Journal of Aerospace Engineering

Abstract

The dispersive modal characteristics of propagating guided waves are useful for efficiently detecting defects in high-speed railway rails due to their advantages. This paper presents an effective method for directly analyzing the effect of defects in rails on the dispersive modal characteristics of propagating guided waves, which can be utilized for the defect detection in rails. By using the semianalytical finite element method and from the characteristic equations of the guided wave propagation in waveguides, a general perturbation method is developed to provide an exact relationship between the perturbations of structural stiffness caused by damage and the perturbations of dispersive modal characteristics of propagating guided waves. Thus, the phase velocity and group velocity for the damaged waveguide can be accurately predicted. From the obtained propagating guided wave modes, a mode selection index used for selecting effective guided waves for damage detection is introduced by evaluating the nodal displacements and energy concentrations on the rail surfaces. Finally, a numerical study for a high-speed rail is given to demonstrate the effectiveness of the proposed method for rail defect detection and propagating mode selection. The results of the numerical study show that the dispersive modal properties of propagating modes are affected by the location and size of defects in the rail, and the proposed method can give useful information for effective detection of defects in rails.

Practical Applications

Defects often occur in rails, which can be detected by using guided waves. Guided waves have advantages of high detection efficiency, long propagation distance, and small attenuation along the propagation path, which is particularly useful for detecting rail defects in practices. In the defect detection process, the group velocity and phase velocity dispersion curves should be analyzed, and the effect of rail defects on these dispersion curves needs to be fully investigated for effective damage identification. Also, these results can be useful for selecting appropriate guided wave modes and effective measuring positions for detecting defects in rails by guided wave testing.

Among all of the offered methods to determine the attitude of a spacecraft, in this research, three optimal algorithms solving the Wahba problem algebraically due to their simplicity of implementation are examined. Furthermore, this study elucidates the sources of attitude determination errors and aims to enable Attitude Determination and Control Subsystem (ADCS) engineers to choose a suitable microprocessor to implement the attitude algorithm. This research tends to show the difference in interpretation of attitude knowledge errors arising from both Model-in-the-Loop (MIL) and Processor-in-the-Loop (PIL) simulations. To this aim, we conduct PIL tests on two different kind of microcontrollers, one of which is a 32-bit and the other is a 16-bit one. We compare the performance of algorithms in MIL with the two PILs resulting from 16- and 32-bit architectures and inspect the source of accuracy reduction. Finally, we conclude the best algorithm regarding implementation and the round-off errors arising from that. In addition, the execution time of methods is checked in order to evaluate prescribed solutions based on this criterion.

Active vibration control approaches are focused on unwanted vibration that cannot be reduced by passive control devices. However, they are hard to apply to vibration-induced structures, especially aerodynamic structures, which are subjected to variable-frequency disturbances. This work includes a new adaptive positive position feedback (PPF) controller methodology using a traditional compensator and an adaptive compensator responsible for variable-frequency disturbances to investigate the effectiveness of the proposed design. A UH1H helicopter auxiliary cargo door equipped with piezoelectric patches and strain gauges was used in the laboratory environment as an implementation of Technology Readiness Levels 4 and 5 to implement of proposed controller methodology. Some challenges affect the traditional PFF controller performance, and the most important one is that independent vibration can occur on structures distinct from the natural frequencies. For overcoming this, two second-order compensators are used to enhance the controller performance; one targets previously determined natural frequencies by experimental modal analysis, and the other targets dominant frequencies obtained from simultaneous measurement. The suggested controller performance was investigated under operationally obtained vibration excitation, and this methodology provides a significant vibration reduction against variable conditions in the laboratory experiments.

Large-scale plate-like structural components have been widely used in engineering fields. The occurrence and accumulation of structural damage will threaten the integrity of plate-like structures. Guided wave-based structural health monitoring (SHM) technology has been demonstrated as an effective tool to detect small-scale damage at a high sensitivity. Due to the advantages of baseline-free, concentrated sensor arrays and rapid inspection of the full area, Lamb wave-phased array SHM systems have been extensively investigated in recent years. However, the inherent multi-mode characteristic of Lamb waves is well-known to have a significant influence on wave propagation and further development of phased array method. In this study, a single Lamb wave mode extraction method combined with phased array SHM system is proposed based on a concentric piezoelectric active transducer (CPAT). First, a brief review of a Lamb wave-phased array system is presented, including working principle, sensor array, and Lamb wave excitation and sensing. Then, the single mode wave extraction method is performed based on the CPAT. The effectiveness of the single mode wave extraction-based phased array SHM system is first examined numerically. Finally, the proposed phased array system is used to monitor large plate-like structures and the availability and validity of the proposed SHM system is confirmed by an experimental study. This work provides an effective and practical tool to detect damage in large-scale plate-like structures.

Abstract

High-angle-of-attack uncommanded roll oscillations are dangerous and can cause significant challenges in flight control. This paper constructs a stability augmented system to suppress roll oscillations with nonzero mean roll angles in a canard-configuration model. To overcome the problem of weak traditional ailerons caused by large-scale flow separations at high angles of attack, spanwise blowing was used as fluid effectors to generate lateral control moments. The control effect and mechanism of spanwise blowing were analyzed through the results of force measurements and experiments using particle image velocimetry (PIV), respectively. Spanwise blowing generates the control moment by changing the trajectory of the leading-edge vortex and delaying vortex breakdown. Subsequently, virtual flight experiment technology was used to train a policy for the stability augmented system based on real-world data using deep reinforcement learning in the wind tunnel. When testing the agent, the transient flow fields around the model were obtained synchronously using time-resolved particle image velocimetry (TR-PIV). The test results showed that the agent learned to keep the model roll at approximately zero by effectively controlling the flow field using fluid effectors.

Practical Applications

The rapid development of artificial intelligence (AI) brings new ideas for various industries. Among various AI technologies, deep reinforcement learning is a self-evolving technique that is suitable for solving complex control and decision-making problems. on the other hand, the complex dynamic characteristics of aircraft at high angles of attack leads to the emergence of uncommanded motions, making flight dangerous. Therefore, this paper focuses on the suppression of uncommanded motion of canard configuration at high angles of attack, using deep reinforcement learning. The technology of jet flow control was used to play the role of aileron. After enough training in wind tunnel, the AI finally learned how to suppress the uncommanded motion of the aircraft and showed interesting behavioral logic. The results of this paper show that deep reinforcement learning can be used for complex control problems in aerospace science; however, the practicality of deep reinforcement learning in real flight needs further verification.

In this paper, a novel magnetohydrodynamic attitude adjustment method is proposed, which uses a spherical magnetohydrodynamic attitude control device for the three-axis attitude adjustment. This paper mainly studies the distribution characteristics of the magnetic field and liquid flow field, and it does not involve attitude control performance and other contents. The magnetohydrodynamic method is used to solve the liquid flow field, which is coupled with the magnetic field. Then, the angular momentum and output torque of the attitude controller are calculated. The calculations aim to investigate the effects of current intensity and frequency on the magnetic induction intensity, flow velocity, angular momentum, and output torque in the spherical magnetohydrodynamic attitude control device. It was found that the velocity, angular momentum, and output torque increased when the current frequency and intensity increased. In addition, to achieve a larger flow rate and more efficient performance of the output angular momentum, priority should be given to increasing the current intensity rather than the current frequency. Thus, an optimal combination of the current intensity and frequency is proposed based on the numerical results. Finally, some implications of the proposed method are further discussed.

To broaden the stable operating range of hypersonic aircrafts, this work examined the mechanisms of the hysteresis phenomenon for hypersonic inlets under acceleration/deceleration process conditions. Guided by unsteady flowfield simulations on two typical 2D hypersonic inlets with different critical unstart modes of hard unstart and soft unstart, two dominant characteristics of the hysteresis phenomenon during the acceleration/deceleration processes were identified. The first characteristic is that, at the same free-stream Mach number in the hysteresis range, the main separation bubble at the cowl-lip station or the internal contraction section of the inlet during the acceleration process is considerably larger than that induced by the cowl shock wave/boundary layer interaction (SWBLI) during the deceleration process. The second characteristic is that the self-starting Mach number during the acceleration process exceeds the unstart Mach number during the deceleration process. Two mechanisms are proposed to explain such hysteresis. First, in the unstarting flowfield at relatively high incoming Mach numbers, the large-scale separation bubble under the unstarted state stems from the self-coupled SWBLI, instead of the throat blockage or the isolated SWBLI, the dimensionless pressure rise through the self-coupled SWBLI during acceleration is much larger than that of the isolated SWBLI during deceleration. Second, the transition between the unstarted state and the started state in the hysteresis range is, in essence, the conversion between the self-coupled SWBLI and the isolated one.

The aerodynamic performance and flow field structure of an inward-turning inlet at a high flying Mach number are attracting more and more attention. The main purpose of this paper is to study the aerodynamic performance and the back pressure characteristics of a circular entrance three-dimensional inward-turning inlet at $Ma=5.0–8.0$, with the combination of wind tunnel test and numerical simulation. This series of experiments is divided into three parts to explore the unstart boundary, self-start ability, and suction slots influence. The experimental results show that at design point $Ma=6.0$, the inlet outlet Mach number is 2.29 and the total pressure recovery coefficient is 0.31. Further numerical simulation results indicate that the opening of suction slots could enhance the maximum capability of back pressure about 7.3%. The circumferential flow in the isolator caused the asymmetry of the shock train and led to the nonuniform distribution of the separation region. The different size of the separation zones on the upper wall and the lower wall eventually leads to different pressure distribution.

The aeroelastic effect of gust generator vanes with a high aspect ratio is more significant in large-scale wind tunnels. High-fidelity numerical results based on the computational fluid dynamics (CFD) and computational structural mechanics (CSM) coupling methodology have an extremely high computational cost. To this end, a data-driven method for rapidly estimating unsteady flexible deformations of gust generators vanes is developed. The aeroelastic rigid-elastic reduced-order model is built by the dynamic mode decomposition with control and prediction error minimization (DMDcP) methods. For all cases validated with various pitching amplitudes and frequencies of gust vanes, the maximum cumulative errors of DMDcP reduced-order model solutions are all under 2%. The efficiency of the presented method has the potential to fast estimate unsteady flexible deformations and reduce computational cost in the early stages of gust generator design while maintaining adequate accuracy as the CFD-CSM coupling methodology.

Thermal analysis of spacecraft involves complex heat flux outside space and heat transfer calculation of equipment inside space, which is the basic guarantee of its long-term stable operation. In this paper, the core module of the Chinese Space Station is taken as an example to conduct hybrid modeling of its complex radiation and heat dissipation network. The external heat flux received by spacecraft is affected by many factors, such as its attitude, orbit parameters, and solar orientation. First, based on the defined reference coordinate system, the whole station is divided into eight quadrants. Then, an intelligent occlusion algorithm is proposed to solve the complex occlusion relationship according to the space quadrant method and fuzzy logic, and then the external heat flux is obtained. Finally, the dynamic response of the station under different conditions is studied by combining traditional mechanism modeling and an intelligent algorithm, and the maximum heat dissipation capacity was obtained. The results show that the temperature control valve opening will seriously affect the main flow of the loop, and the maximum difference can be $0.1\mathrm{kg}/\mathrm{s}$. In addition, the external heat flux changes periodically, so that the flow of each branch and the valve opening also changes periodically. In the sunlight area, the radiator receiving the external heat flux is about 77,000–90,000 W, and the maximum heat dissipation capacity is 12,000 W. Those values are 28,000 W and 22,000 W, respectively, in the shadowed area. Moreover, the hybrid modeling method proposed in this paper has an important reference value for spacecraft thermal analysis and research in other fields.

The main purpose of this paper is to study the flight dynamic mechanism that induced the sudden plunging motion for jet commercial transport in transonic flight during a severe clear-air turbulence encounter. The study of flight dynamic mechanism is to obtain the mitigation concepts for pilot training in loss of control prevention of the IATA program. The comparative analyses of flight environment and pilot control response to severe clear-air turbulence encounter for two similar twin-jet aircraft and two four-jet aircraft are studied. The one with highest dropped-off altitude during the sudden plunging motion among the four will be chosen to construct the flight dynamic mechanism. The nonlinear and dynamic aerodynamic models of the chosen transport are established through flight data mining and the fuzzy-logic modeling based on postflight data. The influences of varying vertical wind and crosswind on loss of control are presented. The source of significant angles of attack to be induced by the crosswind during the sudden plunging motion is proved through the flight dynamic analysis based on model predicted data. The present study is initiated to examine possible mitigation concepts on loss of control prevention in pilot training of IATA.

Abstract

Applying fish bone active camber morphing to the wing-in-ground effect to improve the aerodynamic efficiency was investigated using computational fluid dynamics (CFD) at a Reynolds number of 320,000. Steady-static morphing was first carried out with Reynolds-averaged Navier–Stokes (RANS) equations in two dimensions for morphing start locations off (60%, 80%, and 90% chord), ground clearances ($h/c=0.1$, 0.2, 0.4, 1), and angles of attack (AoAs) 0°, 2°, 3°, 4°, and 12°. A morphing displacement ($w_{te}$) of 0.5% increased the efficiency by 2.8% (compared to non-morphing in the ground effect) for the 3° AoA and 90% start location, and by 62% in comparison to the baseline unmorphed airfoil in freestream. Reducing $h/c=1$ to 0.1 increased the lift between 10% and 17%; the larger gain was with the highest morphing deflection. A key finding was that morphing the airfoil reduced the distance between the trailing edge and ground, enhancing the ground effect. Also, morphing at an earlier start location in the chord direction resulted in a smaller area beneath the airfoil, reducing the total pressure, which reduced the overall lift compared to a later morphing start location. Dynamic morphing at 1 Hz using URANS K-Omega-SST showed a similar amount of lift as static morphing but a slightly higher amount of drag. Reducing the period caused an initial overshoot in drag before settling. The dynamic ground effect showed higher efficiency at low AoAs compared to dynamic morphing in freestream, which is beneficial for aircraft to fly with less pitch. Finally, periodic morphing for $h/c=0.1$ using sinusoidal motion with morphing starting at 25% along the chord and 4° AoA was investigated between 0.05% to 0.15% $w_{te}$ and 0.5 to 3.5 Strouhal number. Periodically morphing at 0.125% $w_{te}$ and Strouhal number of 0.9 using DES simulations increased the efficiency by 5.4%; however, it reduced the lift by 0.7%, the drag reduced by 5.8%, and it showed Kelvin–Helmholtz instability at 9.8 Strouhal number.

Practical Applications

The use of UAVs is increasing in popularity for many missions, which include observation, surveys, and the delivery of supplies, including medical. The use of UAVs typically has lower aircraft and operational costs as well as allowing the craft to carry out dangerous missions without putting the crew in danger. A wing-in-ground effect (WIG) craft typically operates on water due to the large fuel consumption savings as well as allowing the craft to travel at higher speed compared to conventional marine craft. The study focused on applying morphing wings to a UAV WIG effect craft to improve the aerodynamic performance of the craft and allow further fuel efficiency savings compared to a marine craft. The improved performance of the WIG craft and applying morphing translates to improvements in flight time and increased range. Morphing wings also allow the wing to adapt to different flight conditions allowing for optimized aerodynamic performance depending on factors such as cargo weight and weather conditions.

The static bending of rotating two-layer functionally graded material (FGM) beams with shear connectors resting on imperfect elastic foundations was performed for the first time in this study. Timoshenko beam theory and the finite-element method are used to derive finite-element formulations. The precision of the current approach and mechanical model is shown by comparing the calculated findings of this study to those of previous precise papers. A wide variety of parameter studies are carried out to capture the impact of geometric and material characteristics on the structure’s static bending behaviors, such as the distance $d$, rotational speed, volume fraction index, elastic foundation parameters, and boundary conditions. This paper’s theory and mechanical models are fascinating since mechanical systems involving rotational motion are pretty standard in engineering practice. Therefore, the computed results of this work aim to contribute to our understanding of structures of this kind.

This paper proposes an active shape control method for an antenna control system under a limited number of power supplies with a reduction in weight and cost. An analytical model of a grid reflector with embedded piezoelectric actuators was first established, from which an influence coefficient matrix was derived for controller design. A shape controller was designed by synergistically optimizing the actuator group and power supply voltage. The optimization objectives include the root-mean-square (RMS) error, control energy consumption, and a multiobjective combination of both. The design variables include the actuator group and power supply voltage. A dynamic actuator group design with a detached power supply is proposed to improve the control performance, which also significantly increases the design domain of this optimization problem. A hybrid optimization method is proposed based on the general genetic algorithm and least-squares method to reduce the difficulty and complexity of this optimization problem with a large-dimensional design domain. Numerical simulations were carried out to illustrate the effect of the proposed hybrid optimization control method. The results indicate that the RMS error can be reduced by more than 90%, even if only two power supplies are used. The introduction of a group design with a detached power supply can clearly improve the control performance, especially by reducing the energy consumption. By employing multiobjective optimization with a uniform consideration of the RMS error and control energy consumption, a series of Pareto-optimal results can be obtained, and these results can provide meaningful references for engineering designers.

Abstract

In the three-degree-of-freedom virtual flight test, the scaled model has no translational motion, which will result in the model’s dynamics being different from that of free flight, and this difference will directly affect its effectiveness in simulating free flight and the design of the flight control law. In this paper, the longitudinal and lateral-directional dynamic characteristics of virtual flight are studied from the three aspects of modal characteristics, disturbance response characteristics, and control response characteristics. The dynamic characteristic difference between the virtual flight test and free flight is analyzed. In addition, the scaling ratio of the model will also lead to differences in the flight control law design methods of virtual flight and full-size aircraft. The control law design method of the virtual flight test is established by studying the similarity relation of control law parameters between full-size aircraft and the scaled model, and the correctness of similar design of closed-loop control law parameters of the scaled model is verified by a virtual flight test. Finally, the attitude control law parameters obtained from the virtual flight test are applied to the six-degree-of-freedom flight simulation of a full-size aircraft, and the simulation results show that the three-axis attitude control law designed based on the virtual flight test also has good performance in the six-degree-of-freedom flight of the full-size aircraft, and the response of attitude angle and angular rate are approximately the same for both.

Practical Applications

Generally, the steady aerodynamic data of the aircraft are obtained through the traditional static wind tunnel test to construct the aerodynamic model of the aircraft. The aerodynamic model established by this method may not be able to accurately simulate the actual dynamic flight. However, the actual test flight of full-size aircraft has many disadvantages, such as high cost and high risk. To simulate real flight on the ground, the method of placing the scaled model in a wind tunnel for dynamic flight is proposed, which is called wind tunnel–based virtual flight testing (WTBVFT). The scaled model is connected with the strut through a three-degree-of-freedom rotation mechanism and installed in the wind tunnel test section, so that the model has three degrees of freedom of angular motion and no translational motion. This paper analyzes the differences in dynamics between virtual flight and free flight under open-loop and closed-loop control. The results show that the virtual flight test can simulate the angular motion of free flight well and can be used to design, verify, and simulate the attitude control law of full-size aircraft well.

Graphene is recognized as one of the most alluring reinforcements for composite materials due to its exceptional mechanical capabilities. However, the existing models disregarding the compatible requirements of interlaminar stresses in the literature may meet severe challenges for the dynamic analysis of functionally graded graphene-reinforced composite (FG-GRC) laminated thick plates. If transverse shear deformations cannot be described accurately, the prediction of natural frequencies for FG-GRC laminated plates will be seriously and directly affected. Accordingly, an attractive plate theory in conjunction with the analytical approach and finite-element formulation is proposed for free vibration analysis of FG-GRC plates. An improved transverse shear stress field is acquired in terms of the three-dimensional (3D) elasticity equations and the Reissner mixed variational theorem (RMVT). The performance of the suggested model was evaluated using 3D elasticity solutions and the results obtained from other models. Numerical results showed that the suggested model can produce promising results for the thick laminated and sandwich plates. Additionally, a thorough investigation was conducted to study the effects of several graphene parameters, including volume fractions, distribution patterns, span-to-thickness ratios, aspect ratios, boundary conditions, and stacking sequences, on the dynamic responses of FG-GRC laminated plates.

An adaptive controller for the tracking of large-angle maneuvers is proposed for spacecraft equipped with an active pointing ultraquiet platform (AQP). The inertial parameters of the spacecraft and the payload (supported by the AQP) are estimated online, whereas the parameters of the AQP are assumed to be precisely known. To enable the derivation of the proposed controller, a novel dynamic model of AQP spacecraft was derived, which employs relative motion variables for the payload and is linear in the unknown inertial parameters. A Lyapunov-like argument was used to prove the local asymptotical stability of the closed-loop system, and the original controller was simplified further to a physically more intuitive proportional-differential (PD) plus adaptive feed-forward controller. The superior tracking performance of the proposed controller compared with that of a standard PD plus nonadaptive feed-forward controller was shown via numerical simulation. It also was shown that the estimates of a selected set of unknown inertial parameters will converge to their true values if a certain kind of maneuver trajectory is tracked.