Journal of Aerospace Engineering

Bending fatigue, as one of the impact factors to reduce a metal’s service life, could play a critical role in fatigue life estimation of aerospace components and civil infrastructures, such as in machine gears, splined shafts, and wind-turbine blades, for example. However, the loading effects related to fatigue behavior mainly focus on uniaxial or multiaxial loading conditions, in which the tensile or compressive stress is uniformly distributed. Therefore, bending behavior, including the random tensile/compressive stress and shear stress distribution, may lead to different results. In this paper, experimental investigations are performed on the low-cycle bending fatigue of 1018 steel under different loading effects. Constant deflection tests are carried out to determine the specimen fatigue cycles under different deflections. Three types of variable deflection tests, namely, two-step, periodic overload, and combined step, are conducted to analyze the effects of loading amplitude and sequence with the linear damage rule. The final bending cracks are characterized by scanning electron microscopy (SEM) to further explore crack initiation and propagation during the bending test. The mechanical test results show that a change in flexural strain causes a reduction in fatigue life, showing that damage accumulation and crack propagation accelerate during the bending test. Further, a drop in flexural change from high to low can cause more damage than a rise in flexural strain. Microscale observation shows a complex crack pattern of the primary crack growth path. In addition to the cracks initiating from the specimen boundary, several cracks can be observed forming on the edge of the primary crack, which results from resistance during the propagation of the primary crack.

Air-data, recorded in the flight data of civil aviation aircraft, can be disturbed by wind disturbance, thus leading to measurement error. A two-stage air-data estimation in wind disturbance was studied to obtain accurate true airspeed, angle of attack, and sideslip angle based on flight data. To separate the prevailing wind from the wind disturbance, the first stage involved the preliminary air-data optimization by the immune clone algorithm (ICA) and prevailing wind optimization by the Gauss-Newton algorithm. In the second stage, a new filtering system combining air-data and the von Kármán turbulence model was built with the initial value provided by the first stage. A weighted adaptive extended Kalman filtering (WAEKF) algorithm was proposed, in which an exponential weighting in the innovation covariance matrix was used to reduce the estimation error further. Simulation results indicate that the optimized initial value provided by the first stage is fundamental to ensuring the convergence rate and stability. The WAEKF algorithm with initial value (WAEKF-INIT) can improve the estimation accuracy and alleviate the effects of uncertain measuring noise. A further test with flight data shows that weighted adaptive filtering is capable of reducing the estimation error further in uncertain disturbance.

For aerodynamic shape design using gradient-free searching algorithms, which can be applied more flexibly than gradient-based ones, computing costs increase dramatically with dimensionality. Rational design variables are considered vital to elevate design performance in gradient-free optimization. In this paper, the Bèzier surface free-form deformation (FFD) parameterization based on adjoint surface sensitivity analysis is proposed for aerodynamic shape global optimization. Specifically, FFD point lattice is located where a wide variation is identified in adjoint surface sensitivity. In addition, input space has been adjusted accordingly to enhance space coverage due to the smoothness feature of Bernstein polynomial basis in Bèzier surface FFD. The proposed parameterization was applied to a transonic inviscid drag reduction problem for NACA 0012 with thickness constraints in dimensionality from 5 to 11, and 360.5 counts reduction in drag was achieved. In general, compared with the regularly spaced control lattice, the proposed parameterization effectively weakens the strong shock wave, and drag is decreased considerably using a small-scale sample database.

A rocket-based combined-cycle (RBCC) inlet is always required to start at a relatively low Mach number in consideration of the fuel consumption in the ejector mode and the comprehensive engine performance throughout the flight trajectory. However, the practical start of an RBCC inlet is a dynamic process easily and significantly affected by the embedded rocket operating at a high chamber pressure in the ejector mode. Thus, the real start limits, in terms of the Kantrowitz limit and isentropic limit, and the relevant operational characteristics in the practical applications of a central strut–based RBCC inlet were studied through theoretical analysis and numerical simulations. A two-dimensional analysis model of the RBCC inlet fully concerning the expansion and aerodynamic compression of the rocket jet was established, and the theoretical calculation methods for the inlet start limits were carefully validated and corrected through numerical simulations. It was found that the underexpanded rocket jet expanded rapidly outside the nozzle and compressed the aerodynamic configuration of the flow passage, thereby increasing the actual internal compression ratio (ICR) of the RBCC inlet and increasing the difficulty of the start. The influence rules of the operational pressure and the nozzle expansion angle of the embedded rocket on the RBCC inlet start limits were also theoretically studied. As these key parameters kept increasing, the start limits were increased gradually. Moreover, the isentropic limit was more sensitive to the operational chamber pressure, whereas the increasing rate of the Kantrowitz limit was much more sensitive to the enlarging nozzle expansion angle. Finally, a positive control method based on the rocket operation during the inlet start process was promoted, which could not only guarantee the self-start requirement of the RBCC inlet as the designed scheme in the practical operation, but could also improve the compression efficiency and back-pressure resistance effectively in the subsequent ramjet operation.

Onboard computation of a fuel-optimal steering law is a pivotal technology for future lunar and planetary campaigns with pinpoint landings. This paper proposes a method of downrange estimation based on improved powered explicit guidance (PEG) focusing on the existing throttleable engine. The main contribution includes two aspects. First, the conventional PEG algorithm is improved to enforce prediction accuracy. The prediction of lander cutoff state can be used to estimate downrange position, and thus, the thrust switching time is corrected to recalculate optimal steering law. Second, in consideration of the thrust identification error and navigation uncertainty, the thrust switching condition triggered by the state of the lander instead of time is established. The thrust is switched depending on estimated downrange feedback, and thus, a pinpoint landing with near-optimal fuel is achieved. The effectiveness of the proposed algorithm are demonstrated through Monte Carlo simulations in the presence of thrust deviation and navigation error.

The development of nanosatellites is inseparable from micropropulsion systems, and a micropropulsion system cannot be separated from the evaluation of its micropropulsion performance. Although the ground measurement system is relatively mature, it is difficult to simulate accurately the space environment due to the complexity of the space environment. Therefore, it is necessary to estimate the on-orbit microthrusters. This paper presents an on-orbit thrust estimation system for microthrusters based on a closed-loop attitude control system. The magnetorquers and the microthruster are used as attitude actuators, and the torque generated by the microthruster is used to compensate for the environmental disturbance torque and the known disturbance torque generated by the magnetorquers. In addition, the cubature Kalman filter algorithm based on a gyro-free star sensor was used to estimate the satellite attitude information with high precision. In the on-orbit simulation experiment for microthrusters on the order of tens of micronewtons to tens of millinewtons, the error of 100 μN was 5.67%, and the error of 10 mN was reduced to 2.54%. This method solves the problem of the low accuracy of the traditional on-orbit thrust estimation method when estimating the thrust of a microthruster on the order of 100 μN, and it enables a nanosatellite equipped with the microthruster to complete various space application tasks better.

The path following a finless airship is presented in this paper using a nonunified sliding mode-based lateral guidance and control framework. The outer-guidance loop relies on the airship’s kinematics, and the designed sliding surface is based on the leading angle that generates the heading angle reference commands. For the inner-control loop, the sliding surface is designed based on the yaw error angle generating commands by actuating the propulsion arms. The effectiveness of the proposed framework is validated by using a high-fidelity nonlinear simulation along with actuator dynamics and sensor models having measurement noises and external disturbances. The results demonstrate the effectiveness of the proposed design framework compared to the classical techniques using proportional navigation in the outer-guidance loop and a linear quadratic regulator in the inner-control loop.

Because shock absorbers are an important component of high-rise buildings, it is essential to be able to detect damage to them. Viscoelastic (VE) dampers, as a common shock absorber, directly affect the safety and reliability of VE materials over the entire life of structures. This study is aimed at optimizing a high-dissipation VE material with excellent aging resistance. The VE damper was developed after aging of anti-aging and high-dissipation VE materials. To analyze the damping and mechanical properties of the VE dampers in detail, mechanical performance testing was conducted on VE dampers under various loading conditions, demonstrating that dampers had excellent mechanical properties and confirming the anti-aging properties of the VE materials. In addition, a fractional standard linear solid model based on damage modification is proposed, and aging is considered in the mathematical model of VE dampers as an indispensable factor. After the simulated value is compared to the experimental results, it is found that the fractional mathematical model that involves the aging factor precisely describes the real-time mechanical properties of the VE damper.

A moving mass–actuated unmanned aerial vehicle (MAUAV) is controlled by mass sliders installed inside the airframe and has the advantages of high aerodynamic efficiency and good stealth performance. However, designing a controller for it faces severe challenges due to the strong nonlinearity and coupling of its dynamics. To this end, we proposed an attitude controller based on deep reinforcement learning for the MAUAV. It directly maps the states to the needed deflection of the actuators and is an end-to-end controller. For the sparse reward problem, the reward function required for training is reasonably designed through reward shaping to hasten the algorithm’s training speed. In training, random initialization and parameter perturbation are used to strengthen the final policy’s robustness further. The simulation results tentatively demonstrate that the proposed controller is not only robust but suboptimal. Compared with an active disturbance rejection controller (ADRC) optimized by the particle swarm algorithm, our controller still guarantees a 100% success rate in multiple unlearned scenarios, meaning it has good generalization ability.

As research continues for the first crewed missions to and the eventual colonization of Mars, the need for materials for construction of structural and mechanical components remains paramount. The use of in situ resource utilization (ISRU) techniques is critical due to the financial and physical burdens of sending supplies beyond low-Earth orbit. The Bosch process is currently in development as a life-support system at National Aeronautics and Space Administration (NASA) Marshall Space Flight Center (MSFC) to regenerate oxygen (${\mathrm{O}}_2$) from metabolic carbon dioxide (${\mathrm{CO}}_2$) with the by-product of elemental carbon (C). In this study, the use of this novel C source for alloying with iron (Fe) to produce ferrous alloys was studied to determine if the Bosch-sourced C could provide similar metallurgical results to a traditional C source. The Bosch C was produced by MSFC through their C-formation reactor (C-Fr) using mill scale (${\mathrm{Fe}}_3\mathrm{C}$) as the reaction catalyst. Two types of ferrous alloys were manufactured using the Bosch-sourced C and a control source C. Mechanical and microstructural properties obtained from each equivalent grade revealed comparable metallurgical results.

Estimation of aeroelastic damping for a very flexible high-aspect-ratio wing using wind-tunnel test data is presented in this work through the comparison of four different methods: half-power, half-quadratic gain, an adapted version of random decrement, and the least-square complex exponential method. A model with an aspect ratio of 18.8 was tested in the Cranfield $2.4\times 1.8\text{-}\mathrm{m}$ ($8\times 6\text{-}\mathrm{ft}$) low-speed wind tunnel across a Reynolds number range of $2.33\times {10}^5$–$4.66\times {10}^5$ and dynamically excited using the step relaxation technique. Analysis of the results provides a view of the benefits and limitations of each method, which were seen to differ between 3% and 25% relative to the half-power for the dominant first mode. The trends in predicted damping show clear dependency on freestream velocity. Structural damping of the first mode was also obtained using ground vibration test data to separate the aerodynamic and structural contributions to the overall aeroelastic damping. Structural damping was found to be around 5% and the aerodynamic component was found to add approximately 7.5% at one specific test condition. Although the half-power method is widely used in the community for calculating structural damping, this work shows that when aeroelastic damping is higher than 5%, the other three examined methods yield more consistent results.

Unmanned aerial vehicle (UAV) path planning is an essential branch in UAVs research. This paper presents the hardware implementation of the UAV path planning problem using an improved parallel genetic algorithm (GA) in a multi-microcontroller environment. A controller area network (CAN) bus is a robust bus designed to allow microcontrollers to communicate with each other in applications without a host computer. The CAN bus is used to communicate between the microcontrollers and solve the path planning problem with the parallel algorithm. The data exchange on this network is by the multi-master model, so it is possible to implement an asynchronous and multi-master parallel algorithm using CAN bus. Also, we use the 32-bit ARM Cortex-M3 microcontroller (with CPU clock up to 100 MHz) for hardware implementation. The comparison of both single and parallel GA shows that a multi-microcontroller structure produces better results on the CAN bus, and the parallel version experiences significantly faster speeds than the sequential version.

A modular anti-unwinding dynamic scaling–based immersion and invariance (I&I) adaptive control is devised for rigid spacecraft attitude with inertia uncertainties. It is shown that the parametric regressor matrix cannot be integrable in the attitude dynamics, which results in a nonanalytical solution to the partial differential equations in the I&I controller design. First, in order to overcome the integrability obstacle, the proposed method provides a general and simple matrix reconstruction to make the regressor matrix integrable. Second and foremost, by virtue of a novel modified scaling factor involving saturation function, this paper shows that this method does not require any prior knowledge of the spacecraft inertia matrix and can be conducted without a scaling factor in the controller implementation, which achieves a simpler controller structure and lower dimensional dynamic extension. Moreover, the unwinding problem typically arising in attitude quaternion dynamics is addressed just by the initial value of the attitude quaternion. Finally, numerical simulations are carried out to demonstrate the effectiveness and anti-unwinding characteristic of the proposed controller.

Low-speed laminar flow over a slotted circular cylinder was numerically studied to investigate the slit effectiveness on the vortex shedding suppression and also determine the suppression mechanism physically. The base flow is two dimensional and characterized by $\mathit{Re}=150$. The base slit was cut along the cylinder diameter in parallel to the upcoming flow. The effect of the slit width was studied and compared with previous reported results. Obtained results showed that the lift suppression increased with the slit-width increment so that there was no optimum slit width at this particular flow regime. Considering the strength of the structure, the slit with width of 20% of the cylinder diameter approximately suppresses the lift fluctuation completely and decreases the mean drag coefficient by 5%. Computed results showed that the slit induced additional stagnation points on the surface of the cylinder, which made the primary separation bubbles remain attached to the cylinder. These attached separation bubbles form a stable near-wake and a milder periodic vortex shedding at the far-wake regions. Additionally, other slit configurations were proposed and studied to confirm the proposed physical theorem.

This paper investigated a nonlinear control algorithm for flight dynamics of aircraft using an integrator backstepping–based design. The paper focused on two points. Firstly, a systematic procedure was presented for formulating an integrator-backstepping law for a strict-feedback form of nonlinear dynamic system. In this approach, a set of coordinate transformations including a definition of modified tracking error and state feedback laws is provided to transform a strict-feedback model into a state-decoupled linear model. This process of control design is implemented using a backstepping recursive design in which a backstepping control law is formulated for further study. Secondly, successful applications of roll dynamics of L-59 aircraft model and longitudinal dynamics of F-16 aircraft model were restructured in a strict-feedback form of nonlinear dynamic system under a series of less restrictive assumptions. A direct applicability of the proposed control theoretic framework was used to derive backstepping control laws for the achieved models. To show advancement, improvement, and validation of the proposed control method, both simulation and experimental studies were implemented, which used the numerical simulation of the flight path angle control of the longitudinal motion of an F-16 aircraft model and an experimental study of roll angle control of an L-59 aircraft model to validate the proposed control approach.

An aircraft tire is a kind of rubber product with a complex structure. When subjected to impact load, the mechanical properties of rubber material on strain history and strain rate cannot be ignored. According to the previous impact compression experiment data of rubber materials, this paper built a dynamic constitutive model by modifying Mooney-Rivlin constitutive model to describe the rate-dependent material behavior. The calculated stress–strain values were in good agreement with the experimental ones. The proposed constitutive equation was implemented in ABAQUS via a user-defined subroutine (VUMAT) and used to predict the dynamic response of aircraft tire landing impact in the experiments for the first time. The model’s numerical results considering the strain rate effect are closer to the test results than the traditional model. It shows that considering the strain rate effect of rubber material in the case of tire impact can make the results more accurate.

The present study investigates the efficacy of sinusoidal trailing-edge (TE) serrations as a passive means for the reductions of airfoil broadband noise, theoretically and experimentally. Comprehensive parametric studies were conducted to determine the effect of serration amplitudes and wavelengths on the noise reduction performance of a National Advisory Committee for Aeronautics (NACA) airfoil. Initially, the present paper shows the use of the trailing-edge noise (TNO) model for the accurate predictions of the surface pressure spectrum near the TE and hence the far-field noise using the Wiener-Hopf method. The predicted spectra and the noise reduction levels showed good agreement with the measurements for a wide range of frequencies. The present study reveals that the local maximums of the overall noise reductions occur when the transverse turbulence integral length scale is either 1.2 or 0.2 times the serration wavelength, which corresponds to $\lambda /{\mathrm{\Lambda }}_t=0.833$ or 5, where $\lambda$ and ${\mathrm{\Lambda }}_t$ are the serration wavelength and integral length scale. One of the key findings of the paper is that the serration wavelength at which the highest noise reductions occur when the acoustic emissions vary inversely with the modified Strouhal number ${\mathrm{S}}_{\mathrm{thm}}$ [i.e., ${\mathrm{w}}_{\mathrm{sste}}(\omega )/w_{bl}(\omega )\propto 1{/\mathrm{S}}_{\mathrm{thm}}$] for narrow (i.e., small wavelengths) and wider serrations (i.e., large wavelengths), where ${\mathrm{w}}_{\mathrm{sste}}$ and $w_{bl}$ are the acoustic emissions radiated from the serrated and baseline airfoils. Further, the TE serrations are also observed to reduce leading-edge (LE) noise along with the self-noise, which indicates the efficacy of TE serrations in reducing the total far-field noise.

The purpose of this work is to optimize systematically the maneuver required to identify the wind and calibrate the airspeed sensor of a subsonic aircraft using a global positioning system method. The optimization is based on sensitivity analyses that require a considerable number of flight simulations. To face this challenging computational effort, we adapted and parallelized a particle swarm optimization algorithm. We also introduced a new formulation of the sensor model in the Bernstein form. The results show stability using the selected formulation and bring out nonobvious aliasing and precision loss effects that depend on the maneuver configuration. The knowledge of these effects allowed us to fine-tune the maneuver in order to improve the estimation’s precision. Finally, we validated the method using the JSBSim flight simulator under calm and light turbulence conditions.

Inlet separation is one of the primary sources of distortion at fan entrances. High angles of attack (AOAs) are the critical cause of inlet separation. Optimizing the inlet shape to increase the separation angle is of great engineering significance. However, calculating the separation angles of different inlet shapes requires extensive computation. To address this problem, an ensemble indicator, the antiseparation potential index (ASPI), is proposed in this paper. The ASPI is based on the circumferential distortion index (IDC) of the engine under multiple AOAs. The characteristics of both the separation angle and distortion level are comprehensively reflected in the ASPI. The use of this index can avoid the need for an enormous amount of computation and can better describe the ability of the inlet to resist separation at a higher AOA, both of which are more suitable for inlet design. A Kriging–particle swarm optimization (PSO) design model was applied to tackle the high-cost computational fluid dynamics (CFD) simulations and search for an optimal solution. The results show that the critical AOA at which intake separation occurs is successfully increased by approximately 8°, demonstrating the feasibility of the ASPI in engineering design.