Abstract: Interactional aerodynamics of multi-rotor flows has been studied for a quadcopter representing a generic quad tilt-rotor aircraft in hover. The objective of the present study is to investigate the effects of the separation distances between rotors, and also fuselage and wings on the performance and efficiency of multirotor systems. Three-dimensional unsteady Navier-Stokes equations are solved using a spatially 5th order accurate scheme, dual-time stepping, and the Detached Eddy Simulation turbulence model. The results show that the separation distances as well as the wings have significant effects on the vertical forces of quadroror systems in hover. Understanding interactions in multi-rotor flows would help improve the design of next generation multi-rotor drones.

Abstract: Purpose Motivated by the potential of gaining noticeable improvements in vehicular efficiency, this paper aims to investigate the benefits attainable from introducing a more synergistic propulsion/airframe integration. In previous work, the concept of a boundary layer ingesting propulsor encircling the aft section of an axisymmetric fuselage was identified to be particularly promising for the realisation of aircraft wake filling, and hence, a significant reduction of the propulsive power required. Design/methodology/approach After reviewing the theoretical principles of the propulsive fuselage concept, a book-keeping and model matching procedure is introduced, which is subsequently used to incorporate the numerically computed aerodynamic characteristics of a propulsive fuselage aircraft configuration into a propulsion system (PPS) sizing and performance model. As part of this, design heuristics for important characteristics intrinsic to propulsive fuselage power plants are derived. Thereafter, parametric study results of the PPS are discussed, and the obtained characteristics are compared to those of a conventionally installed power plant. Finally, the impact of the investigated PPS on the integrated performance of a propulsive fuselage aircraft concept is studied, and the results are compared and contrasted to previously conducted analyses based on semi-empirical characteristics. Findings It was found that the aircraft-level benefit originally predicted based on semi-empirical methods could be confirmed using the numerically derived PPS design heuristics, specifically an improvement in vehicular efficiency of 10.4 per cent over an advanced conventional reference aircraft. Practical implications The approach presented in the paper may serve as a guideline when incorporating the results of high-fidelity aerodynamic methods into a PPS sizing and performance model suitable for aircraft-integrated assessment of a propulsive fuselage concept. The vehicular efficiency potentials offered through the synergistic PPS integration approach are highlighted. Originality/value The paper contributes to a deeper understanding of the characteristics of a boundary layer ingesting fuselage fan (FF) power plant relative to a conventionally installed PPS. In addition, a set of PPS design correlations are presented allowing for the integrated sizing of a FF power plant.

Abstract: Systems, methods, and devices provide a vehicle, such as an aircraft, with rotors configured to function as a tri-copter for vertical takeoff and landing (“VTOL”) and a fixed-wing vehicle for forward flight. One rotor may be mounted at a front of the vehicle fuselage on a hinged structure controlled by an actuator to tilt from horizontal to vertical positions. Two additional rotors may be mounted on the horizontal surface of the vehicle tail structure with rotor axes oriented vertically to the fuselage. For forward flight of the vehicle, the front rotor may be rotated down such that the front rotor axis may be oriented horizontally along the fuselage and the front rotor may act as a propeller. For vertical flight, the front rotor may be rotated up such that the front rotor axis may be oriented vertically to the fuselage, while the tail rotors may be activated.

Abstract: A tail-sitter aircraft is provided and includes a fuselage having first and second axisymmetric sides, first collectively controllable prop-rotors, which are formed to define a first pair of rotor disks and which are respectively supported at the first axisymmetric side of the fuselage, and second collectively controllable prop-rotors, which are formed to define a second pair of rotor disks and which are respectively supported at the second axisymmetric side of the fuselage.

Abstract: A vertical take-off and landing (VTOL) aircraft is provided and includes a fuselage, wings extending outwardly from the fuselage to define a wing plane and a prop-rotor operably disposed to generate thrust, a flight computer and controllable surfaces disposed on at least one of the fuselage, the wings and the prop-rotor. The controllable surfaces are controllable by the flight computer to position the wing plane in accordance with a predominant local wind direction.

Abstract: Purpose This paper aims to deal with the experimental estimation of both longitudinal- and lateral-directional aerodynamic characteristics of a new twin-engine, 11-seat commuter aircraft. Design/methodology/approach Wind tunnel tests have been conducted on a 1:8.75 scaled model. A modular model (fuselage, wing, nacelle, winglet and tail planes) has been built to analyze both complete aircraft aerodynamic characteristics and mutual effects among components. The model has been also equipped with trailing edge flaps, elevator and rudder control surfaces. Findings Longitudinal tests have shown the goodness of the aircraft design in terms of aircraft stability, control and trim capabilities at typical clean, take-off and landing conditions. The effects of fuselage, nacelles and winglets on lift, pitching moment and drag coefficients have been investigated. Lateral-directional stability and control characteristics of the complete aircraft and several aircraft component combinations have been tested to estimate the aircraft components’ interactions. Research limitations/implications The experimental tests have been performed at a Reynolds number of about 0.6e6, whereas the free-flight Reynolds number range should be between 4.5e6 and 9.5e6. Thus, all the measured data suffer from the Reynolds number scaling effect. Practical implications The study provides useful aerodynamic database for P2012 Traveller commuter aircraft. Originality/value The paper deals with the experimental investigation of a new general aviation 11-seat commuter aircraft being brought to market by Tecnam Aircraft Industries and it brings some material on applied industrial design in the open literature.

Abstract: An aircraft has a bearing structure, the bearing structure having at least one central fuselage and two pylons each situated at a distance laterally from the fuselage. In addition, the aircraft has a wing structure, at least four hub rotors, and at least one thrust drive. Each hub rotor is fastened to the bearing structure, has a propeller having two propeller blades, and produces, through rotation of the propeller, an upward drive force acting in the vertical direction on the aircraft. The thrust drive is produces a thrust force acting in the horizontal direction on the bearing structure. The pylons each have two hub rotors, the hub rotors being configured to arrest respective propeller blades of a hub rotor in a position relative to the pylons. In the arrested position, the propeller blades of a hub rotor do not extend beyond the outer dimensions of the pylons.

Abstract: In Aviation sector, composite materials and its application to each component are one of the prime factors of consideration due to the high strength to weight ratio, design flexibility and non-corrosive so that the composite materials are widely used in the low weight constructions and also it can be treated as a suitable alternative to metals. The objective of this paper is to estimate and compare the suitability of a composite skin joint in an aircraft fuselage with different joints by simulating the displacement, normal stress, vonmises stress and shear stress with the help of numerical solution methods. The reference Z-stringer component of this paper is modeled by CATIA and numerical simulation is carried out by ANSYS has been used for splice joint presents in the aircraft fuselage with three combinations of joints such as riveted joint, bonded joint and hybrid joint. Nowadays the stringers are using to avoid buckling of fuselage skin, it has joined together by rivets and they are connected end to end by splice joint. Design and static analysis of three-dimensional models of joints such as bonded, riveted and hybrid are carried out and results are compared.

Abstract: The invention provides a vertical-takeoff auxiliary system for a fixed-wing aircraft. The vertical-takeoff auxiliary system comprises a multi-shaft rotor-wing aircraft and a fixed-wing aircraft. A plurality of throwing devices are installed in the middle of the fuselage bottom of the multi-shaft rotor-wing aircraft, and a plurality of hanging parts are installed in the fuselage middle of the fixed-wing aircraft; the hanging parts are connected with the throwing devices in a hung mode, and the fixed-wing aircraft is hung below the multi-shaft rotor-wing aircraft accordingly; the multi-shaft rotor-wing aircraft carries the fixed-wing aircraft to vertically take off to the preset height and the preset speed, then the throwing devices release the hanging parts, and the fixed-wing aircraft is separated from the multi-shaft rotor-wing aircraft to autonomously fly. By means of the vertical-takeoff auxiliary system in the mode, vertical takeoff of the fixed-wing aircraft can be achieved under the condition that an existing fixed-wing aircraft does not need to be greatly modified.

Abstract: Crashworthiness is one of the main concerns in civil aviation safety particularly with regard to the increasing ratio of carbon fiber reinforced plastic (CFRP) in aircraft primary structures. In order to generate dates for model validations, the mechanical properties of T700/3234 were obtained by material performance tests, and energy-absorbing results were gained by quasi-static crushing tests of composite sinusoidal specimens. The correctness of composite material model and single-layer finite element model of composite sinusoidal specimens were verified based on the simulation results and test results that were in good agreement. A typical civil aircraft fuselage section with composite sinusoidal specimens under cargo floor was suggested. The crashworthiness of finite element model of fuselage section was assessed by simulating the vertical drop test subjected to 7 m/s impact velocity, and the influences of different thickness of sub-floor composite sinusoidal specimens on crashworthiness of fuselage section were also analyzed. The simulation results show that the established finite element model can accurately simulate the crushing process of composite sinusoidal specimens; the failure process of fuselage section is more stable, and the safety of occupants can be effectively improved because of the smaller peak accelerations that was limited to human tolerance, a critical thickness of sub-floor composite sinusoidal specimens can restrict the magnitude of acceleration peaks, which has certain reference values for enhancing crashworthiness capabilities of fuselage section and improving the survivability of passengers.

Abstract: This paper deals with the design of an observed based optimal state feedback controller having pole location constraints for an active vibration mitigation problem of an aircraft system An eleven-degree-of-freedom detailed full aircraft mathematical model having active landing gears and a seated pilot body is developed to control and analyze aircraft vibrations caused by runway excitation, when the aircraft is taxiing Ground induced vibration can contribute to the reduction of pilot’s capability to control the aircraft and cause the safety problem before take-off and after landing Since the state variables of the pilot body are not available for measurement in practice, an observed based optimal controller is designed via Linear Matrix Inequalities (LMIs) approach In addition, classical LQR controller is designed to investigate effectiveness of the proposed controller The system is then simulated against the bump and random runway excitation The simulation results demonstrate that the proposed controller provides significant improvements in reducing vibration amplitudes of aircraft fuselage and pilot’s head and maintains the safety requirements in terms of suspension stroke and tire deflection

Abstract: An aircraft includes a fuselage, a forward wing assembly, and aft wing assembly, and a propulsion system. The propulsion system includes a first primary thrust propulsor and a first secondary thrust propulsor, the first primary thrust propulsor being different than the first secondary thrust propulsor. Both the first primary thrust propulsor and the first secondary thrust propulsor are mounted to the same one of: a starboard side of the aft wing assembly, a port side of the aft wing assembly, a starboard side of the forward wing assembly, or a port side of the forward wing assembly.

Abstract: Current design concepts for transport aircraft aim at increasing the aircraft efficiency and performance by the introduction of advanced composite materials, such as carbon fibre reinforced plastics (CFRP). These novel transport aircraft designs may show dissimilar dynamic response behaviour due to differences in failure modes and energy absorption characteristics compared with the current transport aircraft designs made of aluminium alloys. For that reason, crash concepts are being developed to utilise the high specific energy absorption of composite materials for predefined load conditions. In the context of this paper, crash concepts for future CFRP transport aircraft were developed in which most of the kinetic energy is absorbed by tension energy absorbers integrated in the cabin and cargo floor, and by crushing energy absorbers integrated in the cabin floor support struts. The developed crash concepts define mainly parallel activation of different crash devices to achieve smooth energy absorption for different crash load scenarios. Crushing of the energy absorbers integrated in the cabin floor support struts is controlled by a novel structural design in this fuselage area. So far, this research is limited to conceptual studies performed on the basis of a generic CFRP fuselage design. Numerical simulations using the explicit finite-element (FE) code Abaqus/Explicit were performed to derive qualitative and quantitative results for an assessment of the crash concepts. A hybrid FE/macro model approach was used that combines typical FE discretisation with macro models for main failure representation. Two different crash kinematics were considered which distinguish between the failure patterns of the frame structure of the lower fuselage shell. The simulation results presented in this paper in terms of energy plots, passenger accelerations, and crash sequences identify favourable crash performance for a load scenario with fully loaded cabin and an impact velocity of 9.1 m/s (30 ft/s). Significant amount of kinetic energy could be absorbed by tension loads. Parallel activation of crash devices resulted in smooth crash kinematics with reduced trigger loads. By utilisation of the cabin floor support strut area as an energy absorption zone, sufficient energy absorption capacity could be provided even for load scenarios with increased impact energies. The results, presented in this paper, are the basis for further detailed research work on this tension crash concept.

Abstract: High-fidelity aerodynamic shape optimization based on the Reynolds-averaged NavierStokes equations is used to optimize the aerodynamic performance of conventional and blended wing-body (BWB) aircraft for a range of aircraft sizes from regional to wide-body classes. Trim-constrained drag minimization is performed, with optimized conventional tube-and-wing (CTW) designs serving as performance references. First, a set of ‘classically’ shaped BWB configurations are optimized across the range of classes. The classically shaped regional and narrow-body-class BWBs offer only a marginal fuel-burn benefit relative to the equivalent conventional designs. The wide-body-class BWB offers up to 10.9% lower fuel-burn than the equivalent CTW. Exploratory optimizations with significant geometric freedom are then performed, resulting in a set of novel shapes with a more slender lifting fuselage and distinct wings. Based on these exploratory results, new lifting-fuselage configurations (LFCs) are designed. The slenderness of the LFC fuselage decreases with aircraft size, such that, for the largest class, the LFC reduces to a classical BWB shape. Due to their lower weight and higher aerodynamic efficiency, the LFC designs burn 6.1% and 9.7% less fuel in cruise than the equivalent CTWs for the regional and narrow-body classes, respectively. In addition, the LFCs are more aerodynamically efficient and burn less fuel than the BWBs. Additional optimizations were performed to determine the aerodynamically optimal cruise altitude of all of the aircraft. Due to their lower wing loading, the resulting increase in cruise altitude is most beneficial for the BWBs, such that the regional and narrow-body-class BWBs burn up to 5.5% less fuel than the CTWs. The LFCs offer up to a 10.3% fuel-burn reduction relative to the CTWs when at their optimal altitude.

Abstract: An unmanned aerial vehicle (UAV), or drone, includes a fuselage, left and right airfoil-shaped wings connected to the fuselage to generate lift in forward flight, a left thrust-generating device supported by the left wing, and a right thrust-generating device supported by the right wing. The UAV further includes a vertical stabilizer, a top thrust-generating device mounted to a top portion of the vertical stabilizer, and a bottom thrust-generating device mounted to a bottom portion of the vertical stabilizer. An onboard power source is provided for powering the thrust-generating devices. The left, right, top and bottom thrust-generating devices provide forward thrust during forward flight and also provide vertical thrust to enable the unmanned aerial vehicle to take-off and land vertically when the fuselage is substantially vertical and further enabling the unmanned aerial vehicle to transition between forward flight and vertical take-off and landing.

Abstract: An unmanned aerial vehicle with deployable components (UAVDC) is disclosed. The UAVDC may comprise a fuselage, at least one wing, and at least one control surface. In some embodiments, the UAVDC may further comprise a propulsion means and/or a modular payload. The UAVDC may be configured in a plurality of arrangements. For example, in a compact arrangement, the UAVDC may comprise the at least one wing stowed against the fuselage and the at least one control surface stowed against the fuselage. In a deployed arrangement, the UAVDC may comprise the at least one wing deployed from the fuselage and the least one control surface deployed from the fuselage. In an expanded arrangement, the UAVDC may comprise the at least one wing telescoped to increase a wingspan of the deployed arrangement.

Abstract: This paper deals with the study and comparison of passive and active landing gear system of the aircraft and dynamic responses due to runway irregularities while the aircraft is taxying. The dynamic load and vibration caused by the unevenness of runway will result in airframe fatigue, discomfort of passengers and the reduction of the pilot’s ability to control the aircraft. One of the objectives of this paper are to obtain a mathematical model for the passive and active landing gears for full aircraft model. The main porpuse of current paper is to design Linear Quadratic Regulator (LQR) for active landing gear system that chooses damping and stiffness performance of suspension system as control object. Sometimes conventional feedback controller may not perform well because of the variation in process dynamics due to nonlinear actuator in active control system, change in environmental conditions and variation in the character of the disturbances. To overcome the above problem, this research designs a controller for a second order system based on Linear Quadratic Regulator. The performance of active system is compared with the passive landing gear system by numerical simulation. The results of current paper in compared with the previous work mentioned in reference, demonstrate 37.04% improvement in body acceleration, 20% in fuselage displacement and 13.8% in the shock strut travel. The active landing gear system is able to increase the ride comfort and good track holding by reducing the fuselage acceleration and displacement and load induced to airframe caused by runway excitation.

Abstract: Longitudinal stretch forming is a vital fabricating process for the aircraft skin components with large dimensions. Loading trajectory of clamping jaws in stretch forming is significant for the qualified skin components fabrication. However, the traditional design methods based on estimation, experience, and trial-and-error are unable to manufacture skin components precisely and efficiently. A precision design method of loading trajectory for longitudinal stretch forming was developed and discussed in this paper. The sectional profile of the stretched sheet was considered to determine an optimal deformation state in the final stretching process. In view of the optimized sectional profile, the extending method of sectional curve was proposed to define the spatial locations of the clamping jaws. Furthermore, the numerical control parameters of the stretch forming press were transformed from the kinetic locus of the clamping jaws through the mechanism-solving algorithm and the computer-aided software was developed to integrate the methodology. For a fuselage skin component, the finite element (FE) simulation and experiment were conducted to verify the validity of the methodology of loading trajectory design and the algorithm of mechanism solving. The experimental and simulative results revealed that the developed methodology and algorithm are accurate and valid for the fabrication of aircraft skin components. By the examination of the formed skin part, the developed methodology thus provided relevant knowledge for loading locus design in longitudinal stretch forming.

Abstract: The invention relates to a vertical take-off and landing aircraft with an attitude controller. The aircraft consists of a fuselage (6), two main thrust devices (7) which provide main thrust for the aircraft, and the attitude controller, wherein the attitude controller comprises a power module, attitude adjusting units (2), motors, speed regulator units and a flight control system; the aircraft adopts a double-main-wing structure, wherein the attitude adjusting units are separately connected to the left end of the first main wing (8a), the right end of the first main wing (8a), the left end of the second main wing (8b), and the right end of the second main wing (8b); the attitude adjusting units are ducted fans; the two main thrust devices are separately connected to the left side and the right side of the middle segment of the fuselage in a symmetrical manner, and the ducted fans are adopted; the air outlet of each duct is provided with a control surface; the main thrust device and the attitude adjusting units tilt towards the front of the aircraft; therefore, conventional complex mechanical components can be omitted, and the mechanical structure of the aircraft is simplified; the attitude controller and the main thrust devices are separated, so that the type of an engine adopted by the aircraft, and the types of thrust devices adopted by the aircraft, are not limited.

Abstract: The utility model provides a heating equipment and culinary art device, this firing equipment includes: a sound sound signal, wherein, microprocessor is connected with sound sensor and heating element electricity respectively for supporting sound sensor and microprocessor are installed to cooking utensil's fuselage and installation cavity heating element and the automatically controlled board that is located the fuselage on the automatically controlled board, and the sound sensor is used for detecting cooking utensil and sends for the sound sound signal who detects according to the sound sensor adjusts heating element's heating power. Among this technical scheme, arrange the sound sensor in the circuit board on, left out the step of sound sensor with the fuselage assembly. Simultaneously, owing to in need not to be embedded into the fuselage with the sound sensor, needn't set up the structure that is used for imbedding the sound sensor on the fuselage to improved the structural strength of fuselage, reduced the processing degree of difficulty of fuselage, in addition, avoided because extra trompil results in the inside into problem of dust of intaking of fuselage.

Abstract: A flying vehicle with a fuselage having a longitudinal axis, a cockpit extending substantially from the center of the fuselage, a left front wing extending from the fuselage, a right front wing extending from the fuselage, a left rear wing extending from the fuselage, a right rear wing extending from the fuselage. Each wing contains a rotor rotatably mounted and a direct drive brushless motor providing directional control of the vehicle. A centrally located ducted fan encompasses the cockpit and provides VTOL capabilities. The central location of the cockpit and central ducted fan aid in balance and stability. The central ducted fan is itself a brushless motor with the stator windings encapsulated in the ducted fan housing and rotor magnets within the fan. All motors and rotatable mounts are controlled by a fly-by-wire system integrated into a central computer with avionics allowing for autonomous flight.

Abstract: A rotary wing aircraft having dual main rotor assemblies, wherein each main rotor is positioned laterally on linkages and are equidistant in a transverse direction from either side of the fuselage. The rotational axis of each rotor is moveable to alter an angle of the rotational axis to control both horizontal and vertical movement of the aircraft. The angle may be altered by rotating the rotational axes in a vertical plane that is parallel and spaced apart from the vertical plane of the longitudinal axis of the fuselage, or the rotational axes may be angled out of a vertical plane that is parallel and spaced apart from the vertical plane of the longitudinal axis of the fuselage. Each rotational axis may rotate independently.

Abstract: A multicopter aircraft with boom-mounted rotors is disclosed. In various embodiments, a multicopter as disclosed herein includes a fuselage; a port side wing coupled to the fuselage; and a starboard side wing coupled to the fuselage. Each of the wings has mounted thereto one or more booms, each boom having a forward end extending forward of a corresponding wing to which the boom is mounted and an after end extending aft of said corresponding wing to which the boom is mounted. The aircraft further includes a first plurality of lift rotors, each rotor in said first plurality being mounted on a forward end of a corresponding one or said booms; and a second plurality of lift rotors, each rotor in said second plurality being mounted on an after end of a corresponding one or said booms. Each rotor produces vertical thrust independent of the thrust produced by the other rotors.