Abstract: In this paper boundary element methods (BEM) are mated with near-field acoustical holography (NAH) in order to determine the normal velocity over a large area of a fuselage of a turboprop airplane from a measurement of the pressure (hologram) on a concentric surface in the interior of the aircraft. This work represents the first time NAH has been applied in situ, in-flight. The normal fuselage velocity was successfully reconstructed at the blade passage frequency (BPF) of the propeller and its first two harmonics. This reconstructed velocity reveals structure-borne and airborne sound-transmission paths from the engine to the interior space.

Abstract: The stability of an aircraft wake model composed of an external vortex pair (modeling the wing tip vortices) and an internal vortex pair rotating in the opposite direction (modeling the vortices generated by the fuselage and the horizontal tail) in a stationary configuration is investigated with the vortex filament stability method used by Crow [AIAA J. 8, 2172 (1970)] and Crouch [J. Fluid Mech. 350, 311 (1997)]. It is shown that this configuration is unstable with respect to two-dimensional and three-dimensional disturbances. For long wavelength three-dimensional symmetric perturbations, the rapid growth observed in the numerical simulations of Rennich and Lele [J. Air. 36, 398 (1999)] is found. Moreover, the analysis allows one to show that without an excitation of the long-wave mode, the system will naturally develop short wavelength instabilities localized within the inner vortices which do not affect the outer vortices. Inspection of the initial value problem shows that the long-wave modes can be efficiently initiated by the introduction of perturbations on the internal vortices.

Abstract: A wind tunnel test of a generic helicopter fuselage model with an independently mounted rotor has been conducted to obtain steady and periodic pressure data on the helicopter body. The model was tested at four advance ratios and three thrust coefficients. The periodic unsteady pressure coefficients are marked by four peaks associated with the passage of the four rotor blades. Blade passage effects are largest on the nose and tail boom of the model. The magnitude of the pulse increases with rotor thrust coefficient. Tabular listings of the unsteady pressure data are included to permit independent analysis. A CDrom containing the steady and unsteady pressure data presented in the report is available from the authors.

Abstract: Ultrasonic guided waves are considered for the testing of various components used in the aircraft industry. Natural wave guides are present on many aircraft components. Subsequently, guided waves can be used to test hard to access areas quite easily. Guided waves are also ideal for applications involving large areas of testing, as well as cases where direct access to a specific component is not possible. A test protocol for crack and anomaly testing in specific parts is developed. Phase velocity and frequency tuning are used to excite specific points on the dispersion curves. By exciting these so called transient resonant modes, it becomes possible to detect anomalies in different mechanical components and structures. Cracks in helicopter blade models were detected with guided waves. Corrosion and cracks in multilayer media and cracks in landing gear models were also studied. Other investigated problems in this work involve the detection of fuselage wall thinning, the integrity of tear strap assemblies, the debonding of skin to core in honeycomb structures, and the debonding of joints, specifically lap splice joints.

Abstract: An aircraft including an airframe having a fuselage which extends longitudinally, and having fixed wings including a main wing, a horizontal tail wing and a vertical tail wing. A propeller-rotor torque transmission has a bevel gear which transmits the rotation of an input shaft simultaneously to a propeller shaft and to a rotor shaft. An engine gearbox supplies the above-mentioned input shaft with rotationalal motive power. The aircraft further includes a propeller collective pitch controller, a rotor collective pitch controller, an engine power controller which controls the output of the above-mentioned engine gearbox for the purpose of changing the rotational speed of the input shaft, and a flight control system having a directional (yaw) control system which controls the flight direction of the aircraft by controlling the positions of the above-mentioned control surfaces.

Abstract: A VTOL-capable tilt-rotor airplane having a single tiltable large-diameter prop-rotor attached to an elongated power pod containing the collective and cyclical pitch mechanism, transmission, and engine. The tiltable power pod is disposed at the airplane's longitudinal axis and is partially buried within the depth of the fuselage and protrudes 3-4 ft (0.915-1.22 m) above the top of the fuselage in the VTOL mode. In the horizontal cruising mode, the power pod perched on top of the fuselage front section with the rotor blades protruding in front of the airplane's nose. A connecting bar is used to connect the power pod to the fuselage, allowing the power pod to travel significantly rearward and forward as required for proper balance as the power pod pivots 90 degrees during transition from VTOL mode to the cruising mode, without the power pod being excessively long and unwieldy. A small anti-torque rotor toward the tail of the aircraft, driven by a hydraulic motor provides the necessary force to overcome the main rotor's torque. A pair of high-aspect-ratio wings supported by wing struts on both sides of the fuselage provide highly efficient lift during cruising flight with very little induced drag. Conventional horizontal and vertical tail planes are used for directional stability in the cruising mode.

Abstract: An integrated air data probe capable of measuring pressure, temperature, relative humidity, turbulence, heading, air speed, the presence of ice, and other relevant measurements for use in building a real time map of the aircraft flying conditions. The probe is attached to the external surface of an airplane such as the wing, fuselage or empannage and is adapted to collect and relay these measurements to a receiving station. By utilizing these probes on numerous airplanes, real time data from numerous data points can be mapped to provide pilots and airlines complete information regarding flying conditions. Additionally, the probe includes a recessed surface for detecting the presence, thickness, and continued buildup of ice by mounting the recessed surface directly into the airflow. One or more light beams are passed over and parallel to the recessed surface, and if ice is present the beams are interrupted.

Abstract: Aircraft landing gears are currently optimised to have optimal performance in the rare case of a hard landing. The resulting suspension layout may lead to unsatisfactory oscillations when the aircraft is taxiing on a rough runway; runway unevenness can excite elastic structural modes, leading to passenger and crew discomfort. Although there are existing modifications of aircraft shock absorbers to reduce the problem, the basic design conflict between the requirements for landing and for rolling cannot be fully overcome by a passive suspension layout. Semi-active suspension techniques promise a solution to this problem. A semi-active suspension, i.e. a damper with a variable, controlled orifice cross-section, is capable of reducing fuselage vibrations effectively while being relatively light-weight and low system complexity. In this thesis, three control laws, a skyhook-type controller, a fuzzy-logic controller, and a state feedback controller are designed for the application to a semi-active suspension for an aircraft nose landing gear. Regarding the aircraft flexibility, the landing gear can no longer be designed independently from the aircraft. The layout of the controllers is therefore undertaken using an integrated design approach. Airframe and landing gear properties are determined taking into consideration models from different engineering disciplines involved in the aircraft development process, making the oleo design part of the concurrent engineering loop. The aircraft model is set up in a multibody simulation environment. The control laws are developed in a control design tool; special consideration is given to the requirements of semi-active actuators. The controllers are exported into the simulation environment and their parameters are optimised by means of multi-objective optimisation. In a further step, the performance of the three control strategies are compared with each other and additionally with passive and fully active approaches. The dependence of the control performance on operational parameters (aircraft weight and speed, runway roughness) is assessed, and limitations due to realistic actuators restrictions are discussed. Finally, the benefits and disadvantages of semi-active nose landing gear control are summarized and open problems are addressed.

Abstract: When stiffening members, such as stringers and frame members, are welded onto a skin sheet to form an aircraft fuselage shell, a crack originating in the skin sheet tends to propagate through the weld joints into the stiffening members. In order to prevent crack propagation into a stiffening member, the stiffening member is reinforced with a web doubler plate or a tension band made of high strength steel or titanium alloys or fiber-reinforced composites. The doubler plate is riveted or adhesively bonded onto a stiffening member web, or the tension band is crimped into the stiffening member web. The resulting fuselage shell structure has crack stopping properties and thus an increased residual strength, so it can be used with welded joints at all areas of the fuselage shell, including the top and sides as well as the bottom of the fuselage.

Abstract: A tilt-rotor aircraft capable of vertical/short take-off and landing. The aircraft utilizes a co-axial counter rotating rotor system to provide vertical and horizontal thrust. The fuselage is positioned over the forward half of the rotor in the vertical take-off configuration and the rotors rotate ninety degrees to provide horizontal thrust in the forward flight mode. The aircraft uses conventional flight control surfaces during forward flight and uses cyclic rotor pitch control or a series of control surfaces located in the rotor wash to control the aircraft during vertical flight. One or more engines powering the rotor system provide thrust for the aircraft. The engine(s) can be located in the fuselage or in the wings. The aircraft uses lifting airfoils during forward flight which can include a wing/tail or a canard/wing configuration and uses a conventional tail for directional stability and control. The rotor transmission allows one rotor to be braked thus increasing the rotation speed of the other rotor during forward flight.

Abstract: A wing (11, 12) in combination with a fuselage (10), the wing having a relatively unswept and sharp leading edge, where the thickness to chord ratio is about 2 % or less beyond a calculated point near the fuselage and the thickness to chord ratio is substantially greater than 2 % near the fuselage, and the fuselage having an indentation (6c') proximate the wing. The selection of the wing and fuselage indentation is characterized by the maximization of various figures of merit such as reduced drag, cost, or weight.

Abstract: A subfloor structure of an aircraft airframe, particularly of a helicopter, includes longitudinal beams and crossbeams that intersect each other and are interconnected to form a grid that is fixedly attached to the floor and the bottom skin of the aircraft fuselage. Structural elements such as pyramid frustums and reinforcements are arranged on the beams. The longitudinal beam and the crossbeam have a trapezoidal cross-section that is preferably open on the wider base side, closed by a spine web along the narrow side, and bounded laterally by inclined leg webs that extend downwardly from the spine web at an angle outwardly relative to each other. The subfloor structure grid effectively absorbs the energy of a crash impact having both axial or vertical as well as non-axial or lateral impact force components.

Abstract: The invention relates to aircrafts, in particular to aeroplanes, which are embodied in a passenger, cargo or cargo-passenger variants thereby making it possible to improve performance characteristics by improving the aerodynamic properties thereof. The inventive aeroplane having a carrying fuselage comprises a wing, an engine, and a fuselage whose width considerably exceeds the height thereof. A lower surface of the basic and the tail parts of the fuselage have a caved-in form in a cross-section thereof. A lower surface of the flattened front part of the fuselage is arranged at an attack angle with respect to the longitudinal horizontal plane of the aeroplane. In addition, the front part of the aeroplane has a curvilinear form in the plan view, which successively passes from the narrow front part to the basic large part of the fuselage. A tail group is embodied in the form of a double tail fins which are mutually spaced at a distance of the fuselage width with respect to the longitudinal vertical plane of the aeroplane and are mounted at an angle with respect to said vertical plane of the aeroplane.

Abstract: A high-speed vertical take-off and land (HSVTOL) aircraft includes a disk-shaped fuselage wit a rotatable fan having a nozzle ring driven by hot jet gasses and fan air from jet engines. Feed ducts and an annular plenum, have a composite duct-in-duct configuration to separate the hot gases from the relatively cool fan air. High efficiency air bearings serve to support the rotatable fan assembly on the fuselage in the vertical direction, and rollers around the perimeter provide horizontal support and stability. A bearing and seal interface hub and other components, are isolated from the hot gases. The combined exhaust is ejected downwardly at approximately 15° for maximum fan spin and provide direct reaction lift. The fan blades are angled at approximately 57° with moveable airfoil portions and stationary trailing flap portions to provide maximum lift and down wash. Various diverter valves and bypass doors are provided for hot gas/air control. Decoupler valves can interrupt all flow for engine out operation. Control of the pitch, roll and yaw is provided by aero flaps providing control surfaces in the path of the down wash. When twisted about a substantially horizontal control axis, the flaps generate lift, as well as being modulated for attitude control. Aerodynamic disk-shape of the fuselage provides high lift, low drag necessary to support the efficient high speed performance. A control circuit includes an on board CPU controller that operates and coordinates all functions through the various actuators, valves, transducers and other control components.

Abstract: The underlying objective of this paper is to describe an approach adopted for investigating the crashworthiness of an airliner fuselage. It is based on detailed Finite Element (FE) modelling and supported by a series of static and dynamic tests on a subfloor section. The detailed FE analysis was undertaken to predict the dynamic collapse mechanism of a fuselage section. A parametric analytical study was performed for one of the dynamic tests modelled here to verify the results by taking into consideration the influence of modelling techniques. The structural response of an aircraft during an impact with the ground is highly dependent upon the terrain encountered. Significant structural response difference are likely to be produced when the impact surface is of a soft soil type as against a rigid surface such as concrete. The methodology to develop ground models was based on extending the problem from a simple to a complex structure/ground interaction.

Abstract: During the period September 1997 to February 1998, the Tupolev 144 Supersonic Flying Laboratory was used to obtain data for the purpose of enlarging the data base used by models for the prediction of cabin noise in supersonic passenger airplanes. Measured were: turbulent boundary layer pressure fluctuations on the fuselage in seven instrumented window blanks distributed over the length of the fuselage; structural response with accelerometers on skin panels close to those window blanks; interior noise with microphones at the same fuselage bay stations as those window blanks. Flight test points were chosen to cover much of the TU-144''s flight envelope, as well as to obtain as large a unit Reynolds number range as possible at various Mach numbers: takeoff, landing, six subsonic cruise conditions, and eleven supersonic conditions up to Mach 2. Engine runups and reverberation times were measured with a stationary aircraft. The data in the form of time histories of the acoustic signals, together with auxiliary data and basic MATLAB processing modules, are available on CD-R disks.

Abstract: A convertiplane having a fuselage with a first axis; and two rotors having respective shafts and fitted to a wing having a fixed portion connected to the fuselage, and a movable portion which supports the rotors, is connected to the fixed portion to rotate about a second axis crosswise to the first axis and to the shafts of the rotors, and is formed in one piece defined by two half-wings located on opposite sides of the fixed portion and each supporting a respective rotor, and by an elongated member extending along the whole wing and connecting the two half-wings; the rotors being powered by a motor via a transmission comprising a transmission shaft connecting the shafts of the rotors and coaxial with the second axis.

Abstract: For ensuring occupant's survivability in case of aircraft crash accident, it is necessary to alleviate the impact load to an allowable level of the human body. This research aims to develop the analytical models for simulation of impact characteristics of aircraft fuselage structure. The numerical analysis on the sub-floor structures was executed by using nonlinear dynamic analysis code LSDYNA3D, and the results of analytical models are evaluated using the results of crash tests of structural models.

Abstract: An aircraft has a fuselage that defines an interior cabin and an exterior; a first housing located along one side of the fuselage exterior; a second housing located along the second side of the fuselage exterior; and an active vibration control system for limiting fuselage vibration, the vibration control system comprising a plurality of sensors located in the cabin; a controller in signal receiving relation with the sensor means; first actuator means located in the first housing and second actuator means located in the second housing and wherein the actuator means are in signal receiving relation with the controller.

Abstract: A computationally efficient scheme for maneuver load analysis, based on nonlinear aerodynamics, is presented. The kernel of the scheme is a computational fluid dynamics (CFD) code for solving the Euler/Navier-Stokes equations for a fixed-shape configuration. A modal structural model is used for elastic-shape updates, and a trim corrections algorithm is used for varying the incidences and control surface deflections until the user-defined maneuver is attained. Computational efficiency is obtained by performing a few elastic-shape changes and maneuver trim updates, all within the fluid dynamics analysis, during the steady-state flowfield convergence. The modal approach, where the structure is represented by a set of its low-frequency vibration modes, greatly simplifies the CFD-structure interface, minimizes the amount of structural data required, and allows simple shape updates of the aerodynamic grid. It is shown that the total computation time required for flowfield convergence for a maneuvering elastic aircraft is typically almost identical to that of the rigid aircraft with given trim variables. The method is demonstrated with a realistic wing-fuselage-elevator transport aircraft model performing symmetric maneuvers at Mach 0.85

Abstract: The ability to extend the valid frequency range for finite element based structural dynamic predictions using detailed models of the structural components and attachment interfaces is examined for several stiffened aircraft fuselage structures. This extended dynamic prediction capability is needed for the integration of mid-frequency noise control technology. Beam, plate and solid element models of the stiffener components are evaluated. Attachment models between the stiffener and panel skin range from a line along the rivets of the physical structure to a constraint over the entire contact surface. The finite element models are validated using experimental modal analysis results.

Abstract: Recent advances in computational speed have made aircraft and spacecraft crash simulations using an explicit, nonlinear, transient-dynamic, finite element analysis code more feasible. This paper describes the development of a simple landing gear model, which accurately simulates the energy absorbed by the gear without adding substantial complexity to the model. For a crash model, the landing gear response is approximated with a spring where the force applied to the fuselage is computed in a user-written subroutine. Helicopter crash simulations using this approach are compared with previously acquired experimental data from a full-scale crash test of a composite helicopter.

Abstract: The Active Structural Acoustic Control (ASAC) technique was applied to reduce propeller-induced noise and vibration in the passenger cabin of the deHavilland Dash-8 turboprop aircraft. Piezoceramic elements were used for structural actuation while velocity feedback of the fuselage was achieved through the use of accelerometers. Three actuators comprised of segmented piezoelectric elements were designed with the objective of reducing the noise and vibration levels at the propeller Blade Passage Frequency (BPF). Twelve accelerometers were grouped to effectively form three sensors. Second order classical compensators were designed for each of the three dominant control loops in order to achieve high gain at the BPF, rendering the closed-loop system insensitive to disturbances at the control frequency. The propeller acoustic field on the port side of the aircraft was simulated in a laboratory using a speaker-ring consisting of four loudspeakers. The control system was tested using this acoustic field, produci...

Abstract: The assembly jig for aircraft fuselage section halves to form a complete section has a frame (2) to support all the required components which are attached to retainers. The frame has hoist transfer frames (3) to position the fuselage sections in the assembly frame.

Abstract: A ree ned coupled rotor/e exible fuselage aeroelastic response model is developed for vibration reduction studies based on active control of structural response (ACSR). The structural model is capable of representing e exible, hingeless rotors combined with a e exible fuselage, and a rigid platform combined with the actuators required for modeling the ACSR system and is combined with a free wake model. The ine uence of rotor/fuselage coupling and improved aerodynamics on the vibratory hub loads is investigated.

Abstract: An aircraft capable of supersonic flight comprises a body portion including a fuselage, a wing, and an engine nacelle mounted below the wing. The aircraft may also include a high-mounted aft, tail. The area/lift distribution of the body portion is tailored to reduce sonic boom disturbance. The body portion further includes a blunt nose and a gull dihedral wing configuration that further reduces sonic boom disturbance and eases constraints on area/lift distribution tailoring. The gull dihedral wing or tail is configured to carry lifting force to its trailing edge to create an expansion at the aft end of the aircraft that reduces aft sonic boom ground shock strength. The volume of the mid-portion of the fuselage can be reduced above the wing to create a sloped surface that generates an expansion fan over the wings. The expansion fan lowers the pressure above the wing in the area covered by the expansion to reduce the pressure required on the upper and lower surfaces of the wing to generate the same lifting force.

Abstract: Acoustic enclosures such as aircraft cabins often display lightly damped modal behavior at low frequencies where passive treatments are impractical due to mass and volume constraints. This work presents a feedback control approach using dynamic H 2 controllers implemented with spatially weighted arrays of collocated pressure sensors and constant volume-velocity actuators. The open-loop system is shaped using spatially weighted transducer arrays to yield increased pole-zero separation, which results in better closed-loop performance. The transducer arrays are weighted to emphasize coupling to a particular acoustic mode or modes, which facilitates global control of the targeted dynamics. This work presents a method to determine the spatial weighting vectors for the transducer arrays from frequency response measurements. The development and implementation of low-order, dynamic H 2 control laws is also discussed. Experimental results are presented for a single-mode and a multiple-mode controller implemented on an aircraft fuselage section, and demonstrate significant reduction of the targeted acoustic modes. To the best of the authors' knowledge, this is the first experimental implementation of a feedback controller of this type capable of achieving such levels of global reduction in a three-dimensional acoustic system.

Abstract: A novel unsteady rotor-fuselage interactional aerodynamics model has been developed. This model loosely couples a Generalized Dynamic Wake Theory (GDWT) to a thin-layer Navier-Stokes solution procedure. This coupling is achieved using an unsteady pressure jump boundary condition in the Navier-Stokes model. The new unsteady pressure jump boundary condition models each rotor blade as a moving pressure jump which travels around the rotor azimuth and is applied between two adjacent planes in a cylindrical, non-rotating grid. Comparisons are made between measured and predicted time-averaged and time-accurate rotor inflow ratios. Additional comparisons are made between measured and predicted unsteady surface pressures on the top centerline and sides of the fuselage.