Journal of Aerospace Engineering

The emergency extension system for nose landing gear (NLG) is vitally important for safe landing in the event of hydraulic source failure. To study the dynamic properties of NLG emergency extension, it is necessary to explore the aerodynamic load in detail, which is a potential adverse factor that hinders the successful deployment of NLG. The emergency extension failure of a door-linkage NLG mechanism was studied from a novel bifurcation perspective in this paper. NLG models (excluding and including aerodynamic loads) formulating the mechanism as a series of steady-state constraint equations were derived. Solutions to these equations were continued numerically, and the NLG retraction–extension cycles in different models were compared. The critical condition for smooth extension was obtained based on the bifurcation analysis. The failure principle was explained, that is, the disappearance of bifurcation points characterizing the successful locking. The variations of bifurcation point as a function of spring parameters and door aerodynamic loads were investigated. An effective solution to the failure was put forward in which the unlocking force is reduced by 33.0%.

The mobility system is an important part of the Mars rover and is one of the key subsystems for the Mars exploration mission. In this paper, the durability performance of China’s Mars rover was tested and analyzed on the simulant Mars terrain. The total traveling distance of the rover for the durability test was 5,040 m, the test results showed that the average power consumption was 23.22 W, and the maximum power consumption was 75.9 W. The average power consumption of the rover increased linearly with the increase of the rover’s velocity. Compared with the power consumption at $50\mathrm{m}/\mathrm{h}$, the power consumption at 100 and $200\mathrm{m}/\mathrm{h}$ increased by 42.71% and 138.54%, respectively. The type of Mars terrain also has a great influence on the power consumption of the rover, the rover has the greatest resistance and the highest power consumption when driving on Terrain type D (Mars soil simulant + rocks), and wheel abrasion mainly occurred in this type of terrain. Finally, there were many scratches on the surface of the wheels, and a small amount of Mars soil simulant was deposited in the drum. However, the mobility performance of the rover did not decrease after traversing 5.04 km. The mobility system and the suspension are in good condition, and the motor and control system works well.

In this work, an implicit finite-volume-element (FVE) method is extended to efficiently simulate the vortical structure of unsteady turbulent flows in domains with moving meshes. The arbitrary Lagrangian-Eulerian (ALE) approach is used to consider the motion of a hybrid mesh distributed in the solution domain. Conventional turbulence models are applied to simply confirm the sample achieved efficiency and accuracy in solving complex turbulent flow domains with moving boundaries. In this regard, the advective terms in the Navier-Stokes equations, including those in the transport equations for the applied turbulence models, are treated in a rather innovative manner. In other words, an advanced physical influence scheme (PIS) is suitably introduced in the context of extended ALE formulations. The accuracy and efficiency of the extended method are carefully evaluated by simulating various turbulent flows, including the fluid flow in stationary domains, separated turbulent flow over a bluff body problem, and the dynamic stall of fluid flow over a flapping airfoil. Comparing the current solutions with experimental data, it is shown that the current PIS-ALE method provides better accuracy and efficiency than those of past numerical methods, which used similar turbulence models in their algorithms.

Many fields, including the aerospace industry, have shown increased interest in the use of plastics to lower the mass of systems. However, the use of plastics in space can be challenging for a number of reasons. Ultraviolet radiation, atomic oxygen, and other phenomena specifically associated with space cause the degradation of polymers. Here we show a path toward creation of space-grade components by combining additive manufacturing (AM) and atomic layer deposition (ALD). Our method produced ALD ${\mathrm{Al}}_2{\mathrm{O}}_3$ coated thermoplastic parts suitable for space applications. The highlight of this work is a significant reduction in outgassing, demonstrated using residual gas analyzer (RGA) sampling. Compared to uncoated parts, the ALD ${\mathrm{Al}}_2{\mathrm{O}}_3$ coating decreased the outgassing of polyether ether ketone (PEEK), acrylonitrile butadiene styrene (ABS), polycarbonate (PC), and nanodiamond-doped polylactide (ND-PLA) by 46%, 49%, 58%, and 65%, respectively. The manufacturing method used in this work enables the use of topology optimization already in the early concept creation phase. The method is ideally suited for spacecraft applications, in which the volume and mass of parts is critical, and could also be adapted for in-space manufacturing.

An ionic wind pump consisting of simple concentric cylinders was developed for future Mars exploration. The ionic wind pump uses gas flow driven by the corona discharge generated at the edge of the inner cylinder electrode. It is suitable for applications such as air circulation in a module and heat transfer of equipment. The relationships between the applied voltage and the corona current, generated gas flow rate, and pressure rise were measured with reduced pressure of air and carbon dioxide gas that simulated the Mars environment. A substantial gas flow was generated at atmospheric pressure, but the pressure rise was less at low pressure. The performance was improved by adopting a multistage configuration. Some effective methods to measure gas velocity in a low-pressure environment were also developed. Numerical calculations were performed for the corona discharge field and the ionic wind induced by the migration of ions generated by the corona discharge. The experimental results agreed well with the numerical calculations. The pump can be effectively applied for Mars exploration because it is extremely simple, consumes less power, and has no mechanical moving parts.

A coupled analysis of impact damage and compression-after-impact (CAI) strength of laminated composite plates was conducted using efficient analytical and semianalytical models with the goal of demonstrating and validating the overall coupled impact/CAI strength model. Within this coupled framework, the impact model determines the transient response of the plate subjected to low-velocity impact, followed by an estimation of impact damage using a combination of strength and energy-based criteria. The predicted impact damage is idealized and used in conjunction with analytical buckling and postbuckling or kink-band analysis to predict the residual compressive strength. The coupled analysis approach was validated by applying it to impact and compression-after-impact experiments conducted on AS4/8552 laminates. It was shown that the coupled analysis model is able to predict accurately the compression-after-impact strength for a range of impact scenarios and material layups in a fraction of the time required by detailed finite-element models.

Thermal protection systems (TPSs) and drag reduction have already garnered considerable attention in the field of hypersonic vehicles, and numerous research efforts have aimed to develop active control technology. In this study, a novel dual-jet strategy is proposed to enhance drag and heat flux reduction capabilities in hypersonic flows. First, an in-house code was validated sufficiently in simulating a hypersonic flow field. Then the flow features dominated by the novel dual-jet and the drag and heat flux reduction mechanism were discussed thoroughly. Based on that, the effects of the spike length and rear jet pressure ratio on the flow field are thoroughly investigated. The obtained results indicated that this novel strategy achieves excellent drag and heat flux reduction performance. The flow structure changes remarkably and the drag on the blunt body changes nonmonotonically once the spike length varies. Larger rear jet pressure can achieve better drag and heat reduction performance. Further, the way in which the rear jet affects the drag reduction and thermal protection is numerically analyzed. Interestingly, the constant mass flow rate of the rear jet, flow structure, and the drag and heat flux reduction capability are highly dependent on the rear jet pressure ratio, and lower rear jet pressure is conducive to drag reduction and thermal protection of hypersonic blunt vehicles.

The response difference between the classical mathematical model of a ducted rocket and a real experiment is compared in this study, and a new revised method is proposed based on the free volume. A virtual free volume that is related to the real free volume is built to eliminate the discrepancy between the model and experiment, and a possible explanation of virtual free volume is carried out. In addition, design constraints and variation characteristics of the ducted rocket control system are analyzed. Furthermore, a radial basis function (RBF) adaptive control method is proposed to deal with the variation. Finally, hardware in a loop simulation device for the ducted rocket control system is presented, and its reliability is proved by analysis and comparison with a direct-connected ground test, and a serial simulation test that consistent with the numerical simulation is carried out to verify the validation of the controller.

This paper is concerned with the flow characteristics and noise generation at the finite wing–wall junction. To characterize junction flow noise, acoustic measurements were taken in the acoustic wind tunnel at the Brandenburg University of Technology in Cottbus, Germany, with a planar 47-microphone array at a chord-based Reynolds number ($R{e}_c$) of $2.3\times {10}^5$. The wings used have an aspect ratio (AR) of 2 and six different section profiles with variations in camber and thickness. The results show that the junction noise dominates at the low-frequency range below 4 kHz (chord-based Strouhal number of $S{t}_c=5.6$), which is particularly obvious for the symmetric wing. On the other hand, the addition of leading-edge bluntness and camber are found to have less of an influence on the junction noise spectra. Both symmetric and nonsymmetric wings near the stall angle exhibit a dramatic increase in the high-frequency junction noise content above 8 kHz ($S{t}_c=11.2$), indicating the existence of a different junction noise mechanism. To provide insight into the junction flow, measurements of the mean streamwise total pressure field within the wing–wall junction were performed in the acoustic tunnel at the University of New South Wales for two wing ARs of 0.2 and 1.0 at $R{e}_c=2.6\times {10}^5$ and several angles of attack (AoA). The contour maps of the mean streamwise total pressure field indicate the existence of the horseshoe vortex around the wing and exhibit variations of the flow structure at different streamwise locations and AoA. Further, the AR influences the upwash and downwash near the junction due to the effect of the tip vortex. These competing effects are observed on different sides of the wing under different lifting conditions and AR. The relatively simple geometries of the test cases and the complex physics they create make this data set particularly valuable for validation of numerical simulations and models.

Extended-duration missions to the Moon require the construction of long-term infrastructure that protects inhabitants and protects life-support systems from the Moon’s hazardous environment. Notable hazards of the lunar environment include ionizing radiation, extreme thermal cycles, and hypervelocity micrometeoroid impacts. Of these hazards, micrometeoroids are a top research priority because they are stochastic in nature and arrive without warning. To design structures that will be impacted by randomly distributed micrometeoroids, a risk-based approach should be used. This paper describes the development of a risk-based design framework termed performance-based impact engineering. This proposed framework allows the risk of micrometeoroid impact to be considered in the design of lunar infrastructure by developing site-specific hazard models and material-specific damage models. The methodologies for micrometeoroid hazard estimation on the lunar surface and material damage prediction are described. A case study for the design of a tiled lunar landing pad using the performance-based impact engineering framework is demonstrated.

Large eddy simulation (LES) with Ffowcs Williams and Hawkings (FW-H) analogy is performed to investigate the flow structure and noise emission of NACA 65(12)-10 aerofoil with and without a serrated trailing-edge at a moderate Reynolds number (${\mathrm{R}}_{\mathrm{ec}}=5\times {10}^5$). Following, the noise reduction mechanisms and the noise source regions are analyzed in detail. Results indicate that trailing-edge serrations reduce the overall sound pressure level by up to 6 dB at a 0° direction angle, and the noise reduction mainly occurs at a frequency lower than 10 kHz. The trailing-edge serrations can act as vortex generators and, thus, reorganize the vortex structure, altering the spanwise vortex to a large-scale streamwise vortex. Besides, serrations reduce the magnitude and modify the distribution of the power spectrum density of pressure fluctuation in the near wake, and the spanwise coherence magnitude is decreased. Additionally, the phase interference of acoustic sources at different spanwise positions is enhanced. Wavelet analysis reveals that the hump of the broadband noise spectrum of the serration aerofoil is mainly contributed by the serration tip region. The conclusion provides inspiration for further noise reduction by trailing-edge serrations.

This paper focused on the aerodynamic optimization and noise reduction of a two-stage series compact fan. The space constraints of compact fans limit the design space, which poses several challenges to aerodynamic and noise performance. The aerodynamic performance and noise characteristics of a baseline series axial fan show that the interaction between the first-stage and second-stage fans is weakened because of the effect of the honeycomb panel. A dual-objective optimization of the rotor blade with the power consumption and static pressure efficiency as the design objectives was carried out on a single-stage axial fan based on a radial basis function (RBF)-enhanced differential evolutionary (RADE) algorithm. The result showed that the dual-objective optimization achieved good results in the aerodynamic design of the two-stage series fan. Compared with the baseline configuration, the static pressure efficiency of the optimized axial-flow fan increased by 6%. The first blade-passing frequency (BPF) noise was reduced by 7 dB. An expansion shroud was designed based on the optimized design of the axial-flow fan. The expansion shroud configuration further increased the static pressure efficiency by 10%.

The coupled problem of aerodynamic heating and structural heat transfer occupies a very important position in the field of aerospace engineering applications because it directly affects the accurate prediction of aerothermal loads and structural deformation. This paper develops a fluid-thermal-structural coupling framework for the investigation of aerothermalelastic problems in hypersonic flow. A loosely coupled analysis strategy equipped with both the constant and adaptive coupling time step size approaches is adopted to integrate an in-house developed computational fluid dynamics (CFD) code using the finite element solver Abaqus and to perform coupling simulations based on CFD/computational thermal and structural dynamics (CTSD). The accuracy, reliability, and capability of the aerodynamic heating and fluid-thermal-structural coupling analysis methods in this framework have been validated by a spherically blunted cone and a cylindrical leading-edge model in a hypersonic environment. A typical low-aspect ratio hypersonic wing is adopted as the computational model to study in detail the impact of sustained aerothermodynamic loads on the aeroheating process, structural deformation characteristics, and aerodynamic performance. The results indicate that the aerodynamic heating effect obviously weakens the structure stiffness and, thereby, directly leads to a significant increment in wing structural deformation. Consequently, the pressure distribution and aerodynamic coefficients of the wing also change significantly after aerothermoelastic deformation. Therefore, the influence of aerodynamic heating on the aerothermoelastic behaviors of a hypersonic wing should be considered seriously in the design stage to avoid unaccepted structural deformation and aerodynamic loss in real flight. Moreover, the loosely coupled analysis strategy equipped with the adaptive coupling time step size approach can be used as a highly efficient simulation method for practical fluid-thermal-structural coupling problems.

In this paper, a systematic approach is considered for the development of a low noise shock tunnel facility. For this purpose, an optimal hypersonic nozzle and test section configuration is presented along with a previously developed low noise shock tube design. In hypersonic experimental studies, one of the most important requirements is a low noise test section with a high-quality uniform flow. The main sources of perturbations are acoustic fluctuations that occur through the turbulent boundary layer and Mach line fluctuations with the passage of turbulent flow through the hypersonic nozzle. The fluctuations are correlated to wall boundary layer thickness, radiated disturbances, and their concentration on the nozzle axis. These parameters can be indirectly controlled via definition of three weighted objective functions—minimum total pressure loss, uniform Mach number distribution, and minimum axial flow deviation—combined to attain the final scalar objective function. Then, a modern optimization strategy is implemented based on a genetic algorithm, parallel CFD solver, and the requirements and constraints from conceptual and preliminary design. In this way, parameterization of the overall nozzle contour is performed with a few control points and Bezier curve that showed good flexibility for generating appropriate nozzle curves. Design objectives are evaluated using a Navier–Stokes solver with a k-$\omega$ turbulence model. Various geometrical and physical constraints such as nozzle length, throat area, inlet and outlet diameters, and inlet boundary conditions are considered. It is concluded that the proposed strategy for tuning the nozzle convergent-divergent contour minimizes the boundary layer effects and shows a significant improvement in the quality of test section flow and consequently a reduction in the noise level of shock tunnel test facility.

Passive air-jet blowing is an effective yet simple technique to control flow-induced noise due to vortex shedding of bluff bodies. The current study investigates the effect of the specific slot angles of passive jets in connection with the suppression capabilities of vortex shedding and the byproduct of vortex-induced noise in the wake region of a circular cylinder. Aeroacoustic tests for a baseline case and 10 different slotted cases with slot angles of $80°\le \theta \le 125°$ are performed for Reynolds numbers $6.6\times {10}^3\le \mathrm{R}\le 3.3\times {10}^4$. This is supplemented by numerical computational fluid dynamics (CFD) analyses to identify the underlying aerodynamic mechanisms. The results obtained reveal that using the current passive control method results in a significant reduction of the vortex shedding tonal noise for slot angles of $90°\le \theta \le 125°$ and high Reynolds numbers. The numerical results showed good agreement with a remarkably reduced kinetic energy for slot-end angles of $115°\le \theta \le 125°$. At low Reynolds numbers, however, the identified aeroacoustic benefits tend to cease.

In this paper, the effect of the resonator cavity shape on a constant volume is investigated for the aeroacoustics performance of a Helmholtz resonator system. This scrutiny is performed for various Mach numbers of the grazing flow in the range of 0–0.4. In this regard, a three-dimensional numerical approach was used. In the work, the resonator shape of sphere, cubic, cone, triangular prisms and cylinders were considered. The numerical approach was first verified with the experimental measurements available in the literature. The assessment of the numerical approach by comparing the numerical results with the available experimental data indicated good accordance, which, in turn, affirmed the validity and competency of the method. The simulation results show that by changing the resonator cavity shape, a maximum change of 24% occurred in ${\mathrm{TL}}_{\max }$. For all cavity shapes, increasing the grazing flow Mach number $M_a$ from 0 to 0.4 led to about a 90% reduction in ${\mathrm{TL}}_{\max }$. Furthermore, a 30% change in the resonant frequency happened with changing the resonator cavity shape. In addition, in $M_a\le 0.1$ and $M_a\ge 0.2$, for all the resonator cavity shapes, the increment trend occurred for the resonant frequency, and in $M_a=0.15$, all cavity shapes had a span for the resonant frequency except cubic. Overall, spheres and cylinders as the resonator cavity had better performance with respect to other considered resonator cavity shapes.

The control of aeroelastic response of a wing section with influence of the fuel sloshing in external tank is examined through analytical studies. Based on the simplified mass-spring-damper equivalent mechanical model of the liquid sloshing and the dynamic modeling method of the wing/store system, the aeroelastic model of the two-dimensional wing/external tank/fuel sloshing system is established in this paper. The aeroelastic instability of the system is studied, and a flutter suppression controller is designed based on linear active disturbance rejection control (LADRC). The stability analysis shows that the tracking error and the estimation error of the linear state observer (ESO) are convergent. It is shown through simulation results that wing flutter suppression under the interference of fuel sloshing and external tank vibrating can be achieved by using two control surfaces.

Heat and moisture influence the interphase properties of polymer matrix composite significantly. The composite samples were exposed to different levels of heat and moisture. The advanced peak force–based quantitative nanomechanics mapping technique with the Derjaguin-Muller-Toropov (DMT) model was used to investigate the interphase at the submicron-scale and nanoscale. The interphase’s width and elastic modulus depend on its location, degree of moisture saturation, the extent of expansion, and shrinkage mechanisms. The interphase’s width is uneven and river-like. The interphase thickness and DMT modulus were collected at $\sim 15$ sites for each condition and $\sim 15$ section lines at each location. The Weibull model linear regression and the moments method with various estimators were employed. Weibull results show that the local moisture content affects the interphase thickness more than elastic modulus. The Weibull analysis shows the enhancement of the interphase thickness from $\sim 9.5$ to 95 nm and reduction of the interphase modulus from $\sim 22.8$ to 19.6 GPa, as the local moisture content reaches $\sim 95\%$.

A multistage aero-engine casing is a typical thin-walled structural component that is required to maintain a stable dynamic performance because it is a key component for transmitting thrust from an engine to a plane. However, the manufacturing and assembly processes will generate initial assembly deviation, influencing the mechanical properties of the casing during service. Previous tolerance design had few considerations of the effect of assembly deviation on the dynamic characteristic stability. In this paper, a novel performance-oriented tolerance design method for multistage aero-engine casing is proposed. Based on the shell theory, an analytical model is constructed to present the bolted jointed multistage casing assembly. By varying the assembly variables, the mapping relationship between the manufacturing precision and the indeterminacy of dynamic characteristic is established. Through the mapping relationship, an acceptable assembly tolerance with a specified stability requirement is obtained and adopted into the finite element model to show the effectiveness of the actual assembly structure. The presented method provides a significant reference for tolerance design in the actual assembly of the multistage casing.

Nowadays, greenhouse gases are a worldwide environmental issue. This is why a significant reduction of pollutant emissions inducing a considerable slowing down of climate change is a matter of primary importance. This topic is dramatically influencing the future of industrial production, and thus, also in the aeronautical field, it is becoming a very important technology driver. Advanced diesel engines are promising solutions to achieve a more fuel-efficient aircraft due to their high efficiency and ability to use widely available and low-cost jet and diesel fuels. In this work, innovative electron beam melting (EBM) additive layer manufacturing (ALM) technologies of titanium alloy were applied to the two-strokes, common rail, GF56 diesel engine, designed and developed by the CMD Company for aeronautic applications. The additive-oriented lattice configuration allowed, through a new design, a substantial saving of weight for the engine, as well as a topological optimization and fabrication of alternative connecting rods of a GF56 engine. After a preliminary investigation of EBM process parameters and mechanical characterization of titanium alloy samples, the obtained engine component prototypes were structurally investigated, topologically optimized, and three-dimensionally (3D) printed. Finally, the overall performance improvement on a typical general aviation (GA) aircraft was estimated by comparing the conventional engine with the optimized GF56 propulsion system.

Space capturing and docking is a key technique for space on-orbit operations. The semiphysical simulator is an essential experimental facility to verify the space capturing and docking task. The simulation divergence problem caused by the various delays in system loops has been studied in the literature, but the simulation convergence problem is not investigated well. In this study, the simulation convergence due to mechanical structure dynamics of the upper motion simulator is investigated. The phase analysis shows that the mechanical structure dynamics of the upper motion simulator is phase led, and therefore simulation convergence happens. To compensate for the simulation error, a mechanical structure dynamics–based convergence compensation method is investigated. The stability condition is analyzed. Simulations and experiments are given to test the convergence compensation method.

In this work, we focus on the attitude control problem of an imaging satellite required to acquire and accurately track a target area on the Earth’s surface. The proposed hybrid methodology combines an open-loop implementation of a near-optimal trajectory for the initial acquisition phase and a quaternion feedback proportional-derivative controller for the subsequent tracking phase. The fast prototyping of the near-optimal attitude trajectory is achieved by integrating a direct method, the inverse dynamics in the virtual domain, and a nonlinear programming solver, the sequential gradient-restoration algorithm. The optimization criteria taken under analysis are the target acquisition time and the maneuver energy associated to the actuation torques. To assess the effectiveness of the proposed approach, we compare its performance against two standard closed-loop attitude controllers, namely, a classical quaternion feedback proportional-derivative controller and a linear quadratic regulator. Simulation results, in a hi-fi environment, show that the proposed controller achieves better tracking accuracy and optimality than the other controllers.

This paper presents the results of experimental investigations on the sound absorption of two porous structures and the far-field noise performance of an airfoil employing sound-absorbing porous structures at the trailing edge. The porous structure consists of a microperforated housing with an air gap (PA structure) or an acoustic foam insert (PF structure). Sound absorption coefficients have been characterized using an impedance tube. Compared with a commercial acoustic foam, PF structures have shown a consistently higher (up to 0.5 higher) sound absorption coefficient over 0.5–6.4 kHz. Moreover, the frequency characteristics of PA and PF structures are predictable, which allows their geometric optimization to suit different applications. Far-field noise of airfoils has been measured using a 64-channel microphone array at various flow conditions. The sound-absorbing microtube structure is found to reduce the trailing-edge (TE) noise by up to approximately 12 dB at lower frequencies (1.5–6 kHz), while increasing high-frequency (over 6 kHz) noise levels due to the roughness elements of porous geometries and unsteady flow permeation. The peak frequency of noise reductions is found to not scale with the flow speed, but remains within a narrow frequency band at different flow conditions ($U_{\infty }=30-50\mathrm{m}/\mathrm{s}$, ${\alpha }_g=0°-10°$). This is consistent with the narrow-band feature of microtube structures’ sound absorption, which indicates the contribution of sound absorption to TE noise reduction. The PF structure, due to its broadband sound absorption and tunable frequency characteristics, is considered a promising device for trailing-edge noise control.