Journal of Aerospace Engineering

In this work, a physics-based first-order reliability method (FORM) algorithm is proposed for the flutter reliability analysis of an aircraft wing in the frequency domain. The limit state function, which is an implicit function of random variables, is defined in terms of the damping ratio of the aeroelastic system in a conditional sense on flow velocity. Two aeroelastic cases, namely, an airfoil section model and a cantilever wing model, are considered for carrying out the studies. These aeroelastic models have well separated mean bending and mean torsional modal frequencies. The geometric, structural, and aerodynamic parameters of airfoil and wing systems are modeled as independent Gaussian random variables. The effects of these on the statistics of frequency and damping ratio, and the cumulative distribution functions (CDFs) of flutter velocity are studied. In the case of the wing, the effects of modeling stiffness parameters as Gaussian random fields on the CDFs of flutter velocity are also studied. Here, spectral stochastic finite element method (SFEM) based on Karhunen–Loeve (K–L) expansion is used to discretize the random fields into random variables. From the study of an airfoil system, it is observed that parameters like torsional stiffness, elastic axis location, free stream density, and mass moment of inertia are more sensitive as compared with other parameters. However, in the case of the wing parameters such as torsional stiffness, free stream density, mass moment of inertia, and mass are observed to be more sensitive. The CDFs of flutter velocity obtained using the proposed algorithm are compared with Monte Carlo simulations (MCS) and found to be accurate. A comparative study of aeroelastic reliability for the wing is also carried out by treating stiffness parameters as random variables and random fields. It is observed that the CDFs of flutter velocity in the tail region are conservative when stiffness parameters are treated as random variables.

Air traffic has developed rapidly in the last decades. The noise generated from aircraft can impair the health of passengers and airport residents, so it is necessary to reduce aircraft noise and relieve the annoyance due to aircraft noise. As one important component of aircraft wings, the high-lift device is a barrier to the noise reduction of the large commercial civil aircraft in the next stage, especially during the approach and landing phases. When the slat and flap of the high-lift device are both stowed, some gaps and cavities will emerge and radiate strong noise. However, little attention has been paid to these geometric structures on real aircraft. This paper presents an experimental investigation on the far-field aeroacoustic characteristics of the stowed 30P30N high-lift configuration in the cruise condition, namely with slat and flap both stowed absolutely. The experiment of acoustic field measurements with a far-field microphone and acoustic microphone array shows that the gap between the main element and stowed flap on the pressure surface is the main noise source and can radiate tonal noise in the low-frequency range. The flow field measurement aided by a hot-wire anemometer reveals that the frequencies of the intense fluctuations near this gap are close to the tonal frequencies in the far-field noise spectra. Finally, with the applications of a gap cover and tripping device, the tonal noise of the stowed 30P30N high-lift configuration can be obviously suppressed because the feedback loop between the boundary layer instability upstream and the pressure wave from the flap–main element gap has been disturbed.

This paper numerically simulated a swirl cooling system to study the influences of the slot area ratio and slot angle on the flow field and heat transfer performance. Numerical simulations were performed for different coolant inlet to outlet slot area ratios (1, 2, 3, and 4) and slot angles (60°, 75°, 90°, and 105°) at different Reynolds numbers. Results indicate that large-scale vortices and small circular or oval vortices are generated in the cooling channel at different slot area ratios and slot angles. At identical Reynolds numbers, the cooling system achieved 30% and 23% increases in global Nusselt number and thermal performance factor, respectively, when the slot area ratio increased from 1 to 4. Although the system obtained only a 7.5% increase in global Nusselt number, it achieved a 29.8% increase in thermal performance factor when the slot angle increased from 60° to 105° at fixed Reynolds number. The applied cooling system is recommended for the internal swirl cooling of a gas turbine blade leading edge at optimum values of design parameters at a slot area ratio of 4 and slot angle 105° at high Reynolds number.

The impact of ground proximity on the aerodynamics and trailing vortices of an inverted delta wing was investigated experimentally. The ground proximity always led to a lift and drag increase compared to their outside ground-effect counterpart. The largest lift increment occurred in the low-incidence regime but decreased nonlinearly with increasing incidence. A 145% lift increment for a ground distance of 0.5% chord at 3° incidence was obtained. The ground effect also led to an enlarged trailing vortex with increased rotational speed and vortex strength but a lowered peak vorticity. The trailing vortices moved further outboard with reducing ground distance. The results suggest that winged ground-effect vehicles with an inverted delta-wing planform can benefit from the ground effect for a ground distance up to 40% chord.

The main objective of this paper is to study the effects of applying a heat source ahead of a supersonic mixed compression air intake. This intake has been designed for the freestream Mach number 2. Four parameters of the heat source, including its location, cross-sectional shape, cross-sectional area, and heat generation rate, are investigated numerically to find a suitable heat source that enhances the performance of the intake. The research is conducted at the freestream Mach numbers of 1.8, 2, and 2.2. Total pressure recovery, mass flow ratio, flow distortion, and drag coefficient are considered as the intake performance parameters. Variation of the critical backpressure is further investigated. The results show that a bow shock is formed around the heat source that affects the intake shock system and can have desirable effects on most of the performance parameters and the critical backpressure at different freestream Mach numbers. It is also observed that the heat source location is the most effective parameter of the heat source, and should be selected meticulously.

A point-based modeling scheme for the mesostructure of a two-step three-dimensional (3D) braided composite is proposed. The yarn path and the section profile are more realistic than the traditional mesh-based model, and the influence of the braiding process on the yarn path is reflected in the model. Owing to the point feature, the discretization of the matrix is very convenient after the reinforced structure is generated. The deformation and the failure of the braided composites can be directly simulated from the mesostructure model. Several examples, which include calculating the macroscopic elastic modulus and the high-velocity impact on the braided composite plate, show nice accuracies. The failure and fragmentation of the mesostructure can be well simulated with the point model and the mesh-free material point method.

A transverse injection into high-speed freestreams is a significant principle applied to realistic engineering applications, such as drag reduction, fuel service in a scramjet combustor, and fluidic thrust vectoring control. Fluidic thrust vector control is quite popular for micro space launcher propulsion systems due to several advantages, including better control effectiveness, few or no moving mechanical equipment, and fast dynamic responsiveness. Shock vector control is a straightforward and direct method in which the secondary flow is injected from the transverse section into the divergent portion of the supersonic nozzle. This work conducted coupled theoretical and numerical analyses to investigate performance variations in a three-dimensional supersonic rectangular nozzle with a slot injector. To verify the reliability and accuracy of the research methodology, numerical results were validated against existing experimental data from the literature. Computational static pressure distributions based on the shear stress transport $k\text{-}\omega$ turbulence model matched well with experimentally measured pressure values. Five control parameters were investigated in detail, involving the nozzle pressure ratio, injection pressure ratio, injector location, slot length, and slot width. Critical performance variations were analyzed quantitatively and qualitatively. Some useful conclusions are offered for fighter jet designers.

In the cutting process of nickel-based superalloys, because of the high strain rate and cutting temperature, the cutting deformation is complex and there exist hardening and softening phenomena. Therefore, the present study developed a physically-based constitutive model, which represents the mechanical response of a material at a given microstructure in terms of dislocation glide to describe the deformation behaviors of Inconel718. The established model also describes evolution equations for internal variables characterizing the microstructure. The internal variables are related to the dislocation density. Comparisons between the experimental results and those predicted employing different models [a physically-based constitutive model has been established in the present work and Johnson–Cook (JC) constitutive model] indicate that the established model can accurately characterize the deformation behaviors for Inconel718 at high strain rates and temperatures.

Aerodynamic shape optimization for the high-subsonic low-Reynolds-number flow regime represents an area of ongoing research. The interaction between supercritical compressible flow and laminar boundary layer separation is not well understood due to the significant challenges associated with setting up relevant experimental work. However, in the design of future fixed-wing aircraft for flight in extraterrestrial atmospheres, such flow conditions might commonly occur. The present study presents a family of single-point and multipoint optimized airfoils designed for high-subsonic flight at a high-lift condition in the Martian atmosphere. A gradient-based optimizer is used, with a second-order finite-volume flow solver and a second-order continuous adjoint solver for determining surface sensitivities with respect to the objective function of minimizing drag. Both fully turbulent and transitional flow are considered to evaluate the impact on the resulting design and to stress the importance of continuing research to develop robust shape optimization, including laminar boundary layer and transition prediction. Both on-design and off-design conditions are evaluated, the airfoils obtained when considering transition effects demonstrating good overall performance.

In this paper, a novel state-propagation extreme learning machine using a Kalman filter (KF-ELM) is proposed. In comparison with the plain extreme learning machine, the proposed algorithm takes the topological parameters as state variables and minimizes the covariance of state estimates to overcome the state conjunction and transformation dilemma in time series. As a result, its topological stability and prediction accuracy are enhanced, and these merits are further proved theoretically. In addition, the computational effort of KF-ELM is on the same order of magnitude as the plain extreme learning machine, while the former possesses a faster convergent speed. Then, several benchmark datasets are utilized to test the effectiveness and soundness of the proposed algorithm. Finally, it is employed to predict the gas path performance of a turbofan engine. The performance prediction accuracy is better than the plain ELM with different input rules in the dynamic process. Particularly under various flight operation conditions, the proposed algorithm performs well and its stability is sufficiently showcased. In a word, the proposed algorithm provides a candidate technique for predicting aircraft engine performance in dynamic behavior.

The stern-mounted propellers of near-space airships affect the aerodynamic characteristics of envelopes. Hence, the influence of propellers on the optimization design of envelopes should be considered. In this study, the computational fluid dynamics (CFD) simulation based on the $\gamma \text{-}{\mathrm{Re}}_{\theta }$ transition model was used as the main research method. An optimization model of an envelope with a propeller was also established. Envelopes with and without a propeller were optimized successively. Optimal results were analyzed using the CFD simulation. Results show that when the axial force coefficient of the propeller is 0.54, the drag coefficient of the optimal envelope considering the propeller is 16.5% lower than that without considering the propeller. The output power of the engine of the optimal airship considering the propeller can be reduced by 10.0% with the same advance speed. A wind tunnel test indicates that the drag coefficient and transition position of the envelope obtained by the CFD method are in good agreement with the test results. Therefore, the accuracy of the CFD method is sufficient, and the effect of optimization is evident.

This paper presents a Volterra series reduced-order model method to evaluate aerodynamic forces of aircraft engine blade designs, whose shape changes during design and are always excited by upstream wakes. The method uses a superposition of the steady and dynamic Volterra series to reach this aim. The steady series represents the influence of the blade shape changes on the blade aerodynamic forces. The dynamic series represents the impact of the unsteady upstream wakes on the blade aerodynamic forces. A T106LPT blade is used as an example to discuss the proposed method. The accuracy of the ROM method is analyzed by the example at the wake excitation condition with or without shape-change. The example shows that the ROM method can predict the wake induced pressure variation of a shape-change blade without repeated computational fluid dynamics (CFD) analyses.

The SpaceX company has introduced a reusable first stage rocket that lands on a floating barge. In the event that the first stage rocket falls off the platform, the investigation of landing location on the seafloor may help to increase the salvage efficiency. This paper deals with the application of experimental and numerical techniques into locating a model rocket freely falling into the towing tank with various initial drop angles. The floor of the towing tank represents the ocean floor with regards to the landing position of the model rocket. The landing position is of interest to recover the fallen rocket. The University of New Orleans Towing Tank facility has been used to release a model first-stage Falcon 9 rocket at initial drop angles from 0° to 90° with a uniform increment of 15°. The landing point distribution at each drop angle is recorded using a landing grid on the bottom of the towing tank. First, a deterministic model based on the Dropped Objects Simulator (DROBS) is used to calculate an ideal landing point, which is compared with the experimental results. It is found that one single simulation by the deterministic model cannot reasonably explain the random landing point distributions observed in the experiment. Therefore, the Monte Carlo method based on a stochastic model is used to numerically consider the effects from various random disturbances. The landing point distribution obtained by stochastic process analysis is close to the experimentally obtained distribution for each drop angle case. The stochastic modeling shows more advantages and advancements in subsequent analysis, such as the calculation of landing velocity/speed. The findings from stochastic modeling of the dropping process is also discussed. Finally, the drop angle is found to significantly affect the trajectory, landing points, and landing speed in both the deterministic and the stochastic model.

The heat load caused by windage heating effect of labyrinth seals seriously influences the sealing of shaft air flow. To reveal the windage heating effect of labyrinth seals to reasonably design an aeroengine rotor, this paper studies the windage heating characteristics of labyrinth seals from theoretical, numerical, and experimental perspectives. First, the theoretical analysis of the windage heating effect of labyrinth is conducted, and the experimental platform of heating labyrinth was established to test the heating labyrinth effect. Then, an $RNGk\text{-}\epsilon$ turbulence model was adopted to investigate the characteristics of flow field, leakage characteristics, and windage heating characteristics of labyrinth seals. Last, the influences of pressure ratio, rotational speed and other factors on the windage heating characteristics of labyrinth seals were investigated by theoretical calculation, numerical simulation, and experiment. Through the comprehensive study, we find that: (1) windage heating is indeed induced when leakage flow absorbs the frictional heat caused by the rub of leakage flow with the rotor; (2) when the speed is lower than $\mathrm{1,000}\mathrm{r}\mathrm{p}\mathrm{m}$, the windage heating effect is small and enhances with the rise of rotation speed from $\mathrm{1,000}-\mathrm{6,000}\mathrm{r}\mathrm{p}\mathrm{m}$, and the maximum increase of temperature reaches to 12.87 K; and (3) the pressure ratio increases from $1:1$ to $1:3$, and the temperature rise decreases by $\sim 7\mathrm{K}$. The results of this paper offer a useful insight for the design of labyrinth seals with the emphasis on windage beating effect, to improve the performance of rotor seals.

In a traditional integrated inertial navigation system (INS) and celestial navigation system (CNS) setup, measurements from both systems are only fused at the data output phase. Navigation star centroid positioning error persists in the calculated celestial measurement, which inevitably affects the entire integrated navigation system’s accuracy. This paper proposes a novel integrated INS and CNS navigation system that includes two filtering processes. The angular velocity derived from the star centroid positioning information provides measurements in the first filter. To correct the star centroid positioning error, the gyroscope’s real-time output is used to fuse this measurement via the nonlinear least-square method. In the second filter, the CNS attitude measurement is calculated from the corrected star centroid positioning information. Then, the INS and CNS measurements are combined via a standard Kalman filter to estimate the spacecraft attitude. Comparing the INS/CNS integrated system with the traditional star centroid positioning method, the simulation results illustrate that the proposed method markedly reduces star centroid positioning error and has strong universal applicability to provide similar high-accuracy spacecraft attitude estimations regardless of star sensor specifications.

In this work the parametric instability regions of an exponentially tapered, pretwisted, and rotating symmetric sandwich beam under a temperature gradient, subjected to a periodic axial load has been studied for clamped-free boundary condition. Pretwist angle has been assumed to vary linearly along the length. The equations of motion along with the boundary conditions have been derived using Hamilton’s principle for coupled bending-bending vibration of the beam and the instability regions for principal resonance and combination resonance have been obtained by using the conditions derived by Saito and Otomi. The static buckling loads have been also obtained. Finally, the effect of pretwist angle, taper parameter, temperature gradient, angular velocity of rotation, and properties of the viscoelastic core on beam dynamic and static stability have been represented graphically. It was observed that beam stability improves with angular velocity of rotation and the pretwist angle seemed to have a very complex relationship with dynamic stability of the beam.

In the present work, the hygro-thermomechanical response on the static characteristics of symmetric and antisymmetric laminated composite and sandwich plates is obtained using the recently developed inverse hyperbolic zigzag theory (IHZZT). The framework of IHZZT inculcates the shear strain shape function assuming a nonlinear distribution of transverse shear stresses. It satisfies the necessary conditions of interlaminar stress continuity at the layer interfaces as well as the condition of zero transverse shear stresses at the top and bottom surfaces of the plates. The theory has seven unknown field variables, which are layer-independent. The displacement-based finite-element approach is employed, using the eight-noded isoparametric serendipity element, accounting for the $C^0$ continuity. A number of numerical examples are solved considering the effect of temperature, moisture concentration, span-thickness ratio, aspect ratio, boundary conditions, loading conditions, and variations in the number of layers. The results obtained in terms of deflections and stresses are validated with the exact results and results available in the existing literature. A few new results are produced here to establish a benchmark study for future research.

Inflatable structures are characterized by being light and easy to manufacture and deploy. Hence, they find many applications in aerospace and aeronautical engineering. In this paper, an inflatable segment with a The National Advisory Committee for Aeronautics (NACA) 0021 airfoil cross-section is designed, fabricated, and tested. The geometrical accuracy of the manufactured inflatable segment is measured using laser scanning. Measurements show that the average normalized error of the chord length and thickness are 2.97% and 0.554%, respectively. The aerodynamic behavior of the inflatable segment is then tested in a wind tunnel at different wind speeds and angles of attack. Lift forces are measured using a six-component balance, while the drag forces are calculated from the wake measurements. The lift and drag coefficients of the inflatable section are compared to those of a standard NACA 0021 airfoil. Finally, flow visualization is examined at different angles of attack using two methods: smoke and tufts. Both methods show that flow separation starts at 15° and full stall occurs at 25°. Results indicate that inflatables can find more applications in the design and construction of aerodynamic structures, such as wings.

Unlike the popularly used approaches (in which the time-varying nature of the system’s dynamics is neglected), a Lyapunov-based nonlinear time-varying framework is provided to handle the longitudinal control of tilt rotor aircraft during conversion flight. The conversion flight control problem is regarded as the trajectory-tracking problem with respect to a desired generalized conversion corridor. Introducing the concept of a virtual plane, the design procedure is boiled down into two steps: the design of the virtual controller and the distribution of the actual control inputs. Firstly, aimed at the nonlinear time-varying control-oriented model, a virtual controller that renders the globally uniformly exponential stability of the origin, together with a constraint to alleviate the input saturation, is given with rigorous theoretical proof using the sum-of-squares (SOS) technique. Next, a practical control allocation strategy is adopted to achieve the deflections of the actual control surfaces. To enhance the robustness of the controller, a nonlinear disturbance observer is proposed to estimate the influence of the uncertainties of the aerodynamic parameters and modeling errors.

Under time-varying extreme dynamic environmental conditions, fatigue damage could gradually accumulate with increased stress cycles at critical details of steel or other metallic structures. As a promising tool, a fatigue indicator parameter (FIP) has been widely used to evaluate the crack initiation and propagation process in a mesoscale model. However, the correlation between the FIP and crack growth rate has not been carefully investigated, especially in the micromechanics level. In the present study, to calculate FIP for nanoscale fatigue crack growth, molecular dynamics simulation is performed. Two cyclic loading regimes are applied to the specimen with a preexisting central crack. In each loading regime, the relationship between the FIP and crack growth rate $da/dN$ is calculated based on the FIP distribution contour plot. In the increasing maximum strain loading regime, cracks propagate linearly with the increase of the numbers of stress cycles, and the FIP distribution is similar to the distribution of the disordered lattice. However, lattices with maximum FIP values are mainly located at the boundary of the disordered lattice due to the dislocation and slip concentration. In the constant maximum strain loading regime, a void forming near the crack tip is observed, and a crack propagates by linking these voids and crack tip. With the increase of loading cycles, the crack growth rate decreases to nearly zero while the maximum FIP decreases slowly, implying a lagging effect in this process. The atomistic simulations in both loading regimes demonstrate a linear correlation between the FIP value and the crack growth rate $da/dN$.

Laminated composite components have been widely used in engineering fields, such as aerospace engineering, automotive industries, marine structures, and civil infrastructure. Due to the inherent properties of laminated composites, the impact of local delamination defects will result in a decrease of assembly dimensional precision and mechanical performance. This paper develops a modeling method to integrate the local delamination defects into the single-station assembly variation simulation procedure. First, the local deformation induced by the delamination defects is obtained by using the flexible joint model of the bimaterial interface. Then, the influence of the local displacement at the crack tip on the assembly variation propagation model is addressed. Based on the variation model, the impact of local delamination length ratio ($a/h$) is further studied for assembly of composite parts, and the comparisons of simulation results show that the delamination has a pronounced effect on the assembly deformation of composite components. Finally, two case studies on variation propagation modeling of a delaminated beam assembly are presented to illustrate and verify the proposed methodology.