Abstract: The disclosure generally pertains to a vertical take-off and landing (VTOL) aircraft comprising a fuselage and at least one fixed wing. The aircraft may include at least two powered rotors located generally along a longitudinal axis of the fuselage. The rotor units may be coupled to the fuselage via a rotating chassis, which allows the rotors to provide directed thrust by movement of the rotor units about at least one axis. By moving the rotor units, the aircraft can transition from a hover mode to a transition mode and then to a forward flight mode and back.

Abstract: Single-lap shear behaviour of carbon–epoxy composite bolted aircraft fuselage joints at quasi-static and dynamic (5 m/s and 10 m/s) loading speeds is studied experimentally Single and multi-bolt joints with countersunk fasteners were tested The initial joint failure mode was bearing, while final failure was either due to fastener pull-through or fastener fracture at a thread Much less hole bearing damage, and hence energy absorption, occurred when the fastener(s) fractured at a thread, which occurred most frequently in thick joints and in quasi-static tests Fastener failure thus requires special consideration in designing crashworthy fastened composite structures; if it can be delayed, energy absorption is greater A correlation between energy absorption in multi-bolt and single-bolt joint tests indicates potential to downsize future test programmes Tapering a thin fuselage panel layup to a thicker layup at the countersunk hole proved highly effective in achieving satisfactory joint strength and energy absorption

Abstract: An aircraft propelled by a turbojet having contrarotating fans, the turbojet being incorporated at the rear of a fuselage of the aircraft and in line therewith and including two gas generators that feed a working turbine having two contrarotating turbine rotors for driving two fans arranged downstream from the gas generators, and distinct air intakes for feeding each gas generator, the air intakes being connected to the fuselage of the aircraft so as to take in at least a portion of the boundary layer formed around the fuselage of the aircraft.

Abstract: An unmanned aerial vehicle, comprising: a fuselage having a first side board and a second side board spaced apart and connected by at least one transverse board; the first side board, the second side board, and the at least one transverse board being printed circuit boards; at least one of the first side board, the second side board, and the at least one transverse board having formed and mounted thereon conductive traces and at least one component, respectively, for controlling and monitoring the unmanned aerial vehicle; first and second wings mounted to the fuselage; and, a tail mounted to the fuselage.

Abstract: In crash simulations of composite aircraft fuselage sections, frame breaking, skin bending and failure of mechanically fastened joints can typically be identified as major contributors to crash energy absorption. In order to generate a database for model validations, corresponding static and dynamic tests have been performed on coupon and structural element level to characterise the rate-dependent failure behaviour and energy absorption. Skin-bending, frame-bending and joint-failure tests under pull-out, bearing and peeling loads were performed on 977-2/HTS carbon fibre/epoxy specimens. On the one hand, effects of loading rate on the frame-bending behaviour could be observed. On the other hand, fastener failure did not appear to be depending on loading rate for the test speeds up to 10 m/s involved in this study. Adequate modelling methods in Abaqus/Explicit were derived and validated, and finally applied to a global aircraft crash simulation model.

Abstract: Conventional commercial aircraft fuselages use all-aluminium semi-monocoque structures where the skin carries the external loads, the internal fuselage pressurisation and is strengthen using frames and stringers. Environmental and economic issues force aircraft designers to minimise weight and costs to keep air transport competitive and safe. But as metal designs have reached a high degree of perfection, extraordinary weight and cost savings are unlikely in the future. Carbon composite materials combined with lattice structures and the use of topology optimisation have the potential to offer such weight reductions. The EU FP7 project Advanced Lattice Structures for Composite Airframes (ALaSCA) was started to investigate this. This article present some of this research which has now led to the development of a new airframe concept which consists of: a load carrying inner skin; transverse frames; CFRP-metal hybrid stiffeners helically arranged in a grid configuration; insulating foam and an additional aerodynamic outer skin.

Abstract: The invention relates to a convertible aircraft comprising a fuselage (F), a pair of wings (A1, A2) arranged one on each side of the fuselage (F), at least one ducted rotor (1) installed in a horizontal position at one of the ends of the fuselage (F) and a first and a second nacelle (N1, N2) arranged respectively at the tip of each wing (A1, A2) and each comprising a ducted rotor (R1, R2) and being pivotably mounted relative to the fuselage (F). The nacelles comprise at least a first and a second movable flap (V1, V2), which flaps are arranged respectively at the outlet of the ducted rotor (R1) of the first nacelle (N1) and at the outlet of the ducted rotor (R2) of the second nacelle (N2). The aircraft according to the invention thus represents an advantageous solution to any applications involving helicopters and airplanes, particularly emergency preparedness missions, rescue missions, and public or private transport.

Abstract: A rotor system for tilt rotor aircraft comprises an engine disposed at a first fixed position on a wing member, and a prop-rotor pylon mechanically coupled to the engine along a drive path extending through the wing member. The engine is disposed adjacent a fuselage of the tilt rotor aircraft, and the prop-rotor pylon is configured to selectively rotate between a vertical position and a horizontal position. The prop-rotor pylon is coupled to a plurality of rotor blades.

Abstract: Methods and apparatus for unmanned long endurance flights are provided herein. In some embodiments, a lightweight solar wing for unmanned aircraft may include at least one airfoil profile, a top surface, a bottom surface, a leading edge, a trailing edge, wing tips, and at least one photovoltaic cell, wherein the surfaces and edges follow an arched bow shape across a span of the wing. In some embodiments, an unmanned solar-powered aircraft may include at least one lightweight solar wing as described above, at least one fuselage, and at least one propeller, wherein the fuselage is placed below the solar wing and contains an electric motor, battery, and electronics.

Abstract: Commercial aircraft are designed to fly but also need to operate safely and efficiently as vehicles on the ground. During taxiing, take-off, and landing the landing gear must operate reliably over a wide range of forward velocities and vertical loads. Specifically, it must maintain straight rolling under a wide variety of operating conditions. It is well known, however, that under certain conditions the wheels of the landing gear may display unwanted oscillations, referred to as shimmy oscillations, during ground maneuvers. Such oscillations are highly unwanted from a safety and a ride-comfort perspective. In this paper we conduct a study into the occurrence of shimmy oscillations in a main landing gear (MLG) of a typical midsize passenger aircraft. Such a gear is characterized by a main strut attached to the wing spar with a side-stay that connects the main strut to an attachment point closer to the fuselage center line. Nonlinear equations of motion are developed for the specific case of a two-wheeled M...

Abstract: A lower hull 403 and an upper hull 402 may form a multi rotor aircraft fuselage system comprising a comprising a thermoformed thin shell monocoque fuselage 500, center console assembly 200, the center console assembly comprising a center console cover 210, an auto pilot housing cover 220 and an auto pilot housing 225. A center fuselage area 300 contains the center console assembly 200. The center fuselage area is part of a plurality of arm assemblies 400. Each arm assembly may comprise a distal end 440 and a proximal end 410, an access panel cover 420 which may expose a access panel opening 419. An access panel opening 419 may lead to an arm void area 430, with the arm void area 430 further defined by the interior sections of the thin shell fuselage 500.

Abstract: An aircraft (10) comprising a fuselage (12) having a longitudinal axis (16), at least one helicopter main rotor (40) operably mounted to the fuselage (12), the at least one helicopter main rotor (40) comprising rotor blades (41) rotatable about a rotation axis (44), wherein the rotor blades (41) can be stopped in flight and adapted to provide symmetrical wing surfaces relative to the longitudinal axis (16); and the aircraft (10) having at least one control surface (62, 82) operable to provide a relative airflow (306) in flight substantially aligned with the rotation axis (44) of the at least one helicopter main rotor (40).

Abstract: This paper presents hardware and data on the use of AWJ for trimming aircraft carbon fiber reinforced plastic (CFRP) parts such as those used on the Airbus 350 and the Boeing 787. Generally, CFRP parts on an aircraft vary in size from large parts such as wings and fuselage sections to small size parts such as clips, brackets, and door stringers. The machinery that is most suitable for these parts is presented. Gantry systems with AWJ and routing end effectors have been the most commonly used machines for large parts while relatively small robotic arms are emerging for trimming small parts. In any of these systems, special sidefire cutting heads have been developed to access tight spaces such as trimming and beveling stringer flanges. Small catcher cups, mounted on the cutting end effector, have also been developed to catch the waste jet. Data are presented on taper, trailback, and surface finish to identify parameters meeting the required accuracy and surface finish.

Abstract: This paper introduces a fully parameterized finite element based tool chain for the structural sizing of transport aircraft. The chain consists of two model generators, a coupling module, a loads module for the computation of aerodynamic, fuel, landing gear and engine loads as well as a structural analysis and sizing algorithm. The finite element models of the wing and the empennage are created by the ELWIS multi-model generator, while the corresponding fuselage model is created using the TRAFUMO model generator. The structural coupling comprises the detailed modeling of all key structural elements of the center fuselage area including a keelbeam, bulkheads, sideboxes and lateral panels. The empennage coupling structure includes a reinforcement framework, reinforced frames and a mounting structure for the horizontal tail plane trim device. To establish suitability in preliminary aircraft design, a knowledge-based approach is chosen that enables a fully automatic model generation and coupling on a minimum set of required input parameters. As a central data format for input and output the DLR aircraft parameterization format CPACS is used. Therefore, the chain can be easily embedded in a wider MDA/MDO approach for overall preliminary aircraft design. Finally, first static sizing results are discussed and different validation methods for the static sizing algorithm are presented, including a comparison with a validated analytical method.

Abstract: NASA sets aggressive, strategic, civil aircraft performance and environmental goals and develops ambitious technology roadmaps to guide its research efforts. NASA has adopted a phased approach for community noise reduction of civil aircraft. While the goal of the near-term first phase focuses primarily on source noise reduction, the goal of the second phase relies heavily on presumed architecture changes of future aircraft. The departure from conventional airplane configurations to designs that incorporate some type of propulsion noise shielding is anticipated to provide an additional 10 cumulative EPNdB of noise reduction. One candidate propulsion system for these advanced aircraft is the open rotor engine. In some planned applications, twin open rotor propulsors are located on the aft fuselage, with the vehicle s empennage shielding some of their acoustic signature from observers on the ground. This study focuses on predicting the noise certification benefits of a notional open rotor aircraft with tail structures shielding a portion of the rotor noise. The measured noise of an open rotor test article--collected with and without an acoustic barrier wall--is the basis of the prediction. The results are used to help validate NASA s reliance on acoustic shielding to achieve the second phase of its community noise reduction goals. The noise measurements are also compared to a popular empirical diffraction correlation often used at NASA to predict acoustic shielding.

Abstract: A composite part, such as a stiffener is formed in place. A composite charge (30) is placed on a tool spanning a mold cavity (54), with the centerline of the charge (32) offset from the centerline of the mold cavity (28). Opposite sides of the charge (70, 74) are held against the tool as the charge (30) is formed into the mold cavity (54). One side of the charge (70) is held against movement on the tool while the other side of the charge (74) is allowed to slip over the tool toward the mold cavity (54).

Abstract: The landing of an aircraft is one of the most critical operations because it directly affects the passenger safety and comfort. During landing, the aircraft fuselage undergoes excessive vibrations that cause the safety and the comfort problem and hence need to be suppressed quickly. A semi-active control system of a landing gear suspension by using Magnetorheological damper can solve the problem of excessive vibrations effectively. In this paper, a switching technique is developed in the simulation of the landing procedure which enables the system to switch from the single degree of freedom to three degrees of freedom system in order to simulate the sequential touching of the two wheels of the main landing gears and the nose landing gear wheels with the ground. A semi-active Magnetorheological damper is developed using two different controllers namely linear quadratic regulator and the H∞. Spencer model is used to predict the dynamic behavior of the Magnetorheological damper. The results of the designed controllers are compared to study the performance of the controllers in reducing the overshoot of the bounce response as well as the bounce rate response. The simulation results validated the improved performance of the robust controller compared to the optimal control strategy when the aircraft is subjected to the disturbances during landing.

Abstract: The present invention relates to an aircraft comprising a fuselage (100) comprising a fuselage axis (101), a first wing arrangement (110) and a second wing arrangement (120). The first wing arrangement (110) is mounted to the fuselage (100) such that the first wing arrangement (110) is tiltable around a first longitudinal wing axis (111) of the first wing arrangement (110) and such that the first wing arrangement (110) is rotatable around the fuselage axis (101). The second wing arrangement (120) comprises at least one propulsion unit (122), wherein the second wing arrangement (120) is mounted to the fuselage (100) such that the second wing arrangement (120) is tiltable around a second longitudinal wing axis (121) of the second wing arrangement (120) and such that the second wing arrangement (120) is rotatable around the fuselage axis (101). The first wing arrangement (110) and the second wing arrangement (120) are adapted in such a way that, in a fixed-wing flight mode, the first wing arrangement (110) and the second wing arrangement (120) do not rotate around the fuselage axis (101). The first wing arrangement (110) and the second wing arrangement (120) are further adapted in such a way that, in a hover flight mode, the first wing arrangement (110) and the second wing arrangement (120) are tilted around the respective first longitudinal wing axis (111) and the respective second longitudinal wing axis (121) with respect to its orientations in the fixed-wing flight mode and that the first wing arrangement (110) and the second wing arrangement (120) rotate around the fuselage axis (101).

Abstract: ‡The aft fuselage of the C-130 is upswept to accommodate the aft cargo ramp. The upswept aft fuselage accounts for as much as 11% of the total vehicle drag at cruise. While concepts to reduce this contribution to the C-130 drag were developed in the past, they interfered with air drop operations and were never integrated into the fleet. Using flow control technology, a relatively simple, structurally benign, microvane concept with minimal air drop impact has been developed. In a collaborative Lockheed Martin, Air Force Research Laboratory program, the microvane concept was matured. Using CFD, the microvane configuration was optimized. Integration and air drop constraints were evaluated and an air drop compliant, retrofittable microvane configuration was identified. Flight testing verified the accuracy of the CFD based drag reduction predictions and confirmed the viability of the design. The flight test configuration saves between 14 and 30 gallons of fuel per hour, resulting in a potential fleet wide savings to the USAF of 2.4 million gallons of fuel per year. With no impact on operational capability and a proven flight test technology readiness level (TRL), this low-risk solution offers benefits of low-cost and rapid deployment to the fleet. The Lockheed Martin (LM) C-130 Hercules tactical military transport incorporates a dual use aft cargo ramp which acts as both a loading ramp and a fairing which closes out the aft fuselage shape. This integrated cargo ramp design is responsible for as much as 11% of the total aircraft drag due to the resulting large aft fuselage upsweep angle. Numerous studies have been performed over the last 40 years to reduce this large drag contribution through the incorporation of relatively large aft fuselage mounted strakes. These strake concepts were developed in the 1970’s – 1980’s through wind tunnel and extensive flight testing. While strakes are effective at reducing aircraft drag, they create integration problems. Results from USAF operational test and evaluations of C-130 aft fuselage strakes in 1981 indicated that “C-130 airdrop capabilities are severely limited with the strakes installed” and that “large container-like loads may require strake removal prior to on or off loading.” 1 They were not incorporated primarily for these reasons. However, current research has indicated that microvanes, devices approximately 10 inches long and 0.5 to 1.2 inches tall arrayed along the breakline of the aft fuselage as shown in Figure 1, can significantly reduce the aft fuselage drag penalty while maintaining compatibility with air drop and loading operations. These devices have been developed using advanced computational fluid dynamics (CFD) methods and their drag reduction performance has been successfully verified with recent flight testing. These tests verified a

Abstract: B ECAUSE the rotor of a helicopter operates in a periodic, asymmetric, unsteady aerodynamic environment, the helicopter fuselage produces a high level of vibration under the strong excitation of the rotor. The effective method for active vibration control of a helicopter fuselage by using the inertial actuators has been used in helicopters, but the inertia actuators have a considerable weight penalty for a better control effect. Meanwhile, the inertia actuators have a limited range of working frequency and a lag response to control signal. The piezoelectric stack actuator has a lot of advantages, such as light weight, large control force, wide range ofworking frequency, and fast response to control signal, and has been used as an actuation element for active control of structural vibration [1,2]. Hence, using the piezoelectric stack actuators is a newway to actively control the vibration of a helicopter fuselage. In an active vibration control system, the locations of actuators have a great influence on the effect of vibration suppression and the power requirement. Many approaches such as the controllability index [3], energy dissipation index [4],H2 norm index [5], and recent index composed of a multi-objective [6] have been developed to find the optimal actuator locations and control parameters. In the investigations of the helicopter fuselage, Hanagud and Babu [7] placed the piezoelectric actuator near the selected control location and investigated the vibration reduction of the helicopter fuselage by using H∞ control. Singhvi and Vennkatesan [8] addressed the piezoelectric stack actuator parallel with the supporting structure between the gearbox and fuselage for a simplified helicopter model. Heverly et al. [9] investigated the optimal placement of piezoelectric stack actuators in the fuselage by using the simulated annealing algorithm. The investigations indicated that the configuration of optimal distributed actuators was capable of greater vibration suppression with less control effort. However, in the existing investigations, the piezoelectric stack actuator was idealized as a force generator. In this case, the characteristic effect of the piezoelectric stack actuator was not included in the optimization process. The optimal locations of actuators may have a lot of selections at many possible positions in an actual structure. The optimal selection of actuator locations cannot uniquely be determined by using the conventional optimization techniques based on the gradient-descent methods. The genetic algorithm as a stochastic search technique has been effectively used to determine the optimal locations. Rao et al. [10] presented a modified binary-coded genetic algorithm to solve the optimal placement of discrete actuator locations in the framework of a zero–one optimization. Liu et al. [5] directly applied the binarycoded genetic algorithm to find optimal locations of actuators and sensors on plate structures. The real-coded genetic algorithm [4] was used to address the optimal locations of the piezoelectric actuators at a continuous spatial coordinate on a beam. Roy and Chakraborty [6] used the integer-coded genetic algorithm to optimize the placement of actuators and simultaneously real-coded genetic algorithm to determine the weighted matrices in the linear quadratic control. In this paper, the active control of a helicopter structural response by using piezoelectric stack actuators has been investigated. In the formulated dynamic model, the piezoelectric stack actuators and fuselage coupled composite structure was decomposed by using the substructure synthesis technique based on frequency response functions. The weighted quadratic of controlled accelerations in the frequency domain was chosen as the optimization index. An improved real-coded genetic algorithm was used to solve the optimization problem with discrete location variables and continuous weighted variables and to minimize the acceleration responses of the targeted locations. The vibration suppression of a simplified elastic helicopter fuselage model was analyzed. The numerical results show that the method proposed in this paper can effectively solve the optimal parameters and improve the control performance.

Abstract: Formation flight is currently being investigated as a means to reduce drag and improve fuel efficiency in commercial aviation. In an attempt to approximate passenger impact in formation flight, a simple aerodynamic model for two aircraft in formation was developed. The wing trailing vortices were assumed to shift in an ideal fashion within atmospheric turbulence resulting in aerodynamic disturbance loads acting on the trailing aircraft. As the sensitivity of the human body to vibrations is frequency dependent, a spectral representation of atmospheric turbulence was incorporated. Monte Carlo simulations were done for various levels of turbulence intensity. The predicted acceleration responses of the trailing aircraft were used to determine the passenger comfort levels by applying the criteria of standard 2631-1 from the International Organization for Standardization and comparing with comfort levels experienced in an aircraft flying in isolation in the same turbulent atmosphere. A significant increase in d...

Abstract: A flying electric generator for obtaining power from wind currents which includes a fuselage having fore and aft portions and an intermediate portion, a rotor assembly including at least two forward rotors mounted on a pair of forward extending support arms extending from the fuselage and at least two rear rotors mounted to a pair of rearward extending support arms extending from the fuselage and at least one first forward wing mounted to a forward portion of the fuselage and extending outwardly on opposite sides of the fuselage and at least one second rear wing mounted to a rear portion of the fuselage and extending outwardly on opposite side of the fuselage.

Abstract: Electroimpact has recently produced a high-speed fuselage panel fastening machine which utilizes an all-electric, CNCcontrolled squeeze process for rivet upset and bolt insertion. The machine is designed to fasten skin panels to stringers, shear ties, and other internal fuselage components. A high riveting rate of 15 rivets per minute was achieved on the firstgeneration E7000 machine. This rate includes drilling, insertion, and upset of headed fuselage rivets. The rivets are inserted by a roller screw-driven upper actuator, with rivet upset performed by a lower actuator driven by a high-loadcapacity ball screw. The rivet upset process can be controlled using either positionor load-based feedback. The E7000 machine incorporates a number of systems to increase panel processing speed, improve final product quality, and minimize operator intervention. The upper riveting head includes automatic tool changers for drills and upper anvils, both designed with very fast tool changes as a primary goal. The machine also includes fastener verification laser curtains to ensure the fastener being inserted is the correct type and length, and oriented properly. The system has automatic calibration functionality to calibrate normality and stringer tracing sensors, as well as upper and lower die lengths. Additionally, an automatic machine vision system provides high-speed, high-accuracy part resynchronization across a wide range of surfaces for improved local accuracy. The E7000 automated fuselage riveter is a fast, robust and flexible system that can be used to fasten a wide range of fuselage panels, while improving reliability and final product quality.

Abstract: A method of automatically triggering an emergency buoyancy system ( 10 ) for a hybrid helicopter ( 20 ) having a fuselage ( 21 ), two half-wings ( 23, 23′ ), and two propulsive propellers ( 24, 24 ′). During the method, said emergency buoyancy system ( 10 ) is primed, and then if a risk of said hybrid helicopter ( 20 ) ditching is detected, two retractable wing undercarriages ( 28, 28′ ) are deployed, each wing undercarriage ( 28, 28′ ) being fastened under a respective half-wing ( 23, 23′ ) and being provided with at least one immersion sensor ( 16 ). Finally, if the beginning of said hybrid helicopter ( 20 ) ditching is detected, at least one main inflatable bag ( 11, 11′ ) 7 B suitable for being arranged under such fuselage ( 21 ) and at least one secondary inflatable bag ( 12, 12′ ) suitable for being arranged under each half-wing ( 23, 23′ ) are inflated so as to ensure that said hybrid helicopter ( 20 ) floats in stable manner.