=0.8$ from the apex of the stationary delta wing with a sweep angle of $\Lambda =40°$ were observed using stereoscopic particle-image velocimetry (stereo-PIV). In addition, the experiments were conducted on three different cross-flow planes such as $x/C=0.6,0.8,$ and $1$ using dye visualization to reveal the development of leading-edge vortices over the delta wing. The angle of attack was varied within the range of $7\le \alpha \le 17°$ and the yaw angle was varied within the range of $0\le \theta \le 8°$. The vortical flow structure and loadings toward the wing surface due to the fluctuations and unsteadiness in the flow structure near the delta wing are investigated using time-averaged parameters such as streamlines, contours of vorticity distributions, Reynolds stress correlations, distributions of turbulent kinetic energy, vertical velocity, and RMS of vertical velocity fluctuations.

The effects of a forward-sweep angle ($\phi$), Reynolds number ($\mathsf{R}$), and angle of attack ($\alpha$) on the boundary-layer flow and aerodynamic performance of a finite forward-swept wing (NACA 0012 airfoil) were experimentally investigated. The Reynolds number and the forward-sweep angle were $2.5\times {10}^4<\mathsf{R}<{10}^5$ and $0°\le \phi \le 45°$, respectively. The chord length was 6 cm, and the semiwing span was 30 cm; therefore, the semiwing aspect ratio was 5. The boundary-layer flow structures were visualized using the surface-oil-flow technique. Seven boundary-layer flow modes were identified by changing $\mathsf{R}$, $\alpha$, and $\phi$: separation, separation bubble, secondary separation, leading-edge bubble, three-dimensional ($\phi =0°$)/bubble-extension flow, and bluff-body wake modes. Furthermore, a six-component balance measured the aerodynamic loadings. The aerodynamic performance related closely to the boundary-layer flow modes. The dimensionless bubble length ($L_b/C$) indicated that a high forward-sweep angle conducted a longer bubble length than that at a low forward-sweep angle. The pseudostall occurred in the bubble-extension mode, while $\phi =30°$, and ${\alpha }_{\mathit{stall}}$ of 19.5° at $\phi =45°$ was about 1.16-fold larger than that of 9° at $\phi =0°$ as $\mathsf{R}=8.0\times {10}^4$. Namely, the forward-sweep angle retards the stall.

This paper describes a prototype and analytical studies of a spherical rolling robot, a new design of a nonholonomic robot system. The spherical robot is driven together by an internal linear motor, continuously changing the center of gravity, and external wind thrust. A mathematical model of the robot’s motion was developed using Maggi’s equations without the Lagrange’s multipliers from a view of the nonholonomic constraint. An obstacle avoidance experiment was carried out to analyze the law of motion of the robot driven three ways: external wind driven, internal motor driven, and a combination of both. The results demonstrate that the proposed dynamic model is valid for the spherical robots driven by the different ways for environmental exploration.

The vibration isolation platform of control moment gyroscopes (CMGs) on satellites is discussed. First, the theory of single-axis vibration isolation is presented and an isolation scheme for CMGs is put forward and applied to a satellite, which can provide an ultraquiet environment for optical payloads. Second, an integrated satellite dynamic model including the vibration isolation platform and pyramid configuration of the CMGs is built. The validity of this dynamic model is then verified by ADAMS. Third, the influence of the vibration isolation system on the attitude control system of the satellite is analyzed in detail and the appropriate parameters of the proportional-integral-derivative (PID) controller are selected. Finally, the attitude stabilization performance with the vibration isolation system is predicted by integrated simulation. The analysis and simulation results show that the vibration isolation platform can attenuate the disturbances to a certain extent—thus, improving the attitude stability—and when the vibration isolation platform is used, the closed-loop control system is conditionally stable, which means that the gain parameter of the PID controller should be within a range according to the root locus diagram.

A laboratory scale experimental setup has been developed to explore an alternative source of solid propellant. A combination of potassium nitrate and ferric oxide embedded in epoxy binder constitutes the ingredients of the propellant used. A scanning electron microscope and nondestructive test (X-ray radiography) were utilized to study the crack surface and internal structure of the cured propellants. The influence of the epoxy mixing ratio on the sample density has been studied. Two solid propellant grain configurations were utilized to study the performance characteristics in terms of the thrust generated. It was found that the manufacturer’s mixing ratio provided the optimum density results for the given set of configurations tested. The multiperforated grain configuration showed higher performance in terms of trust when compared with the single perforated grain configuration.

The axisymmetric peristaltic flow of a viscous compressible liquid through the gab between two coaxial tubes (annulus) was studied. This peristaltic flow is actually a result of the influence of ultrasonic radiation on the flow of a liquid through an annulus, which deforms the walls of the outer tube in the shape of traveling transversal waves exactly like peristaltic pumping. Those traveling transversal waves induce a net flow of the liquid inside the annulus. This problem has numerous applications in various branches of science, including stimulation of fluid flow in the annulus under the effect of elastic waves and studies of blood flow dynamics in living creatures. The wave amplitude is related to the power output of an acoustic source. A perturbation technique was employed to analyze the problem where the amplitude ratio (wave amplitude/outer tube radius at inlet) is chosen as a parameter. In the second-order approximation, a net flow induced by the traveling wave was calculated for various values of the compressibility of the liquid, the Reynolds number, the radius ratio, and the wavelength parameter. The calculations disclose that the compressibility of the liquid, the Reynolds number, and the radius ratio have an apparent influence on the net flow. Furthermore, the net axial velocity at the wall and in the annulus gab is numerically discussed.

A wind-tunnel study has been carried out to investigate the effects of the barrier wall and surface roughness on the simulated atmospheric boundary-layer (ABL) flow generated using the Counihan method, (i.e., the castellated barrier wall, vortex generators, and surface roughness elements). Experimental results indicate that changing the basic barrier height does not influence the mean-velocity profiles. However, turbulence intensity and absolute values of the Reynolds stress increase with an increase in the barrier height, and simultaneously the integral length scales of turbulence decrease. Variations in the height of the barrier castellation influence the simulated ABL flow similarly to the changes in the basic barrier height. Empirical relations were derived to quantify effects of the surface roughness height on the simulated ABL flows. Increasing the height of surface roughness elements results in a larger velocity defect and stronger turbulence in the near-ground region.

Postbuckling behavior of three-dimensional (3D) braided rectangular plates subjected to biaxial compression is presented. The 3D braided composite may be treated as a cell system, and the geometry of each cell is deeply dependent on its position in the cross section of the plate. Based on Reddy’s higher-order shear deformation plate theory and general von Kármán–type equations that include the initial geometric imperfection of the plate, a perturbation technique is employed to determine postbuckling equilibrium paths of 3D braided rectangular plates. The results reveal that the geometric and physical properties have a significant effect on the postbuckling behavior of braided composite plates.