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This paper examines the performance of a 3D printed acoustic metamaterial as an acoustic treatment for aircraft engine nacelles in the Advanced Noise Control Fan. As the level of air travel continues to increase, so too does the demand for better noise-reduction technologies for aircraft. Engines are one of the two main sources of noise generated by aircraft, with fan noise, in particular, being of concern due to its broadband and tonal contributions. Small and lightweight methods of addressing both broadband and tonal noise are necessary due to the limitations presented by the current engine design. Presented in this paper is a novel acoustic metamaterial that has undergone design optimization for broadband noise reduction. The final design was produced using 3D printing and tested using the Advanced Noise Control Fan at the University of Notre Dame. It was found that the material is capable of reducing the first harmonic of the blade passing frequency by up to 18.5 dB, with an overall noise reduction of 3.7 dB. [ABSTRACT FROM AUTHOR]

The structural safety factor is an essential parameter in aircraft design, representing the ratio of the design load to the operating load. Traditional design methods rely on subjective determination of safety factor values based on experience, lacking objectivity in quantifying uncertainty. However, with advancements in aircraft design technology and increasing competition in the commercial space market, new-generation hypersonic aircraft with complex load environments require a more optimal approach. Applying a uniform safety factor to each component within subregions of the aircraft leads to overly conservative results and impacts flight performance. To address this limitation, a design scheme that incorporates subregional, differentiated safety factors is necessary. This approach allows for better material utilization and ensures compliance with safety requirements. This paper utilizes reliability-based design optimization theory to consider uncertainty in structural systems. It establishes a mapping relationship between structural reliability and differentiated safety factors, providing safety under uncertainty while guaranteeing weight reduction. Additionally, this paper develops a subregional, differentiated safety factors distribution program to determine the safety factors of different subregions of the structure. Consequently, a refined subregional differentiated safety factors scheme that balances safety and economy is derived. [ABSTRACT FROM AUTHOR]

Aircraft are subject to small disturbances during taxiing, many of which are accidental and difficult to explain. Although nonlinear factors are considered in traditional analysis, Coulomb friction is generally ignored, and the interaction and common effects of nonlinear factors are mostly not discussed. In this paper, the dynamic model of shimmy is established, and the influence of Coulomb friction on shimmy is studied by using bifurcation theory and the structural mechanics analysis method. The results show that the system with Coulomb friction has subcritical Hopf bifurcation and that the system with square damping has supercritical Hopf bifurcation. Although the two do not change the stability of the system, their cooperation can improve the stability of the system. The Coulomb friction torque of the system has a complex piecewise functional relationship with the stability distance, rake angle, and acceleration. Some combinations will lead to very low Coulomb friction and deteriorate the anti-interference ability of the landing gear. This paper provides theoretical basis and support for the rational design of structural parameter collocation, enhancing the antidisturbance ability of the system during constant-speed or variable-speed taxis and explaining the shimmy phenomenon in the process of variable-speed taxis of some aircraft. [ABSTRACT FROM AUTHOR]

The recent interest in the development of supersonic transport raises concerns about an increase in community noise around airports. As noise certification standards for supersonic transport other than Concorde have not yet been developed by the International Civil Aviation Organization, there is a need for a physics-based scaling rule for supersonic transport takeoff noise performance. Assuming supersonic transport takeoff noise levels are dominated by the engine mixed jet velocity and the aircraft-to-microphone propagation distance, this paper presents a reduced-order model for supersonic transport takeoff noise levels as a function of four scaling groups: cruise Mach number, takeoff aerodynamic efficiency, takeoff speed, and number of installed engines. This paper finds that, as cruise Mach number increases, supersonic transport takeoff noise levels increase while their thrust cutback noise reduction potential decreases. Assuming constant aerodynamic efficiency, takeoff speed, and number of installed engines, the takeoff noise levels and noise reduction potential of a Mach 2.2 aircraft are found to be ∼15.3  dB higher and ∼19.2  dB less compared to a Mach 1.4 aircraft, respectively. This scaling rule can potentially yield a simple guideline for estimating an approximate noise limit for supersonic transport, depending on their cruise Mach number. [ABSTRACT FROM AUTHOR]

This paper focuses on the development of a deep-learning framework for predicting distributed quantities around aircraft flying in the transonic regime. These quantities play a crucial role in determining aerodynamic loads and conducting aeroelastic analysis. Angle of attack and Mach number are chosen as the two independent parameters for the reduced-order models. A comparative assessment is conducted between the proposed deep-learning framework and the proper orthogonal decomposition approach to identify the strengths and weaknesses of each method. The accuracy of the data-driven machine-learning method in modeling steady-state transonic aerodynamics is assessed against three benchmark cases of three-dimensional test cases: Benchmark Super Critical Wing and ONERA M6 wings, and the wing-body Common Research Model configuration. Despite the challenges of the analyzed scenarios, promising results are obtained for each test case, showing the effectiveness of the model implemented. Furthermore, the paper demonstrates the application of the method for aeroelastic analysis and uncertainty quantification. This quantifies the robustness and versatility of the implemented model. [ABSTRACT FROM AUTHOR]

Deployable mechanisms have the characteristics of complex loops and strong coupling. As a result, the solution to the degrees of freedom (DOFs) and kinematic characteristics of the basic mechanism unit are often complicated, and the solution for combined mechanisms is rarely reported in the relevant literature. In this paper, the kinematic characteristics of the rectangular pyramid deployable mechanism (RPDM) are analyzed, and the DOFs of the RPDM are calculated based on the screw theory. In addition, the DOF number and properties of its moving components are analyzed. Based on the condition that the nonhomogeneous linear equations have a unique solution, the singularity analysis of the mechanism is complete, which indicates that the mechanism can be continuously deployed and folded. Based on the geometric characteristics of the mechanism, its folded ratio is calculated. Finally, a ring, a planar, and a parabolic truss deployable antenna mechanism are assembled based on the RPDM, and the DOF variation for the different assembly modes is obtained via the screw constraint matrix. This study provides a good reference for the design and analysis of such multi-closed-loop deployable mechanisms, and the large-scale space-deployable mechanisms constructed in this paper have good application prospects in the aerospace field. [ABSTRACT FROM AUTHOR]

Stowage schemes based on origami wrapping patterns are useful for the design of various deployable spacecraft structures. These origami patterns allow for the compact stowage of flat sheets of small thickness. Typically, some other structure is needed to stiffen this thin sheet when it is fully deployed and flat. This paper describes a class of thin-shell structures that can be compactly stowed using origami wrapping methods but have a corrugated form when deployed. These structures cannot reach a flat state by design. The persistent corrugations provide stiffness without the need for an external structure, thus leading to overall savings in mass and complexity. The geometrical design of the corrugations is highly tunable. This paper describes a method for generating the form of these folding structures; the method guarantees that the structure can be unstrained both when compactly wrapped and when fully deployed. Test articles are fabricated to demonstrate the concept, and stowage experiments and stiffness experiments are conducted. A structural finite element modeling procedure is described to predict the deployed stiffness of these structures, and the stiffness predictions match the experimental results. [ABSTRACT FROM AUTHOR]

Recent studies have shown that semi-aeroelastic hinge devices can enable larger aircraft wingspans. Such a device would be folded on the ground to meet airport width restrictions, locked during cruise for optimal aerodynamic performance, and released during maneuvers to alleviate flight loads. In contrast, this paper uses a wind tunnel experiment to study the aeroelastic behavior of floating wingtip fuel tanks. This device consists of a freely floating wingtip with an additional mass attached in the form of a liquid-filled fuel tank. The static aeroelastic results show that altering the fuel tank's filling level and position allows the wingtip to float at an optimal angle for aerodynamic efficiency across various angles of attack and fuel masses. Additionally, this paper shows that, with careful selection of the mass distribution of the wingtip, dynamic load alleviation comparable to the semi-aeroelastic hinge concept can be achieved during turbulence and one-minus-cosine encounters. Furthermore, the effect of fluid motion is shown to reduce incremental loads during random turbulence encounters by up to 10%; however, it has a negligible impact on the response to one-minus-cosine encounters. Such results are also confirmed by a numerical model incorporating a simple reduced-order fluid sloshing model. [ABSTRACT FROM AUTHOR]

Whirl flutter as an aeroelastic instability can be a limiting factor in the design and certification of turboprop aircraft configurations, especially for the engine suspension. Whirl flutter prediction for these configurations is currently done in the frequency domain using rigid propeller aerodynamic derivatives. Blade flexibility is neglected in this process, although it is known to have an impact on whirl flutter predictions. This paper uses frequency-domain transfer matrices for the propeller hub loads identified from a time-domain multibody simulation model of an isolated turboprop propeller and included into a frequency-domain flutter analysis to study the impact of blade elasticity on propeller whirl flutter. Results demonstrate a significantly stabilizing effect of blade elasticity on propeller whirl flutter due to a reduction of the destabilizing pitch-yaw cross-coupling moment. The method demonstrated in this paper is applicable to arbitrary time-domain propeller models and compatible with standard frequency-domain flutter processes, allowing for increased fidelity in the flutter prediction process. [ABSTRACT FROM AUTHOR]

On 18 February 2021, the Mars 2020 Perseverance rover and Ingenuity helicopter successfully landed inside Jezero Crater. At 1026 kg, Perseverance is the largest, most sophisticated rover ever delivered to another planet. This event marked the ninth successful landing and fifth rover to be delivered at Mars. The Program to Optimize Simulated Trajectories II, a trajectory simulation tool, was the prime entry, descent, and landing performance simulation for Mars 2020. This paper presents comparisons between the flight telemetry and the simulation predictions. In general, approximately 90% of the as-flown values were within ±3⁢𝜎 (standard deviations) of the preflight simulation predictions, and the anomalies are discussed in the paper. These comparisons are important in order to understand how each of the individual models and the integrated simulation as a whole performed. This information is fed forward to future missions, which benefit from knowing where additional resources or studies are needed and where uncertainties may be reduced to enable improved performance. [ABSTRACT FROM AUTHOR]

This paper considers the lift coefficient of an airfoil with nonzero prescribed thickness and camber, along with a prescribed porosity distribution in a steady incompressible flow. In similar prior work, considering a Darcy-type porosity condition on the airfoil surface provides a Fredholm integral equation that can be solved exactly for Hölder-continuous porosity distributions. However, the generated lift predicted by the model diverges from the experimental data for porous airfoils with large values of the porosity parameter. This indicates a fundamental characteristic is missing in the mathematical model. Consequently, in this paper, the (linear) Darcy porosity condition is replaced by the (nonlinear) Forchheimer porosity condition. The Forchheimer porosity condition is decomposed into linear sections and furnishes a modified system of Fredholm integral equations, enabling an approximate solution by superposition for Hölder-continuous porosity distributions. The solution is then verified against the SD7003 airfoil results, and the comparison shows better agreement than considering the Darcy porosity condition. The methodology and results presented here may be combined with previous work on the aeroacoustics of porous airfoils with a Forchheimer boundary condition to address the conflicting aims of improving aerodynamic performance while reducing unwanted aeroacoustic emissions. [ABSTRACT FROM AUTHOR]

Accurate prediction of nonlinear aerodynamics is essential for the transonic aeroelastic analysis of flight vehicles. Though reduced-order aerodynamic models are cheap and reasonable tools, it is still a tough problem to accurately evaluate the unsteady pressure distributions on the surface of an elastic structure. This paper presents a nonlinear data-driven modeling approach based on the high-fidelity simulations in the following three steps. The first step is to compute the dominant modes of unsteady pressure distributions through the proper orthogonal decomposition. The pressure snapshots used for the feature extraction are sampled under a multilevel sine-sweep excitation. The second step is to obtain the low-dimensional temporal dynamics of the coefficients of these modes via polynomial nonlinear state-space identification. The linear estimation implemented by employing the dynamic mode decomposition with control algorithm serves as the initialization of the nonlinear optimization. The third step is to reconstruct the unsteady pressure distributions under arbitrary structural excitation from the temporal coefficients. The paper validates the approach via two numerical examples of the transonic aerodynamic-structural coupling problem. One is an NACA0012 airfoil, and the other is an AGARD 445.6 wing. The examples show that the proposed approach exhibits both accurate and robust performance in the prediction of unsteady pressure distributions, aerodynamic forces, and aeroelastic responses. In particular, the approach well predicts the physical features at the fluid-structure coupling interface, previously neglected in the system identification of aerodynamic systems. Therefore, the approach serves as a promising tool for data-driven aeroelastic analysis. [ABSTRACT FROM AUTHOR]

This paper investigates the influence of tiltrotor blade twist on whirl-flutter stability boundaries. Preliminary evaluations indicate that the whirl-flutter speed can be increased if the blade twist slope is reduced. This positive effect results from the shift in the overall thrust toward the blade tip, increasing the flapwise bending moment at the blade root and the trim coning angle. This, in turn, generates a positive pitch-lag coupling, increasing the whirl-flutter speed. However, the shift of high sectional thrust forces toward the blade tip sections returns a higher induced drag, showing the tendency to increase the power required. The paper shows that, by using blade twist laws based on piecewise linear functions and adding the wing airfoil thickness as a second design parameter, it is possible to identify aircraft configurations that improve the whirl-flutter stability boundaries without penalizing the power required in airplane and helicopter mode flight. This is possible because the blade twist and the wing airfoil thickness have an impact on both power required and whirl-flutter speed, so a simple optimization algorithm can identify good tradeoffs. A detailed tiltrotor model representative of the Bell XV-15 is used to display the effectiveness of the proposed approach. The examples show that increases up to 21% on the whirl-flutter speed are achievable without penalties in the aircraft power required and with the additional benefit of a benign impact on rotor pitch link loads. [ABSTRACT FROM AUTHOR]

Civilian aviation continues to contribute significantly to the total economic and environmental impact of aeronautics. Reduction of the fuel burn and environmental impact of civilian aviation is critical to the overall sustainability of the industry, and it can be accomplished, in part, through the optimization of arrival and approach procedures. This paper proposes the development of a method for measuring the degree of compliance of optimized approach trajectories with air traffic management (ATM) intentions, using an intention compliance level (ICL) indicator. Based on fuzzy logic, this measure reflects the extent to which approach trajectories satisfy the required time-of-arrival constraints. This research demonstrates an approach trajectory strategy that maximizes the ICL, maintains compliance with ATM intent, and achieves efficiency goals inclusive of reduced fuel consumption through selective airspeed changes. Simulations on the Airbus A320 indicate that achieving the optimal trajectory and flight parameters can significantly guide trajectory-based operations to minimize airline economic costs and reduce environmental impact while complying with ATM commands. In this paper we will organize the data as follows. The Introduction will summarize past research as a means of identifying the gaps that this research seeks to bridge and introduce the premise of our findings. Section II proposes the concept of ICL to evaluate the relationship between flight time and the required time of arrival and establishes an en-route descent trajectory model. Section III constructs the optimization strategy based on simulated annealing genetic algorithm (SAGA), evaluates the effectiveness of the algorithm, and verifies the contributions of the ICL in a basic scenario. Section IV analyzes the impacts of various factors on the optimization results in a complex scenario. [ABSTRACT FROM AUTHOR]

The ability to additively manufacture structures on-orbit has the potential to fundamentally alter the traditional paradigm for how large spacecraft are constructed and launched into space. The space environment presents several unique challenges for additive manufacturing, including the need to operate in a vacuum. This paper presents the design, analysis, and test results for a passively cooled fused filament fabrication (FFF) 3D printer capable of manufacturing parts out of engineering-grade thermoplastics in the vacuum of space. Four high-temperature materials were successfully printed in high vacuum, including polyetherketoneketone, carbon-nanotube-polyetherketoneketone, polyetherimide, and carbon-nanotube-polyetherimide. Over 100 test coupons were printed in a vacuum and tested to confirm the feasibility of applying the FFF process in this environment. Lessons learned were documented throughout the vacuum printing test campaigns and are discussed here. This paper is part of a two-part series. Part I presented results for using a low-temperature hotend capable of printing hobby-grade materials in high vacuum and documented initial findings and lessons learned. Part II presents the results for a high-temperature hotend capable of printing engineering-grade plastics that are suitable for on-orbit manufacturing. The combined results of the two papers in this series can be used to inform future on-orbit additive manufacturing. [ABSTRACT FROM AUTHOR]

On-orbit manufacturing and assembly have become major research and development topics for government and commercial entities seeking the capability to build very large structures in space. Additive manufacturing is well suited to this paradigm since the process is highly automated, produces little material waste, and allows for a large degree of geometric freedom. This paper presents a design for a 3D printer that operates in high vacuum. The vacuum 3D printer has completed multiple thermal vacuum test campaigns, with dozens of parts printed to date using low-temperature thermoplastics. Testing of material coupons shows that samples printed in vacuum have strength properties generally within a standard deviation of samples printed at ambient pressure. The overall results from multiple successful tests of the vacuum 3D printer promote the feasibility of on-orbit additive manufacturing while exposed to the space environment. This paper is part one of a two-part series. Part I presents the results using a low-temperature hotend capable of printing hobby-grade materials and documents some initial findings and lessons learned for applying FFF in vacuum. Part II presents the results for a high-temperature hotend capable of printing engineering grade plastics that are suitable for on-orbit manufacturing. The combined results of the two papers in this series can be used to inform future on-orbit additive manufacturing applications as well as potential uses on future moon missions. [ABSTRACT FROM AUTHOR]

This paper will discuss the use of computational fluid dynamics (CFD) for the Space Launch System (SLS) program to model the ascent phase of flight. The ascent phase begins shortly after the vehicle clears the launch tower and extends to the first staging event. To model SLS's ascent, over 1000 numerical solutions of the Navier-Stokes equations were solved, and this analysis has been repeated for five different SLS configurations. To manage this demanding ascent CFD task, the SLS program has developed the Computational Aerosciences Productivity & Execution software. The paper also discusses some of the ways that CFD and high-end computing have advanced in the last decade and offers some comparisons to CFD used in the Space Shuttle Program. [ABSTRACT FROM AUTHOR]

The paper investigates the unpowered descent of a rotor system through the upper atmosphere. Axial and helical trajectories are investigated in the context of fixed points as well as an optimal control problem for maximizing flight time. The mathematical model considered in the paper incorporates the fuselage degrees of freedom, dynamic inflow model, and airfoil characteristics that depend on Mach and Reynolds numbers. Considering a potential application as a descent mechanism, trajectory generation is performed to maximize the flight time. As an example, the performance in the Venusian atmosphere for rotors with different airfoil characteristics is assessed. To delineate the role of constraints, initial conditions, and aerodynamic forces on the optimal descent, the axial trajectory is studied by dividing it into two phases. The first phase corresponds to the trajectory determination through an optimization process wherein control inputs are provided such that states are within bounds. The second phase trajectory (below 70 km), although determined by solving the optimal control problem as in phase-I, is shown to be close to that achieved using control inputs corresponding to fixed points corresponding to each altitude. Apart from the axial flight, helical trajectories and corresponding fixed points are investigated using a rotating constant sideslip frame. Furthermore, optimal helical trajectories are also determined, which could be useful for rotor-based descent mechanisms. A comparison between axial and helical fixed-point solutions is also presented. [ABSTRACT FROM AUTHOR]

In this paper, high-spatial-resolution unsteady pressure-sensitive paint (UPSP) data are utilized to compare two methods for panel buffet forcing function (BFF) estimation for the Space Launch System (SLS). Such methods are based on discrete pressure measurements within a panel but employ coherence factors to account for partially correlated fluctuating pressures across the whole panel. In one method, coherence factors are derived based on the Corcos model, whereas the second method utilizes experimentally derived coherence factors. To simulate discrete measurements using UPSP data, suitable subsets of the data are extracted. When full UPSP resolution is retained, UPSP data provide a benchmark to assess discrete-measurement-based methods. The analysis focuses on the peak SLS buffet environment located downstream of the forward attachment hardware (FAH) between the core stage and solid rocket boosters. Trends of the Corcos-based and experimentally derived coherence factors are in reasonable agreement with the benchmark. However, at certain frequencies, experimentally derived coherence factors are sensitive to the separation distance between pressure measurements utilized to compute coherence lengths. Such sensitivity originates from deviation of the experimentally based coherence function from an exponential decay assumption. On the other hand, the present implementation of the Corcos model fails to capture certain nonturbulent boundary-layer-related environments, such as a subharmonic of FAH vortex shedding. For all methods presented in this paper, at near-transonic conditions, increased pressure coherence and spatial nonuniformity lead to BFF overestimation and sensitivity to the pressure measurement location within the panel. [ABSTRACT FROM AUTHOR]

This work deals with the fundamentally similar dynamic behavior of two apparently different aircraft systems, namely, landing gears and all-movable control surfaces. Both systems typically contain stiffness and damping nonlinearities, leading to shimmy/flutter limit cycle oscillations (LCOs). In a series of earlier papers the authors have developed highly efficient computational methods, based on describing functions, to construct LCO responses for dynamic systems with concentrated nonlinearities in stiffness or damping. These methods are extended in the present paper, and the focus is on the physical insights that can be obtained by using them for systems with multiple nonlinearities. The model parameters are varied to obtain an array of interesting results. For example, each branch of the landing gear response is attributable to a particular nonlinearity. Based on this, a possible reason is advanced for the occurrence of shimmy in the F/A-18 aircraft. On the other hand, the entire flutter response of the all-movable surface can be viewed as being shaped by one nonlinearity and modified by the other. Finally, the well-known Wright Air Development Center wind tunnel flutter test results for an all-movable horizontal tail with freeplay are reinterpreted and explained in the light of the present findings. [ABSTRACT FROM AUTHOR]

A group of satellites, with either homogeneous or heterogeneous orbital characteristics and/or hardware specifications, can undertake a reconfiguration process due to variations in operations pertaining to Earth observation missions. This paper investigates the problem of optimizing a satellite constellation reconfiguration process against two competing mission objectives: 1) the maximization of the total coverage reward, and 2) the minimization of the total cost of the transfer. The decision variables for the reconfiguration process include the design of the new configuration and the assignment of satellites from one configuration to another. We present a novel biobjective integer linear programming formulation that combines constellation design and transfer problems. The formulation lends itself to the use of generic mixed-integer linear programming (MILP) methods such as the branch-and-bound algorithm for the computation of provably optimal solutions; however, these approaches become computationally prohibitive even for moderately sized instances. In response to this challenge, this paper proposes a Lagrangian relaxation-based heuristic method that leverages the assignment problem structure embedded in the problem. The results from the computational experiments attest to the near-optimality of the Lagrangian heuristic solutions and a significant improvement in the computational runtime as compared to a commercial MILP solver. [ABSTRACT FROM AUTHOR]

Effective and reliable slipstream numerical simulation methods are important for propeller-driven aircraft design. This paper presents a high-precision and efficient momentum source method (HPE-MSM) based on a novel actuator disk load prediction model established by the frozen rotor method and blade element momentum theory. The simulation results of two benchmark test cases of an isolated propeller and a typical turboprop airliner show that the accuracy of the HPE-MSM proposed in this paper is close to the time-averaged unsteady results of the sliding mesh method (SMM), with a maximum error of 5% in aircraft lift and drag coefficients compared with experimental values. Meanwhile, the calculation efficiency of the HPE-MSM is comparable to quasi-steady methods, with only about 3.4% of the computation resources of the SMM in the whole aircraft simulation. This novel approach achieves high-precision and efficient simulation of the slipstream, which has the potential to improve the design level of propeller-driven aircraft. [ABSTRACT FROM AUTHOR]

The momentum coefficient is one of the most commonly used parameters to quantify the aerodynamic performance and system output of active flow control (AFC). Estimating the momentum coefficient accurately through experiments is challenging. Several approaches exist to estimate the momentum coefficient in laboratory conditions. However, the question arises as to whether these models hold as ambient and supply conditions change during operation in flight. Therefore, computational fluidic dynamic (CFD) simulations were carried out on two exemplary AFC actuators (i.e., a fluidic oscillator and a steady jet nozzle) in a quiescent environment with supply pressure ratios ranging from 1.5 to 4; supply temperatures at 293, 550, and 800 K; and ambient altitude conditions at sea level and 10 km. When scaled with √Tsup/p∞, the mass flow for different conditions collapses as a function of the pressure ratio. Furthermore, the momentum coefficient collapses when scaled with 1/p∞. The existing models to calculate the momentum coefficient hold for changing ambient and supply conditions. Therefore, these models, in conjunction with scaling parameters, can be employed during the operation of an AFC system in flight with changing conditions in order to provide the desired momentum output. The paper discusses the details of the CFD results and the derivation of the scaling parameters. [ABSTRACT FROM AUTHOR]

There has been a growing interest in utilizing flared folding wingtips as an in-flight load alleviation device to enable increased wing spans that meet airport gate limits but with little increase in wing weight. The semi-aeroelastic hinge (SAH) concept is implemented in high-aspect-ratio wings to enable wingtips to be released during severe load cases such as maneuvers and gusts to alleviate the bending moments while maintaining optimum aerodynamic shape for the rest of the flight. In this paper, scaling methods for wings incorporating the SAH are explored, allowing for the development of equivalent scaled unmanned aerial vehicles or wind tunnel models with similar aeroelastic behavior as full-size aircraft. Three scaling approaches are considered in this study, namely, Iso-Froude, Iso-Frequency, and Iso-Strain, where a set of governing nondimensional quantities and scaling factors are determined. Despite the significant nonlinearities resulting from large wingtip fold angles, it is shown that a linear scaling approach can be appropriate for such a wing configuration. Furthermore, the aeroelastic properties of each scaled model are compared to those of the full-scale model, where the best match was obtained from the Iso-Strain model, although it is challenging to meet the required operational conditions. [ABSTRACT FROM AUTHOR]

Aluminum and boron nanoparticles are added to kerosene in this paper to improve the combustion properties of kerosene. When the particle concentration is 30 g/L, the volumetric heat value of fuel increases by 3 and 5%, respectively, with the addition of aluminum and boron nanoparticles. Meanwhile, combustion experiments in a supersonic combustor are conducted to study the combustion characteristics of Al-kerosene nanofuels and B-kerosene nanofuels. The air flow rates of all experiments are about 1.77 kg/s, the total temperature is 1500 K, and the total pressure is 1.2 MPa. The combustion flow and flame structures of kerosene with different particles and different concentrations are studied, and the results indicate that the addition of nanoparticles to the fuel enhances combustion and heat release enhanced, and the flame stabilization mode is changed from the cavity stabilization to the jet-wake stabilization mode with a higher particle concentration. Meanwhile, the unsteady characteristics of flame are studied. The flame oscillation is intensified, and the fundamental frequency of the flame increases with the increase in particle concentration. Besides, the addition of nanoparticles significantly improves the combustion efficiency of kerosene. When the particle concentration is 30 g/L, the combustion efficiencies of nanofuels are increased by 15% and 17.5%, respectively, with the addition of aluminum and boron nanoparticles. [ABSTRACT FROM AUTHOR]

Uncrewed air vehicles in urban environments will have flights during terminal operations that are dominated by strong transient aerodynamics. These vehicles are not only lighter and smaller than traditional rotorcraft and helicopters, but in many instances they may be hybrid configurations with lifting surfaces similar to those of fixed-wing aircraft. These differences require further understanding of the physics of these transient aerodynamics, specifically large-amplitude transverse gusts and the resulting vehicle response, where classic indicial theory is no longer valid. This is crucial to the safety and certification of these air vehicles near buildings and populations. Prior efforts identified that gust responses depart from traditional linear theory when the leading-edge vortex (LEV) forms as a distinct feature and departs from the lifting surface, resulting in flow nonlinearities. This paper expands our understanding of the interactional physics of these nonlinear transverse gusts with flow separation. LEV and trailing-edge vortex (TEV) behavior are correlated, and these vortex interactions are studied to understand the impact on flow separation and subsequent aerodynamic behavior. Larger gust ratios are observed to increase the LEV normal translation, while flow separation is driven by the location and magnitude of the TEV. [ABSTRACT FROM AUTHOR]

In this paper, we present a wall model based on a modified version of the Spalart-Allmaras turbulence model. Unlike typical wall models, this method avoids the need to query the interior solution by utilizing solution information solely at the boundary, making it well-suited for unstructured grids and mesh adaptation. Furthermore, this wall-modeling method is formulated to lessen the near-wall grid resolution requirements by, below the log layer, making the eddy viscosity approach a constant, nonzero value and the velocity varying approximately linearly with distance from the wall. Using output-based mesh adaptation, this new wall-modeling method is compared to standard Reynolds-Averaged Navier-Stokes (RANS-SA). It is shown that this wall model results in accurate predictions of quantities of interest such as aerodynamic coefficients, surface pressure and temperature, skin friction, and heat transfer compared with RANS-SA, while requiring substantially less near-wall grid. [ABSTRACT FROM AUTHOR]

State feedback is mostly used in the active vibration control of structural systems. However, the most challenging problem is to achieve the whole state of the structure. Another issue in active vibration control is that, in many cases, only some particular modes need to be controlled, while the other modes can remain unchanged in order to save on the costs of the control. To solve the above issues, this paper proposes a state-observing method and conducts a non-spillover partial pole placement control using a multistep method. Using the finite element method, the electromechanical coupling equations of motion of the beam are formulated. The reduced finite element model is then obtained by using the system equivalent reduction expansion process method. Introducing the real measurements by sensors and applying the local corresponding principle, the modes of the structural system are corrected, which are then utilized to predict all the degrees of freedom without measuring by sensors. Then, the non-spillover partial pole placement based on the state feedback is conducted using the multistep method. The optimal locations of actuators are derived by the genetic algorithm to control the specific structural modes effectively. The numerical simulations and experimental studies are also carried out. [ABSTRACT FROM AUTHOR]

Aerodynamic simulations were carried out in the study presented in this paper focusing on the stall performance of the High-Lift Common Research Model obtained from the fourth AIAA High-Lift Prediction Workshop. Various turbulence models of Reynolds-averaged Navier-Stokes simulations are analyzed. A modified version of the transitional 𝑘-̅ν²-𝜔 model was developed to enhance stall prediction accuracy for high-lift configurations with a nacelle chine. The vortex generator, three-element airfoil, and high-lift model are numerically simulated. The results reveal that implementing a 𝑘-̅ν² -𝜔 model with the separation shear layer fixed notably enhances the stall prediction behavior for both the three-element airfoil and high-lift configuration without affecting the prediction of the vortex strength of a vortex generator. Moreover, incorporating rotation correction into the SPF 𝑘-̅ν² -𝜔 model improves the prediction of vortex strength and further enhances stall prediction for the high-lift configuration. The relative error in predicting the maximum lift coefficient is less than 5% of the experimental data. The study also investigated the impact of the nacelle chine on the stall behavior of the high-lift configuration. The results demonstrate that the chine vortex can mitigate the adverse effects of the nacelle/pylon vortex system and increase the maximum lift coefficient. [ABSTRACT FROM AUTHOR]

This paper investigates the combination of a slit at the blunt trailing edge of the flat plate and dielectric barrier discharge plasma to control the vortex shedding of the plate and its associated tonal noise. The noise and flow characteristics of the plate were measured using a far-field microphone array and the particle image velocimetry technique, respectively. The results show that the vortex shedding and the tonal noise can be significantly suppressed by the slit alone, with an average noise reduction of approximately 10 dB in the test Reynolds number. In addition, installing a plasma actuator inside the slit further suppresses the vortex shedding and reduces tonal noise. However, the additional control efficiency of the plasma decreases with increasing wind speeds, with a further 8 dB reduction at a wind speed of U∞=10  m/s (corresponding to an inducing blowing rate BR of 4.5%). However, only an additional 1.5 dB noise reduction is achieved at U∞=20  m/s (BR=2.25%). The particle image velocimetry snapshots were analyzed by proper orthogonal decomposition. The measurements clearly show the variation in vortex shedding at the trailing edge of the plate, revealing the underlying flow mechanisms that lead to the observed noise variations and frequency changes. [ABSTRACT FROM AUTHOR]

This paper presents an improved mathematical expression for semi-empirical wall pressure spectra modeling based on gene expression programming (GEP). The main focus of this work is to obtain a model that applies to a wide range of cases in terms of parameters and the source of data. The dataset comprises flat plate and airfoil cases with adverse and favorable pressure gradients at various Reynolds numbers. First, a characterization of the dataset is performed to understand the low-dimensional distribution of parameters. Then, a feature importance study is conducted to choose the most suitable model input variables from the exhaustive list of nondimensional parameters. The GEP algorithm is modified to ensure that trained models adhere to the basic structure of previously published semi-empirical models. Following training on the diverse database, the new model is compared against existing, best-performing empirical models to quantify the performance improvements. The models are tested on cases with completely different configurations and parameter ranges, unseen during training, and maintain their superior performance. Finally, a comparison is made between models developed with GEP and neural networks in terms of their efficacy, complexity, and interpretability. [ABSTRACT FROM AUTHOR]

Fast and accurate evaluation of aerodynamic characteristics is essential for aerodynamic design optimization because aircraft programs require many years of design and optimization. Therefore, it is imperative to develop sufficiently fast, robust, and accurate computational tools for industry routine analysis. This paper presents a nonintrusive machine-learning method for building reduced-order models (ROMs) using an autoencoder neural network architecture. An optimization framework was developed to identify the optimal solution by exploring the low-dimensional subspace generated by the trained autoencoder. To demonstrate the convergence, stability, and reliability of the ROM, a subsonic inverse design problem and a transonic drag minimization problem of the airfoil were studied and validated using two different parameterization strategies. The robustness and accuracy demonstrated by the method suggest that it is valuable in parametric studies, such as aerodynamic design and optimization, and requires only a small fraction of the cost of full-order modeling. [ABSTRACT FROM AUTHOR]

Accurate real-time prediction of aerodynamic forces is crucial for the navigation of unmanned aerial vehicles (UAVs). This paper presents a data-driven aerodynamic force prediction model based on a small number of pressure sensors located on the surface of a UAV. The model is built on a linear term that can make a reasonably accurate prediction and a nonlinear correction for accuracy improvement. The linear term is based on a reduced basis reconstruction of surface pressure, with the basis extracted from simulation data and the basis coefficients determined by solving linear pressure reconstruction equations at a set of optimal sensor locations, which are obtained by using the discrete empirical interpolation method (DEIM). The nonlinear term is an artificial neural network that is trained to bridge the gap between the DEIM prediction and the ground truth, especially when only low-fidelity simulation data are available. The model is tested on numerical and experimental dynamic stall data of a two-dimensional NACA0015 airfoil and numerical simulation data of the dynamic stall of a three-dimensional drone. Numerical results demonstrate that the machine-learning-enhanced model is accurate, efficient, and robust, even for the NACA0015 case, in which the simulations do not agree well with the wind tunnel experiments. [ABSTRACT FROM AUTHOR]

This paper analyzes ionic wind production and propulsive force in various electrode configurations under atmospheric conditions. By considering the aerodynamic forces in addition to previously considered electric ones, new predictions for steady-state forces and ionic wind flow velocity are successfully compared with experimental measurements, providing convincing quantitative evidence of the predictive capabilities of drift-diffusion modeling associated with one-way Coulomb forcing of Navier-Stokes equations for ionic wind generation. Furthermore, various electrode configurations are analyzed, some of them streamlined, reducing wakes downstream collectors on the one hand and providing additional thrust on the other. The quantification of these additional thrusts is analyzed, physically discussed, and explored in various configurations. [ABSTRACT FROM AUTHOR]

This paper concerns the interaction of an impinging shock wave with a supersonic turbulent boundary layer over several distinct and permanently deformed surfaces, resulting in differences in the shock-boundary-layer interaction and the surface acoustic loading. High-order numerical simulations featuring two-dimensional surface deformations typically encountered in experiments are performed. The deformation amplitudes are up to half the incoming turbulent boundary-layer thickness. The results show that the high-pressure region about the shock impingement is significantly altered and can become narrower or wider depending on the local surface inclination of the deformed panel mode. The surface curvature is found to not significantly affect the separation and reattachment locations of the recirculation bubble. The power spectrum analysis of the pressure fluctuations along the panel's midspan, where the surface attains the largest deformation amplitude, exhibits a rich and varied response. The pressure power spectrum is amplified in all of the surface deformation modes examined, with the magnitude of the amplification varying in the frequency domain, depending on the location and mode. [ABSTRACT FROM AUTHOR]

Under random-interval mixed uncertainties of structures, failure-probability upper-bound function (FPUBF), which varies with the distribution parameters of random inputs, can not only provide the influence of distribution parameters on the failure-probability upper bound (FPUB), but also contribute to decoupling a reliability-based design optimization model. Although FPUBF can be estimated by repeatedly evaluating FPUBs at different distribution parameter realizations, it suffers from unaffordable computational cost resulting from this double-loop framework. To address this issue, this paper proposes a single-loop sampling strategy (SL) to estimate FPUBF at arbitrary realizations in the interested distribution parameter region. Instead of the huge computational cost of a double-loop framework, the SL estimates the entire FPUBF only by one simulation analysis. Moreover, importance sampling (IS) variance reduction technique is introduced, and a single-loop IS probability density function (PDF), or SL-IS-PDF, is constructed to more efficiently estimate FPUBF by reducing the required size of the candidate sample pool. For approximating the optimal SL-IS-PDF and identifying the states of candidate samples efficiently, the double-loop adaptive Kriging model of performance function is introduced to further reduce the number of performance function evaluations. A numerical example and two composite structure examples are employed to verify the accuracy, efficiency, and feasibility of the proposed methods. [ABSTRACT FROM AUTHOR]

The geometrical arrangement of metamaterials controls their mechanical properties, such as Young's modulus and the shear modulus. However, optimizing the geometrical arrangement for user-defined performance criteria leads to an inverse problem that is intractable when considering numerous combinations of properties and underlying geometries. Machine-learning techniques have been proven to be effective and practical to accomplish such nonintuitive design tasks. This paper proposes an inverse design framework using conditional generative adversarial networks (CGANs) to explore and optimize two-dimensional metamaterial designs consisting of spinodal topologies, called spinodoids. CGANs are capable of solving the many-to-many inverse problem, which requires generating a group of geometric patterns of representative volume elements with target combinations of mechanical properties. The performance of the networks was validated by numerical simulations with the finite element method. The proposed inverse design framework vastly improves the efficiency of design exploration and optimization of spinodoid metamaterials. [ABSTRACT FROM AUTHOR]

Accurately predicting the forces and moments acting on trailing-edge devices under different flight conditions is a critical aspect in the design of the kinematics and actuation for high-lift or variable-camber applications. However, accurate modeling without elaborated computational fluid dynamics (CFD) analyses in the subsonic and transonic regimes needs a sophisticated model. Thus, the objective of this paper is to create such a model that accurately predicts the forces and moments acting on flaps during different flight conditions while remaining applicable to the preliminary aircraft design. The target values in this model are the three-dimensional (3D) forces and moments on the flap, which were obtained through 3D CFD simulations. The chosen input values required for the model include two-dimensional airfoil data, and wing geometry data for three different aircraft types: short-, medium-, and long-range, including a high-aspect-ratio configuration. Among several potential approaches, a neural network was deemed to be the most promising for predicting the target values. The neural network was used as a regression tool to accelerate the model development process in the preliminary aircraft design. Consequently, multiple studies were conducted on how the setup of the neural network, including the number of neurons, activation functions, and initialization, influences the results. The results reveal that the developed neural network accurately predicts the flap forces and moments with a mean deviation of under 2% for the vertical force Fz and the lateral force Fx and under 4% for the moment My. [ABSTRACT FROM AUTHOR]

Next-generation aircraft designs often incorporate multiple large propellers attached along the wingspan (distributed electric propulsion), leading to highly flexible dynamic systems that can exhibit aeroelastic instabilities. This paper introduces a validated methodology to investigate the aeroelastic instabilities of wing-propeller systems and to understand the dynamic mechanism leading to wing and whirl flutter and transition from one to the other. Factors such as nacelle positions along the wing span and chord and its propulsion system mounting stiffness are considered. Additionally, preliminary design guidelines are proposed for flutter-free wing-propeller systems applicable to novel aircraft designs. The study demonstrates how the critical speed of the wing-propeller systems is influenced by the mounting stiffness and propeller position. Weak mounting stiffnesses result in whirl flutter, while hard mounting stiffnesses lead to wing flutter. For the latter, the position of the propeller along the wing span may change the wing mode shapes and thus the flutter mechanism. Propeller positions closer to the wing tip enhance stability, but pusher configurations are more critical due to the mass distribution behind the elastic axis. [ABSTRACT FROM AUTHOR]

This paper presents a stochastic approach for modeling the turbulent airwake suitable for real-time simulation of the helicopter-ship dynamic interface. This approach relies on the measurements of unsteady loads collected during a wind-tunnel test campaign with a scaled helicopter operating over the deck of simple frigate shape 1. Power spectral densities of the measured aerodynamic loads combined with the estimated frequency response functions are used to find, through an optimization algorithm, a model of airwake spectra over the range of frequencies which mainly affects the pilot workload during shipboard operations. Then, a set of autoregressive filters is designed for every particular rotor position and wind-over-deck condition, so that when driven by white noise, the spectrum of the output will reproduce those obtained from the optimization. This approach is applied to three different tested wind directions and three rotor positions by implementing the autoregressive filters into the multibody model of the experimental rotor. Frequency response analysis of the aerodynamic loads demonstrates that the turbulent airwake model obtained from the experimental data can predict the unsteadiness of loads comparable to those measured in the wind tunnel across the bandwidth of interest for pilot activities. The identified airwake models could be applied to a full-scale model to simulate the unsteady loads effectively experienced by the helicopter during a ship landing flight. [ABSTRACT FROM AUTHOR]

During aircraft braking, brake-induced frequent wheel slippage may cause coupling vibration between the antiskid brake system and landing gear. Brake-induced longitudinal coupled vibration is also known as "gear walk" vibration. Gear walk vibration not only reduces the structural life of the landing gear but also impairs the efficiency of the aircraft braking system. This paper focuses on the landing gear walk vibration mechanism and suppresses the method by modifying the antiskid brake control law. Firstly, a comprehensive multiwheel landing gear walk vibration model is developed. This model describes not only the longitudinal and vertically coupled vibration characteristics of the landing gear but also the dynamics of the aircraft braking system and the fuselage. Secondly, by simplifying the landing gear walk to the transverse vibration of a cantilever beam with a tip mass, the landing gear walk vibration mechanism is revealed in the frequency domain. Finally, a method to suppress landing gear walk vibration by modifying the pressure-bias-modulated (PBM) antiskid braking control law is proposed, which can solve the serious landing gear walk vibration problem caused by the conventional PBM control law. Simulation and experimental results show that the proposed method can suppress the landing gear walk vibration while improving braking efficiency. [ABSTRACT FROM AUTHOR]

An investigation of the influence of the hydrofoil on load reduction performance during an amphibious aircraft landing on still and wavy water is conducted by solving the unsteady Reynolds-averaged Navier-Stokes equations coupled with the standard 𝑘-𝜔 turbulence model in this paper. During the simulations, the numerical wave tank is realized by using the velocity-inlet boundary wave maker coupled with damping wave elimination techniques on the outlet, while the volume of fluid model is employed to track the water-air interface. Subsequently, the effects of geometric parameters of hydrofoil have been first discussed on still water, which indicates the primary factor influencing the load reduction is the static load coefficient of hydrofoil. Furthermore, the effects of descent velocity, wave length, and wave height on load reduction are comprehensively investigated. The results show that the vertical load reduces by more than 55% at the early stage of landing on the still water by assembling the hydrofoil for different descent velocity cases. Meanwhile, for amphibious aircraft with high forward velocity, the bottom of the fuselage will come into close contact with the first wave when landing in crest position, and then the forebody will impact the next wave surface with extreme force. In this circumstance, the load reduction rate decreases to around 30%, which will entail a further decline with the increase in wave length or wave height. [ABSTRACT FROM AUTHOR]

This paper presents an approach for estimating the power setting, indicated as N1, the rotational speed of the low-pressure shaft, of departing civil jet aircraft to be used as input for accurate noise calculations. The method utilizes the machine learning approach random forest regression and is trained using flight data recorder data divided into three departure sections. Each section uses a unique set of three to four model features based on easily accessible data, such as position and airport meteorological data. Unlike previous methods, neither fixed configuration altitudes, aerodynamic coefficients, or engine coefficients, nor takeoff mass or wind information are required as inputs. To assess the performance of the method, noise calculations for departures of five aircraft types, including narrow- and wide-body jet aircraft, were performed. Comparing overall and aircraft-specific noise levels, calculated from estimated N1 versus recorded N1 values from flight data recorder data, revealed small differences of less than 1 dB. Overall, the present machine-learning-based approach provides reliable estimates of power settings for departing civil jet aircraft, offering improved accuracy and avoiding the need for numerous assumptions and hardly accessible data. [ABSTRACT FROM AUTHOR]

Experimental testing of aircraft configurations requires the use of representative engine units to simulate powered-on conditions. For ultra-high-bypass-ratio engines featuring tighter aircraft integration, the installation effects need to be carefully assessed. This paper presents a computational evaluation of two powered simulator models for wind tunnel tests of an ultra-high-bypass-ratio turbofan, based on a traditional air-driven turbo-powered simulator (TPS) and an electric-powered simulator (EPS). Starting from two-dimensional axisymmetric analyses, the design of the baseline turbofan engine and the choice of the duplication parameters of the simulators are described. Several aspects related to the nonreproducibility of the core flow at wind tunnel level are discussed, in terms of mass flow matching, TPS core nozzle pressure ratio, and plug shape of the EPS. The EPS can provide a more faithful reproduction of the flowfield under installation, with an estimated installation drag from 4 to 2 drag counts lower than the reference geometry, compared to the 6 drag counts higher prediction of the TPS. [ABSTRACT FROM AUTHOR]

A multimesh framework to optimize two-dimensional, flat-steered fiber laminates for a given loading condition is extended to include complex curved laminates in three-dimensional (3D) space. The optimal fiber angle distribution solution is used by a tow path planner to generate (steered) tow paths in curved laminates in 3D space. The framework utilizes two meshes in the optimization step: i) a coarse quadrilateral element-based manufacturing mesh (MM) used to discretize the fiber angle distributions with nodal values; and ii) a fine triangular element-based stress mesh used in the finite-element-based stress analysis to ensure a converged solution. Lagrangian interpolation functions are used to determine the fiber angles at the centroid of each finite element analysis (FEA) mesh element within a given MM. The number of design variables is minimized without jeopardizing the accuracy of the FEA solution. A third mesh, the tessellation mesh, is used to project the generated tow paths onto the curved laminates in 3D space accurately. This paper presents all the steps needed to go from optimized fiber angles to tow paths. In conclusion, a methodology is proposed and implemented to optimize variable-stiffness laminates on complex curved structures. [ABSTRACT FROM AUTHOR]