Abstract: This letter proposes a novel capacitive power transfer system for the unmanned aerial vehicles (UAVs) with lightweight in receivers and robust misalignment tolerance characteristics. The design and analysis of a reconfigurable capacitive coupler are presented at first. The transmitter adopts four rectangular copper plates, which can be reconfigured according to the landing positions and directions. The receiver includes two hollow cylindrical copper-foil sleeves wrapped at the bottom of UAV landing gears, thereby leading to the adaption of the UAV's structure and the tolerance to lateral-, longitudinal-, and angular misalignment. Then, simulation and prototype experiments are carried out to verify the proposal. The results show that the prototype has low leakage electric flux at the UAV fuselage and can deliver 212.1 W with a dc-to-dc efficiency of 82.5%. Especially the proposal allows UAV to be powered within ±0.76 A of output current variation for different landing positions and directions, and the system dc-to-dc efficiency variation is within ±1%.

Abstract: The issue of modeling and fault-tolerant control (FTC) design for a class of flexible air-breathing hypersonic vehicles (FAHVs) with actuator faults is investigated in this article. Different from previous research, the shear deformation of the fuselage is considered, and an ordinary differential equations-partial differential equations (ODEs-PDEs) coupled model is established for the FAHVs. A feedback control is proposed to ensure flight stable and an adaptive FTC method is designed to deal with actuator faults while suppressing the system’s vibrations. Besides, the stability analysis of the closed-loop system is given via the Lyapunov direct method and an algorithm that transfers the bilinear matrix inequalities (BMIs) feasibility problem to the linear matrix inequalities (LMIs) feasibility problem is provided for determining the control gains. Finally, the numerical simulation results show that the proposed controller can stabilize the flight states and suppresses the vibration of the fuselage efficiently.

Abstract: This paper focuses on the quadrotor drag coefficient model and its estimation from flight tests. Precise assessment of such a model permits the use of a quadrotor as a sensor for wind estimation purposes without the need for additional onboard sensors. Firstly, the drag coefficient has been estimated in a controlled environment via wind generator and motion capture system. Later, the evolution of the coefficient is observed for various mass and fuselage shapes. Finally, an estimation method is proposed, based on the least-squares optimization, that evaluates the drag of the quadrotor directly from outdoor flight data. The latter leads the methodology towards easier adoption in other researchers’ systems without the need for complex and expensive flight testing facilities. The accuracy of the proposed method is presented both in simulation, based on a realistic flight dynamics model, and also for real outdoor flights.

Abstract: The implementation of morphing wing applications in aircraft design has sparked significant interest as it enables the dimensional properties of the aircraft to be modified during flight. By allowing manipulation of the 2D and 3D parameters on the aircraft’s wings, tail surfaces, or fuselage, a variety of possibilities have arisen. Two primary schools of thought have emerged in the field of morphing wing applications: the mechanisms school and the smart surfaces approach that uses shape-memory materials and smart actuators. Among the research in this field, the Fishbone Active Camber (FishBAC) approach has emerged as a promising avenue for controlling the deflection of the wing’s trailing edge. This study revisits previous research on morphing wings and the FishBAC concept, evaluates the current state of the field, and presents an original design process flow that includes the design of a unique and innovative UAV called the Stingray within the scope of the study. A novel morphing concept developed for the Stingray UAV, Rear Spar Articulated Wing Camber (RSAWC), employs a fishbone-like morphing wing rib design with rear spar articulation in a cost-effective manner. The design process and flight tests of the RSAWC are presented and directly compared with a conventional wing. Results are evaluated based on performance, weight, cost, and complexity. Semi-empirical data from the flight testing of the concept resulted in approximately a 19% flight endurance increment. The study also presents future directions of research on the RSAWC concept to guide the researchers.

Abstract: A conformal metasurface antenna exhibiting a pencil beam radiation pattern at 10.0 GHz has been designed using the Voronoi partition approach, and fabricated on the Kahu Uninhabited Aerial System (UAS) fuselage. Two manufacturing methods are presented and compared. The first approach utilized a 3-axis Trotec fiber laser to etch the flattened metasurface geometry in copper foil. The etched pattern was then ‘stretched’ over the UAS geometry. The second approach utilized a 6-axis nScrypt (retrofitted with an IDS aerosol jetting tool) to conformally print the metasurface pattern directly on the UAS fuselage. An electroless copper plating step was then utilized to improve the radiofrequency (RF) conductivity of the printed silver. Both manufacturing methods yielded functional metasurface antennas with equivalent performance at the operating frequency. However, the first method is limited to geometries that can be ‘flattened’ with acceptable tolerances, whereas the second approach is amenable to all practical geometries. This demonstration of two manufacturing techniques is a critical step forward in the cost-effective deployment of truly conformal metasurface antennas on realistic geometries.

Abstract: Boundary layer ingesting (BLI) propulsion represents a promising concept that can achieve a great fuel burn reduction compared to conventional aircraft. This study numerically investigates the BLI inflow distortion and its impact on propulsion fan performance using commercial software CFX. First, Reynolds-averaged Navier–Stokes (RANS) simulations were performed to obtain the inflow distortion of a propulsive fuselage aircraft with various geometries and flight conditions. Then, full-annulus URANS simulations were carried out to assess the effect of various inflow distortions on the performance of NASA rotor 67. The results show that the fuselage sweep dominates the BLI inflow distortion, while the effects of the wing and vertical tail are local. As the angle of attack increases, the BLI inflow distortion is alleviated due to the freestream being ingested from the bottom of the propulsor. With the increase in the angle of sideslip, the redistributing and impinging flow significantly aggravates the total pressure and swirl distortion. With regard to the fan performance, various BLI inflow distortions share a similar static pressure distribution upstream of the rotor, as a result of the comparable total pressure distribution, so the cases without swirl distortion show a similar distribution of migration-induced swirl angle. At the peak efficiency condition, a minimum reduction in efficiency of approximately 4.5% and a total pressure ratio reduction of over 2.9% are observed for the investigated inflow distortions, in comparison with the Clean case. Moreover, the findings reveal that the effect of swirl distortion on fan performance is minor for the cruise case, but substantial for the two sideslip cases. These results offer useful insights into BLI propulsion.

Abstract: The analysis of acoustic performance of sandwich panels, designed explicitly for fuselage sidewalls of turbojet aircraft, has been revisited numerous times. However, the layouts and the sound transmission loss of the airborne fuselage sandwich panels have not been made available for public release. Against this idiosyncratic practice, this study envisioned disseminating the know-how on the sound attenuation ability of dissimilar sandwich panels implemented in the fuselage of turbojet aircraft. To efficiently absorb and dampen the acoustic energy, the sandwich panels were systematically configured using a honeycomb core, felts and a closed cell aluminum-foam in the material lay-up. The face sheets of the sandwich panels were composed of either glass fiber-reinforced epoxy or carbon fiber-reinforced epoxy composites. The sandwich panels were tested as acoustic barriers to a diffuse sound field in a sound transmission suite consisting of two adjacent sound reverberation rooms. As found, a rubber damping layer in the material lay-up of a sandwich panel reinforced the sound transmission loss at frequencies beyond 500 Hz. Multiple felts augmented the sound transmission loss in the frequency regime beyond 315 Hz. The panel edges needed to be sealed properly to reproduce the sound transmission loss of a sandwich panel in successive experiments. This study leveraged many-faceted insights into the configuration (constituent materials, number of material layers, thickness, and stacking sequence) of commercially feasible sandwich panels. Besides, this study proposed optimized sandwich configurations based on the sound transmission loss, which will help aircraft manufacturers to apply these sandwich panels without further modifications.

Abstract: Abstract This research applies a numerical study of topology optimization of laminate composite structures by using a finite element method (FEM). In this methodology, the plies orientation is excluded from the optimization. The geometry-based optimization from frames of a MALE UAV fuselage structure is presented. The minimum strain energy with an optimization constraint of 20% of weight reduction is used in the objective function. Before the primary analysis, benchmark studies of topology optimization without considering orientations from previously published literature are performed. The convergence studies were taken to acquire the appropriate mesh size in the FEM technique, which utilized a four-noded shell element. The finite element analysis and optimization results showed that the structural design of the newly framed composite fuselage MALE UAV meets the structural strength requirements specified in the airworthiness standard STANAG 4671.

Abstract: This work aims to design a rotary-wing unmanned aerial system (RUAS) that monitors the pollutants and minimizes their concentration in the atmosphere. This RUAS could be well suited for implementation in cities such as New Delhi and Ghaziabad, where air pollution is a major concern. This RUAV’s well-thought-out design and use would be good for the environment also a step forward in the technology of UASs. Therefore, an advanced approach in design as well as innovative computational composite materials development based on structural analysis of this RUAS has been made. The major components involved in this comprehensive investigation are the fuselage, main rotor and tail rotor of RUAS. The aerodynamic parameters on RUAS have been estimated through the advanced technique adopted computational fluid dynamics approach using ANSYS Fluent 17.2. The finite element analysis (FEA) of the RUAS imposed under two different approaches enforced on lightweight composite materials has been estimated through ANSYS Structural 17.2. Firstly, the advanced computational platform for the development of composite materials has been created through the ANSYS Composite Preprocessor tool 17.2, wherein computational moldings of the fuselages of RUAV are framed. The computational moldings are greatly supported and so the conventional polymer matrix composites, metal matrix based composites, and advanced hybrid composites are well prepared. A ll of these uniquely framed materials have undergone computational structural investigations, and the material suitable for RUAVs has thus been selected. The computational tests are validated with advanced experimental outcomes, which furthermore enhanced the reliability of this proposed work. Additionally, the main rotor and entire RUAV are also computationally investigated under aerodynamic loading conditions through fluid structure interaction (FSI) approach. At last, the suitable lightweight material for all the parts of RUAS is shortlisted through innovative integrated computational engineering analyses.

Abstract: Electric Vertical Takeoff and Landing (eVTOL) aircraft experience complex, unsteady aerodynamic interactions between rotors, wings, and fuselage that can make design difficult. We introduce a new framework for predicting aerostructural interactions. Specifically, we demonstrate the coupling of a finite element solver with Reissner-Mindlin shell theory for computing deflections and a viscous vortex particle for capturing wakes. We perform convergence studies of the aerodynamics and the coupled aerostructural model. Finally, we share some preliminary results of the dynamic aeroelastic response of Uber's eCRM-002 main wing, and share some qualitative observations.

Abstract: The interaction between the slipstream of the propellers and the wing of an aircraft with distributed electric propulsion (DEP) could benefit aerodynamics. A conceptual design and optimization are carried out in order to increase the range of an electric general aviation aircraft without affecting its takeoff and landing velocity in the same fuselage condition. Propellers are modelled using the actuator disk (AD) theory, and the aircraft is modelled using the vortex lattice method (VLM) to obtain DEP aircraft’s aerodynamics in conceptual design. The DIRECT method is used for global optimization. To concentrate on the layout of the propellers and wing, a propeller with the same chord distribution, twist distribution, and number of blades is selected. The design and optimization of DEP aircraft’s range is carried out with the objective of achieving the maximum product of the lift–drag ratio with propeller efficiency under force balance constrains. Additionally, to decrease the takeoff and landing distance, the DEP aircraft’s takeoff and landing performance are optimized with the objective of the smallest velocity at an angle near the tail down angle under the constrains of acceleration bigger than 0 and a Mach number at the tip of blades smaller than 0.7. The CFD simulation was used to confirm the DEP aircraft’s pretty accurate aerodynamics. Compared to the reference aircraft, the improved DEP aircraft with 10 high-lift propellers on the leading edge of the wing and 2 wing-tip propellers may boost cruise performance by 6% while maintaining takeoff and landing velocity. Furthermore, it has been shown that the stall speed of DEP aircraft with smaller wings would rise proportionally when compared to conventional design aircraft, and the power need of DEP aircraft will be increased as a result of the operation of high-lift propellers. The conceptual design and optimal approach suggested in this work has some reference value for the design and research of the fixed-wing DEP general aviation aircraft.

Abstract: Many urban air mobility vehicle designs feature propellers integrated with fuselage, wings, and other appendages. These vehicle designs are based on complex configurations with novel propulsion systems and flight technology. The tightly coupled nature of the systems in these vehicles and the novel technologies within them create opportunities for vehicle analysis and optimization. Furthermore, the geometries being developed are unique, utilizing advanced materials for integral components. Aerodynamic, structural, and acoustic analysis and optimization can provide insight into their performance, and how the propellers, structure, and lifting surfaces can be designed to further improve vehicle efficiency considering operational constraints. This work details an aero-structural-acoustic optimization toolchain built with the MACH and OpenMDAO tools, embedded within the modeling and optimization framework called MPhys. This analysis and optimization framework utilizes multiple model fidelities, including hybrid blade element momentum theory, computational fluid dynamics, finite-element modeling, and an aeroacoustic analogy. With this toolchain, this work presents gradient-based aero-structural-acoustic optimization, minimizing required flight power while ensuring structural integrity and respecting acoustic considerations. The toolchain is applied to the NASA tiltwing concept vehicle to optimize the wing and propeller designs, yielding a 17.8% reduction in required power for cruise flight while considering aerodynamic, structural, and acoustic constraints. This aero-structural-acoustic optimization framework can help make vehicle designs more efficient, lighter, and quieter.

Abstract: This paper reports on the application and challenges of a guided wave-based structural health monitoring (GWSHM) methodology on an industrial 5 m long curved composite fuselage panel, for detecting barely visible impact damage (BVID), based on integrated sensors data. An extensive network consisting of 72 piezoelectric transducers installed on the panel via diagnostic film, a layer including sensors and inkjet printed flexible circuits, that provides uniform bonding quality and repairability. Threshold setting for damage detection is proposed based on statistical distribution of pristine signals, obtained throughout the span of one month. The threshold is determined through an Outlier Analysis and validated through an iterative cross-validation approach. The model validation yields a Gamma distribution with 1.5% false positive rate. A total of 20 impacts at different energies are conducted on the panel for BVID generation and these are respectively categorized based on their location. Methodologies were developed in order to detect, characterize, and localize the BVID generated from the impacts at various locations of the panel. The reliability of the SHM system is quantified by the Probability of Detection for each damage category by computing the damage area with 90% probability and 95% confidence level (a90|95 values). For damages in the skin and at the foot of the stringer a90|95 values of 233 mm2 and 365 mm2 are, respectively, obtained. Finally, the summary of the experimental work is presented through the detection of all imparted damages.

Abstract: The aerodynamic noise of an aircraft leads to vibration fatigue damage to structures. Herein, a prediction method for aircraft surface noise under the comprehensive effect of mixed acoustic sources during flight, primarily surface aerodynamic, air intake, and tail nozzle jet noises, was studied. In the supersonic cruising state, the internal and external flow fields of the aircraft were solved using the Reynolds-averaged Navier–Stokes equations to obtain the statistical average solution of the initial turbulence. The non-linear disturbance equation was used to obtain the surface acoustic load of the aircraft. The calculation results revealed that the main source of aircraft surface noise is aerodynamic noise. The sound pressure level on the fuselage increases gradually from front to rear along the aircraft, and the OASPL at the air intake and tail nozzle is relatively large. The jet noise has little effect on the sound pressure level at the front of the fuselage and only contributes to the OASPL at the tail nozzle of the fuselage. The intensity of pressure pulsations from the engine exhaust in the tail section is 93.3% of the total intensity of pressure pulsations.

Abstract: The integrated infrared suppressor can reduce the infrared radiation signal of a helicopter and is compatible with radar-acoustic stealth. However, the issues that are caused by the integrated infrared suppressor, such as temperature increases on the rear fuselage surface and a lack of shielding at the exhaust port, need to be addressed, in order to further improve the infrared stealth capability of the helicopter. Aiming at this, the effects of the ambient temperature, fuselage surface emissivity, mixing duct shielding, and exhaust port shielding on the infrared radiation characteristics of the helicopter are studied with numerical simulation. The results show that the infrared radiation intensity of the helicopter, in 3–5 μm band and 8–14 μm band, decreases by about 20% and 10%, respectively, for every 6 K reduction in the ambient temperature. As the emissivity of the rear fuselage surface reduces from 0.8 to 0.5, the helicopter’s infrared radiation intensity, in a 3–5 μm band and a 8–14 μm band, decreases by about 6% and by about 4% and 1.3%, respectively, after the mixing duct is equipped with a shielding sheath. Installing deflectors at the exhaust port of the fuselage can prevent the detection rays from detecting the high-temperature components inside the fuselage, and when the emissivity of the deflectors is reduced from 0.8 to less than 0.5, or the deflectors are cooled by more than 80 K, they begin to play a role in suppressing the infrared radiation at the bottom of the helicopter.

Abstract: High aspect ratio wings have been a major topic for research due to their capability to improve the aerodynamic efficiency of modern aircraft. Many numerical studies have shown their flexibility causing nonlinearity through geometric effects and their impact on internal loads and dynamics, such as reduced flutter speed and coupling with the aircraft body causing body freedom flutter. Experimental work is present in the literature for validation of cantilever wing models but only a few have implemented wind tunnel testing on dynamically-mounted full-span aircraft models. The work presented here develops a rigid-flexible coupled numerical model for a wind tunnel test platform known as the 5-degree-of-freedom manoeuvre rig. This model is used for the simulation of a full-span flexible model aircraft constrained by the rig for the investigation of the coupling between rigid body and flexible modes together with geometric nonlinear effects. The modeling of the wing flexibility is based on a reduced order geometrically exact structural method linked with a vortex lattice aerodynamic model. The aircraft fuselage and empennage, and the manoeuvre rig, are modelled as rigid bodies. The findings of the study will aid future experimental wind tunnel explorations.