Abstract: Abstmcf: Flight testing of a fully-instrumented model-scale unmanned helicopter (Yamaha R-SO with loft. diameter rotor) was conducted for the purpose of dynamic model identification. This paper describes the application of CIFER' system identification techniques, which have been developed for full size helicopters, to this aircraft. An accurate, high-bandwidth, linear state-space model was derived for the hover condition. The model structure includes the explicit representation of regressive rotor-flap dynamics, rigid-body fuselage dynamics, and the yaw damper. The R-50 codiguration and identified dynamics are compared with those of a dynamically scaled UH-1H. The identified model shows excellent predictive capability and is well suited for flight control design and simulation applications.

Abstract: A shell component for an aircraft fuselage includes a fuselage skin panel, a plurality of stringers extending in an aircraft lengthwise direction, and a plurality of frames extending crosswise relative to the stringers. The stringers and the frames are respectively welded onto the skin panel with the addition of a weld filler material. Each frame includes a frame root portion and a frame profile portion connected to each other. The frame root portion has cut-out notches receiving the stringers passing therethrough. The frame root portion is welded to the skin panel at the areas between the cut-out notches, and may be welded to the respective stringer in each cut-out notch. The stringers and frames are fabricated from webs and flanges, whereby a premanufactured grid of flanges may be used. The result is a very strong, yet simple shell component structure, that may be manufactured with a simple welding process, at a low cost and with low effort.

Abstract: The objectives were to create a capability to simulate curvilinear crack growth and ductile tearing in aircraft fuselages subjected to widespread fatigue damage and to validate with tests Analysis methodology and software program (FRANC3D/STAGS) developed herein allows engineers to maintain aging aircraft economically, while insuring continuous airworthiness, and to design more damage-tolerant aircraft for the next generation Simulations of crack growth in fuselages were described The crack tip opening angle (CTOA) fracture criterion, obtained from laboratory tests, was used to predict fracture behavior of fuselage panel tests Geometrically nonlinear, elastic-plastic, thin shell finite element crack growth history, and residual strength between measured and predicted results were made to assess the validity of the methodology Incorporation of residual plastic deformations and tear strap failure was essential for accurate residual strength predictions Issues related to predicting crack trajectory in fuselages were also discussed A directional criterion, including T-stress and fracture toughness orthotropy, was developed Curvilinear crack growth was simulated in coupon and fuselage panel tests Both T-stress and fracture toughness orthotropy were essential to predict the observed crack paths Flapping of fuselages was predicted Measured and predicted results agreed reasonably well

Abstract: An improvement to a subsonic passenger aircraft is provided. The passenger aircraft has a fuselage including a main passenger seating deck (14) including forward, middle, and aft regions. The fuselage includes an upper forward deck (16) located above the main passenger seating deck forward region and accessible thereto. Both the main deck and upper forward deck having doors (30), (26) therein for passenger and supplies ingress and egress. The fuselage includes an aft upper deck (18) located above the main passenger seating deck aft region and accessible from the main deck. An upper middle region (20) is located above the main deck middle region. In one embodiment, the upper middle region (20) is of a height elevationally less than either of the forward and aft upper decks. The cross-sectional area of the fuselage at the upper middle region is less than the cross-sectional area of the fuselage at either the forward upper deck region or the aft upper deck region. In another embodiment, the forward upper cabin and the aft upper cabin are separate, longitudinally non-adjacent cabins for accommodating seated passengers.

Abstract: As part of an approach to design fuselage frames for minimum weight, minimum cost, or a combination of the two, the design constraints and the effects of manufacturing process are discussed. Four different fabrication processes are considered: Conventional sheet metal, high speed machined metal, hand laid-up composite, and resin transfer molded composite. For each process, the limitations and applicability are translated to constraints for the geometry of the frame. In addition, the constraints arising from structural requirements are presented and discussed. These constraints are discussed as a necessary foundation for solving fuselage frame cost and weight optimization problems.

Abstract: A 1/5-scale model composite fuselage concept for has been developed to satisfy structural and flight loads requirements and to satisfy design goals for improved crashworthiness. The 1/5-scale model fuselage consists of a relatively rigid upper section which forms the passenger cabin, a stiff structural floor, and an energy absorbing subfloor which is designed to limit impact forces during a crash event. The focus of the present paper is to describe the crashworthy evaluation of the fuselage concept through impact testing and finite element simulation using the nonlinear, explicit transient dynamic code, MSC/DYTRAN. The energy absorption behavior of two different subfloor configurations was determined through quasi-static crushing tests. For the dynamic evaluation, each subfloor configuration was incorporated into a 1/5-scale model fuselage section, which was impacted at 31 ft/s vertical velocity onto a rigid surface. The experimental data demonstrate that the fuselage section with a foam-filled subfloor configuration satisfied the impact design requirement. In addition, the fuselage section maintained excellent energy absorption behavior for a 31 ft/s vertical drop test with a 15 degree-roll impact attitude. Good correlation was obtained between the experimental data and analytical results for both impact conditions.

Abstract: An adaptive ramp mechanism for an aircraft having a fuselage which includes a fuselage skin, an interior cavity, a forward end and an aft end, and at least one door attached to the fuselage and moveable between an open position exposing the interior cavity and a closed position enclosing the interior cavity. The ramp mechanism comprises a ramp structure which is attached to the fuselage between the forward end and the interior cavity thereof. The ramp structure is movable between an undeployed position whereat it is substantially continuous with the fuselage skin such that when the door is in the closed position, a free stream of air flowing along the fuselage is attached to the fuselage skin and the door, and undeployed position whereat at least a portion thereof extends angularly relative to the fuselage skin such that when the door is in the open position, the free stream of air is deflected over the interior cavity and reattaches to the fuselage skin between the interior cavity and the aft end of the fuselage.

Abstract: A system for calculating the bearing of a signal source, with a directional antenna, provides corrections for distortion, such as due to a small fuselage of the monitoring aircraft and the elevation angle of an intruder aircraft with respect to the monitoring aircraft. A correction is applied to the bearing estimate that is based on relevant factors, such as the fuselage size and the elevation angle of the intruder aircraft. The correction can be calculated or applied through the use of a look-up table, which may be either pre-selected or selected after calculation of the elevation angle of the intruder aircraft.

Abstract: A low-sonic-boom design method is developed by combining a three-dimensional Euler computational e uid dynamics code with a least-squares optimization technique. In this design method, the fuselage geometry of an aircraft is modie ed to minimize the pressure discrepancies between a target low-boom pressure signature and a calculated signature. The aircraft cone gurations that generate three types of low-boom pressure signatures, i.e., e attop type, ramp type, and hybrid type, are successfully designed by this method. It is shown that the sonic-boom intensity of the aircraft designed by linear theory is reduced and the e attop-type ground pressure signature is obtained by this method. The results of the study suggest that this method is a useful tool for low-boom design.

Abstract: A control system for aerodynamic vehicles has a fuselage nose segment and at least one strake. The fuselage nose segment is rotationally adjustable with respect to the aircraft fuselage. The at least one strake is mounted on the fuselage nose segment so that the strake can swing out.

Abstract: The invention relates to a device for attaching an aircraft engine onto a pylon fixed to a wing or a fuselage. The device comprises a fitting ( 26 ), fixed to the pylon ( 20 ) by bolts ( 34, 36 ), and at least two connecting arms ( 28, 30 ), linking the fitting to a structural component ( 22 ) of the engine. To allow the installation of engines with larger diameters while preserving the mechanical strength and without increasing aerodynamic disturbance, barrel nuts ( 38, 40 ) are put into position into which the bolts ( 34, 36 ) are screwed, and certain pins connecting the connecting arms ( 28, 30 ) to the fitting ( 26 ) are put into the same bores ( 42, 43 ) of the fitting.

Abstract: A ballistic barrier for protecting an aircraft from damage due to projectile penetration, the aircraft having a fuselage including an outer skin, an inner panel and a structure. The ballistic barrier includes at least one layer of high strength fabric disposed between the outer skin and the inner panel of the aircraft which is substantially fixedly positioned with respect to the fuselage of the aircraft. Preferably, the at least one layer of high strength fabric comprises a plurality of plies. One of the plies can be a felt. Another of the plies can comprise woven fibers. Also preferably, the at least one layer of high strength fabric comprises a polymer material such as one or more of aramid material, polyethylene material, and polybenzoxazole material. The layer of high strength fabric can be attached to at least one of the frames of the structure of the fuselage, to a layer of insulation positioned between the outer skin and inner panel of the fuselage of the aircraft, or be enclosed within an outer covering of a layer of insulation positioned between the outer skin and inner panel of the fuselage of the aircraft.

Abstract: In the case of a helicopter comprising a rotor system ( 3 ) with rotor blades ( 6 and 6 ′) held on a fuselage ( 2 ) and at least one rotor shaft ( 5 and 5 ′) and furthermore comprising a drive system ( 4 ) for causing rotation of the rotor system ( 3 ), the rotor system ( 3 ) together with the drive system ( 4 ) is able to be slid in the longitudinal direction of the fuselage ( 2 ) and to be pivoted around a pivot axis extending along the fuselage ( 2 ).

Abstract: A high speed data network access system for passengers allows access to a global or private data network via data links. In one illustrative embodiment, the system relies on interconnections between aircraft along an established air corridor for interconnecting a passenger with a ground station with data network access. Radiators disposed on the external fuselage of the aircraft allow optical or radio links to be established among aircraft and the ground station. Advantageously, passengers flying on long routes, such as transoceanic flights are able to efficiently access a global data network (e.g., the Internet) with minimal delay.

Abstract: A design-oriented analysis capability for aircraft fuselage structures that utilizes equivalent plate methodology is described. This new capability is implemented as an addition to the existing wing-analysis procedure in the equivalent laminated plate solution (ELAPS) computer code. The wing and fuselage analyses are combined to model entire airframes. The paper focuses on the fuselage model definition, the associated analytical formulation, and the approach used to couple the wing and fuselage analyses. The modeling approach used to minimize the amount of preparation of input data by the user and to facilitate the making of design changes is described. The fuselage analysis is based on ring and shell equations, but the procedure is formulated to be analogous to that used for plates to take advantage of the existing code in ELAPS. Connector springs are used to couple the wing and fuselage models. Typical fuselage analysis results are presented for two analytical models. Results for a ring-stiffened cylinder model are compared with results from conventional finite element analyses to assess the accuracy of this new analysis capability. The connection of plate and ring segments is demonstrated using a second model that is representative of the wing structure for a channel-wing aircraft configuration.

Abstract: A fuselage is substantially simultaneously fabricated, equipped and outfitted in an assembly area including adjacent riveting, equipping and outfitting zones. The riveting zone includes an external riveting apparatus working from the outside of the aircraft fuselage, and an internal riveting apparatus working from the inside of the fuselage to fabricate and join a first fuselage section to an initial structure. The first fuselage section is moved from the riveting zone into an equipping zone, wherein equipment, such as pipes, ducts, hoses, pumps, blowers and structural components and fittings of the aircraft, is installed in the first fuselage section, while simultaneously a second fuselage section is being rivet-joined onto the first fuselage section in the riveting zone. The second fuselage section is moved from the riveting zone into the equipping zone, and the first fuselage section is moved from the equipping zone into the outfitting zone. Equipment is installed in the second fuselage section in the equipping zone, while outfitting components such as electrical cable bundles, insulation blankets, wall paneling, floors, furnishings, and cabin fittings are installed in the first fuselage section in the outfitting zone. Next, the fuselage being formed is shifted so that the second fuselage section moves from the equipping zone into the outfitting zone, while a third fuselage section is joined onto the second fuselage section in the riveting zone.

Abstract: A hydraulic power assembly is comprised of a hydraulic pump being attached to and powered by a ram air turbine. A strut is rotatably attached to an aircraft by having one end mounted on a trunion, and the hydraulic power assembly affixed to the other, distal end. Flexible hoses fluidly communicate the pump with a hydraulic interface located inside of the aircraft. The hydraulic interface fluidly communicates with the aircraft's hydraulic system. The hydraulic power assembly and the strut are normally stored within the fuselage of the aircraft and are deployable into an adjacent airstream in an emergency.

Abstract: The present study addresses the reduction of the sound transmission through an aircraft fuselage test section by means of a combined active noise and vibration control system, so as to reduce the noise level inside the passengers’ cabin. Experiments prove that control loudspeakers and error microphones in the trim cavity yield higher reductions in the radiated sound power than control shakers and error accelerometers on the trim panel. This conclusion remains valid when the skin and the trim panel are mechanically connected by means of four vibration isolators, and when the cavity is filled with thermal insulation blankets. It also appears that, in case of an active cavity noise control system, there exists an almost linear relation between the reduction of the acoustic potential energy in the cavity and the reduction of the radiated sound power. The optimised control system performs much better than a control system with sensors and actuators in arbitrarily chosen positions. * Research Assistant of the Fund for Scientific Research Flanders (Belgium) (F.W.O.)

Abstract: Current research in Active Structural Acoustic Control (ASAC) relies heavily upon accurately capturing the application physics associated with the structure being controlled. The application of ASAC to aircraft interior noise requires a greater understanding of the dynamics of the curved panels which compose the skin of an aircraft fuselage. This paper presents a model of a simply supported curved panel with attached piezoelectric transducers. The model is validated by comparison to previous work. Further, experimental results for a simply supported curved panel test structure are presented in support of the model. The curvature is shown to affect substantially the dynamics of the panel, the integration of transducers, and the bandwidth required for structural acoustic control.

Abstract: : A Short Brothers PLC, Model SD 3-30, airplane was subjected to a vertical impact drop test at the Federal Aviation Administration (FAA) William J. Hughes Technical Center, Atlantic City International Airport, New Jersey. The objective of the test was to determine the impact response of the fuselage, seat tracks, seats, and anthropomorphic test dummies on a high-wing, commuter type airplane. The test was conducted to simulate the vertical velocity component of a severe, but survivable, crash impact. A final impact velocity of 30 feet per second was therefore selected. The airplane was configured in a typical maximum gross weight flight condition, including seats, simulated occupants, fuel, and cargo. The Shorts 3.30 is a twin turboprop, 30-passenger regional transport airplane. The total test weight of the airplane was 21,210 pounds. The internal seating arrangement consisted of pilot and copilot seats, eight rows of standard passenger seats, and two nonstandard seats mounted in the aisle. Twentyone of the 28 seats were occupied by mannequins; the remaining seven seats were occupied by instrumented anthropomorphic test dummies. The Shorts 3-30 fuel system is unique insofar as the two fuel tanks are located on top of the fuselage as opposed to the more conventional location in the wings. During the drop test, a massive amount of simulated fuel spilled into the passenger compartment. The stiff structure of the airplane allowed for only small amounts of airframe crushing. As a result, the fuselage experienced high G(max) levels of approximately 90 g's with an impact pulse duration of 15 ins. The stiff structure also prevented fuselage crushing which allowed the airplane to maintain a protective shell. The seat tracks remained attached to the fuselage. However, 23 of the 26 passenger seats experienced structural failure. The crew seats were undamaged.

Abstract: : This report summarizes the research and full-scale tests undertaken by the Federal Aviation Administration (FAA) to evaluate the fuselage burnthrough resistance of transport category aircraft that are exposed to large postcrash fuel fires. Twenty-eight full-scale tests were conducted in a reusable fuselage test rig to determine the effectiveness of thermal-acoustical insulation improvements in preventing or delaying fuselage burnthrough. The testing showed that the method of attaching the insulation to the fuselage structure had a critical effect on the effectiveness of the insulation material. In addition, the composition of the insulation bagging material, normally a thermoplastic film, was also shown to be an important factor. A number of fiberglass insulation modifications and new insulation materials were shown to be effective in varying degrees. For example, a heat-treated, oxidized polyacrylonitrile fiber (OPF) encased in a polyimide bagging material prevented burnthrough for over 8 minutes. When contrasted with current insulation materials, which were shown to fail in as little as 2 minutes, effective fire barriers such as the OPF insulation offer the potential of saving lives during a postcrash fire accident in which the fuselage remains intact.

Abstract: The aeroelastic response analysis of a coupled rotor/fuselage system is approached by iterative solution of the blade aeroelastic response in the non-inertial reference frame fixed on the hub, and the periodic response of the fuselage in the inertial reference frame. A model of the coupled system hinged with the flap and lag hinges, the pitching bearing which may not coincide with the hinges, and the sweeping-blade configuration is established. The moderate-deflection beam theory and the two-dimensional quasi-steady aerodynamic model are employed to model the aeroelastic blade, all the kinetic and inertial factors are taken into account in a unified manner. A five-nodes, 15-DOFs pre-twisted nonuniform beam element is developed for the discretization of the blade, three rigid-body-motion DOFs are introduced for the motion of the hinges and the bearing. The Hamilton's principle is employed to evaluate the equation of motion of the blade. The derived nonlinear ordinary differential equations with time-dependent periodic coefficients are solved by a modified quasi-linearization method, which is developed for the higher DOF periodic system. The resulting periodic forces and moments exerted on the fuselage by all the blades are evaluated every time, when the converged nonlinear periodic response of the blade is obtained under the consideration of the equilibrium of the blades. The fuselage structure is simplified to be a beam structure, the governing equation is established in the inertial reference frame and a two-nodes beam element is used to discretize the flexible fuselage. The periodic response of the fuselage is solved by a simple shooting method. The iteration of the rotor/fuselage response is continued, until the aeroelastic responses of the blade and the fuselage converge simultaneously. Both the hovering and the forward flight states can be considered. The results of a computed numerical example by the developed program are presented to verify in practice the economy of the modeling as well as the reliability and efficiency of the corresponding solving methods.

Abstract: A hydrodynamic/aerodynamic amphibious aircraft ( 1 ) has a fuselage ( 10 ) with sponsons ( 100 ) extending outwardly and downwardly on either side of the underbelly ( 5 ) to define an inverted channel ( 15 ) having a substantially constant cross section. Each sponson has a forward portion ( 110 ) fixed to the fuselage and a movable aft portion ( 120 ). The aft portion tapers smoothly to a trailing edge ( 122 ). When the aft portion is in the flight position the forward and aft portions form a smooth low-drag symmetrical airfoil aerodynamic shape. When the aft portion is raised a hydrodynamic step ( 112 ) is left on a lower surface of the forward portion, which turns each of the forward portions into a planing hull.

Abstract: Experimental data are presented to show evidence of chaotic response of two adjacent aircraft panels forced by a turbulent boundary layer and pure tone sound. The experiments are a simulation of boundary-layer and fan noise loads on a fuselage sidewall with Reynolds number per meter of 2 :85 £ 10 5 . The response of the panels is purely random and assumed linear when forced by the turbulent boundary-layer e ow and clearly becomes nonlinear with the appearance of the interspersed periodic to chaotic motion when forced by the boundary layer with superimposed pure tone sound. The initial periodic response of two tori of two commensurate frequencies changes with an increase in pure tone sound level. The response of period-doubling bifurcations then makes a transition to chaos, which alternates with quasiperiodic response as the wave loses the spatial homogeneity. The objective is to demonstrate the existence of strong nonlinear effects on the structure response, which is not yet well understood. I. Background M OST studies of nonlinear deterministic and stochastic dynamic problems examine externally excited systems. A typical example of an externally excited system is an aircraft fuselage structure interacting with a turbulent boundary layer and jet engine noise. Periodic, aperiodic, and chaotic responses can occur along the sidewall of the fuselage structure during the acceleration from takeoff, as well as at cruise altitude. One type of load is the socalled buzz-saw noise in high-bypass-ratio turbofan engines. The present experiment is designed to simulate such loads, as well as structural nonlinear responses that result from turbulent boundarylayer e ow and high-intensity sound interaction. Such experiments must be conducted in a wind tunnel with an anechoic test section to prevent standing wave formation between the test panel surface and the opposite sidewall of the tunnel. As a rule, the panel tension and curvaturedependon theloading. Thistension, therefore, constitutes a coupling between the loading and the response. One manifestation of this coupling is the spontaneous surface deformation of the panel, giving rise e rst to the regular and then to the irregular spatial patterns as the load increases. In the previous experiments, panels with periodic nonlinear responses to sound and e ows of constant or accelerated speeds and their active control were considered. 1;2 The present study simulates the abnormal processes of e ow and sound loads, the unsteady loads of the boundary-layer pressure e uctuations coupled with the panel responses,andthesoundradiationbythepanel.Nonlinearbehaviors result fromincreasing levels of pure tone sound as evidenced by the response changes in the panel from periodic motions to broadband chaos. In the past, chaotic signals were not recognized as a physical behavior but were hidden in the broad view given by stochastic processes. At present, the interest is to distinguish between periodic, quasiperiodic,and nonperiodic responses. 3‐6 In the experiments reported herein, the input of the acoustic load superimposed on the turbulent boundary-layer e ow was gradually increased. First, the boundary-layer instability and then responses changing from periodic to quasiperiodic and e nally to chaotic were observed, analo

Abstract: The transport system has a carrier arrangement, esp. a passenger or transport aircraft (2), with a cargo vol. in the form of a module (1) for replaceable mounting in a bay or aperture (89) in a fuselage of a carrier arrangement. The module forms an aerodynamic fuselage contour. An Independent claim is also included for a method of handling a payload with a transport device, esp. a passenger or transport aircraft.