Journal of Aerospace Engineering

This paper presents a new, accurate, and fast methodology, remote sensing satellites coverage analysis (RSS-CA) for determining the coverage area of Earth observation satellites in concurrent design facilities (CDFs) devoted to space mission analysis. The Earth observation areas of interest are discretized by grid points and the coverage surface is computed from the intersection between the grid and the satellite viewing geometry. This geometry is modeled for off-nadir pointing conical and rectangular field-of-view (FOV) sensors and considering a perfectly spherical Earth. To test the RSS-CA methodology, the MARTINLARA mission analysis was selected as a case study. This mission analysis was carried out at the CDF of the Instituto de Microgravedad “Ignacio Da Riva” (IDR/UPM Institute). The results were compared to the ones obtained with the well-known AGI System Tool Kit (STK) software. The high accuracy of the results shows that the developed methodology (RSS-CA) can be simply and effectively applied in CDFs designed for space mission analysis, representing the possibility of a more open and user-friendly environment in relation to other commercial tools.

Fretting wear is a general surface damage form in tightly connected structures. Aimed at the curvic coupling of the disks in combined rotor gas turbine, using the finite-element method and Archard wear theory, the effects of the bolt preload, torque, and transverse vibration on the fretting characteristics of curvic coupling were studied. Additionally, the influences of a single mistuned tie rod and the material properties on the fretting characteristics of curvic coupling have been researched. Finally, a two-disk rotor experiment was carried out to make a comparative study with the simulation. The results showed that with the increase of preload and torque, the contact stress, slide distance, and wear depth gradually increased. Moreover, when the single rod was mistuned, it caused the uneven distribution of fretting characteristics in the circumferential direction. Compared with the Poisson’s ratio and density, the elastic modulus had a notable influences on the fretting characteristics. Comparing the experiment and the transverse lateral load simulation, the simulation results were consistent with the experimental results.

The new generation high-strength steel sheet has been widely used to reduce the vehicle weight. In this study, the mechanical, metallurgical, and fatigue behavior of the resistance spot-welded joints of high strength steel sheets were experimentally investigated. The Dual-Phase (DP) 800 and Twinning-Induced Plasticity (TWIP) 1000 steel sheets were joined together in three different nugget diameters. The welding current was parameterized to create three different nugget diameters. Hence, four groups of welded combining combinations were obtained. The tensile shear test was applied to the samples to study the effects of weld nugget diameter and different welded joint combination in terms of the joint strength. For long fatigue life tests, six test samples were prepared for each test series; in total, 24 tests have been conducted, in which S-N curves of the samples were extracted. Finally, the effect of different nugget diameters on the fatigue life of the welded joint combinations was investigated.

Space robots are widely used for on-orbit capturing. After capturing a rotational non-cooperative target, its momentums produce unpredictable movements for the resultant space robot-target compound system. In order to address this problem, a collision-free detumbling strategy (CFDS) is proposed for space robots with a single 7-degree-of-freedom (DOF) manipulator. First, an acceleration potential field (AccPF) is presented for planning detumbling velocity trajectories, which guide the compound system to terminal rest while guaranteeing collision avoidance in dynamic environments. Energy conservation and acceleration limitations are considered in the trajectory generation. Afterwards, with a radial basis function neural network (RBFNN) approximating the target’s uncertain dynamics, an adaptive control scheme is established to precisely track the planned detumbling velocity trajectory. Furthermore, system stability is certified by a Lyapunov function. The effectiveness of the CFDS is eventually validated by simulations.

This paper presents an adaptive fuzzy sliding mode controller (AFSMC) for the control of a nonlinear aeroelastic wing section. The proposed controller estimates both the model dynamics of the system and the switching control of the sliding mode. Furthermore, the designed controller’s strategy is to enhance quality and ensure flight safety by providing additional damping to the already stable system and increasing the flight speed beyond the open-loop flutter speed. The aeroelastic system describes the plunge and pitch motions of the aircraft wing-section equipped with leading and trailing edge control surfaces. Additionally, the selected two-degree-of-freedom model includes structural stiffness nonlinearities and quasi-steady aerodynamics. The simulation results show the effectiveness of the proposed controller toward suppressing flutter and LCOs and reducing the vibrational level in the subcritical flight speed range. Additionally, the designed control strategy effectively controls and derives the state trajectories to the origin despite uncertainties and gust loads.

The characteristics of unsteady boundary-layer transition on a small-amplitude pitching NACA0012 airfoil are investigated using the $\gamma -\overline{{\mathrm{Re}}_{\theta t}}-A_r$ transition model, which consists of the $\gamma -\overline{{\mathrm{Re}}_{\theta t}}$ transition model and the roughness amplification factor ($A_r$) transport equation. The present transition model is validated by three roughness surface cases, including a zero pressure gradient flat plate, the NACA0012 airfoil, and NREL-S814 airfoil with different roughness surface. The numerical results for these cases are in agreement with experimental data. Three pitching NACA0012 airfoils with smooth surface, fully distributed roughness surface, and roughness leading edge are simulated by the present model. The impact of reduced frequencies ranging from 0.019 to 0.301 is also taken into consideration. The results show that as the reduced frequency increases, the transition location moves downstream in the upstroke and upstream in the downstroke. With the influence of distributed roughness, the transition moves upstream and the time delay of the transition location is decreased. Moreover, the roughness leading edge enhances the asymmetry of the propagation speed and makes it more sensitive to the reduced frequency.

To reveal the effects of random rib assemblage errors caused by misalignment on the performance of umbrella antennas, a novel computational method is proposed to evaluate the effects of random rib assemblage errors on surface accuracy, and is further extended to assess the gain loss of electromagnetic performance. Based on the assumption of equilibrium equation of cable-nets and small amplitude deformation, the computational formula of surface accuracy with respect to random rib assemblage errors is derived. With the help of Ruze’s formula, the computational formula of gain loss with respect to random rib assemblage errors is obtained. Taking a typical umbrella antenna as a simulation example, the effects of random rib assemblage errors on antenna performance are investigated and verified by the Monte Carlo method. Simulation shows that the proposed method can predict the degradation of both structural and electromagnetic performance for umbrella antennas accurately with small amplitude deformation, and improve the computational efficiency with satisfactory accuracy, which will benefit the design of umbrella antennas.

Cable length adjustment is a critical step to minimize the shape error of cable-net reflectors. The adjustment requires tedious repetitive work, but the surface accuracy is often unsatisfactory because of inevitable operating error and modeling inaccuracy. To solve this problem, this work proposes a target-approaching and procedural-learning method. At the initial stage of the adjustment process, the target-approaching method is proposed to ensure the shape error is improved and that adjustment data are generated. Then an online prediction model is built by sensitivity analysis and least-squares support vector machine. On this basis, a multidimensional forward-backward algorithm is applied to calculate the optimal adjustment amounts in order to achieve fast convergence of adjustment accuracy. Finally, the proposed method is verified by both numerical simulation and prototype experimentation. The results show that the proposed method can well reflect the adjustment characteristics of cable-net reflectors, effectively improve shape accuracy, and greatly shorten adjustment time.

In turboprop engines, the lip of the S-shaped intake is behind the propeller rotation plane, so the complex slipstream flowfield has a strong impact on the flowfield in the S-shaped intake and affects engine performance. The present work attempts to improve the flowfield in the S-shaped intake in turboprop engines by using vortex generators. The research was carried out by solving Reynolds-averaged Navier–Stokes equations employed SST $\mathrm{k}\text{-}\omega$ turbulence model. An attempt was made to explore the vortex generators’ effect on S-shaped intake performance in the propeller slipstream. The results show that the total pressure recovery coefficient and distortion coefficient change periodically with time for the S-shaped intake in the propeller slipstream. For S-shaped intake with vortex generators ($s/s_m=0.3$, 12-mm height, 18° with the incoming flow) in the propeller slipstream, the vortices induced by the vortex generators are coupled with the propeller slipstream, which effectively inhibits flow separation in the S-shaped intake for turboprop engine and improves air quality on the aerodynamic interface plane (AIP) section of the S-shaped intake.

Although the increase of the width ratio of a radial diffuser generally leads to a more unstable condition, the effect of volute was seldom considered in previous experimental and numerical studies because of the mismatching between the changed diffuser and the original volute. In this work, vaneless diffusers with different width ratios were theoretically and numerically investigated. In particular, a turbulent stability analysis based on different flow planes was performed using a spectral collocation method. To consider the combined impact of width ratio and volute, the base flow of a diffuser with a numerical volute imposed at the outlet was computed and validated against experimental data. The sensitivity of the eigenvalue was further obtained to localize the wave-maker region of the diffuser by means of a continuous adjoint procedure. The results indicated that the asymmetric flow field could be well recovered by the numerical volute. With a small width ratio of 0.06, the stability margin of the compressor was not significantly widened with a stall first occurring in the impeller, although the diffuser becomes more stable. The asymmetric flow condition led to the regional enhanced global modes in the circumferential direction for the diffuser with middle and large width ratios. The high-sensitivity region of the diffuser with a middle width ratio lay at the 90°–160° position in a scattered form, indicating that the combined effect of the inflow wake and outlet reverse flow near the shroud wall were crucial for flow instability. However, for a diffuser with a large width ratio, the dominant factor in instability is the outlet reverse flow due to the delay of the separation point.

In order to reduce the population of space debris in geosynchronous orbit, a tethered-Coulomb formation (TCF) is proposed that consists of a space tug and a cluster of charged debris connected by elastic massless tethers. The charged debris introduces repulsive coulomb forces to avoid collisions among the pieces of debris. In this way, the system is able to maintain a stable configuration when proper system parameters are given. To determine this relationship, a dynamics model of the formation was derived by using Lagrange’s equation. Then, precise and approximate equilibrium solutions of the debris item charges and tether lengths were derived analytically according to the different number of debris items in pyramid configurations of the TCF. Based on the analytical results, pyramid and double-pyramid TCF configurations were studied via numerical simulations. The results showed that oscillations of the debris had little impact on the motion of the tug. When designing a double-pyramid TCF system, two clusters of debris with large relative positions in different phases would show good performance in a deorbit mission.

Straight-through labyrinth seals are reliable noncontact seal structures widely used in aeroengines. Windage heating and swirl development of labyrinth seals are of great significance to the design of secondary air systems. In the operation process of an aeroengine, a labyrinth seal is supposed to experience uneven temperature distribution, which may cause deformation and resulting nonuniformity along the flow direction. The nonuniformity coefficient was defined to characterize the degree of nonuniformity so that the influence of nonuniform clearance on windage heating and swirl development can be analyzed quantitatively. Windage heating and swirl development were analyzed from the perspectives of different Reynolds numbers, pressure ratios, circumferential Mach numbers, and dimensionless minimum tip clearances. Simulation results showed that when other dimensionless parameters are the same, both windage heating number and exit swirl ratio decrease with the increase of the nonuniformity degree. With the same nonuniform degree, the decrease rate of divergent clearance is larger than that of convergent clearance. The changing trends of the decrease rate of both windage heating number and exit swirl ratio were analyzed in detail. It can be concluded that the decrease rate of the exit swirl ratio reaches the peak when the dimensionless minimum tip clearance is about 0.067. This result makes it clear that the effect of nonuniform clearance is indispensable for the safety assessment of aeroengines.

This paper proposes a sliding mode controller for a novel kind of quadcopter called tilt-rotor quadcopter. The controller is modeled in SIMULINK and implemented on a Pixhawk 2 flight controller board. First, this paper presents the dynamic model of the tilt-rotor quadcopter. Second, the proposed smart sliding mode controller for simulation studies is presented. Third, the hardware platform is discussed in detail with the controller implementation. Finally, numerical simulation results for the smart sliding mode controller for various scenarios (wind disturbances, faulty motors) are presented following which the preliminary hardware tuning results are presented. Simulation results of the tilt-rotor platform hovering at a position with a nonzero attitude are also shown. This paper also presents a discussion on the challenges faced transitioning from a simulation environment to hardware testing and plausible ideas to address the challenges.

Avian flight has inspired scientists and engineers for a long time. Flapping-wing micro air vehicle (FMAV) is a kind of air vehicle mimicking the avian flapping flight that has achieved great success in this century. However, in some complex flight conditions, its flight performance is not as good as that of birds. Birds have evolved many unique features to adapt to different and complex situations. The leading-edge alula is considered a typical structure that can enhance birds’ flight capabilities in slow-speed flight. In this paper, we numerically investigate the effects of bionic alula on the aerodynamic performance of flapping airfoil under different flapping kinematic parameters. The influence of different pitch amplitudes, plunge amplitudes, and mean angles of attack (AoAs) of freestream are considered. Results show that the deflection of the alula will suppress the flow separation and restrain the formation and development of the leading-edge vortex (LEV). This will cause a significant discrepancy in aerodynamic performances between the airfoil with and without bionic alula. Observing their different flow structures, the time-averaged lift coefficient of the single flapping airfoil without alula first decreases then increases with the increasing pitch amplitude, but almost increases continuously with the rising plunge amplitude. However, as for the airfoil with alula, the time-averaged lift coefficient first rises then drops as the pitch amplitude increases, and also first rises then falls with the increasing plunge amplitude. The maximum enhancement of the alula on the time-averaged lift coefficient can reach 82.2% under the optimum kinematic parameters. The bionic alula can also delay stall, and the stall angle postpones by 7° in the scope of our research.

For trajectory planning problem with multiconstraints of free-floating space robots (FFSR), the optimal solution may be unavailable due to strict constraints. In order to avoid identify a solution and obtain an executable joint trajectory, this paper proposes a trajectory planning method based on dual-mode switching. Using this method, the planner can automatically switch between the optimal mode and the effective mode, and obtain a feasible joint trajectory that satisfies the task requirements when the end-effector tracking error constraint and the joint torque constraint are considered. In this paper, the multiconstraints model of FFSR, which mainly consider control and dynamic constraints and the trajectory planning model based on dual-mode switching, are established, and the corresponding solving method is presented. The simulation results demonstrate that the proposed method is effective and feasible.

Guided waves play an important role in practical structural health monitoring (SHM) systems. Due to the complexity of irregular cross sections, there has been little research on guided wave modal control done using excitation and sensing distribution models. In this paper, the semianalytical finite-element (SAFE) method for the quick and reliable solution of guided wave excitation and sensing in irregular cross sections is developed, and modal control method is proposed by optimal transducer design. The free-vibration and forced solutions to a pin-force excitation in an I-type bar are determined up to a frequency of 150 kHz. The effects of four loading and sensing cases are discussed, and the time–transient responses of four cases are calculated. The maximization of the signal amplitude of the desired modal signal and the suppression of other modes can be achieved with different excitation and sensing distributions. The experimental investigation of an I-type bar is shown to validate the reasonability of the modal control method. Finally, a feasible optimal and economical excitation/sensing transducer design method is presented and discussed.

Based on smooth particle hydrodynamics (SPH) method, a model of impact response of tire burst pressure effect model on rear wall panel of landing gear cabin was established. The model was tested for air pressure distribution and airflow impact on the panel according to airworthiness regulations of the breach shape. The air pressure distribution on the panel after tire burst and the bearing capacity of the panel to the airflow impact were studied through test and simulation. Ultimately, it is found that the impact of tire burst airflow can lead to irreversible damage such as deformation of landing gear bulkhead ribs and bolt failure. The SHP simulation method proposed in this paper has a good explanation for the tire pressure effect threat model. Test and simulation can not only meet the requirements of airworthiness certification, but also provide the basis for structural design and strength analysis.

This paper presents the results of experimental and numerical investigations on synthetic jet actuators with various arrangements. First, an actuator with one membrane in the cavity is a standard, reference arrangement; and second, one novel modified arrangement has two membranes in the cavity. Authors use constant temperature-anemometry measurements to identify the characteristic frequencies in which the output jet velocity reaches a maximal value, namely, the resonant frequency of the actuator (Helmholtz frequency) and of the membrane (structural resonance). A numerical model of a synthetic jet actuator has been validated using experimental data for the case with one membrane in a cavity. The simulation results confirm the existence of a toroidal vortex at the actuator exit for the frequencies determined in the experiment. Numerical results of a modified actuator with two membranes in the cavity give promising results in the context of separation control by synthetic jet devices, i.e., increase of output velocity and vorticity from the same occupied space is obtained. Numerical simulations show that in the modified actuator with two membranes in the cavity, the jet output velocity and vorticity of a created toroidal vortex at the actuator outlet can be increased by 80%, compared to the case with one membrane. The new geometry of a synthetic jet actuator with perpendicular membranes is to be implemented into the wind tunnel measurements to prove the use of a modified synthetic jet actuator as a device applicable for active flow control.

The development and calibration of a micronewton space thruster test stand is described. The test stand uses a torsion pendulum-type design. A calibration system consisting of a rarefied-flow, gas-dynamic thruster was used to provide a known reaction force to deflect the torsion pendulum arm. Under-damped oscillation dynamics of the arm are described. The novel displacement-sensing system consists of laser light reflected from the arm and recorded using a linear charge-coupled device (CCD). Calibration thruster pressure is transmitted via pulsed laser to an external receiver. Four experimental tests are discussed characterizing the displacement of the arm during active thrusting periods. The time-response of these displacements are used to experimentally infer the rigidity of the spring system. Micronewton-range thrust values on the order of 1 to 10 micronewtons were estimated to be produced by the calibration thruster. Additionally, the observed torsional rigidity was 1 to 3 times higher than the expected value based on the physical properties (i.e. wire diameter and material) of the system, reducing the precision of the torsion pendulum in its present form. Opportunities for further study and error reduction are also proposed.

This study presents a finite-element (FE) model considering intralaminar damage and interlaminar delamination to predict the biaxial compression behavior of carbon fiber–reinforced composites (CFRCs) with low-velocity impact damage. First, the FE model was validated by a uniaxial compression experiment. The FE predictions in terms of compression after impact (CAI) strength and failure patterns were consistent with the experimental results. Then the validated FE model was used to predict the biaxial compression behavior of the CFRCs. It was found that the residual strength of the CFRCs was dependent on the buckling mode of the laminate due to the coupling effect of the longitudinal and transverse compressive loadings. Correspondingly, the CAI strength under biaxial compression was lower than that under longitudinal or transverse compressive loading.

This paper summarizes the activities performed for improving the understanding of the whirl flutter instability and its most sensitive parameters. Clearance from such aeroelastic instability is required by airworthiness rules for turboprop-powered aircraft and must be cleared for tilt-rotor vehicles. It may be also a required check to be carried out for windmilling propellers. In this article, the design, development, and testing of a small demonstrator are presented and results are discussed with reference to existing literature. The initial focus is on the governing equations and identification of parameters that affect the aeroelastic behavior of the propeller-engine installation. Ground vibration tests, at zero airspeed, have been performed to validate numerical models for the mass/inertia and stiffness distribution. Several measurements have been further carried out at different configurations, changing the airspeed, the stiffness, the mass, and the propeller in order to setup a sensitivity analysis and create a database for further improvements, mainly for a basic computation of unsteady airloads of the propeller. These analyses confirm expected results, allow interesting and confident conclusions and are able to help for showing compliance with current airworthiness rules for propeller-driven aircraft. The results presented in this paper will also represent the basis for asserting the range of validity of scaling effects for realistic applications.

The present paper concerns the introduction of a mixed analytical-numerical method for the estimation of the maximum lift coefficient of box-wing aircraft in unflapped configuration. The analytical aspect is related to the adaptation of the method included in the United States Air Force (USAF) Stability and Control Data Compendium (DATCOM) by means of an approach built on the characteristics of the optimal lift distribution of the box-wing and implemented through simplifying assumptions. Since the formulation depends on parameters of the considered aircraft, numerical simulations are performed through a vortex-lattice method to complete the input data set. The method is first presented and then validated for the case of the 2-seater amphibious PrandtlPlane (PrP) and is then applied to two test cases: a 300 passenger midrange PrandtlPlane and a regional hybrid-electric PrandtlPlane. Results are presented and discussed, and the links between the proposed method and relevant parameters of the box-wing design, such as taper ratios and wing loading repartition among the two wings is analyzed.

A one-equation Reynolds-averaged Navier-Stokes (RANS) closure model has been established for bypass transition and laminar separation bubble (LSB)–induced transition. A new local indicator is proposed to describe the influence of turbulence intensities and pressure gradients, which makes the model Galilean invariant. Based on this new indicator, a novel and efficient transition criterion is formulated. For LSB-induced transition, an empirical correlation is developed to modify the intermittency factor and control the size of separation bubbles. Several flow cases, including flat plates with various pressure gradients, flat plates with LSBs, semicircular leading-edge flat plates, National Advisory Committee for Aeronautics (NACA) 0018 airfoil, and SD7003 airfoil, are employed for the model verification. The predictions show good agreement with the experimental data and large-eddy simulation data for different inlet conditions, which indicates the increment of free-stream turbulence intensity leads to the reduction of the size of laminar separation bubbles.

An electrohydrostatic actuator (EHA) system is one of the key components of electric aircraft flight controls. The EHA system with fixed pump displacement and variable motor (EHA-FPVM) is widely used because of its simple structure, but the power weight ratio of EHA-FPVM is small, although it is very important to an aircraft. A novel EHA with a constant-torque variable-displacement pump (CTVDP) is proposed and simulated; it can substantially decrease the maximum motor-rated torque while reducing the servo-motor volume and weight. First, the CTVDP structure of the EHA system is proposed and a constant-torque variable-displacement pump is designed. Second, mathematical models of the EHA system structure are developed and simplified according to the constant-torque displacement–pressure law. Third, an adaptive backstepping control algorithm is proposed to improve position-tracking performance in the presence of the uncertain viscous damping coefficient. Finally, numerical simulations prove that the EHA system with CTVDP structure can further decrease motor maximum torque by approximately 37.5% and is robust against load impact, pump displacement variation, and parameters uncertainty. The results of this study can serve as the foundation for future application.